an experimental study of dynamic stall on advanced airfoil
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NASA Technical Memorandum 84245 USAAVRADCOM TR-82-A-8
.,o3- 17 5.)__
_,.7 17
An Experimental Study of DynamicStall on Advanced Airfoil SectionsVolume 3. Hot-Wire and Hot-FilmMeasurements
L. W. Carr, W. J. McCroskey: K. W. McAlister,S. L. Pucci, and O. Lambert
December 1982
fU/LS/ National AeronautMcs and
Space Administration
REPRODUCED BYU.S. DEPARTMENT OF COMMERCE
NATIONALTECHNICALINFORMATION SERVICESPRINGFIELD, VA 22161
United States ArmyAviation Research Vand DevelopmentCommand
NASA Technical Memorandum 84245 USAAVRADCOM TR-82-A-8
An Experimental Study of DynamicStall on Advanced Airfoil SectionsVolume 3. Hot-Wire and Hot-FilmMeasurements
L. W. Carr
W. J. McCroskey
K. W. McAlister
S. L. Pucci, Aeromechanics Laboratory
AV RADCOM Research and Technology Laboratories
Ames Research Center, Moffett Field, California
O. Lambert, Service Technique des Constructions Aeronautiques,
Paris, France
NASANational Aeronautics and
Space Administration
Ames Research Center
Moffett Field California 94035
United States ArmyAviation Research and
Development CommandSt. Louis. Missouri 63166
TABLE OF CONTENTS
LIST OF FIGURES .......... . ....................
LIST OF TABLES ..............................
SYMBOLS ...................................
SUMMARY • , • . • • • • • • • • • • • • • • • • • • • • • • • • • * • * • • •
INTRODUCTION ................................
DESCRIPTION OF EXPERIMENTAL PROCEDURES ...................
DATA ANALYSIS AND INTERPRETATION ......................
Skln-Frictlon Gage ...........................Hot-Wire Probe .............................
Reverse-Flow Sensors ..........................
Averaging Techniques ..........................
Example of Signal Analysis .......................
RESULTS ...................................
REFERENCES .................................
TABLES ...................................
FIGURES ...................................
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LESTOFFIGURES
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Diagram showing installation of spar and airfoil shell in tunnel ....
Diagram of hot-film skin-frlction gage .................
Response of hot-film skln-friction gages mounted on Ames A-01
airfoil during airfoil oscillation in pitch (_ = 15 ° + i0 ° sin rot,
k ffi0.i0, M_ ffi0.22) ..........................
Response of hot-film skin-friction gages at surface of NACA 0012
airfoil during airfoil oscillation in pitch (s = 15 ° + i0 ° sin mr,
k = 0.i0, M_ = 0.295) ........................
Diagram of dual-element hot-wlre probe .................
Response of hot-wlre anemometer probes on Wortmann FX-098 airfoil
during airfoil oscillation in pitch (_ ffi15 ° + i0 ° sin _t,
k = 0.i0, M_ = 0.ii) ..........................
Response of hot-wire anemometer probe installed near trailing
edge of the Vertol VR-7 airfoil during oscillation in pitch ......
Results obtained using trlple-wlre flow-reversal sensor:
(a) Typical comparison of flow-reversal sensor and hot-wlre
anemometer signal (from ref. 2); (b) Progression of flow reversal
up airfoil during dynamic stall (from ref. 2) .............
Diagram of three-element, directionally sensitive hot-wire probe
(from ref. 2) ............................
Comparison of lO0-cycle ensemble average and slngle-cycle signals
from hot-wire anemometers for Vertol VR-7 airfoil during oscillation
in pitch: --, i00 cycle average; ...... , slngle cycle ......
Response of hot-film skln-frlctlon gage and hot-wlre anemometer
probes on Vertol VR-7 during oscillation in pitch (a - 15 ° + i0 ° sin _t,
k ffi0.i0, M_ = 0.185) .........................
Phase angle, _t, of flow reversal on NACA 0012 airfoil vs chord
location for a range of Mach numbers at k ffi0.i, _ - 15 ° + I0 ° sin _t --
Mach number effects ..........................
Phase angle, _t, of flow reversal on Ames A-OI airfoil vs chord
location for a range of Mach numbers at k - 0.i, e = 15 ° + i0 ° sin _t --
Mach number effects ..........................
Phase angle, _t, of flow reversal on Wortmann FX-098 airfoil vs chord
location for a range of Mach numbers at k = 0.i, _ ffi15 ° + i0 ° sin _t --
Mach number effects .........................
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Phase angle, _t, of flow reversal on Sikorsky SC-I095 airfoil vs chord
location for a range of Mach numbers at k = 0.i, _ = 15 ° + i0 ° sin _t --
Mach number effects . .........................
Phase angle, _t, of flow reversal on Hughes HH-02 airfoil vs chord
location for a range of Math numbers at k = 0.I, _ = 15 ° + i0 ° sin _t-
Mach number effects .........................
Phase angle, _t, of flow reversal on Vertol VR-7 airfoil vs chord
location for a range of Mach numbers at k = 0.i, e = 15 ° + i0 ° sin _t --
Mach number effects ..........................
Phase angle, _t, of flow reversal on NLR-I airfoil vs chord location
for a range of Mach numbers at k = 0.i, _ = 15 ° + i0 ° sin _t -Mach
number effects .............................
Phase angle, wt, of flow reversal on NLR-7 airfoil vs chord location
for a range of Math numbers at k = 0.i, _ = 15 ° + I0 ° sin _t --Mach
number effects ............................
Phase angle, _t, of flow reversal on NACA 0012 airfoil vs chord
location for a range of frequencies at M_ = 0.295,
= 12 ° + 5° sin _t -- light-stall conditions ..............
Phase angle, _t, of flow reversal on Ames A-01 airfoil vs chord
location for a range of frequencies at M_ = 0.295,
= ii ° + 5 ° sin _t -- llght-stall conditions ..............
Phase angle, _t, of flow reversal on Wortmann FX-098 airfoil vs chord
location for a range of frequencies at M_ = 0.295,
= i0 ° + 5° sin _t -- light-stall conditions ..............
Phase angle, wt, of flow reversal on Sikorsky SC-I095 airfoil vs
chord location for a range of frequencies at M_ = 0.295,
= ii ° + 5° sin _t -- light-stall conditions ..............
Phase angle, _t, of flow reversal on Hughes HH-02 airfoil vs chord
location for a range of frequencies at M_ = 0.295,
= i0 ° + 5 ° sin _t -- light-stall conditions ..............
Phase angle, _t, of flow reversal on Vertol VR-7 airfoil vs chord
location for a range of frequencies at M_ = 0.295,
= 15 ° + 5° sin _t -- light-stall conditions ..............
Phase angle, _t, of flow reversal on NLR-I airfoil vs chord
location for a range of frequencies at M_ = 0.295,
= i0 ° + 5 ° sin _t -- light-stall conditions ..............
Phase angle, _t, of flow reversal on NLR-7 airfoil vs chord
location for a range of frequencies at M_ = 0.295,
e = 15 ° + 5 ° sin _t -- light-stall conditions .........
47
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vl
28 Phase angle, wt, of flow reversal on AmesA-01 airfoil vs chord for arange of frequencies at M_= 0.295, _ _ 15° + i0 ° sin _t -- deep-stall
conditions ...... • ........................ 56
29 Phase angle, _t, of flow reversal on Wortmann W-98 airfoil vs chord
for a range of frequencies at M_ = 0.295, _ = 15 ° + i0 ° sin _t --
deep-stall conditions ........................ 57
30 Phase angle, _t, of flow reversal on Wortmann FX-098 airfoil vs chord
for a range of frequencies at M_ = 0.185, e = 15 ° + i0 ° sin _t --
deep-stall conditions ......................... 58
31 Phase angle, _t, of flow reversal on Vertol VR-7 airfoil vs chord
for a range of frequencies at M_ = 0.295, = = 15 ° + i0 = sin _t --
deep-stall conditions ......................... 59
vii
LIST OF TABLES
i Summary of analyzed flow-reversal data .................
