Scale Effect on Clark Y Airfoil Characteristics from NACA Full-Scale Wind-Tunnel Tests

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NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
REPORT No. 502
SCALE EFFECT ON CLARK Y AIRFOIL
CHARACTERISTICS FROM N.A.C.A. FULL-SCALE
WIND-TUNNEL TESTS
By ABE SILVERSTEIN
1934
For sale b7 the Saperlntendent or Docaments, WashInllton, D.C. • • • • • • • • • • • • • • • • • • • • • • Price 10 cents
AERONAUTIC SYMBOLS
1. FUNDAMENTAL AND DERIVED UNITS
Metric English
Symbol
Abbrevia- Abbrevia-
Unit Unit
tion tion
Length _______ l meter __________________ m foot (or mile) _________ ft. (or mi.)
Time _________ t second _________________ s second (or hour) _______ sec. (or hr.)
Force _________ F weight of 1 kilogram _____ kg weight of 1 pound _____ lb.
-
Power ________ P horsepower (metric) ______ ---------- horsepower ___________ hp.
Speed _________ V {mkielotemrse tperesr pseecr ohnodu _r_ ___________ mk.p.p..hs.. fmeeilte sp epre rs ehcOoUnLd _______________ fm.p..pB..h .
2. GENERAL SYMBOLS
w,
Weight=mg v, Kinematic viscosity
g, Standard acceleration of gravity = 9.80665 p, Density (mass per unit volume)
m/s2 or 32.1740 ft./sec.2 Standard density of dry air, 0.12497 kg-m-'-s2 at
15° C. and 760 mm; or 0.002378 Ib.-ft.-4 sec.2
m, Mass = W
g Specific weight of standard" air, 1.2255 kg/m3 or
II
I, Moment of inertia = mk2 (Indicate axis of 0.07651 lb./cu.ft.
•
radius of gyration k by proper subscript.)
Coefficient of viscosity
JJ.,
3. AERODYNAMIC SYMBOLS
S, Area ~ID' Angle of setting of WlllgS (relative to thrust
SID' Area of wing line)
G, Gap Angle of stabilizer setting (relative to thrust
b, Span line)
Chord Q, Resultant moment
C,
n,
b2 Resultant angular velocity
Aspect ratio
S' Vl
p- , Reynolds Number, where l is a linear dimension
V, True air speed
JJ. (e.g., for a model airfoil 3 in. chord, 100
q, Dynamic pressure = ~p \l m.p.h. normal pressure at 15° C., the cor
{s responding number is 234,000; or for a model
L, Lift, absolute coefficient C. of 10 cm chord, 40 m.p.s. the corresponding
-::s number is 274,000)
D, Drag, absolute coefficient OD Center-of-pressure coefficient (ratio of distance
of c.p. from leading edge to chord length)
D., Profile drag, absolute coefficient OD, ~S Angle of attack
=
Angle of downwash
Induced drag, absolute coefficient OD, - ~ Angle of attack, infinite aspect ratio
Angle of attack, induced
Parasite drag, absolute coefficient OD - DSp Angle of attack, absolute (measured from zero
• q
lift position)
0, Cross-wind force, absolute coefficient 00 = q~ Flight-path angle
R, Resultant force
REPORT No. 502
SCALE EFFECT ON CLARK Y AIRFOIL
CHARACTERISTICS FROM N.A.C.A. FULL- SCALE
WIND-TUNNEL TESTS
By ABE SILVERSTEIN
Langley Memorial Aeronautical Laboratory
72254-34
NATIONAL ADVI ORY COMMITTEE FOR AERONAUTICS
HEADQ ARTERS, NAVY BUILDl G, WASHINGTON, D.C.
LABORATORIES, LA GLEY FIELD, VA.
Created by act of Congress approv d March 3, 1915, for thc supcrvision and direction of thc scielliific
study of the problems of flight Its membcr hip wa increa 'ed to 15 by act approved March 2, 1929. The
memb 1'5 are appointed by the President, and scrvc as slich without compensation.
JOSEPH S. AMES, Ph.D., Chairman, WILLIAM P. MACCRACKEN, Jr., Ph.B.,
President, Johns Hopkins nivcrsity, Baltimorc, Md. Washington, D.C.
DA\'l0 W. TAYLOR, D.Eng., Vice Chairman, CHARLES F. MARVIN, Sc.D.,
Washington, D.C. United States Weather Burcau.
