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NASA Technical Memorandum 107581
NACA0012 BENCHMARK MODEL EXPERIMENTAL FLUTTER
RESULTS WITH UNSTEADY PRESSURE DISTRIBUTIONS
Jos_ A. Rivera, Jr., Bryan E. Dansberry, Robert M. Bennett,
Michael H. Durham, and Walter A. Silva
March 1992
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National Aeronautics and
Space Administration
Langley P,o$oarch Center
Hampton, Virginia 23665
(NASA-TM-]07_,_I) NA£AO012 F_ENCHMARK MOCEL
EXPFRIMENTAL FLUITER RESULTS WITH UNSTEAOY
PRESSURE _ISIRI _,UTIONS (NASA) 13 pCSCL 01A
o_loz
N92-22507
NACA 0012 BENCHMARK MODiI:L EXI_ERIMENTAL FLUTTER RESULTS WITH
UNSTEADY PRESSURE DISTRIBUTIONS
Jos_ A. Rivera, Jr., Bryan E. Dansberry, Robert M. Bennett,
Michael H. Durham, and Walter A. Silva
NASA Langley Research Center
Hampton, VA 23665-5225
The Structural Dynamics Division at NASA
Langley Research Center has started a wind tunnel
activity referred to as the Benchmark Models
Program. The primary objective of the program is to
acquire measured dynamic instability and
corresponding pressure data that will be useful for
developing and evaluating aeroelastic type CFD codes
currently in use or under development. The program
is a multi-year activity that will involve testing of
several different models to investigate various
aeroelastic phenomena. This paper describes results
obtained from a second wind tunnel test of the fast
model in the Benchmark Models Program. This fLrst
model consisted of a rigid semispan wing having a
rectangular planform and a NACA 0012 airfoil shape
which was mounted on a flexible two degree-of-
freedom mount system. Experimental flutter
boundaries and corresponding unsteady pressure
distribution data acquired over two model chords
located at the 60 and 95-percent span stations are
presented.
The development of unsteady aeroelastic
computational fluid dynamic (CFD) codes requires
experimental data to validate computed results and/or
for use as a guide for modification of analyses
methods. The Benchmark Models Program 1 was
initiated by the Structural Dynamics Division at
NASA Langley Research Center to provide such
experimental data and to aid in understanding the flow
phenomena associated with unusual aeroelastic
phenomena.
The Benchmark Models Program (BMP) has
identified several aerodynamic configurations to be
tested in the NASA Langley Transonic Dynamics
Tunnel (TDT). Some configurations are models for
testing on a flexible mount system, referred to as the
Pitch and Plunge Apparatus (PAPA). The NACA
0012 airfoil rectangular wing is the first of these
BMP PAPA mounted models. To date, two
comprehensive wind tunnel tests have been conducted
for this model. During the first wind-tunnel test,
flutter boundaries were defined and wing surface
pressure measurements were obtained for a partial set
of pressure transducers at the 60-percent span station.
Preliminary results from this test are presented in
reference 2. These results were used primarily as a
guide for defining the scope of the second test. The
second wind-tunnel test of this model was conducted
to determine the flutter boundaries while
simultaneously taking surface pressure measurements
at most flutter conditions. For the second test,
additional pressure transducers were installed on the
wing to give more wing surface pressure
measurements at both the 60-percent and 95-percent
span stations. These flutter boundaries and the wing
surface pressure data measured for the conventional
flutter boundary are presented in reference 3 in tabular
format. Reference 3 also contains an extensive set of
wing surface pressure measurements obtained with the
model support system rigidized.
This paper focuses on the flutter and pressure data
available from reference 3 to highlight Mach number
effects on the flutter boundary and to correlate the
measured pressure distributions with the conventional
flutter boundary at transonic Math numbers. The
conventional flutter boundary, a plunge instability
region near M=0.90, and the stall flutter boundary at
M=0.78 are presented. In addition unsteady wing
surface pressure measurements acquired during
conventional flutter are presented in coefficient form
and discussed.
