AC Design Project

8/6/2019 AC Design Project

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THE UNIVERSITY OF ADELAIDE

SCHOOL OF MECHANICAL ENGINEERING

AIRCRAFT DESIGN PROJECT 2009

GROUP 5

AUSTRALIAN FIRE-FIGHTING AIRCRAFT

Kevin Chan 1132668

Rachel Harch 1132827

Ian Lomas 1132921

Simon Mitchell 1132439

Carlee Stacey 1132235


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Aircraft Design Project 2009 Australian Fire-Fighting Aircraft

Page 1 of 139

Table of Contents1 Introduction .......................... ...................... ............................ .................... ......................13

1.1 Background................................................................................................................13

1.2 Aim and Objective..................... ....................... ........................... ...................... .........14

2 Feasibility Study................................. ........................... ........................ .......................... ...15

2.1 Literature Review......................... ......................... ............................ ................... ......15

2.2 Market Evaluation .......................... ...................... ............................ ................... ......16

2.2.1 Prototypes ........................ ...................... ............................ .................... ...........17

3 Conceptual Design............................ ...................... ............................... ................... .........203.1 Technical Task......................... ...................... ........................... .................... ..............20

3.1.1 Standard Requirements................ ......................... ............................ ................. 20

3.1.2 Performance Requirements........................ ............................ ...................... ......21

3.1.3 Technical Level ........................... ...................... .............................. .................... 32

3.1.4 Economical Parameters........................ .......................... ......................... ...........32

3.1.5 Main System Requirements..................... ............................ ....................... ........32

3.1.6 Reliability and Maintainability .................... ............................... ................... ......33

3.1.7 Safety.................................................................................................................34

3.1.8 Unification level ......................... ...................... .............................. .................... 34

3.1.9 Ergonomics ....................... ...................... ............................ .................... ...........34

3.1.10 Cabin Design ........................... ...................... ............................ ................... ......34

3.2 Statistical Analysis............................ ........................... ........................ ....................... 34

3.2.1 Empty Weight versus Takeoff Weight....................... ............................ ..............36

3.2.2 Cruise Speed ........................... ...................... ............................ ................... ......37

3.2.3 Stall Speed ........................ ...................... ............................ ...................... .........37

3.2.4 Rate of Climb................................... ............................ .................... ...................37

3.2.5 Cruise Altitude ........................... ...................... .............................. .................... 37

3.2.6 L/D Estimation ........................ ...................... ............................ ................... ......37

3.3 Mission Profile........................ ...................... ........................... .................... ..............39

3.3.1 Mission Profile Diagram ....................... .......................... ......................... ...........39

3.3.2 Mission Profile Requirements................ ......................... ........................... .........39

3.4 Weight Estimation .......................... ...................... ............................ ................... ......40


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3.4.1 Technical Task Requirements ..................... ............................ ...................... ......40

3.4.2 Statistical Analysis Requirements ..................... ............................ ................... ...40

3.4.3 Remaining Sizing Requirements........................ ............................ ................... ...40

3.4.4 Fuel Fraction Estimates ....................... ........................... ........................ ............41

3.4.5 Takeoff Weight Estimation ......................... ............................ ................... .........42

3.5 Sensitivity Analysis............... ......................... ............................ ................... ..............44

3.6 Aircraft Sizing.................... ...................... ............................ ................... ....................45

3.6.1 Sizing to Stall Speed ........................ ............................ ....................... ................45

3.6.2 Sizing to Takeoff Distance................... ...................... .............................. ............45

3.6.3 Landing Distance Sizing ....................... ........................... ........................ ............46

3.6.4 Sizing to Climb Requirements ..................... ............................ ................... .........46

3.6.5 Corrected Lift Coefficient ......................... ........................... .................... ...........47

3.6.6 Drag Polar Estimate........ ......................... ............................ .................... ...........47

3.6.7 Sizing to Cruise Speed Requirements........ ........................ ............................. .....48

3.6.8 Matching Diagram and Design Point........ ......................... ............................ ......49

3.7 Configuration Selection............. ....................... ........................... .................... ...........50

3.7.1 Concept 1............ ........................ .......................... ......................... ....................51

3.7.2 Concept 2............ ........................ .......................... ......................... ....................52

3.7.3 Concept 3............ ........................ .......................... ......................... ....................53

3.7.4 Concept 4............ ........................ .......................... ......................... ....................54

3.7.5 Concept 5............ ........................ .......................... ......................... ....................55

3.7.6 Design Considerations........................ ............................ ....................... .............56

3.7.7 Concept Selection ........................... ...................... ............................ ................. 56

3.8 Fuselage Design ........................ ...................... ............................ ...................... .........57

3.8.1 Cockpit Requirements ........................ ...................... .............................. ............573.8.2 Overall Design of the Fuselage ......................... ............................ ................... ...58

3.8.3 Visibility Diagram ........................... ....................... ............................ .................60

3.8.4 Fire Retardant Tanks and Distribution System ........................ ............................ 61

3.8.5 Fuselage Structure ....................... ......................... ............................ ................. 62

3.9 Propulsion System Design ....................... ............................ ....................... ................63

3.9.1 Propulsion System Type Selection ....................... ........................... .................... 63

3.9.2 Number of Engines and the Power Required per Engine...... ............................ ...65


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3.9.3 Propeller Sizing .......................... ...................... .............................. .................... 67

3.9.4 Propulsion System Integration.................... ............................ ...................... ......69

3.10 Wing Design................ ....................... ........................... ....................... ......................76

3.10.1 Vertical Position ......................... ......................... ........................... .................... 76

3.10.2 Sweep................................................................................................................76

3.10.3 Aspect Ratio............................ ...................... ............................ ................... ......77

3.10.4 Thickness Ratio .......................... ...................... .............................. .................... 78

3.10.5 Taper Ratio ....................... ...................... ............................ .................... ...........78

3.10.6 Twist ......................... ...................... ............................ ................... ....................78

3.10.7 Dihedral ............................ ...................... ............................ ...................... .........79

3.10.8 Wing Loading .......................... ...................... ............................ ................... ......79

3.10.9 Wing Longitudinal Location ..................... ............................ .................... ...........79

3.10.10 Aerofoil Selection ......................... ........................... ........................ ...............80

3.10.11 Incidence Angle ........................ ......................... ............................ ................. 84

3.10.12 Flap Sizing ........................... ...................... ............................ ................... ......84

3.10.13 Aileron Sizing....................... ...................... ............................ ................... ......84

3.10.14 Spoiler Selection............ ...................... ............................ .................... ...........85

3.10.15 Flow Control Devices ....................... ........................... ........................ ............85

3.10.16 Wing Tips ............................ ...................... ............................ ................... ......85

3.10.17 Centre of Gravity .......................... ........................... ........................ ...............86

3.10.18 Structure........................................................................................................86

3.10.19 Wing Design Summary............................ ............................ ...................... ......87

3.11 Empennage Design ........................... .......................... ........................... .................... 88

3.11.1 Empennage sizing............................ ............................ ....................... ................88

3.11.2 Horizontal Stabiliser Geometry........ ......................... ............................... ...........893.11.3 Vertical Stabiliser Geometry..................... ........................... .................... ...........90

3.11.4 Elevator Sizing and Geometry....... ......................... ............................ ................. 90

3.11.5 Rudder Sizing and Geometry ...................... ............................ ................... .........92

3.11.6 Stabiliser Aerofoils ........................ ........................ ........................... ..................93

3.12 Landing Gear Design .......................... ......................... ........................... ....................95

3.12.1 Landing gear arrangement ......................... ............................ ................... .........95

3.12.2 Landing Gear Sizing Nomenclature ........................ ............................ ................. 96


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3.12.3 Landing Gear Placement Criteria ........................... ............................ ................. 97

3.12.4 Nose Weight Criterion........................ ............................ ...................... ..............97

3.12.5 Height Criterion.................... ......................... ............................ ................... ......98

