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

Traditional Aircraft Design • Single point for the

design • Minimum weight • Maximum CL/CD • Suboptimal elsewhere

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

Configuration Morphing • Change in planform

– Aircraft control

– Aircraft performance

• Change in mission

– High aspect-ratio glide

– Attack mode

4

MFX 1, Span & Sweep

Configuration Morphing

5

Baseline

(Joshi et al., 2004, AIAA)

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)

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

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

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

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

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

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

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

Structural model

Cross section

𝑡𝐿

𝑡𝑅

𝑏𝑐

𝑑

𝑡𝐻

Weight Optimisation Minimize weight

• Consider deflection/stress constraints & minimum gauge of

design variables.

• Use optimality criteria method

Application: Aero design

16

camber ∆Cp Cp√x

Mach 0.74

Mach 0.6

Mach 0.4

Mach 0.25

∆Cp (using M0.74 camber)

17

∆Cp

Cp√x

Mach 0.6 Mach 0.4 Mach 0.25

Spanwise loadings - effect of M

18

Xcp

Enlarged

Non-D by CL

CLL c/cav

0.74M camber. Evaluated at different test cases

Local AoA required to achieve reference loading of M0.74

with M0.74 camber

19

∆Cp Distributions

before --- & after adding twist ---

20

Mach 0.25

-30%

Mach 0.6 Mach 0.4

∆Cp√x

∆Cp

Note -30%

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

Predicted Jig shapes & Twist, different M

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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

Morphing required (Twist only example from half span)

23

GJ Elastic twist ∆GJ

Morphing required (Twist & camber morphing example)

24

GJ Elastic twist ∆GJ

Morphing required (Twist & camber morphing example)

25

GJ Elastic twist ∆GJ

GJ Elastic twist ∆GJ

Twist & camber

Twist

Note Strong Reduction

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

LE morphing effects

• LE Morph: to reduce leading edge suction and flow separation

27

LE morphing

28

Mean camber shape

• Variations from M 0.74 to M 0.25 case

29

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

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

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

Euler results

33

Design M 0.74 Off-design, Mach 0.6