Transcript Slide 1

팔
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ead aeronautics 八
CONCEPTUAL DESIGN OF OPTIMUS –
A SUPERSONIC AIRCRAFT FOR SUPERSONIC AIR-LAUNCH
AE 440-A; PROF. E. LOTH
Nov 28, 2006
3:00 – 4:00pm
ead aeronautics え八팔
TEAM MEMBERS
•
•
•
•
•
•
•
Nathan Jung Her
Calvin Lee
Seiji Matsushita
Phillip Robinson
Janice Quek
Wei Ren Quah
Patrick Woo (T.L.)
(Structures)
(Stability and Control)
(Propulsion)
(Costs & Con/Ops)
(Aerodynamics)
(Config., Weights and Balance)
(Performance)
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INTRODUCTION
• Introduction
- Air-launch for more efficient space access suggested
• Request for Proposal (RFP)
- Air breathing aircraft to air-launch Falcon 1 rocket
- Launch occurs at altitude of at least 50,000 ft
- 2 < Mach no. < 3
- Takeoff from a runway in the U.S.
- Launch occurs at distance of at least 200 miles offshore
- launch angle γ = 25°(3 – M)
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TEAM THEME
• “Balance between Cost and Performance”
• Performance = Payload launch speed
- higher launch speed = higher delta V gain
• Design concepts and selection process
• Specialty areas
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DESIGN CONCEPTS
• 14 design concepts were compared
1
Fuselage
Flying wing,
Swept/Delta
Wing
2
3
4
5
6
7
Cylindrical
Cylindrical
Cylindrical
Cylindrical
blended
Cylindrical
Flying clamp
Top, Delta,
withCanards
Mid,
Swept
/Delta
Low, Delta
with Canards
Mid, Delta
Top,
Variable
Top to Mid
Tail
Twin winglets
Conventional
Twin
winglets
Twin verticle
tail
Twin verticle
tail
Conventional
Conventional
Landing
Gear
Tricyle
Tricycle
Tricycle
Tricycle
Tricycle
Mult-bogey
Multi-bogey
Engines
Turbofan
Turbofan
Turbojet
Turbojet
Turbojet
Turbofan
Turbofan
Payload
Captive on
Top
Captive on
Bottom
Captive on
Top
Captive on
Top
Internal,
Captive on
Bottom
Internal,
Captive on
Bottom
Captive on
Bottom
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DESIGN CONCEPTS
8
Fuselage
Cylindrical
9
10
Blended,
Mid,
Variable
Blended,
Mid, Delta
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Blended,
Mid, Delta
12
Flying wing,
Delta
13
14
Cylindrical
blended
Cylindrical
Low, Swept
Low, Swept
Wing
Top, Swept
Tail
Conventional
Conventional
Twin verticle
tail
Conventional
V-tail
Conventional
V-tail
Landing
Gear
Tricycle
Multi-bogey
Tricycle
Multi-bogey
Tricycle
Tricycle
Tricycle
Engines
Turbojet
Turbojet
Turbojet
Turbojet
Turbojet
Turbojet
Turbojet
Payload
Internal/
Captive on
Bottom
Captive on
Bottom
Captive on
Top
Captive on
Bottom
Captive on
Top
Front
Captive on
Top
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DESIGN CONCEPTS
• Eliminate designs with the following attributes:
- Variable wing geometries
- Internal/external payload carrying method
- Nose forward carrying method
• Justifications:
- penalty of weight
- complexity
- chance of failure
- maintenance
- stability
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DESIGN CONCEPTS
• Group remaining design concepts with morphology
1
Fuselage
Wing
Flying
wing,
Swept/
Delta
2
3
4
5
Cylindrical
Cylindrical
Flying
clamp
Cylindrical
Top,
Delta,
With
Canards
Mid,
Swept/
Delta
Top to
Mid
Low, Delta
With
Canards
6
7
8
Cylindrical
Flying
wing,
Delta
Flying Wing,
Mid, Delta
Low,
Swept
Tail
Twin
winglets
Conventional
Twin
winglets
Conventional
Twin
verticle tail
V-tail
Conventional
V-tail
Landing
Gear
Tricyle
Tricycle
Tricycle
Multi-bogey
