Section an”gfe of attack, &O) deg
314 REPORT 964—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
.
-24 -t6 -8 8 /’6 24 Section an”gfe of attack, &O) deg (a) LiItcharacteriatics.
FlcuRE5.-Acrodwamic chamcterktim of theh'ACA Wl-2l2airfoilwtioL FlnggedsymboIs denote leading.edge roughncaa.
.024 ,020 J .
T .; .0/6 Q L u o 0 .of2 g + ~ .008 .$ t % .004 H -.4 -!6 -12 -,4 0 .4 .8 /.6 -1.2 -.8 0 .4 .8 i.2 1.6 2?0 -.8 Sec+tom I!ft coefficient+, c1 (b) Dragcharacterietios.
~ FKwrRE5. —Conclndod.
EFFECTS OF REYNOLDS NUMBER 03’ THE CHAFL4CTERISTICS OF A’ACA 6-SERIES AIRFOIL SEOTIO?SS 315 (a] Lffc &~a~t~~is~f~~.
Flmggei s~boIs denote Mdirrg-alge roughness.
FIGCRK 6.—Aerod~smk chsrscteristfcs of the X-AC& 6+61S airfoiI section.
(b) Drag characcerfstics.
FLGCEE 6.—Conclndsd.
Section angle of of fack, u;, deg
REPORT 964—NATIONAL ADVISORY COMMIT.1’EE FOR AERONAUTICS -24 -/6 -8 8 16 24 Section angle of of fack, u;, deg (a) Lfft charscterfstfcs.
FIGURE 7.—Aerodynamfc characteristics of the NAOA 6W16 airfoil section. Flagged symboIs denote Ieading-edge roughness.
.036 .032 .028 ~. 024 *.
c u .- ~.
} .020 u m $0/6 * ;.0/2 .u08 .004 A -9 .Se ction Aft coe ftlc;ent, ct (b) Drag characteristics.
FIGURE7.—Concluded.
EFFECTS OF REYNOLDS N~I13ER ON TEE CHARACTERISTICS OF NACA 6-SERIES AIRFOIL SECI1710NS .31.7 ., -2+ -[6 -8 8 !6 24 Secfion angle of uffadr, 42 deg (a) Liit ebaracteristics.
FLGtTP.E &-&ro@namie chmartwistias of the NACA W aicfoii section. FI~ed ayrnbok denote Iaadiig-edge roughneaa.
REPORT 964—NATIONJL ADVISORY COMMITTEE FOR AERONAUTICS -24 /6 24 -/6 . -8 - 8 Sechon angle of af r’ack, Go, deg (n) Lift ehmacteristies.
FIGURE 9.—Aerodynamic charaeteristica of the NTACL4 6.%006afrfoil sedfon. Flagged syrnbrds denote Ic.adinwdgo rorrghncss.
.024 .020 J’ .
t .?.016 u & %_ Q 8 .0[2 ~ % ~ .oo8 .- % u G .004 .4 -1.2.
$.6 -1.2 78 -.4 0 .8 /.2 /.6 -1.6 78 74 0 .4 .8 k2 /.6 Section /if+ coe fficien ~ c1 (b) Drag characteriatiee.
FIGURE9.—Concluded.
EFFECTS OF REYNOLDS N7EWF?ER OX THE CHARACTERISTICS OF 2TACA 6-SERIES AIRFOIL SECTION’S 319.. _z... _ .0080 .0060 P (a] u .0040 .0080 - .0060 u ALA CA 63-006— .0040- & a NACA 63-009 Q NACA 63 L-0!2 ~ NAf24 633-018 .0020 - v NACA 63-209 — b NACA 631-212 e ..
AfACA 64-006 E NACA 64-009 v NACA 65-f?06— v ~ .0060 - *.
.; ? ~ — q + .
$ .0040 - ,, u o c1 [b] F .(7020 - + c -$ .0100 - i m ~ .$ .&380 .
s .0060 ?- .0040 .0020- .0080 [ * .0060 f~ h.
.0040 Reynolds numberj R (a) Efiect of thfchess.
(b) Effect of tbfcknessform.
[c) Effest of csmbw.
??IGCBE 10.—Vsrkdion of seetion mirdnmm d%= coefficient with Reymdds number for nine X-M2A E-serfes Sfrfotk of mrsfng thfclme~ thkknessforrn, snd cember. FIagged symbols denote leading-edgerooghucss.
