Final Design Report: Black Lightningae440a2009.pbworks.com/f/BlackLightning.pdf · Measures of...

Perry OverbeyChad VetterBill Viste

David PerveilerLee HargraveDoug HeizerSteve Moss

Final Design Report:Black Lightning

AIAA RFP

Background– Advanced deep interdiction aircraft– Antecedent aircraft

F-117, F-15E, B-1, B-2

Mission Requirements– Supercruise: Mach 1.6– Stealth exceeding F-117– Range: 3500 nm

Configuration Down-Selection

Key Design Parameters Established:– Stealth Requirement– Supercruise Requirement

Several Configurations Proposed:Several Configurations Proposed:

Config. 9

Config. 11

Config. 12b

Basic Configuration

Diamond Wing– Low aspect ratio

Top-Mounted Inlet2 Engines– Internal– Rear mounted

Internal Stores– 3 Bomb Bays

V-Tail

Signature Control

Design Factors– Radar (RF)– Infrared (IR)– Acoustic– Visible– Electromagnetic

Radar Cross Section (RCS)

Geometric Methods– Number of Planform Angles

– Flat Panel vs. Doubly-Curved

Radar Cross Section

Materials– Radar Absorbent Material (RAM)

Used only on critical areas of aircraftWeight Addition: 1300 lbs

– Radar Absorbent Structure (RAS)Used on leading edge of wingWeight Addition: 775 lbs.

Radar Cross Section

Further RCS Reduction– Cockpit

Metallic Coating added to Windscreen

– Onboard RadarLow-Signature RadarBandpass Resonant Radome

– Active CancellationSends Cancellation Signal

Radar Cross Section

RCS Results:– POFACETS Used for Comparative RCS Measure:

Result: Black Lightning was nearly equivalentBlack Lightning V-Tail (F-117)

Infrared Signature

Engine Exhaust Cooled– Trough Cooling

Glint Reduction– Windscreen with Transparent Coating

Visible Signature

Glint Reduction– Windscreen Coating

Paint Scheme – Dark Grey, Flat Paint

Acoustic Signature

Exhaust– Exhaust Mixing and Cooling

Shock Cone– Inevitable Shock Production

Electromagnetic

Passive Shielding– Attenuated to Overall Aircraft Output

Degaussing Technology

Interdisciplinary Optimization

Interdisciplinary trade studies– Minimize cost by minimizing GTOW

GTOW vs A/C Length

100,000

120,000

140,000

160,000

180,000

200,000

220,000

240,000

260,000

280,000

0 20 40 60 80 100 120 140 160 180

A/C Length (ft)

GTO

W (l

b)

Interdisciplinary Trade Studies

Reduce Gross takeoff Weight to minimize costA/C length trade study– A/C length 100 ft

Trade Studies (cont.)

Wing area– Wing area 2000 ft2

Leading edge sweep– Limited by mach cone– Initial sweep: 51°– Final sweep: 55°

GTOW vs Wing Area

150,000

151,000

152,000

153,000

154,000

155,000

156,000

157,000

1200 1400 1600 1800 2000 2200 2400

Wing Area (ft2)

GTO

W (l

b)

GTOW vs Leading Edge Sweep

145,000

146,000

147,000

148,000

149,000

150,000

151,000

152,000

153,000

154,000

50.5 51 51.5 52 52.5 53 53.5 54 54.5

Leading Edge Sweep (degree)

GTO

W (l

b)


Root Chord– Root chord: 60 ft

GTOW vs Root Chord

148,000

148,500

149,000

149,500

150,000

150,500

151,000

54 56 58 60 62 64 66

Root Chord (ft)

GTO

W (l

b)


GTOW and Root Chord vs Wing Area

121000122000123000124000125000126000127000128000

2200 2400 2600 2800 3000 3200 3400 3600

Wing Area (ft^2)

GTO

W (l

b)