3.12.6 Landing Gear Position ........................ ............................ ....................... .............99

3.12.7 Nose Weight Criterion........................ ............................ ...................... ............101

3.12.8 Height Criterion.................... ......................... ............................ ................... ....101

3.12.9 Roll-Over Criterion ....................... ......................... ............................ ...............101

3.12.10 Over-Turn Angle Criterion ...................... ............................ ................... .......102

3.12.11 Tip-Back Angle Criterion ......................... ............................ ................... .......102

3.12.12 Summary......................................................................................................102

3.12.13 Landing Gear Loads ........................ ...................... .............................. ..........102

3.12.14 Number, Type and Size of Tyres.................... ............................ ................... .103

3.12.15 Tyre Pressure Calculations........... ......................... ............................ ............ 103

3.12.16 Suspension Method and Requirements ......................... ........................... ....104

3.12.17 Length and Diameter of Landing Gear Struts ......................... ....................... 105

3.12.18 Nose-Wheel Steering and Castoring Dimensions ........................ .................. 106

3.12.19 Gear Retraction Geometry ..................... ............................ ...................... ....107

3.13 Isometric Views ........................ ...................... ............................ ...................... .......108

4 Weight and Balance Analysis ...................... .......................... ......................... .................. 109

5 Stability Analysis.................... ...................... ............................ .................... .................... 111

6 Aerodynamic and Performance Analysis.................... ............................... ................... ....114

6.1 Aerodynamic Analysis ........................ ...................... .............................. .................. 114

6.1.1 Zero-Lift Drag Coefficient Calculation ....................... ............................ ............ 114

6.1.2 Required Lift Coefficients in Cruise and Loiter Phases........................ ...............114

6.1.3 Drag Coefficient in Cruise and Loiter Phases..................... ............................ ....115

6.1.4 Lift to Drag Ratio Calculation ...................... ............................... ................... ....1156.2 Final Design Weight Estimate....... ...................... ............................... ................... ....115

6.3 Design Point Analysis ......................... ......................... ........................... .................. 116

7 Conclusion.......................................................................................................................118

8 References ............................ ...................... ............................ .................... .................... 119

Appendix A Fire-fighting Aircraft Statistical Analysis ........................... ...................... ............122

Appendix B Statistical Analysis Relevant Aircraft....................................................................123

Appendix C Calculated Fuel Fractions ........................ ............................ ................... ............124

Appendix D Sensitivity Calculations ..................... ............................ ................... .................. 127

Appendix E MATLAB Code for Takeoff Weight Estimation and Sensitivity Analysis................132


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Appendix F - Honeywell TPE331-14GR Specifications...............................................................135

Appendix G Flap Sizing Data ...................... ........................ ........................... ........................ 136

Appendix H Neutral Point Calculations ......................... ........................... .................... .........137

Appendix I Three View Drawings ......................... ............................ ................... .................. 139


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List of Figures

Figure 1 - Coordinate System (NASA 2009) ........................... ............................ ................... ......12

Figure 2 - Air Tractor 602 (Airliners.net 2009)............................................................................17

Figure 3 - Air Tractor 802 (Airliners.net 2009)............................................................................18

Figure 4 - Canadair CL-215 (Airliners.net 2009)..........................................................................18

Figure 5 - Canadair CL-415 (Airliners.net 2009)..........................................................................19

Figure 6 - Major Australian Airports (Australian Institute of Criminology Website 2004)............23

Figure 7 - Fire Danger Seasons (Australian CSIRO Website 2009) ........................... .................... 23

Figure 8 - Map of the Population Distribution in Australia .......................... .................... ...........24

Figure 9 - Map of land usage in Australia overlayed with areas covered by the aircraft located at

the selected bases. The solid circles indicate most likely bases, and the dashed circles indicate

other possible aircraft bases (Modified from Australian Natural Resources Atlas Website 2008).

.................................................................................................................................................24

Figure 10 - Probability for the Success of a First Attack Success (Plucinski, Gould, McCarthy,

Holis, 2007)...............................................................................................................................25

Figure 11 - Probability for the Success of a First Attack Success (Plucinski et al. 2007) ...............26

Figure 12 Figure showing the regions within Australia which can be reached by the fire-fighting

aircraft within different response times (Modified from the Australian Natural Resources Atlas

Website 2008)...........................................................................................................................28

Figure 13 Figure showing the response time of the fire-fighting aircraft overlayed onto a

population density map (Modified from the Department of Environmental, Water, Heritage and

the Arts Website 2001) ............................ ...................... ............................... ............................ 29

Figure 14 - Australian Runway Lengths ...................... ............................... ................... ..............31

Figure 15 - Graph of Takeoff Weight versus Empty Weight for Statistically Analysed Aircraft.....36

Figure 16 - Mission Profile.........................................................................................................39

Figure 17 - Takeoff and Empty Weight Estimate ........................... ............................ ................. 43

Figure 18 - Matching Diagram with Met Area and Design Point Marked ........................... .........49

Figure 19 - Concept 1 Sketch ....................... ......................... .......................... ........................... 51





Figure 24 - Cockpit Dimensions ...................... ............................ ...................... ......................... 58

Figure 25 - Fuselage Sketch ......................... ......................... .......................... ........................... 58

Figure 26 - Front View of Fuselage Sketches ......................... ............................ ................... ......59

Figure 27 Visibility Diagram ......................... ............................ ................... ............................ 61

Figure 28 - Tank Location in the Fuselage ........................ ........................... ...................... .........61

Figure 29 Engine Selection: Single Engine versus Twin Engine.................................................66

Figure 30 - Propeller Engine Configurations: Tractor and Pusher (Raymer 2006 p.252)..............71


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Figure 31 - Engine Mounting Locations: Fuselage, Wings, Tail or as Upper Fuselage Pod...........72

Figure 32 - Honeywell TPE331-14GR Geometry (all dimension in inches) (Honeywell 2006).......72

Figure 33 - Cooling System Configuration (Raymer 2006, p.256)................................................73

Figure 34 - Empennage Configurations (Raymer 2006) .......................... ........................... .........88

Figure 35 - Horizontal Stabiliser Arrangement ......................... .............................. .................... 89

Figure 36 - Vertical Stabiliser Arrangement ..................... ........................... .................... ...........90

Figure 37 - Elevator Geometry...................................................................................................91

Figure 38 - Elevator Trim Tab Geometry ......................... ............................ .................... ...........92

Figure 39 - Rudder Geometry ......................... ............................ ...................... ......................... 93

Figure 40 - Rudder Trim Tab Geometry......................................................................................93

Figure 41 Landing Gear Configurations (Raymer 2006) ............................ ...................... .........95

Figure 42 - Landing Gear Nomenclature (Roskam 2006) ........................ ............................ ........97

Figure 43 - Over-turn Angle Criterion (Raymer 2006 p. 232) .......................... ................... .........98Figure 44 - Figure Describing Over-turn Criterion.......................................................................99

Figure 45 - Figure Showing Trail and Rake of the Wheel (Raymer 2006)...................................106

Figure 46 - Sliding Bar Linkage (Raymer 2006) ......................... .............................. .................. 107

Figure 47 - Centre of Gravity Envelope ......................... ............................ ................... ............110

Figure 48 - Longitudinal X-plot for the Operational Empty Weight Configuration.....................112

Figure 49 - CG Envelope, Neutral Point and Static Margin for Each Flight Configuration...........113

Figure 50 - Weight Estimate for the Final Design ..................... .............................. .................. 116

Figure 51 - Final Matching Diagram. .................. ........................... ....................... .................... 117


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List of Tables

Table 1 Summary of Response Times for an Aircraft Cruise Velocity of 375 km/h....................27

Table 2: Payload Drop Types. ......................... ............................ ...................... ......................... 30

Table 3 - Aircraft Operating Conditions......................................................................................32

Table 4 - Mission Profile Summary ......................... ............................ ................... ....................39

