NASA 

Technical Paper 1701 

Aerodynamic Characteristics of Three 

Helicopter Rotor Airfoil Sections at 

Reynolds Numbers From Model h!e to Full 

Scaie at Mach nhmbers From 0.35 EG 0.90 8- ,. i. p &” 

- *b 

5.. 

Kcvin W’. Noonan and Gene j. Dingham 

SEPTE.’-IBER I980 

NASA 
NASA 

Technical Paper 1701 

AVRADCOM 

Technical Report 80-B-5 

Aerodynamic Characteristics of Three 

Helicopter Rotor Airfoil Sections at 

Reynolds Numbers From Model Scale to Full 

Scale at Mach Numbers From 0.35 to 0.90 

Kevin W. Noonan and Gene J. Bingham 

Structures Laboratory, A VRADCOM Research aid Technology Laboratories 

Langley Research Center 

Humpton, Virgiuia 

NASA 

iiat i o ri ai Aeronautics 

and Space Administration 

Scientific and Technical 

Information Branch 

1980 
SUMMARY 

An investigation has been conducted in the Langley 6- by 28-Inch Transonic 

Tunnel to determine the two-dimensional aerodynamic characteristics of three 

helicopter rotor airfoils at Reynolds numbers from typical model scale to full 

scale at Mach numbers from about 0.35 to 0.90, The model-scale Reynolds numbers ranged from about 0.7 x 1 O6 to 1 .5 x 1 O6 and the full-scale Reynolds numbers ranged from about 3.0 x lo6 to 6.6 x lo6. 

NACA 0012 (Oo Tab), the SC 1095-R8, and the SC 1095. Both the SC 1095 and the 

SC 1095-R8 airfoils had trailing-edge tabs. 

The airfoils tested were the 

The results of this investigation indicate that Reynolds number effects 

can be significant on the maximum normal-force coefficient and all drag-related 

parameters; namely, drag at zero normal force, maximum normal-force-drag ratio, 

and drag-divergence Mach number. In general, the increments in these parameters 

at a given Mach number owing to the model-scale to full-scale Reynolds number 

change are different for each of the airfoils. 

INTRODUCTION 

The development of new rotors generally includes predictions of the 

full-scale rotor performance based on extrapolation of small-scale-model rotor 

wind-tunnel results and based on rotor-performance analysis programs using twodimensional airfoil data at full-scale Reynolds numbers. Both prediction methods have inherent difficulties. The extrapolation of model-scale data requires 

a knowledge of how to increment the airfoil-section data for Reynolds number 

effects, and data for this purpose are seldom available. Rotor-performance 

analysis programs are already quite sophisticated, but they require a detailed 

knowledge of the rotor flow field, Improvements of these analysis programs can 

best be accomplished by correlation studies which require airfoil-section data 

at both model-scale and full-scale Reynolds numbers. 

In order to provide data for evaluating extrapolation methods and for 

rotor-performance correlation studies, the two-dimensional aerodynamic characteristics of three helicopter rotor airfoils have been determined at Mach numbers 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 1 O6 to 1.5 x 1 06, and 

the full-scale Reynolds numbers ranged from about 3.0 x 1 O6 to 6.6 x 1 06. The 

airfoils investigated were the NACA 0012 (Oo Tab), the SC 1095-R8, and the 

SC 1095. All three airfoils have a tab, which is a flat platelike extension to 

the normal trailing edge of an airfoil, that is usually used to reduce the airfoil pitching-moment coefficient. These airfoils were chosen because some data 

on model rotors and full-scale rotors utilizing these airfoils were available 

(refs. 1 and 2) or were expected to become available. 
SYMBOLS 

The units used for the physical quantities in this paper are given in both 

the International Systems of Units (SI) and U.S. Customary Units. The measurements and calculations were made in U.S. Customary Units, 

C airfoil chord, cm (in,) 

Cd section profile-drag coefficient, c CAE 

C 

Wake 

4 point-drag coefficient, 

section profile-drag coefficient at drag-divergence Mach number 

cdMdd 

Cm section pitching-moment coefficient about quarter-chord, 

cn 

cP 

h 

M 

Mad 

P 

section normal-force coefficient, 1 cpr$) + 1 cpe) 

U.S. L.S. 

p2 - pw stati c-press ure mef f icient, 

height of. wake-survey probe tubes from given reference plane, cm (in.) 

Mach number 

Mach number for drag divergence, 

static pressure, Pa (psi) 

dcd/dM = 0.1 

2 
R Reynolds number based on airfoil chord and free-stream conditions 

t airfoil thickness, cm (in. ) 

V velocity, m/sec (f t/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 airstream 

direct ion, deg 

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

P density, kg/m3 (slugs/ft3) 

Subscripts: 

1 local 

max maximum 

0 zero normal force 

=P separation 

t total 

03 free stream 

Abbreviations: 

L.S. lower surf ace 

U.S. upper surf ace 

APPARATUS AND METHODS 

Airfoils 

The airfoil profiles are shown in figure 1 and the thickness distributions 

and mean lines are shown in figure 2. The SC 1095-R8 airfoil was derived from 

the SC 1095 airfoil by the addition of a drooped nose; the mean line of the 

resultant 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-R8 airfoil forward of 

about 15-percent chord is greater than that of the SC 1095 airfoil. Both airfoils have a trailing-edge tab which is about 3-percent chord in length and is 

deflected upwards about 3O. 

tested is 9.1-percent chord and is located at the 25-percent-chord station; the 

maximum thickness of the SC 1095-R8 airfoil as tested is 9.0-percent chord and 

is also located at the 25-percent-chord station. The maximum camber of the 

SC 1095 airfoil is 0.8-percent chord: the SC 1095-R8 airfoil has the same 

positive camber as the SC 1095, but it has a maximum mean-line ordinate of 

-1.6-percent chord. The addition of the Oo Tab to the NACA 0012 airfoil results 

in a reduction of the maximum thickness from 12-percent chord to 11.7-percent 

chord. The design coordinates for these airfoils are given in tables I to 111. 

The maximum thickness of the SC 1095 airfoil as 

The airfoils were machined from heat-treated stainless-steel blocks and had 

spans of 15.27 cm (6.010 in.) and chords of about 7.87 cm (3.1 in.). The models 

had about 22 orifices located in one chordwise row on each surface; the orifice 

rows were positioned 12.6-percent span on either side of midspan (tables IV 

to VI). Slots were milled in the airfoil surface and tubes were placed in the 

slots and then covered with epoxy. The final airfoil contour had a surface finish of 0.813 pm (0.000032 in.). The orifices were then drilled from the opposite sides of the model so there were no surface irregularities near the orifice 

row. The orifices had a diameter of 0.0508 cm (0.020 in.) and were drilled perpendicular to the local-surface contour. 

Wind Tunnel 

Tunnel description.- The Langley 6- by 28-Inch Transonic Tunnel (ref. 3) 

is a blawdown wind tunnel with a slotted floor and ceiling and is generally 

operated at stagnation pressures fran 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 geanetry is described in detail in reference 4. Mach number is controlled by hydraulically actuated choker doors located downstream of the test 

section. The airfoil model spans the 15.27-cm (6,010-in.) width of the tunnel 

(fig, 3) and is rigidly attached by mounting tangs to two circular end plates 

which are driven by a hydraulic actuator to position the airfoil at the desired 

angle of attack. A test run usually consists of an angle-of-attack sweep at a 

constant Mach number and Reynolds number. 

