Design proposal FB-24 CD Report Final issue version
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Transcript of Design proposal FB-24 CD Report Final issue version
CRANFIELD UNIVERSITY
SCHOOL OF ENGINEERING
MSc THESIS
Academic years 2003 – 2006
GEOFFREY A WARDLE MSc CEng
ADVANCED INTERDICTION AIRCRAFT SYSTEM
CONCEPTUAL DESIGN STUDY (PROJECT NOVA).
INDIVIDUAL RESEARCH PROJECT
Supervisor: Professor J. P. FIELDING
February 2006
This thesis is submitted in partial (30%) fulfilment of the requirements for the degree of Master of
Science in Aircraft Engineering.
Cranfield University 2006. All rights reserved. No part of this publication may be reproduced
without the written permission of the copyright holder.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
i
Theses “Health” Warning
This thesis has been assessed as of satisfactory standard for the award of a Master of
Science degree in Aircraft Engineering. This thesis covers part of the assessment
concerned with the Individual Research Project. Readers must be aware that the work
contained is not necessarily 100% correct, and caution should be exercised if this
thesis or the data it contains is being used for future work. If in doubt, please refer to
the supervisor named in the thesis, or the Aerospace Engineering Group.
All of the views and material contained within this document are the sole research of
the author and are not meant to directly imply the intentions of the Joint Strike Fighter
Project Office, National Security Agency or any contractor, or any third party at this
date. Although the USAF awarded contracts for studies into extending the combat
range and enhancing the capabilities of both the F-35 and the F/A-22 in 2004 this is
not representative of the results or based on the results of any part of that body of
research which is secret.
This thesis is an unsolicited conceptual design study, which has been reviewed by Mr
Robert A. Ruszkowski, Jr Senior Staff Engineer of Lockheed Martin ADP for whose
advice I am eternally grateful.
This document contains no material governed by ITAR restrictions and the
distribution of all information contained within this document is unlimited public
release and has been approval by: - US D o D, UK M o D representatives, Lockheed
Martin (JSF Programme) and BAE Systems (Future Offensive Air Systems). This
document and any part thereof cannot be reproduced by any means for distribution
without written permission form Cranfield University and the author.
None of the data contained within this project is to be used in any format or for any
other research project without consultation with Cranfield University School of
Engineering supervisor named on the front cover.
Acknowledgements
Mr Robert A. Ruszkowski, Jr Senior Staff Engineer at Lockheed Martin ADP for
whose advice I am eternally grateful, Andy Bruce Design Lead for F-35C empennage
for private help with background material on JAST and proof reading the final draft,
and Phil Read JSF BAESYSTEMS security officer for obtaining security clearance
for this project, and finally acknowledgement to the following: - Steven A. Brandt:
John J. Bertin: Randall J. Stiles: and Ray. Whitford: of the UASFA for producing the
AeroDYNAMIC V2.08 analysis software and the AeroDYNAMIC V3 design
resource CD which was a major aid in producing this thesis.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
ii
Abstract.
The objective of this thesis was to produce a conceptual design study for a force
package consisting of a two seat advanced interdiction aircraft and a complementary
UCAV capable of replacing the UK Royal Air Force Tornado, Royal Australian Air
Force F-111, and USAF F-117 and F-15E air assets, both of these aircraft would have
greatly enhanced capabilities in stealth, range, and supercruise capability. The
performance being measured against a representative future mission profile produced
by the American Institute of Aeronautics and Astronautics, and the USAF Academy.
These aircraft represent a fusion of F/A-22A, F-35C, and YF-23 technologies with the
innovative F-120 Variable Cycle Engine (flight proven in the Advanced Tactical
Fighter CDA program), to produce single engine interdictor, with a predicted
supercruise capability, and 900mile combat radius, which was capable of carrying
typical current and near future internal weapons loads, for an estimated unit cost of
$75 million in 2006 dollars. The capability to supercruise combined with a high
degree of stealth was reasoned to severely degrade the capability of enemy defences
reducing their response time to a grater extent than that achievable with F-117 stealth
fighter enabling the aircraft to prosecute a high altitude attack in a greatly reduced
threat envelope, clear of AAA, and MANPADS, which have the fastest response
times, compared with the much larger SV300 and SV400 class weapons.
The initial starting point for these aircraft was the current F-35C USN aircraft carrier
variant of the Joint Strike Fighter, with all naval capability removed, and originally a
growth aircraft with a new wing was considered, using a similar development
methodology as the growth of the F/A-18C Hornet into F/A-18E Super Hornet or F-
16C into the YF-16XL. The requirement for supercruise capability over a large
portion of the mission lead to very major airframe changes: - to reduce wave drag by
increasing the finesse ratio of the fuselage, changing the wing planform, increasing
the wing leading edge sweep angle to keep the wing inside the shock cone produced
by the fuselage, and adoption of duel function twin ruddervators eliminating the
horizontal tail surfaces. The resulting aircraft were an evolution of the F-35C into a
larger aircraft suitable for proposed mission, and although much airframe
commonality was lost provision was made to integrate common systems within the
common airframe of both manned and unmanned variants and these systems represent
a greater percentage of overall weapon system cost than airframe itself.
This conceptual design study was for the modification of an exiting airframe design
and was different to that of a clean sheet of paper design consisting of two phases, the
first of which was configuration design and parametric analysis using both classical
analysis and the Jet306 / AeroDYNAMIC V2.08 analysis tool set, and the second was
major structural component layout of the airframe initial structure with systems
integration (the original intention of using hand calculations with PATRAN /
NASTRAN FEA modelling for structural sizing proved impractical due to time and
training resource constraints which also precluded CFD analysis in phase one). This
final design study for both versions of the AIA aircraft contained herein consists of
parametric analysis, initial optimisation and structural layout and constitutes a
feasibility study proposal to meet the requirements. Recommended further work on
the proposed designs includes CFD, and FEA analysis of the drag, aero and structural
loads.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
iii
Contents Page
Theses “Health” Warning i
Acknowledgements i
Abstract ii
Contents iii
Figures v
Table‟s xiii
Performance and design analysis charts xiv
Glossary xv
1.0 Introduction 1
1.1. Project brief and SOW 3
2.0 Requirements capture 6
2.1 Current F-35 family of aircraft data 10
2.2 Threat analysis 11
2.2.1 Air threats 11
2.2.2 Surface to air threats 13
2.2.3 Stealth requirements 15
3.0 FB – 24 configuration selection and optimisation 16
3.1 Initial conceptual design studies 16
3.1.1 FB - 24 and F-35C common features 20
3.2 FB – 24 and A-24 Configuration concepts 21
3.2.1 Configuration design challenges 36
3.2.2 Fuselage configuration design selection 37
3.2.3 Wing and Empennage configuration design selection 45
3.2.4 Common FB – 24 and A-24 configuration initial sizing 81
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
iv
3.2.5 Common FB – 24 and A-24 supersonic drag analysis 128
3.3 Configuration optimisation by parametric analysis 133
3.3.1 Analysis methodology 133
3.3.2 Analysis of NB1 and NB2 on AIA sizing mission 134
3.3 3 Comparison of the results for NB-1 and NB-2 144
4.0 Structural layout and major system integration 150
4.1 Undercarriage integration 151
4.2 Aircrew integration / AI integration 155
4.3 Propulsion system integration 160
4.4 Weapons systems integration 169
4.5 Structural Layout 176
4.6 Fuel tank integration 198
4.7 Aircraft OML G.A. drawings 199
5.0 Conclusions and recommendations for further work 202
References 205
Appendices A: - Detailed primary threat data 208
Appendices B: - Signature reduction 213
Appendices C: - DSI Overview 225
Appendices D: - F-35 Family Overview 227
Appendices E Design supplement 232
E – 1.1:- Structural amendments and concept completion 232
E – 1.2:- Structural weight estimation as modelled 244
E – 2:- Supersonic range and endurance analysis NB2 248
E – 3:- AeroDYNAMIC V 2.08 analysis methodology 253
E – 4:- Comparison of F-24 and A-24 with the RFP targets 265
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
v
Figures Page
1:- Conceptual strike aircraft RCS pole model 2
2:- UASF Strike Airframe life estimates 3
3:- Statement of work flow chart 4
4:- Jet306 / AeroDYNAMIC analysis examples 5
5:- AIA FB-24 mission profile 8
6:- The MAPO MiG-29M: tactical fighter 12
7:- The MAPO MiG-31B interceptor fighter 12
8:- The Il A-50 AWAC aircraft 13
9:- The S-300V (SA-12) missile system 14
10:- The QFD House of Quality process 18
11:- First tier House of Quality of the AIA 19
12:- Internal weapon options for FB-24 21
13:- Typical performance map for a tactical fighter aircraft 23
14:- General flow regimes encountered by tactical fighters 23
15:- Demonstration wing for twist and camber research NASA 25
16:- Effects of twist and camber on longitudinal aerodynamics 26
17:- Northrop / MDA YF-23 showing ruddervator configuration 27
18:- Boeing X-45 UCAV demonstrator showing tailless configuration 28
19:- Lockheed Martin X-44 showing 2-D vectored tailless configuration 29
20:- F-35A forward fuselage showing extent of the cockpit 32
21:- F/A-22A ACES II ejection seat showing installation angle 32
22:- F-22B Twin seat trainer variant of F/A-22 not built 33
23:- Thrust to weight-v-Wing loading for strike aircraft 35
24:- YF-22 Cross sectional area plot using Jet 306 showing intake effects 37
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
vi
25:- Wing position on combat aircraft fuselages 39
26:- Advantages of the shoulder mounted wing location 41
27:- Effects of area – rule shaping on supersonic aircraft 43
28:- Flow fields around cylindrical and chined bodies 44
29:- Evolution of wing planforms for combat aircraft 46
30:- YF-22 and Eurofighter configuration comparison 47
31:- Pressure and shear forces on an airfoil 43
32:- Nomenclature for aerodynamic forces in the pitch plane 43
33:- Moment balance to trim an aircraft 44
34:- Aircraft reference axes and corresponding aerodynamic moments 45
35:- Characteristics of airfoil sections 53
36:- Airfoil forces and moments 54
37:- Airfoil centre of pressure 54
38:- Aerodynamic centre 1 influence of location on moment 55
39:- Aerodynamic centre 2 location where moment is independent of alpha 55
40:- Special airfoil profiles 57
41:- Types of trailing edge flaps 60
42:- Lift and drag coefficient curves for wings with trailing edge flaps 61
43:- Types of boundary layed control devices 61
44:- Effects of boundary layer control devices on the lift curve 62
45:- NACA 0006 AIA wing root airfoil analysis 63
46:- NACA 64-006 AIA wing tip airfoil analysis 64
47:- Reynolds number effects on lift and drag curves 65
48:- Wing planform RCS spike effects 67
49:- Wing planform influence on design 68
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
vii
50:- Fuselage edge alignment influence on design 68
51:- Surface current scattering 69
52:- Reasoning behind wing configuration (B) 71
53:- Wing geometry 71
54:- Effect of aspect ratio on lift 72
55:- Wing relationship to aircraft Centre of Gravity 74
56:- The effect of sweep angle on MCRIT for a non tapered wing 76
57:- Initial definitions tail sizing 79
58:- F-35C port side tail group showing VT/rudder and flipper HT 80
59:- YF-23 Ruddervator starboard side, showing relative size and shape 80
60:- The effect of reducing wing loading on instantaneous turn speed 87
61:- Initial wing option (A) YF-22 based planform 88
62:- Initial wing option (B) JAST based planform 88
63:- Initial wing option (C) F-117A based planform 89
64:- Determination of the MAC for the initial option (A) wing 90
65:- Determination of the MAC for the option (B) wing 91
66:- Determination of the MAC for the option (C) wing 93
67:- Determination of the baseline F-35C MAC wing 94
68:- Original option (B) wing on FB-24 fuselage with F-35C sized tails 95
69:- Determination of the MAC for the revised option (A) wing 96
70:- Determination of the MAC for the revised option (B) wing 97
71:- Determination of the MAC for the revised option (C) wing 99
72:- Wing positioning on the fuselage option (A) wing 101
73:- Wing positioning on the fuselage option (B) wing 102
74:- Wing positioning on the fuselage option (C) wing 103
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
viii
75:- Empennage location and sizing conventional tail (A) wing (NB1) 111
76:- Empennage location and sizing ruddervator tail (A) wing (NB2) 111
77:- Empennage location and sizing conventional tail (B) wing (NB3) 112
78:- Empennage location and sizing ruddervator tail (B) wing (NB4) 112
79:- Empennage location and sizing conventional tail (C) wing (NB5) 113
80:- Empennage location and sizing ruddervator tail (C) wing (NB6) 113
81:- Baseline F-35C conventional tail sizing 114
82:- Large EHA horizontal tail and ruddervator actuator 117
83:- Small EHA rudder and wing trailing edge control surface actuator 117
84:- Multi segment leading edge lap actuator 120
85:- Flight control surface trade study for NB1 / NB2 wing 120
86:- FB-24 (NB1) Flight control surface layout 121
87:- FB-24 (NB2) Flight control surface layout 122
88:- FB-24 (NB3) Flight control surface layout 123
89:- FB-24 (NB4) Flight control surface layout 124
90:- FB-24 (NB5) Flight control surface layout 125
91:- FB-24 (NB6) Flight control surface layout 126
92:- Fuselage Design Breakdown for Jet analysis 134
93:- Surface model cuts on NB1 OML in isometric view 135
94:- Representative F-136-2 engine modelled in CATIA V5 135
95:- Area distribution for NB1 136
96:- Area distribution for NB2 140
97: Design limitations of the current F-35 aircraft 149
98:- Comparison of F-35 CTOL with F/A-22A 149
99:- F/A-22A internal structural layout illustrating complexity 150
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
ix
100:- FB-24 NB1 Tip back angle analysis 151
101:- FB-24 NB1 Overturn angle analysis 151
102:- FB-24 NB2 Tip back angle analysis 152
103:- FB-24 NB2 Overturn angle analysis 152
104:- Undercarriage retraction common for FB-24 / A-24 154
105:- Undercarriage storage common for FB-24 / A-24 155
106:- Design eye view for aircraft pilot commander FB-24 156
107:- Original crew station layout based on F-22B 157
108:- F-35 Canopy opening arrangement 158
109:- FB-24 Twin crew station integration 158
110:- The A-24 AI unit integration in fwd fuselage 159
111:- PW F-135 cutaway JSF CDA engine 160
112:- PW F-136 undergoing final inspection 161
113:- GE F-120 Schematic illustrating VCE operation modes 161
114:- Longitudinal section showing main components of the F-120 162
115:- Engine integration common to both FB-24 and A-24 164
116:- Common FB-24 / A-24 Intake duct 165
117:- Engine installation 166
118:- Engine airframe mounting 166
119:- Engine transporter 167
120:- Engine mounting and load transfer 168
121:- JASSM on transport mount 170
122:- Installation of SDB‟s 171
123:- Proposed ASRAAM installation 172
124:- Standard F-35 weapons bay with 2,000lb JDAM 173
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
x
125:- FB-24 / A-24 Weapons bay fit study 173
126:- Final FB-24 / A-24 weapons bay integration 174
127:- Future weapons for AIA the LOAAS 175
128:- Future weapons for AIA the SMACM 175
129:- F-35 Continuous wing substructure 177
130:- F-35 Fibre placed continuous wing skin 177
131:- F/A-22A Two separate wing substructure 178
132:- F/A-22A Wing skin / substructure assembly 178
133:-FB-24 / A-24 Phase 1 wing layout 179
134:- Phase 1 wing detailed description 180
135:- FB-24 / A-24 Phase 2 wing layout 181
136:- Phase 2 wing detailed description 182
137:- FB-24 / A-24 Final wing layout 183
138:- Final wing layout detailed description 184
139:- Detailed structural layout of FB-24 forward fuselage top view 186
140:- Detailed structural layout of FB-24 forward fuselage underside view 187
141:- Forward fuselage additional principle structure dimensions 188
142:- Detailed structural layout of centre fuselage top view 189
143:- Detailed structure layout of centre fuselage underside view 190
144:- Centre fuselage and wing integration 191
145:- Differences in centre fuselage frame between FB-24 and F-35 191
146:- Detailed aft fuselage layout top view 192
147:- Detailed aft fuselage underside view 193
148:- Ruddervator structural layout 194
149:- A-24 AI integration into complete forward fuselage structure 194
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
xi
150:- Complete airframe structural integration model 195
151:- F-35C plan view for comparison with figure 149 196
152:- F-35C side view for comparison with figure 149 196
153:- FB-24 Fuel tank integration model 197
154:- A-24 Fuel tank integration model 197
155:- Dimensioned FB-24 GA drawing Plan view 199
156:- Dimensioned FB-24 GA drawing Side view 200
157:- Dimensioned FB-24 GA drawing Front view 201
A.1:- MiG Light Weight Fighter Project aircraft concept 208
A.2:- Chinese J-10 Light Weight point defence fighter 208
B.1:- Tornado GR4 current RAF deep strike aircraft 215
B.2:- USAF F-15E Strike Eagle 216
B.3:- USAF F-111E similar to the RAAF F-111C‟s 216
B.4:- Spike alignment or plan from alignment on the F-35C 217
B.5:- Spike side lobe alignment on the F-35C 218
B.6:- F-117A illustrating facetted approach to RCS reduction 219
B.7:- The F-35 SigMA pole model about to undergo RCS measurement 220
B.8:- The F-35 SigMA model mounted at the test range 221
B.9:- Pratt & Whitney LO nozzle under ground based test 223
B.10:- General Electric AVEN nozzle (LOAN) on show 223
B.11:- F/A-22A 2-D vectoring nozzle installed on an F/A-22A 224
C.1:- Diverterless Supersonic Intake on CDA mock – up 226
C.2:- GFD Model of the flow fields around the DSI 227
D.1:- The F-35A USAF CTOL aircraft three view 228
D.2:- The X-35A in flight 229
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
xii
D 3:- The F-35B USMC / RN /RAF aircraft three view 230
D.4:- The X-35B in hover 230
D.5:- The F-35C USN aircraft three view 231
D.6:- The F-35C maximum stores carriage capability 231
E.1:- The FB-24 MOSC hatch structure 233
E.2:- The FB-24 MOSC hatch integration with the forward fuselage 233
E.3:- Revised A-24 forward fuselage layout top view 234
E.4:- Typhoon refuelling indicating required A-24 field of regard 235
E.5:- Revised structural arrangement common aft fuselage top view 236
E.6:- Revised structural arrangement common aft fuselage underside view 237
E.7:- Revised common centre fuselage structure top view 238
E.8:- Revised FB-24 detailed structural layout model 239
E.9:- Revised A-24 detailed structural layout model 240
E.10:- FB-24 / A-24 revised common wing design 241
E.11:- FB-24 / A-24 metallic wing joint philosophy 242
E.12:- FB-24 / A-24 composite wing joint philosophy 242
E.13:- FB-24 / A-24 composite substructure common wing design 241
E.14:- A-24 Evolution general arrangement model 243
E.15:- Structural weight measurement methodology for frames 244
E.16:- Structural weight measurement methodology for keels and longerons 245
E.17:- Structural weight measurement methodology for wing skins 248
E.18:- Structural weight measurement methodology for fuselage skins 248
E.19:- Airplane geometry for downwash prediction 259
E.20:- AeroDYNAMIC V 2.08 Lifting surface analysis 265
E.21:- AeroDYNAMIC V 2.08 Surface area analysis 267
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
xiii
Tables Page
1:- Mission breakdown for the Advanced Interdiction Aircraft 3
2:- Government Furnished Equipment 34
3:- Predicted variation of MNCRIT with sweep angle 77
4:- Experimental variation of MNCRIT with sweep angle 77
5:- Mass: wing loading: and dry thrust / weight ratios 83
6:- Wing configuration study summary 96
7:- Revised wing sizing study to common wing loading 101
8:- Wing configuration layout 102
9:- Tail sizing results for HT, VT, and RV configurations 109
10:- Exposed tail surface areas 118
11:- Flight control surface sizings relative to the wing area 127
12:- NB1 and NB2 major component weight from AeroDYNAMIC 144
13:- F-24 / A-24 comparison with RFP targets 202
B-1:- Typical Radar Threats 215
E-1:- Wing and Empennage structural component weight table 245
E-2:- Forward fuselage structural component weight table 246
E-3:- Centre and Aft fuselage structural component weight table 247
E-4:- Major component skin weight estimate 249
E-5:- As drawn weight and Jet 2.08 weight prediction comparison 249
E-6:- F-24 / A-24 Definitive weight statement 250
E-7:- F-24 / A-24 Lift curve analysis 266
E-8:- F-24 / A-24 Aerodynamic analysis 266
E-9:- F-24 / A-24 Drag Polar analysis 268
E-10:- F-24 / A-24 CD vs CL analysis 268
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
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E-11:- F-24 / A-24 Lift over Drag vs CL analysis 269
E-12:- F-24 / A-24 CL ^1.5 over CL analysis 269
Performance and design analysis charts Page
1(a):- NB1 Total predicted drag variation with Mach number 136
1(b):- NB1 Drag Polar 136
2:- NB1 Lift over drag v lift coefficient at (a) Mach 0.85 and (b) Mach 1.5 137
3:- NB1 Lift curve CL v 137
4:- NB1 CD v CL at (a) Mach 0.85 and (b) Mach 1.5 138
5:- NB1 Performance analysis Thrust and Drag v Mach number 138
6(a):- NB-1 V-n Diagram 139
6(b):- NB1 Manoeuvre diagram 139
7:- NB1 Specific Excess Power curves 139
8(a):- NB2 Total predicted drag variation with Mach number 140
8(b):- NB2 Drag Polar 140
9:- NB2 Lift over drag v lift coefficient at (a) Mach 0.85 and (b) Mach 1.5 141
10:- NB2 Lift curve CL v 141
11:- NB2 CD v CL at (a) Mach 0.85 and (b) Mach 1.5 142
12:- NB2 Performance analysis Thrust and Drag v Mach number 142
13(a):- NB2 V-n Diagram 143
13(b):- NB2 Manoeuvre diagram 143
14:- NB2 Specific Excess Power curves 143
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
xv
Glossary
ALOSNW Air launched Low Observable Stand off Nuclear Weapon
APU Auxiliary Power Unit
AIA Advanced Interdiction Aircraft
AMRAAM Advanced Medium Range Air to Air Missile
ASRAA Advanced Short Range Air to Air Missile
ATF Advanced Tactical Fighter Programme (now F/A-22A)
B Bomber
BAe British Aerospace now BAESYSTEMS
BCA/BCM Best - cruise - altitude / Best - cruise - Mach
BVR Beyond Visual Range
CDA Concept Demonstrator Aircraft
CFD Computational Fluid Dynamics
CTOL Conventional Take-Off and Landing
CV Aircraft suitable for large aircraft carrier operations
DoD Department of Defence
ECM Electronic Counter Measures
EMPW Electro Magnetic Pulse Weapon
F Fighter
F/A Fighter Attack
FB Fighter Bomber
FEA Finite Element Analysis
FOAS Future Offensive Air System now SUAV (E)
FSAV Future Strategic Air Vehicle
FBW Fly By Wire
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
xvi
FBL Fly By Light
GR Ground attack and Reconnaissance
GE General Electric (aero engines)
IR Infrared (referring to signature)
JAST Joint Advanced Strike Technology now JSF F-35
JSF Joint Strike Fighter
JSOW Joint Stand Off Weapon
LOAN Low Observable Axisymmetric Nozzle
LMTAS Lockheed Martin Tactical Aircraft Systems
MDA McDonnell Douglas Aircraft now Boeing
MoD Ministry of Defence (United Kingdom)
NG Northrop Grumman
PGW Precision Guided Weapons
P&W Pratt & Whitney (aero engines)
PWSC Preferred Weapons System Concept
RAF Royal Air Force
RAAF Royal Australian Air Force
TSFC Thrust Specific Fuel Consumption
SDD System Development and Demonstration
T/W Thrust to Weight ratio
UCAV Unmanned Combat Air Vehicle
USAFA United States Air Force Academy
Wo, We, Wf Aircraft designed takeoff, empty, and fuel weights
W/S Wing loading
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
1
1. Introduction.
This Individual Research Project (IRP) thesis forms the submission by the author for
the IRP component of the Cranfield University MSc in Aircraft Engineering, and in
no way represents the views of the F-35 Joint Project Office, the US D o D, UK M o
D, or any F-35 / F/A-22 Contractor Company.
The mission specification for this project is based on the American Institute of
Aeronautics and Astronautics annual aircraft design project competition of 2001/02.
The Request for Proposal (RFP) published by the AIAA is based on recent industrial
project work. Hence this competition provided a useful source of realistic mission
requirements and operational data that form the basis of this submission, additional
information from public domain material for a hypothetical Future Strategic Air
Vehicle program FSAV was also used to tailor this aircraft to the Tornado, F-15E, F-
117A, and F-111C replacement roles. (Reference 1: - Aircraft Design Projects for
Engineering Students: by Jenkinson L. R. and Marchman J. F. Published by AIAA
Education Series in 2003. ISBN 1-56347-619-3. Pages: - 208-209, and USAFA course
FSAV study included in Aerodynamic Version 3.0 Software: Published by the USAFA
and released by AIAA Education Series in 2004: ISBN 1-56347-689-4.)
The Tornado GR4 is the UK‟s only deep strike aircraft, and like the previous
generations of strike aircraft the Tornado was tailored to high speed ultra low level
interdiction missions. Both Gulf War‟s and the Balkans conflict have demonstrated
that interdiction missions are extremely high risk at low level, and the need for high
level strikes with stealth aircraft in these recent conflicts has been apparent.
In fact the exportation of former Soviet Union and Russian Federation advanced
technology SAM, AWACS, and AAA weapons makes survivability with PGW, even
on the medium altitude missions of 20,000 ft for all non stealthy platforms
questionable, without extensive air defence degradation.
The only stealth tactical strike aircraft in service with any nation is the F-117A which
by 2020 will be entering its airframe retirement age, the same is true for the RAF
Tornado Interdictor Strike (IDS) aircraft and USAF F-15E Strike Eagle and both latter
aircraft require an extensive reduction in the enemy air defence network to survive
long enough to deliver mission ordinance against aggressor nations which have
advanced air defence capability. This in turn requires a larger expeditionary force
with specialised ECM screening and SEAD attack aircraft, which increases the overall
cost of any major air interdiction operation, and eliminates the element of surprise. In
addition to the conventional interdiction role this aircraft will form the common
airframe for both manned and unmanned components of a mixed fleet, and by the
2020 time frame the UK will be seeking a new strategic weapons platform, when the
Trident SLBMS submarines reach hull retirement age therefore every possibility
exists that a new aircraft will have a strategic deterrents role, with tactical nuclear
weapons, and become the Future Strike Air Vehicle (FSAV) platform of choice.
To study the UK‟s future platform requirements a manned stealth technology
demonstration model was produced as the Replica LO pole model figure 1, and a UK
variant of the Lockheed F-117 was also considered, both of these projects have been
replaced by SUAV (E) a strategic UAV research program.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
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Therefore the need exists for a new aircraft which can effectively deliver conventional
tactical PGW‟s or compact stand – off nuclear weapons, over long ranges and which
can rapidly deploy with minimum support to regional conflicts world-wide, and
survive in a heavily defended enemy air space without additional air assets.
Improved threat capabilities posed by the S-300V Gladiator and S-400V Triumf SAM
systems dictate that this new aircraft should have the following attributes: - an RCS of
-20db‟s over a frequency range of 0 to 20 GHz : an IR signature at military power
below a wavelength of 8 microns: the capability to fly at high altitude up to 45,000ft
for sustained periods at high Mach number M1.6 to reduce the defence response time:
and the capability to fly supersonically with reduced emissions: and have long range
on internal fuel, which will allow the aircraft to respond to crises around the world
from out of theatre bases which cannot be used by current strike assets, without
carrying external fuel tanks or air to air refuelling.
Approximately 300 aircraft could be required by the USAF for F-15E / F-117A
replacement, with 200 required for the UK RAF to meet a projected FSAV
requirement, and an additional 50 aircraft for the RAAF as F-111C replacements.
Figure 1: - In support of UK future platform solutions full – scale RCS
qualification model was produced which had attributes of a stealthy aircraft in
terms of outer mould lines and plan - form alignment. Source: - BAE Systems
Warton.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
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AERO 481
Figure 2: - USAF Strike airframes that require replacement by 2020. Source: -
AIAA /USAFA Aerodynamic V3.0 included USAFA AERO 481 lecture material.
1.1 Project brief and SOW.
The objective of this project is to produce a conceptual design and structural layout
for both aircraft of the mixed fleet concept, capable of meeting the requirements of
interdiction and strike missions using both conventional and nuclear weapons outlined
above and detailed in section 2. This aircraft will be based on a fusion of F-35C and
F/A-22A technologies.
This project will not use any material which is not within the public domain. The
designations proposed for this aircraft namely FB-24 for the USAF and AIA-1 for the
RAF and RAAF distinguishes this project from any exiting or near future BAE
SYSTEMS / Lockheed Martin Tactical Aircraft projects and prevents confusion with
other programmes of research.
To meet the AIAA and hypothetical FSAV amended mission requirements the F-35C
is to be used as the starting point for this study, with the fuselage extended to give a
fineness ratio comparable to the F/A-22 and a new wing based on one of the
following configurations: - (1) a new wing based on the F/A-22A planform of
increased size as a conservative option: (2) a derivative version of the JAST cranked
arrow head layout , and as the most radical solution (3) a double delta wing similar in
plan form to the F-16XL, all options would have the stretched fuselage and a reduced
fuselage cross section area. The scope and deliverables for this study are shown in the
flow chart figure 3.
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The maximum GTOW of 71,018lbs (the maximum overload F-35C weight) must not
exceed for the FB-24 in the clean condition (no external stores provision). The
maximum empty weight must not exceed 38,380lbs 18.76% greater than the F-35C,
conceptual weight target breakdown will be as follows:- Pilots plus one M61A
120mm cannon = 2%: engine / nozzle / oils / fluids = 7% (based on the dry weight of
the F110-GE-132 engine which is 5.7% of the GTOW): expendable weapons = 8%:
Structural / avionics / undercarriage / systems = 45% and fuel = 38%. The with a
43,000lb thrust engine the thrust to weight ratio would be 1:0.6 comparable with the
Tornado GR-4, F-15E and the F-117A.
CranfieldUNIVERSITY BAE SYSTEMSDocument No: - WSL2004-IRP PRES 9.
STAGE 1
Configuration generation.
Optimisation analysis and trade study.
Basepoint configuration
selection.
Interim report Sept
2005.
OML Freeze.
Interim oral
report Jan
2006.
STAGE 2
Structural layout and systems integration.
Structural layouts and
trade studies.
Layout Freeze. Systems
integration.
FINAL REPORT.
Feb2006
Figure 16 :- Advanced Interdiction Aircraft Design study flow chart .
Figure 3: - SOW Flow Chart of the FB-24 conceptual design study.
STAGE 1: -
The F-35C public domain data was used to produce the surface model of the F-35C –
230-5 baseline OML configuration to obtain aerodynamic, performance, and
endurance data against the AIAA / FSAV mission requirement. Five proposed aircraft
configurations were then evaluated against the F-35C figures for the same AIAA /
FSAV mission profile to compare, each configuration. The Jet306 Parametric analysis
toolset which is the core of the USAF Academy Aerodynamic V3.0 software was
used for these configuration trade studies.
This tool performs whole aircraft analysis for: - lift: parasite drag: induced drag
supersonic drag: propulsion analysis: weight prediction: constraint analysis: sizing:
optimisation: performance analysis and cost analysis, based on geometry for the
conceptual design and mission profile entered on spreadsheets by the designer with
the fidelity of the analysis depending on the number of section cuts and geometry data
points used this forms the stage 1 trade study. Examples of Jet306 analysis of a test F-
16 are shown in figure 4.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
5
0 10 20 30 40 50
Constraints
0
1
2
3
4
5
6
0 20
40
60
80
100
120
140
160
Wing Loading, Wto/S, psf
Th
rust
Lo
ad
ing
, T
sl/W
to
M xM ach
M idTurn
Ps
Ingress
LoTurn
HiTurn
SloTurn
HiCruise
Takeoff
Landing
Actual
Design Point
Desired
- 4 0
- 3 0
- 2 0
- 10
0
10
2 0
3 0
4 0
5 0
0 10 2 0 3 0 4 0 5 0 6 0
Figure 4: Jet306 / AeroDYNAMIC analysis output examples for an F-16 model.
Measures of Merit
The results of the Jet 306 of the five aircraft configurations were then evaluated using
the following measures of merit to down select a single configuration to meet the AIA
/ FSAV mission requirements as detailed in section 2: -
1.1 Weight summary (GTOW, We, Wf, W/S, T/W, Wf /W) using only
internal fuel.
1.2 Aircraft geometry (wing and control surface area, fuselage size and
volume, frontal cross sectional area distribution, wetted area, inlet
and diffuser, landing gear, weapons carriage, sensor and avionics
locations, crew stations, etc.)
1.3 Mission duration, radius or range, fuel burn by mission segment for
each design mission.
1.4 Take-off and landing distance for each design mission including
standard day and icy runway balanced field length at sea level.
1.5 Performance at maneuver weight with 50% internal fuel and with
two ASRAAM missiles and two 2000lb JDAM design mission
loadings.
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1.5.1 Maximum Mach number at 45,000 ft.
1.5.2 1-g Maximum Thrust Specific Excess Power Envelope
1.5.3 2-g Maximum Thrust Specific Excess Power Envelope
1.5.4 Maximum Thrust Sustained Load Factor Envelope
1.5.5 Maximum Thrust Maneuvering Performance Diagrams
1.5.5.1 34,000 ft
1.5.5.2 45,000 ft
1.6 Flyaway and total life costs.
STAGE 2: -
The down selected configuration was then developed into full CATIA V5 R10 model
for structural analysis and systems integration and this forms the stage 2 design
analysis.
Basepoint structural layouts of the aircraft to meet the systems integration
requirements, for weapons, aircrew, AI system, fuel, and undercarriage were
produced in CATIA V5 R10, and this model was then to be analysed using PATRAN
/ NASTRAN Finite Element Analysis package, however time constraints precluded
this and the structural analysis will differed for a later work package to be perused at a
latter date. This structural analysis will enable initial sizing of the wing substructure
and concluded the conceptual structural layout design freeze.
The final report consists of a complete conceptual design analysis and structural
component layout of the FB-24 Advanced Interdiction Aircraft, although without the
supporting structural analysis, however it forms the basis for the conceptual proposal
submission for the Advanced Interdiction Aircraft as a common manned and
unmanned airframe.
2. Requirements capture.
1. The AIAA and FSAV requirements for the AIA specifies a two-place
advanced deep-interdiction aircraft and ingress and egress portions of the
mission will be flown at supersonic speed at high altitude with target
acquisition and weapons release at high altitude. The mission profile is shown
in figure 5 and broken down in table 1 below, and is flown over a combat
radius of 900 nautical miles. The maximum TOGW of 71,018lbs (the
maximum overload F-35C weight) must not exceed for the FB-24 in the clean
condition (no external stores provision). The maximum empty weight must not
exceed 31,177lbs 3.3% greater than the F-35C.
2. The aircraft must be capable of „all-weather‟ operation including operation
from and on to icy 800 ft runways.
3. The aircraft must operate from forward NATO and allied bases, with the
minimum of support facilities. On these bases the aircraft will be required to
fit into standard NATO shelters.
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4. The design layout should allow for easy maintenance. Minimum reliance on
support equipment is essential for off-base operations.
5. Structural design limit load factors of +7g to -3g (aircraft clean and with 50
percent internal fuel) are required. An ultimate design factor of 1.4 is to be
applied to reduce weight. The structure must be capable of withstanding a
dynamic pressure (q) of 2133 lb/ft2 (i.e. equivalent to (q) at 800kt) and be
durable and damage tolerant.
6. All fuel tanks must be self-sealing. Aviation fuel to JP-8 and JP-5
specifications is to be assumed.
7. Stability and handling characteristics must meet MIL-F-8785B subsonic
longitudinal static margins to be no grater than +10% and no less than -30%.
8. In addition to the basic mission criteria the design specification requires the
following manoeuvring targets be met (specific excess power, SEP, is defined
as Ps in the following equation: - Ps = [(T/W)-(D/W)] V (where weight = W
=Mg)).
SEP (1g) military thrust (dry), 1.6M at 45,000ft = 0ft/s.
SEP (1g) maximum thrust (wet), 1.6M at 45,000ft = 200ft/s.
SEP (2g) maximum thrust (wet), 1.6M at 45,000ft = 0ft/s.
Maximum instantaneous turn rate, 0.9M at 15,000ft = 8.0degrees/s.
9. The design specification calls for five separate internal weapon capabilities as
shown in figure 4, but the deletion of the external weapons carriage options of
the existing F-35 family:
Two Mk-84 Low Drag GP + two ASRAAM.
Two 2000lb JDAM PIP + two ASRAAM.
Two AGM - 154 JSOW + two ASRAAM.
Six 250lb Boeing Small Diameter Smart Bombs + two ASRAAM.
Two ALOSNW (WE177B replacement) + two ASRAAM.
Two Special Elements (EMPW) + two ASRAAM.
10. Signature requirements given in the AIAA specifications are that the front
aspect RCS illuminated by a 1 – 10 GHz GCI, acquisition, and tracking radar
should be less than 0.05 m2. This would be demonstrated by SIGMA full scale
model, as for the F-35 described in appendices B.
11. The flyaway cost for 550 aircraft purchase must not exceed $75M (year 2001
dollars).
12. The IOC for the first operational squadron is 2020.
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Figure 5: - AIA Baseline sizing mission. (This is based on reference 1).
13. The aircraft must be capable of using the following fuels JP-5 (RAF) (JP-5
Mil – spec density of 6.82lb/US gal equal to 51.10lbs/ft3) / JP-8 (USAF) (JP-8
Mil - spec density of 6.80lb/US gal equal to 50.86lbs/ft3) / and JP-4 (Special
fuel) (JP-4 Mil – spec density is 6.55 lbs/US gal equal to 49lbs/ft3) without
reducing range or mission effectiveness.
14. Maximum use is to be made of off platform sensor inputs for mobile target
acquisition (e.g. SCUD‟s, SS-20 IRBM‟s, and SS-25 ICBM‟s).
15. A UACV variant of the FB-24 designated A-24 of common core airframe was
required for the mixed fleet concept and this airframe is covered in this
conceptual design study the detailed systems study will form a subsequent
research project*.
*Covered in full in proposal submission document RFPS – 022007 which is not
part of the Cranfield MSc course.
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Table 1:- Mission breakdown for the Advanced Interdiction Aircraft.
Segment. Description. Height. Speed. Distance/Duration.
1 - 2 Warm-up, taxi
and take-off. Sea – level.
NATO 8000ft icy.
2 - 3
Climb to best
super – cruise
altitude.
3 - 4 Cruise to conflict
area. BCA. BCM 522nm
4 - 5 Climb to
50,000ft.
5 - 6 Dash to target. 45,000ft Mach 1.6 378nm
6 - 7 Turn and
weapon release. 45,000ft Mach 1.6 180 degrees.
7 - 8 Dash out. 45,000ft Mach 1.6 378nm
8 - 9 Descend to super
- cruise altitude.
9 - 10 Cruise return. BCA. BCM 522nm
10 - 11 Descend to
airbase.
11 - 12 Land (with
reserve fuel*). NATO 8000ft, icy.
*Diversion and hold at sea level 30 minutes fuel at economical flight conditions.
(Reference 1(a): - Aircraft Design Projects for Engineering Students: by Jenkinson L.
R. and Marchman J. F. Published by AIAA Education Series in 2003. ISBN 1-56347-
619-3. Pages: - 208-209, and 1(b) USAFA course FSAV study included in
Aerodynamic Version 3.0 Software: Published by the USAFA and released by AIAA
Education Series in 2004: ISBN 1-56347-689-4.)
The A-24 UCAV used the same sizing mission defied above and was considered for
strikes on enemy chemical and biological research facilities with non-nuclear
weapons only, (the arming of UCAV‟s with tactical or strategic nuclear weapons
being politically unacceptable) this variant would be autonomous after control release
from the manned FB-24 (each FB-24 would support four A-24‟s) outside the key treat
zone. The FB-24 manned aircraft would control in-flight refuelling of A-24‟s when
required, conduct systems health checks, and mission management acting as fleet
commander in all mixed fleet operations. As a single strike component the FB-24
would conduct all nuclear strike missions and would be the only nuclear equipped
counterforce component of a mixed fleet operation.
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2.1 Current F-35 family of aircraft.
The current F-35 Joint Strike Fighter family consist of three aircraft variants for three
United States military air arms and the United Kingdom armed forces, which have
common basic configurations and use cousin parts in most cases, are outlined below
in terns of lead dimensions: performance: and weight.
F-35A (CTOL Variant): - This is the USAF variant with 9g manoeuvring
capability and replaces the F-16 and A-10 (current orders are for 1763 aircraft).
Dimensions are: - Wing span 35 ft: length 51.1 ft: and wing area 460 ft2. Empty
weight is: - 27,395 lbs. Internal fuel capacity is: - 18,498 lbs. Combat radius is: -
greater than 590 nautical miles.
F-35B (STOVL Variant): - This is the USMC / U.K. RN & RAF variant to
replace the Harrier AV-8B / Mk - 9 and the F/A - 18 Hornet C/D aircraft (current
orders are for 759 aircraft). Dimensions are: - Wing span 35 ft: length 51.1 ft: and
wing area 460 ft2. Empty weight is: - 30,697 lbs. Internal fuel capacity is: - 13,326
lbs. Combat radius is: - greater than 450 nautical miles.
F-35C (CV Variant): - This is the US Navy aircraft carrier variant which will
replace the F/A - 18 Hornet C/D and the F - 14D Tomcat (current orders are for 480
aircraft). Dimensions are: - Wing span 43 ft: Length 51.4 ft: Wing area 620 ft2.
Empty weight is: - 30,618 lbs. Internal fuel capacity is: - 19,100 lbs. Combat
radius is: - greater than 600 nautical miles. Maximum payload all internal and
external stores stations used: - 21,400lbs.
Currently the F-35 family are to be fitted with either the 43,000lb maximum thrust
Pratt & Whitney F-135 or the GE Aircraft Engines / Roll Royce F-136 engines, which
will be sourced in alternate year USA Department of Defence equipment purchases.
The F-35 family fuselage OML is driven primarily by the need to house the STOVL
variants lift fan / drive shaft combination, and the two 2,000lb JDAM and two AIM-
120 missiles of the CTOL variant requiring a large cross sectional area, additionally
the common fuselage length is driven by deck spot requirements of both the STOVL
and CV variants.
The new FB-24 will not have the same commonality drivers because this airframe is
driven by the requirements of increased range and the ability to super cruise, however
it will use the F-35 systems and many common features developed on the F-35. This
type of reconfiguration follows the precedent set by the F/A-18E and F variants
which although larger than the original F/A-18 family share common baseline
configuration.
For this aircraft study the current F-135 and F-136 Pratt and Whitney, and GE
respective engines of 43,000lbs maximum wet thrust were the initial power plants of
choice which giving commonality with the current F-35 family, although an enhanced
version of the YF-120 VCE was to be considered with a military dry thrust of
30,000lbs if the former engines proved to be unsuitable.
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2.2 Threat analysis.
Although the threat from the former Soviet Union has receded, most potential
aggressor nations have purchased, or inherited similar advanced air-to-air and surface-
to-air threat capabilities. Former Soviet Union fighter missile and airborne early
warning aircraft have been and continue to be sold to threat nations and these will be
assessed below.
The use of lookdown shoot down radars on fighters like the MiG-31 Foxhound
combined with Airborne Warning and Control System (AWACS) aircraft like the
Ilyushin / Beriev A-50 Mainstay, and the high performance of man portable surface to
air missiles like the Stinger, make the terrain masking attacks of the 1970‟s highly
questionable, and direct low level passes for airfield attack, as conducted by RAF
Tornado‟s in the first Gulf War suicidal.
The reaction time of defensive SAM systems, and air defence fighters is reduced if
the aircraft is stealthy, and the significant reduction of an aircrafts radar, infra-red,
visual and acoustic signatures, can greatly enhance its survival, resulting in grater
weapons system effectiveness. Although radar and infra-red stealth are the most
obvious signature features, other aspects of low observability (i.e. acoustic and visual)
cannot be ignored. The lessons learnt by the US air arms during the Vietnam War
emphasised the need for signature control, especially in the areas of Radar Cross
Section (RCS) and Infra – Red (IR) which will be covered below as well as visual and
acoustic signature reduction in general terms.
2.2.1 Air threats:-
For this study the primary air to air threats are anticipated to be MiG-29C‟s: MiG-
31M‟s: and control aircraft of the A-50 Mainstay type because in the future Global
Strike ConOps force in which this aircraft is intended to operate Eurofighter Typhoon,
and F/A-22 Raptor air dominance fighters will be responsible for tying down the Su-
27C / P, and Su-30A enemy air assets, when encountered.
Also with the latter aircraft being the most expensive Russian export platforms
nations capable of procuring them are most likely to employ them for free ranging
attack fighters rather than homeland defence tied combat air patrols. Each of the
probable threat aircraft capabilities are detailed in Appendices A, to the extent to
which published data is available.
The MiG-29 M shown in figure 6 is a major air threat to the AIA and from the list of
end users above the five nations highlighted in dark red are all potential near term
threat nations. Also this aircraft has been used against NATO forces in the Balkan‟s
war of the 1990‟s.
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The MAPO MiG-29M: air defence fighter.
Figure 6:- The MiG-29 M, tactical air defence fighter which is considered to be
one of the best of the fourth generation fighters and is used by: - Belarus:
Bulgaria: Cuba: Czech Republic: Germany: Hungary: India: Iran: Kazakhstan:
Malaysia: Moldova: North Korea: Poland: Romania: Russia: Serbia: Slovakia:
Syria: Turkmenistan: Ukraine: USA(combat training): Uzbekistan: Yemen.
(Ref:-8 and 9).
The MAPO MiG-31B: interceptor fighter.
Figure 7:- The MiG-31B currently Russia‟s principal interceptor fighter as seen
at le Bourget, Paris air show and currently variants are to be offered for sale to
Syria: Iran: North Korea: and Yemen (Ref: - 8 and 9).
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The MiG-31B figure 7 is capable of intercepting airborne targets in all weather
conditions, day or night, in continuous or intermittent radar coverage of Ground
Control Interception (GCI) sites, and is unaffected by the target aircrafts use of
electronic counter measures (ECM) or evasive manoeuvring. For operations outside
continuous GCI radar cover the MiG-31B can use its APD-518 (Apparatura Peredachi
Dannykh / Data Transfer Equipment) data – link system, both for reception of
guidance commands and for the distribution of target information (the air situation)
between aircraft in a formation. This would usually be a flight leader and three
wingmen. The MiG-31B combat radius is limited to a supersonic intercept radius of
389nm due to engine lubrication, and crew endurance, however at subsonic speeds
this is increased to 648nm, without overload tanks and up to 756nm when these tanks
are used.
Ilyushin / Beriev A-50 Mainstay AWAC
Figure 8:- The weathered surface of the rotating antenna of the Shmel
(Bumblebee) radar system is in contrast to the clean lines of this A-50, note the
fin – top fairing is for the Mnk (Poppy) missile approach warning system. Used
by: - Russia: India: and offered to Syria: Iran: North Korea: PRC: (Ref:-9)
The third and possibly the most important asset to the potential enemy nations is the
Ilyushin / Beriev A-50 Mainstay long range airborne warning and control system
(AWACS), shown in figure 8, (developed jointly, as the designation indicates, by the
Ilyushin Design Bureau and the Beriev Aviation Scientific and Technical Complex at
Taganrog on the sea of Azov), as a airborne interception control platform this aircraft
is a priority target for the air launched low observable anti – radiation weapon.
2.2.2 Surface to air threats: -
Russia is selling its latest and best systems on the world market and the two mobile
weapons of choice purchased by the threat nations are outlined below, neither of
theses can easily be destroyed by the X-45 or the X-47 UCAV‟s or indeed sea
launched TOMAHAWK cruise missiles.
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Primary surface to air threats to the FB-24 will come from the mobile S-300V SAM
system, and the highly mobile Buk-M1 (SA-11) replacement for the Kub (SA-6)
SAM that shot down Captain Scott O‟Grady‟s F-16 over Bosnia.
The S-300V (SA-12): - Gladiator SAM system.
Figure 9: - The S-300V (SA-12) missile system which can be deployed or
dismantled and moved in five minutes. Developed in the 1980‟s the full S-300V
system has been exported to the following nations: - Syria: Iran: North Korea:
PRC: (Reference2: -Pages 87-88, F-22 Raptor: by Sweetman B: Published by MBI
Publishing company USA 1998)
The S-300V (SA-12A Gladiator) system shown in figure 9 has two missiles
(developed from the anti-ballistic missile SA-12B Giant SAM), one large and one
small, the smaller one has a peak velocity of Mach 6 and can destroy targets evading
at 8g through clutter and ECM / decoy systems over an effective range of 30nm at
altitudes between 2,000ft and 60,000ft, the second larger missile attains a peak
velocity of Mach 8 and through advanced terminal aerodynamics can destroy targets
manoeuvring at 12g through ECM / decoy systems at altitudes between 12,000ft and
80,000ft. Targeting information can be obtained from the main Almaz NPO family of
sensors or off bored from the A-50 Mainstay or MiG-31 Foxhound airborne
platforms, and are capable of intercepting HARM missiles launched against them.
In addition the Almaz S-400 Triumf (SA-20 Gargoyle) family which are improved
faster derivatives of the S-300V‟s are now available on the world market after
completing field trials. (Reference 3:- pages 284-285, Iron Hand Smashing the
Enemy’s Air Defences: by Thornborough. M. A. and Mormillo. B. F.: Published by
Patrick Stephens Limited an imprint of Haynes Publishing UK 2002)
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The Buk-M1 (SA-11):- SAM system.
The Buk-M1 is the replacement for the Kub (SA-6) and is in full scale production for
the export market, with a single system which is mounted on 11 vehicles a defender
has 36 missiles ready to launch at any one time. The missiles have a reach of some 45
to 105nm, depending on model similar to that of the SA-10 Grumble, with a closing
speed of up to 14,770ft/sec. Also like the SA-10 this weapon system relies on radar
guidance coming from associated F-Band Continuous Wave pulse – Doppler Clam
Shell, tower - mounted Big Bird or 3-D Tombstone long – range surveillance / EW
radars, with I/J-Band Flap Lid phased array radar used for target tracking. As with the
S-300V and S-400 this system can be linked to the A-50 so that target information can
be obtained without the risk of the system being exposed to attack from anti –
radiation missiles.
Short – range Man portable SAM‟s.
The man portable SAM‟s MANPADS of the SA-14/-15/-16 have imaging infra-red
and ultraviolet seekers operating at both ends of the visible spectrum, and the SA-16
Gimlet simply ignores flares altogether, and over 1 million have been sold to date
most to threat nations. The best counter to these is to fly above 20,000ft to avoid them
and anti aircraft artillery (triple –A) fire, and rely on older precision guided munitions
PGM‟s or newer JDAM‟s and J-series weapons in a fast pass lob well away from the
target area. But such tactics place non-stealthy aircraft within the prospective shooting
range of the SA-11 SA-10, and S-300V and S-400 SAM‟s detailed above. (Ref 3)
Against these new threats the time-honoured defence tricks such as terrain masking
and defensive high – g breaks simply would not work against S-300 class SAM‟s
even if the pilot was G-LOC immune: these weapons use gas-dynamic control
systems for super-manoeuvrability during the closing stages of interception, which
allow a 20g acceleration in under 0.025 seconds in response to such tactics, as well as
the ability to engage targets down to treetop level (assuming adequate line-of-sight for
a tower – mounted tracking radar). Commenting on this surface to air threat
environment one pilot put it, “You have to hide from them until you can kill them”.
New fourth generation stealth fighters such as the F-35 Joint Strike Fighter and F/A-
22 Raptor, according to published texts relying on golf ball and marble Radar Cross
Section‟s (RCS) respectively, aim to do precisely that – hide, not evade, and right
over the enemy‟s noses. General John P. Jumper stated “High – altitude attacks
at up-to 40,000ft would help them avoid the operational envelope of infra-red
seeking SAM‟s and triple A, and speeds of Mach 1.3 – to – Mach 1.5 would
reduce the effective envelope of the large radar guided SAM‟s like the S-300V‟s,
S-400‟s, and SA-11‟s” (Ref 2 and 3), which is the most important factor in
determining the mission profile of the FB-24 AIA, and hence the aircraft itself.
2.3.3 Stealth requirements.
This is the ability of an aircraft to attack its target with the maximum amount of
surprise by denying long - range detection and is vital to the FB-24‟s interdiction and
strike missions. This is achieved by reducing the aircrafts visual, radar and infrared
signal strengths of which a detailed treatment of these is given in Appendices B, and
an overview is presented here.
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1 Visual: - Elimination of smoke trail with and without afterburner. Camouflage the
aircraft by painting it in colours predominating in its mission environment, and reduce
the size, number and visibility of identification markings.
2 Radar: - Reduce radar cross – section (RCS) by: avoiding surfaces at right angles to
each other (to limit the number of corner reflections), designing for a minimum of
radar spikes (by minimising the number of airframe angles), shielding the engine
compressor face, using radar absorbent materials, carrying stores internally, treating
the cockpit canopy and radar cavity, minimising energy emissions from the aircraft‟s
own sensors, providing comprehensive electronic countermeasures and decoys.
