1 Team Beta Project Advisor: Dr. James D. Lang, Project Sponsor: Dr. Leland M. Nicolai Laura Brandt...

11

Team BetaTeam Beta

Project Advisor: Dr. James D. Lang,

Project Sponsor: Dr. Leland M. Nicolai

Laura Brandt ~ Team LeadMelissa Doyle ~ Propulsion

Luis Jimenez ~ AerodynamicsNic Seid ~ Structures

L O C K H E E D M A R T I N

22

OutlineOutline

Design ProcessDesign Process

RequirementsRequirements

Mission ProfileMission Profile

Initial ConceptsInitial Concepts

Aircraft ConfigurationAircraft Configuration

WeightWeight

AerodynamicsAerodynamics

PropulsionPropulsion

Stability Stability

Materials and StructuresMaterials and Structures

Cost AnalysisCost Analysis

Final CommentsFinal Comments

33

ScheduleScheduleMarch April May June

1 2 3 4 1 2 3 4 5 6 7 8 9 10 11

Requirements Defined

Schedule and Cost

Historical Research

Initial Weight Estimates

Aircraft Concepts

Initial Sizing

Initial Stability/ Control

Trade Studies

Graphical Modeling

Initial Layout (Location of Internal

Components)

Final Layout (CG Calculations)

Presentation

Initial ScheduleReal Schedule

44

Design ProcessDesign ProcessOutlined Requirements

Assumed Take off Weight

Initial Concepts Selected Concept

Assumed 2 planes for minimum LCC

Sized Fuselage around payload

Airfoil Design and Selection

Assumed W/S, AR, b, engine

Mission Profile

Calculated Component Weights and built up Empty Weight (Wempty),Sized Tail, Performance Calculations, and Aerodynamic Analysis

Fuel Fractions Calculate Empty Weight and compare to Wempty

Trade Studies

Number and Type of Engines

BCM and BCA

Best Patrol Altitude and Mach

Wing Span and Aspect Ratio

Calculate Fuel Required

Evaluate Design

55

RequirementsRequirementsMissionMission

• High Altitude, Long Endurance, Standoff ISR

• Unmanned/autonomous With Man-in-the-loop

• Continuous Coverage (24/7)

• Loiter/ISR Altitude: 50,000 – 65,000 Feet

• Operate From B-52/B-2 Bases

• Mission Radius = 3000 nm

• Low signature side sector

• 50 nm standoff racetrack flight path

66

RequirementsRequirementsSystemsSystems

• 360° Continuous Coverage With GMTI, AMTI and FOPEN

4 X Band Radar Antenna (4.5’x4.5’)

2 UHF Radar Antenna (4’x46’)

• Defensive Payload• On-board jammer (ULQ-21)

• Towed decoy (12 ALE-50 decoys)

• Flares (ALE-40 with 360 flares)

• Threat warning system (RF/EO/IR) part of sensor suite

77

ITEM WEIGHT POWER COOLING (lb) (watts) (watts)

EW/ULQ-21 117 3000 2525

Towed decoy/ ALE-50 193 470 -• 12 decoy unit plus installation • Deploy 2 per RF missile

Flares/ ALE-40 828 80 -

• 12 flare buckets with 360 flares plus installation

• Deploy 60 per IR missile

Total 1138 3550 2525(all air cooled)

Missile Warning System (MWS) and DIRCM part of EO/IR sensor suiteRadar Warning system (RWR) part of ELINT sensor system

RequirementsRequirementsDefensive Payload SystemsDefensive Payload Systems

88


X Band Radar 2227 63200• 4 antennas @ 4.5’x4.5’ 9500 (L)/

located in fuselage, broadside 44000 (A)/

UHF Radar 1900 67500• 2 antennas @ 4’x46’ 25300 (L)

located in fuselage, broadside

(includes FOPEN)

EO/IR 30 600 0

SIGINT 918 1200 1200 (L)

Shared Processor 150 770 770 (L)

Installation (10%) 523 - -

Total (All systems on) 5748 133270 80770

RequirementsRequirementsSensorSensor

99


Vehicle Management System 700 538• Autopilot 203 • Air Data System 45• Radar altimeter 14• Flight Termination System 38• IFF 78

Communications 3100 2125• VHF/UHF (Link 16 & SATCOM) 44• HF/Link 22 130• CDL/Ka SATCOM 18

Installation (15%) 86

Total 655 3800 2663(all air cooled)

RequirementsRequirementsFlight AvionicsFlight Avionics

1010

ITEM WEIGHT POWER COOLING

(lb) (watts) (watts)

Surface Control TBD 500 0

Instruments 137 50 0

Landing Gear TBD 550 0

ECS 414 140 140

Engine TBD• GCCU (engine generator) 2300 2300• Digital Fuel Control 450 0

Fuel System 865 400 0

Electrical System & Control 874 2000 2000• Secondary Power (battery controller) 350 350

Hydraulic 138 50 0

APU TBD TBD TBD

RequirementsRequirementsAirframe SystemAirframe System

1111

ITEM WEIGHT POWER COOLING (lb) (kW) (kW)