Tricycle
Tricycle
Multi-bogey
Tricycle
Engines
Turbofan
Turbofan
Turbojet
Turbofan
Turbojet
Turbojet
Turbojet
Turbojet
Payload
Captive
on
Top
Captive on
Bottom
Captive
on Top
Captive on
Bottom
Captive on
Top
Captive
on Top
Captive on
Bottom
Captive on
Top
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DESIGN CONCEPTS
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
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DESIGN CONCEPTS
3
2
1
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CONFIGURATION,
WEIGHTS & BALANCE
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
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DESIGN CONCEPT 1
Front View
• Delta Wings
- wing's leading edge remains
behind shock wave
- high stall angle
- simplicity
• Canards
- more statically stable
- reduces lift-induced drag
• Captive on top
Top View
Side View
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DESIGN CONCEPT 2
Front View
• Delta Wings
- wing's leading edge
remains behind shock
wave
- high stall angle
- simplicity
• Flying Wing
• Captive on top
Top View
Side View
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DESIGN CONCEPT 3
Front View
• Swept Wings
- reduces drag
- spanwise flow
• Captive on top
Top View
Side View
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INITIAL SIZING
Mission Profile
4 50,000 ft 6
5
30,000 ft
30,000 ft 8
2
3
7
0
1
9
10
0. Start
4.
Climb
8.
Cruise in
1. Warm-up and Take-off
5.
Dash
9.
Descend
2. Climb
6.
Launch
10. Land
3. Cruise out
7.
Descend
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INITIAL SIZING
Design Concept 1
Mach
number
Altitude (ft)
Range (ft)
Wi/(i-1)
0
Start
-
0
-
-
1
Warm-up and Take-off
-
0
-
0.9700
2
Climb (to 30,000ft)
-
30,000
-
0.9850
3
Cruise out
0.8
30,000
1,056,000
0.9299
4
Climb (to 50,000ft)
-
50,000
-
0.9850
5
Dash (for 10 mins)
2.5
50,000
1,640,419
0.9575
6
Payload Drop
2.5
50,000
-
1.0000
7
Descend (to 30,000ft)
-
30,000
-
0.9900
8
Cruise in
0.8
30,000
2,696,419
0.8305
9
Descend
-
0
-
0.9900
10
Land
-
0
-
0.9950
GTOW = 314,086 lbs
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INITIAL SIZING
Design Concept 2
Mach
number
Altitude (ft)
Range (ft)
Wi/(i-1)
0
Start
-
0
-
-
1
Warm-up and Take-off
-
0
-
0.9700
2
Climb (to 30,000ft)
-
30,000
-
0.9850
3
Cruise out
0.8
30,000
1,056,000
0.9484
4
Climb (to 50,000ft)
-
50,000
-
0.9850
5
Dash (for 10 mins)
2.5
50,000
1,640,419
0.9471
6
Payload Drop
2.5
50,000
-
1.0000
7
Descend (to 30,000ft)
-
30,000
-
0.9900
8
Cruise in
0.8
30,000
2,696,419
0.8734
9
Descend
-
0
-
0.9900
10
Land
-
0
-
0.9950
GTOW = 280,576 lbs
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INITIAL SIZING
Design Concept 3
Mach
number
Altitude (ft)
Range (ft)
Wi/(i-1)
0
Start
-
0
-
-
1
Warm-up and Take-off
-
0
-
0.97
2
Climb (to 30,000ft)
-
30000
-
0.985
3
Cruise out
0.8
30000
1056000
0.9187
4
Climb (to 50,000ft)
-
50000
-
0.985
5
Dash (for 10 mins)
2.5
50000
1640419
0.9666
6
Payload Drop
2.5
50000
-
1
7
Descend (to 30,000ft)
-
30000
-
0.99
8
Cruise in
0.8
30000
2696419
0.8053
9
Descend
-
0
-
0.99
10
Land
-
0
-
0.995
GTOW = 336,306 lbs
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WEIGHT SUMMARY
Design Concept 1 Design Concept 2 Design Concept 3
GTOW (lbs)
314,086
280,576
336,306
Empty
Weight (lbs)
109,526
96,627
118,274
Empty
Weight
Fraction
0.431
0.438
0.428
Mission Fuel
Weight (lbs)
84,113
63,624
97,797
Fuel Weight
Fraction
0.331
0.2884
0.3539