3 Infra – red: - Less use of afterburner, with its enormous IR signal, since the very
short afterburners now employed have increased the angular detection range. Shield
and / or cool engine exhaust, and where appropriate use tuned decoys which match the
emission spectra of the aircraft.
As is shown below in the rest of this thesis these requirements have a profound effect
on the design of the FB-24 both externally and internally, as well as systems selection.
3.0 FB-24 and A-24 configuration studies and design selection.
3.1 Initial Conceptual Design Studies.
In order to meet the Advanced Interdiction Aircrafts basic requirements three options
were considered (as per Reference4:- Page-7: Aircraft Conceptual Design Synthesis:
by Howe. D: Published by Professional Engineering Publishing Ltd 2000) and are
detailed below: -
1. Adaptation or a special light version of the existing F-35C by removing all
carrier born equipment and structural requirements from the airframe and
adopting a new larger wing, with a modest forward fuselage extension for the
second crew member, as a low cost, low risk option retaining a high degree of
commonality with the F-35 family. However this conservative approach would
not meet the supercruise capability requirements as the: - fuselage finesse ratio:
wing plan form and sweep angle: and greater wetted area, would induce more
drag and would give a similar performance to the original F-35C which is not
supercruise capable. This in turn would reduce the effect of increasing the fuel
volume on range as large portions of the mission would require afterburner use
to meet the required Mach number for ingress and egress of the target zone.
2. A major modification or direct development of an existing type this option
involved a major redesign of the existing airframe, consisting of: - extensive
fuselage extension to increase the finesse ratio: wing sweep and planform
alignment changes: empennage changes: to reduce drag, and the removal of all
non – land based strike equipment: combined with a reduction in airframe
substructure component weight reflecting the more benign operating
environment.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
17
This was a much more radical approach which was more expensive and sharply
reduced commonality with the rest of the F-35 family in major airframe
components, although internally 40% of the substructure would be cousin parts
with the CTOL variant and the undercarriage would be identical to the F-35C.
The level of systems commonality with the CTOL variant would be
approximately 80% (are estimated to be 45% of total platform costs reference 5
page 21 Fundamentals of Fighter Design 1st Edition: Whitworth. R:
Published:- Airlife 2003), with specialised systems for the UCAV command
and nuclear strike roles. This approach had a higher probability of meeting the
Advanced Interdiction Aircraft requirements than option 1, with a high degree
of commonality in the expensive system components of the platform e.g.
offensive and defensive avionics, EHA‟s, and fibre optic cable data links, as
well as a degree of structural component commonality. However this was a
considerably more expensive and higher risk option.
3. A completely new design this option was to produce a completely new aircraft
using two YF-120 Variable Cycle Engines in a much larger airframe optimised
specifically for the AIA and FSAV missions, incorporating smart structures
and new materials and manufacturing techniques, as well as specific
missionized systems. This would have no commonality with the F-35 family
airframe sub structure, and only some systems commonality. This option could
defiantly meet the AIA and FSAV requirements being specifically designed to
do so but would be too expensive for the production run envisaged and the cost
target would result operational support and logistics problems for F-35 owners
as a new aircraft type. The inherent risks in a completely new aircraft would
also very high and the development cycle would be too long based recent on
legacy projects like the Eurofighter Typhoon, FA-22A, in a changing military
environment. Although for the industry as a whole and national security it is
prudent to maintain an indigenous capability by embarking on a completely
new aircraft design and manufacture using the latest advances in technology
every twenty to thirty years.
On balance the considered decision was taken to pursue option 2 major
modifications to the existing F-35C to develop a platform capable of meeting the
detailed AIA and FSAV requirements within the cost limits and within the timescales,
required for this platform.
To map the customer requirements and determine the level of importance of
individual elements of the overall requirements to the customer and the engineering
solutions available the Quality Function Deployment method was used for this
project. This methodology incorporates the following: -
1. Language understood by all participants:
2. Cross functional cooperation:
3. Focused technology development:
4. Cost / benefit analysis.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
18
The key benefits of QFD are as follows: -
1. Reduction in engineering change:
2. Shorter design cycles:
3. Lower start up costs:
4. Systematic documentation of engineering knowledge:
5. Competitive pricing:
6. A more satisfied customer.
AERO 481 QFD Process – House of Quality
1. Customer needs (whats)
2. Customer priorities
3. Technical solutions (hows)
4. Relationship matrix
5. Technical priorities
6. Target values
7. Correlation matrix
Design Feature
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Figure 10: - The QFD Process illustrating the methodology of House of Quality
construction. (Slide source Reference 1: - USAF Academy lecture AERO 481,
incorporated within AeroDYNAMIC Version 3 software: by: - Brandt. S. A,
Stiles. R. J, Bertin. J. J, Whitford. R.: AIAA Education Series: Pub 2005: ISBN 1-
56347-689-4).
For this project a House of Quality was produced to determine the customer priorities
against each customer need and technical solution and is shown in figure 11 below.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
19
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AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
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3.1.1 FB – 24 and F-35C common features.
The following common features from F-35C were incorporated in to the
configurations modelled are described below:
Primarily the FB – 24 will be a single engine aircraft which will carry all
weapons and fuel internally, it will use the same undercarriage, avionics, and
systems as the current F–35C, but the developed aircraft will carry an internal
M61A1 20 mm Cannon for self protection. Also for all FB-24 missions will
use the ASRAAM as the self protection against air threats instead of the F-35C
AIM-120C advanced medium range air to air missile, because the FB-24 is not
intended to instigate air to air combat relying primarily on stealth for
protection.
The Diverterless Supersonic Inlet DSI developed for the F-35 will be carried
over to the FB-24 AIA, because the DSI effectively eliminates one of the major
RCS contributors in the frontal cone of modern high speed combat aircraft
namely the diverter system as outlined in Appendices C.
The use of foam filled fuel tanks employed on many current fighters including
the F/A-18E/F and proposed for the F-35 will also be employed for the FB-24
AIA rather than basic self – sealing tanks called up in the requirements reduces
the hydrodynamic ram effects caused by the impact and penetration of a
missile fragment through the fuel tanks
The GE F-136 / LOAN combination which is a low cost, light weight means of
achieving signature control in both RCS and IR spectrums, while providing
significant improvements in reliability, maintainability, and supportability,
compared to previous production nozzles. Reduction in RCS and advanced
material technologies allow axisymmetric nozzles to achieve signature levels
previously possible only with two dimensional (2-D) vectoring nozzles.
Advantages of the axisymmetric designs, attributable to inherent structural
simplicity and efficiency, have been employed to achieve substantial weight
and cost reductions compared to 2-D designs. In addition to geometrical
shaping and special materials for signature control, the LOAN also
incorporates an ejector that enhances nozzle cooling.
Durability has been significantly improved, and maintenance – friendlily features
reduce the time required to change components by as much as 80%. These features
strongly influence the selection of this combination for the FB-24 AIA. (Reference 6:-
Lockheed Martin F-35 Joint Strike Fighter: The Universal Fighter: by Harkins H:
Published by Centurion Publishing UK 2004)
The internal weapons carried for primary use will be two of either of the
following: - GBU-31 JDAM PIP Mk – 84 warheads: AGM-154 JSOW glide
bomb: Saber 15-Kiloton ALOSNW (not shown but of similar dimensions to
JSOW): in combination with two AIM-132 ASRAAM missiles requiring a rail
launch trapeze, in place of the AIM-120C AMRAAM‟s of the F-35C.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
21
Figure 12: - Internal weapons carriage options of existing F-35 family retained in part
for the AIAA FSAV concepts. Source: - Reference 2:- Code One Lockheed Martin
Aeronautics publication.
This design study will use the government furnished equipment GFE listed below in
table 3 for weight and airframe OML sizing. Nine configurations were studied from
which a final basepoint configuration was selected the evaluation of these
configurations and how they met the merit measurement criteria is detailed below.
3.2 FB – 24 and A-24 Configuration concepts.
The key design drivers identified from the QFD house of quality analysis (figure 11)
influencing the OML of the new build (NB) AIA concepts considered under the
major modification approach, and the methods of their resolution are detailed
below: -
1. Long range at high speed, to be achieved through supercruise performance:
2. Stealth as defined in section 2 and detailed in appendices B, to be achieved by
planform alignment, and engine emission shielding:
3. Payload capability, to be achieved by stretched weapons bays:
4. Additional crew position, to be achieved by fuselage stretch:
5. Lower cost than a completely new design, to be achieved by retaining a degree
of commonality in systems and airframe with both the F-35 family, and the
F/A-22A.
Design drivers 1 through 5 have a direct impact on the OML configuration of the FB-
24 and are covered below in the initial configuration stage, and some elements of
design driver 5 are incorporated in resolving the first 4 design drivers.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
22
Proposed resolution of Design driver 1:- At supersonic speeds aircraft drag is
composed of: - (a) Skin friction drag: (b) Wave drag due to thickness (or volume drag
also known as zero lift wave drag): and (c) Drag due to lift (a combination of vortex
drag and drag due to lift). The FB–24 configurations attempted to address the wave
drag of the current F-35C configuration by increasing wing sweep angle which raises
Mcrit and reduces wave drag, (but degrades CL and CLmax, and increases induced drag)
and increasing the ratio of the aircrafts maximum cross – sectional area to overall
length which has a larger effect on the configurations supersonic wave drag than dose
the wing sweep or the smoothness parameter EWD also the wing span and sweep angle
was engineered to contain the wing span inside the shock wave cone generated by the
aircraft‟s nose, further reducing the wave drag because Mach number inside the cone
is lower than the freestream Mach number (M), and shock waves are weaker than
they would be if the wing were exposed to M. The fuselage had area rule
configuration to reduce drag at transonic speeds and increase transonic acceleration,
with a blended wing, fuselage interface.
The wing design for Advanced Interdiction Aircraft gives rise to multiple design
points as with all tactical military aircraft (Reference 7: Bertin. J. J: Aerodynamics for
Engineers 4th
Edition: USAFA: Published by Prentice Hall 2002) and this multiple
design point requirement was a major driver, with the aerodynamic requirements for
each point often in conflict with each other. For example the need for rapid
acceleration to supersonic flight and effective supersonic cruise requires a thin wing
section with relatively high sweep and with camber that is designed to trim out the
moments resulting from the aft aerodynamic centre movement at supersonic flight.
However, these requirements conflict with those of efficient transonic manoeuvre
which is better performed with thicker wing sections designed with a camber for high
CL operation and a high aspect ratio planform to provide a good transonic drag polar.
The final design is therefore a compromise solution of a variable camber wing was
considered in preference to variable sweep wings, as on Tornado and F-111 which
have prohibitive volume, weight, complexity, and stealth issues and would be
completely inappropriate for the AIA. A typical performance spectrum for tactical
fighters corresponding to typical mission requirements is shown in figure 13 which
presents a map of lift coefficient versus Mach number. The low Mach end of the
spectrum throughout the CL range is typical of takeoff and landing for the
configuration. The subsonic cruise and supersonic cruise portions are noted in the
moderate lift range. Acceleration to high supersonic speed occurs at low lift
coefficients Sustained manoeuvre takes place in the CL range of less than one for most
fighter configurations, and above this lift coefficient, the aircraft is in the
instantaneous manoeuvre regime. Drag rise occurs depending on the wing geometry,
in the range of Mach 0.8 to 1.2. The particular flow conditions that correspond to the
flow map are shown in figure 14.
At the cruise and acceleration points the primary consideration is attached flow, and
the design objective is to maintain attached flow for maximum efficiency. At the
higher CL values corresponding to instantaneous manoeuvre, separated flow becomes
the dominant feature. Current designs take advantage of the separated flow by forming
vortex flows in this range. Intermediate CL values corresponding to sustained
manoeuvre are usually a mixture of separated and attached flows.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
23
Figure 13:- Typical performance map for a tactical fighter aircraft, (Source
Reference 7 Page 514).
Figure 14:- General flow regimes encountered for tactical fighters, (Source
Reference 7 Page 515).
Consequently, if the aircraft is designed with camber to minimise separation in the
manoeuvre regime, the configuration will have camber drag and may have camber
drag and may have lower surface separation, which increases drag at the low CL values
needed for acceleration.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
24
Therefore the overall wing design for the AIA needed to be a compromise to achieve
optimum flow efficiency through all of the design points of the proposed mission, and
the design was influenced to the greatest extent by the constraints imposed by
supersonic and high subsonic cruise.
For the wing design the ideas of Bradley and Bertin (Reference 7 pages 516 - 517)
were employed as follows, Bradley states in referring to the combination of variable
leading edge flaps with strake configurations: - “We should mention a new concept of
controlled vortex flow for designing military aircraft having emphasis on supersonic
configurations. Wing planforms for supersonic cruise have higher leading edge sweep
and generally lower aspect ratios: these planforms develop vortex flows at relatively
low angles of attack. As a result, the transonic drag characteristics are lacking in the
manoeuvre regime since drag polars generally reflect very little leading edge suction
recovery. Recently, wings of this type have been designed to take advantage of
separated vortex flows rather than to try to maintain an attached flow to higher CL
values.” Small leading edge closeouts incorporated in the YF-22, and both X-35
demonstrator aircraft and the fore body chine of the YF-23 had the potential to create a
vortex flow field, although not anywhere as large as those of the F-16.
Bertin in considering variable – twist, variable – camber wings states: - “Survivability
and mission effectiveness of a supersonic – cruise military aircraft require relatively
high lift / drag ratios while retaining adequate manoeuvrability. The performance of a
moderate – aspect ratio, thin swept wing is significantly degraded at high lift
coefficients at high subsonic Mach numbers because of shock – induced boundary –
layer separation and, at higher angles of attack, because of leading – edge separation
and wing stall. The resulting degradation in handling qualities significantly reduces the
combat effectiveness of such airplanes.” The techniques used to counter leading –
edge stall, are as follows: - leading edge flaps, slats, and boundary – layer control by
suction or by blowing (used on the Blackburn Buccaneer, and Lockheed F-104), which
have been effectively employed in combination with trailing edge flaps for increasing
the maximum usable lift coefficient for both the low – speed landing, and high
subsonic phases of a tactical strike aircraft mission. For supersonic drag and stealth
considerations, boundary layer control was discounted and full span constant hinge
line leading edge flaps were employed in combination with trailing edge flaps for
landing and high subsonic phases of the AIA mission this is common to the F/A-22A
and the F-35 family. Bertin recommends low – thickness – ratio wings incorporating
variable camber and twist for high performance fighters with fixed wing planform,
because the camber can be reduced or reflexed for the supersonic (dash to target)
phase of the mission and increased to provide the high lift coefficients required for
transonic and subsonic manoeuvrability, ideally a smart structure mission adaptive
wing would have been the logical selection, however the research and development is
still at a relatively fundamental stage and this technology was not deemed mature
enough for the FB-24. A test program was conducted by NASA in 1977 (Reference8: -
Ferris, J.C.: “Wind-Tunnel Investigations of Variable Camber and Twist Wing”:
TND-8457: NASA: Aug 1977) to determine the effect of variable camber and variable
twist on the aerodynamic characteristics of a low – thickness – ratio wing. The basic
wing was planar with a NACA 65A005 airfoil at the root and a NACA 65A004 airfoil
at the tip (in effect there was no camber or twist in the basic wing). The section
camber was varied using four leading - edge segments and four trailing - edge
segments, all with span wise, individual hinge lines as shown in figure 15.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
25
Figure 15:- The demonstration wing configuration used to show the effects
variable twist and camber in the 1977 NASA experiments (Source Reference 7
Page 518).
Variable twist was obtained by having greater sweep angels in both the leading – edge
and trailing – edge segments than in the basic planforms leading and trailing edges, as
shown in figure 15. All leading – edge segments were parallel to each other as were all
of the trailing edge segments. Camber and twist could be applied to the wing as shown
in figure 15. The test program (Reference 8) demonstrated that deploying the trailing
edge segments near the wing root crated a cambered section with an effective chord of
increased incidence, where as deploying the leading edge segments near the wing tip
creates a cambered section whose local incidence is decreased. Hence the modified
wing could have an effective twist of approximately 80 of washout. The program
further demonstrated that the use of leading edge camber lowers the drag substantially
up to a lift coefficient of 0.4, and increases the lift / drag ratio over a Mach number
range 0.6 to 0.9, and at lift coefficients above 0.5, the combination of twist and camber
achieved combining both the leading - edge and trailing - edge segments was effective
in reducing drag. Trailing - edge camber was also shown to cause very large
increments in CL with substantial negative shifts in the pitching moment coefficients.
Examples of the data obtained from the NASA research is shown in figure 16, and it
can be seen that the maximum lift / drag ratio for this particular configuration at M =
0.80 is 18 and occurs when CL = 0.4.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
26
Figure 16:- Effect of twist and camber on the longitudinal aerodynamic
characteristics determined from the NASA test program (Reference 8) chart (a) is
lift-to-drag ratio and drag polar, and chart (b) is pitching moment and lift
coefficient.
The airfoil section selected for the AIA was NACA - 0006 which was a symmetric
airfoil with a maximum thickness of 6% of the chord with a sharp leading - edge.
Which although thicker than the F/A-22A airfoil section which has a customised
airfoil thickness of 3.8% optimised for transonic operation, has similar root thickness /
chord ratio of 0.06 compared to the formers 0.592 and a tip thickness / chord ratio of
0.429 (Reference 9:- Miller. J.: Lockheed Martin F/A-22 Raptor: Stealth Fighter:
Published by Midland Publishing 2005) the thickness selected for the AIA wing was
within the transonic and supersonic efficiency thickness range of 5 – to – 8% required
for a multi – role strike / fighter airfoils as stated by S. Kern (Reference 10:- Kern. S.:
Evaluation of Turbulent Models for High – Lift Military Airfoil Flowfields: AIAA96-
0057: presented at the 34th
Aerospace Science Meeting, Reno, Nevada: Jan 1996):-
“Integration of stealth requirements typically dictates sharp leading edges and
transonic and supersonic efficiency dictates thin airfoils on the order of 5-8% chord”
initially the intention was to evaluate the effects of camber and twist variation on the
three wing planforms selected in section 2 with the NACA 0006 airfoil using the Flite
3D CFD software package however the BAE JSF IPT would not to support the authors
training for this activity and this is an area of future study and is now outside the scope
of this thesis. The degrees of twist and camber selected were based on F/A-22A values
from reference 9 and MDA / BAe JAST configuration 9B data (Reference 6:-
Hawkins. H.: Lockheed Martin Joint Strike Fighter (The Universal Fighter):
Published by Centurion Publishing 2004), and F-16XL from (Reference11 / 12:-
NASA Dryden Fact Sheets – F-16XL-1 Testbed Aircraft:
www.dftc.nasa.gov/Newsroom/FactSheets/FS-051-051-DFRC.html).
As illustrated fig 15 As illustrated fig 15
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
27
To further reduce drag four different empennage configurations were investigated in
combination with the three wing planforms which were as follows: - the current F-35
and F/A-22A Four tail layout i.e. Vertical tails canted outward 250 to 27
0 from the
aircraft z, x, plane matching the slope angle of the fuselage side walls with fixed
torsion box and leading edges and attached rudders, and Horizontal tails which were
all moving with leading edge sweeps matching that of the wing, attached either by a
root spigot or integral hinge spar: or the YF-23 all moving Ruddervators canted out
500 from the aircraft z, x, plane matching a localised rear fuselage side wall slope, and
attached with a root spigot. A canard configuration considered without the vertical
tails observed in the Lockheed Martin JAST studies (V- vertical tails) and Typhoon /
Rafale etc (single vertical), similar to the X-36 was also considered. Finally a
completely tailless and canard less configuration with no empennage group at all as
observed on the X-45 and X-47 UCAV‟s.
The first conservative four tail empennage option was readily applicable to large
conventional missionized wing and the cranked arrow head MDA / BAe JAST wing
but not to the double delta X-16 wing and would have had a similar drag to
contribution to the former wing configuration as the F-35C empennage has to the F-
35C, and would be sized proportional to the wing.
The second more radical Ruddervator option was applicable to all three wing
configurations and offered a real drag reduction possibility, on a developed F-35C /
FB-24 configuration. Also this empennage configuration was flight proven in
supersonic, supercruise, and low speed flight regimes by the Northrop YF-23 during
the Advance Tactical Fighter demonstration and validation phase test program in 1991
and is shown in figure 17 below.
Figure 17:- Northrop McDonnell Douglas YF-23 showing 500 canting of the
ruddervators applied to all three of the FB-24 wing configuration. Source:-
Authors private collection.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
28
The third canard configuration considered for the FB-24 AIA, was the short coupled
arrangement with the wing planforms considered, in this arrangement the foreplane
was to be located just ahead of and just above the wing like the X-36, Rafale, and the
original LM JAST configuration pre - F-35. Carefully positioning the canard and wing
relative to each other enables their combined lift effectiveness to exceed the sum of
their individual values. This configuration had the possibility of higher agility than the
first conventional configuration and potentially a lower drag contribution. However
the size (lager than on either the Rafale or JAST configurations, because of not having
any vertical tails) and location of the canards would be detrimental to the frontal RCS
(see appendices B) of the aircraft because of the increased number of reflecting
surfaces, every time the canards are actuated.
Although the fourth completely tailless option theoretically offered the lowest drag,
longitudinal stability was thought to be inadequate due to the short moment arm of the
elevators in all but the delta design, requiring a multi axial vectored nozzle and
complicating the wing flight control surfaces. The X-45 shown in figure 18 has a short
and flattened fuselage shape of 26 feet 5 inches, with the wings of 33 feet 8inches span
located at the rear where the elevators are most effective, where as the X-47 is a
diamond shaped flying wing and incorporates additional control surfaces to elevators
and flaps, buried in the wing top and bottom surfaces. Although only the X-45
employs a yaw thrust vectoring system, neither is supersonic, and there configurations
are totally different to any considered for the FB-24, the only similar configurations to
the FB-24 for which the tailless approach has been advocated i.e. the X-44 MANTA
shown in figure 19 and the FB-22 both use vectored thrust and have two engines to
effect pitch and roll control, therefore this empennage less configuration was not
considered further in this concept design study.
Figure 18:- The X-45 tailless UCAV although not readily apparent from this view
the aircraft has a short fuselage compared to its wing span with pitch, yaw, and
roll being controlled by a total of six trailing edge control surfaces. Source: -
USAFA AeroDYNAMIC Version 3.0 CD-ROM.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
29
Figure 19:- X-44 MANTA F-22 applications project although not built or flown
wind tunnel models and CFD simulations demonstrated the requirement for
continuous vectored thrust augmented stability and to achieve this, a two engine
2-D vectored nozzle was necessary. Source: - Authors private collection.
Therefore in the final analysis only the conventional F-35 and F/A-22 and the YF-23
ruddervator empennage configurations were considered suitable for the FB-24 design
constraints and had the merit for further evaluation in this conceptual design study.
Proposed resolution of Design driver 2:- As stated in Lockheed Martin‟s Affordable
Stealth paper (Reference 13) and reinforced in references 4 through 7 and detailed in
appendices B, In order to develop a low RCS aircraft consideration must be given to
any part of the aircraft that a radar wave can reach. This impacts on the OML design
because the shape of the aircraft is the most critical factor to consider in the design
process for LO capability. A visual study of the F-35 and F/A-22A s reveals that all
hard edges i.e. the wing leading and trailing edges, control surfaces, vertical and
horizontal tail surfaces, intakes, have been aligned to a few common angles, so that the
radar returns from them all point in the same few directions. This is called planform
alignment and is carried over doors, external sensor apertures, inspection hatches, and
even formation and navigation lights, where this is not possible multiple chevrons are
used to break up the radar return. The overall result is a few relatively large but narrow
signature spikes that are difficult to detect and track.
The vertical tails and fuselage sides are tilted the side walls meeting the top surface in
a continuous chine running from the tip of the radome to the rear tip of the fuselage
this avoids a direct radar reflection from side on illumination. The entire surface is
blended smoothly at major component interfaces, enabling electrical surface currents
to flow over the aircrafts surface without interruption, in addition this also has the
effect of reducing drag. The effects of breaks for control surface interfaces are reduced
by radar absorbent seals both on the control surface and on the main plane trailing
edge. This planform alignment was employed in all of the study configurations
proposed in this thesis.
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One of the largest remaining radar reflectors is the cockpit with potential radar returns
coming from the pilots and offensive air systems officers heads and helmets, seats, and
all of the various controls and displays in the cockpit. The most effective way to
reduce the cockpit signature is to prevent the radar energy from entering the cockpit. A
metallic coating is currently applied to all USAF, USN and USMC combat aircraft
which prevents the energy from entering the cockpit, eliminating this concern. The
cockpit frame however is still a potential area of concern and on both F/A-22, F-35 the
frame front follows the planform alignment principal, and in the former aircraft so
dose the frame rear. The cockpit selected for the FB-24 was a two place cockpit based
on the cancelled F-22B training version of the F/A-22A, unlike the F-35 which is
largely driven by the STOVL pilots vision requirements, this cockpit had high sills
enabling smoother blending into the fuselage of the two piece canopy which has an aft
swept bow frame to resist birdstrike this being required because of the increased
length over that of the F/A-22A‟s single piece canopy.
Another major contributor to the aircraft signature would come from the intakes, in the
FB-24 the same intake design features of the F-35 were employed namely the
diverterless supersonic intake, which in the JSF support program was demonstrated to
be suitable for flight in the Mach 2 flight regime (see appendices C for details), and
the characteristic forward sweep to the top of the intake to allow diverted air to spill
out over the local fuselage top surface. The intake ducts are bifurcated as in the F-35
resulting in complete engine face obscuration.
The rear quadrant signature reduction for the FB-24 was achieved by employing the
LOAN nozzle however because the detailed of the nozzle is secret a generic volume
cone of appropriate dimensions was used for all models. The intension was to shield
side view of the engine bay and nozzle with the empennage and this was successfully
accomplished for both four and two piece tail configurations. The engine and nozzle
was situated further aft in relation to the empennage group than is the case with the F-
35 where location is effected by STOVL commonalty requirements.
The final external feature is the Air Data System (which caused an extensive research
work on the worlds first stealth combat aircraft the F-117A), for which the FB-24
adopts the F/A-22A low observable pneumatic air data system (PADS), which
consisted of two small facetted fuselage mounted air data probes, one on either side of
the fuselage located aft of the radome, and four flush - mounted static ports (two on
each side of the fuselage) which were also located aft of the radome above and below
the forward fuselage chine, eliminating the multitude of probes seen on non stealthy
aircraft such as Tornado.
These OML features combined with the structural layout and materials features
covered in section 4 were deemed to resolve design driver 2 for the FB-24 conceptual
design study. (Reference 13: - Lockheed Martin’s Affordable Stealth: paper by Haisty.
B. S.: Published by Lockheed Martin Aeronautics Washington D.C.: November 15th
2000: for National Press Club).
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
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Proposed resolution of Design driver 3:- The weapons bays were positioned so that
its centre of gravity was as close as practical to the whole aircraft centre of gravity. A
single weapons bay was ruled out for two main reasons:- firstly with a single engine
the duct from the bifurcated intakes ran through the aircrafts centre line therefore a
single weapons bay would have to be positioned below this and result in a dramatic
increase in fuselage cross section impacting adversely on the desired finesse ratio:
secondly the air turbulence generated when a single deep weapons bay was opened at
supersonic speeds, resulting in acoustic damage to the internal structure, and adverse
handling. Therefore two weapons bays one on each side of the fuselage, each with
inboard and outboard doors as on the F-35 family was considered to be the best
solution for the FB-24 however the length was increased from approximately 4.5m to
7.5m to accommodate both a GBU-31(v) 3/B (PIP) and a ASRAMM in tandem, in
each by, with the former being released by cold gas ejectors and the latter being
launched from the integrated outboard door launch rail. The rear bulkhead of each
weapons bay was sloped rearwards to improve airflow when the bays are open and to
give better nose down pitching characteristics should the aircraft force land in water
and the weapons bay doors collapse.
The extended bays also enabled integration of the ALOSNW stores for the FSAV
requirement, resolving design driver 3. The impact on OML was a high wing position
as seen on both the F-35 and F/A-22A which had implications for the undercarriage
and wing internal structure.
Proposed resolution of Design driver 4:- The size of the crew compartment for
initial sizing was evaluated to accommodate a volume of 7.5m3 to permit two ejection
ACES II seated crew, one HUD, and one console this resulted in length of 4.35m
width of 1m and a height of 1.2m, from the volumes given in table 2 below for GFE,
this resulted in an increase over the initial size of the F-35 single crew compartment
shown in figure 20, which was estimated at 0.7m width, 1.1m high, and 1.45m long
for this study. The ejection seat pitch and slope angle shown in figure 21
accommodated both crew members with full TLSS tactical life support system suits
(Nuclear / Biological / Chemical and g suits) and was identical to that of the F/A-22A,
and proposed for the F-22B and was designed to give a 99 percentile crew a clear
ejection trajectory, good situational awareness (-150 over the nose, and -45
0 over side
pilot vision angles), access to all interactive control systems MFD‟s, side-stick flight
controller, etc, and good ride quality. This extension in the forward fuselage resolved
design driver 4 and was common to all five configurations evaluated. Like the F/A-
22A noted above in design driver 2 the canopy transparency, of the FB-24 was to be a
3/4 inch thick fusion bonded and drape forged Sierracin Sylmar Corporation unit of
tinted monolithic polycarbonate, 194 inches in length, 45 inches wide, and 27 inches
high, with a rear sloping canopy bow for support in advent of bird strike, this is similar
to that proposed for the F-22B trainer shown in figure 22. The canopy shares the rear
hinge opening method of the F/A-22A, and not the forward hinge currently selected
for the F-35 family, this is due to the weight and size of the canopy, which in
emergency ejection would be separated to clear the aircraft with the aid of the airflow
over the forward fuselage and not have to act against it as is the case with the F-35.
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Figure 20:- F-35A first CTOL SDD aircraft front fuselage nearing completion at
Lockheed Martin Georgia facility. Source Lockheed Martin F-35 general release.
Figure 21:- The ACES (Advanced Concept Ejection Seat) II which has been
proven in the F/A-22A, and therefore would only require minimal compatibility
testing for use in the FB-24. Installation was at 770 tilt back angle to the cockpit
floor. Source: - Reference 14.
770
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G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
33
Figure 22:- The proposed but cancelled training variant of the F/A-22A namely
the F-22B, which could form the base for the Advanced Regional Bomber, and
serves as the basis of the FB-24 canopy design. Source: - Reference 14:- F-22
Raptor (America’s Next Lethal War Machine) by Pace. Steve: part of the Walter J.
Boyne Military Aircraft Series Volume 1: Published by McGraw-Hill 1999.
Proposed resolution of Design driver 5:- This study proposes the use of the F-35C
undercarriage, actuators both: - EHA actuators for the primary flight controls and
rotary actuators for the leading edge flaps, avionics and mission systems (with the
exception of the current radar) to reduce risk and maintain commonality, because the
SDD F-35 programme will validate these systems ahead of the AIA / FB-24 concept
demonstrator PDR (2010), enabling the systems experience to be incorporated at this
early stage. This study also proposes the use of the ACES II proven F/A-22A ejector
seats detailed above.
The F-35C fuel system with a hose and drogue flight refuelling system, instead of the
current USAF system of flying boom refuelling (which is currently under review)
reducing the fuel system complexity. The proposed engines are based on flight proven
and near flight designs namely the F-136 and the YF-120 both offer lower risk then a
totally new engine design and development cycle.
The current design would use the same signature reduction technologies employed by
the F/A-22A, and F-35 aircraft explored in appendices B, but would be readily
adaptable to concepts after these aircraft enter service and mature. Also innovations in
manufacturing which were not at a high state of maturity for these aircraft could be
employed on the FB-24 and these will be covered in section four of this thesis.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
34
Table 2:- Government Furnished Equipment.
Item Volume, ft3 Weight, lb Cost, K$
Avionics
• Base Suite
- ICNIA1 3.0 100 200
- 3 x MFD‟s 1.5 20 60
- Head-Up Display 1.6 35 20
- Data bus 0.5 10 10
• ECM Equipment
- INEWS2 3.0 100 500
Flight and Propulsion Control System
Vehicle Management System 1.0 50 200
Fire Control Systems
• IRSTS3 2.0 50 300
• Active Array Radar 6.0 450 1000
• LANTIRN Navigation System 3.0 350 200
• LANTIRN Targeting System 2.5 350 100
• HARM Targeting System 3.0 150 100
Systems and Equipment
• Electrical System (one engine) 1.0 220 40
• Auxiliary Power Unit (APU) 2.0 100 50
• Ejection Seat 8.0 160 100
• OBOGS5 1.0 35 10
• OBIGGS6 1.0 35 10
Air-to-Air Weapons ASRAAM Missile Launch weight: 192 lb Length: 9.51 ft Max span: 1.6 ft Body diameter: 0.54 ft Launcher rail weight: 50 lb Launcher rail length: 9.5 ft M61A1 20 mm Cannon Cannon weight: 275 lb Length: 74 in Max diameter: 10 in Ammunition feed system (500 rounds) weight: 300 lb Ammunition drum length: 25 in Diameter: 25 in Ammunition (20 mm) 0.58 each Returned casings 0.26 each
Air-to-Ground Weapons JDAM GBU-31PIP Release weight: 2,115lb Length: 12.38ft Max span 2.1ft Cost $20,000 ALOSNW (Projected weapon for FSAV)
7
Release weight: 2,500lb Length: 14ft Max span: 1.5ft Warhead: Evolved W69 (200kts)
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
35
1
Integrated Communication, Navigation, and Identification Avionics 2
Integrated Electronic Warfare Systems 3
Infrared Search and Track System with laser ranging 4
Synthetic Aperture Radar 5
Onboard Oxygen Generation System 6
On-Board Inert Gas Gen. System 7 The ALOSNW is system developed by the author and not a current weapon.
(Reference 1(b): - Aero Engr – 482: Dr Brandt: USAF Academy 1999: reference 15:-
RAF Equipment fact sheet 13 ASRAAM)
For this study the F-110-132 engine is used for integration in the structural layout
because the physical dimensions and weight figures are in the public domain and were
comparable to both the F-136 and YF-120 VCE engines proposed for the FB – 24.
Figure 23 provides a reference of current fighter and strike aircraft loadings which are
expressed as follows:- Thrust loadings are expressed as a ratio of static thrust against
weight and are an indicator of specific excess power during hard manoeuvring, and
also acceleration and initial rate of climb (thrust loadings greater than 1:1 are common
in modern fighters): Wing loading, given in lb/ft2 (kg/m
2) are an indicator of
instantaneous turning ability; combined with thrust loading these indicate sustained
turning ability.
Figure 23: - Comparison of wing and thrust loadings for modern strike aircraft.
STRIK AIRCRAFT WING AND THRUST LOADINGS.
0
0.2
0.4
0.6
0.8
1
1.2
1.4
0 50 100 150 200 250
W/S (lb/ft^2)
T/W
(lb
/lb
)
Series1
Tornado IDS
F-35B
F-35A
F-35C
Jaguar GR.1AF-117A
EF 2000
Tornado F2
F-16C
F/A-18EF-14D
Mirage 2000-5
JAS 39A
Rafale
F-15E
F/A-22
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G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
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3.2.1 Configuration design challenges:
Before detailing the conceptual designs studied for the FB-24 manned and A-24
UCAV derivative it is instructive to highlight the design challenges faced in designing
a supersonic and supercruise capable aircraft, which are very different from those of a
subsonic aircraft, as this will show the reasoning behind the design decisions taken in
the subsequent sections of this thesis. The supersonic military aircraft is essentially
two aircraft in one, optimised for the supersonic mission but also requiring
satisfactory subsonic characteristics to allow safe and reasonable handling at low
speeds. Because the physics of supersonic flow are very different from the physics of
subsonic flow every good supersonic aircraft design is of necessity a careful blend
and compromise between good supersonic and good subsonic characteristics.
The major the major design challenges pertaining to the design of the FB-24 and A-24
were as follows:-
1. Shock waves, which occur at transonic and supersonic speeds, cause a large
increase in drag – supersonic wave drag. Supersonic aircraft need to be
designed with slender bodies (high finesse ratio) and thin wings with sharp
leading edges to reduce wave drag. These same features are not good for
subsonic speeds.
2. Careful consideration must be given as to where the shock waves occur on the
aircraft, and where they impinge on the aircrafts surface. Shock impingements
can cause flow separation and local hot spots of intense aerodynamic heating,
with resulting structural implications.
3. The centre of pressure (therefore the aerodynamic centre ac) for the aircraft
shifts dramatically aft when the aircraft accelerates from subsonic to
supersonic flight. This creates a major design challenge for maintaining the
stability and control characteristics of the aircraft.
4. Airframe – propulsion integration was a design challenge which the author
feels has been resolved by adoption of the current F-35 family intake
configuration which has been flown at Mach 2, but this requires further
aerodynamic analysis which is beyond the scope of this thesis.
5. Aerodynamic heating, which is usually negligible for subsonic aircraft design
became an important design consideration for the FB-24 and A-24, as
aerodynamic heating increases approximately as the cube of the velocity. This
not only has material selection and structural implications, but also has
implications on the infra - red signature of the aircraft, especially in the case of
the wing leading edges and intake lips, and with a desired supercruise speed of
Mach 1.6 this more of a challenge for the FB-24 than on the F/A-22A.
These were the major aerodynamic design challenges faced in the design of the
common airframe for the FB-24 and the A-24 aircraft, which are common to most
other supersonic military aircraft depending on the duration of the time spent in
supersonic flight regime. Attempts to meet these challenges are detailed below in the
following sections.
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3.2.2 FB-24 and A-24 Fuselage configuration design selection:
The common fuselage design was intended to achieve the best streamlined shape with
the minimum surface area and highest finesse ratio capable of supercruise
performance, and lowest weight which was, capable of satisfying the internal volume,
and commonality requirements with the F-35C of the FB-24. Both the drag and
weight of the fuselage are significantly influenced by the surface area, and the grater
the surface area the higher the values of drag and weight become. General
aerodynamic considerations are given in references 4 and 7 which have been adopted
where there was no conflict with the low signature requirements of the aircraft. The
following general approach was adopted: - steps and gaps were avoided: gradual
changes in cross-section: shaping and treatment of all unavoidable protuberances e.g.
blade antennas, air data probes: fairing in the canopy as described above: the
positioning of the maximum cross-section to minimise wave drag due to
compressibility: use of area rule to eliminate the formation of compression waves
which would otherwise originate at the wing root to turn the flow parallel to a straight
fuselage, as Mach number increases (detailed below): and the smooth reduction of the
tail configuration from maximum cross-section to nominally zero, with consideration
being given for the clearance angle between the main undercarriage and the tail of
120.
For supersonic and supercruise capable aircraft, a major consideration in the design is
the distribution of the total volume along the length of the aircraft of with the fuselage
normally being the major contributor to this volume. The distribution was analysed
for both the F-35C and YF-22 using published data, and analysis models
commensurate with conceptual design as detailed below, to enable a target
distribution for the FB-24 to be determined. Ideally the total volume distribution
along the length should be smooth and, somewhat simplistically, close to a sinusoidal
shape in order to minimise volume wave drag, similar to that for the YF-22 shown
below in figure 24.
Figure 24: - YF-22 Cross-sectional area plot using the Jet 306, showing the
overall sinusoidal shape of the volume distribution over the aircraft length,
measurements from 1/72 scale drawings:- Reference 16:- The ATF Contenders:
YF-22 & YF-23 Air Superiority into the 21st Century 3
rd edition: by Sun Andy:
Published by: - Concord Publications Ltd Hong Kong 1991.
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The overall length to effective diameter ratio or finesse discussed in section 3.1.2, is
also of fundamental importance, the higher it is the lower the wave drag, for this
initial work measurements were made from public data of the maximum cross-
sectional area (reference 4) including the intake area from which an equivalent
diameter was calculated for the YF-22 and the F-35C because neither aircraft has a
circular cross-section, and the published respective fuselage lengths were divided by
their respective equivalent diameters as follows: -
YF-22 overall fuselage length = 19.60m (64ft 2in) / core maximum fuselage
diameter = 2.64m (8ft 8in) gave a finesse ratio = 7.42:
F-35C overall fuselage length = 15.69m (51ft 5in) / core maximum fuselage
diameter = 2.41m (7ft 11in) gave a finesse ratio = 6.59.
Therefore for the FB-24 to be supercruise capable a finesse ratio (in combination with
major drag reduction measures covered above) closer to 7.42, of the YF-22 instead of
the 6.59 for the F-35C was required and this could have been approached by retaining
the core fuselage maximum diameter of 2.41m (7ft.11in) and increasing the fuselage
length to 18.89m (61ft.11 1/2in), this gave the finesse ratio of 7.83 from: -
FB-24 overall fuselage length = 18.89m (61ft 11 1/2in) / core maximum fuselage
diameter = 2.41m (7ft 11in) gave a finesse ratio = 7.83.
Although basic assessment indicated that this overall length would could be reduced
to 18.00m and would still obtain a finesse ratio of 7.47 which was closer to that
calculated from the available data for the YF-22 from: -
FB-24 overall fuselage length = 18.00m (59ft) / core maximum fuselage diameter
= 2.41m (7ft 11in) gave a finesse ratio = 7.47.
The fuselage length could be increased further to 18.29m (60ft) without a high
corresponding increases in weight but would require a modest increases in empennage
control surface area over that required for a length increase to 18.00 (59ft) to control
pitch stability. Maintaining the original maximum cross-sectional area while
increasing the fuselage length to 3.05m (10ft) enabled retention of the original F-35C
weapons bay width based on the CATIA V5 F-35C 230-5 OML model, the result of
increasing the fuselage length to 18.29m (60ft) was calculated to give a finesse ratio
of 7.58 from: -
FB-24 overall fuselage length = 18.29m (60ft) / core maximum fuselage diameter
= 2.41 (7ft 11in) gave a finesse ratio = 7.58.
This 3.05m (10ft) extension was considered the optimum fuselage growth the FB-24 /
A-24 study, theoretically for a fixed internal volume subsonic drag is minimised by a
finesse ratio of 3.0, while supersonic drag is minimised by a finesse ratio of about 14
with most aircraft falling between these values. Ideally the local cross-sectional area
of the fuselage should be matched to that of the other volume contributions from the
wing and empennage, to give a smooth overall volume distribution.
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Although this could not have been achieved precisely until the overall conceptual
layout of the aircraft OML had been determined consideration was given to this
requirement from the start of the fuselage design.
The wing – fuselage junction is another critical area requiring careful design to reduce
drag, and decisions taken at this initial layout stage had a major impact on the rest of
the design. The shoulder wing position shown in figure 26 was adopted with the wing
above 65% of the fuselage side wall, this is similar to the F-35C and slightly lower on
the fuselage than the F/A-22A, but unlike the F-35 family the wing was not a single
unit and was joined at the spar ends to the fuselage frames as with the F/A-22A, F-16,
etc, this enable the frames to be single piece and simplified the integration of the
weapons bays. The shoulder wing position was considered to have the least drag, and
introduce the least problems for undercarriage design, compared with wings which
pass totally over or under the fuselage shown in figure 25, e.g. the Boeing X-32,
Harrier, (both restricted by powerplant installation), SEPECAT Jaguar, or the BAE
Systems Hawk and T-45, all requiring large fairings to ensure satisfactory airflow.
(a) High – wing F/A-22 and F-35 JSF.
(b) Mid – wing F-16 / Rafale / Gripen.
(c) Low – wing Eurofighter Typhoon.
Figure 25:- Sketches for the comparison of (a) high – wing, (b) mid – wing and
(c) low – wing configurations with frontal views from Reference 17:-
Superfighters (The next generation of combat aircraft): by Williams. Mel:
Published by: - Airtime Publishing: 2002.
Anhedral
Dihedral
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(a) High – wing position: - This position for the wing is commonly found on large
civil and military cargo transports and small commuter airliners such as the BAe 146,
and enables the fuselage to be placed lower to the ground which simplifies loading in
the case of transports. This configuration enables the whole fuselage section to be
used for cargo stowage without interruption from the wing box, passing through the
cabin this was partly why the Lockheed C-5 was selected over the Boeing 747 as the
USAF‟s heavy lift cargo aircraft.
The high – wing also has greater lateral, rolling stability. For low wing configurations
as in figure 25(c) a dihedral upwards slope is usually built into the wing to increase
lateral, rolling stability.
The reason for this is that when an aircraft rolls the lift vector tilts away from the
vertical, and the aircraft starts to sideslip in the direction of the lowered wing. Where
a dihedral has been incorporated into the wing design of the extra flow velocity
component generated by the sideslip creates an increasing lift on the lowered wing,
hence tending to restore the wings to a level equilibrium position, and this is the basis
of lateral stability for naturally stable aircraft design of low wing aircraft. High – wing
aircraft on the other hand are much more stable in this regard requiring no dihedral,
this is because the extra flow velocity component generated by the sideslip when the
aircraft rolls creates a region of higher pressure in the flow interaction region between
the fuselage side and the bottom surface of the lowered wing at the wing root. This
increased pressure under the lowered wing has the effect of rolling the wings back to
the level equilibrium position.
Indeed the high wing position can be too stable in roll and has to be countered to
improve aircraft manoeuvring as in the case of the Lockheed C-5, C141, and BAe
146, and Harrier where an anhedral or downward slope shown in figure 25(a) was
adopted to reduce over stability in roll manoeuvres.
(b) Mid – wing position: - The mid – wing position used on the Lockheed U-2 spy
plane and the F-16 jet fighter shown in figure 25(b), was originally favoured by the
author because of its low drag, due to the fact that of all three options the mid – wing
configuration has the minimum wing – body interface, and unlike the high – wing and
low – wing positions requires no fillet to decrease wing – body interference, and
neither anhedral or dihedral for stability refinement. However the mid – wing has a
major structural disadvantage namely the bending moment due to wing lift must be
carried through the fuselage, and unlike the case of high – wing, and low – wing
positions where the wing torsion box can be extended across the fuselage, the mid –
wing requires heavy ring frames attached to the leading edge, and trailing edge, and
intermediate wing torsion box spars. The more spars running the length of wing
results in a greater number of fuselage ring frames carrying the bending moment
across the fuselage in the in the case of the F-16 four wing panel root attachment fish
plates form the spar to frame interface on each side of the fuselage, where four frames
are attached to nine spars in each wing. In the case of the high aspect ratio (10.6) U-2
which had only three spars, the bending moment was distributed trough twelve wing
attachment joints (six each side of the fuselage) which mated the wings to the wing
root attachment ribs and hence to four main fuselage frames which carried the
bending moment across the fuselage.
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These heavy carry through frames add considerably to the empty weight of the
aircraft and mid - wing position was considered to be too heavy and complex for the
FB-24 and A-24 configurations.
(c) Low – wing Position: - The major advantages of the low – wing position over
both high – wing, and mid – wing positions is in reducing undercarriage physical size
and weight, and when retractable undercarriage is considered this can be retracted into
the wing torsion box, which for most aircraft is the strongest component of the
airframe structure.
Although for both the high - wing and mid – wing can employ a centreline bicycle
undercarriage for retractable units. Additionally the low – wing configuration requires
dihedral for lateral stability as shown in figure 25(c) and a fillet at the wing body
interface to minimise drag inducing aerodynamic interference, and the reasons for
filleting are covered below.
Until the early 1930‟s most mono-planes were of the high – wing configuration which
was largely the result of the aerodynamic interference at the wing to fuselage junction
which was found to be worst in low – wing configuration. Putting a circular fuselage
on top of the wing has the effect of producing a pair of rapidly diverging surfaces
which steepens an already adverse pressure gradient almost guaranteeing flow
separation and inducing drag, and reduced lift, furthermore the separated flow could
impinge on the empennage horizontal tail resulting in further stability and control
problems for the low – wing configuration (covered in detail in reference 13). It was
only through the discovery of the beneficial aerodynamic effects of mounting a fillet
at the wing body junction at the California Institute of Technology CalTech in the
USA that largely overcame these flow separation problems that enable the low – wing
configuration became widely adopted in modern aircraft designs, and what reaming
inferiority there was in the low – wing configuration could be addressed by dihedral
slopping for roll stability and was compensated for by the reduction in undercarriage
length and weight and the ground cushioning effect on landing with the Wing In
Ground WIG effect.
Figure 26:- Shoulder mounted wing position of the F-35 family adopted for the
FB-24 and A-24 airframes leaving clear weapons bay each side of the fuselage,
ferry tanks would be carried by both the FB-24 and A-24, as is the case with the
F/A-22A.
Unobstructed corner weapon bays.
Shoulder mounted wing. Shoulder mounted wing.
Under - wing clearance for
stores pylons.
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The wing setting on the fuselage at this stage of the configuration was such that the
fuselage was horizontal and the wing had a 0.250 positive angle of attack in common
with the F-35 C, this was a relatively low angle of attack at the bottom of the common
range of 00 – 4
0. The wing root to fuselage joint was a smooth continuous curvature
blend accommodating both aerodynamic and stealth requirements.
For the initial configuration the common fuselage the length was chosen as 18.59m
(61ft) and the core diameter without intakes was chosen as 4.161m (13ft 8in). The
term common fuselage in this thesis denotes that for all wing and empennage
configurations the forward fuselage was identical, and for the trapezoidal, and
cranked arrow wing planforms either the four tail or twin tail empennage aft fuselage
configuration was used, which were created as interchangeable units, for the delta
only the twin tail aft fuselage unit was used. The centre was modified to interface with
each of the three wing planforms studied.
As stated above the common fuselage needed to reduce the peak transonic drag rise,
although currently no closed-form analytical formulas exist to predict the transonic
and even computational fluid dynamics which has been applied the computation of
transonic flows for more than 25 years, dose not always give the right answer,
principally due to uncertainties in the calculation of shock induced separated flow.
There are however two principle design features have been developed reduce the drag
rise itself or delay its effect, namely: transonic area-rule and supercritical aerofoil
(covered in the next section) it is worth using the example given in Dr Anderson‟s
work: - Aircraft Performance and Design (Reference 18: - Aircraft Performance and
Design: by Dr Anderson. D. John. Jr: University of Maryland: Published by: -
McGraw-Hill: International Addition 1999) as a demonstration of the benefits of
transonic area-ruling. Before the start of the 1950‟s designers believed that the way
forward in obtaining high Mach number performance lay in fuselage OML‟s based on
rifle bullets like that of the X-1, and Miles M-52, and they did not realise that kinks in
the cross-sectional area distribution where the wing was added to the fuselage caused
a large transonic drag rise. However by the mid 1950‟s, experimental work conducted
by Richard Whitcomb (based on his personal intuition) who was an aerodynamicist at
NACA (now NASA) Langley Aeronautical Laboratory demonstrated that contrary to
earlier opinion the addition of the wing had a profound effect on the performance of
transonic and supersonic aircraft. These results showed that the best performance
could only be obtained where the cross-sectional area distribution was smooth, this
could be achieved in part by decreasing the cross – sectional area of the fuselage in
the wing region to compensate for the cross-sectional area increase due to the wings.
The resulting fuselage OML has been likened to the old American Coke-a-Cola bottle
of the 1950‟s in shape, and pictures of pre area ruling (YF-102 prototype) and post
area ruling (F-106) on aircraft are shown in figures 27(a) and (b). Figure 27(c) shows
a schematic representation of the effect of area in reducing the peak transonic drag
rise. The first application of area-rule was to the YF-102 when flight testing
demonstrated that the prototype with a bullet shaped fuselage could not exceed Mach
1, this resulted in a major redesign effort which incorporated Whitcomb‟s area rule to
reduce transonic drag. The modified YF-102A achieved supersonic performance in its
first flight providing the USAF with its first operational supersonic delta wing fighter
subsequently this technique was adopted on all supersonic military aircraft.
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Figure 27(a) Non area ruled F-102. Figure 27(b) Area ruled F-106
Figure 27(c) Schematic of the drag – rise properties of an area-ruled and non-
area-ruled aircraft.
CD
1.0 M 0.0
Without area rule
With area rule
Wasp Waist
Area Distribution. Amax = 41 ft2Area Distribution. Amax = 45 ft2
Note: Both aircraft have the same internal volume
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Another feature of the fuselage which had both aerodynamic and signature reduction
benefits were the chines which ran from the radome to the end of the aft fuselage, and
which decreased the travel of the neutral point as M is increased which was
important for static longitudinal stability where the neutral point must be located
behind the aircrafts centre of gravity. The normalised distance between the centre of
gravity (cg) and the neutral point is called the static margin, this is positive when the
neutral point is behind the centre of gravity, and static longitudinal is achieved, and
hence the higher the positive static margin value the more stable the aircraft will be.
However, too much of a positive static margin is not desirable, because the aircraft
will be too stable for manoeuvring requiring a large elevator deflection to trim the
aircraft because of the large distance between the neutral point and the centre of
gravity. The net result is a trim drag which is unacceptably large. When cg is
positioned to give a good subsonic static margin it will be too large at supersonic
speeds, and conversely if the cg is located to give the correct static margin at
supersonic speed, the subsonic static margin is likely to be negative (with the neutral
point ahead of the centre of gravity), thus making the aircraft unstable in subsonic
flight. The chines offer a potential solution and the chine concept has been used on the
SR-71, YF-23, and a lesser extent on the F/A-22, and F-35 so empirical application
support this view. The additional aerodynamic benefit observed on both the SR-71,
and YF-23 at supersonic speeds was the increased directional stability (yaw stability)
imparted by the chines. For example consider fuselage of cylindrical cross – section at
a small yaw angle to the flow, this body will experience cross flow separation with a
resultant large side force, as shown in figure 28(a). However when the fuselage is of a
blended chine cross – section at the same small yaw angle to the flow no separation
occurs and the flow remains attached subjecting the body to a much reduced side
force, shown in figure 28(b). Therefore the chines are beneficial imparting additional
directional stability with the consequence that the vertical tails can be reduced in size
or combined with the elevators to produce ruddervators, resulting in reductions in
weight, skin – friction drag and systems complexity.