Defensive Payload Systems 1138 3.6 2.5

Sensors 5748 133.3 80.8

Flight Avionics 655 3.82.7

Miscellaneous Airframe Systems 2428* 6.8** 4.8**

Total 9969 147.4 90.7

*Control Surfaces, Engine, Landing Gear, and APU not included**APU not included

RequirementsRequirementsWeight/Power/Cooling TotalsWeight/Power/Cooling Totals

1212

Radar antennas:Radar antennas:

Radar electronics:Radar electronics:

EO/IR:EO/IR:

SIGINT:SIGINT:

Central HPPS processor for radar, EO/IR and SIGINT: 4 cubic feetCentral HPPS processor for radar, EO/IR and SIGINT: 4 cubic feet

Communications:Communications:

ALE-50 Towed RF decoy:ALE-50 Towed RF decoy:

ALE-40 chaff/flare dispenser: ALE-40 chaff/flare dispenser: 10.6”x8.5”x6.7” each dispenser carries 30 flares10.6”x8.5”x6.7” each dispenser carries 30 flares

ECM/ULQ-21: ECM/ULQ-21: 1671 cubic inches1671 cubic inches

RequirementsRequirementsSensor, Communications and ECM/Decoy Sensor, Communications and ECM/Decoy

Dimensions/VolumesDimensions/Volumes• X Band: 4 @ 4.5’x4.5’x1’X Band: 4 @ 4.5’x4.5’x1’• UHF: 2 @ 4’x46’x1’UHF: 2 @ 4’x46’x1’

• X Band: REX and front end processor 1240 cubic inchesX Band: REX and front end processor 1240 cubic inches• UHF: REX and front end processor 1720 cubic inchesUHF: REX and front end processor 1720 cubic inches

• Sensor apertures: 5 @ 3” diameter x 8” depthSensor apertures: 5 @ 3” diameter x 8” depth• Electronics: 420 cubic inchesElectronics: 420 cubic inches

• Antennas: 8 blade antennas 4’x1’x1.25” (retract when not in use)Antennas: 8 blade antennas 4’x1’x1.25” (retract when not in use)• Receivers: 808 cubic inchesReceivers: 808 cubic inches

• VHF/UHF (Link 16, voice, SATCOM): 12.5”x10”x7.5”VHF/UHF (Link 16, voice, SATCOM): 12.5”x10”x7.5”• Microwave (CDL/SATCOM): 15”x6”x5”Microwave (CDL/SATCOM): 15”x6”x5”• HF: RX/EX: 25”x25”x10”, PA: 20”x10”x10”, Coupler: 20”x6”x6”HF: RX/EX: 25”x25”x10”, PA: 20”x10”x10”, Coupler: 20”x6”x6”

• Electronics: one box @ 7.5”x5.7”x14”Electronics: one box @ 7.5”x5.7”x14”• Each decoy canister: 2.74”x2.74”x20.25”Each decoy canister: 2.74”x2.74”x20.25”

1313

Mission EnduranceMission EnduranceTime on Station (TOS)

2*(Transit Time) + Turnaround Time = 20 hours

Minimum Endurance Time

2*(Transit Time) + Time on Station = 36 hours

Turnaround Time = 4 hours Total Transit Time = 16 hours (Assumed based on U-2) (Calculated based on Mach .65 at 50,000 ft)

Transit TOS Turnaround

1st AircraftTOSEndurance

2nd Aircraft

1414

Patrick

Lakenheath

UAE

KadenaCyprus

Mission Theater Mission Theater

Minimum time on station = 20 hoursMinimum number of planes per station = 2 planesRequire 4 planes in reserve at Patrick, Lakenheath, UAE, and KadenaTotal number of planes = 14 planes

1515

Mission ProfileMission Profile

Takeoff

Climb to Altitude

Loiter 20 hr

Cruise Out

8 hrDescend to Sea Level

Sea Level Loiter for 30 min

3,000 NM

3,000 NM

Altitude

50,000 ft

Cruise Back8 hr

Distance

1616

Radar GeometryRadar Geometry

X Band Radar (4) : 4.5’ x 4.5’

UHF Radar (2) : 4’ x 46’

Field of Regard is +/- 70 degrees in the vertical and horizontal

1717

Radar GeometryRadar Geometry

Angle Antennas are to be placed: 1.85 degrees

Altitude

LOS

5 degrees

GA

Standoff Distance

Horizon

Altitude 50,000 ft (9.052Nm) LOS = 250 NmGrazing Angle = 5.2 degrees Standoff Distance = 100 Nm

