AY g1 827 NATIONAL AKOONUATICS AND SPACE ADMNI4STRATIONA moorET--IETC FIG 20'AN EXPERIMENTAL STUDY OF DYNAMIC STALL 0ON ADVANCED AIRFOIL SECT--ET,JL02 V .1 MCCROSKEY. K v MCALISTER. L 9 CARR OL CTIT()
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NASA Technical Memorandum 84245 USAAVRADCOM TR-82-A-8
An Experimental Study of DynamicStall on Advanced Airfoil SectionsVolume 1. Summary of the Experiment
W. J. McCroskey, K. W. McAlister, L. W. Carr, andS. L. Pucci
July 1982 DTIC8 cD OCT4 19823
_____________________ United States AryE
Aviation Researhand Development
National Aeronautics and -- -"Command
Space Administration 82 1o 0 O 0 w• ._ ,:, . . ._. , . .,. ,,.9 .. Z ,,0 5 6. :
NASA Technical Memorandum 84245 USAAVRADCOM TR-82-A-8
An Experimental Study of DynamicStall on Advanced Airfoil SectionsVolume 1. Summary of the ExperimentW. J. McCroskeyK. W. McAlisterL. W. CarrS. L. Pucci. Aeromechanics Laboratory
AVRADCOM Research and Technology LaboratoriesAmes Research Center, Moffett Field, California
National Aeronautics and United States ArmySpace Administration Aviation Research and
Ames Remrch Cente Development Command
Moffett Field. California 94035 St. Louis, Missouri 63166
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TABLE OF CONTENTS
Page
LIST OF TABLES .. .. ....................... ........ v
LIST OF FIGURES. .. ....................... ....... vii
SYMBOLS .. ............ ......................... ix
SUMMARY. .. ....................... ............. 1
1. INTRODUCTION .. ........................ ....... 1
2. DESCRIPTION OF THE EXPERIMENT .. ............ ........... 2Test Apparatus. .. ....................... .... 2Instrumentation. ............. ................ 3Data Analysis and Measurement Accuracy .. ................. 4Test Conditions. .............. ............... 7
3. GUIDE TO THE DATA. .. ...................... ...... 8
4. RESULTS AND DISCUSSION. ............... .......... .... 10Static Data. .............. ................ 10Dynamic Data .. ............. ................ 12Comments on Wind-Tunnel Effects .. ................... 14
5. SUMMARY AND CONCLUSIONS. .. .......................... 14
REFERENCES. ............. ...................... 16
TABLES. .............. ........................ 19
FIGURES .. ............. ........................ 55
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LIST OF TABLES
Page
1 Harmonic C-fficients of the Oscillation Mechanism .. ........... ... 19
2 Airfoil Coordinates: NACA 0012 and Ames A-01 Airfoils .......... ... 20
3 Airfoil Coordinates: Wortmann FX-098 and Sikorsky SC-1095Airfoils ............. ............................... .... 21
4 Airfoil Coordinates: Hughes HH-02 (-5o Tab) and Vertol VR-7(-30 Tab) Airfoils .......... ........................... .... 22
5 Airfoil Coordinates: NLR-l and NLR-7301 Airfoils .. ........... ... 23
6 Transducer Locations on the Airfoils ......... ................. 24
7 Static Drag Coefficients at H. - 0.30 based on Wake Surveys .. ..... 25
8 Summary of the Measured Static Airfoil Characteristics atM. - 0.30, including Wind Tunnel Wall Corrections .. ........... ... 25
9 List of Test Points with Unusual Zero Drift of PressureTransducers ............ ............................... .... 26
10 Coefficients of Linear Curve-Fit of Static Lift Data,without Wind-Tunnel Corrections ....... .................... ... 27
11 List of Data Frames .......... .......................... ... 28
12 List of Static Data ........ .......................... ... 44
13 Mach Number Sweep at a = 150 + 10* sin wt, k - 0.10 ............ .. 45
14 Frequency Sweep at M. - 0.29, a - 15* + 10* sin wt ..... ........... 45
15 Frequency Sweep at Mw - 0.30, a - 100 + 10* sin wt ..... .......... 46
16 Frequency Sweep at M. = 0.30, a - 15* + 50 sin wt ..... ........... 46
17 Frequency Sweep at M. " 0.30, a - 10* + 5* sin wt ..... ........... 46
I, Stall Onset at M. - 0.30, a - a o + 10* sin wt, k - 0.10 .......... ... 47
19 Stall Suppression at M. - 0.30, a - ao + 10* sin wt ..... .......... 47
20 Stall Suppression at M, - 0.18, a -a o + 10' sin wt ..... .......... 47
21 Pitch Damping Studies at M. - 0.30, a ao + 2* sin wt .... ......... 48
22 No Separation: M. - 0.30, a - 50 + 50 sin wt ....... ............ 50
23 Dynamic Boundary-Layer Trip Data ................. 50
V
. . * . t e o e ~ o o e o ea'
24 Hiscellaneous Dynamic Data. .... ...................... 51
25 Test Cases for Numerical Analysis (ref. 1) .. ................. 54
vi
LIST OF FIGURES
Page
1 Airfoils tested in the experiment ...... ................... ... 55
2 Model installation in the test section ..... ................. .... 56
3 Photograph of the oscillation mechanism ..... ................ ... 57
4 Sketch of the wooden model shells surrounding the steel spar ... ..... 58
5 Pressure transducer and hot-wire installation: view from
inside the upper-surface shell ....... ..................... .... 59
6 Coordinate axes for the airfoils ...... .................... .... 59
7 Sketch of the shadowgraph system for visualizing the leading-edge region ........... .............................. .... 60
8 Representative shadowgraphs before (upper) and during (lower)dynamic stall: Sikorsky SC-l095 airfoil, Mw - 0.30,a - 10 + 100 sin wt, k = 0.10 ....... ..................... .... 61
9 Static lift and moment data on the NACA 0012 airfoil atM. = 0.3; shaded bands represent uncertainty limits of datacorrected for wind-tunnel-wall effects ..... ................. .... 62
10 Static lift and moment data on the Wortmann FX-098 airfoilat M. = 0.11 ........... ............................. .... 63
11 Static lift and moment data on the Vertol VR-7 airfoilat M. - 0.30 ........... ............................. .... 64
12 Comparison of measured lift-drag polars for the NACA 0012airfoil at M. - 0.30, including wind-tunnel-wall corrections ..... . 65
13 Comparison of lift-curve slopes on the NACA 0012 and SC-1095airfoils, Including wind-tunnel-wall corrections ... ............ ... 65
14 Typical data presentation from volume 2; no wall corrections ...... ... 66
15 Typical data presentation from volume 3 ..... ............... ... 67
16 Static characteristics of the NACA 0012 airfoil atM. - 0.30, including wind-tunnel-wall corrections .. ........... ... 67
17 Static characteristics of the Ames A-0l airfoil atM - 0.30, including wind-tunnel-wall corrections .. ........... ... 69
18 Static characteristics of the Wortmann FX-098 airfoil atMm - 0.30, including wind-tunnel-wall corrections .. ........... ... 71
vii
Page
19 Static characteristics of the Sikorsky SC-1095 airfoil atK. - 0.30, including wind-tunnel-wall corrections .. ........... ... 73
20 Static characteristics of the Hughes HH-02 airfo!, atM. - 0.30, including wind-tunnel-wall correctior's .. ........... ... 75
21 Static characteristics of the Vertol VR-7 airfoil atM, = 0.30, including wind-tunnel-wall corrections ........... 77
22 Static characteristics of the NLR-l airfoil atM. = 0.30, including wind-tunnel-wall zorrections ........... 79
23 Static characteristics of the NLR-7301 airfoil atM. = 0.30, including wind-tunnel-wall corrections .. .......... .. 81
24 Comparison of maximum static lift on the NACA 0012 airfoil ....... ... 83
25 Comparison of maximum static lift on the Ames A-01 airfoil ... ....... 83
26 Comparison of maximum static lift on the Wortmann IX-098airfoil ............... ................................ 84
27 Comparison of maximum static lift on the Sikorsky SC-1095
airfoil ............. ............................... ..... 84
28 Comparison of maximum static lift on the Hughes HH-02 airfoil ..... 85
29 Comparison of maximum static lift on the Vertol VR-7 airfoil ... ...... 85
30 Comparison of maximum static lift on the NLR-l airfoil .... ......... 86
31 Comparison of maximum static lift on the NLR-7301 airfoil. . . . . . .. 86
32 Maximum unsteady lift on the eight airfoils:solid symbols - stall onset; open symbols = deep stall ........... ... 87
33 Comparison of maximum lift on the eight airfoils at M. - 0.30 .. ..... 88
34 Comparison of maximum lift on the NACA 0012 airfoil under deep-dynamic-stall conditions: a - 150 + 10* sin wt, k - 0.10 ....... . 89
35 Comparison of the lift hysteresis on the NACA 0012 airfoil:M. 9 0.1, a - 15* + 10* sin wt, k - 0.10 ................ ........ 89
36 Comparison of maximum airloads on the NACA 0012 airfoil atM0, - 0.30 and cik 2 2 constant ....... .................... ... 90
37 Comparison of maximum airloads on the Sikorsky SC-1095 airfoilat M. - 0.30 and auk2 a constant ...... .............. . . . 91
38 Comparison of maximum airloads on the NLR-l airfoil atM(- 0.3 and 0max - 20. ....................... 92
viii
- - -
SYMBOLS
A static lift coefficient at u 0 (see table 10)
B static CL, Ifii. (see table 10)
CC chord force coefficient
CD form drag coefficient derived from surface pressure measurements
CDW total drag coefficient derived from wake survey (see table 7)
CL lift coefficient
CL0 lift:-curve slope at low a, per deg
CM quarter-chord pitching moment coefficient
CMO static pitching-moment coefficient at zero angle of attack
CN normal force coefficient
C p pressure coefficient
c airfoil chord, m
k reduced frequency, wc/2U,,.
L/D ratio of lift to drag
M. free-stream Mach number (also M in table 11 and fig. 14) IMmax maximum local Mach number on the airfoil
q. free-stream dynamic pressure, N/m2 (also Q, psi, in table 11)
Re Reynolds number based on chord and free-stream conditions
ro leading-edge radius, m
t time, sec
U. free-stream velocity, m/sec
Xa.c. chordwise location of the aerodynamic center of pressure at zero lift
x chordwise coordinate, m (see fig. 6)
y normal coordinate, m (see fig. 6)
a angle of attack, deg
Cimn angle of attack for maximum negative chordwise force, deg
ix
aLmax angle of attack for maximum lift, deg
aMmax angle of attack for maximum local Mach number, deg
ao mean angle, deg (also AO in computer printouts); also angle for zero lift intable 8 and figs. 9-11
ass static-stall angle, corresponding to CLmax, deg
aI amplitude, deg (also Al in table 11 and fig 14)
a2 magnitude of second harmonic of a, deg
aerodynamic pitch damping coefficient, - 4 CM da
0 2 phase of second harmonic component of a, deg
w circular frequency, rad/sec
xN
AN EXPERIMENTAL STUDY OF DYNAMIC STALL ON ADVANCED AIRFOIL SECTIONS
VOLUME 1. SUMMARY OF THE EXPERIMENT
W. J. McCroskey, K. W. McAlister, L. W. Carr, and S. L. Pucci
U.S. Army Aeromechanics Laboratory (AVRADCOM), Ames Research Center
SUMMARY
-,The static and dynamic characteristics of seven helicopter sections and a fixed-
wing supercritical airfoil were investigated over a wide range of nominally two-dimensional flow conditions, at Mach numbers up to 0.30 and Reynolds numbers up to4x106. Details of the experiment, estimates of measurement accuracy, and test condi-tions are described in this volume, (the first of three volumes). Representativeresults are also presented and comparisons are made with data from other sources.The complete results for pressure distributions, forces, pitching moments, andboundary-layer separation and reattachment characteristics are available in graphicalform in volumes 2 and 3.
The results of the experiment show important differences between airfoils, whichwould otherwise tend to be masked by differences in wind tunnels, particularly insteady cases. All of the airfoils tested provide significant advantages over theconventional NACA 0012 profile. In general, however, the parameters of the unsteadymotion appear to be more important than airfoil shape in determining the dynamic-stall airloads.
1. INTRODUCTION
Retreating-blade stall limits the high-speed performance of most modern heli-copters. In the past decade, numerous new airfoils have been designed in attempts toimprove the stall characteristics of rotors without compromising the advancing-bladeperformance. Only a few of these have been tested under unsteady conditions, andsome have not been tested at all. Furthermore, there is almost no overlap betweenthe existing data sets with regard to the important parameters of oscillatory motion.
The motivation of the present experimental investigation was the obvious need fora standard data base for a series of modern rotor-blade sections. The primary objec-tive was to measure the unsteady airloads, over an extensive matrix of test conditions,on the eight profiles shown in figure 1. Other investigations were also overlapped asmuch as possible. The NACA 0012 served primarily as a standard reference section; thesix modern helicopter sections were chosen as representative of contemporary designsfrom several different companies and research organizations. A modern fixed-wingsupercritical profile was also included to extend the range of leading-edge geometriesand to provide a basis for comparison with oscillating-airfoil results obtained inother wind tunnels.
Secondary objectives were to investigate the type of stall and boundary-layerseparation characteristics for each profile, to provide guidelines for estimating thedynamic-stall characteristics of new airfoils in the future, to supplement the con-ventional lift and pitching-moment measurements with unsteady drag data and
stall-flutter boundaries, and to determine the effeLts of leading-edge roughnessthat is comparable to the erosion of blades in service or in incipient icingconditions.
