Conceptual
Design
Review
presented by:
• XG International
Gihun Bae - Joe Blake - Jung Hoon Choi - Jack Geerer - Jean Gong – Sang Jin Kim Mike McCarthy - Nick Oschman - Bryce Petersen - Lawrence Raoux - Hwan Song
Outline of Contents
I. Mission Statement
II. Design Mission/
Requirements
III. “Best” Aircraft Concept
IV. Sizing, Carpet Plots
V. Design Trade-offs
VI. Aerodynamics
VII. Performance
VIII. Propulsion
IX. Structure
X. Weights
XI. Stability/Control
XII. Noise
XIII. Cost
XIV. Summary
2
Mission Statement
Develop an environmentally-sensitive aircraft which
will provide our customers with a 21st-century
transportation system that combines speed,
comfort, and convenience while meeting NASA’s N+2
criteria.
3
Design Requirements
• Noise (dB)
– 42 dB decrease in noise
• NOx Emissions
– 75% reduction in emissions
• Aircraft Fuel Burn
– 40% lower TSFC
• Airport Field Length
– 50% shorter distance to
takeoff
**Values for NASA N+2 protocol are found in the Opportunity
Statement**
NASA ‘s Subsonic Fixed Wing Project Requirements.
4
Previous vs. Final Models
Previous
Final
5
Previous vs. Current
Previous
Current
•
•
•
•
• 3 Turbofans
• No Canard
• Cruciform Tail
2 Turboprops/UDF
1 Turbofan
Canard
T-Tail
Previous
lbf
ft 2
Current
97.5
lbf
ft 2
Wing Loading
82
Aspect Ratio
9
7.8
Thrust-to-Weight
0.3
0.33
Wing Sweep Angle
35°
28.13°
6
Previous vs. Current - Justification
• Removal of UDF: Lack of historical data
Noise will exceed regulations
• Turbofan vs. Turboprop: Faster speed
• Cruciform vs. T-tail: Reduce structure weight
• Engine placement: Reduce structure weight
(pylons, nacelles)
• Removal of Canard: Weight increase overrides
the benefits
7
“Best” Aircraft Concept
Cruciform
Tail
Turbofan
Engine
Duct
Winglet
Solar Films
8
“Best” Aircraft Concept
3rd Engine
9
“Best” Aircraft Concept
•
•
•
•
3 Turbofan engines
2 Outer engines for cruise
Cruciform Horizontal Stabilizer
Dropped canard configuration
Important Specifications
lbf
ft 2
Wing Loading
97.5
Aspect Ratio
7.8
Thrust-to-Weight
0.33
Wing Sweep Angle
28.13°
10
“Best” Aircraft Concept
Advanced Concepts
a. Solar Panels – Powers cabin electronics
b. 3 Engines – Maximizes fuel efficiency during cruise
– Reduces takeoff distance
– Safer for 1-engine-out condition
c. Closable duct – Reduces drag of the duct that might
be produced when the engine is not
used.
11
Sizing Code
• Used Cargo/Transport
Weights from Raymer’s
• Used Excel Spreadsheet
• 6 Different Sections
a)
Main
i. Fuselage
ii. Wing
iii. Engine
b)
c)
d)
e)
f)
Geometry
Constraint Diagram
Weight
Airfoil
Mission Detail
12
Sizing - Assumptions
Performance Specs
Value
CLmax
1.6
(L/D)max
9.3
We/WO
0.714
SFCcruise
0.5 /hr
SFCloiter
0.4 /hr
t/c
0.0158
Sweep Angle (Λ)
28.13°
Taper Ratio (λ)
0.5
e
0.8
Vcruise
710 ft/s
Vstall
223 ft/s
Vtake-off
245 ft/s
Vapproach
280 ft/s
13
Sizing – Drag Prediction
•
•
•
•
•
CD = CDP + CDi + Cmisc + Cw
CD = Parasite Drag Coefficient +
Induced Drag Coefficient
CDmisc and CDw are assumed to be zero.