1818

Initial ConceptsInitial ConceptsDesign ADesign A

UHF Antenna

UHF Antenna

MGMG

NGX-Band Antenna

Engine

Inlet

1919

Initial ConceptsInitial ConceptsDesign BDesign B

UHF Antenna

MG

Inlet

X-Band Antenna

Engine

InletX-Band Antenna

X-Band Antenn

a

UHF Antenn

a

X-Band Antenn

a

Engine

2020

Initial ConceptsInitial ConceptsDesign CDesign C

UHF Antenna

UHF Antenna

Avionics

Nozzle Engine Inlet

X-Band Antenna

LGLG

2121

Initial ConceptsInitial Concepts

Concept AConcept A Concept BConcept B Concept CConcept C

DragDrag

Mission Mission DurationDuration

Payload Payload PlacementPlacement

ControlControl

A modified version of Concept B was Selected

Exceed Compliance Satisfactory Compliance Short of Compliance

2222


TOGW = 50,900 lb

W/S = 60

Span = 184 ft

MAC = 4.89 ft

Wing Area = 850 ft^2

Aspect Ratio = 40

Wing Sweep = 0 deg

Airfoil: Custom Airfoil (Modified CAST 10-2/DOA 2 Transonic Airfoil)60% Laminar Flow

Engine: Two AE3007 Turbofan

Fuselage Length = 60 ft

Fuselage Height = 6 ft

Fuselage Mean Cross Section Area = 32.5 ft2

T/W = .32

2323


2424

Aircraft ConfigurationAircraft Configuration3-View Drawing3-View Drawing

184 ft.

60 ft.

6 ft.

6.6 ft.

2.83 ft.

14.7 ft.

5.9 ft.

3.5 ft.

2525

Aircraft ConfigurationAircraft ConfigurationInternal ComponentsInternal Components

APU

Avionics

X-Band

Rear Landing Gear

UHFAntenna

Front Landing Gear

X-Band Antenna

Inlet

AE-3007’s

FuselageFuel Cell

V-Tail

Bicycle Landing Gear

C.G.

Aerodynamic Center

2626

WeightWeight

Take-Off Weight

Empty Weight

Fuel Weight

Fuel Fraction

Fuel Volume

50,900 lb*

26,800 lb

24,100 lb

0.47

3,500 gal

*Takeoff weight is estimated to be within 500 lb of Actual Gross weight

2727

Weight FractionsWeight Fractions

1. Start up/Take-Off

2. Climb to Cruise Alt

3. Cruise Out

4. On Station

5. Cruise Back

6. Descend to Sea Level

7. Loiter 30 min

0.970

0.994

0.890

0.752

0.883

0.993

0.993

Best Cruise Altitude = 50,000 ftBest Cruise Mach = .65Best Loiter Altitude = 50,000 ft Best Loiter Mach = .62

2828


Aspect Ratio = 40

Span = 184 ft

Reference Area = 850 ft2

Taper Ratio = .4

MAC = 4.89 ft

Croot = 6.6 ft

Ctip = 2.83 ft

Wing Sweep = 0 deg

Airfoil: Custom Airfoil (Modified CAST 10-2/DOA 2 Transonic Airfoil)

60% Laminar Flow

t/c = .15

e = .63 *

K = 0.013

CLmax = 1.2

* Estimated ‘e’ using Wing Efficiency vs Aspect Ratio Graph provided by Lockheed Martin

2929

AerodynamicsAerodynamicsAt 50,000 ft and M=.62At 50,000 ft and M=.62

L/Dmax = 50.58 @ CL=.8

CLmax = 1.2

Zero Lift Drag = 0.008

Clalpha = 10.4 rad-1 = .18 deg-1 at Mach .62

Vstall = 479 ft/s (at half fuel weight)

3030

CLalpha v Mach

0.11

0.13

0.15

0.17

0.19

0.21

0.23

0.25

0.27

0.29

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

Mach

CL

alp

ha

(deg

^-1

)AerodynamicsAerodynamics

3131

Sears HaackSears Haack

3232

AerodynamicsAerodynamics50,000 ft at Mach .6250,000 ft at Mach .62

L/D versus CL

0

10

20

30

40

50

60

0 0.4 0.8 1.2 1.6 2 2.4

CL

L/D

Begin Loiter at CL = .8

3333

AerodynamicsAerodynamics50,000 ft at Mach .6550,000 ft at Mach .65

Drag PolarMach 0.55

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0.000 0.010 0.020 0.030 0.040 0.050

CD

CL

L/Dmax = 50.58at CL of .8

3434

AerodynamicsAerodynamicsDrag Divergent Mach NumberDrag Divergent Mach Number

Mdd = Mdd(L=0)*LFdd – 0.05*ClMdd = Mdd(L=0)*LFdd – 0.05*ClMdd is higher than loiter and range mach numbers.Mdd is higher than loiter and range mach numbers.Mdd(L=0) was taken from Boeing chart for uncambered airfoils, which Mdd(L=0) was taken from Boeing chart for uncambered airfoils, which typically resulted in Mdd of about 0.08 mach above critical mach number.typically resulted in Mdd of about 0.08 mach above critical mach number.LFdd taken from plot intended to adjust Mdd to actual lift coefficient.LFdd taken from plot intended to adjust Mdd to actual lift coefficient.

Drag Divergent Mach

0.640.650.660.670.680.690.7

0.710.72

-4 -2 0 2 4 6 8 10

AOA

Md

d

(deg)

3535


60% laminar flow airfoil was desired 60% laminar flow airfoil was desired

Computer programs such as Profili and Computer programs such as Profili and XFOIL were used in order to simulate XFOIL were used in order to simulate fluid flow across different airfoils.fluid flow across different airfoils.

