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NASA Technical Paper 1701

Aerodynamic Characteristics of Three Helicopter Rotor Airfoil Sections at Reynolds Numbers From Model h ! e to Full Scaie at Mach n h m b e r s From 0.35 EG 0.90

8- ,. i. p &” - *b

5..

Kcvin W’. Noonan and Gene j . Dingham

SEPTE.’-IBER I980

NASA

https://ntrs.nasa.gov/search.jsp?R=19800023826 2020-03-21T03:15:41+00:00Z

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 a i d Technology Laboratories Langley Research Center Humpton, Virgiuia

NASA i i a t i o r i ai Aeronautics and Space Administration

Scientific and Technical Information Branch

1980

SUMMARY

An i n v e s t i g a t i o n has been conducted i n t h e Langley 6- by 28-Inch Transonic Tunnel to determine t h e two-dimensional aerodynamic c h a r a c t e r i s t i c s of t h r e e h e l i c o p t e r rotor a i r f o i l s a t Reynolds numbers from t y p i c a l model scale to f u l l scale a t Mach numbers from about 0.35 t o 0.90, The model-scale Reynolds num- b e r s ranged from a b o u t 0 .7 x 1 O6 to 1 . 5 x 1 O6 and t h e f u l l - s c a l e Reynolds num- b e r s ranged from abou t 3.0 x l o 6 t o 6 . 6 x l o6 . NACA 0012 (Oo Tab) , t h e SC 1095-R8, and t h e SC 1095. Both t h e SC 1095 and t h e SC 1095-R8 a i r f o i l s had t r a i l i n g - e d g e t a b s .

The a i r f o i l s t e s t e d were t h e

The resu l t s of t h i s i n v e s t i g a t i o n i n d i c a t e t h a t Reynolds number e f f e c t s can be s i g n i f i c a n t on t h e maximum normal-force c o e f f i c i e n t and a l l d rag - re l a t ed parameters; namely, d rag a t zero normal force , maximum normal-force-drag ratio, and drag-divergence Mach number. I n gene ra l , t h e increments i n t h e s e parameters a t a g iven Mach number owing t o t h e model-scale t o f u l l - s c a l e Reynolds number change are d i f f e r e n t f o r each of t h e a i r f o i l s .

INTRODUCTION

The development of new r o t o r s g e n e r a l l y inc ludes p r e d i c t i o n s of t h e f u l l - s c a l e rotor performance based on e x t r a p o l a t i o n of small-scale-model rotor wind-tunnel resu l t s and based on rotor-performance a n a l y s i s programs us ing t w o - d imensional a i r f o i l d a t a a t f u l l - s c a l e Reynolds numbers. Both p r e d i c t i o n methods have i n h e r e n t d i f f i c u l t i e s . The e x t r a p o l a t i o n of model-scale d a t a r e q u i r e s a knowledge of how t o increment t h e a i r f o i l - s e c t i o n d a t a f o r Reynolds number e f f e c t s , and d a t a f o r t h i s purpose are seldom a v a i l a b l e . Rotor-performance a n a l y s i s programs are a l r e a d y q u i t e s o p h i s t i c a t e d , bu t t hey require a d e t a i l e d knowledge of t h e rotor flow f i e l d , Improvements of t h e s e a n a l y s i s programs can b e s t be accomplished by c o r r e l a t i o n s t u d i e s which r e q u i r e a i r f o i l - s e c t i o n d a t a a t both model-scale and f u l l - s c a l e Reynolds n u m b e r s .

I n o rde r to provide d a t a f o r e v a l u a t i n g e x t r a p o l a t i o n methods and f o r rotor-performance c o r r e l a t i o n s t u d i e s , t he two-dimensional aerodynamic charac- t e r i s t ics of t h r e e h e l i c o p t e r rotor a i r f o i l s have been determined a t Mach numbers from 0.35 t o 0.90 and a t Reynolds numbers from model scale t o f u l l scale. The model-scale Reynolds numbers ranged from about 0 .7 x 1 O6 to 1 . 5 x 1 06, and t h e full-scale Reynolds numbers ranged from about 3.0 x 1 O6 to 6 . 6 x 1 06 . The a i r f o i l s i n v e s t i g a t e d were t h e NACA 0012 (Oo Tab), t h e SC 1095-R8, and t h e SC 1095. A l l t h r e e a i r f o i l s have a t a b , which is a f l a t p l a t e l i k e e x t e n s i o n t o t h e normal t r a i l i n g edge of an a i r f o i l , t h a t is u s u a l l y used t o reduce t h e a i r - f o i l pitching-moment c o e f f i c i e n t . These a i r fo i l s were chosen because some data on model rotors and full-scale rotors u t i l i z i n g t h e s e airfoils were a v a i l a b l e ( r e f s . 1 and 2) or were expected t o become a v a i l a b l e .

SYMBOLS

The u n i t s used for t h e p h y s i c a l q u a n t i t i e s i n t h i s paper are g iven i n both t h e I n t e r n a t i o n a l Systems of U n i t s (SI ) and U . S . Customary Un i t s . The measurements and c a l c u l a t i o n s were made i n U . S . Customary U n i t s ,

C a i r f o i l chord, cm ( i n , )

Cd s e c t i o n p ro f i l e -d rag c o e f f i c i e n t , c C A E C

Wake

4 point-drag c o e f f i c i e n t ,

s e c t i o n p ro f i l e -d rag c o e f f i c i e n t a t drag-divergence Mach number cdMdd

Cm s e c t i o n pitching-moment c o e f f i c i e n t a b o u t quar te r -chord ,

cn

cP

h

M

Mad

P

s e c t i o n normal-force c o e f f i c i e n t , 1 cpr$) + 1 c p e ) U.S. L.S.

p2 - pw s t a t i c-press u r e m e f f i c i e n t ,

he igh t o f . wake-survey probe t u b e s from given r e fe rence plane, cm ( i n . )

Mach number

Mach number f o r drag divergence,

s t a t i c p res su re , Pa ( p s i )

dcd/dM = 0.1

2

R Reynolds number based on a i r f o i l chord and f ree-s t ream c o n d i t i o n s

t a i r f o i l t h i ckness , cm ( i n . )

V v e l o c i t y , m/sec ( f t/sec)

X a i r f o i l a b s c i s s a , cm ( i n . )

z a i r f o i l o r d i n a t e , cm ( i n . )

ZC o r d i n a t e of a i r f o i l mean l i n e , cm ( i n . )

a ang le of at tack, ang le between a i r f o i l chord l i n e and airstream d i r e c t ion , deg

a C ang le of attack c o r r e c t e d for l i f t - i n t e r f e r e n c e e f f e c t s , deg

P d e n s i t y , kg/m3 ( s l u g s / f t 3 )

