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Final Design Report: Black Lightning

David Perveiler Lee Hargrave Perry Overbey Doug Heizer Chad Vetter Steve Moss Bill Viste

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:

Config. 9 Config. 12b

Config. 11

Basic Configuration

Diamond Wing

­ Low aspect ratio

Top-Mounted Inlet 2 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 aircraft Weight Addition: 1300 lbs

­ Radar Absorbent Structure (RAS)

Used on leading edge of wing Weight Addition: 775 lbs.

Radar Cross Section

Further RCS Reduction

­ Cockpit

Metallic Coating added to Windscreen

­ Onboard Radar

Low-Signature Radar Bandpass Resonant Radome

­ Active Cancellation

Sends Cancellation Signal

Radar Cross Section

RCS Results:

­ POFACETS Used for Comparative RCS Measure:

Result: Black Lightning was nearly equivalent

Black 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

Interdisciplinary Trade Studies

Reduce Gross takeoff Weight to minimize cost A/C length trade study

­ A/C length 100 ft

GTOW vs A/C Length

280,000 260,000 240,000

GTOW (lb)

220,000 200,000 180,000 160,000 140,000 120,000 100,000 0 20 40 60 80 100 120 140 160 180

A/C Length (ft)

Trade Studies (cont.)

Wing area

157,000

GTOW vs Wing Area

­ Wing area 2000 ft2

GTOW (lb)

156,000 155,000 154,000 153,000 152,000 151,000 150,000 1200

1400

1600

1800

2000 2

2200

2400

Wing Area (ft )

Leading edge sweep

­ Limited by mach cone ­ Initial sweep: 51° ­ Final sweep: 55°

GTOW (lb)

GTOW vs Leading Edge Sweep

154,000 153,000 152,000 151,000 150,000 149,000 148,000 147,000 146,000 145,000 50.5

51

51.5

52

52.5

53

53.5

54

54.5

Leading Edge Sweep (degree)

Trade Studies (cont.)

Root Chord

­ Root chord: 60 ft

GTOW vs Root Chord

151,000 150,500

GTOW (lb)

150,000 149,500 149,000 148,500 148,000 54 56 58 60 62 64 66

Root Chord (ft)

Trade Studies (cont.)

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

GTOW and Root Chord vs W ing Area

128000 127000 126000 125000 124000 123000 122000 121000 2200 100 80 60 40 20 2400 2600 2800 3000 3200 3400 0 3600 R o o t C h o rd (ft)

G T O W (lb )

Wing Are a (ft^2)

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

Initial Design Point

1.2

Thrust-to-Weight Ratio

Cruise Out

1

Dash Out

0.8

Dash In

0.6 0.4 0.2 0 0 50 100 150

Cruise In SEP = 0 Initial Design Point Afterburner Turn -2g SEP = 200 ft/s

Wing Loading, lb/ft^2

Performance Optimization

Trade Studies

­ Interdisciplinary trade studies ­ Loiter velocity ­ 318 ft/s

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

30

hrs (L/D)*(1/c)

25 20 15 10 5 0 200.0 250.0 300.0 350.0 400.0 450.0 500.0 550.0

Velocity (ft/s)

Optimization Constraint Diagram

Post-optimization design point

­ T/W = 0.50 ­ W/S = 62 lb/ft2

1.2

Design Point Cruise Out Dash Out

Thrust-to-Weight Ratio

1 0.8

Dash In

0.6 0.4 0.2 0 0 50 100 150

Cruise In SEP = 0 Design Point Afterburner Turn -2g SEP = 200 ft/s

Wing Loading, lb/ft^2

Onx/Offx Engine Data

Change to Onx/Offx engine data from RFP engine data equation Excess thrust calculated to find most constraining design point

­ Mach 1.6 at 55,000 ft

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

Results

All performance requirements met Mission duration ­ 4.17 hours Mission radius ­ 1,750 nm Mission fuel required ­ 56,497 lb Balanced 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

80,000 70,000 60,000

altitude (ft)

50,000 40,000 30,000 20,000 10,000 0 0.0 0.5 1.0 1.5 2.0

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

Stall Limit q Limit Engine Design Limit Max Thrust, SEP=0 Military Thrust, SEP=0

M

Measures of Merit

1-g Maxim um Thrust Specific Excess Power Envelope

80,000 70,000 60,000

altitude (ft)

50,000 40,000 30,000 20,000 10,000 0 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

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

Stall Limit q Limit Engine Design Limit SEP=0 ft/s SEP=500 ft/s SEP=1000 ft/s

M

Measures of Merit

2-g Maxim um Thrust Specific Excess Power Envelope

80,000 70,000 60,000

altitude (ft)

50,000

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

Stall Limit

40,000 30,000

q Limit

20,000 10,000

Engine Design Limit SEP=0 ft/s SEP=500 ft/s SEP=1000 ft/s

0 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

M

Measures of Merit

5-g Maximum Thrust Specific Excess Power Envelope

80,000 70,000 60,000

altitude (ft)