CHARLES G. ABBOT, c.D., HENRY C. PRATT, Brigadicr General, United Statcs Army,
ecretary, mithsonian Institution. Chicf, Materiel Division, Air Corps, Wright Ficld, Dayton,
LYMAN J. BRIGGS, Ph.D., Ohio.
Dircctor, National BlIl'cau of Standards. EUGE E L. VIDAL, C.E.,
BENJAMIN D. Fo LOIS, Major General, nited States ArUlY, Dircctor of Aeronautics, Dcpartment of Commcrce.
Chief of Ail' Corps, War Department EDWARD P. WAR ER, M .. ,
HARRY F. G GGE HElM, M.A., Editor of Aviation, Jcw York City.
Port Washington, Long Island, . Y. R. D. WEYERBACllER, Commander, niled I::>tatos uvy,
ER E T J. KING, Rear Admiral, nited States Na\'y, Bureau of Aeronautics, Navy Department
Chief, Bureau of Ael'Onautics, Navy Department. ORVILLE WRlGllT, Sc.D.,
CHARLES A. LINDBERGH, LL.D., Dayton, Ohio.
ew York City.
GEORGE W. LEWIS, Director oj Aeronautical Research
JOHN F. VICTORY, Secreta7'y
H";NRY J. E. REID, Engineer in Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.
JOHN J. IDE, 'Technical Assistant in Burope, Pari~, France
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REPORT No. 502
SCALE EFFECT ON CLARK Y AIRFOIL CHARACTERISTICS FROM N.A.C.A. FULL
SCALE WIND-TUNNEL TESTS
By ABE SILVERSTEIN
SUMMARY indica Led, however, thaL it is unusual to obtain the
same results from several tunnels, even when these
Te ts were conducted in the N.A.G.A. full-scale wind
fundamental similitude requirements arc satisfied.
tunnel to determine the aerodynamic chamcteristics of
Some of the more important sources of experimental
the OZark Y airfoil over a large range of Reynolds Num
discrepancies are wind-tunnel boundary interference,
bers. Three O1'rfo1ls of aspect ratio 6 and with 4-, 6-,
airfoil-support interference, and air- tl'eam irregulari
and 8:foot chotds were tested at velocitie between .~5 and
tiC's and asymmptl'irs.
118 miles pel' hour, and the ckaracterist1'cS wete obtained
As a result of the failure of wind-tunnel te ting to
Jor Reynold::; Numbers (based on the airfoil chonZ) in
fulfUl the exacting requirements of similarity in both
the range between 1,000,000 and 9,000,000 at the low
the flow and the test procedure, disagreements occur
angles of attack, and between 1,000,000 and 6,000,000
in published results purporting to give the experi
at maximum lift. With increasing Reynolds Number' the
mentally obtained characteristics of airfoils of the same
airfoil characteristics are affected in the following
section. These conflicting results from tests in numer
manner: The drag at zero l~ft decreases, the maximum
ous wind tunnels confront the designer with an arduous
lift increa es, the slope of the lift curve increases, the angle
task. The variety of data must not only be analyzed
of zero lift occurs at smaller negative angles, and the
and interpreted for application to the particular design
pitching moment at zero lift does not change appreciably.
problem, but it must al 0 be extrapolated to flight
The Gla?'!r Y airfm'l characteristic obtained from the
Reynolds umboI'. This extension of the data has
tests in the f11ll-scale tunnel are compated with those from
usually been nece sary because experimental informa
the variable-density and the propeller-7'esearch tunnels,
tion ha not been available ,tbove a Reynolds Number
and with the theoretical value. An analysis of the com
of about 3,000,000, whereas the night range lies between
parat1've experimental data indicates that the air stream
2,000,000 and 25,000,000. There is no exact and ra
of the full-scale tnnnel has a relatively low turbulence.
tional method for making a transformation from the
This injerence is substantiated by the close agreement
best wind-tunnel information to the desired flight
obtained between the characteristics oj airplanes measured
characteristics, although e:xperience serves a a useful
in the jull-scale tunnel and those jrom flight tests, and by
guide.
sphere drug measurements that show the t1lnnel has a
With the idea of belping the designer to span this gap
turbulence s1'milar to jree air. It is the'refore believed
between small-tunnel information and :£light conditions
that the effects of turbulence on the characteristics of an
the study of airfoil characteristics has been continued
airfoil tested 1"n the full- cale tunnel are small, and may be
in the .A.C.A. full-scale wind tunnel. lIere unique
neglected in applying the data to design.
equipment .is available for testing large size airfoils at
Reynolds umbers comparable with those of flight.
