Document
ACR July 1942
NA llONAl ADVISORY COMMITTEE FOR AERONAUTICS
,
"'<
ORIGINALLY ISSUED J~ 1942 as Advance Confidential Report PRB:L!MINARY INVl!STIGATION IN THE NACA U>lJ....11URBUl roNNEL OF UM-DRAG AIRFOIL SECTIOIf3 SUITABLE FOR AJ:NITJ:'Im AIR AT 'mE IEADllG EDGE By Albert E. von Doenhof! and Elmer A. Horton Langley Memorial Aeronautical laboratory "'-
Langley Field, Va. '- ~
· ~ . A.
JP ~ 4
WASHINGTON NACA WARTIME REPORTS ar e rep r int s of pap ers ori ginally issued to provide rapid distribution of advance r ese ar ch results to an author ized gr oup req ui r i ng them for the war effort. They were pre- viously held under a security st a tus but ar e now un cl assifie d. Some of these reports were not tech - IDeally edited . All have been reproduced wit hout change in order to expedite general distribution.
L- 694 J
. - - -- -----------~-~-------~~ ~~ ,-~- .... -.----. -- --.,---- -- NATIONAL ALVISORY COMMITTEE FOR AE::~ONAUTICS ADlrANCE CONFIDEN TIAL REPORT 'TI JNN ZL OF L OU - DR\G . AIRwort SEC rl' IOND SmrrA:B LE B3' Albert E . 'yon ].ben h'J:t'f ,mel El me r A . Forton ill;' invet>Mgat ion W 0.8 c J .rr ie(l. 0 t in the NP..CA 'l ow- turbu l enco tmIDel to d.e velop l m·r - fu'ag C' .l r fo.i. 1 30 c ticns ::;u i tab le for' adm:l ttins 3.i:c .. tt. the le a(Ung 80.ge . !.'>. thic: neGS Ole t.ri b l1.t i o'n havi n g the deoired t y }} e :)" ]?:t. ' 038lE 'O d.:Lstrj __ Dut : ~O!l 1.- ",,8 found f r on tests of a flex i bl e ,nadel . Ot her aJ.rl'c-11 ,s}lO,pes I-, ~re deri vea. from this origina l sh a pe by var y ing the t, h1.c]D1e ;::." t.he canbeY.', the leaLling-edge radius, and the s iz o of ·(,116 le a~:i. ),w ' e ,1c~ e op'3ning .
t'at::t ar e IJres en ted 8i vj.n.g the chJ.r s.c terlstlcs of the airf'oH shapes i.n t h e I' 8nge of li n. coeff;Lcien t ·· for high - spee .:: l and cruis-ing f l i< . S} l apes Daxe been "l . eveloped \ ·l b. ich show no ,u b stant il'l.l incre a.o e in clre.g 0'181' thE.. t of nonaal I f) ,.,- drac; tY}l6 86ctions having mini mum -r, reGG Ire a.t t. he same posi ti en al cllg the chord . Many cf the s3 sh.s,pes ap pear to have h:igh er c rL, l cal compredsibili ty speeds th'ln pl ~:i.n oj. doils of the arune thiclmess . Low- ill' vB airfoil sections h (t ve b een clevf:llope(l wi th openi ng s in t h9 le ading e og e as la.rge as 41 .5 pel'Gent of che maxinnun. th ickne os . The range of l ift coefficients for l ow Ql" ag in several ca se s i s ne ar l y ·a.s 13.1'ge as that of the corre s p ond:'L ng l)la.:-n 2.i rf011 s ec t io ns .
Measurement] of rn;-;.xirm.>-m li ft Che.rd .cteristics 'vere made for onl y a f ew co nf i gurat ions 8IHl ~lO co nclu s ions c o ul d. be c.rc:m as to what effect [:' he le ad.i nG - er lee cp er.. i ng3 wo uld have on the maxin1l.ml lift char a c tertst ic s of the com pl et e Wi : .l g .
IN'L'ROIUC1.'IOJiI 'l 'h e leadin g edge of the ' ,"""lng h HS :pro'.red to be a conyenient l ocation fo r the 'ent r an ce to a ir duc ts . 'This l oc ation io p ote ntially . ,.
efficient b ecause t he atr can · b·e prought . to rest at this point
vli thout l oss of total pres , sure . 'The · pl a cing of :Ju ch opening :3 in the l eading edge of airfo il s can l e :3. d, however, to serio u s increases in the extern B J. drag an(L to pre Ss ure pe ak s near the le ading eage th o.t ca,n seriously reduce the critica l compressipility s peed .
Even a ve ry s ma l l peak , of course , eliminates the possibility of m aintai n 1ng B....'1 y extensive regions of l a m.:i.n F:l. r f lm ·T, Previ ou s tests in the NACA l ow- tu.rbulence tunnel ( unreporteo .)
showed the po o sibili ty of :1 dmi t ting air ef f icien t ly at t he lend.ing e dg e of l ow- drag - airf o il sections wi thout disturbing the I o. m.1.nar l aye r . These tests , hmvever , dea l t v ri th rel a tive ly smal l op en in gs ab out 10 perc en t of the ma.. .. dmum thiclmes s on illJ, 8 .ir foil section of 21 - per6ent thickness . In order ' to lim:.i. t the 8p an of the o pening and to reci.uce O.uct 1 08se s , it is desirab le to have as l arge an opening a s pos8 ibl e and to a,d.mi t the ail' at as low an intali:e- vel oci ty r a t:Lo as pO f3 sib l e . '1'he purp0 3e of tbe present inves tiga tio n is the deve l opment of l ovl- d .rag - ail'f o il s e cti on s having l arge o penings in the l eading e~ ~e .
In the deve l o pment of the basic sh a pe s J a model consistlng of two . flexib l e me ta l s neet l3 fitted. vrl th pressure o rifices vTa S used , The model w as ni o unte lt in the t e s t s 8 ctl on in such a manner that its" shape could be 13.1 te red from out a :L cle the t UIlnel vl"h:). 18 the effe ct on the pressu r e dist:d . bu t ion coul o_ be o b se ~ C' ve d on a mu1 ti tu b e ' Il1.'llometer . The entr a nce - fIm v 1' at ·e · llT a s con t r o lled. by main t.a inin g a fixed r a tio 0 :' nose - to - t ail openin gs . ~ v hen the sh ap e h a ving t he de s ired t y-pe of pre ss ure dj.stribution \- las ob t ainecl , the ordin a te s of the . s/lape we re me as ured an-i the pre s sure distribution ,-l as recorded • . The or d:inate s of the s ymmetric al section obtained in this m a.. '1ne rYlere plo tt ed and faj.re d; t he fa i red or 6 in a tes were th en used · in ~h? .c on ; 3truction of a lvoo d en model . A mo re deta il ed investig a tion of the char a cteri s tics of the airf o il secti on an d the effect of various c hc m ge B i n ' shape vlaS carried. out wi th vl ooden mode ls.
The pre sent inv estigat ion deals p;rj.ma rily vll t1) the de t erm in atio n .
of s0ct10n ch o. J;'acteris ti cs in the r an.ge of lift coeffic ie nts f or h igh'·speed . and cruisi ng fli g ht . Although t he i mp or t ance of dete rmi ning the ef 1 fect of the u s e of the s e se ctions on the maximum l ift of t he wi ng is rea l ized , it i s fe l t th at thls effect c aIl best be found from test of a complete ai r pl ane mod.el rather t han from tests of a t vl O- dimens 1 0n al mo de l of th e no se - air intalc e section . Because the o penings in the l eadin g ea .g e may extend ove r only 8.. relat . ively small po rti on o:f the s pan , meas urem en ts of t he maximum lift of the n os e - opening sections alone '. lOul,Q. not give rolia'b l e information concerni.ng the effect of the use of the secti on s on the comple t e ai rplane.
For this reason systemati c me asu r eme n ts of section /llDximum lilt coefficient havo not be en made ; although some ini'orma tion on this s ubject has b een ob tai ned fo r a fe" c .onditj.ons .
SYMBOLS The symb ols used ar e defined as folioi-Ts: V f r ee ~ stream ve l o city Vn velocity of air ente r ing the nose opening
. q free-st re am dynamic pressure (~pv?
p l oca l static pr essu re H free - st ream tota l pre ssu re
S pre ssure coefficient (! _ ~ p)
tota l pr essure at exit l oss of total pressure through duct ( rr - TIt)
section profile.-drag coefflcient ( do \
, gc) sectio~ l ift coeffic~ent (q Zc).
angle of attack , degre es flap d. eflection , degrees area of tr ::d l i~g ~ edge e::d t ar ea of .; le ading - edg e entrancze x distan ce al ong chord from ' leadine; edge of al r foi l y d.;i.stance perpendicul ar . to · chor d chord c density p mass drag section do 1, lift section nu m ber Reynolds R DS AND ME THO RATUS A PPA sections - d rag a irfoil o f low tigation The in v es -t urbulence NACA low in the e d ou t wa s carri herein re -p orted in iwo~dimensional mode ls to test designed which i s tu~nel,
B nd 7i
t wide 3 fee te st section ha s a s t unnel w . T hi flo Ol e in t air , s tream of the lev el turbulence g_ . The feet h i a effients wi t h m~a s~r bulence lo w . Tur xt r em el y el is e t unn o f ati ons fluctu t ha t the indic < te mo met er wi re ane h o t - ent .