50,000 40,000 30,000 20,000 10,000 0 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

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

Stall Limit q Limit Engine Design Limit SEP=0 ft/s SEP=500 ft/s SEP=1000 ft/s

M

Measures of Merit

Maximum Thrust Sustained Load Factor Envelope

80,000 70,000 60,000

altitude (ft)

50,000 40,000 30,000 20,000 10,000 0 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

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

n=1 Stall Limit, n=1 n=2 Stall Limit, n=2 n=5 Engine Design Limit

M

Measures of Merit

Maximum Thrust Maneuvering, Sea Level

Load Factor, n

40 2 3 4 5 6 7

Stall Lim it Turn Radius = 1000 ft

35

Turn Radius = 500 ft SEP = 0 ft/s Corner Velocity Turn Radius = 2000 ft

30

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

Turn Rate (deg/s)

25

Turn Radius = 3000 ft

20

15

Max Load Factor Lim it

Engine Des ign Lim it

10

5

q Lim it

0 0.0 0.2 0.4 0.6 M 0.8 1.0 1.2 1.4

Measures of Merit

Maximum Thrust Maneuvering, 15,000 ft

Load Factor, n

2 40 3 4 5 6 7

Turn Radius = 1000 ft

35

Turn Radius = 500 ft

30

Turn Radius = 2000 ft Stall Lim it Corner Velocity SEP = 0 ft/s

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

Turn Rate (deg/s)

25

Turn Radius = 3000 ft

20

15

Max Load Factor Lim it

Engine Des ign Lim it

10

5

0 0.0 0.2 0.4 0.6 M 0.8 1.0 1.2 1.4

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 Data Similar to Airfoil on F-22 Raptor Thickness = 6%

Wing Dimensions

Planform Area Root Chord Tip Chord Taper Ratio Span Aspect Ratio Leading Edge Sweep Trailing Edge Sweep Quarter Chord Sweep Wetted Surface Area 3000 77.3 1.5 0.02 76.1 1.9 55.0 29.4 42.9 8255 ft2 ft ft ft Degrees Degrees Degrees ft2

Parasite Drag

Drag Buildup

1.80E-02 1.60E-02 1.40E-02

Landing Gear Flaps W ave Fuselage W ing Leaks & Protuberances Tail

Clean Subsonic Supercruise Takeoff

Parasite Drag

1.20E-02 1.00E-02 8.00E-03 6.00E-03 4.00E-03 2.00E-03 0.00E+00

Aerodynamic Results

Flight Condition Take-Off Clean Subsonic Supercruise Landing CL 1.158 0.057 0.049 1.158 CD 0.243 L/Dmax 9.76 Angle of attack 11 1 1 11

0.00631 16.10 0.0104 0.238 8.91 7.17

Internal Configuration

Diffuser Inlet Face Fuel Tank

Cockpit

Nozzle and Trough

Avionics Bay

Wing Tanks

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Internal Configuration

Electronics Bay Radar Starter JDAM Bays Engine

Nose Gear

AMRAAM Bay

Main Gear

Weight Calculations

Various Methods Used to Calculate Empty Weight Methods for Major Components:

­ Roskam Fighter and Bomber Equations

USAF US 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 W ings 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

130000 120000 110000 Normal Mission 100000 90000 80000 70000 60000 35 40 45 50 55 60 Feet from Nose of Airplane Aborted Mission Forward Limit Aft Limit

Weight of Airplane lbs

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.

15

Maximum Load Factor with Safety Factor

10

· Design limit load · Safety Factor: 1.5 · Maximum Dynamic

Pressure: 2,133 psf

L o a d F a c to r , n

Maximum Load Factor

5

factors: +7 and-3 g's

C Nmax+

0 0 250 500 750

Maximum Velocity

1000

Minimum Load Factor

-5

Minimum Load Factor with Safety Factor

C Nmax-

-10

Velocity (kts)

Lift Distribution

Computed at 50% Fuel Load Used Proportional Distribution for Spanwise Lift Used to Construct Shear and Bending Moments Spanwise Lift Distribution

25000

-1.4

Chordwise Lift Distribution

20000

-1

L if t ( lb /f t )

15000

-0.6

10000

-0.2

5000

0

0.2

0.4

0.6

0.8

1

1.2

0 0 10

0.2

Half-Span (ft)

20

30

Chord (ft)

Shear and Bending Moments

Used to Size Wing Spars

Shear Force Distribution

4000000

Moment Distribution

Bending Moment (ft-lb)

3000000

270000

Shear Force (lb)

180000

2000000

90000

1000000

0 0 10 20 30

0 0 10 20 30

Half-Span (ft)

Half-Span (ft)

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

Engine Design

­ Designed with OnX / OffX ­ SFC key design driver

Propulsion Configuration

Inlet Configuration

Top Mounted For Stealth Two RampVariable Inlet

Propulsion Configuration

Diffuser and Nozzle Configuration

Cooling Air Bypass Single Expansion Ramp 2-D Ejector Nozzle Sensors S- duct for Stealth

