MORPHING CHARACTERISTICS OF CHIRAL CORE AIRFOILSruzzene.gatech.edu/LabWeb/research/chiral_applications/Presentation_auxetics_2006.pdfAtlanta GA - USA Chrystel Remillat, Fabrizio Scarpa,

Auxetics 2006 4 - 6 September 2006, University of ExeterExeter, UK 1

Alessandro Spadoni, Massimo RuzzeneSchool of Aerospace EngineeringGeorgia Institute of Technology

Atlanta GA - USA

Chrystel Remillat, Fabrizio Scarpa, Kevin PotterDepartment of Aerospace Engineering

University of BristolBristol, UK

MORPHING CHARACTERISTICS OF CHIRAL CORE AIRFOILS


Objective:Application of chiral geometry for the design of an airfoil with morphing characteristics;

Motivation:

- Morphing is an effective way to enhance performance of wings and rotor blades:

• improve flow conditions, • minimize drag,• eliminate the need for flap mechanisms,• improve handling and control of aircraft.

- A chiral structure provides compliance and allows continuous deformation of airfoil.

Objectives & Motivation


• Negative Poisson’s ratio:

Estimated νxy ~-0.9;

• High in-plane shear modulus:

As a result of νxy being close to -1;

• Unique deformation mechanism:

Allows large deflections, while material remains in elastic range;

• Design flexibility:

Property of the assembly strongly depends on characteristic parameters of chiral geometry (r,R,L,θ);

)1(2 xyx

xy

EGυ+

=

r

θ

R

Lt

Chiral Geometry


Objective:Passive changes of mean camber line of the wing in response to changing incident airflow speeds

Reduction of the TAB ANGLE in response to increase in car speed;Corresponding elastic deformations are recovered when the speed decreases, so that the wing tip moves back upward;

Increase in MAXIMUM SPEED and better HANDLING

Previous Work:Racecar wing with passive adaptive capabilities(*)

EPPLER 420 Airfoil with 300 mm chord

5 mm

(*) Bornengo, D., Scarpa, F., Remillat, C. "Morphing airfoil concept with chiral core structure,“IMechE J. Aer. Eng., 2005.

Homogeneous material with homogenized chiral properties


Rationale

x

y

z

Τ = GJdxdθ

M = EIxx 22

dxwd

• Given a Poisson’s ratio of ≈ -1, G ≈ ∞ Wing does not require close section to carry torsional loads;

• Large decambering deformations can be sustained within the elastic range of the constitutive material;

• The core allows for continuous deformations which are important to maintain aerodynamic efficiency;

• The airfoil core can be tuned to achieve different functionalities by changing core geometric parameters;


Outline

U. Of BRISTOLGATECH

• Present two designs resulting from parallel developments; • Discuss results and provide recommendations for future research;• Show related research on chiral structures.


Development of GATech’s chiral airfoil


Eppler 420 series

• Eppler 420’s camber provides significant lift at low speed;

• Assembly is intended to conform as dictated by flow conditions camber decreases as velocity and lift increase

• Lift-induced drag decreases with velocity;

• Configuration can be modified for active morphing applications (active camber control).

Overview• The airfoil hosts a MACROSCOPIC chiral structure;• Evaluation of compliance characteristics:

– Numerical analysis with steady aerodynamic loads;– Experimental investigation using static loads.


Design• Chiral core is MAPPED into airfoil;• Influence of number of cells and L/R ratio is investigated;• All other parameters are kept constant.