Two-dimensionality of flm.- The results of a previous investigation of 

rotorcraft airfoils in the Langley 6- by 28-Inch Transonic Tunnel (ref. 5) 

have shown that the indicated maximum normal-force coefficient is reduced by 

tunnel-wall boundary-layer influences. This reduction is characteristic of 

two-dimensional wind tunnels without proper sidewall boundary-layer control 

and occurs because the tunnel-wall boundary layer is thicker than that of the 

airfoil: therefore, initial separation begins at the tunnel wall. Efforts are 

under way to correct this difficulty, but the solution was not available for 

the investigation described in this paper. 

Although it is not possible to determine precisely the affected Mach 

number range or the loss in maximum normal-force coefficient of the airfoils 

reported herein, a camparison of the NACA 0012 data measured in this facility 

4 
with 0.125 open slots (ref. 5) with unpublished data from two other facilities 

has been useful in indicating the magnitude of these losses. The maximum 

normal-force coefficients measured in the Langley Low-Turbulence Pressure 

Tunnel and the United Technologies Research Center 8 foot tunnel at similar 

Reynolds numbers and at a Mach number of 0.36 are higher than that fran the 

Langley 6- by 28-Inch Transonic Tunnel by about 0.15. The difference between 

the data fran the Langley 6- by 28-Inch Transonic Tunnel and the United Technologies data decreased to 0.10 at a Mach number of about 0.55. Incremental 

values for other airfoils may vary slightly because of specific configuration 

influences. 

An investigation conducted in the Office National d'Etudes et de Recherches 

A6rospatiales (ONERA) R1 Ch wind tunnel (ref. 6) has shown that the: tunnel sidewall boundary layer can affect the normal-force coefficients at all angles of 

attack (that is, with either attached or separated boundary layers). In this 

investigation, the sidewall boundary-layer thickness was varied by applying 

sidewall suction upstream of the model while the Mach number and Reynolds number were held constant. Generally an increase in sidewall boundary-layer thickness resulted in a decrease in the normal-force coefficient at a given angle of 

attack; the trend reversed at Mach numbers greater than 0.85 with a supercriti- I 

cal airfoil. 

Apparatus 

Wake-survey probe.- A traversing wake-survey probe is cantilevered fran one 

tunnel sidewall to measure the profile drag of the airfoils. The probe vertical 

sweep rate, which was selected after experimental determination of acceptable 

lag time in the pressure measurements, was about 2.54 cm/sec (1 .OO in/sec) . 

The probe was located 3.69 chords (based on the 7.87-cm (3.10-in.) chord 

model) downstream of the airfoil trailing edge and has a maximum vertical travel 

of about k27.9 cm (kll. 0 in. ) from the tunnel center line (fig. 3). Data are 

acquired with four total-pressure tubes, which are made of stainless-steel 

tubing, with a 1.53-mm outside diameter and a 1.02-mm inside diameter (0.060 in. 

by 0.040 in,); the tubes are spaced 0.953 cm (0.375 in.) apart laterally as 

shown in figure 4. 

Instrumentation.- All measurements made during the test program were 

obtained with the use of a high-speed, computer-controlled digital data acquisition system and were recorded by a high-speed tape-recording unit (ref. 3). 

All free-stream conditions were determined fran stagnation and static pressures. 

All airfoil surface pressures and all wake pressures were measured with precision capacitive potentianeter pressure transducers. The electrical outputs 

from each of these transducers were connected to individual autoranging signal 

conditioners which have seven available ranges. 

four signal conditioners measuring the wake pressures were filtered with 

20-Hz low-pass filters before input to the data acquisition system; the range 

of frequencies to be passed was experimentally determined during a previous 

investigation. The geanetric angle of attack was determined fran the output 

of a digital shaft encoder attached to a pinion engaging a rack on one model 

support end plate. 

The output signals fran the 

5 
Tests and Methods 

All of the testing was conducted with smooth model surfaces. Tests were 

made at stagnation pressures fran 121 kPa (17.5 psia) to 138 kPa (20 psia) to 

obtain Reynolds numbers typical of model-scale rotors at Mach numbers fran about 

0.35 to 0.90. This range of stagnation pressures was belaw that normally run in 

the Langley 6- by 28-inch Transonic Tunnel, and a tunnel calibration at these 

pressures indicated that a new calibration should be used to reduce these data. 

An additional new calibration was used to reduce data at stagnation pressures 

from 165 kPa (24 psia) to 193 kPa (28 psia). This range of stagnation pressures 

was 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 pressures fran 531 kPa (77 psia) to 621 kPa (90 psia) at the lowest and highest test 

Mach numbers, respectively. Geometric angles of attack ranged from -4O to 18O 

at increments of 20 at the lower test Mach numbers: this range was decreased at 

the higher test Mach numbers. 

Section normal-f orce and pitching-manent coefficients were calculated from 

the airfoil surface pressures by a trapezoidal integration of the pressure coefficients. The pressure coefficient at the most rearward orifice on each surface 

was applied frun that station to the airfoil trailing edge in the integration 

because the mall model size did not allow installation of orifices in the airfoil tab. Each of the pressure coefficients represents the average of five measurements obtained in a 1.0-second interval. A form of the equation described 

in reference 7 was used to calculate the point-drag coefficients frm the measured wake pressures, and a trapezoidal integration of the point-drag coefficients was used to calculate the drag coefficient. The static pressures used in 

the wake drag calculation were measured with tunnel sidewall orifices located at 

the same longitudinal tunnel station as the tips of the tubes on the wake-survey 

probe. All of the drag coefficients presented in this paper represent the mean 

of the measurements made with three total-pressure tubes on the wake-survey 

probe in one sweep through a wake. The corrections for lift interference, which 

have been applied to the angles of attack, were obtained fran references 4 

and 8. The basic equations for the correction (see ref. 8) are 

= cc + ACC 

where 

and a is the slot spacing and h is the semiheight of the tunnel. The 

slotted-wall boundary-condition coefficient k for the present tunnel configuration is 0.4211K. A value of 3.5 was selected for the slotted-wall prform6 
ante coefficient K, based on the data and discussion presented in reference 4. 

This substitution results in a correction given by the equation 

Results 

where c is in centimeters, a is in degrees, and the constant is in degrees 

per centimeter. 

Airfoil Figure 

PRESENTATION OF RESULTS 

Cn against a,; % and Cd 

against cn 

The results of the investigation have been reduced to coefficient form and 

are presented in the following table: 

NACA 0012 (Oo Tab) 

SC 1095 

SC 1095-R8 

Cp against M SC 1095-R8 

Cd against M SC 1095-R8 

SC 1095 

21 

22 

5 

6 

7 

c~,~~~ against M NACA 001 2 (Oo Tab) 

SC 1095-R8 

SC 1095 

8 

Cd10 against M NACA 001 2 (Oo Tab) 

SC 1095-R8 

SC 1095 

9 

NACA 001 2 (Oo Tab) 

SC 1095 

SC 1095-R8 

10 

NACA 001 2 (Oo Tab) 

SC 1095-R8 

SC 1095 

11 

against cn 

CdMdd 

NACA 001 2 (Oo Tab) 

SC 1095-R8 

SC 1095 

Cp against x/c NACA 001 2 (Oo Tab) 