Dynamic stall depends on a large number of parameters. Consequently, a verylarge number of unsteady test points (more than 600) plus 44 sets of static data wererequired to fulfill the objectives of this investigation. As a result, the completereport consists of three volumes. The present volume summarizes the experiment andsome of the principal results, including comparisons with data from other sources.It also contains a comprehensive index of the individual unsteady data points. Vol-ume 2 (Pressure and Force Data) contains the pressure, force, and moment data ingraphical form. These data are also available upon request on digital computertapes, one tape for each airfoil, as explained in volume 2. In addition, there isa single tape containing only the 10 test cases that were discussed in reference 1for the NACA 0012, Vertol VR-7, and NLR-7301 airfoils. Boundary-layer transition,flow reversal, and reattachment results appear in volume 3 (Hot-Wire and Hot-FilmMeasurements).
This report is primarily intended to assist ie users of the data; therefore,the results are not discussed at length. Th pt ival results have been publishedin references 1 and 2.
2. DESCRIPTION OF THE - RIMENT
Test Apparatus
The experiment was performed in the 2- by 3-m atmospheric-pressure, solid-wallWind Tunnel at the U.S. Army Aeromechanics Laboratory. The tests were conducted inessentially the same manner as those in a previous experiment (refs. 3,4), exceptthat the free-stream Mach number was extended to 0.3, the model chord c was reducedto 0.61 m (except for the Hughes HH-02 airfoil, c = 0.69 m), the frequency of oscil-lation was extended to 11 Hz, and the data processing was refined considerably. Themodels spanned the 2.13-m vertical dimension of the wind tunnel, as indicated infigure 2, and were oscillated sinusoidally in pitch about the quarter chord. A gapof approximately 2 mm existed between the ends of the model and the wind-tunnel walls.
The drive mechanism used (fig. 3) was the same one described in references 3and 4, with some notable improvements. In some cases, the connecting push rod wasfitted with a remotely controlled jackscrew mechanism that allowed the mean angle,ao , to be varied continuously while the tunnel was operating. Discrete amplitudes ofoscillation of 20, 50, 60, 80, 10, or 140 could be set between runs. The motion ofthe airfoils was given by a = ao + ai sin wt, with maximum higher harmonic distor-tion approximately 2% of ai. Table 1 gives the harmonic content of the mechanismfor various values of ao and aI. The frequency of oscillation could be variedbetween approximately 0.02 and 12 Hz.
The models of the eight airfoils (fig. 1) consisted of interchangeable shellsconstructed of wood and fiberglass. These shells surrounded a stainless steel sparthat contained the instrumentation and wiring, as indicated schematically in fig-ures 2 and 4. The shells contained special fittings for the pressure transducersand hot-wire or hot-film sensors (fig. 5) that facilitated model changes without dis-connecting the instrumentation.
2
Each set of shells was precision-machined, while mounted on the spar, to adesign accuracy of ±0.1 mm. However, measurements after the test revealed that therms standard deviation of the coordinates from the design values was about 0.4 mm,or 0.06% of chord, and that the maximum error was about 0.8 mm. The nominal designcoordinates of the airfoils are given in tables 2-5, referred to the standard coordi-nate system sketched in figure 6. The coordinates were taken originally from refer-ences 5-9 and from Amer (K. Amer, private communication, 1977).
A limited amount of static and dynamic data were obtained on each airfoil atM = 0.185 and 0.29 with a boundary-layer trip, consisting of a 3-mm-wide band of0.10-mm-diam glass spheres glued to the leading edge. The purpose of the trip was toeliminate the laminar separation bubble that would normally form near the leadingedge as the stall angle was approached. It also approximately simulated surfaceabrasion on helicopter blades operating under severe field conditions, as well asroughness caused by incipient icing conditions.
Instrumentation
The primary data were obtained from 26 Kulite differential pressure transducers,types YCQH-250-1 and YCQL-093-15. Those of the latter type were used in the leading-and trailing-edge regions, because of their smaller size. The locations of the trans-ducers for each airfoil are given in table 6. The back side of each transducer wasreferenced to the total pressure of the wind tunnel; total pressure was measuredabout 1.5 m upstream of the model. The measuring side of the transducers mated withthe fittings shown in figure 5, which had 0.79-mm-diam orifices. The transducersthus installed had flat amplitude versus frequency responses of 250 Hz or better andtypical cavity resonance frequencies of about 850 Hz.
Special on-line analog computers that calculated and displayed the instantaneousnormal force, pitching moment, pitch damping, and pressure distributions proved to beextremely valuable in assessing the dynamic-stall behavior, as well as the perfor-mance of the instrumentation, while the tests w-e in progress. These devices alsoenabled the unsteady parameters to be adjusted until some desired result was obtained,such as the maximum lift condicion in the absence of moment stall or neutral aero-dynamic damping in pitch.
Boundary-layer transition, flow reversal, separation, and reattachment werestudied with a variety of surface hot films and hot-wire sensors (single-, double-,and triple-element probes), using the techniques described in references 4, 10, and11. Six sensors were used on the upper surface of each airfoil, at the locationsgiven in table 6. In addition, a hot-wire probe protruding just outside the boundarylayer was mounted near the leading edge of the NLR-1 profile to aid in diagnosing thelocal supersonic zone that wis frequently inferred at high incidence.
The leading-edge region was also examined with a shadowgraph flow visualizationsystem (fig. 7). The high-intensity strobe light was fired at selected phase anglesduring the oscillation, and the pattern that developed on the Scotchlite high-gainreflective sheeting on the floor of the tunnel was photographed by the pulse cameraabove the test section. A representatire photograph is shown in figure 8.
Finally, a traversing pitot-static probe was used to survey the wake behind eachairfoil under steady-flow conditions. The steady drag of the airfoils at M. = 0.30was derived from these measurements; these drag coefficients are listed in table 7.
3
Data Analysis and Measurement Accuracy
For quantitative purposes, the pressure transducer and hot-wire signals were
amplified and recorded on a 32-channel analog tape recorder with 2500-Hz flat fre-
quency response. In addition, the average free-stream dynamic pressure, the instan-taneous angle of attack of the model, and l/cycle and 200/cycle timing indicatorswere recorded simultaneously. Calibrations of the pressure transducers were recordedat the beginning and end of each analog tape. The unsteady data tapes were digitizedand ensemble-averaged off line. At least 50 cycles of data were normally sampled200 times per cycle; however, for the NACA 0012 airfoil at very low frequencies, thatis, k < 0.002, only about 10 cycles were recorded. Reference and calibration signalsand the steady pressure data were acquired with the same system and were digitallysampled 100 times over a 5-sec interval. The averaged pressure data were then pro-cessed and integrated numerically by trapezoidal rule to determine the unsteady lift,moment, and pressure drag.
End-to-end checks of the data acquisition and processing system indicated thatthe pressure signals were reproduced to within an rms error of approximately 70 N/m
2
(0.01 psi), and that the transducer calibrations were reliable to better thant150 N/m2 (0.02 psi) or ±3% of the reading, whichever was greater, over the range of
tunnel speeds and temperatures. The model temperature, measured inside the shells,
was closely monitored and not allowed to vary more than 3C between records ofno-flow pressure readings. Transducer zero drift was normally controlled to withinthe greater value of either ±150 N/m 2 (0.02 psi) or ±5% of free-stream dynamic pres-sure. However, some exceptions are noted later in this suction.
The hot-wire and hot-film signals were recorded as consecutive, separate dataframes, and individual cycles of the analog records were examined to determine the
boundary-layer characteristics, as discussed in references 4, 10, and 11. For thesedata, the results from three to eight cycles were averaged to obtain the relative
times within the cycle, wt, at which the various boundary-layer events occurred.
The instantaneous angle of attack was measured ,ith a potentiometer attached tothe tubular portion of the model spar (fig. 3). The angle-of-attack signal was cali-
brated for each data point based on the value of al, which was set by the oscilla-tion linkage, and physical measurements of amax and amin that were obtained fromthe trailing-edge position relative to the centerline of the tunnel with the wind off.The maximum absolute error in a was estimated to be ±0.20, with a relative uncer-tainty of ±0.05 ° over the cycle. The maximum torsional deflection of the model atthe centerline was calculated to be ±0.3 ° . Table 1 gives the amplitude and phase ofthe second harmonic component of a for various nominal values of al. The frequencyof the oscillation was maintained and measured to an estimated accuracy of ±0.03 Hz.
The tunnel dynamic pressure was measured with a conventional pitot-static probemounted approximately 1.5 m upstream of the model and connected to a pressure trans-ducer and amplifier system with a net accuracy of approximately ±14 N/m 2 (0.002 psi)
under steady conditions. The measured values ranged from 90 N/m2 (0.013 psi) atM. - 0.04 to 6200 N/m 2 (0.90 psi) at M,,. - 0.3. The output of this transducer wasrecorded by hand and on the 32-channel analog tape recorder. An average of these twovalues, which rarely differed by more than 2%, was used to compute q., except in afew cases in the early stages of the test program in which the tape-recorded valuewas obviously in error and was therefore ignored. The 25-mm-thick ground plane shown
in figure 2 caused a 1% reduction in tunnel cross-sectional area between the pitot-
static tube and the model; this was ignored except as noted in connection with thesteady lift results presented in section 4 under the heading Static Data.
4
A detailed examination of the digitized data revealed that the 200/cycle sam-pling of the analog signals was not always synchronized perfectly with the 200/cycletiming indicators. That is, the effective time base of the digitized data was inerror, the cumulative effect of which was either to leave a small gap in the data atthe end of the cycle or to overlap the 200th sample of a given cycle with the firstsample of the next cycle. Consequently, a corrected time base for the digital dataarrays was obtaint by least-squares curve-fitting a first- and second-harmonic sinewave to the angle-of-attack signal, a. All of the pressure data were then linearlyinterpolated onto the new time base at 200 even intervals per cycle and stored innew arrays, with the first data point in each array corresponding to wt - 0. Theend result is that the final data appear at the desired times, but suffer an effec-tive "smearing" that would be, at worst, equivalent to sampling at a rate of100 points per cycle instead of 200 per cycle.
Experimental uncertainty of the airloads- For the purposes of comparing thestatic and dynamic-stall characteristics of the eight airfoil sections, the absoluteaccuracy of the measurements and the consequences of wind-tunnel blockage, circula-tion interference, and sidewall boundary-layer interference are less important thanthe random experimental errors outlined above. However, an attempt was made toassess all of these, as described below.
The total measurement uncertainty in the pressure, force, and moment coeffi-cients depends on the operating conditions. For example, the probable error in Cpbased on the instrumentation characteristics quoted above varies from less than±0.07 at M. - 0.3 and a = 0 to about ±0.4 near the leading edge at M. = 0.11and a approaching the stall angle. For most of the static data at M. - 0.3, themeasurement uncertainty is estimated at ±0.03 for CLmax, ±0.005 for CM, and
±0.0005 for CD derived from the wake measurements. However, the uncertainty inthe SC-1095 lift and moment data is thought to be at least twice as large, becauseof some unresolved difficulties with the pressure measurements. These valuesincrease with decreasing Mach number, rising by a factor of about 5 in the extremecase M. = 0.035, where the pressure signals were very small.
Some representative examples of static CL and CM versus a are given infigures 9-11, and the primary characteristics of each airfoil at M. = 0.30 are pre-sented in table 8. The symbols in the figures indicate the individual uncorrecteddata points, as presented in volume 2 of this report; the shaded bands denote theestimated bounds of the airfoil characteristics. The bounds of the airfoil charac-teristic include static wind-tunnel-wall corrections according to Allen and Vincenti(ref. 12) and a 1% correction due to the reduction in test-section area at the modelcaused by the steel plate on the floor of the tunnel (fig. 2). (This wall correctionmethod is only valid below stall, where the corrections are about 1% for a and 1.5%for CL.) These boundaries were derived based on the measurement uncertaintiesdescribed above, on data that were obtained with the on-line analog computers, andon the dynamic data obtained at k S 0.01. It should be noted that the scatter inthe data and the uncertainty bounds increase considerably for conditions above thestall angle. The last line in table 8 indicates the experimental uncertainties forthe various quantities listed. The static data are discussed further in section 4.
A novel feature of the present experiment was the determination of unsteadypressure drag, CD - CC cos a + CN sin a, where CC and CN are the chordwise andnormal force coefficients derived from the upper and lower surface-pressure distri-butions. The two terms in this expression for CD are approximately equal and oppo-site at high angles of attack below stall, so that the probable percentage errors of
.... ' , .. . ',L. _ HI in I I . ll ll5
CD are much greater than for CC, CN, CL, or CM. Figure 12 shows a typical staticlift-drag polar based on pressure measurements and on the more accurate wake surveyof the total drag (table 7). The measured pressure drag, which neglects the contri-bution due to skin friction, is less than the total drag at low lift coefficients,but it incorrectly exceeds the wake measurements by as much as 0.02 near the stallangle, that is, by as much as 100%. (It may be noted that Woodward (ref. 13)reported similar, unexplained discrepancies between measured pressure drag and CDbased on wake surveys.) However, the percentage errors are much less in the stallregime, where the magnitude of CC decreases considerably and the maximum drag coef-ficient becomes of the order of CL tan a (i.e., of the order of unity) for the deep-dynamic-stall cases studied.