Laminar separation points and Pressure Laminar separation points and Pressure coefficients were obtained using coefficients were obtained using various lift coefficients.various lift coefficients.

3636

XFOIL was used to determine the Separation BubbleXFOIL was used to determine the Separation Bubble

Profili and Excel were used to modify different Profili and Excel were used to modify different characteristics of airfoils such as camber, leading characteristics of airfoils such as camber, leading edge radius, and percent thickness.edge radius, and percent thickness.

A trade study involving how thickness affects the range A trade study involving how thickness affects the range of laminar flow along the airfoil, was done through of laminar flow along the airfoil, was done through viscous operations on XFOILviscous operations on XFOIL


3737

Final AirfoilFinal AirfoilModified CAST 10-2/DOA transonic airfoil Modified CAST 10-2/DOA transonic airfoil

SpecificationsSpecificationsMax. Thickness 14.98%Max. Thickness 14.98%

Max. Thickness at 45.6% of ChordMax. Thickness at 45.6% of ChordMax. Camber 1.90%Max. Camber 1.90%

Max. Camber at 69.5% of ChordMax. Camber at 69.5% of ChordLaminar Separation at 60.34% of chordLaminar Separation at 60.34% of chord


3838Graph using a Lift Coefficient of 0.8Graph using a Lift Coefficient of 0.8

Velocity Profile for modified CAST 10-2/DOA 2 transonic airfoil

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 0.2 0.4 0.6 0.8 1 1.2

X (%)

U/U

inf.


3939

AerodynamicsAerodynamicsKarman-Millikan MethodKarman-Millikan Method

Alternate method of determining laminar separation.Alternate method of determining laminar separation.

Relates pressure coefficients from XFOIL to velocity Relates pressure coefficients from XFOIL to velocity profiles.profiles.

CCpp = 1 – (U/U = 1 – (U/Uoo))22

Iteration of velocity gradient with pressure Iteration of velocity gradient with pressure coefficients resulted in separation point using plot coefficients resulted in separation point using plot of Uof Uss/U/Uoo versus F. versus F.

4040

Used to verify laminar separation through velocity Used to verify laminar separation through velocity profile.profile.

U/UU/Uo o = 1 for 0 < x/x= 1 for 0 < x/xe e < 1< 1

U/UU/Uo o = 1 + F * ((x-x = 1 + F * ((x-xee)/x)/xee) for 1< x/x) for 1< x/xee

x = distance along surface from leading edge.x = distance along surface from leading edge.

xxee = length from leading edge to center of pressure. = length from leading edge to center of pressure.

U = velocity outside boundary layer.U = velocity outside boundary layer.

UUoo = maximum velocity. = maximum velocity.

F = velocity gradient.F = velocity gradient.


4141

UUss = velocity at laminar separation point. = velocity at laminar separation point.

Laminar Separation

y = -6.2868x3 - 4.2885x2 - 1.0802x + 0.8998

0.88

0.9

0.92

0.94

0.96

0.98

1

-0.3-0.25-0.2-0.15-0.1-0.050

Velocity Gradient F

Ve

loc

ity

Ra

tio

Us

/Uo


4242

Trade study involving same airfoil and camber as a Trade study involving same airfoil and camber as a control basis, then thickness as a variation.control basis, then thickness as a variation.

Percent thickness versus laminar flow was plotted Percent thickness versus laminar flow was plotted to depict a trend or tendency.to depict a trend or tendency.

Plot shows ideal thickness of about 15% in order to Plot shows ideal thickness of about 15% in order to achieve a relative maximum laminar flow of achieve a relative maximum laminar flow of about 60% (for CAST 10-2/DOA 2 transonic about 60% (for CAST 10-2/DOA 2 transonic airfoil).airfoil).

AerodynamicsAerodynamicsThickness to Chord VariationThickness to Chord Variation

4343


Airfoil thickness trade studyAirfoil thickness trade study

t/c selected was .15t/c selected was .15

CAST 10-2/DOA 2 transonic airfoil (cl 0.80)

0

10

20

30

40

50

60

70

13 14 15 16 17 18 19

% Thickness

% C

ho

rd L

eng

th o

f L

amin

ar F

low

4444

AerodynamicsAerodynamicsCp DistributionCp Distribution

4545

AerodynamicsAerodynamicsCdo at half fuel weightCdo at half fuel weight

Cdo v Mach

0.0090

0.0095

0.0100

0.0105

0.0110

0.0115

0.0120

0.0125

0.4 0.45 0.5 0.55 0.6 0.65 0.7 0.75 0.8

Mach

Cd

o

50,000 ft 55,000 ft 60,000 ft 65,000 ft

4646

AerodynamicsAerodynamicsfor Sfor Srefref = 850 ft = 850 ft22

Aircrafts wing span is 184 ft with an aspect ratio of 40

4747

Cruise for different Altitudes and Mach

Two AE

One F118

5.50E+07

7.50E+07

9.50E+07

1.15E+08

1.35E+08

1.55E+08

1.75E+08

0.55 0.575 0.6 0.625 0.65

Mach

(L/D

)*(V

/tsf

c)