SC 1095-R8 

SC 1095 

SC 1095-R8 

12 

13, 16, 20 

14, 17 

15, 18 

19 

7 
DISCUSS ION 

Normal Force 

Maximm normal-force meff icient. - The maximum normal-force meff icients of the NACA 0012 (0° Tab), the SC 1095-R8, and the SC 1095 airfoils have 

been 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 figure 8. Figure 8 shows that the effect of Reynolds number on Cn,max is different for the three airfoils. The maximun normal-force coefficient increases with 

increasing Reynolds number for the range of Mach numbers presented for both the 

NACA 0012 (Oo 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 in 

c~,~~~ for the SC 1095 airfoil. In addition, the increment in cnlmax resulting from the same increment in Reynolds number 

NACA 0012 (Oo Tab) and the SC 1095-R8 airfoils. These results are not surprising since the flow field on these airfoils at 

regions of supercritical flow, and it has been shown that the type and magnitude 

of the scale effect on for the incompressible case do not vary in any 

consistent manner (ref. 9). These trends can be explained by the normal-force 

curves and the pressure distributions (figs. 13 to 18). The increases in 

Cn,max of the NACA 0012 (Oo Tab) and the SC 1095-R8 airfoils are the result of 

either one or both of the following: (1) Increased suction (more negative C ) 

on the upper surface near the leading edge at the same angle of attack (figs.'13 

and 14); and (2) a delay of the turbulent boundary-layer separation to a higher 

angle of attack at the higher Reynolds number. (See figs. 5 (a), 13, 16, 6 (a), 

14, and 17.) The values of %,max for the SC 1095 airfoil generally decrease 

with increasing Reynolds number because of the occurrence of substantial 

boundary-layer separation at a lower angle of attack at the higher Reynolds number (figs. 7(a) and 15). This trend of c~,~~~ is possibly caused by a forward 

movement of a laminar separation bubble with increasing Reynolds number which 

would result in a greater length of turbulent boundary layer at the higher 

Reynolds number. The presence of a laminar separation bubble may be responsible for the related tendency of Cic for c~,~~~ to be generally lower at the 

higher Reynolds numbers (M 2 0.54). 

A%/& is not the same for the 

c,,,,, generally includes 

%,max 

At the full-scale Reynolds numbers the values of Cn,max for the 

SC 1095-R8 airfoil are higher than those for the SC 1095 airfoil at all Mach 

numbers presented, but at the model-scale Reynolds numbers the values of Cn,max 

for the SC 1095-R8 airfoil exceed those for the SC 1095 airfoil only for Mach 

numbers up to about 0.40 (fig. 8). This trend at the highest Reynolds number is 

the expected result of the increased leading-edge camber and greater thickness 

in the leading-edge region of the SC 1095-R8 airfoil. Figure 8 also shows that 

the airfoil with the highest values of 

sensitivity to increasing Mach number for M < 0.55. The decrease in Cn,max 

of the SC 1095-R8 airfoil is the result of either supercritical flow-induced 

separation occurring at a lower angle of attack or a more extensive separation 

(separation point farther forward) occurring at the same angle of attack with 

increasing Mach number (figs. 17 and 19). Since the SC 1095-R8 airfoil is more 

highly loaded in the leading-edge region than the other two airfoils and the 

boundary-layer separation reduces the leading-edge suction, the reduction in 

Cn,max with Mach number is greater for this airfoil. 

cnlmax, the SC 1095-R8, has the greatest 

8 
For the range of Mach numbers presented in figure 8, the stall of both the 

NACA 0012 (Oo Tab) and the SC 1095-R8 airfoils becomes less abrupt as a result 

of increasing Reynolds number from model scale to full scale (figs. 5(a) and 

6(a)). The stall of the SC 1095 airfoil becomes less abrupt with increasing 

Reynolds number for Mach numbers of about 0.39 and 0.44 but not at the other 

Mach numbers (fig. 7(a)). The pressure distributions presented in figures 16 

to 18 indicate that the stall of these airfoils is of the trailing-edge type by 

the characteristic loss of pressure recovery (more negative Cp) on the upper 

surface near the trailing edge. 

Slope.- The slopes of the normal-force curves of both cambered airfoils 

increase slightly with increasing Reynolds number at almost all Mach numbers 

presented. The slopes of the curves of the NACA 0012 (Oo Tab) airfoil are 

essentially insensitive to Reynolds number changes except at the highest test 

Mach number. Figure 5(a) shows a large and unusual change in the slope of the 

normal-force curve of the NACA 0012 (Oo Tab) airfoil with increasing Reynolds 

number at a Mach number of about 0.88. The test runs at this Mach number and 

a Reynolds number of 2.1 x lo6 were made twice and the repeatability of the data 

was excellent. The pressure distributions presented in figure 20 for an angle 

of attack of about -2.0° indicate that the shock position on the upper surface 

moves forward as a result of increasing Reynolds number from 1.5 x 1 O6 to 

2.1 x lo6 and then moves rearward again at the highest Reynolds number. This 

forward 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 lo6, 

thus giving rise to a steeper slope. 

The SC 1095-R8 and SC 1095 airfoils both have near-zero normal force at 

aC = Oo, although they are cambered airfoils. 

is the result of both the camber-line geometry and the trailing-edge tab. 

This near-zero normal force 

Pitching Moment 

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

Cn = 0) is essentially unchanged by increases in Reynolds number at all 

Mach 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 the 

nose-down break to higher normal-force coefficients for all the airfoils at Mach 

numbers up to about 0.78. This result is expected where the maximum normalforce coefficients increase with increasing Reynolds number. At a Mach number 

of about 0.88, the trend of the pitching moment of the NACA 0012 (Oo Tab) airfoil is substantially different at a Reynolds number of 2.1 x lo6 than at the 

lowest and highest Reynolds numbers. The reason for this difference is the 

shift of the shock position with Reynolds number mentioned previously. The 

pitching-moment coefficient about the aerodynamic center (Cm at Cn = 0) of 

the SC 1095-R8 airfoil displays the most sensitivity to increasing Mach number. 

Analysis of the pressure distributions (figs. 17 and 18) indicates that the 

SC 1095-R8 airfoil develops supercritical flow on the lower surface at a lower 

free-stream Mach number than the SC 1095 airfoil, and it is the growth of this 

supersonic zone with increasing Mach number that causes the greater Mach number 

sensitivity (fig. 21). 

9 
It is interesting to note that the data of figures 6(b) and 7(b) suggest 

that both of the cambered airfoils were designed for zero pitching-manent coefficient, although sane authors have mentioned a pitching-manent level of up to 

10.021 as acceptable (ref. 10). The zero pitching-manent level of the SC 1095 

and the SC 1095-R8 airfoils is due to the trailing-edge tab. It is also interesting to note that the leading-edge modifications made to the SC 1095 airfoil 

which resulted in the SC 1095-R8 configuration did not change the pitchingmanent coefficient about the aerodynamic center by more than about 0.01 for Mach 

numbers 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 of 

increasing Reynolds number from model scale to full scale is to reduce 

for both cambered airfoils for all Mach numbers presented and to reduce Cd,o 

for the NACA 0012 (Oo Tab) airfoil for Mach numbers up to about 0.65 (fig. 9). 