CDi = Induced drag coefficient =
Parasite drag calculated from sizing code
14
Sizing – Tail
The rudder and ailerons are based on conventional
business jet values (Raymer).
Rudder Dimensions
Aileron Dimensions
Span
4.95 ft
Span
9.6 ft
Chord
2.12 ft
Chord
1.33 ft
Planform Area
10.5 ft2
Planform Area
12.73 ft2
Aspect Ratio
2.33
Aspect Ratio
7.22
15
Sizing - Validation
Bombardier Challenger 300 Specification (XG Endeavour)
•
•
•
•
•
Range : 3560 nmi (3700 nmi)
Passenger number: 9 (9)
Crew Number : 2 (2)
Cruise Mach Number : 0.8 (0.8)
Service Ceiling : 45000 ft (45000 ft)
16
Sizing - Validation
• Weights based on the sizing code
a) Empty Weight = 17500lb
b) Fuel Weight = 14000lb
c) Total Weight = 34400lb
• Actual Weights of Bombardier Challenger 300
a) Empty Weight = 18500b
b) Fuel Weight = 14100lb
c) Total Weight = 35400lb
• Fudge Factor
17
Design Trade-offs
Carpet Plot
• Based off of calculations in the constraint diagram
• Constraints vs. Wing Loading
1. Gross Weight
2. 2g Maneuver
3. Takeoff Ground roll
4. Landing Ground roll
18
Design Trade-offs
Carpet Plot
Carpet Plot
22000,00
21800,00
21600,00
Gross Weight
21400,00
T/W=0.3
T/W=0.4
21200,00
T/W=0.5
W/S=70
21000,00
W/S=80
20800,00
W/S=90
2g maneuver
20600,00
takeoff ground roll
20400,00
Landing ground roll
20200,00
20000,00
70
75
80
85
90
95
100
105
110
115
120
Wing Loading
19
Design Trade-offs
Pros
Cons
Cruciform Tail
Aft fuselage engine
Increase weight
Solar Film
More Engine Efficiency
Increase empty weight
3 Engines
Safer 1 engine-out situation
Heavier empty weight
3 Engines Cont’d Better fuel efficiency at cruise
More maintenance cost
20
Design Trade-offs
Cabin Layout
7 ft Cabin
Increase in Drag
More head Room
6.5 ft Cabin
Less Drag
21
Three Views
Dimensions
22”
Wing Leading Edge
36”
Tail leading edge
37”
Vertical Stabilizer
38”
Tail Mounted Engine
40”
Center Engine
50”
Total aircraft length
Internal Layout
Avionics
compartment
and nose
landing gear
housing
Fuel
Tank
Enlarged
equipment
compartment:
Fuel pump and reservoir
Duct
Engine
Equipment (APU, AC,
etc.)
Wheel
housing
Equipment
compartment
Cabin Layout
Airfoil Selection
NACA 2414
Drag Polar
Shape
www.worldofkrauss.com
26
Airfoil Selection
NACA 2414
Parameter
Values
CLmax
1.276
Angle CLmax
15°
(L/D)max
48.157
Angle (L/D)max
6.5°
Angle Stall
6.5°
Angle Zero-lift
-2°
www.worldofkrauss.com
27
Drag Polar
Drag Polar
1,5
Coefficient of Lift
1
0,5
Cruise
0
0
0,02
0,04
0,06
0,08
0,1
0,12
Takeoff
Landing
-0,5
-1
-1,5
Coefficient of Drag
28
Performance
•
Diagram provides visual n
V-n Diagram
understanding of wing
loading with increasing
velocity.
Created V-n diagram
using maximum wing
loading of +3.333Gs and -1G (using a 1.5 SF).
V is velocity represented in ft/s.
n is load factor in Gs.