Predefined layout

Mapped to conformTo airfoil profile

Note:• Core mapping facilitates meshing;• Core periodicity is lost


STRUCTURAL MODEL

Numerical AnalysisCFD-FEA Coupled models

• Steady-state aerodynamic loads are defined by specified flow conditions;

• Airfoil performance is investigated through weakly-coupled structural and computational fluid dynamics (CFD) models;

• Air-loads and corresponding displacements are iteratively passed to the structural and fluid codes respectively until convergence is achieved

x,u2u1

y,v2

φ2

v1

-φ1zz

Timoshenko beam element

u1v1

u2

u3

v3

φ3

Isoparametric planar element

E = 71 GPadensity = 2700 Kg/m3ν = 0.33 wall t = 0.76 mmOut-of-plane t = 2.54 cm

E = 710 MPadensity = 2700 Kg/m3ν = 0.33 wall t = 0.76 mmOut-of-plane t = 2.54 cm

Soft skin

Fixed


Unstructured triangular mesh, using finite-volume Galerkin NSC2KE(*)

Inflow points

Outflow pointsCharacteristic technique(*)

Wall boundary, 0=⋅nu rr Tangency condition

• Minimum element size 0.001 chord, at leading and trailing edges

• Wake elements near trailing edge have a size of 0.001 chord

• Airfoil element size is linearly relaxed away from LE and TE by 5

• Wake element size linearly relaxed up to outflow boundary

• Inflow and Outflow boundaries are at least 6 chords away from airfoil

• Total of approximately 11,500 fluid elements.•Mohammadi, B., "Fluid Dynamics Computation with NSC2KE," INRIA Report 0164, 1994.

Steady-state Fluid Model

http://www.inria.fr/rrrt/rt-0164.html

http://www.inria.fr/rrrt/rt-0164.html


0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

1.3

1.4

1.5

1.6

1.7

1.8

1.9

Cl

Number of Iterations 0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

10-5

10-4

10-3

10-2

10-1

100

log 1

0(N

orm

aliz

ed L

2 N

orm

)Number of Iterations

Analysis of convergence of fluid model

8000 iterations are used, as reasonable residual reduction target is 4 orders of

magnitude

• Sea level conditions• Mach 0.45• angle of attack 2°

• No gravity • Euler time stepping (steady-state solution)

Lift Coefficient Residual norm

Steady-state Fluid Model


Structural mesh of airfoil and chiral core

are obtained

ANSYS

Beginning of convergence iterations MATLAB

Flow field region is discretized with an

unstructured triangular-element mesh

Flow field is solved forpressure, density and

velocityNSC2KE

Equilibrium is solvedimposing aerodynamicloads on chiral-core

airfoil

MATLAB

Deformed airfoil is splinedand new profile is

computed MATLAB

Initial Iteration

Beginning of next Iteration

Solution

YesNo

L/RNumber of core cellsMaterial propertiesFree-stream conditionsAngle of attack

New Iteration converges with previous one?

uuL T ⋅=

31 101

−− ×


TE displacement 1.3 cm

L/R = 0.6 L/R = 0.9

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 10.006

0.007

0.008

0.009

0.01

0.011

0.012

0.013

0.014

L/R

Trai

ling-

Edg

e D

ispl

acem

ent [

m]

Results: 3 cellsInfluence of L/R ratio


Results: 2 cells Influence of L/R ratio

L/R = 0.6 L/R = 0.95

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 10

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08Tr

ailin

g-E

dge

Dis

plac

emen

t, [m

]

L/R, [m]

3 cells


L/R = 0.70 L/R = 0.90

Large node radius facilitates bending deformation of the ligaments,which is a main contributor of overall deformation

Results:Influence of L/R ratio


Lift vs. Mach number

0.15 0.2 0.25 0.3 0.35 0.4 0.451

1.1

1.2

1.3

1.4

1.5

1.6

1.7

1.8

1.9

2

M

Cl

Lift coefficient

L/R = 0.60

L/R = 0.90

0.15 0.2 0.25 0.3 0.35 0.4 0.450

0.5

1

1.5

2

2.5x 104

L/R = 0.60

L/R = 0.90

M

Lift

Ll

2nd CHIRAL-CORE AIRFOIL CONFIGURATION


Experimental ValidationStatic compliance tests

L/R = 0.602-y cells

L/R = 0.603-y cells

L/R = 0.943-y cells

Water-jet manufacturing


Out-of-plane thickness = 2.00 cm

0.7 m

2.54 cm

0.36 m

0.7 m

2.54 cm

0.36 m

0.7 m

2.54 cm

0.36 m

r

t

r

t

t

r

r = 0.67 cm t = 0.65 mm

r = 1.07 cm t = 0.65 mm

r = 0.3 cm t = 0.65 mm

E = 71 GPadensity = 2700 Kg/m3ν = 0.33

Material: Aluminum 6061 T651

Out-of-plane thickness and chord dimensions were chosen given manufacturing restrictions



Strain gages

Clamped b.c.