The difference in of the NACA 0012 (Oo Tab) airfoil at model-scale and 

full-scale Reynolds numbers is small at the lowest test Mach number (0.34), and 

this difference gradually disappears with increasing Mach number. The incremental decrease in at a given Mach number is generally different for each 

of the airfoils; the ACd,o for the SC 1095-R8 airfoil is the largest, and that 

for the NACA 0012 (Oo Tab) airfoil, the smallest. These trends are the result 

of the well-known reduction in laminar and turbulent boundary-layer skin friction with increasing Reynolds number (ref. 11 1. 

cdl0 

Cd,o 

cdr0 

The insensitivity of Cd, 0 of the NACA 0012 (Oo Tab) airfoil to Reynolds 

number is consistent with low-speed data of the NACA 001 2 airfoil (refs. 9 

and 12). for the NACA 001 2 (Oo Tab) airfoil at a Mach number of 0.43 and a Reynolds number of 3.9 x lo6 is about 0.0005 higher than that 

measured for the NACA 0012 airfoil in the Langley Low-Turbulence Pressure Tunnel 

at a Mach number of 0.36 and the same Reynolds number (unpublished data). The 

difference in drag level may be due to a higher turbulence level in the Langley 

6- by 28-Inch Transonic Tunnel at stagnation pressures above about 517 kPa 

(75 psia). Unpublished NACA 0012 airfoil data measured on two different chord 

models at the same Mach number and Reynolds number in the Langley 6- by 28-Inch 

Transonic Tunnel indicate a higher Cd,O for the smaller chord model, thus 

implying a higher turbulence level at the higher stagnation pressure (552 kPa 

(80 psia)). Therefore the increments in Cdl0 frm model-scale to full-scale 

Reynolds numbers presented in this paper are believed to be smaller than those 

which would be measured in free air. 

The value of cd,o 

Analysis of -the pressure distributions indicates that the increase in 

Cd,o with increasing Mach number for each of the airfoils at both model-scale 

and full-scale Reynolds numbers is the result of the developnent of supercritical flow and shock waves. 

Maximum normal-force-drag ratio.- The values of (Cr/cd)max have been 

determined from the drag data shown in figures 5(c), 6(c), and 7(c) and are 

presented as a function of Mach number in figure 10. 

the values of 

Reynolds number frm model scale to full scale for all Mach numbers presented. 

At a constant Mach number, 

(Cn/cd)max for all the airfoils increase with an increase in 

10 
The increases in this parameter with Reynolds number are the result of both a 

lower skin-friction drag for a given normal-force coefficient and a delay of 

turbulent boundary-layer separation to higher normal-force coefficients at the 

higher Reynolds numbers. The increments in (Cn/cd)max due to Reynolds number 

at a given Mach number less than 0.70 are generally quite different for each 

of the airfoils. The increases in (Cn/cd)max due to Reynolds number are generally small for all the airfoils at Mach numbers above about 0.70 where supercritical flow effects predominate over the viscous effects. This predominance 

can be illustrated by showing that the knee of the drag curve is controlled by 

the developnent of supercritical flow and not by separation. By drawing a 

tangent to the drag curve of the NACA 001 2 (Oo Tab) airfoil at a Mach number 

of 0.77 and a Reynolds number of 6.5 x lo6 (fig. 5(c)), 

to 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 flaw 

on the upper surface from about 3- to 50-percent chord and no characteristics of 

separation. (Minor separation could exist aft of 93-percent chord. ) 

(Cdcd)max is shown 

Drag divergence.- The drag coefficients at constant values of were 

cross plotted to obtain the drag-divergence Mach numbers. The normal-force 

coefficients corresponding to the drag-divergence Mach numbers at model-scale 

and full-scale Reynolds numbers are presented in figure 11. The effect of 

Reynolds number on Mdd 

0.7 for all the airfoils. The drag divergence at high normal-force coefficients 

(Cn > 0.7) may be controlled more by shock-induced boundary-layer separation 

than by sonic flow (ref. 13) moving aft of the airfoil crest. An analysis of 

the pressure distributions for these airfoils at model-scale Reynolds numbers 

suggests that this is the case. This explanation for drag divergence would be 

consistent with the generally larger effect of Reynolds number on at the 

high values of 

divergence resulting from flow separation. The drag-divergence Mach number at 

zero normal-force coefficient of the NACA 001 2 (Oo Tab) and the SC 1095 airfoils 

is unchanged by the increase in Reynolds number, but that of the SC 1095-R8 airfoil is reduced by about 0.02 because of the increase in Reynolds number. Figure 11 also indicates that the drag-divergence Mach numbers corresponding to all 

normal-force coefficients between about -0.1 and 0.3 of the SC 1095-R8 airfoil 

are lower at the full-scale Reynolds numbers than at the model-scale Reynolds 

numbers. A study of the curves of Cd against M for this range of normalforce coefficients (fig. 22) indicates that the drag coefficients at a Mach number of 0.78 at the full-scale Reynolds numbers would have to be lower for the 

to be as high as that at the model-scale Reynolds numbers. Although 

Reynolds number has essentially no effect on Mdd of the NACA 0012 (Oo Tab) 

airfoil at normal-force coeff icients up to about 0.75, the drag coefficient 

at Mdd is reduced at the full-scale Reynolds number for normal-force coefficients greater than 0.2 (fig. 12). Figure 12 shows that decreases 

with increasing Reynolds number for almost all normal-force coefficients for 

both cambered airfoils. 

is greatest for normal-force coefficients above about 

Cn because boundary-layer thickness would be crucial to drag 

CdMdd 

11 
CONCLUSIONS 

An investigation has been conducted in the Langley 6- by 28-Inch Transonic 

Tunnel to determine the two-dimensional aerodynamic characteristics of three 

helicopter rotor airfoils at Reynolds numbers from model scale to full scale at 

Mach numbers from about 0.35 to 0.90, The airfoils included in this investigation were the NACA 0012 (Oo Tab), the SC 1095-R8, and the SC 1095. Analysis of 

the test data has resulted in the following conclusions: 

1. The maximum normal-force coefficients of the NACA 0012 (Oo Tab) and the 

SC 1095-RE airfoils increased with increasing Reynolds number at all Mach nunbers presented. The maximum normal-force coefficients of the SC 1095 airfoil 

at full-scale Reynolds numbers were about the same or slightly lower than those 

at model-scale Reynolds numbers. 

2. The pitching-moment coefficients about the aerodynamic center of these 

three airfoils were essentially unchanged with increases in Reynolds number at 

all Mach numbers presented. 

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

Cdl0 

all test Mach numbers. At a constant Mach number up to about 0.65, the values 

of Cdl0 of the NACA 0012 (Oo Tab) airfoil decreased with increasing Reynolds 

number. The difference in Cdl0 of the NACA 0012 (Oo Tab) airfoil at modelscale and full-scale Reynolds numbers was small at a Mach number of 0.34. This 

difference gradually disappeared as the Mach number increased to 0.65. 

of both cambered airfoils decreased with increasing Reynolds number for 

4. For all test Mach numbers presented, the maximum normal-force-drag 

ratios of these three airfoils at the full-scale Reynolds number were higher 

than those at the model-scale Reynolds number. 

5. In general, the effect of Reynolds number on drag-divergence Mach number 

was greatest for normal-force coefficients above about 0.7 for these three airfoils. The drag-divergence Mach number at zero normal-force coefficient of the 

NACA 0012 (Oo Tab) and the SC 1095 airfoils was insensitive to Reynolds number 

changes, but that of the SC 1095-R8 airfoil was reduced by about 0.02 because 

of the change from model-scale to full-scale Reynolds number. 