Subsc r ip t s :

1 local

max maximum

0 ze ro normal f o r c e

=P separation

t to t a l

03 f r e e stream

Abbreviat ions:

L.S. lower s u r f ace

U.S. upper s u r f ace

APPARATUS AND METHODS

A i r f o i l s

The a i r f o i l p r o f i l e s are shown i n f i g u r e 1 and t h e t h i c k n e s s d i s t r i b u t i o n s and mean l i n e s are shown i n f i g u r e 2 . The SC 1095-R8 a i r fo i l was de r ived from t h e SC 1095 a i r f o i l by t h e a d d i t i o n of a drooped nose; t h e mean l i n e of t h e r e s u l t a n t a i r f o i l forward of about 20-percent chord is drooped r e l a t i v e to t h a t

3

of t h e SC 1095 a i r f o i l , and t h e t h i c k n e s s of t h e SC 1095-R8 a i r f o i l forward of a b o u t 15-percent chord is g r e a t e r t h a n t h a t of t h e SC 1095 a i r f o i l . Both airfoi ls have a t r a i l i n g - e d g e tab which is about 3-percent chord i n l e n g t h and i s d e f l e c t e d upwards about 3 O . tested is 9.1-percent chord and i s l o c a t e d a t t h e 25-percent-chord s t a t ion ; t h e maximum th ickness of t h e SC 1095-R8 a i r fo i l as t e s t e d is 9.0-percent chord and is also l o c a t e d a t t h e 25-percent-chord s t a t i o n . The maximum camber of t h e SC 1095 a i r f o i l i s 0.8-percent chord: t h e SC 1095-R8 a i r f o i l has t h e same p o s i t i v e camber as t h e SC 1095, but i t has a maximum mean-line o r d i n a t e of -1.6-percent chord. The a d d i t i o n of t h e Oo Tab t o t h e NACA 0012 a i r f o i l r e s u l t s i n a reduct ion of t h e maximum t h i c k n e s s from 12-percent chord t o 11.7-percent chord. The des ign c o o r d i n a t e s f o r t h e s e a i r f o i l s are g iven i n t a b l e s I t o 111.

The maximum t h i c k n e s s of t h e SC 1095 a i r f o i l as

The a i r f o i l s were machined from h e a t - t r e a t e d s t a i n l e s s - s t e e l blocks and had spans of 15.27 cm (6 .010 i n . ) and chords of about 7 .87 cm (3.1 i n . ) . The models had about 22 or i f ices located i n one chordwise row on each s u r f a c e ; t h e o r i f i c e rows were p o s i t i o n e d 12.6-percent span on e i t h e r s i d e of midspan ( tab les I V t o V I ) . Slots were m i l l e d i n t h e a i r f o i l s u r f a c e and t u b e s were placed i n t h e slots and then covered wi th epoxy. The f i n a l a i r f o i l contour had a s u r f a c e f i n - i s h o f 0.813 pm (0.000032 i n . ) . The o r i f i c e s were then d r i l l e d from t h e oppo- s i t e sides of t h e model so t h e r e were n o s u r f a c e i r r e g u l a r i t i e s near t h e o r i f i c e row. The o r i f i c e s had a diameter of 0.0508 cm (0.020 i n . ) and were d r i l l e d per- pendicular t o t h e l o c a l - s u r f a c e contour .

Wind Tunnel

Tunnel d e s c r i p t i o n . - The Langley 6- by 28-Inch Transonic Tunnel ( r e f . 3) is a blawdown wind t u n n e l wi th a s l o t t e d f l o o r and c e i l i n g and i s g e n e r a l l y o p e r a t e d a t s t a g n a t i o n p r e s s u r e s f r a n about 207 kPa (30 psia) t o 621 kPa (90 psia) and a t Mach numbers from 0.35 t o 0.90. The s e l e c t i o n of t h e 0.05- open-slot g e a n e t r y is d e s c r i b e d i n d e t a i l i n r e f e r e n c e 4 . Mach number is con- t ro l led by h y d r a u l i c a l l y a c t u a t e d choker doors l o c a t e d downstream of t h e t es t s e c t i o n . The a i r f o i l model spans t h e 15.27-cm (6,010-in.) width of t h e t u n n e l ( f i g , 3) and i s r i g i d l y a t t a c h e d by mounting t a n g s to t w o c i r c u l a r end plates which a r e d r i v e n by a h y d r a u l i c a c t u a t o r to p o s i t i o n t h e a i r f o i l a t t h e d e s i r e d a n g l e of a t t a c k . A test run u s u a l l y c o n s i s t s of an angle-of-at tack sweep a t a c o n s t a n t Mach number and Reynolds number.

Two-dimensionality of f lm.- The r e s u l t s of a prev ious i n v e s t i g a t i o n of r o t o r c r a f t a i r f o i l s i n t h e Langley 6- by 28-Inch Transonic Tunnel ( r e f . 5 ) have shown t h a t t h e i n d i c a t e d maximum normal-force c o e f f i c i e n t is reduced by tunnel-wall boundary-layer i n f l u e n c e s . T h i s r e d u c t i o n is c h a r a c t e r i s t i c of two-dimensional wind t u n n e l s wi thout proper sidewall boundary-layer c o n t r o l and occurs because t h e tunnel-wal l boundary l a y e r is t h i c k e r t h a n t h a t of t h e a i r f o i l : t h e r e f o r e , i n i t i a l separation begins a t t h e t u n n e l w a l l . E f f o r t s are under way to correct t h i s d i f f i c u l t y , b u t t h e s o l u t i o n was n o t a v a i l a b l e for t h e i n v e s t i g a t i o n d e s c r i b e d i n t h i s paper.

Although i t is not possible to determine p r e c i s e l y t h e a f f e c t e d Mach number range or t h e loss i n maximum normal-force c o e f f i c i e n t of t h e a i r f o i l s r e p o r t e d h e r e i n , a camparison of t h e NACA 0012 data measured i n t h i s f a c i l i t y

4

w i t h 0.125 open s lots ( r e f . 5) w i th unpublished d a t a from t w o o t h e r f a c i l i t i e s has been u s e f u l i n i n d i c a t i n g t h e magnitude of t h e s e losses. The maximum normal-force c o e f f i c i e n t s measured i n t h e Langley Low-Turbulence P r e s s u r e Tunnel and t h e Uni ted Technologies Research Center 8 f o o t t u n n e l a t similar Reynolds numbers and a t a Mach number of 0.36 a r e h igher than t h a t f r a n t h e Langley 6- by 28-Inch Transonic Tunnel by about 0 . 1 5 . The d i f f e r e n c e between t h e d a t a f r a n t h e Langley 6- by 28-Inch Transonic Tunnel and t h e Uni ted Tech- no log ie s d a t a decreased to 0.10 a t a Mach number of about 0.55. Incrementa l va lues f o r o t h e r a i r fo i l s may v a r y s l i g h t l y because of s p e c i f i c c o n f i g u r a t i o n i n f l u e n c e s .