LVDT

Strain gage conditioner and amplifierVishay Measurement Group 2100 System

Iotech 2000 series acquisition board

Experimental set-up

Strain-gage locations are chosenbased on a preliminary FE analysis



Static compliance testsL/R = 0.60, 3 unit cells across airfoil thickness

1

2

34

56

0 5 10 150

10

20

30

40

50

60

T.E. disp, [mm]

App

lied

load

, [N

]

P

Four cuts have been used to core and skin compliance

0 5 10 150

100

200

stra

in g

auge

1 μ ε

T.E. disp, [mm]0 5 10 15

0

1000

2000

3000

stra

in g

auge

2 μ ε

T.E. disp, [mm]

0 5 10 150

200

400

600

stra

in g

auge

3 μ ε

T.E. disp, [mm]0 5 10 15

0

50

100

stra

in g

auge

4 μ ε

T.E. disp, [mm]

0 5 10 150

1000

2000

stra

in g

auge

5 μ ε

T.E. disp, [mm]0 5 10 15

-50

0

50

100

stra

in g

auge

6 μ ε

T.E. disp, [mm]


1 2 35

4

0 2 4 6 8 10 12 14 16 18 200

5

10

15

20

25

T.E. disp, [mm]

App

lied

load

, [N

]

0 5 10 15 200

500

1000

stra

in g

auge

1 μ ε

T.E. disp, [mm]0 5 10 15 20

0

500

1000

1500

stra

in g

auge

2 μ ε

T.E. disp, [mm]

0 5 10 15 200

500

1000

1500

stra

in g

auge

3 μ ε

T.E. disp, [mm]0 5 10 15 20

0

500

1000

stra

in g

auge

4 μ ε

T.E. disp, [mm]

0 5 10 15 200

2000

4000

stra

in g

auge

5 μ ε

T.E. disp, [mm]

P



Non-linear FE models suggest more compliance within elastic range

0 10 20 30 40 50 60 70 80 90 1000

10

20

30

40

50

60

70

80

90

T.E. disp, [mm]

App

lied

load

, [N

]Experimental results

Numerical results



4

1

32

5

0 2 4 6 8 10 120

10

20

30

40

50

60

70

80

90

100

T.E. disp, [mm]

App

lied

load

, [N

]

0 5 10 150

1000

2000

stra

in g

auge

1 μ ε

T.E. disp, [mm]0 5 10 15

0

500

1000

stra

in g

auge

2 μ ε

T.E. disp, [mm]

0 5 10 150

1000

2000

stra

in g

auge

3 μ ε

T.E. disp, [mm]0 5 10 15

0

500

1000

stra

in g

auge

4 μ ε

T.E. disp, [mm]

0 5 10 150

1000

2000

stra

in g

auge

5 μ ε

T.E. disp, [mm]



0 2 4 6 8 100

2

4

6

8

10

12

14

16

18

20

Applied Load, [N]

Trai

ling-

Edg

e di

spla

cem

ent,

[mm

]

3y cells, L/R = 0.942y cells, L/R = 0.603y cells, L/R = 0.60

• The core can be designed to achieve different compliance through a change in a limited

number of geometric (L/R);

• Significant decambering deformations can be sustained within the elastic range of the

constitutive material;

Static compliance testsSummary


Development of U. of Bristol’s chiral airfoil


Design Constraints

• Selective Laser Sintering for core manufacturing– Maximum core length is 0.2 m, chord 0.3 m;– Minimum thickness is 1mm for any part of the structure:

• Ligaments thickness is 1 mm;• Opted for solid nodes (nodes are stiff compared to the ligaments)• Only a two cell deep core would fit into the aerofoil whilst

maintaining a reasonable ligament aspect ratio.