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Speed of sound, ft/sec
Mean pressure coefficient during flutter
Wing streamwise local chord length, 16-inches
Frequency, Hz
Wind-off pitch frequency, 5.20 Hz
Flutter frequency, Hz
Flutter frequency ratio
Strucuwal damping
Reduced frequency,k=(c/2)co/V
Wing spanwise length,32 inches
Leading edgu
Calculated moving mass of wing/PAPA
mechanism, 5.966 slugs
Free-stream Mach number
Phase angle referenced to pitch
displacement, degrees
Free-stream dynamic pressure, psf
Reynolds number based on chord length
Trailing edge
Free-stream velocity, t/see
Flutter speed index, VI=V/(c/2)_r_
Distance from wing leading edge, inches
Fraction of local chord
Vertical (plunge) displacement, inches
Wing angle of attack (also alpha), degrees
Pitch displacement, degrees
Mass ratio, _t= mhtpl(c2/4)
Density, slugs/ft 3
Circular freqtmncy, rad/sec
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The wind-tunnel tests were conducted in the
Langley Transonic Dynamics Tunnel (TDT). 4 The
TDT is a continuous flow, single return wind tunnel
with a 16-foot square test section (with cropped
comers) having slots in all four walls. It is capable of
operating at Mach numbers up to 1.2 and at
stagnation pressures from near vacuum to
atmospheric. The tunnel is equipped with four quick-
opening bypass valves which can be used to rapidly
reduce test-section dynamic pressure and Mach
number when an instability occurs. Although either
air or a heavy gas can be used as a test medium, only
air was used for the present tests.
Model
The model is a semispan rigid wing mounted on a
flexible mount system referred to as the Pitch and
Plunge Apparatus (PAPA).5, 6 A photograph of the
model mounted in the TDT test section is shown in
figure 1. A planform view of the model is shown in
figure 2. The model has a NACA 0012 airfoil section
and a rectangular planform with a span of 32 inches
and a chord of 16 inches. The mount system is
attached to a turntable which provides for angle-of-
attack variation. Transition strips made up of No. 30
carborundum grit were applied to the model
approximately one inch back from the leading edge
(approximately 6-percent chord) on both the upper and
lower surfaces.
The model was designed to allow installation of
80 in-situ pressure transducers for measurement of
unsteady wing surface pressures. These pressure
transducers were referenced to wind-tunnel static
pressure. Forty of the transducers are located at the
60-percent span station, and forty at the 95-percent
span station. The span locations for these pressure
measurements are indicated in figure 2. The physical
locations of orifices and corresponding pressure
transducers on the airfoil cross section are available in
reference 3 and illustrated in figure 3.
Details of the model construction can be seen in
the photographs of figure 4. The lower photograph
shows that the model was fabricated in three sections.
Each section was machined from solid aluminum
stock. The sections were bolted together after the
pressure transducers, reference pressure tubes, and
wiring were installed. In the upper left photograph is
an expanded view of a portion of the mid section
which shows holes drilled in the edge of the section.
These holes were used for insertion of the pressure
transducers. Two pressure transducers are shown next
to the model. One of the pressure transducers is
shown mounted in a brass tube. The brass tube is
used to protect the transducer when it is inserted and
removed from the model. The associated orifice holes
for the pressure transducers are located about one inch
from the inboard edge of the mid section and tip
section. When the pressure transducers and sleeves
are inserted, the measurement face of the pressure
transducer is within 0.2 inch of the orifice location on
the wing surface where the pressure measurement is
being made. Exceptions are the trailing edge pressure
transducers which are approximately 0.7 inch from
the orifice location.
There are four accelerometers in the model, one
near each comer, used to assist in identifying model
dynamic characteristics during testing. These
accelerometers are mounted in pockets, one of which
is shown in the photograph in the upper right of
figure 4.