Langley Research Center 

National Aeronautics and Space Administration 

Hampton, VA 23665 

July 3, 1980 

12 
1. Weller, William H.: Experimental Investigation of Effects of Blade Tip 

Geometry 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 Main 

Rotor - Investigation of Tip Mach Number Effects and Comparisons of Four 

Tip Shapes. Preprint No. 03, Proceedings of the 34th Annual National 

Forum, American Helicopter Soc., Inc., May 1978. 

3. Ladson, Charles L.: Description and Calibration of the Langley 6- by 

28-Inch Transonic Tunnel. NASA TN D-8070, 1975. 

4. Barnwell, Richard W.: Design and Performance Evaluation of Slotted Walls 

for Two-Dimensional Wind Tunnels. NASA TM-78648, 1978. 

5. Noonan, Kevin W.; and Bingham, Gene J.: Two-Dimensional Aerodynamic 

Characteristics of Several Rotorcraft Airfoils at Mach Numbers From 

0.35 to 0.90. NASA TM X-73990, 1977. 

6. Bernard+uelle, Re&: Influence of Wind Tunnel Wall Boundary Layers on 

Two-Dimensional Transonic Tests. NASA TT F-17,255, 1976. 

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

8. Davis, Don D., Jr.; and Moore, Dewey: Analytical Study of Blockage- and 

Lift-Interference Corrections for Slotted Tunnels Obtained by the Substitution of an Equivalent Hmogeneous Boundary for the Discrete Slots. 

NACA FtM L53E07b, 1953. 

9. Loftin, Laurence K., Jr.; and Smith, Hamilton A.: Aerodynamic Characteristics of 15 NACA Airfoil Sections at Seven Reynolds Numbers From 

0.7 x lo6 to 9.0 x lo6. NACA TN 1945, 1949. 

10. Wortmann, F. X.; and Drees, Jan M.: Design of Airfoils for Rotors. 

CAL/AVLABS Symposium Proceedings: 

V/STOL Aircraft, Volume I - Rotor/Propeller Aerodynamics, Rotor Noise, 

June 18-20, 1969. 

Aerodynamics of Rotary Wing and 

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.: Summary of Airfoil Data. NACA Rep. 824, 1945. (Supersedes NACA WR L-560.) 

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

13 
TABLE 1.- DESIGN COORDINATES FOR 

NACA 001 2 (Oo Tab) AIRFOIL 

and ordinates given in percent 

L airfoil chord 1 

Stat ion 

0 

.244 

.488 

.977 

1.465 

2.441 

3.418 

4.395 

5.859 

7.324 

9.766 

11.719 

14.648 

16.602 

19.531 

24.41 4 

29.297 

34.1 80 

39.063 

43.945 

48.828 

53.71 1 

58.594 

63.477 

68.359 

73.242 

78.125 

83,008 

87.891 

92.773 

97.137 

100.000 

Upper surface 

0 

.851 

1.193 

1.664 

2.015 

2.553 

2.972 

3.318 

3.748 

4.101 

4.573 

4.871 

5.220 

5.399 

5.603 

5.802 

5.861 

5.809 

5.667 

5.450 

5.1 70 

4.836 

4.456 

4.036 

3.578 

3.086 

2.562 

2.004 

1.414 

.788 

.195 

.195 

Lower surf ace 

0 

-. 851 

-1.1 93 

-1.664 

-2.01 5 

-2.553 

-2.972 

-3.31 8 

-3.748 

-4.101 

-4.573 

-4.871 

-5.220 

-5.399 

-5.603 

-5.802 

-5.861 

-5.809 

-5.667 

-5.450 

-5.1 70 

-4.836 

-4.456 

-4.036 

-3.578 

-3.086 

-2.562 

-2.004 

-1 .414 

-. 788 

-.195 

-. 195 

14 
TABLE 11.- DESIW COORDINATES FOR SC 1095-R8 AIRFOIL 

and ordinates given in 

percent airfoil chord 1 ~~ 

Station 

0 

.086 

.230 

.470 

.71 0 

.950 

1.1 89 

1.429 

1.669 

1 .go9 

2.148 

2.388 

2.628 

2.868 

3.108 

3.347 

3.587 

3.827 

4.307 

4.786 

5.266 

5.745 

6.225 

6.704 

7.664 

8.623 

9.582 

10.541 

L J 

Upper 

s ur face 

-1.621 

-1.113 

-.681 

-.182 

.201 

.51 8 

.796 

1.055 

1.285 

1.496 

1.707 

1.890 

2.072 

2.235 

2.388 

2.532 

2.657 

2.782 

3.002 

3.194 

3.367 

3.520 

3.654 

3.779 

4.000 

4.182 

4.335 

4.479 

Lower 

surf ace 

-1.621 

-2.129 

-2.427 

-2.71 4 

-2.887 

-3.002 

-3.088 

-3.156 

-3.204 

-3.251 

-3.290 

-3.328 

-3.357 

-3.386 

-3.41 5 

-3.443 

-3.462 

-3.482 

-3.520 

-3.549 

-3.578 

-3.597 

-3.606 

-3.626 

-3.654 

-3.664 

-3.674 

-3.683 

Station 

1 2.939 

15.337 

17.735 

20.132 

22.530 

24.928 

27.326 

29.724 

34.51 9 

39.31 5 

44.111 

48.907 

53.702 

58.498 

63.294 

68.089 

72.885 

77.681 

82.477 

87.272 

92.068 

94.466 

95.137 

95.521 

96.864 

98.782 

100.000 

upper 

surface 

4.748 

4.949 

5.1 03 

5.208 

5.275 

5.304 

5.304 

5.294 

5.208 

5.064 

4.882 

4.642 

4.364 

4.028 

3.654 

3.232 

2.762 

2.264 

1.726 

1.180 

.614 

.31 7 

.230 

.192 

.249 

.345 

.403 

Lower 

surf ace 

-3.693 

-3.693 

-3.693 

-3.693 

-3.693 

-3.693 

-3.693 

-3.683 

-3.626 

-3.530 

-3.395 

-3.232 

-3.031 

-2.801 

-2.532 

-2.235 

-1 .go9 

-1.554 

-1.1 89 

-.806 

-. 422 

-. 221 

-.173 

-.153 

-. 067 

.038 

.086 

15 
TABLE 111.- DESIGN COORDINATES FOR SC 1095 AIRFOIL 

[Stations and ordinates given in percent airfoil chord] 