• It was chosen not to have ligaments attached to the aerofoil skin, to avoid high point loads.

• A custom chiral core was created, joining only nodes onto the skin and following the curvature of the aerofoil.


Configuration

Nose:Stiff, made of pine wood

Skin:• 0.5mm glass fibre composite • Flexible but stiff enough to maintain the aerofoil shape and prevent surface

buckling.• Non-symmetric lay-up [0º,90º,+45º,-45º] facilitates conforming to airfoil shape

Chiral Honeycomb Core:• Material Polyamide Duraform. • Design attempts to keep L/r and R/r ratios uniform (uniform E)• Node radii (r) are decreased along the chord as the ligament length (L) was constrained by the taper of the aerofoil.

- Some of the internal cell angles, θ, differ from 30°- Initial chiral geometry has been used as a guideline.

Rubber Strip:Added to induce the chiral rotation(*)

(*) Bornengo, D., Scarpa, F., Remillat, C. "Morphing airfoil concept with chiral core structure,“ IMechE J. Aer. Eng., 2005.


Numerical analysis

Static Analysis:– Structural model is coupled with two flow solvers– An iterative process was required to achieve

convergence of the solution.– Ansys FE is used for the structure part;– Structural FE model is validated experimentally– Flow solvers:

• Inviscid– vortex lattice panel method coded in Matlab, assumes two-

dimensional (2-D) inviscid flow, and does not consider flow separation.

• Viscous (XFoil)– combines a vortex panel method with a boundary layer model to

provide a viscous analysis. It includes boundary layer growth, producing a more realistic pressure distribution, especially at higher angles.


Example of Results

• Viscous analysis predicts a much lower deflection:– pressure forces acting close to the trailing edge are reduced

due to the predicted boundary layer at high angles.

Graph of Tip Vertical Deflection Against Velocity at an Angle of 15º

0.0

0.5

1.0

1.5

2.0

2.5

3.0

0 10 20 30 40 50 60 70

V elo cit y ( m/ s)

Tip

Vert

ical

Def

lect

ion

(mm

)

ViscousInviscid

Airfoil tip deflection for increasing velocity, at 15ºInviscid and viscous predictions


Analysis of modified Airfoil

• Effects of:– Higher number of cells;– Reduction of ligament thickness (0.2 mm) and length (6 mm);

• Changes produce a 200% increase in tip deflection at 80 m/s.

Effect of chiral cell density on tip displacement

0

2

4

6

8

10

0 20 40 60 80 100

Airspeed (m/s)

Tip

Dis

plac

emen

t (m

m)

5 Cell Deep

2 Cell Deep


Aerodynamic testing:• Wind tunnel:

– Maximum speed ~70m/s. – Equipped for lift, drag and pitching moment measurements

• Visual Measurement Technique:– The motion of the aerofoil is tracked by a digital camera system, connected to a

laptop for real-time processing. – The position of a number of “targets” can be tracked at a rate of 7.50Hz.– Two targets could be used at once to monitor the strains.

Wing Prototype

Prototype mounted in low turbulence wind tunnel


Examples of Results

Numerical (FEA) Experimental

Lift Force against Displacement at Varying Incidence for the Experimental Analysis

0

50

100

150

200

250

300

350

0.0 0.5 1.0 1.5 2.0 2.5

Vertical Deflection (mm)

Lift

(N)

15º

10º

5º

0º

Deflection vs Speed at varying angles of incidence

0

0.5

1

1.5

2

2.5

0 10 20 30 40 50 60 70 80

Speed (m/s)

Def

lect

ion

(mm

)

15º EXP

10º EXP

5º EXP

15º FEA


Comments

• Deflections increase as velocity increases. • The relationship is non linear with respect to the lift force on the aerofoil:

– At low velocities, the rate of increase of deflection is lower than at higher velocities.