The measurement uncertainty of the unsteady data is probably comparable to thatof the static data, but fewer independent checks were available to assess the randomexperimental errors and the wind-tunnel interference, especially in the post-stallregime. Fromme and Golberg (ref. 14) have indicated that unsteady wall correctionscan be greater than the corresponding static corrections, but it is not clear to whatextent their potential flow analysis can be applied to the present measurements.Likewise, it is not possible to estimate reliably the post-stall tunnel sidewalleffects nor how these vary from one airfoil to another, but tuft flow visualizationand experience suggested that these problems became less important as the frequencyof oscillation is increased. It is the authors' judgment that for M. 0.2, theunsteady data in the deep-dynamic-stall regime should be in error by no more than±0.2 for CL, ±0.05 for CM, and ±0.10 for CD, except as noted in the next section.The results are thought to be about twice this accurate below stall and in lightstall, whereas the accuracy was seriously degraded for M. < 0.1 because of the smallvalues of the pressure signals.
Special cases of questionable accuracy- Despite efforts to monitor the perfor-mance of the pressure instrumentation during the test and to control and minimize themeasurement uncertainties, various problems sometimes arose that only became evidentduring the post-test reduction and analysis of the data. In most cases, it was pos-sible to correct these problems on an individual basis, using redundant informationor by interpolating in time or space between neighboring values, without signifi-cantly compromising the accuracy of the results. In other instances, the measure-ments appeared to be qualitatively correct, but the experimental uncertainty waslikely to have been outside the normal bounds discussed in the previous section.These cases are identified below by data-point or "frame" number.
Frame 10202 for the NACA 0012 airfoil had an unusually large number of randomirregularities, a total of 44 in the 5,200 pressure data samples. These were elimi-nated by linearly interpolating between data at preceding and succeeding time incre-ments. Because some of these irregularities occurred during rapid fluctuations ofthe flow, the time-histories of part of the pressure data for this particular framemay have been degraded. However, the effect on the integrated force and moment coef-ficients was probably small.
Table 9 lists the frames for which the "zero" drift of one or more of the trans-ducers appeared to have exceeded by a significant amount the nominal values quoted inthe previous section. Also included are the low Mach-number cases for which theno-flow pressure readings taken before and after recording data varied by more than50% of free-stream dynamic pressure, even though this drift amounted to less than thenominal measurement uncertainty of 150 N/m 2 (0.02 psi). It should be mentioned thatin all cases the differences between these pretest and post-test zeros were linearlyinterpolated with respect to elapsed time to obtain effective zeros for the individual
data frames. In principle, this should have reduced the effects of the transducerdrift; however, the actual improvement in the measurement accuracy because of thistechnique remains unknown.
For the Hughes HH-02 airfoil, the responses of pressure transducers No. 1 (lead-ing edge) and No. 25 (x/c - 0.0081, lower surface) were rather sluggish, possiblybecause the orifices were partially clogged. Therefore, the unsteady data from thesetwo transducers are suspect. In calculating the force and moment data for this air-foil, transducer No. 25 was ignored and the pressure integrals
CN - P p dx/c etc.
were replaced by
CN = -2 fCpE dE etc.
where = x7, thereby eliminating the influence of transducer No. 1, sinceCp 1 A=o 0. Another problem with the HH-02 force and moment data is that the
trailing-edge transducers were at x/c = 0.925 instead of 0.98, so that the errorin extrapolating to x/c = 1.0 is greater for this airfoil. The net effect of thesemodifications is difficult to assess, but it probably increased the experimentaluncertainties for the lift, pressure drag, and pitching moment data by no more than50%.
The NLR-7301 airfoil had a large amount of concave curvature on the lower sur-face downstream of x/c = 0.5, which produced larger pressure gradients there thanexisted on the other airfoils. Therefore, the relatively sparse distribution ofpressure transducers in that region may have led to larger errors in determining theforces and moments than the nominal values quoted in the preceding section.
The reduced data for the Sikorsky SC-1095 airfoil under static conditions andat low frequencies consistently exhibited values of maximum lift coefficient andlift-curve slope that appeared to be about 5% too large, based on comparisons withthe other airfoils and with the results obtained from the special on-line analog com-puter described above under Instrumentation. In particular, the comparison with thepresent NACA 0012 data (fig. 13) contrasts significantly with the steady results ofNoonam and Bingham (ref. 15) and Jepson (ref. 16), who found CL, to be approxi-mately the same for both airfoils. A detailed examination of the present data andthe transducer calibrations revealed somewhat erratic performance in a few cases, but
no systematic behavior emerged that could explain the apparent problem. Therefore,the conclusion is that the SC-1095 results should be viewed with caution, even thoughthey appear to be qualitatively correct. :1
Test Conditions
The primary reference conditions for the initial comparisons of the various air-foils were static and deep-dynamic stall at M,, = 0.3, with the nominal unsteadymotion given by a - 100 + 100 sin wt and k - wc/2U. - 0.10. Limited but system-atic variations in Mach number and the unsteady parameters were explored for all air-foils as indicated below and in section 3, where the specific test points are indexedand cross-referenced.
7
Static data- Pressure measurements were recorded at discrete values of abetween -5* and 200 for M, - 0.11, 0.185, 0.25, and 0.30 for all airfoils except theNACA 0012. In the latter case, static data were recorded only at M. - 0.30; quasi-steady data were obtained for a continuous range of a = ao + 100 sin wt fork z 0.001 for nine values of M. between 0.035 and 0.30. A number of the staticconditions were repeated with a boundary-layer trip at the leading edge. Wake sur-veys for static drag were obtained at M. -c-0.3 for a between -50 and the staticstall angle.
Unsteady data- The parameters that were varied under dynamic-stall conditionswere Mach number, reduced frequency, mean angle, and amplitude of the oscillation.The effect of Mach number was studied between M. = 0.035 and 0.30, primarily in thedeep-stall regime for a = 158 + 100 sin wt and k = 0.10. In these cases, theReynolds number also varied, proportional to Mach number, according to the relationRe = 14x10s M6.
The principal ranges of reduced frequency, mean angle, and amplitude were0.01 : k S 0.20, ao - 10* and 15, and a, = 2, 50, and 10', respectively; theeffects of these parameters were studied primarily ac M. = 0.30. Additional varia-tions in k and ao were effected to achieve specific dynamic effects, such as nostall, stall onset, stall suppression because of unsteady effects, and neutral aero-dynamic damping in pitch.
Finally, additional test points were selected that duplicated some of the condi-tions of references 3 and 17-19 as closely as possible. A complete list of theunsteady test conditions and descriptions of the parametric variations are given inthe following section.
3. GUIDE TO THE DATA
A very large data base was generated in this investigation. As mentioned in theIntroduction, summary graphs of the pressure, force, and moment coefficients andselected results from the boundary-layer studies are contained in separate volumes.The airloads data are also stored on digital computer tapes, one for each airfoil, asexplained in volume 2. This section describes briefly the data presentations to befound in the subsequent volumes and indicates by test point, or "frame number," thevarious types of data that are available.
Figure 14 illustrates the format of volume 2 for the unsteady pressure, force,and moment coefficient data, that is, CL, CM, and CD versus a and wt, and theupper-surface pressure distributions throughout the cycle. Additional informationis listed at the top if the graphs. Following the airfoil name is the identificationnumber for each test point. As explained in volume 2, these frame numbers comprisedata at a single angle of attack for the steady data, and data at 200 evenly spacedtime intervals throughout the cycle for the unsteady cases. The quantities AO andAl are the mean value and the first-harmonic amplitude, respectively, of theinstantaneous angle of attack, a; Mmax is the estimated maximum value of the localMach number at any time in the cycle, calculated from the classical gas-dynamic equa-tions for steady isen:ropic flow and the measured pressure coefficient, -Cpmn
(cf. ref. 2); aLmax, %min, and aMmax are the angles of attack corresponding to
maximum lift, minimum chord force (cf. ref. 3), and Mmax, respectively; and c is
8
the aerodynamic damping in pitch. The asterisk on the ordinate of the pressure-coefficient graph represents sonic conditions.
The dotted line in the CL vs a curve in figure 14 is an approximation to the
quasi-static lift behavior for this flow condition, according to the relation
BaCL -A +
where a is in degrees and A and B were obtained from the relevant steady and verylow-frequency data, that is, for k 1 0.01. The values of A and B are given intable 10. Finally, it should be mentioned that in contrast to the data in table 8and the static results presented in section 4 under the heading Static Data, wind-tunnel wall corrections have not been applied to A and B, to the data in volume 2,nor to the numerical data tapes.
Figure 15 shows two representative examples of the boundary-layer "flowreversal" information contained in volume 3. The abscissa in the figures show theposition on the airfoil where the surface instrumentation first indicated a break-down of the attached boundary-layer flow at the beginning of dynamic stall, asexplained and discussed in volume 3 and in references 4, 10, and 11. This eventeither signifies or is closely associated with the separation that accompanies thebeginning stages of dynamic stall. The ordinate indicates the nondimensional timein the cycle, wt, at which this event occurred.
Tables 11-24 provide a comprehensive summary and index of the entire experi-mental program. Table 11 lists the frame numbers of all the pressure data, in thesequence in which they appear on the data tapes. The airfoil and pertinent test con-ditions are also listed, and the conditions for which boundary-layer data wererecorded are indicated in the last column. The letter "Y" in the "TRIP" column indi-cates the use of the boundary-layer trip; "N" denotes the standard smooth condition.The notations "ST" and "US" denote steady and unsteady data, respectively, and thefrequency of oscillation in Hertz is given in the column labeled "FREQ."
Table 12 is an index of the steady-data sets, arranged by airfoil and Mach num-ber. The use of a boundary-layer trip is indicated by the letter "T." The notation"Quasi-steady" indicates the data that were acquired on the NACA 0012 airfoil asunsteady data, but at very low frequency, k 5 0.002.
A cross-reference index that groups the unsteady data by types for each of theeight airfoils is given in tables 13-24. There are some duplicate entries in thesetables, in order to facilitate the identification of data sets with variations in theindividual parameters of the unsteady motion. There are also blank entries, sincenot all conditions were recorded for all airfoils. The principal types of unsteadyconditions are outlined below.
Variations in Mach number- Table 13 lists the test points concerned with theeffect of Mach number on deep dynamic stall, for a - 150 + 100 sin wt and k - 0.10.Although the NLR-7301 airfoil was only tested at three values of M. with ao W 150,it was also tested with ao - 10° at M. - 0.11, 0.18, 0.22, and 0.30; these framesare given in table 24. Stall-suppression conditions, tables 19 and 20, and theeffects of leading-edge trips, table 23, were studied at M. - 0.18 and 0.30 forvarious values of mo and k. As stated in section 2 under Test Conditions, the vari-ation of Reynolds number with Mach number was Re a 14x105 M.
9
it
Reduced frequency sweeps- The test points concerned with the effect of frequencyon dynamic stall are given in tables 14-17. These data cover the range0.01 5 k S 0.20 at M. = 0.3, with mean angles of 100 and 150 and amplitudes of 5*and 10*. In addition, the t,\CA 0012 airfoil was tested over an extensive range ofother values of ao (table 24).
Stall onset- This condition, defined in references 1 and 2 as obtaining themaximum possible lift without moment stall occurring at any time throughout the cycleof oscillation, was studied at M,. = 0.30, k - 0.10, a, - 100 , and variable meanangle, as indicated in table 18.
Stall suppression caused by unsteady effects- With a. fixed at 10, ao wasvaried so that 0amax was slightly greater than the static-stall angle. Data werethen recorded (tables 19 and 20) at various reduced frequencies to study whetherstall would diminish or increase with increasing k.
Pitch damping boundaries- Stall conditions relevant to small-amplitude flutterboundaries are listed in table 21, at a. = 2° and M.. - 0.30. Mean angle andreduced frequency were varied to obtain approximate boundaries of neutral aerodynamicdamping in pitch and to obtain the maximum negative value of pitch damping, -Cmin"However, no data of this type were recorded for the NACA 0012 airfoil.
No separation- A limited number of test points were recorded at M. - 0.30 anda = 5 ° + 50 sin wt, as indicated in table 22. Some additional conditions for the
NLR-l and NLR-7301 profiles without separation are given in table 24.
Boundary-layertrLip- Data with the leading-edge trip were obtained staticallyfor a between 0° and 20* and dynamically for a - 15* + 100 sin wt at two valuesof Mach number, 0.18 and 0.30. The values of k for the dynamic data are given intable 23; the static data with trip are so indicated in table 12. An exception wasthe NLR-7301 section at M. - 0.30, for which a - 10* + 50 sin ut (table 24). Inaddition, the NLR-l section with trip was studied with ao - 2.5' (table 24).
Miscellaneous- These test points are included in table 24. In addition to thecases mentioned above, the unsteady test conditions of references 3 and 17 for theNACA 0012, of reference 18 for the Sikorsky SC-1095, and of reference 19 for theNLR-l airfoil were reproduced insofar as possible. Also, for the Vertol VR-7 air-foil, k was varied from 0.01 to 0.25 at MO - 0.18 with ao - 100 and 15* anda1 = 100. Finally, dynamic stall on the NLR-l profile at negative incidence wasstudied at M. = 0.30 for a - -2* + 100 sin wt and 0.01 5 k 5 0.10.
Selected test cases- Finally, table 25 lists the unsteady data that were pro-posed in reference 1 as specific test cases for evaluating unsteady viscous flowtheories and computational methods. These data were obtained on the NACA 0012,Vertol VR-7, and NLR-7301 airfoils. They include conditions of no-stall, stall-onset,light-stall, and deep-dynamic-stall, all at M. - 0.3.
4. RESULTS AND DISCUSSION
Static Data
The measurements performed under steady or quasi-static flow conditions providea frame of reference for the dynamic-stall results and a basis for comparison with
10
... .. . . .. . .. ... .7 -
data from other wind tunnels. Some of the highlights of the static data are pre-sented below, with particular reference to the force and moment coefficients atM. - 0.3. With the exception of the drag data listed in table 7, wind-tunnel-wallcorrections have been applied to all of the static results presented in this section,using the formulae of reference 12.