2 Phase angle of flow reversal: NACA 0012 airfoil ............
3 Phase angle of flow reversal: Ames A-01 airfoil ............
4 Phase angle of flow reversal: Wortmann FX-098 airfoil .........
5 Phase angle of flow reversal: Sikorsky SC-I airfoil ..........
6 Phase angle of flow reversal: Hughes HH-02 airfoil ..........
7 Phase angle of flow reversal: Vertol VR-7 airfoil ...........
8 Phase angle of flow reversal: NLR NL-I airfoil ............
9 Phase angle of flow reversal: NLR-7301 airfoil ............
10 Error-bound for flow-reversal measurements (deg): NACA 0012
airfoil ................................
ii Error-bound for flow-reversal measurements (deg): Ames A-01
airfoil ................................
12 Error-bound for flow-reversal measurements (deg): Wortmann FX-098
airfoil ................................
13 Error-bound for flow-reversal measurements (deg): Hughes HH-02
airfoil ................................
14 Error-bound for flow-reversal measurements (deg): Vertol VR-7
airfoil ...............................
15 Error-bound for flow-reversal measurements (deg): NLR-I
airfoil ................................
16 Error-bound for flow-reversal measurements (deg): NLR-7301
airfoil ................................
17 Notes pertaining to tables 18 to 25 ..................
18 Catalog of recorded data: NACA 0012 airfoil ..............
19 Catalog of recorded data: Ames A-01 airfoil ..............
20 Catalog of recorded data: Wortmann FX-098 airfoil ...........
21 Catalog of recorded data: Sikorsky SC-I095 airfoil ..........
22 Catalog of recorded data: Hughes HH-02 airfoil ............
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Catalog of recorded data:
Catalog of recorded data:
Catalog of recorded data:
Vertol VR-7 airfoil .............
NLR-I airfoil ...............
NLR-7301 airfoil ............
30
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34
x
SYMBOLS
C chord, m
CM moment coefficient
CN normal force coefficient
FR flow reversal
HF hot-film
HW hot-wire
k reduced frequency
LS lift stall
M free-stream Mach number
MS moment stall
NFR no flow reversal detected
R tea ttachment
TI transition from turbulent to laminar flow
T2 transition from laminar to turbulent flow
t time, sec
u local velocity, m/sec
x distance along the chord, m
angle of incidence, deg
rotational frequency, rad/sec
xl
AN EXPERIMENTAL STUDY OF DYNAMIC STALL ON ADVANCED AIRFOIL SECTIONS
VOLUME 3. HOT-WIRE AND HOT-FILM MEASUREMENTS
L. W. Carr, W. J. McCroskey, K. W. McAlister, and S. L. Pucci
U.S. Army Aeromechanics Laboratory (AVRADCOM), Ames Research Center
and
O. Lambert
Service Technique des Constructions Aeronautiques, Paris, France
SUMMARY
Detailed unsteady boundary-layer measurements are presented for eight airfoils
oscillated in pitch through the dynamic-stall regime. The present report (the third
of three volumes) describes the techniques developed for analysis and evaluation of
the hot-film and hot-wlre signals, offers some interpretation of the results, and
tabulates all the cases in which flow reversal has been recorded.
INTRODUCTION
The study of dynamic stall of oscillating airfoils has demonstrated the need for
obtaining detailed boundary-layer data during the stall process. Results from the
present experiment show that boundary-layer characteristics can be signifi=antly
altered by airfoil shape, and that the boundary-layer behavior is sensitive to many
parameters associated with the airfoil motion. These conclusions are based on analy-
sis of signals from hot-wire and hot-film probes mounted near or at the surface of
the various airfoils. However, evaluation of hot-wire data is very subjective, and
presents a formidable analytical task. The present report describes the techniques
developed for analysis and evaluation of the hot-film and hot-wlre signals, offers
some interpretations of the results, and tabulates all the cases in which flow-
reversal data have been recorded. An overview of the experiment has been presented
in reference I; a detailed summary of this test and the experimental conditions that
were studied is presented in volume 1 of the present report; details of the pressure
distribution results, along with lift and moment data are presented in volume 2.
The present report presents the corresponding details of the viscous flow measure-
ments that were obtained.
DESCRIPTION OF EXPERIMENTAL PROCEDURES
The experiment was designed to allow accurate testing of various airfoils under
virtually identical operating conditions. Therefore each airfoil profile wasmachined into a shell which could be attached to the metal spar that contained all
the instrumentation. After each airfoil profile was tested, the instrumentation was
removed from the shell; it then remained with the spar, ready for installation of
the next shell. In this way, the various profiles could be tested using identical
instrumentation and oscillation mechanisms; details of this system are presented in
reference i; figure I is a diagram of the spar with a shell installed. Instantaneous
single-surface pressure measurements were obtained for a wide range of test condi-
tions. Hot-wire, hot-film measurements, or both, were made near the airfoil surface
to determine the flow-reversal characteristics for each test condition. Three dif-
ferent types of hot-wlre anemometer sensors were used during the oscillating airfoil
test: hot-film surface skin-friction gages, dual hot-wire probes, and trlple-wlre
flow-reversal sensors. The most common configurations had either six hot-films along
the airfoil upper surface, or one hot-film at the leading edge (x/C = 0.025) and five
hot-wires distributed along the upper surface. The data were recorded on 32-channel
analog tape, with a timing code that allowed comparison of hot-wire data and the
pressure data, which were recorded separately for each test condition.
DATA ANALYSIS AND INTERPRETATION
Skin-Fric tion Gage
The skin-friction gage that was used during a major portion of the test program
consisted of an alumina-coated platinum surface element epoxied into a metal sleeve
(see fig. 2). This sensor, which was very resistant to damage, was used for much of
the oscillating airfoil test program. However, the characteristics of this probe
design must be taken into account when analyzing the output signals.
The output from the hot-film probe is related to the shear stress; when flow
reversal occurs, the instantaneous value of shear stress passes through zero, and
there is a local minimum in the resultant signal. Unfortunately, a significant part
of the energy supplied to the probe element is transmitted from the element to the
substrate of the gage. This heat transfer results in a relatively high dcmoffset in
the output voltage of the probe. In addition, this heat transfer causes the minimum
value of the hot-film signal to decrease slowly with time, even when the flow is
fully separated (with a nominal shear-stress value = 0). These effects can make the
interpretation of the signal somewhat difficult.
Figure 3 presents an example of the output from skin-friction gages mounted
near the leading edge of the Ames A-01 airfoil during oscillation. At the marker
"TI," the flow has passed through transition from turbulent to laminar flow, with a
resultant reduction in shear stress and decrease in fluctuation intensity. The flow
remains laminar during the low-angle portion of the cycle; as the angle increases,
transition to turbulent flow occurs (at "T2"), and the skln-frlction gage shows a
corresponding increase in signal magnitude, as well as an increase in fluctuation
amplitude. The next major event, marked by "FR," is the occurrence of flow reversal;
this results in a drop in the magnitude of the shear stress. Note that the signal
does not remain constant, even though the airfoil flow has separated; this continuing
decrease is associated with the heat-transfer effects outlined earlier. Finally,
marker "R" indicates the point when flow reattaches to the airfoil (during the down-
stroke), beginning the oscillation cycle once more.
Unfortunately, the relatively crisp delineation of flow conditions that appears
in figure 3 is not always present. Figure 4 shows an example of a less clear case
of leading-edge flow: here, the development of flow reversal is relatively slow,
and the decreasing of thesignal to its minimum is difficult to separate from the
decreasing of the minimum itself. The estimated flow-reversal points are marked by"FR."
Hot-Wire Probe
Hot-wire anemometermeasurementswere performed using a dual-wire probe (seefig. 5); this dual-wlre approach was chosen to reduce the chance of interruption ofthe test as a result of wire breakage; since both wires were being recorded, the
loss of either wire would not mean the loss of flow-reversal information at that
x-statlon. The output signal from a hot-wlre probe is a nonlinear function of the
local velocity; therefore, the signals were linearized and scaled such that the
resultant signal was approximately proportional to the associated velocity. Figure 6
shows a representative example of hot-wlre data for flow near the leading edge of the
FX-098 airfoil.