Figure 28: - Schematic showing cross flow streamlines over: - (a) a cylindrical
body: and (b) a blended body with chines.
Round
body
Blended
chine-body
Flow
separation.
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The same common fuselage was slated for the A-24 UCAV variant of the FB-24
which was analysed with the FB-24 in this thesis, as an external configuration and
initial structural layout conceptual study only.
3.2.3 FB-24 and A-24 Wing and Empennage configuration design selections:
The wing is the most important major component of an aircraft as it gives the aircraft
its characteristic external appearance and determines its flight characteristics. The
many variously configured wing shapes available emphasises that to all intents and
purposes the ideal wing dose not exist and that like all aircraft design wing design is a
compromise where a given wing design will possess specific properties which can
only be attained with that configuration, while the other performance requirements are
met to a grater or lesser degree, the preference for a particular wing characteristic
having been laid down in the design requirements. Wing design has been progressing
as with all other areas of aircraft technology and this development is on going for both
military and civilian aircraft. To aid the reader in the rest of this section the author
will give a brief overview of these developments based on Reference 19:- Modern
Combat Aircraft Design: by Huenecke Klaus: Published by Airlife Publishing: 1987
as follows:-
During the early 1940‟s the general wing shapes for piston engine aircraft of the
period were dominated by the elliptical and trapezoidal wing planforms of relatively
high aspect ratios. The Spitfire typifies the elliptical wing with curved leading and
trailing edges, while the Messerschmitt Me 109 and P-51 Mustang were typical
examples of the trapezoidal planform with straight leading and trailing edges. These
wing shapes were optimal for the speeds which could be attained by these aircraft
which was in the order of 700km/h (435mph) corresponding to Mach = 0.55, because
of their favourable drag characteristics, as shown in figure 29 below. With the
introduction of the turbojet engines in the late 1940‟s, and the greater speeds these
made possible the limitations of these early wing planforms became apparent, and in
order to achieve the speed potential of the new engines the transition to swept back
wings became essential for high subsonic aerodynamics. The first application of wing
sweep was in the Messerschmitt Me 262 with a slight sweep back of the leading edge
by an angle of 18.50 and the radical rocket powered Messerschmitt Me 163 with a
wing sweep of about 500, swept wings were not a feature of the first allied jets the
Meteor and the P-80 Shooting Star which had classical straight wings.
By the 1950‟s the USA and the USSR had absorbed the original German research and
the North American F-86 Sabre and the MiG-15 went on to confirm the principle of
the swept back wing (shown in figure 29). In fact the F-86 it became possible for the
first time to just exceed the speed of sound in a shallow dive. Over this decade as
engines became more powerful operational speeds dramatically increased and by
1960 combat aircraft routinely flew at twice the speed of sound Mach = 2.0, and as
operational speeds increased other planforms were found to be more suitable, for
example the delta wing on the MiG-21 and Mirage, the cropped delta of the F-4 and
the trapezoidal wing of the F-104 Starfighter.
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Figure 29: - Evolution of the modern combat aircraft wing planform source
reference 19.
By the 1970‟s the new high speed wings were found to be difficult to fly at low
speeds for landings, and landing speeds were becoming unacceptably high for the
Mirage 111 delta for example the landing speed was 200mph, and in Europe the
venerability of long airfields needed to cope with these factors was becoming
apparent also the aircraft carrier operation of new larger fighters was becoming
difficult due to high landing speeds. In an effort to overcome the low speed handling
problems variable sweep wings were developed which could translate from a high
speed low aspect ratio delta planform to a low speed high aspect ratio planform as
required imparting greatly improved low speed performance, however these have a
high weight and complexity penalty, examples included the General Dynamics F-111,
the BAE Systems Tornado, Rockwell B-1 bomber, and the Su-24. Also in this decade
aircraft were developed to fly for extended periods at three times the speed of sound
Mach = 3 i.e. the Mig-25 Foxbat and SR-71 Blackbird see figure 29 above.
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While the MiG-25 featured a conservative swept back wing which had a clearly
separated wing and fuselage configuration, the SR-71 exhibited a delta-like wing
which was blended into the fuselage both reducing drag and RCS.
During the 1980‟s the demand for grater manoeuvrability for close combat led to the
development of hybrid planform wings composed of several simple planforms, such
as in the case of the F-16 and F-18 above, where the main wing is trapezoidal with the
leading edge swept back at between 300 and 40
0 and with a very slim delta portion
installed in front as a strake or leading edge extension (LEX).
In the early 1990‟s the development of thrust vectored nozzles, and advanced fly by
wire control systems enabled high agility from a verity of unstable and stealthy
planforms the most extreme example being the F-117 Nighthawk who‟s multi facetted
design is completely unstable, however the latest 5th
generation United States fighters
the F/A-22A and F-35 appear quite pedestrian compared with the European Typhoon
and Rafale but the former rely on advanced stealth shaping, materials, and thrust
vectoring to make the dog-fights the European aircraft are designed for largely
unnecessary, figure 30 shows a comparison of YF-22 with Eurofighter Typhoon.
Figure 30:- YF-22 and Eurofighter configuration comparison.
YF-22
Eurofighter
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As outlined above in sections 1.1 and 3.2, three planforms with high lift devices were
evaluated in this thesis. In order to obtain the best initial configurations to go forward
into the Jet306 study phase the wing designs were formulated on the basis of three
study areas:-
3.2.3.1:- Aerofoil section, including the use of high lift devices:
3.2.3.2:- Planform shape and geometry:
3.2.3.3:- Empennage shape and geometry.
Basically: - the first study area was undertaken to obtain the best compromise
between all of the aerodynamic, structural and mission requirements for the aircraft:
the second study area was undertaken to determine the optimum shape governed
primarily by the requirements of high Mach flight conditions (need to supercruise),
which was also influenced by aerofoil shape: and the third study area was conducted
to the optimum overall size based on operational requirements for given values of the
first two items.
Before starting the detailed study of these three areas it is necessary to discuss the
characteristic parameters for aerofoil and wing aerodynamics, starting with the
characterisation of aerodynamic forces and moments. Fundamental aerodynamics
show the motion of air around the aircraft produces pressure and velocity variations
through the flow, and although the viscosity of the medium is a fluid property, which
acts throughout the flow field, the viscous forces acting on the vehicle depend on the
velocity near the surface as well as the viscosity itself.
The pressure forces which act normal to the surface and the shear forces which act
tangentially to the wing surface are the result of the motion of the air around the
aircraft and are illustrated on a NACA 0006 aerofoil section in figure 31, below.
These pressure and shear forces can be integrated over the surface, on which they act
in order to yield the resultant aerodynamic force (R), which acts at the centre of
pressure (cp) of the aircraft.
For convenience, the total force vector is usually resolved into components. Body –
orientated force components are used for applications concerned with the aircrafts
response (for example: - aerodynamics: and structural dynamics). Consider the forces
and moments in the plane of symmetry (i.e. the pitch plane), as shown in figure 32
below, the body – orientated components are the axial force, which is the force
parallel to the aircraft axis (A), and the normal force, which is the force perpendicular
to the aircraft axis (N).
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Figure 31: - Normal (or pressure) and tangential (or shear) forces on an aerofoil
surface.
Figure 32: - Nomenclature for the aerodynamic forces in the pitch plane.
Shear force
Pressure Force
Shear force
Pressure Force
Shear force
Pressure
Force
A
N
R
D Thrust
CP
L
U
Weight
cg
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In the case of applications such as trajectory analysis, the resultant force is divided
into components taken relative to the flight path or velocity vector, therefore for these
applications the resultant force is divided into a component parallel to the flight path
which is drag (D), and a component perpendicular to the flight path which is the lift
(L), as shown in figure 33. For a powered aircraft its motion through the air is
determined by its weight, the thrust produced by the engine, and the aerodynamic
forces acting on the aircrafts surface. Taking the case of steady, unaccelerated level
flight in a horizontal plane (or steady level cruise), as the simplest condition the
requirements for this condition are, that the sum of the forces along the flight path are
zero and that the sum of the forces perpendicular to the flight path are also zero.
Considering only the cases where the angles are small (for example the component of
the thrust parallel to the free – stream velocity vector is only slightly less than the
thrust itself). Summing the forces along the flight path (parallel to the free-stream
velocity), the equilibrium condition requires that the thrust must equal the drag acting
on the aircraft. Summation of the forces perpendicular to the flight path leads to the
conclusion that the aircraft weight is balanced by the lift.
Consider the case common to most modern combat where the lift generated by the
wing and body configuration acts ahead of the aircraft centre of gravity as shown in
figure 33. In this configuration the lift produces a nose – up (positive) pitching
moment about the centre of gravity (cg). Therefore to trim the aircraft and make the
sum of the moments about the cg equal to zero (i.e. Mcg = 0) a force from a control
surface located aft of the cg (i.e. the empennage) is required to produce a nose – down
(negative) pitching moment about the cg, which should balance out the positive
pitching moment produced by the wing and body lift. The empennage lift force is
illustrated in figure 32 below.
Figure 33: - Moment balance required for trimming the aircraft.
cg
Weight
(W)
Lift
(L) Lift
(LT)
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The orientation of the tail surface which produces the lift force illustrated in figure 33
also produces a drag force, which is known as trim drag, and typically ranges between
0.5% and 5% of the total cruise drag of the aircraft. Note that the trim drag is
associated with the lift generated to trim the aircraft, but dose not include the tail
profile drag (which is included in the total drag of the aircraft at zero lift conditions).
In addition to the force components acting in the pitch plane (i.e. lift acting upwards
and perpendicular to the undisturbed free-stream velocity, and drag, which acts in the
same direction as the free stream velocity) there is a side force, which is the
component of force in a direction perpendicular both to lift and drag. This side force
is positive when acting toward the aircrafts starboard or right wing.
As stated above the resulting aerodynamic forces do not usually act through the
aircrafts cg (centre of gravity). In fact the moment produced by the resultant force
acting at a distance from the origin can be divided into three components, referred to
the aircrafts reference axes (or cg) these three moment components are:- (1) the
pitching moment: (2) the rolling moment: and (3) the yawing moment, as shown in
figure 34.
Figure 34:- Reference axes of the aircraft and the corresponding aerodynamic
moments.
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1. Pitching moment: - The moment about the lateral axis (the y axis of the
aircraft-fixed coordinate system) is the pitching moment. The pitching
moment is the result of the lift and the drag forces acting on the vehicle. A
positive pitching moment is in the nose up direction.
2. Rolling moment: - The moment about the longitudinal axis of the aircraft (the
x axis) is the rolling moment. A rolling moment is often created by differential
lift, generated by aileron and / or spoiler deployment. A positive rolling
moment causes the starboard or right hand wing tip to move downward.
3. Yawing moment: - The moment about the vertical axis of the aircraft (the z
axis) is the yawing moment. A positive yawing moment tends to rotate the
nose to starboard.
The magnitude of these forces and of these moments acting on the aircraft depends on
the combined effects of many different parameters and the major ones are listed
below:-
1. Configuration geometry:
2. Angle of attack (i.e. the aircraft attitude in the pitch plane relative to the
flight direction:
3. Aircraft size (CFD full-size model):
4. Free-stream velocity:
5. Density of the undisturbed air (hence altitude):
6. Reynolds number (as it relates to viscous effects):
7. Mach number (as it relates to compressibility effects).
The calculation of these aerodynamic forces and moments usually requires data from
a range of flow conditions rather than just the one of principle interest, therefore data
from wind tunnel test scale models as well as full scale CFD models is often used or
data from different flow conditions. In order to correlate the data for various free-
stream conditions, and configuration scales, the measurements are usually presented
in dimensionless form, and in this form, the results are independent of all but the first
two parameters listed above, i.e. configuration geometry and angle of attack. However
in practice flow phenomena, such as boundary – layer separation, shock – wave /
boundary layer interactions, and compressibility effects, limit the range of flow
conditions for which dimensionless force and moment coefficients would remain
constant. For these cases, parameters such as the Mach number and the Reynolds
number appear in the correlations for the force and moment coefficients.
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3.2.3.1:- Airfoil selection: -
The characteristics of the airfoil section are defined by several shape parameters of
which the most significant are shown in figure 35 and include: -
1. The maximum thickness to chord ratio and its chordwise location:
2. The nose radius, which should be relatively large to give good maximum lift
coefficient CLmax:
3. The degree and distribution of camber, if employed, some degree of camber is
common for wing sections to enhance lift characteristics:
4. The trailing edge angle, which is usually made as small as possible within
handling and manufacturing constraints.
The NACA nomenclature is used to describe a wide range of airfoil sections in use
today these were developed in the 1930‟s through to the 1950‟s and below is the
descriptive nomenclature for the four digit airfoil section series:-
4 digit code used to describe airfoil shapes:
1st digit - maximum camber in percent chord:
2nd digit - location of maximum camber along chord line (from leading
edge) in tenths of chord:
3rd and 4th digits - maximum thickness in percent chord:
For example: NACA 2412 with a chord of 4 feet:
A max camber: 0.08 ft (2% x 4 ft):
Location of max camber: 1.6 ft aft of leading edge (0.4 x 4 ft):
Max thickness: 0.48 ft (12% x 4 ft).
AERO 315
Airfoil Characteristics
Mean camber line
Chord line
Chord
x=0 x=c
Max thickness
Max camber
Leading edge Trailing edge
x
z
Figure 35: - Characteristics of an airfoil section
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Figure 36: - Airfoil forces and moments.
Figure 37: - Airfoil centre of pressure.
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Figure 38: - Aerodynamic centre 1 illustrates how the moment changes with
location.
Figure 39: - Aerodynamic centre 2 illustrates the point on the airfoil where the
moment is independent of the angle of attack.
For the conceptual design phase the most critical design parameters for this high
speed aircraft were the maximum lift coefficient, the drag coefficient, and the
moment coefficient, which were obtained from the NACA airfoil data charts,
although consideration was given to the following characteristics as advised in
reference 4: - Aircraft Conceptual Design Synthesis: Dr Howe. Denis: Published by: -
Professional Engineering Publishing Ltd: 2002, namely: -
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A. The maximum lift coefficient both at low and high Mach numbers.
B. The stalling characteristics where a gentle loss of lift is preferable,:
C. The airfoil drag especially in aircraft climb and cruise conditions, when the
lift to drag ratio should be as high as possible, and at high Mach numbers.
Most airfoil sections operate with the greater part of the chordwise flow in a
turbulent state but some sections are suitable for laminar flow applications,
this can be difficult to achieve for practical applications but has the potential
for drag reduction and these supercritical sections are described below:
D. The airfoil pitching moment characteristics which may be particularly
important at high speeds causing a significant drag penalty:
E. The depth and shape of the airfoil with respect to the effect on structural
design, ease of manufacture, and possible fuel storage:
F. The slope of the lift curve as a function of incidence in that it affects the
overall aircraft attitude, especially at high values of lift coefficient, such as
required on landing.
Lift coefficient: - (3.1)
Drag coefficient: -
(3.2)
Moment coefficient: -
(3.3)
Note: no dimensional coefficients!
Where: - L, D, and M are the actual lift, drag, and moment (positive nose up) acting
on the airfoil respectively, S is the airfoil reference area and c is mean chord (S
divided by the span b), V is the flight velocity, and is the local air density.
The choice of airfoil section is broadly based on the need to obtain the best
aerodynamic efficiency in the primary operating conditions of the aircraft which in
the case of the FB-24 was supercruising flight. The airfoils selected for supersonic
aircraft are often adapted from basic biconvex (symmetrical) sections to which a small
nose radius and possibly some degree of camber has been added.
2
2
2
2
2
2
VcS
MC
SV
DC
SV
LC
pM
pD
pL
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Maximum lift coefficient (CLMAX):-
From reference 4, the maximum lift coefficient of a basic 2-D airfoil can vary over a
wide range and is heavily influenced by the nose radius of the airfoil section, and
decreasing with the nose radius. In cruising flight buffet margin considerations could
limit the usable lift coefficient to as low as 40% of its maximum value. Therefore the
desire for a supersonic small nose radius has to airfoil had to be balanced with low
speed performance for landing and loitering.
Thickness to chord ratio (t/c):-
This is an important parameter and has an effect on CLMAX and this value is heavily
influenced by structural design requirements. Under incompressible flow conditions
relatively high thickness to chord ratios of up to 0.2 are acceptable at the root of the
wing and give a good structural depth with a small profile drag penalty. The value at
the tip is typically about two-thirds of that at the root.
At higher Mach numbers in the transonic range where compressibility effects become
important, it is usual to use somewhat thinner airfoil sections and root values in the
range of 0.10 to 0.15 are common, with the tip value again being in the order of about
two-thirds of that of the root, but the spanwise variation is not necessarily linear
especially where the wing trailing edge is cranked.
For supersonic and supercruise aircraft such as the F/A-22A and the FB-24 and A-24
the need to reduce wave drag at supersonic speed is paramount therefore the adoption
of much thinner airfoil sections was necessary with thickness to chord ratios rarely
exceeding 0.06 and some as low as 0.03 to 0.02, with any spanwise variation being
very small if any.
This lead the author to select the NACA 0006 section for the wing root with a
thickness to chord ratio of 0.06000 and the NAC 64-006 section for the wing tip with
a thickness to chord ratio of 0.05813, although not as thin as some sections suggested
in reference 4, in the authors view these sections provided the right balance between
aerodynamic properties and practical structural integrity without the wing becoming
too heavy.
Critical Mach number (M CRIT) (2-D airfoil):-
When aircraft are flying in an the high subsonic flight case at or close to their critical
Mach number the rate of drag increase due to compressibility becomes unacceptable,
and a simple definition of this condition from reference 4 is that the critical Mach
number is the one at which the wave drag due to compressibility results in an
increment of 20 drag counts (0.002) to zero lift drag coefficient. It is possible
according to reference 4 to design an airfoil such that the critical Mach number is
unchanged, or even increased, as lift coefficient is increased, although this would be
achieved at the expense of a lower zero lift value of critical Mach number and it is
more usual for the critical Mach number to reduce as lift increases. Increasing the
thickness to chord ratio also results in a reduction of the critical Mach number.
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Reference 4 presents a simple formula for evaluating 2-D airfoil critical Mach number
originally attributed to Korn (and presented by Boppe, C.W., in AGARD-FDP Special
Course notes, May 1991, equation 25), and this was used by the author for the MNCRIT
evaluation of the 2-D airfoil sections selected, an was as follows: -
M CRIT = AF – 0.1 CL – (t/c) = Af – (t/c) (3.4)
Where: - M CRIT = the critical Mach number for a given form of 2-D airfoil:
CL = the lift coefficient:
(t/c) = the thickness to chord ratio.
AF is a numerical value dependant on the design standard of the airfoil section, in the
case of old the value would be in the order of 0.8 – to – 0.85, but for modern
optimised advanced airfoils values of 0.95 are possible, and this latter value was used
by the author in this thesis. The value of Af is equal to (AF – 0.1 CL).
Hence for the purpose of this design study: -
M CRIT = 0.95 – 0.1 CL – (t/c) (3.5)
The cruise lift coefficient for a highly manoeuvrable combat aircraft in reference 4 is
0.3 therefore the critical Mach number for the 2-D airfoils considered for this
conceptual design were:-
For NACA 0006 section: - M CRIT = 0.92 – (0.06) = 0.8600 (3.6)
For NACA 64-006: section: - M CRIT = 0.92 – (0.05813) = 0.8619 (3.7)
The effect of sweep which was of major importance to this design study is covered in
the next section on wing planform.
Lift curve slope:-
The theoretical value of the lift curve slope for thin airfoils is given in reference 4 as:-
dCL / d = 2 per radian (3.8)
Practical 2-D airfoils have a rather higher value but this decreases with the reduction
in aspect ratio and sweep, which is covered in the next section, but an approximate
value for swept low aspect ratio wings could be determined from: -
dCL / d = A/[(0.32+0.16A/cos1/4){1-(MNcos1/4)2}
1/2] (3.9)
Where: - A = the aspect ratio:
MN = the fight Mach number:
1/4 = the sweep of the quarter chord line.
The lift curve slope is only marginally affected by the deployment of leading and
trailing edge flap high – lift devices, unless their deployment grossly increases wing
area.
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Special airfoil sections: -
The general range of combat aircraft airfoils are based on the NACA - 6 series the
Lockheed Martin F-16 family used the NACA-64A204 modified airfoil section where
the second digit (4) denotes the position of minimum pressure (maximum speed) in
tenths of the profile chord, i.e. 40% of chord. The letter A (instead of the standard
hyphen) already represents a modification of the original profile so that 80% of the
upper and lower chord surface contour has been removed producing a flat upper and
lower surface. The third from last number identifies the lift coefficient at which
minimum drag was expected (this value has to be multiplied by 0.1 to obtain the
designed lift coefficient (CL = 0.1 x 2 =0.2)), and the last two digits give the profile
thickness which for the F-16 airfoil is 4%. Manufactures employ either develop their
own airfoil sections or as in the F-16 case adapt NACA profiles to the special needs of
the aircraft, and the co-ordinates of either of these airfoils are military secrets and not
in the public domain. for the purposes of this conceptual design project standard
NACA thin airfoil sections were used with a thickness to chord ratio of 6% which lay
in the general range for supersonic combat aircraft namely 3 to 7% thickness, time
and resources permitting the author would have devised his own sections, but these
were deemed adequate for the level of analysis presented here.
Alongside the classical airfoils which were originally developed for subsonic speeds,
there are further types which have been developed especially for transonic and
supersonic speeds, as shown in figure 40 below.
Figure 40:- Special airfoil sections source reference 19.
Among these was the double – wedge airfoil section used on the F-104 Starfighter,
and the parabolic airfoil section, these had the common characteristic of a sharp
leading edge which effectively fixed the shock position of the supersonic flow.
However these sharp leading edges had disadvantages in subsonic flow, separation
accrued even at low which increased drag and only slight drag-reducing suction
could form at cruising speed.
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In the 1980‟s attempts have been to apply the newly developed supercritical airfoils to
combat aircraft wing design, because these sections provided high lift at low drag
values. However these properties were only manifest within a narrow Mach range
around M = 0.8, and at higher Mach numbers, especially in the supersonic range, a
considerably greater drag than produced by conventional sections was experienced.
This was particularly due to the large nose radius of the supercritical airfoil section as
shown in bottom section in figure 40.
This type of airfoil was considered unsuitable for the FB-24 and A-24 aircraft
configurations because of the large range of speeds which had to be covered from
sustained Mach 1.6 supercruise, BCA / BCM, to an approach speed in the order of
120kts. The principal application of this airfoil type would be large transport aircraft,
with a single design point in terms of a primary Mach number and flight altitude.
High lift devices: -
As stated in section 3.1.1 the FB-24 and A-24 employed high lift devices in the form
of a continuous leading edge plain flap, and a trailing edge plain flap, these were
incorporated into the wing design in order to obtain maximum lift at low speeds,
which could not be supplied by the clean wing.
The plain trailing – edge flap was considered a simple form of flap with an interface
which could be easily aligned and obscured using bull-nose seals on the flap and
blade seals on the wing (see appendices B), being created by rotating the rear part of
the airfoil section around one point within the section, as shown in figure 41 below.
Figure 41:- Types of trailing edge flaps, source AERO 315 USAFA Lecture notes.
The flap deflection increases the camber, so that a greater lift is created for the same
wing angle of attack. At deflections of 100 to 15
0, the flow on the upper surface of the
flap begins to separate, but this separation zone would be confined to the flap itself.
The lift is increased with increasing flap deflection and reaches its maximum value
just before the entire wing flow breaks down, when the flow separation on the flap
jumps forward from the flap on to the wing.
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With a plain trailing – edge flap a lift increase of around CL = 1.0 would be
possible, provided that the initial section is only slightly curved or not curved at all,
and these conditions were fulfilled in the case of the selected NACA sections for the
FB-24, A-24 configurations in common with most combat aircraft. The plain trailing
edge flap represents an effective low – cost, light weight, and low risk common usage
high lift solution which was the reason for its selection for the Advanced Interdiction
Aircraft System.
Figure 42:- Lift and drag coefficient curves for wings with flaps, source AERO
315 USAFA Lecture notes.
Several devices could be used to control the boundary layer flow and generating lift
without unacceptable increases in drag. However, their effectiveness begins only at
high angles - of - attack in the maximum lift range where separation is occurs. These
leading edge devices shown in figure 43 contribute to a reduction in the negative
pressure peaks, and can permit increases in maximum lift of up to CL = 0.2.
The front part of leading - edge plain flap shown in figure 43, would be rotated
around a point within the airfoil section to increase the camber of the wing and move
the point of minimum pressure farther aft on the upper surface of the airfoil at high
angles of attack. This aft movement of the point of minimum pressure extends the
region of favourable pressure gradient and delays separation. The plain flap is simple
and robust requiring little servicing, also like the plain trailing edge it could be
integrated into the planform alignment of the wing and the interface obscured using
blade seals on the flap trailing edge, and bull nose seals on the wing leading edge.
Basic Wing Section
Wing with Flap
CL
Basic Wing Section
Wing with Flap
CL
CD
(a): Affect of trailing edge flap on CLmax. (b): Affect of trailing edge flap on CD.
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Figure 43:- Leading - edge flaps and boundary - layer control devices, source
AERO 315 USAFA Lecture notes.
The effect of such leading – edge flaps and boundary - layer devices on the lift
coefficient is shown below in figure 44 and as can be seen the general effect is to raise
CLmax as the angle of attack increases.
Figure 44:- Effect of leading – edge flaps and boundary – layer control devices
on lift coefficient curves.
The leading – edge flaps effectiveness, ease of integration, low maintenance
requirement, and simplicity, were the reasons for its selection for the FB-24 / A-24
configuration.
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Figure 45: - Aerodynamic characteristics of the NACA 0006 AIA wing root
airfoil.
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Figure 46: - Aerodynamic characteristics of the NACA 64-006 AIA wing tip
airfoil
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For a specific angle of attack and Reynolds number the following coefficients can be
determined from the NACA data charts for the two selected airfoil sections shown in
figures 45 and 46: -
Lift (3.10)
Drag (3.11)
Moment (3.12)
N.B.:- No dimensional coefficients.
Reynolds number effects:-
As the Reynolds number increases the transition from a laminar to turbulent boundary
layer flow, occurs closer to the leading edge of the airfoil, which causes more skin-
friction drag, but delays separation reducing pressure drag. At lower angles of attack
this change in the relative magnitude of skin friction and pressure drag may result in
either higher or lower total drag at higher Reynolds numbers. At higher angles of
attack, where separation and pressure drag dominate, the reduction in pressure drag
due to delayed separation generally results in less total drag at higher Reynolds
numbers. This is shown graphically in figure 47 below which shows an airfoil that for
higher Reynolds numbers has almost the same drag at low angles of attack, yet less
drag at higher angles of attack.
Figure 47:- Airfoil lift and drag coefficient curves for two different Reynolds
numbers.
10
10
CL
.01
10
CD
Cl
Re = 9,000,000 Re = 3,000,000
Re = 9,000,000 Re = 3,000,000
cSq
mc
Sq
dc
Sq
lc
m
d
l
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The airfoil section selection for this conceptual design study was therefore concluded
with the choice of the NACA 64-006 and 0006 sections as being the best compromise
between structural depth and aerodynamic performance on the basis that the whole
wing would be within the supersonic wave cone. The use of leading edge plain flaps
and trailing edge plain flaps as high lift devices was justified in terms of their
simplicity, effectiveness ease of obscuration both in planform alignment and interface
concealment, and this was supported by their current use on the F/A-22A, F-35, and
B-2 bomber, all of which are stealthy platforms.
3.2.3.2:- Wing geometry selection:-
At supersonic speed a Mach wave is formed which surrounds the aircraft, forming a
cone with its apex at the tip of the aircrafts nose, and the angle of this cone relative to
the aircrafts longitudinal axis is known as the Mach angle (). This angle is a function
of the aircraft Mach number as follows (reference 1): -
= sin-1
(1/M) (3.13)
Where M = the Mach number of flight (1.6 from the design requirements in section 1)
Therefore the Mach cone angle () = 38.70
In order to keep all of the airframe within the Mach wave cone requires the wing
sweep back angle to be greater than 90 - 0 = 51.3
0.
Because the air flow in this region is much lower than the free stream velocity the
wave drag would be effectively reduced. The wing leading edge sweep angle selected
for there wing options were as follows:-
Option (A) Large trapezoidal wing LE sweep angle = 550:
Option (B) Arrow head wing LE sweep angle = 550:
Option (C) Swept delta wing LE sweep angle = 600.
The original option C was an F-16XL double delta but this was dropped in favour of
the swept delta based on the F-117A which enabled a larger wing to be constructed
with better RCS spike capabilities, and greater fuel volume. The key drivers for these
planforms were reduced RCS as shown in figure 48, reduced wave drag, and increase
fuel volume, with the reduction in RCS being based on values for RCS spikes
outlined by the USAFA Department of Aeronautics in Reference 20: - lecture notes
AERO-481 Lesson 12: - Survivability Propulsion Integration and Systems.
For the configurations RCS spikes were predicted as follows:-
Option (A) with a conventional four component empennage and a shaped fuselage as
shown in figure 50 and described in appendices B, 8 RCS spikes would be produced,
however with a two component ruddervator empennage this would be reduced to 6
RCS spikes.
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Option (B) with a conventional four component empennage and a shaped fuselage as
shown in figure 50 and described in appendices B above 8 RCS spikes would be
produced this could be reduced to 4 RCS spikes with no empennage but would
present lateral stability issues, however these could be resolved with a two component
empennage although this in turn increases the number of RCS spikes to 6.
Option (C) with a conventional four component empennage or indeed a two
component empennage this wing planform creates 8 RCS with a low RCS planform
aligned fuselage, however this could be reduced to four spiked by opting for a no
empennage configuration.
Figure 48:- The relative effects of wing geometry and empennage configuration
on RCS spikes for a low observable fuselage with all apertures and doors
planform aligned. Source authors LO design presentation.
The initial sizing of these three configurations was based on the requirement for the
largest possible volume for fuel storage to be available within the airframe without
compromising structural integrity or aerodynamic performance. So using empirical
data for: - wing loading: empty mass to take off mass: fuel mass to take off mass: dry
thrust to weight: and payload ratios from published data for fighter, interdiction, and
bomber aircraft as bench mark ranges as shown in tables 5 the author was able to
produce realistic initial sizes for the wings using the target take off weight given by
the requirements in section one.
Figures 49 and 50 below show how planform alignment works and influences design
choices.
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Figure 49:- F-35C planform alignment of RCS spikes to illustrate the behaviour
of an illuminated aircraft. Source authors LO design presentation.
Figure 50:- The contribution of fuselage and empennage alignment to RCS
reduction of an illuminated aircraft. Source authors LO design presentation.
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Having reduced the forward direct specular returns by planform alignment of the
leading edges of the aircraft OML the surface current scattering contributions needed
to be considered at this configuration design stage. This surface current scattering
results from electromagnetic currents which build up on the skin when the aircraft is
illuminated by radar. These currents flow across the skin until they meet a
discontinuity such as a sharp trailing edge, a wing tip, a control surface, or a gap
around a removable panel or door. At a discontinuity, the currents “scatter,” or radiate
electromagnetic energy, as sure as night follows day some of this will be detected
back at the radar, this is shown in figure 51 below.
Figure 51:- Surface current scattering, source reference 20.
The surface current scattering effect is of lower intensity than the specular return, but
is still sufficient for detection, in fact a radar operator only needs two spike returns in
quick succession to obtain an attacking aircrafts direction and velocity and compute a
firing solution which at worst results in dead pilots, or at least some very unhappy
ones. The effect is strongest when the discontinuity is straight and perpendicular to
the radar beam, therefore the wing and empennage trailing edges needed to be swept
as well as the leading edges as was the case with the three planforms studied for the
FB-24 and A-24, to minimise detection from the front, note the large negative
(forward sweep) angles on the trailing edges of the wing and horizontal tails of the F-
35C in figure 49 its not just aerodynamics. Good LO design also calls for diamond or
chevron doors on every access panel as on the B-2, F-117, F/A-22A, and F-35.
The scattering of surface currents shown in figure 51 actually represents three
different types of radar returns: -
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(1) Illuminated surface discontinuities cause diffraction, which analogous to diffracted
light through a prism, and this diffraction not only occurs at a physical edge, such as
the wing or control surface trailing edges discussed above, but at any location that has
a sharp corner this mitigated against the use of the double wedge airfoil section (see
figure 40) for the FB-24 and A-24 wing, this also occurs where there is a shadow edge
where the airfoil section curves away from the illumination source in the transition
from wing leading to trailing edge, this mitigated further against the use of a
supercritical airfoil section (see figure 40) (in addition to its high Mach aerodynamic
performance): (2) Travelling waves are another radar return product of the scattering
of surface currents and occur when a sharp discontinuity is reached and the energy
released (which cannot be destroyed) travels back to the front where it reradiates.
Both edge diffraction and travelling wave radar returns mean that for successful
signature reduction straight trailing edges perpendicular to the illuminating radar
which the original option C choice of a of a planform based on the F-16XL: (3) The
third radar return is from creeping waves which occur when the backside of the
illuminated body is smoothly curved enabling the energy to creep all the way around
the body, slightly radiating as it goes around, this cannot be directly controlled by
shaping and radar absorbent materials covered in section 4 and appendices B would
be used to suppress creeping wave effects on the FB-24 and A-24 airframes. (The
above paragraph was based on material presented by Daniel Raymer which was also
covered in reference 20: - Reference 21: - Page198: Aircraft Design A Conceptual
Approach 3rd
edition: Raymer. Daniel. P.: Published by AIAA: 1999).
Although the choice of the option (B) wing planform was founded upon a similar
planforms consideration for the Northrop Grumman / McDonnell Douglas / BAe
Mach 1.6 JAST submission, and the use of a sub-sonic version on the Northrop
Grumman / Boeing B-2 bomber, Daniel Raymer adds some further supporting
augments for this planform from both a signature and aerodynamic view – point
which are included here for completeness. As will be seen in the next sub-section a
non-tapered wing with a taper ratio of 1 shown in figure 52(a) or a diamond wing with
a taper ratio of 0 shown in figure 52(b) are aerodynamically the worst possible
configuration. The former would have excessive outboard lift especially when swept,
and the diamond wing would have insufficient outboard lift to form an elliptical lift
distribution as desired for minimum drag due to lift. However by combining these two
aerodynamically ineffective planforms as shown below in figure 52(c) and apply twist
and camber to the resulting combination, a fairly good aerodynamically efficient wing
planform would be created. The basic non-tapered wing shown in figure 52(c) is
similar to the original Northrop Grumman / Boeing B-2 bomber configuration which
was revised later on in the design cycle to the current configuration to crate a more
balanced design. The combined planform has beneficial RCS characteristics as well as
improved aerodynamic qualities in so much as when illuminated from the front aspect
the two return spikes are angles away from the receiver, and when illuminated from
the rear the trailing edge also produces two return spikes which are angled away from
the receiver as shown in figure 52, hence the four spike classification of this planform,
however when a empennage is added the number of possible reflectors is increased
rising to eight for a conventional four component tail. As can be seen from figure 48
option (B) for the FB-24 and A-24 configuration studies had a taper in the outboard
wing sections to reduce tip loading and return the lift distribution to a more desirable
elliptical form.
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Figure 52:- Wing planform combination of planform (B):based on reference 21.
Geometric and physical dimensions of the preliminary design:
Aerodynamics Specialists were not available to contribute to the activity of planform
geometry selection and provide assistance on the basic wing geometry, or assist in any
analysis of the three planforms selected, and this was solely conducted by the author
based on reference material accumulated in the course of this project, and on the
results of the Jet306 analysis tool.
Figure 53: Wing geometry major parameters, source reference 19.
Diamond
(b)
Non-tapered but
swept wing (a)
Combined
(c)
Spike 3
Spike 1 Spike 2
Spike 4
c 1/4
S
b / 2
cr
b = full span
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The major drivers for sizing these planforms were as follows: -
Supercruise aerodynamics:
Front and rear spar locations:
Undercarriage location to be aft of the Centre of Gravity (C of G):
RCS (discussed above):
Weight and sizing:
Usable fuel –fuel storage volume within the wing structure:
Type sizing and location of control surfaces:
Location and sizing of high lift devices.
A summary of the influence of the wing parameters given above in figure 53 on the
aircrafts performance is given below:-
1. Aspect Ratio: - The effect of increased Aspect Ratio is to improve the Lift and
Drag ratio (L/D) as shown in figure 54, and is beneficial when the useable take off
incidence is restricted by ground clearance. However, for low altitude, high – speed
fight, (as specified in the requirements document) profile drag is dominant and little
benefit is derived from high Aspect Ratio in respect of radius of action.
Lower Aspect Ratio not only reduces profile drag but gives a smoother ride to crew or
in our case Flight Systems Computer, and reduces structural loading in turbulent
conditions because „g‟ due to gust is proportional to lift slope, and the lower the slope
the more attenuated is the response.
Figure 54:- Effect of Aspect ratio on lift, source reference 19.
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2. Sweep: - Sweep gives higher lift dependent drag and requires increased roll control
of cross wind take offs. However, it delays drag rise „M‟ and reduces lift curve slope.
In this case the sweep angle range of 55 – 60 degrees was dictated by the speed
requirements of Mach 1.6, and for options A and B the leading edge sweep angle of
55 degrees was selected and for option C 60 degrees was selected, see below for wing
sweep analysis.
3. Taper: - The taper transfers load from the wing tip towards the wing root, thus
increasing the likelihood of tip stall (which gives wing drop and pitch up on a swept
wing). Usual taper ratio values lie between 0 and 0.5. A wing with a constant profile
chord i.e. no sweep has a taper ratio of 1; a typical delta on the other hand has a taper
ratio of 0 because of its non-existent wing tip chord. By means of the taper ratio the
load distribution of the wing outwards can be influenced and a quasi – elliptical lift
distribution produced. For a swept wing increased taper gives lower trailing edge
sweep, which enhances the effectiveness of trailing edge flaps and controls (giving
reduced take off and landing speeds and improved controllability in cross winds).
4. Thickness: - Thick section wings incur a Profile Drag Penalty. Increasing
thickness, dose however, give increased maximum lift, eases mechanisation of flaps
and slats, generates a lighter internal structure and presents a greater internal volume
for fuel carriage, so there is a trade off between profile drag, mass and range.
5. Camber: - Chamber is added to enhance lift, however the configurations were
judged to have adequate lift and therefore no camber was employed on the wing, and
chamber is detrimental to low speed performance and was not really considered
worthwhile.
6. High lift devices: - These are of primary benefit on thin swept wings at supersonic
speeds. All of the wing configurations studied use high lift devices as considered in
detail in section 3.2 above.
In summary therefore sweep and low Aspect Ratio (lowest Profile Drag) give a low
lift slope, which coupled with high wing loading gives, rise to good ride quality in
turbulence.
Having considered the basic wing geometry and the wing location the aircraft C.G.
could now be discussed.
(1.) The Mean Aerodynamic Chord (MAC) may be obtained graphically from
the intersection of a diagonal line (constructed as shown above in figure 55), with
the mid chord line of the wing.
(2,) The aircraft centre of gravity can be estimated as a percentage of the MAC
for a given aircraft configuration e.g. 25 to 30% for a stable aircraft with aft tail:
40% (approx) for an unstable aircraft with aft tail as in the case of the FB-24
and A-24 configurations, and 15 to 20% for an unstable aircraft with foreplanes
such as Eurofighter.
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Figure 55: - The wing relationship to the Aircrafts Centre of Gravity.
The effect of wing sweep on MCRIT for finite wings:-
As determined above the leading edge of the wing planforms needed to be swept in
order to keep the wing within the Mach 1.6 wave cone this also had the effect of
raising the wings critical Mach number which was additional to reducing airfoil
thickness.
The method by which this could be achieved is demonstrated in figure 56, by
considering the non-tapered swept wing, where the effective chord length increased
because the chord is measured in the stream wise direction, due to the airfoil shape the
air must flow around is a stream wise cross – section of the wing. From the geometry
shown in figure 56, the relationship between the chord of the unswept wing and the
chord of the swept wing is determined from:-
C (swept wing) = C (non-swept wing) / COS LE (3.12)
So that:-
(t max /c)(swept wing) = (COS LE) (t max /c) (unswept wing) (3.13)
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In the cases of the airfoil sections selected for this conceptual design (e.g. NACA 64-
006) the swept wing chord expression (3.13) could be substituted into the curve fit for
M CRIT data for NACA 64-series airfoils i.e.:-
M CRIT =1.0 – 0.065[100(t max /c)] 0.6
(3.14)
There by producing an expression the critical Mach number for 3-D swept wings:-
M CRIT =1.0 – 0.065 cos0.6
LE [100(t max /c)] 0.6
(3.15)
Or for the unswept wings M CRIT:-
MCRIT = 1.0 – cos0.6
LE (1.0 - MCRIT (unswept)) (3.16)
For tapered wings as in the planforms types evaluated in this conceptual design study,
the effect was modelled by using 0.25c, the sweep angle of the line connecting the
quarter-chord points of the wings airfoils and using the maximum value of (t max /c) on the wing i.e.:-
MCRIT = 1.0 – cos0.6
0.25c (1.0 - MCRIT (unswept)) (3.17)
The effective critical Mach number MCRIT for the leading edge sweep angles used in
this conceptual design study were determined from the methodology presented in
USAFA Aero Lecture notes WIN-12 as reproduced below and Reference 22: -
Introduction to Aeronautics A Design Perspective: by Brandt S. A: Stiles R. J: Bertin
J. J: Whitford R.: Published by AIAA: 1997, in the configurations design and analysis
section 3.2.4 below.
This analysis assumes a wing with a taper ratio of 1.0, so we don‟t have to deal with
different leading edge, quarter chord, and trailing edge sweep angles. The wing as
drawing with the total Mach number over the swept wing equal to 1.0, as shown in
figure 56. Assuming that the velocity component perpendicular to the leading edge of
the swept wing speeds up as it flows over the wing by the same ratio as the flow going
over the unswept wing speeds up, so that:-
unsweptswept
unsweptswept
critcrit
critcrit
MMM
MM
M
/cos
1
cos
wingover thelar perpendicu
wingover thelar perpendicu
(3.18)
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Figure 56:- Unswept and Swept Wings at Their Critical Mach Numbers source
USAFA Aero lecture WIM – 12 supplied within the AeroDYNAMIC V3.0
software cd Rom reference 1(b).
Then, assuming the total Mach number over the wing is 1.0, using the Pythagorean
Theorem the following equation can be derived: -
2
22
2
222
222
22
cossin
1
cossin1
/cossin11
/cossin1
unswept
swept
unswept
swept
unsweptsweptswept
unsweptsweptswept
crit
crit
crit
crit
critcritcrit
critcritcrit
M
M
MM
MMM
MMM
(3.19)
This was a relatively complex equation, but it gave useful results, and was worth the
extra complexity. Table 3 shows the critical Mach numbers this equation would
predict for a wing which has an unswept Mcrit = 0.7. Note that Mcrit does not go to 1.0
until the wing sweep goes to 90 degrees, which is a much more reasonable result.
1m
LE =450
1m
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Table 3 Variation of Critical Mach number with Wing Sweep
MCRIT Non-Swept 0.7
MCRIT Swept Sweep Angle
0.7 0
0.70544532 10
0.721863822 20
0.749402818 30
0.787920521 40 0.836230244 50 0.890798665 60 0.944172829 70
0.984667738 80
1 90
Now, this derivation is not perfect. As Figure 56 shows, for the perpendicular
component to speed up by the same ratio as on the unswept wing, if the parallel
component does not speed up at all, then the flow must turn as it goes over the wing.
The flow does, in fact, turn as it goes over swept wings, but it turns the other way,
toward the wingtip.
However if the flow is assumed not to turn at all, then the derivation presented in
above applies:
M CRIT =1.0 - cos0.6
LE (1.0 – M CRIT (unswept))
Table 4:- Variation of Predicted and Measured Critical Mach
number with Wing Sweep based on F-111 experimentation.
Sweep Angle, deg
Mcrit, Classical
Mcrit, Modified
Mcrit, BSBW
Mcrit, F-111 Flight Test
0 0.75 0.75 0.75 10 0.76 0.75 0.75 16 0.78 0.76 0.76 20 0.8 0.77 0.76 26 0.83 0.78 0.77 0.75 30 0.87 0.79 0.77 35 0.92 0.81 0.78 0.8 40 0.98 0.83 0.79 50 1.2 0.87 0.81 0.85 60 1.5 0.91 0.84 0.9 70 2.2 0.96 0.87 72 2.4 0.96 0.88 0.95 80 4.3 0.99 0.91 90 infinite 1 1 1
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The analytical method results were compared with those measured on actual aircraft.
For this purpose, flight test data for the F-111 with several different wing sweep
settings was used, and Table 4 compares the F-111 flight test data with values of Mcrit
predicted using the various methods described in lecture WIM-12. This gave
confidence in this prediction method over the classical method publicised in some
textbooks.
Having determined the airfoil section choice: the wing planform choices options (A),
(B) and (C): and the wing leading edge sweep angles for these planforms, the next
stage in the wing design was to size the wings in conjunction with the rest of the
airframe OML and analyse their aerodynamic properties in term of the lift and drag,
characteristics of each option and as part of the full configuration analysis was
undertaken in section 3.2.4 below.
3.2.3.3:- Empennage geometry selection: -
The principle role of the tail is to counter the moments produced by the wing. Thus
the tail size is directly proportional to the wing size and this was determined from the
moment equations examined in section 3.4. It would be expected that the tail area
divided by the wing would show a consistent relationship for different aircraft
empirical data, if the effects of the tail moment arm could be accounted for. A method
for doing this was adopted from reference 21, and was termed the “tail volume
coefficient” method for the initial estimate of tail size, which was based on the fact
that the force due to the tail lift is proportional to the tail area times the tail moment
arm.
For the vertical tail, the wing yawing moments have to be countered which are most
directly related to the wing span b W and from this relationship the “vertical tail
volume coefficient” was derived as defined in equation 3.20:-
c VT = LVT SVT / bW SW (3.20)
For the horizontal tail the pitching moments which must be countered are most
directly related to the wing mean chord C-W and from this relationship the
“horizontal tail coefficient” was derived as defined in equation 3.21:-
c HT = LHT SHT / C-W SW (3.21)
As recommended in references 19 and 21, the moment arm (L) was approximated as
the distance from the tail quarter – chord (25% of mean chord length measured back
from the leading edge of the mean chord) to the wing quarter chord. The tail moment
arm definition is shown in figure 57 below along with the definitions of tail area. Note
that the horizontal tail area was measured to the aircrafts centreline which is standard
practice according to reference 21. In the case of both the FB-24 and A-24 as part of
the low RCS design twin vertical tails were adopted and for this configuration the
vertical tail calculated represented the sum of the areas of both tails.
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Figure 57:- The definition of tail moment arm and tail area definitions for initial
sizing.
Using historical data from Table 4.1 (Configuration data for modern combat aircraft
F-15 to MiG-25) page 59 of reference 19, an average value for the horizontal tail
volume was determined as follows: - c HT = 0.476 and a value for the vertical tail
volume was taken from Table 6.4 (Tail volume coefficient) page 125 of reference 21
c VT = 0.07 to use in equation 3.22 to determine horizontal tail areas and equation
3.23 to determine vertical tail areas respectively:-
SHT = c HT C-W SW / LHT (3.22)
SVT = c VT b W SW / LVT (3.23)
The moment arm was approximated by a percentage of the fuselage length and from
reference 21 for an aircraft with an aft mounted engine a tail moment arm range of
between 45% and 50% of the fuselage length was recommended. This was verified by
examining historical combat aircraft data from reference 19 as well as measurements
taken from public domain data for the F/A-22A and the F-35C.
For the YF-23 type ruddervator tail group configuration studies the tails were sized to
provide the same surface area as a conventional four tail group, and as stated in
section 3.2 these were given a dihedral angle of 500 based on the flight proven YF-23
aircraft.
LHT
LVT
CL
SW = wing area:
bW = wing span:
C-W = wing mean chord.
bV
bH
b/2
c1/4
c1/4
c1/4
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Figure 58:- F-35C port – side conventional tail layout with fixed VT and all
moving HT although the flip tail was supplanted by a pivoted tail, source
authors‟ private collection.
Figure 59:- YF-23 Ruddervator all moving V – tail for FB-24 and A-24
configuration sizing the ruddervator area equalled the conventional four tail
group source authors‟ private collection.
All moving (flipping) HT
area = 5.78m2
Fixed VT (dihedral 270) total area
from rudder + tail = 4.547m2
Rudder area = 1.332m2
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3.2.4 Common FB-24 and A-24 configuration initial sizing.
This stage of the conceptual design was intended to produce complete models of the
proposed modifications to the original F-35C aircraft to produce the definitive OML‟s
for the FB-24 and A-24 for evaluation with the Jet306 toolset, in section 3.3.
These models contain following key attributes:-
1. A common 3.048m (10ft) extended two crew position fuselage for all of the
FB-24 wing and empennage variants:
2. A common 3.048m (10ft) extended no canopy fuselage for all of the A-24
UCAV wing and empennage variants:
The following FB-24 aircraft surface models were produced:-
1. New Build 1 (NB-1):- Common two crew position fuselage with option (A)
520 leading edge sweep angle trapezoidal wing, and conventional four
component empennage:
2. New Build 2 (NB-2):- Common two crew position fuselage with option (A)
520 leading edge sweep angle trapezoidal wing, and ruddervator two
component empennage
3. New Build 3 (NB-3):- Common two crew position fuselage with option (B)
550 leading edge sweep angle cranked arrow wing, and conventional four
component empennage:
4. New Build 4 (NB-4):- Common two crew position fuselage with option (B)
550 leading edge sweep angle cranked arrow wing, and ruddervator two
component empennage:
5. New Build 5 (NB-5):- Common two crew position fuselage with option (C)
600 leading edge sweep angle swept arrow wing, and conventional four
component empennage:
6. New Build 6 (NB-6):- Common two crew position fuselage with option (C)
600 leading edge sweep angle cranked arrow wing, and ruddervator two
component empennage:
The following A-24 UCAV aircraft surface models were produced:-
1. New Build UCAV 1 (NBU-1):- Common no canopy fuselage with option (A)
520 leading edge sweep angle trapezoidal wing, and conventional four
component empennage:
2. New Build UCAV 2 (NBU-2):- Common no canopy fuselage with option (A)
520 leading edge sweep angle trapezoidal wing, and ruddervator two
component empennage:
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3. New Build UCAV 3 (NBU-3):- Common no canopy fuselage with option (B)
550 leading edge sweep angle cranked arrow wing, and conventional four
component empennage:
4. New Build UCAV 4 (NBU-4):- Common no canopy fuselage with option (B)
550 leading edge sweep angle cranked arrow wing, and ruddervator two
component empennage:
5. New Build UCAV 5 (NBU-5):- Common no canopy fuselage with option (C)
600 leading edge sweep angle swept arrow wing, and conventional four
component empennage:
6. New Build UCAV 6 (NBU-6):- Common no canopy fuselage with option (C)
600 leading edge sweep angle cranked arrow wing, and ruddervator two
component empennage:
The intention was then to analyse these configurations against the F-35C 230-5 OML
model using the hand calculation whole aircraft analysis techniques given in reference
22 section:- 4.7 to obtain initial comparison of:- Lift: Parasite drag: Induced Drag:
and Supersonic drag, to determine which of the 6 FB-24 configurations and 6 A-24
configurations gave the highest improvement in complete aircraft drag polar over a
Mach number range of Mach 0.3 – Mach 1.6, over the baseline F-35C analysed by
the same methods over the same Mach range, and only these configurations were put
forward for Jet306 analysis, this however had to be re-scoped to meet the constraints
of this design study and a less extensive comparison was made.
To produce accurate models for initial sizing required estimates of the main aircraft
parameters and as an initial starting point the values associated with existing aircraft
of similar types to the intended Advanced Interdiction Aircraft System were
investigated. Although the A-24 UCAV was unique in function i.e. supercruise long
range unmanned combat aircraft no strictly comparable weight data exists, however
because the A-24 was basically a common airframe with the manned FB-24 the use of
manned aircraft was deemed appropriate, and if the A-24 empty weight was
significantly less than the FB-24 this would be used for additional fuel storage. To
this end a list was compiled for existing military aircraft in the: - Fighter: Interdictor:
and Bomber categories (excluding subsonic aircraft) table 5 below, from published
data using references: - 19: 23:-Jet Bombers (From the Messerschmitt Me262 to the
Stealth B-2): by: - Gunston. B. and Gilchrist. P.: Published by: - Osprey Aerospace:
1993: and 24:-Modern Fighters: by: - Spick. M.: Publishes by: - Salamander Books
Ltd: 2000.
The production of these accurate models of the study variants enabled estimates to be
made of the component weights, drag, and lift. The predictions for thrust requirement
derived below allowed down selection of the engine capable of providing the required
performance over all fight conditions. With weight, aerodynamic, and propulsion data
it was possible to perform the initial performance analysis and produce the initial
constraint diagram.
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Table 5:- Comparison of equivalent military aircraft parameters.