As noted earlier, table 8 gives a summary of the primary static characteristicsof each airfoil at M. - 0.30, and figures 16-23 show the basic variations of lift,pitching moment, and drag coefficients for the eight sections. The dashed lines inthe "a" parts of figures 17-23 represent curve-fits of the lift data in the linearCL - a regime. The drag data derived from the wake surveys are listed in table 7.In the following discussions, some comparisons are made for each airfoil between thepresent measurements and data obtained elsewhere.
NACA 0012 airfoil- This profile has been tested by many investigators, with swide range of results. Figure 24 shows the variation in CLmax with Mach number,
including results reported or sumarized in references 3, 5, 15-17, and 20-24 over awide range of Reynolds numbers. The present values of C1 nax increase with increas-
ing Mach number for M. < 0.22, probably because of the effects of increasingReynolds number, whereas compressibility effects are thought to be responsible forthe decrease in CLmax for H. > 0.22. The boundary-layer trip was found to be
relatively unimportant for this airfoil at the Mach and Reynolds numbers of the test.
The present CLmax data tend to lie near the upper range of the values from
other sources. The same is true for the lift-curve slopes in the linear regime,
CL., which is not shown.
Ames A-01 airfoil- Figure 25 compares the data from the present I.est with mea-surements made in a transonic wind tunnel at somewhat lower Reynolds numbers (ref. 6)for the A-01 airfoil. Although the lift-curve slopes for CL < 1.0 were not sig-nificantly different in the two tests, the airfoil stalled at lower angles of attackin the transonic tunnel. Consequently, lower values of maximum lift coefficient weremeasured and reported in reference 6 at M. - 0.2 and 0.3, which was near the loweroperating limit of that facility.
Wortmann FX-098 airfoil- Maximum-lift data from several investigations (refs. 8,24-26) are compared with the present data in figure 26 for the FX-098 airfoil. Allof the data agree reasonably well over the Mach-number range of the present test.However, there are marked differences at higher Mach numbers.
Sikorsky SC-1095 airfoil- Steady results for this section are shown in figure 27,where the comparison is generally unfavorable. The suspicious nature of the presentlift data was mentioned earlier in section 2 under Data Analysis and MeasurementAccuracy; here the open circles indicate the present data analyzed in the normal wayand the solid symbols represent what are thought to be the true values. The latter,somewhat lower, values are based primarily on the on-line measurements. It shouldbe mentioned that the data of Noonan and Bingham (ref. 15) were obtained on a modi-fied profile with a reflex training edge that reduced CMo to approximately zero,
compared with the present value of -0.027 at H. - 0.3 (cf. table 8). Also, the dataof Jepson (ref. 16) in figure 27 came from a slotted-wall tunnel with 12.5% porosity,which was thought to yield somewhat lower values of CL than comparable tests insolid-wall tunnels. Furthermore, the Reynolds numbers in references 15 and 16 were
11
lover than those of the present tests. Nevertheless, the discrepancies in figure 27seem to be -.oo large to be attributed to these factors or to measurement uncertain-ties. It will be shown later that dynamic data on the SC-1095 section are generallyin better agreement.
Hughes Hl1-02 airfoil- Figure 28 shows the measured maximum lift coefficients forthe present HH-02 airfoil, in comparison with !ata from a section that is almostidentical except for a slightly smaller leading-edge radius (ref. 27). Although theMach number range does not overlap, the two sets of results seem consistent.
Vertol VR-7 airfoil- Results from four sources are plotted in figure 29 for theVR-7 profile. The present data are somewhat higher than those of Coulomb (ref. 28),primarily because the stall occurred at slightly higher angles of attack, but thelift-curve slopes (not shown) and the effect of a boundary-layer trip were approxi-mately the same. The value of CLmax at M. - 0.3 is slightly lower than that ofDadone (ref. 5), whose measurements at higher Mach numbers exceed considerably thoseof Bingham et al. (ref. 29).
NLR-l airfoil- Figure 30 shows the good agreement of the present measurementswith those of Dadone (ref. 19) for the NLR-I airfoil. It should be mentioned, how-ever, that the details of the pitching-moment behavior in the vicinity of CLmax (not
shown) were somewhat different. As in the previous example, the data of Noonan andBingham (ref. 24) for Clmax at M. 2 0.35 tend to be lower than the data of Dadone
(ref. 19). This airfoil appears to be more sensitive to Mach number than any of theother modern helicopter sections.
NLR-7301 airfoil- As shown in figure 31, the maximum static lift for theNLR-7301 airfoil exceeded that of the other sections by a considerable margin; how-ever, CM was -0.083 (cf. table 8). The values of CLmax shown are also greater
than those obtained at NLR under virtually identical conditions (ref. 30). This wasobtained at a significantly larger stall angle, more than 1* larger at M. - 0.18,than in the NLR experiments, apparently because of different boundary-layer separa-tion characteristics and sidewall interferences.
Dynamic Data
Although the static data described above comprised an essential part of theinvestigation, the primary objective was to obtain a comon data base of unsteadycharacteristics for helicopter applications. In this section some representativeexamples are presented and comparisons made with other investigations. More completediscussions of the basic phenomena and of the results obtained are given in refer-ences 1 and 2.
The unsteady stall-onset and dynamic-stall counterparts of the static CLmax
results discussed above are shown in figures 32 and 33, reproduced from reference 2with some minor corrections. The dashed lines in figure 33 indicate the estimateddeep-stall CLmax for the NLR-7301 airfoil; data were not obtained for this condi-
tion for M. > 0.25. These results have not been corrected for wind-tunnel-wallinterference.
12
Figures 32 and 33 illustrate an important general result of the investigation:the parameters of the unsteady motion tend to be more important than the airfoilgeometry. For example, the differences in the values of CLmax for the Wortmann,
Sikorsky, and Hughes airfoils can hardly be discerned within the experimental uncer-tainty, but the unsteady stall-onset and deep-stall results are much higher than thestatic values shown ir. figures 26-28 and 33. It is also interesting to note that atleast for M S 0.25, the deep-stall CLmax values for the NLR-l and NLR-7301 airfoils
are almost identical. In contrast, the static and unsteady stall-onset results forthese two very different profiles are considerably different and represent the lowerand upper bounds, respectively, of all the airfoils tested.
In view of the aforementioned scatter in the static results from different windtunnels, it is logical to inquire how different sets of dynamic data might compare.Because of the large number of parameters that affect dynamic stall and the tendencyfor past investigators to select different combinations of these parameters, the pos-sibilities for direct comparison of unsteady results are much more limited. However,some examples are given below.
NACA 0012 airfoil- The first comparison for this profile is shown in figures 34and 35, where data from reference 3 were obtained in the same wind tunnel as the pres-ent results, but with a model whose chord was twice as large. Figure 34 shows thatthe large values of CLmax reported in reference 3 were not realized in the present
experiment. Figure 35 shows CL versus a, where the two results are seen to differby approximately 10% during the portion of the cycle when a is increasing but beforedynamic stall begins. This is approximately the same as the difference in the lift-curve slopes for the corresponding static data, and it is consistent with the differ- tences that would be predicted for static wind-tunnel-wall corrections (ref. 12) forthe two chord-to-height ratios. However, it can be inferred from the differences inthe peaks of the lift curves in figure 35 that the organized vortex-shedding phenom-enon was more pronounced on the larger model after stall began. Also, reattachmentof the boundary layer on the downstroke occurred earlier. These do not seem to besolely Reynolds-number effects; rather, it is suspected that in the earlier teststhere was excessive interference between the boundary layers on the upper and lowerwalls of the tunnel and the unsteady viscous flow on the ends of the verticallymounted airfoil.
St. Hilaire and Carta (ref. 17) have reported on dynamic-stall tests of theNACA 0012 airfoil at UTRC under conditions similar to those in the present experiment.Figure 36 compares some of the data from the two investigations. The format andchoice of unsteady parameters is based on an extension of the observation in refer-ence 2, that for sinusoidal pitching oscillations the values of amax and the prod-uct alk 2 seem to be particularly important in determining the detailed time-historyof the unsteady airloads during dynamic stall. In order to compare as many testpoints as possible, data were selected that satisfied the criterion0.0014 < alk 2 < 0.0022, where a, is in radians. The variations in CLmax and CMmin
ir figure 36 are seen to correlate reasonably well on this basis, and the resultsfrom the two sources are in fairly good agreement. Some of the Cmax data from the
UTRC wind tunnel are slightly higher than the present measurements.
SC-l095 airfoil- Gangwani (ref. 18) has reported data that were obtained on theSC-1095 section in the same facility that was used by St. Hilaire and Carta (ref. 17)to obtain the NACA 0012 data described in the preceding paragraph. The results are
13
compared with the present data in figure 37, following the same format as above.Fewer data points are available, but the degree of correlation is approximately com-parable to that of the NACA 0012 results in figure 36. In contrast with that figure,however, the present values of CLmax tend to be slightly higher than the UTRC data
(ref. 18). In any case, the discrepancies generally appear to be within the measure-ment uncertainty, and the agreement is better than for the static results (fig. 27).
NLR-l airfoil- This profile was tested by Dadone (ref. 19) over a wide range ofMach numbers, mean angles, and amplitudes. Based on the considerations outlinedabove regarding amax and 1Ik
2I his results are compared with the present data infigure 38 as functions of alk at a constant value amax = 200, where a3 is alsoin degrees. The lift data are in better agreement than in the previous examples,but more scatter appears in the pitching-moment results than before.
No unsteady results from other sources are presently available from othersources for comparison with the data obtained on the Wortmann FX-098, Ames A-01,Hughes HH-02, Vertol VR-7, and NLR 7301 airfoils.
Comments on Wind-Tunnel Effects
It is well known that testing the same airfoil in different wind tunnels oftengives different results, especially for the static-stall characteristics. This isborne out in figures 24-31. In fact, if the results from these eight figures wereoverlaid, the real differences between the individual airfoils would be almost com-pletely obscured by the differences attributable to the test facilities.
Although more limited in scope, the comparisons of dynamic-stall data shown infigures 36-38 are more encouraging than the static results. Since all of these datacame from tests with either high aspect-ratio models or sidewall boundary-layer con-trol, this suggests that the present dynamic data may be relatively free of wind-tunnel-wall contamination and other three-dimensional effects. A detailed examina-tion of the complete time-histories of the unsteady airloads and further studies onmodels of various aspect ratios would be required to confirm this speculation.
A special feature of the present experiment is that a large number of airfoilswere studied over a wide range of unsteady flow conditions in the same facility.This provides the basis for meaningful comparisons, even though wind-tunnel inter-ference effects were not completely negligible. However, as stated in reference 1,it is recommended that the wind-tunnel walls be included or considered in any quan-titative uses of the data.
5. SUMMARY AND CONCLUSIONS
A large amount of steady and unsteady data has been obtained on eight airfoilsections over a wide range of test conditions, at Mach numbers up to 0.30. Thedetails of the experimental arrangements, estimates of the measurement accuracy, andthe test conditions are described in this volume. Some comparisons are also madewith data from other sources. Volume 2 (Pressure and Force Data) presents theresults in graphical form and describes the digital computer tapes that contain theextensive numerical data. Volume 3 (Hot-Wire and Hot-Film Measurements) describesthe boundary-layer studies performed with surface-mounted hot wires and hot films.
14
_ _ t
The results of the experiment show important differences between airfoils, dif-ferences that would otherwise tend to be masked by differences in wind tunnels, par-ticularly in steady cases. All of the airfoils tested offer significant advantagesover the standard NACA 0012 profile. In general, however, the parameters of theunsteady motion appear to be more important than airfoil shape in determining thedynamic-stall airloads.
15
7!
REFERENCES
1. McCroskey, W. J.; and Pucci, S. L.: Viscous-Inviscid Interaction on OscillatingAirfoils in Subsonic Flow. AIAA J., vol. 20, no. 2, Feb. 1982, pp. 167-174.
2. McCroskey, W. J.; McAlister, K. W.; et al.: Dynamic Stall on Advanced AirfoilSections. J. American Helicopter Soc., vol. 26, no. 3, July 1981, pp. 40-50.
3. McAlister, K. W.; Carr, L. W.; and McCroskey, W. J.: Dynamic Stall Experimentson the NACA 0012 Airfoil. NASA TP-IIO0, 1978.
4. Carr, L. W.; McAlister, K. W.; and McCroskey, W. J.: Analysis of the Develop-ment of Dynamic Stall Based on Oscillating Airfoil Experiments. NASATN D-8382, 1977.
5. Dadone, L. U.: U.S. Army Helicopter Design Datcom. Vol. I - Airfoils. NASACR-153247, 1976.
6. Hicks, R. M.; and McCroskey, W. J.: An Experimental Evaluation of a HelicopterRotor Section Design by Numerical Optimization. NASA TM-78622, 1980.
7. Balch, D. T.: Helicopter Blade. L.S. Patent 3,728,045, 1973.
8. Kemp, L. D.: An Analytical Study for the Design of Advanced Rotor Airfoils.NASA CR-112297, 1973.
9. Barche, J., ed.: Experimental Data Base for Computer Program Assessment. AGARDAdvisory Report 138, Advisory Group for Aerospace Research and Development,Neuilly-sur-Seine, France, 1979.
10. Carr, L. W.; and McCroskey, W. J.: A Directionally Sensitive Hot-Wire Probe forDetection of Flow Reversal in Highly Unsteady Flows. International Congresson Instrumentation in Aerospace Facilities, 1979 Record, Sept. 1979,
pp. 154-162.