Station 

0 

,242 

.484 

.726 

.968 

1.210 

1.453 

1.695 

1.937 

2.1 79 

2.421 

2.663 

2.905 

3.389 

3.873 

4.358 

4.842 

5.326 

5.81 0 

6.778 

7.747 

8.71 5 

9.683 

12.104 

14.525 

upper 

surf ace 

0 

.668 

.988 

1.259 

1.501 

1.714 

1.908 

2.092 

2.256 

2.41 1 

2.556 

2.682 

2.808 

3.031 

3.225 

3.399 

3.554 

3.689 

3.81 5 

4.038 

4.222 

4.377 

4.522 

4.793 

4.997 

Lower 

surface 

0 

-. 533 

-. 804 

-1.036 

-1.230 

-1.385 

-1.540 

-1.666 

-1.772 

-1.879 

-1,975 

-2.063 

-2.150 

-2.305 

-2.450 

-2.576 

-2.692 

-2.789 

-2.866 

-3.002 

-3.108 

-3.196 

-3.273 

-3.428 

-3.544 

Station 

16.946 

19.367 

21.788 

24.208 

26.629 

29.050 

33.892 

38.733 

43.575 

48.41 7 

53.258 

58.100 

62.942 

67.783 

72.625 

77.467 

82.309 

87.150 

91.992 

94.41 3 

95.091 

95.478 

96.834 

98.770 

100.000 

upper 

surf ace 

5.152 

5.258 

5.326 

5.355 

5.355 

5.345 

5.258 

5.113 

4.929 

4.687 

4.406 

4.067 

3.689 

3.263 

2.789 

2.285 

1.743 

1.1 91 

.620 

.320 

.232 

.194 

.252 

* 349 

.407 

Lower 

surface 

-3.622 

-3.680 

-3.709 

-3.728 

-3.728 

-3.71 8 

-3.660 

-3.563 

-3.428 

-3.263 

-3.060 

-2.828 

-2.556 

-2.256 

-1.927 

-1.569 

-1.201 

-.813 

-. 426 

-. 223 

-.174 

-. 155 

-. 068 

.039 

.087 
TABLE IV.- STATIC-PRESSURE ORIFICE LOCATIONS 

FOR NACA 001 2 (Oo Tab) AIRFOIL 

[Locations given in percent airfoil chord] 

Upper-surf ace 

station 

0 

1.10 

2.47 

4.96 

7.25 

9.82 

14.67 

19.54 

24.59 

29.52 

34.1 5 

39.02 

43.90 

48.84 

53.74 

58.52 

63.52 

68.32 

73.17 

77.98 

83.01 

88.00 

92.92 

Lowe r-s urf ace 

station 

0 

1.01 

2.50 

5.06 

7.04 

9.76 

14.68 

19.55 

24.62 

29.34 

34.28 

38.95 

43.96 

48.80 

53.70 

58.48 

63.50 

68.40 

73.23 

78.19 

83.1 1 

87.95 

92.83 

17 
TABLE V.- STATIC-PRESSURE ORIFICE LOCATIONS 

FOR SC 1095-R8 AIRFOIL 

[Locations given in percent airfoil chord] 

Upper-surface 

station 

0 

1.12 

2.06 

3.24 

5.70 

7.90 

10.40 

15.24 

20.1 0 

24.88 

29.67 

34.46 

39.29 

44.09 

48.88 

53.67 

58.50 

63.28 

68.07 

72.87 

77.66 

82.49 

87.28 

Lower-surf ace 

station 

0 

1 . 31 

3.69 

6.11 

8.52 

12.85 

20.03 

24.84 

29.62 

34.41 

39.20 

44.00 

48.81 

53.60 

58.39 

63.22 

68.02 

72.80 

77.59 

82.37 

87.1 9 

----- 

----- 

18 
TABLE VI.- STATIC-PRESSURE ORIFICE LOCATIONS 

FOR SC 1095 AIRFOIL 

[Locations given in percent airfoil chord] 

Upper-surf ace 

station 

0 

7.05 

2.40 

4.66 

7.10 

9.56 

14.34 

19.33 

24.21 

29.07 

33.89 

38.77 

43.59 

48.46 

53.30 

58.16 

63.00 

67.85 

72.68 

77.50 

82.29 

87.25 

Lower-s urf ace 

station 

0 

1.20 

2.33 

4.63 

7.05 

9.56 

14.51 

19.34 

24.1 6 

29.05 

33.87 

38.72 

43.58 

48.41 

53.25 

58.08 

62.98 

67.71 

72.66 

77.53 

82.37 

87.19 

19 
NACA 0012 (Oo Tab) 

SC 1095-R8 

SC 1095 

Figure 1.- Airfoil profiles. 

20 
21 
22 

d 

I 

il 

b 

A 

II 

II 

II 

I1 

V 

3r 

a 

I 

W 

h* 

a,- 
,953 (.375) 

I > 

a. 

> 

9.398 (3.700) 

5.080 ( 2 .O 0 0) 

2 -540 ( 1.000) 

.635( .250) 

1 T f 

Figure 4.- Wake-survey probe used in Langley 6- by 28-Inch Transonic Tunnel. 

All dimensions are in cm (in.). 

23 
U 

0 

0 

0 

0 

m 

5 

4 

0 

C 

h 

2 

W 

0 

U 

u 

aJ 

W 

W 

w 

I 

VI 

aJ 

LI 

3 

P 

-4 

E 
.2 

.I 

0 

0 

0 

0 

0 

3 

cm 

0 

0 

0 

0 

0 

0 

-. I 

-. .~ 2 

-. 8 -. 6 -. 4 -. 2 0 .2 .4 .6 .8 I. 0 I. 2 1.4 I. 6 I. 8 

n 

c 

(b) Section pitching-moment coefficients. 

Figure 5.- Continued. 

25 

v) 

JJ 

c 

aJ 

-4 

0 

-4 

w 

w 

aJ 

8 

8 

E 

5 

aJ 

u 

w 

I 

d 

la 

0 

c 

-4 

JJ 

u 

aJ 

v1 

I4 

-d 

0 

VI 

U 

4 

la 

co 

In 

QI 

0 

9: 

5: 

F 

W 

0 

v) 

u .A 

JJ 

v) 

.4 

U 

aJ 

e, 

u 

la 

U 

la 

c 

u 

u 

-i 

c 

$ U 

9 c 

aJ 

la 

8 

U 

aJ 

27 
C 

m 

2 
I 

29 
ul 

JJ 

c 

aJ 

-4 

V 

-rl 

W 

W 

i! 

8 

8 

E 

E 

W 

I 

4 

a 

30 
I 

31 
. 10 

.w 

.08 

.07 

.ff 

. 0: 

.01 

.o: 

.a 

. 01 

Cd 0 

0 

0 

0 

0 

0 

0 

0 

0 

C 

[ 

32 
c 

U 

-4 

3 

U 

C 

aJ 

-4 

o 

-4 

w 

w 

aJ 

8* 

rn 

-? 

Ej 

LIO 

ow 

wLI 

I -4 rca 

m 

E2 

00 el3 

m 

h 

\o 

'f. 

u5 

oc 

wu 

om 

e. 

0- 

-4 P 

ua 

4E-r 

.-I 

&O 

$E 

ccv 

07 0 

LIO 

rnw 

50 

w 

w 

r V 

33 
.03 

.02 

d, 0 

C 

.01 

0 

I 34 

.03 

.02 

d, 0 

C 

.01 

0 

.03 

.02 

d, 0 

C 

.01 

0 

--- 

Full scale - - - 

3 .4 .5 .6 .7 .8 .9 

M 

Figure 9.- Effect of Reynolds number on section drag coefficient at zero normal 

force of NACA 0012 (Oo Tab), SC 1095-R8, and SC 1095 airfoils. 
c 

U 

m 

m 

h 

‘9s 

.n 

9 

m 

7 

E 

.d I 

XLA 

mm 

Eo 

W 

ou 

VI 

c 

Or 

u a 

aa 

-4 t3 LI 

do 

c 

0-4 

35 
I 

P 

a 

LI 

'D 

c 

JJ 

-4 

3. 

rn 

. CUoa 

a 

CB 

0 

-4 0 

36 
.03 

.02 

d C 

Mdd 

.01 

0 

.03 

.02 

d C 

Mdd 

.01 

0 

.03 

.02 

d C 

Mdd 

.01 

0 

- 

6 R 

Model scale ~0.7 - 1.5 x IO 

6 R 

Model scale ~0.8 - 1.5 x IO 

6 R 

Model scale ~1.0 - 1.4 x IO 

Figure 12.- Effect of Reynolds number on section drag coefficient at dragdivergence Mach number of NACA 0012 (00 Tab), SC 1095-R8, and SC 1095 

airfoils. 