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CONFIGURATION OF OPTIMUS
C.G. of Engines = 128.1 ft
660
Span = 95.71 ft
C.G. of Fuel = 75 ft
Overall C.G. = 87.1 ft
C.G. OF Empty Weight of Aircraft
= 80.75 ft
C.G. of Falcon 1 = 87.1 ft
Length of Aircraft = 154.79 ft
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AERODYNAMICS
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
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NUMERICAL COMPARISONS
Property
Design Concept Design Concept Design Concept
1
2
3
Description
Strategic Bomber
Flying Wing
Conventional
Configuration
Maximum Speed
Mach 3.1
Mach 0.67
Mach 2.21
Wing Type
Delta
Delta
Aft Swept
Wing Area
6296
5000
6200
Wing Span
105
172
39.9
Sweep Back Angle
66˚
33˚
26.6˚
Aspect Ratio
1.75
5.9168
8
Fuselage Length
185ft
69ft
58.67ft
L/D
8.33
7.03
12.8
CD0
0.013
0.027
0.023
e
0.66
0.7
0.6
CD
0.03
0.03
0.04
CL
0.25
0.211
0.5
Aircraft in Industry
XB-70
B-2 Spirit
C-5A
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WING GEOMETRY
• Delta Wing Geometry versus Aft-Swept Wing
Geometry
• Performance Characteristics
• Theme and Team Goals
Balance between COST and PERFORMANCE
Wing Type
Delta
Delta
Aft-Swept
CD
0.03
0.03
0.04
CL
0.25
0.211
0.5
Lift
494485
331437
973892
Drag
59338
47123
77911
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ASPECT RATIO
•
-
Importance of Aspect Ratio
Wing tip Vortices
Reducing Induced Drag
Reducing Wave Drag
Key: Optimizing
Aspect Ratio of
Wing
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TRADE STUDY: EFFECT OF AR ON CD
• Speeds at Mach 0.8 and Mach 2.5
Trend:
• At Mach 0.8, CD decreases as aspect ratio increases.
• At Mach 2.5, CD increases as aspect ratio increases
0.04
CD(wing)
0.035
0.03
Subsonic
0.025
Supersonic
0.02
0.015
1.4
1.9
2.4
Aspect Ratio
2.9
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NUMERICAL ANALYSIS
Using the component
build-up method,
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MISSION MODEL
Drag
Drag Model of Mission
Take-off:
59,032lb
250000
Dash
Drag (lb)
200000
Subsonic Cruise:
58,800lb
After
launch
150000
Take Off
100000
Subsonic Cruise
After Launch:
209,468lb
50000
Land
0
0
10000
Dash:
217,695lb
20000
30000
40000
Altitude (ft)
50000
60000
Land:
43,246lb
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MISSION MODEL
Lift Model of Mission
CD at different Mach Numbers
0.5
0.1
0.3
0.08
CD
CL
0.4
0.06
0.04
0.2
0.02
0.1
0
0
0
0
10000
1
20000
2
Mach
Num bers
30000
40000
3
50000
60000
Height (ft)
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FUSELAGE DESIGN
• At supersonic speeds, one of the greatest challenges is to
minimize wave drag (pressure drag due to formation of
shocks)
• Related to total cross-sectional area of aircraft
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CONCLUSION & FUTURE
CONSIDERATIONS
• Preliminary analysis was performed on all 3 aircraft
design concepts.
• Detailed numerical analysis was conducted of the
Optimus.
FUTURE CONSIDERATIONS
• Methods to reduce drag.
• A more refined lift & drag model.
• Airfoil Selection.
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PERFORMANCE
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
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FUEL CONSUMPTION
Mission Profile
4 50,000 ft 6
5
30,000 ft
30,000 ft 8
2
3
7
0
1
9
10
0. Start
4.
Climb
8.
Cruise in
1. Warm-up and Take-off
5.
Dash
9.
Descend
2. Climb
6.
Launch
10. Land
3. Cruise out
7.