11. McCroskey, W. J.; McAlister, K. W.; and Carr, L. W.: Dynamic Stall Experimentson Oscillating Airfoils. AIAA J., vol. 14, no. 1, Jan. 1976, pp. 57-63.
12. Allen, H. J.; and Vincenti, W. G.: Wall Interference in a Two-Dimensional FlowWind Tunnel with Consideration of the Effect of Compressibility. NACAReport 782, 1944.
13. Woodward, D. S.: The Twodimensional Characteristics of a 12.2% Thick R.A.E.Aerofoil Section. RAE Technical Report 68303, Royal Aircraft Establishment,Farnborough Hants, England, Jan. 1969.
14. Fromme, J. A.; and Golberg, M. A.: Unsteady Two-Dimensional Airloads Acting onOscillating Airfoils in Subsonic Ventilated Wind Tunnels. NASA CR-2914,1977.
15. Noonan, K. W.; and Bingham, G. J.: Aerodynamic Characteristics of Three Heli-copter. Rotor Airfoil Sections at Reynolds Numbers from Model Scale to FullScale at Mach Numbers from 0.35 to 0.90. NASA TP-1701, 1980.
16
i
16. Jepson, W. D.: Two Dimensional Test of Four Airfoil Configurations with anAspect Ratio of 7.5 and a 16 Inch Chord up to a Mach Number of i.i. ReportSER-50977, Sikorsky Aircraft, Stratford, Conn., Apr. 1977.
17. St. Hilaire, A. L.; and Carta, F. 0.: The Influence of Sweep on the Aerody-namic Loading of an Oscillating NACA 0012 Airfoil. Vol. II - Data Report.NASA CR-145350, 1979.
18. Gangwani, S. T.: Prediction of Dynamic Stall and Unsteady Airloads for RotorBlades. American Helicopter Society Paper 81-01, May 1981.
19. Dadone, L. U.: Two-Dimensional Wind Tunnel Test of an Oscillating Rotor Airfoil.NASA CR-2915, 1977.
20. Lizak, A. A.: Two-Dimensional Wind Tunnel Tests of an H-34 Main Rotor AirfoilSection. TREC Technical Report 60-53, U.S. Army Transportation ResearchCommand, Ft. Eustis, VA, 1960.
21. Prouty, R. W.: A State-of-the-Art Survey of Two-Dimensional Airfoil Data.J. American Helicopter Soc., vol. 20, no. 4, Oct. 1975, pp. 14-25.
22. Bevert, A.: Essais Comkaratifs en Courant Plan des Profils "G.1" et NACA 0012.ONERA Doc. No. 76/1157.AN, Office National d'Etudes et de Recherches 1ro-spatiales, ChAtillon, France, Mar. 1970.
23. Harris, C. D.: Two-Dimensional Aerodynamic Characteristics of the NACA 0012Airfoil in the Langley 8-Foot Transonic Pressure Tunnel. NASA TM-81927,1981.
24. Noonan, K. W.; and Bingham, G. J.: Two-Dimensional Aerodynamic Characteristicsof Several Rotorcraft Airfoils at Mach Numbers from 0.35 to 0.90. NASATM X-73990, Jan. 1977.
25. Bingham, G. J.; and Noonan, K. W.: Low-Speed Aerodynamic Characteristics ofFive Helicopter Blade Sections at Reynolds Numbers from 2.4xi06 to 8.4xi06 .NASA TM X-2467, 1972.
26. Wortmann, F. X.: Design of Airfoils with High Lift at Low and Medium SubsonicMach Numbers. Paper No. 7, AGARD Conference Proceedings CP-102, AdvisoryGroup for Aerospace Research and Development, Neuilly-sur-Seine, France,1972.
27. Prouty, R. W.: Airfoil Section Data Report. Report No. 150-A-1012, HughesHelicopters, Culver City, Calif., Mar. 1978.
28. Coulomb, J.: Caractdristiques Stationnaires du Profil VR 7. Procs-verbalNo. 102 B/SC, Centre d'Essais Adronautique de Toulouse, Toulouse, France,June 1979.
29. Bingham, G. J.; Noonan, K. W.; and Jones, H. E.: Results of an Investigationof Several New Rotorcraft Airfoils as Related to Airfoil Requirements.Paper No. 8, NASA Conference Publication 2046, Mar. 1978.
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30. Joosen, C. J. J.; and Kho, C. G.: Two Dimensional Low-Speed Wind TunnelInvestigation on a NLR 73-108-10 Airfoil with Fowler Type Flap, Part 1:Text, Tables, and Figures. NLR TR 74058 C, National Lucht- enRuimtevaartlabotorium, Amsterdam, The Netherlands, Sept. 1975.
18
TABLE 1.- HARMONIC COEFFICIENTSOF THE OSCILLATION MECHANISM
a= + a,1 Sin Wt + Q2 Sin(Wt + 02)
5 10 1.00.05.2 (a)
10 10 9.90 .20 260015 10 9.90 .20 (a)
15 14 14.10 .38 2000
aNot measured.
19
TABLE 2. - AIRFOIL COORDINATES: NACA 0012 AND AMES A-01 AIRFOILS
x/c NACA 0012, y/c AMES A-01, y/c
upper lower upper lower
0.0000 0.00000 0.00000 0.00000 0.000000.0005 0.00395 -0.00395 0.00377 -0.003380.0010 0.00556 -0.00556 0.00541 -0.004720.0020 0.00781 -0.00781 0.00766 -0.006510.0035 0.01027 -0.01027 0.01013 -0.008440.0050 0.01221 -0.01221 0.01214 -0.009940.0065 0.01386 -0.01386 0.01388 -0.011200.0080 0.01531 -0.01531 0.01543 -0.012270.0100 0.01704 -0.01704 0.01732 -0.013500.0125 0.01894 -0.01894 0.01945 -0.014810.0160 0.02127 -0.02127 0.02214 -0.016340.0200 0.02360 -0.02360 0.02490 -0.017770.0250 0.02615 -0.02615 0.02801 -0.019220.0350 0.03043 -0.03043 0.03335 -0.021370.0500 0.03555 -0.03555 0.03991 -0.023650.0650 0.03966 -0.03966 0.04523 -0.025490.0800 0.04307 -0.04307 0.04961 -0.027100.1000 0.04683 -0.04683 0.05421 -0.029020.1250 0.05055 -0.05055 0.05829 -0.031040.1500 0.05345 -0.05345 0.06098 -0.032770.2000 0.05737 -0.05737 0.06344 -0.035510.2500 0.05941 -0.05941 0.06431 -0.037270.3000 0.06002 -0.06002 0.06446 -0.038280.3500 0.05949 -0.05949 0.06409 -0.038660.4000 0.05803 -0.05803 0.06316 -0.038480.4500 0.05581 -0.05581 0.06154 -0.037820.5000 0.05294 -0.05294 0.05924 -0.036650.5500 0.04952 -0.04952 0.05623 -0.035010.6000 0.04563 -0.04563 0.05249 -0.032970.6500 0.04132 -0.04132 0.04792 -0.030560.7000 0.03664 -0.03664 0.04246 -0.027850.7500 0.03160 -0.03160 0.03600 -0.024860.8000 0.02623 -0.02623 0.02860 -0.021530.8500 0.02053 -0.02053 0.02064 -0.017860.9000 0.01448 -0.01448 0.01260 -0.013740.9250 0.01132 -0.01132 0.00899 -0.011440.9500 0.00807 -0.00807 0.00598 -0.008880.9750 0.00472 -0.00472 0.00392 -0.006030.9900 0.00265 -0.00265 0.00322 -0.004211.0000 0.00126 -0.00126 0.00299 -0.00300
ro/C - 0.0158 r /c - 0.012
20
TABLE 3. - AIRFOIL COORDINATES: WORTMANN FX-098 AND SIKORSKY SC-1095 AIRFOILS
x/c WORTMANN FX-098, y/c SIKORSKY SC-1095, y/c
upper lower upper lower
0.0000 0.00000 0.00000 0.00000 0.000000.0005 0.00293 -0.00249 0.00307 -0.002570.0010 0.00426 -0.00343 0.00443 -0.003680.0020 0.00619 -0.00471 0.00640 -0.005350.0035 0.00837 -0.00609 0.00865 -0.007240.0050 0.01017 -0.00717 0.01054 -0.008800.0065 0.01175 -0.00807 0.01221 -0.010160.0080 0.01319 -0.00886 0.01374 -0.011380.0100 0.01494 -0.00978 0.01560 -0.012850.0125 0.01692 -0.01079 0.01771 -0.014500.0160 0.01944 -0.01202 0.02041 -0.016570.0200 0.02204 -0.01321 0.02320 -0.018650.0250 0.02501 -0.01451 0.02635 -0.020920.0350 0.03021 -0.01664 0.03140 -0.024540.0500 0.03681 -0.01913 0.03677 -0.028420.0650 0.04234 -0.02111 0.04070 -0.031080.0800 0.04705 -0.02277 0.04374 -0.032950. 1000 0.05222 -0.02464 0.04680 -0.034640.1250 0.05714 -0.02658 0.04963 -0.036190.1500 0.06073 -0.02819 0.05174 -0.037390.2000 0.06491 -0.03059 0.05447 -0.038840.2500 0.06650 -0.03198 0.05548 -0.039330.3000 0.06630 -0.03251 0.05524 -0.039180.3500 0.06515 -0.03242 0.05437 -0.038580.4000 0.06336 -0.03184 0.05299 -0.037600.4500 0.06097 -0.03096 0.05105 -0.036220.5000 0.05798 -0.02982 0.04854 -0.034460.5500 0.05445 -0.02843 0.04555 -0.032340.6000 0.05040 -0.02678 0.04212 -0.029850.6500 0.04586 -0.02487 0.03819 -0.027020.7000 0.04085 -0.02273 0.03375 -0.023840.7500 0.03543 -0.02034 0.02887 -0.020340.8000 0.02962 -0.01768 0.02362 -0.016580.8500 0.02337 -0.01473 0.01808 -0.012650.9000 0.01642 -0.01134 0.01235 -0.008650.9250 0.01253 -0.00932 0.00943 -0.006640.9500 0.00856 -0.00702 0.00642 -0.004540.9750 0.00476 -0.00423 0.00328 -0.002330.9900 0.00255 -0.00237 0.00132 -0.000931.0000 0.00110 -0.00110 0.00000 0.00000
ro/c 0.007 ro/c 0.008
21
TABLE 4. - AIRFOIL COORDINATES: HUGHES HH-02 (-50 TAB) AND VERTOL VR-7 (-30 TAB) AIRFOILS
x/c HUGHES HH-02, y/c VERTOL VR-7, y/c
upper lower upper lower
0.0000 0.00000 0.00000 0.00000 0.000000.0005 0.00283 -0.00284 0.00337 -0.003300.0010 0.00405 -0.00388 0.00483 -0.004600.0020 0.00594 -0.00532 0.00696 -0.006330.0035 0.00819 -0.00683 0.00943 -0.008000.0050 0.01009 -0.00800 0.01149 -0.009190.0065 0.01176 -0.00895 0.01330 -0.010100.0080 0.01327 -0.00978 0.01494 -0.010860.0100 0.01510 -0.01072 0.01695 -0.011720.0125 0.01717 -0.01172 0.01923 -0.012630.0160 0.01975 -0.01290 0.02213 -0.013670.0200 0.02237 -0.01404 0.02512 -0.014670.0250 0.02531 -0.01524 0.02846 -0.015750.0350 0.03029 -0.01714 0.03423 -0.017510.0500 0.03640 -0.01943 0.04144 -0.019660.0650 0.04137 -0.02127 0.04759 -0.021540.0800 0.04553 -0.02276 0.05299 -0.023200.1000 0.05012 -0.02432 0.05922 -0.025160.1250 0.05468 -0.02575 0.06565 -0.027090.1500 0.05828 -0.02675 0.07091 -0.028550.2000 0.06328 -0.02793 0.07887 -0.030550.2500 0.06608 -0.02843 0.08378 -0.031860.3000 0.06738 -0.02834 0.08592 -0.032730.3500 0.06750 -0.02755 0.08574 -0.033080.4000 0.06640 -0.02600 0.08365 -0.032710.4500 0.06391 -0.02377 0.07984 -0.031480.5000 0.06008 -0.02104 0.07451 -0.029520.5500 0.05504 -0.01797 0.06781 -0.027120.6000 0.04891 -0.01482 0.05996 -0.024640.6500 0.04174 -0.01176 0.05171 -0.022070.7000 0.03344 -0.00952 0.04322 -0.019290.7500 0.02403 -0.00851 0.03442 -0.016390.8000 0.01436 -0.00889 0.02527 -0.013460.8500 0.00481 -0.00984 0.01575 -0.010500.9000 -0.00431 -0.01041 0.00558 -0.007440.9250 -0.00394 -0.00777 0.00117 -0.006090.9500 -0.00203 -0.00583 -0.00016 -0.005120.9750 -0.00006 -0.00387 0.00115 -0.003800.9900 0.00112 -0.00269 0.00194 -0.003001.0000 0.00190 -0.00190 0.00247 -0.00247
ro/c- 0.008 r0 /c 0.011
22
TABLE 5. - AIRFOIL COORDINATES: NLR-1 AND NLR-7301 AIRFOILS
x/c NLR-1, y/c NLR-7301, y/c
upper lower upper lower
0.0000 0.00000 0.00000 0.00000 0.000000.0005 0.00359 -0.00288 0.00730 -0.007480.0010 0.00499 -0.00388 0.01051 -0.010200.0020 0.00687 -0.00518 0.01518 -0.013730.0035 0.00890 -0.00643 0.02030 -0.017350.0050 0.01053 -0.00730 0.02424 -0.020160.0065 0.01194 -0.00799 0.02756 -0.022520.0080 0.01321 -0.00858 0.03043 -0.024550.0100 0.01475 -0.00929 0.03375 -0.026880.0125 0.01648 -0.01006 0.03729 -0.029350.0160 0.01868 -0.01101 0.04140 -0.032250.0200 0.02097 -0.01196 0.04514 -0.035020.0250 0.02358 -0.01301 0.04873 -0.037940.0350 0.02799 -0.01477 0.05372 -0.042640.0500 0.03328 -0.01688 0.05920 -0.048060.0650 0.03750 -0.01859 0.06321 -0.052290.0800 0.04093 -0.02007 0.06636 -0.055760.1000 0.04435 -0.02179 0.06985 -0.059620.1250 0.04701 -0.02363 0.07347 -0.063580.1500 0.04905 -0.02522 0.07648 -0.066890.2000 0.05200 -0.02775 0.08115 -0.071940.2500 0.05386 -0,02958 0.08441 -0.075270.3000 0.05489 -0.03082 0.08649 -0.077130.3500 0.05528 -0.03154 0.08755 -0.07763
0.4000 0.05511 -0.03185 0.08764 -0.076720.4500 0.05443 -0.03176 0.08678 -04074120.5000 0.05327 -0.03126 0.08495 -0.069340.5500 0.05164 -0.03025 0.08206 -0.062370.6000 0.04948 -0.02882 0.07789 -0.053860.6500 0.04677 -0.02707 0.07212 -0.043970.7000 0.04348 -0.02503 0.06458 -0.033160.7500 0.03892 -0.02276 0.05551 -0.022270.8000 0.03172 -0.02028 0.04523 -0.012210.8500 0.02368 -0.01756 0.03415 -0.004090.9000 0.01562 -0.01427 0.02269 0.001080.9250 0.01179 -0.01199 0.01696 0.002280.9500 0.00811 -0.00903 0.01129 0.002460.9750 0.00454 -0.00511 0.00577 0.001530.9900 0.00244 -0.00253 0.00258 0.000421.0000 0.00103 -0.00103 0.00055 -0.00055
r/c- 0.007 r/C 0.0550 0
23
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24
TABLE 7. - STATIC DRAG COEFFICIENTS AT M - 0.30 BASED ON WAKE SURVEYS
a, deg N-0012 AMES-O W-098 SC-1095 HH-02 VR-7 NLR-l NLR-7301
-5.0 0.00843 0.00851 0.00886 0.00739 0.00846 0.00899 0.02602 0.00952-2.0 0.00729 0.00832 0.00771 0.00713 0.00719 0.00759 0.00743 0.007800.0 0.00711 0.00794 0.00683 0.00708 0.00679 0.00723 0.00710 0.009682.0 0.00718 0.00662 0.00664 0.00670 0.00655 0.00707 0.00745 0.008915.0 0.00865 0.00767 0.00755 0.00807 0.00816 0.00800 0.00831 0.010118.0 0.01031 0.00965 0.01142 0.01013 0.01112 0.01059 0.01086 0.0130510.0 0.01190 0.01248 0.01405 0.01127 0.01382 0.01353 0.01322 0.0156912.0 0.01711 0.01600 0.01773 0.01586 0.01849 0.02156 0.02006 0.0202213.0 .. .. .. 0.02015 0.02236 .- -.