An i n v e s t i g a t i o n conducted i n t h e Of f i ce Na t iona l d 'E tudes e t de Recherches A6rospatiales (ONERA) R1 Ch wind t u n n e l ( r e f . 6 ) has shown t h a t the: t u n n e l s ide - w a l l boundary l a y e r can a f f e c t t h e normal-force c o e f f i c i e n t s a t a l l ang le s of attack ( t h a t is, wi th e i t h e r a t t a c h e d or sepa ra t ed boundary l a y e r s ) . I n t h i s i n v e s t i g a t i o n , t he s i d e w a l l boundary-layer t h i ckness was v a r i e d by app ly ing s i d e w a l l s u c t i o n upstream of t h e model while t h e Mach number and Reynolds number were h e l d c o n s t a n t . Gene ra l ly an inc rease i n s i d e w a l l boundary-layer t h i ck - n e s s r e s u l t e d i n a dec rease i n t h e normal-force c o e f f i c i e n t a t a g iven a n g l e of a t tack; the t r e n d r eve r sed a t Mach numbers greater than 0.85 wi th a supercriti- I

cal a i r f o i l .

Apparatus

Wake-survey probe.- A t r a v e r s i n g wake-survey probe is c a n t i l e v e r e d f r a n one tunne l s i d e w a l l to measure t h e p r o f i l e drag of t h e a i r f o i l s . The probe v e r t i c a l sweep rate, which was s e l e c t e d a f t e r exper imenta l de t e rmina t ion of a c c e p t a b l e l a g t i m e i n t h e p r e s s u r e measurements, was about 2.54 cm/sec (1 . O O i n / s e c ) .

The probe was l o c a t e d 3.69 chords (based on t h e 7.87-cm (3 .10- in . ) chord model) downstream of t h e a i r f o i l t r a i l i n g edge and has a maximum v e r t i c a l t r a v e l of a b o u t k27.9 cm ( k l l . 0 i n . ) from t h e tunnel c e n t e r l i n e ( f i g . 3 ) . Data a re acqu i red wi th fou r total-pressure tubes , which are made of s t a i n l e s s - s t e e l t ub ing , w i th a 1.53-mm o u t s i d e diameter and a 1.02-mm i n s i d e diameter (0.060 i n . by 0.040 i n , ) ; t h e tubes are spaced 0.953 cm (0.375 i n . ) apart l a t e r a l l y as shown i n f i g u r e 4 .

I n s t rumen ta t ion . - A l l measurements made du r ing t h e tes t program were o b t a i n e d wi th t h e use of a high-speed, computer-controlled d i g i t a l d a t a acqui- s i t i o n system and were recorded by a high-speed tape- record ing u n i t ( r e f . 3 ) . A l l free-stream c o n d i t i o n s were determined f r a n s t a g n a t i o n and s ta t ic pressures. A l l a i r f o i l s u r f a c e pressures and a l l wake p res su res were measured w i t h preci- s i o n c a p a c i t i v e p o t e n t i a n e t e r pressure t r ansduce r s . The electrical o u t p u t s from each of t h e s e t r a n s d u c e r s were connected t o i n d i v i d u a l au toranging s i g n a l c o n d i t i o n e r s which have seven a v a i l a b l e ranges. four s i g n a l c o n d i t i o n e r s measuring t h e wake p r e s s u r e s were f i l t e r e d w i t h 20-Hz low-pass f i l t e r s be fo re i n p u t t o t h e da t a a c q u i s i t i o n system; t h e range of f r equenc ie s to be passed was exper imenta l ly determined du r ing a p rev ious i n v e s t i g a t i o n . The geanetric a n g l e of a t t a c k was determined f r a n t h e o u t p u t of a d i g i t a l s h a f t encoder a t t a c h e d t o a p in ion engaging a rack on one model s u p p o r t end plate .

The o u t p u t s i g n a l s f r a n t h e

5

T e s t s and Methods

A l l of t h e t e s t i n g was conducted wi th smooth model s u r f a c e s . Tests were made a t s t agna t ion pressures f r a n 121 kPa (17.5 psia) t o 138 kPa (20 p s i a ) t o o b t a i n Reynolds numbers t y p i c a l of model-scale rotors a t Mach numbers f r a n a b o u t 0.35 to 0.90. Th i s range of s t a g n a t i o n pressures was belaw t h a t normally run i n t h e Langley 6- by 28-inch Transonic Tunnel, and a t u n n e l c a l i b r a t i o n a t t h e s e pressures i n d i c a t e d t h a t a new c a l i b r a t i o n s h o u l d be used to reduce t h e s e data. An a d d i t i o n a l new c a l i b r a t i o n was used to reduce d a t a a t s t a g n a t i o n pressures from 165 kPa (24 p s i a ) t o 193 kPa (28 p s i a ) . T h i s range of s t a g n a t i o n pressures w a s used to o b t a i n data a t Reynolds numbers between model scale and f u l l scale. The f u l l - s c a l e Reynolds number data were ob ta ined by t e s t i n g a t s t a g n a t i o n pressures f r a n 531 kPa ( 7 7 psia) to 621 kPa (90 p s i a ) a t t h e lowest and h i g h e s t t es t Mach numbers, r e spec t ive ly . Geometric ang le s of a t tack ranged from -4O t o 18O a t increments of 20 a t t h e lower tes t Mach numbers: t h i s range was decreased a t t h e higher test Mach numbers.

Sec t ion normal-f orce and pitching-manent c o e f f i c i e n t s were calculated from the a i r f o i l su r f ace p re s su res by a t r a p e z o i d a l i n t e g r a t i o n of t he p r e s s u r e coef- f i c i e n t s . The pressure c o e f f i c i e n t a t t h e most rearward o r i f i c e on each surface was appl ied f r u n t h a t s t a t i o n to the a i r f o i l t r a i l i n g edge i n t h e i n t e g r a t i o n because t h e mal l model s i z e d i d not allow i n s t a l l a t i o n of o r i f i c e s i n t h e air- f o i l tab. Each of the p r e s s u r e c o e f f i c i e n t s r e p r e s e n t s the average of f i v e measurements ob ta ined i n a 1.0-second i n t e r v a l . A form of t h e equa t ion described i n re ference 7 was used t o c a l c u l a t e t h e point-drag c o e f f i c i e n t s f r m the measured wake pressures, and a t r a p e z o i d a l i n t e g r a t i o n of t h e poin t -drag c o e f f i - c i e n t s was used t o c a l c u l a t e t h e drag c o e f f i c i e n t . The s t a t i c pressures used i n t h e wake drag c a l c u l a t i o n were measured wi th tunne l sidewall o r i f i c e s located a t t h e same l o n g i t u d i n a l tunnel s t a t i o n as the t ips of the t u b e s on t h e wake-survey probe. A l l of t h e drag c o e f f i c i e n t s p re sen ted i n t h i s paper r e p r e s e n t t h e mean of the measurements made wi th t h r e e total-pressure t u b e s on t h e wake-survey probe i n one sweep through a wake. The c o r r e c t i o n s f o r l i f t i n t e r f e r e n c e , which have been app l i ed to the ang le s of attack, were ob ta ined f r a n r e f e r e n c e s 4 and 8 . The b a s i c equations f o r t h e c o r r e c t i o n (see r e f . 8 ) are