Table 5 - Parameters Estimated from Prototypes and Literature ........................... .................... 40

Table 6: Estimated Fuel Fractions (Roskam 2005)......................... ............................ ................. 41

Table 7: Sensitivity Analysis Results...........................................................................................44

Table 8 - Aircraft Sizing Results..................................................................................................49

Table 9 - Design Considerations.................................................................................................56

Table 10 - Fineness Ratio as Specified by Roskam (2004) ......................... .......................... ........59

Table 11 - Comparison of the Fineness Coefficient for the Designed Aircraft Compared with the

Recommended Values as Specified by Roskam (2004)...............................................................60

Table 12 - Recommended Frame and Longeron Spacing, and Frame Depth for a Small

Commercial Aircraft as specified by Arjomandi (2009)...............................................................62

Table 13 - Suggested Engine Models (Jackson 2008)..................................................................67

Table 14 - Statistical Analysis of Relevant Engines (Roskam III 2002)..........................................68

Table 15 - Aerofoil Candidates...................................................................................................82

Table 16 - 2D Aerofoil Comparison Table...................................................................................82

Table 17 - 3D Aerofoil Comparison Table...................................................................................82

Table 18 - Wing Tip Table ......................... ......................... ........................... .......................... ...85

Table 19 - Wing Design Summary ...................... ........................... ....................... ......................87

Table 20 - Tyre Selection Table................................................................................................103

Table 21 - Suggested Weight Distribution as Percentages (Eger 1983; Arjomandi 2009) ..........109

Table 22 - Aircraft Weight Breakdown and Centre of Gravity Locations .......................... .........109

Table 23 - Centre of Gravity Locations for Various Payload and Fuel Configurations ................ 110

Table 24 - Longitudinal Stability in Each Flight Configuration ......................... ...................... ....112

Table 25 - Comparison of Assumed and Estimated Lift to Drag Ratios......................................115

Table 26 - Fire-fighting Aircraft Statistical Analysis...................................................................122

Table 27 - Honeywell TPE331-14GR Specifications (Jackson 2008) and (Honeywell TPE331-14

2006) ........................... .................... .............................. .................... ............................ .........135

Table 28 - Flap Sizing Table......................................................................................................136

Table 29 - Aileron Sizing Table.................................................................................................136


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Nomenclature

Acronyms

2D Two -Dimensional

3D Three-Dimensional

CAD Computer Aided Design

CASA Civil Aviation Safety Authority

CFS Country Fire Service

FAR Federal Aviation Regulations

MAC Mean Aerodynamic Chord of wing

NACA National Advisory Committee for Aeronautics

NAFS National Aerial Firefighting Centre

NASA National Aeronautics and Space Administration

UIUC University of Illinois at Urbana-Champaign

Symbols

a Speed of sound

Vclimb Climb Velocity

(L/D)aircraft Aircraft (L/D)

(L/D)aerofoil Aerofoil (L/D)

(L/D)max Maximum L/D

(L/D)wing Wing (L/D)

(t/c)wing Wing thickness ratio

(W/S) Wing loading

A exhaust Area of Exhaust

Aintake Area of Intake

ARwing Wing aspect ratio

B The distance between the nose and the main landing gears

bwing Wing span

Cd Aerofoil drag coefficient

CGaft The distance from the nose of the aircraft to the most aft CG

CGfore The distance from the nose of the aircraft to the most forward CG


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CGwing Wing centre of gravity

CL Wing lift coefficient

Cl Aerofoil lift coefficient

CL max Maximum wing lift coefficient

Cl max Maximum aerofoil lift coefficient

CL max, L Maximum wing lift coefficient at landing

CL max, TO Maximum wing lift coefficient at takeoff

CL Lift-curve slope

Cm Aerofoil pitching moment coefficient

CM Wing pitching moment coefficient

Cwing Wing chord

D Drag

Dfuselage The diameter of the fuselage

DPPropeller Diamater

g Acceleration due to gravity (32.2 slugs/ft3)

Hlg The height of the landing gear (from the ground to the bottom of

the fuselage)

Htail The height of the tail above the bottom of the fuselage

HW The half-width of the main landing gear, i.e. the lateral distance

between a main landing gear and the centre-line of the aircraft

L Lift

l characteristic length

Lfuselage The length of the fuselage [ft]

M Mach number

Ma The distance between the main landing gear and the most aft CG

Mf The distance between the main landing gear and the most forward

CG

Na The distance between the nose landing gear and the most aft CG

Nf The distance between the nose landing gear and the most forwardCG

npNumber of propeller blades

PblBalde Power Loading

PmaxMaxium power output per engine

S Platform area

Sflapped Flapped surface area

SHP Uninstalled Engine Power

Sref Reference surface area


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Swet Wetting surface area

THP Installed Engine Power

Vcr Velocity at cruise altitude

Vfoam Foam Volume

Vfuel Fuel Volume

VrotRotational Speed of Engine

VtipPropeller Tip Speed

VTO Takeoff velocity

WE Engine weight

WE Installed Installed Engine weight

Wlanding The total weight of the aircraft at landing

WTO Takeoff weight

xnosegear The distance between the nose of the aircraft and the nose landing

gear


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Coordinate System Designation

Figure 1 below shows the coordinate system used throughout this report.

Figure 1 - Coordinate System (NASA 2009)


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1 IntroductionThe purpose of this report is to detail the design of an Australian fire-fighting aircraft.

1.1BackgroundBushfires present a significant risk to Australia and its people, land and resources. Recently, 210

people died when the 2009 Victorian bushfires destroyed over 400,000 hectares of land (WA

Today 2009). It is imperative that there be systems in place to suppress and control suchbushfires to minimise the risk to human life. One of the most effective methods of containing a

bushfire is through aerial fire-fighting, which is the use of an aircraft for releasing fire fighting

chemicals onto a fire. Both fixed wing and rotary wing aircraft are capable of aerial fire-fighting,

with possible chemicals including water, foams, gels and fire retardants. The key characteristics

of a fire-fighting aircraft include a high useable payload weight and a high cruise speed.

Several aircraft designs have demonstrated excellent aerial fire fighting effectiveness, including

those specially modified for aerial fire-fighting purposes. For large fires, modified commercial

airliners or military transport aircraft have been used with great success. In the past, Australia

has considered using larger aircraft for aerial fire-fighting, but this has proven to be both

expensive and unnecessary. Small companies contracted by state and commonwealth

governments use modified agricultural aircraft, such as the Air Tractor 802, Air Tractor 602 and

M18 Dromader aircraft, for aerial fire-fighting (Dunn Aviation Australia 2009). Agricultural

aircraft often have poor aerodynamic efficiency, but posses improved manoeuvrable over larger

aircraft.

A market analysis was performed to compare existing fire-fighting, agricultural and twin-engine

regional turboprop aircraft. Different configurations were examined, and the most optimal

aircraft were selected. The aerodynamics, stability and performance of the aircraft were

investigated, before a final design was proposed and documented using CAD models and

engineering drawings. A description of manufacturing, maintenance and through-life support is

beyond the scope of the project.


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1.2Aim and ObjectiveThe aim of this project is to design an Australian fire-fighting aircraft. A design tailored for

unique Australian conditions would give the aircraft an advantage in performance and mission

effectiveness compared with fire-fighting aircraft currently used in Australia. The project will

focus on the conceptual phase of the design process. The primary purpose of the project is to

teach undergraduate students the aircraft design process.


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2Feasibility StudyThe feasibility study was conducted at the beginning of the project to determine the viability of

the project concept and scope. The feasibility study consists of a literature review of texts

pertaining to aircraft design, and a market evaluation and benchmarking study to investigate

similar aircraft

2.1Literature ReviewThe conceptual design of the fire-fighting aircraft required research of current prototypes and

design techniques through a literature review. A comprehensive investigation was carried out,

which yielded a number of useful references, including textbooks, published reports, databases

and websites. These sources will be discussed in the following sections, and include those used

for the design of the aircraft structure, configuration and sizing. During the feasibility study and

statistical analysis, numerous aircraft were referenced for statistical data. Aircraft primarily

designed for aerial fire-fighting did not provide adequate data, so agricultural aircraft were also

considered. Of particular interest were the Air Tractor series of aircraft.

