Information on unsteady separation and dynamic stall is being obtained from two experimental programs that have been underway at United Technologies Research Center since 1984. The first program is designed to obtain detailed surface pressure and boundary layer condition information during high amplitude pitching oscillations of a large (17.3 in. chord) model wing in a wind tunnel. The second program involves the construction and testing of a pressure-instrumented model helicopter rotor. This presentation describes some of the results of these experiments, and in particular compares the detailed dynamic stall inception information obtained from the oscillating wing with the unsteady separation and reverse flow results measured on the retreating blade side of the model rotor during wind tunnel testing

Carta, Franklin O.

Lorber, Peter F.

English

NASA Technical Reports Server

N94- 34967
UNSTEADY SEPARATION EXPERIMENTS ON 2-D AIRFOILS, 3-D WINGS,
AND MODEL HELICOPTER ROTORS
Peter F. Lorber and Franklin O. Carta
United Technologies Research Center
East Hartford, CT 06108
ABSTRACT FOR NASA/AFOSR/ARO WORKSHOP ON
PHYSICS OF FORCED UNSTEADY SEPARATION
APRIL 17-19, 1990
Information on unsteady separation and dynamic stall is being obtained from
two experimental programs that have been underway at United Technologies
Research Center since 1984. The first program is designed to obtain detailed
surface pressure and boundary layer condition information during high amplitude
pitching oscillations of a large (17.3 in chord) model wing in a wind tunnel.
The second program involves the construction and testing of a pressure-
instrumented model helicopter rotor. This presentation describes some of the
results of these experiments, and in particular compares the detailed dynamic
stall inception information obtained from the oscillating wing with the unsteady
separation and reverse flow results measured on the retreating blade side of the
model rotor during wind tunnel testing.
An inital, two-dimensional oscillating wing experiment was performed in
1986 under AFOSR sponsorship, and has been documented in Refs. 1 and 2. Surface
pressure and hot film data were acquired for constant pitch rate ramps and
sinusoidal oscillations in the range of _ = 0 to 30 deg, for M = 0.2, 0.3, and
0.4, and Re = 2,000,000 to 4,000,000. Figure 1 shows typical results for an
M = 0.2, A = _c/2U = 0.01 ramp. This figure is similar to those in Refs. 1 and
2, and shows time histories of the ensemble-averaged pressures at each of the 18
transducers on the airfoil surface. A negative pressure spike (caused by the
dynaic stall vortex) forms near _ = 0.47, and moves back along the airfoil.
Figure 2 (not previously published) shows chordwise pressure distributions at
several values of • during this process. The passage of the vortex is shown by
the pressure bulge on the upper surface. The references discuss the effects of
pitch rate, pitching waveform, and Mach number on the stall process. Compres-
sibility effects were very significant, as a small supersonic bubble forms near
the leading edge at H = 0.4, and the peak suction pressures and the unsteady
increments to the airlods are much weaker. Reference 3 describes a Navier-
Stokes simulation of the 2-D experiments. Good agreement was obtained up
through the formation of the dynamic stall vortex, while many of the
quantitative aspects of the periodic vortex shedding after stall were missed.
This study is now being extended under ARO and AFOSR sponsorship to include
three-dimensional measurements on a finite tip model. In addition to obtaining
information on how the presence of the wing tip affects the dynamic stall
process, this experiment is intended to study sweep and compressibility effects.
The model, shown in Fig. 3, consists of a square wing with the same (17.3 in)
chord and airfoil section (SSC-A09) as the 2-D wing. The instrumentation
consists of chordwise arrays of pressure transducers at 5 spanwise stations (112
https://ntrs.nasa.gov/search.jsp?R=19940030461 2020-06-16T13:01:29+00:00Z
transducers) and arrays of surface hot film gages at 3 spanwise stations (16
total gages) to determine transistion and separation information. The model
will be tested at 3 sweepangles: A = O, 15, and 30 deg, and at Machnumbers
between 0.2 and (structural loads permitting) 0.6. The experiment is scheduled
to be completed in 1990.
In addition to large amplitude rampsand sinusoids, information will be
obtained on small amplitude (±0.5 to 2 deg) oscillations near the static stall
angle. This program will be sponsored by NASALewis and ARO,and is designed to
study the incipient stages of stall flutter, with particular application to
aircraft propellors. Results of an earlier, smaller scale experiment were
reported in Ref. 4. The aerodynamic damping was found to be substantially more
negative for very small amplitude oscillations, allowing a rapid growth to a
limit cycle motion.
The helicopter rotor program involves the construction and testing of a
heavily instrumented, 9.5 ft diameter scale model of a current-technology main
rotor (Fig. 4). The model contains 176 miniature pressure transducers, as well
as strain gages, temperature sensors, and surface hot film gages. Hover testing
was described in Ref. 5, and aerodynamic results from a 1989 wind tunnel test
are given in Ref. 6. A great deal of information has been obtained using this
model rotor. Of current interest is the behavior of the inboard portion of the
retreating blade at moderately high advance ratios (U = U /QR ~ 0.28-0.36).
This region is subject to rapid increases in angle of attack and rapid
reductions in relative velocity. Figure 5 shows chordwise pressure
distributions for r/R = 0.4 at four azimuths on the retreating side. The flow
appears highly loaded but attached at _ = 190, shows leading edge separation at
at _ = 220, has a very large aft loading at _ = 270, and is beginning to
reattach at , = 320. Time histories of the ensemble averaged pressures at
r/R = 0.225 and 0.4 are shown in Fig. 6. Sharp negative pressure spikes are
present (on the upper surface only) near , = 175 at r/R = 0.225 and near _ = 210
at r/R = 0.4. The flow appears to separate after the spikes have passed, as
shown by flat ensemble averaged pressures between _ = 180 and 315. This
phenomenon is similar to the shedding of the dynamic stall vortex on the
oscillating 2-D airfoil (Fig. i). The non-dimensional convection speed of the
spike (0.25 times the local relative velocity) is also similar.
The rotor flow field has many complexities not present with the 2-D
airfoil. The sequence observed on the retreating blade side at a particular
radial station may include: forming and shedding a leading edge vortex, entering
the region of reverse relative velocity, shifting from positive to negative
lift, shedding a vortex from the trailing edge that moves towards the leading
edge, resuming positive relative velocity and lift, and interacting with the
wake of the rotor hub. Additional complications include radial velocity and
twist gradients and aeroelastic deflections. With all of these factors present,
it is encouraging to see some similarities to the simpler, oscillating 2-D
results, but one must not forget how complex the rotor flow field actually is.
This observation is lent particular weight by the many references to the
helicopter stall problem in the introductory sections of oscillating airfoil
papers.
References
I. Lorber, P.F., and Carta, F.O., "Airfoil DynamicStall at Constant pitch Rate
and High reynolds Number," AIAA J. Aircraft, Vol 25, pp. 548-556, June 1988.
2. Lorber, P.F., and Carta, F.O., "Unsteady Stall Penetration Experiments at
High Reynolds Number," AFOSR TR-87-1202, April 1987.
3. Patterson, M.T., and Lorber, P.F., "Computational and Experimental Studies
of Compressible Dynamic Stall," 4th Symposium on Numerical and Physical Aspects
of Aerodynamic Flows, Long Beach, CA, Jan 1989 (accepted for publication in the
J. Fluids & Structures, 1990).
4. Carta, F.O., and Lorber, P.F., "Experimental Study of the Aerodynamics of
Incipient Torsional Stall Flutter," AIAA J. Propulsion and Power, Vol. 3,
pp.164-170, March-April, 1987.
5. Lorber, P.F., Stauter, R.C., and Landgrebe, A.J., "A Comprehensive Hover Test
of the Airloads and Airflow of an Extensively Instrumented Model Helicopter
Rotor," 45th Annual Forum of the American Helicopter Society, Boston, HA,
May 1989.
6. Lorber, P.F., "Aerodynamic Results of a Pressure-Instrumented Model Rotor
Test at the DNI;," to be presented at the 46th Annual Forum of the American
Helicopter Society, Washington, DC, May 1990.
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Unsteady separation experiments on 2-D airfoils, 3-D wings, and model helicopter rotors

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