14.0 0.02901 - - 0.08922 - - - -...
-4.0 ....- - - -. 0.00773 0.00843-1.75 - - 0.00874-1.0 .. .. ...- - 0.009621.0 - - 0.00738 .......... 0.009731.5 .. .. ...- - 0.009102.5 - - 0.008963.0 - - 0.00702 . .. .. .. .. .. .4.0 - - 0.00712 . .. .. .. .... .6.0 - - 0.00791 . .... .... .. .9.9 - - 0.01218 . .. .. .. .. .. .
TABLE P.- SUMMARY OF THE MEASURED STATIC AIRFOIL CHARACTERISTICS AT M. = 0.30,INCLUDING WIND TUNNEL WALL CORRECTIONS
Airfoil CLa a° CMo CDmin Xa'c" CLmax Oss (L/D)max
NACA 0012 0.109 -0.10 -0.007 0.0072 0.24 1.33 13.70 90Ames-Ol .111 -.6 -.005 .0070 .25 1.45 13.6 100FX-098 .109 -1.3 -.02b .0066 .24 1.43 13.1 94SC-1095 (.11O)a -.9 -.027 .0073 .245 (1.4 6)a 13.5 (98)a
HH-02 .114 -.6 -.002 .0066 .255 1.42 13.2 92VR-7 .117 -1.6 -.016 .0071 .26 1.51 12.6 107NLR-1 .102 -1.0 -.025 .0071 .22 1.29 12.4 87NLR-7301 .117 -1.9 -.083 .0078 .25 (1 .8 3)a (1 7 .2)a 89Nominal ±.003 .2 .005 .0005 .005 .03 .3 5
uncertainty
auncertainty larger than nominal value in table.
25
TABLE 9.- LIST OF TEST POINTS WITH UNUSUAL ZERO DRIFT OF PRESSURE TRANSDUCERS
Airfoil Frame M. Type Problem Airfoil Frame M. Problemtransducers TyPe transducers
NACA 0012 8019 U.035 U All Wortmann8021 FX-098 18414 .11 S 20,228023 19401 .25 2,3,48102 19402 .258104 19405 .258106 19406 .258114 .07 23 20103 .25 2,38116 .07 23 20104 .258118 .07 23 20122 .308210 .11 4 20123 .3012118 .26 Q.S. 3 20203 .3013107 .11 1,4,20 20204 .3013115 .07 Many Sikorsky13120 .07 1,3,4,18, SC-]095 33022 .07 U 1,17,18,25
24,26 33106 .11 U Many13205 .033 Many 33110 .11 U Many13217 .035 Many 34409 .29 U 2,314104 .18 U 3,8 35021 .30 S 1114106 .18 U 3,8 35023 .30 1114108 .18 U 3,8 35100 .30 11
Ames A-01 26306 .30 S 2,3 35102 .30 1126307 .30 2,3 35103 .30 1128019 .11 1,20 36209 .11 1,20,2228021 1,20 36210 .11 1,20,2228023 1,20 35211 .11 1,20,2228101 1,20 35212 .11 1,20,2228106 All 35213 .11 1,20,2228107 Hughes28109 HH-02 42309 .22 U 62.8110 42313 .25 628115 43308 .30 1328116 43-09 .30 1328117 Vertol28119 VR-7 47213 .18 1,4,2428120 + 47217 .22 1,4,2429317 .035 U 5,12,14,23 47301 .25 3,24
Wortmann 1 47305 .28 3,24FX-098 16019 .035 U Many NLR-1 62020 .07 1,16,18
16200 .18 U 4,11 63018 .30 217220 .30 U 2 63019 .30 218102 .18 S 2,3,4 63020 .30 218106 .18 2,3,4 63021 .30 218108 .18 2,3,4 65207 .20 2,3,418410 .11 20,22 65209 .30 2,3,418411 .11 20,22 NLR-7301 66616 .11 S Many18413 .11 I 20,22 NLR-7301 66617 .11 S Many
aS - steady; U - unsteady; Q quasi-steady, k 0.002.
26
TABLE 10.- COEFFICIENTS OF LINEAR CURVE-FIT OF STATIC LIFT DATAWITHOUT WIND-TUNNEL CORRECTIONS
CL A + Ba
Airfoil A = CL(O) B - OCLa
NACA 0012 0 0.110Ames 01 .15 .108Wortmann FX-098 .07 .111Sikorsky SC-1095 .11 .110Hughes HH-02 .07 .116Vertol VR-7 .19 .117NLR-1 .11 .102NLR-7301 .24 .116
I
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TABLE 13.- MACH NUMBER SWEEP AT a - 150 + 10° sin wt, k = 0.10
M NACA 0012 A-01 FX-098 SC-1095 HH-02 VR-7 NLR-l NLR-7301
0.035 8102 16019 58102.07 8114 24323 16105 33022 42121 47123 62020
.11 8214 24314 16114 33106 42321 |4711 62104 67120124217 I47213
.18 8220 2317 16200 33110 42302 4721 62112 67220131209 58121
.18T 141 29117 17103 34321 42110 47112 64109 67021
.20 62114.20 165207
.22 9202 24209 16300 33205 42309 47217 62208
.25 9203 24201 16308 33207 42313 47301 62210 67305
.28 9208 24117 22208 33215 42218 47305 62218
.29 |9217 24105 22201 33300 42210 45023 f62307114220 165209
.29T 14208 29106 17200 34308 42100 47100 64023
aT = trip.
TABLE 14.- FREQUENCY SWEEP AT M . 0.29, a - 15* + 100 sin wt
ka NACA 0012 A-01 FX-098 sC-1095 HH-02 VR-7 NLR-1 NLR-7301
0.01 9210 30019 21100 38300
.025 9213 24022 22023 33217 42206 45019 6230214218
.025T 14117 29023 17117 42019 47020 6401914200
.05 9214 24100 22103 33222 42208 45021 623041421914119
.05T 14202 29101 17119 34306 42021 47022 64021
9217 62307.10 14220 24105 22201 33300 42210 45023 65209
• 1420
.10T 29106 17200 34308 42100 47100 64023114210[42212
.15 9218 24109 22206 34409 4 45101 62309
aT - trip.
45
TABLE 15.- FREQUENCY SWEEP AT M. = 0.30, a - 100 + 100 sin wt
k NACA 0012 A-01 FX-098 SC-1095 HH-02 VR-7 NLR-1 NLR-7301
0.01 9221 30105 21107 38306 43019 45109 62317 69019125022
.025 9222 |31102 22216 37023 43106 45111 62320 69100131102
.05 9223 125102 22217 37101 43108 45113 62322 69102131104
.10 9302 25104 22218 37107 43112 45117 62400 69105
.12 62403525109 143114
.15 9307 31110 22219 37109 431 45119 62405 69107131112
TABLE 16.- FREQUENCY SWEEP AT M = 0.30, a = 15* + 5* sin wt
k NACA 0012 A-01 FX-098 SC-1095 HH-02 VR-7 NLR-1 NLR-7301
0.01 10113 30110 21112 39104 45203 63018 68019.025 10114 25204 23021 38021 43303 45205 63019 68100.05 10117 25205 23022 38022 43304 45207 63020 68102.10 10118 25208 23023 38102 43305 45209 63021 68104.12 63100.15 10120 25209 23100 38103 43308 45211 63101 68109.20 10123 25210 23101 38104 43309 45213 63102 68111
TABLE 17.- FREQUENCY SWEEP AT M. 0.30, a = 100 + 50 sin wt
k NACA 0012 A-01 FX-098 SC-1095 HH-02 VR-7 NLR-1 NLR-7301 NLR-7301T
0.01 10202 30119 21200 39107 44019 68119S12 4 221618 682170
.025 17112 25117 22307 37207 144029 45221 63108 68121 67108
.05 10204 25118 22308 37208 44023 45223 68123 67110
.075 10207.1 713f44104
.70 7113 25119 22309 37210 44104 45300 63112 68201 6711210208 24118
.15 700 (25121 22311 37213 44106 45302110211 144112
.20 I7114 25123 22312 37215 44120 45303 63114 68203
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3 %0
o 0 -.i * s0 0 0* 0
*~1- -. 0% 0 II CJN (' ' N
ID 0 C) Ni (D 0 1- r-1 -7 -. II
0 T- t t00l
0 0;
C! - -4 a 0 00 1 C ,4 N % C -4 C4-4 11
I ciE-4 ~C-0
5-; -4 -4 M 4 n A0
Ito 0 14 11 1 0 I r4U) ) 00 0 U, 1 a - r, r -
a% IT 0- 4C4( 1 n n -1I 0 1 % c(n V)Al 0-U -It(N-,IC..-,4 0- u' 4 0-4 u-0
0 C1 CA 0
~e InIif
-4 coIN 4 m n M0 -0 m 04 0- o4 C1 4 N 4C
0 0 -7 0 .4 - 4
0H 0 % 0En 0 1. LM
^- . n M r - - n c)0I 4L 4L1 % l
4 -4 ) -4 w4.1 4M 40 N 0 aeen u0 0 ' G CN U 0
V))
0 0 C1 G
C) 0 o -4 -4 47
0 00 0 0 00
oD ON0 r-'. %00o0 %0 -4 rt 0 ts- 4 003-t 'o e Ir--Icn 4 4 -10 o--4 -I 4N '-4 0 r-4 C4 -I0 -IC r-4 -4
r-4 C- cn .-I Nq N m N117% 110 ON0 l 0% 110m 110 ON ' NO a,
0 0 ~ 0 0 % 0 0 0C5 d
4133 0 0 0
'-4 C n0 0- e
I -4 ' N - N
0. en0 11M
0 0 0
H C4 0 C 0 0 'D 0 0 0n 04 0n r 40 -4~ 0 r1*1 0 04 P-4 q 0 04C4 -
6n In C' -4 CO) 1-4 CN m 4 Nq
0 0 0 0 0 0 0 0 0
0 0 0 0 0 0 0 0 0
04~ -1 U1 C4 LNC1 L3 -.t N~ &I 0M V)TC49 .