As the angle of attack increases, transition to turbulent flow occurs at
x/C = 0.025; this is observed at "T2" in figure 6 for hot-wlre probe HWI. Note that
there is no dramatic change in the output signal magnitude. Transition on airfoils
occurs at low angles of attack, for conditions where the boundary layer is thin. In
these conditions, the hot-wlre probe is often near or at the edge of the boundary
layer. Therefore, the change of the velocity profile during transition has little
or no effect on the value of U; transition will mainly be marked by changes in the
fluctuation level. The next major flow phenomenon is marked by "FR"; at this point
the flow has separated from the airfoil, causing an abrupt decrease in the local
velocity. Note that the hot-wlre signal changes abruptly to zero, and then contin-
ues at a well-defined constant value (compare with the hot-film output of fig. 3).
Later, reattachment occurs (at "R"); as the minimum angle is approached, the flow
becomes laminar again, and the cycle repeats.
As was noted for the hot-film, the hot-wire results are not always clearly
delineated. Figure 7 shows a hot-wire signal measured near the trailing edge of the
VR-7 airfoil which was difficult to evaluate. The turbulence level in this signal
is very high, and is masking the development of the periodic component of the signal.
Because this turbulent component is superimposed on the periodic part of the signal,
the instantaneous value of the signal reaches zero long before and after flow
reversal of the ensemble-averaged flow (marked as "FR" in the figure) would have
occurred. Therefore, the error band for signals measured near the trailing edge is
significantly larger than those associated with leading-edge, or mldchord locations.
Reverse-Flow Sensors
A specially designed hot-wlre probe was developed for evaluation of the flow
reversal on the VR-7 airfoil. This airfoil has trailing-edge flow reversal during
almost all unsteady flow conditions, and a better method was needed for determining
the reversal point under these conditions. The probe is described in detail in ref-
erence 2; operation is based on the use of a highly heated center wire, with two
additional wires, one upstream and one downstream of this heater, operated at low
overheat ratio. These additional wires detect the heated wake of the center wire,
and a comparison circuit is used to determine the instantaneous flow direction. This
probe system can detect both the magnitude and the direction of the local flow, and
is especially effective in regions of hlgh-turbulence, low-veloclty flow. Examples
of the output from this probe are presented in figure 8; a diagram of the probe is
presented in figure 9.
Averaging Techniques
Ensemble-averaging is often used to extract determinate signals from unsteadyturbulent flow data, and this approach was applied to the present hot-wire data.Figure i0 presents the results of an ensemble-average of i00 cycles of the hot-wiresignals on the VR-7 airfoil. It is evident in this figure that cyclic averagingsmears the flow-reversal signal (to the point where no approach to zero voltage isobservable in the averaged signal). In contrast, note the data for the last cycledigitized (shownas dotted in fig. i0). In this case, there are several instancesof zero velocity; there are also indications of vortex motion on the airfoil (in the40, 60, and 80 percent x/C wire outputs), which cannot be observed in the averageddata. There were small but significant variations in the angle at which flowreversal occurred between one cycle and the next; therefore, averages based onmechanical timing marks were not always able to capture the flow phenomena. In fact,this variation was sufficient in the present case to completely obscure the flow-reversal point in the data (in order to properly correlate these data, a true condi-tional ensemble-averaging technique would be needed, possibly triggered by a changein the character of the leadlng-edge pressure). Therefore, although someof the hot-wire and hot-film data were digitized and cyclically averaged, the analysis presentedin this report has been based on visual evaluation of the analog signals for each ofseveral cycles, after which the values of _t associated with flow reversal for agiven sensor were averaged.
Example of Signal Analysis
Figure ii showsan example of a set of hot-wire and hot-film analog signalsobtained during one period of oscillation. The first three signals are the angle ofattack, the lift coefficient, and the momentcoefficient, showing the llft stall(LS) and the momentstall (MS). The next six signals comefrom anemometersensors:one hot-film near the leading edge (HFI), and five hot-wire probes (HWI to HW6).The markers on these signals refer to the various events that have an effect on thehot-wlre and hot-film readings: FR-- initiation of reversed flow; R -- reattachmentof flow; TI -- transition from turbulent to laminar flow; T2 -- transition from laminarto turbulent flow (as determined from hot-film signals).
RESULTS
Results similar to these have been analyzed for all eight airfoils. In particu-lar, the phase angle wt, at which flow reversal first appears at the x/C locationof each hot-wire or hot-film probe, has been documentedfor a range of Machnumbers,frequencies, and stall severity for each airfoil. Thesephase angles, determined bythe techniques outlined earlier, have been recorded in degrees measured through theoscillation cycle, referenced to the meanangle, for ds/dt > 0. Table i presents asummaryof the analyzed flow-reversal data. The Machnumber studies were performedfor e = 15° + i0 ° sin _t, k = 0.i, and cover Mach number conditions that range from
incompressible values (M_ _ 0.035) to ones that include small regions of supersonic
flow near the leading edge (M_ = 0.30). The "llght-stall" frequency studies present
data for a range of frequencies at M = 0.30, where the amplitudeand mean angle have
been chosen to cause a slight overshoot of the static stall angle associated with
each airfoil during the oscillatory motion. The "deep-stall" study presents data for
a range of frequencies at M_ = 0.30, _ = 15 ° + i0 ° sin _t (deep stall has been
defined in ref. i as a condition in which a fully developed vortex is formed during
4
the oscillation cycle). The experimental data in deep stall were less amenable toanalysis -- the results were more subjective and in somecases inconclusive. There-
fore, the results for only three airfoils are reported.
The results of these surveys are presented graphically in figures 12 to 31.
Figures 12 to 19 present Mach number effects for deep-stall conditions; figures 20
to 27 present frequency effects for llght-stall conditions; and figures 28 to 31
present frequency effects for deep-stall conditions. These data are also presented
in tabular form in tables 2 to 9. The error bounds for these surveys are presented
in tables i0 to 16. Finally, a catalog of all the hot-film and hot-wlre data that
were recorded is presented in tables 17 to 25, tabulated according to the correspond-
ing pressure data (stored in digital form, as explained in vols. i and 2).
5
REFERENCES
lo
,
McCroskey, W. J.; McAlister, K. W.; Carr, L. W.; Pucci, S. L.; Lambert, O.; and
Indergand, R. F.: "Dynamic Stall on Advanced Airfoil Sections," J. of the
American Helicopter Society, July 1981.
Carr, L. W.; and McCroskey, W. J.: "A Directionally Sensitive Hot-Wire Probe
for Detection of Flow Reversal in Highly Unsteady Flows," in International
Congress on Instrumentation in Aerospace Facilities 1979 Record, Sept. 1979,
pp. 154-162.
6
TABLEi.- SUMMARYOFANALYZEDFLOW-REVERSALDATA
Airfoil
NACA0012,A-01FX-098SC-I095HH-02
VR-7
NLR-I
NLR-7301
Mach No. a
Film d
Film d
Wireg
Film d
Film d
Comb. e
Film d
Film d
Light stall c
Film d
Film _
C omh. e
Film d
Comb .e
Comb. e
Deep stall b
Comb .e
Wireg
Comb .f
aMach number sweep a = 15 ° + i0 ° sin _t, k = 0.I.
bFrequency sweep, a = 15 a + i0 ° sin _t, M = 0.295.
CFrequency sweep, a = ao + al sin _t, M = 0.29.dHot-film shear-stress gage.
"Hot film at x/c = 0.025; hot wire at all other
locations.
fHot wire at 0.025, 0.i0, 0.25; reverse-flow sensors
at x/c - 0.4, 0.6, 0.8
gHot-wire velocity probe.
TABLE 2.- PHASE ANGLE OF FLOW REVERSAL: NACA 0012 AIRFOIL
Mach
No.
