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Optimization Framework for Design of Morphing Wings
Jian Yang, Raj Nangia & Jonathan Cooper
Department of Aerospace Engineering, University of Bristol, UK
&
John Simpson Fraunhofer IBP, Germany
AIAA Aviation 2014 Atlanta 16-20 July 2014
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Traditional Aircraft Design • Single point for the
design • Minimum weight • Maximum CL/CD • Suboptimal elsewhere
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Aircraft Morphing Morphing, when applied to aerospace vehicles, is a technology or set of technologies applied to a vehicle that allow its characteristics to be changed to achieve better performance or to allow the vehicle to complete tasks it could not otherwise do.
Jason Bowman, AFRL/VSSV
3
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Configuration Morphing • Change in planform
– Aircraft control
– Aircraft performance
• Change in mission
– High aspect-ratio glide
– Attack mode
4
MFX 1, Span & Sweep
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Configuration Morphing
5
Baseline
(Joshi et al., 2004, AIAA)
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Performance Morphing
• Change in structural properties
– Stiffness
– Camber
– Leading / trailing edge shape
• Aircraft control
• Aircraft performance
– Lift / drag
– Roll control
– Loads
6 NASA, Active Aeroelastic Wing (AAW)
NASA, Active Flexible Wing (AFW)
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Variable Stiffness Morphing
• Range of different adaptive stiffness methodologies to be explored
• Vary EIx, EIy, GJ, flexural axis using changes in internal structure
• Changes to spar orientation, rib position, spar cap position – used in several previous EU projects
– 3AS, SMorph, NOVEMOR
7 7
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Aims • Develop an inverse design approach
- optimise internal wing stiffness distribution - morphing capability.
• Meet optimal aerodynamic shape throughout the
flight envelope • Define the required changes in stiffness / elastic axis • Minimize the amount of morphing • Minimize the weight • Satisfy any other constraints – to be defined
progressively
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Inverse Design Approach
9
• Define required aerodynamic distributions
• Define required wing shapes – Twist / camber / bending
• Optimise weight to determine stiffness distributions
• Evaluate required stiffness changes
• Determine morphing concept(s) to achieve stiffness changes
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0 0.5 1 0
2
4
6
8
10
0
0.5
1
y
x/cchor
dwis
e lo
adin
g Wing Aerodynamic shape
Via Lifting Surface VL or DL (lattice) (Lamar) mean camber for minimum drag
wing
Aspect ratio (AR) 12
Taper ratio 0.5
Sweep angle 10 degrees
Semi-span 20 meters
• Planform dimensions
0 0.2 0.4 0.6 0.8 1
0
0.05
0.1
x/c
z/c
Mean camber lines
0 5 10 15 20
0
5
x
y
vortex panels
Camber lines required to achieve M0.74 loading
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Aerodynamic shape design
- Lifting Surface, Mach Effect
0 0.25 0.4 0.6 0.74 10
0.5
1
1.5
2
Mach number
Required C
L
CL vs Mach
• ΔCp distributions
0
0.5
1 0 0.2 0.4 0.6 0.8 1
0
0.5
1
1.5
2
2.5
3
y/sx/c
C
p M=0.74
00.5
1 0
0.5
10
1
2
3
y/sx/c
C
p M=0.6
00.5
1 0
0.5
1
0
1
2
3
y/s
x/c
C
p
M=0.4
Design for 4 different test cases – thin lifting surface
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Aerodynamic shape – Super-Critical (twist & Camber)
- Panel & Euler (AR=12, Ma=0.74)
0 5 10 15 20
0
5
x
y
vortex panels Non-D
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0
0.5
1 0 0.2 0.4 0.6 0.8 1
0
0.5
1
1.5
2
2.5
3
y/sx/c
Cp
Euler
Panel
Lifting
Surface x/s
Aerodynamic shape Different approaches, 0.5 CL
01
23 0
5
10
-2
-1
0
1
y
M0.74, pressure distribution
x
Cp
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Structural model
Cross section
𝑡𝐿
𝑡𝑅
𝑏𝑐
𝑑
𝑡𝐻
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Weight Optimisation Minimize weight
• Consider deflection/stress constraints & minimum gauge of
design variables.
• Use optimality criteria method
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Application: Aero design
16
camber ∆Cp Cp√x
Mach 0.74
Mach 0.6
Mach 0.4
Mach 0.25
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∆Cp (using M0.74 camber)
17
∆Cp
Cp√x
Mach 0.6 Mach 0.4 Mach 0.25
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Spanwise loadings - effect of M
18
Xcp
Enlarged
Non-D by CL
CLL c/cav
0.74M camber. Evaluated at different test cases
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Local AoA required to achieve reference loading of M0.74
with M0.74 camber
19
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∆Cp Distributions
before --- & after adding twist ---
20
Mach 0.25
-30%
Mach 0.6 Mach 0.4
∆Cp√x
∆Cp
Note -30%
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Structural Optimisation, 0.74M solution process
21
0 50
2
4
6
8
10
12
14
16
18
20
x
y
(b) (a)
1 2 3 4 5 6 7 8 9 103000
3500
4000
4500
5000
Iteration number
weig
ht
(kg)
Convergence
Initial design
Flex. axis varies as function of tR / tL
Weight Convergence - fast
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Predicted Jig shapes & Twist, different M
22
Mach 0.6
Mach 0.25
Mach 0.74
Mach 0.4
Using M0.74 camber Achieving M0.74 pressure dist using wingbox twist only
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Morphing required (Twist only example from half span)
23
GJ Elastic twist ∆GJ
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Morphing required (Twist & camber morphing example)
24
GJ Elastic twist ∆GJ
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Morphing required (Twist & camber morphing example)
25
GJ Elastic twist ∆GJ
GJ Elastic twist ∆GJ
Twist & camber
Twist
Note Strong Reduction
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Camber morphing case • TE Morph: to achieve desired spanwise loading and reduce
RBM
26
0 5 10 15 20
0
2
4
6
x
y
LE area
TE area
Wingbox area
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LE morphing effects
• LE Morph: to reduce leading edge suction and flow separation
27
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LE morphing
28
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Mean camber shape
• Variations from M 0.74 to M 0.25 case
29
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Conclusions • Multiple point design useful for enhancing aircraft
efficiency
• Adaptive wing / morphing concepts reviewed
• An inverse design approach to optimise the internal stiffness distribution & morphing capabilities
• Application to a regional aircraft wing
• Weight reductions achieved via structural optimisation
• Required morphing examined to meet the multi-design points
• Need a lot of twist morphing to achieve required shape over flight envelope
• Use of camber morphing reduces the amount of required twist
30 30
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Ongoing Work
• Continue with aerodynamic shape determination – Configuration / flight conditions (Panel, Euler)
– Winglet
– Controls
• Implement inverse approach on coupled-beam model, using more advanced Aero – Panel & Euler, Bodies
• Determine required stiffness variations in terms of morphing approaches
• Investigate reasonable morphing requirements
31 31
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0
1
2
30
510
-0.2
-0.1
0
0.1
0.2
yx
z
PANAIR Results
• M0.25, varying twist
0
1
2
3 0
5
10
-3
-2
-1
0
1
yx
Cp
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Euler results
33
Design M 0.74 Off-design, Mach 0.6