– This could be due to a certain amount of force required to ‘activate’ the honeycomb deformation mechanism.

– These forces appear to be different for the varying angles of incidence, – This could be explained by the varying pressure distribution around the

aerofoil (with respect to angle of incidence), and – the anisotropy of the chiral structure (i.e. the main bulk of the pressure force

acts on a different area of the aerofoil at different angles, giving a different deflection response).

• This deflection at 15° incidence corresponds to:– Camber change of 0.3%, – A reduction in CL of 0.05, – A reduction of CD of ~0.02.– Whilst this change is small, it proves that the morphing concept works.


Comments

• The measured strains between nodes 19 and 24, and 21 and 22 were found to be positive;– Computed values giving a Poisson’s ratio of approx. -0.9.

e1e2


Future developmentsAerolastic tailoring through wings with span-wise graded properties

Spars with different L/R ratios(Different chord-wise compliance)

L/R

L/RWing with continuous variation of

L/R ratio


Related researchDynamic shape control

DAQ &

Signal processing

Post-Processing

Shaker & F. Transducer

MATLAB

10-lb PCB Piezotronics

LDS V203

Polytec PSV-400 M2

Scanning head (Polytec PSV400 M2)


Related research


Comparison with numerical results

ω = 1744 Hz ω = 2250 Hz


Dynamic Shape Control:Motivations

Control of boundary layers and flow-separation phenomena;

Vibrating airfoil skins have been found to produce similar results to synthetic jets [Munday]:

• postpone stall or airflow separation, • reduce pressure drag,• reduce wave drag.

Oscillatory camber concept by D. Munday, J. Jacob and G. Huang


• Dynamic deformed shapes could be useful for reducing shock strength and thus wave drag.

Hogawa H., babinsky H., “Evaluation of wave drag reduction by flow control”, Aerospace Science and Technology, 10, pp. 1-8, 2006

Dynamic Shape Control:Motivations


Wave propagationin chiral networks

O A B O0

500

1000

1500

2000

2500

k-space position

Ω

Band-gap structure

-300 -200 -100 0 100 200 300-250-200-150-100-50

050

100150200250

Cg X [m/s]

Cg

y[m

/s]


Related research:Chiral Honeycombs

-30 -20 -10 0 10 20 30

500

600

700

800

900

1000

1100

1200

1300

, [deg]

( �e l) 3/�

* /�s,

[KP

a]

0.75 0.8 0.85 0.9 0.95 1

Global Buckling Chiral

Hexagonal/Auxetic Honeycombs

Co

lla

ps

es

tre

ng

thp

er

un

itw

eig

ht

�

�

L/R

-30 -20 -10 0 10 20 30

500

600

700

800

900

1000

1100

1200

1300

, [deg]

( �e l) 3/�

* /�s,

[KP

a]

( �e l) 3/�

* /�s,

[KP

a]

0.75 0.8 0.85 0.9 0.95 1

Global Buckling Chiral

Hexagonal/Auxetic Honeycombs

Co

lla

ps

es

tre

ng

thp

er

un

itw

eig

ht

�

�

L/R

Flat-wisestrength

Thermal behavior


THE END

RationaleOutline DesignResults: 2 cells �Influence of L/R ratioDesign ConstraintsConfigurationNumerical analysisExample of ResultsAnalysis of modified Airfoil�Aerodynamic testing:�Examples of ResultsCommentsCommentsFuture developmentsRelated researchRelated researchComparison with numerical resultsWave propagation�in chiral networksRelated research:�Chiral Honeycombs

Documents

MORPHING CHARACTERISTICS OF CHIRAL CORE AIRFOILSruzzene.gatech.edu/LabWeb/research/chiral_applications/Presentation_auxetics_2006.pdfAtlanta GA - USA Chrystel Remillat, Fabrizio Scarpa,