Document
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$ 3 1176005002606 I Iw COMMIT’IEE FOR AERONAUTICS ., ..
NATIONAL ADVISORY ... .
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w’iur’’IBIlI lUWOKT
ORIGINALLY ISSLED October 1944 as ~ Bulletion L4H10 TESTS OF THE NACA 653-018 AIRFOIL SECTION WITH BOUNDARY-LAYER CONTROL BY SUCTION By John H. Quinn, Jr.
Langley Memorial Aeronautical Laboratory Langley Field, Va.
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.:, , WASHINGTON F1 A C A IJ13WRY M3U3X/Ff MEMORIAL AERONAUTIC ,W NACA WARTIME REPORTS are reprints ofpapersoriginally issuedtopro~%&~8&ribution of ,, advance researchresults to an authorized group requiring them for thew~_Fi~e~ere pre- viously heldunder a security status but sre ~ Some ofthesereportswere not tech- nically edited.All have been reproduced withoutchange in order to expeditegeneral distribution.
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NACA CB No. I@O . . . .
NATIONAL ADVISORY COMMITTEE FOR AEROIVAUTIES ,.
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.- C@ilkmENT14L B~IJj’Tm . .
A------- “ TESTS OF THE I!ACA 653-018 AIk.FOILSECTION WITH BOUNDARY-”LA”YER CONTROL BY SUCTION By John H. Quinn, ill?. “ .
slJmM1-mY “ . Tests of the NACA 655-013 atrfoil section with .boundary-layeu control’ by suction have been made In the Langley-two=dirensional ~ow-turbulence and Langley two- dimensional Iow’-turbulsnce pressure tunnels. Slots were tested at 30 and 75 percent and at 45 and 75 percent of the airfoil chord at Reynolds numbers of 1.9 and 6.0 x 106.
An attempt was made to remove only a moderate amount of alr through the slots and to locate the slots so that the low-dra~ ~ropertles of the airfoil could be reallzed. The results of these tests were compared with results for a plain NACA 6J3-OIS airfo~l section.
A maximum section lift coefficient of 1.85 at a Reynolds number of 6.o x 106 was obtained on the YACA 653-018 airfoil section with boundary-layer control when the total amount of air removed corresponded to a flow having free-stream veloclty.through an area equal to approximately 1.2 percent of the wing area.
ml s lift coeffiolent was found at approximately the same angle of maximum lift as for the plain airfoil and with suction slots at 45 and 75 percent of the airfoil chord.
The surface discontinuity, which would be found with a flush-t~e sliding door placed at 45 percent of the air- foil chord,would not impair the low-drag properties of ‘this airfoil section.
INTRODUCTION Extensive investigations have been made to develop various types of’devioe to increase the maximum lift of airfoils. The most common high-lift detices are the tratling-edge flap and the leading-edge slat.
Both i .
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., NO. d@10 -.~ . NACA CB .
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these devices have d~sadvantages; the flap produces high pitching moments, whereas a slat definitely llmlts -theregion of lamlnar flow and thus results in high drag even when retracted.
The purpose of the present investigation was to determine the increase in msximum lift coefficient that could be obtained with the arrangement of the NACA 65q-G18 airfo?.1section presmted herein by using bohdary-1-ayer control antiremoving only a moderate amount of air. Locating tbs slots so that the low-drag properties of the airfoil could be reallzed was given primary consideration in the design. By sucking low- energy air off the upper surface of the airfoil, separation of the flow at high lift cogf?iciants may be greatly dela~ed flxithe straight nortion of the llft curve may be extended to higher angles of attack.
The arpl!catlon of boundary-la:rercontrol to :ncrease maxtmrcnItft seem advantageous for use with tailless air~jl.afies.
The hl~~.plichln.~ moments associated with flaps, which would be prohibitive on such a design, and the high drag of leading-edge slats &re avoided.
653-018 sym- In the Present lnvast~gatton, an NACA metrical low-drag a~rfoil secticn was tested in the Lan:ley two-dimens+.onal iow-turbulence and the Langley two-dimensional low-turbulence prassure tunnels (desig- nated LV17 and TDr, respectively) at Reynolds numbers of approximately 1.9 x 10C and 6.o x 104.