4
3,5
3
2,5
2
1,5
1
V
0,5
•
0
-0,5 0
200
400
600
800
1000
1200
-1
-1,5
•
•
29
Performance
Performance Specification
Values (units)
Take Off Distance
4000 (ft)
Landing Distance
2500 (ft)
Best Range
3700 (nmi)
Best Endurance Velocity
710 (ft/s)
Stall Speed
220 (ft/s)
Stall Speed @ Max. (+) Load Factor
400 (ft/s)
Stall Speed @ Max. (-) Load Factor
220 (ft/s)
Dive Speed
1020 (ft/s)
Cruise Speed
710 (ft/s)
30
Propulsion
Engine Description
•
•
•
For the final design 3 turbofan engines will be used, one capable of producing
6,800 pounds of thrust, and two that produce 2,000 pounds of thrust.
These engines are modeled from the HF120 turbofan which is manufactured
by GE Honda Aero Engines.
Below are a picture of the engine, and a schematic showing dimensions. Both
are for the 2000 pound thrust version.
31
Propulsion
• The 2000 pound thrust model has the following
characteristics:
– Bypass Ratio=
– Takeoff Thrust=2050 lbs
– Compressor pressure ratio=24
• The 6800 pound thrust model has the following
characteristics:
– Bypass Ratio=
– Takeoff Thrust =6800 lbs
– Compressor pressure ratio=26
32
Propulsion
• Assumptions for Engine Modeling:
– The baseline model was scaled to meet the mission’s
thrust requirements using an Excel sizing routine.
– Technological improvement factors were used to
determine performance in 2020.
– Since the 2000 pound thrust model did not need to be
scaled, available data was used in calculations and no
efficiencies were needed. To scale the larger engine the
sizing routine was used to determine the appropriate
weight given the thrust required.
33
Propulsion
1. The following graphs show the thrust available from the
engines and the thrust required to power the aircraft
versus velocity for several altitudes :
Thrust vs Velocity at Takeoff
12 000
10 000
Thrust (lb)
8 000
6 000
Thrust Available
Thrust Required
4 000
2 000
0
0
20
40
60
80
100
120
140
Velocity (mph)
34
Propulsion
12000
10000
8000
6000
4000
2000
0
Thrust vs. Velocity
(35000 ft, climbing)
Thrust Available
Thrust (lb)
Thrust (lb)
Thrust vs. Velocity
(15000 ft, climbing)
Thrust Required
0
200
400
12000
10000
8000
6000
4000
2000
0
600
Thrust Available
Thrust Required
0
100
200
Velocity (mph)
Thrust Available
Thrust Required
400
Velocity (mph)
500
600
Thrust vs. Velocity
(45000 ft, cruise)
600
Thrust (lb)
Thrust (lb)
12000
10000
8000
6000
4000
2000
0
200
400
Velocity (mph)
Thrust vs. Velocity
(25000 ft, climbing)
0
300
12000
10000
8000
6000
4000
2000
0
Thrust Available
Thrust Required
0
200
400
600
800
Thrust Available (2
engines)
Velocity (mph)
35
Load path overview
Load path estimation
Load path overview
A closer inspection
• Main formers, ribs,
stringers and
longerons made of
TiAl
• Additional
components
added to reenforce strength of
the structure.
Main
supports
Additional
components
Wing intersection
• Wings
– Common Low Mount
– Through fuselage for stability
– Uses two main aft formers of the aircraft
• Stabilizers
– High mount
Engines
• Innovative locations of engines
• Tail mounted Engines.
– Requires that the tail be mount
to the fuselage
• The ‘3rd’ Engines
– Placed in line with the
centerline of the aircraft
to avoid pitching moment.
Landing Gear
• Retracts inward and is stored
under the fuselage and wing
when in flight.
• Placed on wings to increase yaw
stability during taxi
Side retracting landing gear.
Landing Gear
• Located on the intersection of
the main stringer and a rib.
• Stringer is supported on the
frame of the craft where the CG
is located.
Far right, side view of
landing gear relative to
location of center of
gravity. Near right,
view from below the
craft.
41
Material Selection
•
•
•
•
Fiber Glass
Composites
Thermoplastics
Aluminum based
Alloys
GE’s, GEnx engine currently uses
an lightweight Aluminum based
alloy, Gamma Titanium Aluminide.