37 
Figure 1 

of NA 

lower 

38 

4.0 

-7 .95 .39 .I 

-5.6 

-5.2 

-4.8 

-4.4 

-4.0 

-3.6 

-3.2 

-2.8 

Cp -2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

1.2 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 

x/c 

(a) aC - loo. 

3.- Effect of Reynolds number on chordwise pressure dist 

CA 0012 (Oo Tab) airfoil. Symbols with plus sign inside 

surface. R given in millions. 

r i but ion 

indicate 
-~ 

-6.4 

-6.0 

-5.6 

-5.2 

-4.8 

-4.4 

-4.0 

-8.6 

-8.2 

-2.8 

cP 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

0 1.2 

(b) = 11.7O. 

Figure 13.- ConclGded. 

39 
-4.8 

-4.4 

-4.0 

-3.6 

-3.2 

-2.8 

-2.4 

-2.0 

cP 

-1.6 

-1.2 

-.8 

-.4 

( 

.4 

.8 

1.2 

x/c 

(a) aC = 9.6O. 

Figure 14.- Effect of Reynolds number on chordwise pressure distribution of 

SC 1095-R8 airfoil. Symbols with plus sign inside indicate lower surface. 

R given in millions. 

40 
-5.6 

-5.2 

-4.8 

-4.4 

-4.0 

-3.6 

-5.2 

-2.8 

-2.4 

cP 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

TPI I I I I I I I I1 I I I I I I I I I El 

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

XIC 

1.2 

(b) aC * 12O. 

Figure 14.-, Concluded. 

41 

cP 

1.2; ' ' I' .2 " .3 ' .4 ' ' .5 .6 .7 .8 .9 1.0 .1 

x/c 

(b) ac 12O. 

Figure 1 5 .- Concluded. 

43 
cP 

-4.8 

-4.4 

-4.0 

-8.6 

-3.2 

-2.8 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

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

1.2.. . '. . * " " " ' ' ' ' ' ' J 

x/c 

(a) M = 0.34; R - 2.9 x lo6. 

Figure 16.- Effect of angle of attack on chordwise pressure distribution of 

NACA 0012 (Oo Tab) airfoil. Symbols with plus sign inside indicate lower 

surface. 

44 
45 
-4.8 

0 0.0 0.0 .01 .LfLf 

A 2.1 2.0 .26 .Lf3 

-4.4 h Lf.1 3.8 .Lf9 .Lf3 

n 6.1 5.6 .70 .~3 

-4.0 

-8.6 

-8.2 

-2.8 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

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

x/c 

1.2 

(c) M = 0.43; R = 3.9 x lo6. 

Figure 16 .- Continued. 
cP 

-4.8 

0 0.0 0.0 -02 -98 

,A 2.0 1.8 -2'4 -98 

-4.4 h 9.0 3.7 .'47 .98 n 6.1 5.7 .73 .w 

0 8.2 7.7 .92 .Y8 

O 10.2 9.6 -96 -Y8 

-4.0 0 12.3 11 -7 -98 .98 

-9.6 

-3.2 

-2.8 

-2.4 

-2.0 

-1.6 

-1.2 

-.a 

-.4 

0 

.4 

.a 

.2 .3 .4 .5 .6 .7 .a .9 1.0 1.2' ' ' ' ' ' ' ' ' ' E 8 ' a I " 0 .1 

XJC 

(a) M = 0.48; R 4.3 X 10'. 

Figure 16 .- Continued. 

47 
-4.8 

-4.4 

-4.0 

-5.6 

-5.2 

-2.8 

-2.4 

-2.0 

cP 

-1.6 

-1.2 

-.8 

-.4 

(I 

.4 

.8 

1.2 

cM 9 n U 

-.1 -.l 0.00 .5Y 

2.1 1.9 .27 -5’4 

4.1 3.8 .51 .5Y 

12.3 11.7 .92 .5’4 

(e) M = 0.54; R - 4.9 x lo6. 

Figure 16.- Continued. 
Figure 16 .- Continued. 

49 
(9) M = 0.64; R = 5.5 x lo6. 

Figure 1 6 .- Continued. 

50 
-4.8 

-4.4 

-4.0 

-5.6 

-5.2 

-2.8 

-2.4 

-2.0 

cP 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

1.2 

I 

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

x/c 

(h) M 0.68; R = 5.6 x lo6. 

Figure 16 .- Continued. 

51 
52 

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

x/c 

Figure 16.- Continued. 
x/c 

(1) M = 0.77; R S, 6.5 x lo6. 

Figure 7 6 .- Continued. 

53 
-4.8 

-4.4 

-4.0 

-3.6 

-3.2 

-2.8 

-2.4 

-2.0 

cP 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

10 

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

x/c 

(k) M = 0.83; R = 6.6 x lo6. 

Figure 1 6 .- Continued. 

54 
55 I, 
(a) M - 0.33; R *( 3.0 X lo6. 

Figure 17. 

SC 1095 

- Effect of angle of attack on chordwise pressure distributi 

-RE airfoil. Symbols with plus sign inside indicate lower s 

on of 

, ur f ace. 

56 
-4.8 

0 4.1 -3.8 -.Y7 -38 

0 -2.1 -1-9 -.26 -38 

0 0.0 0.0 -e03 -38 

A 2.1 1.9 -21 .38 

h Ll-1 3.8 .Ll3 .38 

b 6.Y 5.9 -69 -38 

0 12.3 11.5 1.28 -38 

-4.4 

0 8.3 7.7 e90 -38 

0 10-3 9.6 1.10 -38 

0 1Ll.q 13.6 1-29 -38 

0 16-Ll 15-6 1.31 -38 

-4.0 

-3.6 

-3.2 

-2.8 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

0 .1 .2 .3 .4 .5 .6 .8 .9 1.0 

x/c 

1.2 

(b) M = 0.38; R 3.5 x 106. 

Figure 17.- Continued. 

57 
58 

MC 9 cn -4.8 (I 

PrOniCl 

M 
-4.8 

-4.4 

-4.0 

-3.6 

-3.2 

-2.8 

-2.4 

-2.0 

I cP 

I -1.6 

I -1.2 I 

-.8 

-.4 

( 

.4 

.a 

1.2 

x/c 

(d) M = 0.48; R = 4.4 x IO6. 

Figure 17.- Continued. 

59 
C 

n 4.8 a M 

x/c 

(e) M 4 0.54; R - 4.8 x lo6. 

Figure 17.- Continued. 

60 
WC a C -4.8 

P*onicl = c a W 

x/c 

(f) M = 0.58; R - 5.2 x IO6. 

Figure 17.- Continued. 

61 
-4.8 

-4.4 

-4.0 

I I I I I I I I I I I I I8 

a a C M 

C n 

-9.2 -3.9 --Li8 -63 

-2.0 -1.8 -.26 *63 

0.0 0.0 -SO2 .63 

2.1 1.9 -26 -63 

9.0 3.7 -51 -63 

6.2 5.7 -7’4 e63 

8.3 7.7 -88 *63 

10.2 9.6 093 -63 

12.9 11.8 *96 -63 

-8.6 

-8.2 

-2.8 

-2.4 

-2.0 

cP 

-1.6 

-1.2 

-.8 

-.4 

(I 

.4 

.8 

1.2 

x/c 

(9) M = 0.63; R S, 5.5 x lo6. 

