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
t
s
h
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
0.8
0.84
0.88
0.92
0.96
0.4
0.5
0.6
0.7
0.8
0
2
4
6
8
x 10
6
M∞ω*
|p′
|2
/f
s
[Pa2
/Hz]
Pressure
fluctuations in
empty test
section
M∞
x2
z2
Experimental setup:
Wind tunnel facility

200 300 400 500 600 700 800
Wavelength [nm]
Intensity[arbitraryscale]
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,
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,
fs = 1500 Hz

 Skewing of the boundary layer velocity
profile in the rear of the wing
 Incipient separation in the shock foot region
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*
log
10
(|x′
/c
1
|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
-1
0
1
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
-1
0
® µ
ω*
log
10
(|θ′
|2
/f
s
[1/Hz])[1/Hz]
■ Weak SBI test case for
harmonic forcing experiments

1
0.5
0
-0.5
-1
-1.5
cp
Strong supersonic field
Time-resolved pressure
distribution on wing upper
surface from AA-PSP with
a-priori calibration,
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
on wing upper surface from AA-PSP
with a-priori calibration,
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
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*
log
10
(|x′
/c
1
|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
-1
0
1
®
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
-1
0
®
ω*
log
10
(|θ′
|2
/f
s
[1/Hz])[1/Hz]
■ 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

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

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

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