0

20

40

60

80

100

Roo

t Cho

rd (f

t)

Wing Area and Root Chord Revisited– Performance and Weight analysis improved– Maintain a 30° Trailing Edge– Limited by root chord– Root chord: 77.3 ft– Wing area: 3000 ft2

Performance

Primary objectives of Performance– Meet or exceed all Performance requirements of RFP– Determine thrust required– Determine fuel required

Constraint Diagram

Initial design point– T/W = 0.45– W/S = 110 lb/ft2

0

0.2

0.4

0.6

0.8

1

1.2

0 50 100 150

Wing Loading, lb/ft^2

Thru

st-t

o-W

eigh

t Rat

io

Initial Design Point

Cruise Out

Dash Out

Dash In

Cruise In

SEP = 0

Initial Design Point -AfterburnerTurn -2g

SEP = 200 ft/s

Performance Optimization

Trade Studies– Interdisciplinary trade studies– Loiter velocity – 318 ft/s

(L/D)*(1/c) vs. Velocity

0

51015

2025

30

200.0 250.0 300.0 350.0 400.0 450.0 500.0 550.0

Velocity (ft/s)

(L/D

)*(1

/c)

hrs

Optimization Constraint Diagram

Post-optimization design point– T/W = 0.50– W/S = 62 lb/ft2

0

0.2

0.4

0.6

0.8

1

1.2

0 50 100 150

Wing Loading, lb/ft^2

Thru

st-t

o-W

eigh

t Rat

io

Design Point

Cruise Out

Dash Out

Dash In

Cruise In

SEP = 0

Design Point -AfterburnerTurn -2g

SEP = 200 ft/s

Onx/Offx Engine Data

Change to Onx/Offx engine data from RFP engine data equationExcess thrust calculated to find most constraining design point– Mach 1.6 at 55,000 ft

T/W = 0.53W/S = 42 lb/ft2

Results

All performance requirements metMission duration – 4.17 hoursMission radius – 1,750 nmMission fuel required – 56,497 lbBalanced field length– Standard day – 2,814 ft– Icy runway – 3,524 ft

Landing distance –– Standard day – 3,488 ft– Icy runway – 7,494 ft

Measures of Merit

Flight Envelope

0

10,000

20,000

30,000

40,000

50,000

60,000

70,000

80,000

0.0 0.5 1.0 1.5 2.0

M

altit

ude

(ft)

Stall Limit

q Limit

Engine Design Limit

Max Thrust, SEP=0

Military Thrust, SEP=0

Clean ConfigurationManeuver Weight W=98,757 lbs 50% Internal Fuel AIM-120 (2) 2,000 lb JDAM (4)

Measures of Merit

1-g Maximum Thrust Specific Excess Power Envelope

0

10,000

20,000

30,000

40,000

50,000

60,000

70,000

80,000

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

M

altit

ude

(ft)

Stall Limit

q Limit

Engine Design Limit

SEP=0 ft/s

SEP=500 ft/s

SEP=1000 ft/s


Measures of Merit

2-g Maxim um Thrust Specific Excess Power Envelope

0

10,000

20,000

30,000

40,000

50,000

60,000

70,000

80,000

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

M

altit

ude

(ft)

Stall Limit

q Limit

Engine Design Limit

SEP=0 ft/s

SEP=500 ft/s

SEP=1000 ft/s


Measures of Merit

5-g Maximum Thrust Specific Excess Power Envelope

0

10,000

20,000

30,000

40,000

50,000

60,000

70,000

80,000

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

M

altit

ude

(ft)

Stall Limit

q Limit

Engine Design Limit

SEP=0 ft/s

SEP=500 ft/s

SEP=1000 ft/s


Measures of Merit

Maximum Thrust Sustained Load Factor Envelope

0

10,000

20,000

30,000

40,000

50,000

60,000

70,000

80,000

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0M

altit

ude

(ft)

n=1

Stall Limit, n=1

n=2

Stall Limit, n=2

n=5

Engine Design Limit


Measures of Merit

Maximum Thrust Maneuvering, Sea Level

0

5

10

15

20

25

30

35

40

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4

M

Turn

Rat

e (d

eg/s

)