Figure 17.- Continued. 

62 
.L .2 .s .4 .5 .6 .7 .8 .9 1.0 

x/c 

(h) M 4 0.68; R S, 5.9 X lo6. 

Figure 17.- Continued. 

63 

XIC 

(j) M = 0.78; R = 6.6 x lo6. 

Figure 17 .- Continued. 

65 
cP 

66 

-5.2 - 
cP 

-4.8 

0 -2.q -2.2 -.2q .3q 

0.0 0.0 0.00 .3Lt 

-1 2.0 a23 .3’4 

-4.4 

-4.0 

-3.6 

-3.2 

-2.8 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1 1.2 

(a) M = 0.34; R - 2.9 x iG6. 

Figure 18.- Effect of angle of attack on chordwise pressure distribution of 

SC 1095 airfoil. Symbols with plus sign inside indicate lower surface. 

67 
68 

cP 

(I M c cn (I -4.8 

x/c 

(b) M - 0.38; R - 3.4 x lo6. 

Figure 18.- Continued. 
-4.8 

-4.4 

-4.0 

-8.6 

-5.2 

-2.8 

-2.4 

-2.0 

cP 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

1.2 

(C) M = 0.43; R = 3.8 x lo6. 

Figure 18.- Continued. 

69 
x/c 

(a) M = 0.48; R = 4.3 x lo6. 

Figure 18.- Continued. 

3 
+s8 I I I I I I I I I Cp,txiicl a ‘c C n M 

(e) M = 0.54; R = 4.7 X lo6. 

Figure 18.- Continued. 

71 
cP 

72 

(f) M = 0.58; R = 5.1 x lo6. 

Figure 18.- Continued. 
x/c 

(9) M = 0.63; R - 5.4 x 106. 

Figure 18.- Continued. 

73 
74 
-4.8 

-4.4 

-4.0 

-8.6 

-8.2 

-2.8 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

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

x/c 

1.2 

(i) M = 0.73; R = 6.1 x 106. 

Figure 18 .- Continued. 

75 
-4.8 

-4.4 

-4.0 

-8.6 

-8.2 

-2.8 

-2.4 

-2 .o 

cP 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

1.2- ' I I I 1 I I I I I ' I 

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

x/c 

Cj) M - 0.78; R - 6.4 x lo6. 

Figure 1 8 .- Continued. 

76 
-4.8 

-4.4 

-4.0 

-9.6 

-9.2 

-2.8 

-2.4 

-2 .o 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

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

x/c 

1.2 

(k) M - 0.83; R = 6.6 x lo6. 

Figure 1 8 .- Continued. 

77 
78 
SeP (a,) 

( x IC 

16 

12 

8 

4 

1. 0 

.8 

sep .6 

.4 

R ~3.0 - 4.8 x 10 6 

pQ 

II 

0 

I 

I 

R 3.0 - 4.8 x 106 

Q 

\ 

\ 

\ 

0 

\ 

\ 

‘4 

Figure 19.- Effect of Mach number on angle of attack at which boundary-layer 

separation first occurs and on separation point of SC 1095-R8 airfoil. 

79 
-4.8 

-4.4 

x/c 

M R a,@ ac 'n 'p,sonic 

--'O 0.88 1.5 -2.0 -2.0 -0.06 -0.23 

0 0.88 2. I -2. I -1.9 -0.20 

'00.88 6.9 -2.1 -2.1 -0.04 

Figure 20.- Effect of Reynolds number on shock wave location on NACA 0012 

(Oo Tab) airfoil. 

R given in millions. 

Symbols with plus sign inside indicate lower surface. 

80 
- 3.6 

-3.2 

-2.8 

-2.4 

- 2.0 

cP 

-1.6. 

-1.2 

- .8 

-.4 

0 

' 

2 .3 .4 .5 .6 .7 .8 .9 I .o 

M 

\ I 

Cp, SC 1095-R8 SC 1095 sonic 

/ xlc x IC 

I 

\/ 0 0.0131 A 0.0120 

. 

\ 0 0.0369 I 0.0463 

I 0 0.0852 I b 0.0956 

1 

n 

0 W 

00 

Q 0 

-----~ ~__~

0 

\ U 

0 0 

- 

U b 

. b 

b 

- 

U 00 0 b by 

I 

0 L h \ 

I I .' b 

e** 

AA A AA A 

- mo 8 \ 

\ 

D A 

A a 

- 

C = 0.0 R z2.9 - 6.6 x 106 

Figure 21.- Effect of Mach number on sane lower-surface pressure coefficients 

forward of 10-percent-chord station of SC 1095-R8 and SC 1095 airfoils. 

81 
C 

d 

" .3 .4 .5 .6 .7 .a .9 I. 0 

M 

Figure 22.- Effect of Reynolds number on variation of section drag coefficient 

with Mach number of SC 1095-R8 airfoil. 

82 
1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. 

NASA TP-1701 AVRADCOM TR 80-B-5 

4. Title and Subtitle 5. Report Date 

AERODYNAMIC CHARACTERISTICS OF THREE HELICOPTER ROTOR September 1980 

AIRFOIL SECTIONS AT REYNOLDS NUMBERS FROM MODEL SCALE 6. performing Organization &de 

TO FULL SCALE AT MACH NUMBERS FROM 0.35 to 0.90 

7. Author(s) 8. Performing Organization Report No. 

Kevin W. Noonan and Gene J. Bingham L-13139 

~ 10. Work Unit No. 

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

11. Contract or Grant No. 

NASA Langley Research Center 

and 

Structures Laboratory 

AV~DCOM Research and Technology Laboratories 

Hampton, VA 23665 

12. Sponsoring Agency Name and Address 

13. Type of Report and Period Covered 

Technical Paper 

lL161102AH45 

National Aeronautics and Space Adminfstration 

Washington, DC 20546 14. Army Project No. 

and 

U.S. my 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 Transonic 

Tunnel to determine the two-dimensional aerodynamic characteristics of three 

helicopter rotor airfoils at Reynolds numbers from typical model scale to full 

scale at Mach numbers from about 0.35 to 0.90. The model-scale Reynolds numbers 

ranged from about 0.7 X lo6 to 1.5 x lo6 and the full-scale Reynolds numbers 

ranged from about 3.0 X lo6 to 6.6 x lo6. 

(0' Tab), the SC 1095-R3, and the SC 1095. 

airfoils had trailing-edge tabs. 

Reynolds number effects can be significant on the maximum normal-force coefficient 

and all drag-related parameters; namely, drag at zero normal force, maximum 

normal-force-drag ratio, and drag-divergence Mach number. In general, the 

increments in these parameters at a given Mach number owing to the model-scale 

tc full-SC?le Feynnldn nurnher chanae are different for each of the airfoils, 

The airfoils tested were the NACA 0012 

Both the SC 1095 and the SC 1095-R8 

The results of this investigation indicate that 

17. Key Words (Suggested by Author(s)) 1 18. Distribution Statement 

Airfoils 

Helicopter s 

Rotors 

Reynolds number effects 

Unclassified - Unlimited 

Subject Category 02 

I 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. NO. of Pages 22. Price' 

I unclassified Unclassified 82 A0 5 

For sale by the National Technical Information Service, Springfield, Virginia 221 61 

NASA-Langley, 1980