Parameter Fighters radius
<700miles
Interdictors radius
1000miles
Bombers radius
>1000miles
Empty weight ratio (ME / MTO) 0.46-0.72 0.49-0.57 0.40-0.59
Fuel weight ratio (MF / MTO) 0.22-0.45 0.19-0.33 0.33-0.64
Payload ratio (MPAY / MTO) 0.21-0.28 0.09-0.32 0.08-0.134
Wing loading (MTO / S) kg/m2 261-818 225-1050 204-1194
Thrust / Weight ratio (dry) 0.65-1.48 0.44-0.56 0.26-0.52
Aspect ratio (b2 / S) 2.36-7.28 1.64-7.08 1.75-9.58
To make the data as representative as possible only three subsonic aircraft were
included in the above table studies and these were both interdictors namely the
Lockheed Martin F-117A, the British Aerospace Buccaneer S-2, and the BAe /
Boeing Harrier, all others being supersonic capable aircraft.
The highest wing loadings for all categories came form swing wing types i.e.: -
Tornado F-3 Fighter: Tornado GR-4 Interdictor: and the B1B bomber, and the lowest
wing loadings came from the Mirage 2000 Fighter: F-117 Interdictor: XB-70 Valkyrie
Bomber. The highest non variable geometry values being: - MiG-31 =666kg/m2
Fighter: Buccaneer = 550kg/m2 Interdictor: B-58 Hustler = 560kg/m
2 Bomber.
The lowest fuel weight ratios for all categories came from the: - AIDC A1 Ching Kuo
Fighter: Tornado GR-4 Interdictor: and the TSR-2 Bomber, and the highest came
from the: - F/A-22A Fighter: F-117A Interdictor: and F-111F Bomber.
The highest aspect ratios were all from swing wing aircraft understandably i.e.:-
Tornado F-3 Fighter: Tornado GR-4 Interdictor: and the B-1 bomber, the highest
values for non variable geometry aircraft were: - F/A-18E Hornet = 4.00 Fighter:
Buccaneer S-2 = 3.52 Interdictor: and A-5 Vigilante = 3.75 Bomber.
Applying this approach to the AIA system is valid for both the manned FB-24 aircraft
variant as well as the A-24 as both are conceived as extended range interdictors with
the common airframe approach for the A-24 UCAV.
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However the AIA system dose not follow the „fighter‟ category of aircraft because of
the need to a longer distance with a heavier payload than is normal for fighters,
neither dose the AIA system fit into the „bomber‟ category of aircraft because of the
requirement for sustained high speed and lighter payload, over a modest range. The
most realistic category is long range interdictor but the current generation of
interdictors are not expected to supercruise to the targets for long periods or deliver
weapons from 45,000ft undetected, they do have greater manoeuvrability and are
expected to fight their way to and from the target at medium or low level, resulting in
a higher stressed 9g and hence heavier airframe than either the FB-24 or A-24.
From Table 5, it is clear that wide variations exist in the aircraft studied, excluding the
extreme values from variable geometry aircraft which were a niche phase in military
aircraft in the 1970‟s, it is possible to assess the variations in some design parameters,
although this only serves as a crude guide for initial sizing and the refined method of
whole aircraft drag polar analysis against the F-35C using reference 22 methods, was
used airframe refinement and for selection for Jet306 toolset submission. To enable
the selection of suitable starting values reflection on the key requirements of the AIA
system aircraft was necessary, and the following considerations raised from the
requirements: -
All but one of the aircraft in the authors survey were not supersonic in dry /
military thrust therefore the AIA system aircraft would require a higher thrust
to weight ratios than the values for interdictors and bombers.
Travelling for longer distances at supersonic speed obviously requires more
fuel than is seen in the fighter and interdictor categories in table 5 but less than
max bomber value from the Valkyrie (Mach 3).
The fuel capacity required was larger than for an equivalent sized aircraft
which favoured a larger wet wing for increased fuel storage, which unlike the
naval F-35C aircraft with its need for a folding outboard wing section, could
be wet to the tip, which could also produce bending relief.
The larger wing with the resultant lower wing loading would also help the
aircraft meet the icy runway landing requirement.
The weapons load based on the weights supplied in table 2 section 3.2 above,
and defined in section 2 constituted a relatively low payload ratio to the Max
TO weight. Where as the range to be flown at supersonic speed gave a fuel
weight ratio.
The empty weight ratio of the aircraft compared to the F-35C would be
reduced due to the de-navalisation of the airframe by removing structural
weight associated aircraft carrier landing, catapult launch loads, from
reference 21 this would be a weight saving in the order of 12%. This weight
saving could be eroded by the requirements of materials which would be
required for the aircraft to endure relatively high temperature long duration
Mach 1.6 flight, at 45,000ft like BMI composites, ceramics, and titanium.
Therefore a reduction of 10% was considered more realistic for the AIA
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system aircraft. The use of low observable coatings (see appendices B, and
reference 13) is no longer considered a high airframe weight driver and other
methods have been adopted to eliminate structural steps and gaps (which
cannot be presented here under ITAR clearance rules) that do not
demonstratively add weight to the airframe.
Taking the above considerations into account the initial estimates for the AIA system
aircraft were made as follows and compared to calculations made for the F-35C 230-5
using OML publicised data:-
Empty weight ratio = 0.40: - this was low compared with the fighters and
interdictors surveyed in table 5 but just within the supersonic capable bomber
category and was below the navalised F-35C empty weight ratio of 0.43
with all internal and external weapons stations filled. This took into account a
non navalised 10% structural weight saving and the weight growth with
fuselage growth.
Fuel weight ratio = 0.51: - This was higher than any of the fighters or
interdictors surveyed in table 5 but was mid range for the supersonic bombers
surveyed, and was 1.8 times that of the F-35C which in its current form has
an internal fuel weight ratio of 0.27.
Payload weight ratio = 0.09: - Based on the published release weights of
ASRAAM, and JADAM – GB31 PIP weapons from table 2, and the projected
weight for the ALOSNW, and for the mission requirements weapons fit
options the actual payload weight ratio required was 0.076, however the
inclusion of the two crew and potential weapons weight growth increased this
value of 0.09. This value falls just below the lowest range points for the
interdictor range form the F-117A value and the lowest range point for
bombers from the clean F-111F.
Wing loading range = 221kg/m2 to 352kg/m: - The top of this range was
about mid range for all non variable geometry aircraft surveyed in the three
categories in table 5 and was considered the highest permissible value for all
three wing planform options, in view of the need to reduce structural weight of
the airframe. The factors influencing the final wing loading and the analysis
supporting this selection are detailed below. Compared to the wing loading
of the F-35C at MTO = 560kg/m2 the top of this range was 0.63 times the
calculated F-35C figure from model data, however the maximum vale in
the range was just above the published figure for the F/A-22A =
348.7kg/m2.(79.43lb/ft
2) the only true combat aircraft supercruise in
existence.
Thrust loading = 0.48: - This was low for fighters, in the lower range for
interdictors but higher than bombers surveyed in table 5. This was also higher
than the navalised F-35C which had a Tdry / WTO (Thrust loading) value =
0.40 calculated from published data.
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It was now possible to use these assumed values and the fuselage length and cross-
section estimate from section 3.2.2, and the methodology of determining tail areas
from section 3.2.3.3, and the requirements data from section 2 to make the first rough
predictions of the size of the aircraft, and create accurate models for analysis.
From the requirements definition in section 2 the value of MTO = 32,213.23kg
(71,018.16lbs) a payload of 2 x ALOSNW + 2 x ASRAAM = 2442.20kg
(5,384lb) = 0.076 of MTO therefore based on a payload weight ratio of 0.09 x
MTO a further 450.98kg (994.25lb) was available to accommodate the weight
of the two crew members and weapons growth for the FB-24 manned aircraft
and for the AI system for the A-24 UCAV.
The empty weight ratio of 0.40 x MTO = 12,885.28kg (28,407.20lb) which
was below the do not exceed target empty weight set out in the requirements
detailed in section 2 of 14,141.65kg (31,177lb)
The fuel weight ratio of 0.51 x MTO = 16,428.74kg (36,219.18lb) which was
1.89 times that of the F-35C value of 8,663.61kg (19100lb).
The wing loading range of 220kg/m2 to 350kg/m
2 resulted in gross wing
areas of 146.42m2 to 92.04m
2 respectively, which were applied as target
ranges for practical wing sizing of options (A), (B), and (C). The benefits of a
low wing loading are good supersonic performance and acceleration, and rapid
roll rates, but this is at the cost of excessive speed loss during hard turns and
an increase in the speed required for maximum instantaneous rate turns as
Mach number increases as shown in figure 60 below.
The thrust loading of 0.48 x MTO (sea level static dry) = 15,462.34kgs
(34,088.63lbs) of thrust required from the engine, this dictated the selection of
the YF-120 derived F-136 (Variable Cycle Engine) engine with a
maximum thrust of 43,000lbs and a projected dry thrust of 34,400lbs over
conventional F-119 derived F-135 engine with a maximum thrust of 4300lb
but a dry thrust of only 28,810lbs, see also appendices D.
An aspect ratio between 2.00 and 3.5 was considered as reasonable for the
AIA system aircraft, and was driven by the desire to keep the wing within the
Mach wave cone, so that the shockwave drag due to volume (that part of wave
drag due to the bulk of the aircraft, which is independent of lift) was
minimised, and to keep the wing within manageable proportions with respect
to the weight of the wing, and to maintain control surface effectiveness, as the
aerodynamic centre shifts aft with Mach number.
The fuselage length and maximum cross section were determined above to
give a comparable finesse ratio to the YF-22 as length = 18.29m (60ft) and
maximum cross-section = 4.47m2.
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Figure 60:- Effect of reducing wing loading and increasing thrust to weight
ratios on the speeds for maximum instantaneous and maximum sustained turn
rates respectively with increasing Mach number. (Chart based a diagram from
reference 22).
From the wing loading and corresponding areas calculated above and after several
iterations the most practical configurations for the three wing options were modelled
based on aerodynamics, structural and signature requirements, and these are detailed
below in figures 61 through 63 and the key dimensions were as follows:-
Option (A) figure 61:-
Total span = 13.58m: Total area = 91.76m2: Wing loading = 351kg/m
2: Aspect
ratio = 2.0: Reference chord = 11.97m: Tip chord = 1.43m: Leading edge sweep =
550: Trailing edge sweep = - 20
0.
Option (B) figure 62:-
Total span = 19.02m: Total area = 110.76m2: Wing loading = 290kg/m
2: Aspect
ratio = 3.29: Reference chord = 11.45m: Tip chord = 3.18m: Leading edge sweep
= 550: Trailing edge sweep inboard = -14.98
0 / outboard = 48
0.
Option (C)figure 63:-
Total span = 19.02m: Total area = 145.64m2: Wing loading = 221kg/m
2: Aspect
ratio = 2.48: Reference chord = 11.97m: Tip chord = 3.12m: Leading edge sweep
= 600: Trailing edge sweep = 38.7
0.
Mach Number
Ra
te o
f T
urn
Lift limit curve
Increasing lift or
reducing W/S
Increasing T/W or L/D
Speed for max instantaneous rate of turn
Speed for max sustained rate of turn
Sustained manoeuvre boundary
Structural limit (independent of aerodynamics and geometry
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Figure 61:- Wing option (A) final initial sizing iteration.
Figure 62:- Wing option (B) final initial sizing iteration.
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Figure 63:- Wing option (C) final initial sizing iteration.
In order to determine the aerodynamic moment about the aircrafts aerodynamic centre
for the as drawn planforms the mean aerodynamic chord for each was determined
using both the geometric method, and the calculation method (as a verification of the
geometric analysis) presented by Raymer reference 21 and shown in figure 55. The
mean aerodynamic chord c- is defined as the chord length that when multiplied by the
wing area, the dynamic pressure, and the moment coefficient about the aerodynamic
centre, yields the value of the aerodynamic moment about the aircrafts aerodynamic
centre, and can be calculated from:-
b/2
c- = 1/S ∫-b/2 c
2 dy (3.24)
Where c is the local value of chord length at any spanwise location, the spanwise
location and c- is the mean aerodynamic chord (MAC) determined by the geometric
analysis based on CATIA V5 model geometry of the wing planform options A, B, and
C, and the baseline F-35C is shown below in figures 64 through 67. As a geometry
check the numerical method was applied to each wing planform option and given
below each respective geometric analysis below.
The location and length of the mean aerodynamic chord is important when locating
the wing with respect to the fuselage, because the wing is located so that some
selected percentage of the MAC is aligned with the aircrafts centre of gravity, which
provides a first estimate of wing position to attain the required degree of stability.
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Figure 64:- Determination of the mean aerodynamic chord for option (A).
As a verification check the mean aerodynamic chord length and location was
calculated for wing option A from: -
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 11.975 (1 + 0.11 + 0.11
2) / (1 + 0.11) = 8.067m
y- = (b/6) (1 + (2 / 1 + ) (3.26)
y- = (13.58/6) (1 + (2 x 0.11)) / (1 + 0.11) = 2.488m
The calculated values were lower than those produced by the geometric method by
17mm for c- chord length, and 13mm for y
- chord location from the aircraft
centreline and therefore in this case the geometric analysis gave a valid
approximation.
CRoot = 11.975m
CTip = 1.426m
C- = 8.084m
F-35C Outline to scale
for reference.
MAC
50%
CL
CRoot
CTip
y-= 2.501
b/2
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For non trapezoidal wings or combinations of two trapezoidal planforms as was the
case with option (B) the approaches of splitting the wing in two was used for this
analysis of the wing mean aerodynamic chord and this is shown below, in figures 65.
Figure 65:- Determination of the mean aerodynamic chord for option (B) by
breaking the wing into Inboard (Inbd) and Outboard (Outbd) sections.
As a verification check the mean aerodynamic chord length and location for both
inboard and outboard sections were calculated for wing option B from: -
Inboard:-
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 11.445 (1 + 0.43 + 0.43
2) / (1 + 0.43) = 8.617m
y- = (b/6) (1 + (2) ) / (1 + )
(3.26)
Outbd
Inbd
F-35C Outline to scale
for reference.
CL
C Inbd Root
CTip
C- = 8.640m
MAC Inbd
MAC Outbd
C- = 4.149m
y-= 2.640m
y-= 1.656m
b/2
CRoot = 11.445m
CTip = 4.989m
CRoot = 4.989m
C Outbd Root
50%
50%
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y- = (7.62/6) (1 + (2 x 0.43)) / (1 + 0.43) = 1.652m
For the Inboard wing calculated the value for c- chord length was lower than those
produced by the geometric method by 23mm, and for y- chord location from the
aircraft centreline the calculated value was lower by 4mm therefore in this case the
geometric analysis gave a valid approximation.
Outboard:-
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 4989 (1 + 0.64 + 0.64
2) / (1 + 0.64) = 4.157m
y- = (b/6) (1 + (2 / (1 + )
(3.26)
y- = (11.40/6) (1 + (2 x 0.64)) / (1 + 0.64) = 2.641m
The calculated values were higher than those produced by the geometric method by
5mm for c- chord length, and 1mm for y
- chord location from the aircraft centreline
and therefore in this case the geometric analysis gave a valid approximation.
In order to determine the length and location of the mean aerodynamic chord for the
combined wing the outboard chord length was subtracted from the inboard chord
length to give c- = 4.460m and the values of chord location were added together y
- =
4.293m, mapping these values onto the geometry gave a c- value = 4.835m or =
375mm more than the calculated value which was attributed to the 70 difference in
sweep angle between the leading and trailing edges of the outboard wing.
Therefore for the purposes of this conceptual design study for the option B wing the
geometric value for c- and the calculated value for y
- were used for all further
analysis and empennage sizing, therefore for the combined option B wing planform:-
c- = 4.835m
y- = 4.293m
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Figure 66:- Determination of the mean aerodynamic chord for option (C).
As a verification check the mean aerodynamic chord length and location was
calculated for wing option C from: -
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 11.974 (1 + 0.26 + 0.26
2) / (1 + 0.26) = 8.411m
y- = (b/6) (1 + (2 / 1 + ) (3.26)
y- = (19.02/6) (1 + (2 x 0.26)) / (1 + 0.26) = 3.824m
The calculated values were lower than those produced by the geometric method by
1mm for c- chord length, and 2mm for y
- chord location from the aircraft centreline
and therefore in this case the geometric analysis gave a valid approximation.
CRoot
CL
50%
CTip CRoot = 11.974m
CTip = 3.120m
F-35C Outline to scale
for reference.
MAC
C- = 8.412m
b/2
y-= 3.826
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The values from all three wing options lead particulars were compared with the
baseline F-35C in table 6 below.
Figure 67:- Determination of the mean aerodynamic chord for F-35C.
As a verification check the mean aerodynamic chord length and location was
calculated for the F-35C wing from: -
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 7.412 (1 + 0.18 + 0.18
2) / (1 + 0.18) = 5.077m
y- = (b/6) (1 + (2 / 1 + ) (3.26)
y- = (13.11/6) (1 + (2 x 0.18)) / (1 + 0.18) = 2.518m
50%
F-35C outline to 230-5 OML.
CRoot = 7.412m
CRoot
b/2
CL
CTip
CTip = 1.358m
C- = 5.081m
MAC
y-= 2.522m
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The calculated values were lower than those produced by the geometric method by
4mm for c- chord length, and 4mm for y
- chord location from the aircraft centreline
and therefore in this case the geometric analysis gave a valid approximation.
Table 6: - Wing configuration summary.
Parameter F-35C wing Option A
wing Option B
wing Option C
wing
Span (m) 13.11 13.56 19.02 19.02
CRoot (m) 7.412 11.973 11.445 11.973
CTip (m) 1.358 2.036 3.176 3.119
0.18 0.17 0.28 0.26
Sweep angle 340 520 550 600
Area m2 57.478 94.984 109.176 143.572
Loading (kg/m2)
560 339 295 224
Aspect ratio 2.99 1.94 3.31 2.52
MAC (m) 5.077 8.179 4.835 8.411
Location of MAC (m)
2.518 2.588 4.293 3.824
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Figure 68:- Original option (B) on FB-24 fuselage with F-35C empennage.
From the analysis conducted to this point the wing sizes in terms of span required to
meet the lower wing loading ranges in the initial sizing estimates were in the authors
view becoming too large to remain stiff enough to combat the effects of flutter and
aeroelaticity even with advanced composite aeroelastic tailoring at the speeds of
Mach 1.6 without becoming unacceptably heavy i.e. greater than 7% of MTO, for
example the option (B) wing to meet a wing loading of 290kg/m2 had a span of
19.02m (62ft), shown above in figure 68. The option (C) wing had also grown in span
and chord to achieve its target of 221kg/m2 to a size that would not permit a four tail
empennage layout on the optimum stretched fuselage and would foul the ruddervator
layout at the wing trailing edge.
The cantilever wing, because of its limited spar depth tends to be inherently flexible,
and early low speed wings were made as thick as possible, particularly at the root and
had cantilever ratios (semispan to root thickness) in the order of 10, and wings of this
low cantilever ratio were relatively stiff in bending but their torsional stiffness was
still low. By the 1940‟s cantilever ratios of the order of 15 were common and serious
consideration had to be given to aeroelastic effects in the design of fighter aircraft
wings, the option B and option C wings had a cantilever ratio of 26.8 in their original
form, therefore aeroelastic effects would be a serious issue. Sweepback also induced
additional torsion because the applied loads on the wing act significantly aft of the
wing root, so that the stiffness requirements would be greater than for a trapezoidal
wing of the same span.
Wing twisting in response to aileron deflection decreases the available rolling
moment as a function of dynamic pressure that is the speed2, because as speed
19.02m
18.29m
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doubles the associated loads are quadrupled. In the worst case aileron reversal can
occur, this is where the control surface loads deflect the wing rather than the air.
Under these conditions the rolling moment obtained is in the opposite direction to that
commanded. This raised issues undesirable wing bending and possible reduced tip
AoA (angle of attack) as well as control reversal manoeuvring the planforms at Mach
1.6 for both planforms B and C if the low wing loading values proposed in the initial
sizing were to be achieved. Moreover to achieve the required stiffness the wing skins
would need thick and relatively heavy even in CFC, even the F/A-18A composite
wing suffered from a marked loss in roll rate in transonic flight due to its relative
flexibility, in spite of having a differential tailplane, and inset ailerons (Reference:-
25:- Fundamentals of Fighter Design 2nd
Edition: Whitford. R: Published by Airlife
Publishing 2005) although this was resolved by software re-writes. Also the
undercarriage layout would require a high degree of modification to support these
wing sizes during ground handling.
In order to address these issues at this early point in the design process a single point
wing loading was selected to enable high torsional stiffness, high sweep, high taper –
ratio, low aspect ratio versions of all three planforms to be devised as shown below.
The final wing loading selected after detailed assessment of comparable fighter /
interdictor aircraft was 388kg / m2 which reduced the wing area for all three
planforms to a standard of 83m2.
Another point of concern with respect to the option A wing was the very low taper
ratio of 0.11, which was considered as a means of reducing wing structural weight,
because as decreases from 1.0 (for a rectangular wing) to 0 (for a triangular wing)
the preponderance of the lifting force shifts inboard, closer to the wing root i.e. the
centroid of the lift distribution (c.p. centre of pressure) moves closer to the wing root.
This in turn reduces the moment arm from the root to the centre of pressure and
consequently the bending moment at the root decreases, but the lift remains the same,
and as a result the wing structure could be made lighter which would clearly benefit
the FB-24 and the A-24 aircraft. (This was supported by references 18, 21, and 22.)
However as highlighted in reference 18, as the taper ratio is decreased, the region
where flow separation first develops moves outward towards the tip, and if is
reduced to 0 the separated flow region occurs at the wing tip area, which would result
in a total loss of aileron control, this would be unacceptable for either the manned FB-
24 or the A-24 UCAV, or indeed any other aircraft. Therefore after analysing both the
F-35C and the F/A-22A which have similar control requirements to the AIA system
aircraft and have taper ratio‟s of = 0.18, and = 0.17 respectively the author
considered increasing the option A taper ratio to = 0.18 at this stage as a
precautionary measure against later redesign was incorporated into the planform A
wing maturation and reduced the impact on the amount of work required at a later
stage in the design.
The redesign was affected by reducing the forward sweep angle of the leading edge to
520 and the trailing edge to -15.32
0 from the original 55
0 and -20
0 respectively, and
this combined with cropping the tip at 900 to the leading edge achieved an increase in
tip chord to CTip = 2.030m from the original CTip = 1.430m giving a = 0.17. This
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reconfigured wing was used for the rest of the conceptual design study, and the MAC
analysis is given below in figure 69.
As a verification check the mean aerodynamic chord length and location was
calculated for the F-35C wing from: -
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 11.473 (1 + 0.18 + 0.18
2) / (1 + 0.18) = 7.859m
y- = (b/6) (1 + (2 / 1 + ) (3.26)
y- = (13.72/6) (1 + (2 x 0.18)) / (1 + 0.18) = 2.636m
The calculated value for c- was 455mm higher than those produced by the geometric
method because of the right angle tip crop, and the value for y- was 18mm higher
than the geometric analysis and the geometric values were used for this conceptual
design, see figure 69 below therefore the geometric values were used throughout.
Figure 69:- Determination of the mean aerodynamic chord for the revised wing
option A.
MAC
C- = 7.404m
CTip CRoot = 11.473m
CTip = 2.030m
CRoot
50%
F-35C Outline to scale
for reference.
CL
y-
= 2.588m
b/2
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Figure 70:- Determination of the mean aerodynamic chord for the revised wing
option B.
As a verification check the mean aerodynamic chord length and location for both
inboard and outboard sections were calculated for wing option B from: -
Inboard:-
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 11.445 (1 + 0.46 + 0.46
2) / (1 + 0.46) = 8.736m
y- = (b/6) (1 + (2 / (1 + )
(3.26)
y- = (6.80/6) (1 + (2 x 0.46)) / (1 + 0.46) = 1.490m
For the Inboard wing calculated the value for c- chord length was lower than those
produced by the geometric method by 31mm, and for y- chord location from the
aircraft centreline the calculated value was lower by 4mm therefore in this case the
geometric analysis gave a valid approximation.
F-35C Outline to scale
for reference.
CRoot = 5.351m
CTip = 5.351m
CRoot = 11.445m
CTip = 2.289m
CL
CRoot
C- = 4.025m
C- = 8.767m
MAC
MAC
y-
= 1.494m
y-
= 1.494m b/2
50%
50%
CTip
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Outboard:-
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 5.351 (1 + 0.42 + 0.42
2) / (1 + 0.42) = 4.010m
y- = (b/6) (1 + (2 / (1 + )
(3.26)
y- = (6.92/6) (1 + (2 x 0.42)) / (1 + 0.42) = 1.494m
The calculated values were lower than those produced by the geometric method by
13.5mm for c- chord length, and 3.5mm for y
- chord location from the aircraft
centreline and therefore in this case the geometric analysis gave a valid
approximation.
In order to determine the length and location of the mean aerodynamic chord for the
combined wing the outboard chord length was subtracted from the inboard chord
length to give c- = 4.726m and the values of chord location were added together y
- =
2.984m, mapping these values onto the geometry gave a c- value = 6.098m which
was 1.372mm more than the calculated value which was attributed to the 28.50
difference in sweep angle between the leading and trailing edges of the outboard
wing.
Therefore for the purposes of this conceptual design study for the option B wing the
geometric value for c- and the calculated value for y
- were used for all further
analysis and empennage sizing, therefore for the combined option B wing planform:-
c- = 6.098m
y- = 2.984m
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Figure 71:- Determination of the mean aerodynamic chord for the revised wing
option C.
As a verification check the mean aerodynamic chord length and location was
calculated for wing option C from: -
c- = (2/3) cRoot (1 + +
2) / (1 + (3.25)
c- = (2/3) 10.100 (1 + 0.20 + 0.20
2) / (1 + 0.20) = 6.958m
y- = (b/6) (1 + (2 / 1 + ) (3.26)
y- = (13.72/6) (1 + (2 x 0.20)) / (1 + 0.20) = 2.668m
The calculated values were identical to those produced by the geometric method for
c- chord length, and y
- chord location from the aircraft centreline and therefore in
this case the geometric analysis gave a valid approximation.
F-35C Outline to scale
for reference.
50%
CRoot
CTip = 2.020m
CTip CRoot = 10.100m
MAC
C- = 6.958m
b/2
CL
y-
= 2.668m
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The values from all three revised wing options lead particulars were compared with
the baseline F-35C in table 7 below.
Table 7: - Wing configuration summary.
Parameter F-35C wing Option A
wing Option B
wing Option C
wing
Span (m) 13.11 13.72 13.72 13.72
CRoot (m) 7.412 11.473 11.445 10.100
CTip (m) 1.358 2.030 2.289 2.020
0.18 0.18 0.20 0.20
Sweep angle 340 520 550 600
Area m2 57.478 83.0 83.0 83.0
Loading (kg/m2)
560 388 388 388
Aspect ratio 2.99 2.27 2.27 2.27
MAC (m) 5.077 7.404 6.098 6.958
Location of MAC (m)
2.518 2.618 2.984 2.667
From the Mean Aerodynamic Chord (MAC) data obtained graphically and by
calculation the location of the wings Centre of Pressure (C of P) or Aerodynamic
Centre (a.c.) could be calculated which in subsonic flight is 25% of the MAC, but in
supersonic flight this moves back to along the wing to up to 48% of the wings MAC.
The location of the wing relative to the aircraft centre of gravity could now be
estimated as a percentage of the MAC for a given aircraft configuration e.g. 25 to
30% for a stable aircraft with aft tail: 40% (approx) for an unstable aircraft with aft
tail and 15 to 20% for an unstable aircraft with foreplanes (reference 21).
Table 8: - Wing configuration layout.
Parameter F-35C wing
Option A wing
Option B wing
Option C wing
a. c. Mach < 1.0.
1.269m 1.851m 1.525m 1.740m
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a. c. Mach > 1.0.
2.437m 3.554m 2.927m 3.340m
Location relative to
aircraft C of G. 2.031m 3.144m 2.439m 2.783m
¼ Chord sweep
23.90 43.90 47.60 55.20
Aircraft C of G position
8.798 10.111m 11.042 11.200
From this analysis the wings could be located on the fuselage relative to the aircraft
estimated C of G.
At this stage in the conceptual design in order to reduce fuselage drag still further the
concept of the twin tandem canopy was dropped in favour of positioning the UCAV
commander in a virtual enclosure directly behind the pilot occupying 2/3rds of the
volume devoted to the lift fan engine in the F-35B and the auxiliary fuel tank in the F-
35A and F-35C variants between the intake ducts. Ingress and egress would be via a
chevron hatch immediately above the enclosure. This had four major benefits for the
FB-24 configuration which were: - no net increase in drag from extra crew station:
original F-35 common forward hinging canopy retained: reduce risk of canopy
delamination and cracking as experienced on the larger F/A-22A canopy in SDD
phase, and greater durability in bird strike: reduced risk of EM emissions, and reduce
risk of flash blinding from special stores. Current ITAR restricted indicate that the
UCAV commander could fly both the FB-24 and the A-24 in the common virtual
environment.
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Figure 72:- Determination of the wing position relative to the aircraft C of G for
the revised wing option A mounted on the NB1 aircraft.
Figure 73:- Determination of the wing position relative to the aircraft C of G for
the revised wing option B mounted on the NB3 aircraft.
18.29m
10.111m
CL
CL
C of G
a.c.
MAC
1.851m
¼ Chord = 43.90 sweep
MAC
18.29m
¼ Chord = 47.60 sweep
1.525m
11.042m C of G
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Figure 74:- Determination of the wing position relative to the aircraft C of G for
the revised wing option C mounted on the NB5 aircraft.
The aerodynamic centre location for the wing configurations at subsonic speeds was
determined from: -
x ac = y- tan LE + 0.25 c
- (3.27)
And at supersonic speeds the aerodynamic location for the wing configurations was
determined from: -
x ac = y- tan LE + 0.48 c
- (3.28)
Where the leading edge of the reference wing root chord was taken as x = 0, and the
values quoted are actual length on of the MAC.
CL
MAC
¼ Chord = 55.20 sweep
18.29m
C of G 10.200m
a.c.
1.740m
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Empennage sizing: - From this data the empennage was sized for each configuration
using the methodology detailed above in section 3.2.3.3 as follows:-
Using the total area equations 3.22 and 3.23 with the historical cHT and cVT values
giving:-
SHT = 0.476 C-W SW / LHT (3.29)
SVT = 0.070 b W SW / LVT (3.30)
Where:-
LHT = 29.9% for F-35C of fuselage length (model measurements):
LVT = 28.5% for F-35C of fuselage length (model measurements):
SW = Planform area of each option:
bW = Wing span of each option:
C-W = Wing mean aerodynamic chord for each option.
(This was determined by geometric method shown in figure 55.)
For F/A-22A the moment was approximately 30% for the vertical tails LVT and
approximately 36% for the horizontal tail LHT were as for the YF-23 the approximate
value for the ruddervators was LRV = 40% from model measurements the moment
arm. Therefore the calculated value for each planform based on an initial as draw tail
the taking the distance between the 0.25% wing chord interaction point with the MAC
and the intersection of the 0.25% tail chord intersection with the reference tail MAC,
and the same method was used for the vertical tails which gave the following values:-
Option A: - LHT = 7.316m = 40% fuselage length:
LVT = 6.350m = 35% fuselage length.
Option B: - LHT = 6.526m = 36% fuselage length:
LVT = 5.082m = 27% fuselage length.
Option C: - LRV = 7.816m = 43% fuselage length:
LHT = 5.372m = 29% fuselage length.
These values were below the range of 45 – 50% given in reference 21 but were
consistent with F-35C and F/A-22A values but below that of the value measured for
YF-23 from both model and 1/48 scale drawings supplied by the Collectaire accurate
resin model company USA, with the YF-23 kit, also supported by drawing
measurements from reference 16.
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For the conventional four tail versions of the FB-24 and A-24 aircraft namely NB1:
NB3: and NB4 the tail sizings were as follows:-
NB1 wing option A:-
SHT = 0.4046 C-W SW / LHT (3.29)
SHT = 0.4046 (7.404 x 83) / 7.316 = 34.0m2
Volume coefficient reduced by 15% was all moving (reference 21) and area
reduced by 10% as aircraft have “active” FBL computerised control system
(reference 21).
Therefore total tail horizontal area SHT = 30.6m2
SVT = 0.070 b W SW / LVT (3.30)
SVT = 0.070 (13.72 x 83) / 6.350 = 12.6m2
Therefore total vertical tail area SVT = 12.6m2
NB3 wing option B:-
SHT = 0.4046 C-W SW / LHT (3.29)
SHT = 0.4046 (6.098 x 83) / 6.526 = 31.4m2
Volume coefficient reduced by 15% was all moving (reference 21) and area
reduced by 10% as aircraft have “active” FBL computerised control system
(reference 21).
Therefore total horizontal tail area SHT = 28.3m2
SVT = 0.070 b W SW / LVT (3.30)
SVT = 0.070 (13.72 x 83) / 5.082 = 15.7m2
Therefore total vertical tail area SVT = 15.7m2
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NB5 wing option C:-
SHT = 0.4046 C-W SW / LHT (3.29)
SHT = 0.4046 (6.958 x 83) / 7.816 = 29.9m2
Volume coefficient reduced by 15% was all moving (reference 21) and area
reduced by 10% as aircraft have “active” FBL computerised control system
(reference 21).
Therefore total horizontal tail area SHT = 26.9m2
SVT = 0.070 b W SW / LVT (3.30)
SVT = 0.070 (13.72 x 83) / 5.372 = 14.8m2
Therefore total vertical tail area SVT = 14.8m2
The values for all three options for total HT and VT area and their ratios to the wing
area are shown in table 9, together with the individual tail areas and the ruddervator
area.
As a reality check the respective SHT / SW and SVT / SW ratio values for the three
options were compared with the respective SHT / SW and SVT / SW ratio values for
similar size four tail combat aircraft recorded in table 4.1 of reference 19.
For the V tail versions of the FB-24 / A-24 aircraft namely NB2, NB4, and NB6 the
tail sizings were as follows using the cumulative size method from reference 21 page
125:-
NB2 wing option A:-
For a V tail the HT and VT areas were estimated as above then added together to give
the same total surface area for the two ruddervators as would have been required for
two separate horizontal tails, and two separate vertical tails: -
SRV = SHTc + SVTc (3.31)
SRV = 30.6 + 12.6 = 43.2m2 – 4.3m
2 = 38.8m
2
Total area reduced by 10% as aircraft have “active” FBL computerised control
system, and HT volume coefficient reduced by 15% because the ruddervator
surfaces are all moving (reference 21).
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NB4 wing option B:-
For a V tail the HT and VT areas were estimated as above then added together to give
the same total surface area for the two ruddervators as would have been required for
two separate horizontal tails, and two separate vertical tails: -
SRV = SHTc + SVTc (3.31)
SRV = 28.3 + 15.7 = 44.0m2 – 4.4m
2 = 39.6m
2
Area reduced by 10% as aircraft have “active” FBL computerised control
system, and HT volume coefficient reduced by 15% because the ruddervator
surfaces are all moving (reference 21).
NB6 wing option C:-
For a V tail the HT and VT areas were estimated as above then added together to give
the same total surface area for the two ruddervators as would have been required for
two separate horizontal tails, and two separate vertical tails: -
SRV = SHTc + SVTc (3.31)
SRV = 26.9 + 14.8 = 41.7m2 – 4.2m
2 = 37.5m
2
Area reduced by 10% as aircraft have “active” FBL computerised control
system, and HT volume coefficient reduced by 15% because the ruddervator
surfaces are all moving (reference 21).
Comparing the total tail and total wing area ratios calculated for the three options in
table 9 within those of the F-15: F/A-18: and MiG 25 showed that the tails were about
average for both the HT and VT for all options. These were deemed suitable for the
initial layout, but were possibly conservative when compared with the F-35C and
F/A-22A, as detailed below.
The values for F-35C reference horizontal tails were 11.017m2 per tail equal to
22.034m2 for total horizontal tail area, and the F-35C reference vertical tails had a
value of 4.547m2 per tail equal to 9.094m
2 based on model measurements of the 230-5
OML.
The values for the F/A-22A reference horizontal tails were 12.630m2 per tail equal to
25.26m2 for total horizontal tail area, and the F/A-22A reference vertical tails had a
value of 16.65m2 per tail equal to 33.300m
2 based on public domain data.
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Table 9: - Tail sizing results for HT / VT and RV configurations.
Option.
Total tail areas and ratios to total wing areas.
HT (m2) SHT / SW VT (m2) SVT / SW RV (m2) SRV / SW
A 30.6 0.37 12.6 0.15 38.8 0.47
B 28.3 0.34 15.7 0.18 39.6 0.48
C 26.9 0.32 14.8 0.17 37.5 0.45
NB1 wing option A:-
Individual reference horizontal tail surface area = 15.3m2:
Individual reference vertical tail surface area = 6.3m2.
NB3 wing option B:-
Individual reference horizontal tail surface area = 14.2m2:
Individual reference vertical tail surface area = 7.9m2.
NB5 wing option C:-
Individual reference horizontal tail surface area = 13.5m2:
Individual reference vertical tail surface area = 7.4m2.
NB2 wing option A:-
Individual reference ruddervator tail surface area = 19.4m2.
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NB4 wing option B:-
Individual reference ruddervator tail surface area = 19.8m2.
NB6 wing option C:-
Individual reference ruddervator tail surface area = 18.8m2.
F-35C baseline reference aircraft:-
These results were consistent with those from the F-35C (230-5) reference model
which had values of: -
SHT / SW = 0.38:
SVT / SW = 0.20:
Individual reference horizontal tail surface area = 11.02m2:
Individual reference vertical tail surface area = 5.61m2:
Wing reference area = 57.5m2.
From the above empennage sizing exercise the two layouts investigated namely twin
vertical tails and twin horizontal tails, or twin ruddervators was a wise choice not only
for reasons of signature and drag reduction. There were also structural and control
effectiveness implications, this was because a tall swept fin loses a large portion of its
effectiveness at high speed as a result of aerodynamic loads twisting the tip. Typical
values (quoted in reference 25) of fin efficiency loss from rudder distortion and
spanwise twist resulting from bending were 20 to 25%. The use of twin fins can
overcome this problem of tall flexible single fins, provided that they are set far
enough apart laterally to overcome mutual biplane interference.
The main supporting argument for the implementation of twin fins or ruddervators for
the FB-24 and A-24 empennage was that there would always be one vertical surface
in relatively clean flow to provide directional stability regardless of the aircrafts AoA
/ sideslip combination. The twin fin layout was also best suited for high supersonic
speeds, when mutual interference is eliminated as the Mach lines from each fin pass
behind its neighbour, this was one of the reasons for their wide spread adoption on
high speed fighter aircraft such as the F-15, F-18, F-14, F/A-22A, F-35, and Su-27 /
30, MiG-25 / 29. Although at low Mach numbers, the twin fin layout would not be the
most efficient method of improving directional stability. Additionally as explained
above when canted outwards to match the fuselage side slope angle their signature
contribution would be considerably reduced, this was exemplified by the F/A-22A,
YF-23 and F-35 aircraft, although alternatively they could be canted inward as on the
Have Blue, and the SR-71 which could have the effect of shielding the engine exhaust
as in the case of the Sabre long range bomber conceptual design the author is working
on as a private study, outside both Cranfield and BAE Systems.
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The full aircraft OML models resulting form the above empennage sizing are shown
below in figures 75 through 81 below with their reference sizing data.
Figure 75: - NB1:- FB-24 wing option A, illustrating complete OML model with
sized vertical and horizontal tails and final wing planform, ready for control
surface sizing.
Figure 76: - NB2:- FB-24 wing option A, illustrating complete OML model with
sized ruddervators and final wing planform, ready for control surface sizing.
Vertical tails area = 6.3m2 each including
rudder. 520 sweep leading edge and trailing
edge sweep of 650.
Spigot mounted horizontal tails area = 15.3m
2 each.
520 sweep leading edge
and 150 tailing edges.
Total reference wing area = 83m2.
Total reference wing area = 83m2.
Spigot mounted „Ruddervator‟ tails exposed
area = 5.500m2 each. 55
0 leading edge sweep.
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Figure 77: - NB3:- FB-24 wing option B, illustrating complete OML model with
sized vertical and horizontal tails and final wing planform, ready for control
surface sizing.
Figure 78: - NB4:- FB-24 wing option B, illustrating complete OML model with
sized ruddervators and final wing planform, ready for control surface sizing.
Spigot mounted horizontal tails area = 14.2m
2 each. Leading edge
sweep 550 and 28.5
0 trailing edge
sweep.
Vertical tails area = 7.9m2
each including rudder. 550
leading edge sweep and 840
trailing edge sweep.
Total reference wing area = 83m2.
Total reference wing area = 83m2.
Spigot mounted „Ruddervator‟ tails exposed area = 5.129m2 each.
Leading edge sweep 400.
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Figure 79: - NB5:- FB-24 wing option C, illustrating complete OML model with
sized vertical and horizontal tails and final wing planform, ready for control
surface sizing.
Figure 80: - NB6:- FB-24 wing option C, illustrating complete OML model with
sized ruddervators and final wing planform, ready for control surface sizing.
Vertical tails area = 7.4m2 each including rudder.
Leading edge sweep 600 and 83
0 trailing edge
sweep.
Total reference wing area = 83m2.
Spigot mounted horizontal tails area = 13.5m
2 each. Leading edge
sweep 600 and trailing edge
sweep of 290.
Spigot mounted „Ruddervator‟ tails exposed area = 5.445m2 each.
Leading edge sweep 550.
Total reference wing area = 83m2.
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Figure 81: - F-35C baseline illustrating complete OML model with reference
sized vertical and horizontal tails and wing planform, for comparison with the
NB aircraft sizing above.
Conventional horizontal tail layouts: - The horizontal tail surface leading edges of
NB options 1, 3, and 5 shown above in figures 75, 77, and 79 respectively were
planform aligned with the wing leading edges for frontal signature reduction, as were
the trailing edges of NB 3 and NB5.
However trailing edge alignment was not possible on option 1 because this would
have resulted in a large portion of the tail overhanging the spigot with an impact on
spigot loads and on the sizing of the actuator required to drive the tail surface (current
EHA shown in figure 84 was at the limits contained within the boom IML), increasing
rear fuselage weight, attempts were made to move the surface forward into a clipped
flap as in the F-35C, and F/A-22A but to achieve a structurally safe root chord to tip
chord for the exposed surface the trailing edge spar of the wing torsion box would
have to be kinked which in the authors view was unacceptable. Therefore NB1 failed
on one count of empennage planform alignment although inverse alignment was
achieved with the trailing edge of the port horizontal tail being a sweep continuation
of the starboard wing and vice versa for the starboard horizontal tail trailing edge
which may have some RCS merit.
None of the horizontal tails were given a dihedral slope like that which has been
applied to the current F-35C in publicly released drawings, or that applied to the X-
35C Concept Demonstrator Aircraft, instead they were aligned with the wing to
reduce drag and potential radar reflections. The exposed areas of these surfaces were
consistent with F-35C and F/A-22A as well as historical data from reference 19 table
4.1, and are presented in table 10 and compared with the F-35C measurements, and
F/A-22A published values.
Vertical tails area = 5.61m2 each including rudder.
Flipper mounted horizontal
tails area = 11.02m2 each. Total reference wing area = 57.5m
2.
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Conventional vertical fin layouts: - The vertical tails of NB options 1, 3, and 5
shown above in figures 75, 77, and 79 respectively were all canted out by 250 to the
aircraft Z, X, plane in common with the fuselage side skins to retain side aspect
planform alignment and to overcome any risk of mutual biplane interference at
subsonic speeds. This represented a decrease in the side skin cant angle from that of
the F-35C, and was a fuselage cross sectional area reduction measure to reduce wave
drag and represented the maximum reduction possible without adversely impacting on
the volume of the internal weapons bays or fuel volume and is covered in section 4 of
this thesis.
The NB1 vertical tails shown in figure 75 above were of a swept planform like that of
the F-35C, but with a leading sweep angle of 520 to match the wing leading edge
sweep angle, and a trailing edge angle of 650 which again were departures from the
sweep angles employed on the F-35C, which impacts on all major side door and
access panel chevron lines, but retains mirror commonality with the intake lips. These
changes were driven by drag reduction, exposed tail surface area given in table 10,
and structural weight considerations.
The NB3 vertical tails shown in figure 77 above were of a trapezoidal planform
similar to that of the F/A-22A, but with a leading edge sweep of 550 to match the
leading edge sweep angle, and a trailing edge sweep angle of 840 this latter angle was
driven by the need to keep the exposed tail volume on the aft fuselage booms with
adequate volume in the boom itself below the rudder to house the rudder EHA
actuator system (shown in figure 84). The trapezoidal planform was selected for NB3
vertical tails to reduce the risk of flutter which could have occurred if the larger tail
area of NB3 had used a swept surface planform similar to NB1.
The large vertical tail area calculated for NB5 shown in figure 79 raised the same
issues as those raised for NB3 which resulted in adoption of the trapezoidal planform
with a leading edge sweep of 600 common to the leading edge sweep angle, and a
trailing edge sweep angle of 830 to meet rudder actuator housing requirements.
Ruddervator layouts: - All of the ruddervator layouts NB2, NB4, and NB6 shown in
figures 76, 78 and respectively were of a similar planform to that employed on the
YF-23 (shown in figure 59 above), in having a continuous leading edge sweep and
chevron trailing edge, of two sweep angles, a lower section positive angle and an
upper section negative sweep angle. All of the ruddervators were canted outwards at
500 to the aircraft Z, X, plane to give the greatest pitch control authority and to avoid
acute corners or right angles in side elevation or front view. These all-flying
ruddervator tails were to be rotated about a single spigot by the ruddervator EHA
actuators, and could be driven differentially for yaw and roll control. These canted
surfaces because of their location on the aft fuselage booms would also act as shield
for the engine exhaust in all angles except immediately above, behind or below the
aircraft. On layouts NB2, NB4, and NB6, the wing trailing edge controls also provide
roll control and lift augmentation, and in combination with the ruddervators function
as speed brakes and rudders. For straight line deceleration, the FCS commands the
outer ailerons to deflect up and the inboard flaps to deflect down, thus producing a
decelerating force but creating no other moments.
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Yaw control would be provided by doing this on one side only. This augmentation
permitted a reduction in the exposed area of the ruddervator control below that which
would be required if the ruddervators had no augmentation, for example the exposed
ruddervator area for the YF-23 which employed the same wing control surface
augmentation to that described above was 4.739m2 for each ruddervator and the YF-
23 had a reference wing area of 87m2 according to public domain data (Reference 26:
- Advanced Tactical Fighter to F-22 Raptor Origins of the 21st Century Air
Dominance Fighter: by Aronstein D. C., Hirschberg M. J., and Piccirillo A. C.:
Published by American Institution of Aeronautics and Astronautics 1998). If no
augmentation had been used from the cumulative area calculations based on reference
21 methodology the ruddervator total reference area would have been 25m2 resulting
in a larger exposed area. Similar wing control surface augmentation was used on the
F-117A to reduce ruddervator tail size, (Reference 27): - Have Blue and The F117A
evolution of the “Stealth Fighter”: by Aronstein D. C. and Piccirillo A. C.: Published
by American Institute of Aeronautics and Astronautics 1997). Subsequent assessment
of straight line deceleration and positive yaw control indicated that the effect of the
ailerons could be incorporated into single unit flaperons and toeing in the ruddervators
(i.e. both ruddervators turning in towards the centreline) in combination with
downward deflection of the single unit flaperons could produce the same breaking
effect, therefore the need for the ailerons for the NB1 and NB2 configurations was
questioned, as described below.
The NB2 ruddervator geometry was driven by the need to reduce to a minimum
control surface overhang on the aft fuselage boom in order to position the diving
spigot as close to the centre of the ruddervator as possible to reduce the loads on the
spigot and the driving forces required from the actuator, (with the corresponding
actuator size requirement shown in figure 82). In order to achieve this the leading
edge sweep angle of 550 was selected along with a trailing edge sweep of -48
0 for the
top section and 260 for the lower section imparting chevron trailing similar to that of
the YF-23 ruddervators, and although alignment with F-35C features was not
maintained for the new build aircraft all the features would be changed to match the
ruddervator geometry. The exposed planform area was 5.5m2 for each tail, which was
considered adequate with the previously mentioned augmentation.
The same drivers were applied to the ruddervators for NB4, which resulted in a
leading edge sweep of 400 and a trailing edge chevron with a -52
0 sweep for the upper
section and a 270 for the lower section. The exposed planform area for this surface
was 5.1m2 for each tail.
For NB6 the same planform as NB2 was adopted to elevate the issues discussed above
in terms of actuators and overhang, and this geometry was modified to give an
exposed planform area of 5.4m2 for each tail surface.
In order to prevent clashes between the ruddervators and the wing trailing edge
control surfaces a clearance separation of 0.4m aft of the reference wing trailing edge
was established to permit free rotation of the ruddervators and the trailing edge flaps,
without the need for flap trailing edge cutouts as used on the F/A-22A and F-35C.
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Figure 82: - EHA actuator for horizontal tail and ruddervator actuation. Source
authors private collection.
Figure 83: - EHA actuator for rudders actuation on conventional vertical tails,
and trailing edge flap and aileron actuation on all aircraft configurations. Source
authors private collection.
Length = 6.94m.
Height = 3.45m.
Height = 1.12m.
Length = 2.75m.
Weight = 116kg.
Weight = 32kg.
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Table 10: - Exposed tail areas for all configurations.
Configuration Individual
Horizontal tail area in m2.
Individual Vertical tail area
in m2.
Individual Ruddervator tail
area in m2.
NB1 6.7 5.8 N/A
NB2 N/A N/A 5.5
NB3 5.8 7.3 N/A
NB4 N/A N/A 5.1
NB5 5.7 6.6 N/A
NB6 N/A N/A 5.4
F-35C (ref) 5.8 4.6 N/A
Control surface and high lift device sizing: - The next stage before full aircraft drag
polar analysis was to size the wing control surfaces (and the rudder for the
conventional four tail aircraft layouts), which was addressed below, using the F-35C
sizes to determine a wing area to control surface area relationship, as an initial starting
point for analysis. The following control surfaces were sized in this manner: - rudder
based on tail area: trailing edge flaps based on wing area: ailerons based on wing area:
and finally the leading edge flap. The results in terms of the ratio of control surface
area to wing reference area were as follows for the F-35C: -
Total trailing edge flap area to reference wing area = 0.10 = 5.76m2:
Total aileron area to reference wing area ratio = 0.05 = 2.70m2:
Total leading edge flap to reference wing area ratio = 0.08 = 4.672m2:
Total rudder area to total vertical tail reference area ratio = 0.24 = 2.724m2.
The final sizing of these control surfaces would be based upon dynamic analysis of
control effectiveness, including structural bending and control system effects,
however these ratios taken from a real aircraft of similar type and compared with data
from table 4.1 in reference 19 were considered by the author as being adequate for
this initial concept design study with constitutes a proposal submission.
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The reference wing area of the F-35C 230-5 OML represented 69.25% of the
reference wing area of all Advanced Interdiction Aircraft configurations i.e. 57.478m2
for the F-35C compared to 83m2, therefore for the purposes of this study the sizes of
the reference F-35C control surfaces actual sizes should have increased by a factor of
0.3075 = 30.75% to yield the same effectiveness on the larger AIA wing. However
when the above ratios were applied the sizes were larger than predicted as follows: -
Ailerons = 83 x 0.05 = 4.15 = 2.075m2 (per aileron):
Trailing edge flaps = 83 x 0.10 = 8.30 = 4.150m2 (per trailing edge flap):
Leading edge flap = 83 x 0.08 = 6.64 = 3.32m2 (per leading edge flap).
For the vertical tailed configurations rudder size determination used the same scaling
ratio method applied to the exposed vertical tail area each four tail configuration as
follows:-
NB1 rudder area = 5.8 x 0.29 = 1.682m2 (per rudder):
NB3 rudder area = 7.3 x 0.29 = 2.117m2 (per rudder):
NB5 rudder area = 6.6 x 0.29 = 1.914m2 (per rudder).
From these results the control surfaces were modelled onto the exposed wings of
surface models of each configuration with the leading edge flaps occupying the
leading edges from root to tip and the trailing edge flaps occupying the inboard 50%
of the wing trailing edge, and the ailerons occupying the outboard 50% of the wing
trailing edge, as shown in figures 85 to 90. This enabled location of the leading and
trailing edge spars for each configuration and defined the outer boundaries of the
wing torsion box. The ailerons and trailing edge flaps were to be driven by a single
MOOG EHA actuator each, of the type shown in figure 83. Were as the leading edge
flaps were driven by four segmented MOOG rotary actuator of the type shown below
in figure 84.
The rudders were driven by same type of EHA as used for the wing trailing edge
control surfaces. The rudders began at 10% of the exposed trailing edge of the vertical
tails for configurations NB3 and NB5 to clear the horizontal tail spigot bulge in the aft
fuselage booms, and extended to the tip of the tail. Because of the twin tail layout loss
of rudder efficiency at high speeds with this rudder span was not considered an
important issue for reasons explained above, as it would have been on a single tail
aircraft probably would have required an all moving vertical tail like the TSR-2, or
Vigilante. The NB1 rudders began at 5% of the trailing edge chord to prevent OML
clash and extended over the full span of the tail.
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Figure 84: - The multi segment rotary actuator selected to drive the leading edge
flaps with four actuators located at ¼ span intervals along the leading edge, for
even load distribution. (Note assembly represents a single actuator reference
MOOG).