AE 50,000 ft AE 55,000 ft AE 60,000 ft AE 65,000 ftF118 55,000 ft F118 60,000 ft F118 65,000 ft


BCA 50,000 ft at BCM 0.65

i

i

W

W

D

L

tsfc

VR 1ln

4848

Loiter for different Altitudes and Mach

TwoAE

30

35

40

45

50

55

60

65

70

0.55 0.575 0.6 0.625 0.65Mach

(L/D

)/ts

fc

AE 50,000 ft AE 55,000 ft AE 60,000 ft AE 65,000 ft

F118 55,000 ft F118 60,000 ft F118 65,000 ft

OneF118


Best Loiter Altitude 50,000 ft at Best Loiter Mach 0.62

i

i

W

W

D

L

tsfcE 1ln

1

4949

PerformancePerformance

Stall Velocity at Sea Level 205 ft/s

Take-Off Velocity 225 ft/s

Take-Off Distance 3995 ft

Landing Distance 3770 ft

5050

Limit Load FactorLimit Load Factor 3.53.5

Ultimate Negative Load FactorUltimate Negative Load Factor 3.853.85

Ultimate Positive Load FactorUltimate Positive Load Factor5.255.25

Sustained Turn Load Factor Sustained Turn Load Factor 3.523.52

Sustained Turn Rate Sustained Turn Rate 9.84 deg/s9.84 deg/s

Dynamic Pressure Limit Dynamic Pressure Limit 99.6 99.6 lb/ftlb/ft22

PerformancePerformanceRobust Structural Design - Taking Into Account Gust Robust Structural Design - Taking Into Account Gust

LoadingLoading

5151

PerformancePerformanceRobust Structural Design Robust Structural Design

Taking Into Account Gust LoadingTaking Into Account Gust Loading

V-n Diagram

High Limit Load Factor

Lower Limit Load Factor

Ultimate Load Factor

Ultimate Load Factor

Vcruise

Positive High AOA

-5

-3

-1

1

3

5

7

0 100 200 300 400 500 600 700 800

Velocity (ft/s)

Lo

ad

Fa

cto

r "n

"

Structural Failure

Structural Failure

Structural Damage

Structural Damage

Negative High AOA

Positive Low AOA

Negative Low AOA

5252

PerformancePerformanceFerryFerry

Takeoff Climb to Altitude

Cruise Out Descend to Sea Level

Sea Level Loiter for 30 min

16,000 NM

Altitude

50,000 ft

Distance

Best Cruise Altitude = 50,000 ftBest Cruise Mach = .65

Ferry Distance = 16,000 NM Endurance = 43 hr

5353

PerformancePerformance Two Allison AE3007 Turbofan EngineTwo Allison AE3007 Turbofan Engine

Climbing from Sea Level to BCA for BCM Fuel to Climb = 960 lb

Climb TrajectoryFuel Specific Energy

fs = 0

fs = 25

fs = 50

fs = 60fs = 70

he 60,000

he 50,000

he 40,000

he 30,000

he 20,000

he 10,000he 5,000

0

10000

20000

30000

40000

50000

60000

70000

0.000 0.100 0.200 0.300 0.400 0.500 0.600 0.700

Mach

Alt

itu

de

(ft)

5454

PropulsionPropulsionAllison AE3007 Turbofan EngineAllison AE3007 Turbofan Engine

Bypass Ratio 5.3

Overall Pressure Ratio 23

Thrust (Static Sea Level) 8813 lb

Thrust (at 50,000 feet) 1062 lb

Dimensions:

Length 9.59 ft

Fan Diameter 3.21 ft

Max Diameter 3.43 ft

Engine Weight 1588 lb

System Weight 3084 lb (One engine)

5458 lb (Two engines)

Max. Power Extraction 26 kW (52 kW for two engines)

* assuming an 87% efficiency extraction *

5555

PropulsionPropulsionF118-GE-101 EngineF118-GE-101 Engine

Bypass Ratio 0.72

Overall Pressure Ratio 32.8

Thrust (Static Sea Level) 15940 lb

Thrust (at 50,000 feet) 2576 lb

Dimensions:

Length 8.38 ft

Fan Diameter 3.17 ft

Max Diameter 3.92 ft

Engine Weight 3150 lb

System Weight 5671 lb (One engine)

10038 lb (Two engines)

Max. Power Extraction 45 kW (90 kW for two engines)

* assuming an 87% efficiency extraction *

5656

AE3007 (per an engine)AE3007 (per an engine)

Pitot Inlet Size: 10.64 ftPitot Inlet Size: 10.64 ft2 2

Fixed Convergent Nozzle: 4.01 ftFixed Convergent Nozzle: 4.01 ft22

F118-GE-101 (per an engine)F118-GE-101 (per an engine)