0.036
.076
.ii0
.145
.185
.220
.250
.270
.280
.290
.295
0.025
i0.0
50.0
59.5
67.0
60.5
43.5
21.5
14.5
10.5
8.0
8.5
I O. i00
x/c
0.250 0.400 O.6OO O. 80O
a = 15 ° + i0 ° sin _t, k ffi0.I
1.0 3.0
40.0 35.5
44.5 40.0
50.5 50.5
45.0 41.5
38.0 36.5
26.0 29.0
18.0 21.0
21.0 21.5
16.0 20.5
13.5 16.5
0.0
46.5
54.5
61.5
53.0
39.0
24.5
16.5
15.0
13.0
10.5
6.0
23.0
35.5
47.0
36.5
35.5
29.5
28.0
23.0
24.0
22.0
12.5
15.0
19.5
35.0
30.0
27.5
33.5
28.5
24.0
20.5
20.5
Reduced x/c
freq. 0.025 0.i00 0.250 0.400 0.600 0.800
Ref.
frame
8013
8115
2320
2314
2310
2208
2204
2202
2200
2103
2101
Ref.
frame
a - 12 ° + 5 ° sin _t, M = 0.295
0.025
.050
.i00
.200
NFR
NFR
NFR
35.5
55.5
32.5
34.0
44.0
48.0 37.0 32.5
37.0 38.0 33.0
42.5 45.0 47.5
54.0 59.0 64.0
26.5
31.0
41.0
71.0
7201
7204
7206
7208
ORIGINAL PAGE IS
OF POOR QUALITY
TABLE 3.- PHASE ANGLE OF FLOW REVERSAL: Ames A-01 AIRFOIL
Mach
No.0.025
x/c
0.I00 0.250 0.400 0.600 0.800
Ref.
frame
= 15 ° + i0 ° sin _t, k = 0.I
0.076
.ii0
.185
.220
.250
.280
.295
Reduced
freq.
48.5
56.5
56.5
53.5
29.5
18.0
12.0
48.5
47.5
53.0
46.5
29.0
19.5
16.0
32.5
35.5
31.5
32.5
26.0
19.5
17.5
26.5
33.5
34.0
33.0
29.5
23.0
19.5
25.5
37.5
38.0
39.0
32,0
27.0
23.0
22.5
43.5
44.5
28.5
33.5
31.5
28.5
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
= ii ° + 5 ° sin _t, M = 0.295
24400
24316
24219
24210
24202
24118
24108
Ref.
frame
0.010
.050
.010
NFR 63.5 59.5 1 59.5 59.0 I 55.5NFR 96.0 72.0 68.5 65.5 56.5
Data too irregular to be analyzed
30202
25215
25217
a = 15 ° + i0 ° sin _t, M = 0.295
0.010
.025
.05
.i00
.150
NFR
12.5
12.0
14.5
23.0
12.0
15.5
16.0
17.5
28.5
6.5
11.5
12.0
17.5
23.5
5.0
ii.0
14.5
19.0
28.0
5.0
Ii.0
18.5
27.5
33.5
2.0
ii.0
24.5
31.0
38.5
30021
31016
31018
31019
31020
ORIGINAL PAGE IS
OF POOR QUALITY
TABLE 4.- PHASE ANGLE OF FLOW REVERSAL: Wortmann FX-098 AIRFOIL
Mach x/c
No.0.025 . 0.i00 0.250 0.400 0.600
= 15 ° + i0 ° sin _t, k = 0.i
Ref•
frame0.800
0.036
.076
•ii0
•185
.220
.250
•280
.295
Reduced
freq.
2.5
36.5
43.0
37.0
22.5
14.5
9.0
6.5
-1.2
34.5
39.5
37.0
24.5
15.5
12.0
12.5
-3.6
27.0
32.5
36.5
25.0
18.0
18.0
15.5
-2.0
18•5
24.5
33.5
26.5
18.0
20•0
16.5
-4.6
14.5
16.5
31.0
24.0
17.5
17.5
18.5
-8.6
4.5
10.5
24.0
21.5
21.5
15.5
21.0
16022
16106
16115
16 201
16301
16 309
22209
22202
x/c Ref.
frame0.25 0.i00 0.250 0.400 0.600 0.800
= i0 ° + 5 ° sin _t, M = 0.295
0.010
.025
•050
.i00
•150
.200
NFR
NFR
NFR
NFR
68.0
64.0
NFR
NFR
NFR
72.0
76.0
69.5
67.0
95.0
69.0
77.5
82.0
79.0
67.0
93.5
66.0
75.5
76.0
68.5
66.5
82.0
61.5
70.0
81.0
75.0
63.0
49.0
57.0
66.0
85.0
83•0
21201
22223
22300
22301
22302
22303
= 15 ° + 10 ° sin _t, M = 0•295
0.010 -99.9
.025 0.0
•050 0.5
.i00 i0.0
37•5
3.5
1.5
12.5
4.5
3.5
4.5
14.5
2.5
3.5
6.5
15.0
2.5
3.5
8.0
19.0
a = 15 ° + I0 ° sin tot, M = 0•185
0.0 21102
5.5 17118
9.5 17123
20.5 17201
0. 050
.i00
.150
14.0
20.5
28.0
15.5
21.5
30.0
17.5
25.0
32.0
16.0
24.0
32.0
I0.0
21.0
33.5
6.5
19.0
26.5
17102
17108
17110
TABLE5.- PHASEANGLEOF
O_'_,,_n' PAGE ;g
OF POOR QUALIT'Y
FLOW REVERSAL: Sikorsky SC-I AIRFOIL
Mach
No.0.025
x/c
0.i00 0.250 0.400 0.600 0.800
= 15 ° + i0 ° sin wt, k = 0.i
Ref.
frame
0.076
.ii0
.185
.220
.250
.280
.295
Reduced
freq.
33.5
43.5
42.0
32.0
22.0
15.0
9.0
30.5
41.0
38.0
28.5
18.5
14.5
12.0
24.5
28.0
33.0
26.5
22.5
18.5
15.0
21.0
28.0
35.0
24.5
26.0
20.5
18.0
15.5
36.5
36.5
28.5
29.5
23.5
22.5
23.5
42.5
48.5
35.5
34.5
27.5
16.5
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
a = Ii ° + 5 ° sin _t, M = 0.295
33023
33107
33111
33206
33208
33216
33303
Ref.I
frame
0.050 -99.9
.i00 66.0
70.0 61.0
62.5 61.5
52.0 65.0 67.5 37220
63.5 65.5 67.0 37222
TABLE 6.- PHASE ANGLE OF FLOW REVERSAL: Hughes HH-02 AIRFOIL
Mach
No.O. 030
x/c
0.120 0.250 0.380 0.560
= 15 ° + i0 ° sin _t, k -- 0.i
Ref.
frame0.750
0.076
.ii0
.185
.220
.250
.280
.295
Reduced
freq.
40.0
48.5
52.5
25.0
15.0
7.0
5.0
40.0
45.0
42.0
25.0
16.0
9.0
9.5
32.5
4O.5
40.0
28.0
17.0
11.5
15.1
28.0
36.5
38.5
31.5
19.5
13.0
18.5
17.5
30.5
37.0
36.0
24.5
14.5
13.0
x/c
0.025 0.i00 0.250 0.400 0.600
= i0 ° + 5° sin mr, M = 0.295
11.5 42112
13.5 42322
32.4 42303
15.5 42310
18.0 42314
5.8 42319
13.0 42211
Ref.
frame0.800
0.010
.025
.050
.i00
.150
.200
NFR
NFR
53.5
58.5
56.0
57.5
72.0
78.5
60.0
67.0
67.0
67.0
68.5
74.5
64.5
78.0
80.0
79.0
59.0
60.0
62.5
79.0
83.5
86.0
47.5
49.0
57.0
84.0
94.0
94.0
20.5
33.0
36.5
50.0
54.0
58.0
44020
44022
44100
44105
44107
44113
i0
OF FSOR ,_U_'_-7_
TABLE7.- PHASEANGLEOFFLOWREVERSAL:Vertol VR-7 AIRFOIL
MachNo.
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
= 15 ° + I0 ° sin _t, k = 0.i
Ref.
frame
0.076
.ii0
.185
.220
.250
.280
.295
Reduced
freq.