936646-?il-22 REPORT 964—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS > The greater extent of the Iami.mtrboundary layer which The addition of standard roughness to the NACA 63-009 results as the point of minimum pressure is moved rearward section (@. 10 (c)) causes a large increase in the minimum is evidenced by the progressively lower minimum drag co- drag at all Reynolds numbers, but increasing the Reynolds efficients of the NACA 63–006, NAC!A 64-006, and NACA number has a favorable effect in reducing the drag. These 65-006 airfofl sections at a Reynolds number of 3.0X10! results are to be expected from a consideration of bounclary- (fig. 10 (b)). In general, moving the point of minimum layer theory for a fully developed turbuIent boundary lmycr.
pressure rearward has little effect on the sequence in which (See reference 4.)
the boundary-layer effects occur. The values of the drag Low-drag range,—Increasing the Re.ynoMs number from coefficient for these airfoils appear to be relatively insensitive 90X 10Gto 15.0X 10s resulted in the almost complete dis- to variations in the Reynolds number until a Reynolds num- appearance of the low-drag range of all the airfoils exccp t ber of the order of 15.0X 10s is exceeded. At higher Reyn- that of 18 percent thickness (figs. 1 to 9). The previously olds numbers, the rate of forward movement of transition d~cussed predominating influence of forwmd movement of appears to be reduced as the point of minimum pressure is transition at the higher Reynolds numbers, together with moved from 30 percent to 40 percent chord. Further rear- the influence of pressure gradient upon the Reynolds number ward movement of the position of miuimum pressure has at which this forward movement begins to predominant e, little efl’ect on the rate of the forward movement of tran- explains these drag results.
sition, at least for these thin airfoils. The data for the 9- Brag data outside the low-drag range,—The drag poltirs percenMhick-d3- series and 64-series airfoils show the same for the different airfoils (figs. 1 to 9) indicate that, for n given trends. lift coefficient outside the low-drag range, the clrag decreases An inspection of figure 10 (c) shows that the addition of as the Reynolds number is varied from 3.0X 10° to 9.0X 10°.
a small amount of camber to the 9-percent-thick and 12- Further increases in the Reynolds number, however, C1O not percent-thick 63-seriessections does not have any consistent seem to have any appreciable effect upon the chg. Vmia- effect upon the value of the minimum drag between Reynolds tions in the airfoil design parameters appcm to have no con- numbers of 3.0 X’108 and 9.0X 10s. Increases in the Reyn- sistent influence upon the effect of R.eynolds number on the oIds number beyond 9,0X 10B, however, appear to cause drag outside the low-drag range. Although roughness in- more rapid forward movement of transition for the cambered creases the drag greatly in this region, the vtdue of the drag airfoils than for the symmetrical airfoik. Only two cam- for the rough-surface condition seems to be rcl~tive]y insen- bered sections were tested, however, and this trend is there- sitive to Reynolds number as shown by the clata for the fore not very well established. NACA 63-009 section (fig. 2).
‘4 .100 ,120 (!$ o NAcA 63-006 b (J u IVACA 63-009 “; ./00 0 ~ACA 631-012 — A AIACA 633- 0/8 : NACA 63-20.9 NACA 64-006 .— f . fzo ~ NACA 64-009 w b v A(ACA 65-006 y \ ~ — — ~ — - * — [b] ~ .100 k * $ ,.120 ‘ t % .100 % .120 6.0 9.0 /5.0 20.0 250 30.0 40.0 50.0 60.0 70.OX 106 Reynolds number, R (a) Effect of thicknms.
(b) ENect of thicknessform; (c) l?ffeut of camber.
FIGURE11,—Varietmn of ecctlon lift-curve slope.with Reynolds number for nine N’AC.4 6-eories airfoilsof varying Chlcknws,thicknessform. m.1 c~mbw f?ltggc 1 symbols denote leading-edgeroughness.
EFFECTS OF REYXOLDS 3S~ER ON THE @RACTERLSTICS OF NACA 6-SERIES AIRFOIL SECTIONS .— LIFT variation -ivit.hReynolds number and, therefore, are not presented as a cross plot against Reynolds number. The The important cha.r~cteristics associated tith the Mt.
values of the section lift-curve slope and ma.xinmm section cur-r-e are the angle of zero Iift, Iift-curve slope, and the ma.ti- lift coefficient are presented as functions of Reynolds num- mum lift coefficient. In order to facilitate the analysis of ber in figures 11 and 12.
Lift data presented in figures 1 to 9, va.lues of these pmam- Lift-curve slope.—The lift-cur-i-e slopes mere obtained eters -weredetermined from the test data at the six Reynolds from tie best representative straight line through the numbers between 3.0 X 10sand 25.0X 10G. The values of the e.sperimental-data points in the m@e-of-attnck range of 4° angle of zero lift of the cambered airfoils showed almost no 1.8 0 ALAcA 63-006 D ALAcA 63-009 O AfACA 63%-012 1.6 ~ & NACA 633-018 ; .
s / c LL :’ ;$ /.4 —< ~ k u) o u / t ~ !s e z u 1.2 u ; G ..
=s 1.0 (a] o .8 L 40.0 9.0 15.0 20.0 250 30.0 50.0XI06 6.0 la? 3.0 Reynolds numberj E (a) Effect of thickness.