Descend
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FUEL CONSUMPTION
Design Concept 1
Mach
number
Altitude
(ft)
Range
(ft)
Wi/(i-1)
Fuel burned
(lb)
0
Start
-
0
-
-
-
1
Warm-up and Take-off
-
0
-
0.97
7,624
2
Climb (to 30000ft)
-
30,000
-
0.985
3,697
3
Cruise out
0.8
30,000
1,056,000
0.9299
17,020
4
Climb (to 50000ft)
-
50,000
-
0.985
3,387
5
Dash (for 10 mins)
2.5
50,000
1,640,419
0.9575
9,452
6
Payload Drop
2.5
50,000
-
1
0
7
Descend (to 30000ft)
-
30,000
-
0.99
2,129
8
Cruise in
0.8
30,000
2,696,419
0.8305
35,732
9
Descend
-
0
-
0.99
1,751
Land
-
0
-
0.995
867
10
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FUEL CONSUMPTION
Design Concept 2
Mach
number
Altitude
(ft)
Range
(ft)
Wi/(i-1)
Fuel burned
(lb)
0
Start
-
0
-
-
-
1
Warm-up and Take-off
-
0
-
0.97
7,908
2
Climb (to 30000ft)
-
30,000
-
0.985
3,835
3
Cruise out
0.8
30,000
1,056,000
0.9299
17,655
4
Climb (to 50000ft)
-
50,000
-
0.985
3,513
5
Dash (for 10 mins)
2.5
50,000
1,640,419
0.9575
9,804
6
Payload Drop
2.5
50,000
-
1
0
7
Descend (to 30000ft)
-
30,000
-
0.99
2,209
8
Cruise in
0.8
30,000
2,696,419
0.8305
37,065
9
Descend
-
0
-
0.99
1,816
Land
-
0
-
0.995
899
10
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FUEL CONSUMPTION
Design Concept 3
Mach
number
Altitude
(ft)
Range
(ft)
Wi/(i-1)
Fuel burned
(lb)
0
Start
-
0
-
-
-
1
Warm-up and Take-off
-
0
-
0.97
15,498
2
Climb (to 30000ft)
-
30,000
-
0.985
7,517
3
Cruise out
0.8
30,000
1,056,000
0.9299
34,601
4
Climb (to 50000ft)
-
50,000
-
0.985
6,885
5
Dash (for 10 mins)
2.5
50,000
1,640,419
0.9575
19,214
6
Payload Drop
2.5
50,000
-
1
0
7
Descend (to 30000ft)
-
30,000
-
0.99
4,329
8
Cruise in
0.8
30,000
2,696,419
0.8305
72,641
9
Descend
-
0
-
0.99
3,559
Land
-
0
-
0.995
1,762
10
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FUEL CONSUMPTION SUMMARY
Design
Concept 1
Design
Concept 2
Design
Concept 3
Fuel burned
(lbs)
94,264
84,704
166,006
Mission Fuel
Weight (lbs)
97,092
87,245
170,986
• Assuming no payload drop so the Falcon 1 rocket can be
safely returned
• Although Design Concept 2 consumes the least amount
of fuel, Design Concept 1 is chosen
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CONSTRAINT ANALYSIS
• Take-off with 50ft clearance from 15,000 ft runway at sea
level
• Landing distance of 3,000 ft
• Cruises at M = 0.8 at 30,000 ft
• Dashes at M = 2.5 at 50,000 ft
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CONSTRAINT ANALYSIS
Takeoff
Landing
Cruise at M=0.8 at 30kft
Dash at M=2.5 at 50kft
2
1.8
1.6
(T/W)o
1.4
1.2
1
(60, 0.9)
0.8
0.6
0.4
0.2
0
20
30
40
50
(W/S)o
60
70
80
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CONCLUSION AND
FURTHER ANALYSIS
• Thrust to Weight ratio of 0.9 is required for dash
constraint
• Enough thrust must be provided!