Engine Des ign Lim itMax Load Factor Lim it

q Lim it

Corner VelocitySEP = 0 ft/s

Stall Lim it

Turn Radius = 500 ft




Load Factor, n2 3 4 5 6 7

Measures of Merit

Maximum Thrust Maneuvering, 15,000 ft

0

5

10

15

20

25

30

35

40

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4

M

Turn

Rat

e (d

eg/s

) Turn Radius = 3000 ft


Load Factor, n2 3 4 5 6 7

Engine Des ign Lim it

SEP = 0 ft/s

Max Load Factor Lim it



Turn Radius = 2000 ftStall Lim it

Corner Velocity

Aerodynamics

FLUENT Static Pressure Gradient at Trailing Edge of NACA 64-206 at Flight Conditions

Airfoil Selection

NACA 64-206 Airfoil

Selected Based on Historical DataSimilar to Airfoil on F-22 RaptorThickness = 6%

Wing Dimensions

ft28255Wetted Surface AreaDegrees42.9Quarter Chord SweepDegrees29.4Trailing Edge SweepDegrees55.0Leading Edge Sweep

1.9Aspect Ratioft76.1Span

0.02Taper Ratioft1.5Tip Chordft77.3Root Chordft23000Planform Area

Parasite Drag

Drag Buildup

0.00E+00

2.00E-03

4.00E-03

6.00E-03

8.00E-03

1.00E-02

1.20E-02

1.40E-02

1.60E-02

1.80E-02

Clean Subsonic Supercruise Takeoff

Para

site

Dra

g

Landing Gear

Flaps

W ave

Fuselage

W ing

Leaks &ProtuberancesTail

Aerodynamic Results

117.170.2381.158Landing

18.910.01040.049Supercruise

116.100.006310.057Clean Subsonic

119.760.2431.158Take-Off

Angle of attack

L/DmaxCDCLFlight Condition

Cockpit

Fuel Tank

Avionics BayWing Tanks

Nozzle and Trough

Diffuser

Inlet Face

Internal Configuration

Radar

Nose Gear

JDAM Bays

AMRAAM Bay Main Gear

Starter Engine

Electronics Bay


Weight Calculations

Various Methods Used to Calculate Empty WeightMethods for Major Components:– Roskam Fighter and Bomber Equations

USAFUS Navy

– Raymer Fighter and Transport Equations– Torenbeek

Method for Smaller Components: – Raymer Fighter Equations

Empty Weight: 61,534 lbs.Take-off Gross Weight: 127,005 lbs.

Empty Weight: 61,354 lbs.

Engines 31% 19020 lbs

Fuselage 22.7% 13987 lbs

Wings 19% 11631 lbs

Misc. 7.4% 4531 lbs

Landing Gear 6.3% 3861 lbs

Fuel Systems 5.1% 3140 lbs

Stealth Systems 3.4% 2075 lbs

Air induction 3.2% 1978 lbs

Empenage 1.8% 1127 lbs

CG Movement

High Fuel Weight– CG Movement Throughout Mission

60000

70000

80000

90000

100000

110000

120000

130000

35 40 45 50 55 60

Feet from Nose of Airplane

Wei

ght o

f Airp

lane

lbs

Normal Mission

Aborted Mission

Forward Limit

Aft Limit

Structures

Objectives:– Construct V-n Diagram– Generate Lift, Shear, and Bending

Moment Distributions– Design Preliminary Structures– Design and Place Landing Gear– Select Aircraft Materials

V-n Diagram

• Illustrates load limits of aircraft as function of velocity.