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c ai.rfollchord “ CL section llft coef’i’flcient section angle of attack ‘O section profile-drag coefficient cd o . .
af.rfoi?. span ..’ b free-sbrsam w.loc~.ty - Uo free-stream dynamic pressure qQ .’L $ . .
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“\ ,“ ) NACA CB NO. L@10 \~D 3 H free-stream total pressure -. . I . .---0., .
— .. , slot qrwntity rate of flow through Q Hb total nressure inside duct blower drag coefficient: that is, profile-drag
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coeffic~ent equivalent to power required to discharge at-free-stream t~tal pressure alr withdrawn from turbulent-boundary .
total section drag coefficient cd. + Cdob ( ) & flow coef’flcient () static pressure on airf’oilsurface P u velocity inside boundar~ layer local velocity outside boundary layer u Ho-p pressure coefficient —— s qo () 6 total thickness of.boundary layer perpendicular diste.nceabove airfoil surface e momentm thickness of boundary la~r ~:,1-;,d] displacement boundary-layer thlckness .
[~’c +y] &k shape parameter H T
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R Reynolds number b ..
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NACA CB No. I)+H1O
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Subscripts: 30 at 30 peroent of’the alrfoll chord at 45 percent of the airfoil chord at 75 pe~cent of the ai.rfollchord MODEL Ah9 TESTS The alrfotl used in the present investigation was a 5&inch-chord wooden model of tha NACA 653-018 airfoil section, which was painted and prenared for-testing by tha methods described in reference 1. The slots and ducting arrangement are shown in figure 1. The slots were used in pairs; either the slots at 0.30c and 0.75c or the slots at G.~5c and 0.75c were used togsther.
Air was sucked off the u~per surface through the slots into the ducts in the model and out through the side of the tunnel. Cutside the tunnel, the air from each slot was piped through an individual Venturi to ‘11.e quantity of air flow was the inlet cf a blower.
dstorminod by mcas-aringtotal and static pressures in the Venturi throat. Tobal-pressure tubes wem placed ln9i?e the win~ ducts to determine tireloss in total pressurs incurred In suckkg the air through tie slots.
In order to immure this loss for the slots at 0.30c and 0.45c, ona total-pressure tube was fastenad to che downstream wall of the forward duet; for tke glot at 0.75c, one total-pressure tube was fastened to the upstream wall of the rear duct, Approximately twice as much air was sucked from the wing through the rear slot as through the front slot. Frelimhary tests were made with tufts to determino the corract proportion of the total flew to be removed through each slet for affective operation.
Prasaure-distributionmeasurements were made by placing a small stat:c-~ressure tube at a series of chordwiso stations about 3/32 Inch above the airfoil surface. At aach station, the tube was bent to approxi- mate the contour of’the airfoil. Lifts obtained from integration of pressure-distribution diagrams obtained NACA CB No. L@O . .
by this method ha+ been found to be in good agreement .
with the llf%s- obtained--from fcmo6- teats.. -Lif.ti was determined by integrating ths pressures along the floor @ oelllng of the tunnel teat ssotion. External drag was meawureilby the wal%-survey method.- Both lift and drag oopfflcients hava been corrected for tunnel-wall Interference.
Boundary-layer measurements were made by the method described in reference 2.
.were calculated on the assumption The values of od% that the air removed fr~ the boundary layer was exhausted at free-stream total pr9ssure. When the power required for boundary-layer control is calculated on this basis, there are no additional power effects due to an excess or defeot of total Fressure at the point where the boundar~la~r air is exhausted. The required power was furnished by a machine that was assumed to be 100 percent efficient.
RESULTS AND DISCUSSION The lift characteristics of the NACA 655-018 airfoil section with and without boundary-layer control are presented in figure 2.
The characteristics for the plain airfoil at a Reynolds number of 6.o x 106 were taken from revious tests of a ~-inch-chord model of the NACA 853-018 airfoil section in the TDT (unpublished). A maximum section lift coefficient of 1.85 was obtained in the TDT at a Reynolds number of 6.o x 106 for a flow coefficient Or 0.0120. For”this flow coefficient, the total amount of air removed from the boundary layer corresponded to a flow with free-stream velocity through an area equal to approximately 1.2 percent of the”wing area.
Figure 3 shows that, at an &gle of attack near maximum lift and for a given lift coefficient below a flow coefficient of 0.0120, more air was required for the slots at 0.450 and 0.75c than for the slots at 0.300 alla0.75C.
This phenomenon may be explalned by the fact that the boundary layer was thinner at 0.30c than at 0.L5c.and, consequently, less air was required to control the boundary layer with the slot at 0.300 than with the slot farther downatrbam. Figure 3 also NACA CB No. I@Ilo lndloates that. at an amzle of attack near maximum lift and for a givefilift coe~ficient, a,lower flow coef’fl-- clent was required at R = 6.o X-l(3b than at R = 1.9 x 106. Since the boundary layer was thinner at the higher Reynolds number, relatiwly less air was required t“ocontrol the boundary layer at a Reynolds nmber of 6.0 X 10~ than was required at a Reynolds - number of 1.9 x 106.
No improvement could be obtained other than a straightening of the lift curve; therefore, greater Increments were possible at the lower Reynolds number. At an ahgle of attack of 12.J0, the favorable effect of Reynolds number Is clearly shown and the difference between the cur%ms for the slots at 0.30c and 0.75c and at 0.45c and 0.75c has disappeared.
Little scale effect on maximum llft Is evident in figure 2 at the msximum flow rate for the test eynolds t number range. At a Reynolf: number of 1.9 x 10 , the angle of maximur.lift was ~ higher fcr tineairfoil with slots than for the plain airfoil. At a“Re~olds number of 6.0 x 106, the angle of naximum llft was approxi- mately 3° lowe? for the airfoil with slots than for the plain a rfoil. In the LTT, at a Reynolds number i the maximum llft coefffcienz at the highest of 1.9 x 10 flow rate wag found to be limited by stalling at the leadlng edge; the position of the stall was determined f’romtuft studies. From the similarity of’the break at maximum lfft for the conditions tested, maximum lift at the highgst flow rate for a Reynolds number of 6.o x 106 seemed also to be limited by stalling at,the leading edge. A few explo story measurements at a Reynolds z “ number of 6.o x 10 indicated that little”further gain In maximun lift could be obtained by increasing the flow rate above C.0120.
The increases in maximum lift coefficient, whloh are made possible by boundary-layer control, are due to the increased slope of the lift curve in the range of high lift coefficient - that is, to the ex~enslon of the strai@t nortion of th6 lift cur-%eto higher angles or attack. As long as separation moves fbrward from . .
the trailing edge without moving forward of the suction slots, boundarv-1.ayercontrol is effectim. When ttle maximum lift stalling at the leading edge limits present Investigation, coefficient, as was the case in tha o .—— .—— — l?ACACB No. I)@10 ~ a“slot has to be plaoed extremely oloae to the leading . ..
e~ge -of’the airfoil to be eff’eotiys.. ,. Since -laminar flow WOUUL .bq~intained with difficulty in the high-speed “cohdit~~ oveP a slot so placed, no attempt was made in the preqent investigation to place a slot near the leading.edge. “ “ . .
me. amotit o’f”alr:fiemoved with boundary-layer control,may”be presented In terms of the displacement bmnddry-@yer .thiclmeda- immediately tipstreamof the slot. A convenient nondimensional fora-ma.y be given by the expression Q/U&b. When G@6~ reached a value of 1, tie slots were operating near maxinium effe~ti.yeneaa. ~creaslng the flw above this value” had no noticeable ef’feeton further dela$lng separation.
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. Drag characteristics of the NACA 653-o18 airfoil section with and without boundary-layer control are presented in figure ~~. The model with the slot at 0~45c se~l~d With “scotch” cellulose tape (fig. 4(b)) approximates the plain airfoil, because transition on the NACA 65 -018 airfoil section normally occurs at approximate y 0.45c.
1 At a Reynolds number of 1.9 x 106, the dra of the airf’oll with the slot at 0.30c sealed the s-e as that of the (fig. 4?a) ) is practlcall airfoil wtth the slot at” - 0 sealed; therefore, the K45 airfoil characteristics with the slot at 0.30c sealed are also thought to approximate those of the plain.
airfoil.” The messured values of external-drag coeffi- cient may be.obtained by deducting the blower drag coefficient given in figure 5 from me corresponding value of’total drag coefficient given in figure 4.
“An envelope curve, if drawn outside the polars for different flow rates in figure 4$ would indicate that at the higher lift coefficients boun@y-.layar control resultq in a.net reduction of drag.
m figures 4(b) ati 4(0), at .l~ftcoefficients ~loy approximately 0.4, higher drags are found wdth than.withoti.bqundary-layer .
control because.of pressure “lossesin the internal ~tem; Since no-separated flow occurs “h thfi-range of lift - .ccef’flbieqt., boti-ar~laysr control is npt required ~ some means of sealing the”slot ~ for-example, a slid= door - should be provided.. Figure”4(c] Indicates that, if the flow system is not sealed whemno air is being .
removed, serious drag increases may be;e~cte~ ---- “% - ..
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—- . =—11=1 1- Ilmn 1 mmmm 1 1- —l-- ml n -mm n u inn-m . mm—— ——— Total.drag coeff tihiltp depend to a great.extent on v”;~e way atr is .ductpdjn’toand tlyrou@ thg mode.1’i:.: No amttempt was “made 5M tbe prosent Investigation, liowq ver, to design more .effiotentducts “and slots because lifgh lifts de end”more on W.e quantity of dir thkdn in” Wan !!
on the s ot design. Further tests m-ight” be”“devotedto the development of slots and wing ducts that minimize the .~ower- required for bo~dary-layer control.
;: .-, A Cbiparison of the- values of .odob in f i gum 5 indidatds” that mudh higher drag “Mqaes were enco~tered with slots at 0.50c and 0.75c”0.than a~.O.45c and 0.75c.
This “co+ition is du~ primarily to the larger pres,stie difference against which t~e slot must operate atiO.30c.
.“’ ““Bechwe o.?these ”hl~er drag losses and because of the larger amount of lanl.narflow thet can be obtained with the glot at O.~~c than with the $lot at,O.3C)c,it appears more stitable fnom drag considerations to keep the.slot as far:downstream as posslbl? without losing e~.%ct.i.ve lift control.
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Values of the ratio of’the amount of’air flow t~ough the rp.arslot to that.through the front slot are presented U. In.fiWe 69 Tuftsti i&es,were.made &uring the test~ for which data .are.sh.o~. in figure 6(a), The tuftq “ Indipated that, at.”t~ higher angles of attack and highefi flbw coefficients, a flow “ratioof’2 appeared to yield the optimum lift effects. . ~ the rest ofothe tests; .therefore, an attpmpt was made to maintain a flow ratio of approximately 2 (figs: 6(b) and 6(c).). The”curyes of figWe”6 do-not represent a ccm!.stent variation of the aiq=floy.ratio but merely present the ratio at seyeral angles of attack for a range of flow coefficient. “ . . ..
. . Pressure-dls~~~bution.diagr“ at maxiinumlift and r at’a “Reynoldsrr@bpti.of1.:9. x 10 for both slot configu- rations are Freseqted in figure ~. There appears to be little if my separated flow .even at maximum lift!” ,..
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Considerable pressure i.s-r.ecovered in passing over the .
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slots. This recovery is due-primarily to the so-called Hi.nkeffeot qa’qsed by supklng air from the boundary layer.
:’rntegratlon of:,tilesq diagrqms.cabined with the Integration ‘ of”di?graqs of<notia~ pressuz@s plotted en a base line .. perpendicular:tq ‘the. chofidl@ flel~ed a value of”-O.058 “for-the seot~on .p+.$chlhg-mqentcoeffic~ent. This @ue of the pltohing-moqent- coe$f}bient ind$cates that the center of ?ressure is approximately 3 percent nearer the .. . .
.1: “, NACA CB-No. 4H10 ~ 9 trailing edge of the airfoil for a lift ooefficlant of .1..85 ~ for the zero-lift oondltion. - . . . ..
In the regton in whioh lift coeffi.cfents are lower than maximunland correspondhgly less air flow 1s required, It Is thought that pitching moments approximating those of the plaln alrfoll may be reallzed. A pressure distribution on the plaln airfoil at a Reynolds number of 6.o x 106 and at maximum llft Is presented In figure 7 for comparison.
The variation of the bo~dary-layer shape parameter H and the momentum thickness e/o over the upper surface of the airfoil at maximum lift is shown in f’lgure 8. me shape parameter hes an average value of 1.5 and at no point approaches the value of 2.6, at whloh separation was fot~d to be imminent in the analysfs of referenCe 3.
A discontinuity In H is found just downstream of the slot at 0.75c. The boundary-layer profile at this point indicated that some tnter.ference effects from the slot were present.
CONCLUSIONS ~ NACA 653-c18 airfoil section equlnped with slots for boundary-layer control was tested in the Langley two-dimensional low-turbulence and Langley two- ressure tunnels at Reynolds dimensional low-turbulence g nLm?+~.~s of 1.9 and 6.o x lC . Slots were tested at 30 s~:d75 percent and at 45 and 75 percent of the airfoil chord. Approximately twice as much alr w~s remcved through the rear slot as through the forward slot. A comparison of the results of these tests with the results for a plain NACA 653-o18 airfoil section indicated the fullo~fn.qconclusions: 1. A rmximum section lift coefflcien~ rf 1.85 was .
obtained at a Reynolds number of 6.o x lC~ .Llththe NACA 653-018 airfoil section by using boundary-layer control. This lift coefficient was obtained with suction slots at 45 and 75 percent of the airfoil ohord. The low-drag characteristics of this airfoil sect:m co’ald be realized with this arrangement when the clobs we~e net operating by covering them with flush-tJpe doors.
The total amount of air removed at this lift coefficient NACA CB No. L4H1O Fig. 1 I i I 1- I +/w, –“-r-- – — I
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I I A A I / t _— -t ---1- , 7oTAL- m/?E;uRL —— _ —— —— / — .76-C — .4* ..wC -.ocwwc + 06# c / ,> ~ SECT/ON A-A CHORD =36 //v, ALL /?AD//?oo6vc NATIONAL AMSJRy WMMlllF.E FORAEROtWITIC$ Figure 1.- Model of MACA 653-018 airfoil section with dots ~or boundary-layer oontrol.
Slots sealed o 0 Slotscsealed l-!
Q -Q 0.00s0 0.0G20 — — .0040 —— .0040 ———— —- .0080 — ———— . . .0080 —- — .0115 --— .0111 / / L /’ ,/ / )/ : ( , -.
e -. 8< 8 16 8 16 o 24 0 24 j 8 16 2& -8 -8 0 Section angle of l ttack, ao, deg Section angle of attack, ao, deg Section.mgle of sttnok, IZO, % (c) slots at 0.h.5c and 0.75c ~ mT.
(b) Slots at 0.450 and 0.75. “n LTT.
(a) slots at 0.30 R, 1.9 X 106. R, 6.0 X 106.
R ~.9x101=’d o”75cln LTT” .
, Figure.Z .- Lift characteristics of NACA 653-018 airfoil with and without boundary-layer control.
?0 I NACA CB NOO L4H10 Fig. 3 -a Szot-looation .Tunncl .. ...-. R (dog) o 12.JJ O. Oe and 0.750 1.9 x 106 + 12.4 O. 50 and 0.75e ;~T} t x 12.4 6.0 00450ud 0.750 0 16.4 0. Oo and 0.750 Lm~T 0 16.4 O. 50 and 0.75e LTT} t /-------”” A 15.9 0.450and 0.750 6.0 TDT 2.0 F / 1.6 x L r .8 .4 ‘ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
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I o I o .002 ,004 .006 .008 .010 .012 Flow coefficient, C(-J Figure 3.- Variation of section lift coefficient with flow coefficient for the NACA (j53-018 airfoil section with boundary-layer control.
Fig. 4a NACA CB No. L4H1O l O* --.., .
‘o 510ta--9ealed / .032 ‘Q — 0.0020 —— .0040 ---— .
.0080 —— — l 0111 l / I I .028 \ .024 , / / / / ) .004+ NATIONAL ADVKXJRY COMMITTEE FOil AERON o -. B -.
4 0 .8 .4 1.2 1.6 2.0 Section lift coefficient 9 C$ (a) Slot? at 0.30c and0.75c in LTT. R, 1.9x 106.
Flg&e4 .- Drag charactertetlcs of NACA 653-018 airfoil eectlon with and withoutboundary-layer control.
NACA CB No. L4H1O Fig. 4b .036 X-A. .
“ .OSlots maled CQ l 032 — ‘0. 0020 .— .0040 -—___ .0080 —-— .0115 .028 / + \ I ~- . O* ./----- -+== .