Pitot Inlet Size: 10.36 ftPitot Inlet Size: 10.36 ft2 2

Fixed Convergent Nozzle: 3.91 ftFixed Convergent Nozzle: 3.91 ft22

PropulsionPropulsionInlet and Nozzle Size*Inlet and Nozzle Size*

*Based on Historical Data

5757

PropulsionPropulsionThrust vs Mach at Various AltitudesThrust vs Mach at Various Altitudes

Thrust vs Mach for Both the AE3007 and F118-GE-101 Engines

0500

1000150020002500

0.55 0.6 0.65 0.7 0.75

Mach Number

Thru

st [l

b]

AE3007 at 55K feet AE3007 at 60K feet AE3007 at 65K feet AE3007 at 70K feet

F118-GE-101 at 55K feet F118-GE-101 at 60K feet F118-GE-101 at 65K feet F118-GE-101 at 70K feet

5858

PropulsionPropulsionSFC vs Mach at Various AltitudesSFC vs Mach at Various Altitudes

SFC vs Mach for Both the AE3007 and F118-GE-101 Engines

0.6

0.7

0.8

0.9

1

1.1

1.2

0.55 0.6 0.65 0.7 0.75Mach Number

SF

C [

1/h

r]

AE3007 at 55K feet AE3007 at 60K feet AE3007 at 65K feet AE3007 at 70K feet

F118-GE-101 at 55K feet F118-GE-101 at 60K feet F118-GE-101 at 65K feet F118-GE-101 at 70K feet

F118-GE-101 at 50K feet AE3007 at 50K feet

5959

PropulsionPropulsionThrust vs. Altitude at Mach 0.6

for (2) AE3007H Engines and (1) F118-GE-101 Engine

50,000

52,000

54,000

56,000

58,000

60,000

62,000

64,000

66,000

68,000

70,000

500 700 900 1100 1300 1500 1700 1900 2100

Thrust [lbf]

Alt

itu

de

[ft]

(2) AE3007H engines (1) F118-GE-101 engine

6060

PropulsionPropulsionThrust vs. SFC at Mach 0.6

for (2) AE3007 and (1) F118-G-101 Engines

0.6

0.8

1

1.2

1.4

1.6

0 200 400 600 800 1000 1200 1400 1600 1800 2000Thrust [lbf]

SF

C [

1/h

r]

(2) AE3007 at 55K feet (1) F118-GE-101 at 55K feet (2) AE3007 at 60K feet(2) AE3007 at 65K feet (2) AE3007 at 70K feet (1) F118-GE-101 at 60K feet(1) F118-GE-101 at 65K feet (1) F118-GE-101 at 70K feet

6161

ITEM POWER COOLING (kW) (kW)

Defensive Payload Systems 3.6 2.5

Sensors 133.3 80.8

Flight Avionics 3.8 2.7

Miscellaneous Airframe Systems 6.8** 4.8**

Total 147.4 90.7

Total Wattage = 238 kW

**APU not included

PropulsionPropulsionRequired Power and CoolingRequired Power and Cooling

6262

PropulsionPropulsionPower and CoolingPower and Cooling

Max Power ExtractionMax Power Extraction

No. of enginesNo. of engines 11 22 33

AE3007AE3007 2626 5252 7878 kWattskWatts

F118-GE-101F118-GE-101 4545 9090 135135 kWattskWatts

Power Required From APU for Powering an Power Required From APU for Powering an Cooling SubsystemsCooling Subsystems

No. of enginesNo. of engines 11 22 33

AE3007AE3007 212212 186186 160160 kWattskWatts

F118-GE-101F118-GE-101 193193 148148 103103 kWattskWatts

6363

Auxiliary Power Unit (APU)Auxiliary Power Unit (APU)Turbocharged 310 hp Continental TSIO 520-LTurbocharged 310 hp Continental TSIO 520-L

Total Weight535 lbAPU Fuel Weight 450 lbTotal Weight985 lb

Height 1.67 ftWidth 2.8 ftLength 4.22 ft

6464

Auxiliary Power Unit (APU)Auxiliary Power Unit (APU)Turbocharged 310 hp Continental TSIO 520-LTurbocharged 310 hp Continental TSIO 520-L

Using Two AE3007sUsing Two AE3007s

Required Total Power of Subsystems 147.4 kW

Required Total Cooling Power of Subsystems 90.7 kW

Assuming 87 % Efficiency 102.5 kW

Total Power and Cooling for Subsystems 249.9 kW

Power Extracted from Engines 52 kW

Additional Power Required from APU 197.9 kW

APU Provides 231.2 kW

Available Excess Power 33.3 kW

6565

Air Intake

(Bleed Air From Inlet)

Coolant SupplyCoolant Return

Air Exhaust to Engine Compartment

Heat Exchanger

X-Band 2.4 kW Each

(9.5 kW Total)

Pump Cooled w/

Internal Coolant

(Assume 87% efficiency)

Heat Exchanger

47 kW Total Cooling

Assuming 87% EfficiencyAPU and

Turbochargers Cooling

From APU and Turbochargers

Cooling

Cooling SystemLoiter (Liquid Cooling of System)Loiter (Liquid Cooling of System)