= cc + ACC

where

and a is the s lo t spac ing and h i s the semiheight of the tunnel . The s lo t t ed -wa l l boundary-condition c o e f f i c i e n t k for t h e p r e s e n t t unne l config- u r a t i o n is 0.4211K. A value of 3.5 was selected for t h e slotted-wall p r f o r m -

6

ante c o e f f i c i e n t K, based on t h e data and d i s c u s s i o n p r e s e n t e d i n reference 4. T h i s s u b s t i t u t i o n r e s u l t s i n a c o r r e c t i o n g iven by t h e e q u a t i o n

R e s u l t s

where c is i n cent imeters , a is i n degrees, and t h e c o n s t a n t i s i n degrees per cent imeter .

A i r f o i l F i g u r e

PRESENTATION OF RESULTS

Cn a g a i n s t a,; % and Cd a g a i n s t cn

The r e s u l t s of t h e i n v e s t i g a t i o n have been reduced to c o e f f i c i e n t form and are presented i n t h e fo l lowing table:

NACA 0012 (Oo Tab)

SC 1095 SC 1095-R8

Cp a g a i n s t M SC 1095-R8

Cd a g a i n s t M SC 1095-R8

SC 1095 21

22

5 6 7

c ~ , ~ ~ ~ a g a i n s t M NACA 001 2 (Oo Tab) SC 1095-R8 SC 1095

8

Cd10 a g a i n s t M NACA 001 2 (Oo Tab) SC 1095-R8 SC 1095

9

NACA 001 2 (Oo Tab)

SC 1095 SC 1095-R8

1 0

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

1 2

13, 16, 20 14, 1 7 15, 1 8

1 9

7

DISCUSS ION

Normal Force

Maximm normal-force m e f f i c i e n t . - The maximum normal-force m e f f i- c i e n t s of t h e NACA 0012 (0° Tab), t h e SC 1095-R8, and t h e SC 1095 a i r f o i l s have been determined from t h e normal-force curves p re sen ted i n f i g u r e s 5 (a ) , 6 ( a ) , and 7 ( a ) , r e spec t ive ly , and are p l o t t e d as a f u n c t i o n of Mach number i n f i g - u re 8. Figure 8 s h o w s t h a t t he e f f e c t of Reynolds number on Cn,max is d i f f e r - e n t f o r t he t h r e e a i r f o i l s . The maximun normal-force c o e f f i c i e n t i n c r e a s e s with i n c r e a s i n g Reynolds number f o r t h e range of Mach numbers p re sen ted f o r both t h e NACA 0012 (Oo Tab) and t h e SC 1095-R8 a i r f o i l s but n o t f o r t h e SC 1095 a i r f o i l . I nc reases i n Reynolds number r e su l t i n l i t t l e change or a small r educ t ion i n c ~ , ~ ~ ~ f o r t h e SC 1095 a i r f o i l . I n add i t ion , t h e increment i n cnlmax result- i n g from the same increment i n Reynolds number NACA 0012 (Oo Tab) and t h e SC 1095-R8 a i r f o i l s . These results are n o t s u r - p r i s i n g s ince the f l o w f i e l d on t h e s e a i r f o i l s a t r eg ions of s u p e r c r i t i c a l flow, and it has been shown t h a t t h e type and magnitude of the s c a l e e f f e c t on f o r t h e incompress ib le case do not vary i n any c o n s i s t e n t manner ( r e f . 9 ) . These t r e n d s can be exp la ined by t h e normal-force curves and t he p re s su re d i s t r i b u t i o n s ( f i g s . 13 t o 18) . The i n c r e a s e s i n Cn,max of t h e NACA 0012 (Oo Tab) and t h e SC 1095-R8 a i r f o i l s are t h e resu l t of e i t h e r one or both of the fol lowing: (1) Increased s u c t i o n (more nega t ive C ) on the upper s u r f a c e near t h e l e a d i n g edge a t t h e same ang le of attack (figs.'13 and 1 4 ) ; and (2) a d e l a y of the t u r b u l e n t boundary-layer s e p a r a t i o n t o a higher ang le of a t t a c k a t t h e h igher Reynolds number. (See f i g s . 5 (a) , 13, 16, 6 (a) , 14 , and 17 . ) The v a l u e s of %,max f o r t he SC 1095 a i r f o i l g e n e r a l l y d e c r e a s e with i n c r e a s i n g Reynolds number because of t h e occurrence of s u b s t a n t i a l boundary-layer s e p a r a t i o n a t a lower ang le of at tack a t t h e higher Reynolds number ( f i g s . 7 ( a ) and 1 5 ) . This t r e n d of c ~ , ~ ~ ~ is p o s s i b l y caused by a forward movement of a laminar s e p a r a t i o n bubble w i t h i n c r e a s i n g Reynolds number which would resul t i n a g r e a t e r l e n g t h of t u r b u l e n t boundary l a y e r a t t h e higher Reynolds number. The presence of a laminar s e p a r a t i o n bubble may be respons i - b l e f o r t h e r e l a t e d tendency of Cic f o r c ~ , ~ ~ ~ to be g e n e r a l l y lower a t t h e h igher Reynolds numbers (M 2 0.54).