• Further analysis in the next semester
- Maximum dash speed
- Maximum altitude
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PROPULSION
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
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COMPARISONS
• Design Concept 1 (BEST)
- Able to carry more engines
• Design Concept 2 (WORST)
- Limit of engine size and
numbers
- Very high thrust engine
needed
• Design Concept 3 (GOOD)
- Limit of engine size and
numbers
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TURBOJETS
• Concorde
(Rolls-Royce/SNECMA
Olympus 593 Mk 602
turbojets )
• XB-70
(General Electric J-93
afterburning turbojets
Peter, St. James, “The Histroy of Aircraft Gas Turbine Engine Development in the United
States … A Tradition of Excellence” 1 st ed., The International Gas Turbine Institute of
ASME., 1999, pp.430-569
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TURBOFANS
• F-15
(Pratt & Whitney F100220 afterburning
turbofans )
• F-111F
(Pratt & Whitney TF30111 afterburning
turbofans)
Peter, St. James, “The Histroy of Aircraft Gas Turbine Engine Development in the United
States … A Tradition of Excellence” 1 st ed., The International Gas Turbine Institute of
ASME., 1999, pp.430-569
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ENGINES DATA
Turbojets and Turbofans with Afterburner (AB) at Sea Level Condition
Olympus 593
Max/Normal
Thrust (lb)
πc
38,000/32,000
11
Length Diameter Weight
(in)
(in)
(lb)
α
280
47.75
7,000
NA
28,000/17,700 13.85
236.3
54.2
5,220
NA
F110-220
23,830/14,670
25
191.2
46.5
3,200
0.6
TF30-P-111
25,100/14,560
21.8
241.7
49
3,999
0.73
J93
Mattingly, D. Jack, “Elements of Gas Turbine Propulsion,” 1 st ed., McGraw-Hill, Inc.,
1996, pp.240-265
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SFC vs. Mach
Cruise Altitude (30,000ft)
1.3
1.2
SFC [(lbm/hr)/lbf]
1.1
1
0.9
0.8
0.7
Olympus 593 AB off
J93 AB off
F110-220 AB off
TF30-P-111 AB off
M = 0.8
0.6
0.5
0.4
0
0.5
1
1.5
M
2
2.5
3
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SFC vs. Mach
Dash Altitude (50,000ft)
2
Olympus 593 AB on
J93 AB on
F110-220 AB on
TF30-P-111 AB on
M = 2.5
1.9
SFC [(lbm/hr)/lbf]
1.8
1.7
1.6
1.5
1.4
1.3
1.2
0
0.5
1
1.5
M
2
2.5
3
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SFC vs. Altitude
M = 0.8 with Altitude = 30,000 ft
Olympus 593 AB off
J93 AB off
F110-220 AB off
TF30-P-111 AB off
Altitude = 30,000 ft
1.5
SFC [(lbm/hr)/lbf]
1.3
1.1
0.9
0.7
0.5
0.3
0
10000
20000
30000
Altitude (ft)
40000
50000
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SFC vs. Altitude
M = 2.5 with Altitude = 50,000 ft
2.5
Olympus 593 AB on
J93 AB on
F110-220 AB on
TF30-P-111 AB on
Altitude = 50,000 ft
SFC [(lbm/hr)/lbf]
2.3
2.1
1.9
1.7
1.5
1.3
0
10000
20000
30000
Altitude (ft)
40000
50000
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WHICH IS BEST?
•
Turbojets with AB
- Higher thrust
- SFC is low in dash
with AB
• Turbofan with AB
- Lower thrust
- SFC is low in cruise
without AB
Turbojets with AB are better
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FINAL DECISION
Max/Normal Thrust
(lb)
Olympus 593
J93
38,000/32,000
28,000/17,700
SFC
[(lbm/hr)/lbf ]
Cruise
1.169
0.934
Dash
1.4
1.45
J93 Turbojet with Afterburner is BEST!