FIG= IZ-Vsrfatfon of rnsximum sestion Uft eoeffictit with ReynoIde nnmber for nine A’AC& 6-serfes akfoii ofwwing thickness,thicknessform. and csmber. Flogged symbols — denote Iesding+dge rouehness.
/.4 O NACA 63-006 q NACA 63-009 N ; V NACA 64-O(.W G v NAGA 65-.906 >- $ /.2 $ u z k * ~ ; Lo % g & ..
(w ‘i’ 4 .8 6.0 so 15.0 20.0 25.0 30.0 40.0 3.0 50.0XI06 2.0 Reynolds numhe~ R (b) EJTectof tbiskncssform.
FIGuRE12.-Contfnusd.
322 REPORT 964—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
on each side of the design lift coefficient. Throughout the percent thickness or Iess, the maximum lift remains relatively range of Rey~olds number of this investigation, the values of constant over the lower range of Reynolds number. lMend- the lift-curve slope (fig. 11) for the smooth sections tested are ing the Reynolds number beyond this range, however, causes a rapid increase followed by a leveling off or slight decrease of very close to that predicted by thin-airfoti theory (2u per the maximum lift. The results obtained for the 18-percent- radian or 0.110 per degree). The lift-curve sIopes of some thick section, however, show an entirely different typo of of the sections show a sIight tendency to increase with scale effect as evidenced by a relatively steady increase in Reynolds number but, for design purposes, this slight fiect l’or the airfoils under considera- maximum lift over the Reynolds number range.
is probably unimportant.
The detailed differences in the flow mechanism responsib~e tion, the section lift-curve slope varies only slightly with the for the observed differences in the type of scale effect shown airfoil thickness form but increases with thickness. This trend was noted in the data of reference 1 for all NACA by the thick and thin sections are not entirely clem. Un- 6-series airfoils. The addition of leading-edge roughness to published data at a Reynolds number of 6.0X 10° show that the ~ACA 63-009 section does not affect appreciably the 63-series airfQils, of 12 percent thickness and less, stall m a section lift-curve slope in the range of Reynolds number of result of abrupt laminar sepwation of tho flow near the this investigation. This result should not, however, be leading edge, whereas 63-series airfoils of 18 percent thickness taken to apply to airfoils of all thickness ratios. The data stall as a result of a gradual separation of the turbulent layer of reference 1 show the values of the lift-curve slope of the moving forward from the trailing edge. By the use of smooth and rough airfoils to diverge appreciably as the these results as a starting point, a qualitative flow mechanism thickness ratio is increased above 10 to 12 percent. These can be traced which offers a possible explanation for tho type data- are for a Reynolds number of 6.0 X 10° but a somewhat of scale effect shown by the thick and thin sections. The simih-ir trend might be expected at higher Reynolds imrnbers. basic id.~s presented in the following discussion of the flow Maximum Iift.-The effects on the maximum lift of increase mechanism are those of Jacobs and Sherman (reference 5) in the Reynolds number from. 3.0X 10° to 25.0X 10° foIIow in a some-what extended form.
either of two general trends, depending upon the order of mag- Consider first the airfoils of 12 percent thickness or less nitude of the airfofi thickness ratio (fig. 12). For airfoils of 12 which are known to stall as a result of laminar separation at n N..cA 63-009 t.8 ~ V M4CA 63-209 1/ b NA CR 631-2f2 L N.4CA 63’009 (ftop) J $“ > + < / A $1,6 b : k * g / 4 ‘- c > * — — . .
~ L4 @ f Q % s 1.2 / (c] 1,$0 3. u 6,0 9.0 15.0 20.0 25.0 30.0 40.0 50.0%106 Reynoldo numbe~ R (c) Effect of camber.
FIGURE12.—ConrIudcd.
EFFECTS OF REYXOLDS NUMBER ON TEE CHWCTERISTICS OF X’ACA 6-SERIES AIRFOIL SECTIONS
3.23
the leading edge. The point at which laminar separation adverse effect upon the possibility of flow reattachment. On .
occurs and the magnitude of the pressure recovery -which the other hand, because of the increased -reloc.itiesover t-he may be withstood before the h-imina.r layer separates are not surface, the linear distance corresponding to the Reynolds influenced by the -due of the Re.yuolds number. For number R’ required for turbulence to begin in the separated airfoils which stall by separation of the laminar layer near layer decreases, and this decrease has a favorable effect upon flow attachment. For a given angle of attack and bubble the leading edge, the Reynolds number would not, therefore, be expected to have any effect upon the maximum lift if the sise, further increases in M at the same Reynolds number ‘-- possibility of the separated layer reattaching itself to the would seem to depend upon the relative strength of these surface were disregarded. Since the data of figure 12 show two effects. The data of figure 12 (a), which show the maxi- no sca.le effect on the maximum lift of the thk airfoils over mum lift of the thinner m“rfoilsto increase rapiclly over a t-helower range of Reynolds number and since these airfoik relatively short range of Reynolds number, would seem to are kuovm to stal.Iby la.mimwseparation tit-hiu this range, indicate that ai a gi~en angle of attack and depending upon it n@ht be assumed that,,once the flow is completely sepa- the inititd bubble size, which in turn depends upon the wing rated, increasing the Reynolds number does not result in its Reynolds number, appreciable iocrease in lift is possible reattachment within this lower range of Reynolck number. before forwarcl movemen~ of sepa.ration becomes the pre- The subsequent rapid increase in maximum Iift o-rer a dominant effect and causes the flow to separate permanently.