The literature used for the project is based on information and equations contained in a range of

texts pertaining to different aspects of aircraft design. For the general embodiment design,

several textbooks and reference books were used. These were namely the Airplane Design

series (Roskam, 2004) andAircraft Design: A Conceptual Approach (Raymer, 1992). The Roskam

series provides an incremental approach to the design of an aircraft, which can be adapted to

suit the requirements specific to the fire-fighting aircraft. In contrast, Raymer offers a classical

approach to aircraft design with detailed theory and equations.

Aerofoil selection was aided with the use of the UIUC Aerofoil Coordinate Database (UIUC 2008).

This database provides a considerable selection of aerofoils designed and recommended for

aircraft. In addition, Javafoil aerofoil analysis online software was used to compare and select

the most appropriate and suitable aerofoils for the aircraft. Introduction to Aeronautics: A

Design Perspective (Brandt et al. 2004) was used for stability calculations and determination of


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landing gear location. Other references have also been used throughout the project, and are

cited where applicable.

2.2Market EvaluationA market evaluation of existing fire-fighting aircraft was undertaken in order to gain knowledge

regarding fire-fighting aircraft. The market evaluation was conducted in parallel with the

literature review, and provided the group with invaluable knowledge regarding fire-fighting

aircraft and valuable benchmarking from which design work could be compared.

Initial investigation focused on fire-fighting aircraft. Properties such as take-off weight, empty

weight, payload capability, cruise speed, range and wing area were determined for over twenty

aircraft that had fire-fighting capabilities. These aircraft included the following:

Bronco OV-10

TBM Avenger

Douglass DC-3

Grumman F7F-3 Tigercat

Grumman S2-Tracker

Grumman CDF S-2A Tracker

Bombardier Canadair 415

Bombardier Canadair CL-215

Consolidated PB4Y-2 Privateer

Boeing B-17 Flying Fortress

Alenia C-27J Spartan

Douglas DC-4

Fairchild C-119 Boxcar

Beriev Be-200 Altair

Shinmaywa US-1A

P3-Orion

McDonnell Douglas DC-7

C-130 Hercules

JRM Mars

McDonnell Douglas DC-10-10

Boeing 747

Antonov An-2 'Colt'

ROKS-Aero T-101 Grach


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The capabilities of these aircraft, tabulated in Appendix A, exhibited significant variation. The

investigated aircraft included both amphibious and non-amphibious aircraft, converted jet transport

aircraft and small single piston engine aircraft. The confliction in the data meant that it was not

possible to determine a defining relation between takeoff weight and empty weight. However,

several conclusions could be drawn from this data as outlined below:

Both amphibious and non-amphibious aircraft are used for fire fighting. Amphibious aircraft

demonstrate great payload capability relative to takeoff weight. However, the design

complexity and limitation of suitable landing locations in Australia meant that the

amphibious aircraft were considered unfavourable.

Large aircraft with fire-fighting capabilities are often produced as single models. These

appeared to represent heavily modified transport aircraft rather than specially designed

fire-fighting aircraft. Consequently, they exhibit comparatively reduced payload capability

compared to smaller aircraft that are intentionally designed for fire-fighting capacities.

2.2.1 PrototypesThe selection of these prototypes was based on the following:

Similar physical size to the expected fire-fighting aircraft size

Similar weight to the expected fire-fighting aircraft size

Similarity of mission requirements and applications

The Air Tractor 602 is a single engine turboprop agricultural aircraft. It has a maximum takeoff

weight of 12,500 lb and has a payload capacity of 630 gallons (2,380 L). The first flight of the Air

Tractor 602 occurred in 1995, with production currently continuing. (Air Tractor 2009)

Figure 2 - Air Tractor 602 (Airliners.net 2009)


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The Air Tractor 802F is a single engine turboprop aircraft primary designed for fire-fighting

applications. It has a takeoff weight of 1,600lb and a payload capacity of 820 gallons (3,100L). The

Air Tractor 802F is a modified version of the Air Tractor 802 agricultural aircraft. The 802 is the

largest existing agricultural aircraft, and as such, defines the boundaries of agricultural aircraft

design. Both models are popular as they offer high efficiency and similar performance compared

with larger twin-engine aircraft. The first flight of the Air Tractor 802 occurred in 1990, and

production of both the 802 and 802F models is currently continuing. The 802F can also be fitted

with Wakeri Floats to create an amphibious aircraft (Air Tractor 2009).

Figure 3 - Air Tractor 802 (Airliners.net 2009)

The Canadair CL-215 is a twin engine amphibious fire-fighting aircraft. It has a take-off weight from

land of 43,500 lb and a payload capacity of 1,400 gallons (5,455 L). The first flight occurred in 1967

and production ceased in 1998 with 121 aircraft built. The CL-215 has a flying boat configuration,

and hence, offers significant aerodynamic advantages when compared with the Air Tractor 802F

fitted with floats. The CL-215 was designed for Canadian conditions, where large lakes provide still

flat surfaces where rapid water collection can occur. (Airliners.net 2009)

Figure 4 - Canadair CL-215 (Airliners.net 2009)


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The Canadair CL-415 was developed from the CL-215, and first flew in 1993. The CL-415 offers

advantages such as an increased takeoff weight of 43,850 lb and a payload capacity of 1,620 gallon

(6,120 L). Other design improvements include an updated cockpit, improved water release system

and corrosion resistance. The CL-415 has been popular in Canada, France and Italy. However, as the

aircraft refills by scooping water from larger rivers or lakes, it does not meet Australian requirements

(Airliners.net 2009).

Figure 5 - Canadair CL-415 (Airliners.net 2009)


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3Conceptual DesignThe conceptual design process aimed to generate, select and develop the most feasible concepts

that could meet all the design requirements. This process was conducted using a classical approach

involving multiple design iterations. Each iteration led to further development of the concepts until

design decisions were made based on sound knowledge and calculations. The following section

outlines the conceptual design process, from initial configuration design through to planform design,

aerofoil and control surface selection, fuselage sizing and propulsion system selection. The resultant

design is brought together in three view drawings.

3.1Technical TaskThis section outlines design requirements for the aircraft. Requirements due to standards,

performance, technological level, economics, main sub-systems and reliability are used to define the

overall constraints on the aircraft.

3.1.1 Standard RequirementsThe aircraft must be compliant with regulations defined by the Civil Aviation Safety Authority (CASA)

and the National Aerial Firefighting Centre (NAFC). CASA regulations outline required performance

parameters, construction, testing and operational procedures. Civil aircraft operating in Australia

must receive CASA certification, and hence, it is necessary that the aircraft satisfies all relevant CASA

requirements. This design will be engineered and constructed in Australia, and hence, must adhere

to the Type Certificate (Australian Manufactured) and be manufactured by a CASA approved

company (CASA 2008). CASA regulations frequently refer to requirements defined by the Federal

Aviation Regulation (FAR). A fire-fighting aircraft will need to satisfy components of Part 25

(Airworthiness Standards: Transport Category Airplanes) and Part 91 (General Operating and Flight

Rules). FAR does not outline specific requirements for fire-fighting aircraft. Consequently, Part 137

(Agricultural Aircraft Operations) will be utilized for additional requirements.

NAFC is an Australian Commonwealth government organisation that coordinates the procurement of

fire-fighting aircraft, and defines the required capabilities of fire-fighting aircraft, specifying the

required payload capabilities and delivery systems.


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3.1.2 Performance RequirementsAircraft Base Location and Range

The aircraft is being designed to supplement the existing aerial fire-fighting capabilities of Australia.

The location at which the aircraft would potentially be based is an important consideration when

determining the range of the aircraft. Once possible bases are identified, the range can be

determined by identifying distances that the aircraft would be required to travel to the site of a fire.