Figure 85: - Flight control and high lift device integration on the NB1/NB2 wing
planform illustrating (a) F-35C based layout and (b) Revision for FB-24 / A-24.
T/E Flap
L/E Flap
Tip Aileron
L/E Flap Flaperon
(a) Three surfaces.
(b) Two surfaces.
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During design maturation studies of the „Option A‟ wing control surfaces as sized
above the aileron could only be accommodated as a spigot mounted moving wing tip
device as a result of the large size of the trailing edge flap, on the relatively short span
exposed wing as shown above in figure 85(a). This lead to concerns about: - the
effectiveness of the aileron as an aerodynamic control: the difficulty of integrating an
actuator: and structural weight penalty of reinforcing the torsion box to withstand high
torque at the wing tip rib and spar tips resulting from the ailerons actuation. The
trailing edge flap was judged by the author to extend for enough outboard and to be of
sufficient surface area to perform both the functions of a flap and an aileron for the
„Option A‟ thereby being a flaperon with the same area ratio to the reference wing as
the flaperons of the F-35A, F-35B and F-16 aircraft and located over the same extent
of the trailing edge exposed span, as shown in figure 85(b). Therefore the aileron for
this wing option was dropped for both the NB1 and NB2 configurations as shown in
figures 86 and 87. This had the benefits of reducing the weigh of the wing both in
terms of structural weight and systems weight (through elimination of two wing
actuators), although this had to be traded against provision of two small fuselage
mounted ventral airbrakes for straight line deceleration, on landing which had
individual areas of one quarter of that for an aileron as sized above, and their
actuators, for the NB1 configuration.
Figure 86: - FB-24 NB1 Configuration CATIA V5 solid model control surface
layout, illustrating the size and location of Flaperon control surfaces and
Leading edge flap high lift device in the cruise condition i.e. not deflected for
manoeuvre. The high swept wing angle to keep within the Mach wave cone
reduced the trailing edge available necessitating the adoption of Flaperons.
These surfaces are crudely sized using comparative area assessment and require
a full – scale analysis as described above which was recommended as further
work.
L/E Flap area = 3.515m2 each
T/E Flaperon area = 4.150m2 each
Rudder area = 1.682m2 each
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Figure 87: - FB-24 NB2 Configuration CATIA V5 solid model control surface
layout, illustrating the size and location of control surfaces and high lift devices
in the cruise condition i.e. not deflected for manoeuvre. Wing control surfaces
and high lift devices are identical to NB1. These surfaces are crudely sized using
comparative area and require a full – scale analysis as described above which
was recommended as further work.
The proposed Control Surface Authority of these devices was estimated based on
similar legacy aircraft to be as follows: -
L/E Flaps: - 3.00 Up / 40.0
0 Down (Normal to Hinge Line):
T/E Flaperons: - 350 Up / 35
0 Down (Normal to Hinge Line):
Horizontal Tails (NB1): - 300 Up / 30
0 Down (Normal to Spigot Centre Line):
Rudders (NB1): - 300 Inboard / 30
0 Outboard (Normal to Hinge Line):
Ruddervators (NB2): - 250 Inboard (limited by structure) / 30
0 Outboard
(Normal to Spigot Centre Line).
L/E Flap area = 3.515m2 each
T/E Flaperon area = 4.150m2 each „Ruddervator‟ tails exposed
area = 5.500m2 each.
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Figure 88: - FB-24 NB3 Configuration CATIA V5 solid model control surface
layout, illustrating the size and location of control surfaces and high lift devices
in the cruise condition i.e. not deflected for manoeuvre. These surfaces are
crudely sized using comparative area and require a full – scale analysis as
described above which was recommended as further work.
As shown in figures 88 and 89 the NB3 and NB4 configurations high wing leading
edge sweep of 550 reduced the effective trailing edge span available for the larger
control surfaces sized above, as was the case for configurations NB1 and NB2 and
this was further compounded by the novel cranked arrow wing planform of the wing,
which constrained the flap to the inboard wing. This resulted in a trailing edge flap
area only 58% that of the estimated value based on the ratio of flap size to wing
reference area. However the aileron area was increased by 13% over target and the
leading edge flap area was increased 10% to compensate in rotation and landing, and
these size increases were judged by the author as being the maximum values that the
outboard wing structure could withstand without large structural weight penalties.
The larger than estimated aileron area was also used to compensate for the fact that
the rudders were 1% smaller than estimated the target value due to the vertical tail
geometry and structural requirements of flutter as well as the systems to be mounted
on the tip rib under the tip rib fairing.
The NB3 configuration would in the view of the author require the addition of ventral
airbrakes as in the case of the NB1 configuration for straight line deceleration, where
as the NB4 configuration would employ the same breaking technique as the YF-23
and NB2 configuration described above, and would therefore not require the addition
of airbrakes.
L/E Flap area = 3.673m2 each
Aileron area = 2.376m2 each
T/E Flap area = 2.444m2 each
Rudder area = 2.091m2 each
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Figure 89: - FB-24 NB4 Configuration CATIA V5 solid model control surface
layout, illustrating the size and location of control surfaces and high lift devices
in the cruise condition i.e. not deflected for manoeuvre. These surfaces are
crudely sized using comparative area and require a full – scale analysis as
described above which was recommended as further work.
The proposed Control Surface Authority of these devices was estimated based on
similar legacy aircraft to be as follows: -
L/E Flaps: - 3.00 Up / 40.0
0 Down (Normal to Hinge Line):
T/E Flaperons: - 350 Up / 40
0 Down (Normal to Hinge Line):
Aileron s: - 250 Up / 25
0 Down (Normal to Hinge Line):
Horizontal Tails (NB1): - 300 Up / 30
0 Down (Normal to Spigot Centre Line):
Rudders (NB1): - 300 Inboard / 30
0 Outboard (Normal to Hinge Line):
Ruddervators (NB2): - 250 Inboard (limited by structure) / 30
0 Outboard
(Normal to Spigot Centre Line).
L/E Flap area = 3.673m2 each
Aileron area = 2.376m2 each
T/E Flap area = 2.444m2 each
Ruddervator tails exposed area = 5.129m2 each
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Figure 90: - FB-24 NB5 Configuration CATIA V5 solid model control surface
layout, illustrating the size and location of control surfaces and high lift devices
in the cruise condition i.e. not deflected for manoeuvre. These surfaces are
crudely sized using comparative area and require a full – scale analysis as
described above which was recommended as further work.
Once again the penalty for a high leading edge sweep angle required to reduce wave
drag was a reduction in trailing edge span available for high lift and flight control
surface, and the 600 sweep of configurations NB5 and NB6 was the most severe. Also
attempting to extend these surfaces chordwise to recover the estimated incurred
structural weight penalties due to the resulting reduction in torsion box chord leading
to the torsion and bending loads being carried by fewer structural members which in
turn would need to be increase in section to retain the same overall stiffness, and this
would increase the wings structural weight. After much iteration the layout shown in
figures 90 and 91 was developed which in the authors view was the most practical
compromise between, control and high lift device requirements and the stiffness and
fuel capacity requirements for the wing torsion box.
In this layout the ailerons were modelled at 77% of the estimated target value, and the
trailing edge flaps were modelled at 81% of the estimated target values, although
some small redress was achieved by increasing the leading edge flap area to 109%,
and increasing the rudder area to 102% of the estimated target values. However
longitudinal stability a key requirement in a bombing platform could still be an issue
and requires further investigation.
In the authors opinion both NB5 and NB6 would require ventral air breaks which in
addition to simultaneous deployment for straight line deceleration, would be deployed
differentially to aid the ailerons in roll control.
T/E Flap area = 3.388m2 each
L/E Flap area = 3.651m2 each
Aileron area = 1.617m2 each
Rudder area = 1.950m2 each
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Figure 91: - FB-24 NB6 Configuration CATIA V5 solid model control surface
layout, illustrating the size and location of control surfaces and high lift devices
in the cruise condition i.e. not deflected for manoeuvre. These surfaces are
crudely sized using comparative area and require a full – scale analysis as
described above which was recommended as further work.
The proposed Control Surface Authority of these devices was estimated based on
similar legacy aircraft to be as follows: -
L/E Flaps: - 3.00 Up / 40.0
0 Down (Normal to Hinge Line):
T/E Flaperons: - 350 Up / 40
0 Down (Normal to Hinge Line):
Aileron s: - 250 Up / 25
0 Down (Normal to Hinge Line):
Horizontal Tails (NB1): - 300 Up / 30
0 Down (Normal to Spigot Centre Line):
Rudders (NB1): - 300 Inboard / 30
0 Outboard (Normal to Hinge Line):
Ruddervators (NB2): - 250 Inboard (limited by structure) / 30
0 Outboard
(Normal to Spigot Centre Line).
T/E Flap area = 3.388m2 each
L/E Flap area = 3.651m2 each
Aileron area = 1.617m2 each
Ruddervators exposed
area = 5.445m2 each
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The author had reservations with both the option B and C wing control surfaces
abilities to counter any „Mach Tuck‟ (nose down attitude at transonic acceleration,
and extreme nose up attitude on deceleration which is a phenomenon high wing
sweep short wing span long fuselage configurations reference 25) which may occur
on acceleration to penetration speed, and deceleration on exiting the threat area.
Detailed sizing analysis was required to establish the capabilities of the wing and trail
control surfaces for theses configurations which could not be undertaken within the
time scale of this thesis.
Table 10 below brings together all of the control and high lift device modelled areas
for all configurations and compares this with the estimated target values.
Table 11: - Installed control surface sizings single component areas.
Control Surface.
Wing Option A (NB1 and NB2)
in m2
Wing Option B (NB3 and NB4)
in m2
Wing Option C (NB5 and NB6)
in m2
Target areas in m2
Aileron N/A 2.376 1.617 2.075
L/E Flap 3.515 3.673 3.651 3.320
T/E Flap 4.150 2.444 3.388 4.150
Rudder NB1 1.682 N/A N/A 1.682
Rudder NB3 N/A 2.091 N/A 2.117
RudderNB5 N/A N/A 1.950 1.914
From table 10 and the above it is clear that the high wing sweep and root chord
constraints had adverse effects on the trailing edge areas of each wing planform which
could be devoted to these surfaces and high lift devices. However the target areas
could be met or exceeded for the leading edge flap high lift device for all planforms.
The rudder areas of the three conventional tail options met the target for NB1, and
exceeded the target for NB5, but NB3 was just below target.
As stated above the validity of these target values needs to be established by dynamic
analysis of control effectiveness, including structural bending and control system
effects which would form part of a further more in depth study than undertaken in this
thesis, and this concluded the configuration OML sizing.
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3.2.5: - Common FB-24 and A-24 supersonic drag analysis for Jet306 analysis
selection.
For the specified mission both manned and unmanned configurations of the Advanced
Interdiction Aircraft would spend a large proportion of flight time at supersonic
speed, therefore it was important that the aerodynamic design concentrated on
reduction of the wave drag of the aircraft.
The initial aerodynamic estimation concerned the prediction of aircraft drag and lift
and for the three configurations investigated the main focus of drag was on the
supersonic wave drag (CDw) estimation. Using the supersonic drag analysis
methodology from reference 22 pages 132-133 and Swet values obtained from
measurements taken from the CATIA V5 surface models, the values for MCrit, CDw,
and k1, for the NB1, through NB6 configurations were determined for MCrit, M = 1.4,
and M = 1.6.
Supersonic drag analysis for NB1 / NB2:-
Step 1: - MCrit, from equations 3.14 and 3.17:-
MCrit (unswept) = 1.0 – 0.065[100(t max /c)] 0.6
= 1.0 – 0.065(5.9)0.6
= 0.82
MCrit = 1.0 – cos0.6
0.25c (1.0 - MCrit (unswept)) = 1.0 – cos0.6
43.90 (1 – 0.82) = 0.852
MCDo max = 1 / (cos0.2
LE) = 1 / (cos0.2
520) = 1.10
Step 2: - CDw, from supersonic zero lift drag equation 3.32:-
CDw = 4.5 / S (Amax / l) 2 EWD (0.74+0.37 cos LE) [1-0.3M-MCDo] 3.32
Where Amax = 4.47m2:
l = 18.29m
LE = 520
S = 83m2
MCDo = 1.10
EWD = 1.4
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CDw=4.5 /83 (4.47/18.29)2 EWD (1.4) [1.0 – 0.31.10-1.10] = 0.01424
CDw=4.5 /83 (4.47/18.29)2 EWD (1.4) [1.0 – 0.31.40-1.10] = 0.01190
CDw=4.5 /83 (4.47/18.29)2 EWD (1.4) [1.0 – 0.31.60-1.10] = 0.01122
Step 3: - k1 from the supersonic drag due to lift equation 3.33:
k1 = [AR (M2 – 1) / (4ARM
2 – 1) – 2] cos LE 3.33
k1 = [2.27(1.102 – 1) / (4 x 2.27 (1.10
2 – 1) – 2] cos 52
0 = 0.13581
k1 = [2.27(1.402 – 1) / (4 x 2.27 (1.40
2 – 1) – 2] cos 52
0 = 0.13701
k1 = [2.27(1.602 – 1) / (4 x 2.27 (1.60
2 – 1) – 2] cos 52
0 = 0.23700
Supersonic drag analysis for NB3 / NB4:-
Step 1: - MCrit, from equations 3.14 and 3.17:-
MCrit (unswept) = 1.0 – 0.065[100(t max /c)] 0.6
= 1.0 – 0.065(5.9)0.6
= 0.82
MCrit = 1.0 – cos0.6
0.25c (1.0 - MCrit (unswept)) = 1.0 – cos0.6
47.60 (1 – 0.82) = 0.823
MCDo max = 1 / (cos0.2
LE) = 1 / (cos0.2
550) = 1.00
Step 2: - CDw, from supersonic zero lift drag equation 3.32:-
CDw = 4.5 / S (Amax / l) 2 EWD (0.74+0.37 cos LE) [1-0.3M-MCDo] 3.32
Where Amax = 4.47m2:
l = 18.29m
LE = 550
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S = 83m2
MCDo = 1.10
EWD = 1.1
CDw=4.5 /83 (4.47/18.29)2 EWD (1.1) [1.0 – 0.31.00-1.00] = 0.01230
CDw=4.5 /83 (4.47/18.29)2 EWD (1.1) [1.0 – 0.31.40-1.00] = 0.00906
CDw=4.5 /83 (4.47/18.29)2 EWD (1.1) [1.0 – 0.31.60-1.00] = 0.00859
Step 3: - k1 from the supersonic drag due to lift equation 3.33:
k1 = [AR (M2 – 1) / (4ARM
2 – 1) – 2] cos LE 3.33
k1 = [2.27(1.002 – 1) / (4 x 2.27 (1.00
2 – 1) – 2] cos 55
0 = 0.00000
k1 = [2.27(1.402 – 1) / (4 x 2.27 (1.40
2 – 1) – 2] cos 55
0 = 0.18124
k1 = [2.27(1.602 – 1) / (4 x 2.27 (1.60
2 – 1) – 2] cos 55
0 = 0.21745
Supersonic drag analysis for NB5 / NB6:-
Step 1: - MCrit, from equations 3.14 and 3.17:-
MCrit (unswept) = 1.0 – 0.065[100(t max /c)] 0.6
= 1.0 – 0.065(5.9)0.6
= 0.82
MCrit = 1.0 – cos0.6
0.25c (1.0 - MCrit (unswept)) = 1.0 – cos0.6
55.20 (1 – 0.82) = 0.821
MCDo max = 1 / (cos0.2
LE) = 1 / (cos0.2
600) = 1.04
Step 2: - CDw, from supersonic zero lift drag equation 3.32:-
CDw = 4.5 / S (Amax / l) 2 EWD (0.74+0.37 cos LE) [1-0.3M-MCDo] 3.32
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Where Amax = 4.47m2:
l = 18.29m
LE = 600
S = 83m2
MCDo = 1.10
EWD = 1.2
CDw=4.5 /83 (4.47/18.29)2 EWD (1.2) [1.0 – 0.31.04-1.00] = 0.01148
CDw=4.5 /83 (4.47/18.29)2 EWD (1.2) [1.0 – 0.31.40-1.00] = 0.00989
CDw=4.5 /83 (4.47/18.29)2 EWD (1.2) [1.0 – 0.31.60-1.00] = 0.00937
Step 3: - k1 from the supersonic drag due to lift equation 3.33:
k1 = [AR (M2 – 1) / (4ARM
2 – 1) – 2] cos LE 3.33
k1 = [2.27(1.042 – 1) / (4 x 2.27 (1.04
2 – 1) – 2] cos 60
0 = 0.02689
k1 = [2.27(1.402 – 1) / (4 x 2.27 (1.40
2 – 1) – 2] cos 60
0 = 0.16235
k1 = [2.27(1.602 – 1) / (4 x 2.27 (1.60
2 – 1) – 2] cos 60
0 = 0.18955
The wave drag analysis above was inconclusive and in order to focus on a single
configuration for further analysis other factors than supersonic drag were considered
in order to down select the PWSC configuration without extensive analysis of all
three configurations; these were usable fuel volume (based on the fact that the
fuselage volumes were almost identical the wing volume became the deciding volume
factor), predicted structural weight based on an initial layout study and the predicted
number of structural members.
Although NB3 and 4 had the best predicted wave drag at M 1.6 the wing had a
fuel tank volume of 6.39m3 compared with 8.2m
3 for NB1 and NB2
configurations.
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The NB5 and NB6 configurations although marginally better wave drag at M1.6
than the NB1 and NB2 configurations they had the smallest wing fuel tank
volume 5.63m3 compared to 8.2m
3 for NB1 and NB2 configurations.
This was considered a major risk if the NB3 / 4 or NB5 / 6 configurations were
required for endurance rather than high speed missions resulting from a mission
profile change.
The exposed wing of the of NB3 and NB4 configurations were marginally smaller
than the NB1 and NB2 configurations at 18.122m2 compared to 18.628m
2 which
would offer some degree of benefit in parasitic drag reduction, but the cracked arrow
layout would not offer the same even load distribution which should be possible with
the trapezoidal wing of the latter configuration. Loads would need to be concentrated
into a small number of relatively heavy members to obtain a high degree of structural
stiffness. Also the need to back – up the aileron control surface attachment lugs with
ribs, and house the another actuator in a relatively thin portion of the wing would
further increase structural weight.
The exposed wing area of configurations of NB5 and NB6 were larger than those of
configurations at 19.425m2 compared with 18.628m
2 for the NB1 and NB2
configurations, which would have had an adverse affect on parasitic drag values
compared with the latter configurations. The structure would also need to be heaver
outboard to maintain structural stiffness in view of the wings high sweep angle the
lower taper ratio, and the influences of the aileron loads on the wing structure, as well
as the incorporation of the actuator in the outboard wing. Also lateral stability could
be an issue with the NB5 and NB6 configurations requiring a re-examination of the
empennage design and sizing.
Also all four of these configurations would have a higher approach and landing speed
than configurations NB1 and NB2 in view of their higher wing sweep, therefore the
landing field requirements may be harder to achieve.
On balance and in view of the workload to investigate all three configurations to the
fullest extent the author selected the NB1 and NB2 configurations as the best
candidates for further study in view of being the most versatile configurations with
the greatest potential for role change and possible systems weight growth, even if the
supercruise capability was diminished, the overall requirements were versatility range
and endurance capabilities, rather than point design.
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3.3 Configuration optimisation by parametric analysis.
3.3.1 Analysis methodology:
This design study was a detailed analysis of the down selected NB1 and NB2
configurations using analysis models created for Jet306. The following Conceptual
design technologies available within the AeroDYNAMIC V3.0 tool set were used for
the configuration analysis:-
Customer Focus – Needs, House of quality.
Design Synthesis – Aircraft configuration modeling.
Geometric Modeling – Areas and Volumes.
Aero Analysis – Parametric aerodynamic analysis.
Propulsion Modeling – Parametric based on F-135 / F-136 published data.
Constraint Analysis – Design Point.
Mission Analysis – Better Mission Fuel Fraction.
Structural / Weight Prediction – Weights analysis.
Sizing – Sized Wing Area, WTO, and TSL.
Performance Analysis - Ps
Cost Analysis – Acquisition and operating costs, life cycle.
Sensitivity / Optimization – “Best” Design, Cost trade.
With the configuration generation stage complete and initial down selection through
drag polar analysis and range analysis using the methodologies in pages 126 to 133 of
reference 22, (as detailed in appendices E) which were used to compare the six FB-
24 manned (worst case with canopy drag) configurations against the F-35C, the
detailed analysis of the two variants of the PWSC wing option was undertaken: -
1. This started with the construction and of the NB1 and NB2 aircraft
configuration analytical models in Jet306 from measurements taken from the
detailed OML models constructed in the configuration generation stage. These
models were crude by comparison with the CATIA V5 models due to the
limits of the Jet306 geometry generation tool, for example the aft fuselage
booms could not be modelled as separate items on the aircraft and had to be
incorporated as extensions to the empennage aft of the engine nozzle. Also the
intakes could not be modelled as chevrons in plan view but only swept, as in
the side elevation, and aircraft with very extensive compound curvature would
be difficult the model with this tool. The Jet306 exhibited numerous run time
errors in operation and a replacement was going to take three months to arrive
and the author is still awaiting delivery. So for the analysis in this study the
older and geometrically less sophisticated AeroDYNAMIC V2.08 was used.
2. These configurations were then analysed using the AeroDYNAMIC toolset,
and compared with the baseline aircraft and the weight targets given above.
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3. The results of these studies were used to determine the best final basepoint
configuration, and this basepoint configuration was developed into a higher
fidelity CATIA V5 solid model for structural layout and analysis.
AeroDYNAMIC 2.08 was also a Whole Aircraft Analysis tool capable of determining
and reporting the following aerodynamic behaviour of a given configuration: - Lift:
Parasite Drag: Induced Drag: Supersonic Wave Drag: Mcrit: supporting the following
analysis: - constraint analysis: manoeuvre analysis: performance and specific excess
power analysis: stability and control analysis: sizing: weight prediction: optimisation:
and cost analysis. The above are based on geometry, mission and engine data as
detailed below. The initial data required in order to construct the Jet306
AeroDYNAMIC analysis model in the Main spreadsheet consist of data parameters
defining the fuselage, and parameters defining the wing, and empennage geometry for
determining the initial wetted area the configuration, and the fuselage is then further
defined in the geometry spreadsheet with the 20 specific geometry data points at 20
frame stations, and this much more accurate data is used for the Swet.
The core fuselage design is broken into three cylindrical sections (note cross –
sectional shapes of the items shown as cylinders and half cylinders in figure 92 may
be circular elliptical, rectangular or any other shape) in Jet306 each frame is defined
by specifying the length, midline width and centreline height of each section with the
exception that the dimensions where two adjacent sections connect must match. If the
height and width of a section are equal, the section will have a circular cross section.
If the height and width of a section differ, the section will have an elliptical cross
section. If the height and width of a section are zero, the section will effectively
become a cone.
Figure 92: - Fuselage Design Breakdown for 3 – section fuselage.
3.3.2:- NB1 and NB2 Analysis model construction.
Form the NB1 and NB2 surface models created above and the as drawn engine based
on YF-120 dimensions the input data was obtained as shown in figure 93 and 94.
Then section cuts were taken through the CATIA surface model at 20 equal distance
stations and breaking the OML into outline contour cuts, from which the width and
height for each section was measured. This data was then entered into the
AeroDYNAMIC 2.08 geometry data spreadsheet to create a basic definition of the
Fwd Mid Aft h1
w1
h1
w1 w2
h2
h2 w2
Fwd len Mid len Aft len
Cone # 1 Cylinder # 2 Cylinder # 1
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fuselage. The airframe centre line was determined to be on the engine thrust line and
all point co – ordinates were measured relative to this, at each frame station.
Figure 93: - NB1 Section cuts for Jet306 analysis (note what appear to be surface
blemishes in the cockpit are white reflections as capture used show white as
black feature to show the cuts).
Figure 94: - Analysis engine for Jet306 data from YF-120 data used for NB1 and
NB2 evaluation constructed in CATIA V5 by the author.
This data was used with the location and size of the crew stations, fuel weight, engine
weight and C of G location, control surface sizing and location, undercarriage layout,
payload weight and volume, fuel weight and volume, to evaluate the NB1 and NB2
configurations for which the results are presented below.
Airframe mate joints denoted by orange planes.
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Chart 1:- (a) NB1 Total predicted drag variation with Mach number: (b) NB1
Drag polar to Mach 0.85
Figure 95:-NB1 Cross-sectional area distribution which showed a marked
similarity to that of the F/A-22A as shown in figure 24.
(a)
(b)
Wave drag
Profile drag
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Chart 2:- NB1 Lift over drag verses lift coefficient at Mach 0.85 and Mach 1.5
curves: - (a) L/D –v- CL at Mach 0.85, and (b) L/D – v – CL at Mach 1.5.
Chart 3:- NB1 Lift curve CL v
(a)
(b)
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Chart 4:- NB1 CD – v – CL at Mach 0.85 (a) and Mach 1.5 (b).
Chart 5:-NB1 Thrust and Drag –v-Mach number.
(a)
(b)
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Chart 6:- (a) NB1 Vn Diagram: (b) NB1 Manoeuvre Diagram.
Chart 7:- NB1 Specific Excess Power over a range of Mach numbers and
altitudes.
(a)
(b)
Stall - Limit
q - Limit
NB-1
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Chart 8:- (a) NB2 Total predicted drag variation with Mach number: (b) NB1
Drag polar to Mach 0.85
Figure 96:-NB2 Cross-sectional area distribution which showed a marked
similarity to that of the F/A-22A as shown in figure 24.
(a)
(b)
Wave drag
Profile drag
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Chart 9:- NB2 Lift over drag verses lift coefficient at Mach 0.85 and Mach 1.5
curves: - (a) L/D –v- CL at Mach 0.85, and (b) L/D – v – CL at Mach 1.5.
Chart 10:- NB2 Lift curve CL v
(a)
(b)
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Chart 11:- NB2 CD – v – CL at Mach 0.85 (a) and Mach 1.5 (b).
Chart 12:-NB2 Thrust and Drag –v-Mach number.
(a)
(b)
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Chart 13:- (a) NB2 Vn Diagram: (b) NB1 Manoeuvre Diagram.
Chart 14:- NB2 Specific Excess Power over a range of Mach numbers and
altitudes.
(a)
(b)
Stall - Limit
q - Limit
NB-2
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Table 12:- Comparative component weights: GTOW / Landing Weight: Costs
for NB1 and NB2 based on the RAND Corporations DARCO IV model in
AeroDYNAMIC V2.08.
COMPONENT NB1:-kg / (lb) NB2:-kg / (lb)
Fuselage 2,837.6 / (6,255.7) 2,837.6 / (6,255.7)
Verticals 333.6 / (735.4) 323.2 / (712.6)
Wings 1,232.4 / (2,716.9) 1,232.4 / (2,716.9)
Undercarriage 693.1 / (1528.0) 590.4 / (1301.7)
Horizontals 753.4 / (1661.0) N/A
GROSS WEIGHTS NB1 NB2
GTOW 23,182.8 / (51109.4) 22,327.4 / (49223.5)
Landing Wt 10% fuel 12,063.45 / (26,595.4) 11,208.0 / (24,709.5)
x CofG TO 5.334m / 17.50ft 4.898m / 16.07ft
x CofG Landing 10.250m / 33.62ft 9.754m / 32.00ft
COSTS* NB1 NB2
Fly Away Cost 1st Estimate
$51,109,360 $49,223,450
Total fly away costs $27,745,394 $25,967,104
Total Q & M costs $19,856,443 $19,856,443
Total Life Cycle costs $47,601,837 $45,823,547
Stability NB1 NB2
MAC 25.21123ft 25.21123ft
xMAC 22.69141ft 22.69141ft
yMAC 7.99775 7.99775
NPSub 0.283562 0.160700
NPSup 0.517118 0.407723
SMTO Sub 0.723202 0.670459
SMlnd Sub -0.15007 -0.20884
*N.B.:- Costs are based on production run of 500 aircraft with 3 flight test
aircraft.
Based on the above values NB2 had better weight and cost values than NB1
therefore NB2 went forward to the structural layout phase detailed in section 4.
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For the FB-24 configuration the maximum endurance is achieved at the speed for
(L/D)max which can be determined by first calculating the required value of CL, then
solving for the speed required to achieve L=W at that CL therefore for FB-24:-
CL = CDo / k = 0.014361 / 0.167051 =0.293
L = W = CLqS, q = W / CLS = 53,187.50 / 0.293 (893ft2) = 203.27lb/ft
2
And using at 34,400ft = 0.000767 slug/ft3 obtained from standard atmospheric
tables (reference 22) and the definition of q,
V = 2q / = 2 x (203.271lb/ft2) / 0.000767 slug/ft
3 = 728.05ft/s
for maximum endurance. Note that this is only the initial velocity for maximum
endurance and that as fuel is burned the velocity for best endurance will decrease. In
order to calculate the maximum endurance time, it is first necessary to determine the
magnitude of (L/D) max using equation 3.34: -
(L/D) max = (CL / CD) max = 1 / 2 k CDo = 1 / 2 0.167051(0.014361) = 10.2
The TSFC is also predicted using equation 3.35 with a = 977.5ft/s at 34,400ft and
aSL = 1116.4ft/s obtained from the standard atmospheric table (reference 22):-
ct = cSL (a / aSL) = (0.89[(lb/hr)/lbf])(977.5ft/s / 1116.4ft/s) = 0.78[(lb/hr)/lbf]
Then the endurance can be calculated using equation 3.36 with W1 = 49,223.5 and
that for W2
W2 = W1 – Wf = 53,187.46lb – 21,126lb = 31,061.46lb
E = 1/ ct CL / CD ln (W1 / W2) = 1 / 0.78 (10.2) ln (53,187.46 / 31,061.46) = 7.03h
Therefore for FB-24 the maximum endurance at 34,400ft at Best Cruise Mach
BCM is 7.33hours.
Similarly the value for maximum range is obtained by solving equation 3.36 for CL
and equation 3.37 for q:-
CDo = 3k CL2, CL = CDo / 3k = 0.014361 / 3(0.167051) = 0.239
q = W / CLS = 53,187.46 / 0.239(893ft2) = 249.20lb/ft
2
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V = 2q / = 2 x (249.20lb/ft2) / 0.000767 slug/ft
3 = 806.11ft/s
for maximum range. As with the velocity for maximum endurance, the velocity for
best range will decrease as fuel is burned. The vale calculated for CL is now used to
calculate CD after which the maximum range is predicted using equation 3.38.
CD = 0.014361 + 0.167051 CL2 = 0.014361 + 0.167051(0.239)
2 = 0.0239
R = 2/S / ct CL2 / CD (W1
1/2 – W2
1/2)
R= 2 / (0.000767 slug/ft3) (893ft
2) 2 / 0.78 (0.239)
1/2 / 0.0239
x ((53,187.46)1/2
-(32,061.46lb)1/2
)
R = 15.315 (ft2 / lbs
2) (2.564h)(20.45)(10.563lb
1/2) = 2,167.49
Therefore for the range in nautical miles divide through by 1.69(ft/s/kn)
Therefore the FB-24 would have a range value of R = 1,282.54nmiles.
As stated above the drag values from AeroDYNAMIC were pessimistic being based
on the a oval cross section throughout the length of the fuselage and the booms having
to be treated as complete sections of the fuselage, both of theses factors have lead to
an overly pessimistic parasitic and wave drag values. The A-24 has a larger fuel
capacity than the FB-24, as will be seen in section 4, also a single seat version of the
FB-24 would have the same fuel capacity as the A-24 and with modern systems
becoming more autonomous the need for a second crew station may well diminish
before the FB-24 configuration is finalised.
Of the two configurations studied using AeroDYNAMIC V2.08 the NB2
configuration was demonstrated to have the best weight and hence lowest cost figures
therefore this configuration was further developed in the structural layout and systems
integration section below.
Further subsonic and supersonic analysis of range and endurance for both the
FB-24 and A-24 is given below in Appendices E at the end of this report pages
251 to 256, also it is worth noting that a single seat FB-24 would have the same
fuel capacity as the A-24 with the corresponding greater range and endurance,
and with enhanced single crew operation this should be considered for further
study.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
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148
The characteristics of the F-136-2 engine selected for this conceptual design study
were determined below from published data and installed thrust was calculated using
the methodology advised in reference 22.
There needs to be a much more comprehensive and detailed aerodynamic analysis
conducted on the NB2 configuration to enable more realistic assessment of the aircraft
configurations range and performance and ability to supercruise to be determined and
this is addressed as further work required in section 5 below.
For the F-136-2 engine with a published uninstalled maximum sea level thrust of
43,000-lbs would have a maximum installed thrust of 34,400-lbs.
The F-136 engines with an estimated uninstalled military power thrust of 34,400-lbs
at sea level would have a military power installed thrust of 27,520-lbs.
Using the uninstalled static sea level TSFC figures based on the experimental engine
i.e. TSFC at maximum thrust equal to 1.40[(lb/hr)/lbf] and at military power equal to
0.74[(lb/hr)/lbf]. Using the 20% factor for an installed engine these TSFC figures
become 1.68[(lb/hr)/lbf] for maximum thrust and 0.89[(lb/hr)/lbf)] on military
power.
Therefore for this project the installed sea level performance of the F-136 engine
selected for the Advanced Interdiction Aircraft FB-24 and A-24 derivative are as
follows:-
Maximum thrust = 34,400-lbs
TSFC at Maximum thrust = 1.68[(lb/hr)/lbf]
Military Power thrust = 27,520-lbs
TSFC at Military Power thrust = 0.89[(lb/hr)/lbf]
Taking the values for Maximum thrust / TSFC at maximum power and the Military
thrust / TSFC at military power calculated above, the fuel consumption for the
proposed mission can be determined for (1.1) the best cruse altitude (BCA) targeted
as 34,000ft and best cruse Mach (BCM) targeted as Mach 0.8, and (1.2) the high
speed Dash in weapon release and Dash out targeted at Mach =1.6 at an altitude of
45,000ft phases of the mission.
The S/L rate of climb rate envisaged for the AIA in the clean condition i.e. no external
stores and Maximum power is 50,000ft/min (254m/sec).
Using the equations for low bypass ratio turbofan performance modelling (Reference
22:- Introduction to Aeronautics: A Design Perspective: by Bradt. S: Stiles. R. J.:
Bertin. J. J: Whitford. R: Published by the American Institute of Aeronautics and
Astronautics, Inc: 1997) the TSFC over the mission portions of interest can be
determined as follows:-
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
149
TA = TSL ( / SL) (equation 1.1)
TA = TSL ( / SL) (1+0.7M) (equation 1.2)
Where: - TA = Thrust at altitude.
TSL = Thrust at sea level.
= Density at altitude.
SL = Density at sea level.
NB: - equation 1.1 is only valid for M < 0.9
ct = cSL (a / aSL) (equation 1.3)
Where; - ct = TSFC at altitude.
cSL = TSFC at sea level.
a = Speed of sound at altitude.
aSL = Speed of sound at sea level.
The installed sea level values being:-
Maximum thrust = 34,400lbs
TSFC at Maximum thrust = 2.51[(lb/hr)/lbf]
Military Power thrust = 27,520lbs
TSFC at Military Power thrust = 0.89[(lb/hr)/lbf]
Figures 97 and 98 below indicate some of the short comings of the F-35 when applied
to roles other than tactical interdiction and as will be seen below this conceptual
design has gone some way to address these issues, although its capabilities as a
bomber are an improvement over both the F-35 the F/A-22A and planned UCAV‟s
the latter being high subsonic aircraft it is not a fighter as typified by the F/A-22A in
figure 98 below, and lacks the agility for that role, which was not requested in the
request for proposals.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
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150
Figure 97: - Design limitations of the current F-35 family reflecting the need for
the AIA design divergence from the original baseline F-35C design.
Figure 98: - Comparison between F/A-22A and CTOL F-35 (published data),
contrast with figure TBD of AIA vs F/A-22A: JSF: Tornado: F-15E.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
151
4.0 Structural layout and systems integration.
Figure 99: - F/A-22A internal structural layout illustrating structural density
source reference 17.
The objective of this section is to propose structural integration solutions for the
installation of the crew / AI system, propulsion system, offensive and defensive
weapons systems, undercarriage, and fuel systems within the FB-24 and A-24 OML
envelope, by designing a structure capable of accommodating these systems and
capable withstanding the aerodynamic and inertia loads to which these aircraft would
be subjected during the sizing mission. This section proposes the materials from
which the common FB-24 / A-24 airframe would be produced and highlights some
possible manufacturing routes. Each system is dealt with individually in sections 4.1
to 4.5, culminating in full structural integration modelling in section 4.6, which covers
layout, materials and weight.
Figure 99 above illustrates the structural density of advanced stealth combat aircraft
and the issues involved in systems integration. The level of definition in this figure is
much higher than could be achieved within this part – time MSc thesis and represents
an airframe post detail design which was beyond the scope of this work, however a
level of detail commensurate with the PWSC proposal submission phase was
produced here using notional skin thicknesses, and sub – structure thicknesses, and
the profile e.g. „I‟-section: „C‟- section are stated for each component or component
class. The complete airframe was solid modelled based on key datum face models
designed by the author which were intended for structural analysis using PATRAN /
NASTRAN FEA toolset however this analysis was not possible within the time scale
of this conceptual design study. Analysis of this structure was used to give a very
provisional indication of major structural component weight to support the proposal
submission. As stated above the NB-2 configuration was considered by the author to
have the best possibility of successfully meeting the design mission requirements and
was the only configuration carried forwards, although NB1 was used to draw some
comparisons with the NB-2 configuration.
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4.1 Undercarriage systems integration for FB-24 and A-24.
Figure 100: - FB-24 and A-24 Common four tail airframe tip back angle = 18.70
for a wheel base of 8.086m and an aircraft C of G positions most fwd Frame
Station 9.19, (9.19m) and most aft Frame Station 10.11, (10.11m) along the a/c
axis.
Figure 101: - FB-24 and A-24 Common four tail airframe overturn angle = 75.30
for a wheel track of 3.325m, aircraft height of 4.71m and an aircraft C of G
positions most fwd Frame Station 9.19, (9.19m) and most aft Frame Station
10.11, (10.11m) along the a/c axis.
LG = 8.086m
CoG Most Fwd = FS 9.19 CoG Most Aft = FS 10.11
A/C height = 4.71m
18.70
Ground line
CoG Most Aft = FS 10.11
CoG Most Fwd = FS 9.19
75.30
W = 3.328m
53.50
420
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Figure 102: - FB-24 and A-24 Common ruddervator airframe tip back angle =
18.70 for a wheel base of 8.086m and an aircraft C of G positions most fwd
Frame Station 9.19, (9.19m) and most aft Frame Station 10.11, (10.11m) along
the a/c axis.
Figure 103: - FB-24 and A-24 Common four tail airframe overturn angle =
75.250 for a wheel track of 3.325m, aircraft height of 3.79m and an aircraft C of
G positions most fwd Frame Station 9.19, (9.19m) and most aft Frame Station
10.11, (10.11m) along the a/c axis.
18.70
CoG Most Aft = FS 10.11
CoG Most Fwd = FS 9.19
LG = 8.086m
Ground line
A/C height = 3.79m
W = 3.328m
CoG Most Aft = FS 10.11
CoG Most Fwd = FS 9.19
71.90
53.50
420
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The undercarriage used in this study and layed out above in figures 100 through 105
for NB1 and NB2 was a high Flotation / high Sink Rate type based on public released
data for the F-35C and measurements taken from X-35C drawings, and the main
characteristics of which were as follows and these were used to model the
undercarriage as shown above: -
Main undercarriage tiers sized at = 86.36 x 27.94 x 43.18cm:
Main undercarriage stroke = 33.17cm:
Nose undercarriage tire sized at = 59.69 x 19.05 x 25.40cm:
Nose undercarriage stroke = 52.07cm.
Analysis of the proposed undercarriage layout with a wheel base of 8.086m and a
wheel track of 3.328m demonstrated its suitability for both the NB1 four tail and the
NB2 ruddervator configurations, imparting a common tip back angle of 18.70 above
the minimum of 160 recommended in reference 21, and an overturn angle of 75.3
0 for
NB1 and 71.90 for NB2, both above the minimum angle of 62
0 recommended in
reference 21. The angle from the most fwd C of G position to the main undercarriage
oleo centre line was 53.50 for both NB1 and NB2, and the angle from the most aft C
of G position to the main undercarriage centre line was 420 for both NB1 and NB2.
The single nose wheel folds forwards into a nose wheel bay ahead of the oleo without
rotation about its axis, where as the main undercarriage also folds forwards into the
wing root blend but also rotates through 900 to stow into the wing blend wheel bay aft
of the wing leading edge spar as shown in figures 104 and 105 below, which is the
same mode of retraction used in the F-35 family of aircraft and thereby retains
functional commonality. In the event of a mechanical failure hanging up the
undercarriage the doors can be blown open with inert gas canisters and the
undercarriage can be lowered using a combination of inert gas and the airflow over
the aircraft underside into the locked position, thus enabling recovery of the aircraft in
such an emergency. The undercarriage bays were sized as: - nose wheel bay = 167cm
x 36cm x 54cm: and the main wheel bay = 288cm x 110cm x 52cm respectively.
The accommodation of the main undercarriage outboard of the fuselage in the wing
root gave a wide and stable wheel track enabling the possibility of deployment to
relatively austere forward airbases within NATO and allied nations, as well as good
handling in ice and poor weather which were primary mission requirements from
section 2. This layout did produce challenges however both aerodynamic in terms of
the extensive lower wing skin root to fuselage blend, and structural in terms of
incorporating the wheel bay and oleo within the root, and the design of the
undercarriage doors which needed to be both stiff and of complex curvature. There
could also be RCS issues arising from the need to eliminate any steps and gaps which
could arise from the articulated doors operation, although this would be addressed by
shaping the door with planform aligned chevrons and seals at the door periphery.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
155
No
se
un
de
rca
rria
ge
retr
acts
forw
ard
th
rou
gh
94
0 in
to th
e
forw
ard
wh
eel
bay.
Main
u
nd
erc
arr
iag
e
retr
acts
fo
rward
thro
ug
h 9
20 in
to t
he w
ing
wh
eel b
ay
.
At
45
0
forw
ard
ro
tati
on
an
gle
th
e
main
wh
eel
itself
is r
ota
ted
th
rou
gh
90
0 a
rou
nd
th
e o
leo
to
lay f
lat
in t
he
wh
eel b
ay
.
Fig
ure
104:
- N
B2 U
nd
erc
arri
age
ret
ract
ion
an
d d
eplo
ym
ent
wh
ich
foll
ow
s th
e sa
me
met
hod
olo
gy a
s th
e F
-35 f
am
ily.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
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156
Figure 105: - Main undercarriage stowage in the wing root blend of the NB2
configuration, based on the F-35 family methodology.
4.2 Aircrew and AI integration for the FB-24 and A-24.
The as designed FB-24 NB-2 aircraft which was the final manned configuration
selected for this proposal submission enables the accommodation of two fully suited
flight crew representing the average 95th
percentile pilot in conditions required for
their completion of the desired mission. The original intention was to accommodate
these crew members in 1m high Martin Baker Mk-16 ejection seats, under a single
bow 2.6m canopy as envisaged for the stillborn F-22B trainer version of the F/A-22A,
however this would involve a complete redesign of the forward fuselage build module
and eliminate any commonality with the F-35 family of aircraft from which the FB-24
was derived, sharply increasing costs of manufacture, as well as increasing the drag
penalty of a canopy, and increasing structural weight. Therefore an alternative
solution was required, as a minimum change alternative and this was the direction
followed in this section for the manned airframe. In the region of the canopy there
were the conflicting requirements of drag reduction for supercruise flight favouring a
smooth cross-section distribution over the forward fuselage, and the need to provide
pilot visibility equal to that of the F-35 family for the Aircraft Commander (front seat
pilot) providing at least 150 of vision over the nose in level flight and ground
handling. Consequently there was a strong desire to retain the original F-35C layout
and substructure which gave the front seat pilot a published (reference 6) over nose
view angle greater than 160.
Wheels rotate through 900
to lay flat in the wheel bay, reducing storage envelope to a minimum in the x, z,
plane.
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For the FB-24 an angle of 16.250 was obtained after some maturation and the design
eye location is shown in figure 106 below, in terms of FS (frame station) and WL
(water line) position on the aircraft axis.
Desig
n
ey
e
of
pilo
t air
cra
ft
co
mm
an
der
FS
367
0.3
WL
32
76.6
(Wate
r L
ine P
lan
e)
16.2
50
Vie
w a
ng
le
WL
21
35.7
(Wate
r L
ine P
lan
e)
Sta
tic g
rou
nd
lin
e.
Fig
ure
106:
- N
B2 D
esig
n e
ye
of
pil
ot
(air
cra
ft c
om
man
der
) w
ith
was
a 1
6.2
40over
nose
vie
w a
s in
th
e F
-35 f
am
ily.
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This left the issue of the location of the second crew member the Mission Operation
System / UCAV Commander (MOSC) within the airframe, after dropping the tandem
canopy option shown in figure 107, and to provide this crew member with a means of
ingress / egress and emergency escape, as well as situational awareness.
The latter could be addressed by artificial vision and computer controlled imagery, the
MOSC would receive all information from the aircrafts onboard sensors, off – board
platforms e.g. AWAC, Recon UAV‟s / Global Hawk, and the synthetic real-time
environment of the A-24 UCAV‟s, and all of this information would be available in a
choice of single aspect or multi window formats on the 20” flat panel (God‟s Eye
View) touch sensitive display driven by a 400Gb 4.0GHz EMP hardened CPU
currently projected to be the size of a large notebook pc.
The former however proved more challenging, although after evaluation of the fuel
tank layout required to house the mission fuel the author determined that the MOSC
could be accommodated in the F-1 fuel tank location between the intake ducts, by
moving this fuel into the extended F-2 and F-3 tanks resulting from the fuselage
extension and removal of the internal gun provision (reversion back to the
missionized gun pod of the F-35C / F-35B variants), and the adoption of hose and
drogue refuelling as opposed to the USAF traditional boom method, covered in
section 4.4 of this thesis.
Figure 107:- The original crew accommodation layout under a large F-22B type
canopy, with height adjustable seating would have increased drag from the
canopy, and increased overall weight (generic aircrew and seats reference MB).
This location was supported by the fact that in the F-35 family the canopy uniquely
hinges forward just below the windshield as shown in figure 108, permitting the
fitting of a single chevron clam shell door behind the canopy for the second crew man
without risk of fouling the canopy hinge mechanism. The resulting structural
arrangement required to this is shown below in figure 109. The door would remain
sealed with the same type of pressurisation sealing as the canopy for the duration of
the mission, although pyrotechnic locking bolts would provide quick door jettison into
the airflow down stream over the fuselage permitting safe ejection of the MOSC in an
emergency.
2.6m
THE ORIGINAL TWO CREW STATIONS
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Figure 108: - The F-35 families‟ unique forward pivot canopy hinge mechanism
which is shown here on the F-35A airframe mock-up model (reference 28).
Figure 109: - Crew accommodation integration in the FB-24 showing extent of
frame cut out to house second crew station in space envelope of F-1 fuel tank,
modelled by the author in CATIA V5. (Note substructure is complete consisting
of frames and longerons thickness modelled is 50mm (flange width) where as
actual frame pocket thickness will be 4mm in carbon PMR-15.)
Orange frame segments are part of clam shell door.
Pilot (aircraft commander)
MOSC (UCAV commander)
16.250
Door hinge frame
3.5m
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The seat pitch in the FB-24 was set at 220 as in the current F-35C aircraft design, as
there was no requirement for the crew to endure loads over 7g throughout the duration
of the mission. The seat dimensions were 1.20m high by 0.50m by 0.54m
commensurate with the published dimensions for the Martin Baker Mk-16 seat
designed for the F-35 family. The 0.61m3 volume requirements for the crew systems
i.e. OBOGS, and OBIGS could also be contained within the as designed structure
shown in figure 109 above outboard of the MOSC crew station, within the DIS IML
boundary, although definitive positioning would undertaken as further work under
detailed systems integration. The design of the clam shell MOSC door was also left to
follow on work due to the time constraints of this study, but the author envisages a
cover which maintains OML continuity with the airframe. This concluded the FB-24
air crew integration study.
The A-24 AI (artificial intelligence system) integration was a simpler task, as this
system was envisaged as a complete self contained module with its own ECS as in the
X-45C and X-47B UCAV‟s requiring only a plug and play power interface, and fibre
optic sensor data link. The module installation did not adversely effect the provision
of the F-1 fuel tank and as such the A-24 fuel capacity was greater than that of the
FB-24 as will be covered below. Figure 110 illustrates the AI module installation as
modelled by the author in CATIA V5.
Figure 110: - AI installation integration in the A-24 showing extent of frame cut
out to house system space envelope without affecting the F-1 fuel tank, modelled
by the author in CATIA V5. (Note substructure is incomplete consisting of
carbon PMR-15 frames only full structure is shown in section 4.6).
However the height of the module i.e. 1.25m did break the fuselage OML if the
canopy envelope was to be replaced with a fairing to continue the curvature of the
surrounding skin.
AI Module Optical
tracking system
AI Module with ECS incorporated
F-35 Standard radar
mounting frame
1.74m
Carbon PMR-15 all frames
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Therefore with the module in its current form the canopy shape will need to be
retained with its optical properties. However the drag penalty will be off - set by the
fuel capacity of the F-1 tank, as discussed in section 4.4. The canopy could therefore
serve as the primary access and servicing aperture, and because it is already designed
for low RCS reflections and emissions there will be no adverse affects on the aircraft
signature from the canopies retention. This concludes the section on the aircrew and
AI system integration, in which the ability to accommodate two aircrew within the
substructure and skin IML was established, although further work is required to
define the MOSC access door. The ability to accommodate the AI system in the
systems current preliminary design form was also established and no adverse impact
on fuel capacity was determined.
4.3 Propulsion systems integration for the FB-24 and A-24.
The F-135 Propulsion system: - The F-135 propulsion system has been further
developed over that of the JSF119 shown in figure 111 and installed in the X-35
Concept Demonstrator Aircraft (CDA) currently the F-135 figure 112, has been rated
at 43,000-lb uninstalled thrust with afterburner, and has bypass ratio in the order of
0.30 and with 67% of maximum thrust available in military power the dry thrust was
calculated to be 28,810-lb. The engine integrates the proven F-119-100 core and has a
three – stage fan with hollow first stage blades, and composite fan inlet guide – vanes.
Behind this fan section is a six – stage high pressure bladed blisk compressor, a single
– stage high pressure turbine, and a two – stage low – pressure turbine, which has a
new low – pressure spool replacing the single - stage of the F-119 engine. In addition
the propulsion system features advanced prognostic and on – condition management
systems that provide maintenance awareness, automatic logistic support, and
automatic field data and test systems. This engine also has 40% fewer parts combined
with 50 % lower infrastructure support requirements compared with current in service
engines, according to the manufacturer.
The variant considered for the FB-24 and A-24 would be the F-135-400 CV variant
engine as this engine is proposed to have an installed thrust of 43,000-lb with
afterburner.
Figure 111:- Pratt & Whitney JSF119 low by – pass ratio turbofan cut – away,
showing the very high degree of F-119 commonality source P&W press office.
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Figure 112:- The Pratt & Whitney F-135-100 JSF CTOL engine under final
assembly at Pratt & Whitney‟s plant, and like the F-119 it was designed to be
maintained with six common hand tools and the elimination where possible of
wire lock fasteners. (Source: - Reference 6.)
The F-120 / F136 Propulsion system: - The General Electric F-120 engine program
was selected by the JSF Joint Project Office as the baseline from which to develop an
Alternative Engine Program (AEP) to the Pratt & Whitney F-135.
Figure 113:- General Electric F-120 Variable Cycle Engine Schematic illustrating
the engines two modes of operation (Source reference 17).
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163
Figure 114:- Longitudinal Section showing main component layout of the General
Electric F-120 Variable Cycle Engine (Source reference 26).
This selection was based on the F-120 engines core thrust capability to meet the JSF‟s
multi - service aircraft requirements, including STOVL capability without scaling up,
and for its growth potential. Originally designated the YF-120FX for the JSF, the
engine is based on the GE F-120 developed and successfully demonstrated in flight in
the early 1990‟s in the USAF Advanced Tactical Fighter engine competition.
Development of the F-120 engine shown above in figures 113 and 114 has continued
since the ATFE competition under the Advanced Research Projects Agency‟s
advanced short take-off and vertical landing program which has seen the engine run
for 21 hours, and since 1995 funding has come from the JSF Program Office, for: -
Phase 1 engine definition (1995-1997): Phase 2 Critical Design Review (1997-2001):
Phase 3 Detail design (2002-2005) and 12,000+ ground test hours. Currently an F-136
derivative F-120 engine planned to fly in 2009 with first production engine delivery
anticipated 2011 and the engine was formally designated the F-136 in 2005, however
with the on going drain on the US defence budgets this was reassessed and after the
Quadrennial Defence Review of February 2006 the possibility exists that this
alternative engine may be delayed or cancelled. This thesis assumes that funds will be
found for this engine within the development time frame of the FB-24 and A-24
aircraft i.e. 2015 through 2020.
The technical characteristics of the F-136-2 are as follows:-
Front fan (Rolls-Royce): Long wide - chord, titanium, three–stage blisk: stage one
has hollow core blades: while stages two and three have solid blades. (Currently two
builds have been tested, which has verified the fan flow, as well as the efficiency of
linear friction welding for blade attachment).