relatively shori range of Reynolds number (fig. 12) is b elievecl The preceding discussion is based on the assumption that ““” to indicate that the separated la-mina.rlayer is reattaching ma~um Iift is a function only of phenomena occur@ at itself to the surface as a turbulent layer. Data. showing such the leading eclge. The changes in the flow field near the a reattachment wit-ha “bubble” or “dead air” region misting leading eclge, however, cannot be considered as affecting between the points of laminar separation and turbulent only local conditions at that point but must also be considered reattachment are presented in references 6 and 7. These in relation to the flow over the rear of the airfoil. The decrease in sise of the la.minar-separation bubble near the results aIso show that the bubble decreases in size as the leading edge has a beneficial effect upon the hrrbulerd layer Reynolds number is increased for an aidofi at a giveri angle “f’hisbeneficial effect depends on t-he of at-tack. A qwditative speculation is acbrancedin reference near the trailing edge.
6 as an e.xplanatiofifor the reattachment aRd decrease in size fact that the init-icilconditions of the turbulent layer as it of the bubble with increasing Reyuolds n~ber under given begins near the leading edge me so altered that more pressure conditions of pressure gradient. According to these ideas, a recovery may be withstood before separation begins near The increased negative pressure peaks defite Reyuolcls number R’ should exist between the point the trailing edge.
at which laminar separation occurs and the point of transition near the leading edge which the decrease in size of the along the separated la.mimm layer at which turbulence lamina.r-separation bubble permits, however, have a dis- begins. If the assumption is made that t-he turbulence tinctly adverse effect upon the tendency of the turbulent spreads from the transition point tit a given a.n@e, reattach- layer to separate at the rear of the a“doil.
ment will occur when this spreading turbulent flow strikes & the process of increasing maximum lift with increasing Reynolds number continues, a situation may be imagined t-he surface and estabIiehw itself as a turbulent boundary layer. For a given airfoil shape at a given angle of attack, in which the turbulent layer near the traiIing edge becomes increasing the wing Reynolds number w-ill decrease the critical and starts to separate. The effect of this separation distance corresponding to the Reynolds nu~ber R’ necessary on the flow field around the airfoil is of the same type as that for the separated laminar layer to break up-into turbulence. produced by the small negative deflection of a pla@, flap.
The size of the bubble, therefore, decreases with increasing The beginn.@ of turbulent separation at the rear of the Re-ynolds number. airfoil thus results in higher negative pressure peaks near By application of the ideas.just discussed to the phenome- the .leding edge for a given lift coefficient (reference 8).
non of laminar separation of the flow near the leading edge The effect of these higher peaks is to increase the size of the of an a.irfofi, the point of reattachment may be seen to d- hmnin.ar+eparat-ion bubble- which, together with the higher penalupon the pressure gradient., the Reynolds number, a.ncl pressure recoveries, tends to cause more turbulent separation the curvature of the airfoil surface. hsume that the Reyn- at the rear of the airfofl. A regenerative process could thus olds number of one of the thin airfoils (fig. 12 (a.)) is such be established which would quickly limit t-hemaximum lift.
t-hat.the flow just reattaches itself to the surface at a.nangle Such a process is believed to be responsible for the experi- of attack corresponding to maximum M at a somewhat lower mentally observed fact (fig. 12 (a)) that the maxi.qygn Iift Reynolds number. Increasing the angle of attack under of the thin airfoils, after a rapid rise o-ma a relatively ,short such circumst-ant= will have the following effects. The rarge of Reyaolds number, rather suddenly ceases to increase.
pressure gradient at the leading edge will become more ad- A consideration of these ideas indicates t-hat, even within verse and the negative pressure peak, higher. The lamimr that rarge of Reynolds number where laminar separation at separation point will then move forward around the curved the leading edge is known to Iimit the Iift &in the first flat lea.dingedge of the airfoil. On the assumption that the sep- portion of the scale-effect curves (fig. 12 (a)), the tenclency arated kninar layer flows away from the surface in a direc- toward turbulent separation “at the rear of the airfoil may tion tangential to the surface at the poink of separation have a controlhg effect upon the observed phenomenon of forward movement of the separation point has a defu&Iy kmina.r separation at the leading edge.
