aerodynamic characteristics of three helicopter...

AERODYNAMIC CHARACTERISTICS OF THREE HELICOPTER ROTOR AIRFOIL S-CETC(U)SEP 80 K W NODNAN, 6 J BINGHAM

UNCLASSIFIED MASA-L-13139 NASA-TP-1701 M

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iTechnical Paper 1701 Technical Report 80-B-5

6)kAerodynamic Characteristics of ThreeHelicopter Rotor Airfoil Sections atReynolds Numbers From Model Scale to Full

Scale at Mach Numbers From /0.35 to 0.90.

S LEfSEP 3 0 1980O

19 Kevin W, tNoonan M Gene J./Bin ham-iiit-rei- Latoratory, AVRADCON Research anid Technology Laboratories.

Langley Research Center .Hamptot, Virginia . ,

National Aeronauticsand Space Administration

Scientific and TechnicalInformation Branch

1980

SUMMARY

An investigation has been conducted in the Langley 6- by 28-Inch TransonicTunnel to determine the two-dimensional aerodynamic characteristics of threehelicopter rotor airfoils at Reynolds numbers from typical model scale to fullscale at Mach numbers from about 0.35 to 0.90. The model-scale Reynolds num-bers ranged from about 0.7 x 106 to 1.5 x 106 and the full-scale Reynolds num-bers ranged from about 3.0 x 106 to 6.6 x 106. The airfoils tested were theNACA 0012 (00 Tab), the SC 1095-R8, and the SC 1095. Both the SC 1095 and theSC 1095-R8 airfoils had trailing-edge tabs.

The results of this investigation indicate that Reynolds number effectscan be significant on the maximum normal-force coefficient and all drag-relatedparameters; namely, drag at zero normal force, maximum normal-force-drag ratio,and drag-divergence Mach number. In general, the increments in these parametersat a given Mach number owing to the model-scale to full-scale Reynolds numberchange are different for each of the airfoils.

INTRODUCTION

The development of new rotors generally includes predictions of thefull-scale rotor performance based on extrapolation of small-scale-model rotorwind-tunnel results and based on rotor-performance analysis programs using two-dimensional airfoil data at full-scale Reynolds numbers. Both prediction meth-ods have inherent difficulties. The extrapolation of model-scale data requiresa knowledge of how to increment the airfoil-section data for Reynolds numbereffects, and data for this purpose are seldom available. Rotor-performanceanalysis programs are already quite sophisticated, but they require a detailedknowledge of the rotor flow field. Improvements of these analysis programs canbest be accomplished by correlation studies which require airfoil-section dataat both model-scale and full-scale Reynolds numbers.

In order to provide data for evaluating extrapolation methods and forrotor-performance correlation studies, the two-dimensional aerodynamic charac-teristics of three helicopter rotor airfoils have been determined at Mach num-bers from 0.35 to 0.90 and at Reynolds numbers from model scale to full scale.

The model-scale Reynolds numbers ranged from about 0.7 x 106 to 1.5 x 106, andthe full-scale Reynolds numbers ranged from about 3.0 x 106 to 6.6 x 106. Theairfoils investigated were the NACA 0012 (00 Tab), the SC 1095-R8, and theSC 1095. All three airfoils have a tab, which is a flat platelike extension tothe normal trailing edge of an airfoil, that is usually used to reduce the air-foil pitching-moment coefficient. These airfoils were chosen because some dataon model rotors and full-scale rotors utilizing these airfoils were available.(ref . 1 and 2) or were expected to become available.

SYMBOLS

The units used for thej physical quantities in this paper are given in boththe international Systems-M<Units (SI) and U.S. Customary Units. ThS measure-ments and calculations vere made in U.S. Customary Units.

[c airfoil chord, cm (in.)

cdj section profile-drag coefficient, cd-c

Wake

CA point-drag coefficient,

6/ (tp 2/7.. r 1(pt /1 7 (Pt/P) 2 /7 1/2j

S/7(Pt CP ) -/ L p'1p)/P (t,/.) 1

-cc~ (-4 4~ 2

cd Md section profile-drag coefficient at drag-divergence Mach number

CM section pitching-moment coefficient about quarter-chord,

S[CP (0. 25 - + (p()(x + C [ .25 x)(&-

U.S. L.S.

Cn ~ static-pnrssu rce coefficient, - AE + C(X

h height of vake-survey probe tubes from given reference plane, on (in.)

M Mach number

Mdd Mach number for drag divergence, dcd/d34 0.1

p static pressure, Pa (psi)

2

.. .... ...

q dynamic pressure, Iv 2, Pa (psi)

R Reynolds number based on airfoil chord and free-stream conditions

t airfoil thickness, cm (in.)

v velocity, m/sec (ft/sec)

x airfoil abscissa, cm (in.)

z airfoil ordinate, cm (in.)

zc ordinate of airfoil mean line, cm (in.)

a angle of attack, angle between airfoil chord line and airstreamdirection, deg

ac angle of attack corrected for lift-interference effects, deg

p density, kg/m 3 (slugs/ft3)

Subscripts:

1 local

max maximum

o zero normal force

sep separation

t total

Gfree stream

Abbreviations:

L.S. lower surface

U.S. upper surface

APPARATUS AND METHODS

Airfoils

The airfoil profiles are shown in figure I and the thickness distributionsand mean lines are shown in figure 2. The SC 1095-R8 airfoil was derived fromthe SC 1095 airfoil by the addition of a drooped nosel the mean line of theresultant airfoil forward of about 20-percent chord is drooped relative to that

3

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

of the SC 1095 airfoil, and the thickness of the SC 1095-RS airfoil forward ofabout 15-percent chord is greater than that of the SC 1095 airfoil. Both air-foils have a trailing-edge tab which is about 3-percent chord in length and isdeflected upwards about 30. The maximum thickness of the SC 1095 airfoil astested is 9.1-percent chord and is located at the 25-percent-chord station; themaximum thickness of the SC 1095-RB airfoil as tested is 9.0-percent chord andis also located at the 25-percent-chord station. The maximum camber of theSC 1095 airfoil is 0.8-percent chord; the SC 1095-R8 airfoil has the samepositive camber as the SC 1095, but it has a maximum mean-line ordinate of-1 .6-percent chord. The addition of the 00 Tab to the NACA 0012 airfoil resultsin a reduction of the maximum thiqkness from 12-percent chord to 11.7-percentchord. The design coordinates for these airfoils are given in tables I to II.

The airfoils were machined from heat-treated stainless-steel blocks and hadspans of 15.27 cm (6.010 in.) and chords of about 7.87 am (3.1 in.). The modelshad about 22 orifices located in one chordwise row on each surface; the orificerows were positioned 12.6-percent span on either side of midspan (tables IVto VI). Slots were milled in the airfoil surface and tubes were placed in theslots and then covered with epoxy. The final airfoil contour had a surface fin-ish of 0.813 im (0.000032 in.). The orifices were then drilled from the oppo-site sides of the model so there were no surface irregularities near the orificerow. The orifices had a diameter of 0.0508 am (0.020 in.) and were drilled per-pendicular to the local-surface contour.

Wind Tunnel

Tunnel description.- The Langley 6- by 28-Inch Transonic Tunnel (ref. 3)is a blowdown wind tunnel with a slotted floor and ceiling and is generallyoperated at stagnation pressures from about 207 kPa (30 psia) to 621 kPa(90 psia) and at Mach numbers from 0.35 to 0.90. The selection of the 0.05-open-slot geometry is described in detail in reference 4. Mach number is con-trolled by hydraulically actuated choker doors located downstream of the testsection. The airfoil model spans the 15.27-am (6.010-in.) width of the tunnel(fig. 3) and is rigidly attached by mounting tangs to two circular end plateswhich are driven by a hydraulic actuator to position the airfoil at the desiredangle of attack. A test run usually consists of an angle-of-attack sweep at aconstant Mach number and Reynolds number.

Two-dimensionality of flow.- The results of a previous investigation ofrotorcraft airfoils in the Langley 6- by 28-Inch Transonic Tunnel (ref. 5)have shown that the indicated maximum normal-force coefficient is reduced bytunnel-wall boundary-layer influences. This reduction is characteristic oftwo-dimensional wind tunnels without proper sidewall boundary-layer controland occurs because the tunnel-wall boundary layer is thicker than that of theairfoil; therefore, initial separation begins at the tunnel wall. Efforts areunder way to correct this difficulty, but the solution was not available forthe investigation described in this paper.

Although it is not possible to determine precisely the affected Machnumber range or the loss in maximum normal-force coefficient of the airfoilsreported herein, a comparison of the NACA 0012 data measured in this facility

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with 0.125 open slots (ref. 5) with unpublished data from two other facilitieshas been useful in indicating the magnitude of these losses. The maximmnormal-force coefficients measured in the Langley Low-Turbulence PressureTunnel and the United Technologies Research Center 8 foot tunnel at similarReynolds numbers and at a Mach number of 0.36 are higher than that from theLangley 6- by 28-Inch Transonic Tunnel by about 0.15. The difference betweenthe data from the Langley 6- by 28-Inch Transonic Tunnel and the United Tech-nologies data decreased to 0.10 at a Mach number of about 0.55. Incrementalvalues for other airfoils may vary slightly because of specific configurationinfluences.