A%/& is not t h e same f o r t h e

c,,,,, g e n e r a l l y inc ludes

%,max

A t t h e f u l l - s c a l e Reynolds numbers t h e va lues of Cn,max f o r t h e SC 1095-R8 a i r f o i l are higher than t h o s e f o r t h e SC 1095 a i r f o i l a t all Mach numbers presented, but a t t h e model-scale Reynolds numbers t h e va lues of Cn,max f o r t he SC 1095-R8 a i r f o i l exceed t h o s e for the SC 1095 a i r f o i l on ly f o r Mach numbers u p to about 0.40 ( f i g . 8 ) . This t r e n d a t t h e h i g h e s t Reynolds number is t h e expected result of the inc reased leading-edge camber and g r e a t e r t h i c k n e s s i n t h e leading-edge r eg ion of t h e SC 1095-R8 a i r f o i l . F igure 8 also shows t h a t the a i r fo i l with t h e h i g h e s t va lues of s e n s i t i v i t y to i n c r e a s i n g Mach number f o r M < 0.55. The dec rease i n Cn,max of the SC 1095-R8 a i r f o i l is the r e s u l t of e i t h e r supercr i t ical flow-induced sepa ra t ion o c c u r r i n g at a lower ang le of at tack or a more e x t e n s i v e s e p a r a t i o n ( sepa ra t ion po in t f a r t h e r forward) o c c u r r i n g a t t h e same a n g l e of at tack w i t h i n c r e a s i n g Mach number ( f i g s . 1 7 and 1 9 ) . S ince t h e SC 1095-R8 a i r f o i l is more h igh ly loaded i n the leading-edge r eg ion than t h e o t h e r t w o a i r f o i l s and t h e boundary-layer s e p a r a t i o n reduces t h e leading-edge s u c t i o n , t h e r e d u c t i o n i n Cn,max wi th Mach number i s g r e a t e r f o r t h i s a i r f o i l .

cnlmax, the SC 1095-R8, has the g r e a t e s t

8

For t h e range of Mach numbers presented i n f i g u r e 8, t h e s t a l l of both t h e NACA 0012 (Oo Tab) and t h e SC 1095-R8 a i r fo i l s becomes less a b r u p t as a r e s u l t of i n c r e a s i n g Reynolds number from model s c a l e to f u l l scale ( f i g s . 5 ( a ) and 6 ( a ) ) . The s t a l l of t h e SC 1095 a i r f o i l becomes less a b r u p t wi th i n c r e a s i n g Reynolds number for Mach numbers of about 0.39 and 0.44 b u t n o t a t t h e o t h e r Mach numbers ( f ig . 7 ( a ) ) . The p r e s s u r e d i s t r i b u t i o n s p r e s e n t e d i n f i g u r e s 16 to 1 8 i n d i c a t e t h a t t h e s t a l l of t h e s e a i r f o i l s is of t h e t r a i l i n g - e d g e type by t h e c h a r a c t e r i s t i c loss of p r e s s u r e recovery (more n e g a t i v e Cp) on t h e upper s u r f a c e near t h e t r a i l i n g edge.

Slope.- The slopes of t h e normal-force curves o f bo th cambered a i r f o i l s i n c r e a s e s l i g h t l y wi th i n c r e a s i n g Reynolds number a t almost a l l Mach numbers presented . The slopes of t h e curves of t h e NACA 0012 (Oo Tab) a i r f o i l are e s s e n t i a l l y i n s e n s i t i v e to Reynolds number changes e x c e p t a t t h e h i g h e s t test Mach number. F i g u r e 5 ( a ) shows a l a r g e and unusual change i n t h e slope of t h e normal-force curve of t h e NACA 0012 (Oo Tab) a i r f o i l wi th i n c r e a s i n g Reynolds number a t a Mach number o f about 0.88. The test r u n s a t t h i s Mach number and a Reynolds number of 2.1 x l o 6 were made twice and t h e r e p e a t a b i l i t y of t h e d a t a was e x c e l l e n t . The p r e s s u r e d i s t r i b u t i o n s presented i n f i g u r e 20 for an angle of attack of about -2.0° i n d i c a t e t h a t t h e shock p o s i t i o n on t h e upper s u r f a c e moves forward as a r e su l t of i n c r e a s i n g Reynolds number from 1.5 x 1 O 6 to 2.1 x l o 6 and then moves rearward a g a i n a t t h e h i g h e s t Reynolds number. This forward s h i f t o f t h e shock p o s i t i o n r e s u l t s i n s u b s t a n t i a l l y more down l o a d (negat ive l i f t ) on t h e rear of t h e a i r f o i l a t a Reynolds number of 2.1 x lo6, t h u s g i v i n g r i s e to a steeper slope.

The SC 1095-R8 and SC 1095 a i r f o i l s b o t h have near-zero normal force a t aC = Oo, a l though t h e y are cambered a i r f o i l s . is t h e r e s u l t of both t h e camber-line geometry and t h e t r a i l i n g - e d g e t ab .

T h i s near-zero normal force

P i t c h i n g Moment

The pitching-moment c o e f f i c i e n t about t h e aerodynamic c e n t e r (Cm a t Cn = 0) is e s s e n t i a l l y unchanged by i n c r e a s e s i n Reynolds number a t a l l Mach numbers presented f o r t h e s e t h r e e a i r f o i l s ( f i g s . 5 ( b ) , 6 ( b ) , and 7 ( b ) ) . These f i g u r e s also show t h a t t h e o n l y e f f e c t of Reynolds number is to change t h e "knee" of t h e curve; t h a t is, i n c r e a s e s i n Reynolds number g e n e r a l l y move t h e nose-down break to h igher normal-force c o e f f i c i e n t s f o r a l l t h e a i r f o i l s a t Mach numbers up to about 0.78. This r e s u l t i s expected where t h e maximum normal- force c o e f f i c i e n t s i n c r e a s e wi th i n c r e a s i n g Reynolds number. A t a Mach number of about 0.88, t h e t r e n d of t h e p i t c h i n g moment of t h e NACA 0012 (Oo Tab) air- f o i l is s u b s t a n t i a l l y d i f f e r e n t a t a Reynolds number of 2.1 x l o 6 than a t t h e lowest and h i g h e s t Reynolds numbers. The reason f o r t h i s d i f f e r e n c e is t h e s h i f t of t h e shock p o s i t i o n w i t h Reynolds number mentioned previous ly . The pitching-moment c o e f f i c i e n t about t h e aerodynamic c e n t e r (Cm a t Cn = 0) of t h e SC 1095-R8 a i r fo i l displays t h e most s e n s i t i v i t y to i n c r e a s i n g Mach number. Analys is of t h e p r e s s u r e d i s t r i b u t i o n s (f igs. 17 and 18 ) i n d i c a t e s t h a t t h e SC 1095-R8 a i r f o i l develops s u p e r c r i t i c a l flow on t h e lower s u r f a c e a t a lower free-s t ream Mach number than t h e SC 1095 a i r f o i l , and i t is t h e growth o f t h i s s u p e r s o n i c zone wi th i n c r e a s i n g Mach number t h a t causes t h e g r e a t e r Mach number s e n s i t i v i t y ( f ig . 21).