• Design limit load factors: +7 and-3 g’s

• Safety Factor: 1.5

• Maximum Dynamic Pressure: 2,133 psf

-10

-5

0

5

10

15

0 250 500 750 1000

Velocity (kts)

Load

Fac

tor,

n

Maximum Load Factor

Maximum Load Factor with Safety Factor

CNmax+

Minimum Load Factor

Minimum Load Factor with Safety FactorCNmax-

Maximum Velocity

Lift Distribution

Computed at 50% Fuel LoadUsed Proportional Distribution for Spanwise LiftUsed to Construct Shear and Bending Moments

Spanwise Lift Distribution Chordwise Lift Distribution-1.4

-1

-0.6

-0.2

0.2

0 0.2 0.4 0.6 0.8 1 1.2

Chord (ft)0

5000

10000

15000

20000

25000

0 10 20 30Half-Span (ft)

Lift

(lb/

ft)

Shear and Bending Moments

Used to Size Wing Spars

0

90000

180000

270000

0 10 20 30

Half-Span (ft)

Shea

r Fo

rce

(lb)

Shear Force Distribution

0

1000000

2000000

3000000

4000000

0 10 20 30

Half-Span (ft)

Ben

ding

Mom

ent (

ft-lb

)

Moment Distribution

Wing Structural Layout

• Five Main Spars• Modeled as Cantilevered I-Beams• 15 Ribs placed at 24” Intervals• Analyzed using ANSYS to Find Max. Stresses

Fuselage Structural Layout

• 8 Main Bulkheads• 3 Secondary Bulkheads• 4 Longerons

Landing Gear• Tricycle layout • Two wheels for each gear• Tire selection: Nose Gear: Type VII 30 x 7.7

Main Gear: Type VII 40 x 14

Height of Landing Gear (ft) 8.8Main Gear Distance From Fuselage Centerline (ft) 10.5Distance From Nose to Main Gear (ft) 62Distance From Nose to Nose Gear (ft) 7Tipback Angle (degrees) 14.1Overturn Angle (degrees) 56.3 - 50.3Percent of Load on Nose Gear 19.5 - 15.7Maximum Load on Nose Gear (lbs) 24708Maximum Load on Main Gear (lbs) 102297

Material Selection

Aluminum 7075-T6 – Ribs, Spars, Bulkheads, Longerons

Aluminum 7050-T7351– Aircraft Skin

High-Strength Carbon Fiber-Epoxy– Control Surfaces

Titanium Ti-13V-11Cr-3Al– Structural Elements near Engines

Aircraft Steel (5 Cr-Mo-V)– Landing Gear

Propulsion

Engine Type– Low Bypass Supercruise Capable Augmented

Turbofan EngineEngine Design– Designed with OnX / OffX– SFC key design driver

Propulsion Configuration

Inlet Configuration

Two RampVariable Inlet

Top Mounted For Stealth

Propulsion Configuration

Diffuser and Nozzle Configuration

Single Expansion Ramp2-D Ejector Nozzle S- duct for Stealth

Cooling Air Bypass

Shut-Off Doors

Sensors

5% Duct Oversize

Engine Design

Engine design specifications

Inlet / Diffuser Loss– Determined from Inlet / Design– Dependent on Flight Speed

ECS / Avionics HPC Bleed 1.50%HPT Cooling 5%LPT Cooling 5%HPX 150 HpMax T4 3200 deg RJP-8 18750 BTU/LbMCompressor Pressure Ratio 30

Engine Design

Engine Design Trade Studies– Bypass Ratio vs SFC

– Fan Pressure Ratio Maximized to 3.5Constrained by Model Converge - Operability Limit

Mil thrust at Mach 1.6, 50,000 ft altittude

1.2201.2251.2301.2351.2401.2451.2501.2551.260

0.80 0.90 1.00 1.10 1.20 1.30

BPR

SFC

Engine Design

Final Engine Design Performance

Thrust (LBF) SFCSLS Dry 46109 0.7366Cruise Dry 8000 1.1123SLS Wet 78337 1.7212Cruise Wet 20099 1.8562