High-Pressure compressor (General Electric GE): Five – stage, all – blisk system
consisting of three rotors in stage one and stage two, and stages three to five being
inertia – welded together.
No
zzle
rep
laced
by
LO
AN
on
FB
-24
an
d
No
rth
rop
n
ozzle
on
A-2
4.
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164
The compressor features forward swept airfoils, and robust blade tips, bowed / swept
stators designed from 3-D aerodynamic codes, and high stage loading to support
40,000-lb thrust class operation.
Combustor (Rolls-Royce): Single annular simplified design, fabricated from
Lamilloy cooling material, and based on the IHPTET (Integrated High Performance
Turbine Engine Technology) program research into higher turbine operating
temperatures.
Turbine (GE and Rolls-Royce): Single-stage high – pressure turbine (HPT) and a
three stage low-pressure turbine (LTP), the HPT and stage one LPT forming a
coupled, vaneless counterrotating system (with single crystal HPT blades). This
turbine system has been successfully rig - tested at the time of writing.
Augmentor (GE): Radial, non-stage, variable flow control, based on GE YF-120:
F414-GE-100: F110-129 and F110-132 engines.
Technical information from the following General Electric Aircraft Engine web page:
- http://www.geae.com/engines/military/f136/background.html accessed on the 1st
February 2005.
The original YF-120 design employed a fan and airflow size 12% larger than that of
the XF-120 experimental ATF engine. This size increase was in response to the
requirement for more cooling air for the metal exhaust system and the increased thrust
requirements. Overall pressure ratio was 22. Durability was addressed using thermal
barrier coatings (TBC) and tailored cooling air distribution.
As a result of the engines short and compact hot section vaneless high – pressure / low
– pressure turbine concept, the hot section cooled surface area was 30% less than in
the General Electric F-110 engine. Variable cycle engine (VCE) features were
simplified from the XF-120 design, and maintainability requirements were rigidly
enforced during the design process. At time of writing this thesis the F-136 has been
rated at 43,000-lb uninstalled thrust with afterburner, and has bypass ratio in the order
of 0.32. For the original YF-120 engine dry thrust was 28,000-lb using VCE
technology as shown in figure 113 were the engine can be operated as either a turbofan
or a turbojet when the maximum thrust was 35,000-lb or 80% of maximum thrust was
available in the dry condition. Translating this into the current figures for the F-136 if
the same VCE technology was employed then for a 43,000-lb maximum thrust a dry
thrust value of 34,400-lb could be achieved.
For both the F-135 and F-136 engines all components, harnesses, and plumbing were
located on the bottom of the engine for easy access, and all line replaceable units
(LRU‟s) were located one layer deep (i.e. units were not located on top of each other),
these features and a reduction of the tools required to maintain the engines to just six
standard tools increases the supportability of the engines in service.
The basic engine dimensions provided by GE and PW press offices that were
used in this thesis are given below:-
For both engines the fan face diameter was = 1.10m (3ft 7in):
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For both engines the overall length including nozzle was = 6.06m (19ft 10in):
The nozzle length for both installations was = 0.72m (2ft 4in):
The nozzle exit face diameter was = 0.79m (2ft 7in):
The nozzle base diameter was = 1.27m (4ft 2in)
The overall engine weight including nozzle and equipment was = 3,402kg
(7,500lb) based on the F-119 engine weight figures.
Figure 115: - Proposed FB-24 and A-24 Propulsion system integration in which
the standard bifurcated intake duct of the F-35C is integrated with the General
Electric YF-120 derived Variable Cycle Engine (F-136-2) and GE Axisymmetric
Vectoring Nozzle (AVEN). This installation was modelled by the author in
CATIA V5, and the structural integration was modelled in section 4.6.
The FB-24 and A-24 utilise the F-35 family bifurcated composite intake ducts and the
diverterless supersonic intake DSI (see appendices C) as shown above in figure 115.
The fuselage extension for wave drag reduction had the additional benefit of enabling
a longer flow recovery stage to be introduced between the intake convergence point
and the engine face reducing the risk of turbulent airflow and shockwaves causing
flow separation and pressure loss reaching the fan face, thereby permitting better
pressure recovery at Mach 1.6. That is the flow recovery stage of 3.7m for the FB-24
and A-24 compared with 1.7m for the F-35 family.
6.06m
6.45m
2.93m
YF-120 derived F-136-2
Variable Cycle Engine
AVEN Nozzle
Bifurcated intake duct
Aircraft Centre Line and Engine Thrust Line.
Fixed trunnion
Free sliding trunnion
Fwd link
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An example of the bifurcated made from CFC without the use of mechanical fasteners
is shown below in figure 116. The bifurcated intake results in a 100% line of sight
blockage of the engine face and thus eliminates all return energy leaks substantially
reducing the FB-24 and A-24 RCS which was in common with the F-35 family of
aircraft.
Figure 116:- F-35 / FB-24 / A-24 common intake duct design with the supporting
stiffeners being co-bonded and z pinned to the duct wall thus obviating the need
for mechanical fasteners thereby reducing weight and potential RCS reflection
sources. (Source: - Lockheed Martin JSF media relations).
Engine installation and removal would be via the titanium engine bay doors located
on the underside of the rear fuselage, as per the current F-35 family as shown in figure
119 below. The engine locates onto an integral rail machined into the top of the rear
fuselage frames which with side mounting bolts transfer the thrust loading to the
airframe, and the engine will be screened by refractory coated Niobium alloy or
carbon BMI nacelle skins from the rear fuselage substructure, as shown in figure 117,
details of this structural arrangement for the F-35 are ITAR restricted and therefore
cannot be elaborated on within this thesis. The engine bay will have a full fire
suppression system. For transportation to and from the aircraft the engine the common
A/M32M-34 trailer used for both F-35 and F/A-22A aircraft shown in figure 118,
which is 4.26m (14ft) in length and 1.8m (6ft) wide, and 1.5m (5ft) high with a
maximum payload of 3,402kg (7,500lb).
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Figure 117:- CGI modelling of the installation of the F-136 engine into the F-35C
from the common A/M32M-34 engine transport trailer which can be raised and
lowered as shown. The identical system would be used for the FB-24 and A-24
aircraft. (Source Lockheed Martin F-35 Project Media relations).
Figure 118:- Proposed engine mounting for FB-24 / A-24 based on F-35 family
(Source reference 6).
Engine attach rail.
Engine free - sliding trunnion mount.
Engine fixed trunnion mount.
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Figure 119:- The Boeing developed A/M32M-34 trailer which greatly eases the
installation and removal of the F-136 / F-119 / and F-135 engine by permitting
extremely precise vertical and lateral movement of the engine. (Source: -
Reference 14)
Engine mounting within the airframe was achieved with the addition of a minimal
amount of interconnecting structures, but the engine mount had to be able to prevent
deflections in the main aircraft fuselage structure from inducing loads into the engine
itself, and also had to allow for the thermal expansion of the engine both along its axis
and around its circumference. The basis for the method used for the FB-24 and A-24
and shown above in figures 115 and 118 was based on the Lockheed F-104 Starfighter
fighter aircraft mounting system described in reference 28: - Airframe Structural
Design (Practical design information and data on aircraft structures): by Niu.
Michael. C. Y.: Published by Hong Kong Conmilit Press Ltd 1997, because the F-104
was a single engine fighter in which the engine was mounted in the rear of the
fuselage, and was integrated with bifurcated ducts.
The major portion of the vertical loads shown below in figure 120 would be carried
by the trunnions located close to the engines centre of gravity. Side (or lateral) loads
would be taken out on the fixed starboard trunnion only, with the port trunnion being
free to move laterally to allow for any radial thermal expansion of the engine. The
forward mount would be a universal joint capable of carrying vertical loads only.
Because the would be located near the C of G the major forces experienced by the
forward support link mount would come from the gyroscopic couple crated by the
angular velocity in yaw and the inertia moment caused by angular acceleration in
pitch which would be relatively small.
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Unlike piston engines the moving components of the gas turbine engine have a simple
rotary motion and the combustion process is continuous as apposed to intermittent as
in the case of a piston engine thus the unbalanced forces which could result in
mechanical vibrations would be small in size and few in number, therefore
mechanical vibration should not be a major issue.
Figure 120: - The F-136-2 engine / fuselage mounting arrangement based on that
of the Lockheed F-104 Starfighter, illustrating the loads carried by the engine
mounts, as explained above.
Acoustic vibration on the other hand resulting from the jet plume could have had a
major impact on the down stream empennage booms which had to have inboard faces
angled outboard by 150 to avoid the nozzle acoustic wave cone, because prolonged
exposure would result in acoustic fatigue. This angling outboard of the inboard boom
faces was used on the F-35 family aircraft to reduce potential for acoustic fatigue, for
the short stock booms to which the horizontal are mounted, by comparison the FB-24
/ A-24 booms were nearly twice as long and tapered so the risks were perceived to be
grater for this aircraft.
Although the AVEN nozzle was a thrust vector nozzle the potential for acoustic
damage resulting from its use off the aircraft centre line was considered slight because
of the very short dwell times involved (fractions of a second on average) to invoke a
manoeuvre, therefore the current layout was deemed adequate to withstand any
impingement.
This concludes this thesis coverage of the engine system integration, and the
integration into the fuselage substructure is shown in section 4.6 below (to the extent
to which this can be reported).
Free - sliding trunnion mount for engine radial
expansion.
Fixed trunnion mount.
Link – mount for engine
axial expansion.
Vertical
Forward Lateral
F-136-2 Engine
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4.4 Weapons systems integration for the FB-24 and A-24.
This section covers the accommodation of the primary and probable future weapons
within the FB-24 and A-24 airframes and the resultant sizing of the internal weapons
bays, as well as the installation of the avionic systems required for target acquisition
(within ITAR restrictions) and the proposed defensive systems for these aircraft.
The primary weapons for these aircraft called up in the requirements document were
as follows:-
JDAM GBU-31PIP
Release weight: 959.3kg (2,115lb) Length: 3.77m (12.38ft) Max span 0.6m (2.1ft) Cost $20,000
ALOSNW (Projected weapon for FSAV) Release weight: 1133.9kg (2,500lb)
Length: 4.26m (14ft) Max span: 0.45m (1.5ft)
Warhead: Evolved W69 (200kt‟s)
Air-to-Air Weapons
ASRAAM Missile Launch weight: 87kg (192 lb) Length: 2.8m (9.51 ft) Max span: 0.48m (1.6 ft) Body diameter: 0.16m (0.54 ft) Launcher rail weight: 22.8kg (50 lb) Launcher rail length: 2.8m (9.5 ft)
The M61A1 20mm cannon requested in the requirements documentation was to be
fitted as an optional weapon in the form of a missionized pod external to the airframe
structure as per the F-35C, and the F-35B aircraft, in order to increase internal fuel
provision and simplify the structural design of the forward fuselage reducing
structural weight. This was considered as acceptable by the customer as the internal
provision for this weapon was solely based on legacy air superiority aircraft and had
no specific relevance to either the FB-24 or A-24‟s primary mission.
The ALOSNW was comparable in size weight and shape to the JASSM, figure 121 a
conventional weapon and this latter weapon was used to size the weapons bay volume
for the former as a public domain system which can be openly discussed within this
proposal.
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171
Figure 121: - JASSM mounted on a storage / transportation fixture as a
conventional low observable stand – off attack munition with a range of 200nm
but the same OML and dimensions as the ALOSNW this was used sizing and as
a future store for the FB-24 / A-24 weapons system. (Source Lockheed Martin
ADP).
AGM-158A / B JASSM is a 2,250lb turbojet powered cruise missile with a range of
more than 200nm, and a 1,000lb warhead multi-purpose warhead effective against
hardened targets, and a light weight composite airframe. This weapon uses GPS to
navigate to the vicinity of the target, where upon an IR sensor in the nose images the
target and the weapons computer system compares it to a visual template loaded
before the mission and based on reconnaissance images. This weapons small size,
planform aligned shape, and selective use of RAM, combine to make this a very
effective low RCS weapon which is reasonably cheap at $500,000 each. The JASSM
would be carried internally in common with all other FB-24 / A-24 expendable
weapons because as stated above neither of these aircraft have external weapons
provision as a result of the need to reduce airframe drag and maintain a low radar
cross section, both of these requirements could not be met from the external carriage
of stores. (Reference 29: - Ultimate Fighter Lockheed Martin F-35 Joint Strike
Fighter: Sweetman. Bill: Zenith Press USA: 2004).
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The Boeing GBU-39/B Small Diameter Bomb low cost precision strike weapon
shown in figure 122 below, was also required to be carried by the FB-24 / A-24
aircraft in order to carry a greater number of weapons internally than two JDAMS,
and sizing information for this was obtained from Boeing (Reference 30:- Marguerite
Ozburn Global Strike Systems, Boeing IDS Business Support, Communications and
Community Affairs, P.O. Box 516 St Louis MO 63166 e-mail), and given below: -
Small Diameter Bomb GBU-39/B
Length: 1.8m (5ft.9”)
Width 19cm (7.5”)
Weight 130kg (265lbs)
Warhead Penetration 3ft of steel reinforced concrete
Range >11.12km (60nm) with wings
BRU-61/A “smart” pneumatic carriage
Payload capacity four GBU-39/B Small Diameter Bombs:
Weight empty = 145kg (320lb): Max weight = 664kg (1460lb):
Dimensions: - length = 3.6m (143”): width = 40.6cm (16”) height = 40.6cm (16”):
Figure 122: - Installation of the BRU-61/A “smart” pneumatic carriage with four
GBU-39/B Small Diameter Bombs into a weapons bay possibly a B-1 Lancer.
(Source Boeing IDS Global Strike Systems)
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Figure 123: - Proposed ASRAAM launch method from the FB-24 / A-24 weapons
bay by retractable launch rail concept was based on the side weapons bay AIM -
9M/X ejector system of the F/A-22A using swing arms to deploy a stub rail
mounted inboard of the outer weapons bay hinge line (authors CATIA V5 model
based on references 14 and 29 no components sized except ASRAAM).
The ASRAAM infra red seeking short range missile would be carried for self defence
and was not considered as an offensive system. The weapons location within the
weapons bay rather than the external carriage as is the case on Tornado resulted in a
limited field of regard for the seeker head, being bound on one side by the weapons
bay door and above by the FB-24‟s lower fuselage, this being a different case from
the F/A-22A where the AIM-9 infer red seeker is swung out from fuselage side
weapons bays in clear air with a much better field of regard for the seeker head.
Therefore in the case of both the FB-24 and A-24 the ASRAAM‟s would be launched
in “Lock-on After Launch” (LOAL) mode which was the only option from the
weapons bay location. There would be issues with this approach because in short
range combat the target aspect could change dramatically and unpredictably, and by
the time the missile clears the airframe the target may not be where it was at firing
initiation, and something else may have assumed that location e.g. the FB-24‟s
wingman, so time could be lost in trying to re-acquire the target. These issues would
be addressed to some degree by off board tracking and data feed and further research
would be required to build confidence in this deployment.
As explained above in figure 123 the ASRAAM would be swung out of the weapons
bay on a stub launch rail using actuated swing arms mounted on the weapons bay
internal longeron / outboard side wall, this was in preference to using a door mounted
rail as the door would have become too heavy if stiffened to carry the rail, which
would have increased the size of the door actuators and surrounding structure. With
this system the ASRAAM exhaust plume is clear of the weapons bay and little
additional stiffening of the bay structure would be required.
Door
Bay longeron
ASRAAM
Swing arms
Actuator
Bay door hinges
Launch rail
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Figure 124:- The port side weapons bay of the F-35A common to the F-35 family
with one 2,000lb JDAM installed (Source reference 29).
Figure 125:- Illustrates the results of the first test in the FB-24 / A-24 redesigned
F-35C which shows that the proposed lager weapons bay of the FB-24 could not
be accommodated in the positions indicated for the F-35C weapons bay doors on
the Scalecraft model which appear to be incorrect. The conclusion was to move
the new bays aft which was closer to the aircraft C of G and beneficial.
Initial 4.5m long 0.63m wide 0.5m high weapons bay in F-35 model weapons bay door location impinges on intake duct, and therefore was rejected.
STOVL model 3.5m long 0.63m wide
0.5m high weapons bay is a snug fit.
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Figure 126:- The final weapons bay integration aft and clear of the bifurcated
intake duct with the doors starting at Frame Station FS-5905.9 and ending at
Frame Station FS-1155.6
After sizing studies using the data in table 2 page 34 and data from reference 15, as
well as fit checks in CATIA V5 (figure 125) and with 1/48 scale kit models of the F-
35 the weapons bay was sized at: - length = 4.55m: width = 0.64m: height = 0.51m
giving a volume of 1.48m3 for each bay which was sufficient to accommodate one
ASRAAM plus one JDAM GBU-31 PIP, or one ASRAAM plus one JASSM, or one
ASRAAM plus one ALOSNW per bay. Also up to six GBU-39B Small diameter
bombs per bay could be accommodated without the BRU-61 rack or four per bay with
the rack but without the ASRAAM missile or launch rail. These weapons bays met
the operational requirements captured in section 2 of this thesis, and their integration
into the extended fuselage is shown above in figure 126, achieved with analysis of
public release data which sites the weapons bay doors too far forward on the baseline
F-35C in relation to the bifurcated duct leading to a clash which was resolved by
moving the bays to start at Frame Station FS-5905.9, and close out at FS-1155.6.
In addition to the primary weapons currently in service or approaching service entry
the author was asked to consider two future weapons which will enter service by the
entry date of the FB-24 / A-24 airframes, which were the LOAAS battlefield loiter
attack weapon figure 127 and the SACM small cruise missile figure 128. From the
data provided both weapons can easily be accommodated within the weapons bay and
expelled using the same cold gas ejector system as would be used for the FB-24 / A-
24 primary air to ground weapons.
In all cases of air to ground weapons release the author proposed that a drop down and
retractable flow disruptor plate was installed at the front of each weapons bay to aid
clean separation of the store from the aircrafts flow field, so that the stores fall away
and are not captured in the weapons bay.
10.91m
1.78m
0.64m
4.55m Volume 1,48m
3
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Figure 127:- One of the future weapon considered was the LOAAS battlefield
loiter GPS guided weapon: length 0.91m (3ft): weight 45kg (100lb): This would
be ejector launched from the weapons bay, and is powered by a 13kg (30lb)
thrust turbojet engine, LOAAS has a straight line range of 100nm and a
15minute duration at 750ft and 200knts covering a foot print of 25nm2. (Source
Lockheed Martin ADP Skunk Works)
Figure 128:- The other future weapon considered was the SMACM small attack
cruise missile which can be used both as a reconnaissance or attack platform:
length 1.7m (5.8ft): weight 64kg (142lb): This would also be ejector launched
from the weapons bay, and is powered by a J45G turbojet imparting a range of
200nm at BCA/BCM. (Source Lockheed Martin ADP Skunk Works)
This concludes the weapons bay integration study for this thesis, and further work
required is highlighted in section 5.
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4.5 Structural layout of the FB-24 and A-24.
The objective of this section was to present the airframe structural layout maturation
for the NB2 configuration and the reasoning behind the design selections for the final
PWSC FB-24 / A-24 conceptual design submission.
This section covers structural layout description, structure to structure joint definition,
major component attachment selection, build joints, selection, as well as the material
selection to be used in the manufacture of the primary structural members. The
development of each major structural component was described in turn starting with
the wings, followed by the forward fuselage, centre fuselage aft fuselage and
ruddervators and culminating in the full airframe structural model.
In order to establish an inner mould line IML the author used notional constant skin
thicknesses based on the public domain values for legacy aircraft namely the F/A-22A
and F/A-18E Super Hornet (Reference 31: - Lessons Learned from the F/A-22 and
F/A-18E Development Programs RAND_MG276: by Younossi. Obaid: Stem. David.
E. et al: Prepared by for the USAF by RAND Project Airforce), and from the work of
Alan Baker et al in Composite Materials for Aircraft Structures (Reference 32: -
Composite Materials for Aircraft Structures second edition: by Alan Baker et al:
Published by the American Institute of Aeronautics and Astronautics 2004) nominal
skin thickness values of 5.08mm for the fuselage and 20mm wing root tapering to
4mm at the wing tip were used for all structural modelling. Due to the time constraints
involved in this conceptual design all frames, spars, and ribs were modelled with
thicknesses representative their respective flange widths not their web thicknesses
therefore all spars are modelled as 76.2mm, ribs are modelled as 63.4mm, frames as
50mm, longerons as 30mm, and keels as 50mm thickness. Thus the structure looks
over designed however the intention is to impart layout and not detail design to the
reader. Also due to the time constrains the structural joints between members were not
modelled but their type is indicated as either “Bath-tub” or “Tag on stiffener” were
possible no Cleats were to be used to reduce parts count.
The airframe structure was conventional for a modern interdiction aircraft employing
high-speed machined aluminium, near net forged titanium, and HIP processed
titanium, carbon BMI skins throughout, unlike the F-35 in which most of the wetted
area was carbon epoxy, and some aluminium honeycomb core stricture. The only
fabrications envisaged are the engine bay doors and the land based arrestor cover both
would be made from SPF/DB titanium. The structure was to be mechanically fastened
using the same build philosophy as the F-35C family thereby reducing cost of
ownership by ease of maintenance in service. Where possible handed parts have been
avoided and common non-handed parts developed for ease interchangeability of
components and structure. This extensive use of metallic substructure is due to two
key factors covered in references 30 and 31, namely the relatively high cost of
composite substructure compared to high-speed machined aluminium, and the formers
limited tolerance of ballistic damage. The major component build joint philosophy
was soft mate where skins, longerons and keels land on a single interface frame this
reduces structural weight compared with the more traditional hard frame to frame
mate joints employed on some legacy aircraft.
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4.5.1:- Wing structural layout and undercarriage integration.
Figure 129:- F-35 original single component aluminium and titanium wing
structure which was mated at a water line joint with the centre fuse frames,
resulting in a heavy and complex joint and this has been changed on all SDD
aircraft after AA-1. (Source Lockheed Martin Public affairs office)
Figure 130:- F-35 original single component carbon epoxy thermoset wing top
skin produced by fibre placement, the wing skin also incurred weight growth
with maturation although the concept was good from a signature reduction stand
point. This has also been revised in all SDD aircraft after AA-1. (Source
Lockheed Martin Public affairs office)
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Figure 131:- F/A-22A wing torsion box illustrating a two component wing
structural layout with spars connected to single piece frames, and is composed of
titanium and carbon composite spars, a lighter solution than that originally
selected for F-35. (Source Lockheed Martin Public affairs office)
Figure 132:- F/A-22A Carbon BMI wing skins which have proven resilient in
service to the supercruise environment, and these are lighter by area than the
original single component wing after the latter design had matured. (Source
Lockheed Martin Public affaires office)
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Figure 133:- Phase 1 initial FB-24 / A-24 wing structural layout overview
detailed in the text below and in figure 136.
The design choices available for wing major structural component of the FB-24 / A-
24 common airframe as shown the figures 129 through 132 above and were a
continuous wing torsion box with a centre fuselage waterline frame break shown in
figures 129 and 130 favoured for the F-35 - 230-5 PWSC design and built as F-35
SDD aircraft AA-1, or two separate wing components (port and starboard) attached to
continuous centre fuselage without any frame breaks shown in figures 131 and 132
which has been successfully used on the F/A-22A now in service with the USAF.
Three factors drove the author to select the latter wing design approach, which were: -
(1) Complexity of the split frame joint: (2) Weight implications on the fuselage
frames: (3) Weight implications for the continuous wing substructure. Unlike the
BAE Systems / Boeing Harrier where the wing sits on top of the frames attached by
six pick-up fixtures, the F-35 wing mate joint was at 40% of the fuselage depth, and
although distributed over a lager number of frames would still require substantial
thickening in these members to transfer the wing bending and shear loads from the
spars and the fuselage bending loads from the longerons and keel. Additionally apart
from the two Concept Demonstration Aircraft X-35A/B and X35C this depth of joint
had no in service reference on legacy aircraft and was considered high risk by the
author.
Therefore the more conventional design philosophy shown in figures 133 and 134 was
selected for the common FB-24 / A-24 wing, with a metallic substructure of titanium
peripheral members and an advanced aluminium alloy core substructure. The skins
were carbon BMI (Bismaleimide matrix) high temperature thermoset with Tg values
ranging from 1750C to 235
0C although normal operational temperatures for long
periods in the hot wet condition would usually be limited to 1500C (reference 31).
This material was also a proven material for the F/A-22A wing skins which operates
in the same environment envisaged for the FB-24 / A-24 airframe. The wing structure
is detailed below in figure 134.
Ti boundary structural members
Ti root rib
Al core sub -structure
Main undercarriage bays
Carbon BMI skin 5.52m
9.39m
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min
ium
rib
1
bath
tu
b jo
ints
to
sp
ars
3 p
art
Alu
min
ium
rib
2
bath
tu
b jo
ints
to
sp
ars
Alu
min
ium
cap
rib
3 b
ath
tu
b jo
ints
to
sp
ars
Wh
eel
bay
A
ctu
ato
r b
ay
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
182
Figure 135:- Phase 2 maturation FB-24 / A-24 wing structural layout overview
detailed in the text below and in figure 138.
The initial Phase 1 layout shown in figures 133 and 134 formed the basis for all
subsequent wing maturation studies. The author had concerns that the Phase 1 spar
pitch was too wide leading to large skin panels which were susceptible to buckling,
therefore three additional sub spars were added in Phase 2 maturation to break up the
Phase 1 skin panels into thinner panels which would be more resistant to buckling
(reference 31), as shown above in figure 135 and detailed below in figure 136. The
number of ribs was not increased as their function was to transfer the flaperon loads
and leading edge flap loads into the spars and the Phase 1 layout was considered
capable of achieving this.
During Phase 2 the wing attachment philosophy was selected based on reference 28,
chapter 8. From the types of wing root joint described in this work the most suitable,
and that with which the author has had first hand experience of was the shear lug type,
which were easily assembled, least costly to manufacture compared to splice plates,
and generally more economic for military aircraft with thin airfoil sections such as the
FB-24 / A-24 compared to transports. This system was also used on the F/A-22A so
the author considered that the weight penalty alluded to in reference 28 should not be
sever enough to exclude this type from consideration for the FB-24 / A-24 aircraft.
In the design presented in this thesis both top and bottom lugs are in the same plane,
and both take axial loads, with the vertical load being shared between them (although
this is difficult to predict, and the small moment arm produces high lug axial loads).
The major advantages were a strong fitting at relatively low machining cost.
New Ti structural members
New Al spars
Moment fittings
Fwd shear fitting
Aft shear fitting
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
183
All of these fittings were to be titanium so a logical solution was to make all spars
below Rib 1 titanium and split the Root rib into sections which would have bath-tub
joint into the spars as shown in figure 136 below.
Fig
ure
136:-
Det
ail
ed d
escr
ipti
on
of
Ph
ase
2 w
ing s
tru
ctu
ral
layou
t.
Ti
Lead
ing
ed
ge
sp
ar
wit
h
inte
gra
l sh
ea
r att
ach
men
t.
Al
/ T
i R
ib
1
Fla
pero
n
att
ach
men
t b
ac
ku
p r
ib a
nd
als
o
the f
uel
tan
k b
ou
nd
ary
wit
h t
he
wh
eel b
ay
bro
ken
by r
ear
sp
ar
Al
Rib
2 F
lap
ero
n c
en
tra
l att
ach
men
t
rib
sp
lit
by
sp
ars
4, 5,
an
d 6
Ti
Stu
b
sp
ar
1
wit
h
inte
gra
l
sh
ear
/ m
om
en
t att
ach
me
nt.
Al
/ T
i S
tub
sp
ar
2 w
ith
in
teg
ral
sh
ear
/ m
om
en
t att
ach
me
nt.
Ti S
tub
Rib
Al sp
ar
3
Al
/ T
i S
tub
sp
ar
4 w
ith
in
teg
ral
sh
ear
/ m
om
en
t att
ach
me
nt.
T
i S
tub
Rib
T
i T
ip R
ib
Al sp
ar
5
Al / T
i R
ib 3
Al
/ T
i S
tub
sp
ar
6 w
ith
in
teg
ral
sh
ear
/ m
om
en
t att
ach
me
nt.
Ti
Aft
sp
ar
wit
h
inte
gra
l
sh
ear
att
ach
men
t.
Actu
ato
r b
ay
Wh
eel
bay
T
i 4 p
art
Fw
d R
oo
t R
ib
Ti 2 p
art
Fw
d R
oo
t R
ib
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
184
Figure 137:- Final FB-24 / A-24 wing structural layout overview detailed in the
text below and in figure 140.
The final stage of wing maturation saw the deletion of the aluminium substructure
core to reduce weight for higher structural stiffness during relatively long periods of
high speed and high temperature flight and enhance damage resistance as shown in
figures 137 and 138. One concern the author had was the possibility differential
expansion between the aluminium and titanium components and the carbon BMI skin
with the aluminium likely to expand to an extent where cracking around fastener
sleeves could become an issue. Therefore the solution would be to move to an all
titanium substructure unlike that of the F/A-22A, (references 2, and 14), which
although originally planed as composite core wing had to retrospectively incorporate
titanium spars, this new wing would have no other substructure material. Although
more expensive to produce this would be traded – off over time by much higher
durability of the structure in terms of fatigue and thermal stress resistance, and
titanium has been proven to be fully compatible with carbon BMI skins in high
temperature sustained supercruise conditions by the F/A-22A EMD test phase, weight
would also be traded – off with reductions in the thickness of structural members.
In this final phase the undercarriage attachment and support structure was designed
although only in initial form and was quite possibly over engineered as time did not
permit the structural analysis the author had intended to undertake and a
rationalisation of this structure is proposed for further work undoubtedly it could
withstand the weight of the FB-24 or A-24 but could also by employed on an SR-71.
That having been said the undercarriage loads are carried out of the lugs on the two
ribs to the support stub spar and the rear torsion box trailing edge spar which appears
to be a satisfactory load distribution. Note although the flaperon actuator bay appears
shallow it is sited directly above the wing root faring which offers 20cm3 of growth if
Ti structural members
Undercarriage attachment lugs
Moment fittings
Fwd shear fitting
Aft shear fitting
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
185
required by extending out to the lower fairing skin without resorting to additional
blister fairings as in the case of the F-35 an F/A-22A. As will be described in the next
subsection the skins extend inboard of the fairing so that the root rib fastener line is
not exposed, on the top. This blending accommodates both RCS and aerodynamic
requirements without compromising either.
Fig
ure
138:-
Det
ail
ed d
escr
ipti
on
of
fin
al
win
g t
ors
ion
box s
tru
ctu
ral
layou
t.
Ti S
tub
Rib
Ti F
lap
ero
n a
ttac
hm
en
t R
ib 1
Ti F
lap
ero
n a
ttach
men
t R
ib 2
Ti S
tub
Rib
T
i T
ip R
ib
Ti K
ick
ed
Sp
ar
2
Ti S
tub
Sp
ar
1
Lead
ing
ed
ge T
i S
pa
r
Ti S
tub
Sp
ar
3
Ti K
ick
ed
Sp
ar
4
Ti In
term
ed
iate
Sp
ar
5
Ti U
nd
erc
arr
iag
e
att
ach
men
t S
par
5
Ti F
lap
ero
n a
ttach
men
t R
ib 3
Tra
ilin
g e
dg
e T
i S
par
Fla
pero
n a
ctu
ato
r b
ay
Fla
p a
ctu
ato
r b
ay
Un
derc
arr
iag
e a
ctu
ato
r b
ay
T
i R
oo
t R
ib
Un
derc
arr
iag
e a
ttach
men
t lu
gs a
nd
su
pp
ort
stu
b r
ib a
nd
stu
b s
par
str
uctu
re Ti S
tub
Sp
ar
6
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
186
4.5.2:- Fuselage layout and propulsion and weapons bay integration.
4.5.2.1:- Forward fuselage structure.
The forward fuselage retains the same basic OML shape as the F-35C including the
DSI intake however internally the structure was stiffened to accommodate the cut-out
for the second crew member as shown in figure 139, and this was aided by adopting a
single nose wheel undercarriage for land based operations rather than the larger twin
wheel catapult launch capable / carrier landing nose undercarriage of the F-35C. This
permitted substantially larger continuous keel beams, which do not have to diverge
around a broad wheel bay as shown in figure 140. The chine longerons extend the full
length of the forward fuselage sweeping upward at the intake to become thickened
shoulder longerons either side of the second crew station and additional intake chine
longerons were added outboard of the intake ducts. More substantial canopy landing
longerons were also incorporated for the aircraft commander (pilot) station with the
option of extending these further aft to form a hatch frame landing for the second
crew position, as currently this was to be incorporated into the forward fuselage top
skin. The radar array provision was also doubled over that of the F-35C by
incorporating a chevron mount swept at 520 to match the wing sweep with the radome
mounting line, giving a practical constant field of regard of 2400 which was intended
to resolve one of the issues raised in figure 97.
Initially high speed machined aluminium alloy was considered for the forward
fuselage substructure however issues of weight and thermal expansion differences
within the carbon BMI skin so alternatives were reviewed and at an early stage even
titanium was considered as in the Mach 3 SR-71. However reference chapter 4
revealed polyamide resin based carbon composites of the PMR family to the author,
which are compatible with carbon BMI up to temperatures of 1700 over long exposure
times, and have been used on advanced military aircraft substructure. Although these
structures have been produced by the pre-preg route which is a costly manufacturing
method, efforts are being made to develop a resin transfer moulding process for this
material by reducing its viscosity, by dissolving the resin in solvent, or in a low
viscosity reactive polymer, or chemical engineering to introduce twists into the
backbone of the polymer, and the author envisages an RTM route being available
early in the design cycle of the FB-24 / A-24 airframe. The major advantages of this
material are stability at high temperatures high stiffness and resistance to most
chemicals; however the cost of the forward fuselage structure would increase
substantially if RTM failed to mature as the main processing route for this material.
The principle dimensions of crew station size and longeron separation in this structure
are shown in figure 141 below, and the largely identical structure for the A-24 is
shown in figure 149 below.
4.5.2.2:- Centre fuselage structure.
The centre fuselage being less susceptible to kinetic heating than the forward fuselage
or the wings could be produced using high speed machined aluminium for the
substructure with the exception of the last frame which was the engine trunnion
attachment frame and must withstand close proximity to the higher temperature
section of the engine than the frames ahead of this engine mount frame. This frame
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
187
was also the mate joint frame for the aft fuselage and the wing torsion box trailing
edge spar attachment frame.
Fig
ure
139:-
Det
ail
ed s
tru
ctu
ral
layou
t m
od
el o
f fr
on
t fu
sela
ge
of
the
FB
-24 t
op
vie
w.
Ch
evro
n A
ES
A r
ad
ar
mo
un
t fr
am
e t
ota
l
mo
un
tin
g a
rea =
1.2
7m
Sta
rbo
ard
Fo
rwa
rd
ch
ine lo
ng
ero
n
Sta
rbo
ard
In
tak
e
ch
ine lo
ng
ero
n
Sta
rbo
ard
Fo
rwa
rd f
usela
ge
an
d h
atc
h s
up
po
rt lo
ng
ero
n
Sta
rbo
ard
Can
op
y
su
pp
ort
lo
ng
ero
n
Po
rt F
orw
ard
ch
ine
lon
gero
n
Po
rt C
an
op
y
su
pp
ort
lo
ng
ero
n
Po
rt F
orw
ard
fu
sela
ge
an
d
hatc
h s
up
po
rt lo
ng
ero
n
Po
rt In
tak
e c
hin
e
lon
gero
n
MO
SC
hatc
h h
ing
e f
ram
e
MO
SC
hatc
h f
ram
e s
eg
men
ts
Inta
ke s
up
po
rt a
nd
att
ach
men
t fr
am
es
N.B
.:-
All s
tru
ctu
ral m
em
be
rs a
re
RT
M C
arb
on
PM
R-1
5
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
188
Fig
ure
140:-
Det
ail
ed s
tru
ctu
ral
layou
t m
od
el o
f fr
on
t fu
sela
ge
of
the
FB
-24 u
nd
ersi
de
vie
w.
Bu
ild
jo
int
fram
e
582.0
mm
Tw
in c
on
tin
uo
us k
eel
bea
ms
No
se w
hee
l b
ay
308.8
mm
No
se w
hee
l att
ach
fra
me
Ch
evro
n A
ES
A r
ad
ar
mo
un
t fr
am
e L
O c
oate
d
field
of
reg
ard
240
0
Bif
urc
ate
d in
take d
ucti
ng
4.8
m
0.3
5m
N.B
.:-
All s
tru
ctu
ral m
em
be
rs a
re
RT
M C
arb
on
PM
R-1
5
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
189
Figure 141: - Principle structural dimensions for the spacing of the longerons,
and the nose wheel bay. Also shown are the dimensions of the MOSC crew hatch.
The angle of the chevron frame mounted AESA radar is also shown (the
mounting area of 1.27m2 compared to 0.682m
2 for the conventional F-35C
mount).
All of the wing attachment lugs interface onto titanium clevis lugs attached to the
aluminium frames (which there was not time to model) and all of the aluminium
structure has class packers where it interfaces with the carbon BMI skin to prevent
galvanic corrosion. Two substantial full length longerons and two full length keel
beams resist fuselage bending loads, and the weapon bays internal walls form
additional partial length keel beams, as shown in figures 142 through 144 below. The
frames consist of a major load baring core frame for the first four frames being closed
out by the weapons bay, and by fairing close out stub frames on the rear four frames
as shown in figure 145 (a), which differs considerably from the F-35 centre fuselage
frames as shown in figure 145 (b), where all three segments were primary load
bearing structure and consequently had to be of heaver construction.
4.5.2.3:- Aft fuselage structure.
The aft fuselage structure was all titanium from the outset and the approach was to
make the structure as ridged as possible, because of the two booms which house the
ruddervator actuators, and will move relative to each other and to this end a large
number of frames longerons and keel beams are employed. This structure was also
intended for structural analysis which is highlighted for further work, to determine if
additional keel beams are required in the F-5R and F-5L tanks running out to link
with the boom actuator mechanism beams which the author suspects or would the
20mm closure skins be capable of maintaining rigidity. The current structure is shown
in figures 146 and 147, the Niobium coated engine bay / tank nacelle skin was
3.4m
1.2m
1.5m
2.0m
520
0.9m
Nose wheel bay
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
190
replaced by carbon BMI on grounds of cost weight and effectiveness because none of
the surrounding structure could withstand a fire of 13700C should it occur. The land
based arrester hook is attached to the bottom of first frame of the aft fuselage, and the
rest of the frames split at the base to allow engine removal.
Fig
ure
142:-
Det
ail
ed s
tru
ctu
ral
layou
t m
od
el o
f ce
ntr
e fu
sela
ge
of
the
FB
-24 t
op
vie
w.
En
gin
e a
ttach
men
t fr
am
es A
l fo
rward
of
max t
em
pera
ture
zo
ne a
nd
Ti at
tru
nn
ion
att
ach
men
t / m
ate
fra
me
Win
g a
ttach
men
t fr
am
es A
l ex
cep
t fo
r re
ar
fram
e w
hic
h i
s
Ti an
d is t
he e
ng
ine / b
uil
d m
ate
jo
int
fram
e.
Al fu
ll len
gth
co
nti
nu
ou
s s
ecti
on
lon
gero
ns 7
0m
m b
y 5
0m
m
Ti
sid
e
walled
w
eap
on
s
bays
att
ach
ed
to
th
e
fram
es
form
ad
dit
ion
al lo
ad
path
En
gin
e t
run
nio
n
att
ach
men
ts
En
gin
e lu
g a
tta
ch
men
t
Un
derc
arr
iag
e lo
ad
carr
yin
g f
ram
es
So
ft b
uild
jo
int
1.7
9m
1.3
4m
2.2
4m
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
191
0.8
5m
Weap
on
s b
ay w
alls f
orm
ad
dit
ion
al lo
ad
path
s
Tw
o c
on
sta
nt
0.3
2m
dep
th K
icked
ke
els
So
ft b
uild
jo
int
Fig
ure
143:-
Det
ail
ed s
tru
ctu
ral
layou
t m
od
el o
f ce
ntr
e fu
sela
ge
of
the
FB
-24 u
nd
ersi
de
vie
w.
0.4
7m
7.6
m
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
192
Figure 144:- FB-24 / A-24 common centre fuselage structural arrangement plan
view, showing the propulsion wing integration.
Figure 145:- Comparison of the FB-24 / A-24 common centre fuselage structural
frame philosophy (a), with the F-35 frame philosophy of the 230-5 OML
configuration (b).
The aft fuselage bending loads are resisted by 100mm deep lower longerons and
70mmm by 50mm shoulder longerons as shown below. The ruddervator spigots are
Shear attachments
Moment attachments
Undercarriage load
carrying frames
Core fuselage
frame
Close out frames Split frames
FB-24 / A-24
Wing spar /frame
F-35
Wing joint
Shear attachments
(a) (b)
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
193
mounted through bearings in the boom spigot mounting plates, and are driven by the
actuator mechanism housed and mounted between the actuator drive bay keels shown
in figure 147 below. The structure of the ruddervators is shown below in figure 148
and is elaborated on there.
Fig
ure
146:-
Det
ail
ed s
tru
ctu
ral
layou
t m
od
el o
f aft
fu
sela
ge
of
the
FB
-24 t
op
vie
w.
Sta
rbo
ard
Ti fu
ll l
en
gth
co
nti
nu
ou
s s
ecti
on
sh
ou
lder
lon
ge
ron
s 7
0m
m w
ide b
y 5
0m
m d
eep
Sta
rbo
ard
Ti fu
ll l
en
gth
aft
fu
sela
ge c
hin
e lo
ng
ero
n
Po
rt T
i fu
ll l
en
gth
co
nti
nu
ou
s s
ecti
on
sh
ou
lder
lon
gero
ns 7
0m
m w
ide b
y 5
0m
m d
eep
Po
rt T
i fu
ll l
en
gth
aft
fu
se
lag
e c
hin
e lo
ng
ero
n
F-1
36-2
En
gin
e
Sta
rbo
ard
Ca
rbo
n B
MI
Nacell
e s
kin
Po
rt C
arb
on
BM
I
Nacell
e s
kin
Sta
rbo
ard
bo
om
actu
ato
r
sp
igo
t b
eari
ng
pla
te
Po
rt b
oo
m a
ctu
ato
r
sp
igo
t b
eari
ng
pla
te
N.B
.:-
All s
tru
ctu
ral
me
mb
ers
are
Ti
Ru
dd
erv
ato
r
Sp
igo
ts
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
194
Fig
ure
147:-
Det
ail
ed s
tru
ctu
ral
layou
t m
od
el o
f aft
fu
sela
ge
of
the
FB
-24 u
nd
ersi
de
vie
w.
N.B
.:-
All s
tru
ctu
ral
me
mb
ers
are
Ti
F-1
36-2
En
gin
e
Po
rt b
oo
m a
ctu
ato
r b
ay
s
Sta
rbo
ard
bo
om
actu
ato
r b
ays
Po
rt b
oo
m a
ctu
ato
r d
rive
bay k
eels
Sta
rbo
ard
bo
om
actu
ato
r
dri
ve b
ay k
eels
Po
rt C
arb
on
BM
I
Nacell
e s
kin
Sta
rbo
ard
Ti fu
ll l
en
gth
co
nti
nu
ou
s s
ecti
on
lo
wer
lon
gero
ns 3
0m
m w
ide b
y 1
00m
m d
eep
Po
rt T
i fu
ll l
en
gth
co
nti
nu
ou
s s
ecti
on
lo
we
r
lon
gero
ns 3
0m
m w
ide b
y 1
00m
m d
eep
Sta
rbo
ard
Ca
rbo
n
BM
I N
acell
e s
kin
F
R-5
Fu
el ta
nk b
ou
nd
ary
fra
mes
FL
-5 F
uel ta
nk b
ou
nd
ary
fra
me
s
Lan
d
based
a
rreste
r
ho
ok a
ttach
men
t fr
am
e
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
195
Figure 148:- Ruddervator internal structural layout (Port Ruddervator shown),
the ruddervator skin material would be a quartz single wave pass composite to
permit the radar emissions to enter the structural RAM core material beneath,
this structural RAM core would be used on all control surfaces.
Figure 149:- Final AI installation into completed structural forward fuselage
layout, structurally identical to the FB-24 layout only without second crewman
occupancy provision or fuselage hatch.
Unitised Ti substructure frame machining
from initial near net forging
Ti spigot mating fixture machining from
initial near net forging
Structural RAM honeycomb
filled with absorbent foam
Same structural members as forward fuse of FB-24 for manufacture and assembly commonality and cost reduction
All structural members
are RMT carbon PMR-15
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
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Figure 150 above shows the full integration of the major systems with the major
structural build modules within the carbon BMI skin. Shown in figures 151 and 152 is
the F-35C internal structural layout for comparison with that of the FB-24.
Fig
ure
150:-
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Figure 151: - F-35C side view internal structural layout 230-5 configuration.
Figure 152: - F-35C plan view internal structural layout 230-5 configuration.
The substructure of the FB-24 / A-24 common airframe has been described above to
the extent of its current maturation, and during the preparation of this work the F-35C
has also matured beyond the 230-5 configuration used above for comparison, but the
matured structure is ITAR restricted. Therefore when this work is used in the future
the reader is reminded that the capabilities quoted for the F-35C represent only the
230-5 configuration and not any final SDD or production aircraft. This concludes the
structural description section of this thesis, and general arrangement views with key
dimensions are presented in figures 156 through 158 below.
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4.6 Fuel system integration and tank layout of the FB-24 and A-24.
The fuel tank layouts for the two crew FB-24 and the A-24 UCAV shown below in
figures 153 and 154 respectively the airframe contours and ruddervators are shown for
special orientation. The fuel tank system for both aircraft was identical except for the
provision of the F-1 tank in the forward fuselage of the A-24, which could also be
installed in a single seat version of the FB-24 if that were adopted as a direct F-35
replacement. The dimensions were determined from the measurement of CATIA V5
solid models produced by the author, and the capacities were calculated from the data
provided in reference 28and presented below: -
JP-5 (RAF) (JP-5 Mil – spec density of 6.82lb/US gal equal to 51.10lbs/ft3).
JP-8 (USAF) (JP-8 Mil - spec density of 6.80lb/US gal equal to 50.86lbs/ft3).
JP-4 (Special fuel) (JP-4 Mil – spec density is 6.55 lbs/US gal equal to 49lbs/ft3).
From these values and the volumes measured from the CATIA models the total fuel
capacity using JP-8 as the standard fuel selected for both aircraft:-
For the FB-24 in two crew configuration of 22,238lb reduced to 21,126lb with foam
filling:
The A-24 UCAV capacity of 26,458lb also reduced with foam filling to 25,135lb.
These values compare favourably with the standard F-35C which in 230-5 OML form
had a publicly quoted internal fuel capacity of 19,100lb of JP-8 and also indicate that
a single seat version of the FB-24 has the potential for substantially longer range
missions. The fuel tanks specified were of the self – sealing integral foam filled type,
common in modern combat aircraft which results in 2.5% of the fuel volume being
lost to displacement of fuel by the foam and a further 2.5% of the fuel being absorbed
by the foam. Consideration in the fuel tank layout was given to the method of in-flight
refuelling and the placement of the Integrated Power Pack Unit (IPPU) and the units
associated exhaust ducting.
Unlike the current USAF fighter inventory in-flight refuelling of this aircraft was to
be conducted using the „probe and drogue‟ method instead of the „flying boom‟ which
although representing a lower fuel transfer rate, would enable more aircraft to refuel
from a single tanker and commonality with the USN / USMC / and NATO allies who
all use the „probe and drogue‟ system. This system removes complexity from the
centre fuselage and enabled the enlargement of the F-2 and F-4 tanks and the resulting
increased internal fuel capacity. The provision of the systems envelope and the weight
balance required by the IPPU reduced the scope for extending the F-4 tank below the
level of the weapons bays which form the lower boundary of the F-2 tank in the fuel
system layout, because the IPPU location was a critical systems placement for the
following reasons: - structural support of the IPPU, systems routing, minimise exhaust
gas signature, and impingement of the exhaust gas on the left ruddervator skin, the
exhaust would be re-routed to exit lower left below the aircraft fuselage instead of
above it as in the F-35 family ( there was not time to model this and it fall into the
scope of future work). This complete the fuel system integration section of this
conceptual design study within the constraints of this thesis.
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Figure 153:- The FB-24 fuel tank layout and sizing in volume and capacity with a
total capacity for JP-8 = 22,238lb equal to 21,126lb when the reduction for foam
filled tanks is made. Note airframe boundary and ruddervators for reference.
Figure 154:- The A-24 fuel tank layout and sizing in volume and capacity with a
total capacity for JP-8 = 26,458lb equal to 25,135lb when the reduction for foam
filled tanks is made. Note airframe boundary and ruddervators for reference.
F-3R tank volume = 4.1m3
capacity = 5,905lb JP-8 fuel
F-2 tank volume = 3.08m3
capacity = 4,439lb JP-8 fuel
F-3L tank volume = 4.1m3
capacity = 5,905lb JP-8 fuel F-4 tank volume = 2.66m3
capacity = 3,827lb JP-8 fuel
F-5R tank volume = 0.751 m3
capacity = 1,081lb JP-8 fuel
F-5L tank volume = 0.751 m3
capacity = 1,081lb JP-8 fuel
F-3L tank volume = 4.10m3
capacity = 5,905lb JP-8 fuel
F-4 tank volume = 2.66m3
capacity = 3,827lb JP-8 fuel
F-5L tank volume = 0.751 m3
capacity = 1,081lb JP-8 fuel
F-5R tank volume = 0.751 m3
capacity = 1,081lb JP-8 fuel F-3L tank volume = 4.10m
3
capacity = 5,905lb JP-8 fuel
F-2 tank volume = 3.08m3
capacity = 4,439lb JP-8 fuel
F-1 tank volume = 2.93m3
capacity = 4,221lb JP-8 fuel
IPPU Location
IPPU Location
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Fig
ure
155:-
FB
-24 P
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Fig
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156:-
FB
-24 S
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Fig
ure
157:-
FB
-24 G
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5.0 Conclusions and further work recommendations.
The objective of this thesis was to produce a conceptual design study for a force
package consisting of a two seat advanced interdiction aircraft and a complementary
UCAV capable of replacing the UK Royal Air Force Tornado, Royal Australian Air
Force F-111, and USAF F-117 and F-15E air assets, both of these aircraft would have
greatly enhanced capabilities in stealth, range, and supercruise capability. The
performance being measured against a representative future mission profile produced
by the American Institute of Aeronautics and Astronautics, and the USAF Academy.
Table 13: - Comparison of RFP with F-24 / A-24 Capabilities.
Customer Needs RFP Requirements F-24 / A-24
Capabilities Meets Need
Maximum Mach number 45K ft
1.6 1.6 / 1.6 Yes
Weight GTOW 71,000lbs 53,187lbs / 56,897lbs Yes
Max Span Not > 65ft 45.02ft Yes
Accommodation 2 crew / AI unit 2 crew / AI unit Yes
Weapons internal load
4,614lbs 4,194lbs / 3,974lbs No*
Max Range internal fuel.
1800nm 1,283nm / 1,453nm No*
SEP, M=0.9, mil thrust, h=0.
0ft/s 200ft/s Yes
SEP, M=0.9, max thrust, h=0, n=1.
300ft/s 800ft/s Yes
SEP, M=0.9, h=0,n=5 max
thrust 50ft/s 670ft/s Yes
Maximum Instantaneous turn rate h=0
8.0 deg/s 18 deg/s Yes
Cost 500units $75,000,000.00 $ 45,823,547.00 Yes
*Capability can be met within weight limit and engine thrust limits and cost
limits.
The configuration design and parametric analysis using both classical analysis and the
AeroDYNAMIC V2.08 analysis tool set resulted in the analysis models being of a
much cruder level of definition, which made the drag results pessimistic because the
fuselage had to be defined as a ellipse and therefore the blending and side wall
sloping which reduced drag as well as RCS but this could not be evaluated. This
phase was originally intended to include CFD analysis of the CATIA V5 surface
models produced enabling precise drag analysis but this could not be incorporated
because of the time scales involved in the project. As a result the two best
configurations NB1 and NB2 went forward for analysis with AeroDYNAMIC V2.08.
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The NB1 being a trapezoidal wing planform with a conventional empennage
configuration with two fixed vertical tails and rudders, and two all moving horizontal
tails. The NB2 configuration had the same trapezoidal wing and fuselage, but two all
moving ruddervators instead of the conventional empennage arrangement of NB1.
From the AeroDYNAMIC analysis there was a drag penalty for the conventional
tailed NB1 configuration which the author expected, and a weight penalty as can be
seen in table 11 in section 3 of the thesis, which contains weights and cost data for the
two configurations based on the RAND Corp DAPCOIV model within
AeroDYNAMIC. These results also indicate that both aircraft could be built below
the target values in the USAFA / AIAA Request For Proposals which is more a
reflection of the level of sophistication and fidelity of modelling tool and a much
more accurate tool is needed to determine if this assertion is close to reality. As a
result of completion of this phase of the design study the NB2 ruddervator
configuration was selected to be matured into a full structural concept layout in phase
two, the NB1 concept configuration was held in reserve and not developed further.