An investigation conducted in the Office National d'Etudes et de RecherchesAerospatiales (ONERA) Ri Ch wind tunnel (ref. 6) has shown that the tunnel side-wall boundary layer can affect the normal-force coefficients at all angles ofattack (that is, with either attached or separated boundary layers). In thisinvestigation, the sidewall boundary-layer thickness was varied by applyingsidewall suction upstream of the model while the Mach number and Reynolds num-ber were held constant. Generally an increase in sidewall boundary-layer thick-ness resulted in a decrease in the normal-force coefficient at a given angle ofattack; the trend reversed at Mach numbers greater than 0.85 with a supercriti-cal airfoil.

Apparatus

Wake-survey probe.- A traversing wake-survey probe is cantilevered from onetunnel sidewall to measure the profile drag of the airfoils. The probe verticalsweep rate, which was selected after experimental determination of acceptablelag time in the pressure measurements, was about 2.54 cm/sec (1.00 in/sec).

The probe was located 3.69 chords (based on the 7.87-am (3.10-in.) chordmodel) downstream of the airfoil trailing edge and has a maximum vertical travelof about ±27.9 cm (±11.0 in.) from the tunnel center line (fig. 3). Data areacquired with four total-pressure tubes, which are made of stainless-steeltubing, with a 1.53-mm outside diameter and a 1.02-n inside diameter (0.060 in.by 0.040 in.); the tubes are spaced 0.953 cm (0.375 in.) apart laterally asshown in figure 4.

Instrumentation.- All measurements made during the test program wereobtained with the use of a high-speed, camputer-controlled digital data acqui-sition system and were recorded by a high-speed tape-recording unit (ref. 3).All free-stream conditions were determined from stagnation and static pressures.All airfoil surface pressures and all wake pressures were measured with preci-sion capacitive potenticmeter pressure transducers. The electrical outputsfrom each of these transducers were connected to individual autoranging signalconditioners which have seven available ranges. The output signals fram thefour signal conditioners measuring the wake pressures were filtered with20-Hz low-pass filters before input to the data acquisition system; the rangeof frequencies to be passed was experimentally determined during a previousinvestigation. The geametric angle of attack was determined from the outputof a digital shaft encoder attached to a pinion engaging a rack on one modelsupport end plate.

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Tests and Methods

All of the testing was conducted with smooth model surfaces. Tests weremade at stagnation pressures from 121 kPa (17.5 psia) to 138 kPa (20 psia) toobtain Reynolds numbers typical of model-scale rotors at Mach numbers from about0.35 to 0.90. This range of stagnation pressures was below that normally run inthe Langley 6- by 28-inch Transonic Tunnel, and a tunnel calibration at thesepressures indicated that a new calibration should be used to reduce these data.An additional new calibration was used to reduce data at stagnation pressuresfrom 165 kPa (24 psia) to 193 kPa (28 psia). This range of stagnation pressureswas used to obtain data at Reynolds numbers between model scale and full scale.The full-scale Reynolds number data were obtained by testing at stagnation pres-sures from 531 kPa (77 psia) to 621 kPa (90 psia) at the lowest and highest testMach numbers, respectively. Geometric angles of attack ranged from -40 to 180at increments of 20 at the lower test Mach numbers; this range was decreased atthe higher test Mach numbers.

Section normal-force and pitching-mcnent coefficients were calculated fromthe airfoil surface pressures by a trapezoidal integration of the pressure coef-ficients. The pressure coefficient at the most rearward orifice on each surfacewas applied from that station to the airfoil trailing edge in the integrationbecause the small model size did not allow installation of orifices in the air-foil tab. Each of the pressure coefficients represents the average of five mea-surements obtained in a 1.0-second interval. A form of the equation describedin reference 7 was used to calculate the point-drag coefficients from the mea-sured wake pressures, and a trapezoidal integration of the point-drag coeffi-cients was used to calculate the drag coefficient. The static pressures used inthe wake drag calculation were measured with tunnel sidewall orifices located atthe same longitudinal tunnel station as the tips of the tubes on the wake-surveyprobe. All of the drag coefficients presented in this paper represent the meanof the measurements made with three total-pressure tubes on the wake-surveyprobe in one sweep through a wake. The corrections for lift interference, whichhave been applied to the angles of attack, were obtained from references 4and 8. The basic equations for the correction (see ref. 8) are

ac = a3 + Act

where

-n c 1\10

8 \3.9/k+

ak -K

h

and a is the slot spacing and h is the semiheight of the tunnel. Theslotted-wall boundary-condition coefficient k for the present tunnel config-uration is 0.4211K. A value of 3.5 was selected for the slotted-wall perform-

6

ance coefficient K, based on the data and discussion presented in reference 4.

This substitution results in a correction given by the equation

Aa. = -CnC(O.0800)

where c is in centimeters, a is in degrees, and the constant is in degreesper centimeter.

PRESENTATION OF RESULTS

The results of the investigation have been reduced to coefficient form andare presented in the following table:

Results Airfoil Figure

cn against (1c; cm and cd NACA 0012 (00 Tab) 5against cn SC 1095-R8 6

SC 1095 7

cn,max against M NACA 0012 (00 Tab) 8SC 1095-R8SC 1095

cd,o against M NACA 0012 (00 Tab) 9SC 1095-RBSC 1095

(cn/cd)max against M NACA 0012 (00 Tab) 10SC 1095-RBSC 1095

cn against Mdd NACA 0012 (00 Tab) 11SC 1095-R8SC 1095

CdM against c n NACA 0012 (00 Tab) 12SC 1095-R8SC 1095

Cp against x/c NACA 0012 (00 Tab) 13, 16, 20SC 1095-R8 14, 17SC 1095 15, 18

(lac) sep and (x/c) sep SC 1095-R8 19against M

Cp against M SC 1095-R8 21SC 1095

cd against M SC 1095-R8 22

7

DISCUSS ION

Normal Force

Maximum normal-force coefficient.- The maximum normal-force coeffi-cients of the NhCA 0012 (00 Tab), the SC 1095-RB, and the SC 1095 airfoils havebeen determined from the normal-force curves presented in figures 5(a), 6(a),and 7(a), respectively, and are plotted as a function of Mach number in fig-ure 8. Figure 8 shows that the effect of Reynolds number on Cn,max is differ-ent for the three airfoils. The maximum normal-force coefficient increases withincreasing Reynolds number for the range of Mach numbers presented for both theHACA 0012 (00 Tab) and the SC 1095-R8 airfoils but not for the SC 1095 airfoil.Increases in Reynolds number result in little change or a small reduction inCn,max for the SC 1095 airfoil. In addition, the increment in cn, max result-ing from the same increment in Reynolds number Acn/AR is not the same for theNACA 0012 (00 Tab) and the SC 1095-R8 airfoils. These results are not sur-prising since the flow field on these airfoils at Cn,max generally includesregions of supercritical flow, and it has been shown that the type and magnitudeof the scale effect on Cn,max for the incompressible case do not vary in anyconsistent manner (ref. 9). These trends can be explained by the normal-forcecurves and the pressure distributions (figs. 13 to 18). The increases incn,max of the NACA 0012 (00 Tab) and the SC 1095-R8 airfoils are the result ofeither one or both of the following: (1) Increased suction (more negative Cp)on the upper surface near the leading edge at the same angle of attack (figs. 13and 14); and (2) a delay of the turbulent boundary-layer separation to a higherangle of attack at the higher Reynolds number. (See figs. 5(a), 13, 16, 6(a),14, and 17.) The values of Cn,max for the SC 1095 airfoil generally decreasewith increasing Reynolds number because of the occurrence of substantialboundary-layer separation at a lower angle of attack at the higher Reynolds num-ber (figs. 7(a) and 15). This trend of cn,max is possibly caused by a forwardmovement of a laminar separation bubble with increasing Reynolds number whichwould result in a greater length of turbulent boundary layer at the higherReynolds number. The presence of a laminar separation bubble may be responsi-ble for the related tendency of ac for cn,max to be generally lower at thehigher Reynolds numbers (M 0.54).

At the full-scale Reynolds numbers the values of Cn, max for theSC 1095-R8 airfoil are higher than those for the SC 1095 airfoil at all Machnumbers presented, but at the model-scale Reynolds numbers the values of cn, maxfor the SC 1095-R8 airfoil exceed those for the SC 1095 airfoil only for Machnumbers up to about 0.40 (fig. 8). This trend at the highest Reynolds number isthe expected result of the increased leading-edge camber and greater thicknessin the leading-edge region of the SC 1095-R8 airfoil. Figure 8 also shows thatthe airfoil with the highest values of cn,max, the SC 1095-R8, has the greatestsensitivity to increasing Mach number for M < 0.55. The decrease in cn,maxof the SC 1095-R8 airfoil is the result of either supercritical flow-inducedseparation occurring at a lower angle of attack or a more extensive separation(separation point farther forward) occurring at the same angle of attack withincreasing hach number (figs. 17 and 19). Since the SC 1095-R8 airfoil is morehighly loaded in the leading-edge region than the other two airfoils and theboundary-layer separation reduces the leading-edge suction, the reduction incn,max with Mach number is greater for this airfoil.

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For the range of Mach numbers presented in figure 8, the stall of both theNACA 0012 (00 Tab) and the SC 1095-RS airfoils becomes less abrupt as a resultof increasing Reynolds number from model scale to full scale (figs. 5(a) and6(a)). The stall of the SC 1095 airfoil becomes less abrupt with increasingReynolds number for Mach numbers of about 0.39 and 0.44 but not at the otherMach numbers (fig. 7(a)). The pressure distributions presented in figures 16to 18 indicate that the stall of these airfoils is of the trailing-edge type bythe characteristic loss of pressure recovery (more negative Cp) on the uppersurface near the trailing edge.