324 REPORT 964—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
If the preceding discussion is assumed to depict a reason- somewhat with further increases in Reynolds number. An ably accurate qualitative picture of the mechanism by which indication that this type of scale effect would actiudy occur maximum lift is reached at the upper end of that small may be found in the results for the hTACA 8318 airfoil range of Reynolds number over which the maximum lift which are discussed in reference 5.
increases rapidly, the lack of further appreciable scale effect Although the characteristic shape of the cu.rvcof maximum would seem to indicate that separation of the turbulent layer ~ift against Reynolds number is essentially t.hc same for the is little affected by variations in the Reynolds number. The airfoils of 12 percent thickness and hm, the wducs of the work of Voti Doenhoff and Tctervin (reference 9) on turbulent Reynolds number at which the different cflccts occur vary separation indicates that; if the initial conditions of the somewhat with the airfoil thickness and thickness form turbulent iayer are not altered, increasing the Rcynolds (figs. ]2 (a) and 12 (b)). One effect upon tb.c varimtion of number actually has a slightly adverse effect upon the amount maximum ~ift with Reynolds number of incrmsing the of pressure recovery which may be withstood before turbu- airfoil thickness ratio seems to be a decrease of t-heVIIIUC of lent separation occurs. The lack of adverse scaIe effect the Reynolds number at which the maximum lift begins to shown by most of the data of figure 12 (a) can possibly be increase rapidIy with Reynolds number (fig. 12 (a)). An explained by variations in tbe condition of the short laminar increase in airfoil thickness ratio causes tk severity of the layer near the leading edge which change the initial conditions surface curvature near the leading edge to be rcduccd which of the turkndentlayer a sufficient amount to mask the ex- in turn decreases the magnitude of the adverse prcssLwc pected adverse eflect. gradient just behind the leading edge. When consiclcrcd in The hwge differences in the type of stall and scale effect relation to the previous qualitative discussion of the mech- of. the thinner sections as compared with those of the 18- “ anism of maximum lift, these two effects of increasing percent-thick airfoil have already been pointed out. The thickness would tend to explain the experimental results.
data obtained in previously mentioned unpublished stall The data of figure 12 (a) also show the magnitude of the studies show gradual separation of the turbulent boundary favorable scale effect to decrease somewhat with airfoil layer near the trailing edge to limit the lift of the 18-percent- thicl&es up to thickness ratios of 12 percent of the chord, thick section. The character of the lift-curve peak of the A change in the relative strength of the tendency toward NACA 633-018 airfoil (fig. 6) as compared with that of the laminar separation at the leading edge aid turbulent scpma- thinner sections also gives some indication that turbulent tion at the trailing edge is probably responsible for this separation is limiting the lift of the 18-percent-thick section. behavior.
In view of the preceding discussion of the effect of Reynolds The data pertaining to the effect of thickness form upon number on turbulent separation, however, the only explana- the maximum lift are restricted to movement of the position tion for the large scale effect shown by this airfoil would of minimum pressure on the basic thickness form at zero lift seem to be associated with rapidly changing initial conditions from 30 percent to 50 percent chord and from 30 percent to of the turbulent layer near the leading edge as the Reynolds 40 percent chord for airfoil-thickness ratios of 6 and 9 percent of the chord, respectively. For these thickness ratios, the.
number is varied. For an explanation of the variation of these initial conditions, the behavior of the short laminar position of minimum pressure does not appear to have a very layer near the leading edge must again be examined. powerful effect upon the maximum lift (fig, 12 (b)). Between The pressure gradients near the. leading edge of the Reynolds numbers of 15.OX 106and 25.0X 10°, the data for 1~.-percent-thicksection, although not sutlicient.lyadverse to the airfoils of 6 percent thickness seem to indicate that moving cause complete separation a.t the Reynolds numbers .of this the position of minimum pressure rearward decreases the maximum lift and delays the rapid rise in maximum lift with investigation, might be great enough to produce a laminar- sepamtion bubble of the type previously described. A steady, Reynolds number. The results, however, are not entirely Moving the position of minimum prcssuro decrease in size of this bubble with increasing Reynolds consistent.
number could probably cause a favorable change in the rearward has somewhat the same effect upon the surface initial conditions of the turbuleit layer of such magnitude curvature and the resultant pressure gradients ncm the that turbulent separation at the rear would be delayed to leading edge as decreasing the thickness ratio for a given higher lift coeillcients. Such. a phenomenon would account position of minimum pressure. Rearward movement of the for the variation of the maximum lift with Reynolds number position of minimum pressure would, therefore, be expectcd for the 18-percent-thick section. It seems reasonable to to shift to higher values the Reynolds number at which the suppose, however, that, at some higher value of the Reynolds rapid rise in maximum lift with Reynolds number begins.