9

It i s i n t e r e s t i n g to note t h a t t h e data of f i g u r e s 6 ( b ) and 7 ( b ) s u g g e s t t h a t both of t h e cambered a i r f o i l s were designed f o r zero pitching-manent coef- f i c i e n t , al though sane au thor s have mentioned a pitching-manent l e v e l of up t o 10.021 as accep tab le (ref. 1 0 ) . The zero pitching-manent l e v e l of t h e SC 1095 and the SC 1095-R8 a i r f o i l s is due to t h e t r a i l i n g - e d g e tab. I t is also i n t e r - e s t i n g t o note t h a t t h e leading-edge m o d i f i c a t i o n s made t o t h e SC 1095 a i r f o i l which r e s u l t e d i n the SC 1095-R8 c o n f i g u r a t i o n d i d n o t change t h e p i t ch ing - manent c o e f f i c i e n t about t h e aerodynamic cen te r by more than about 0.01 f o r Mach numbers as high as 0.73 b u t d i d i n c r e a s e t h e va lues of Cn,max s i g n i f i c a n t l y .

Drag

Drag a t zero normal force.- A t a cons t an t Mach number, the e f f e c t of i nc reas ing Reynolds number from model scale to f u l l scale is to reduce f o r both cambered a i r f o i l s f o r a l l Mach numbers p re sen ted and to reduce Cd,o f o r t h e NACA 0012 (Oo Tab) a i r f o i l f o r Mach numbers u p to about 0.65 ( f i g . 9 ) . The d i f f e r e n c e i n of the NACA 0012 (Oo Tab) a i r f o i l a t model-scale and f u l l - s c a l e Reynolds numbers is small a t t h e lowest tes t Mach number (0.34) , and t h i s d i f f e r e n c e g radua l ly d i sappea r s wi th i n c r e a s i n g Mach number. The incre- mental decrease i n a t a g iven Mach number is g e n e r a l l y d i f f e r e n t f o r each of the a i r f o i l s ; t he ACd,o f o r t he S C 1095-R8 a i r f o i l is the l a r g e s t , and t h a t f o r t h e NACA 0012 (Oo Tab) a i r f o i l , t h e smallest. These t r e n d s are t h e resu l t of the well-known reduc t ion i n laminar and t u r b u l e n t boundary-layer s k i n f r i c - t i o n w i t h i n c r e a s i n g Reynolds number ( r e f . 11 1.

cdl0

Cd,o

cdr0

The i n s e n s i t i v i t y of Cd, 0 of the NACA 0012 (Oo Tab) a i r f o i l to Reynolds number is c o n s i s t e n t with low-speed data of t h e NACA 001 2 a i r f o i l ( r e f s . 9 and 1 2 ) . f o r t he NACA 001 2 (Oo Tab) a i r f o i l a t a Mach number of 0.43 and a Reynolds number of 3.9 x l o 6 is about 0.0005 h igher t han t h a t measured f o r t h e NACA 0012 a i r f o i l i n t h e Langley Low-Turbulence P r e s s u r e Tunnel a t a Mach number of 0.36 and t h e same Reynolds number (unpubl ished da t a ) . The d i f f e r e n c e i n drag l e v e l may be due to a higher t u rbu lence l e v e l i n t h e Langley 6- by 28-Inch Transonic Tunnel a t s t a g n a t i o n p res su res above about 517 kPa (75 p s i a ) . Unpublished NACA 0012 a i r f o i l data measured on t w o d i f f e r e n t chord models a t t h e same Mach number and Reynolds number i n t h e Langley 6- by 28-Inch Transonic Tunnel i n d i c a t e a higher C d , O f o r t h e smaller chord model, t hus implying a higher t u rbu lence l e v e l a t t h e higher s t a g n a t i o n p r e s s u r e (552 kPa (80 p s i a ) ) . Therefore the increments i n C d l 0 f r m model-scale to f u l l - s c a l e Reynolds numbers p re sen ted i n t h i s paper are be l i eved to be smaller than t h o s e which would be measured i n f r e e a i r .

The value of cd,o

Analysis of - t h e p r e s s u r e d i s t r i b u t i o n s i n d i c a t e s t h a t t h e i n c r e a s e i n Cd,o with i n c r e a s i n g Mach number f o r each of the a i r f o i l s a t both model-scale and f u l l - s c a l e Reynolds numbers is t h e resul t of t h e developnent of s u p e r c r i t i - cal flow and shock waves.

Maximum normal-force-drag ratio.- The va lues of (Cr/cd)max have been determined from the drag data shown i n f i g u r e s 5(c), 6 ( c ) , and 7 ( c ) and are presented as a f u n c t i o n of Mach number i n f i g u r e 10 . t h e values of Reynolds number f r m model scale to f u l l scale for a l l Mach numbers presented .

A t a c o n s t a n t Mach number, (Cn/cd)max for a l l the a i r f o i l s i n c r e a s e w i t h an i n c r e a s e i n

1 0

The i n c r e a s e s i n t h i s parameter wi th Reynolds number are t h e r e su l t of both a lower s k i n - f r i c t i o n drag f o r a given normal-force c o e f f i c i e n t and a de lay of t u r b u l e n t boundary-layer s e p a r a t i o n t o higher normal-force c o e f f i c i e n t s a t t h e h ighe r Reynolds numbers. The increments i n (Cn/cd)max due t o Reynolds number a t a given Mach number less than 0.70 are g e n e r a l l y q u i t e d i f f e r e n t f o r each of t h e a i r f o i l s . The i n c r e a s e s i n (Cn/cd)max due t o Reynolds number are gen- e r a l l y small fo r a l l the a i r f o i l s a t Mach numbers above about 0.70 where supe r - c r i t i ca l flow e f f e c t s predominate over the v iscous e f f e c t s . This predominance can be i l l u s t r a t e d by showing t h a t t he knee of the d rag curve is c o n t r o l l e d by t h e developnent of s u p e r c r i t i c a l flow and n o t by s e p a r a t i o n . By drawing a t angen t to t h e drag curve of the NACA 001 2 (Oo Tab) a i r f o i l a t a Mach number of 0.77 and a Reynolds number of 6.5 x l o 6 ( f i g . 5 ( c ) ) , t o occur a t a normal-force c o e f f i c i e n t of about 0.26. The p res su re d i s t r i b u t i o n ( f i g . 1 6 (j)) corresponding to a higher Cn (0.36) i n d i c a t e s supercr i t ical f law

on t h e upper s u r f a c e from about 3- to 50-percent chord and no c h a r a c t e r i s t i c s of sepa ra t ion . (Minor s e p a r a t i o n could e x i s t a f t of 93-percent chord. )