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The model mounting system is composed of two
basic parts. They include a flexible support and a
large splitter plate. The model is mounted outboard of
the splitter plate.
The flexible support, which allows pitch and
plunge motion of the model, is located behind the
splitter plate. A description of the flexible mount
system, referred to as the PAPA (Pitch and Plunge
Apparatus),5,6 is presented in figures 5, 6, and 7.
Figure 5 is a photograph which shows a moving
plate supported out from the tunnel wall by a system
of four rods and a centerline flat plate drag strut all
with fixed-fixed end conditions. At the tunnel wall
the rods and drag strut are attached to a mounting
plate attached to a turntable so that the model angle of
attack can be varied.
The rods and flat plate drag strut provide linearly
constrained motion so that the model can oscillate
sinusoidally in pitch and plunge. The oscillations are
functions of the stiffness of the rods, the mass
properties of the moving apparatus, and the
aerodynamic forces on the model. The structural
properties of this simple mount system can be well
defined mathematically and can be easily measured for
flutter calculations. This makes the PAPA mount
system a valuable tool for obtaining experimental
model flutter data for correlation with analysis
because disagreement between theory and experiment
can be primarily attributed to aerodynamics. The
PAPA is instrumented with two strain gage bridges
oriented to measure bending and torsional moments
from which wing model instantaneous plunge
position and pitch angle can be obtained. These are
located on the flat plate drag strut near the mounting
plate.
The PAPA splitter plate, shown in figure 6, is
suspended out from the test-section wall by struts
which are about 40 inches long. The splitter plate is
12 feet long and 10 feet high. The centerline of the
model and the PAPA support system is 7 feet
rearward from the leading edge of the splitter plate.
The PAPA mount system rods and drag strut are
enclosed in a fairing behind the splitter plate. The
wing model and end plate are the only parts of the
apparatus that are exposed to the flow in the test
section. The splitter plate serves to separate flow
over the model from flow around the mount system
fairing which is located between the splitter plate and
the test section wall.
A top view sketch which shows how the wing
model, the PAPA apparatus, the splitter plate and
other components fit together is presented as figure 7.
The model is attached to a short pedestal or spacer
which protrudes through the opening in the splitter
plate, all of which attaches to the moving plate. The
moving plate has provisions for the addition of
ballast weights (indicated in figure 7) to adjust the
mount system structural dynamic characteristics. The
opening in the splitter plate is covered by a thin
circular end plate attached to the root section of the
model to prevent flow through the splitter plate. The
circular end plate has a diameter equal to the model
chord length. The circular plate can be seen in the
photograph of figure 6. The gap between the end plate
and the splitter plate was less than one-tenth of an
inch, but sufficient so that the end plate did not rub
against the splitter plate.
Structural Dynamic Characteristics
The first two wind-off natural modes of vibration
for the NACA 0012 model/PAPA mount system
assembly are the wing-model rigid-body plunge and
rigid-body pitch modes respectively. Inertia coupling
between these two modes was eliminated by
positioning ballast weights on the PAPA system
moving plate so that the system center of gravity was
on the PAPA elastic axis (centerline). Therefore the
rigid-body plunge mode consists only of vertical
translation of the wing model and the rigid-body pitch
mode consists only of rotation of the wing model
about the mid-chord. The measured frequencies,
damping and stiffnesses for these two modes are
presented in table 1. Modal displacements for
corresponding, unit-generalized-masses are presented
in table 2.
Data Acauisition and Reduction
Wing model and mount system transducer time
history data were acquired at the conventional flutter
boundary test conditions with the TDT data
acquisition system. The data were acquired
simultaneously (not multiplexed) for all transducers at
a rate of 100 samples per second for 40 seconds and
recorded in digital form on disk.
For each differential pressure transducer (the
pressure transducers were referenced to wind-tunnel