0 . 1 p erc e s th ~ n are le vel oc ity 0 . 024 - of t wo co n sisted which mode l, e f le xible Th of 2 a c ho rd had umin um alloy, ts o f a l ic k shee i n c h - t h nted on were mo u sheets The of 3 feet.
a spa n feet and the throu gh xte nded that e s tringers ise - inch s panw ght 1/4 ei could m odel e of t h e t he shap g es i n . C han n el walls tun - the po si chan g in g 'g by was r unnin tu n nel the mad e \'7hi1e be each On the tunnel.
outside ers fr o m s t rin g n of the tio 1 5 were p o siti o n , midspa n , at the model ce of the surf a to a multitube co nnected t were e s t ha orific pressure g ute 1.
in fi is given th e ' model etch of . A s k manometer y for determining isf a c t or l was s at ble mode flexi g h the Althou wa s not i t the section, par t of g reater of the o ut li nes t h e the ne a r of ~hang~~ effects of the a s tudy le for s uitab mak e d i d not t he r m ore s hap e fur l The interna g ed g e .
le a din n- was co g ation the investi ng l y , . Ac cordi duct a suitable models.
w ooden use of w ith the tin ued roxi- o f app s p an and of 3 - foot ls we re en mo de The w ood in were mude . They ch chord
Or 60 -in
c hord y 24 - i nch mat el tr a iling m et al f l ex ible a having each section sections, t wo for the exit size of the e d to adjust b e us t hat c ould ed g e t o p and f~ ap . The split and as a w r ate g the flo v ary in end plywood 1/2-inch w ith to gether we~ e held sectio n s botto m g r ap h A photo s pa cers.
st ee l in ter nal sever a l lat e s a nd p s hown in odel is wo oden m h -c h ord 60 -inc a typical of 2.
figure ',: : Pressure-distribution me a sure men ts were made on the flexible model by the use of the pre ss ure erifices and ' on the we , e , d , e~ , medels , by mean ,s of sm al ). " , stat ic " tubes of 0.04 0- inc h ' out " si.de diame t er, mounted en su p p or ts appreximately " Q '~25 inch , abeve , the model sur .f ace . , P r e ssur e dist , ribut , io ns ~~e presen~ed ~ s ' c u r~es of the p r ess~r e ~o~ f fici~nt S plotted , against cho r dwis ~ e ,p ositie n. ,,, ~ I t ,i s t o. , be noted that. in 1ermini this ciefficie n t , free - ~tream total p , ressur ,e i:8 u~ , e d a s t ,h ,e re f ere n ce p r e, ssnr ,a r at.her than ,,,
free- ' st ,: reamstatic p ressure ,' " Su r v e y s in a ,v, artical plan e
at mid ,s p an indicated t :h at t :h e .- fl ,o w was mo re uniform at t he exit than ; " at t ):l e, e, ntrance. M ea snre.Jllents , ef fl e w were the~~f~re ma~~ ' by m e8~ u ~in g st a t ic p ressure and tetai p r , e s ~ ~,r eat , the c e n t e r 0. f t b, e, e:x it . ' Dra g ,wa s " meas l lred , py t h e ,\va,ke - , sur v ey me thod. ' The inte g ralo.f t h e lessef tet e, l pr essure in the wa~e ' . a fairJ y clo.ie aFp ro. x i mat io.n to. t h e drag, was measur~d ~ ith an inte g rati ng m ano.~et~r. ' Co.~r6c tioni to. this , valu~ ~er e ebt",d.n , $db'y ' a me t ho 'd su t st an ' :ti :a l1 Y eguiv9.1ent to th a t of B. M: " ' Jones g iven in re 'f ' ere n 'c e 1. ' The lift was de'tf:')r, min ed fro m " me a sure in ' en 't s of p res sure s' a l on g" the ' floer and roef of t h e tu nn el. Bec au s e the ent ire l ift was not' tra~sferred to t ~ e . tu n nel walls wi t hin .. t he di stance cov e rsd by the o'r if i c 'e: s, aco l' r 'e c t i 0 rJ ' , ,d a t e r mi 'i1ed t heo 'rCet'j 'c a ll; r ., ' was a pp li~d ~i t h ~ ' me a sure~ r e s ult~ ' t o. obtai~ the tot.l lift .
The , aata : pr~s e i~ed h~ re i n " h av ~ " be en c orr ' act , ed " fo1 ' thnnel - wa 11 ef f e c t s :. ' " D~VEiOPM:ENT OF 'N OSE-Op :mUNG AIRFOIL SEAPES
, ~~~£~il_~h~~I~~ : Th~ me asu r ed ~rdinat e s of the
tlexible ' m edel w er e f ~ ~r e d to o b t ain a sy mm etr~pal sha~e.
The t h ic k ne , s s aft hi s sy mm ~ 't ri c a 1 a hap ew e,s '" .t llen r e. <?-u Cl3d
to 16. 900perce 'n t ,' :' c. Or d in at as f o r this , thic~ne , ss , , diS , tribution, . c a lle d a irfo 1' l s h a pe 7. ,: ~re g iven in table 1.
The sect ion '\Vas ' combined with a n , a = () . 5 ', type ef mean
line havin 'g ' ~ "' desi g n l ift co eff ic ient of 0 . 2 to obtain the ordi n ates o f the m odel. (See r ef~rences 2 and ; 3.)
The chor d " of t ~ e m 6del ~ as ,, 24 i n ~ h e s~ The model w as " first t est ed wi t h i;l. ,. sharp leading edg e.
The pre ' ss t ae distri but i on fer t h i s condit ,i , on is g iven in fi g ure ' 3~ T h e ' sli g ht p ea k in t h e press~ri distribution en t h e lower su rface near t h e l ead i n ~ ed g e. together with rather h ig h val u es of the dr ag co efficient , indicated th e ·6 desirability of making some ' ~odifications to the le a din g ed g e. Rounding t he leading edge to 1/32~inch radius resulted in the im p roved g haracteris t ics s· ho wn i n fi g u res 4 and 5.
In order to . check . the operation of t h e ' airfoil sec- tion in climb .w ith an internal resistance to sim u late a radiator, screens w ere i n s talled th a t had a pressure drop Was equ a l to 0.9. Tests were equ a l to " , ~q when ma d e to determine whether this flow rate could be obtained .at a lift coefficient " of .· 0.8. For · t hi s series of ' m easure m ents, the sheet - metal tr a i l in g ed ge on the lower surface was · bent down, ~orming a 0.1 5c s plit flap. The characteristics w ere meas u red f or fl ap deflections of o lli , 15°; an~ 20 °. The resu ~ ts a r e g iven in fi g ure 6 .
Tests to deter mi n e the maximum lift coeff i cient of the section wh e n f i tted wi t h a 0.20e s p lit flap de f lected 60 w ere made int~e NACA t w o-dime n sionai tunnel. The maximum lift ·; co e ffici e nt s h owe d li · tt · le v a riation with Re y nolds number. Re m ovin ~ the sc~ee n s al s o ha d little ef fect . A ty p ical lift 6u rve showi n g t h e peak is g iven in figur e 7.
~i~[~il_~£~~_~~- Ai rfoil s hap e 8 was derived from an imp~ov c ment in the fairin g of the ordinates of t h e flexible mo del used in d e rivin g airfoil s h a p e 7. No r educti o n was made in the thi c knessl however , w hich was the same as that o f the flexible mo d el, 18.892 percent c.
The ordin at e s ' for the s ymm etrical m od e l are given in table II. Figure 8' ' shows the s h a p e · of the a irfoil section.
Tests of s ba p e . 8 . wit~ the shar p leadin g ed g e ga ve results similar to t h e initial res ul ts ob ta ined fo .r sha p e 7, indic a ting t h at t h e sh a r p le a di n g e.S , ge wa s too critical .• Th e leadin g ed g e was ther e fore roun d ed t~ ~ p p ro x imately 1/32 r inch radius (fi ~ . 8). Fi g ure 9 s h ows the p ressure distribution for , the mode l i .n this condition. ' Lift, dra g , d u c t loss, and int a ke ve ' locit y w er e t h e n m ea.sur e d. Th ese res u lts are g iven in figure 10 in nondim ~ ~sional for m.
In an effort to incre a se t h e lo w- dra g ran g e, the le a din g e d g e wa s cut ba c k 2 .4 89 percent c and was faired to a lar g e radius (fi g . 8) . Or d i n ates are g iv en in t a ble III . Alt h ou g h t~ is chan g e i m~ roved the section characte r istics (f ig s. 11 and 12). ~t le a st at low Reynolds numbers, it . affected the p . ressure distrib'\, ~t ' ion '.
adversely · near the · l.e . ad i r.. f- ·. ed g ~; . as. i:s. seen in ,f.i g :ure 9.
For . ~ucceedin g . ~ od e ls, ' ~ ' so mewh ~t . smal~er . leading~ed g e · .
radiul;i . was : tneref , ore chosen '. . ' . ' . ' ~i!:'£Q_i1.._[£@~_~!.;- ,A,1.rfoilshape · ~ ;, i ,s the sa ,me as , .'.
airfoil shape .8 . exce p t . for t,h~ , le ad~ng-edge radius, which is so mew ha~ smaller than the ~arge radius t . eat ed 9? shape 8 .
The ordinates for shape 9 are g iv en in table IV. In <;>rder to obtain results at highe r ?,eyn91ds numbers, the chord of t l"li s .and of · succee d in g models Vias 'inc reased to 60. inches.
Lift, drag, duct loss, intake veloc ity, and pressure ~ d~stribut . ion were ~easured for thr ee different , widths of t he tail ; o p ~nin g . These res ult s . are given in coefficient form in fi gu res 13 to 15 •
. " . "
~ , -!:tr.f.Q..i1.._[h@~-1Q..!..- Ai rf 0 i i s hape 10 r e s111 ted frol}} . ", an e f for t to . fa ira n 0 pen i n g 0 fa g i v ens i z e i n t 0 anN.A..C.A.. : 6 . 5. ,2-21 ~ airfoil ~ection ' wiz 'h mean line a, = · 0.8 (refer- " ence 3) without c hangin g the oFdinates of , the orig~nal sectio n bac J;~ of tne 0 . G5 c position. In order to , aVOid :· ch anging t he s~ape ' of . the . mean li ne, a new symmetrical airfoil ~h ap e with th e desired nose opening was derived, an~ this sh ap e wa s camb ered . to the original mean line.
This o pera tion was performed by the use of shape 9, reduced so m ew hat in size, as a guid e fo r the fairin g in the neighborh oo d of the lea d ing e dge ; this portion o~ the s e ctio n wa s t h en faired into th e NACA 65,2-015 ' se~t ' ion.