The major structural component layout of the complete aircraft and systems
integration, without detailed structural sizing of the substructure beyond notional sizes
based on legacy aircraft and academic texts due to time constraints. The primary
objective was to determine a structural layout capable of accommodating the principal
systems i.e. the aircrew, AI (UCAV) system, propulsion system, weapons system, and
fuel system. The structural arrangement contained within section 4 of this thesis
accomplishes this objective incorporating advanced high temperature composite
materials namely carbon BMI for all aircraft skins and carbon PRM-15 for the
forward fuselage, high speed machined aluminium, and titanium for the centre
fuselage and wing structure, and titanium exclusively for the rear fuselage, with
ceramic skinned structural RAM used for the control surfaces and high lift devices.
The weapons bay size was increased over that of the current F-35C aircraft to
accommodate JASSM internally as well as all of the weapons contained in the
Request for Proposals. The aircrafts ability to supercruise will be dependent on the
development of the F-136-2 YF-120 derived engine, however the two crew FB-24
aircraft has a fuel capacity of 21,126lb and the unmanned A-24 has a fuel capacity of
25.135lb both exceeding the current F-35C value of 19,100lb. The current F-35 radar
frame was replaced by the author‟s chevron design swept at the same angle as the
wings this doubles the area available for AESA array mounting and substantially
increases the field of regard, and this could be incorporated into the current F-35
airframe.
The final FB-24 the manned and A-24 unmanned Advanced Interdiction Aircraft
contained herein has common forward fuselage skin OML with the F-35 aircraft in
order to retain a degree of commonality, the rear fuselage top skin OML is also
similar. The build modules are different consisting of three fuselage modules instead
of four for the F-35, this new arrangement being more structurally efficient, and the
wing attachment philosophy is the conventional fuselage side interface with a two
piece wing instead of the current continuous wing with split frames of the F-35.
Although the FB-24 and A-24 airframes evolved into a larger configuration and differ
in structural materials and layout from the F-35C they can be seen as a next
generation long range Joint Strike Fighter and as a follow on in the same way as the
F/A-18E and F have grown out of the F/A-18 Hornet program, and these aircraft
would still be less expensive than a clean sheet of paper design. Although at their
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current weights they are below the RFP targets in payload and range there is more
than sufficient weight growth margins for increased fuel and payload capabilities.
Future work required to move this design proposal forward into a preliminary design
is outlined below: -
1. Detailed parametric analysis, using a more sophisticated analysis tool.
2. Computational Fluid Dynamics investigation of the true OML drag values
through the following Mach numbers: - 0.8: 0.9:1.0:1.4:1.5:1.6 and 1.8 as well
as 1800 turns at Mach 1.6 with control surfaces deployed at angles throughout
their range of motion given in section 3, to determine control effectiveness and
drag contribution.
3. Optimisation of the structural layout by using hand calculations and the results
of the analysis in further work activity 1 to create a Vn diagram for each major
build component, and then create a Finite Element model of the structure
based on the initial layout models created by the author who would be willing
to supply them through Cranfield University School of Engineering. This
analysis would start with the wing structure to rationalise the main
undercarriage interface, and then explore the kicked spare to rib interface
joints and the control surface and high lift device interfaces, and thereafter
develop a skin model to determine pad up requirements, then moving on to
address the issues raised in section 4 with respect to the stiffness of the rear
fuselage booms.
4. Develop a conceptual design for the MOSC hatch and determine the need for
any additional landing support longerons within the forward fuselage
structure, over the provisions which could be made in the skin.
5. Model the wing to centre fuselage to wing interface joints on the frame side
using the wing fittings as a base point, to determine if the aluminium frames
could take the wing interface loads distributed through Ti bolt on fittings as
the author feel would be possible.
6. Determine more accurate component weight values for adoption as targets for
the as draw structure.
These are the main issues which need to be addressed before PDR can be
contemplated and could form the basis for another Airframe Engineering students
study at Cranfield University.
6.0 References and literature review:-
1) (a) Aircraft Engineering Projects for Engineering Students: by Jenkins.
L. et all: Published by the AIAA Education Series 2003, as referenced
in section 1 introduction and 2 requirements capture.
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1) (b) Aerodynamic V3.0 software produced by the USAFA and released
through the AIAA Education Series 2004, as referenced in sections 1,
introduction and 2 requirements capture.
2) F-22 Raptor pages 87-88: by Sweetman. B.: Published by MBI in the
USA 1998, as referenced in section 2 requirements capture.
3) Iron Hand (Smashing the Enemy‟s Air Defences), pages 284-285:
Published by Patrick Stephens Ltd 2002, as referenced in section 2
requirements capture.
4) Aircraft Concept Design Synthesis page 7: by Howe. D.: Published by
Professional Engineering Publishing 2000, as referenced in section 3
FB-24 and A-24 concept design and selection.
5) Fundamentals of fighter design 1st edition: Whitford R: Published by
Airlife 2003, as referenced in section 3.
6) Lockheed Martin F-35 Joint Strike Fighter: The Universal Fighter: by
Harkins, H.: Published by Centurion Publishing UK 2004, as
referenced in section 3.
7) Aerodynamics for Engineers 4th
Edition USAFA: by Bertin. J. J:
Published by Prentice Hall 2002, as referenced in section 3.
8) “Wind - Tunnel Investigations of Variable Camber and Twist Wing”
TND-8457: NASA: Aug 1977, as referenced in section 3.
9) Lockheed Martin F/A-22A Raptor: Stealth Fighter: by Miller. J.
Published by Midland Counties Publishing 2005.
10) Evaluation of Turbulent Models for High – Lift Military Airfoil
Flowfields: by Kern. S: AIAA96-0057, as referenced in section 3.
11) NASA Dryden Fact Sheet –F-16XL-1 Testbed aircraft as referenced in
section 3.
12) www.dftc.nasa.gov/Newsroom/FactSheets/FS-051-051-DFRC.html, as
referenced in section 3.
13) Lockheed Martin‟s Affordable Stealth paper by Haisty B.S, Published
by Lockheed Martin Aeronautics Washington D.C. 2000, for the
National Press Club.
14) F-22 Raptor (Americas next lethal war machine) by Pace. Steve:
Volume 1 of Walter J. Boyne Military Series. Published by McGraw
Hill 1999, as referenced in section 3.
15) The F-22 web site www.F-22.com.html as referenced in section 3.
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16) The AFT Contenders YF-22 and YF-23 Air Superiority into the 21st
Century 3rd
edition. By Sun Andy: Published by Concord Publications
Ltd. Hong Kong 1991, as referenced in section 3.
17) Superfighters (The next generation of combat aircraft) by Williams
Mel. Published by Airtime publications 2002.
18) Aircraft Performance and Design: by Dr Anderson. D. John. Jr
University of Maryland: Published by McGraw Hill 1999, as
referenced in section 3.
19) Modern Combat Aircraft Design: by Huenecke Klaus: Published by
Airlife Publishing 1987, as referenced in section 3.
20) Lecture notes AERO-481 Lesson 12: Survivability Propulsion
Integration and Systems, USAFA on reference 1 software title, as
referenced in section 3.
21) Aircraft Design A Conceptual Approach 3rd
Edition: by Raymer,
Daniel. P.: Published by the AIAA 1999, as referenced in section 3.
22) Introduction to Aeronautics a Design Perspective: by Brandt S. A. et
al: Published by AIAA 1997, as referenced in section 3.
23) Jet Bombers (From the Messerschmitt 262 to the Stealth B-2): by
Gunston B. et al: Published by Osprey Aerospace 1993, as referenced
in section 3.
24) Modern Fighters: by Spick. M.: Published by Salamander Books ltd
2000, as referenced in section 3.
25) Fundamentals of Fighter Design 2nd
Edition: by Whitford R.:
Published by Airlife Publishing 2005, as referenced in section 3.
26) Advanced Tactical Fighter to F-22 Raptor Origins of the 21st Century
Air Dominance Fighter: by Aronstein. D. C. et al Published by AIAA
1998, as referenced in section 3 and appendices B.
27) Have Blue and the F-117A evolution of the Stealth Fighter: by
Aronstein. D. C. et al: Published by AIAA 1997, as referenced in
section 3.
28) Airframe Structural Design (Practical design information and data on
aircraft structure): by Nui. M.: Published by Hong Kong Conmilit
Press Ltd 1997, as referenced in section 4.
29) Ultimate Fighter: the Lockheed Martin F-35 Joint Strike Fighter: by
Sweetman. Bill. : Published by Zenith Press 2004, as referenced in
section 4.
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30) Marguerite Ozburn Global Strike Systems, Boeing IDS Business
Support, Communications and Community Affairs, P.O. Box 516 St
Louis MO 63166 e-mail) as referenced in section 4.
31) Lessons Learned from the F/A-22 and F/A-18E Development
Programs RAND_MG276: by Younossi. Obaid: Stem. David. E. et al:
Prepared by for the USAF by RAND Project Airforce, as referenced in
section 4.
32) Composite Materials for Aircraft Structures second edition: by Alan
Baker et al: Published by the American Institute of Aeronautics and
Astronautics 2004, as referenced in section 4.
33) Aeronautical Engineers Data Book: by Clifford Matthews: Published
by Butterworth Heinemann 2002, as reference in appendices E.
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Appendix:
Appendices A: - Details of primary air and surface threats.
For this study the primary air to air threats are anticipated to be MiG-29C‟s:
MiG-31M‟s: and control aircraft of the A-50 Mainstay type because in the future
Global Strike ConOps force in which this aircraft is intended to operate
Eurofighter Typhoon, and F/A-22 Raptor air dominance fighters will be
responsible for tying down the Su-27C / P, and Su-30A enemy air assets, when
encountered. Also with the latter aircraft being the most expensive Russian
export platforms nations capable of procuring them are most likely to employ
them for free ranging attack fighters rather than homeland defence tied combat
air patrols. Although more capable point defence fighters such as the MiG Light
Weight Fighter Project figure A1 or the Chinese J-10 figure A2, could pose long
term treats to both the FB-24 and A-24 components of the mixed fleet, however
the J-10 could be detected and avoided and the Russian LWFP may never
appear as a real aircraft.
The MAPO MiG-29M: air defence fighter.
The statistics of this aircraft from published data (References 8:-Pages 60-76: Modern
Fighters – The Ultimate Guide to In Flight Tactics, Technology, Weapons and
Equipment: by Spick Mike: Published by Salamander books ltd London 2000 and
Reference 9:- Pages 15-67: Russian Air Power – 21st Century Aircraft, Weapons and
Strategy: by Gordon Yefim and Dawes Alan: Published by Airlife Publishing Ltd
2002) are given below:-
Dimensions: - Span = 37ft 3in: Length = 57ft 0in: Height = 15ft 6in: Wing Area =
409ft2
Aspect Ratio: - b2/S = 3.40
Weight: - Empty WE = 24,250lbs: Take off WTO = 37,038lbs
Power: - 2 x Klimov RD – 33K engines delivering a maximum installed thrust of
19,400lbs, and a military thrust of 12,125lbs each.
Fuel: - Internal = 9,978lbs: fuel fraction = 0.27
Loading: - Thrust to weight (maximum) = 1.05: Wing Loading = 91lbs/ft2
Performance: - Vmax high = Mach 2.30: Vmax low = Mach 1.23: Vmin = 110kt:
Operational Ceiling = 55,777ft: Climb Rate = 64,945ft/min
Weapons: - One Gryazev / Shipunov GSh 301, 30mm cannon with 100 rounds:
Two R-27R (AA-10a Alamo-A) medium range active radar homing missiles: and
up to Six R-73 (AA-11 Archer) short range infra – red homing missiles or Six R-
60M (AA-8 Aphid) short range infra – red homing missiles.
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Figure A.1:- MiG LWFP proposal highly manoeuvrable light weight fighter to
rival the F-35 and defend against UCAV‟s, a questionable project, source
Aviation Week & Space Technology (AW&ST) August 2003.
Figure A.2:- The Chinese J-10 comparable with the BAE Systems / SAAB
Gripen, this late 4th
generation aircraft could be heavily exported to potential
enemy combatants over the next two decades because it is projected to be
cheaper than Russian MiG‟s but have more modern systems however it has no
stealth features and could be avoided, source AW&ST January 2006.
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Air to Air missile capabilities: -
R-27R (AA-10a Alamo-A) intended to intercept and destroy aircraft, helicopters,
UAV‟s, UCAV‟s and ALCM‟s at medium ranges, day and night in all weather
conditions, from any angle, in ground or sea clutter and in the presence of a wide
range of countermeasures and defensive actions by a target.
The missile can attack targets at heights between 66ft and 88,583ft, and at speeds of
up to 1,890kts regardless of its initial position, within a field of view of +/- 50o. The
launch aircraft can pull 5G at the moment of missile release and the maximum height
differential between the launch aircraft and the target is 32,808ft above or below. For
the MiG-29 the AA-10 Alamo is mounted on two special underwing pylons.
R-73 (AA-11 Archer) short range heat seeking missile fitted with the Mayak (Beacon)
infrared detector head, is intended to intercept and destroy highly manoeuvrable
piloted and unmanned air vehicles, day and night from any angle in the forward and
rear hemispheres, in ground clutter and in the presence of intense ECM activity. This
missile has outstanding agility with the capability to destroy targets manoeuvring at
up to 12G. This missile can attack targets flying at heights between 66ft and 65,617ft
and speeds up to 1,350kts, from any initial position within a field of view of +/- 45o
and with an angular velocity of up to 60o/sec off boresight. Target designation for the
Archer‟s missile seeker head can be accomplished through the MiG-29‟s helmet –
mounted sighting and targeting system. All variants of the MiG-29 can employ the
AA-11 Archer missile which can be mounted on any of six underwing hard points on
the aircraft.
R-60M (AA-8 Aphid) this is an earlier short range heat seeking missile fitted with the
Komar-M (Mosquito) seeker head, and like the Archer it is intended for the
destruction of highly manoeuvrable targets but is limited to those which are within
visual range of the launch aircraft. The low launch weight of the missile and its
advanced aerodynamic layout give the Aphid outstanding agility and the capability to
destroy targets manoeuvring at up to 12G. This missile can engage targets flying at
heights between 98ft and 65,617ft and speeds of 1,350kts from any angle within the
pilot‟s field of view +/- 20o and at angular speeds of up to 35
o/sec off boresight. Like
the AA-11 the Aphid can be used for forward hemisphere engagements at short range
and the launch aircraft can pull 7G at launch without compromising the engagement
profile. The AA-8 Aphid has no equal in terms of weight and size in the west.
The MAPO MiG-31B: interceptor fighter.
The statistics of this aircraft from published data (References 8:-Pages 60-76: Modern
Fighters – The Ultimate Guide to In Flight Tactics, Technology, Weapons and
Equipment: by Spick Mike: Published by Salamander books ltd London 2000 and
Reference 9:- Pages 15-67: Russian Air Power – 21st Century Aircraft, Weapons and
Strategy: by Gordon Yefim and Dawes Alan: Published by Airlife Publishing Ltd
2002) are given below:-
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Dimensions: - Span = 44ft 2in: Length overall = 74ft 5in: Height = 20ft 2in: Wing
area = 663ft2
Aspect Ratio: - b2/S = 2.94
Weight: - Empty WE = 48,104lbs: Take off WTO = 90,389lbs
Fuel: - Internal = 36,045lbs: fuel fraction = 0.40
Power: - 2 x Soloviev D-30F6 afterburning bypass turbojet engines delivering a
maximum installed thrust of 38,580lbs and a military installed thrust of
20,944lbs each.
Loading: - Thrust to weight (maximum) = 0.85: Wing loading = 136lb/ft2
Performance: - Vmax high = Mach 2.83: Vmax low = Mach 1.23: Vmin = 140kt:
Operational Ceiling = 67,589ft: Climb Rate = 41,000ft / min
Weapons: - One Gryazev / Shipunov GSh 6-23, 23mm six barrelled cannon with
260 rounds: Four R-33 (AA-9 Amos) long range (162nm) active radar homing
missiles developed exclusively for the MiG-31: or Four R-40 (AA-6 Acrid)
medium range active radar homing missiles: and Two R-60 (AA-8 Aphid) short
range infra-red homing missiles. Additional configurations offered include up to
Six R-37 advanced long range active radar homing missile, and Four advanced
R-77 (AA-12 Adder) active radar homing missiles capable of sustaining 12G
while manoeuvring and having a range of 54nm under the wings on individual
pylons in overload condition.
All variants of the MiG-31 possess the capability of countering mass air attacks by
both manned aircraft and cruise missiles, by being able to track several targets
simultaneously and to launch all their missiles against individually identified targets
in salvo. Four MiG-31‟s are for example, capable of covering a 540nm linear front,
with each aircraft simultaneously tracking ten targets and able to engage four. Unlike
current Western fighters which have to engage their targets in comparatively narrow
azimuth and range limits, the MiG-31 has an engagement zone of +/- 70o of boresight.
The Ilyushin / Beriev A-50 Mainstay AWAC aircraft
The A-50 differs from the basic transport Il-76 Candid, from which it was derived, by
having a slightly lengthened fuselage and more noticeably, by the presence of the 29ft
6in diameter rotodome of the Shmel (Bumblebee) AWAC suite (known by NATO as
Squash Dome). The usual operating height of the A-50 is around 32,808ft and the
normal operating endurance, without air – to – air refuelling AAR is four hours.
The A-50 carries a fifteen – man crew, of whom five are fight – deck crew (pilot, co-
pilot, two navigators and a flight engineer), plus a rear crew of ten to operate the radar
consoles and other mission equipment.
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The entire mission complex of the A-50 is called the Shmel, developed by NPO
Vega-M, and includes the radar itself, Identification Friend or Foe IFF, the
information processing system and operators‟ consoles and an Electronic
Countermeasures ECM resistant digital communications suite, linking the A-50 with
both ground command posts, SAM‟s, ships, and fighter aircraft control networks. The
coherent pulse Doppler 3-D radar has a 360o circular scan and permits detection and
tracking of airborne targets including low flying cruise missiles against local terrain
(including water and any other type of surface relief, such as steppe, forest or
mountain).
The system is also capable of detecting and plotting surface shipping. Airborne –
controlled intercept (ACI) are achieved both by use of automated data – links and in
manual mode, using normal radio channels and plain – voice radio commands.
Information on the movements of targets of interest is transmitted to the Command
Posts of the Automatic Control System (KP ASU / Komandnyy Punkt
Avtomatizirovannykh Sistyem Upravleniya) of Russia‟s armed forces by digital data-
link through special relay centres. The transmission range of such information to the
KP ASUs is 189nm to 1,080nm.
When working at extreme ranges from any of the command posts satellite
communications links are used to transmit target data where available. Although the
threat nations may not have the satellite coverage capability a lower specification
Automatic Control System is offered with the A-50 package as well as French system
upgrades. The A-50 has colour displays which give alphanumeric data, including
track number, course, altitude, and speed of friendly interceptors, plus their remaining
fuel state, in panoramic format. Data can also be recorded in a documented text
format for onward transmission.
The aircraft is capable of AAR and is equipped with chaff and flare dispensers for self
protection. This asset will be heavily screened by three MiG-29 aircraft in the
immediate kill zone following Russian defence training patterns therefore the AIA
signature must be maintained in all flight phases especially manoeuvre and weapons
release of either Advanced Anti Radiation Weapon or the ALOSNW stores.
Development of the A-50 commenced in 1965, but the first aircraft only entered
service 1984 and by 1992 there were around twenty five in service in total, although
with the end of the Cold War only ten are in front line units in Russia, the remainder
are possible exports. Licensed production is being sought by the PRC and India, also
Iran with Russian assistance has produced an indigenous derivative of the A-50 from
stock Il-76 airframes, and North Korea intends to purchase surplus Russia A-50‟s
through an export company. (Reference10: -Combat Aircraft Magazine issues: -
Volume 3 No4 and Volume 4 No 5 Published by Ian Allan 2001 and 2002
respectively).
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Primary surface to air threats come from the mobile S-300V SAM system, and
the highly mobile Buk-M1 (SA-11) replacement for the Kub (SA-6) SAM that
shot down Captain Scott O‟Grady‟s F-16 over Bosnia.
The S-300V (SA-12): - Gladiator SAM system.
The S-300V (SA-12a Gladiator) system has two missiles (developed from the anti-
ballistic missile SA-12B Giant SAM), one large and one small, the smaller one has a
peak velocity of Mach 6 and can destroy targets evading at 8g through clutter and
ECM / decoy systems over an effective range of 30nm at altitudes between 2,000ft
and 60,000ft, the second larger missile attains a peak velocity of Mach 8 and through
advanced terminal aerodynamics can destroy targets manoeuvring at 12g through
ECM / decoy systems at altitudes between 12,000ft and 80,000ft. Targeting
information can be obtained from the main Almaz NPO family of sensors or off bored
from the A-50 Mainstay or MiG-31 Foxhound airborne platforms, and are capable of
intercepting HARM missiles launched against them.
The on – board initial long range tracking and target acquisition radar system
currently exported is the frequency – agile F-Band Bill Board radar, which is backed
up by the Grill Pan phased array for aircraft tracking and launch / interception
solution generation, which is able to track up to twelve targets and control six SAM‟s
against them simultaneously. The S-300V is a mature system and has been fully
operational in Russia since 1986.
The S-300PMU-1 and the longer range S-300PMU-2 (SA-10 Favorit) variants are
specifically designed for the export market and are at time of writing operational in
North Korea, the Balkans, the Indo-Pakistan border area and Syria-all potential threat
regions.
In addition the Almaz S-400 Triumf (SA-20 Gargoyle) family which are improved
faster derivatives of the S-300V‟s are now available on the world market after
completing field trials. These use the same GAZ-66 and MAZ-543 6x6 or 8x8 chunky
wheeled self propelled launchers which can disperse from the central radar nervous
system along with survey, mess, dormitory and power generator modules fielded on
MAZ-543 chassis vehicles, allowing them to pop up at unexpected locations, and all
have A-50 and MiG-31 target data links. (Reference 12:- pages 284-285, Iron Hand
Smashing the Enemy’s Air Defences: by Thornborough. M. A. and Mormillo. B. F.:
Published by Patrick Stephens Limited an imprint of Haynes Publishing UK 2002)
The Buk-M1 (SA-11):- SAM system.
The Buk-M1 is the replacement for the Kub (SA-6) and is in full scale production for
the export market, with a single system which is mounted on 11 vehicles a defender
has 36 missiles ready to launch at any one time. The missiles have a reach of some 45
to 105nm, depending on model similar to that of the SA-10 Grumble, with a closing
speed of up to 14,770ft/sec. Also like the SA-10 this weapon system relies on radar
guidance coming from associated F-Band Continuous Wave pulse – Doppler Clam
Shell, tower - mounted Big Bird or 3-D Tombstone long – range surveillance / EW
radars, with I/J-Band Flap Lid phased array radar used for target tracking.
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As with the S-300V and S-400 this system can be linked to the A-50 so that target
information can be obtained without the risk of the system being exposed to attack
from anti – radiation missiles.
Short – range Man portable SAM‟s.
The man portable SAM‟s MANPADS of the SA-14/-15/-16 have imaging infra-red
and ultraviolet seekers operating at both ends of the visible spectrum, and the SA-16
Gimlet simply ignores flares altogether, and over 1 million have been sold to date
most to threat nations. The best counter to these is to fly above 20,000ft to avoid them
and anti aircraft artillery (triple –A) fire, and rely on older precision guided munitions
PGM‟s or newer JDAM‟s and J-series weapons in a fast pass lob well away from the
target area. But such tactics place non-stealthy aircraft within the prospective shooting
range of the SA-11 SA-10, and S-300V and S-400 SAM‟s detailed above. (Ref 12)
Commenting on this surface to air threat environment one pilot put it, “You have to
hide from them until you can kill them”. New fourth generation stealth fighters such
as the F-35 Joint Strike Fighter and F/A-22 Raptor, relying on golf ball and marble
Radar Cross Section‟s (RCS) respectively, aim to do precisely that – hide, not evade,
and right over the enemy‟s noses. High – altitude attacks at up-to 40,000ft would
help them avoid the operational envelope of infra-red seeking SAM‟s and triple
A, and speeds of Mach 1.3 – to – Mach 1.5 would reduce the effective envelope of
the large radar guided SAM‟s like the S-300V‟s, S-400‟s, and SA-11‟s.(Ref 11
and 12), which is the most important factor in determining the mission profile of
the FB-24 AIA, and hence the aircraft itself.
Appendices B: - Signature control.
B.1 Radar Signature Control.
The threat to aircraft of radar detection and tracking comes from a variety of sources,
as shown in Table 6. Since radar range is a function of the fourth root of the radar
cross-section (i.e. RCS1/4
) an order of magnitude reduction in RCS, for example will
give a 44% reduction in detection range:
R1 / R2 = [RCS1 / RCS2]1/4
= [1 / 10]1/4
= 0.56
Similarly, the search area of the radar will be reduced to 32% and search volume to
18%. Hence a very large reduction in RCS not one tenth but one thousandth is
necessary to achieve a tactically useful effect (e.g. a reduction of 82% in an enemy
systems useful detection envelope).
RCS depends on the following features: - aircraft shape: aspect angle or orientation
with respect to the enemy tracking radars line of sight (LOS): ratio of radar
wavelength to target size: polarisation of transmit and receive antennas of the enemy
radar: surface quality: and constitution of the target. Methods required to control RCS
depend critically on the size of the electrically conducting component being
illuminated compared with the wavelength of the radar signal illuminating it.
If the wavelength of the signal is much less than the physical size of the component,
and if it is smooth enough, radar waves reflect much as a mirror reflects light.
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Table B.1: - Typical radar threats.
Radar system. Frequency (GHZ). Wavelength (cm).
Early Warning. 0.15 - 0.2 150.0 – 200.0
3.0 - 4.0 7.5 – 10.0
Ground control
interception (GCI) 2.0 – 3.0 7.5 – 15.0
Height finders 2.0 – 7.0 4 – 15.0
Aircraft 8.0 – 20.0 1.5 – 4.0
Air – to – air missiles 10.0 – 20.0 1.5 – 3.0
SAM transportable.
Acquisition 0.15 – 3.0 10.0 – 200.0
Tracking 5.0 – 10.0 3.0 – 6.0
SAM mobile.
Acquisition 2.0 – 6.0 5.0 – 16.0
Tracking 5.0 – 13.0 2.3 – 6.0
Radar – guided anti
aircraft artillery AAA 14.0 – 16.0 1.8 – 2.0
(Reference 25: - page 137, Fundamentals of Fighter Design: by Whitford Ray:
Published by Airlife Publishing Limited UK 2000).
The RCS of an aircraft is determined by the magnitudes of two distinctly different
contributions: -
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(a) The aircrafts shape (Outer Mould Lines – OML) both overall and in detail
(including surface quality). (Realistic reductions in combat aircraft size have an
insignificant effect on their RCS): (b) The electromagnetic properties of the
airframe materials.
The major contribution features to RCS for conventional non-stealthy aircraft are as
follows: - engine compressor faces (forward) and turbines (aft) in radar LOS due to
their Doppler signature: engine air intakes and diverter plates: external stores,
including missiles seeker heads: wing leading edges: corner reflections at
intersections of fins and tailplanes: wing planform viewed from above or below:
radome and bulkhead, if transparent to illuminating radar: cockpit, including cavity
effect due to the very large number of corner reflectors: engine nozzle if viewed from
rear: flat slab-sided fuselage when viewed from side elevation. All of these features
are present on the BAE Systems Tornado GR-1A figure B1 below; Boeing F-15E;
and General Dynamics (now Lockheed Martin) F-111‟s rendering them holey
inadequate, except against the most primitive of enemy defences. (Ref 25)
Figure B.1: -Tornado GR-1A, externally similar to the GR-4 is the only RAF
deep strike aircraft and the lack of any LO capabilities contributed to aircraft
losses in the first Gulf Wars. Source: - BAE Systems.
The same features are equally evident on the F-15E Strike Eagle figure B.2, the
USAF‟s current strike asset, and the RAAF F-111C shown in figure B3.
Active sensors and
ECM systems
detectable emissions
External weapons, and
fuel high RCS signature
Afterburning engines high IR
signature and turbine face visible
Right angle sides and has
no planform alignment
Engine compressor
face visible to radar Tall fin at right angle to
horizontal tails
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Figure B.2: -F-15E Strike Eagle the USAF equivalent to the RAF Tornado
developed from the F-15D air superiority fighter. Source: - USAF Academy.
Figure B.3: -F-111E the USAF equivalent to the RAAF F-111C from which it was
developed, note the F-111‟s were retired from USAF service in the mid 1990‟s.
Source: - USAF Academy.
Smaller contributions to RCS for a non stealthy aircraft are as follows: - fuselage in
head on view: wing leading edge and control surface gaps which cause scattering: local
air inlets e.g. for cooling and air conditioning: local surface protuberances, even the
Two seat non coated
canopy
External stores only under wing
and under fuselage
No blending and no planform alignment
Twin fins at right angle to
horizontal tails
No blending and no planform alignment
Two seat non coated
canopy
External stores, under wings
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smallest protuberances cannot be ignored and each may become resonant at a different
frequency: long thin fairings including missiles fins and tailplanes.
Note if the major contributors to RCS are carefully dealt with then the previously minor
ones become important, therefore stealth is difficult to achieve but easy to lose through
lack of attention to detail. Aircraft shaping is useful over a wide range of radar
frequencies but over a limited range of aspect angles. Typically, for fighter aircraft, the
forward cone of angles is of greatest interest and, hence large returns can be shifted out
of this sector into the broadside directions. Aircraft can be shaped to ensure that most
radar waves will be scattered and not reflected back to the transmitter. Leading and
trailing edges of the wings, control surfaces, inlet lips, door gaps, etc., can be aligned to
ensure that the energy that is unavoidably, reflected back to the transmitter is
concentrated into a few spikes, as shown in figures B4 and B5 below.
This will give the enemy radar one good return when the alignment is ideal, but a
much weaker return on subsequent scans. (Ref 25)
Figure B.4:- Spike alignment or planform alignment on the F-35C, the alignment
of the leading and trailing edges of the wings, control surfaces, inlet lips, door
gaps, etc., are aligned to ensure that the energy that is reflected back to the
receiver is concentrated to 8 spikes away from the forward cone, also the canopy
is coated with a tin iridium oxide to make it opaque to radar, and the radome is a
selective wave pass design.
Spike
Alignment
Spike
Alignment
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Figure B.5:- The same spike alignment or planform alignment is applied to the
fuselage sides and vertical tails on the F-35C, which have the same common
alignment angle of 270 to ensure that the energy that is reflected back to the
receiver is concentrated away from the illumination source.
Two different approaches to aircraft shaping to reduce RCS have been developed
which primarily resulted from the amount of computing power available for analysis
when they were formulated: -
The first approach was the faceted configuration which used flat panels to minimise
normal reflections back to the illuminating radar. The major work on faceting for RCS
reduction was undertaken at Lockheed Skunk Works by Denys Overholser who
applied a work on diffraction published in 1962 by the Russian physicist Ufimtsev, to
the problem of RCS reduction. This lead to the development of a computational
method of predicting the RCS of two dimensional shapes built up from a series of flat
surfaces. This eventually gave rise to the Have Blue and F-117A Nighthawk aircraft
shown in figure B.6 below. (A detailed description of this approach is given in
appendix b of Reference 27).
The second approach was the use of a compact, smoothly blended external geometry
to achieve a continuously varying curvature, as employed in the Northrop Grumman
B-2 and the Lockheed Martin F/A-22A and the F-35 Joint Strike Fighter, this required
far grater predictive ability and enormously increased computational capacity. (A
detailed description of this approach is given in the appendix of Reference 26).
Spike
Alignment
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Figure B.6:- F-117A Nighthawk illustrating facetted construction to reduce RCS
and slotted engine exhaust to reduce IR signature as well as rear quarter RCS.
By the late 1950‟s engineers in the field of radar signature reduction realised that the
very large RCS reductions necessary to achieve any operational benefit could not be
accomplished simply by coating an otherwise conventional aircraft with radar
absorbent materials (RAM) (a lesson seemingly forgotten outside the USA for
example in the Dassault Rafale), Many physically small and apparently insignificant
features of an aircraft generate radar returns that are still quite detectable. Indeed a
very low RCS must be designed into an aircraft from the outset, with rigorous
attention paid to all three elements of RCS control: overall shaping of the airframe and
its components, special detail treatments, and RAM. (Ref 13)
The Advanced Interdiction Aircraft being derived from the F-35 JSF family of aircraft
will capitalise on the experience gained within Lockheed Martin and Northrop
Grumman, from development and production of their legacy stealth aircraft namely
the: - Lockheed Martin F-117A Nighthawk: F/A-22A Raptor: and the Northrop
Grumman B-2A Spirit. Additional experience from in service evaluation of these and
other low-observable UCAV projects within the F-35 team have been utilised to make
the F-35 more affordable and more easily supportable in front line units. The
Advanced Interdiction Aircraft will employ the same low observable technologies as
the F-35 family demonstrated on the high – fidelity SigMA model shown in figure B7.
A critical objective of the F-35 and the AIA program is to produce a stealthy aircraft
that stays stealthy in severe combat conditions. This goal is best achieved by building
an aircraft that is hard to damage requiring durable battle damage tolerant structure
and making sure that anything except the most severe damage will not significantly
degrade the RCS. The RCS of the FB-24 AIA would be validated by a similar test
programme to that detailed above for the F-35.
Slotted exhaust nozzle
Grill covered intakes
High drag facetted OML
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Figure B.7: - The SigMA model about to undergo RCS tests at the Helendale,
California facility. Source: - Lockheed Martin Press Release Office.
The design must also ensure that any required repairs are rapid and can be performed
at front line units on the flight line. The SigMA testing program demonstrated this and
the AIA will build upon the experience, as part of the F-35 family.
The SigMA models shown in figures B7 and B8 were full scale aircraft models
complete with a representative internal structure, and high – fidelity features including
removable doors and access panels, canopy transparency, cockpit interior, external
lights, air data probes, engine components, edges, re-positional control surfaces,
antenna apertures, radar array, and a flight capable nose – cone. Testing began in
February 2000 and included measuring aircraft RCS and the performance of various
antennas on the SigMA model aircraft. Tests also demonstrated the robustness of
supportable low – observable (LO) materials and their repair.
After baseline testing, several doors and panels were intentionally damaged and later
repaired, then installed on the model and RCS measurements were taken to determine
the impact of the defects and the effectiveness of the repairs.
The inflicted damage, more than three dozen significant defects represented in types
and frequency the cumulative effect of more than 600 flight hours of military aircraft
operations.
When engineers overlaid the RCS curves of the undamaged configuration on the
damaged and repaired configurations they found it extremely difficult to detect any
changes between the two configurations. These repairs were conducted quickly and
easily with engineers repairing all of the damage within a single eight hour shift. A
similar repair on a legacy B-2A would be estimated to require more than 72 hours.
RCS testing was completed by early October 2000, validating the F-35 OML, the
resilience of the LO materials, and previewing its cost – savings potential.
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Figure B.8: - SIGMA Model on test pole. Source: - Lockheed Martin Press
Release Office.
In further efforts to reduce maintenance and supportability costs and maintain
signature continuity in October 2000, Lockheed Martin demonstrated a new
generation of “paint – less” aircraft covering film on a fully covered F-16 which made
a series of successful test flights including one at Mach 1.8 which is the highest speed
ever reached in an appliqué flight test. Aircraft appliqués consist of paint –
replacement adhesive films designed to bring savings in production costs, support
requirements, disposal costs and importantly savings in aircraft weight. (They also
offer significant environmental advantages since military painting operations are a
significant source of hazardous – material emissions).
(Reference 6:- Lockheed Martin F-35 Joint Strike Fighter: The Universal Fighter: by
Harkins H: Published by Centurion Publishing UK 2004)
The Advanced Interdiction Aircraft will use this appliqué coating technology to enable
frequency tuned material application at reduced weight, and ensure signature
continuity over the airframe.
2.3.4 Infra-red Signature Control.
Passive (non-emitting) IR detection targeting devices rely on contrasts between hot
parts on an airframe such as jet pipes and surfaces subjected to kinetic heating, and the
background radiation. Many IR devices operate in the 8 to 13 micron band because
this is the most IR transparent band in the atmosphere. In engine exhaust, carbon
dioxide produces most of the IR signature at 4.2 microns, so most modern IR sensors
can “see” at two different wavelengths (medium:- 3 to 5 microns and long:- 8 to 14
microns) to provide good target discrimination.
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The radiated IR energy is proportional to the fourth power of absolute temperature (i.e.
T4). With engine turbine entry temperatures currently in the order of 1,627
o C, and
rising, the rear fuselage of a fighter is the greatest source of IR radiation. With
afterburner on, IR emissions are vastly increased. Moderate stagnation temperatures
(which vary with Mach number squared) are inevitable due to kinetic heating on
leading edges of fighters flying at high Mach numbers.
As the stealthiness of fighters improves, the exhaust plume of the missiles launched by
the fighters becomes a contributor to IR detection. Low – visibility plumes will
minimise detection of both launch platform and missile (vitally important for the
ALOSNW and the AIA). There are several ways of reducing the IR signature of a
fighter aircraft which are as follows:-
The ability to supercruise (cruise at supersonic speeds without afterburner)
limited to aircraft with a thrust to weight ratio of 1:1 or above like the F/A-
22 Raptor and YF-23, this reduces the temperature of the nozzles: moreover,
supercruising allows the pilot to engage on his terms, increases weapons
envelopes, minimises exposure to SAM threats and not only stretches
combat radius but forces an adversary to expend his own fuel in order to get
his aircraft to an energy state where he can engage it.
Although not currently available to the F-35 family, the JSF 119-614 engine
has demonstrated installed thrust of 50,000lbs and has further growth
potential, and also variable cycle engines like derivatives of the GE YF-120
ATF engine could offer the potential of thrust to weight ratios of 1:1 in the
future. The Advanced Interdiction Aircraft would also benefit from such
developments and an engine based on the GE VCE technology should be
explored for supercruise capability.
Use of a high bypass-ratio (BPR) engine to mix in cold air to reduce exhaust
temperature, however a BPR of more than 0.4 conflicts with the requirement
of high dry thrust to achieve supercruise.
Use of a curved jet pipe to mask the hot turbine stages.
Use of two – dimensional nozzles (which increase the surface area of the
exhaust plume) or ejector nozzles (which mix in ambient air) to increase the
rate of cooling.
Increase cooling of the outer skin of the engine bay or insulate to reduce the
temperature of the engine bay outer skin.
Use a curved air intake (the F-35 and AIA use a bifurcated duct) to mask
forward emissions from the engine.
Limit maximum supersonic speed to reduce IR signature due to kinetic
heating (one beneficial side effect of reducing the AIA dash speed).
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Figure B.9: - The Pratt & Whitney LON nozzle for use on the FB-24, undergoing
ground test during the Concept Demonstration Phase of the JSF program.
Source: - Pratt & Whitney Press Release Office.
Figure B.10: - The General Electric AVEN nozzle design of the type to be
employed on the A-24, not shown in this picture are the chevron trailing edge
tips the nozzle segments which reduce rear illumination radar returns
(Reference 15).
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The FB-24 in common with the F-35 family will employ the Pratt & Whitney Low
Observable convergent / divergent exhaust nozzle on the F-135 engine (ground tested
as shown in figure B9 below), or the GE Low Observable Axisymmetric Nozzle
(LOAN) on the F-136 engine (ground tested on a Lockheed Martin F-16C for over
500-hours) (References 15 and 16) as will the A-24, although employment of the GE
Aircraft Engines AVEN axisymmetric vectoring nozzle with exhaust vectoring in any
direction up to 20 degrees from the centre axis shown in figure B10 was also
considered.
The two different nozzle types considered for the FB-24 and the A-24 would be
tailored to each airframe rear fuselage design to optimise the performance and
survivability of each aircraft, but these nozzles would be common from the turbine aft
face up to the throat (the point of minimum cross-sectional area). Although the Pratt
& Whitney nozzle would not have a thrust vectoring capability the GE LOAN /
AVEN would have and the A-24 would require such a nozzle due to its direct
penetration role, and greater need for threat avoidance, the FB-24 would also benefit
from increased manoeuvrability.
Figure B.11: - The 2-D thrust vectoring nozzles of the F/A-22A shown here to
good effect, the top and bottom chevron tipped flaps move up and down in
unison to direct the engine thrust imparting high agility to this large aircraft
(Source: - Lockheed Martin Press Office).
The 2-D vectoring nozzle system as employed on the F/A-22A shown in figure B11
above was considered too heavy and complex and inappropriate for the missions
envisioned for the FB-24 and A-24, which would evade rather than engage the enemy
and therefore would not require the agility of an air dominance fighter.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
227
Appendices C: - Details of the F-35 Diverterless Supersonic Intake.
A fighter inlet must provide the engine or engines with high-quality airflow over a
wide range of speeds, altitudes, and manoeuvring conditions while accommodating the
full range of engine airflow from idle to maximum military or afterburning power.
The inlet designer must also consider the constraints imposed by configuration
features, such as nose landing gears, weapon bays, equipment access panels, and
forebody shaping. The designer must produce the lowest drag, lowest weight, lowest
cost, and highest propulsion performance solution. It must also meet stringent low
observable requirements.
Historically, inlet complexity has been a function of the top speed for a fighter aircraft.
Higher Mach numbers require more sophisticated devices for compressing supersonic
airflow to slow it down to subsonic levels before it reaches the face of the engine. (Jet
engines are not designed to handle the shock waves associated with supersonic
airflow.) These compression schemes involve the conversion of the kinetic energy of
the supersonic air stream into total pressure on the compressor face of the engine.
Speeds over Mach 2 generally require more elaborate compression schemes like those
of the F-14, F-15, and MiG-25 Foxbat MiG-31 Foxhound.
The F-15 inlet, for example, contains a series of movable ramps and doors controlled
by software and elaborate mechanical systems. The ramps move to adjust the external
and internal shape of the inlet to provide optimum airflow to the engine at various
aircraft speeds and angles of attack.
Doors and ducting allow excess airflow to bypass the inlet. The entire system results
in high radar returns and a high weight penalty for the aircraft, signature reduction was
not a priority when the “teen series” of aircraft was designed and built.
Fighter inlet designers must also account for boundary layer air which is a layer of low
- energy air that forms on the surface of the fuselage at subsonic and supersonic
speeds. (These layers also form on the inlet compression surface). This layer of
relatively slow moving, turbulent air, can create havoc when disturbed by the shock
waves created by the inlet. This results in unwanted airflow distortions at the engine
face. If the shock wave / boundary layer interaction is severe enough, the engine will
stall.
An important consideration for the Advanced Interdiction Aircraft is that this
boundary layer increases in thickness with speed and forebody length, as the AIA
seeks a 3ft increase in forebody length for the second crew member, later CFD
analysis should be directed to investigate the impact on DSI size and effectiveness, the
intake position may need to change as on the F/A-22 or the original X-35 / 230-2
OML intake configuration may have to be adopted with its inherent weight penalty (to
be addressed in future Trade Studies).
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
228
Fighter designers‟ deal with this boundary layer phenomenon by redirecting the layer
before it reaches the engine by placing the inlet away from the boundary layer in the
free stream, where airflow is unaffected by the boundary layer phenomenon. On the F-
16, a structure called a diverter provides a 3.3-inch gap between the fuselage and the
upper lip of the intake. The size of the gap is directly proportional to the thickness of
the boundary layer at maximum speed of the F-16. Other fighters remove boundary
layer airflow with combinations of splitter plates and bleed systems. The latter
redirecting the unwanted airflow through small holes in the compression ramps to
bleed ducts within the inlet. All of these systems result in radar waves being either
returned from the splitter plate / forebody cavity or from the compression ramps.
The Diverterless Supersonic Inlet bump shown in figure C1 functions as a
compression surface and creates a pressure distribution that prevents the majority of
the boundary layer air from entering the inlet at speeds up to Mach 2 as described in
the Aerodynamic V3 slide shown in figure C2.
In essence, the DSI dispenses with complex and heavy mechanical systems as well as
significantly reducing the RCS of the aircraft.
The DSI concept was introduced to the F-35 (then JAST / JSF) program as a trade
study item in 1994, and was compared with the traditional conical translating inlet
seen on F-111‟s. After CFD analysis supported by wind tunnel testing full-scale flight
testing was conducted using an F-16 at Mach 2, full scale testing culminated in the X-
35 Concept Demonstration Phase with both CDA X-35 aircraft built with the DSI
configuration and demonstrating its capabilities throughout the complete test program.
Figure C.1: - F-35 SDD mock up aircraft showing the DSI and a cut back intake
lip compared with the X-35 aircraft as a weight reduction measure. Source: -
Code One magazine reference 17.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
229
WaveriderWaverider--like like ““BumpBump”” diverts diverts
boundary layer using pressure boundary layer using pressure
gradientgradient
CFD tool advances allowed for CFD tool advances allowed for
integration into todayintegration into today‟‟s vehicless vehicles
STEPS:STEPS:
Define 3Define 3--D Compression Surface D Compression Surface
From From ““Virtual ConeVirtual Cone”” CFD SolutionCFD Solution
•• Early:Early: Traditional ConeTraditional Cone
•• SOA:SOA: Isentropic Cone at Isentropic Cone at
Develop Develop ““CenterlineCenterline”” GeometryGeometry
•• Compression Surface / Compression Surface /
Shoulder / Diffuser Fairing Shoulder / Diffuser Fairing
IntegrationIntegration
•• Cowl / ACowl / Aii / A/ A
tt IntegrationIntegration
Integrate Complete Inlet / Integrate Complete Inlet /
Forebody Forebody
•• Forebody / Aperture / Duct Forebody / Aperture / Duct
IntegrationIntegration
•• Real Aircraft ConstraintsReal Aircraft Constraints
* Protected by U.S. Patents 5,749,542 & 5,779,189* Protected by U.S. Patents 5,749,542 & 5,779,189
ADVANCED INLET INTEGRATION
Diverterless Inlet Technology*
Conic
al S
hock
Conic
al S
hock
Stream Lines
Stream Lines
CompressionSurface
CompressionSurface
FLOWFLOW
Virtual ConeVirtual Cone
1
ConeCone
Comp Surface
Comp Surface
TransitionShoulder
TransitionShoulder
DiffuserFairing
DiffuserFairing
CowlCowl
2 3
NACA RM E56L19 (1957)NACA RM E56L19 (1957)
Compression OnlyCompression Only Compression + BL DiversionCompression + BL Diversion
Figure C.2: - The operational principals of the DSI and CFD representation of
the boundary layer flow fields around DSI with the original X-35 intake lip
configuration. Source AIAA AeroDYNAMIC V3.
A more detailed description of the flight test DSI program on the F-16 can be found in
reference 17 which was the source of the material presented in this section.
(Reference17:- Pages 9 to 13, Code One magazine: Volume15 number 3: by Hehs
Eric: Published by Lockheed Martin Aeronautics Company 200)
Appendices D: - Current F-35 family of aircraft.
The current F-35 Joint Strike Fighter family consist of three aircraft variants for three
United States military air arms and the United Kingdom armed forces, which have
common basic configurations and use cousin parts in most cases, are outlined below
in terms of lead dimensions: performance: and weight and are shown in figures D1 -
D6.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
230
Figure D.1: - F-35A USAF (CTOL Variant). Source: - Reference 3:- Document
number 17521 Briefing by Brig Gen J. Hudson.
F-35A (CTOL Variant): - This is the USAF variant with 9g manoeuvring capability
and replaces the F-16 and A-10 (current orders are for 1763 aircraft). Dimensions
are: - Wing span 35 ft: length 51.1 ft: and wing area 460 ft2. Empty weight is: -
27,395 lbs. Internal fuel capacity is: - 18,498 lbs. Combat radius is: - greater than
590 nautical miles.
Figure D.2: - X-35A CDA for F-35A in fight 2001:-Source Authors Private
collection.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
231
Figure D.3: - F-35B USMC / UK RAF and RN (STOVL Variant). Source: -
Reference 3:- Document number 17521 Briefing by Brig Gen J. Hudson.
F-35B (STOVL Variant): - This is the USMC / U.K. RN & RAF variant to replace
the Harrier AV-8B / Mk - 9 and the F/A - 18 Hornet C/D aircraft (current orders are
for 759 aircraft). Dimensions are: - Wing span 35 ft: length 51.1 ft: and wing area
460 ft2. Empty weight is: - 30,697 lbs. Internal fuel capacity is: - 13,326 lbs.
Combat radius is: - greater than 450 nautical miles.
Figure D.4: - X-35B CDA for F-35B Edwards Air Force base tests 2001:-Source
Authors Private collection.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
232
Figure D.5: - F-35C USN (CV Variant). Source: - Reference 3:- Document
number 17521 Briefing by Brig Gen J. Hudson.
F-35C (CV Variant): - This is the US Navy aircraft carrier variant which will
replace the F/A - 18 Hornet C/D and the F - 14D Tomcat (current orders are for 480
aircraft). Dimensions are: - Wing span 43 ft: Length 51.4 ft: Wing area 620 ft2.
Empty weight is: - 30,618 lbs. Internal fuel capacity is: - 19,100 lbs. Combat
radius is: - greater than 600 nautical miles. Maximum payload all internal and
external stores stations used: - 21,400lbs.
Figure D.6: - F-35C in maximum stores configuration: - Source Reference 3:-
Document number 17521 Briefing by Brig Gen J. Hudson.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
233
Appendices E: - Design Supplement.
The objective of this appendices was to cover aspects of the FB-24 and A-24 design
which could not be covered in the bulk of the report and the following aspects of the
design are covered herein: -
1) Structural design amendments and concept completion with weight analysis:
2) Supersonic range and endurance and mission stage fuel burn for NB2:
3) Aerodynamic and performance methods of the AeroDYNAMIC toolset:
4) Comparison of as designed aircraft with the RFP targets.
E-1.1:- Structural design amendments and concept completion
The author was not satisfied with the initial proposal for a large MOSC ingress /
egress hatch of 2.0m wide and 1.2m in length, he felt that this large cut out would
weaken the front fuselage and would be difficult to seal effectively and ample
provision for MOSC emergency ejection could be provided by a smaller aperture
which would land on additional shoulder longerons framing the MOSC compartment.
This new hatch would have the advantage of requiring a smaller and lighter actuation
system, a lighter structure, and less risk of impacting the airframe on release which
could result in debris impacting the aircrew during emergency decent after ejection.
This new hatch is shown in figure E-1 below and was 0.88m wide by 1.6m in length,
with all round pressurization sealing. The hatch was hinged at the rear and rotated
around a single hinge bolt designed to fail under hatch emergency release loads i.e.
when the hatch was opened into the airflow over the forward fuselage over the canopy
the bolt would shear and the hatch would separate from the aircraft. The ejection
sequence would be MOSC first followed by the pilot. The hatch would be secured at
the front face by two explosive release automatic locking bolts, of the same type used
to attach the actuator to the hatch, and the front face would seal on the pilots canopy
frame. The integration of this hatch with the forward fuselage is shown in figure E-2
below were the hatch is shown open to 450 in normal operation this would open to 80
0
to enable ingress and egress of the MOSC in full NBC fight clothing, as well as
conventional Night Vision Goggles and the Advanced Virtual Environment Helmet,
proposed for the FB-24. The original proposal to chevron the forward face of the
hatch in the same manner as the F-117A canopy frame was dropped because this face
would be completely obscured from radar illumination by the pilots canopy obviating
the need for treatment beyond a RAM loaded „p – seal‟. This concluded the forward
fuselage design proposal for the FB-24 airframe. The A-24 requirements are covered
in figures E-3 and E-4 below.
The second major design issue reflected in the common fuselage structural design of
both the FB-24 and A-24 airframes was the load paths in the rear fuselage in which all
of the boom loads were carried by the single fuselage closure frame. This was clearly
unsatisfactory as the resulting frame would be massive and very heavy so to resolve
this, the author proposed a kicked inboard keel boom attaching to the shoulder
longeron and lower longeron of the aft fuselage, as well as a continuous outboard keel
as shown in figures E-5 and E-6 below.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
234
Figure E-1: - Proposed MOSC hatch as described above made from RTM
substructure and fibre placement skin all in Carbon PMR-15 high temperature
composite.
Figure E-2: - Proposed MOSC hatch integration with forward fuselage as
described above with frame stations (FS) numbered in metres.
0.88m
1.6m
Co-Bonded Stub frames Hatch sealing and
landing longerons
Hatch attachment longerons
Hatch sealing and
landing longerons
Frame back-up stiffeners and
hatch attachment fixtures.
Hatch open at 450
FS2.51
FS2.92
FS3.35
FS3.71
FS4.09
FS4.46
FS4.88 FS5.35
FS5.70 FS6.12 FS6.43
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
235
AI
Mo
du
le
Op
tical
tra
ckin
g
syste
m
180
0 f
ield
of
reg
ard
po
rt /
sta
rbo
ard
an
d -
16
0 t
o 1
20
0 in
ve
rtic
al p
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e.
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ctr
o
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tical
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ng
S
yste
m
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ac
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ap
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ire
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od
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h E
CS
inco
rpo
rate
d
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ve
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fre
qu
en
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nd
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en
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LO
rad
om
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n A
ES
A r
ad
ar
mo
un
t fr
am
e
LO
co
ate
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ield
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reg
ard
240
0
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rbo
ard
Can
op
y
su
pp
ort
lo
ng
ero
n
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orw
ard
ch
ine lo
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ero
n
Po
rt C
an
op
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ero
n
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rt In
tak
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‟ secti
on
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rt F
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ith
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l sti
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: -
F1
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el
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un
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fra
me
s:
mati
ng
fra
me:
inte
rface
co
nn
ecto
r su
pp
ort
fram
e.