48.0
51.5
54.0
38.0
26.0
24.5
16.5
0.025
46.0
49.0
49.5
40.5
26.5
25.5
19.0
0.010 I
a = 15 °
37.0 30.0
44.0 32.5
45.5 37.8
39.5 36.0
29.0 29.5
30.0 33.0
26.5 26.0
xlc
0.250 0.400
I0.5
15.0
25.0
25.0
25.5
23.5
19.0
0.600
+ 5 ° sin _t, M = 0.295
-4.0
-6.0
3.5
4.5
7.0
7.0
2.0
0.800
47200
47207
47214
47218
47302
47306
45100
Ref.
frame
O. i00
.025
.050
.i00
.150
.200
NFR
NFR
NFR
NFR
41.5
27.5
NFR
NFR
31.0
36.0
44.5
32.5
-3.0
15.0
26.5
36.0
49.5
48.0
-Ii .0
8.0
23.0
30.0
41.5
44.0
-14.0
-Ii.0
2.5
17.5
39.5
30.0
-63.0
-39.0
-35.0
-23.0
2.0
9.5
45204
45206
45208
45210
45212
45214
= 15 ° + i0 ° sin wt, M = 0.295
0.025
.050
.i00
.150
NFR
17.5
22.0
26.0
14.5 18.0
18.5 20.0
28.0 30.5
37.0 43.0I I
6.5
14.0
27.0
29.0I
-4.5
0.0
8.0
9.5Z
50021
50019
50017
50015
ii
ORIGINAL PAGE _3
OF POOR QUALITY
TABLE 8.- PHASE ANGLE OF FLOW REVERSAL: NLR-I AIRFOIL
Mach
No.
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
a = 15 ° + i0 ° sin mt, k = 0.i
Ref.
frame
0.076
.ii0
.185
.200
.220
.250
.280
.295
Reduced
freq.
17.0
29.0
36.0
30.5
20.5
9.5
1.5
0.0
17.5
26.0
32.0
30.5
17.5
12.5
ii.0
6.0
18.0
23.0
26.0
27.0
18.5
15.5
14.5
9.5
21.5
26.0
28.5
33.5
21.0
21.0
18.0
12.5
25.5
30.0
33.0
41.0
21.5
24.0
21.5
17.5
32.5
36.0
38.0
41.0
29.5
24.5
27.0
24.0
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
62021
62105
62113
62115
62209
62211
62218
62308
Ref.
frame
= i0 ° + 5 ° sin _t, M = 0.295
0.025
•I00
.200
NFR
45.0
52.0
43.5
50.0
55.0
44.0
47.0
54.5
42.0
49.0
55.0
36.5
55.0
61.0
35.5
59.0
66.5
63109
63113
63115
TABLE 9.- PHASE ANGLE OF FLOW REVERSAL: NLR-7301 AIRFOIL
Mach
No.
xlc
0.025 0.i00 0.250 0.400 0.600 0.800
a = 15 ° + i0 ° sin wt, k = 0.i
Ref.
frame
0.ii0
.185
.250
Reduced
freq.
84.0
98.5
69.5
78.5
93.5
58.5
75.0 66.5
82.5 76.0
55.0 52.5
x/c
56.0
50.5
48.0
24.0
35.0
38.5
0.025 0.i00 0.250 0.400 0.600 0.800
62105
62113
62211
Ref.
frame
a = 15 ° + 5 ° sin _t, M = 0.295
0. 010
.025
.050
•i00
.150
.200
NFR
NFR
NFR
NFR
NFR
NFR
56.5
64.0
68.5
34.0
37.5
35.0
54.5
57.5
60.0
43.0
46.0
53.0
51.0
53.5
56.5
44.5
51.0
64.5
48.5
48.0
43.5
43.0
61.0
44.0
40.5
16.0
2.0
Ii.0
24.0
23.0
68020
68101
68103
68105
68110
68112
12
TABLEi0.- ERROR-BOUNDFORFLOW-REVERSALMEASUREMENTS(deg):
NACA 0012 AIRFOIL
Mach x/c Ref.
No. frame0.025 0.i00 0.250 0.400 0.600 0.800
Corresponds to table 2: _ = 15 ° + i0 ° sin _t, k = 0.i
0.035
.073
.ii0
.145
.185
.185
.220
.250
.270
.280
.290
.295
Reduced
freq.
4.0
2.0
1.5
5.0
0.5
2.5
3.0
0.0
2.0
2.0
0.5
0.5
0.0
0.0
0.5
2.5
2.0
3.0
1.0
0.0
2.0
2.0
2.5
1.5
1.5
1.5
0.5
4.0
1.0
2.5
0.5
1.5
2.0
1.0
1.5
1.5
1.0
2.5
1.0
3.5
1.0
1.5
2.5
2.0
2.5
2.0
2.5
0.0
1.5
5.0
3.0
3.5
3.5
4.0
2.5
0.5
1.5
2.0
1.0
0.0
2.0
3.0
3.0
3.5
2.0
9.0
3.0
1.5
0.0
3.0
1.5
2.5
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
8103
8115
2320
2314
8221
2310
2208
2204
2202
2200
2103
2101
Ref.
frame
Corresponds to table 2: a = 12 ° + 5 ° sin mr, M = 0.295
0.025
.050
.i00
.200
NFR
NFR
NFR
5.0
5.0
0.0
2.0
1.0
4.0
2.0
2.0
3.5
2.0
5.0
2.5
5.0
2.0
2.0
4.0
2.0
1.5
2.0
4.6
1.5
7201
7204
7206
7208
13
ORIGINAL PP--,GE |2
OF POOR QUALITY
TABLE ii.- ERROR-BOUND FOR FLOW-REVERSAL MEASUREMENTS (deg):
Ames A-01 AIRFOIL
Mach
No.
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
Corresponds to table 3: a = 15 ° + i0 ° sin _t, k = 0.i
Ref.
frame
0.076 1.5
.ii0 1.0
.185 1.5
.220 2.0
.250 1.0
.280 0.0
.295 0.5
Reduced
freq. 0.025
1.5
0.5
2.5
3.0
1.5
1.5
1.5
6.0
2.0
5.0
3.0
1.0
1.5
0.5
3.0
2.0
4.0
3.5
0.0
1.5
1.5
0.5
2.0
1.2
6.5
4.0
2.0
1.5
6.0
3.0
3.0
5.0
2.0
4.0
3.0
x/c
0.i00 0.250 0.400 0.600 0.800
24400
24316
24219
24210
24202
24118
24108
Ref.
frame
Corresponds to table 3: _ = Ii ° + 5° sin _t, M = 0.295
0. 010
.050
•i00
NFR
NFR
I3.0 4.0 4.0 3.5 I
6.5 2.5 2.5 2.0 IData too irregular to be analyzed
2.5
5.5
30202
25215
25217
Corresponds to table 3: a = 15 ° + I0 ° sin _t, M = 0.295
O. 010
.025
•050
•i00
.150
NFR
2.0
1.0
1.3
1.5
3.0
2.5
1.0
1.0
3.0
2.0
3.0
0.5
1.0
2.5
3.0
1.0
0.0
1.5
1.0
1.0
1.0
2.5
4.0
1.5
1.5
1.0
3.5
2.0
0.5
30021
31016
31018
31019
31020
14
OF POOR QU&L]3"_'
TABLE I2.- ERROR-BOUND FOR FLOW-REVERSAL MEASUREMENTS (deg):
Wortmann FX-098 AIRFOIl,
Mach
No.
x/c
0.025 0.I00 0.250 0.400 0.600 0.800
Corresponds to table 4: cx = ]5 ° + i0 ° sin ,_t, k = 0.I
0.036
.076
.110
.185
•220
.250
Reduced
freq.
0.5
2.0
1.0
1.0
1.0
1.5
].5
3.0
2.0
1.0
1.0
1.0
0.0 l .5
1.5 ] .0
3.0 ] .0
1.0 3.0
] .5 2.0
] .0 0.5
x/c
1.0
l. 0
l. 0
3.0
2.0
1.5
1.0
2.0
1.0
2.0
3.0
2.0
0.025 0.i00 0.250 0.400 0.600 0.800
Corresponds to table 4: a = i0 ° + 5 ° sin wt, M = 0.295
16022
16106
16115
16201
16301
16309
Ref.
frame
0.025
.050
.i00
.150
.200
NFR
NFR
NFR
2.5
3.0
NFR
NFR
2.0
0.5
1.5
3.0
2.0
3.0
1.0
0.0
3.5
2.0
1.0
1.5
0.0
7.0
2.0
1.0
1.0
1.0
2.5
2.0
1.0
1.0
3.5
22223
22300
22301
22302
22303
Corresponds to table 4: a = 15 ° + i0 ° sin wt, M = 0.295
0.010 NFR
.025 0.0
.050 1.0
.i00 1.0
2.0
0.0
1.5
2.0
1.0
0.0
1.0
0.5
2.0
0.0
0.5
2.0
2.0
0.0
1.5
2.0
2.5
0.0
0.5
5.0
22102
17118
17123
17201
Corresponds to table 4: a = 15 ° + i0 ° sin _t, M = 0.295
0.050
.I00
•150
14.0
20.5
28.0
15.5
21.5
30.0
17.5
25.0
32.0
16.0
24.0
32.0
i0.0
21.0
33.5
6.5
19.0
26.5
17102
17108
17110
15
ORIGINAL PAGE I_
OF POOR QUALITY
TABLE 13.- ERROR-BOUND FOR FLOW-REVERSAL MEASUREMENTS (deg):
Hughes HH-02 AIRFOIL
Mach
No.
x/c ! Ref.
I l i i0.030 0.120 0.250 0.380 0.560 0.750 frameI
Corresponds to table 6: a = 15 ° + iO ° sin cat, k = 0.1
0.076
.Ii0
.185
•220
.250
.280
.295
Reduced
1.0
1.5
3.0
1.0
1.0
0.0
1.0
1.0
2.0
2.0
1.0
1.5
3.0
1.0
3.0
3.0
3.0
1.0
2.0
2.0
7.0
x/c
1.5
7.0
2.0
1.0
1.5
2.0
1.0
3.0
8.0
3.0
1.5
1.0
3.0
1.0
4.0 42122
2.0 42322
3.0 42303
4.0 42310
4.0 42314
1.0 42319
i .0 42211
Ref.
freq. 0.050 0.i00 0.250 0.400 0.600 ] 0.800 frame
Corresponds to table 6: _ = i0 ° + 5 ° sin _t M = 0.295
0.010
.025
•050
•i00
.150
•200
NFR
NFR
1.0
2.0
1.0
1.0
3.5
6.5
1.5
0.5
3.0
1.0
5.0
6.5
1.5
2.5
3.0
0.0
4.5
3.5
1.5
2.0
3.0
1.0
2.5
3.0
2.0
2.0
6.5
2.0
2.0
3.0
2.0
2.0
0.0
5.5
44020
44022
44100
44].05
44107
44113i
16
ORIGINAL PAGE IS
OF POOR QUALITY
TABLE 15.- ERROR-BOUND FOR FLOW-REVERSAL MEASUREMENTS (deg):
NLR-I AIRFOIL
Maeh
No.
x/e
0.025 0.i00 0.250 0.400 0.600 0.800
Corresponds to table 8: a = 15 ° + i0 ° sin u_t, k = 0.i
Ref.
frame
0,076
.ii0
.185
•200
.220
•250
.280
.295
Reduced
freq.
0.5
0.5
1.0
1.0
1.0
0.5
1.0
0.0
2.0
4.0
3/01.0
0.0
1.0
2.0
1.0
1.0
2.0
3.0
2.0
0.5
1.5
1.0
2.0
2.0
2.5
1.0
2.0
1.0
1.0
0.0
1.0
2.5
4.0
3.0
2.0
3.0
1.5
0.5
3.0
5.0
1.0
2.0
2.0
1.0
1.0
1.0
1.5
xlc
0.025 0.i00 0.250 0.400 0.600 0.800
Corresponds to table 8: _ = 10 ° + 5 ° sin ut, M = 0.295
62021
62105
62113
62115
62209
62211
62218
62308
Ref.
frame
0.025 NFR
.i00 0.0
.200 2.0
3.5
0.5
0.5
3.0
5.0
5.5
3.5
2.0
2.0
0.5 O.5
2.5 2.0
0,0 2.5
63109
63113
63115
TABLE 16.- ERROR-BOUND FOR FLOW-REVERSAL MEASUREMENTS (deg):NLR-7301 AIRFOIL
Mach
No.
x/c
0.025 0.i00 0.250 0.400 0.600 0.800
Corresponds to table 9: _ = 15 ° + I0 ° sin wt, k = 0.i
Ref.
frame
0.ii0
.185
.250
Reduced
freq.
4.0
5.0
2.5
4.0
6.0
2.5
i0.0 13.0
7.0 7.0
1.0 2.0
x/c
ii.0
4.0
5.0
0.025 0.i00 0.250 0.400 0.600
Corresponds to table 9: a = 15 ° + 5 ° sin mr, M = 0.295
6.0 67121
1.5 67221
I.i 67306
Ref.
frame0.800
0.010 NFR
.025 NFR
.050 NFR
.i00 NFR
.150 NFR
.200 NFR
1.0
2,0
3.5
1.0
1.5
0.5,.J
1.5
3.0
3.5
1.0
1.0
4.0
0.5
2.0
5.5
2.0
4.5
4.0
1.5 2.0 68020
1.5 4.0 68101
O. 5 2,5 68103
0.5 5.0 68105
2.5 5.5 68110
Ii,0 9.0 68112
PRECEDING PAGE BLANK NOT FILMED18
r',_" POO.q Qo_.I,'TY
TABLE 17.- NOTES PERTAINING TO TABLES 18 TO 25
DATA LISTED IN ORDER A FRAMES STORED ON DIGITAL TAP[B FRAMES ARE ON ANALOG TAPE ONLY
A FRAME - CATALOG ENTRY FOR PRESSURE DATATRIP - TRIP %S PRESENT - (YIES. OR (RIOTYPE - TEST CONDITIONS (STIEADY. OR IUNISTEADY
AO - MEAN ANGLE OF OSCILLATION. DEGREESA1 - AMPLITUDE OR OSCILLATION, DEGREESQ - FREE STREAM DYNAMIC PRESSURE. PSIM - FREE STREAM MACN NUMBERRE - FREE STREAM REYNOLDS NUMBER
FRE@ - DIMENSIONAL FRE@UENCY. HERTZB FRAME - CATALOG ENTRY FOR HOT-FILM AND HOT-WIRE DATA
19
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21
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34
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35
ORIGINAL PAGE ',S
OF POOR QUALITY
OUTLINE OF /OUTER MODEL SHELL
AXIS OF ROTATION _ /
OUTLINE OF
I
• • • • a%_
CONNECTED TO DRIVE SYSTEM
PLEXlGLASS WINDOWIN TUNNEL CEILING
/Y,/,/z$
/ PRESSURE ORI FICES
HOT WIRE SENSORS
III 25ram GROUND PLANE
"', t__7_ I ON TUNN/_L FLOOR
.K_\\\\\\\\\\\\\\\\\\\\\\\\I. | I I:{_\" _\\\\_\\\\\\\_'_
_///////////////////////_ _1 \ _/////////////////////_
Io
i
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IIIIII
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!
Figure I.- Diagram showing installation of spar and alrfoil shell in tunnel.
36
ORIG_i'_:._".-. :'_.,'._ ''.j
OF POOR _L_.c_...TT_'
3" TEFLON-COATEDLEAD Wl RES
ALUMINA-COATED PLATINUM FILMSTRIP _ //
(0.12) (0.0005) WIDE BY (1.0) (0.4) LONG _J//
OVERALL LENGTH J
NOTE: PROBE MODIFIED FROM TSI MODEL 1237
FLUSH SURFACE SENSOR
Figure 2.- Diagram of hot-film skin-frlction gage.
NF_-J/T1 T2 -_-_R_ ""_'_ - _ _ I
HF2 _ ' V--"I'_ -.... _ I
21r
Figure 3.- Response of hot-film skln-frictlon gages mounted on Ames A-01
airfoil during airfoll oscillation in pitch (_ = 15 ° + i0 ° sin _t,
k = 0.i0, M_ = 0.22).
37
PAGE _:3
ORIGINALOF POOR QUALITY
x/c = 0.025
FR
Figure 4.- Response of hot-film skin-friction gages at surface of NACA 0012
airfoil during airfoil oscillation in pitch (_ = 15 ° + I0 ° sin mr,
k = 0.I0, M= = 0.295).
=22 AWG SOLID COPPER WIREOR SIMILAR
6 in. PIGTAIL LEADS
ALL DIMENSIONS IN INCHES
DESIGN BASED ON TSI PROBE 1244
3/4 -_ _'-_ 0.030 -+0.005
FLATFACE / (3.2) diam/
0+040 APPROX.
Figure 5.- Diagram of dual-element hot-wire probe.