Frontal Dimension and Weight– Comparison with P&W F100– NASA EngineSim

Physical Engine Sizing

BRP 1.1OnX

BRP .72OnX

BRP .72F100 Scaled

M dot Scaled M dotMatchM dot

Get Area

BRP 1.1F100 Rescaled

Scaled Area

Physical Engine Sizing

Length– F101-GE-102 Frontal Area / Length Ratio

EngineLength 15.774 ftDiameter 5.354 ftWeight 9510 lbs

Inlet Design

Diffuser Mach Estimates– M 0.8 Entering Subsonic Diffuser– M 0.4 Entering Entering Engine– Pressure Recovery of 0.98

Inlet Shock Performance– 0.1 Mach Margin of Safety in Design– Mach Reduces From 1.7 to 0.8 Over Two Oblique

Shocks and a Normal Shock

Inlet Design

Ramp Angle Effects

– Pressure Recovery Maximum of 0.9782– Theta 1 = 6 deg, Theta 2 = 6.25 deg

0

2

4

6

8

10

12

14

0 1 2 3 4 5 6 7 8 9 10 11

theta 1 deg

thet

a 2

deg

0.962

0.964

0.966

0.968

0.97

0.972

0.974

0.976

0.978

0.98

tota

l pre

ssur

e ra

tio

theta2

MIL-E-5008BPt ratio

Inlet Design

Free-Stream Capture Area– Determined from Raymer Equation

– Massflow Breakdown

M dot Engine From OffX analysisM dot Bypass 20%M dot Hydraulic Cooling 1%M dot Oil Cooling 1%M dot Nacelle Cooling 4%M dot BL Bleed Area ratio estimate

Components of M dot Total

Inlet Design

Bleed Area Ratio

Abl / Acap = .02

Diffuser Sizing

Diffuser Optimized for Minimum Separation

5

10

15

20

25

30

35

0 2 4 6 8 10 12 14 16 18 20

L/H

2 th

eta

minimun seperationMinium Separation

Diffuser Sizing

Diffuser Sizing

Diffuser Dimensions

DiffuserSupersonic Capture Area 34.55 ft 2̂Diffuser Throat Area 28.37 ft 2̂Diffuser Duct Length 14.94 ft

Installed Thrust

NPF determination– Nozzle CD=.0075 Assumed Constant

CD Based on 120 square ft Cross-sectional Area

– Inlet Drag Dependent on Flight Condition

Bypass OnBypass Off Bypass On

Subsonic Drag Supersonic DragSubsonic Drag

Installed Thrust

Inlet DragSubsonic

Supersonic

Stability and Controls

Primary Objectives of S&C– Obtain tail size and geometry– Determine control surface size and arrangement– Determine dynamic stability characteristics

Restrictions and Goals

Conform to Signature Requirements– Minimize total planform angles

Horizontal projection must match

– Minimize dihedral angleTwin tail effect minimizes vertical projection

– Minimize tail sizeAll-moving tail reduces total tail area

Meet MIL-F-8785C Level 1 Requirements

Tail Size and Geometry

Horizontal Size Based on Static Longitudinal Stability– CG obtained from W&B– Subsonic static margin between –30% and 10%

Vertical Size Based on OEI Takeoff Condition– Rudder must provide sufficient yawing moment

Root chord cannot extend past fuselage– Breakpoint required in tail to increase area

Control Surface Sizing

Pitch and yaw controlled by all moving “ruddervator”Roll controlled by ailerons– Must counter adverse rolling moment caused by tail

at OEI TO condition– Must meet MIL-F-8785C roll rate requirement

S&C Sizing

Tail Geometry:

Aileron Size:– Extend from 75% to 95% span

ca/c sa/s sa

--- --- ft15% 20% 6.5

Area Half span

Half span to break point

Root chord

Chord at break point Tip chord

Dihedral angle

Leading edge sweep

Trailing edge sweep

S s sb co cb ct Γ ΛLE ΛTE

ft2 ft ft ft ft ft deg deg deg203.53 10.48 2.94 14.6 14.6 1 25 52.31 27.02

Static Margin Range

60,000

70,000

80,000

90,000

100,000

110,000

120,000

130,000

-26% -24% -22% -20% -18% -16% -14%Static Margin (% MAC)

Tota

l Wei

ght (

lbs)

Aborted

NormalTransonic

Transonic

Payload Release

Takeoff

Landing

Turn

Dynamic Stability AnalysisDetermination of Subsonic Nondimensional Stability Derivatives – Roskam’s method Based on USAF Datcom

Assumed an effective static margin to account for the stability augmentation system (No equivalent lateral assumption made)

Takeoff Clean LandingShort Period 1 1 1

Phugoid 1 1 1Dutch Roll Below 3 Below 3 1Phugoid --- --- 3

Spiral Mode 1 1 ---Rolling Mode 1 1 ---

1 1 1

MIL-F-8785C Compliance Level

Roll Rate

Longitudinal

Lateral-Directional

Cost Analysis

3 Main Components– Research, Testing, Development, and Evaluation– Production Cost– Operation Cost

From Data:– Operating Cost Per Hour Flight Time– Unit Price– Price Per Pound of Empty Weight

Cost Trade Study Performed – Cost Then Compared to Current Aircraft

Total RTDE: $5.72 Billion

Flight Test Airplanes Cost 35.9%

Test and Simulation Facilities Cost20.1%

Airframe Engineering and DesignCost 14.6%

RDTE Profit 10%

Cost to Finance 10%

Development Support and TestingCost 5.1%

Flight Test Operations Cost 4.2%

Total Production Cost: $10.31B

Cost Breakdown for 200 Aircraft Purchase

Manufacturing Labor 39%

Avionics and Engines 36.7%

Manufacturing Material 9.6%

Tooling 9.6%

Quality Control 5.1%

Acquisition Cost: $13.7B

Cost Breakdown for 200 Aircraft Purchase

Airplane Production Cost 75.3%

Finance Cost 9.1%

Profit 9.1%

Airframe Engineering and DesignCost 4.9%

Production of Flight Test operationsCost 1.5%

Operations Cost: $27.93B

Cost Analysis for 200 Aircraft12,000 Hour/AC Flight Time over 30 Years

Direct personnel 38.8%

Depot 17%

Indirect Personnel 14%

Spares 14%

Fuel Oil and Lubricant Cost 9.7%

Maintenance, consumablematerials 4.5%

Misc. 2%

Cost Analysis

Unit Price for 200 Aircraft Sold:– $97.08 Million

Operation Cost per Hour: – $13,139 /hr

Price per Pound of Empty Weight:– $1,582 /lb

Cost Analysis, cont.

Cost per Aircraft vs Number of Airplanes sold

0

20

40

60

80

100

120

140

0 200 400 600 800 1000 1200

# of Black Lightning

Cost

in M

illio

ns

Cost Verification

Cost Verification Through Comparison– B-2 Spirit– F-117 Nighthawk– F-15 Eagle

Cost/Empty Weight Black Lightning ComparisonAircraft $/lb $/lbB-2 8296 6620

F-117 1499 3366

F-15 883 1169

Same Number of Aircraft Purchased for Each Case

Conclusion

Meets or Exceeds all RFP RequirementsMajor Design Drivers– Low Observability– Supercruise

Final Design Report: Black Lightningae440a2009.pbworks.com/f/BlackLightning.pdf · Measures of...

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Transcript of Final Design Report: Black Lightningae440a2009.pbworks.com/f/BlackLightning.pdf · Measures of...