Fig
ure
E-3
:- A
-24 F
orw
ard
fu
sela
ge
stru
ctu
ral
lay
ou
t an
d p
osi
tion
ing o
f A
I m
od
ule
.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
236
AI Track and scan module.
The down selected configuration for the modular AI optical tracking system was
based on the requirement to give the operator a pilot eye view with a full 1800 field of
regard and a 160 over nose view angle, imparting greater awareness on landing, take-
off, and in fight refuelling shown below for EFA in figure E-4, of the A-24 UCAV.
This OTS field of regard combined with the chin mounted Electro Optical Targeting
System (EOTS), all-digital infra-red system, offers a much grater passive situational
and target awareness than the Predator for example which has been compared to
flying through a drinking straw. The OTS is shrouded by a multi layer tin iridium
oxide coated opaque to radar optically perfect bowless canopy.
Figure E-4: - A Spanish Eurofighter Typhoon refuelling from a C-130 tanker
using the hose and drogue technique advocated for both FB-24 and A-24
aircraft.
Although autonomous air to air refuelling should be fairly simple with the boom
method in which the recipient aircraft is passive station keeping role, and the boom is
flown from the tanker into the receptacle, it limits the number of aircraft which can be
refuelled to one per tanker. The hose and drogue method is much more demanding
with the recipient aircraft in the active role of achieving hook on and station keeping,
but permits up to three aircraft to be refuelled at the same time (all be it with a
reduced fuel transfer rate) therefore a large field of regard is essential for this
operation to be undertaken by an off-board pilot. The AI interface with the aircraft
systems would be through a common (manned / unmanned) optical FBL data bus
connection housed in the lower forward fuselage, enabling plug and play
interchangeability of AI units.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
237
Fig
ure
E-5
:- F
B-2
4 /
A-2
4 C
om
mon
sim
pli
fied
aft
fu
sela
ge
stru
ctu
ral
layou
t to
p v
iew
.
Sta
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ard
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i fu
ll
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nti
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„T‟
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n
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ou
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ge
ron
s 7
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y 5
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m d
eep
Sta
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I‟ s
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ull d
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rt T
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i „I‟
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-2 E
ng
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AF
T
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
238
Fig
ure
E-6
:- F
B-2
4 /
A-2
4 C
om
mon
sim
pli
fied
aft
fu
sela
ge
stru
ctu
ral
layou
t u
nd
ersi
de
vie
w.
Aft
fu
se
bo
w
fram
es
a
s
a
rem
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le c
en
tre s
ecti
on
of
the
main
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am
e
for
en
gin
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acti
on
FR
-5 F
uel ta
nk b
ou
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om
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ato
r b
ays
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-2 E
ng
ine
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-5 F
uel ta
nk b
ou
nd
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s
Po
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m a
ctu
ato
r b
ay
s
NB
:- A
ft f
usela
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ed
esig
n p
rovid
es
co
nti
nu
ou
s
load
p
ath
s
inte
rfa
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wit
h t
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en
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us
ela
ge
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.
FS
14.9
2
FS
15.4
1
FS
14.3
9
FS
15.8
9
FS
16.3
7 F
S17.0
8 R
F
S17.9
4 R
FS
17.0
8 L
FS
17.9
4 L
A
FT
UP
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
239
Weap
on
s b
ay l
oad
path
keel
En
gin
e a
ttach
men
t fr
am
es A
l fo
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em
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full l
en
gth
co
nti
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C‟
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cti
on
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gero
ns 7
0m
m b
y 5
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So
ft b
uild
jo
int
NB
:-
Cen
tre
fus
ela
ge
o
nly
m
ajo
r
ch
an
ge
was
ad
dit
ion
o
f w
eap
on
s
bay
inb
oard
w
all
load
p
ath
keel
wh
ich
acts
as
a
fusela
ge
ben
din
g
load
dis
trib
uto
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oard
of
the m
ain
keel
be
am
s an
d m
ate
s w
ith
th
e aft
fusela
ge in
bo
ard
bo
om
lo
wer
ke
els
.
Fig
ure
E-7
:- D
etail
ed s
tru
ctu
ral
layou
t u
pd
ate
of
the
com
mon
cen
tre
fuse
lage
of
the
FB
-24 /
A-2
4. T
op
vie
w s
how
n.
Lan
d
based
a
rreste
r
ho
ok a
ttach
men
t fr
am
e
FS
8.3
8
FS
9.5
2 F
S10.6
2
FS
11.6
0
FS
7.1
2
FS
12.5
8
FS
13.1
5
FS
13.9
7
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
240
Fig
ure
E-8
:- D
etail
ed s
tru
ctu
ral
layou
t m
od
el o
f re
vis
ed F
B-2
4 c
on
figu
rati
on
.
FW
D f
usela
ge b
uild
mo
du
le
RT
M c
arb
on
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5
Weap
on
s b
ay
s T
i S
PF
/DB
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mic
/ S
tru
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ral
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lap
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fu
sela
ge b
uild
mo
du
le in
Ti
Ti / S
tru
ctu
ral R
AM
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derv
ato
rs
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on
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ing
sk
in t
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s
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re T
i
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tre f
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an
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sela
ge
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ess
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ot
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ow
n
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10
149 –
to
–F
S15
893.
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
241
Fig
ure
E-9
:- D
etail
ed s
tru
ctu
ral
layou
t m
od
el o
f re
vis
ed A
-24 c
on
figu
rati
on
wit
h 7
5%
com
mo
nali
ty w
ith
FB
-24.
FW
D f
usela
ge b
uild
mo
du
le
RT
M c
arb
on
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AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
242
Fig
ure
E-1
0:-
Det
ail
ed s
tru
ctu
ral
layou
t m
od
el o
f re
vis
ed c
om
mon
FB
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A-2
4 w
ing.
Rib
1A
Rib
1B
Rib
1C
Rib
1D
Rib
1E
Rib
2A
Rib
2B
Rib
2C
Rib
2D
Rib
3A
Rib
3B
Rib
R (
A)
Rib
R (
B)
Rib
R (
C)
Rib
R (
D)
Rib
R (
E)
Rib
R (
F)
Tip
Rib
Au
x R
ib 1
Au
x R
ib 2
(a)
/ (b
)
Au
x R
ib 3
Fw
d S
par
Sp
ar
1
Sp
ar
2
Sp
ar
3
Sp
ar
4
Sp
ar
4
Sp
ar
5
Sp
ar
5
Aft
Sp
ar
Stu
b S
pa
r
AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.
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243
Figure E-11: - Structural joint philosophy for rib to spar metallic joints.
Figure E-12: - Proposed structural joint philosophy for Ti spar to carbon PMR-
15 rib joints to be employed in the carbon composite wing torsion box
substructure trade study illustrated below in figure E-13.
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244
Figure E-13:- Proposed weight reduction trade study to using carbon PRM-15
for wing torsion box substructure using the layout defined in figure E-10.
Figure E-14:- The A-24 estimated total fuel capacity 25,135lbs equal to 16%
greater fuel capacity of the two manned FB-24, and estimated GTOW of
48,200lbs.
Ti wing boundary and carbon PMR-15 sub-structure with multi spar layout to resist buckling of
skins with long thin panels.
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E-1.2:- Structural design weight estimation
The structural weight measurement for each major build component of the final
design was determined by measurement of the CATIA V.5 models in the as drawn
configuration i.e. as geometric blanks sized to the generalised IML and then
subtracting the delta percentage difference between the blank weight and the as drawn
weight of a nominally sized structure as shown in figures E-15 and 16. The densities
for the structural materials used were taken from references 33 which were listed in
the reference section 6 on page 207 of this thesis.
The densities applied were as follows:-
2024-T351 Aluminium plate = 2768 kg/m3 (0.1 lb/in
3)
2104-T6 Aluminium forgings = 2768 kg/m3 (0.1 lb/in
3)
BMI / graphite = 1522.4 kg/m3 (0.055 lb/in
3)
6Al-4V Titanium plate = 4428.8 kg/m3 (0.16 lb/in
3)
6Al-6V Titanium forgings = 4539.52 kg/m3 (0.614 lb/in
3)
Al alloy 2024-T351 plate frame :-
8mm flanges: 4mm web: 4mm
stiffener thicknesses.
Ti alloy 6Al-6V-2Sn forging frame :-
8mm flanges: 3mm web: 4mm
stiffener thicknesses.
BMI carbon RTM frame :- 7mm
flanges: 4mm web: 25mm top hat
stiffener thicknesses.
Weight comparison.
As drawn Block Frame with weight
measured in Al: Ti: and BMI carbon.
As drawn Design intent Flyaway Part
Al Block frame = 183.3kg Pocketed frame = 20.6kg Actual weight % = 11
Ti Block frame = 300.5kg Pocketed frame = 27.4kg Actual weight % = 9.1
BMI Block frame = 100.8kg BMI RTM frame = 19.4kg Actual weight% = 19.2
Figure E-15:- Weight comparison between the representative „As drawn frames‟
and „Notional thickness detailed design intent frames‟ to determine the
percentage of the as drawn frame weight an actual „Flyaway frame‟ would equal.
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246
Al alloy 2024-T351 plate Keel :- 4mm
flange and 4mm web thicknessBMI Carbon RTM Keel :- 4mm
flange and 4mm web thickness
Ti alloy 6Al-4V plate keel :- 4mm
flange and 4mm web thickness
As drawn Block Frame with weight
measured in Al: Ti: and BMI carbon.
Weight comparison.
As drawn Design intent Flyaway Part
Al Block keel = 369.7kg Pocketed keel = 46.2kg Actual weight % = 12.5
Ti Block keel = 591.4kg Pocketed keel = 73.9kg Actual weight % = 12.5
BMI Block keel = 203.3kg BMI RTM keel = 25.4kg Actual weight% = 12.5
Figure E-16:- Weight comparison between the representative „As drawn keels‟
and „Notional thickness detailed design intent keels‟ to determine the percentage
of the as drawn keel weight an actual „Flyaway keel‟ would equal.
Table E.1:- Ti Wing & Empennage structural weight measurements.
Wing component Weight in kg
Port Wing boundary structure 281.12
Stbd Wing boundary structure 281.12
Port Wing core structure 305.28
Stbd Wing core structure 305.28
Port leading edge ribs 10.93
Stbd leading edge ribs 10.93
Total wing weight 1,194.64
Empennage components Weight in kg
Port / Stbd boundary structure 91.08
Port / Stbd core structure 69.02
Total empennage weight 160.10
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Table E.2:- Forward fuselage structural weight measurements.
Component Frame Station in m*
FB-24 Weight in kg A-24 Weight in kg
Chevron Frame 11.15 11.15
Frame 2.51 13.37 13.37
Frame 2.92 8.87 8.87
Frame 3.35 7.85 7.85
Frame 3.71 10.48 10.48
Frame 4.09 17.30 17.30
Frame 4.46 20.67 20.67
Frame 4.88 17.42 32.26
Frame 5.35 25.52 44.00
Frame 5.70 44.08 59.58
Frame 6.12 50.36 62.95
Frame 6.43 64.60 64.60
Keels (Port / Stbd) 41.80 41.80
Chine longerons (Port / Stbd)
4.14 4.14
Outboard Longerons (Port / Stbd)
1.78 1.78
Canopy landings (Port / Stbd)
2.36 2.36
Hatch landings (Port / Stbd)
1.16 N/A
Frame 6.43 stiffeners 3.32 N/A
Hatch 92.77
Total weight 439.00 403.16
* SEE FIGURE E-2 FOR FRAME STATION LOCATIONS.
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Table E.3:- Centre and Aft fuselage structural weight measurements.
Centre Fuselage Component*
Weight in kg
Aft Fuselage Component*
Weight in kg
Frame 7.12 106.45 Frame 14.39 79.65
Frame 8.38 76.98 Frame 14.92 74.90
Frame 9.52 80.62 Frame 15.41 67.16
Frame 10.62 79.42 Frame 15.89 58.08
Frame 11.60 89.87 Frame 16.37 46.36
Frame 12.58 88.29 Frame 17.08 (x2) 11.88
Frame 13.15 87.85 Frame 17.94 (x2) 5.48
Frame 13.97 113.42 Kicked Keel (Port /
Stbd) 43.20
Lower Keels (Port / Stbd)
98.18 Shoulder longerons
(Port / Stbd) 20.94
Shoulder longerons (Port / Stbd)
21.74 Lower Longerons
(Port / Stbd) 38.76
Mid aft Keels (Port / Stbd)
65.14 Chine longerons
(Port / Stbd) 11.04
Total weight 907.96 Keel beam
(Port / Stbd) 80.8
I/B Fairing chine (Port / Stbd)
3.96
Boom closure (Port / Stbd)
0.74
Total weight 636.01
*COMMON STRUCTURAL COMPONENTS FRAME STATIONS IN
METRES SEE FIGURES E-6 AND E-7 FOR FRAME STATION LOCATIONS.
The weights measured for the aircraft skins were taken directly from solid models
created by the author (as were all models contained herein). The wing skin lay up was
a simplified proportional representation without skin ramps and consisted of stepped
wing skins from 20mm at the root to 4mm at the tip as shown in figure E-17. The
fuselage skins were constant 5.08mm OML off-set solids without pad-up areas, as
shown in figure E-18, and the same method was used to generate the ruddervator
skins. The weights of these skins are given in table E-4 below.
The total measured weights of the major structural components are compared with the
estimated component weights in lbs and kgs in table E-5 below, feeding into the
definitive weight statement in table E-6 below.
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Figure E-17:- Weight of wing skins to a notional balanced ply lay-up
configuration.
Figure E-18:- Weight of fuselage skins to a notional balanced ply lay-up
configuration.
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Table E.4:- Major component skin weight measurements.
Skin component Weight in kg
Port Wing skin 944.00
Stbd Wing skin 944.00
Forward fuselage skin 185.12
Centre fuselage skin 305.52
Aft fuselage skin 123.41
Port ruddervator skin 72.50
Stbd ruddervator skin 72.50
Total component weight Weight in kg
Total Wing skin weight 1,888.00
Total Fuselage skin weight 614.05
Total Ruddervator skin weight 145.00
Total skin weight 2,647.05
Table E.5:- Major component weight measurements compared with
AeroDYNAMICTM
estimated values.
Major airframe component
Measured weight in kg / lb
Estimated weight in kg / lb
Wings 3,082.64 / 6,796.06 1,232.40 / 2,716.90
FB-24 Fuselage 2,597.02 / 5,725.45 2,837.60 / 6,255.70
A-24 Fuselage 2,561.18 / 5,646.44 2,837.60 / 6,255.70
Verticals 497.96 / 1,097.8 323.2 / 712
Undercarriage Main 583.64 / 1,286.70 590.40 / 1,301.70
Undercarriage Nose 243.40 / 536.60 In Main U/C value
Engine weight 3,402.00 / 7,500.00 3,402.00 / 7,500.00
GTOW FB-24 19,989.24 / 44,068.73 22,327.4 / 49,223.5
GTOW A-24 21,771.86 / 47,998.73 22,327.4 / 49,223.5
Landing weight with 10% fuel FB-24
10,406.66 / 22,942.76 11,208.00 / 24,709.50
Landing weight with 10% fuel A-24
10,484.83 / 23,115.09 11,208.00 / 24,709.50
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The measured structural weights to notional sizings are higher for the wings and
ruddervators, than the AeroDYNAMIC values and yet significantly lower for the
fuselage, this was attributed to the generalised thick skins used for the wings and
reduction in internal structure in the fuselage. This leads to the definitive weight
statement below.
Table E.6:- Definitive as designed weight measurements statement.
Major aircraft component Measured weight in kg / lb
Wings 3,082.64 / 6,796.06
FB-24 Fuselage 2,597.02 / 5,725.45
A-24 Fuselage 2,561.18 / 5,646.44
Verticals 497.96 / 1,097.8
Undercarriage Main 583.64 / 1,286.70
Undercarriage Nose 243.40 / 536.60
Propulsion weight 3,402.00 / 7,500.00
Equipment weight* 2,233.81 / 4,924.70
Design empty weight FB-24 12,640.47 / 27,867.46
Design empty weight A-24 12,604.63 / 27,788.45
Useful load weight* F-24 11,484.96 / 25,320.00
Useful load weight* A-24 13,203.62 / 29,109.00
GTOW FB-24 24,125.43 / 53,187.46
GTOW A-24 25,808.25 / 56,897.45
Landing weight with 10% fuel FB-24
13,598.73 / 29,980.06
Landing weight with 10% fuel A-24
13,744.74 / 30,301.95
*Employs releasable figures from F-35 Joint Project Office (no individual
weights to be identified ITAR restrictions).
N.B.:- Fuel capacity of the F-24 = 9,582.59kg = 21,126.00lb
N.B.:- Fuel capacity of the A-24 = 11,401.04kg = 25,135lb
Equipment: - Flight controls: Instruments: Hydraulics: Electrical: Avionics: GFE
(ejection seats): Air conditioning: Handling gear: Engine heat shield: etc.
Useful load: - Crew: Fuel: Oil: Stores (weapons and defensive aids Chaff / flares):
mission critical devices NVG‟s: etc.
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252
This concluded the structure and weight study for the F-24 and A-24 PWSC
designs.
E-2 The Subsonic and Supersonic range and endurance values for the FB-24 and
A-24 aircraft NB-2 configurations.
E-2.1:- The maximum endurance and range for A-24 at BCM at 34,400ft.
For the A-24 configuration the maximum endurance is achieved at the speed for
(L/D)max which can be determined by first calculating the required value of CL, then
solving for the speed required to achieve L=W at that CL therefore for A-24:-
CL = CDo / k = 0.014361 / 0.167051 =0.293
L = W = CLqS, q = W / CLS = 56,897.45 / 0.293 (893ft2) = 217.45lb/ft
2
And using at 34,400ft = 0.000767 slug/ft3 obtained from standard atmospheric
tables (reference 22) and the definition of q,
V = 2q / = 2 x (188.131lb/ft2) / 0.000767 slug/ft
3 = 700.4ft/s
for maximum endurance. Note that this is only the initial velocity for maximum
endurance and that as fuel is burned the velocity for best endurance will decrease. In
order to calculate the maximum endurance time, it is first necessary to determine the
magnitude of (L/D) max using equation 3.34: -
(L/D) max = (CL / CD) max = 1 / 2 k CDo = 1 / 2 0.167051(0.014361) = 10.2
The TSFC is also predicted using equation 3.35 with a = 977.5ft/s at 34,400ft and
aSL = 1116.4ft/s obtained from the standard atmospheric table (reference 22):-
ct = cSL (a / aSL) = (0.89[(lb/hr)/lbf])(977.5ft/s / 1116.4ft/s) = 0.78[(lb/hr)/lbf]
Then the endurance can be calculated using equation 3.36 with W1 = 56,897.45 and
that for W2
W2 = W1 – Wf = 56,897.45lb – 25,135lb = 31,762.45lb
E = 1/ ct CL / CD ln (W1 / W2) = 1 / 0.78 (10.2) ln (56,897.45 / 31,762.45) = 7.62h
Therefore for A-24 the maximum endurance at 34,400ft at Best Cruise Mach
BCM is 7.62hours.
Similarly the value for maximum range is obtained by solving equation 3.36 for CL
and equation 3.37 for q:-
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CDo = 3k CL2, CL = CDo / 3k = 0.014361 / 3(0.167051) = 0.239
q = W / CLS = 56,897.45 / 0.239(893ft2) = 217.45lb/ft
2
V = 2q / = 2 x (217.45lb/ft2) / 0.000767 slug/ft
3 = 753.00ft/s
for maximum range. As with the velocity for maximum endurance, the velocity for
best range will decrease as fuel is burned. The vale calculated for CL is now used to
calculate CD after which the maximum range is predicted using equation 3.38.
CD = 0.014361 + 0.167051 CL2 = 0.014361 + 0.167051(0.239)
2 = 0.0239
R = 2/S / ct CL2 / CD (W1
1/2 – W2
1/2)
R= 2 / (0.000767 slug/ft3) (893ft
2) 2 / 0.78 (0.239)
1/2 / 0.0239
x ((56,897.45lb)1/2
-(31,762.45lb)1/2
)
R = 15.315 (ft2 / lbs
2) (2.564h) (20.45) (12.568lb
1/2) = 2,263.97miles
Therefore for the range in nautical miles divide through by 1.69(ft/s/kn)
Therefore the A-24 would have a range value of R = 1,457.97nmiles.
E-2.2:- Supersonic wave drag and critical Mach number analysis.
This performance appeared to be suspiciously low and can be attributed to the high
drag values based on the crude modelling in AeroDYNAMIC V2.08. For supersonic
range and performance values were based on the CATIA V5 model geometry to
predict MCrit C D wave and k1 using the methods in reference 22 as follows:-
MCrit (unswept) = 1.0 – 0.065 [100 (tmax / c)] 0.6
= 1.0 – 0.065 (5)0.6
= 0.82
MCrit = 1.0 – cos 0.6
0.25 (1.0 - MCrit (unswept))
= 1.0 – cos 0.6
43.90 (1 – 0.82) = 0.87
MCDo max = 1 / cos 0.2
LE = 1 / cos 0.2
520 = 1.0
C D wave = 4.5 / S (Amax / l) 2 EWD (0.74 + 0.37 cos LE) [1 - 0.3 M - MCDo max]
C D wave = 0.0158 x (24.46 / 60) 2 x 1.355 x 0.379 = 0.0134 at Mach 1.6
k1 = [AR (M2 – 1) / (4AR M
2 – 1) – 2] cos LE
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254
k1 = [2.27 (1.62 – 1) / (4 x 2.27 1.6 – 1) – 2] cos 52
0 = 0.208
The parasitic drag was estimated using the equivalent skin friction coefficient of 0
0025 (representative of a smooth fast jet transport aircraft) from reference 1 and the
CATIA V5 surface wetted area of 161m2 = 1,732ft
2 therefore the parasitic drag value
was:-
C D min = 0.0025 (1,732 / 893) = 0.00485
This gave a total clean aircraft zero lift coefficient of 0.01825 in supersonic cruise
condition
E-2.3:- The maximum endurance and range for A-24 at Mach 1.6 and 45000ft.
Maximum supersonic endurance at Mach 1.6 (L/D)max was determined by first
calculating the value of CL, then solving for the speed required to achieve L=W at that
CL therefore for A-24:-
CL = CDo / k = 0.01825 / 0.208 = 0.6495
L = W = CLqS, q = W / CLS = 56,897.45 / 0.6495 (893ft2) = 98.10b/ft
2
And using at 45,000ft = 0.000462 slug/ft3 obtained from standard atmospheric
tables (reference 22) and the definition of q,
V = 2q / = 2 x (98.10lb/ft2) / 0.000462 slug/ft
3 = 651.67ft/s
For maximum endurance: Note that this is only the initial velocity for maximum
endurance and that as fuel is burned the velocity for best endurance will decrease. In
order to calculate the maximum endurance time, it is first necessary to determine the
magnitude of (L/D) max using equation 3.34: -
(L/D) max = (CL / CD) max = 1 / 2 k CDo = 1 / 2 0.208(0.01825) = 8.3
The TSFC is also predicted using equation 3.35 with a = 968.1ft/s at 45,000ft and
aSL = 1116.4ft/s obtained from the standard atmospheric table (reference 22):-
ct = cSL (a / aSL) = (1.68[(lb/hr)/lbf])(968.1ft/s / 1116.4ft/s) = 1.45[(lb/hr)/lbf]
Then the endurance can be calculated using equation 3.36 with W1 = 48,200 and that
for W2
W2 = W1 – Wf = 56,897.45lb – 25,135lb = 31,762.45lb
E = 1/ ct CL / CD ln (W1 / W2) = 1 / 1.45 (8.3) ln (56,897.45 / 31,762.45) = 3.34h
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Therefore for A-24 the maximum endurance at 45,000ft at Mach 1.6 is
3.34hours.
Similarly the maximum range value for the FB-24 is obtained by solving equation
3.36 for CL and equation 3.37 for q:-
CDo = 3k CL2, CL = CDo / 3k = 0.01825 / 3(0.208) = 0.1710
q = W / CLS = 56,897.45 / 0.1710(893ft2) = 372.601b/ft
2
V = 2q / = 2 x (315.645lb/ft2) / 0.000462 slug/ft
3 = 1,168.94ft/s
for maximum range. As with the velocity for maximum endurance, the velocity for
best range will decrease as fuel is burned. The vale calculated for CL is now used to
calculate CD after which the maximum range is predicted using equation 3.38.
CD = 0.01825 + 0.208 CL2 = 0.01825 + 0.208(0.1710)
2 = 0.0243
R = 2/S / ct CL2 / CD (W1
1/2 – W2
1/2)
R= 2 / (0.000462 slug/ft3) (893ft
2) 2 / 1.45 (0.1710)
1/2 / 0.0243
x ((56,897.45lb)1/2
-(31,762.45lb)1/2
)
R = 4.84 (ft2 / lbs
2) (1.378h) (17.02) (12.568lb
1/2) = 648.48miles
Therefore for the range in nautical miles divide through by 1.69(ft/s/kn)
Therefore the value for R = 383.72nmiles.
E-2.4:- The maximum endurance and range for FB-24 at Mach 1.6 and 45000ft.
Maximum supersonic endurance at Mach 1.6 (L/D)max was determined by first
calculating the value of CL, then solving for the speed required to achieve L=W at that
CL therefore for FB-24:-
CL = CDo / k = 0.046832 / 0.349518 = 0.6495
L = W = CLqS, q = W / CLS = 53,187.46 / 0.6495 (893ft2) = 91.70lb/ft
2
And using at 45,000ft = 0.000462 slug/ft3 obtained from standard atmospheric
tables (reference 22) and the definition of q,
V = 2q / = 2 x (91.70lb/ft2) / 0.000462 slug/ft
3 = 630.05ft/s
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For maximum endurance: Note that this is only the initial velocity for maximum
endurance and that as fuel is burned the velocity for best endurance will decrease. In
order to calculate the maximum endurance time, it is first necessary to determine the
magnitude of (L/D) max using equation 3.34: -
(L/D) max = (CL / CD) max = 1 / 2 k CDo = 1 / 2 0.208(0.01825) = 8.3
The TSFC is also predicted using equation 3.35 with a = 968.1ft/s at 45,000ft and
aSL = 1116.4ft/s obtained from the standard atmospheric table (reference 22):-
ct = cSL (a / aSL) = (1.68[(lb/hr)/lbf])(968.1ft/s / 1116.4ft/s) = 1.45[(lb/hr)/lbf]
Then the endurance can be calculated using equation 3.36 with W1 = 49,223.5 and
that for W2
W2 = W1 – Wf = 53,187.46lb – 21,126lb = 32,061.46lb
E = 1/ ct CL / CD ln (W1 / W2) = 1 / 1.45 (8.3) ln (53,187.46 / 32,061.46) = 2.90h
Therefore for FB-24 the maximum endurance at 45,000ft at Mach 1.6 is
2.90hours.
Similarly the maximum range value for the FB-24 is obtained by solving equation
3.36 for CL and equation 3.37 for q:-
CDo = 3k CL2, CL = CDo / 3k = 0.01825 / 3(0.208) = 0.1710
q = W / CLS = 53,187.46 / 0.1710(893ft2) = 348.31lb/ft
2
V = 2q / = 2 x (348.31lb/ft2) / 0.000462 slug/ft
3 = 1,227.93ft/s
for maximum range. As with the velocity for maximum endurance, the velocity for
best range will decrease as fuel is burned. The vale calculated for CL is now used to
calculate CD after which the maximum range is predicted using equation 3.38.
CD = 0.01825 + 0.208 CL2 = 0.01825 + 0.208(0.1710)
2 = 0.0243
R = 2/S / ct CL2 / CD (W1
1/2 – W2
1/2)
R= 2 / (0.000462 slug/ft3) (893ft
2) 2 / 1.45 (0.1710)
1/2 / 0.0243
x ((53,187.46lb)1/2
-(32,061.46lb)1/2
)
R = 4.84 (ft2 / lbs
2) (1.378h) (17.02) (10.563lb
1/2) = 541.86miles
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G. A. Wardle MSc CEng. MSc Individual Research Project Thesis
257
Therefore for the range in nautical miles divide through by 1.69(ft/s/kn)
Therefore the value for R = 320.63nmiles.
*Therefore the maximum range and endurance of the FB-24 two place manned
aircraft was:-
At 34,400ft and Mach 0.85: -
Endurance = 7.03 hours
Range = 1,282.54nmiles
At 45,000ft and Mach 1.6: -
With A/B Endurance = 2.90hours and without A/B Endurance = 5.80hours.
With A/B Range = 320.63nmiles and without A/B Endurance = 640nmiles.
* Therefore the maximum range and endurance of the A-24 UCAV unmanned
aircraft was:-
At 34,400ft and Mach 0.85: -
Endurance = 7.62hours
Range = 1,457.97nmiles
At 45,000ft and Mach 1.6: -
With A/B Endurance = 3.34hours and without A/B Endurance = 6.68hours.
With A/B Range = 383.72nmiles and without A/B Endurance = 767.44nmiles.
N.B. Mach 1.6 values for both aircraft are given with and without afterburner
engaged the endurance and range of the YF-120 in dry supersonic flight
condition is twice that with afterburner with 80% of the thrust based on USAF
released data reference 26.
E-3:- A description of the AeroDYNAMICTM
V2.08 toolset analysis codes.
As explained at length above the Jet306 tool could not be used for this conceptual
design study due to numerous run – time errors, and a replacement was not available
in the required time frame therefore the older less sophisticated AeroDYNAMICTM
V2.08 tool was used and it is this which is described below.
A complete aircraft will frequently generate significantly more lift than its wing
alone. AeroDYNAMICTM
estimated the whole aircraft‟s lift by summing the lift
contributions of its various components. The methodology employed and described
below was a simple means of making an initial estimate of an aircraft‟s aerodynamic
capabilities and was suitable for use in the early conceptual phases of design, and was
deemed by the author as applicable to this design proposal submission.
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E-3.1:- Lift analysis.
For most aircraft including the FB-24 / A-24 the majority of the lift is generated by
the wing with the pitch and trim surfaces contributing additional lift.
AeroDYNAMICTM
used the non-elliptical finite wing expression to determine the lift
curve slope per degree of angle of attack CL and CLt for the wing and the pitch and
trim surfaces as shown as equations E.1 and E.3 below. The Span efficiency factor
was determined for the wing and the pitch and trim surfaces using empirical
expression for e shown in equations E.2 and E.4 below.
Wing
0.052019 per degree
where:
0.659126
Pitch Trim Surface
0.052019 per degree
where:
0.659126
Cc
c
e AR
L
l
l
157 3.
eAR AR t
2
2 4 12 2( tan )max
t
l
l
tL
ARe
c
cC
3.571
eAR AR t
2
2 4 12 2( tan )max
Equation block 3.1(a): - AeroDYNAMICTM
Aerodynamic lift analysis codes.
Where:-
AR = aspect ratio of the wing or pitch and trim surface:
t max = the sweep angle of the line connecting the point of maximum thickness
on each airfoil section of each wing or pitch and trim surface:
cl = lift coefficients at the angle of attack
5.73 = section lift – curve slope, per radian.
The data generated from this analysis was used by AeroDYNAMICTM
in the
generation of the following charts: - 1: 2(a) and,2(b): 3: 4: 8: 9(a) and (b): 10: and 11
shown in pages 136 to 142 inclusive.
(E.1)
(E.2)
(E.3)
(E.4)
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An aircraft fuselage is usually long and slender and therefore dose not produce much
lift. In the region of the horizontal lifting surfaces, however, the lift being generated
by those surfaces carries over the fuselage. In order to model this effect the wing is
treated as if it extends all the way through the fuselage without any changes in airfoil,
sweep, or taper. In fact, the fuselage shape is significantly different from the wings
airfoil shape and may be less effective at generating lift. However, because the
fuselage lifting area is generally larger than the portion of the wing in the fuselage, the
two effects may be treated as cancelling each other out, at least for early conceptual
design. For a design with strakes or LEXs, the effect is modelled in
AeroDYNAMICTM
as shown in equations E.5 and E.6 below (equation block 3.1(b),
based on wind tunnel testing, where S strake. includes only the exposed area of the
strake, and not any portion inside the fuselage. Because an angle of attack = 150 is
usually the maximum useable equations E.5 and E.6 are adequate for the useable
range (reference 22).
Equation block 3.1(b): - AeroDYNAMIC
TM Aerodynamic lift analysis codes
continued.
To determine the horizontal tail or canard surfaces contribution to the whole aircraft
lift curve slope it is first necessary to determine the rate at which downwash (or
upwash, as appropriate) changes with changing aircraft angle of attack.
AeroDYNAMICTM
estimated the rate of change of downwash with angle of attack
using the empirical curve fit given above as equation E.7 in equation block 3.1(b),
(based on wind – tunnel testing rather than theory). Where cavg is the mean geometric
chord of the wing: lh is the distance from the quarter-chord point of the average chord
of the main wing to the quarter – average – chord point on the horizontal surface, as
shown below in figure E-19: Zh is the vertical distance of the horizontal surface above
the plane of the main wing, as shown in figure E-19: and CL is in degrees.
Whole Aircraft
0.065639 +
where:
0.055487
and:
0.457872
C CL L ( ) ( )whole aircraft with strake CLt
1
S
S
t
C CS S
SL L
strake
( ) ( )with strake without strake
21 10 3
71
0 725
o C
AR
c
l
z
b
L avg
h
h .
zh
lh
.25 croot
.25 croot
lc
.25 croot
(E.5)
(E.6)
(E.7)
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Once the value of / has been predicted, the horizontal surfaces contribution to
the aircrafts CL can be approximated as: -
CL (due to horizontal tail) = CLt (1 - / St / S E.8
were the subscript t denotes parameters for the horizontal tail.
The common CL (due to horizontal tail) values vary from almost zero to 35% or
more of CL.
Figure E-19:- Aircraft geometry for downwash predictions.
For horizontal surfaces ahead of the wing, also known as canards AeroDYNAMICTM
another empirical equation based on wind – tunnel experimentation for predicting the
rate of change of upwash with angle of attack, for wings with a quarter chord sweep
angle 0.25 < 350 expression used was: -
u / = (0.3AR0.3
– 0.33) (lc / c)-(1.04+6AR^-1.7)
E.9
Where u was the upwash angle and lc was the distance from the wings quarter –
chord to the canards quarter – chord as shown in figure E-19 above.
Once u / had been estimated, the canards contribution to the aircrafts CL was
approximated as:-
CL (due to canard) = CLc (1 - u / Sc / S E.10
Where the subscript c identifies quantities related to the canard. Contributions of
canards to the total aircraft lift curve slope are typically larger than those for the
horizontal tails. This is partly due to the canard being in an upwash field rather than
the downwash field surrounding most horizontal tails. Once the contributions of
canards or horizontal tails have been estimated, the whole aircraft lift curve slope is
generated from:-
CL (whole aircraft) = CL (wing + body + strakes) + CL (due to horizontals) +
CL (due to canard) E.11
zh
lh
.25 croot
.25 croot
lc
.25 croot
0.25croot
0.25croot
0.25croot
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E-3.2:- Drag and Critical Mach number analysis.
The drag for the complete aircraft is identified as either parasite drag or drag due to
lift. Parasite drag is all of the drag on the aircraft when it is not generating lift. This
includes both skin – friction and pressure drag, as well as several additional types of
zero-lift drag that are associated with the complete aircraft configuration. The drag
due to lift includes all types of drag that depend on the amount of lift that the aircraft
is producing which includes induced drag due to downwash, the pressure drag, which
increases with lift due to forward movement of the separation point, induced and
pressure drag from canards and horizontal tails, and additional drag, e.g. vortex drag
due to leading-edge vortices on strakes and highly swept wings. All of the above
types of drag can be approximated by the simple expression for the drag coefficient
given below as equation E.5 which was used by AeroDYNAMICTM
for Drag Polar
determination, shown in figures 1 and 8 on pages 136 and 140 respectively.
Drag Polar:
Mach CDmin CDo k1 k2
0.1 0.01792 0.01800 0.1168 -0.0061
0.872737 0.01792 0.01800 0.1168 -0.0061
1.054749 0.05253 0.05261 0.2419 -0.0046
1.6 0.04486 0.04494 0.3045 0.0000
2 0.04244 0.04251 0.3670 0.0000
0.872737 M crit = 1.0 - 0.065 cos0.6
LE
add B-1 data to WIM-13
C C k C k CD Do
L L 1
2
2
100
0 6t
c
max
.
Equation block 3.2(a): - AeroDYNAMIC
TM drag and critical Mach number
analysis codes.
Equation E.12 where:-
k1 = 1 / ( e0 AR) (subsonic): - e0 = Oswald‟s efficiency factor: AR = Aspect
Ratio = b2 / S were b
2 is the wing span squared and S is the wing area:
k1 = AR (M2 – 1) / (4ARM
2 – 1) – 2 (supersonic): - AR = Aspect Ratio: M
= Mach number:
CDo = the parasite drag coefficient:
k2 = - 2 k1 CL min D (negative value subsonic and zero value supersonic):
CL min D = the CL value for which CD is a minimum:
k1CL2 = induced drag.
(E.12)
(E.13)
0.6
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The value of k2 is chosen to allow modelling of wings with airfoils that generate
minimum drag at some nonzero value of lift.
Parasite drag: - The first step in determining Parasite drag in AeroDYNAMICTM
was
to determine the wetted area by approximate the aircraft as a set of simple geometric
shapes, as described in section 3.3.2 page 134. The equations for the surface areas of
these simple shapes were well known, and by taking care to avoid counting areas
where two shapes touch, it was relatively easy to determine Swet the shapes available
to define the fuselage were: - cylinders: half - cylinders: circular: elliptical:
rectangular: cone: and half - cone. The surface areas of these shapes did not include
ends which butted up against another shape because such ends would not be wetted.
When the longitudinal flat face of a half cylinder or half - cone touches another body,
twice the surface area of that face was subtracted, because it and an equal area of the
other body were in contact with each other, and were therefore not wetted. The
margin of error within AeroDYNAMICTM
for relatively simple geometry aircraft e.g.
F-15, F-16 was 5% depending on the number of data points and hence the fidelity of
the model.
The skin - friction drag on a complete aircraft configuration is generally much greater
than on the wing alone because the wetted area is much greater. To make the initial
estimate of subsonic parasite drag AeroDYNAMICTM
used the concept of equivalent
skin – friction drag coefficient, Cfe employing data values from large numbers of
similar aircraft types (see table 4.1 in equation block 3.2(b) bellow), which is as
follows: -
Cfe = CDo S / Swet E.14
These values are based on historical data and are a function of such diverse factors as:
- aircraft skin materials, shape, paint, typical flight Reynolds number, number of
additional air scoops for ventilation, type size, number, and location of engine intakes,
and attention to detail in sealing doors, control surface gaps, etc. Using Cfe to predict
CDo for an aircraft that generates minimum drag when it is generating zero lift only
requires selecting a Cfe for the appropriate category of aircraft and estimating the total
wetted area of the aircraft concept as described above. The value of CDo could then be
obtained by solving
CDo = Cfe S / Swet E.15
As shown in equation block 3.2(b).
Drag due to lift:- AeroDYNAMICTM
predicted Oswald‟s efficiency factor with a
curve fit of wind tunnel data for a variety of wing and wing – body combinations and
the equation for this curve is shown as equation E.16 in equation block 3.2(b). Note
that increasing wing sweep decreases the value of eo. This is because for high aspect
ratio wings, that part of the airfoil profile drag that varies with lift is a larger part of
the total drag due to lift that eo must model.
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subsonic, supersonic
0.017921 0.0037 from Table 4.1
1.054749
0.116774 k 1 = 1/( e o AR) subsonic
where:
0.908619
supersonic
Table 4.1 Common C fe Values
-0.006074 Type C fe
subsonic Jet Bomber and Civil Transport 0.003
Military Jet Transport 0.0035
0.026009 Air Force Jet Fighter 0.0035
Carrier-Based Navy Jet Fighter 0.004
supersonic Supersonic Cruise Aircraft 0.0025
Light Single Propeller Aircraft 0.0055
Light Twin Propeller Aircraft 0.0045
clean 1.0502 Assume max = 15 degrees Propeller Seaplane 0.0065
Jet Seaplane 0.004
takeoff 1.2538 DCLmax =2/3 of DCLmax for Landing, CLmax = CLmax clean + DCLmax
landing 1.3587 DCLmax = 0.5 * S h /S * l h /(0.5*c root ) = max trimmable DCL
e ARo LE 4 61 1 0 045 310 68 0 15. ( . )(cos ) .. .
k k CLminD2 12
CS
A
lE M MD
maxWD LE C maxwave Do
4 50 74 0 37 1 3
2.
. . cos .
MC
LEDo max .cos
1
0 2
k
AR M
AR MLE1
2
2
1
4 1 2
cos
2
0L
LminD
L CC
S
SCC wet
fD emin
C C k CD D Lo min minD 1
2
efC
100
0 6t
c
max
.
wavesubsonico DDD CCC 0
02 k
C C CLmax L amax L max L ( )0
Equation block 3.2(b): - AeroDYNAMIC
TM drag and critical Mach number
analysis codes.
Effects of Camber:- AeroDYNAMICTM
determined profile drag resulting from:-
cambered airfoils: fuselage shape and orientation, which can result in minimum drag
being generated at a positive (non zero) value of lift coefficient, by using the
additional k2 CL term in equation E.12 in equation block 3.2(a) above. For example,
the minimum drag coefficient for an aircraft occurs at a lift coefficient signified by the
symbol CL min D then the necessary value of k2 is given by: -
k2 = - 2 k1 CL min D E.18
The value of CL min D is determined from an approximation which assumes that the
airfoil generates minimum drag when it is at zero angle of attack and that the effect of
induced drag is to move CL min D to a value half way between zero and the value of
CL when = 0. The value of CL when = 0 being given by: -
CL min D = CLa () = CLa (- L = 0) E.19
Because = - L = 0 and = 0 equation E.19 becomes
(E.16)
(E.17)
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CL min D = CLa (- L = 0 / 2) E.20
The equation is shown above in AeroDYNAMICTM
equation block 3.2(b). This value
of CL min D was used for the entire aircraft. This is done because the assumption was
made that aircraft fuselage and strakes etc were designed so that they also had their
minimum drag at the angle of attack that places the wing at its CL min D. When this is
done, the minimum value of C D which is given the symbol C D min must not be any
lower than C Do predicted by equation E.15 above. Recalling that C Do is the aircrafts
zero – lift drag coefficient. For aircraft with minimum drag at non zero lift, this
resulted in the following revised predictions: -
CD min = Cfe S / Swet E.21
C Do = CD min + k1 C 2L min D E.22
Critical Mach number:- AeroDYNAMICTM
uses equation E.13 in equation block
3.2(a) to determine the critical mach number of the configuration under analysis, this
substitutes swept wing chord equation (E.23) into the curve fit of Mcrit data for the
NACA 64-seies airfoils (used when actual airfoil data is not available in conceptual
design analysis).
Therefore: -
( t max / c ) (swept wing) = ( cos LE ) ( t max / c ) (unswept wing) E.23
Becomes: -
Mcrit = 1.0 – 0.065 cos0.6
LE [100 ( t max / c )]0.6
E.24
There by producing an expression the critical Mach number for 3-D swept wings, for
the unswept wings AeroDYNAMICTM
uses:-
Mcrit = 1.0 – cos0.6
LE (1.0 – Mcrit (unswept)) E.25
For tapered wings as in the planforms types evaluated in this conceptual design study,
AeroDYNAMICTM
uses 0.25c, the sweep angle of the line connecting the quarter-
chord points of the wings airfoils and using the maximum value of (t max /c) on the
wing i.e.:-
Mcrit = 1.0 – cos0.6
0.25c (1.0 – Mcrit (unswept)) E.26
*See also Reference 22: -page 123: Introduction to Aeronautics A Design
Perspective: by Brandt S. A: Stiles R. J: Bertin J. J: Whitford R.: Published by AIAA:
1997
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Supersonic Drag Due to Lift:- AeroDYNAMICTM
uses equation E.17 in equation
block 3.2(b) to calculate the supersonic value of k1. At supersonic speeds all airfoils,
regardless of shape generate zero lift at zero angle of attack. Practical supersonic
airfoil shapes also generate minimum drag at zero angle of attack, and so in the
supersonic regime, k2 = 0. The supersonic value of k1.is given by:-
k1.= [AR (M2 – 1) / (4AR M
2 – 1) – 2] cos LE E.27
Where:-
AR = Aspect Ratio:
M2 = Mach number:
LE = Sweep angle of the leading edge.
For a well designed supersonic aircraft, the transition from subsonic to supersonic
values of k1 and k2 is gradual, so that the variation of these parameters through the
transonic regime can be approximated with a smooth curve, as shown in charts 1 and
8 in the main body of this report on pages 136 and 140 respectively for the NB1 and
NB2 configurations of the F-24 / A-24 configurations studied in this report.
Total Drag:- In summary, the total drag on an aircraft is the sum of the profile drag
(the subsonic drag not due to lift), wave drag, and the drag due to lift or induced
drag:-
CDo = CDp + CD wave E.28
And:-
CD = CDo + k1 CL2 + k2 CL E.29
And AeroDYNAMICTM
uses these equations as shown in equation blocks 3.2(a) and
(b) above to determine total drag for the configuration under study, for the NB1 and
NB2 configurations investigated in this report the results are shown graphically as
charts 1(a) / (b) and 8(a) / (b) respectively on pages 136 and 140 of this report.
This concludes the overview of the aerodynamic codes and predictions used by
AeroDYNAMICTM
analysis tool, as can be seen this is a basic conceptual design tool
employing Microsoft Excel spreadsheets to process classical aerodynamic analysis
equations and although not very sophisticated it is the equal of a similar older MS.Dos
tool developed by Daniel Raymer and is deemed adequate for use by the United States
Airforce Academy as an academic design analysis tool. Unlike some conceptual
design approaches the user dose have to generate a concept rather than pages of
algebra, which is one reason the design approach taken in this report may appear odd
to some academics.
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E-4:- The application of AeroDYNAMICTM
V2.08 toolset to the NB2 / F-24 and
A-24 configuration.
The analysis of the F-24 / A-24 final airframe configuration conducted using
AeroDYNAMICTM
V 2.08 resulting in the charts 8 through 14 pages 140 – 143 used
applied the aerodynamic analysis methods and codes described above to the aircraft
configuration data inputs as described in the main report and below to generate the
following intermediate analysis stages.
Figure E-20:-Lifing surface analysis data for the common F-24 / A-24
configuration.
This data was used to calculate the aspect ratio AR of the wing and the ruddervators
and from this the value of e was determined for both surfaces and the two
dimensional lift curve slope for the NACA 64-0006 was approximated allowing CL
to be determined from equation E.1 for each surface as presented in the previous
section above. This configuration had no strakes so the CL(with strakes) term made no
contribution to the whole aircraft lift coefficient. The distance from the quarter chord
of the main wings mean chord to the same point on the aircrafts ruddervator l h the
wings taper ratio and the distance of the ruddervators above the wing Zh was used
to determine from equation E.7 described in the previous section above.
Finally using the wing and ruddervator areas S and S r the whole aircraft coefficient
CL(whole aircraft) was determined from equation E.5 as described in the previous
section above.
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Form these calculations the lift curve and aerodynamic analysis data shown below in
tables E-7 and E-8 were produced.
Table E-7:- Lift curve analysis data for the F-24 / A-24.
alpha CL CLalpha CLmax CLmaxTO CLmaxL
1 -2 -0.090175 0.0451 0.68 1.04 1.23
2 -0 0 [alpha at L = 0] Sref Swet Amax Length
3 3.5 0.157806 906.98 2835.41 72.56 60.00
4 7 0.315611
5 10.5 0.473417
6 14 0.631223
7 16 0.6763 [alpha at Clmax]
8 18 0.631223
Table E-8:- Aerodynamic analysis data for the F-24 / A-24.
Horiz Surf AR Taper Sweep S Section clAlpha e k
1 2.2327 0.07094 52 906.975 NACA0006 2.535512 0.853437 0.167051 L= 60
2 2.5002 0.23283 52 332.20814 NACA0006 2.567374 0.826795 0.153983 sWet= 2835.41321
3 f= 11.3416528
4 cfe= 0.004
5 cdo= 0.01250492
6 clat= 2.5833153
7 clMax= 0.67631036
8 clMaxTO= 1.04220329
9 clMaxL= 1.22514975
10 sRef= 906.975
11 clopt= 0
12 amax= 75.9676318
13 amax - A0= 72.557718
14 vol= 2653.8041
15 clAlpha= 2.583315
16 fuelVolume = 0
17 systemVolume = 0
18 expPayloadVolume = 0
19 seatVolume = 9
20 gearVolume = 86.5956engineVolume = 200.4529
The drag of the final F-24 / A-24 common configuration was calculated as follows
from the input data starting with the parasite drag. The first step in determining
Parasite drag in AeroDYNAMICTM
was to determine the wetted area by approximate
the aircraft as a set of simple geometric shapes, as described in section 3.3.2 page 134,
and shown below in figure E-21. The equations for the surface areas of these simple
shapes were well known, and by taking care to avoid counting areas where two shapes
touch, it was relatively easy to determine Swet the shapes available to define the
fuselage were: - cylinders: half - cylinders: circular: elliptical: rectangular: cone: and
half - cone. The surface areas of these shapes did not include ends which butted up
against another shape because such ends would not be wetted. When the longitudinal
flat face of a half cylinder or half - cone touches another body, twice the surface area
of that face was subtracted, because it and an equal area of the other body were in
contact with each other, and were therefore not wetted.
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The margin of error within AeroDYNAMICTM
for relatively simple geometry aircraft
e.g. F-15, F-16 was 5% depending on the number of data points and hence the fidelity
of the model.
CONE
HALF CONE
HALF CYLINDER
HALF CONE
SURFACE #1
SURFACE #2
SURFACE #4
SURFACE #3
SURFACE #6
SURFACE #5
SURFACE #8
SURFACE #9
OVAL CYLINDER #1
OVAL CYLINDER #2
OVAL CYLINDER #3
OVAL CYLINDER #4
CYLINDER #1
Figure E-21:- F-24 / A-24 configuration geometry approximation by simple
shapes.
The skin - friction drag on a complete aircraft configuration is generally much greater
than on the wing alone because the wetted area is much greater. To make the initial
estimate of subsonic parasite drag AeroDYNAMICTM
used the concept of equivalent
skin – friction drag coefficient, Cfe employing data values from large numbers of
similar aircraft types (see table 4.1 in equation block 3.2(b) above), and the fighter
aircraft value of 0.0035 was used in equations E.15 above to determine CDo.
Total drag was calculated using equation E.12 through calculation of the Oswald‟s
efficiency eo and thereby calculating the values of k1 for both the subsonic and
supersonic cases as described in the pervious section for the range of Mach numbers
considered. From this the drag polar values were calculated as given below in table E-
9 below.
From this data the CD vs CL: Lift over Drag vs CL: and CL ^ 1.5 over CL: values
were generated as shown below in tables E-10 through E-12.
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Table E-9:- Drag Polar analysis data for the F-24 / A-24.
Mach Cdmin Cdo k1 k2 Mach = 1.5
0.10 0.012505 0.012505 0.16705 -0 CL CD
0.85 0.012505 0.012505 0.16705 -0 1 -0.090175 0.049675
1.10 0.053862 0.053862 0.22418 -0 2 0 0.046832
1.50 0.046832 0.046832 0.34952 0 3 0.157806 0.055536
2.00 0.042685 0.042685 0.49731 0 4 0.315611 0.081648
5 0.473417 0.125168
6 0.631223 0.186095
7 0.67631 0.2067
8 0.631223 0.186095
Table E-10:- CD vs CL analysis data for the F-24 / A-24.
Mach = 1.5
CL CD
1 -0.090175 0.049675
2 0 0.046832
3 0.157806 0.055536
4 0.315611 0.081648
5 0.473417 0.125168
6 0.631223 0.186095
7 0.67631 0.2067
8 0.631223 0.186095
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Table E-11:- L over D vs CL analysis data for the F-24 / A-24.
M=Mcrit M=1.5 L/D vs. CL
CL L/D L/D Plot Mach No. (L/D)max
1 0 0 0 1 Mcrit 10.83
2 0 0 0 2 1.5 3.87
3 0.157806 9.469333 2.8414841
4 0.315611 10.82903 3.86550806
5 0.473417 9.47878 3.78226356
6 0.631223 7.983598 3.39193757
7 0.67631 7.606415 3.27193701
8 0.631223 7.983598 3.39193757
Table E-12:- CL^1.5 over CL analysis data for the F-24 / A-24.
M=Mcrit M=1.5 CL^1.5/CD vs. CL
CL CL^1.5/CD CL^1.5/CD Plot Mach No. (CL^1.5/CD)max
1 0 0 0 1 Mcrit 6.52
2 0 0 0 2 1.5 2.69
3 0.157806 3.7616711 1.12877308
4 0.315611 6.08367624 2.1716158
5 0.473417 6.52190518 2.60239867
6 0.631223 6.34293187 2.69487892
7 0.67631 6.25537056 2.69077853
8 0.631223 6.34293187 2.69487892
The data generated from the aerodynamic analysis of the common F-24 / A-24
airframe configuration was then employed in the performance: constraint: Vn and
manoeuvre: and weight and stability analysis based on the programmed mission
which was entered through the mission builder interface in the AeroDYNAMICTM
toolset.
This concludes discussion on the aerodynamic analysis of the F-24 / A-24 final
configuration using the AeroDYNAMIC V2.08 toolset.