dglr kongress 2010-1

59. Deutscher Luft- und Raumfahrtkongress Vortrag Nr. 1341

Experimental investigation of transonic fluid-structure interaction phenomena

at a high aspect ratio swept wing P. C. Steimle*

Aerodynamisches Institut, RWTH Aachen *HE Space Operations GmbH

Steimle - Experimental investigation of transonic fluid-structure interaction phenomena at a high aspect ratio swept wing 2

■ Dynamic shock - boundary layer interaction is major reason for aeroelastic instabilities in transonic flight

■ Accurate prediction of local and global interaction between the wing structure and transonic flow by numerical simulation necessary for future highly elastic wing designs

■ Transport type wings exhibit structural response to unsteady aerodynamic loads in their first bending – torsion mode

■ Acquisition and analysis of 3D time-resolved flow data to contribute to the understanding of aerodynamic unsteadiness in transonic flow

■ Simulation of wing flutter response to the unsteady aerodynamic field by harmonic oscillations in pure pitch and heave DOF

Introduction

Flutter stability limit in transonic flight regime


Introduction

)sin(10 tαωαα +

)sin(1 tsh

hω

Harmonic pitch oscillation

Harmonic heave oscillation

■ Reduction of aero-structural interaction to focus on the aerodynamic problem

■ Harmonic oscillations of a swept wing model in pitch and heave to simulate wing flutter


Experimental setup: Swept wing model

■ Highly stiff wing model to uncouple structural and flow response

■ Usage of UHM carbon fiber composite sandwich structure

■ Pressure sensors incorporated one wing section in the area of highest aerodynamic loading

■ Flow analysis tools □ Oil flow visualization

□ Time-resolved pressure distributions cp(t) from in-situ pressure sensors in one pressure tap section in y = 80mm

□ Pressure-sensitive luminescent paint on open anodic aluminum binder

□ Photogrammetric wing deformation measurement

M∞

optical markers

Supercritical wing section BAC 3-11/RES/30/21

Mean aerodynamic chord = 74.3 mm

Pressure tap section length c1 = 82.71 mm

LE sweep 34° TE sweep 22° and 26°

Aspect Ratio 10.3

c


Trisonic wind tunnel of the RWTH Aachen University

Test section 0.4m x 0.4m Mach number 0.4 - 3.0 Testing time 2 - 5s Unit Reynolds number 1.5 x 107m-1

2-D adaptive test-section 0.80.84

0.880.92

0.96

0.40.5

0.60.7

0.8

0

2

4

6

8x 106

M∞ω*

|p′ |2 /f s [P

a2 /Hz]

Pressure fluctuations in empty test section

M∞ x2

z2

Experimental setup: Wind tunnel facility


200 300 400 500 600 700 800Wavelength [nm]

Inte

nsity

[arb

itrar

y sc

ale]

ambient pressure vacuum

PSA

Sample image of anodized porous surface A1050

Voltage 20V Current density 15mA/cm2

Temperature 18°C

aluminum foil 47μm

adhesive tape 70μm

Luminescence signal spectrum of PSA on porous A1050

■ Aluminum tape attached to carbon fiber wing, anodization provides micro-pores Ø25 to 40nm

■ 1.96% to 5% of airfoil thickness added locally to wing geometry

Experimental setup: AA-PSP


wing model

Optical setup for AA-PSP measurements at the test section side wall

○ UV-light source flicker-free mercury vapor lamp Osram HBO 500W with band pass filter Schott UG-11 and focusing lens

○ Images recorded with Photron Fastcam 1024 PCI CMOS device with Leica Noctilux M 1:1/50mm lens and band pass filter combination Schott KV-408 + Lee V28 Blueberry 8

○ Images acquired at fs = 1.5 and 2kHz

○ Camera set in Scheimpflug condition to focus entire wing surface

Experimental setup: AA-PSP


Instantaneous distribution of PSA luminescence intensity corresponding to the local pressure on the upper wing surface, image acquisition rate fs = 1500 Hz

[α0 M∞] = [0°,0.86]

Fixed wing aerodynamics: Weak supersonic field


Time-resolved pressure distribution on wing upper surface from AA-PSP with a-priori calibration, image acquisition rate fs = 1500 Hz [α0 M∞] = [0°,0.86]

1

0.5

0

-0.5

-1

-1.5

cp



Aluminum foil causes slight change in time-averaged pressure distribution

Weak and highly dynamic shock wave in η = 0.286

Strong fluctuations in shock position and strength


Time-averaged pressure distribution on wing upper surface from AA-PSP with a-priori calibration, image acquisition rate fs = 1500 Hz


Skewing of the boundary layer velocity profile in the rear of the wing

Incipient separation in the shock foot region

Time-averaged pressure distribution and pressure fluctuation quantities


Surface flow pattern on wing upper side

Skin friction line

Separation line

separation


Spectral analysis of oblique shock parameters

■ Shock buffet at ω* = 0.72 ■ Marginal separation acts as

sound source due to shedding of vortex structures at the sharp TE

■ Buffet originates from sensitivity of weak shock wave to sound waves travelling upstream

0 0.5 1 1.5 2 2.5-12-11-10-9-8-7-6-5-4-3-2

® s

ω*lo

g 10(|x

′ /c1|2 /f s [1

/Hz]

) [1/

Hz]