(Cdcd)max is shown

Drag divergence.- The d rag c o e f f i c i e n t s a t cons t an t va lues of were cross p l o t t e d to o b t a i n t h e drag-divergence Mach numbers. The normal-force c o e f f i c i e n t s corresponding t o t h e drag-divergence Mach numbers a t model-scale and f u l l - s c a l e Reynolds numbers are presented i n f i g u r e 11 . The e f f e c t of Reynolds number on Mdd 0.7 for a l l t h e airfoils. The drag divergence a t high normal-force c o e f f i c i e n t s (Cn > 0.7) may be c o n t r o l l e d more by shock-induced boundary-layer s e p a r a t i o n than by s o n i c f l o w ( r e f . 1 3 ) moving aft of t h e a i r f o i l crest. An a n a l y s i s of t h e pressure d i s t r i b u t i o n s for t h e s e a i r f o i l s a t model-scale Reynolds numbers s u g g e s t s t h a t t h i s is t h e case. This exp lana t ion f o r d rag d ivergence would be c o n s i s t e n t w i th the g e n e r a l l y l a r g e r e f f e c t of Reynolds number on a t t h e high va lues of divergence r e s u l t i n g from flow sepa ra t ion . The drag-divergence Mach number a t zero normal-force c o e f f i c i e n t of t h e NACA 001 2 (Oo Tab) and t h e SC 1095 a i r f o i l s i s unchanged by t he i n c r e a s e i n Reynolds number, but t h a t of the SC 1095-R8 a i r - f o i l is reduced by a b o u t 0.02 because of t h e i n c r e a s e i n Reynolds number. Fig- u r e 11 also i n d i c a t e s that t h e drag-divergence Mach numbers corresponding t o a l l normal-force c o e f f i c i e n t s between about -0.1 and 0.3 of t h e SC 1095-R8 a i r f o i l are lower a t t h e f u l l - s c a l e Reynolds numbers than a t t h e model-scale Reynolds numbers. A s t u d y of t h e curves of Cd a g a i n s t M for t h i s range of normal- f o r c e c o e f f i c i e n t s ( f i g . 22) i n d i c a t e s t h a t t h e drag c o e f f i c i e n t s a t a Mach number of 0.78 a t t h e f u l l - s c a l e Reynolds numbers would have to be lower f o r t h e

to be as high as t h a t a t t h e model-scale Reynolds numbers. Although Reynolds number has e s s e n t i a l l y no e f f e c t on Mdd of t h e NACA 0012 (Oo Tab) a i r f o i l a t normal-force coeff i c i e n t s up to about 0.75, t he d rag c o e f f i c i e n t a t Mdd is reduced a t t h e f u l l - s c a l e Reynolds number f o r normal-force coef-

dec reases f i c i e n t s g r e a t e r t h a n 0.2 ( f i g . 12 ) . Figure 12 shows t h a t

wi th i n c r e a s i n g Reynolds number f o r almost a l l normal-force c o e f f i c i e n t s f o r both cambered a i r f o i l s .

is g r e a t e s t f o r normal-force c o e f f i c i e n t s above about

Cn because boundary-layer t h i ckness would be crucial to d rag

CdMdd

11

CONCLUSIONS

An i n v e s t i g a t i o n has been conducted i n t h e Langley 6- by 28-Inch Transonic Tunnel to determine t h e two-dimensional aerodynamic c h a r a c t e r i s t i c s of t h r e e h e l i c o p t e r rotor a i r f o i l s a t Reynolds numbers from model scale to f u l l scale a t Mach numbers from about 0.35 t o 0.90, The a i r f o i l s inc luded i n t h i s i nves t iga - t i o n were t h e NACA 0012 (Oo Tab) , t h e SC 1095-R8, and t h e SC 1095. Ana lys i s of t h e t es t da ta has r e s u l t e d i n t h e fo l lowing conclus ions :

1 . The maximum normal-force c o e f f i c i e n t s of t h e NACA 0012 (Oo Tab) and t h e SC 1095-RE a i r f o i l s i nc reased wi th i n c r e a s i n g Reynolds number a t a l l Mach nun- b e r s presented . The maximum normal-force c o e f f i c i e n t s of t h e SC 1095 a i r f o i l a t f u l l - s c a l e Reynolds numbers were about t h e same or s l i g h t l y lower than t h o s e a t model-scale Reynolds numbers.

2. The pitching-moment c o e f f i c i e n t s about t h e aerodynamic c e n t e r of t h e s e t h r e e a i r fo i l s were e s s e n t i a l l y unchanged wi th i n c r e a s e s i n Reynolds number a t a l l Mach numbers presented .

3 . A t a c o n s t a n t Mach number, t h e drag c o e f f i c i e n t a t z e r o normal f o r c e Cdl0 a l l tes t Mach numbers. A t a c o n s t a n t Mach number up to about 0.65, t h e v a l u e s of Cdl0 of the NACA 0012 (Oo Tab) a i r f o i l decreased wi th i n c r e a s i n g Reynolds number. The d i f f e r e n c e i n Cdl0 of t h e NACA 0012 (Oo Tab) a i r f o i l a t model- scale and f u l l - s c a l e Reynolds numbers was small a t a Mach number of 0.34. Th i s d i f f e r e n c e g r a d u a l l y d isappeared as t h e Mach number inc reased to 0.65.

of both cambered a i r f o i l s decreased wi th i n c r e a s i n g Reynolds number f o r

4 . For a l l test Mach numbers presented , t he maximum normal-force-drag ra t ios of t h e s e t h r e e a i r f o i l s a t t h e f u l l - s c a l e Reynolds number were h igher than those a t t h e model-scale Reynolds number.

5 . I n gene ra l , t h e e f f e c t of Reynolds number on drag-divergence Mach number was g r e a t e s t f o r normal-force c o e f f i c i e n t s above abou t 0.7 f o r t h e s e t h r e e a i r - f o i l s . The drag-divergence Mach number a t zero normal-force c o e f f i c i e n t of t h e NACA 0012 (Oo Tab) and t h e SC 1095 a i r f o i l s was i n s e n s i t i v e to Reynolds number changes, bu t t h a t of t h e SC 1095-R8 a i r f o i l was reduced by a b o u t 0 . 0 2 because of t h e change from model-scale to f u l l - s c a l e Reynolds number.

Langley Research Center Nat iona l Aeronaut ics and Space Adminis t ra t ion 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 Wake- Survey 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- and Lift-Interference Corrections for Slotted Tunnels Obtained by the Substi- tution of an Equivalent Hmogeneous Boundary for the Discrete Slots. NACA FtM L53E07b, 1953.

9. Loftin, Laurence K., Jr.; and Smith, Hamilton A.: Aerodynamic Charac- teristics 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.: 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

TABLE 1.- DESIGN COORDINATES FOR

NACA 001 2 (Oo Tab) AIRFOIL

and ordinates given i n percent a i r f o i l chord 1 L

S t a t 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 S C 1095-R8 AIRFOIL

and ord ina tes given i n percent a i r f o i l chord 1

~~

S t a t i o n

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 f ace

-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 su r f 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

S ta t ion

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 su r face

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 su r f 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

1 5

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 O R I F I C E LOCATIONS

FOR S C 1095-R8 A I R F O I L

[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

[Loca t ions given i n percent a i r f o i l chord]

Upper-surf ace s t a t i o n

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 s t a t i o n

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

22

d I

il

b

A I I

I I I I I 1

V

3r a I

W

h * a,-

I

,953 (.375) >

a.