38
HW1 ,_
x/c = 0.025
HW2 mm
x/c = 0.10
72
FR R
Figure 6.- Response of hot-wlre anemometer probes on Wortmann FX-098 airfoil
during airfoil oscillation in pitch (_ = 15° + i0° sin _t, k = 0.10,
M= = O.li).
HW6x/c = 0.80
FR
Figure 7.- Response of hot-wlre anemometer probe installed near trailing edgeof the Vertol VR-7 airfoil during oscillation in pitch.
39
ORIGINAL PACE IS
OF pOOR QUALITY
_c/2V = 0.15, c_= 15 ° + 10 ° sin _t M = 0.185
(A) INITIAL FLOW REVERSAL
(B) VORTEX-INDUCED UPSTREAM FLOW
(C) STAGNATED FLOW
(D) UPSTREAM MOVEMENT OF FLOW
(E) REATTACHMENT
REVERSE FLOW SENSOR
_ _/ -- zerovoltz
I HOT WIRE ANEMOMETER
ii zero voltsI IAB C DE
i '
j_ HEATER-OFF_ SIGNAL
(a) _ ALPHA
_cl2V = 0.15
M - 0.185
¢z • 15 ° + 10 = sin(_t
FR - FLOW REVERSAL
R - REATTACHMENT
(b)
RFS AT x/c - 0.40
R
R_l _ _ RFSATx/c=0,60
RFS AT x/c = 0.80
T AT x/c = 0.80
ALPHA
Figure 8.- Results obtained using trlple-wire flow-reversal sensor:(a) Typical comparison of flow-reversal sensor and hot-wire anemometer
signal (from ref. 2); (b) Progression of flow reversal up airfoil duringdynamic stall (from ref. 2).
4O
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0.051
I_ o.osol_1_._1 _O.lg'
DIMENSIONS IN mm (ft)
(0.02)
25 ,u (0.001) diamRODAR
HEATER WIRE
(SPOT WELDED)
0.02 (0.008) diam
TERMINALS
Figure 9.- Diagram of three-element, directionally sensitive hot-wire probe
(from ref. 2).
41
ORIGINAL PAGE 13OF POOR QUALITY
1.5
ALPHA
0
CM
x/c - 0.025
0
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0
x/c = 0.25
0
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0
100 CYCLE AVERAGE
..... SINGLE CYCLE
'-'% A,,!_. !L ,.. _,,
/ ! ix/¢ - 0,60 L . a &,',, ! I _ I. ._, .... - - -- --- --'- - -- -'" - ,._,.__',.. - ."_L._/,L -IV "" ' '- i: _- -- - - _
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I'11" I I I I I
0 20 40 60 80 100 120 140 160 180 200
PHASE ANGLE IN CYCLE, _t, deg
Figure i0.- Comparison of 100-cycle ensemble average and single-cycle signals
from hot-wire anemometers for Vertol VR-7 airfoil during oscillation in
pitch: --, I00 cycle average; , single cycle.
42
OF _,_ "-
CM
LS
HF1 x/c = 0.025
HW2
0.10
HW3 0.25
0.40
HW4
HW5 0.60
0.80
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0 36O_
_ot
Figure II.- Response of hot-fllm skln-frlctlon gage and hot-wlre anemometer
probes on Vertol VR-7 during oscillation in pitch (= = 15 = + i0 = sln ut,
k - 0.10, M= = 0.185).
43
ORIGI, NAL PAGE |SOF POOR QUALITY
7O
g,"ID
40
)-C_
Z
ua 30
z
N 20
60
50
10
0
-10
0
NACA-0012 MACH NO
= 0,036© = 0.076A = 0.110+ = 0.145× = 0.1850 = 0.220
= 0.250[] = 0.270
= 0.295
.2 .4 .6 .8 1.0
x/c
Figure 12.- Phase angle, _t, of flow reversal on NACA 0012 airfoil vs chord location
for a range of Mach numbers at k = 0.i, u = 15 ° + I00 sln _t --Mach numbereffects.
44
OF FC'E:_ QU#,L!TY
7O
50-
40
_z30
z
20=.=
lO
AM ES-01 MACH NO
O '= 0.076© - 0.110
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0 ,, 0.2800.295m
-10o . .4 .6 .8 1.0
x/©
Figure 13.- Phase angle, mr, of flow reversal on Ames A-01 airfoil vs chord location
for a range of Mach numbers at k = 0.i, _ = 15 ° + I0 ° sin _t --Mach number
effects.
45
ORIG;,_AL p.'_GE i_OF POOR QU/-,Liz"_'
70
6O
" 40
z
_ 30
2o
10
0
-10
MACH NOWORTMANN FX-098
O ,, 0.036© = 0.076
Z_ = 0.110+ = 0.185× - 0.2200 = 0.250
= 0.280[] = 0.295
I
1 I I I
0 .2 .4 .6 .8
x/c
1.0
Figure 14.- Phase angle, _t, of flow reversal on Wortmann FX-098 airfoil vs chordlocation for a range of Mach numbers at k = 0.I, u = 15 ° + I0 ° sin _t --Mach
number effects.
46
OF POOR Q!-.";,..i_'(
7O
6O
5O
40
30
20
10
MACH NOStKORSKY SC-1095 [] - 0.076
© = 0.110A - 0.185+ - 0.220x - 0.25Oo = 0.280
... 0.296
0 .2 .4 .6 .8 1.0x/c
Figure 15.- Phase angle, _t, of flow reversal on Sikorsky SC-i095 airfoil vs chord
location for a range of Mach numbers at k - 0.i, _ - 15 ° + i0 ° sin _t -- Mach
number effects.
47
0RiG;NAt F,_,GE_S_i= pOOR QU_,L_,'P/
7O
.°150
40,..i
>,.
Z
ua 30,,.i
Z,<M.I
_ 20-I-Q,.
10
HUGHES HH02 MACH NO
[] ,, 0.076© = 0.110/_ = 0.185+ = 0.220× = 0.250
= 0.280_7 = 0.295
-10 ! i0 .2 .4 .6 .8 1.0
x/c
Figure 16.- Phase angle, _t, of flow reversal on Hughes HH-02 airfoil vs chord loca-
tion for a range of Mach numbers at k = 0.I, s = 15° + I0° sin _t --Mach numbereffects.
48
OF ?OOR _v ......
7O
6O
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uJ 30.=I
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lO
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= 0.285
-I0
0 .2 .4 .6 .8 1.0x/c
Figure 17.- Phase ansle, _t, of flow reversal on Vertol VR-7 airfoil vs chord loca-
tion for a range of Mach numbers at k = 0.i, _ = 15 ° + I0" sin ut -- Mach number
effects.
49
ORIGINAL PAGE _S
OF POOR QUALITY
7O
6O
50
NLR-1
4O
¢J
z30
2o
10
0
MACH NO
E_ = 0.076© = 0.110Z_ = 0.185+ = 0.200× = 0.2200 = 0.250
= 0.280
[] - 0.295
-10 . , . , • , . , .
0 .2 .4 .6 .8 1.0
x/c
Figure 18.- Phase angle, _t, of flow reversal on NLR-I airfoil vs chord location for
a range of Mach numbers at k = 0.I, u = 15 ° + 10 ° sin ,._t -- Mach number effects.
50
100
9O
4O
30
20
NLR-7
I MACH NO
[] = 0.110© 0.185A 0.250
\
0 .2 .4 .6 .8 1.0
x/c
OF _JOi "_,QU:_,LFYY
Figure 19.- Phase angle, mr, of flow reversal on NLR-7 airfoil vs chord location for
a range of Mach numbers at k = 0.i, s = 15° + I0 ° sin _t --Mach number effects.
100
9O
8O
-J 70¢J>-
,U
Z
w 60,,-I
z,,<ua 50
-r"
40
3O
2O
NACA-0012 RED FREQ
= 0.0250 = 0.050L_ = 0.100-,- = 0.200
I • ! • i
0 .2 .4 .6 .8 1.0
X/C
Figure 20.- Phase angle, mr, of flow reversal on NACA 0012 airfoil vs chord location
for a range of frequencies at M_ = 0.295, _ = 12 ° + 5 ° sin mt -- light-stall
conditions.
51
OF POOR QUALFi'Y
100
9O
-D 80=,l.J
O
Z
,,, 70.J
oz
MJ
60-ra.