The fire-fighting aircraft being designed will be larger than the existing aircraft currently used by

Australia, which will enable a greater amount of fire retardant to be released upon arrival. To enable

a more economical usage of these aerial fire-fighting resources, it is intended that these aircraft will

operate out of major Australian airports, where existing maintenance facilities and personnel can be

utilised. By centralising the fleet, it is hoped that placing fire-fighting aircraft on standby during

extreme fire hazard days will be more easily accommodated.

Operational costs of the aircraft will also be significantly less, and allows for the set up of specialised

facilities to assist with the loading, maintenance of the aircraft, and to reduce the number of aircraft

(and, because of this, the cost) of placing aerial fire-fighting aircraft on standby. Although aerial fire-

fighting aircraft cannot be used in populated areas due to the hazard of the falling fire retardant,

generally, populated areas are the most central location about which regional areas, where fixed

wing aerial aircraft are most effective, are located.

By examining Figure 7, Figure 8, and Figure 9, it is likely that aircraft would need be based out of, or

nearby, the following airport:

Perth

Adelaide

Melbourne (Tullamarine)

Mildura

Sydney (KSA)

Canberra (The region surrounding Canberra could be covered by aircraft based out of

Melbourne and Sydney. Due to political reasons and public perception, it is likely that an

aircraft would be based at Canberra regardless).

Tamworth


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Hobart (Unlikely to warrant its own aircraft due to the climate. If the range of the aircraft is

sufficient, Tasmania could be covered by an aircraft based out of Melbourne.)

Mackay (Unlikely, as the population density near Cairns is small. This would not warrant a

first attack aircraft. Since the fire season for the north of Australia is during winter, it is

possible to locate the aircraft stationed in the southern regions during summer and in the

northern regions during winter).

Examining the fire danger seasons from Figure 7, the largest number of populated regions within

Australia are exposed to the fire danger seasons during summer. To enure sufficient coverage of all

fire danger areas, the following minimum aircraft bases are recommended to provide sufficient

coverage throughout the summer:

Perth

Adelaide

Melbourne

Sydney

Tamworth

It is also recommended that aircraft be stationed at the locations listed below for additional

coverage, faster response times to all areas, and to ensure that there is a degree of contingency

should aircraft from one location be unable to be deployed to a nearby fire:

Canberra

Hobart

Mildura

During other seasons, it would be possible to relocate aircraft from the above bases to other

locations. Using Figure 8 and Figure 9, the distances between these bases, and the potential regions

requiring aerial fire-fighting assistance, were determined. The selected range was selected to be a

minimum of 500km (one way), as this provides sufficient coverage of most regions within Australia.

Consequently, the aircraft should be capable of flying in cruise configuration for up to 1000km. The

coverage provided by an aircraft with these capabilities is shown in Figure 8 and Figure 9.


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Figure 6 - Major Australian Airports (Australian Institute of Criminology Website 2004)

Figure 7 - Fire Danger Seasons (Australian CSIRO Website 2009)


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Figure 8 - Map of the Population Distribution in Australia

(Modified from the Department of the Environmental, Water, Heritage and the Arts Website 2001)

Figure 9 - Map of land usage in Australia overlayed with areas covered by the aircraft located at the selected

bases. The solid circles indicate most likely bases, and the dashed circles indicate other possible aircraft

bases (Modified from Australian Natural Resources Atlas Website 2008).


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Desired Response Time and Cruise Velocity

As the aircraft is being designed primarily as a first attack aircraft, the response time has a direct

impact upon the success of the first attack. The sooner the aerial fire-fighting aircraft arrives at the

scene, the higher the probability of the first attack being successful. A successful first attack refersto an occasion where the contribution of a first attack aircraft contributed to controlling the fire.

The desired response time of the fire-fighting aircraft can be determined by considering the

probability of success of a first attack by a fixed wing aircraft. This is shown graphically in Figure 10

and Figure 11. From these figures, it can be seen that the probability of success is greater if the time

to first attack is reduced.

Figure 10 - Probability for the Success of a First Attack Success (Plucinski, Gould, McCarthy, Holis, 2007)


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Figure 11 - Probability for the Success of a First Attack Success (Plucinski et al. 2007)

It can be seen from Figure 11, that as the time increases, the probability of a successful first attack is

reduced. For an immediate first attack (i.e. a time of zero), the probability of success for FFDI (Forest

Fire Danger Index) values 50), the effects of the response time on the probability of success

are more pronounced. For the ideal, zero time to first attack, the probability of a successful first

attack is approximately 50%. After one hour, this has dropped to approximately 45%, and after 2

hours, the probability has dropped to approximately 40%. It is fires on high FFDI days such as those

recently experienced in Victoria, which threaten to cause the most harm to people and property.

Any advantage to assist with the control of these fires would be desirable. As a result, it is desirable

to achieve the fastest response time possible.

To design an economical aircraft to meet Australias fire-fighting needs, some compromise is

required. Although it would be desirable to have the first attack aircraft reach every possible

location of a fire within 30 minutes, this is not feasible. It was decided that the aircraft have a


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response time of no less than 2 hours, including the time from when the first call is received to when

the aircraft takes off from the runway.

For the purpose of this report, it will be assumed that the time between receiving notification of the

fire and takeoff is 30 minutes. The aircraft is therefore required to travel a minimum of 500km in 1.5

hours. This requires a cruise velocity of 333.33km/h. The aircraft will therefore be designed with a

375km/h (or 189 knots) cruise speed.

Table 1 Summary of Response Times for an Aircraft Cruise Velocity of 375 km/h

Response Time Distance of Fire

Approximate

probability of

success FFDI 24

(high)

Approximate

probability of

success FFDI 50

(extreme)

0.5 hours 0 km

1.0 hours 175 km 82% 50%

1.5 hours 350 km

1.9 hours 500 km

2.0 hours 525 km 78% 40%

The above coverage is shown graphically in Figure 12 and Figure 13.


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Figure 12 Figure showing the regions within Australia which can be reached by the fire-fighting aircraft

within different response times (Modified from the Australian Natural Resources Atlas Website 2008)


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Figure 13 Figure showing the response time of the fire-fighting aircraft overlayed onto a population

density map (Modified from the Department of Environmental, Water, Heritage and the Arts Website 2001)

Although the aircraft is designed to return to base if necessary, for extended aerial suppression

campaigns, it is intended that the fire retardant is transported to a closer regional airport and the

aircraft can use this as a base to reduce the turnaround time and fuel costs. It is hoped that the

larger payload capacity and faster response time of the fire-fighting aircraft will allow increased

suppression of the fire, and hence, a more effective first attack.

Payload Weight


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Aerial fire-fighting aircraft standards require that fixed wing aircraft drop retardant or water

payloads in an effective zone which is no less than 40 m long and 15-20 m wide, and that no more

than 15% of the release falls outside of this effective zone (NAFC 2004). The standards require a

minimum coverage of 0.2 L/m2. However, coverage up to 4.0 L/m

2is required to suppress the

heaviest bushfires (Plucinski et al. 2007). Standards also require a leakage loss rate of no more than

15 L/hr. To provide 4 L/m2

coverage to an effective zone of 40m by 20m and assuming a total time

between payload delivery and filling of 140 minutes (20 minutes between filling and takeoff, 100

minutes to target and 20 minutes on target), the volume of fire retardant required is calculated as

follows:

Equation 1: Required Payload Volume

Long-term fire retardants, such as Phos-Chek D-75-R, are up to three times more effective in

containing bushfires than water (Plucinski et al. 2007). The payload of the fire-fighting aircraft can be

assumed to have a similar density to Phos-Chek D-75-R of 1.067 kg/L (USDA Forest Service 2006).

The payload mass is then 3,966 kg, which was rounded up to 4,000 kg as a conservative estimate to

allow for possible density variations. A payload of 4,000 kg of Phos-Chek allows the payload drop

types seen in Table 2. A three-drop configuration may be possible, depending on the payload

delivery system, but is not required by aerial fire-fighting aircraft standards.

