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 ...................................

V

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3

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7

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PRECEDING PAGE BLANK NOT FILMED

lli

LESTOFFIGURES

i0

II

12

13

14

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 .........................

36

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42

43

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25

26

27

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

48

49

50

51

51

52

52

53

53

54

54

55

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 ............

7

7

8

9

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12

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25

Catalog of recorded data:

Catalog of recorded data:

Catalog of recorded data:

Vertol VR-7 airfoil .............

NLR-I airfoil ...............

NLR-7301 airfoil ............

30

32

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

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ORIGINAL PAGE ',S

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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

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_///////////////////////_ _1 \ _/////////////////////_

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Figure I.- Diagram showing installation of spar and alrfoil shell in tunnel.

36

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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 ,_

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HW2 mm

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72

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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

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_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

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FR - FLOW REVERSAL

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(b)

RFS AT x/c - 0.40

R

R_l _ _ RFSATx/c=0,60

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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).

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DIMENSIONS IN mm (ft)

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HEATER WIRE

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Figure 9.- Diagram of three-element, directionally sensitive hot-wire probe

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41

ORIGINAL PAGE 13OF POOR QUALITY

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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

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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

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= 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

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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"_'

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= 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_'(

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... 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/

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x/c

Figure 16.- Phase angle, _t, of flow reversal on Hughes HH-02 airfoil vs chord loca-

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48

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-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