Slope.- The slopes of the normal-force curves of both cambered airfoilsincrease slightly with increasing Reynolds number at almost all Mach numberspresented. The slopes of the curves of the NACA 0012 (00 Tab) airfoil areessentially insensitive to Reynolds number changes except at the highest testMach number. Figure 5(a) shows a large and unusual change in the slope of thenormal-force curve of the NACA 0012 (00 Tab) airfoil with increasing Reynoldsnumber at a Mach number of about 0.88. The test runs at this Mach number anda Reynolds number of 2.1 x 106 were made twice and the repeatability of the datawas excellent. The pressure distributions presented in figure 20 for an angleof attack of about -2.00 indicate that the shock position on the upper surfacemoves forward as a result of increasing Reynolds number from 1.5 x 106 to2.1 x 106 and then moves rearward again at the highest Reynolds number. Thisforward shift of the shock position results in substantially more down load(negative lift) on the rear of the airfoil at a Reynolds number of 2.1 x 106,thus giving rise to a steeper slope.

The SC 1095-RB and SC 1095 airfoils both have near-zero normal force atac = 00, although they are cambered airfoils. This near-zero normal forceis the result of both the camber-line geometry and the trailing-edge tab.

Pitching Moment

The pitching-moment coefficient about the aerodynamic center (cm at

cn = 0) is essentially unchanged by increases in Reynolds number at allMach numbers presented for these three airfoils (figs. 5(b), 6(b), and 7(b)).These figures also show that the only effect of Reynolds number is to change the"knee" of the curve; that is, increases in Reynolds number generally move thenose-down break to higher normal-force coefficients for all the airfoils at Machnumbers up to about 0.78. This result is expected where the maximum normal-force coefficients increase with increasing Reynolds number. At a Mach numberof about 0.88, the trend of the pitching moment of the NACA 0012 (00 Tab) air-foil is substantially different at a Reynolds number of 2.1 x 106 than at thelowest and highest Reynolds numbers. The reason for this difference is theshift of the shock position with Reynolds number mentioned previously. Thepitching-moment coefficient about the aerodynamic center (cm at cn - 0) ofthe SC 1095-R8 airfoil displays the most sensitivity to increasing Mach number.Analysis of the pressure distributions (figs. 17 and 18) indicates that theSC 1095-RB airfoil develops supercritical flow on the lower surface at a lowerfree-stream Mach number than the SC 1095 airfoil, and it is the growth of thissupersonic zone with increasing Mach number that causes the greater Mach numbersensitivity (fig. 21).

9

It is interesting to note that the data of figures 6(b) and 7(b) suggestthat both of the cambered airfoils were designed for zero pitching-mament coef-ficient, although some authors have mentioned a pitching-moment level of up to10.021 as acceptable (ref. 10). The zero pitching-moment level of the SC 1095and the SC 1095-RB airfoils is due to the trailing-edge tab. It is also inter-esting to note that the leading-edge modifications made to the SC 1095 airfoilwhich resulted in the SC 1095-RS configuration did not change the pitching-moment coefficient about the aerodynamic center by more than about 0.01 for Machnumbers as high as 0.73 but did increase the values of cn,max significantly.

Drag

Drag at zero normal force.- At a constant Mach number, the effect ofincreasing Reynolds number from model scale to full scale is to reduce cd,ofor both cambered airfoils for all Mach numbers presented and to reduce cd, ofor the NACA 0012 (00 Tab) airfoil for Mach numbers up to about 0.65 (fig. 9).The difference in Cd,o of the NACA 0012 (00 Tab) airfoil at model-scale andfull-scale Reynolds numbers is small at the lowest test Mach number (0.34), andthis difference gradually disappears with increasing Mach number. The incre-mental decrease in cd,o at a given Mach number is generally different for eachof the airfoils; the Acd,O for the SC 1095-R8 airfoil is the largest, and thatfor the NACA 0012 (00 Tab) airfoil, the smallest. These trends are the resultof the well-known reduction in laminar and turbulent boundary-layer skin fric-tion with increasing Reynolds number (ref. 11).

The insensitivity of cd,o of the NACA 0012 (00 Tab) airfoil to Reynoldsnumber is consistent with low-speed data of the NACA 001 2 airfoil (refs. 9and 12). The value of cd,o for the NACA 0012 (00 Tab) airfoil at a Mach num-ber of 0.43 and a Reynolds number of 3.9 x 106 is about 0.0005 higher than thatmeasured for the NACA 001 2 airfoil in the Langley Low-Turbulence Pressure Tunnelat a Mach number of 0.36 and the same Reynolds number (unpublished data). Thedifference in drag level may be due to a higher turbulence level in the Langley6- by 28-Inch Transonic Tunnel at stagnation pressures above about 51 7 kPa(75 psia). Unpublished NACA 0012 airfoil data measured on two different chordmodels at the same Mach number and Reynolds number in the Langley 6- by 28-InchTransonic Tunnel indicate a higher cd,o for the smaller chord model, thusimplying a higher turbulence level at the higher stagnation pressure (552 kPa(80 psia)). Therefore the increments in cd,o from model-scale to full-scale

Reynolds numbers presented in this paper are believed to be smaller than thosewhich would be measured in free air.

Analysis of the pressure distributions indicates that the increase inCd,o with increasing Mach number for each of the airfoils at both model-scaleand full-scale Reynolds numbers is the result of the development of supercriti-cal flow and shock waves.

Maximum normal-force-drag ratio.- The values of (cn/cd)max have beendetermined from the drag data shown in figures 5(c), 6(c), and 7(c) and arepresented as a function of Mach number in figure 10. At a constant Mach number,the values of (cn/cd)max for all the airfoils increase with an increase inReynolds number from model scale to full scale for all Mach numbers presented.

10

,...' --.

• I

The increases in this parameter with Reynolds number are the result of both alower skin-friction drag for a given normal-force coefficient and a delay ofturbulent boundary-layer separation to higher normal-force coefficients at thehigher Reynolds numbers. The increments in (cn/cd)ma x due to Reynolds numberat a given Mach number less than 0.70 are generally quite different for eachof the airfoils. The increases in (cn/cd)max due to Reynolds number are gen-erally small for all the airfoils at Mach numbers above about 0.70 where super-critical flow effects predominate over the viscous effects. This predominancecan be illustrated by showing that the knee of the drag curve is controlled bythe development of supercritical flow and not by separation. By drawing atangent to the drag curve of the NACA 0012 (00 Tab) airfoil at a Mach numberof 0.77 and a Reynolds number of 6.5 x 106 (fig. 5(c)), (cn/cd)max is shownto occur at a normal-force coefficient of about 0. 26. The pressure distribution(fig. 16(j)) corresponding to a higher cn (0.36) indicates supercritical flow

on the upper surface from about 3- to 50-percent chord and no characteristics ofseparation. (Minor separation could exist aft of 93-percent chord.)

Drag divergence.- The drag coefficients at constant values of cn werecross plotted to obtain the drag-divergence Mach numbers. The normal-forcecoefficients corresponding to the drag-divergence Mach numbers at model-scaleand full-scale Reynolds numbers are presented in figure 11. The effect ofReynolds number on Mdd is greatest for normal-force coefficients above about0.7 for all the airfoils. The drag divergence at high normal-force coefficients(cn > 0.7) may be controlled more by shock-induced boundary-layer separationthan by sonic flow (ref. 13) moving aft of the airfoil crest. An analysis ofthe pressure distributions for these airfoils at model-scale Reynolds numberssuggests that this is the case. This explanation for drag divergence would beconsistent with the generally larger effect of Reynolds number on Mdd at thehigh values of cn because boundary-layer thickness would be crucial to dragdivergence resulting from flow separation. The drag-divergence Mach number atzero normal-force coefficient of the NACA 0012 (00 Tab) and the SC 1095 airfoilsis unchanged by the increase in Reynolds number, but that of the SC 1095-R8 air-foil is reduced by about 0.02 because of the increase in Reynolds number. Fig-ure 11 also indicates that the drag-divergence Mach numbers corresponding to allnormal-force coefficients between about -0. 1 and 0.3 of the SC 1095-R8 airfoilare lower at the full-scale Reynolds numbers than at the model-scale Reynoldsnumbers. A study of the curves of cd against M for this range of normal-force coefficients (fig. 22) indicates that the drag coefficients at a Mach num-ber of 0.78 at the full-scale Reynolds numbers would have to be lower for theMdd to be as high as that at the model-scale Reynolds numbers. AlthoughReynolds number has essentially no effect on Mdd of the NACA 0012 (00 Tab)airfoil at normal-force coefficients up to about 0.75, the drag coefficientat Mdd is reduced at the full-scale Reynolds number for normal-force coef-ficients greater than 0.2 (fig. 12). Figure 12 shows that cdMdd decreases

with increasing Reynolds number for almost all normal-force coefficients forboth cambered airfoils.

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ODlNCLUSIOt4S

An investigation has been conducted in the Langley 6- by 28-TInch TransonicTunnel to determine the two-dimensional aerodynamic characteristics of threehelicopter rotor airfoils at Reynolds numbers from model scale to full scale atMach numbers from about 0.35 to 0.90. The airfoils included in this investiga-tion were the NACA 0012 (00 Tab), the SC 1095-R8, and the SC 1095. Analysis ofthe test data has resulted in the following conclusions:

1. The maximum normal-force coefficients of the MACA 001 2 (00 Tab) and theSC 1095-RB airfoils increased with increasing Reynolds number at all Mach num-bers presented. The maximum normal-force coefficients of the SC 1095 airfoilat full-scale Reynolds numbers were about the same or slightly lower than thoseat model-scale Reynolds numbers.