A smooth curve was drawn by e ye , joining the forward ·oor- · tion of the section with the NA CA 65,2-015 section,' In ,.
or. de r to c heck the fairness o f this curve , a measure or t he c urva ture at several po i nts along the surface was found, and thi s. quantity was pl ott ed against chordwise p o ~i tio n . The, lIleas ur e o,f the cur vature was comput ' ed acco~ding to the f ollowing formula y(n-l ) + y(n+l) = . -------------- --- wh~re Yn is the ordin a te a t the chordwise pa.itian x ' n
T, p.-e ya rious c ho rd w ise po sitions Xl" xa ••.. xn must
be equally , sp a ced '~ ' The ori gina l curve of " h against x was n,oi; smooth. It was found n ece ssary to make this curve smooth in order to obtain satisfactory pressure distributions. ~he curve of h against x Was made smooth by successive a rbitrary c h~nges in , the or~i~ates.
The ' trailin g ed ge was cut " off a t 0.910 to form the rear ~ pening . The re su lti ng sy m mett i c a lsection, desi g na ted ai rfoil s h a p e 10 , fot whic h , tbe ord i nates a re g i ve n in t able V wa s t hen cambered ~ bout an a ~ 0.8 t yp e mean t
li~e w ith a desigri lift co e fficieni , 61 o.~ to o btain the "
o rd , inate s of the mode l tes t ed. ' The c' ha r acte ristics of t his s e ction were measu r ed , f or th.ree " d.ifferen ,t W'idGh~ of th e tr a ilin g- ed 'g e op ening . ' These results , are g iven in fi g ure 1 6 .
, In or d er' to det ermine ' the , effe c't o'f : ch angin-g the angle between the line joi n in g t h e urp er- and lo w er- s u rf a ce , l~ a d i p g e dge s ' and t he c h ord , line, tests w ere mad~ wit h t h e u pp er Cr nd 1 0 we'r surf a c e s sh if ted with res pe ct to each o th ei~o g ive variou~ ~ mount s o f st a g g er . . The or :~i ~al ' sta gg er, 1 ueto the c em b ~r, wa s 0.265 ind h .
Te ~{s were a lso made w ith sta gge rs of 0 . 53 inch and 1.10 inc hes. Data fo r the te sts with increased sta gg e~ ' a~e g i v en in ' f i gur e s ] ," 7 and 1 8. T 11 eli f tc 0 e f f i c i en t a s a f unction of t h e an g le of a tta c k fQr the vario1ls test ' con '- di{{~~~ , is given in fi C ure 1 9 .
' The r esults of ' preGsure - di stri ~u tion me asu re men ts for shape 10 a r e giT en i n f i gu r e 2 0. Fi gure 21 gi v e s a co mp ri~o .n. bet w een the theo re t ical p -r e s su r e dist ri bu tion for th e NA qA 65,2 - 0 15 ~lrfoil sec t i ~ n ~nd the bas ic sy rL1me tricaJ, p re s su r e dis t ri 'b ut i on derived fr o!'!l fi gu re 20 .
~i!:£QJ!.l_~p._ ~Q.~_J-,- 1...!.. - Ai rf 0 i 1 s h a p e 11 j, s au a ir fo i 1 sectio n o ~ , app r6 x im a tely 0.25c maximum t h ic kness . The o rdine tes w ere derived fro m t h ose of air foil s hape 9 by incre a si ng the ordi nat es fo r ~h ape 9 in the r a tio o f th e t h ic k nesses of the sha pe s. T he l ead in~ ~ edge radius wa s also incre a sed by th is r a tio. Or dina tes for a ir foi l , s hap e 11 are g iven i n t ab l e VI . The , usual t e st res u lts for t l isa i r f 0 i 1 sect ion are i i v e.:Q. in , f i go . r 8 s 22. 2 3 • .
a.n d 24 .
4.i!:foil _!i£~_~_l~.!.. - Airfoil s hap e 1 2 was derived to s tudy the effect of va ri a tio n s in the size of the op ening in t h e le a din ~ e dg e . Sh ape 1 2 has th e s ame maximum thic k ness as s hap~ g', b ~ t the l ea d ing -e dge o p ehin g has been ,~ educed from approrJ,inately 32.5 percent of the maxi- mum thic k ness t o 23 pe~cent of t he maximuM t h io kn ess.
Or d in a tes fQ~ this sh ape ar~ g ive n in table VII. Th e tes t re::;ults a r e , gi ven in .' fi gu r es 25, 2 6 , an d 27 .
.ihir.f.Q..i~_~h@~ _ 1~.!..- ' Airfoi1 s hape l3re p r esent s an e ffort to ob ta in an a irf~i1 sect ion ha vin g a very lar g e op enin g in th e le a din g' e dg e . It was obt a ined by simply ." . ~ spread ing apart tpe upp eT . '!and l:owe r s ur · fae ·e a . of a irfQ 11 shape 9. The re sutt ing "ect: iO .n ha Q. . ; ~ I;l ~xl mum . t'h ickn,.es s' of appro xima t ely 21. 7 perc ent c an d an op en 1 n·, in th e 1~ading edge of about 41.0 . percen t of t he max imum thlek • • • • Ordinates for shape 13 ar, · g~ven i n t ab le ~ II I. the t e at ..::t results a~e . pr,sented in , figure~ 28 t o 3~ • .'
!li
I . : ..
H . ' ; '.. . ' ~ 'r
. ... .. ~~
·DISCUSSION , • " . .. t . . . .. ' :"
, \ . ~ ". ' ~~~~1.~~'!:.~c!g,,~_raC!i~~~- As stat ed. j)1 oeVi ouf? lY. ,.: lIa\la~ '
factori : results were not · obtained wlth . t~e ,\ s h arp le~Q.i,g edge. Comparison of figure~ 3 , and 4 sh ows " t h ~t . tlle ", Etffe c t on the pressure distribution of s light l y ro un ~{~g ~h~ ~ ~ leading edge is to eliminate the peak o n the lower surfa c e .
, ..
. . . ; Tests through a rang-e of an gle ' 0'£ atta c k'~ ho w evefl , showed that the range of lift . coe ffici e nt~ - f o rl o w ~r~~ Was very smalL (See fig. 5.) I n or de r ·· t-o incr e ~se the ran g e of lift ' coeffibiepts fo~ lo w drag, , the lea~i~g ~dg e was cut back considerably ' and rou nded to a large . radtu s as' shown in figure 8. Although t his ch ange improved tbe low..,drag range, as is · seen in f i g:ur e 11 t , ,i t s eems pro ba - "! ble ' that this radius is too lar ge be~aus~ ' of ! it ·s ~dver s e
eft'ect , on the pressure distribut i on sh o wn . ln , figure 9~
It i 's ,: , bel ieved that the low-dra g ran g e at I, higher ~eynold s numbers would be · considerabl y s mal l er than t h~t : sh~wn ~ 1n fi g ure 11.
An intermedi a ~e value o ~ t he leading-~dge r a di ~ s ', wa s therefore ch os ' en for ' airfoil s.h ape 9. Altho u gh this , '.
value , of the leading~ed g e ra ~ i?s , may n o t be precisel ~ th e optimum, th .e dat a indicate tha.t s o me what larger or s m aller radii lead to ch a racteristic s le ~ ~ satisiac~ o ry tha n tho s e for the intermediate radius. ' E.~Q.'i!._!:.~t~.!..- The effect o.f v ar i ~ti o ns in , the rate of air int a ke has . been studied f or a num be r o£ the airfoil sect~ons. Air , must be " adm~~t~d ' ~t , the lea d i ng e d ge in ord~r , t~ obtain satisf~ct9ry ch aract eri stics. The , min~ . mum rate of intake to obtain lo w , drag , h~ w e y, e r, depen , ~s ,u p on th~ p, artic-q.lar s~ction. For ' air f ·o i-l , s hape 9 w ith a leading- edg e · opening 0 f a bout 32 pe rc en t . 0 f the ". rnaximum ' , • . . V t . hickne~s ,,. . this : minimum rate is a valu e of -B. of ?- pp r o x i - V . mat~ly , O.38; for shape 12 with a l ea di n g -edg e o pening , of ._------ _ . _--
'10
abo ' ut ""2'3 'p' 91'cerrf : of the ma~i-mum t _ hicknes!:!. it is less . than 0.~7. ·: (S '~e ': figs ' ~ ' 13( ' b) • . :1. 3( 'G) ·, ~~d q5(C):'~ ' ) : " . .: • t : ' . , : ~~~ : .' . '::' · -' In general. · the Cha~ ac teJ~i ~ti~s of the '. s~ct'ions ' i mp rove ~ ' ; with ' increase in the · flow ·, rate · up ' to th~ p oi :(1t " w hei~ the internal duct ' losses begin to b~ seriQ~s; tha~ :' i~; ~ th~ ' low-drag range is increas ed a nd the v a. lue of ' th e "' min1 m-l lm p ressure cb efficient is red uc ed slig h tly as the f lo w rate is increased. It is noted tha~1 a lt h ou gh the low-dra g r ang e at first increases rapidl y w it h incr ea se in flow rat~ abo v e the minimum necessar y to obtain low dr a g as s , ~ ' ~n ' frb m'the data for a-irfoil sha p es 9 Q , nd 10 . (f~ g s , 13(b), .l .3· (c), ' and " 16) : , ' further -l'Tlc:r.e a s e in tlie ·,.f , low 'r 'a Fe '" h 8. S · , ' li : ttle effe ' ct a,s indicated by the da t a fo :r shap.e ·· · 12 · : ...
(fig ', 2 ' 5) '. , .. " " . ' '.
, .. " " - ": .
. '.
' . In all c as es the loss of total pres ' sure ' " in ·t · he internal " f1 'ow w: a:·s 'n : ifgl'i' gia l e ' fora · ran g e o.f li ·ft ,,, co~ffic i e . n t s so me - whi t'fri e· xc · es s· ·· o'f· :·t"h e ,l o ·w - dra g r .an,g e. • .. ,. F . U:r.'tl ~ e' I-' ': in c re a s e in · .. .. 'the1 : i{t : coefffc t e:n·t tesulteq in a gl" a d- ia11y ' incr: 'e a: : sing . 1Qssassoci a-e ed . .w it ·h local sep arat+6r~ ofthe " int :' er na l ' ~low . ~~ · ~~e leadin g edge~ ~ . ..
. ;, .
Aitfbil " thicknes s.- Th e effect of c h~ n ~ i n~ the ' · th i. c k ;~;;-; ~ ti;-~~; · -b;;- · s~ . ~n fro m a co mp?- ri : so~ of ': ·t li e da t a fl?r a irfoil :. s .h a pes 9 : (fi , gs · ~ - : "p 3 t.9 15) i3.nd li (fi~s, , : 22 to 24)'; ··· · In · creasi n g · the : · t . :r,:t · i .c ~ n e ,' . 9 , s results in . a 'n . ill'crease of · t h·e lowld .r ag r a ng . e. .... f ·oI" : ?- :. · g.i y 'e. P: . r~ . tio of o 'p Em.i~ ·g: to maximum t h ickness. Althou g h t h e minimum p res8u~ : ~~ e~ k of ... .. sha p e 11 was hi gh er than t h at of sh a pe 9 , th 'e ir {d "e a- se is . n 'o·t .. ~ so " mu ch ' as · ·tv o·u"1 o. . l: re ; e ' ~p ec · t : e . d : froTI' ..; .a cor . r~,spo.nding , .' .. i? ~ re : a se in th e t hi c-kne$ s oi' " ';a :,p, laJ :n a i . rfo i 1 ~ , e dt ion . . .
.. , ' ' . . ,' In ·: th ' i's' ' con ne ction i :t s h oul ,d b e' ' not ' ed ,;t;hat bot - h she. ne ' 9 :: , >' . ~rt.d ':' ~ ha:p ' e 11 have · co . :nsidera.bly lo wer peak p r · ess il r es : ·t han ,.' wOUl"d' be ' found on " ,plain air:fo. .i ·l .. s e , ctio~ s of 't h e' -s. ame .
thickness, As p revio- u sly stated, ' s l;l a pe ll , :w a s . deri -v e.d from s hape 9 sim p l;9"b y mu1ti p. lyi . ng thE3 'ord, in ates o r "' s ' hap e 9 by the desired "r a tio of ~ thic k~ . e;:;ses, Ano th ~r m et h od of incre~si ng t he ' thickness ~ ~ · ill ustr ~ t e d ~i.~hape : 1 3 . In this Case the uppe r and lQ w~ r sur f a ces .w er e , s ep ar a ted by a Gonstant amount • . The dat a .. for ,· sh a.p e 13 (fi g s , ' 2 8 to 30 ) are very similar to t~ose ' for ~ ha p e 9 .• ~n spit ~ ' of ' the fact t ha ~ the thic k ness has b e e n ' i ' n~rea s ed fro m a bout 19 to 22 p e' rce:nt and the ·r a t .. i~ , ~f . -the . s ize of the 1e a din g - ed g e o pening to the maximu m t hi c k ' nes$ ' ha s b ' ee 'n ' in creased from a pp roximi3.tely 32:5 . to 41 , ~ 5' . perce . nt .. I t is , s ign ific a nt that the peak pressure for ~h~pe 13 is practically the s am e as th a t for sh ap e 9.
§"i&~ _Q..:f._l. E?_ ~9-.iQ.f~ ::.. 9. !& ~_9 ]2.. ~ IlL~.&.!.. - The e f f e c t 0 f. Va r yin g the size of the op enin g in the leading ed~e while the maximum thic k n e ss is k ept const~nt can be seen froill .. B co mpa riso i1 of th e d a t a for a irf oil shapes 9 and 12 (fi g s. 13 to 1 5 and 25 to 27). Shape 9 has a leadi ng -ed ge qpenjng approxim a tel y 32.5 percent of the maximum thic k~e~s. and t he ope ni n g in sha p e 12 is approximately 23perce~t ' of the maximum thickness. The d ata indicate that the sm~ll~r o pentn g i s muc h les 'S critic al to flpw · rate . and ct-ange of ang le of attack th a n th e l a r ge r opening. It is ' felt that shape 9 h as about the lar geBt - Biz~ opening iri . the leaaing e dg e th at c e n be pla ced i n a s e ct ion of its thicknes~ while still main t aini n g favor aJ le aerodynamic characteristi cs.
T here is soree indi c at ion that the lo w -dra g range is stro ng l y inf lu en ced by t h e slo pe of the external contour in the n~ighborhood of the leading edge. Decreasing . the S~2e of the op~nin~ Q nd i n cr eas i ng the thickness of . the airfoil se ct ion both hav e the eff . ect of increasing . tbe slop~ n ear t h e le ading ed ge . This le~ger slope has a te nden c y to increase the lo w- d ra g range . The conclusion s hould no t be draw n, h owever. that this slop~ c~n be ind e fin it ' el y inc re a sed, ·· b e c au se it becomes difficult to · . f a i~ t h e forwa rd po rti on of the section into a shape of re aso na b le t h ickn~ss . it h 0 1.1t cau s Ine,; pressure peaks to oc cu r a s hort · . dist ~ .. n ce fro · ~n . t h e l eading edge: .- . r..!:~§' £."9:.!:.~ , _!i .§.~ !_ LQ.~~i9_!±'~- Co mp a. r is 0 n 0 f the 'r res S1.1 r e ' distri but ions for ' the various shapes with those for . plai n airfoils o~ corresponding thic knesses show~-th3t the values of . the minimum pressure coet"f icie'nt for' , man 'y of the nose - o pe ning s hape s are.considerab~ylo wei . than those for the plain a irfoils. As a n example , ' ai~foil shape 13, which is .2 1.77 4 . p erce n t thic k , has approximat~ly , the same valu e of the IDinim"U!U -p re s s u r e coe fficient a$ ·. tlie NicA 66,2-016
air f o.i 1 sec t ion a t z e r 0 ]. i ft. · . A.1 0 w e r ·v Ii 1 u e 0 f the . pea k
p r e ssure is af importancs ' b~c a u se . it indicates an.increas e in the crit . icB.l comp ressibilit. y s peed : of the . sec . t"ion.
This , increase enables t · he des igner to use .e.. t~ic k er section
t han w o uld~ther w i s e · prove feasihle .. '.
Th e theoretical pre ssur ~ di s tri but ions . gi . ven for co mpar iso n wit:.11those f or .t he ve.riousno~ e~op , e ·£ ling shapes are t he pressure dis tri. . bu ' t · io p~ ;fo 'r ' N,ACA . low~drag airfoils h a Vi ng the s am e .· t1+ic 'l-; ness ratio :s.;;., thos ·e of the nose- op ening shapes, except in the ca . s~ .:.of shape 10. ,' Figure 1 2 21 gi ves a comparison bet w een ai rf o il shape 10 8..'1.d the NACA 65 , 2- 015 airf o il section . I n t h1s figL l re t he tra il ing edge of s hape 10 c o rre s ponds to a value of 0 . 91 fo r x/c, and , the trailing edg e of the N ACA 65 , 2 -015 airfoil c or resp o nds to a va lu e of 1 .0 for x/ c .
The actual thickness ra t i o of shape 10 is , of co ur se , greater t h an 0 . 15 bec ause the c ho rd has been de cre as ed by 9 percen t . This c omparison shm.,8 t h at ' t he minim um pre ssure c oeffi ci ent for , a nos 'e - ope ning shape is very near ly ~ he same as ' that of the plain sect i on in to v7hi ch it fairs ; th a t is , no consid.e~c a bl e incr eases in critical c ompress1 bili ty sp ee ds a re t o b e expecte d fr om 1. ose - o penin g sections der i ved by modifyi n g o nly th e l eading e'dge of the or i gin81 pl a in airfoi l s ection .
l"Jaxi;..nmm lif t .- As stat~d i n the In t r od uc tion , i t i s fel t tp at the ef f ect of nO :;l e - op ening , sect io ns 'on maximum lif t c an be ' st b e found f r om te sts of a comp let e a irplan e mod .el rather th8.J.J. fro m tests of a tw o ~d.i. menr :J io nal m od el of ' the nos e- air in take se c tton .
S uch tests have n ot yet b een IUade . ' Some prelim:lnary checks , ho wever , indic ate that the p os si ble de creases , in maximum li ft shov ~d n ot b e l a rg ~ . The nnxiprum lift of a i r fo il s hape '7 i-Th en £.i tte d wi th a 20 - percent -ch ord ap li t f l cip def l ected ' 60 i s see n fro m figure 7 to b e 2 . 15 . Measure me n t s of t h e max i mum l ift of sh a De 11 cambe red
for a design cz of 0 .4 I·rL th an a = 1.0 ' ty ' pe me an lin e : gave a
, 6 ,.
v 8 ~u e of 1 .41 at a Rey no l ds . numb er of 6 X 10 as compa re d wj, th 1.42
at the same Reyn olds number fo r a, n. 'N A CA 65 , 2-)+22 airfo il se ction
'IT i th an a = 1 .0 type me arl line . , Ti1(3 tnaxinrum lift of 3n
NACA 65 , 2 - 215 , a = 0 .8 , airfoil ' .sec tion ' v7 as me asu r ed Ivith nose -
operdng shap e 10 ex t en ( ling o ver a pproxima te ly 11 perc en t of the span of the model . No c hang e i n t he madmum lif t vTaS ob s erv eo. in t his c ase . Such elata ,holvever, are t oo incompl 'e t e t o orai'T any , co nclusi on s as to the p os sib l e effect of le ading - e dge ' ' openings on the illC'..ximum li ft char a cteri s tics of ' the comp le t. e i'T ing .