N* -IN " -I-4 N .-N-4 qP-IN" -4 - '-4 0 '-401 N N Ldn N Nq N4 m' N N
1- 0 t 11'T - It -It 0r -N 'T'-E-4 ~ ~ I 't 0 't o * a o
E4 0 0 0 0 0 0 0 0 6
0-4
in4 04LC7% --4 -40
'-4
z co
00
044
I 08
00 0 0
r,-)4 4 ,L -o- - -40 .- q -4 ,-1 "
0 0 083 Z5 80 0 0 0-4 0 It 0
r-i0 r-4~ C -4 0 -4r4
0 0 0 0
o a 0 0 0
4 0 00 0 ,
-0 -0 r-40 -- -In eC-1 C4 C, C,4
11W 00 t0 11 0 11W co f-- t -It -zr '
0 0 0 0 0
0 0 0 0 0 0Lr) L11 L1 Lfn 114
ej40 -4 -4 0 0 -44 r4 0uJ 0 C,4 enJ C,4 cq (N
C.) Z * 0 0 * 0 0 0 0
-I ( 0 0
Ln, C1 -I -T %
0 C" -4 -4
E-4 U.. 0 0 0e3 C5 8
0 0 0W (0 -0 .0
0% -4 -404C -4 -
J 00 0-4 C4 1 I 4 y
1-4 4 - 4 C-4 r- 0 0 4
0 0 00
-40
49
TABLE 22.- NO SEPARATION: M = 0.30, = 50 + 50 sin wt
k NACA 0012 A-01 FX-098 SC-l095 HH-02 VR-7 NLR-Ia NLR-7301a
0.01 10218.10 10221 25301 23107.20 10222 25303 23109 68211
aSee table 24.
TABLE 23.- DYNAMIC BOUNDARY-LAYER TRIP DATA
M. k NACA 0012 A-01 FX-098 SC-1095 HH-02 VR-7 NLR-1 NLR-7301
0.18 0.05 14104 29115 17100 34318 42108 47110 64107 6701914021
.18 .10 114021 29117 17103 34321 42110 47112 64109 67021114106
.18 .15 114108 29119 17109 34323 42113 47114 64111
.18 .20 67023f14 117
.30 .025 j14200 29023 17117 42019 47020 64019a (a)114119
.30 .05 11422 29101 17119 34306 42021 47022 6 40 21a (a)30 i0 1142(12
.30 .10 14210 29106 17200 34308 42100 47100 64 0 23a (a)
aSee table 24.
50
,iMZ 7
TABLE 24.- MISCELLANEOUS DYNAMIC DATA
Airfoil Frame MW ao al k Remarks
N-0012 8019 0.035 10.0 10.0 0.10 Low Reynolds number, 0.5x106
8021 .035 10.0 10.0 .158023 .035 10.0 10.0 .258104 .035 15.0 10.0 .158106 .035 15.0 14.0 .108116 .07 15.0 10.0 .15 Match reference 3
8118 .07 15.0 10.0 .258123 .07 15.0 14.0 .10 Match reference 38203 .07 10.0 10.0 .25
8210 .11 10.0 10.0 .258222 .18 15.0 10.0 .15 Match reference 38306 .18 15.0 14.0 .10 Match reference 39022 .18 15.0 6.0 .24 Match reference 39101 .18 15.0 5.0 .299106 .18 10.0 10.0 .257108 .30 8.0 5.0 .025 Variable ao7110 8.0 .107111 8.0 .207216 8.8 .057214 8.8 .107212 8.8 .157104 9.0 .0257019 9.0 .057021 9.0 .107101 9.0 .157023 9.0 .20
10.0 See table 177117 11.0 .0257118 11.0 .057119 11.0 .107120 11.0 .15
7121 11.0 .207200 12.0 .0257202 12.0 .057205 12.0 .107305 12.0 .157207 12.0 .20
15.0 See table 1610309 2.8 10.0 .1010305 3.810303 5.09302 10.0
10022 12.09217 .29 15.014220 .29 15.010101 .27 20.010104 .30 12.0 8.0 .05 Match reference 1710105 .30 12.0 8.0 .10 Match reference 1710108 .30 12.0 8.0 .13 Match reference 1715218 .29 15.0 10.0 .10 Pressure orifices closed
51
TABLE 24.- Continued.
Airfoil Frame MW ao a1 k Remarks
N-0012 Many Variable Variable 10.0 0.001 Quasi-static; see table 12W-098 23117 0.30 5.0 10.0 .10Ames-Ol 30201 11.0 5.0 .01Ames-0l 25214 jI .05Ames-O1 25216 .10SC-1095 39110 .01
37219 .0537221 .1037304 12.0 8.0 .05 Match reference 1837305 12.0 8.0 .10 Match reference 1837306 12.0 8.0 .13 Match reference 18
HH-02 43314 11.0 5.0 .025HH-02 43315 11.0 5.0 .05HH-02 43316 11.0 5.0 .10VR-7 54019 .18 10.0 10.0 .025
54022 10.0 .0554101 10.0 .1054 10 10.0 .1554113 10.0 .2054116 10.0 .2549023 15.0 .0149110 .02549117 .0549120 .1058121 .1049203 .1554216 .1557018 .1558018 .1558120 .1549206 .20
NLR-I 65223 .11 7.0 5.0 .025 No separation65300 .11 7.0 5.0 .20 No separation62114 .20 15.0 10.0 .1065207 .20 15.0 10.0 .1062121 .20 10.0 10.0 .17 Match reference 1962202 .20 15.0 5.0 .1762201 .20 15.0 5.0 .2862403 .30 10.0 10.0 .1263100 15.0 5.0 .1263122 12.0 8.0 .1265309 7.0 5.0 .01 No separation65311 7.0 5.0 .20 No separation65121 -2.0 10.0 .01 Stall at negative a65122 .025 Stall at negative a65123 .05 Stall at negative a65200 .10 Stall at negative aNLR-1T 64212 .01 Trip; stall at negative aNLR-IT 64213 .025 Trip; stall at negative a
NLR-lT 64214 .05 Trip; stall at negative a
52
TABLE 24.- Concluded.
Airfoil Frame M. ao al k Remarks
NLR-lT 64215 0.30 -2.0 10.0 0.10 Trip; stall at negative aNLR-IT 64119 .30 2.5 .01 Trip; stall suppressionNLR-lT 64121 .30 2.5 .025 Trip; stall suppressionNLR-lT 64202 .30 2.5 .05 Trip; stall suppressionNLR-lT 64204 .30 2.5 .10 Trip; stall suppressionNLR-7 67201 .11 10.0 .10
67208 .18 10.0 .02567210 .18 10.0 .1067212 .18 10.0 .2067218 .18 15.0 .02567220 .18 15.0 .10
67222 .18 15.0 .2067310 .25 10.0 .1068219 .30 12.0 2.0 .05 No separation68221 .30 12.0 2.0 .10 No separation68304 .30 12.0 2.0 .20 No separation
NLR-7T 67108 .30 10.0 5.0 .025 TripNLR-7T 67110 .30 10.0 5.0 .05 TripNLR-7T 67112 .30 10.0 5.0 .10 Trip
I-i
53
TABLE 25.- TEST CASES FOR NUMERICAL ANALYSIS (ref. 1)
Case Frame Airfoil Co a, k Case Frame Airfoil Co L k
1 10222 NACA 0012 5 5 0.20 7 10212 NACA 0012 10 5 0.202 68211 NLR-7301 5 8 9302 10 10 .103 7111 NACA 0012 8 9 10113 15 5 .014 68203 NLR-7301 10 j 10114 .0255 7023 NACA 0012 9 1 10117 .056 45221 VR-7 10 .025 10118 .10
45223 .05 10120 .1545300 .10 10123 .20
45302 .15 10 45203 VR-7 .0145303 .20 45205 .025
7 10202 NACA 0012 .01 45207 .0510203 .025 45209 .1010204 .05 45211 .1510208 .10 4521.3 .2010211 .15
54
NACA 0012 AMES-01
SIKORSKY SC-1095 NLR-1
VERTOL VR-7 WORTMANN FX-098
HUGHES HH-02 NLR-7301 SUPERCRITICAL
Figure 1.- Airfoils tested in the experiment.
51
55
- __... .. .. .. .,_. . .._-
_ -
AXIS OF ROTATION 'CONNECTED TO DRIVE SYSTEM
PLEXIGLASS WINDOWOUTLIE OFIN TUNNEL CEILING
STAINLESS STEEL SPAR
OUTER MODEL SHELL
I PRESSURE ORIFICES II
I * * * '-HOT WIRE SENSORS III
2mm GROUND PLANE
LOWER BEARING ASSEMBLY
Figure 2.- Model installation in the test section.
56
La 0 9-
uJ ;U cL
CL~
0
ccc
444
m LU> 0
00
57x
* - ---, ~- .
T .*2.11m
FLOWFL0.61m'l TYPICAL CROSS SECTION
WOODEN UPPER SHELLWITH FIBERGLASS SKIN
PRESSURE TRANSDUCERAND HOT WIRE FITTINGS
HOLLOW STEEL .S "AR
WOODEN LOWER SHELLSWITH FIBERGLASS SKIN
Figure 4.- Sketch of the wooden model shells surrounding the steel spar.
58
Figure 5.- Pressure transducer and hot-wire installation: view frominside the upper-surface shell.
y/c N C M A I E
I] xlc00
Figure 6.- Coordinate axes for the airfoils.
59
PLANE MIRROR PULSE CAMERA
COLLIMATING MIRRORAPERTURE
STROBE LIGHT
PLEXIGLASS WINDOW
LIGHT BEAM
7~MODEL
TUNNEL TEST SECTION(SIDE VIEW)
SCREEN ,
Figure 7.- Sketch of the shadowgraph system for visualizing the
leading-edge region.
60
Figure 8.- Representative shadowgraphs before (upper) and during (lower) dynamiic
stall: Sikorsky SC-1095 airfoil, M,,, 0.30, a. 100 + 100 sin wt, k 0.10.
61
18i
1. NACA 0012
16 M ,=0.3016 CLO 0.109±0.004
1. a 0.1' ± 0.20' o
C =1.33 ±0.031.2- MAX
CM0 = -0.005 j0.010
1.0 ~CM =-0.08 ±0.02M~iN
0
CL EXPERIMENTAL
.6- UNCERTAINTY
-. 2 -
-.4~
_.62
1.8 -WORTMANN 098M 0.11
1.6 CL =0.108 t 0.0061.4- CL a~ .o0
CLM = 1 .3' 001.2
12 CM 0 = -0. 2 5 ±O0,010
1.0 -CMM, -0.09 ±0.03
.8-CL
.6
.4
.2
0 ___
-. 2
-.4
CM
a, dog
Figure 10.- Static lift and moment data on the Wortmann FX-098 airfoil
1.8- VR-7
h =0.301.6 = 0.117 ± 0.003
1.4- a0 -1.6'o±0.2'
1.2 LMAX1.51 ±0.03
cMi = -0.015 ±0.00510
CMMIN =-0.10 ±00
.8-
CL .
.4-
.2
-. 2
-. 4
CM
-. 2-10 -5 0 5 10 15 20
a, deg
Figure 11.- Static lift and moment data on the Vertol VR-7 airfoilat M. = 0.30.
64
___________________________________________________________________t.7_
.04 0 STATIC WAKE SURVEYPRESSURE DRAG
a= 5 + 100 sin wt; k = 0.001
.03 0
CD
.02
0
0
0 i ,
-1 -. 5 0 .5 1 1.5CL
Figure 12.- Comparison of measured lift-drag polars for the NACA 0012 airfoilat M., = 0.30, including wind-tunnel-wall corrections.
.12 0 NACA 0012
c SC-1095SOLID SYMBOLS - TRIPPED
0
CLa . . . . . 0.106
0~~~-/ 0-__M1 -__
.10
L L0 .1 .2 .3
Figure 13.- Comparison of lift-curve slopes on the NACA 0012 and SC-1095airfoils, including wind-tunnel-wall corrections.
Sft65
0 03 P
o~0 0,U
a0 0 'n.
o a
0- ......0 0 N - . *
U.au 0 ta 00
a 0 0
44
00
00
4
o (
C4O
C; 0 0
* iiSj
66U
60 70BOEING VERTOL VR-7 RED FRED MACH NO
0 - 0.010 BOEING VERTOL VR-7 - - 0.0760D - 0.025 o - 0.110
- 0.050 60 A - 0.185+ "0.100 + - 0.220
40. x - 0.2500 -0.200o - 0.2110
30
a 20- 40 V=0.
z z,- w 30
101
-3030
0 0-20-
-10-
-50 "-10
0 .4 .6 .8 1.0 0 .. .2 .4 .6 .8 1.0x/C x/c (
(a) Reduced frequency sweep: (b) Mach number sweep:
light stall, deep stall.
Figure 15.- Typical data presentation from volume 3.
0 STATIC
a - 5* + 10° sin wt; k 0.0011.5
1.0 2CL . E1
.5
0
-.5
-10 -6 0 5 10 15 20 25a, dog
(a) CL VS ct.
Figure 16.- Static chracteristics of the NACA 0012 airfoil at M. - 0.30,including wind-tunnel-wall corrections.
67
- .
- - 1
.15 0 PRESSURE DRAG0 WAKE DRAG
.10
CD
.05
0
-. 051 aa
.1 0 STATIC- ot= 5+ 10°sin wt; k 0.001
CM 00
.1 0
-. 2
-. 3 1 , . ,i
-10 -5 0 5 10 15 20 25
(b) CD and CM vs a.
Figure 16.- Concluded.
68
/0.080 + 0.112
/
1.5 - CLMAX =1.451
1.0 in/CL
//
/
-.5 /
0 I I I p
-10 -5 0 5 10 15 20 25a, dog
(a) CL vs a.
