final design report: black lightningae440a2009.pbworks.com/f/blacklightning.pdf · measures of...
TRANSCRIPT
![Page 1: Final Design Report: Black Lightningae440a2009.pbworks.com/f/BlackLightning.pdf · Measures of Merit Maximum Thrust Maneuvering, Sea Level 0 5 10 15 20 25 30 35 40 0.00.2 0.40.6 0.81.0](https://reader036.vdocuments.us/reader036/viewer/2022070709/5ebc758239b9d9738f052f22/html5/thumbnails/1.jpg)
Perry OverbeyChad VetterBill Viste
David PerveilerLee HargraveDoug HeizerSteve Moss
Final Design Report:Black Lightning
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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
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Configuration Down-Selection
Key Design Parameters Established:– Stealth Requirement– Supercruise Requirement
Several Configurations Proposed:Several Configurations Proposed:
Config. 9
Config. 11
Config. 12b
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Basic Configuration
Diamond Wing– Low aspect ratio
Top-Mounted Inlet2 Engines– Internal– Rear mounted
Internal Stores– 3 Bomb Bays
V-Tail
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Signature Control
Design Factors– Radar (RF)– Infrared (IR)– Acoustic– Visible– Electromagnetic
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Radar Cross Section (RCS)
Geometric Methods– Number of Planform Angles
– Flat Panel vs. Doubly-Curved
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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.
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Radar Cross Section
Further RCS Reduction– Cockpit
Metallic Coating added to Windscreen
– Onboard RadarLow-Signature RadarBandpass Resonant Radome
– Active CancellationSends Cancellation Signal
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Radar Cross Section
RCS Results:– POFACETS Used for Comparative RCS Measure:
Result: Black Lightning was nearly equivalentBlack Lightning V-Tail (F-117)
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Infrared Signature
Engine Exhaust Cooled– Trough Cooling
Glint Reduction– Windscreen with Transparent Coating
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Visible Signature
Glint Reduction– Windscreen Coating
Paint Scheme – Dark Grey, Flat Paint
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Acoustic Signature
Exhaust– Exhaust Mixing and Cooling
Shock Cone– Inevitable Shock Production
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Electromagnetic
Passive Shielding– Attenuated to Overall Aircraft Output
Degaussing Technology
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Interdisciplinary Optimization
Interdisciplinary trade studies– Minimize cost by minimizing GTOW
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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
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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)
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Trade Studies (cont.)
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)
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Trade Studies (cont.)
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
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Performance
Primary objectives of Performance– Meet or exceed all Performance requirements of RFP– Determine thrust required– Determine fuel required
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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
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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
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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
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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
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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
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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)
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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
Clean ConfigurationManeuver Weight W=98,757 lbs 50% Internal Fuel AIM-120 (2) 2,000 lb JDAM (4)
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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
Clean ConfigurationManeuver Weight W=98,757 lbs 50% Internal Fuel AIM-120 (2) 2,000 lb JDAM (4)
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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
Clean ConfigurationManeuver Weight W=98,757 lbs 50% Internal Fuel AIM-120 (2) 2,000 lb JDAM (4)
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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
Clean ConfigurationManeuver Weight W=98,757 lbs 50% Internal Fuel AIM-120 (2) 2,000 lb JDAM (4)
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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
)
Clean ConfigurationManeuver Weight W=98,757 lbs 50% Internal Fuel AIM-120 (2) 2,000 lb JDAM (4)
Engine Des ign Lim itMax Load Factor Lim it
q Lim it
Corner VelocitySEP = 0 ft/s
Stall Lim it
Turn Radius = 500 ft
Turn Radius = 1000 ft
Turn Radius = 2000 ft
Turn Radius = 3000 ft
Load Factor, n2 3 4 5 6 7
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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
Clean ConfigurationManeuver Weight W=98,757 lbs 50% Internal Fuel AIM-120 (2) 2,000 lb JDAM (4)
Load Factor, n2 3 4 5 6 7
Engine Des ign Lim it
SEP = 0 ft/s
Max Load Factor Lim it
Turn Radius = 500 ft
Turn Radius = 1000 ft
Turn Radius = 2000 ftStall Lim it
Corner Velocity
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Aerodynamics
FLUENT Static Pressure Gradient at Trailing Edge of NACA 64-206 at Flight Conditions
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Airfoil Selection
NACA 64-206 Airfoil
Selected Based on Historical DataSimilar to Airfoil on F-22 RaptorThickness = 6%
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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
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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
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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
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Cockpit
Fuel Tank
Avionics BayWing Tanks
Nozzle and Trough
Diffuser
Inlet Face
Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Internal Configuration
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Radar
Nose Gear
JDAM Bays
AMRAAM Bay Main Gear
Starter Engine
Electronics Bay
Internal Configuration
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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.
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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
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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
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Structures
Objectives:– Construct V-n Diagram– Generate Lift, Shear, and Bending
Moment Distributions– Design Preliminary Structures– Design and Place Landing Gear– Select Aircraft Materials
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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
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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)
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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
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Wing Structural Layout
• Five Main Spars• Modeled as Cantilevered I-Beams• 15 Ribs placed at 24” Intervals• Analyzed using ANSYS to Find Max. Stresses
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Fuselage Structural Layout
• 8 Main Bulkheads• 3 Secondary Bulkheads• 4 Longerons
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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
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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
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Propulsion
Engine Type– Low Bypass Supercruise Capable Augmented
Turbofan EngineEngine Design– Designed with OnX / OffX– SFC key design driver
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Propulsion Configuration
Inlet Configuration
Two RampVariable Inlet
Top Mounted For Stealth
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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
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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
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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
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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
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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
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Physical Engine Sizing
Length– F101-GE-102 Frontal Area / Length Ratio
EngineLength 15.774 ftDiameter 5.354 ftWeight 9510 lbs
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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
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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
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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
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Inlet Design
Bleed Area Ratio
Abl / Acap = .02
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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
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Diffuser Sizing
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Diffuser Sizing
Diffuser Dimensions
DiffuserSupersonic Capture Area 34.55 ft 2̂Diffuser Throat Area 28.37 ft 2̂Diffuser Duct Length 14.94 ft
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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
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Installed Thrust
Inlet DragSubsonic
Supersonic
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Stability and Controls
Primary Objectives of S&C– Obtain tail size and geometry– Determine control surface size and arrangement– Determine dynamic stability characteristics
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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
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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
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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
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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
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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
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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
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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
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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%
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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%
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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%
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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%
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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
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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
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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
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Conclusion
Meets or Exceeds all RFP RequirementsMajor Design Drivers– Low Observability– Supercruise