number, the bubble would be nonexistent and, at an even For the very thin airfoik, however, the effect dots not a~pear higher Reynolds number, the laminar layer would be so thin to be important. On the other hand, the data of rcfercncc 1 that further decrease in its thickness resulting from increasing show that, at Reynolds numbers between 3.0X 10° and Reynolds number would have littIe effect on the initial 9.0 X 1(Y,moving the position of minimum pressure rearward conditions of the turbulent layer. When such a condition is has a defmitely adverse effect upon the maximum lift of the thicker airfoils.
reached, the maximum lift, would presumably decrease EFFECTS OF REYNOLDS NUMBER ON TEE CHARACTERISTICS OF NACA 6-SERIES AIRFOIL SECTIONS s% ----- laminar layer to the surface and the conditions of t-he tur- TIM effect upon the masirnum lift of the addition of a srnaLIamount of the uniform load type of camber to the bulent layer folIowing reattachment are necessary. Should 63-series airfoils of 9 percent thickness and 12 percent a general investigation of these problems yield fruitful thickness is shown in fi=gure12 (c). The camber increases results, it- is believed that, -with the aid of the relations for “““ - the mmimum Lift of both airfoils at all Reynolds nnmbers turbulent separation pre.viowdy developed by the hTACA, but does not materially change the general character of an intelligent approach to the ‘calculation of the mtiunv the scale-effect curves. The value of the Reynolds number lift coefficient for difTerent airfoils at different Reynolds -- – numbers could be made.
at which the maximum Liftrises rapidly, ho-wever, is lowered when camber is addecl to the 9-percent-thick section. Since UntiI such time as calculations of t-hisnature are possible, -” the most important. conclusicm to be drawn from the ma-xi- .
camber so charges the curvature of the airfo~l surface near mum Iift results of this investigation, from a consideration the leading eclge that the separated kuninar layer may of airpkme design, relates to the comparison of the airfoils attach itacdf to the surface more readily, t-hisresuh is not at different Reynolds numbers. Although the airfoik of surprising.
12 percent thickness and less had the same type of scale- The results obtained for the NACA 63-009 airfoil section effect curves, the Reynolds numbers at -which the dif!ierent equipped with a 0.20c sirmdated split flap deflected 600 effects predominate -raried. The H3-percent-thick section are also presented in figure 12 (c). These data show the had a type of ma.xirnum-hft variation with the Reynolds scale-effect curve for the airfoil with split flap to parallel number that was entirely clithrent from the thinner sections.
that for t-heplain airfoiI throughout the range of Reynolds Any comparison of airfoiI ma-tium-hft characteristics can number. This result would seem to indicate that the rela- tionship between the -rarious parameters -which ha-re been be made only if the data for the group of airfoils under suggestedas cont,roll~the maximum M is uncha~aed by consideration are available at the same Reynolds number.
the deflection of a spfitiflap. SufEcient data are not avail- The choice of an optimum airfoil for maximum lift for a able, however, to show the generaI validity of th$ remdt. given application, therefore, must be determined from data The fact should be remembered that the discussion of the correspondi~m to the operating Reynolds number of the application.
effects of camber is based on tests of thin NACA 6-series sections having smaH amounts of the uniform load type of Accordingly, the conclusion cannot be made that camber.
the effect of different types and a.monnts of camber in LANGLEY AERONAUTICAL LABORATORY, combination wiih d.iflerent types of basic thickness forma ~ATIOXAL ADVISORY COMMITTEE FOR hROMUTICS, -would be the same as that show-n by the present tests.
LANGLEY AIR FORCE B..wE, VA., October 13, 19.48.
Siiarly, the results obtained for the 9-percent-thick section with split flap are not necessarily results that might be REFERENCES obtained with other teypesof flaps on other airfoils.
1. Abbottl Ira H., Von Doenhoff, Albert E., and Stivers, Louis S., Jr.: Tests of the NACA 63-009 airfoiI with a roughened lead@ Summary of &rfoii Data. NACA Rep. 824, 1945.
edge (fig. 12 (a)) show that the maximum lift rema.inarela- 2. Von Doenhoff, AIbert E., and Abbott, FranJrT., Jr.: The LangIey tively constant throughout the Reynolds number range of Two-Dimensional Low-Turbulence Pressure Tunnel. NACA TN the tests. The roughness at the leading edge, of course, 1283, 1947.
causes the boundary-layer flow to be turbulent over t-he 3. Schfichting, H., and UIrich, A.: Zur Bereclmung des Umschleges laminar/turbulent. Jahbr. 1942 der deutsehen LuftMwtfor- entire airfoil. From a consideration of this fact in relation schung, R. Oldenbourg (Munich), pp. I S-I 35.
to the previous discussion of turbulent separation, the absence 4. Prandtl, L.: The Mechanics of Viscous Fluids. Turbulent Flow of scale eflect for the rough condition might have been along a Wall with Special Reference to the Frictional Resistance expected.
of Plates. Vol. III of Aerodynamic Theory, div. G, sec. 23, W.