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FURTHER ANALYSIS
• Find more recent engines
- Turbojets and Turbofans with AB
• Design Fuel System and Fuel Tanks
• General Propulsion System Integration Losses
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STABILITY & CONTROL
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
11
14
17
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DESIGN CONCEPT 1
• Cons
-Canards are not as
common as aft tails
-May need high lift airfoil
for canard
-Difficulty in using flaps
• Pros
-Canards can be made to
stall before wing
-Enhanced roll rate from
elevons
Rank : 2nd
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DESIGN CONCEPT 2
• Cons
-Flying wings inherently
unstable
-Requires complex
reflexed trailing edge for
static stability (inefficient)
-May require automatic
flight control systems
-Complicated wing
planforms with varying
chords and twist to
achieve restoring moment
Rank : 3rd
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DESIGN CONCEPT 3
• Cons
-V-Tail requires a more
complex control system
-Significant flight testing
needed to program V-Tail
• Pros
-Aft tail is a time tested
design
-Plenty of historical data to
compare design space
Rank : 1st
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INITIAL SIZING
• Auxiliary Lifting Surfaces
- Used historical tail volumes of Large cargo/transport
aircraft
- Fin and canard airfoil is NACA 0012
• Control Surface Sizing
- MILSPEC roll rate for Class III aircraft is 30 degrees in
1.5 seconds
- Initial sizing based on historical data and guidelines
• Final Sizing based on dynamic analysis
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STATIC MARGIN
• Neutral point calculation completed to determine
acceptable CG range
• 5-10% Static Margin for Large Bomber/Cargo Aircraft
• Canard experiences upwash as opposed to tail which
experiences downwash
Static Margin
Cruise (M=0.8) Dash (M=2.5)
Neutral Point
90.7 ft
91.9 ft
5% SM
87.1 ft
88.3 ft
10% SM
83.5 ft
84.7 ft
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LONGITUDINAL STATIC STABILITY
• Longitudinal stability most vital to airplane
• Placing CG ahead of neutral point satisfies one of two
conditions for stability
• Must check to see that CM0 is greater than zero
• Assumed virtually no CG shift from rocket release
Longitudinal Static Stability
CM0
CMa
CG
Cruise
0.006
Dash
0.006
-0.12 per rad -0.1154 per rad
87.1 ft
87.1 ft
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TRIM ANALYSIS
• Used graphical method rather than iterative
computational process.
• Trim analysis shows aircraft can be trimmed for many
different CL.
• Subsonic and supersonic trim very similar due to
comparable CMa .
• Positive elevon deflection produces upload on canard.
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TRIM ANALYSIS
Trim Analysis for Subsonic Cruise
0.08
de = 0 deg
0.06
de = -5 deg
0.04
de = 20 deg
CMcg
0.02
0
-0.5
-0.02
0
0.5
1
1.5
2
-0.04
-0.06
-0.08
-0.1
CL Total
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LATERAL STATIC STABILITY
• Coupled analysis on roll and yaw
• Meets the typical yaw moment derivative values as
described by Raymer.
Lateral-Directional Stability Derivatives
CNb (per rad)
Cruise
0.094
Dash
0.127
Clb (per rad)
-0.137
-0.091
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FURTHER ANALYSIS
• Must examine dynamic stability and control
characteristics
• Investigate high lift airfoils for canard
• Flexibility Effects
• Engine out analysis
• Ground effects
• Adverse yaw and differential control surface inputs
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STRUCTURES
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
11
14
17
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STRUCTURAL REQUIREMENT
• Speed – Mach 2.5
• Altitude – 50,000 ft
- The speed and the altitude requirements yield:
- Kinetic heating ranges from -25ºF to 450ºF
- Thermal cycling under moisture and radiation impact
• Payload – 59,960 lb
Must have strong mounts for the payload
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STRUCTURAL SELECTION
Selection criteria:
• Feasibility of new concepts
• Structural strength
• Minimum Weight
• Ease of manufacturing
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STRUCTURAL SELECTION
Design Concept 1
Pros
• Delta wing
• straight linear structures
(spars)
• enough room for fuel,
landing gear, and structure
Cons
• Curved longerons for the
fuselage
• Difficulty placing the rocket
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STRUCTURAL SELECTION
Design Concept 2
Pros
• Simple geometry for spars
and ribs (straight path)
• Smaller structure (light
weight)
• Weight is distributed along
