Document
., ., ,,., > .’i “ NATIONALA AQVISORY COMMITTEE FOR AERONAIJT~~._ -— ..--— .- . .
WAIUWIl? REPOIW
?
/’ ORIGINALLY ISSUED Septenber 1944ae Advanoe Confidential ReportL4122 SCAIJ2+EFFECT TEsrS IN A~TUNKELOF THE NACA 653-418, a= 1.0 AJRF03Z SEUl!ION WITHo l 20-KCRF033.FCHORD sPIzTFJxJ By WarrenA. TuckeremdArthur R. Wallace Langley Memorial. Aeronautical La330ratory Langley Field, Va.
.
,.
.. ”-” .,-- ,..
., ,! . .
NAtML : ,..
,,.
‘,. - ,.
LN A C A L.I?NL4RI” ..
LANGLEY MEMORIAL..PL2RONATJTI(XL WASHINGTON L.ABoRtlrroRY Land6y Field, Va NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified.
Some of these reports were not tech- nically edited. AU have been reproduced without change in order to expedite general distribution.
L - 128
.--, ”
./< .: w ,, .* .
31176001876698 4.,. .% .
P
..
. .
NACA ACR NO. 422 : ‘.
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS SCALE-EFFECT TESTS IN A TURBULENT TUNNEL OF 1.0 AIRFOIL SECTION .
TEE NACA 65+8, a = WITH 0.20-ATRFOIL-CHORD SPWI’ FLAP By Tarren A. Tucker and Arthur R. Wallace SUMMARY The effect of Reynolds number on the aerodynamic characteristics of a low-drag airfoil section tested under conditions of relatively high stream turbulence was determined b tests in the LVAL 7- by 10-foot tunnel of the NACA 653- 1S, a = 1.0 airfoil section with a $ split flap having a chord 20 percent of the airfoil chord. e Peyn~lds number ranged from 0.19 to 2.99 x 10 ; the Xach number attained was never greater than 0.10.
The data are presented as curves of section angle of attack, section profile-drag coefficient, and section pitching-moment coefficient against section li~t coefficient for various flsp deflections.
The maximum lift coefficient increased with Reynolds number. Deflecting the flap added an Increment of maximum lift coefficient that seemed to be almost con- stant at all Reynolds numbers. The slope of the section lift curve with flap deflected showed no cons~stent variation with Reynolds number, although the slope of the section lift curve for the ~laln Mrfgil Increased up to a Reynolds number of about 1.0 x 10 and then remained nearly constant up to a Reynolds number of ths llmit of the tests.
about 3.0 x 10 For flap deflections abo~e 15°, the slope of the section lift curve decreased with increase in flap deflection.
The section drag coefficient with flap deflected remained almost constant with Reynolds number, although t~ section drag ooefflcient for the plain airfoil .
decreased up to a Reynolds number of about C.S x 10b and then remalned,nearly constant to a Reynolds number of about 3.0 x 10°.
I?ACAACR No. 4122
2 t me pitching-momer~t-coefficient slope with flap deflected was erratlo, but the pitching-moment-coefficient slope for the plain alrfoll became slightly more negative with inoreaslng Reynolds number.
INTRODUCTION Scale effect on low-drsg airfoils has re ularly been % determined at ??eynolds numbers above 3.0 x 10” in the liACAtwo-dimensional low-turbulence Fressure tunnel (designated TDT). Tests were recently made in the TDT, In the NACA two-dimensional low-turb~lence tunnel, and In the LMAL 7- by 10-foot tunnel to determine scale and turbulence effects on the lift and drag characteristics of a ty~ical low-drag airfoil section over a tide rmge of Reynolds number !refercnoe 1).
!l!he object of the present investigation was to find the effect of Reynolds number on the aerod’ynamio characteristics of a typical low-drag fla~ped airfoil section tested under conditions of relatively high stream turbulence. The NACA 65 -111.8, a = 1.0 airfoil section equipped with a split flb having a chord 20 percent of the airfoil chord (0.20c) was tested in the LMAL 7- by 10-foot tunn 1 over a range of Reynolds number from 0.19 z tO 2.99 x 10 .
TWO models of ?-foot span with chords of 1 foot and )-L feet were tested. Both models were built af laminated wood ~’flth suitable steel reinfcmcemsnts and were shaped to the NACA 65z-~1.18 profile.
Ordinates follthis section were derived b; the methods of reference 2 end are given in table 1. Both models were carefully finished end were ~olished just before testin~.
A 0.20c split flap was tested on each model. lhe flaps were made of sheet steel and were fmmed to the airfoil contour.
The airfoil section with the flaF is shown in flgurel.
I — NACA ACR No. L4122 ~. ~~j ..
.
.. V.-. -----
Teata .
. .
The models were mounted vertically in the tunnel i so that the test section was spanned completely except for a small clearanoe at each end. The models were rigidly attached to the balance frame by torque tubes extending through the tunnel walls. The angle of attack was set by rotating the torque tubes by means of a calibrated electric drive. This Installation is thought to approxi- mate closely two-dlmenstonal flow, thus making it possible to detemine the seotion oharaoteristics of the models being tested. This setup is desoribed In reference 3.
I?achmodel was tested at dynamio pressures of 1.02, 4..09, 9.21, and 16.37 pounds per square foot, which correspond to tunnel airspeeds of approximately 20, 40, 60, and 80 miles per hour, respectively. These air- speeds correspond to test Reynolds numbers of 0.19, 0.37, 0.56, and 0.75 x 106, respectively, for the mgdel of l-foot ;hord and 0.75, 1.50, 2.2&I, and 2.99 x 106, respectively, for the model of k-foot chord. The by 10-foot tunnel is 1.6.
turbulence factor of the lXAL 7- Although the data are presented for various test Reynolds numbers, the corresponding effective Reynolds numbers can be obtained by multiplying the test R6ynolds numbers by the turbulence factor. The highest Mach number reached was 0.10, so that no effect of Mach number on maximum lift coefficient is thought to be present (reference 4).
At eaoh tunnel airspeed, each model was tested both as a plain airfoil and wlt~ the flap attached and deflected 15°, 30°, and 60 . The flap deflections were set by means of templets and were checked after each test. The flap was sufficiently braced so that no perceptible deflection occurred under load.
Balance readings were used to measure llft, drag, and pitching moment, except for the drag of’the plain airfoil. Because of t- insensitivity of the tunnel balance system, particularly at low speeds, the drag of ,.
the plain airfoil was obtained from wake-survey tests.
The angle of attack ranged from a negative angle through the stall for each test. In most cases, readings were taken at 2° intervals, with 1° increments near the stall.
— .—- — .- -—- -- .
\~J NACA ACR NO. 4122 PRESENTATION OF RESULTS Coefficients and Symbols The test results are presented In the form of standard nondimensional sectton coefficients.
me coef- ficients and symbols used are defined as follows: section lift coefficient (L/qc)
Cz
cd sectfon profile-drag coefficient (do/qc) o seotion pitchl -moment coefficient about quarter-
c%/b
Tm/qc2) chord point maximum section lift coefficient c 2- where 2 section 11.ft do section profile drag m section pitching moment about quarter-chord point free-stream dynamic pressure q $m (?
c airfoil chord, feet v airspeed, feet per second mass density of air, slugs per cubic foot P“ and R Reynolds number (Fvc/p) M ?Jachnumber (V/a) a speed.of sound (1129 fps) v5scosity of air, pound-seconds per square foot P a.
an~le of attack for infinite aspect ratio .6*..
---–f-lap--defleo tl-on-, -measured from - flap-retracted ..
“ position do &iao slope of lift ourve for infinite aspeot ratio Precision test results.- The experimental errors Aocuracy of In the results presented herein are believed to be within the limits Indioated in the following table: n unlit of ~oouraoy R Ohord ox cd. at ol =
max ‘%/l+
(ft)
I
I x 0.19 20, 10 *(?.05 10 to.o15 *.08 *.C3 *.O1O .3 2 *.06 ?.G2 too *.015 .75 took t.G15 *75 1.50 i.o t.C12 2.2L *.O ii *.oc)i t.009 *.03 *.G06 t.0006 I The averue errors are much smaller. With flap deflected, errors ma; be as muoh as three times the value~ given. - angle-of attack and flap deflection were held within The the following limits of aocuracy: a., degrees. . . . . . .“. . . . . . . . . . . . to.2 *O-2 degrees. . . . . . . . . . . . . . . . . . .
afs Wind-tunnel corrections.- The ltft coefficients are correoted Interference effects (reference 3).
me drag coeffi%en~~or the ~lain airfoil. whloh were obtalne~ from wake-survey test~, were corrected for blooklng as In reference 1.
No corrections to the drag and pitohing-moment ooefflolents have been detemnined for * two-dimensional foroe tests In the IJ4AL 7- by 10-foot “ tunne1.
-1 .6 DISCUSSION The curves of section angle ot attack, section profile-drag coefficient, and section pitching-moment coefficient against sectjon lift coefficient, for the various Beynolds numbers investigated, are presented in figure 2.
?itft. - Tha angle of attack at the maximum lift coeff~nt seems to Increase progressively with There is no scale effect on the angle Reynolds number.
of attack for zero lift, although there is an unexplained difference between the angles for ~~ro litt of the modols of l-foot and ~-foot chord. As shcvn by the curves of maxl?mumsection lift coefficient .~einst Reynolds number (fig. 3), the scale eff’ecton CZ is of the usual K.az form; that 1s, c1 increases with increasing R.
max Moreover, deflecting the split flap adds an almost con- stant increment of through the Reynolds number c2m= range. This effect ls.usual for a split flap (refer- ence 5). The scale effect on the slope of the lift curve wlthln the low-drag range Is given :n f5gure 4..
me slope af the lift wrve for the plain airfoil increases up to a Rey.mlds number of about 1.0 x 106 and then ~meins a-imst constant up to a Reynolds number , the limit of the tests.
of about 3.0 Y 10 With flap deflected, the slo~e IS erratic but approxinmtely con- stant with ReTnalds nm.ber. For flap deflections above 15°, the slope of the lift curve decreases with increase in flap deflection.
.- Th9 effect of Reynolds number on the section I)ra A profi e- rs& coefficient Is shown in figure 5.
The drag coefficients for the plain airfoil were obtained from wake-survey tests: the drag cosfficlents for the airfoil with flap deflected were abtained from force tents.
All drags were taken &t the angle of attack corresponding to the design lift coefficient (0.4) of the plain airfoil; this value corresponds to an angle of attack of about 1°.
For the plain airfoil, th9 dr~~ decreases sharply with increasing Reynolds number below a Reynolds number of about c.S Y 10~. Abcwe this Reynolds nurrber,the dreg remains nearly constant.
.
.- 1: NACA ACR NO. 4122 ‘f~j 7.
... -, - - For t~ ‘at rfof-l ‘wfth flap dafld’dtddj “W’la -results show no consistent variation of section profile-drag ooefficlent with Reynolds number.
In faot, It may be conoluded from these results that the sedti.on profile-drag ooeffiolent with flap deflected is, to a first approximation, inde- pendent of Reynolds number.
.Pitohingmoment.- The somewhat irregular curves of section pltchlng-moment coeffiolent at the lowest Reynolds numbers appear to be caused by the inaccuracy of the tunnel balanoe system at the low speeds.
ml s inadouraoy Is also shown by the large difference between the original and check testsat R = 0.19 x 106 (fig. 2(a))7 Accuracy at Reynolds numbers higher than 0.19 x 10 is ach better, as shown by the table in the section entitled “Frecislon.” The slope of the pitching-moment- ooefficient curve of the plain airfoil becomes slightl more negative with increase in Reynolds number (fig. 6~.
The pitching-moment-coefficient slope for the airfoil with flap defleoted varied with lift coefficient in such a way that presentation of the slopes”was not practicable.
CONCLUSIONS Scale-effect tests of the NACA 653-418, a = 1.0 air- foil section with a split flap having a chord 20 percent of the airfoil chord have been made In the LMAL 7- by 10- foot tunnel. The Reynolds number ranged from 0.19 to 2.99 X 106; the Mach number attained was never greater than G.1O.
From these tests, the following conclusions have been drawn: 1. The maximum lift coefficient Inoreased with Reynolds number. Deflecting the flap added an increment of maximum lift coefficient that seemed to be .al.most “ oonstant at all Reynolds numbers.
2. The slope of the section lift curve with flap deflected showed no consistent variation with Reynolds number, although the slope of the seotion lift curve for the plain airf~il increased up to a Reynolds number of about 1.0 X 10 and then remained nea?ly constant up to a Repolds number of about 3.0 x 106,-the limit of the tests.
I .—.
NACA ACR No. 4122 “~ 3. For flap deflections above 15°, the slope of the section lift curve decreased with increase In flap deflection.
4. The sectIon proffle-drag coefflclent w~th flap deflected remained almost constant with Reynolds number, although the section profile-drag coefftclent for the plain airfoil decreased up to a Reynolds number of about 0.8 x 106 and then rema~ned nearly constant to a Reynolds number of about 3.0 X 10 .
5. The slope of the Fitching-moment-coef ficient curve of the plain elrfoll becama slightly more negative The pitching-moment- with increase in Reynolds number.
coefficient slope for the alrfoll with flap deflected varied with lift coefficient In such a way that presen- tation of the slopes was not practicable.
Langley Memorial Aeronautical Laboratory National Advisory Committee For Aeronautics Langley Field, Va.
,.
.— . . . . . REFERENCES -.
1. Quinn, John H., Jr., and Tucker, Warrpn A,: Scale and Turbulence Effec!tson the Idft and Drag Charaoter- istios of the NACA 65 -418, a = 1.0 Airfoil Section.
NACA ACR ~Oa I@lJ 19 .
?4 2. Jacobs, Eastman N., Abbott, Ira H., and Davidson, l?? lton: Preliminary Low-Drag-Airfoilland Flap Data from Tests at Large Reynolds Numbers and Low Turbulence, snd Supplement. NACA ACR, March 194.2.
3. Wenzinger, Carl J., and Harris, Thomas A.: Wind- Tunnel Investigation Of an N.A.c.A. 23012 Airfoil tith Various Arrangements of Slotted Flaps.
NACA Rep.
No. 664, 1939.
4. Staok, John, Fedzluk, Eenry A., and Cleary, Harold E.: Preliminary Investigation of the Effect of Compres- sibility on the Maximum Lift Coefficient. NACA ACR, Feb. 1943.
5. Jacobs, Eastman N., and .Sherman,Albert: Alrfoll Section Characteristics as Affected by Variations of the Reynolds lhzmber. NACA RCP. NO. 5S6, 1937.
.,
!
HACA ACR NO. 4122 \~\ 10 a TABLE I OI?DINATES OF NACA 655-418, a = 1.0 AIRFOIL SECTION ---- -=..-. —.-. .. ------- -. .
,- [Statlons and ordinates in percent airfoil chord] Upper surfaoe Lower surface Station Ordinate Station Ordinate o 0 0 .28 1.L2 ~.g: -:.22 .50 1973 -1”4 2.21 l 9 -1.7 a 1: 3 x a 2.1 .10 2. 2 -2.36 5.z6 z h.& ..l$e -3.22 7.~2 . 7.~8 - .37 ? ii ?.62 10.38 7 L.41 7:9 14.6~ 15.36 :5s:3 -.
k 1-.67 9.0 20.5 25.2 i 2 9.91 ? ; l 72 - :;5 10.54 3;.: -6.65 .7 ?1 ? 10.94 -6.82 3;.~~ 11.4 0:12 -6.66 11.09 )L5.06 ?L 4 ,.. -6.71 “9 ~o 10,77 50 -6. 6 ~;. c)? m .20 z -5. 2 54.95 5 ~Cl+ .41 l 12 8 65:13 Jb -z .45 7;.:; 69.85 j: : :?
u i 74.:5 z 0:15 -1:74 ;.~~ 85.13 X5 ?:E7 -.9 w . Og 2:35 8 .91 i g.05 1.12 % 9 .95 t o 100 0 L.E. radius: 1.96 Slope of radius through end of chord: 0.168 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS .
% ?3 — z .
30° \ \ NATIONM AW1.YM 60” COMMllTEE FOR~ \ Figure e~-418, u=LO sechon with O.20c split F/op.
‘% l-.
.
IJ I 1111111111 kl I I I z 0, I I I I I l I II I I I I I I 1“[1 I I & “ IL /6 1 1 1 1 1 1 1 1 1 I I I I I UI I I J I I I I I-A! VJ I I I I I I I I I I I I I I 1A~ , I ! 1 1 1 1 1 1 I I 1 K 1 ,.” t [b) O/E -fiwt-chom’ MO&/. R= 0.37./06; M. 0.05.
(u) One-Foot - chord mode/. R= O.19./0 ~ h’= 0.03, F/gffre 2.. - Aerod numic F_gure 2.- Conifmued.
aectlon characterlstjcs OF NACA 65*-4/ u/rfo# WIi% o 0.20 C sD//t ,40a # -, --- -/ z o .
.
(cJ One-?bot-chord mcde/. R= 0.56 x/0’;t7= 0.08.
Section Itft coefficient,cz Figure 2.- Continued.
(d) Of@-Foot-c40rd mode/. R= 0.75 xIOfM= 0.10, F!quP6’2. - Cbnt/nued.
o > (-2 :2 +?/ ?J z o .
-.4 r’ * l-l NJ co I I I I I I /6 H-t-t-t I -4 -8 (e) Four- foot - chord model. R= 0.75” 10? M =0.03. CF) Fow+bot-chord model. R- 1.50 x/o”; M= 0.05.
Figure 2.- @ontlnuea!
Figure 2.- CWimued.
u z o .
l r’ — Is H —
R
i’
‘8
!
N) (g) Four- fbot-chom’ mode/. R= 2.24$ lo? P?=O. 08. (h) Four-F& -cho& z?wii?/. R= 2.W X/O’i w. O./0.
m .
Figure z .- Com%ued. Rqu* 2.- Coduded.
b, D- z o .
w NATIONAL ADVISORY COMMITTEi FORAERONAUTICS.
~
Peyhofds fiwnbe~p , -
NAC!A ACR No. L4122 Fig.
4a,b @)P/a/h airfoli F/gure 4- Scff/e effect m fi~+-curve slope of the NACA 653-4/8 ujrfulf Secf)on.
NATtONM ADVLSOQY COHH17TEE FM MRDNNJTICS.
.
NACA ACR No. L4122 Fig. 5a,b . . .. .
?J@6 b) P/u/h u/kfo/f fwuke-surve te,s(sj Q&?
F/gUr$P:;SCQfe effect on. drag c effic/e/)t ut the / llft CW?ff/c/COt of the h!A[A 65s -4/8 Y? a..rf /’ sectmn.
NATIONAL ADVISORY COMMITTEE fca AERONAUTICS.
,20 ./5 ./0
:$
$
::6 .05 .04 .03 .02 .0/ ./ .2 .4 .6 .8 LO 20 304.0 ~lo’ Reynolds numieGU (b)flup deflected (force test s]jcqal?
F/gU/-e 5- C@nc[uded.
!2 o .
J’ NATIONAL ADVISORY .
COHNITTfE PM AERONAUTICS.
J
2,0
30 4.0ds96
@y/7G’12?s Oumkq R
.— -.-—— —- - ~ ‘-.fgf ,-- .,-- .
,,. .
., ,.
,, . .