5O
40
(_ AMES-01 _ED FREO- 0.010© " 0.050
I ! • i i
.2 .4 .6 .8 1.0
x/c
Figure 21.- Phase angle, mr, of flow reversal on Ames A-01 airfoil vs chord location
for a range of frequencies at M_o = 0.295, _ = ii ° + 5 ° sin mt -- light-stall
conditions.
100
9O
=='_ 80
_1U
U
Z
., 70
z
,_ 60"r"a.
5O
4O
WORTMANN FX-098
X
RED FREO
[] = 0.010© = 0.025
= 0.050+ = 0.100X = 0.150O =0.200
0 .2 .4 .6 .8 1.0x/c
Figure 22.- Phase angle, mt, of flow reversal on Wortmann FX-098 airfoil vs chord
location for a range of frequencies at M= = 0.295, _ = i0 ° + 5° sin _t -- light-stall conditions.
52
100
9O
M,I.J
o>.oz
w 70_I
z<iu
_ 6o-ra,.
ORi"JrNqL P;_C:-F:_S
OF POOR QUALITY
Sl KORSKY SC-1095
I RED FREQ[] - 0.050O " 0.100
[3
• I I I I
0 .2 .4 .6 .8 1.0
x/c
Figure 23.- Phase angle, _t, of flow reversal on Sikorsky SC-I095 airfoil vs chord
location for a range of frequencies at M_ = 0.295, _ = ii ° + 5° sin _t -- light-stall conditions.
100
HUGHES-HH02
90 ...........
4O
RED FREQ
[] - 0.010O " 0.025Z_ = 0.050+ = 0.100× = 0.1500 = 0.200
Figure 24.- Phase angle, _t, of flow reversal on Hughes HH-02 airfoil vs chord loca-
tion for a range of frequencies at M_ = 0.295, _ - 10 ° + 5 ° sin _t -- llght-stall
conditions.
53
OR,=J',I,r_L Ff.: _ "_
OF POOR QUALITY
60
BOEING VERTOL VR-7
50
40
3O
20-4
¢J>-u 10z
U3.J
0z
MJ
-10
o.
-3O
-40
RED FREQ
[] = 0.010© = 0.025A - 0.050+ = 0.100× = 0.150O = 0.200
-50
0 .2 .4 .6 .8 1.0
x/c
Figure 25.- Phase angle, _t, of flow reversal on Vertol VR-7 airfoil vs chord loca-
tion for a range of frequencies at M= = 0.295, _ = 15 ° + 5 ° sin =t -- light-stall
100
90
8o
-_ 70LP
>-L)Z
,, 60.J
z
50
40
30
conditions.
20
NLR-1 RED FREQ
[] = 0.025© = 0.100L_ = 0.200
f
o
I " i • i • I •
.2 .4 .6 .8 1.0
x/c
Figure 26.- Phase angle, _t, of flow reversal on NLR-1 airfoil vs chord locationfor a range of frequencies at M= = 0.295, _ = 10 ° + 5° sin _t -- light-stallconditions.
54
_. tq
OF POOR QUALITY
,.,,,I¢,.)
>.ljz
l.u,-I
z
I,u
_czo.
70
60
5O
4O
3O
20 .....
I0
0
-10 -- •
0
NLR-7
0
RED FREQ
[] = 0.0100 '= 0.025A = 0.050+ = 0.100X = 0.1500 = 0.200
I ! t I
.2 .4 .6 .8 1.0
x/c
Figure 27.- Phase angle, _t, of flow reversal on NLR-7 airfoil vs chord location
for a range of frequencies at M_ = 0.295, e = 15 ° + 5 ° sin _t -- light-stall
conditions.
55
ORIGINAL PAG_ _
OF POOR QUALITY
5O
4O
_30
-I
>.¢j,zm 20.J
_3Z
u.I
z 10D.
-10
IAMES-01 | RED FREQm
I [] = o.oloI o = 0.025I _ = 0.05o
.... I + = o.'mo
.jsJ x = 0.150
+
o " .;, " ._ " ._, .e 1.ox/c
Figure 28.- Phase angle, _t, of flow reversal on Ames A-OI airfoil vs chord for a
range of frequencies at M_ = 0.295, _ = 15 ° + 10 ° sin _t -- deep-stall conditions.
56
5O
40
WORTMANN FX-098
RED FREQ
O " 0.0100 = 0.025A = 0.050+ = 0.100
+
.2 .4 .6 .8 1.0
x/c
Figure 29.- Phase angle, _t, of flow reversal on Wortmann W-98 airfoil vs chord for a
range of frequencies at M_ = 0.295, _ = 15 ° + i0 ° sin _t -- deep-stall conditions.
57
ORIGINAL P._G_ [<"_LJ
OF POOC_ _"" ' ",'.p
5O
4O
•_ 30
,--I
r,.,)
z,1, 20--I
z.<MJ
<•I- 10o.
-10
WORTMANN FX-098 RED FREQ
[] =0.05O© = 0.1_
= 0.15O
[]
i i i I
0 .2 .4 ,6 .8 1.0
x/c
Figure 30.- Phase angle, _t, of flow reversal on Wortmann FX-098 airfoil vs chord for
a range of frequencies at M= = 0.185, _ = 15 ° + i0 ° sin _t -- deep-stall conditions.
58
OHIL_,,AL PAGT- ,:_
OF POOR QUALITY
50
40
BOEING VERTOL VR-7
+
RED FREQ
[7 = 0.0250 = 0.050
= 0.100+ = 0.150
._ 3O
--I
).¢Jz
w 20
oz<uJ(n<:]: 10a.
-10 o _ " 4 ._ " ._ ,ox/c
Figure 31.- Phase angle, _t, of flow reversal on Vertol VR-7 airfoil vs chord for a
range of frequencies at M= = 0.295, ¢ = 15 ° + i0 ° sin _t -- deep-stall conditions.
59
I. Report No. NASA TM-84245 2. Go_rnment A¢cmsion No. 3. R_,ipkmt'$ Catalog No.
US AAVRADCOM TR-82-A-8
54. Title and Subtitle
AN EXPERIMENTAL STUDY OF DYNAMIC STALL ON ADVANCED
AIRFOIL SECTIONS
VOLUME 3. HOT-WIRE AND HOT-FILM MEASUREMENTS
7. Aurar(s)
L. W. Carr, W. J. McCroskey, K. W. McAlister,
S. L. Pucci, and O. Lambert*
9. Performing Organi_tion Name and _ NASA Ames Research Center,
Moffett Field, Calif. 94035, and U.S. Army Aero-
mechanics Laboratory (AVRADCOM), Ames Research
Center, Moffett Field, Calif. 94035
12. S_nsoring Agency Name and AddressNational Aeronautics and
Space Administration, Washington, D.C. 20546, and
U.S. Army Aviation R&D Command, St. Louis, MO 93166
Report _ta
December 1982
6. Performing Orglnizltion Code
8. PIwformiog OrgBnizetion Report No.
A-8938
1(). Work Unit No.
K-1585
11. Contract or Grant No.
13. Type of Report and PIriod Covered
Technical Memorandum
14. Sponsoring Agency Code
15. _p_ementary Notes
*Service Technique des Constructions A_ronautiques, Paris, France.
Point of Contact: L. W. Carr, Ames Research Center, MS 215-1, Moffett
Field_ Calif. 94035. (415) 965-5892 or FTS 448-5892.16. Abstract
Detailed unsteady boundary-layer measurements are presented for eight
airfoils oscillated in pitch through the dynamic-stall regime. The present
report (the third of three volumes) describes the techniques developed for
analysis and evaluation of the hot-film and hot-wire signals, offers some
interpretation of the results, and tabulates all the cases in which flowreversal has been recorded.
17. Key Wor_ (Suggutad by Author(s))
Dynamic stall Maximum lift
Oscillating airfoils Airfoil data
Boundary layer measurements
Unsteady pressure distributions
19. Security Oamif. (of this _rt)
Unclassified
18. Distribution Statement
_, Security C_mf. (of this _)
Unclassified
Unlimited
Subject Category - 02
21. No. of Pages 22. Price"
67 A04
"For sale by the National Technical Information Service. Springfield. Virginia 22161
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