UHF 12.7 kW Each (25.3 kW Total)

ECS(0.14 kW)

GCCU(2.3 kW)

SIGINT & Shared

Processor 2 kW Total

Electrical System & Control(2 kW)

Battery Controller(0.35 kW)

6666

Cooling SystemCooling SystemLoiter (Liquid Cooling of APU and Turbochargers)Loiter (Liquid Cooling of APU and Turbochargers)

APU Air Intake

(Bleed Air from Inlet)

APU

Turbocharger (Boost Pressure to 3.75 psi)

Turbocharger (Boost Pressure to 14.7 psi)

Exhaust

Heat Exchanger

Heat Exchanger

Coolant Supply

Coolant Return

Power from

Engine

Power from

Engine

6767

Cooling SystemLoiter (Air Cooling)Loiter (Air Cooling)

Air Intake(Bleed Air From Inlets)

Air Cooling System

55.5 kW Total Cooling

X-Band

(44 kW Total)Defensive Payload

(EW/ULQ-21)(2.5 kW)

Vehicle ManagementSystem

(0.54 kW Total)

CommunicationSystems

(2.1 kW Total)

Air Exhaust to Engine Compartment

Pump (Assume an 87% efficiency)

6868

Cooling SystemLoiter (Using Two AE3007s)Loiter (Using Two AE3007s)

Total Liquid Cooling 47 kWTotal Air Cooling 55.5 kWTotal Overall Cooling 102.5 kW

Available Cooling APU Provides83.8 kWPower Extracted from Engines52 kWTotal Power Available for Cooling 135.8 kW

Excess Cooling Power 33.3 kW

6969

Power and Cooling Cruise (Using Two AE3007s)Cruise (Using Two AE3007s)

Total Power 14.14 kWTotal Cooling9.98 kW

Defensive PayloadFlight AvionicsAirframe System

Total Overall Power and Cooling24.1 kW

Available Extracted Power from Engine52 kWExcess Power27.9 kW

**APU not required during Cruise

7070

Fuselage 1,170 lb

Wing 7,316 lb

Surface Controls 432 lb

Tail 384 lb

Landing Gear 1,535 lb

Propulsion System 5,458 lb

Flight Avionics 655 lb

Fuel System/Tanks 865 lb

ECS 414 lb

Hydraulic System 138 lb

Electrical System 874 lb

Defensive Payload 1,138 lb

Take-Off Weight 50,800 lb

Empty Weight 26,800 lb

Fuel Weight 24,000 lb

Fuel Fraction 0.47

Fuel Volume 3,500 gal

Stability and Control Stability and Control Component WeightsComponent Weights

Sensor Payload (installed) 5,748 lb

APU 535 lb

Paint 121 lb

7171

Stability and ControlStability and ControlCenter of Gravity Location at TakeoffCenter of Gravity Location at Takeoff

Aircraft Weight Distance from Moment about

Component (lb) Nose (ft) Nose (ft-lb)

Fuselage 1170 25 35100

Payload/Subsystems/Engine 17930 32.64 585275

Wing 7316 33.1 242159.6

Tail 384 52 19968

Fuel Tanks in Wings 14861 32.8 491905

Fuel Tanks in Fuselage 7431 32.6 242237

Gross Weight Moment

50873 lb 1616644 ft-lb

CG LocationFrom Nose of Aircraft Static Margin

32.93 ft 0.014

Aerodynamic Center

33.1

7272

Fuel Tanks

Main Spars

Stability and Control Stability and Control Fuel TanksFuel Tanks

7373

Stability and ControlStability and ControlCenter of GravityCenter of Gravity

Center of Gravity Travel

Xac

0%

20%

40%

60%

80%

100%

32.80 32.90 33.00 33.10

CG Location (ft)

% F

ue

l

100% 80% 60% 40% 20% 5%

% Fuel in Wings 100% 70% 70% 35% 40% 10%

% Fuel Fuselage 100% 90% 50% 45% 0% 0%

100% 80% 60% 40% 20% 5%

CG Location 32.931 32.920 32.943 32.923 32.960 32.938

SM 0.014 0.016 0.012 0.016 0.008 0.013

7474

Stability and ControlStability and Control

Ailerons

Span = 50 ft

Area = 42.5 ft2

MAC = .85 ft

Aileron Chord: 16 % of Wing MAC

Aileron Span: 25 % of Wing Span

Flaps – not required The runway distance requirements are met without the assistance of flaps

7575

Stability and ControlStability and ControlV-Tail at 45 degreesV-Tail at 45 degrees

with .6 taper ratiowith .6 taper ratio

Vertical Projection

Svt = 71 ft2

Cvt = 0.0145**

Height = 14.7 ft

Mean Cord = 5.9

Horizontal Projection

Sht = 142 ft2

Cht = 0.34**

Length = 29.4 ft

Mean Cord = 5.9

Rudder/Elevator*

Area = 22.69 ft2

MAC = 1.72 ft

Span = 13.23 ft

*Values for rudder/elevator are for one half of the V-tail only.