INTRODUCTION
The full-scale tunnel has a further advantage over
The aerodynamic characteristics of airfoils ascer smaller tunnels in that the full-seale-tunnel data on air
tained from different wind-tunnel investigations are planes may be directly compared with tho e obtained
frequently not in agreement. The reasons for these in night tests, thus di closing any disturbing tunnel
discrepancies are generally understood, having been effects and checking the wind-tunnel testing conditions
revealed partly by theory and part;ly through experi and technic.
ment. The compJete force equation, which includes Tests were therefore mnde in the tunnel to determine
the terms expressing dynamic imilitude, show theo the aerodynamic characteri tic of the Clark Y airfoil
retically that comparable wind-tunnel results hould over a large range of Reynolds umber. By tests of
be obtained when airfoils having similar surfaces are airfoils with the same a pect rntio and chords of 4, 6,
tested at the same Reynolds umber in wind tunnels and 8 feet at velocities from 25 to 118 miles per hour,
with like turbulences. Experimental research has the characteristics were investigated over a Reynolds
3
4 REPORT NA'l'lONAL ADVISOHY COMMI'lvl'EE FOR AERONAUTICS
LImber range from about 1,000,000 to 9,000,000, EQUIPME TA D AIRFOILS
although datt1, were not ecured above a Reynolds The .A.C.A. full-scale wind tunnel and equipment
umber of about 6,000,000 at maAimum lift. A por are described in reference 2. ince the general equip
tion of the e 1'e ult was u ed in an experimental veri- ment and apparatus used in these te t were essentially
the same a reported in the aforementioned reference, a
further description will not be given.
During the tests the airfoils were mounted in the jet,
a shown in figure 1, on supports that attach to the air
foils at the one-quarter-chord point, and tra.nsmit the
forces to the balance below. The mall diagonal
treclmline arms connected to the rear of the airfoil
erve to change the angle of attack by pivoting it
nbout the main support pin. The lower ends of these
diagonal arms arc attached to screw mechani ms by
mean of which the angle is adjuste 1 to within ± 0.05°,
The fairings over the airfoil supports arc not connected
to the balance but are independently upported at the
balance-house roof. The short exposed upper por
tions of the main supports have Navy no. 1 strut sec
tions, and taper to a cross section of about 1 by 3
inches where they connect to the airfoil.
Three melal lark Y airfoils with 4-, 0-, and 8-foot
t:bords and of aspecL rcltlo 6 were used. The airfoil
covering of }{6-inch aitUllillUJ11 sheeL was aLlacbed to a
rigid internal sLrucLure by means of flush cOlUltersllnk
screw . The spars were teel barns and the profile was
formed by aluminum ribs paced at 12-inch interval.
Acces to the airfoil support pins was provided by
FIGURE I.- The 6 by 36 airCoil mounled in the Cull-scale tunnel.
removable plates which were screwed flush with the
fication of the theoretical jet-boundary cOlTeetion for sUJ-face during the tests. Tapped opening for fitted
the elliptical-jet wind tunnel which has been reported eyebolts were spaced over the airfoil for a Wiachmen ts
in reference 1. when taking tare measurements. F lush screw plugs
FIGUItE 2.-.\ lare·Coree seL·up wiLb invcrled-6 by 36 airfoil.
SCALE EFFECT 0 CLARK Y AIRFOIL FROM N.A.C.A. FULL-SCALE WIND-TUN EL TESTS 5
were inserted in these opening during the regular force when applying the result to the airfoil. Interference
tests. The smooth aluminum surfaces of the airfoil drag for the 6 by 36 airfoil wa interpolated f['Om dn.ta
were covered with a. protective coa.t of varnish. The on the other two airfoils.
airfoils were manufactured under careful in pection so Static and dynamic pressure surveys were made
a to maintain the specified ordinates, H,nd were accu everal chord lengths ahead of the 4 by 24 and by
rately measured just before testing. The specified and 4 airfoils to determine the blocking effect of the air
measured ordinates are given in table 1. 0 appreci foils upon the tunnel stream. These surveys were
a.ble twists, deformations, or local irregularities changed made at a number of angle of attack between zero
the a.irfoil accura.cy during the period of the tests. and maximum lift. For the 6 by 36 airfoil the block
ing efrect was int.erpolat.ed f!'Om data on the other two
TESTS
aidoils.
CORRECTION OF DATA
The lift, dmg, a.nd pitching moments were measured
at ix speeds between 25 n.nd 118 miles per hour over a The uncorrected lift a.nd drag forces on the airfoils
range of angles of attack from - 8° to 24 These tests were measlll'ed on recording scales, and the pitching
0.
were made with the airfoils in an upright position in moment was computed by multiplying the lift and
the tunnel, and then repeated through an angle range drag forces by the proper lever arms. The observed
of -8° to 5° with the airfoils inverted. wind-tunnel data were then corrected in the following
Tare force on the supports were measured with the manner:
airfoils in the test position but supported independently (a) The first process in correcting the data was to
of the regular support and rigidly held in place by adjust the mea ured dynamic pre sures. The dy
aLL\"iliary cables (fig. 2). The tare-force mea urements namic pressure of the wind-tunnel jet is measured with
therefore include the in te rf erence of the airfoils 1I pon a manometer, which indicates the pres ure difference
the upports. Tare forces were mea ured for all Lhe bet.ween the return ptlSsage and the tesL chamber
airfoils a,t five angles of attack and at all I.e t speed . (reference 2). The dynamic pressure in the jet is
The interference or the supports upon the 8 by 48 obt.ained by a calibration. PrevioLls study has shown
airfoil was ascertained by adding duplicate support that this indicated velocit.y head, obtained from It
ing struts to the normal instaUa.tion (fig. 3). A these calibration with no body ill the jet, is in error owing
dummy strut were not connected to the airfoil or to t.he blocking action of the body in the air stream.
balance, any change in the mea.sured characteri~tic The blocking increases with the angle of attack; th
with the truts in place could be attrihuted to tbeir Reynold umber of the test are therefore slightly
interference. A imilar method was employed for the difl' rent at the low and high angle of attack. A
tesLs of Lhe 4 by 24 airfoil II ing, however, only a ingle full di cu ion of the correction it applied to the air
dummy support and dOli bling the interference enect foil data is giYen in reference l. The magnitude of
FIGURE 3.-Dummy supports added to the by 4 airfoil set·up Cor interCerence tests.
6 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
the blocking effect of the three airfoils is shown in drag values at zero lift, with and without the dummy
figure 4. support struts, showed that the supports exerted i:l.
(b) Tare force and moment coefficients were then large unfavorable interference in the upright tests,
computed and deducted from the gross force coeffi and a slightly favorable one when the airfoil was
cien ts to obtain net values. The tare drag is about inverted. In all cases for the upright tests the effects
2 percent of the minimum drag for the 8 by 48 airfoil became very small on both drag and lift above a lift
coefficient of 0.3.
+-
1 1,1. )1 (d) Upright and inverted tests on the airfoils indi
1_ (m~asLre~
--0-4 by;:4 Cla.. rk ) cated that the air stream had an initial downflow
- - 6by36 .. (interpolated)
- x- 8by48 (measured) r- angle; it was necessary to correct the characteristics
-I
- for this effect. In order to determine the magnitude
r- -x_
A of the air-stream angle, plots were made of the DjL
against OL for the upright and inverted airfoil tests
(fig. 6). The DjL ordinate between the two curves
is equal to 2 sin /3, where /3 is the air-stre.1m angle.
A check on the air-stream angle is possible by noting
-4 o 4 8 12 16 20 the separation of the upright and inverted lift curves.
Anqle of attock. deqrees
FIGURE 4.-Blocking corrections for the tbree airfoils tested in the full-scale tunnel.
./0
and 10 percent of the minimum drag for the 4 by 24 I~Lectlon
airfoil. The tare lifts and moments are negligible. .09 p-, ~
(c) Interference effects of the struts on the airfoils -Tunnel axis
1/
were then included. Figure 5 illustrates the inter
ference cau ed by two stru ts on the lower mface of ~1'-l·08 Air foil tested upright.-_. V
li§.07
the 8 by 48 airfoil. The effect on the drag i quite V
./'
V
~/
.5 I 05
/1 I I I I V
.4 ). I 1 I 04 .06 2 si/n3 =/30 =.4 03.0° 150 / « Airfoil fesied
/ "'No correc- . inverted - i"-
~V
/ lhioftn ctuor v~e t= .05 I 1 I
(S.3 / / rfeoqr uiinrteedr -r--. 03(5 o .I .2 .3 .4 .5 .6 .7 .8
/ V ference LiTt coeffiCient. CL
V J I FIGURE 6.- Metbod of obtaining air-stream angles from uprigbt and inverted t ts.