Material Selection
Aluminum based Alloys
• Nickel Aluminide
– Extremely high strength to weight ratio
– Ductile
– Common in gas turbines and get engines
• Titanium Aluminide
– Intermetallic chemical compound
– Resistant to oxidation and heat
– Low ductility
• Gamma Titanium Aluminides
– Currently focused on use in engines.
– Can withstand temperatures from 600oC and higher
– Half the density of steel or nickel based alloys
Weights
Structures
Weight (lb)
Loc (ft)
Mom (ftlb)
Wing
1150
31.1
36044
Horizontal Tail
440
40.7
Vertical tail
400
Fuselage
Weight(lb)
Loc (ft)
Mom (ftlb)
Flight Cont
460
11
5016
18180
APU Installed
200
43
8510
38.8
15950
Instruments
110
11
1160
4100
44.5
184664
Hydraulics
100
31.1
3200
Main Landing
800
34.1
27250
Electrical
590
31.1
18260
Nose Landing
210
11
2355
Avionics
1200
7.3
8470
Propulsion
Weight (lb)
Loc (ft)
Mom (ft-lb)
Furnishing
270
31.1
8395
A/C
200
40.7
8470
1800
43
80060
Anti-ice
60
31.1
1860
Engine Cont
40
43
1740
Handling Gear
9
31.1
280
Starter
85
43
3640
Cargo / Seats
700
31.1
21800
Fuel System
280
31.1
8830
Total Wempty
16,800
Weight (lb)
Loc (ft)
Mom (ftlb)
Loc (ft)
Mom (ftlb)
Pilots
240
20
4680
Passengers
1600
N/A
N/A
Engine
Loads
Equipment
Loads
Luggage
TOGW
Weight (lb)
240
29,800
44
CG Travel
CG Travel
27000
26000
25000
24000
Weight
23000
22000
21000
20000
19000
18000
17000
25
27
29
31
33
35
37
39
Location
Take Off Gear up
CG
Weight
27.97
26163
28.13
26011
Fuselage
Tank
Wing
Tank
Gear
Down
Land
31.18
23468
31.89
22951
31.93
22916
31.93
22916
Reserve Passenger
Add
Crew Off Add Fuel Add Crew
Fuel
Off
Passenger
35.41
20669
36.51
18680
36.97
18446
28.49
23940
28.21
24174
27.97
26163
Min
Max
27.97
18446
36.97
26163
45
Stability/Control
• Control surfaces are sized to minimize weight and
drag while ensuring stability of the aircraft.
• Static Longitudinal Stability:
– 4% static margin calculated from the sizing code. This
makes the aircraft more responsive to pilot inputs.
– The center of gravity was determined to be positioned at
33 feet from the nose of the fuselage.
– The neutral point is thus 0.266 feet behind the c.g. (wing’s
mean chord length is 6.644 ft.)
46
Stability/Control
• Based on conventional business jet sizing values
(Raymer), we designed the elevator to be about 90%
of the tail span and 32% of the tail chord. Each
elevator thus has a chord length of 1.43 ft, a span of
10 ft, planform area of 14.3 ft2, and an aspect ratio of
about 7.
47
Stability/Control
Trim Diagrams
48
Stability/Control
• Potential Issues:
– ‘One-engine out’: In case one of the two aftfuselage engines were to go out, the turbofan at
the end of the fuselage can be turned on to
provide enough thrust to maintain cruise flight.
– ‘Cross-wind landing’: Sideslip technique used (i.e.
rudder/ailerons adjust aircraft’s heading in order
to keep the aircraft lined up with the runway until
touchdown).
49
Noise
• Smaller HF120 turbofan engines designed to
be fully stage IV compliant
50
Noise
• Larger 3rd engine housed in the aircraft body
reduces external noise
• Future propfan integration to feature Active
Vibration Control System, reduces internal
noise
– Deemed unnecessary for turbofan platform
• External noise is estimated using
combination of scaled engine and historical
data
51
Noise
• Aerodynamic noise comparable to similarlysized current aircraft
• Decreased time
to climb reduces
ground signature
during flyover
stage
Noise Certification Values according to ICAO Annex 16,
Volume I, Chapter 3
52
Cost Prediction
•
Calculated using Rand DAPCA IV model along with
information from Raymer’s text.
Type
Price
RTD&E + Flyaway Cost
$2.4 Billion
Production Cost
$15 Million
Profit Per Aircraft
$750,000
Breakeven Point
20 aircrafts
Production Run
160 in 5 years
53
Cost Prediction
Rates
Values
Depreciation
6.6% / year
Insurance
$30,000 / year
Crew
$230 / block hour
Fuel/Oil
$1158/flight hour
Maintenance
$764 / flight hour
DOC
$2274 / flight hour
DOC/Seat-Mile
$0.22
54
Cost Prediction
•
Miscellaneous Customer Costs
Type
Cost
Hangar
$80,000 / year
Training
$40,000 / year
Landing
$386 / landing
55
Summary
• Trans-Atlantic flight
• 12 Passenger luxury cabin
• 3 Turbofan Engines
56
Summary
Requirement
Maximum Mach Number
Empty Weight (lb)
Gross Weight (lb)
Takeoff Distance (ft)
Maximum Range (nmi)
Design Mission Range (nmi)
Noise (dB)
Seats
Volume Per Passenger (ft^3)
TSFC (% of avg)
N0x Emissions (% of avg.)
Target
0.85
18,500
28,000
3,400
3,700
3,700
42
10
65
55
25
Threshold
0.8
20,000
32,000
3,800
3,600
3,600
50
8
60
65
50
Endeavour XG
0.8
11,714
22,116
4,000
3,700
3,700
<48
9
60
65
50
Compliant
Yes
Yes
Yes
No
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Charge Time - 220V 80A* (hr)
2
4
1.5
Yes
Charge Time - 125V 15A** (hr)
3
5
4
Yes
Internal Systems Power (kWh)
5
6.5
8
No
57
Summary
Environmentally-sensitive business aircraft concept is a
plausible opportunity. However:
a) Meeting 40% reduction of fuel consumption is still a big
challenge
b) Difficult to meet all of NASA’s N+2 goals at once
c) With further research on UDF conducted to meet the
noise requirement , 40% reduction in fuel consumption
may be possible.
58
Summary
Areas in need of further research:
a) Catalytic reduction technology on the aircraft
b) Shorten takeoff distance
c) Reduce empty weight to increase fuel efficiency
59
Thank you
60
Sources
•
___ Gas Turbine Engines. Aviation Week & Space Technology Source Book 2009. p 118
•
___ GE Aviation. The GEnx Engine Family. Available online [http://www.geae.com/engines/commercial/
genx/combustor.html], 2010
•
___ GE Honda Aero Engines. Available online [http://www.gehonda.com], 2010
•
___ Calculating noise ICAO Annex 16, Volume I, Chapter 3.
•
Campbell, G.S., and Lahey, R.T.C., A survey of serious aircraft accidents involving fatigue fracture, Vol. 1
Fixed-wing aircraft, National Aeronautical Establishment, Canada. 1983
•
Christensen, R.M. Mechanics of Composite Materials. New York, John Wiley & Sons, 1979
•
Hoskin, B.C., and Baker, A.A., eds. Composite Materials for Aircraft Structures, New York: American Institute
of Aeronautics and Astronautics, Inc., 1984.
•
Martin, Christopher L.; Goswami, D. Yogi (2005). Solar Energy Pocket Reference. International Solar Energy
Society
•
Megson, T.H.G. Aircraft Structures for engineering students. Burlington MA: Butterworth-Heinemann. 2001
•
Kroo, Ilan. Stanford university. “Aircraft Structural Design”. Available online
[http://adg.stanford.edu/aa241/structures/structuraldesign.html] 2010.
61