Drag .- 'rhe values of the dra g co efficient of pose - o peh ing s ec tio ns in the lov 7- dr ag ran ge a re pr a c tlca lly t h e same as , t h os e , of t he cQ rre sp onding 10W - dra 8 se c tions . Figu re 31 g ive s a c omp ari s on be t i'r een ai rf o il shape 10 and t h ~ ' NACA 65, 2 -215 ai r foil section .
It i s s een that t he l Q i'7 ' -dra g r ange is somew hat l ess than t h at of t h e or ig in a l s ec tion and, t h a t the drag outside of t he l mr-drag ran ge incre ases at a greater r at e vr i th l :if t co ef f icie nt than fo.r . t1l0 pl a in secti on . A l arge :p art of ' th:i:s' i nc r e a8 ein dr ag is due to the i nte r nal l osses t h at : occur at an g le s of attac k outside of the l ov - drag :r.<an (30 . ' TIle dat a ind lca t e that the l Oiv -dr ag range increases (a ) wi t h in cre asing flo i -T rate , ( b) ivi th elecreasing en tran ce si ze for,sections of a gi ven thiclme ss ; ' all, cl (c ) ,dth incr easing thickness if t h e r at io of the \"ld t h of th~ o penin g to 't he maximum thickness .
" , i s lTJa in ta inod con stant . Of the shapes tes te d, the larges t l o w· l ' ag ran ge "l as s hmm by Ghape 11 ,\, Ti t h A, li ft - co e fficie n t :canE3s for 1 01 .". dr a g of 0 .4 . 1'11 th a of 0 . 48 ( fig , 22 (a )j, Applic at i on ,- It ap pe l'l.rs fr om t he pre se nt data that 'ch e proper u se of th e nose - o penjng sect i ons pr es ented_ in this report can lead to c oo ll . ng i.n st allation ,., h a vin g pr actically no adeli t i onal ex +:.crnal dr ag in the r ange of lif t c oef fici ents for high - speod Mel c::."nistng fli ght . Alth oug h most of t he airf,il . shapes for wh:].ch data. .... re given a;re s ymme trical , thes e s ha p es can b e treated . i n the same ma nn e r as Cllry oth er 1 0i' T- dr a g type symmetrical sections ; th.at 1 .3 , the s ynun et ri cal shapes cnn b e comb ine d \'11 th a me::tn lin e having the d eai r 0d. C.Gs ign l ift in or cle r t o .sh:U't t he range of lift coefficlents for 10 vl dr ag a nd ef f ici ent lntema l fl ow , as is indi.catecl by the data for shape 1 0 . ( Se e fig , 16 ( a) .) Stagger in g the opening lad. -'n e ffect sim;i.l a . to t he ef fec t of lID increase in the camber , 01~Y' smal le r , Resu l ts showi n g the ef fe ct of varioU3 amounts of sta.gger ar e given in I'~.[;,ures 16 (b ) , 17 , eno_ 18 , If it i s d.e si red to fair p ose - open ing shape8 into exis tj.ng aj . rfoil s e ctions J it is re cGDJ.iUen(;.e d t: l at a procedure atmilJ.r to tha.t usoe!. in the cleriva ti on of shape 10 be enr.Qloyecl . In pal'tlculu r, it i s u sua lly desi r ab le not ·~ o .:0_ tel" the shape of th e or;_ginal mean I l ne an d to b e c e r tain t h at the variation oJ. cu:~va ture a10nG the s urf ace is smooth and continuou s .
Be ca use t he amount of air re q uireu for cooling in the c~imb condi tion is pe al' l y as much as is r equired in the high - speea.
co nd1 tion , the i ntal{:e -v elocity r at i o in the c li mb cont3i'cion must 1) e considerabl' grea-er th:m for the "igh-speed cO:lditionc. Test3 of a irf oi l sha p e 7 11 i t l all in terna l resi ut81 ce havj . ng a pressure
- ~
drop of 2.11 a t an in tak e- veloc i t y ra tio of 0.9 S1101.,e (1 the
8 V
possi bili t y of ob taining high in take - v el oei ty ratios at lligh lift coefficients , Interpolation of the resuJ_ts given in figu re 6 Sh01-Tl3 Vn that a f l ovT ra t e of 0 .9 c an be o bt ained at a lift coefficient if of 0 .8 'Hi t h a c ODl b iI~ e d co ol ing con t rol and .:~ sp l it flap d.eflected 17 . 2 • A:rl anal 'sis of the drag d at a o -b ta inecl from 'Chis series of' tea t::.
indicates t ha t the oxterna l drag caused by deflec tion of the flap is much l ess t han that or dinarily associated with the def l ection of a s pl it f l ap ; in fact , the increase in tota l drag is the 'lncrease that co uld be asso cia t e d ,\, li th t he int ernal l osses . 'l'h is res'i.ll t is r easonab l e b ecause th e flo w o ve r the u pper surface of t'le flal? vIas no t sta lled .
A fevr sho rt t est:] vTere made to de termine su i ta ble m et h eds of en ding the ope nin g an d f airing the n ose - o penin g s hape int o t he w ing in the s pamTi sE) cJJ. r e ct .i cn . T he se te sts inrlic;at ed t h B .t the opening s ho ul d be cl osed gr adu a ll y i n a l en6th eq 1..1 .al to at l e ~st t Hi c e the maxi m um height of the ope nin g . Sem ic i rc ul ar or e llip ti cal en (18 vlere un sat i sfactor y .
C ONCIJUSI mm 1 . Ai r foH se c tio n s 0 1' the l aw- dr ag type , sui table fOl~ ad.rJ1..i tU ng air at th o le adlng edge - withou t. fl ubs t nti<:. l increase in cl r ag , ha v e been dev eloped .
2 . Ma.n. y of the sectj.ons test er.. appear t o hav e h ighe r cri ti c o. l compre ssi bility s p eedfJ than ph.in sections of t he S8.L 1e thickness .
3 . LO I-T- c1 r !1.g sect :i.o nG h av e be en d.evelopecJ. that h av e ope ninGs in t he l ea d.ing ed "'c ao l 3. r ge aa 41 .5 peYC6nt of th e ma ximum t hickness .
I t- . '1'h 0 r ange of l ] .ft coe:f:'f:!.ciGnt o for 1m', ox ag in seve r a l c as e s i s ne a rly as l arge [tEl that of the cor~ .' esp6ndlng pla.in airfo il s e ctio n.
5 . 'r he meas ur oment a of' maximum l ift ch a racteristics ivere too inc o mplet e to d:cavr a 1Y C OnC l 11.8i on 8 re garding the effe ct of l e ading - edge openings 0l1. the maxi ra uIa l ift char3.c t eri3 t::.cs of the c om plete wine; .
Lan gl ey Memori a l Aero _ au U. caJ. TJaboratory , Na t i onaJ. 1\.(1 v :i.801:"J · C om mi ttee for Ae ronautics , L ang l ey F i el d, Va .
-- - - - _._ - , _. --- - REFERENCES i. The Cambridge University Aeronautics Laboratory: The Measurement of Profile Dra g by the Pitot-Traverse M ethod. R. & M. No. 168 8 . British A.R.~., 1936.
2, Jacobs. Eastman N •• Abbott. Ira H .• and Davidson.
Milton: Preliminary Low-Drag-Airfoil and Flap Data from Tests at Large Re y nolds Numbers and Low Turbulence. NACA A.C.R .• March 1942.
3. Jacobs. Eastman N .. Abbott. Ira H. . and Davidson, M ilto n : Supplement to Advance Confidential Report.
Preliminary Lo w -D ~ag - A irfoi l and Flap Data from Te st s a t L arg e Re yn olds Numbers and Low Turbulence.
Bl a c k loo se- le af notebook dated March 1942.
Ib NAC A TABLE I TABLE II " ..
THICKNESS ORDIN ATES, NOSE-oPENING THICKNESS ORDIN ATES, SHARP - LEAnING-EDGE NOSE-OPENING AIRFOIL SHAPE 8 AIRFOIL SHAPE 7 x y x.. y , ' '(~ercent c) (percen t c) (pe r cent c) (p~ r c . ent cJ.
" -- 0 , ~ . ElOO , 0, - , ' 2~ , 7.5; , ~ " " . , .5 3. 45 1 . 5 , . 15 4 .75 , , .625 .'
'.75, ,.2 92 ..
1.2 '5 - 3· 9,5 : 1,2 5 , ,, · 559 ' ' , 2.~ 4. 500 2.5 4·110 5.0 5 ., 52 5· 0 .. 4. 925 7,. 5 5;'978 " , 7·5 ; 5.440 10 .. 6· 505 " 10 5. 880 f,5 ' ' 15 6. 908 20 ',- . 000 20 ..
7.1 88 25 , 8. 478 " 25 7. 612 30 8. 902 ,0 9. 222 ' 7,9}5 ' 35 8.222 40 9. ,86 ' ~5 ' 9-.4L6 " 40 '8. }88 45 : 9'. }48 ' 50 45 8. 450 55 9.159 50 8.410 /' 8. 913 8.210 55 ' ..
65 ' , " 8.484 60 7. 808 7.228 7.793 6~ .
20 75 -' 6.843 70 6.5 5.~07 75 5.677 80 4.770 4.745 : 85 ,.8 72 }.701 2.960 2.402 90 95 2.040 1.}39 10 0 1.250 Nose opening in percent of maximum N ose opening in p ercen t of maxiIIIUIII. thickn .s.: 31. 760 , thiclme •• : }2 . 580 TABLE IV TAB LE III THICKNESS O RDINATES, NOSE-oPENING THICKNESS ORD I N ATES, LARGE LEADING - EDGE - RADIUS AIRFOIL SHA PE 9 NOSE - QPENIIG AI RFOIL SHAPE 8 x y (pe r cent c) (percent c) ,-- - --- y , x o , . 343 (pe rce n t (pe rcent c) c) ,.835 · 5 2.L99 , . 41} . 75 , . 976 2. 60 }. 772 1.25 4. 228 2.8 4. 038 2. 5 4 .7 45 ,.0 4· 2L5 5. 0 5· 532 I 4. 0 h. 908 7.5 6. 1,7 5. 0 5. 337 10 6. 652 5. 978 7· 5 15 7. 1+67 10 I 6. 505 20 8. 098 15 7· ,5, 25 8. 593 20 8. 000 30 8. 965 I I 8. 478 9. 224 25 35 8. 902 40 30 9. 379 9. 435 ,5 9. 222 45 I 50 9. }91 40 9. ,86 I 9. 40 L5 9. 4L6 55 60 8. 9 66 50 9. ,48
I
65 8. 5 10 55 9. 1 59 70 7. 804 60 8.9 1, 75 6. 878 8. 484 5·816 70 7. 7 9, 85 4.679 6.84, 90 ,.5 22 , 80 5. 80 7 95 2. 387 4. 7 45 100 1.,14 90 , . 701 Leading - edge radiu. : 0.251 percent c 2. 402
95 I
1 00 1. 339 Location of leading - edge radius center: I ra d ius : Location of leadin g - edge 0.251 3. 343 , . 413 Location of fairing po i nt i n ooe nin :: ~ . 772 radi us : 0.28, percent c _ _ , ~,:ad i ng - edg e 0.407 3.067
I
Nose opening in p er cen t of m ax 1m um
No.e opening in pe rcent of maximum l
thicknes.: 33. 1 30 thickness: 32 ~ . ~ 5~0~ 7 _ ______ __ ~ r -~--~ TABLE V RACA TABLE VI 17 THICKNESS ORDIN A TES , NOSEoOPENING THICKN ESS ORDINATES, NOSE-OPENING A IRFOIL SHAPE 10 AIRFOIL SHAPE 11 y
r x
x "1 (percent c J (percent cJ (percent oj (percent cJ 2.001 0 4.404 2. 409 . 5 5.052 . 5 2.546 .7 5
I 5.238
.75 1. 25 2.785 1.25 5. 570 2. 5 3.264 6.251 2.5
I
3. 979 5. 0 7. 287 5. 0
I
7. 5 4.552 8.084 7.5 10 5. 064 8.763.
, 10 9 .8,6 15 5.944 15 I 10.668 6. 660 20 11.320 7.235 25 11. 8 10 7.678 30 I 12.151 7. 99 3 35 I 12.355 8.1 80 40 12.429 8.240
45 I
12.371 50 8.163 50 , 12.172
I
7. 906 55 11.811 7.h39 11. 2 10 6. 79 8 I 70 10 .280 6. 030 9.060 5 .1 82 80 7.661
I
4.286 6.164 85 3.369 4.6qO 2.u52 3.1 uq 1.616 1 00 1.731 , 100 .877 • radius: 0. 331 p ercent c Lea ding-edge Leading - edge radius: 0.151 p ercent c Location of leading-edge radius center: Location of leading-edge radius center: 0.??1 4. 404 0. 151 2.001 - - Loca tion of fair1n g o o1nt in op en1ng: Location of fairing point in opening: 0.536 4.040 0. 244 1.836 I - in perc ent of maximum Nose opening Nose o pening in percent of maximum t h ioknes. : 3 .505 thic kn ess: 22.282
I
TABLE VIII TABLE VII THICKNESS ORD INATES, 1I0SE-OPENING THICKNESS ORDINATES, NOSE-OPENING AIRFUIL SHAPE 13 AIRFOIL SHAPE 12 x Y x Y c) (percent (p ercent cJ (percent cJ I (percent c J 2.378 4.79 5 5. 7 ·3.163 . 5 .5 5. 428 . 75 3.352 .75
I
5.680 1.25 3. 66 7 1. 25 6.197 26 2.5 u. 4 2.5 6.984
5.141 5. 0 I
5. 0 7.589 5. 8 19 7· 5 7·5 8.10u 6.392 10
I
8. 919 7. 291 15 15 I 7.982 9.550 10.045 8. 521 25 10.417 8.925 ,0 30 10.676 9. 2 06 35 35 10.821 9.,75 10.887 9. 435 u5 45 10.843 9.391 50 10.692 9.240 10.w.8 8.966 60 60 8.510 9.9 65 65 7.804 9. 2 56 6.878 8.330 75 75 816 7. 2 68 5. 80 6.131 4.679 22 4.974 3 .• 5 9C 2.387 3. 839 95 95 2.766 1.314 100 100 0.179 percent 0 0.251 perc e nt c Leading - edge radius: Leadin g- edge radius: Location ot lead i n g-e dge radius cente r : Locetion of leading-edge radius center 2.378 0.179 4.795 0.251 toeation of fairing point in opening: Loeation of tair1n! poi n t in opening : 2.182 .
4.519 0.290 0.407 ot maximUIII 1n percent Nose opening 1I0ie openiDfl in pe rcent of .an."", thicknes. : 23.122 thiolm.... 41.508 Fig. 1 NACA \ \ \ \ \ I I I I I I I
I
I I I Z.
'- "- -.....
z > (") > "%j ..... tv .
OQ
L-694
model.
nose-opening 60-inch-chord typical showing View 2.- Figure Figa. 3,4 NAC!
I I
~ I I I I_
I 2.0 r-- en I 0 Up pe r s ur f ace W I I 6 Lower s ur f ace r---- H , I 1.8 I .
, r-- t -- 1.6 - , ; I ......, I ----- .
f-- . ~ i ~ I h--- !
!
.~ ~ I
I
d i s ---1.
"if ~ . '\ I
t
P
-- -j- 1.2 &-<
II ~ I
I
If
, ~ I I I : I
~ \ I
1.0 I
I
i I I
I I
I I
I
I
I I
I .8 1.0 o .1 .2 .4 . 5 .6 .8 x/ c
Figure 3.- Pressure distribution for airfoil sh ape 7 c ambered for c. = 0.2 w i~
sharp leading edge. a. 0 ; ~n /V . 0. 426 ; At /An. 0. 4 3 9; R. 2.02 x 1 00.
2.0 I I I I-r-T--'~ I -I I 0
' Upp ~ r su r f~ ce . ~ Lo w er su r f ace ~ __ ~ __ ~~ __ 41 ____ -+i __ .- __________ ~______ 6 I I I I I
i I :' I
1.8 I I I
i I I I
I
i
I I 1.6
-"I- ---/)..... - I I I
~ I! I
!
~ -IT
s
I
, 1.2 + -f- --i I I 1.0
1 I
I
.8 1 .0 .8 .6
.2 . 4
o .1 x/c I _ Figure 4.- Pressure distribution for air f oi l s ha pe 7 camb er ed for c. = 0.2.
Leading-edge radius. 1/32-incb ; a. 0 ; v n/V. 0.4 26 ; At/An. 0.439; R. 2.02 x 106.
rlga. 5,' NACA
o 1 0do t-
.
- ~ A VnfV -
~
...:I (;] AH/q -
/
.020 1.0
/
1 \
II
\
.8 .016
\ !
vn/V /
1 \
.012 .6 V
\
/ 6B/q °do 1.
'\
.008 .4 &> I~IJ .2 iii,
V
'\ V
'"
./ o o
-.2 0 . . .6 1. o
-. -. ~
Figure 5.- Bection charaoteristios ~or a1r~oi1 sh ape 7 oambered ~or 0, = 0.2. Small
le.ding-edge radius. «. 0 ; AtlAn. 0. 4 ~ 9; R. 2.25 • 106.
I T ~ >- - ~ I ~ 2.0 I
\
I
/
j )- 1.6 I
/
- I I : .8
!
.4 f-- --
1 i
o -16 o 8 16 -8 Figure 7.- Section li~t coe~ticie n ts for air~oil shape 7 oambered for c, = 0. 2 wlth 0. 20e s pilt flap de~leeted 60°.
Two ~O-mesh sereens to sim ulate cooling reslstanoe.
R. ~.l~ • 106.
NACA Fig. 6 o e do /). VnfV [!] toH/q ' 1.0 .05 I .8 ~ ~ ~ ~ .6
"'"
I
~
toH/q I . 1 I ~~ .4 ~l ' . ; c l( a) 0.15c split flap deflected 11_1/2 ; 1.045 At/An.
I I I I I I 1.0 .05 I !
!
I I I
-- -- I i!
I I
1 ~ ~
.8 (;) I I I !
I
- l- I
~/ ~
[;J
I I
-- .6 I
i I
toH/q I I
I I 11
(b) 0.15c split flap deflected 15 ; 1.408 At/~.
I . I I
I
I
i I 10 I
-.
1.2 .06 I I I i I
I
I I
i
-f- I J
I
I
I
I I
I
I I l- 1.0 .05 f---
I
I ;~V ~
: I
t- : I
~ , - I
I
;K I I
I .- - .8 I::: I f-- - Vn/V ~ 'r.
.6 I
I I
f - - toH/q I I I : . - l-
I
.2 I Figure 6.- Section characteristics for airfoil shape 7 cambered for cL=0 . 2 . Two 30-megh screens to 001 t 2 25 x 10 • simulate c ing resis ance. R. • __ _ 00 .2 .4 .6 .8 1.0 e, (e) 0.15c split flap deflected 20°; At/An. 1.697. ' ~ -- .
... (J) <0
... ..
• t; -=
j n/n
L-i94
ojJe Exit 1.0 duct.
.9 I 0.465; internal radius .8
~
16 vn/V,
_ ~-
in. 1~2 ; ~ typical I 0 .7
and I '"
a, '-..., Lefdin~-ed~e I 8.
C .6 shapes ~ shape x/c - .5 airfoil • leading-edge .4 or f x n s I several 2.00 .3 R, ~ I sbowing distributio .2 -~ 0.536; ure s
--
______________________ sbape .1 Pres At/An, -
Y
9.- &-'.
i.
- If
o Airfoil inch 8 . - Figure · .-:.h.~76'~/~·n~C~h::.....--," 1.2 1. 0 1.8 1.6 1.4 ~2 s ) Figure .rad/"S~I rad/t/s edge ing-ed:;e d /~ac(il79' ea L eadln!1-C'dgl3
L ~
Shar'p NACA lI'i ga . 10, 11 o Cd O 8. Vn /V I!l AH /q 1.2 .024
?
/ 1.0 .020
/
i
II / . 8 .016
I I
Vn /V
/
.012 .6 I I
I f
l> H/ q Iir
-
/ I , - .008
+-
I
t
--.- - I I
r"
v
0-' -- +- - r-- - .2
- 1·- -'- .004
I
I f--EJ
, ~ _ .-
rl-
-I--
, r-r I
% "' I
I
J
o o -.6 o -.4 -.2 .8 .6 .2 .4 1.0 Fi g ure 10.- Secti on chara c t e ristics for a i rfoil shap e 8 w ith small nose radius.
At/An. 0.536; R. 2.27 x 10 • -- 1.0 .0 20 I i
I
I I
, - -- I I I
I I -
.8 . 016 - .
.--1 --- -i-
1 I ~
-.
- - - - .6
t .0 12
r
n
, j
/
l> H /q -
I
11:>- I ~'" ~/ I I r.
. .L
I
- - '- - T .008 I I
If I
I
- -
.~ ~
r ~
I
I
I .- .2 I
I -
- -
-~
I i
J J J
o o -.6 -.2 o -.4 .2 1.0 .4 .6 .8 c J Figure 11.- Sectio n ch aracteristics for air f oil shape 8 with lar g e nose radlu6.
4 /An. 0.536; R; 2.27 x 10 •
t :z= ~ ~ .... lU '" CD ~ ~ ".
('l for .460 .4o~ At/An 0.584
L-694
shape .~77 .~40 0.47~ vn/V 106.
El 0 Ii> airfoil ; x ~- for
I I I
- 6.43 --, -- I I R.
- - - , -
-i--l- -t a
--t--
I L coefficients rates.
o ,
J
lift flow , 1"---
i
-4
-'-I
I
Section various I
:t
~
-8 14.- ,--t- o
.8 .6 .4 .2
-.2 -.4 -.6
1.0 c, Figure with I
I
---j -
r
shape I I radiu.
12 0.536;
-+--+-1
-r
1,6 1~2 1 airfoil ,I At/An.
-h---+'--
'in7;~~~' for " radii.
I I I
a -, nose coefficients
l o
~ ~ 6• large 10 T , lift x - and ~ 1
-4
--+- 2.27 --j-- 1- Section small R, --.;.---t-
~--~--~~---i-;t-~I -8
12.- --t ----r---j ---+-+-
1-1 1 ~ t-I r-.-I '
1-+- L-
IH--=
O
.8 .6
.4 .Z
1.0, -.2 -.4 -.6 Figure c, NACA 1ig. 13 1. .020 (;) ;;> vn/V
\
/
.012
/
1'\
~H/q I--
r--
l!.- ~
K r--
V
~ - .008
yz
~ ) ~
V
j
/
t
1"'- . .004 ~
"'
V
V
(a) At/An. 0.584.
'"
o -.6 1.0 .6 .8 -.2 o .2
.4 -.4
""
.016 .012 r.
J Ir
V
K r
~
2 .004 [7 El.., I~ V ~ (b) AtI~. 0.460.
V
..
"'-
0 o -.6 .6 .8 -.2 o .2
-.4
.016 .8 I
I
I I .6 .012 AH/q 1..0 .008 0-<:l
~ V
Gl.
.2 .004
V
t--...
/'
/B'
(c) At/An. 0.40;.
""
~ o o -.6 o -.2 .2
-.4 .6 .8 1.0
0, 6 • Figure 13.- Seotion oharacteristics for airfoil shape 9. R, 6.43x 10 ' J'ige. 15 , 19 NACA 1.6 1- - -.....
-
-
~ f--
-
"- ,
--
--
'>- !==--' ~ ~
..A ~ ~ ~
s ~ .
~
JI ~
1.2 , ~ f\
it
~I ~ I'\.
1.0 '\ ~I
J ~ V n/V At/Ali t\
0 0.473 . 0. 58 4
"\
I
: ~
~ .8 6. . a77 .4 60 (;l .340 .4 03
I
~
- - - Theoretical pr es sur e di strib u ti on
""
for an NACA 66 ,2 -018. 87 airf o il - '- .6 sectioJl.
"
I .2 , I
I
I
I I
I o
o .1 .2 .4 . 6 .8 1.0
x /o Figure 15.- Pressure distributions ror air r o i1 sh ap e 9. a. 0 °; R, 6.43 x 106.
r-- I
I
I
I
I co I ~. ~ ~
~ &
~ ~ f..
~
~ ~ I
o ~
~
~ § ~ ~ t---..
..t 0 -=t .., ..t ~ .
I () ~ C) as O'l >
•
.
co 0-.
.-t I CI 0 co '-0 -=t C\I C\I -=t '-0
. . . . . . . . ~
I I .-t I
... r;!
() NA CA ;) , r , i
~ I
I
m
to
i I I
I ~
I
I
, I
I / ~
I i
.020 1.0 I I I I ,
'/ i
I
I
.- , I
r
I
\
/: I I
-, .016 .8 I
I
I
I
I
'\
. ~- vn/V , ,
I
\ /
-- .012 .6 i I
\J ~ ~
~/
~ 6H/q .-'
\
V
I. ~
- - - .008 .4 I
j i-
~ I r
1 --1 .....l- ~- - I I
A
~
I
~
.
.- r- - .004 I .2 I I I ~ - I I - -i------L- I--f- ! - I r I I
N
1 I I _.1 ~ o o 1.0 .6 -.6 .8 -.2 o .2
-.4
(a) At/An. 0.655.
Figure 16.- Section characteristics for airfoil sha pe 10 at R = 6.4~ x 10 and with a (a,b,c) stagger of 0.265 inoh.
l-T -:- _0
,---I
__ ,_ ___ 8 ~-~- +1--
I I I 1 I
.020 .016 .012 .008 .004 o 1.0 Fig • . 16c -Lf ~~_I~l~~ fd .
-+- ) - --+- . . - - ---T . -- .R-- - ·----r---r--'T
1.0 i ,-- ' I -rl-r---lf --i ·
I 'I Iii
c--\--l--~ - +- -r-.---t- - I if l
\ . I 1 1 I I I
. 81 -- \ I--lJ+r-+-t--t-- [ . 016
o
Vn/V i \ J I -I-~+-+t l cd
. 6 -- \ --l---r- I jv r- . 012
tili / q \ t I I 71 ~--
4 ~-+----+ -- t~ k - - I I I E / +----1-- ··- - ---- . 008
• 1 ---.......... 1\ A ! / I\ :
~ - I-- ~ ! "0 1 19 I -<f---t----a
+ 1! j I V
. 2
'~ + I t~I-++ '~ / < -- . oo ·~
.~ it- . I :r/ ~-+--i
o I _--L- _ ..L_......I..-~-'--- __ ,~Q I 'rJ - ~ : LLJ 0
-. 4
-. 2 o
. 2 . 4 . 6 . 8.
Cl, (c) A t/ An , O . 4~9.
Fi gu re 16 .- Co ncluded .
r-
I
Fig •• 17, 18 no!
~ en ~ ~ 1.0 .020 .8
/
v n/V / .6 .012
V
t.H/q .A/ ~ ~
-
.4 .008 ~
-
""
\
r ~
J
1 \
/
.2 .004
/
P\
/ "'- -= .Fl...m
'" o
o
-.2 0 .2 . 1.0
-.4 . .
-.
Figure 17.- Section characteristice for air f oil sh ape 10 a t R = 6 . 4~ ~ loP and with a etagger ot O.5~ inoh. At/An. O.5}8.
I . rp
I
GJ
/
I
1.0 .020
/
/
\
.8
1\ /
.\
.6 .012 / Ii
I'b
\ '/
~ ...A.
./
.4 .008
!F
~ . V
1\
er-
j
\
~
.2 .004
V
"'-
~
/
o o o .2
-.6 -.4 .6 .8 -.2 1.0
° 1
Figure 18.- section charaoterietice tor air t oil .h ap. 10 at R = 6.43 x loP and with a
stagger of 1.10 inoh. At/An, O.5~8.
IA.eA ~.
s 1.2 . ~~ p- ! "~~ fJif.l''-t'JI:f'--- --f----j ----+--- -- -- --l---+--+------ -I- -t- ~~ _t_~I___+___{ i.o c
~f ~ ~ vn/v '" 0.556; AtA n 0. 6 55 i ""'f',., '\~
I'IrH W '-t- -+--t---,-----jI- I 1 1 I /' -+- ', --t--t--t--r-"-d;;---t--I
I
i 0 Upper surface --t--+- -+ --+- - - , -+ - -I----t--+--+--t--
.8
~ -+ ; -+ I --t- "l Lower surface I ! "
V n/V '" 0 . 440; At An = O. 538 - -- 1-,-+--I--I___+-,/ ---j I i
.6 f--- , '----+'- -t- 8. Upper surface I I I I i -1-- - - -t- --t---;- I ---j
i I 'if' Lower surface I i I
vn/v", 0.36 4, At An = ' 0.449 --' I --:- -t--+--+-- '-- -1 1----'-- +--+ - G Upper surface +I_-f'- -t-I-+ -+-~- -'--+--+--+--t-- --I
.4
; (> Lower surface I I !