Figure 17.- Static characteristics of the Ames A-01 airfoil at M. - 0.30,including wind-tunnel-wall corrections.
69
i _ I Il il l i l l .. .
.15 0 PRESSURE DRAG
0 WAKE SURVEY 0
8.10
CD
.05
0000
.1
0 U 93 0
-. 1 0
-. 2
3 ii i i i
-10 -5 0 5 10 15 20 25a, dog
(b) CD and CM vs a.
Figure 17.- Concluded.
70
//
1.5 CLMAX = 1.444 0.142 + 0.109 c
1.0A 1.44
CL 1/
.5 /
0
-.5 /
-10 -5 0 5 10 15 20 25a dog
(a) CLV .a L vs a.
Figure 18.- Static characteristics of the Wortmann FX-098 airfoil at M. = 0.30,including wind-tunnel-wall corrections.
71
, , , IlL II I
.15 0 PRESSJRE DRAG 50 WAKE SURVEY o
.10 -
0CD
0.05 - 0
131
0 n
-.500
0 0
-. 2
-10 -5 0 5 10 15 20 25
a dog
(b) CD and CM vs a.
Figure 18.- Concluded.
9' 72
/ -0.089 + 0.118 oC
1.5 - 1.529CLMAX =
1.0 8 0
CL P'~/11
C///
-.5I I I 1 I I I
-10 -5 0 5 10 15 20 25a, deg
(a) CL vs (1.
Figure 19.- Static characteristics of the Sikorsky SC-1095 airfoil at M' - 0.30,
including wind-tunnel-wall corrections.
7
73
15 0 PRESSURE DRAG
0 WAKE SURVEY
.10
CD
.05
0
0 0 0
-. 05
.1
0
CM 8-. 1 0
-. 2
-. 3 -I I-10 -F 0 5 10 15 20 25
a, deg
(b) CD and CM vs a.
Figure 19.- Concluded.
74
0.081 + 0.114 a
1.5 CLMAX = 1.417
1.0 - 03
CL
/
.5 /
0/oYA'
I /
I /-. 5
-10 -5 0 5 10 15 20 25cr, deg
(a) CL vs a.
Figure 20.- Static characteristics of the Hughes HH-02 airfoil at M. - 0.30,including wind-tunnel-wall corrections.
75
w o
.15 0 PRESSURE DRAG
0 WAKE SURVEY
B.10
CD
.05
3 0 00 00
0 0 0~ 0 ULJ0
CMU U
-. 05
.I-
00
Cu m
-. 1 01
-. 2
-. 31-10 -5 0 5 10 15 20 25
a, deg
(b) CD and CM vs a.
Figure 20.- Concluded.
76
/ 0.188 +0.117o1.5 -1.516 '
CLMAX=l.SlS
1.0 /1
CL.5 /
/
/
-.5 /II I I , I I I -t
-10 -5 0 5 10 15 20 25a, deg
(a) CL Vs a.
Figure 21.- Static characteristics of the Vertol VR-7 airfoil at Ml, = 0.30,including wind-tunnel-wall corrections.
77
O "
.15 0 PRESSURE DRAG
0 WAKE SURVEY ]
.10 r0
CD
.05
0003
0 -
-. 051
CM EI 10 1 21 0 n rE, r0 0-
CM
0-.1 03
0
.3
-10 -5 0 5 10 15 20 25a, dog
(b) CD and CM VS a.
Figure 21.- Concluded.
78
1.5 CLMAX = 1.302 ,,,,. 0.090+ 0.103 a
41.0 0 0
CL 0
.5
0
-.5 /
-10 -5 0 5 10 15 20 25(Ydeg
(a) CL Vs a.
Figure 22.- Static characteristics of the NLR-l airfoil at M, = 0.30,including wind-tunnel-wall corrections.
79
.15 0 PRESSURE DRAG
0 WAKE SURVEY
.10
CD
.05 o0 0
0 0 o] t 0
-. 05 ' I I
."1
0 nE3800
CM[3[1.3 0
-. 1 0 0
-. 2
-. 3-10 -5 0 5 10 15 20 25
a, dog
(b) CD and CM vs a.
Figure 22.- Concluded.
80
/1
/' 01
/ 01.5 CLMAX 1.821 A A
1.0 , 0.231 + 0.118u
CL //
0
/
/
-. 5
-10 -5 0 5 10 15 20 25Q, dog
(a) CL vs .•
Figure 23.- Static characteristics of the NLR-7301 airfoil at M. - 0.30,4including wind-tunnel-wall corrections.
81
-- lo
.15 0 PRESSURE DRAG
0 WAKE SURVEY
.10
CD
Go OOGM0 o 0 0 00
-. 05 I I I I I I
.1
0
C- 31 0 0 13013 E
-. 2
-. 31-10 -5 0 5 10 15 20 25
a, dog
(b) CD and CM vs ct.
Figure 23.- Concluded.
82
2.0
CLMAX
1.0
.5* PRESENT, NO TRIP
o PRESENT, WITH TRIP
~ REFS. 3, 5, 15-17, 20-24I I I
0 .1 .2 .3 .4 .5 .6M_
Figure 24.- Comparison of maximum static lift on the NACA 0012 airfoil.
2.0 -
1.5 * . -
CLMA x 0 0 0
1.0
Re/M.5 0 PRESENT 14 X 106
* REF. 6 5-9 X 106
FLAGGED SYMBOLS - TRIPPED
., |I I 4.1 .2 .3 .4 .5 .6
Figure 25.- Comparison of maximum static lift on the Ames A-01 airfoil.
83
2.0-
1.5 * *
CLMAX
1.0
Re/M.
* PRESENT 14 X 106
* REF. 8 10 X 106
.5 I-3 REF. 25 11-40 X 106-* REF. 24 13 X 106
o REF. 26 15 X 106FLAGGED SYMBOLS - TRIPPED
I I I I
0 .2 .3 .4 .5 .6Mo
Figure 26.- Comparison of maximum static lift on the Wortmann FX-098 airfoil.
2.0 -
1.5 - 0 0 0
CLMAX
1.0 o 0
Re/Moo
.5 0 PRESENT, ON-LINE ANALYSIS
o PRESENT, NORMAL ANALYSIS< REF. 15 8 X 106
A REF. 16 4-9 X 106I I I I
0 .1 .2 .3 .4 .5 .6Moo
Figure 27.- Comparison of maximum static lift on the Sikorsky SC-1095 airfoil.
84
- Be . -- - - -
2.0
1.5 *
CLMAX .E
1.0
.5 O PRESENT
O PROUTY (HH-01) [27]
FLAGGED SYMBOL - TRIPPEDi p I I I
0 .1 .2 .3 .4 .5 .6M..
Figure 28.- Comparison of maximum static lift on the Hughes HH-02 airfoil.
2.0 I
1.5 * VV.
CLMAX oo,, O
1.0
Re/Mw* PRESENT 14 X 106
.5 I-0 C O U LO M B 128 1 X 106
V DADONE 15] 18 X 106o a o0,0B IN G H A M [29117 X 106
FLAGGED SYMBOLS - TRIPPED
0 .1 .2 .3 .4 .5 .6M..
Figure 29.- Comparison of maximum static lift on the Vertol VR-7 airfoil.
85
2.0
1.5 -
CLMAX
1.0 0 0 0 13
Re/Mo.5- 0 PRESENT 14 X 106
* REF. 19 16 X 106
0 REF. 24 3 X 106
FLAGGED SYMBOLS TRIPPEDII I I
0 .1 .2 .3 .4 .5 .6Moo
Figure 30.- Comparison of maximum static lift on the NLR-l airfoil.
2.0
0 0 0
1.5
CLMAX
1.0
.5 Re/M oo
O PRESENT 14 X 106
* NLR 1301 14× 106
FLAGGED SYMBOLS - TRIPPED
I I I
0 1 .2 .3 .4 .5 .6Moo
Figure 31.- Comparison of maximum static lift on the NLR-7301 airfoil.
t 86iI _ __
0 to
0
(A -
0
a1 - w 4
M + 0a Co4-4 U
00 + 4U +
3 U 4c.U LU '* +
w1 00
e370 CuoJ
-i4
C? C 80 a
00
87t
AN EXPERIMENTAL STUDY OF DYNAMIC STALL 0ON ADVANCED AIRFOIL SECT--ETC(iJULLA 82 v .J NCCAPSKET. K V MCALISCR. L v CARR
ACASIFICO NAAA8924VO-1 NASA-TM-64245-VOL-1 NL2,2 ffflll
2
CLMA
Z:'X. -r
*A''~-
III ~ CC cc
0.
STTI YNMCTALL RE DYNAMIC, DEEP STALL
Figure 33.- Comparison of maximum lift on the eight airfoils at MH,= 0.30.
88
3
2
CLMAX
'I
o PRESENT* McALISTER, et al [31
FLAGGED SYMBOLS - TRIPPED
.1 .2 .3M..
Figure 34.- Comparison of maximum lift on the NACA 0012 airfoil under deep-dynamic-stall conditions: a - 150 + 100 sin wt, k - 0.10.
ao HI2.0
PRESENT - , -
CL
1.0 -------- o~m
0 5 10 15 20 26a, dig
Figure 35.- Comparison of the lift hysteresis on the NACA 0012 airfoil:M. a 0.1, a - 15* + 100 sin wt, k = 0.10.
89
STALL A mONSET o ( SOLID SYMBOLS- REF.17
CLMAp SYM. a1 , deg k
0 5 0.15
v 8 0.10
STATIC & 8 0.13
CL Vs a 0 10 0.10I MEASUREMENT
UNCERTAINTY
0
.4 0
13
00
i CMMN. * V
I ASUREMENT
UNCERTAINTY
II I=15 20 25 3
a MAX, d*9
Figure 36.- Comparison of maximum airloads on the NACA 0012 airfoil atM- 0.30 and cik 2 constant.
90
2 - A
CLMAX . CA -SOLID SYMBOLS - REF.18
SYM. a, k
A S0 0 5 0.15A N STATIC 8 0.10
CL vs( A 8 0.13
E0 10 0.10I MEASUREMENTUNCERTAINTY
0 1 IV 10 15 20 25
a MAX
A
-CMMIN Y
.20
CP MEASUREMENTUNCERTAINTY
1 0 16 20 25*MAX. dog
Figure 37.- Comparison of maximum airloads on the Sikorsky SC-1095 airfoil at- 0.30 and aik 2 a constant.
91
o0SOLID SYMBOLS - REF. 19
(A 0SYM. oil, dog
CLMAX 0 5
1 0 810
< STATIC
I MEASUREMENTUNCERTAINTY
SI p -I
.4
-CMMIN L a
.2 0
IMEASUREMENTUNCERTAINTY
0 .25 .50 .75 1.0
Figure 38.- Comparison of maximum airloads on the NLR-1 airfoil atM. 0.3 and aiax 20.
92
1. Report No.NASA TM 84245 2. Goverment Acusoe No. 3. Flpstsa cawog No.USAAVRADCOH TR-82-A-8 ,p,-llq R 41-_
4. Title &W Subtitle . Report DowAN EXPERIMENTAL STUDY OF DYNAMIC STALL ON ADVANCED July 1982AIRFOIL SECTIONS 6. Peformiig Organization codaVOLUME 1. SUMMARY OF THE EXPERIMENT
7. Authorls) & Performing Orgenization Report No.
W. J. McCroskey, K. W. McAlister, L. W. Carr, A-8924and S. L. Pucci 1o. work Unit No.
9. Perforring Organization Nam.e and Addrs NASA Ames Research Center, K-1585
Moffett Field, Calif. 94035, and U.S. Army Aero- 11. ContractorGrantNo.mechanics Laboratory (AVRADCOM), Ames ResearchCenter, Moffett Field, Calif. 94035 13. Typeof ReporendPeriodCovred
12. Sponsoring Agency Name and Address National Aeronautics and Technical MemorandumSpace Administration, Washington, D.C. 20546, and 14 SponingAgncyCU.S. Army Aviation R&D Command, St. Louis, MO 93166
15. Supplmentary NotesPoint of Contact: W. J. McCroskey, Ames Research Center, MS 202A-1
Moffett Field, Calif. 94035(415) 965-6428 or FTS 448-6428
16. Abstract
The static and dynamic characteristics of seven helicopter sections anda fixed-wing supercritical airfoil were investigated over a wide range ofnominally .-.o-dimensional flow conditions, at Mach numbers up to 0.30 andReynolds numbers up to 4xlO 6. Details of the experiment, estimates of mea-surement accuracy, and test conditions are described in this volume (thefirst of three volumes). Representative results are also presented and com-parisons are made with data from other sources. The complete results forpressure distributions, forces, pitching moments, and boundary-layer separa-tion and reattachment characteristics are available in graphical form involumes 2 and 3.
The results of the experiment show important differences between air-foils, which would otherwise tend to be masked by differences in wind tun-nels, particularly in steady cases. All of the airfoils tested providesignificant advantages over the conventional NACA 0012 profile. In general,however, the parameters of the unsteady motion appear to be more importantthan airfoil shape in determining the dynamic-stall airloads.
17. Key Words (Suggested by Author(s)) 18. Oistnibution SwtementDynamic stall Maximum lift UnlimitedOscillating airfoils Airfoil data
Boundary layermeasurements
Unsteady pressure distributions Subject Category - 0219. Security Oaf. lof ths irport) 20. Security loesaif. (of this pae) 21. No. of Pages 22. Peice
Unclassified Unclassified 103 A06
'For Weby the Notional Technical Information Service, Springfeld, Virginia 22161