2. The pitching-moment coefficients about the aerodynamic center of thesethree airfoils were essentially unchanged with increases in Reynolds number atall Mach numbers presented.

3. At a constant Mach number, the drag coefficient at zero normal force

cd 0o of both cambered airfoils decreased with increasing Reynolds number forall test Mach numbers. At a constant Mach number up to about 0.65, the valuesOf Cd 'o of the MACA 0012 (00 Tab) airfoil decreased with increasing Reynoldsnumber. The difference in cd,o of the MACA 0012 (00 Tab) airfoil at model-

scale and full-scale Reynolds numbers was small at a Mach number of 0.34. Thisdifference gradually disappeared as the Mach number increased to 0.65.

4. For all test Mach numbers presented, the maximum normal-force-dragratios of these three airfoils at the full-scale Reynolds number were higherthan those at the model- scale Reynolds number.

5. In general, the effect of Reynolds number on drag-divergence Mach numberwas greatest for normal-force coefficients above about 0.7 for these three air-foils. The drag-divergence Mach number at zero normal-force coefficient of theNRCA 0012 (00 Tab) and the SC 1095 airfoils was insensitive to Reynolds numberchanges, but that of the SC 1095-RB airfoil was reduced by about 0.02 becauseof the change from model-scale to full-scale Reynolds number.

Langley Research CenterNational Aeronautics and Space AdministrationHampton, VA 23665July 3, 1980

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REFERENCES

1. Weller, William H.: Experimental Investigation of Effects of Blade TipGeometry on Loads and Performance for an Articulated Rotor System.NASA TP-1303, 1979.

2. Stroub, Robert H.: Full-Scale Wind Tunnel Test of a Modern Helicopter MainRotor - Investigation of Tip Mach Number Effects and Comparisons of FourTip Shapes. Preprint No. 03, Proceedings of the 34th Annual NationalForum, American Helicopter Soc., Inc., May 1978.

3. Ladson, Charles L.: Description and Calibration of the Langley 6- by28-Inch Transonic Tunnel. NASA TN D-8070, 1975.

4. Barnwell, Richard W.: Design and Performance Evaluation of Slotted Wallsfor Two-Dimensional Wind Tunnels. NASA TM-78648, 1978.

5. Noonan, Kevin W.; and Bingham, Gene J.: Two-Dimensional AerodynamicCharacteristics of Several Rotorcraft Airfoils at Mach Numbers From0.35 to 0.90. NASA TM X-73990, 1977.

6. Bernard-Guelle, Rene: Influence of Wind Tunnel Wall Boundary Layers onTwo-Dimensional Transonic Tests. NASA TT F-17,255, 1976.

7. Baals, Donald D.; and Mourhess, Mary J.: Numerical Evaluation of the Wake-Survej Equations for Subsonic Flow Including the Effect of Energy Addi-tion. NACA WR L-5, 1945. (Formerly NACA ARR L5H27.)

8. Davis, Don D., Jr.; and Moore, Dewey: Analytical Study of Blockage- andLift-Interference Corrections for Slotted Tunnels Obtained by the Substi-tution of an Equivalent Homogeneous Boundary for the Discrete Slots.NACA RM L53E07b, 1 953.

9. Loftin, Laurence K., Jr.; and Smith, Hamilton A.: Aerodynamic Charac-teristics of 15 NACA Airfoil Sections at Seven Reynolds Numbers From0.7 x 106 to 9.0 x 106. NACA TN 1945, 1949.

10. Wortmann, F. X.; and Drees, Jan M.: Design of Airfoils for Rotors.CAL/AVLABS Symposium Proceedings: Aerodynamics of Rotary Wing andV/STOL Aircraft, Volume I - Rotor/Propeller Aerodynamics, Rotor Noise,June 18-20, 1969.

11. Schlichting, Hermann (J. Kestin, transl.): Boundary-Layer Theory.Sixth ed. McGraw-Hill Book Co., Inc., c.1968.

12. Abbott, Ira H.; Von Doenhoff, Albert E.; and Stivers, Louis S., Jr.: Sum-mary of Airfoil Data. NACA Rep. 824, 1945. (Supersedes NACA WR L-560.)

13. Bingham, Gene J.: An Analytical Evaluation of Airfoil Sections for Heli-copter Rotor Applications. NASA TN D-7796, 1975.

13

TAL .- DESIGN COORDINATES FOR

[tations and ordinates given in percentL airfoil chordJ

Station Upper surface Lower surface

0 0 0.244 .851 -.851.488 1.193 -1.193.977 1.664 -1.664

1.465 2.015 -2.0152.441 2.553 -2.5533.418 2.972 -2.9724.395 N~3318 -3.3185.859 .1748 -3.7487.324 4.101 -4.1019.766 4.573 -4.573

11.719 4.871 -4.87114.648 5.220 -5.22016.602 5.399 -5.39919.531 5.603 -5.60324.414 5.802 -5.80229.297 5.861 -5.86134.180 5.809 -5.80939.063 5.667 -5.66743.945 5.450 -5.45048.828 5.170 -5.17053.711 4.836 -4.83658.594 4.456 -4.45663.477 4.036 -4.03668.359 3.578 -3.57873.242 3.086 -3.08678.125 2.562 -2.56283.008 2.004 -2.00487.891 1.414 -1.41492.773 .788 -.78897.137 .195 -.195

100.000 .195 -. 195

14

TABLE II.- DESIGN COORDINATES FOR SC 1095-RB AIRFOIL

[tations and ordinates given in

Sain Upper Lower Sain Upper LowerStaion surface surface Sain surface surface

0 -1.621 -1.621 12.939 4.748 -3.693.086 -1.113 -2.129 15.337 4.949 -3.693.230 -. 681 -2.427 17.735 5.103 -3.693.470 -. 182 -2.714 20.132 5.208 -3.693.710 .201 -2.887 22.530 5.275 -3.693.950 .518 -3.002 24.928 5.304 -3.693

1.189 .796 -3.088 27.326 5.304 -3.6931.429 1.055 -3.156 29.724 5.294 -3.6831.669 1.285 -3.204 34.519 5.208 -3.6261.909 1.496 -3.251 39.315 5.064 -3.5302.148 1.707 -3.290 44.111 4.882 -3.3952.388 1.890 -3.328 48.907 4.642 -3.2322.628 2.072 -3.357 53.702 4.364 -3.0312.868 2.235 -3.386 58.498 4.028 -2.8013.108 2.388 -3.415 63.294 3.654 -2.5323.347 2.532 -3.443 68.089 3.232 -2.2353.587 2.657 -3.462 72.885 2.762 -1.9093.827 2.782 -3.482 77.681 2.264 -1.5544.307 3.002 -3.520 82.477 1.726 -1.1894.786 3.194 -3.549 87.272 1.180 -.8065.266 3.367 -3.578 92.068 .614 -.4225.745 3.520 -3.597 94.466 .317 -.2216.225 3.654 -3.606 95.137 .230 -.1736.704 3.779 -3.626 95.521 .192 -.1537.664 4.000 -3.654 96.864 .249 -.067

8.623 4.182 -3.664 98.782 .345 .0389.582 4.335 -3.674 100.000 .403 .08610.541 4.479 -3.683

____________________ ____________________ ____________________ ____________________ ____________________'

TABLE III.- DESIGN COORDINATES FOR SC 1095 AIRFOIL

(Stations and ordinates given in percent airfoil chord]

Station Upper Lower Sain Upper Lowersurface surface Sain surface surface

0 0 0 16.946 5.152 -3.622.242 .668 -.533 19.367 5.258 -3.680.484 .988 -.804 21.788 5.326 -3.709.726 1.259 -1.036 24.208 5.355 -3.728.968 1.501 -1.230 26.629 5.355 -3.728

1.210 1.714 -1.385 29.050 5.345 -3.7181.453 1.908 -1.540 33.892 5.258 -3.6601.695 2.092 -1.666 38.733 5.113 -3.5631.937 2.256 -1.772 43.575 4.929 -3.4282.179 2.411 -1.879 48.417 4.687 -3.2632.421 2.556 -1.975 53.258 4.406 -3.0602.663 2.682 -2.063 58.100 4.067 -2.8282.905 2.808 -2.150 62.942 3.689 -2.5563.389 3.031 -2.305 67.783 3.263 -2.2563.873 3.225 -2.450 72.625 2.789 -1.9274.358 3.399 -2.576 77.467 2.285 -1.5694.842 3.554 -2.692 82.309 1.743 -1.2015.326 3.689 -2.789 87.150 1.191 -.8135.810 3.815 -2.866 91.992 .620 -.4266.778 4.038 -3.002 94.413 .320 -.2237.747 4.222 -3.108 95.091 .232 -.1748.715 4.377 -3.196 95.478 .194 -.1559.683 4.522 -3.273 96.834 .252 -.06812.104 4.793 -3.428 98.770 .349 .03914.525 4.997 -3.544 100.000 .407 .087

16

----------------.......-- -rC.