the span of the wing
Cons
• Large cutouts for landing
gear (not enough space for
both spars and cutout)
• Extra bulkhead needed for
engine mounts
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STRUCTURAL SELECTION
Design Concept 3
Pros
• Engine inlet structure
supports the wing loading
• Bulkhead in the aft fuselage
shares the load with engines
and landing gears
Cons
• Excessive structure
(fuselage)
• Wing loading concentrated
at the smaller wing root than
delta wing
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MATERIALS
Titanium
Stainless Steel
Honeycomb
Steel
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BULKHEADS AND LOAD PATH
Rocket Mounts
Forward
Bulkhead
Landing Gear
Mounts
Engine Mounts
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V-n DIAGRAM
8
Upper Stall
Lower Stall
Upper Gust
Lower Gust
6
4
n
Vdive = 3136fps
n=+7
2
0
0
500
1000
1500
2000
2500
3000
3500
-2
n=-3
-4
V (fps)
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CONCLUSION AND
FUTURE ANALYSIS
• Conclusion
- Delta wing structure was chosen
- No complex composites were used to lower the
cost
• Future Analysis
- Finite Element Method (FEM) analysis should be
conducted
- Investigate the stress of the rocket attachment fittings
- Design landing gear
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COSTS
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CONCEPT SELECTION
Concept Selection Criteria
Criterion
Design Concept 1 Design Concept 2 Design Concept 3
Structure
1
3
2
Aerodynamics
1
2
3
GTOW
2
1
3
Stability
2
3
1
Fuel Consumption
2
1
3
Propulsion
1
3
2
Costs
2
1
3
Total
11
14
17
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COST ANALYSIS
• Used RAND DAPCA IV Model
- Find approximate unit price
• Hours needed: Engineering, Tooling, Manufacturing
• Cost estimated: Develop, Flight Test, Manufacturing,
Material, Engineering production
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DESIGN CONCEPT 1
• We: 109,526 lbs
• Velocity: M2.5 at 50,000 ft
– 1433 Knots
• Number Produced (Q): 5
• FTA:3
• Neng:40
• Thrust max: 28,000 lbs
– GE J93 Turbojet w/ AB
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DESIGN CONCEPT 2
• We: 96,627 lbs
• Velocity: M2.5 at 50,000 ft
– 1433 Knots
• Number Produced (Q): 5
• FTA:3
• Neng:40
• Thrust max: 28,000 lbs
– GE J93 Turbojet w/ AB
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DESIGN CONCEPT 3
• We: 118,274 lbs
• Velocity: M2.5 at 50,000 ft
– 1433 Knots
• Number Produced (Q): 5
• FTA:3
• Neng:40
• Thrust max: 28,000 lbs
– GE J93 Turbojet w/ AB
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ESTIMATED COST
Costs
Design Concept
1
Design Concept
2
Design Concept
3
Total Cost
(5 AC)
$8,954,080,073
$8,619,522,958
$9,851,267,581
Total Cost
(Individual)
$1,790,816,015
$1,723,904,592
$1,970,253,516
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DESIGN 1 - OPTIMUS
•
•
•
•
Moderate expensive aircraft to build.
Best performance capabilities for the best price
Based off a similar design
Marketable
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CONCEPT/OPS
•
•
•
•
•
Base: Kennedy Space Center, Cape Canaveral, FL
Fly out east of KSC
Captive on top delivery
γ = 25 °(3-M)
γ = 12.5 °
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CONCLUSION
• Cost Estimation seems reasonable
• Moderately expensive out of the three
• Much cheaper than the traditional launches from earth
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CONCLUSION & FUTURE
PLANS
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CONCLUSION & FUTURE PLANS
• Highly versatile delivery aircraft and favored in terms of
Structure & Stability
• Optimus meets the RFP
• Optimus will provide an alternative method to deliver rockets
into orbit
• Promote this idea to potential buyers, hopefully expanding
the market for this innovative method to launch rockets into
space.
• Decrease thrust requirement by reducing drag. Lower cost.
Will be looked into next semester.
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ead aeronautics え八팔
REFERENCES
• Jenkinson, L., Civil Jet Aircraft Design, AIAA, 1999
• Mattingly, D. Jack, “Elements of Gas Turbine Propulsion,” 1 st
ed., McGraw-Hill, Inc., 1996, pp.240-265.
• McCormick, B.W., “Static Stability and Control,”
Aerodynamics, Aeronautics, and Flight Mechanics, 2nd ed,
Wiley, New York, 1995, pp. 473-534.
• Peter, St. James, “The History of Aircraft Gas Turbine Engine
Development in the United States … A Tradition of Excellence”
1 st ed., The International Gas Turbine Institute of ASME., 1999,
pp.430-569
• Raymer, D.P., Aircraft Design: A Conceptual Approach, AIAA
Education Series, 2002
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Questions?
93