**Historical Data

7676

Materials and StructureMaterials and StructureMaterial SelectionMaterial Selection

Carbon Fiber

Primary and Secondary Structures

Wings

Control Surfaces

Fiberglass

Fuselage

Steel

Landing Gear

7777

Materials and StructureMaterials and StructureInternal Fuselage StructuresInternal Fuselage Structures

7878

Materials and StructureMaterials and StructureWing Load Distribution (G airload of 4)Wing Load Distribution (G airload of 4)

Wing Loading (Schrenk's Approximation)

050

100150200

250300350400

0 10 20 30 40 50 60 70 80 90

Chord (ft)

Lo

ad (

lb/f

t^2)

Root Tip

7979

Materials and StructureMaterials and StructureShear Distribution Shear Distribution (G airload of 4)(G airload of 4)

Shear Load Distribution

0.0E+00

5.0E+04

1.0E+05

1.5E+05

2.0E+05

0 10 20 30 40 50 60 70 80 90

Chord (ft)

Sh

ea

r (l

b)

Root Tip

8080

Materials and StructureMaterials and StructureMoment Distribution Moment Distribution (G airload of 4)(G airload of 4)

Moment Distribution

-2.E+07

-2.E+07

-1.E+07

-5.E+06

0.E+00

0 10 20 30 40 50 60 70 80 90

Chord (ft)

Mo

men

t (f

t-lb

)

Root Tip

8181

Materials and StructureMaterials and Structure Moments of Inertia**Moments of Inertia**

Ixx = 25.0E6 slug*ft2

Iyy = 5.95E6 slug*ft2

Izz = 36.9E6 slug*ft2

** Values Based on Geometrical and Historical Data

Ixx Izz

x z

Iyyy

8282

Future WorkFuture WorkAerodynamics: Research ways to make the fuselage more aerodynamically efficient Perform Trade Studies that look into W/S and Wing Sweep

Propulsion: Perform Trade Study cruising at lower altitudes Refine inlet and nozzle designs

Cooling System: Design actual heat exchanger, pumps, and piping for cooling system

Materials and Structures: Calculate the weights of the aircraft components based on weights of material and actual sizes

Stability and Control Analyze high lift devises (Mission Adaptive Wing)

Signature Research further ways to reduce signature

Cost Compare the benefit light weight materials will have on overall weight and cost of aircraft

8383

Cost AnalysisCost AnalysisModified DAPCA IV Cost Model (in constant United Modified DAPCA IV Cost Model (in constant United

States 2004 dollars)States 2004 dollars)Engineering, Tooling, and Manufacturing hours and Manufacturing Material Costs:

•Empty weight

•Velocity •Quantity produced in 5 years

Flight Test Costs: •Empty weight •Velocity •Number of flight test aircraft

Quality Control hours: •Historical data•Manufacturing hours

Development Support Cost: •Empty weight •Velocity

Engine and Avionics Cost Provided By:

•Lockheed Martin

8484

Cost AnalysisCost Analysisper 100 Aircraft & 4 Flight Test Aircraftper 100 Aircraft & 4 Flight Test Aircraft

Hours

Engineering 10,550,000

Tooling 6,376,700

Manufacturing 19,750,000

Quality Control 1,951,300

Total Costs

Development Support $115,310,000

Flight Test $44,690,000

Manufacturing Materials $383,980,000

Engine $206,700,000

Avionics $1,590,000,000

2004 Adjusted Labor Rates

Engineering $96

Tooling $98

Manufacturing $81

Quality Control $90

8585

RDTE cost per aircraft = $35,731,000

Flyaway cost per aircraft = $23,986,000Flyaway plus RDTE cost per aircraft = $59,717,000

Acquisition cost per aircraft = $67,913,000

Estimated RDTE + Flyaway Cost: per 100 aircraft = $5,970,000,000per aircraft = $59,700,000

Total Life Cycle cost:per 100 aircraft = $8,293,600,000per aircraft = $82,936,000

Cost AnalysisCost Analysisper 100 Aircraft & 4 Flight Test Aircraftper 100 Aircraft & 4 Flight Test Aircraft

8686

Cost AnalysisCost Analysis

Engineering Labor Hours

0

5

10

15

20

25

30

1 2 3 4 5 6 7 8 9 10

Week Number

Co

st

(hr)

Laura Melissa Luis Nic Average

8787

AcknowledgementsAcknowledgements and References and References

Acknowledgements

Dr. James D. Lang, UC San Diego, Project Advisor

Dr Leland M. Nicolai, Lockheed Martin, Project Sponsor

References

Aircraft Design: A Conceptual Approach, Raymer, D.P., 3rd Edition

Fundamentals of Aircraft Design, Nicolai, L.M., Revised 1984

8888

Thank YouThank You

1 Team Beta Project Advisor: Dr. James D. Lang, Project Sponsor: Dr. Leland M. Nicolai Laura Brandt...

Documents

Transcript of 1 Team Beta Project Advisor: Dr. James D. Lang, Project Sponsor: Dr. Leland M. Nicolai Laura Brandt...