Reynolds umber, 6.12X10'; 8 by 48 airfoil.
~ b- - -- --7 ~ ~ ---0--Nsoermt-aul p_ I -I--. Since the separation of the upright and inverted lift
- x- Two dumn;r-
/ supports curves, when plotted as values of OL against (x, is due
added J to the air-stream angle, the value of the air-stream
---Correc ted
/
for inter-- deflection is equal to one-half the angle between the
(erence
two CLlrves. If the interference effects are not prop
-.1 V I I
erly accounted for, the value of the air-stream angle,
/
1 1 from the two methods, will not agree. The angles
-.2
-8 -6 -4 -2 0 2 4 6 determined by these two methods generally agreed
Angle of attock. degrees
within about 0.1 The average value was taken as
D.
FIGURE 5.-The effect of strut interference on the characteristics of the 8 by 48 the true air-stream angle, although no rational excuse
Clark Y airfoil wben tested uprigbt. Reynolds Number. 6.12XlO'.
can be offered for this practice, except that the prob
large in the region of zero lift, but decreases and be able percentage of error is reduced.
comes negligible at higher lift coefficients. The inter (e) The limited boundaries of the wind-tunnel jet
ference effect on the lift is negligible and within the are a source of error in ascertaining the characteristics
e>rperimen tal errol'. of any body tested therein. A correction for this
The support interference on the 4 by 24 airfoil h~td boundary interference was therefore applied to the
an effect similar to changing the camber of the au·foil. airfoil angle of attack and the drag coefficient. For
The angle of zero lift was changed by the interference these tests the correction factor was determined
when the airfoil was tested both in the upright and experimentally by an extrapolation of the airfoil data
inverted positions. A comparison of the measured to free air values. A complete description of this
SCALE EFFECT ON CLARK Y AIRFOIL FROM N.A.C.A. FULL-SCALE WIND-TUN EL TESTS 7
method with the values of the experimental and aspect-ratio characteristics by the following formulas:
theoretical corrections 1 is given in reference 1.
.. ao= a- ~.R(1+7) 57.3
:::.:,.22 - ,./ 08 OL2
u V / ODO= OD- 7rR(l + 0-)
.E \;)
..o.... V 1/ g. where
...... 20 07
o / V {; ao IS the angle of attack in degrees at which an
/
Q) V V 'tJ airfoil with infinite span would give the same
Q§, .18 / V 06~Qj lift coefficient as the airfoil tested in the
0·'tgSQJ) . 16 II- Ii-- - / / V/ ~ os.;·"2SI\.5. ODRO,' tthheet u panrsnopefeiLlce t- rdartaigo .c oefficient.
.;l2. r-- - V / / ~llJ 7, a ftoa ctaollro wco rfroerc ttihneg tchhea ningde ufcreodm a negllliep toicf aalt tsapcakn,
l." /4 '-- "-/ V 04(2
o loading to one resulting from the use of an
..... r::.
lJ 1/ V ~ airfoil with rectangular plan form.
~
lJ
r::. .12 / V 03 Qt:j IJ, a factor correcting the induced drag, to allow
..0;:: for the change from elliptical span loading to
lJ V <3
Q) one resulting from the use of an airfoil with
l.lo 02
8 rectangular plan form.
3 4 5 B 7 8 9 and where a, OL, and OD, are the corrected character
Aspect rotio, b~S istics for finite aspect ratio. The angle of attack, a,
FIOIIRF. 7. Correction factors for transforming rectaogulnr airfoils from finite to is in degrees. Values of 7 and (J are taken from figLU'e 7,
infinite aspect ratio. and are based on the assumptions of a theoretical
rectangular loading, and a value of 0.101 for the slope
(j) The corrected characteristics for the airfoils with
of the infinite-aspect-ratio lift CLU've. Experimentally
aspect ratio 6 were then transformed into infinite-
the rectangular airfoil did not have a loading identical
1 The corrections reported in tbis reference were from the results of tests at a to the theoretical, owing to jet-boundary effects and
Reynolds Number of 2,000,000. When tbe complete results of the airfoils at all
Reynolds Numbers were analyzed it was found tbat values were obtained for the velocity asymmetrie . This variation would require
Jet-boundary correction which were slightly difIerent from those reported, aDd the use of values for and slightly larger than tho e
approached even more closely the theoretical values given iu reference I. 'l'bese (J 7
corrected factors have heen applied to the present data. in figure 7. Since results were not available to indicate
1
-loa ./0 40
I I
C.p. I
-80 .44 .09 36~
g
1) <!'
I
.1co.. -60 2.0 .40 -c:: .08 I 32t
2248 ~~:C~Io:IIl-' -2240000 1I-II r-- I- I- q-.-p- --')l f-' ~1--:9,u: 1: -/- .... 1111...2.684 ~....23234286 .:" ~i'fQ.: -~.Ul~g0Q3~) ' ....00004576 ~---1---.~~- - ![Ii =-_t_-+ I ,.t~i -- -W- -LIIi a.tI. ~1II \\ \ l It ~-o- - - 222/6408 ~"-~QL1c\0151 i.
.~~~"~~."iIagl- 5..f. 2114B0206 . 'E"t¢~10u~LL~at -.:: ; 1 86400000 --II-----rTPr Ir]DI-IfI. , V-j j;"'-I.. .iI.Jc. --1I/-A' -irfCoI-DI;"i " l ~: C ~la r.kI ."\lIY\ ' .-..-...-. rS-ize: 81' x I4 8' I.....o462O8 s-:-~u.~~~:J. . .....00//2260480 C~80~l> C~~'la t ~ ...- -000..03212/ f~--+IL -,tIIII- - tI ,-/+'tI - 1IIII -III+ -I( '" tI /"" -/ ~£1 iI --"1V'VC/1~iII1I • I -I8448EO-'' <~~.'0-~o~IUgC.ISoC::.::..::ll: -
-4 <~c..:i '~'- _. - r-RTRe.NNs toaetdf mz ineax rFoim. 5lui.mfrt.: H6.R tle:e 5xs 13u0l8t-s6 x 10co-'r - -.2 ..'cgu.":. --.A3 ~ - _.1I I !II III /2
-8 , rected for tunnel-wall effect -.4 f-+- . -t-+- -/6
-8 -4 o 4 8 /2 16 20 24 28 32 ~ -;4 ~2 o .2 .4 .6 .8 lO /,2 1.4 /.6 1.8
Angle o( o/Iock, deqrpes. d Lift coefficient, C
L
FIGURE 8.-Characteristies of the 8 by 48 Clark Y airfoil at a Reynolds Number of about 6,000,000.
8 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
the pressure distribution over the airfoils in the tunnel, covers the change in the angle of attack for zero lift
this effect, which is small in magnitude, is noLi ncluded. with Reynolds Number. Results from the val'iabl<'
den ity and propeller-research tunnel , a well as a
RES LT
'd. I 10 +
I I I I
The corrected results are tabulated giving values of "'d -- I-
1
OL, Ci, OD, LID, and c.p. for the Olark Y airfoil with r.:J.IOO ~ 4 b; 24 dlar~ ~
'd o 6 by 36 :: I
aspect ratio 6, and values of Cio, ODO, and Om for the /), 8 by 48
Q)
airfoil with infinite a pect ratio. The e data for the (a.0 90 (Theoretical JJ II
three airfoils at all Reynolds umbers te ted are ~ ~ ~
presented in tables II to XX, inclusive. Value of ~..:::. 080 I- variabIJJen~if~ )u;LI ....
C.p. are given in percent chord. A typical pIoL of thE' ..o... 1 t I: 1 ' ~ 1 I. t" •
Q).070 t- ~~--f ~. ! -
data from table XVIII is given in figure & -I' Propeller -research funnel I
The curves ummarizing variations of the principl1l ti) 0 2 3 45 6 7-
ReynOlds Number
airfoil characteristics with Reynolds um ber are of
FIGUIlE ll.- Variation with Reynolds Number 01 the slope of the lift curve for the
I.B Clark Y airfoil (slope for an airloil of aspect ratio 6; at in degrees). Propeller-researcb
tunnel value Irom reference 3. Variable-density-tunnel value from reference 6.