—--- --- --—” / I I CA)MM117PE FORAERO AUTICs o
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-.8 -0 0 .8 4 1.2 1.6 2.0 Section lift coefficient, cil (b) Slote l t 0./+5cand 0.75c in LTT. R, 1.9 x 106.
Figure 4,- Oontinue&, NACA CB No* L4H1O Fig. 4C .oy6 . . . . . .- 0 Plain airfoil ,.
, aQ .032 0.0020 —— .0040 -—-- -..
.0080 —. — / --t .0120 t / siOts open, boundary-lay.r A i028 / oontrol lnoperativo { IH / 1, $ / \ .024 I u“ / TA / I . Oal / .
\ J \ 1’ /’ \ / ‘1 \ \ / ,7] c \ \ / \ .01,6’ : 4J o / .
/—\ m \_ L --——. _.— / \ * .012 Q / ‘ \f ;/ h / “ .008 .004 NAT:ONAL ADVISORY COMMITTEE FOR AEfiONAUTICS
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-, -e 0 4 .8 l 4 1.2 1.6 2.0 Section llft coefficient, c1 (c) slOtS at ().4SC and ().75c ~~ ~Te R, 6.0 ~ ~06e Figure 4.- Concluded.
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Fig, 5a NACA CB NO. L4H1O .032+ ,...
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.Cq 0.0020 .028 —— .00 0 * -—- —..
8 .00 0 / —. — .0120 / / .024‘ / .020 =dob / / / .016 / /- .- / / / / /- .012 . / / “ .008 F .004 NATIONAL ADVISORY COMMITTEE FOR AERONAIJTICfb 0t -. 8 -.
o 4 J-l .8 1.2 2.0 1.6 SectIon llft coefflclent, c1 (a) Slots at 0.30c and 0.75c In LTT. R, 1.9 x 106.
Figure 5.- Profile-drag coefficient equlvale>t to power required to discharge at free-streaq total pressure air withdrawn from turbulent boundary layer cd on NACA 653-018 airfoil section ‘b with boundary-layer control.
I II Fig. 5b ‘ NACA CB No. L4H1O .032 I .
CQ ,... —- . ..
I I
0.0020 ——.000 ———.
!) .00 0 —-— .0120 .028 ;/ / \ / .024 / r .020 cd ‘b .016 / “ / ./” /’ .- - /-- _.
.012 _— -- .008 - .004 — ~ NATIONAL AWISORY “’ Cf)MMITTEE FOR AERONAUTIC:
I I I
c 1.2 1.6 2.0 0 .8 -.8 -e 4 .4 Seczion lift coefficient, c1 (b) Slots at 0.450 and 0.75c in LTI!. R, 1.9 x 106. ‘ Figure 5.- Continued.
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NACA CB No. L4H1O ‘-.
Fig. sc .032 >- . . . . . .
CQ 0.0020 —— .00 0 ------- .- .00 0 $ —-— .0120 I .028 / / .024 / .
.020 , / NATIONAL ADVISORY ‘d.
COMMITTEE FOR AERONAUTS b .016 i / / /“ /.~ ---- .012 /“-- -/ /- - - ---- .008 ~ - .004 0 1 -.0 -. u 4 .5 .4 1.2 .1.6 2.0 Section lift coefficient l CZ (c) Slots at 0..45cand 0.75c in TDT.
R, 6.0 X 106.
Figure 5.- Concluded.
NACA CB Iio. L4Hlo Fig. 6a .W .
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e h= 16.5° Y q- 12.4 + 0’ $ Q30 4’ NATIONALmvi.y)Ry Cf)MMl~EE FORAERONAUm 4’ 2 — —— — — ~ — . . —- * d,= 4.1° ~- ~ t- ~ 0+ A 0+.002 .004 .006 .Ooe l cm l o12 Flow coefficient, CQ’ (a) Airfoil with slots at 0030c ~d 0075C in LTT. R = 1.9 ~ 106.
Figure 6.- Ratio of air flow through rear slot to air flow through forward slot at several angles of attack for NACA 653-018 airfoil section with boundary-layer control.
Fig. 6b NACA CB No. L4H1O Cze = -4.1° 2 — --t- - — — — -=+ “ — — -~ T —+ — “ 0“ @16.5° A A +– 2 - +-- -+\ & , 4 “ I I a.= =.4° J.
2 % 0’ %45 l 47 I I ~0= 13Jj0 k .
+- 2 +— — ~- — — — 4. 1 c%, = 4.1° + 2- , +— — - 0 “ NATIONALADVISORY IXIMMITIEE FOR AERONAUTIC + 4- I I of?
a. s K + o 0 .002 .004 l 006 .008 l 010 l 012 Flow coefficient, CQ (b) Airfoil with slots at 0.45c and 0.75c in LTT. R= 1.9X 106.
Figure 6.- Continued.
I ---- .-.
I NACA CB No. L4H1O Fig. 6C a .= 6.2° +- ~ -+ + -t 2 . + .
cc~= 4.1° +— ~ %52 + , n * CQ45 o coo= 2.1U -k =+—l—— h +-t + ltATiONAL AOVISORY CO!’ilMtTiEE FORAERONAIJTiCS 4 r f (XO‘S(f - * 2 +w- — 1- +–+- — + — =++- o~ .002 .m4 .00$ l 008 .010 .012 Flow coefficient, CQ (c) Airfoil with slots at 0.45c and 0.75c in TDT.
R= 6.0 X 106.
Figure 6.- Continued.
. ..—.
NACA CB No. L4H1O Fig. 6c-Cent.
, +- - —— + A 2 ~~ I &;. 16=00 “ - ~ ; I .,.., :. ’,,,..
,., ,, +. ~ ~ ~ “+ +- “ “ : “h 4 2 r :, “, . .
G%= 1s.5° 0- .
+.
\ L >2 ~ 1 , +-A a.= 14.s0 Q45 o 4 “ I ) +~ 1.
2 I C6,=12.4 A o’ NATIONAL_ COMMllTEE FOR AEROWnM I +- - * + +-~ ~. A.
d,= 10.3° !
I l OCE .004 .006 .008 ,010 l o12 Flow coefficient, CQ (c) Airfoilwith slots at 0.45cand 0.75c in TDT.
R=,6.0X 106. ‘Continued.
Figure 6.- Continued.
NACA CB 140. L4H1O Fig.
6c-Cone.
.
, 4 8 C6a”-6.2° —- — + 2 — - - — t + < () I Cl!@ -2.1” +-..._ _ &52 +— + --t Q45 A , 4= 1 / C&= 17.5° +/ 2 + ‘l- 0“ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS.
\ 4“ / GG@= 16.5° +’ 2 +— - - - — *- -+ .010 .012 l OCQ .006 .oCe 0 l 004 Flow coefficient, C(J (c) Airfoil with slots at 0.45c and 0.75c in TDT.
R = 6.0 X 106. Concluded.
Figure 6.- Concluded.
NACA CB No. L4H1O Fig. 7 =...- .. —- --
I I
surface -upper Slot~ at 0.30c and 0.75c T : Iower CLo= 17.00; CQ= O.olu: R = 1.9 x 106 11, x Upper Slots at 0.45c and 0.75c a Lower Cto= 17.6°; CQ= cl.o~~; R = 1C9X # 10.
. . .
6.c 5.0 J.L. o MTi IN AL ADvls cOMMITT E FOR~Eif( 3.0 2.0 .
I 1.0 L & .
tlmse 8tati0n, Xjc ~gure 7 .- Pne~IIurOdistributions on-the NACA 655-018 airfoil S.ction with and without boundary-layer control.
~’”
,, ..
~ o El/c Slots at O ~Oc and 0.75c +H = 17.00 eQ= 0.0111 } =0
T
;.
x e/c Slots at 0.&5cand 0.75c ,..
q H } = 17.6°; CQ= 0.0115 ‘O .-..
.< .:- -- :-s.
-’ ,.
; “.: <&.
,“.
..: .,’ ,.
. .
.- .. !
u - ,.
., ..,, NATIO!
lADVISOR Ct)MMITIEE OR AEROii TICS I + I 1 I I I I 1 t .2 .1 J1 .6 •~ .5 l 7 * 1 chordw18e 8tat%on. XIE %1 P P’lgure~ .- Variation of b&dary-la er shapeparameter andmomentum w thickness on uppersurface of IVACA 53-018 airfoil section with boundary.
z “- .
.
layer control for two slot configurat~ona. R, 1.9 x lob.
(KJ l 3 1176005002606 I m – . .. —-——-. —. . –.. - .Jll