TABLE IV.- STATIC-PRESSURE ORIFICE LOCATIONS

FOR NACA 0012 (00 Tab) AIRFOIL

(Locations given in percent airfoil chord]

Upper-surface Lower-surfacestation station

0 01.10 1.012.47 2.504.96 5.067.25 7.049.82 9.76

14.67 14.6819.54 19.5524.59 24.6229.52 29.3434.15 34.2839.02 38.9543.90 43.9648.84 48.8053.74 53.7058.52 58.4863.52 63.5068.32 68.4073.17 73.2377.98 78.1983.01 83.1188.00 87.9592.92 92.83

17

TABLE V. - STATIC-PRESSURE ORIFICE LOCATIONS

FOR SC 1095-R8 AIRFOIL

[Locations given in percent airfoil chord]


0 01.12 1.312.06---3.24 3.695.70 6.117.90 8.52

10.40 12.8515.24---20.10 20.0324.88 24.8429.67 29.6234.46 34.4139.29 39.2044.09 44.0048.88 48.8153.67 53.6058.50 58.3963.28 63.2268.07 68.0272.87 72.8077.66 77.5982.49 82.3787.28 87.19

----- ---....... ...18 .4

I

TABLE VI.- STATIC-PRESSURE ORIFICE LOCATIONS

FOR SC 1095 AIRFOIL

[Locations given in percent airfoil chord]


0 01.05 1.202.40 2.334.66 4.637.10 7.059.56 9.56

14.34 14.5119.33 19.3424.21 24.1629.07 29.0533.89 33.8738.77 38.7243.59 43.5848.46 48.4153.30 53.2558.16 58.0863.00 62.9867.85 67.7172.68 72.6677.50 77.5382.29 82.3787.25 87.19

19

NACA 0012 (00 Tab)

S C 1095- R8

S C 1095Figure I.- Airfoil profiles.

20

.20- .

.16 f- _-

SC 1095-R8 NACA 0012 (00 Tab)

tic -- -- -- - - ------ - - -

.08------------------

.~04 SC 109

0-

/ SC 1095-R8

4.0

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0x/C

Figure 2.- Airfoil thickness distributions and mean lines.

21

A 0

0-OD4

0 '44

C9.

.0

-

I~ '4

II 0

4

-4

L0

'4

22

9.398 (3.700)

.953(.375)Typ.

5.080(2.000)

2.540( 0.000)

.635(.250)

T'Figure 4.- Wake-survey probe used in Langley 6- by 28-Inch Transonic Tunnel.

All dimensions are in cm (in.).

23

I , ~~ ~~ -. .,. -w~ - ' .-,.:-, , -. -:i --. - .. ....--...--..

-4

C40 8B

4.4

<3 <s

0 (B G

... ..... . 44

00 0

o0 4S S

44

o 0044

-4

24

.2.-

0 j.

6

E 0M 073 88 -D R 1.5 x1

o. M1 069.

00 06R 1.59

C O~M 63 06.6.n 0b 0647

cr M =0.57 OR 1.4 6

0.54,6-5

0 400M 07 1.4 x1

-- 7 A 1.3- ~ .3

0~0

0 OM =0.39 ?RI.- 1.310

e0 0.368 1~R -. 79 o

-5.6

-. - .64 @. . 2 . 6 . . . . 1.6 8

5.0A

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

.10L R 1. 5 10 6

2.1

.09 04ILL .10 6.9

A .08E M 0 89

.07 1B 0 0'8806. -.8 -.4 0 .4 .8

lit :1Hit:;it

7 w1r pi iii"it: i4 i I M.:T

1: 1 is :1 :

I it 1113 g -,l:llRinii 1:

it i

it iii Millil it 11!'it I: :N li .. ...... ..till I i It "i gi;:f 77M.it H...03 it t.

:imi ;:qz. Hil %tm;I Nit !.i tl ;i;: iziji;j:iim; ii;R 1Mh liliiiii HIM11i i::::::02 1H.11111-11N I I Iti ii 11111 IN !lilii iiiii, H. I It i ::M a ii:; 4. 1 :i!:i il :M l

i M.It i WitW : i; I J! iiiji :1:::!:: 11 NV

.011zit it fill R 1.5 x 10

+0.83 NJ H III!

Cd 0 it 1:1 MW k I:;.

R 1.4 106

m 0. 78 :M=tm: V. :M0.7 ) 7 -8 7

M.N. ttm! :.::.T .. R IA 10'

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

4:i,*OM 0 F3 6 30

+ 1.9

Q m 0.69R 1.3 106

+00 0.69

L

"Allow R 1. 10,ffi 1.6

1h;m 0.63 0 5.5V 0.640 oft N&

... ........ ...... R 1.0 10

OM 0.59' ... .........

0.58 j : 0

wo R 1.0 106.. ..... ...... .... ..... . 4

mvm 0.5C Q

-*M 0 49 R 0. Q x 10,

...... ... .. . ....... ....... . A, 1.30:48 4 4.3

Now. 6..... ......... ...... ...... R 0.8 it 10

O+M 0. 44 .. ... ......... ... .... ... -- -- I + 1.1

0 k it .. 1 0.43 - A 3.91 - - I - I= i R = 0.7 x 106

...... ....... ........ ......... ........ ...... .......... .0 m m 0. 39 1.00 0 % 3.3

0 .Wwm 6R = 0. 7 x 10... ....... . ..... ....... ... ..... ..... 0, Q

m 0.

.9 - b -. 4 2 0 2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8

cn

(c) Section drag coefficients.

Figure S.- Concluded.

26

- ........ ....

-4.4

1< > <9

aj Ai

o m

--- --- 7 7i.. ~- . ...

.... ..... .. . ............. . d

00 0 0

GsD

44.

ig fa Lo

_ _ _ _go

___~~t-, I __ _ _ __ _G

L Li

ON 0

27

.26.

.. ~ ... .5... .

___- O_ R =1.3 x 106

L 0R =1.1 x io0 6.6

Ll, ~ ~ -M 0. 78f .'- SR . 0

7 : 77,

0 0

28. .... ... .... ... .... ..

.10.. .1l - - ... ....... ...... I.

.08 .10

.07 -

.0 06

- -T

0

.03

.029

______01__________ -- ____-___----___

i .n

__:, CT

__--~---~-- -- I - -f . - ON-r-

U)

cj&j

'4.4 0

00

0 40

-J. 080 '

II 4.

_ _ _~14 _ _ _ _

I.-

.0 ' C~ ~ '_ _ _ ~ : ~ L

0< go 0 ~

1 i' t. 44

0

It 6.00

0 0 R 1.6 x 106WK 6.4

0t0

I O R 1. 1~ o

R~ 10.4 1

jo. .......~ 6.4 - 1

0 .... .... .. 0 .

.. -....6 ... . 4 ....... 2 . ... . 6 .10 .24. 80C . . .. .. .... .... .. ...

131

if~~ I 6.0

.05,

.0 . W 087..

-.8 -. 1 A~ .58

.6.4

.6.1

A..

Cd

OOM 0.73

0I

-8 -6OR = 6 8 1 1 1. 1.8

DOM .63, j- 7 7. 5.

(C eto rg ofiins

Fiur 1.- Conlued

32m 9 .

0 .58 ...................... ......1

'A7S:~v-. 1:7

00

oo

.'.4

M'D -00

00~

Ix2. ... ........ I

uE

.I .. ....

-A 4t4oO*~ 9 .~t~e.J*OaO*'*

.~~~* . . . .

.03 RModel scale 0.7 - 1.5 x 16

0 0.9 - 1.6Fu tIu &-#-- 2.9 - 6.6....

.. .. .. . .. . .. .. . .. .. . . .. ... .::..I.

.0 1 ... . . . ... .

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

.. ~p ~:2 j~SC 1095

03 Modelscale 0 50 6

... 0.9J7 0 1.9-3.2FullI scale 3.0 -6.6

.01

.3 4... . . 8.

343

r-i

0 0

win

00 00

q

-

4 J40

~44

I;IV

V44

- - t -- -

F 0D

3 35

-w~~ 0 a -.,4 44I

4 OLa-- -----

..... ...EI ...I4 GLO

i~U.'M E-H-HE

0

CE-.F.l....... < i 1**. ....... .. ....

A

-44

0eel.

36

.3Model scale --0.7 R . x 10U 6 ______

FullI scale -3.8 - 6.6

dMd

K 't-oU S C 1095RB_______

Model scale --0.8 1.4 x 106.03.42.

-Full scale - - 3.9 - 6.5

02cd

.L1............. .4

01

2oe 0 cl 2. .4 -.6 .8 1.0.

Figure 12. Efec ofRyod2.0rnscinda oefceta rgdivergencel Machnum e of4. -AC 002(0Ta).c5 0 ,an c19

airfoils

027

46.0 --- 14 C 1 a a C 9 iP'SOnCC C fi

.38 -4.12 0 10.2 9.7 .95 .39 .7K-5.6 :38 -3.88 0 10.7 10.0 1.0l8 .38 3.3

-5.2

-4.8 - - - - - - -

-4.84-- -

-4.0 - - --

-3.2 - - - -

C P-2.4 -- -

-2.0

-1.6---- --

-.4-

0-

.4 -- _ _

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0X/c

(a) ac 100.

Figure 13.- Effect of Reynolds nuber on chordwise pressure distributionof NACA 0012 (00 Tab) airfoil. Symbols with plus sign inside indicatelower surface. R given in millions.

38

M Aflc C a.38 -4 12 0 12.2 11.7 .80 .3S .7

-6.0 .39 -3-88 0] 12.3 11.7 1.03 .38 3.3

-5.6

-5.2

-4.8

-4.4

-4.0 --

-$.6

-3.2

-2.8

.2.4 4--- ------

-2.0

-1.6 - -

-1.2 -

00-.4- -t

oD

* .3m 3333

.4--

.8 -- - - - - - - - - -

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0

X/C

(b) ac = 11.70.

Figure 13.- Concluded.