CONCLTJDING REMARKS F. Durand, cd., Julius Springer (Berlin), 1935, pp. 145-154.
5. Jacobs, Eastman N., sad Sherman, Albert: Airfoil Section Charac- Results are presented of an investigation made to deter- teristics as AHected by Variations of the Reynolds ,h’umber.
mine the two-dimensional lift and drag characteristics of NACA Rep. 586, 1937.
nine hTACA 6-series airfoiI sections at Reynolds numbers 6. Von DoenhoE, Albert E.: A Preliminary Investigation of Boundary- of 15.OX 10G,20.0 X106, and 25.OX 10G. Also presented are Layer Transition along a Flat P1ate with Adverse Pressure Gradient. NACA TN 639, 1938.
data from NACA Rep. 824 for the same airfoils at Reynolds numbers of 3.0X 10G, 6.0X 10G,and 9.0X 10C.. Qualitative 7. Von Doenhoff, A1be~ E., and Tetervin, Neah Investigation of the Ve.riat.ion of Lift Coe5cient with Reynolds Number at a lIoderate explanations in terms of flow behavior are advanced for the Angie of Attaclr on a Low-Drag Airfoil. NACA (IB, N’ov. 1942.” observed types of scale effect.
8. Pinkerton, Robert Xl.: CalcuIated and lIeasured Pressure Distri- The discussion of the phenomena “at maximum Iift is butions over the M.Mspan Section of the N.A.C.A. 4412 Airfoil.
particularly speculative aud indicates that much more re- NACA Rep. 563, 1936.
search is necessary before this problem can be analyzed 9. Von Doenhoff, Albert E., and Teterti, NeaL Determination of quantitatively. In particular, quantitative data relating to General Relations for the Behavior of Turbulent Boundary h7ACA Rep. 772, 1943.
the mechanism controlling the reattachment of the separated Layers.
TABLE I.-ORDINATES OF AIRFOIL SECTIONS
NACA 631-012 NACA. 638-018 NACA 63-009 NACA 63-209 [Stations and ordiites given in pcrscrrt of [Stations and ONWWCSgiven fn percent of [Stations and ordinates given in pereent of [Stations and ordmtes givm in perrvmtof [Stations and ordimtw zivcrr in pcnxnt oi rdrfoilohord] airfoil chord] oirfoil chord] airlofi chord] afrfoil chord] LOwcr surfacb Lower surfmo Upp2r surface Lower snrfaeo Upper snrface Upper surface Lower snrfacn Upper surface Upper surfrme Lower arrrface - Station Station Ordinat Drdiiatf Station Oi-dinait Ordhmta Station Ordfrratc Station Wdinat( Station Station Minatl Station Station Station hdinim —— o o o o 0 0 0 o 0 o 0 o 0 0 –. 935 .5 –! 404 L 404 .5 -.503 .6 .749 .5 –. 740 .5 .5 -.696 –1. 713 -1. KM 1.713 ::5 .75 .75 .75 .906 .76 -.206 .76 -.323 : ;5 .75 ~. HI –1. 519 L 25 1.2.5 -2.217 2.217 1.s6 1..7.5 1.151 –1. 161 1.25 1.25 –1. 041 %5 2.G –3. 104 -2.102 3.104 2.6 :5 –1. 057 2.5 L 5sf 2? -1.582 -1.892 2.5 N –2. 925 :5- –4. 362 .4362 5 5 .“ .. -1.462 5 ::: --2196 b –1. 878 5 ?.i 7.5 -3. 64=2 6.808 7.G ;. 5 -2055 7.G -5.308 -2. 27X 7.5 7.5< -1..760 7.5 –402!2 10 -6. 06!? 0.083 10 10” y -2,010 3:024 10 –3. 024 10 –2. 505 10 –4. 7911 M 7.223 :! –2. 3S.6. :! “ 3.591 15 -3.591 –7. 225 -2.917 15 —5,342 20 g’ -8. 04S & 048 H :: : –2 656 20 3.097 20 –2 997 !2$ –3. 200 20 -4.275 25 -6.712 3.200 –2. 341” 4.275 –8. 800 -3.379 26 –6. 020 : 30 –8. 913 8.013 :: 20 –2 4%4 : 4,442. :; -4.442 30 –3. 470 % 35 -4.500 36 -R. m 9.mo 35 35 35 .4.500 ‘ 35 –9.,000 -3.470 35 36 –3. 000 —5,224 40 40 -s. 345 8345 40 40 –z 971 4447 ~ 40 -4.447 40 -3.376 40 40. !, km 45 46 –L 296 -6.704 8.432 45 46 45 45 45 –8. 432 -3.201 45 –% 877 –5, 370 50 50 -7.942 7.942 50 50 -2.723 50 4.056 50 -4.”056 .50 -2.438 30 66 –3. TdQ –7. 256 -4.635 7.256 55 55 –2, 517 56 3.739 55 55 -2.644 55 63 –4. 420 60 00 -6.435 0,455 60 ml -2267 60 : 3.358 60 -3.353 CO –2. 2a7 60 65 05 -2,028 -3.340 5.567 83 –1. 982 05 2928 05 -5. 5f37 –1. 898 65 65 –3. 210 70 70 -4.622 4.622 ;. 70 -1.670 70 2458 m –Z 458 m -1.430 76 75 –L 966 –2. 654 3.660 75 –1. 342 75 L 066 75 –3. 360 –1. 071 : –1. 902 30 -2.601 2,691 1? -; 0~8 30. L 471 80 -1.471 30 30 -.676 30 g -.896 -L 274 l:%? % 35 35 .990 86 85 -1.737 –. 317 ~ 85 –. 085 –. 707 90 90 -.332 w .550 90 -,550 90 -.033 .