I__ -+- , -- --j.. - - r - - - , -I' - I- .2 I - .. -I-- -+- -1
I---- I I---- ---t--+- --t-~ -- I i- . I L __ I~_ '1"-_---'-- ~-'----'
o .8 1 .0 o .1 .2
.4 .5 . 6
x/o Figure 20.- Pressure distributions fo r airf oil shape 10. a , 0.55°; R, 6.43 x 10 •
J _ - ~ -= -1: _ + ~, iT- --r ~-+-- --+---1
s 1.2 I
>/" : !: ±;' I ~ I
, ,
i I I '-'~ - __ '~_:: - - t- ~ - ~- r ~- _--'- ---'
, 1.0 ,
~ 0 Airfoil shape 10; V Iv '" 0. 44 0; A.tA n '" 0.53 8 I 'i ~r"r~ '{
I ITT--+- - t- ' I I I n , !I I I I I~ , - i ~ -- - The l o ret l ica ~ pr J ss ui e d { s tr ibu t ~ or!.n ! I , .8
I NACA 65 . 2-015 r rf , il sec ti on 1 1- I ~
~ ,
~~-+ -. 1 --t-~~ ~~~ ~--~
.6 I; i -+! I ~ :! 1-' L
1'---+- - 1+-1 __, I _, I- I
! , 1
11- 1 ~ I I---+- -+- __ i -C ~~_ . _ l
.2 o o .1 .2 1. 0
.4 .5 .6 .8
x/c Figure 21.- Basio symmetrioal pressure distribut i on f or airfoil shape 10, derived fr om figure 20, oompared with the t heoret i c al pressure dis t ribution for an NACA 65,2-015 airfoil section.
rig . 22 IACA.
Cd 8-
vn/V
[] AH/Q 1.0 .020 .8 (!'...
\
/
.6 .012
\ /
6H/q ~
V-
I \ ,
.4
.008
j
I
I
i
\
I
.v
--< ~. i
I I
.2 , i _.
r '" I I I
! V
i .l.
""" ~.
o o -.6 0
-.li -. .2 . .6
8 1 .0 (a> At/An, 0.594.
Figure 22.- Section characteristics for airfoil shape 11.
R, 6.43 )( 106.
I I
I
I I I I I !
: L 1.0 .020 I : !
! I : I .016 .8 I ~ .--- - --
/
' 0- .6 .012
~ I
/
b--.
6H/Q -
!---d
,<
I
.4 .008
l!r
l'
I
..JU.. - LP
.2 .004 ~ ,..- - -- EJ.., /f-" o o 1.0 -.6 -.2 o .2 .6 .8
-.4
(b) AtlAn. 0.459.
.. ~ c.o .,
III ~ ~ .. : ~
I w -
\'
I
"~
6, \: .9
~lI-
-+--+-~ x _
\ -
~.
I 1\
II II
L-694 \
~
6.4~
1 ~ r
.B ~~ R, r\ ;
T
\
__
~ r
.7 II.
~ r \ 11.
or --"".
I I •
- ---1" on I r shape x/c I seotion Ion I
I I I I I I I I I I
+--+--+--+--+--+--+--+--+- airfoil .459 d1stribut At/ 0.594 -- • I tor airfoil -r
~
I
r- j
pressure --i--
~
I
~11
.360
-
Vn/V 0 . 480 66,2-025
17" I II
-- 0 6 distribution --~-r 1- HAC.4- I -'- --- an Theoretioal
~p~~ v
I I I I I I I I
Pressure 11/+1 v/ I I , -;.- -
~-'-'~~~~r-r-r-~r-~~~~~
, 1/ (/V ! ' 24.-
T
J
,. I
7l I
: I I I I
I r 0.1.2.~.4.5.6
2rl--r--r--r--r 0' sl .
l.~~~ I" 1.4 l2V l.Cl~
Figure 11.
.459
AtlAn 0.594
shape .~58 0.500 Vn/V airfoil I , I ! , o 6 for
t
19/ II
j
Ii
VI
II
~
coefticients
!
o
j
106.
tr
lift x
J
II
\ I '"
+ Ii!
-4
6.4~ section R, -8 2~.- o .6 .41 .2 .8
_.2 -.4 -.6
1.0 c Pigure lUCA o o~g· 25 A Vn/V
~
U) I I!I oUI/q H 1.0 .020 e..
/~ ~ .8 .016
" V
vn/V
'"
'" '\.
/
.6 .012
\
I!I \ I
6H/q ~ ~ ~
~ /a
~ ~ ..,..k ~ ...........
. 4 t V 4r
~ /
\
!
"'~
""
~ "\ '~l",
A V
.2 .004 . ~
1 /
'"
~ ,.
(II.) At/An, 0.642.
,.r-(' o o -.2 o .2
-.6 -.4 .6 .8 1.0
r.
.8
, j/
-
+-~~
V '
I .
r.< .6 .012
/
~
AH/q °do
/
/
.008
.4
.---- / ~~ ~ --y
\ I
--
,
"" /
G1
dO
;1 \ ~ .004 .2 '" ./
'\J
/
~ (b) At/An, 0.496 .
,/
t'---a o o o .6 .8 1.0 -.2 .2
-.6 -.4
.016 .8 / ~ .6 .012
V
I, AH/q V "'\.
V
r;J cI /
~
~ Ib- ~.
~;.p ~ .2 .004
,t
~ ~ (c) At/An, 0.439.
/
'S.. ~ o o '" '" o .2 .8 1.0
-.6 -.2 .4 .6
-.4
J'igure 25.- Section characteristics for airfoil shape 12. R, 6.43 x 10 • NACA Figs. 26,31 1.0
I Vn/V
At/1L 0 0.500 0.64 .380 .496 (; .270 .439 o.
• 8 '"
If
';.
.6 /;
~
.~ /;
l.f
.2
I
l!
o
j
If
-.2
/I
.~
-.L ;
I I
-.6 -8 ! -4 8 12 o a Figure 26.- Section lift coefficients for airfoil shape 12.
R, 6.!~3 x 10 • • 020
I
I I I I
I
I' o Airfoil shape 10 6 NACA 65,2-815 airfoil .016 section (TDT test 3)- ?
/
.012
/
'"
~
\
/
/
.008
V
1\
L!
/- Y
~ ~ \
L/
~ ~ .004 • o -.6 .6 -.2 o . 2
-.4 .8 1.0
Figure 31.- Comparison of low-drag range for airfoil shape 10 from figure 16 and 6• NACA 65,2-215 airfoi l section, R, 6.7 x 10 NACA a7,a 8 F1gs.
Vn/V At/An 0.500 0.642 A .~BO .496 El .270 .4~9 -- - - The oretical. preaaure diatributio n for an NACA 56.2-01B.B7 ai r foi l aeotion.
1. 6 .- - .......
. -
F= 1- -
-
~ - I- - \ ./ /
M ~
I
!Y
\
~~
1 J
1.2 I I ~ .
I
I
r
I
i I
1 I
i
\~
1.0 I
~ ,,"\ ~
I
1 i I l
I : I 1 \
I
I I .8
I
I
", '\
I I
'" I
I
I I
I
.6 I" " I I
I I
I I
I
!
. I
i !
, I I I I I
i I ~
.4
I I I I
I I
i
I ~
, I I I
! I
I I I
l
L~
.2 I I I I i
I
! i
I J --' I I
i I I 1 I
1 J J
o .6
.2 . 8 1. o
.4 .~ o .1 5 7 .
• x/o Figure 27.- Pressure distributions tor ai rf oil shape 12. a. 0 ; R, 6. 4 ~ x 106.
-
+--+-- -+1---1 1--- 1- - _-+----'_ 0 C do ~ _+-I f" I----+--+--+----l
.8 .016
I i A V n/V /
~---- ' _- ~ ~. ~+_ --~-- +__r --~~--r-~~~ \ ! 0 lI H/ q / I I 1\ \ .6 . 012
I
. 008
.4 r----
. 2 .004 I o o -.6 -.2 o .2
-.4 .4 .6 .B 1 .0
0, Figur e 28. - Sec t io n ch ara cteri st i os for airtoil shap e l~. At/An, 0. 671; R. 6 .4~ x 106• .
ru
IZI ..... ... til ~ ?:l
~ ~ oq I 0.671~w .
.9 ~ surtace lurtaoe l'\.
AtlAn, -694
'"
~ L uPper Lower .8
~
f\ I
o /),.
0.560;
r\
~
.7 vnIV,
"" ~
......
- .
l~ . 6 x/a shape r .5 I--
-
a1rtoil
.4
tor n s ;r-- .~ I i j...---- .
° .
d1stributio .2 """~ x 1.092 t- a, 6.4~
t-- V
.1 (a) Pre.sure R,
V I
v"
I
~o.- ~
I !9 ; j
o 0) , b , I I , ~ ~ ) ~ .8 .6 ( a Figure s 1.4 1.2 1.
2.0 1.8 1.6 .
l~.
shape 106.
x a1rto11 6.4~ R, tor ; ~
I
r5 0.67i
/ ·
/
f:> / Ij coetticient, At/An, a o J
I
'P ~ 560; 11tt .., 0.
V
, Gl -4 Sect10n Vn/V - -8 .
29.- 0 8 6 4 2 0 ·2 ·4 ,6 1.
C Figure IIACA' Fig. 3 0b ,c - I-- --" 1.6 _:---f.---
--
s 1.2 1.0
" ""\ -
~ I I _+~I--r_+'~ ~1--'--~!-4I-----~-~- ~ ! --r-+-~-- ' \-~~~~
.8
I I I I I I '\
.6
ft- ~ -__'_ t- _ ~_=~-~---~- ~~ - .~ ± h + -~ c-~~_+_I--__t_~ ! . ~ ~
o Upper surface I I t I : I I
6 Lower surface , .4
t--- --- ----- - - - Theoretical pressure distribution 1 I I
I , for an NACA 66,2-021.77 airfoil I
It--- ' --t section. " 1 - -----
.2 ____ mil I!-+_I Illi
f---- -- -- -
'-1 -j-- ~ t -~ I: 1 (b)
o 1.6 s 1.2 1.0 .8 .6
i I I I I!: I '\
1++' I 1 I ; I I
.4 I j:
: =: ' =:=-r-+ --- ====:=: I ~ I r-+ ! ~_ ~~~:~:I~~r!I ~::~:~~ I :~:~: , ==
I I I' ~ I I ! I
.2
I I I I
iii ( c)
I J I !
I o o .1 .2 • ~ .4 .5 .6 .8 1.0 ( b) 0,0°; (c) 0,0.55°.
x/c I .: ... ~re 3 0. - Co nc l ud ed .
I J