Shut-Off Doors

5% Duct Oversize

Engine Design

Engine design specifications

ECS / Avionics HPC Bleed HPT Cooling LPT Cooling HPX Max T4 JP-8 Compressor Pressure Ratio 1.50% 5% 5% 150 Hp 3200 deg R 18750 BTU/LbM 30

Inlet / Diffuser Loss

­ Determined from Inlet / Design ­ Dependent on Flight Speed

Engine Design

Engine Design Trade Studies

­ Bypass Ratio vs SFC

Mil thrust at Mach 1.6,

1.260 1.255 1.250 1.245 1.240 1.235 1.230 1.225 1.220 0.80

50,000 ft altittude

SFC

0.90

1.00 BPR

1.10

1.20

1.30

­ Fan Pressure Ratio Maximized to 3.5

Constrained by Model Converge - Operability Limit

Engine Design

Final Engine Design Performance

Thrust (LBF) SFC 46109 0.7366 8000 1.1123 78337 1.7212 20099 1.8562

SLS Dry Cruise Dry SLS Wet Cruise Wet

Physical Engine Sizing

Frontal Dimension and Weight

­ Comparison with P&W F100 ­ NASA EngineSim

M dot Scaled M dot MatchM dot Get Area Scaled Area

BRP 1.1 OnX

BRP .72 OnX

BRP .72 F100 Scaled

BRP 1.1 F100 Rescaled

Physical Engine Sizing

Length

­ F101-GE-102 Frontal Area / Length Ratio

Engine Length Diameter Weight

15.774 ft 5.354 ft 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

14 12 10 0.974 theta 2 deg 8 6 4 0.966 2 0 0 1 2 3 4 5 6 7 8 9 10 11 0.964 0.962 0.972 0.97 0.968 0.98 0.978 0.976

total pressure ratio

theta2 MIL-E5008B Pt ratio

theta 1 deg

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

Inlet Design

Free-Stream Capture Area

­ Determined from Raymer Equation

­ Massflow Breakdown

M M M M M M dot dot dot dot dot dot Components of M dot Total Engine From OffX analysis Bypass 20% Hydraulic Cooling 1% Oil Cooling 1% Nacelle Cooling 4% BL Bleed Area ratio estimate

Inlet Design

Bleed Area Ratio

Abl / Acap = .02

Diffuser Sizing

Diffuser Optimized for Minimum Separation

35

30

25 2 theta

20

15

10

5 0 2 4 6 8 10 L/H 12 14 16 18 20

Minium Separation

minimun seperation

Diffuser Sizing

Diffuser Sizing

Diffuser Dimensions

Diffuser Supersonic Capture Area 34.55 ft^2 Diffuser Throat Area 28.37 ft^2 Diffuser 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 Off Subsonic Drag

Bypass On Subsonic Drag

Bypass On Supersonic Drag

Installed Thrust

Inlet Drag

Subsonic

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 angle

Twin tail effect minimizes vertical projection

­ Minimize tail size

All-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:

Area S ft 203.53

2

Half span to break Root chord Half span point s sb co ft 10.48 ft 2.94 ft 14.6

Leading Chord at break Dihedral edge sweep point Tip chord angle cb ct LE ft 14.6 ft 1 deg 25 deg 52.31

Trailing edge sweep TE deg 27.02

Aileron Size:

­ Extend from 75% to 95% span

c a/c --15% s a/s --20% sa ft 6.5

Static Margin Range

130,000

Takeoff

120,000

Transonic

Total Weight (lbs)

110,000

Payload Release

100,000

Turn

90,000

Landing

80,000

70,000

Aborted Transonic Normal

60,000 -26% -24% -22% -20% -18% -16% -14%

Static Margin (% MAC)

Dynamic Stability Analysis

Determination 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)

MIL-F-8785C Compliance Level Takeoff Clean Landing Short Period 1 1 1 Longitudinal Phugoid 1 1 1 Dutch Roll Below 3 Below 3 1 Phugoid ----3 Lateral-Directional Spiral Mode 1 1 --Rolling Mode 1 1 --1 1 1 Roll Rate

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 Cost 20.1% Airframe Engineering and Design Cost 14.6% RDTE Profit 10% Cost to Finance 10% Development Support and Testing Cost 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 Design Cost 4.9% Production of Flight Test operations Cost 1.5%

Operations Cost: $27.93B

Cost Analysis for 200 Aircraft 12,000 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, consumable materials 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

140 120 100 80 60 40 20 0 0 200 400 600 # of Black Lightning 800 1000 1200

Cost in Millions

Cost Verification

Cost Verification Through Comparison

­ B-2 Spirit ­ F-117 Nighthawk ­ F-15 Eagle

Cost/Empty W eight Black Lightning Comparison Aircraft B-2 F-117 F-15 $/lb 8296 1499 883 $/lb 6620 3366 1169

Same Number of Aircraft Purchased for Each Case

Conclusion

Meets or Exceeds all RFP Requirements Major Design Drivers

­ Low Observability ­ Supercruise

Information

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