Table 2: Payload Drop Types.

Drop type Coverage

One drop 4 L/m2

Two drops 2 L/m2

Four drops 1 L/m2


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Crew Weight

NAFC outlines a pilot weight of 190 lb (86kg), with 15kg of baggage. The aircraft should only provide

accommodation for one crew member. No additional crew members are required to operate theaircraft. Hence, controlling the aircraft and releasing the fire retardant are both performed by the

pilot.

Takeoff and Landing Capabilities

Due to the mission of the aircraft, it is desirable for the aircraft to be operated from all airports in

Australia. Runway lengths for airports are shown graphically in Figure 14.

Figure 14 - Australian Runway Lengths

The presence of several short personal runways significantly skews the data. Consequently, it was

decided that the aircraft should operate from the upper 75th

percentile of Australian runways. This

suggests a take off and landing length of 4000ft.

Operational Conditions

The operating conditions of fire-fighting aircraft were researched. However, no overriding

documents or guidelines were found. Consequently, the meteorological conditions of the ten worst

bushfires in Australia's history were investigated. From investigation, the extreme of the aircraft

operating conditions were determined.


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Table 3 - Aircraft Operating Conditions

Air Temperature (C) 46 2009 Victorian Bushfires

Maximum Temperature in Fire (C) 2000 1983 Ash Wednesday Bushfires

Temperature Recommended byBuilding Codes for Bushfire Prone

Areas (C)

1300 2009 Victorian Bushfires

Wind Speed (km/hr) 120 Mount Lubra Bushfires

Humidity 6% 2009 Victorian Bushfires

Air Pollution

1500 g of small

particles per cubic

meter

2009 Victorian Bushfires

Speed of Burning Front Forest

(km/h)11 Otways Bushfires

Speed of Burning Front

Grassland (km/h)22 Otways Bushfires

The above conditions outline an extreme bushfire normally classified as a firestorm. The height of

the fire front can be over 15m (50ft). The formation of Pyrocumulus cloud can lead to serve

turbulence.

3.1.3 Technical LevelThe aircraft is designed to replace existing aircraft, and hence, should demonstrate improved

technologies. In particular, increased fuel efficiency, improved materials and better manufacturing

processes are desirable. The cockpit should also benefit from superior instrumentation. It is

intended that this aircraft will be flown by a single pilot with high-level skills and appropriate

certification.

3.1.4 Economical ParametersThe aircraft should be affordable by small companies as well as larger organisations and government

bodies. It is intended that the aircraft should be more affordable than competing aircraft, in initial

purchase cost, running costs and maintenance costs.

3.1.5 Main System RequirementsPropulsion System Requirements

Propulsion requirements are outlined in FAR 25 Subpart E. Particular reference should be made to

Section 25.961 (Fuel System Hot Weather Operation). No specifications regarding engine number or

engine type exist.


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Landing Gear Subsystem

Rural operation requires that the aircraft must be able to operate from paved and unpaved runways.

Amphibious landing capabilities are not required. FAR 25 Section 25.473 requires the following:

Maximum descent velocity of 10ft/s at the design landing weight

Maximum descent velocity of 6ft/s at the design takeoff weight

The coefficient of friction between the tires and the ground should be less than 0.8

Fuselage Requirements

The fuselage design is required to accommodate the fire retardant release system.

Fire Retardant Release System

NAFC specifies the following requirements:

The fire retardant release system must be able to produce a full dump with a minimum

flow rate of 1000 litres per second under typical dumping conditions.

The system must be capable of dropping fire retardants at rates less than the maximum flow

rate. It is recommended that the system is capable of at least four flow rates. Flow rates of

500 litres per second, 1000 litres per second and 1500 litres per second are recommended.

The systems must be capable of splitting the load into more than one drop. Systems with

capacity greater than 3000L must be able to drop the load in four parts.

The system should be well constructed and include appropriate sealing mechanisms to

prevent leakages. During sixty minutes of static ground testing, losses should be less than

two litres. During a twenty minute turnaround, mission losses should be less than five litres.

The systems should have the capability to inject the water payload with a measured amount

of foam concentrate.

3.1.6 Reliability and MaintainabilityNAFC recommends the following:

Systems should be simple, robust and reliable

Systems should have an appropriate level of redundancy. In the event of partial equipment

failure, it must be possible to continue the firebombing mission.

The use of specialised parts should be avoided

The aircraft should be field maintainable


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3.1.7 SafetyFAR 91 Section 91.107 states the requirements of one shoulder safety belt as a minimum

requirement for all aircraft. FAR Part 137 requires that agricultural aircraft be fitted with a bird proof

windshield, wire cutters and wire deflectors due to their low altitude operation. The criteria will alsobe applied to the aircraft.

3.1.8 Unification levelThe vehicle should incorporate both new and existing design components. Inherited design elements

include the wing and empennage aerofoil, the propulsion system, and the flight deck

instrumentation. New designs will occur for the fuselage and fire retardant release system. Iterative

design of the aircraft aerodynamics and the fire retardant release system will be required to reach

the optimal design solution.

3.1.9 ErgonomicsNAFC recommends that the aircraft should be controllable without excessive strength or movement

by the pilot. In particular, fire retardant release should not result in large pitch movements or

excessive trim changes.

3.1.10Cabin DesignTo achieve high accuracy when releasing the fire retardant, the pilot visibility pattern must be

considered. The cockpit should be designed such that the over-nose angle is a minimum of ten

degrees. The pilot should have over-the-side vision of 35 degrees, with 70 degrees of head

movement. The pilot should have completely unobstructed upward vision angles. The cockpit

windscreen should have a minimum angle of 30 degrees to prevent mirroring effect of sunshine

angles.

3.2Statistical AnalysisStatistical analysis of relevant data is required to produce the technical diagram and suggest base

parameters for design. The technical task outlined a payload capability of 8,820 lb and a range of

584nm. These definitions were used to determine the relevance of aircraft data. Only aircraft

currently in use were considered.

The statistical analysis was limited by relevant fire-fighting aircraft. Consequently, additional data

points were obtained by using agricultural aircraft and small regional turboprops. The investigated

aircraft included the following:


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Bombardier Canadair 415 (Fire-fighting Aircraft)

Bombardier Canadair CL-215 (Fire-fighting Aircraft)

Air Tractor AT602 (Fire-fighting Aircraft)

Air Tractor AT802 (Fire-fighting Aircraft)

PZL-Mielec_M-18_Dromader (Agricultural Aircraft)

Antonov An-2 (Agricultural Aircraft)

G-164B Super B Turbine (Agricultural Aircraft)

Pac Cresco (Agricultural Aircraft)

CASA C-212 (Regional twin turboprop)

Saab 340B (Regional twin turboprop)

Sukhoi Su-80 (Regional twin turboprop)

Convair CV-240 (Regional twin turboprop)

Embraer EMB 110 Bandeirante (Regional twin turboprop)

Embraer EMB 120 Brasilia (Regional twin turboprop)

Handley Page Jetstream (Regional twin turboprop)

Grumman G-159 Gulfstream I (Regional twin turboprop)

CASA C-235 (Regional twin turboprop)

Antonov An-140 (Regional twin turboprop) Dornier 328 (Regional twin turboprop)

Properties that were investigated included:

Weights (takeoff, empty and payload weights)

Speed (maximum, cruse and stall speed)

Rate of climb

Range

Ceiling

Geometrical properties (wing area and wing span)

The full data set for these aircraft can be found in Appendix B.


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3.2.1 Empty Weight versus Takeoff WeightA technology diagram was created to determine the relationship between takeoff weight and empty

weight. The diagram is shown in Figure 15 below.

Figure 15 - Graph of Takeoff Weight versus Empty Weight for Statistically Analysed Aircraft

Three data sets were used to determine a relationship between takeoff weight and empty weight.