39

-48M C. a a C M 3

.38 -4.12 0 10.2 9.6 1.05 .39 .8-4.4 ------ - .33 -3.88~ 0 10.3 9.6 1.10 .38 3.5

-4.0 - -

-3.6

-3.2

-2.8-

-2.4-- --

-2.0---------- - - - - - - -

-1.2--

-. 8 - - -

00

.4 - -

.8 - - - - -

1.2 --------------- __0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0x/c

(a) ac =9.60.

Figure 14.- Effect of Reynolds number on chordwise pressure distribution ofSC 1095-R8 airfoil. Symbols with plus sign inside indicate lowier surface.R given in millions.

40

. . . . .. .. . . ..... ....A.-

-5. M C. a a e 3

.38 -4.12 0 12.3 11.8 .82 .39 .8

-5.2 .39 -3.88 .0 12 11.5 1.28 38 3.5

-4.8+

-4.4--------

-3.6

-3.2-

-2.8-- - - - - -

P --

-2.0-

-1.6--

-1.2- - - -- - -8-0-0 -o - - -- ---

-. - 0 0

0, -40

.4 - - -

.8~---02 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

K/c

(b) etc 120.

Figure 14.-, Concluded.41

P-.onic c ac % M.38 -4.12 0 10.2 9.6 .39 .39 .7

44 .39 -3.88 0 10.3 9.6 1.04 .38 3.4

-4.0 I

-3.6

-3.2

-2.8-

-2.4 -

-2.0-

CP -

p \

-1.2 ---

-. 8 - - -

0 7 0

1.2 - - - - - - - - - - -

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0I/C

(a) ac = 9.60.

Figure 15.- Effect of Reynolds number on chordwise pressure distribution ofSC 1095 airfoil. Symbols with plus sign inside indicate lower surface.R given in millions.

42

m Ca a C m-4.8 M C C a c 4

.38 -4.12 0 12.3 11.7 1.05 .39 .74.4 .39 -3.88 0 12.2 11.6 .95 .38 3.4

-4.0 ]-

-3.6

-3.2

-2.8

-2.4 --

-2.0

-1.6 -

-1.2

.8N

-.4 - -- - -- -0

00

.4 .. .. - - - -_| -l - -0 =- =-' -= - - -

.8

1.20 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

X/c

(b) OIC 120.


43

,' • .. ,. J--> ."

pAic C.34 -5.29 K0 0.0 0.0 .01 .34

444 2.1 2.0 .24 .34

L, 6.3 5.9 .68 .340 8.5 7.9 .85 .34

4.0 + 0 10.3 9.7 1.06 .344.0 12.2 11.5 1.09 .34

- -- - - 7 - 1q.4 13.7 1.13 .34

-8.6 -

-.2 i ~ - -

-2.4

-.0 E3 . 3 . 5 .6 . 8 . .

A/

(a M A\ 0.42 .9x16

444

.4 -

.8,-

4. a~ a a

442.1 2.0 .25 .38

- - -~ 6.q 5.9 .71 .3808.2 7.6 .90 .38

-4.00 10.7 10.0 1.08 .380 12.3 11.7 1.09 .38

- - - - - - __0_14.5 13.8 1.fl7 .38_

-3.2-5

-2.8 - C

- - - ~-4 - - - - --

-2.4

-1.6-

0-1.2 - - -

-.4 -~=

.8

x/c

(b) M4 - 0. 38; R -,3.3 x 106.

Figure 16.- Continued.

45

jx*

4.8 a 6PAO- a

.43 -3.09 ' 0.0 0.0 .01 .44-4.4 .41 -2.9-33 2.1 2.0 .26 .43

- ~ 4.1 3.8 .49 .43b 6.1 5.6 .70 .43a 8.1 7.5 .30 .439 10.2 9.5 1.01 .434.0 0 12.2 11.5 1.03 .43

~ 4.3 13.6 1.03 .43

-8.2 - - - - - - -

-2.8-

-2.4 --p

-2.0 - - - -

C2p

08 . 2 . 4 . 6 . 8 . .

46.

4.8 ca a cPicnic e n

.48 -2.37 0 0.0 0.0 .02 .48-4.4 2.0 1.8 .24 .48

08.2 '7.7 .32 .484.0, Q-- -- - 10.2 3.6 .36 .48

-. 26

-.82

-2.4 - - -

-1.6-

*.8 ~ -~

.. 0

.4 ~ -

.8

10 .1 . .3 .4 . .6 .7 .8 .9 1.0K/c

(d) N - 0. 481 R ft4. 3 x 106.


47

...........

-4.8 m a cx M

.54 -1.74 -o -. 1 0.00 .54-. 4 - - - - - - - - 2.1 1.3 .27 .54

S4.1 3.8 .51 .54- - -- - - -~6.5 6.0 .78 .54

0 8.2 7.6 .89 .54-4.0 -- - -- - -- 1-- 0 0.2 9.7 .92 .54

S12.3 11.7 .32 .54

-8.2

-2.8

-2.4-

-2.0-

cp-1.6

-1.2

0 1 . .. 48 . 7 .8 . .

-. 4x/c

(e) M * .54; R 49 X 10g

Figur 1.- Cnnud

48 ~ ~ . T---_ _ _ _ - ~I

4.8 m Qa a c.58 -1.'43 '0 0.0 0.0 .01 .58

-4.4 -- - ~ - - - - - 2.0 1.9 .27 .58q .2~ 3.9 .53 .58

10 8.1 7.5 .89 .58-4.0 - - - - - - - - -0 10.3 9.7 .90 .58

1 1 12.5 11.9 .88 .58

-2.8

-2.4- - - -

-2.0

-1.6

-1.2

-. 44

0 1 2 . .4- M6 .7 .&- ., 1.

(f)W 14w = 0.58;; R- = 5. 16

.49

4.8 m a a c MPlaonic c a.63 -1.12 o m m .01 .644.4 .64 -1.06 ,_ 2.1 1.9 .28 .64

r, 4.1 3.7 .54 .64b3 6.2 5.7 .79 .640 8.2 7.6 .84 .64

4.0 -0 10.2 9.7 .88 .630 12.5 11.9 .88 .63

-8.6

-8.2

-2.8

-2.4

-2.0 ,

cp

-1.2

-. 8 .7 4. .4 0-. A --- 0

Q. . --- O--- --

.4

.8

1/c

(g) M - 0.64; R " 5.5 x 106.


50

.68 -. 87 0 0.0 0.0 .01 .6844 -- - - - - - .63 -. 82 A 2.1 1.9 .30 .63

-4. 4.3 4.0 .61 .63r~6.4 5.9 .77 .630 8.3 7.8 .81 .68

-400 10.2 .7 2.87 .68

-2.8 - - - - - - - - -

-2.4 - - - - - - - - - - - - - - - - - - -

-2.0 -

-1.6

-1.1

4.8 Ii C. a a

.73 -66 o 0.0 0.0 0.00 .734. 4.3 3.9 .62 .73

S6.3 5.8 .71 .73------------------------ '~8.2 7.7 .73 .73

-400 10.4 3.5 .86 .73

-2.8 - - - - - - - - - - - - - - - - - -

-2.4 - - - - - - - - - - - - - - - - - -

-2.0 - - - - - - - - - - - - - - - - - - - -

-1.62

-.4

-.4

1.2

10 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0x/c

Mi M - 0.73; R -6. 3 x 106.


52 -

-4.8 a a c 1

.7 -. 5 c 0.0 -1 .01 .7*7-4.4 ._4 0 1.5 .36 .78

4~L.0 3.7 .56 .77rl 6.1 5.7 .68 .770 8.2 7.7 .76 .77

-4.0- - -

-3.6 - - - - - - - - - - - - - - - - - - -

-3.2-------

-2.8----------------

-2.4---------------

-2.0----------------

cp-1.6------------------

-1.2

0

.4

~0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0K/c

Mj - 0. 77; R -6.5 x 106.


53

.........

-3. 4-239 q 8

-1.6

-2.2

cx/c

__A_____________

-. a a CP'sonic C nl

.88 -. 23 0 -4.0 -3.3 -. 20 .884.411 2.1 -2.1 -. 04 .88

0 0.0 0.0 0.00 .88

-. 6---

-3.2 - - - - - - - - - - - - - - - - - - - -

-2.48

-2.0~

p . ./c

(1.1.086R . x16

Fiue16-.ocldd

.54

- C,

(1) M -. ....88. 0..

4.8 a a cP..a C n

-- 33 -5.65 0 -4.1 -3.8 -. 43 .3444+-- .34_-5.23 0 -2.2 -2.1 -. 27 .34

-4. 0 0.0 0.0 -. 02 .34S2.1 2.0 .20 .341~4.1 3.8 .42 .34

D, 6.1 5.7 .64 .33-4.0- -- - -- - 8.4 7.8 .87 .34

IQ- 10.2 5.5 1.05 .341 10 -- -~ 12.3 11.5 1.27 .33

0 14.3 13.4 1.33 .34-8.6 L--- -b- b1.5 /. 1.25 .34

1- -- _ 018.5 17.8 1.04 .3q

-8.2-

-2.8

-2.4 -C

-2.0 -

-1.6

-1.2 x C

-.8

-.4

01.2 .2 .3 .4 .5 .6 .7 .8 .9 1.0K/c

(a) M -0. 33; R m3.0 x 106.

Figure 17.- Effect of angle of attack on chordwise pressure distribution of

SC 1095-RB airfoil. Symbols with plus sign inside indicate lowier surface.