0 0.5 1 1.5 2 2.5-12-11-10-9-8-7-6-5-4-3-2

®

ω*

log 10

(|(p 2/p

1)′ |2 /f s [1

/Hz]

) [1/

Hz]

0 0.5 1 1.5 2 2.5

-9-8-7-6-5-4-3-2-101

ω*

log 10

(|β′ |2 /f s [1

/Hz]

) [1/

Hz]

0 0.5 1 1.5 2 2.5-10-9-8-7-6-5-4-3-2-10

® µ

ω*lo

g 10(|θ

′ |2 /f s [1/H

z]) [

1/H

z]


■ Weak SBI test case for harmonic forcing experiments


1

0.5

0

-0.5

-1

-1.5

cp

Fixed wing aerodynamics: Strong supersonic field

Time-resolved pressure distribution on wing upper surface from AA-PSP with a-priori calibration, image acquisition rate fs = 1500 Hz [α0 M∞] = [0°,0.92]


Aluminum foil causes slight displacement of time-averaged shock position downstream

Strong shock wave with λ-configuration in η = 0.286

Flow field dynamics focused on area of shock wave

Time-averaged pressure distribution on wing upper surface from AA-PSP with a-priori calibration, image acquisition rate fs = 1500 Hz



Skin friction line

Separation line Outer stream line

■ Shock-induced full scale TE separation ■ Performance boundary of this

configuration

separation

Surface flow pattern on wing upper side

Time-averaged pressure distribution and pressure fluctuation quantities



■ Shock oscillation at ω* = 0.73 and higher harmonics

■ Still distinct oscillation in flow deflection due to pulsation of separated area

■ No fluctuation with ω* = 0.42!

Spectral analysis of oblique shock parameters

0 0.5 1 1.5 2 2.5-12-11-10-9-8-7-6-5-4-3-2

ω*lo

g 10(|x

′ /c1|2 /f s [1

/Hz]

) [1/

Hz]

0 0.5 1 1.5 2 2.5-12-11-10-9-8-7-6-5-4-3-2

ω*

log 10

(|(p 2/p

1)′ |2 /f s [1

/Hz]

) [1/

Hz]

0 0.5 1 1.5 2 2.5

-9-8-7-6-5-4-3-2-101

®

ω*

log 10

(|β′ |2 /f s [1

/Hz]

) [1/

Hz]

0 0.5 1 1.5 2 2.5-10-9-8-7-6-5-4-3-2-10

®

ω*lo

g 10(|θ

′ |2 /f s [1/H

z]) [

1/H

z]


■ Strong SBI test case for harmonic forcing experiments


Harmonic forcing experiments: Weak interaction test case

Amplitude effect on 1st harmonic pressure distributions on upper surface in η = 0.286

■ Distinctive harmonic response of the shock wave ■ Reduction of flow field response with increasing amplitude


Harmonic forcing experiments: Weak interaction test case

Frequency effect on 1st harmonic pressure distributions on upper surface in η = 0.286

■ Strongest flow field response on heaving wing at smallest frequency

■ Significant reduction with increasing frequency


Harmonic forcing experiments: Strong interaction test case

Amplitude effect on 1st harmonic pressure distributions on upper surface in η = 0.286

■ Strong shock configuration generally more sensitive to structural motion due to harmonic motion of the separation line in phase with the wing oscillation



Frequency effect on 1st harmonic pressure distributions on upper surface in η = 0.286



■ Fluid structure energy exchange determines the development of aeroelastic instabilities potentially occurring in cruise flight conditions

■ Averaged local energy exchange estimated based on the time-resolved pressure and synchronously measured wing motion data in η = 0.286

■ Work coefficient cw describes the work of a fluctuating pressure cp‘(t) exerted on a wing surface element dAi corresponding to the normal vector n


Local energy exchange: Weak SBI flow

Time-averaged work coefficient in η = 0.286 on the upper surface,

weak interaction flow

■ Weak shock wave present at decreasing aeroelastic stability boundary

■ Amplification of excitatory effect of the fluid-structure interaction

■ Cross-flow region in the rear of the wing illustrates damping nature of boundary layer flow with skewed velocity profile

■ Heave amplitude effect would have general potential to drive destructive flutter amplitude increase at pure bending motion

■ Structural excitation reduced when pitch DOF is activated and heave DOF suppressed


Time-averaged work coefficient in η = 0.286 on the upper surface, strong interaction flow

Local energy exchange: Strong SBI flow


Summary

● Results from experimental test campaign with supercritical high aspect ratio swept wing can be used to analyze unsteady transonic flow behavior

● Single shock wave with incipient separation close to trailing edge: highly dynamic flow with self-induced periodic shock oscillation

● Lambda shock system with full scale 3D separation: lower level of unsteadiness, but still with active acoustic feedback mechanism; unsteadiness reduced by steady character of the separation line

● Development of shock wave motion along the wing span shows synchronous shock motion in the inner halfspan region

● Trailing edge kink region identified as major source of disturbances with ω* = 0.72

● AA-PSP is a valid measurement tool for this unsteady flow, despite the high noise level contained in the images as result of weak dynamic intensity signal

● AA-PSP coating able to visualize the unsteady pressure field with high degree of reliability regarding frequencies contained in the dynamic flow process

● Experiments were performed within the Collaborative Research Center SFB 401 of the German Research Foundation (DFG)


Thank you for your attention

dglr kongress 2010-1

Documents