The data sets were chosen to match the desired aircraft demographic as closely as possible.

Sufficient data on fire-fighting aircraft were not available, so data on large agricultural aircraft and

regional twin turbo-prop aircraft were used to supplement the statistical analysis. All aircraft used a

turboprop engine for propulsion, and were all designed within the last thirty years. The relationship

between takeoff weight and empty weight is best described using a logarithmic equation. The outlier

(Bombardier Canadair CL-215) was not considered in the analysis. The following resulting

relationship was used as part of the matching diagram:


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3.2.2 Cruise SpeedThe technical task outlines a cruise speed of 375km/h (202 knots). Agricultural aircraft exhibit

substantially lower speeds than that required, whilst regional aircraft exhibit speeds higher than the

design requirement. The difference in trends between the three data sets shows that the statistical

analysis is attempting to define an aircraft that is not simply classified. The aircraft required by the

technical task has the roles of a fire-fighting aircraft, and operates similarly to an agricultural aircraft.

The aircraft is heavier than an agricultural aircraft, and lighter than a twin turboprop aircraft.

3.2.3 Stall SpeedThe aforementioned statistical analysis was used to determine an appropriate stall speed. For the

aircraft sized in Section 3.4, the stall speed from the statistical analysis was determined to be 82.5

knots.

3.2.4 Rate of ClimbThe rate of climb from the statistical analysis was determined to be 850 ft. This was influenced by

the Air Tractor AT-802F fire-fighting aircraft. As discussed in the technical task, FAR 25 requirements

dictate the minimum rate of climb as 300ft, which is much lower than the rate of climb from the

statistical analysis. The difference is due to the agility and manoeuvrability required in order to fightfires effectively.

3.2.5 Cruise AltitudeThe cruise altitude from the statistical analysis was based on the Air Tractor AT-802F, which was

deemed to have the same altitude requirements for fire fighting. The altitude from prototyping in

the statistical analysis was 14,000ft.

3.2.6

L/D EstimationData on L/D statistics are not readily available. For the statistical analysis, the L/D was calculated

from other statistics using the Breguet Range equation. Usage of this equation is likely to be

accurate to within 30%, due to the following assumptions:

The aircraft is cruising for the entire flight

The aircraft has a constant L/D at all times

The aircraft has a constant cruise speed at all times

The aircraft has a constant fuel consumption at all times


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From these assumptions, the L/D was calculated for each aircraft by using the following formula,

where CD is approximated to be 0.137 for each aircraft:

A mathematical model was made from this data, and the relation is as follows:

For the design weight, the L/D for cruise is 12.7. The L/D for loiter is 0.866(L/D cruise) (Raymer 2006).

Thus,


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3.3Mission ProfileThe following section outlines the mission profile and its associated requirements.

3.3.1 Mission Profile DiagramFigure 16 below diagrammatically illustrates the mission profile for the fire-fighting aircraft.

Figure 16 - Mission Profile

3.3.2 Mission Profile RequirementsThe phases of the mission profile and associated relevant details are given in Table 4.

Table 4 - Mission Profile Summary

Phase Details

1 Engine start and warm-up

2 Taxi

3 Takeoff

4 Climb Climb to 14 000 ft

5 Cruise 540 km (335.54 sm) at 375 km/h

6 Descent To assumed payload drop altitude of 70 ft

7 Loiter and Payload drop 20 minutes (E=0.33 hrs) at 1.3 Vstall

8 Climb Climb to 14 000 ft

9 Cruise 540 km (335.54 sm) at 375 km/h10 Descent To sea level

11 Landing, taxi and shut down


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3.4Weight EstimationThe takeoff weight and empty weight of the fire-fighting aircraft can be estimated from the mission

profile, the requirements of the technical task (Section 3.1) and the results of the statistical analysis

(Section 3.2). The requirements from each of these sections are summarised below.

3.4.1 Technical Task RequirementsThe technical task requirements are summarised below:

Payload: 4000 kg (8818.49 lbs)

Single pilot and baggage design weight: 86kg + 15kg = 101 kg

Cruise speed: 375 km/h = 341.7542 ft/s Radius: 540 km

Loiter time for payload drop: 20 minutes

3.4.2 Statistical Analysis RequirementsParameters that were not specified by the technical task were determined using a statistical analysis.

The values of some parameters were weight dependent. Hence, an iterative process was used to

determine the requirements. The results of the statistical analysis are presented below.

Stall speed, Vstall=82.5 knots = 139.2443 ft/s = 94.9393 sm/h

Cruise altitude, hcr = 14,000 ft

Technology diagram: A = -0.8126 and B = 1.2966

3.4.3 Remaining Sizing RequirementsSeveral parameters were not defined by the stages above, and were estimated from prototypes and

literature. Values for these parameters and the corresponding prototypes are shown in Table 5.

Table 5 - Parameters Estimated from Prototypes and Literature

Parameter Value Source

Rate of Climb 850 fpm = 14.167 ft/s Air Tractor 802F (Air Tractor 2007)

Propeller Efficiency 0.82 (Raymer 2006)

Cruise Power SFC 0.471 lbs/hp/hr (Honeywell 2009)

Loiter Power SFC 0.571 lbs/hp/hr cp(loiter) = 0.1 + cp(cruise) (Raymer 2006)

Reserve Fuel Fraction 0.06 (Roskam 2005)


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Unusable Fuel Fraction 0.005 (Roskam 2005)

3.4.4 Fuel Fraction EstimatesFuel fractions for phases 1-4, 6, 8, 10 and 11 were estimated using statistics for agricultural aircraft.

Fuel fractions for phases 5, 7 and 9 were calculated based on mission profile requirements. The

mission fuel fraction was then calculated from the individual phase fuel fractions. The results are

shown in Table 6 and the corresponding calculations in Appendix C. Whilst the start and finish

altitudes for the climb and decent of phases 4 and 10 differ from the altitudes for phases 8 and 6, it

is reasonable to assume that these phases have equivalent base fuel fractions as this difference in

small.

Table 6: Estimated Fuel Fractions (Roskam 2005)

Phase Fuel fractionEngine Start and Warm-Up (Phase 1)

Taxi (Phase 2)

Takeoff (Phase 3)

Climb (Phase 4)

Cruise (Phase 5)

Descent (Phase 6)

Loiter and Payload Drop (Phase 7)

Climb (Phase 8, Corrected for Payload

Drop)

Cruise (Phase 9)

Descent (Phase 10)

Landing, Taxi and Shutdown (Phase 11)

Mission Fuel Fraction

*Indicates a base value that must be corrected for payload drop at a later stage.


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3.4.5 Takeoff Weight EstimationThe takeoff weight of the aircraft is estimated from a takeoff weight component breakdown and the

technology diagram. This is achieved by solving Equation 2 and Equation 3 simultaneously for takeoff

weight.

Equation 2 - Takeoff Weight Component Breakdown

Equation 3 - Technology Diagram Equation for Takeoff and Empty Weight

Fuel weight is calculated as a percentage of takeoff weight, and consists of useable and trapped fuel.

Useable fuel consists of mission fuel and reserve fuel. The technical task stated no specific

requirements for trapped fuel or reserve fuel. Hence, conventional fuel fraction estimates of 0.005

and 0.06 respectively, were used. The fuel weight is calculated in Equation 4.

Equation 4 - Fuel Weight

Substituting Equation 4 into Equation 2 and rearranging for WTO gives Equation 5.

Equation 5 - Empty Weight Equation

Equation 5 and Equation 3 were solved graphically using Figure 17, resulting in a takeoff weight of

19,735.3 lbs and an empty weight of 8,697.9 lbs.