56

-48 a a c IP"Utlc C

.38 -q. 12 0 -4.1 -3.8 -. 47 .38.40 -2.1 -1.9 -. 26 .38

- 0.0 0.0 -. 03 .38- ---- A 2.1 1.5 .21 .38

N~ 4.1 3.8 q43 .384.bD 6.4 5.9 .69 .38

-4.0- --- 0 8.3 7.7 .90 .38- 0 10.3 9.6 1.10 .38

0 12.3 11.5 1.28 .38-86-- 14.4 13.6 1.29 .38--6o 16.4 15.6 1.31 .38

6 1-~ -0 18.5 17.7 1.21 .38

-3.2-

-2.8-4

-2.4 - __ ---p

-2.0-

p -1.6

0 /c

(b)2 W X I .8 35xi

Fiur 17. IoInud

-. 87

pionk c.43-305 0 -4.5 -4.2 -. 48 .43

-4 0 - .- 44 -- 3 1 -2.2 -2.0 -. 27 .43c~0.0 0.0 -. 03 .43- --- -- A 2.4 2.2 .25 .43

L 4.1 3.8 .45 .44-4.0 - - - - - - - - - - 6.4 6.0 .71 .4i3

08.3 7.8 .92 .43--- ------ 0 10.3 9.6 1.10 .43

-8.2

-2.8

-2.4

-2.0

-1.6b

-.8

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0I/c

(c) M o 0. 43; R -3.9 x 106.


58

4. - a a cpAoflAc n

.48 -2.37 C0 -4. 1 -3.8 -. 46 .48

-4.4----- ---~2.1 -1.3 -. 25 .480 0.0 0.0 -. 02 .48A 2.1 2.0 .23 .48L. 4.1 3.8 .46 .48

O 8.1 7.6 .92 .48- -- -- - -- - 10.2 9.5 1.05 .48

0 12.3 11.6 1.11 .48-3.6 - 14.3 13.6 1.09 .48o 16.4 15.7 1.08 .48

-3.2

-2.8

-2.4

-2.0- - - - - - - - - - - - - - - -

c p

0 1 . .14 . . 7 .8 . .

-. 8c

(d)~z~s? M 0.41R . =16

Figure ~ ~ F 17-Cniud

.49

r73- J 14_4_

0A

[ ~4. 8-8 a a cp~1

*53 -1.83 0 -4.3 -41.0 -. 47 .534.4 .54 -1.74 0~ -2.1 -1.9 -. 26 .54

J 0.0 0.0 -. 03 .54L 4.2 3.3 .49 .5q

4.0 6.2 5.7 .73 .540 8.6 8.0 .94 S54

- -- ~ - 0.2 9.6 1.00 .54

-. 2- liii-

-2.0

-1.2

-.8

.4 - _

00

.4

0.o .1 . .8 .4 .5 .6 .7 .8 .9 1.0X/c

(e) Mb* 0. 54; Rbw 4.8 x 106.


60

-4.8IIn-143 0 -4.2 -3.3 -. 45 .58

-4.4 - - - -2.0 -1.3 -. 25 .580 0.0 0.0 -. 02 .58

2.1 1.3 .25 .58b, 4.1 3.7 .50 .58-4.0 11ra 6.2 5.-7 .74 .58o 8.1 7.5 .88 .58Q 10.2 3.6 .34 .588 12.3 11.6 .37 .58

-8.2

-2.8 ---

-2.4--

-2.0

cp

A4 ' °=" "I - ,I"" '-

-1.6 114

-1.2 A

0, .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0

z/.

0I(f M 0.58; R 5.2 x 106.


61

Ma"

-4.8 m c a a c mpionic c n

.63 -1.12 0 -4.2 -3.3 -. 48 .634.4 -- -- --- 2~- .0 -1.8 -. 26 .63

00.0 0.0 -. 02 .63A-- - - - - 2.1 1.3 .26 .63

S4.0 3.7 .51 .634.0 - - - - - - -~6.2 5.7 .74 .63

0 8.3 7.7 .88 .63------- 1-0 0.2 3.6 .93 .63

-3. 12.4 11.8 .96 .63 1

-8.6-

-2.8 -- - - - - - - - - - - - - - - - -

-2.4-

-2.0 1-

c8

-.6

-.8

1.0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0I I/C

(g) M - 0.63; R -5.5 x 106.


62 ____

4. c a ai c mpmoflic n

.67 -.31 0 -4.1 -3.8 -. 48 .674.4 .68 -. 87 0 -2.0 -1.8 -. 26 .68

0 0.0 0.0 -. 02 .68

r. 4.2 3.3 .57 .68-4.0 -b- - 6.4 5.9 .76 .68

0 8.2 7.7 .84 .68-- 0--- - - 10.2 3.7 .30 .68

-3. 12.4 11.8 .94 .68

-3.2 - -- - - -- -- -

-2.8 - - - - - - - - - - - - - - - - - - -

-2. ---

-1.6

-. 2

*0 . 2 N48 . 7 .8 . .

1/c 0

Fig2re 17-Cniud

63J

4. m a a cpAonic c n

.72 -. 70 0 -4.1 -3.8 -. 51 .7344.73 -. 66 __0 -2.0 -1.3 -. 28 .72

-4.4 0.0 0.0 -. 01 .72A 2.1 1.3 .30 .73S4.1 3.7 .60 .73

-4. 6.0 5.6 .72 .7340- 8.1 7.6 .82 .73

0 10.2 3.7 .8~3 .73

-3.2 - - - - - - - - - - - - - - - - - - -

-2.8---

-2.0 .

Cp-

-1.2

-.2

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0Kilo

(i) M - 0.73; R - 6.3 x 106.


64

-4.8 m ---- g--c a a c mP.Mnk c

- 78 -. 4 0 -4.2 -3.8 -. 54 .78-44---------------2.1 -1.8 -. 38 .78

0 0.0 0.0 -. 01 .78A 2.0 1.8 -34 .78L~ 4.1 3.7 .53 .78

-4.0 6-- - - - - ' .0 5.6 .'70 .780 8.3 7.8 .81 .78

-8.2 - - - - - - -

-2.8 - - - - - - - - - - - - - - - - - - -

-2.4 - - - - - - - - - - - - - - - - - -

-. 0

-1./2

Fiur 17. C-nied

L-65

-4 8 - - - - a a~ c P

.83 -. 35 11 -2.2 -1.9 -. 48 .834.4 - - - - - - 0.0 0.0 0.00 .83

S2.2 1.9 .43 .83

-4.0-

-3.2

-2.4 - - - - - - - - - - - - - - - - - - -

-2.0 - - - - - - - - - - - - - - - - - -

p-1. ---

-1.2

-.8 - A

120 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0x/c

Wk M = 0.83; R -6.6 x 106.


66j

-4.8-- M1 CT a a cp~sonic n

-34 -5.23 0 -4.2 -3.9 -. 43 .34.40 ----- - 0-2.4 -2.2 -. 24 .34-44) 0.0 0.0 0.00 .34

S2.1 2.0 .23 .34S4.1 3.8 .44 .34

-4. 6.1 5.7 .65 .34-40 -- -- -- .- ~8.1 7.6 .84 .34

0 10.2 9.7 .99 .341012.4 11.7 1.02 .34

-3.6 I _ q142 13.7 .81 .34

-2.8 /

-2.4-

-1.2

-.

-4 /Ij -.-

0 /' -

.8

K/c

(a) M = 0. 34; R -2. 9 x 106.

Figure 18.- Effect of angle of attack on chordwise pressure distribution ofSC 1095 airfoil. Symbols with plus sign inside indicate lower surface.

67

...... .. ..I

4.8 a a cp.103w c

.38 -4.12 0 -4.3 -4.0 -. 48 .3844 -- - -. 39 -3.88 0 -2.1 -2.0 -. 241 .38

-4.4 - K 0.0 0.0 0.00 .38& 2.0 1.9 .23 .39L 4.2 3.9 -47 .38

4. -- 8.4 7.9 .92 .38

0 10.3 9.6 1.04 .380 12.2 11.6 .95 .38

-8.6 -- - 1.4 13.9 .80 .38

-3.2

-2.8

-.4

c/

(b M-18.63. 16

688

X/AC

pnc a a l

.'43 -3.09 0 -4. -3.8 -. 47 .43

40 61 . 8 .4 . 8 .2 5.8 .0 .3

(C). M. .9 O.3 433. 16

Fiur 18.- Conti.0ued.

c p

mca a 'apiOniC

- --. 48 -2.37 0 -q4.3 -4.0 -. 51 .484.4 _ 433A5 -2.2 -2.1 - .27 .48

o~ 0.0 0.0 0.00 .49A -- ' 2.1 1.9 .25 .48

~. 4.0 3.7 .47 .q8-4.0 - - - - n 6.1 5.6 .71 .48

0 8.2 7.6 .93 .48----------- 10.2 9.6 1.01 .48

S12.3 11.7 .90 .48

-8.2

-2.8 - - - - -- - ---

-2.4- - - - - - - -

-2.0

op

-. 2

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0K/C

(d) M -0. 48; R ,4. 3 x 106.