>

9.398 (3.700)

5.080 ( 2 .O 0 0) 2 -540 ( 1.000)

.635( .250)

1 f T

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

I 4 -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

ul JJ c aJ -4 V

-rl W W

i! 8 8

E E

W I 4 a

30

. 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

-?

E j L I O o w w L I

I -4 r c a m E2 0 0 el-

3

m

h

\o

'f.

u 5 o c

w u o m e . 0 - -4 P u a 4E-r .-I

& O $ E ccv 0 7 0

LIO

r n w 5 0

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.- E f f e c t of Reynolds number on s e c t i o n drag c o e f f i c i e n t a t z e r o normal force of NACA 0012 (Oo Tab), SC 1095-R8, and SC 1095 a i r fo i l s .

c U

m

m

h

‘9s

.n

9

m

7 E

.d I XLA m m E o W o u

VI c O r

u a a a -4 t3 LI d o

c 0-4

35

I P a LI 'D c JJ -4 3 . rn

. C U - o a a C B 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



Figure 12.- Effect of Reynolds number on sec t ion drag c o e f f i c i e n t at drag- divergence Mach number of NACA 0012 (00 Tab), SC 1095-R8, and SC 1095 airfoi ls .

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.- E f f e c t of Reynolds number on chordwise p re s su re d i s t CA 0012 (Oo Tab) a i r f o i l . Symbols with p lus s i g n i n s i d e

su r face . R given i n mi l l i ons .

r i b u t ion i n d i c a t e

-~

-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 1 3 . - 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.

F igu re 14.- E f f e c t of Reynolds number on chordwise pressure d i s t r i b u t i o n of SC 1095-R8 a i r f o i l . Symbols wi th plus s i g n i n s i d e i n d i c a t e lower su r face . R given i n mi l l i ons .

4 0

-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 I 1 I I I I I I I I I El

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

X I C

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 . 3 4 ; R - 2.9 x l o 6 .

Figure 16.- Effect of angle of attack on chordwise pressure d i s t r i b u t i o n of NACA 0012 (Oo Tab) a i r f o i l . Symbols with plus s i g n ins ide ind ica te lower surface.

44

-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 l o 6 .

Figure 1 6 .- 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

c M 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 l o 6 .

Figure 16.- Cont inued .


49

(9) M = 0 .64; R = 5.5 x l o 6 .


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 l o 6 .


51

52

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


x/c

(1) M = 0.77; R S, 6.5 x l o 6 .

Figure 7 6 .- Cont inued .

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

1 0

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

(k) M = 0 . 8 3 ; R = 6.6 x l o 6 .


54

(a) M - 0 .33; R *( 3.0 X l o 6 .

Figure 17 . SC 1095

- Effect of angle of attack on chordwise pressure d i s t r i b u t i -RE a i r f o i l . Symbols with p l u s s i g n i n s i d e i n d i c a t e 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 e 9 0 -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.


57

58

M C 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 I O 6 .


59

C n 4.8 a M

x/c

(e) M 4 0.54; R - 4.8 x l o 6 .

F igu re 17.- Continued.

60

W C a C -4.8 P*onicl = c a W

x/c

(f) M = 0.58; R - 5.2 x I O 6 .


61

-4.8

-4.4

-4.0

I I I I I I I I I I I I I 8

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 . 6 3 ; R S, 5 . 5 x l o 6 .


62

.L .2 .s .4 .5 .6 .7 .8 .9 1.0 x/c

(h) M 4 0.68; R S, 5.9 X l o 6 .


63

X I C

(j) M = 0 . 7 8 ; R = 6.6 x l o6 .


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 1 8 . - E f f e c t of angle of a t tack on chordwise pressure d i s t r i b u t i o n of SC 1095 a i r f o i l . Symbols with p l u s s ign i n s i d e i n d i c a t e lower sur face .

67

68

cP

(I M c cn (I -4.8

x/c

(b) M - 0.38; R - 3 . 4 x l o 6 .


-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 l o 6 .


69

x/c

(a) M = 0.48; R = 4.3 x l o 6 .

F igu re 18.- Continued.

3

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

(e) M = 0.54; R = 4.7 X l o 6 .


71

cP

72

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


x/c

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


73

-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.

F i g u r e 18 .- C o n t i n u e d .

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 l o 6 .


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 l o 6 .


77

SeP (a,)

( x IC

16

12

8

4

1. 0

. 8

sep . 6

. 4

R ~ 3 . 0 - 4.8 x 10 6

p Q I I

0 I I

R 3.0 - 4.8 x 106

Q \

\ \ 0 \ \ ‘4

Figure 19.- E f f e c t of Mach number on angle of attack a t which boundary-layer separation f i r s t occurs and on separation point of SC 1095-R8 a i r f o i l .

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

' 0 0 . 8 8 6.9 -2 .1 -2.1 -0.04

Figure 20.- Effect of Reynolds number on shock wave l o c a t i o n on NACA 0012 (Oo Tab) a i r f o i l .

R g iven i n m i l l i o n s . Symbols with plus s i g n i n s i d e i n d i c a t e lower sur face .

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

SC 1095-R8 SC 1095 Cp, sonic

/ xlc x IC I

\ / 0 0.0131 A 0.0120 .

\ 0 0.0369 I 0.0463 I b 0.0956 I 0 0.0852

1

n

0 W

0 0

Q 0 -----~ ~ _ _ ~ ~

0 \ U

0 0

- U b . b

b

- U 0 b b y 0 0

I 0 L h \ I I .' b

e** A A A A A A

- m o 8 \

\

D A A

a

-

C = 0.0 R z 2 . 9 - 6.6 x 106

Figure 21.- Effect of Mach number on sane lower-surface pressure c o e f f i c i e n t s forward of 10-percent-chord s t a t i o n of SC 1095-R8 and SC 1095 a i r f o i l s .

81

C d

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

Figure 22.- E f f e c t of Reynolds number on v a r i a t i o n of s e c t i o n d rag c o e f f i c i e n t w i th Mach number of SC 1095-R8 a i r f o i l .

82

1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. NASA TP-1701 AVRADCOM TR 80-B-5

5. Report Date 4. Title and Subtitle 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 1 1 . 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 a l l 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 f u l l - S C ? l e 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 o f 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 2 2 1 61 NASA-Langley, 1980

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