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Figure 17 - Takeoff and Empty Weight Estimate


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3.5Sensitivity AnalysisA sensitivity analysis provides information about the consequences of changing design parameters

on the aircraft takeoff weight. It is a useful tool for determining which parameters have the greatest

effect on the aircraft design. A sensitivity analysis also provides guidance on where to focus weight

reduction efforts. The sensitivity of takeoff weight was calculated to the following:

Payload weight

Crew weight

Empty weight

Power specific fuel consumption

Propeller efficiency

Lift to drag ratio

Range

Endurance

Loiter speed

Cruise speed

Sensitivity results are shown in Table 7 and the calculations are shown in Appendix D. Takeoff weight

has the greatest sensitivity to power specific fuel consumption, lift to drag ratio and propeller

efficiency during cruise. A reasonable change in power specific fuel consumption or propeller

efficiency of 0.01 can result in changes in takeoff weight of 29 lbs and 17 lbs respectively, whilst a

change in lift to drag ratio of one results in a 108 lbs change in takeoff weight. Large increases in

mission profile requirements (cruise radius and endurance) will also have a significant effect on the

takeoff weight of the aircraft.

Table 7: Sensitivity Analysis Results

Parameter Takeoff Weight Sensitivity

Payload 1.79 lbs/lbs

Crew 1.79 lbs/lbs

Empty weight 2.94 lbs/lbs

Cruise radius 4.10 lbs/sm

2924 lbs/lbs/hp/hr

during cruise -1680 lbs

(L/D)cruise -108 lbs

Endurance 532 lbs/hr

310 lbs

during loiter -216 lbs


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1.44 lbs/sm/hr

(L/D)loiter -12.1 lbs

3.6Aircraft SizingThe aircraft has a takeoff weight of 19,735 lbs and must be sized according to FAR25 requirements.

FAR25 includes requirements for takeoff, landing and climb phases of flight. The technical task

specifies a cruise speed requirement and the statistical analysis provides a reasonable stall speed. A

matching diagram method was used to ensure that all requirements were met simultaneously.

3.6.1 Sizing to Stall SpeedThe statistical analysis indicated that a stall speed of 82.5 knots ( 139 ft/s) is appropriate for a fire-

fighting aircraft of this size. Stall speed sizing was required for the clean configuration at cruise

altitude as this was the limiting case due to lower lift coefficients and air density. The aircraft was

sized to the stall speed requirement at cruise altitude using Equation 6.

Equation 6 - Stall Speed Equation.

3.6.2 Sizing to Takeoff DistanceTakeoff distance requirements for FAR25 state that the aircraft must clear a 35 ft obstacle at the end

of its takeoff field length. The technical task requires that the takeoff field length be less than or

equal to 4,000 ft. It is assumed that takeoff occurs at 1.1Vstall, and hence, a lower takeoff lift

coefficient is required as shown in Equation 7.

Equation 7 - Takeoff Lift Coefficient

The FAR25 takeoff parameter, shown in Equation 8, is used in to calculate the relationship between

wing loading and thrust loading as suggested by Roskam (2005). The appropriate conversion, seen in

Equation 9, between thrust and static shaft power can then be made to determine the power

loading. The relationship between wing loading and power loading for takeoff requirements is given

in Equation 10, and assumes takeoff at sea level.


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Equation 8 - FAR25 Takeoff Parameter

Equation 9 - Correction between Thrust and Static Power

Equation 10 - Limiting Relationship between Wing Loading and Power Loading for FAR25 Takeoff

3.6.3 Landing Distance SizingFAR25 landing requirement state that the aircraft must clear a 50 ft obstacle at the start of the

landing distance. It is desired that the aircraft be able to land with full payload and fuel. Hence, no

weight correction will be necessary to the wing loading or power loading. Statistical data is used to

size aircraft to FAR25 landing requirements. The approach speed (in knots) is related to the landing

field length by Equation 11.

Equation 11 - FAR25 Relationship between Approach Velocity and Landing Field Length

The stall speed in the landing configuration is given by , which gives the

limiting wing loading for landing in Equation 12.

Equation 12 - Limiting Wing Loading for Landing

3.6.4 Sizing to Climb RequirementsThe Air Tractor 802F, the prototype aircraft for this analysis, only requires a single turboprop engine.

This aircraft will be initially sized assuming a single engine. However, if the required power is in

excess of what can be provided by a single engine, the aircraft will be resized for two engines. A


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FAR25 aircraft with a single engine must only be sized to the FAR25.119 (AEO) climb gradient

requirement. The drag polar and corrected lift coefficient must be calculated for the FAR25.119

configuration and requirements.

3.6.5 Corrected Lift CoefficientFAR25.119 (AEO) required a speed of 1.3VSL, and hence, the corrected lift coefficient is given by

Equation 13.

Equation 13 - Corrected Lift Coefficient for FAR25.119 Requirements

3.6.6 Drag Polar EstimateThe drag polar is estimated from the wetted area ratio , equivalent skin friction coefficient

and the estimated effect of landing gear. The wetted area ratio of a fire-fighting aircraft of this size

will be similar to that of a Cessna Skylane. Hence, is a reasonable assumption

(Raymer 2006). The equivalent skin friction coefficient for this fire-fighting aircraft may be assumed

to be similar to a single engine lift aircraft, and hence, (Raymer 2006). Roskam (2005)

suggests that landing gear add an additional 0.015 0.025 to the zero-lift drag coefficient. Assuming

well-designed landing gear with fairings, is a reasonable estimate. It

was also assumed that approach flaps were equivalent to landing flaps with

. The zero-lift drag coefficient for the FAR25.119 (AEO) condition is

calculated in Equation 14. The drag polar is then given by Equation 16, where Oswalds efficiency

factor was calculated for the clean configuration in Equation 15 and landing flaps were assumed to

reduce Oswalds efficiency factor by 0.05.

Equation 14 - Zero-Lift Drag Coefficient for the FAR25.119 Configuration

Equation 15 - Oswald's Efficiency Factor for the Clean Configuration


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Equation 16 - FAR25.119 (AEO) Drag Polar

The FAR25.119 climb gradient requirement of 3.2% is met by Equation 18, where the climb gradient

parameter (CGRP) is given by Equation 17. The power loading must be corrected for temperature

and humidity effects. Roskam (2005) suggest that a correction factor of 0.85 is appropriate.

Equation 17 - Climb Gradient Parameter

Equation 18 - FAR25 Climb Gradient Limiting Relationship between Power and Wing Loading

3.6.7 Sizing to Cruise Speed RequirementsCruise speed sizing for propeller aircraft uses the power index, I

P. Roskam (2005, p. 163) suggests

that for a cruise speed of 375 km/h (233.0142 mph), a power index of I P=1.32 is required. The

density at cruise altitude, 0.001546 slugs/ft3, gives a density ratio of

. The limiting relationship between power loading and wing

loading is given by Equation 19. A correction factor of 0.7 was required to convert the cruise power

loading at cruise altitude to a takeoff sea level power loading (Roskam 2005).

Equation 19 - Limiting Relationship between Wing Loading and Power Loading for Cruise Requirements


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3.7Configuration SelectionFire-fighting aircraft can be classified by their payload capability, propulsion system and landing

system. Payload capacity for the aircraft was specified by the technical task as 8,820 lb. This

payload is heavier than that carried by agricultural or existing single engine turboprop aircraft.

However, the payload is much less than that carried by twin-engine aircraft. Consequently, both

configurations were investigated.

Common propulsion systems include jet, turboprop, piston or radial engine. Aircraft that use a jet

propulsion system are significantly faster than those powered by radial or piston engines. However,

large aircraft have reduced aerobatic capabilities and are hence, rarely used for fire-fighting aircraft.

Turboprop and piston engines are regularly used for fire-fighting aircraft. Both propulsion methods

are further investigated.

Possible landing configurations include seaplane (water only), amphibious (both water and land) and

normal landing (land only) arrangements. Seaplanes and amphibious aircraft offer significant speed

advantages for water refilling. However, Australia lacks the large still bodies of water required for

the refilling process. Hence, water landing capabilities are not seen as advantages. Furthermore,

both seaplanes and amphibious aircraft have reduced aerodynamic performance.


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