70

M poi c ca

.54 -1 .741 0 -. 1 -.3.8 -. 50 .54

0 0.0 0.0 .01 .54A 2.1 1.9 .25 .54~. 4.0 3.-7 .49 .514

0 8.6 8.0 .92 .54- - - - Q- - 10.7 10.1 .93 .54

S12.2 11.7 .87 .54-8.6 - - - - - - - - - - - - - - - - -

-. 2--

-2.4

-. 0

cp-1.6

Z.4 -0 -0 -0--A

0 ~ ~ ~ ~ ~ ~ ~ -.1 .8 . 5 .6 . 8 . .-. 4c

(e)H- .5; Rink" 06

.i4e1. Cniud

A__________ 71

Paoinc CA

+ .58 -1.43 0 -4.1 -3.8 -. 52 .58

A4-- 2.1 1.3 .26 .584. 4.2 3.8 .53 .58

D 6.2 5.7 .75 .580 8.1 7.5 .8? -.58

-4.0 10...- 1 .2 3.6 .92 .58-40 1 12.3 t11.7 .96 .58

-3.2 -

-2.8--

-2.4

-2.0

-1. I - -----

(f M -05; R-51x16

72.

4. ~~'~-- - PAiac "6 ci

.63 -1.12 0 -. 0 -3.7 -. 52 .634 - .-. 611 -1.06 _1 -2.1 -1.9 -. 27 .63

-44 0.0 0.0 0.00 .63- - - - - ---- & 2.0 1.8 .28 .63

L. 4.0 3.7 .53 .634.0 ti ----- ! 6.6 6.1 .79 .63

Q 8.2 7.7 .85 .64- -- -- -- Q0 10.4 9.8 .88 .64

-861 12.3 11.8 .30 .63

-2.8 -- - -

-2.4 -- - - --

-2.0 ---

-1.2

I-I

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0I/C

(g) M 0. 63; R 5 . 4 x 106.


73

.68 -. 87 0 -4.2 -3.8 -. 56 .69-4.4 - - - - - - - - .65 -. 82 13 -2.1 -1.3 -.2-3 .69

Z0 0.0 0.0 .01 .68

0 8.2 7.7 .84 .68

-8.2 - - - - - - - - - - - - - - - - - - -

-2.8 - - - - - - - - - - - - - - - - - - - -

-2.4 - - - - - - - - - - - - - - - - - - -

-2.0

c p-1,6 m-\1

-1.2

-.4__ _ _

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0K/c

(h) M -0.69; Ro 5.8 x 106.


742

.72 -. 0 0-4-40-6 7-4-_ 73 -. 66 0 -2.2 -2.0 -. 32 .73.74 -. 63 00.0 0.0 .01 .74

-A 2.0 1.8 .30 .73L- 4.1 3.7 .62 .73

-4. - 6.1 5.6 .74 .73108.3 7.7 .83 .72

0 10.2 3.7 .8 .72

-3.2 - - - - - - - - - - - - - - - - - -

-2. 0- - -

op

-1.2 - - - -

-40 1

-.4

0I

.41

0 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0x/c

Mi M w0.73; R -6.1 x 106.


75

4.8 m ca a c m

.77 - :53 0 -q. -3.8 -. 65 .78-4.4 - '_ 78 - "4 0 -2.3 -2.0 -. 40 .78

75 -~ -46 0.0 0.0 .02 .79S1.9 1.7 .37 .78L4.0 3.6 .61 .78

4.0 - ----- - 6.1 5.6 .71 .7708.2 7.7 .82 .7

-3.2--

-2.8

-2. - - - - -----

-2.0-

-.2

-.4 o -

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.01/C

dj M4 - 0. 78; R -6.4 x 10o6.

Figure IS.- Continued.

76

.4- ~l

-4.8 m -- a a c-

.82 -. 38 0 -4.1 -3.7 -. 62 .83-4. .83 -. 35 01 -2.2 -1.9 -. 46 .8244 - - - - .84 -. 33 O0 0.0 0.0 .04 .83

A 2.1 1.9 .42 .84I~4.0 3.-7 .54 .83

-4.0-------------------i 6.2 5.8 .67 .83

-3.2---------------

-2.0 - - - - - - - - - - - - - - - - - -

cp-1.6

-1.2 - -

-.4

1.8

0.2 .1 .2 .8 .4 .5 .6 .7 .8 .9 1.0

z/C

(k) M ft0. 83; R * 6.6 x 106.


77

- - .. A- IA

-4.8 a 6 c.87 -. 25 0 -4.5 -4.2 -. 55 .88

-4.4 - 1~! -2.0 -1.8 -. 41 .87c~0.0 0.0 .06 .8-7

-4.0

-8.2

-. 8 2---- --

-2.4--- ---

-2. -----

Up

13- --

-q -C -4- -0 -E - -

.4

1 .2 r - ~ 7 - -. 4 7- - - - - - - -

.1.2 8 .4 . .6 .7 .8 .9 1.0z/c

(1) M 0.87; R -6.7 x 106.


78

me"-

16 R ~3.0-4.8 x 106

12 '

sep11 9

8

4

10R 3. 0 -4. 8 x 106

.80

(xICsep .6

0

.4

.2 ac 11.5-11.7.2.3 .4 .5 .6 .7

M

Figure 19.- Effect of Mach nuber on angle of attack at which boundary-layerseparation first occurs and on separation point of SC 1 095-R8 airfoil.

79

Who"'

-4.8 - -_ _ _ _ _ _ _ _ _ _ _ _ _- -I-M R a.deg a c C~ n , oi-440 0.88 1.5 -2.0 -2.0 -0.06 -0.23

0 0.88 2. 1 -2. 1 -1.9 -0.200 .88 6.9 -2. 1 -2. 1 -0.04

-3.6--3.2 ~--~i1

-2.8 -- - - - - - _ -

-2.4---------------

-2.0__

op-1.6_

-1.2 _

800

.4 ,.

-3.6 iI

-3.2_

Cp. sonic SC 1095-R8 SC 1095

-2.8 xIc x/c- 00.0131 A 0. 0120

- 0.0369 L 0. 046300.0852 & O. 0956

-2.4

I 0

-2.0 0 0Cp o 00

L0 _

-- lie -IV,- -0 -

-. 2 .- 4 5 .1.0

\

- - - - - - -

0 - - - - - *'- _

4% C- 0.0 R 2.9 - 6.6 x i06

2 .3 .4 .5 .6 .7 .8 .9 1.0M

- zffect of Mach number on some lower-surface pressure coefficients*,, yI )f 10-percent-chord station of SC 1095-RB and SC 1095 airfoils.

81

LMU I_ _

.06

05..............>..7

P., T...

.04

Cd" -- 4-- ..1.............

*..... ......

C ~ 1 00 526.

c 03 5.4-6.6 Ful scale".. *-

3 ... 5. .6 .... .dM

........ ..

1. Report No. 2. Government Accession No- 3. Recipient's Catalog No.NASA TP-1701 AVRADCOM TR 80---5 -,9,L./1 711__

4. Title and Subtitle 5. Report Date

AERODYNAMIC CHARACTERISTICS OF THREE HELICOPTER ROTOR September 1980AIRFOIL SECTIONS AT REYNOLDS NUMBERS FROM MODEL SCA E 6. Performing Organization CodeTO FULL SCALE AT MACH NUMBERS FROM 0.35 to 0.90

7 Authorls) 8. Performing Organization Report No.

Kevin W. Noonan and Gene J. Bingham L-1313910. Work Unt No.

9 Performing Organization Name and Address 505-31-33-02

NASA Langley Research Center

and 11. Contract or Grant No.

Structures LaboratoryAVRADCOM Research and Technology Laboratorie.-

Hampton, VA 2366513. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Technical Paper

National Aeronautics and Space Administration

Washington, DC 20546 14 Army Project No.

and 1L161102AH45

U.S. Army Aviation Research and Development Command

St. Louis, MO 63166

15. Supplementary Notes

Kevin W. Noonan and Gene J. Bingham: Structures Laboratory, AVRADCOM Research

and Technology Laboratories.

16. Abstract

An investigation has been conducted in the Langley 6- by 28-Inch TransonicTunnel to determine the two-dimensional aerodynamic characteristics of threehelicopter rotor airfoils at Reynolds numbers from typical model scale to fullscale at Mach numbers from about 0.35 to 0.90. The model-scale Reynolds numbersranged from about 0.7 x 106 to 1.5 x 106 and the full-scale Reynolds numbersranged from about 3.0 x 106 to 6.6 x 106. The airfoils tested were the NACA 0012(00 Tab), the SC 1095-RO, and the SC 1095. Both the SC 1095 and the SC 1095-R8airfoils had trailing-edge tabs. The results of this investigation indicate thatReynolds number effects can be significant on the maximum normal-force coefficientand all drag-related parameters; namely, drag at zero normal force, maximumnormal-force-drag ratio, and drag-divergence Mach number. In general, theincrements in these parameters at a given Mach number owing to the model-scaleto full-scale Reynolds number change are different for each of the airfoils.

17 Key Words (Suggested by Author(s)) 18. Distribution Statement

Airfoils Unclassified - UnlimitedHelicoptersRotorsReynolds number effects Subject Category 02

19 Secu, ty Class,f (of this reportl 20 Security Classif (of this pagel 21 No of Pages 22 Price'

Unclassified Unclassified 82 A05

For sale by the National Technical Information Service. Springfield. Virginia 22161NASA-Langley, 1980

n ...... . .. ..... .. ll

Bpad AW M0WuuionA.roe~W It.C54 9M-8tAUA45r

PA'

2 1 1 U, A, 081280 S02672DUJDEPT OF DEFENSEDEFENSE TECHNICAL INFORMAT~ION CENTER

O~ATTN: DDC-DDA-2CAM.ERON STATION BLDG 5. . .

ALEXANDRIA VA 223 14

E W~Do PAOMA5M

, 4

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aerodynamic characteristics of three helicopter...

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