"~'1 .7 ~ 41 b; 2~ ClarA Y 'l'heoretical \'alue from reference 5.
o 6 by 36 " theoretical value from reference 5, are also included on
~'I.6 /), 8 by 48 , "
tills figure. In a similar ml1nner, figure 11 pre ents
<;:: Variable - tjensity tunnel the change in slope of the lift curve with scale. The
'8Q ) /.5 , X I f.--
V I , -H-
/ V g/ v: V0 r-----Full scale tunnel Ic\.v ) 0110 - - !~ ooi 6·1 1 bbyyi 32~64 Cil a..r AI YI
'V 0 ~O J) c. ...: _ " 8 Ib-Ly -4J8 .. I
~ -Propeller-research tunnel 'C~_.OIOO I r Vanoble -denSity tunnel
/V 8'~ ........ ::::-- 1"1 I I I I I
2 3 4 5 6 G, ~.0090 - q, r
Reynolds Number a\c.1 ._ - PropelIlte.r.-.rIe-seLa1rc hJ ~t unnnel-
FIGURE 9.-Variation with Reynolds Number of maximum-lift coefficients for the (J0080 - 1--
Clark Yairfoil. Propeller-research-tllnnel value frolll reference 3. Variable-den l 1/1/11/1
sity-tunnel data from reference 4. ll. o ? J ./ 5 6 7
He'ynolds Number
particular intere t. Figure 9 shows the variation of FIGURE 12.-Variation witb Reynolds 1 umber 01 tbe Clark Y profile-drag coeffi-
the maximum lift coefficient for the lark Y airfoil cient at zero lift. Propeller-research-tunnel value from reference 3. Variable
density-tunnel value from reference 6.
over a Reynolds N um bel' range from 1,000,000 to
6,000,000. In this figure the results of Olark Y tests rtirfoil profile-drag coefficient at zero lift is shown on
figure 12 over a Reynolds umber range from 1,000,000
"-."..':- -6.B 1b ; 21 +-- to 9,000,000, and values from the variable-density and
:.::: 01 4 dlar~ ~ -
0~ -6.4 o", 86 , bbyY 34~8, f:'.. 'O~~~~~~r-.--.~I~.II-;_~l~l-d~LI.I~J~l~I_J~I~.l~l~
'l".. "" Full-scale" tunnel ,r=-=:::::.. n -j :1• P[ roflClela-drkr oqY oaltr rz:oe~rso 11ft
~ ~-6.0
I.'< .004 0 LLLL
g::o:: ~~ -5.6 ., ~o --:;::::: Z . . .. -ThIe oreticol fI---S-pk:l:taln:t e-- finw~c t-tf,hl-o -n_tu 0r-fborurt l-e+fnl-aIt -t--.·. --f-
..o... -5.2 Propellfr-research tunnel 0 '002rI---bro-u--n-dro-rfy-+ -laryl-e+r -~++~--H-r-+Ht-i~-~++H+H-IH+--Il---+-I--r~~l+-l4--W~ +
t ,.:' -Variable-density tunnel f---+-t--I-+-+-+-H-+-+-++l-H-l/), 8 by 48 Clark Y
~c: 3 456 7 8x10' t--+~I-t-+-H-++t+++++1 o0 46 bbyy2 346 .... ILIl
Reynolds Number O~~~~~~LL~LU~uiDIII~I~LI~I~I~.
I 2 3 4 5 6 7 8 9xl06
FIGURE IO.- Variation with Reynolds Number of the Clark Y angle of attack at ReynoldS Number
zero lift. Propeller-research-tunnel value from reference 3. Variable-density
tunnel value from reference 4. 'l'heoreticnl value from referenro 5. FIGURE 13.-Comparison of the Clark Y profile-drag coefficient at zero lift with the
skin-friction drag coefficient for a Oat plate having a completely turhulent hOlln<1Rry
in the I.A.O. . vari<tble-dellsity wind tUllnel OYer rt layer. C, for uil'foils based Oil actual surfat'tl area.
runge from 1,000,000 to 3,000,000 arc also given. A the propeller-research tunnels are again included. In
single point give the maximum lift obtained on the figure 13 the profile-drag coefficient rtt zero lift for the
Olark Y airfoil in the propeller-research tunnel at a airfoil i compared with the skin-friction drag coeffi
Reynolds umber of about 2,000,000. Figure 10 cient for a flat plate with turbulent boundary layer.