-.196 -.346 ;254 .348 –. 138 .196 . E4ff 05 # 1% 0 1% 1o11 0 1$ 0 1% 0 1% 0 o 1% L.Il. radius: 1.037 L.E. radius: 2.120 L.E. rndius: 0.031 L.13. rdrrw 0.631 L.E. rsdiux 0.307 Slope of rndlue through L.E.: 0.0842 NACA 64-4)09 NACA 65-006 [Stations aqd ordinatm g4van in porcsnt of [Ststimrs and ordinatesgiven in pm.?entof [Stations and ordhmtns givm- fn poreent of ,[Stdions nnd ordinates given in pormnt of ; airfoil ohord] : airfoil ohord] II airfoif chord] airfoil chord] ,, Uppwsurface Upper surfrvm Lower snrfmm Lowor surface Upper snrfoco upper Surflleo Lowor aurfrmo
===7
—1 Station Minsk Station )rdiiti )rdinat Stntion Ordiits Drdhsata Station” %ntion ordinate o o o 0 o o o o 0 o o 0 .6 –. 739 .5 .464 .5 -.494 -.932 .739 .5 .6 –. 476 .392 .75 .523 .75 -.596 –1. 120 .75 -.892 : ;5 .75 .75 -.574 1.f15 1.25 1.128 .1.25 -.754 –1. 408 –1. 128 1.25 .754 1.25 1.23 -.717 2.5 L 623 2.5 2.5 L 024 2.6 -L 024 –1. 012 -1.623 2.5 2.5 -.956 ~ .3 ~ 2.109 b 1.405 –1. 406 –2 IMG -2.109 5 –1. 310 7.5 2542 7.5 L 692 ?. 6 -1.692 –3, 115 -2543 7.5 7.5 7.5 -1.539 10 lC 1.928 10 -1.923 –3, 520 2.308 ~: y9 10 10 10 –1. 824 3.455 15 2.298 -2.298 -4124 -2.197 # 3.36s 2.572 -2672 —4,545 –3: 863 2 % : x –2. 432 4.lm 25 % ; 7J: 26 -2.772 -4.816 –4. 170 -2.697 % 30 –2. 907 –4. 957 4.373 -4373 30 20 3 # -2.852 35 4.479 35 35 2:931 35 -2.031 –4. 9m -4.479 36 35 –2. 952 40 4434 40 -2 S96 -4,349 40 -4. 49) 40 2.6$5 40 j: -2998 45 4.364 45 45 2.919 45 -z 919 -4.006 –4, 364 —2.233 50 50 -2776 -4267 4.136 -4.136 50 2 i15 xl 50 : -2 fslo 53 3.3-M 55 55 2.575 55 -2.675 –3. Sto –3. 826 55 55 –2 741 60 3.452 51 2.331 –2 321 –% 349 –3. 452 60 60 -2518 65 :$: 65 2060 -2. Om –2. 810 -3.026 % 65 % -2. Z41f 70 70 1.740 –1. 746 -2.238 70 –2, 581 70 % 70 -L 235 2.069 75 75 1.412. 76 –1. 412 –1. 661 -2039 –1. 594 R 1:$lii -1.072 –1. 106 % L 584 –1. 534 so : -1.232 36 L 009 36 $’ i% –. m?
–. 601 -’ g 85 –. 805 -.120 60 .611 96 .423 -. 4B 00 ‘al 20 ’30 -.510 .227 .157 –. 157 .626 -.227 95 -.195 1% 0 1% 1% 0 1!4! 0 0 [00 0 12 L.E. radhrx 0.579 L,E. radfrrs: 0.258 1,.E. radius: 1.087 L.E. radiw 0.240 Slope of radha thmnglr L.E.: &0342 1: