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‘ NATIONALADVISORY COMlhTEE ‘- -FOR AERONAUTICS —..

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~Oo 976 ..-= -=.--,---

TESTS QF AIRFOILS DESIGNED TO UCLAY THE

COM?FU3SSIBILITY BURBLE.-. .;—-—.

By John Stack

Iam?lev Memorial Aeronautical. Langley Field, Va.

.’

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Laboratory...

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December 1944 “.-,-=--7,.. , .,...-.:—.,-—-” “-. -..-’-L-

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(Reprint of ACR, June .1939) ------.:- _;,..-.. . ,. -. ---- - ..--.-:.-

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https://ntrs.nasa.gov/search.jsp?R=19930081766 2020-04-16T20:53:45+00:00Z

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NATIONAL ADV_ISORY COMiKtTTE33FOR AESONAtfTICS-- .. .. ~ ___-.”.

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TECHNICAL NOTE NO.

. . . .

976 :-&--: . .. . ..---- -- . .. .... . .

.. . -----... .@~~Ay ~~ “- >’. -. -=,

..=. .-,-: .,. -..

UmPRESSI”BILITY WJRBLE : . ‘: ~:= .T::.~<. .

By John Stack “ ‘_ ‘-- ‘--’=---.”’-‘-.-

..._...____--

SUX!MRY -- -. —

-- --- - .--“-.:.Fundamental investigations of compressibility

phenomena for airfoils have shown that serious tidversechanges of aerodynamic characteristics occur- as the - .local speed over the surface exceeds the local speed of - -:‘

...-

Sound s These adverse changes have been delayed to higherfree-stream speeds by development of suitable airl’oil

.

shapes. The method of deriving such air~oil shapes is...—

described, and aerodynamic data for a wide range of Mach.—-_

numbers obtained from.tests of these airfoils in the . ‘: .-”_:Langley 2L-inch high-speed tunnel are presentad. Theseairfoils, designated the NACA 16-series, ha-ve increased - -– .:critical Mach number. The.same methods by which these -airfoils have been developad are aairplane components.

~l~c~ble to other .- ‘“>-~$ “~.---- -, . -,=---- .::.~._-:-,-----c .. .--T-: .-.;..<— ..-.

+. e ~ 7. .--.,--—----~--..Il~TRODUCTION ~ ...-~;. .=;;------- - --:# ------------- ‘-=-=-”+:::.:;=~--:--=- -.., - ....

Development of airfoil. sections. suitable for high-speed applications has generally been difficult because- -little was known of the flow phenomenon thstoccurs “athigh speeds. .A definite critical speed has been found ‘ ---at,which ssriaus.:”detrimentalflow ch~ges occur thatlead to serious losses i-nlift and large-increases indrag. This flow ~henomenon, called,the compressibility *burble, was originally a propeller problem.but, with thedevelopment of high-spe~-d Rircraft; serious consid~ration - --–has to be given to other parts of’the airplane. It iS . .

important to realize,-— .

however, that the.prdpeller Willcontinue to offer the most serious compressibility ;,.= -: ‘-problems for two reasons: first, because propeller.section sp~sds.:arehigher than tinespeed of.the qitiplaneand, secondj :be.cause-:slmucturalrequirements lead tothick sections near the root.

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2. . ,, NACA TN No. 976..” ,.. -

.,-,. ...Fundamental Investigations of high-speed air-flow

phenomena recently completed (references 1 to ~} haveprovided much new information.” From practical considera-tions an impdrtant conclusion of tliE3B~ investigationshas been the determination of the .cri.ticalspeed, thatis, the speed at which the compressibility burble occurs.The critical speed was shown to be the translationalvelocity at which the sum of the tral~slstionalvelocityantithe maximum local induced velocity at the surfaceof the airfoil’or other body equals the local speed of9ound ● Obviously, then, hi@.er critical speeds can beattained through the development of airfoils that haveminimum induced velocity for any given value -of’the lfftcoe~f’icient.

Presumably, the highest critical speed will be

attained b’yan airfoil that has uniform chordwise distri-bution of’induced Velocity or, in othsr words, a flatpressure-distribution curve. All conventional aitifollstend to have liighnegative pressqres tid col’despondingly.high induced velocities near the nose, which graduallytaper off to theair-stream conditions at hh”erem ofthe airfoil.’ If the same lift coe~fictent cafibe obtainedby de.creasi~”the induced velocity z-.isa.rthe no$e and “inc]?easing the induced velocity over the rear portion of’the airfoil, the critioal speed will be increased by anamount proportional to the decrease obtained in themaximum induced velocity. The ideal airfoil for anygiven high-speed application is, then, that shape whichat its operating lift coe~ficient has uniform chordwlsedistribution of induoed velocity. Accordingly, an ana-lytical search for such a~rfol.lshas been conducted bymembers of’the staff’of the L~”le~ Memorial AeronauticalLabcwatory and these airfoils have been investigatedexperimentally in the Langley 2)+-inchhigh-speed tunnel.

.“The 11’rst airfoils investig-a%ed showed m~rked

improvement over those shape”salready available; notonly was the critical speed increased but also the dragat low ‘speeds was decreased considerably. Because ofthe marked improvement achieved, it was considereddesirable to extend-the thickness and the li.ft-coeffic’fentranges for which the original airfoils had been designedto obtatn data of’immediate practical value before rurtherextending the investiga-tio”nof the fundamental aspects _Of the-~prob~emo —-.—

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NACA-TN’Noo 976

SYMBOLS “ ‘

3

.. -,

x abscissa of camber line,..

Y~ ordinate of camber line .

t thickness, percent of chord,-

C airfoil chord

e defined by ~ = ~(1 - cos e)

CL lift coefficient

CD drag coefficient

‘Dminminimum drag coefficient

%C,L Pitj~~-~-moment coefficient about

1? pressure coefficient

M

.>

... . ..::

...

.——

.-.

---—

=—.:.., --.—.

.,.

quarter-chord

.-

M’ critical Mach number..

cr

R Reynolds number .,.-.... ‘:”:-..+—.

a angle of attack, degrees .,..’ .. .>..=...’.,..

. .. . .. -

DEVELOPMENT OF AI~OIL SERIES “ .,. ,- ......-’i::--,...

The aerodynamic character~s’tics of any airfoil ”ar-e~in general, dependent upon the air~oil ctiber -lineand -the thickness form. Mean ctiber lines W=Pe derived. ~analytically to obtain a uniform ch.ordwise distributionof induced velocity or pressure for certain designated .lift coefficients, an~ an analytical search for a thick-ness form that likewise has low and uniform chordw~seiindn~ed-velocity dist~ibution was then undertaken. .

Derivation of the camber line.- Glauert tr&fe$en5&L)has derived expressions for the local induced velofiityat a point on an ai.rfoi.l(zero thickness assumed) in

....-.—... . -

terms of’the circul~tion .aidund.an aizfoil correspondingto a certain .distribution of vorticity along the .aIrfoilsurface. .Eyassuming the distrlbuti-on of vorti.city to beconstant, a line airfoil is determined ‘that givas uniformchordwise p“re”ssuredistribution. TEs form oftton go derived is

.,. -.

where T-l :s the ordinate of t.h~mean cm~cr

(1). ._ ,.=---

tha abs~~ssa, and the ckcrd is taken as unity. Theidealized form described by this equation has discontl-nui.ties at tke nose and.at the tail:, .ThiS difI’iculty ~s.circumvented by assuming very slight gradients in theChordwise load distribution just at the.mse and just atthe tail. This form, derived by wshg the Fourier seriesmethod, is given by the equatl~n

.-,...’-. ;.J~..--:. .=.=-..

yc CL—= @O.3833 - 0.3333 Cos 2f3= 000333 Cas b? . - :c

- 0.0095 C09 ~fl - Jei30@ COS 86.. ,.. i- .—-,.4 .-..

xwher~> - = *(1 - 00s e) and c is the airfo~~ chord.c

Equation (2) expresses thQ mean cmbgr ‘line choi% forairfoils of the series developed. Load or i.nduced-veloci.tygrad@s derived from both equations (1) and (2)are actually identical for all practical .pilrposes. Mean-camber-line ordimates are given in table I Ior CL = 1.0. -

In.order to.ob&aln the rfleancamber,line.giving uniformcb.ordwlile,dis.trib.uti.onof. inc?udedvelocl$~ for,a~her

,-

ve.luesof tij+e:,l.l$tcoefficient; the”values gl$5nYintable I are r?m~tfplietf,b~the value of the desirbd.lift ~ ....coefl~ctent. .’. -.,.

,.. . . .

IW5ivatidti,of the.-~hib.l{nessforh.-,The deriva”~ionof th=.thicknesstforzn is .no~.q.q..”s~~eor ,direct.as thederivation of the,meati camber llne. The theoreticalpressu:re.dl.stributiou.wg~;,computsdby the znetkio@s ofrefere]i-tie,:~::fo~,,0acQ..Q~;”t@6;.”s-Gveralth$ckgess-$cpms .,investlgated,in reteqence~... .S6m& of “theseifor~+, .

NACATN NO. 976

approached the deskti~d shape but further modificationswere investigated analytically and~ finally~ two shapeswere chosen for tests. These shapes, the NACA 07-009and the NACA 16-009, an~ the thpor~t~c?,} pres$:re distri--bution for each ar~ shown in fi~.e k?. .The completeairfoil profile is derived by first c.alculati~ the meancamber line for the.desired Iift,c,oef”ficient“and thenlaying out the thickness ordinates glveh in table 11 fromthe camber line along perpendiculars to this line:

-..Airfoil designation.- Because the ideal se~iqs of

.-._

airfoils requires an extremely large variation of shape,it becomes practically impossible- to use previous num-bering systems and, further, because this new series ofairfoils is designed to obtain a specific pressure.diagram,these airfoils are designated by a new series of-numbersthat is related to the flow and the operating character-istics of the airfoil. The first ntiber-is a serialnumber that describes the class of pre$sure distribution,the second number gives the-location of the i@ximun”nega-tive pressure in percent of chord from the leading edge,.the first number following the dash giveg the lift coef~ficieritfor which the airfoil was designed to.o~erate, ““..andthe last two ntlmbers give the airfoil thickness inpercent of chord. Thus the NACA16-509 airfoil has-theshape of the NACA 16-009 disposed about the uniform.chord-wise load camber’line designed for”-a lift c@5ff’ic$ent of!O*5”

Airfoils investigated,- As previously stated, ‘two ‘-basi~ airfoils were investigated. (See fig. l? ) TheNACA 07-009 airfoil should, theoretically, give highercritical speed than the NACA 16-009 but =earlier inves-tigation (reference 3). indicated that, for pressures

# occurring near the leading edgej tiie’”tnqreas.gifithepressure coefficient as a result of compressibilityeffects was gretiter thsn that for pressures oc”ctiringfarther back on the airfoil. Consequently, it was believedthat, at speeds as high as the critical speed, the NACA07-009 airfoil ’might have, as a result.of compr~ssibility

., effects, a pressure peak near the leading edge~ ~~ “’”““-.NACA 16-009 airfoil was therefore develo.ped in afi~ttemp~.to achieve the”uniform c~ordwi$e lo:d distribution at .high speeds. Botlnforms were tested and the resultsshowed higher drag and lower critical speed for the NAC.A ::07-009 airfoil. Accordingly, the NACA 16-009 airfoilwas chosen as the basic form for a series of airfoils

, ..-...-,.,“-..

. -,.. ..,. .,,LZ-..: .:-----,t ,:,’.- :,....:G-

——

. ——

.,

----—.—--——

.C”.-=,...”

t-- -

6 NACA TN NO. 9’/6”

designed to operate. at various lift cae.$ficients.:.~.or,+.one value of the li.~tcoefficient the effect o.f..thic,knessvariation was also ,investigated. The airf,oi$s tested,of’which prol’iles are shown in figure 2; are as follo”ws:6

NACA 16-009NACA 16-109HACA 16-2o9NACA 16-509NACA 16-709NACA 16-1009NACA 07-009

APPtiATUS

NA~~ l&@ ‘-“NACA 16-5w,NACA 16+15NACA 16-521NACA 16-530NACA 16-106NACA 07-509

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AND METHOD

The tests were conducted in the ~-&gley 24_inchhigh-speed t~el, in which velocities.a~proachlng thespeed of sound can be obtained. A brief descr$~tion ofthis tunnel is given in reference 3. The balance meas-ures lift, drag, and pitching noment snd, except rorimprovements .thet permit a more accurate determinationOF the forces, is simil&r in principle to the balanceused in the Langley n-inch high-speed tunnel. Themethods of operation are likewise similar to thoseemployed in the operation of-the Langley Ii-inch high-speed tunnel (reference 7).

The models were of ~-inch chord and 30-inch spanand were medd gf*c@alumin. A complete descrfptiion orthe method of constructing the models is given in ref-erence 8. The model mounting ia similar to that usedin ttLe.LangIey n-inch high-speed tunnel (rqference 7).The model ext+nds actiossthe tunnel aridthrough holes,which are of the same shape as but slightly larger thanthe model, cut in flQxible brass end plqtqs .~hat preservethe c~ntour of t-hetunnel walls. The model ends aresecured in th8 balance, w~ich,extends halfway around thetest section and is enclosed ~n the airtight tunnelchamber @Qlar to the insta-llation in the Langley n-inchhigh-speed tunnel (reference 7). .

The speed range .oyerwhich ”rne~.su.r~mentswere madeextended, in general, from 25 percent of the speed ofsound to values in excess of the critical speed, Thecorresponding Reynolds number range wqsj,&orn approxi-mately 700,0’00 t’onearly 2,000$000. The lift-coefficient

NACA TN No. 976 ,... .. .t 7

range for wh~ch tests were made eTt@nded from zero liftfor each airfoil to values approaching maximum lift.

.

= - -__= -.,.,.PR%CI”SION, ,’-.,:.”,,’.-;. . -1.,

.,,,. . .... ..

. Acciden”tai‘e.rror.sare indicated by the scatter onthe plots showing the measured test data (rigs. 3 and23 tO 25). These errors are, in general, rather Snlalland affect neither the application nor the comparison ofthe data. Tunnel effects arising from end lea~age,”restrictions and the usual type of’tunriel-wall’effectare important. F~act k~owledge 0~ these various effectsis incomplete at the present tim~:. T& ‘laYgGst‘effectsappear to arise from air leakage through the c-learancebetween the model ~nd the brass end plateS in the tunnelwall through which the model passes. Investigations of

-

. -. .-.

—.——

. .

.

the leakag; effects have been made “for the NACA ~0~2 air- .foil with a special .type of’internal gap or clearance “ --that pem.its wide varistion of the gap. .Data-obtainedwith various gap Settings of 0.01 inch and larger ‘extra-polate:d to zero gap were used to evaluate the-leakage --- - -:----correction for the standard type Of no~ttng=. ‘These

..

corrected data -ware then checked by means of.wake-survey - ~drag measurements witinend le~age ~l~i~t~~ b~%bbe~- .

.-.

seals. Because the balance chamber is ~fi.rtt$ht,the. ‘.:end-leakage condition is related to,the pres~ur~ ‘distri- .bution around the’model. It was thefefo”re“coti$>de%ed .-v.advisable to check the method of correction for end”leakage by-wake-survey tests with end leaka@” eltii~t>@ ”’”” .by rubber seals for these new @.rfoils, which.have ~~f= ,:..tally different pressure distributions from the elder “airfoils such as the NACA 0012. Some of thes,e*ta. are ‘

.

shown in figure 3. In general,- the agreement”is excel-lent. The data have accordingly been c~~e”cted for e~d-leakage effects. ,,- ..,’” .,.-. .

..

Other tunnel ef~ects have not been-completely inves-~” “‘ tigated and-the data have not been corrected for such

effects as.restriction or the more- usual type of walleffect: As presented, the data are therefore conserva-”tive, .lhasmuch as investigations made thus far” indicat~that the coefficients are high and the critical speeds” ‘ “ ‘“may be low, Strictly comparable data for two oldey air-

I foils,: the 3c8 and the NACA 2J09A3L, forT~tiio-Mach-~bersare included so thati~comparis~n~ can b~”made. ‘-”’”’-I-- .::j:

,. -..,, ..;.... !?. ..:.. ..-.

—

.—

El

DTSCUS310Y r .,”,. ., .. :—

&erodyn&.ic “charecteristic$ for several of’the N4CA16-series a~rfoils are given irifigures ~ to 18, Exami-nation of’these figures’ indicates two important discrep-ancies between the theoretical design cor~ditions and tk.edata obtained from the tastss First, none of the’airfoilsattains the design lift coe~~ici~nt +t the desi.&n angleof’ eltt&ck (Oo) and,”second, the departure incred~esma]?~edly with the design lift coef’~}cient. The departures??@.be importan~ If .v~riatlo”nfrom the ideal pressuredistribution is rapidwith change in lift coefficient.This effect, if great, would t~nd to cause lGwer drag :,ap,dhigher critical speed fir’a narrow region near the”clbsign,condition than are shown by these data. ~he~e !:

departures also increase with the airfoil thickness.

The .diff’erencesbetween the designconditions andthe actual test results.r,aybe expected because of the ““simpl’Lfyingassumptions of the thin-airfoil theory..Theclretically, i.td-sassumed thet the i,nduced?velocities&orene-gligiblysmall as compared witi> the stream v~locity.For thin airfoils at low lifts, this approximation isvalid. ,W’ithincreases of lift or thickness, however, theinduced velocities approach aridsometimes .exceeci”thestream velocity. Study of these effects. appears to-bevery “important in order to obtain the proper ai.rl?oils-for high lift coefficients and large thickness ratios.DeviaH.onqshown by the airfoils in this series havitighigh lift and high thickness ratiiosappear to indicatethat,the use o~--asin@e basic ~hspe Is..unwarraratedifit $S+’.desiredto obtain optimum ,alr<oilsfor a wide rangeofr”IY”~.~.~ef’ficiemt .

.,..and ‘hich4W ‘@distribution”’”

. .

,,‘?%eoret”icalpressur~-d+str~~utio~, di~grams for th~

th$bk~r airfotls.showed much greater slope.af the p~eg-sur~’”curve tk,anis shown by the basic NACA 16-oa9 airfoil.Preliminary study indica~d that.,@cJr.easing the leadi’ng-edge rad+uq .~d,.~the“fullness of .th,eairfoil b.e~e”r” thel~adin~ edge.and..,themaximum or@@tit@ may-lead to con-slderaljiejmprovplnentiv”efi ths”thicka,r airfoils hereinrepo-rte-do ::.

,... ,., .----., .—

.. . . . . .-. -”.comparis.~fi.ofaii+f’’oils~-~l@res 19.and 20<i.llustrate

the.dfi~erences in ~erodyri~~c ”char.acteristicsbetweenolder propeller-blede sect~ons “and the NACA 16-series -.

.—

NACA TN No=’976 9

airfoils. At lower speeds (~ = 0..45;.fig. 19) the5c8 airfoil appears to attain a much higher maximw liftcoefficient than the new airfoils.” This result ‘is ifipoF-tant in that the wider u:sefizlangle-of’-attack rtige nisyfrequently be required to prevent stalling af a propell~~during take-off. Over the n“ormslfli’@htrange,.however, “-and in most cases for which rational choice of sectioncan be mlade, the lower drag-of th”e”newsec_t&Q.nsoffersconsiderable opportunity to achieve higher efficiencies.The LOW drag attained by the l?ACA2409-34 airfoil developedfrom earlier tests in the Langley n-inch high-speed tu”n-nel may appear surprising. Actually the type of flow forthis airfoil approaches the flow that might be expectedfor the NACA 16-3o9 airfoil. -----. —.

The low drag common to most. of the NACfi 16-seriesairfoils is associated with more extensive regions oflaminar ‘flow in the boundary layer resulting from therearward position of the point of msximum negative pres-sure. Unfortunately, however} the Reynolds number is solow that effects of laminar separation may appear andsome pressure drag might occur. The small differencesin drag between the.envelope polar for the new atrfoilsand for the NACA 2409-34 airfoil are prabably a resultof this phenomenon. Actually the point of_maximum nega- ‘tive pressure for the new airfoils is considerablyfarther baqk than the corresponding paint for the NACA2)+09-34.airfoil but, if laminar separation occurs early,nearly equal drag coefficients might be expec-ted.

. . . . ..-~+ :--.+. ----At high speeds (M = ‘“7’5sfig- 20), the pegion”~--- .’-

which the NACA ~6-series airfoils were designed, thesuperiority of the new airfoils is clear-.”__TF&’eA21ieronset of the compressibility effects .f?r the older air-foils leads to early drag increases and lowered maximum

, lift coefficients. At speeds above M = 0.75 the use ofthe blder sections appears unwarranted for any purpose.

... - -~..___Crtt.ical speed.- The variat~on=of tie “critical’” ‘“--

speed with lift coefficient and with tlnickness is givenin figures 21 and 22, respegtivsly. -These curves indicatethat critical speeds exceeding the theoretical values - ‘“”were attained in the tests. In the choice of the test ‘ ‘“critical speeds, the values were selected on the basis ‘“ ‘~of earlier experience that indicated some rise in drag ‘“’;before large Flow disturbances occurred. If these sp6edswere chosen as tk.e highest values reached before any - ‘:’”appreciable ”drag:inerement occurred, the agreement with ‘ : -

. .-.._ _

.

.-

.-

.,

- -_‘“”- “–——

10 NACA TN Na ; 976

,,tke theoretical .curves,”V,ouldbe $e’r$.gtio~”.FcJroornpenlsonthe critical speed of’the 5c8 airfoil is p’lotted.im fig-ure 21. The.dit’feuen.cebetween.the,new and the olderairfoils 1S greater than shown by the curves bec”ause the3C{3is 8 percent thick, or 1 percent” of’’thechord thinnerthan the airfoils of’the NACA 15-009 series. #*..:L. ,,

.

Minimum “.drag.-Coefficients of minimum drag plottedEi~&liIISt Reynolds number are given “inf’i&m?es,23 and Z’&The lowest drag coefficient was obtained for the NACA16-106 airfoil; this coefficient tk appr~ximately 0,0026at low speeds and increases tcnapproximatel~r 0.0032immediately below the critical speed. Cf the”~-percent-thi:k airfoil series designed to operate at various lift

coe:Yficient~ ~he.NACA 16-109 airfoil appears to havethe lowest drag. This result is contrary to expectationbecause the symmetrical or basic form of the.NACA 16-W9would normally have the lowest minhHJ.M drag aoeffici.%nt.The difference may be due to some irregularity of theairfoil surface. .

The comparison of the.minimum dra%

coefficientsfor the 3c8 and the NACA 2409-3.4 and”l -?09 airfoils Isshown in fi~re 25. The high critical speed,f’or the NACA16-2o9 airfoil is apparent,. The comparison as glverl.di.rectl.yb~ figure”25 is a little misleading because ofthe mmller thickness ratio for tke 3C8 airfoil. Forequal.thickness rstios, the differences between theC-series and the NM2A 16-series airfoils will be greaterthan shown. ,

,USeof the “data.- The envelope polars that hay bedrawn fur the NACA 16.series airfoila represent a “newand much lower dtia.ga“swell as higher c“r~tical speed ‘-attainable for the design of propeller-blade sect~ons.Ev&n though the angle-of-attack range is less than for,,the older sections, there will be numerous designs furwhich sufficient angle-af-.a’ttackrange is given by thefiewsections. For high-speed, hi@~-aJtitude aircraft,the advantagesof the low drag an”dhigh critical speedare of paramount importance and, in the~e designs, rationalchoice or section is of increasi~”i”mportance .

,.In many

designs the diameter is fixed by considerations otherthan pl?opeller eff.icl.ency. Thus the induced losses a“refixed and propell~rs of hi@est efficiency can be developedonly by operating and designing the blade secticin”stooperate on the envelope polfir,s. Another import~nt con-sideration in using new blad6’sections to achieve highest

.,NACA TN NO- 9.76 11

efficiency concerns the adaptation of the sections toolder propeller designs. Optimum efficiency csnnot beachieved by simply substituting the new sections for theold on a given design”. The use of better blade. s&Ctionspermits the use of larger diameter and necessitates some”plan-form changes. ,..A11these factors should be consideredin a desi&n for best e’~ficiency with the new blade-&4ctions...,. ..

CONCLUSION.- .

By a new approach to airfoil design based uponfindings of fundamental flow studies, a new series ofairfoils, the NACA 16 series, have been developed whichhave increased critical Mach number and at low speeds “-

.-

—

.. --~

.-

reduced dr~g. ---—.--:;---~~.. -..._. ...=

a

.—>. —

Langley Memorial Aeronautical Laboratory..—;

National Advisory Committee for AeronauticsLangley Field, vs., June 4, 1959 .-

.—

—..

.--—

._.—

12 “ ~ACA TN NO. “976 ‘“’”

-.,.

1. Stack, Jokn.: The.C”ompressibility Burble. NACA TNNO* 5’!!5’,1935. “ ._.._ ._...-——...,’— — -..

2.-Jacobs, Eastman N. : Methods Enployed in America forthe Experimental Investigation of Aerodynamic

.--,

Phenomena at High Speeds. NACA Misc. Peper No. G, ......->1956. ...—.

39 St~~k, John, Lindsey, W.j“ “* “

“F., and Littell, Robert E.::?‘T& Compressibility Burble and the Ef’i’ectof Com--pressibility on Pressures and ~orces Acting on anAirfoil. NACA Rep. “No. @6, 1938.

~, Glauert, “H.: The Elements of Aerofoil and AiPscrewTheory. C~mbridge Univ. Press, 1926, PP. 87-93.

~. Theodorsenj Theodore: Theory of ‘LingSections of’Arbitpary Shape. NAC.4Rep. No. @l, 1931. -. # i, ,,

6. Stack, John, and von Doenhoff’, Albert E. :.>

Tests of16 Related Airfoils at High Speeds.NO. .@2, 1934.

X4CA %3p.

7. Stack, John: The N.A.C.A. Hi@.-Speed Wind Tunnel andTests of’six Propeller Sections. N/lCARGp.1933.

No. 463,__—.—-—..- —.——.-.--—.------- ---

so Jacobs, JJa~t~~~q~a, ~{~~~, Ke~*~~h ~ml””~n~ ~~~erton,i%obert M. : The Characteristics of 78 RelatedAirfoil Sect”lD5nsf’romTests ~n the Varf&ble-DensityWind Tunnel. NACA Rep. No. l&O, 1933.

*-

NACA TN No. 976

TABLE I

.

. .

CNmR-LINE ORDINATES FOR NACA 16- AND 07-SZRIES

AIRFOILS WITH CL = Ls~.

cAll values measured in percent chord

Jfrom chord line

.

,

.

.

I Station I Ordinate I Slope

o1.252.5fi

/ 5.~67.50

,10.00.15.00.20:00,25.00,~().oo.40-0050.0060-60

i(j.00“(j.m,ya.oO,95-OhLW).60

0.. I..-

3

.

.

●

.

.●

o“----.- ““9

.-

13

____ . ...-

*.tJaz3y

. .

-.

NATIONAL ADVISC!WComITTm FOR AERONAUTICS - “-”- ““”: -’”vI

.

A

.

. . _. _. _.

.

—

!-—=- ..-

I

1!

+- .:,-@-* =----- . .

r

‘ NATIONALADVISORY COMlhTEE ‘- -FOR AERONAUTICS —..

/___________________ --- ---- ____

I...-.—---—.----— .-. . ,—

MOO 976 ..-= -=.--,---

TESTS QF AIRFOILS DESIGNED TO D~LAY THE

COM2RXSSIBILITY BURBLE.-. .;—-—.

By John Stack

Iam?lev Memorial Aeronautical. Langley Field, Va.

.’

--A==

Laboratory...

+-<=.—-=—

---- ..

..., ..- .’+’- ... ... . . . . :

‘“”‘;’-’”““--wlii! ‘.‘“””~_“-‘:”:..-_:.--.—.a .%

.- r,

. .- --. , ----.:< -. : ~,-->” --- -%--4 ,4=..-

wa~hingt~f’” “’-,, ~‘:, .-,:~ .-----”::.-.::--=.—— ------.?J.:

December 1944 “.-,-=--7,.. , .,...-.:—.,-—-” “-. -..-’-L-

------ .—-: - , - - . . ... . . .: .._-:

(Reprint of ACR, June .1939) ------.:- _;,..-.. . ,. -. ---- - ..--.-:.-

-.

—

..

‘. . ..

.,

NATIONAL ADV_ISORY COMIiKtTTZEFOR AESONAtfTICS-- .. .. ~ ___-.”.. . . .

976 :-&--: . .. . ..---- -- . .. .... . .

.. . -----. .. .@~~Ay ~~ “- >’. -. -=,

..=. .-,-: .,. -..

COMPRESSIBILITY WJRBLE ~ . ‘Y ~=* .T;;.~<. .

By John Stack “ ‘_ ‘-- ‘--T:---.”:-‘-.-

..._...____--

sUXWiRY -- -. —

-- --- - .--“-.:.Fundamental investigations of compressibility

phenomena for airfoils have shown that serious tidversechanges of aerodynamic characteristics occur- as the - .local speed over the surface exceeds the local speed of - -:‘

...-

Sound s These adverse changes have been delayed to higherfree-stream speeds by development of suitable airl’oil

.

shapes. The method of deriving such air~oil shapes is...—

described, and aerodynamic data for a wide range of Mach.—-_

numbers obtained from.tests of these airfoils in the . ‘: .-”_:Langley 2L-inch high-speed tunnel are presentad. Theseairfoils, designated the NACA 16-series, ha-ve increased - -– .:critical Mach number. The.same methods by which these -airfoils have been developad are aairplane components.

~l~c~ble to other .- ‘“>-~$ “~.---- -, . -,=---- .::.~._-:-,-----c .. .--T-: .-.;..<— ..-.

+. e ~ 7. .--.,--—----~--..Il~TRODUCTION ~ ...-~;. .=;;------- - --:# ------------- ‘-=-=-”+:::.:;=~--:--=- -.., - ....

Development of airfoil. sections. stitable for high-speed applications has generally been difficult because- -little was known of the flow phenomenon thstoccurs “athigh speeds. .A definite critical speed has been found = ---at,which ssriaus.j”detrimentalflow ch~ges occur thatlead to serious losses i-nlift and large-increases indrag. This flow ~henomenon, called,the compressibility *burble, was originally a propeller problem.but, with thedevelopment of high-spe~-d ~ircraft; serious consid~ration - --–has to be given to other parts of’the airplane. It iS ..important to realize,

-— .however, that the.prdpeller Will

continue to offer the most serious compressibility ;,.= -: ‘-problems for two reasons: first, because propeller.section sp~sds.:arehigher than tinespeed of.the qitiplaneand, secondj :be.cause-:slmucturalrequirements lead tothick sections near the root.

--— -._-.””--:—._ 1-..-+

--+ >.-. ---

.—_ —-----

2. . ,, NACA TN No, 976..” ,.. -

.,-,. ...

Fundamental Investigations of high-speed air-flowphenomena recently completed (references 1 to 5} haveprovided much new information.” From practical considera-tions an impdrtant conclusion of tlieaeinvestigationshas been the determination of the .cri.ticalspeed, thatis, the speed at which the compressibility burble occurs.The critical speed was shown to be the translationalv(?locity at which the sum of the tral~slstionalvelocityantithe maximum local induced velocity at the surfaceof the airfoil’or other body equals the local speed of9ound ● Obviously, then, hi@.er critical speeds can beattained through the development of airfoils that haveminimum induced velocity for any given value -of’the lfftcoe~f’icient.

Presumably, the highest critical speed will be

attained b’yan airfoil that has uniform chordwise distri-bution of’induced Velocity or, in oth~r words, a flatpressure-distribution curve. All conventional aitifollstend to have liighnegative pressqres tid col’despondingly.high induced velocities near the nose, which graduallytaper off to theair-stream conditions at hh”erem ofthe airfoil.’ If the same lift coe~fictent cafibe obtainedby de.creasi~”the induced velocity zisa.rthe no$e and “inc]?easing the induced velocity over the rear portion of’the airfoil, the critioal speed will be increased by anamount proportional to the decrease obtained in themaximum induced velocity. The ideal airfoil for anygiven high-speed application is, then, that shape whichat its operating lift coe~ficient has uniform chordwlsedistribution of induoed velocity. Accordingly, an ana-lytical search for such a~rfolls has been conducted bymembers of’the staff’of the L~”le~ Memorial AeronauticalLaba$atory and these airfoils have been investigatedexperimentally in the Langley 2)+-inchhigh-speed tunnel.

.“The 11’rst airfoils investig-ated showed m~rked

improvement over those shape”salready available; notonly was the critical speed increased but also the dragat low ‘speeds was decreased considerably. Because ofthe marked improvement achieved, it was considereddesirable to extend-the thickness and the li.ft-coeffic’fentranges for which the original airfoils had been designedto obtatn data of’immediate practical value before rurtherextending the investiga-tio”nof the fundamental aspects _Of the-~prob~emo —-.—

.-..... .-.,,.-.. .:: ,.’,-.-

._.—

-

i

-—

.. ....— ——

..-

.. ,

NACA-TN’Noo 976

SYMBOLS “ ‘

3

.. -,

x abscissa of camber line,..

Y~ ordinate of camber line .

t thickness, percent of chord,-

C airfoil chord

e defined by ~ = ~(1 - cos e)

CL lift coefficient

CD drag coefficient

‘Dminminimum drag coefficient

%C,L Pitj~~-~-moment coefficient about

P pressure coefficient

M

.>

... . ..::

...

.——

.-.

---—

=—.:.., --.—.

.,.

quarter-chord

.-

M’ critical Mach number..

cr

R Reynolds number .,.-.... ‘:”:-..+—.

a angle of attack, degrees .,..’ .. .>..=...’.,..

. .. . .. -

DEVELOPMENT OF AI~OIL SERIES “ .,. ,- ......-’i::--,...

The aerodynamic character~s’tics of any airfoil ”ar-e~in general, dependent upon the air~oil ctiber -lineand -the thickness form. Mean ctiber lines w=re derived. ~analytically to obtain a uniform ch.ordwise distributionof induced velocity or pressure for certain designated .lift coefficients, an~ an analytical search for a thick-ness form that likewise has low and uniform chordw~seiindn~ed-velocity dist~ibution was then undertaken. .

Derivation of the camber line.- Glauert tr&fe$enG&L)has derived expressions for the local induced velofiityat a point on an ai.rfoi.l(zero thickness assumed) in

....-.—... . -

terms of’the circul~tion .aidund.an ai~foil correspondingto a certain .distribution of vorticity along the .aIrfoilsurface. .Eyassuming the distrlbuti-on of vorti.city to beconstant, a line airfoil is determined ‘that givas uniformchordwise p“re”ssuredistribution. TEs form oftton go derived is

.,. -.

where T. :s the ordinate of t.h~mean cm~cr

(1). ._ ,.=---

tha abs~~ssa, and the ckcrd is taken as unity. Theidealized form described by this equation has discontl-nui.ties at tke nose and.at the tail:, .ThiS difI’iculty ~s.circumvented by assuming very slight gradients in theChordwise load distribution just at the.~~se and just atthe tail. This form, derived by wshg the Fourier seriesmethod, is given by the equatl~n

.-,...’-. ;.J~..--:. .=.=-..

xwher~> - = *(1 - 00s e) and c is the airfo~~ chord.c

Equation (2) expresses thQ mean cmbgr ‘line choi&n forairfoils of the series developed. Load or i.nduced-veloci.tygrad@s derived from both equations (1) and (2)are actually identical for all practical .pilrposes. Mean-camber-line ordimates are given in table I Ior CL = 1.0. -

In.order to.ob&aln the rfleancamber,line.giving uniformcb.ordwlile,dis.trib.uti.onof. inc?udedvelocl$~ for,a~her

,-

ve.luesof tij+e:,l.l$tcoefficient; the”values gl$~n;intable I are r?m~tfplietf,b~the value of the desirbd.lift ~ ....coefl~ctent. .’. -.,.

NACA TN NO. 976

approached the deskti~d shape but further modificationswere investigated analytically and~ finally~ two shapeswere chosen for tests. These shapes, the NACA 07-009and the NACA 16-009, an~ the thpor~t~ca,} pres$:re distri--bution for each ar~ shown in fi~.e k?. .The completeairfoil profile is derived by first c.alculati~ the meancamber line for the.desired Iift,c,oef”ficient“and thenlaying out the thickness ordinates glveh in table 11 fromthe camber line along perpendiculars to this line:

-..Airfoil designation.- Because the ideal se~iqs of

.-._

airfoils requires an extremely large variation of shape,it becomes practically impossible- to use previous num-bering systems and, further, because this new series ofairfoils is designed to obtain a specific pressure.diagram,these airfoils are designated by a new series of-numbersthat is related to the flow and the operating character-istics of the airfoil. The first ntiber-is a serialnumber that describes the class of pre$sure distribution,the second number gives the-location of the i@ximun”nega-tive pressure in percent of chord from the leading edge,.the first number following the dash giveg the lift coef~ficieritfor which the airfoil was designed to.o~erate, ““..andthe last two ntlmbers give the airfoil thickness inpercent of chord. Thus the NACA16-509 airfoil has-theshape of the NACA 16-oc9 disposed about the uniform.chord-wise load camber’line designed for”-a lift c~aff’ic$ent ofO*5”

Airfoils investigated,- As previously stated, ‘two ‘-basi~ airfoils were investigated. (See fig. l? ) TheNACA 07-009 siz&Oil should, theoretically, give highercritical speed than the NACA 16-009 but afie~rlier inves-tigation (reference 3).indicated that, for pressures

# occurring near the leading edgej tiie’”tnqreas.gifithepressure coefficient as a result of compressibilityeffects was gretiter thsn that for pressures oc”ctiringfarther back on the airfoil. Consequently, it was believedthat, at speeds as high as the critical speed, the NACA07-009 airfoil ’might have, as a result.of compr~ssibility

., effects, a pressure peak near the leading edge~ ~~ “’”““-.NACA 16-009 airfoil was therefore develo.ped in afi~ttemp~.to achieve the”uniform c~ordwi$e lo:d distribution at .high speeds. Botlnforms were tested and the resultsshowed higher drag and lower critical speed for the NAC.A ::07-009 airfoil. Accordingly, the NACA 16-009 airfoilwas chosen as the basic form for a series of airfoils

, ..-...-,.,“-..

. -,.. ..,. .,,LZ-..: .:-----,t ,:,’.- :,....:G-

——

. ——

.,

----—.—--——

.C”.-=,...”

t-- -

6 NACA TN NO. 9’/6”

designed to operate. at various lift cae.$ficients.:.~.or,+.one value of the li.~tcoefficient the effect o.f..thic,knessvariation was also ,investigated. The airf,oi$s tested,of’which prol’iles are shown in figure 2; are as follo”ws:6

NACA 16-009NACA 16-109HACA 16-2o9NACA 16-509NACA 16-709NACA 16-1009NACA 07-009

APPtiATUS

NA~~ l&@ ‘-“NACA 16-5w,NACA 16+15NACA 16-521NACA 16-530NACA 16-106NACA 07-509

..-----

AND METHOD

The tests were conducted in the ~-&gley 24-inchhigh-speed t~el, in which velocities.a~proachlng thespeed of sound can be obtained. A brief descr$~tion ofthis tunnel is given in reference 3. The balance meas-ures lift, drag, and pitching noment snd, except rorimprovements .thet permit a more accurate determinationOF the forces, is simil&r in principle to the balanceused in the Langley n-inch high-speed tunnel. Themethods of operation are likewise similar to thoseemployed in the operation of-the Langley Ii-inch high-speed tunnel (reference 7).

The models were of ~-inch chord and 30-inch spanand were medd gf*c@alumin. A complete descrfptiion orthe method of constructing the models is given in ref-erence 8. T&I model mounting ia similar to that usedin ttLe.LangIey n-inch high-speed tunnel (rqference 7).The model ext+nds actiossthe tunnel aridthrough holes,which are of the same shape as but slightly larger thanthe model, cut in flQxible brass end plqtqs .~hat preservethe c~ntour of t-hetunnel walls. The model ends aresecured in th& balance, w~ich,extends halfway around thetest section and is enclosed ~n the airtight tunnelchamber @Qlar to the insta-llation in the Langley n-inchhigh-speed tunnel (reference 7). .

The speed range .oyerwhich ”rne~.su.r~mentswere madeextended, in general, from 25 percent of the speed ofsound to values in excess of the critical speed, Thecorresponding Reynolds number range wqsj,&orn approxi-mately 700,0’00 t’onearly 2,000$000. The lift-coefficient

NACA TN No. 976 ,... .. .t 7

range for wh~ch tests were made eTt@nded from zero liftfor each airfoil to values approaching maximum lift.

.

= - -__= -.,.,.PR3CI”SION, ,’-.,:.”,,’.-;. . -1.,

.,,,. . .... ..

. Acciden”tai‘e.rror.sare indicated by the scatter onthe plots showing the measured test data (rigs. 3 and23 tO 25). These errors are, in general, rather Snlalland affect neither the application nor the comparison ofthe data. Tunnel effects arising from end lea~age,”restrictions and the usual type of’tunriel-wall’effectare important. F~act k~owledge 0~ these various effectsis incomplete at the present tim~:. T& ‘laYgGst‘effectsappear to arise from air leakage through the c-learancebetween the model ~nd the brass end plateS in the tunnelwall through which the model passes. Investigations of

-

. -. .-.

—.——

. .

.

the leakag; effects have been made “for the NACA 00~2 air- .foil with a special .type of’internal gap or clearance “ --that pem.its wide varistion of the gap. .Data-obtainedwith various gap Settings of 0.01 inch and larger ‘extra-polate:d to zero gap were used to evaluate the-leakage --- - -:----correction for the standard type Of no~ttng=. ‘These

..

corrected data -ware then checked by means of.wake-survey - ~drag measurements witinend le~age ~l~i~t~~ b~%bbe~- .

.-.

seals. Because the balance chamber is ~fi.rtt$ht,the. ‘.:end-leakage condition is related to,the pres~ur~ ‘distri- .bution around the’model. It was thefefo”re“coti$>de{red .-v.advisable to check the method of correction for end”leakage by-wake-survey tests with end leaka@” eltii~t>@ ”’”” .by rubber seals for these new @.rfoils, which.have ~~f= ,:..tally different pressure distributions from the elder “airfoils such as the NACA 0012. Some of thes,e*ta. are ‘

.

shown in figure 3. In general,- the agreement”is excel-lent. The data have accordingly been c~~e”cted for e~d-leakage effects. ,,- ..,’” .,.-. .

..

Other tunnel ef~ects have not been-completely inves-~” “‘ tigated and-the data have not been corrected for such

effects as.restriction or the more- usual type of walleffect: As presented, the data are therefore conserva-”tive, .lhasmuch as investigations made t%us far” indicat~that the coefficients are high and the critical speeds” ‘ “ ‘“may be low, Strictly comparable data for two oldey air-

I foils,: the 3c8 and the NACA 2J09A3L, forT~tiio-Mach-~bersare included so thati~comparis~n~ can b~”made. ‘-”’”’-I-- .::j:

,. -..,, ..;.... !?. ..:.. ..-.

—

.—

El

DTSCUS310Y r .,”,. ., .. :—

&erodyn&.ic “charecteristic$ for several of’the N4CA16-series a~rfoils are given irifigures ~ to 18, Exami-nation of’these figures’ indicates two important discrep-ancies between the theoretical design cor~ditions and tk.edata obtained from the tasts: First, none of the’airfoilsattains the design lift coe~~ici~nt +t the desi.&n angleof’ eltt&ck (Oo) and,”second, the departure incred~esma]?~edly with the design lift coef’~}cient. The departures??@.be importan~ If .v~riatlo”nfrom the ideal pressuredistribution is rapidwith change in lift coefficient.This effect, if great, would t~nd to cause lGwer drag :,ap,dhigher critical speed fir’a narrow region near the”clbsign,condition than are shown by these data. ~he~e !:

departures also increase with the airfoil thickness.

The .diff’erencesbetween the designconditions andthe actual test results.r,aybe expected because of the ““simpl’Lfyingassumptions of the thin-airfoil theory..Theclretically, i.td-sassumed thet the i,nduced?velocities&orene-gligiblysmall as compared witi> the stream v~locity.For thin airfoils at low lifts, this approximation isvalid. ,W’ithincreases of lift or thickness, however, theinduced velocities approach aridsometimes .exceeci”thestream velocity. Study of these effects. appears to-bevery “important in order to obtain the proper ai.rfoils-for high lift coefficients and large thickness ratios.Deviati.onqshown by the airfoils in this series havitighigh lift and high thickness ratiiosappear to indicatethat,the use o~--asin@e basic ~hspe Is..unwarraratedifit $S+’.desiredto obtain optimum ,alr<oilsfor a wide rangeofr”IY”~.~.~ef’ficiemt .

.,..and ‘hich4W ‘@distribution”’”

. .

,,‘IT:eoret”icalpressur~-d+str~~utio~, di~grams for th~

th$bk~r airfotls.showed much greater slope.af the p~eg-sur~’”curve tk,anis shown by the basic NACA 16-oa9 airfoil.Preliminary study indica~d that.,@cJr.easing the leadi’ng-edge rad+uq .~d,.~the“fullness of .th,eairfoil b.e~e”r” thel~adin~ edge.and..,themaximum or@@tit@ may-lead to con-slderaljiejmprovplnentiv”efi ths”thicka,r airfoils hereinrepo-rte-do ::.

,... ,., .----., .—

.. . . . . .-. -”.comparis.~fi.ofaii+f’’oils~-~l@res 19.and 20<i.llustrate

the.dfi~erences in ~erodyri~~c ”char.acteristicsbetweenolder propeller-blede sect~ons “and the NACA 16-series -.

.—

NACA TN No=’976 9

airfoils. At lower speeds (~ = 0..45;.fig. 19) the5c8 airfoil appears to attain a much higher maximw liftcoefficient than the new airfoils.” This result ‘is ifipoF-tant in that the wider u:sefizlangle-of’-attack rtige nisyfrequently be required to prevent stalling af a propell~~during take-off. Over the n“ormslfli’@htrange,.however, “-and in most cases for which rational choice of sectioncan be mlade, the lower drag-of th”e”newsec_t&Q.nsoffersconsiderable opportunity to achieve higher efficiencies.The LOW drag attained by the NACA 2409-34 airfoil developedfrom earlier tests in the Langley n-inch high-speed tu”n-nel may appear surprising. Actually the type of flow forthis airfoil approaches the flow that might be expectedfor the NACA 16-3o9 airfoil. -----. —.

The low drag common to most. of the NAC~ 16-seriesairfoils is associated with more extensive regions oflaminar ‘flow in the boundary layer resulting from therearward position of the point of msximum negative pres-sure. Unfortunately, however} the Reynolds number is solow that effects of laminar separation may appear andsome pressure drag might occur. The small differencesin drag between the.envelope polar for the new atrfoilsand for the NACA 2409-34 airfoil are prabably a resultof this phenomenon. Actually the point of_maximum nega- ‘tive pressure for the new airfoils is considerablyfarther baqk than the corresponding paint for the NACA2)+09-34.airfoil but, if laminar separation occurs early,nearly equal drag coefficients might be expec-ted.

. . . . ..-~+ :--.+. ----At high speeds (M = ‘“7’5sfig- 20), the pegion”~--- .’-

which the NACA ~6-series airfoils were designed, thesuperiority of the new airfoils is clear-.”__TF&’eA21ieronset of the compressibility effects .fgr the older air-foils leads to early drag increases and lowered maximum

, lift coefficients. At speeds above M = 0.75 the use ofthe blder sections appears unwarranted for any purpose.

... - -~..___Crtt.ical speed.- The variat~on=of tie “critical’” ‘“--

speed with lift coefficient and with tlnickness is givenin figures 21 and 22, respegtivsly. -These curves indicatethat critical speeds exceeding the theoretical values - ‘“”were attained in the tests. In the choice of the test ‘ ‘“critical speeds, the values were selected on the basis ‘“ ‘~of earlier experience that indicated some rise in drag ‘“’;before large Flow disturbances occurred. If these sp6edswere chosen as tk.e highest values reached before any - ‘:’”appreciable ”drag:inerement occurred, the agreement with ‘ : -

. .-.._ _

.

.-

.-

.,

- -_‘“”- “–——

10 NACA TN Nd ; 976

,,tke theoretical .curves,”~,ouldbe $e’ry.gtio~”.FcJroornpenlsonthe critical speed of’the 5c8 airfoil is p’lotted.im fig-ure 21. The.dit’feuen.cebetween.the,new and the olderairfoils 1S greater than shown by the curves bec”ause the3C{3is 8 percent thick, or 1 percent” of’’thechord thinnerthan the airfoils of’the NACA 15-009 series. #*..:L. ,,

.

Minimum “.drag.-Coefficients of minimum drag plottedag&linStReynolds number are given “inf’i&ures,23 and Z’&The lowest drag coefficient was obtained for the NACA16-106 airfoil; this coefficient tk appr~ximately 0,0026at low speeds and increases tcrapproximatel~r 0.0032immediately below the critical speed. Cf the”~-percent-thizk airfoil series designed to operate at various liftcoe:~ficient~ ~he.NACA 16-109 airfoil appears to havethe lowest drag. This result is contrary to expectationbecause the syriimetricalor basic form of the.NACA 16-U09would normally have the lowest minhNJ.M drag aoeffici.%nt.The difference may be due to some irregularity of theairfoil surface. .

The comparison of the.minimum dra%

coefficientsfor the 3c8 and the NACA 2409-3.4 and”l -?09 airfoils Isshown in fi~re 25. The high critical speed,f’or the NACA16-2o9 airfoil is apparent,. The comparison as given.di.rectl.yb~ figure”25 is a little misleading because ofthe mmller thickness ratio for tke 3C8 airfoil. Forequal.thickness rstios, the differences between theC-series and the NM2A 16-series airfoils will be greaterthan shown. ,

,USeof the “data.- The envelope polars that hay bedrawn for the NACA 16.series airfoila represent a “newand much lower dtia.gB“Swell as higher c“r~tical speed ‘-attainable for the design of propeller-blade sect~ons.Ev&n though the angle-of-attack range is less than for,,the older sections, there will be numerous designs furwhich sufficient angle-af-.a’ttackrange is given by thefiewsections. For high-speed, hiah-a~titude aircraft,the advantagesof the low drag an”dhigh critical speedare of paramount importance and, in the~e designs, rationalchoice or section is of increasi~”i”mportance .

,.In many

designs the diameter is fixed by considerations otherthan pl?opeller eff.icl.ency. Thus the induced losses a“refixed and propell~rs of hi@est efficiency can be developedonly by operating and designing the blade secticin”stooperate on the envelope polfir,s. Another import~nt con-sideration in using new blad6’sections to achieve highest

.,NACA TN NO- 9.76 11

efficiency concerns the adaptation of the sections toolder propeller designs. Optimum efficiency csnnot beachieved by simply substituting the new sections for theold on a given design”. The use of better blade. s&Ctionspermits the use of larger diameter and necessitates some”plan-form changes. ,..A11these factors should be consideredin a desi&n for best e’~ficiency with the new blade-&4ctions...,. ..

CONCLUSION.- .

By a new approach to airfoil design based uponfindings of fundamental flow studies, a new series ofairfoils, the NACA 16 series, have been developed whichhave increased critical Mach number and at low speeds “-

.-

—

.. --~

.-

reduced dr~g. ---—.--:;---~~.. -..._. ...=

a

.—>. —

Langley Memorial Aeronautical Laboratory..—;

National Advisory Committee for AeronauticsLangley Field, vs., June 4, 1959 .-

.—

—..

.--—

._.—

12 “ ~ACA TN NO. “976 ‘“’”

-.,.

1. Stack, Jokn.: The.C”ompressibility Burble. NACA TNNO* 5W, 1935. “ ._.._ ._...-——...,’— — -..

2.-Jacobs, Eastman N. : Methods Enployed in America forthe Experimental Investigation of Aerodynamic

.--,

Phenomena at High Speeds. NACA Misc. Peper No. G, ......->1956. ...—.

39 St~~k, John, Lindsey, W.j“ “* “

“F., and Littell, Robert E.::?‘T& Compressibility Burble and the Ef’i’ectof Com--pressibility on Pressures and ~orces Acting on anAirfoil. NACA Rep. “No. @6, 1938.

~, Glauert, “H.: The Elements of Aerofoil and AiPscrewTheory. C~mbridge Univ. Press, 1926, PP. 87-93.

~. Theodorsenj Theodore: Theory of ‘LingSections of’ArbitPary Shape. NAC.4Rep. No. @l, 1931. -. # i, ,,

6. Stack, John, and von Doenhoff’, Albert E. :.>

Tests of16 Related Airfoils at High Speeds.NO. .@2, 1934.

X4CA 2(3p.

7. Stack, John: The N.A.C.A. Hi@.-Speed Wind Tunnel andTests of’six Propeller Sections. N/lCARGp.1933.

No. 463,__—.—-—..- —.——.-.--—.------- ---

80 Jacobsj JJa~tm~~q~t, ~{~~~, Ken*~~h ~.l”” ~n~ ~~~erton,itobert M. : The Characteristics of 78 RelatedAirfoil Sect”lD5nsf’romTests ~n the Varf&ble-DensityWind Tunnel. NACA Rep. No. l&O, 1933.

*-

NACA TN No- 976

TABLE I

.

. .

CNmR-LINE ORDINATES FOR NACA 16- AND 07-SZRIES

AIRFOILS WITH CL = Ls~.

cAll values measured in percent chord

Jfrom chord line

.

,

.

.

I Station I Ordinate I Slope

o1.252.5fi

/ 5.~67.50

,10.00.15.00.20:00,25.00,~().oo.40-0050.0060-60

i!(j.06“(j.m,ya.oo,95-OhLW).60

0.. I..-

3

.

.

●

.

.●

o“---

-.- ““9

.-

13

____ . ...-

*.tJaz3y

. .

-.

NATIONAL ADVISC!WComITTm FOR AERONAUTICS - “-”- ““”: -’”vI

.

A

.

. . _. _. _.

.

—

!-—=- ..-

NACATN No* 976

.

.

.d

;

. .4

..:

.

. .

.

“P

*

THICKNESS ORDINATES FOR AIRFOILS WITH

THICKNESS ~P+CENT OF CHC@l

EAll~values measured in peycent chord fromi and perpendicular to camber lin~

Ordinates

Station NACA 16-series NA::r::i:09.:- ,..--.>“ alrfcils

Slope of radiusthrou&h ,y&i Of chord =,. so. [.. ...-

L.E. radius ofNACA 16-series airfci~s = o*396(t/o.og)2.“. :.,

1--, 7

hor other thicknesses (t., in perpent)multiply ordinates for NACA 1~-series air-foils by t.{o.og.:

.,- -- ..- ....— ___!

-:

---— -=

.

- -.-- -

NATIONAL ADVISORYk. .

COMMITTZZ FOR A&lON/!!UTICS “““’vA. . .

*---—%_ ..—— -:--

.●

,

,-.

●

- NACA TN No. 976 Figs. 1,3

< ---’<:-NA CA 07-009

.

NA CA 16-009

\/.0””

- ,,

I.io

Figure l*. - Basic airfoils and theoretical pressuredistributions.

NATIONAL ADVISORY 00MMITTEEFOR AERONAUTICS

Figur

A.fuchnumbw,M L .-_.

.

.

--—

e 3.- Comparison of minimum drag obtained byvarious methods.

.4:..-—.:,

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.

. -.<.:s <.:

.:. . . . ..

—

:--

.-.

.

-.

NACA TN No. 976 Fig. 2

NATIONALADVISORY 00MMITTEE3’ORAERONAUTICS

N A C A /?409-34

NA CA 07-009

N A C A /6-009,

N A C A /8-/09

N A C A /6–209

NA CA 07-509

A/A C A /6–509

IVACA /6–709 .

A/A C A /6–/009

IVACA 00/2

< m

N A C A /6-/06

IV A C A /6-506’

NACA ;6-5/2 ‘“ .— -.

N A C A /6’-5/5

NA CA /6-52/ - ‘“

N A c A /6–53’Q .... -..

,

Figure 2.- Profiles for airfoils having high critical speeds.

\.-

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,Fig.4

I?ACA

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.

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8

b.

TN

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.. .--

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(a) Polar plots.

Ornsnt- ‘data”, ..

Figure 4*- Aerod~mio oharaeteristics of NACA 16-009-seriesairfOi.ls. ~ z 0.300

.

..

●

✎ ✎

-.

.*

A TN No. 976 “ Ei&i

E!“Sl

%00OHmEs.mor33

Z&~.@4a

(a) Polar plots. (b) Lift and moment

Figure 5.- Aerodynalnic ~h~acc~ristics of NAC~ 16-009-airfoils. .

-.

.

.

. . --,.,

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...

.=” .-..–-

.-

....

.,-.. ....+

-— ..---- .

.

data:

series

NACA TN NO. 976 Fig. 6

..

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.

.

.

. . . . . .

11 , ,

I II1/ /4A f

Y/ I .

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{a) Polar plots. (b) Lift and moment

Figure 5.- Aerodynalnic ~h~acc~ristics of NAC~ 16-009-airfoils. .

. . . . ___

....

.,-.. ....+

-— ..-,--.’.

.

data.:

series

I

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“.“%”?.< . . ..is~

. Angle of.o{iockao?,deg Kc,..

(a) Polar plots. (b) Lift and ‘momefitdata-.” - ‘

Figure 7.- Aerodynamic oharaoteristics of NACA 16-009-seriesairfoils. M = 0.70.

NACA

..

.*

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(a) Polar plots.. (b) Lift and motient”

FigUre 8.- Aerodynmic oharaoteristios of NACA 16-009-airfoilS. M = 0-750

.8----–

.

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series

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IIACATN No. 976 Fig. 9

. .

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Angle of af&ck,a,u’eg “– (&c,. -.

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-... .- . . . ... .

“’-.%7:.

..” +.—.

.- “- - .—. .,--- ‘.-

) Polar plots. (b) Lift ‘and moment d~ta.

Figure 9.- Aerodynamic characteristics of NACA 16-500-seriesairfoils. M= 0.30. ●

NACA TN lb. 976 Fig. 10

.

i-b

/.oH-+- 16-5218 r 1, , . ,

f 4’ /

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1.2

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Angle of a%ck, et,deg -c“+”

.. ..—

.

....—

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... .-.: : ‘.:”

.- .-. .. .. ‘-.

.-

(a) Polar plots. (b) Lift and moment data.

Figure 10.- Aerodynamic j&=raoo5:istios of NACA 16-500-seriesairfoils.

NA(YATN HO, 976 Flgo “11.08

.

&.06 .

$“,. .

$.04

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(a) Polar plots.

Figure 11. - Aerodynamicairfoils.

(b) Lift -d

characteristics of NAOAM = 0060.

./----—-

moment data..-

16-500-series

.

NACA TN No* 9’?6 wig..08

c5.06

$-.$3.040L1g)

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0:2 0 .2 .4“ .6 .8 LO~ if~Coefficient-cj

-4 0 4 -.2.-J --0 ./Angle of u}}ock,a!,deg _ .LCmc[.

. —..—.._

(a) Polar plots. (b) Lift and

Figure 12.- Aerodynamic :hfi::+::istics of l?ACAairfoils.

.

—1

.

.. . . . .,:.-

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—.

,.

-,.—..—‘--”

..—. —._. _

.- ___

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—.. . . . . .. —moment data.

16-500-series

—.

I

I?ACATN No. 976-.,,‘-,L.-.....-

I .-.

Fig. 10..- -.

,. -. . :.. ,

.— ——

————

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Angle ofoffqck, d,deg~

me/4 ,

(a) Polar plots.

Figure 13.- Aerodynamio c!haracteristiosof NACAairfoils. .M=O.75.

.,—

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—

—

.,

—. ..”. ..-

“.. .

/’— . ..—

--.—

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mo”ment data. .-

16-500-aeriei .—.

NACA TN ~Oo 976 . Fig. 14.. ... ...”--..

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Angleofafiuck,a,deg Crndd ... .........—

(a) Polar plots. (b) Lift and moment data. “--........____

Figure 14.- Aerodynamlo characteristics of NACA 16-000-series .airfoils. M= 0.30.

~AOA TN No. 976F$g. W

..-=----T–

...-..—------

.

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(a) Polar plots.

Figure 15.- Aerodynamicairfoils.

.(b) Lift and.moment data. ...r.—

characteristics of l?ACA16-000-seriesM = 0.45.

NAOA TN No. 976 Figs 16 -.—-.08

QQ.06

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.

Angle of ati%ck,a, deg cme/4

(a) Polar plots.

Figure 16.- Aerodyn=i~airfoils.

(b) Lift and

characteristics of’NACAM= 0,60-

.-

—

. ... ... ..—.. . .

.-—

.. -.=.. . ...—.

..-..-

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.—

moment data.

16-000-series

I

MACA TN No. 976 Figs-...08

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Angle of uttack,Q, deg c%/4

[a) Polar plots. (b) Lift and moment data.

Figure 17.- Aerodynamio oharaoteristics of NACA 16-000-seriesairfoils. M= 0.70. .

NACA TN No. 976 Fig. 18

.-

C.?;“ ““.u$

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8“ . . .bc1& ..

——.

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. . ..

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.

.,.

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Angle ofoflock. ~,deg -;! ,;“Cmc,4- ‘“- ‘:””. _—

(a) Polar plots. (b) Lift and moment data.

Figure 18.– Aerodynamic characteristics of NAGA 16-000-seziesairfoils. M = 0.75.

*

NACA TN NO. 976 Figs . 19,20

.

L iffcoefficientCl. -

.“.”:..:. .-

Figure 19.- Comparison of airfoils. M = 0c45c

NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS .

.-

.. ,

I

— 1I I I I I I I I I 1 1 I I

0. .2 .4 ..6 .. ...Li[/ coeflicien( CL

..——

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1

Figure 20.– Comparison of airfoils. 1!!=0.75.

--

. .

-—. ..

—

NACA TN No. 9’?6 Figs . 21,22

Lif+ coefficient, C=

.—

Figure 21.- Variation of airfoil orittoal Mach number withlift coefficient.

/.0

.8L

Airfoi/ fhickness-chord ro+io, per~enf

.- - --—

.

Figure 22.- Variation of airfoil critical Mach number withthickness-chord ratio.

. .--_—

H--tHATIOHAL AOV120RY 00HMTTEEHUR AERONAUTICS t-l-l

111111‘m’3 456 8/0 20 30 40 60%10s

Reynolds number

Figure 23. - Hinimm drag for the RACA16-009-seriesairfoils.

too

.080

.060

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.

Figure 24. -Hlnimumdrag for thel?AOA“$16-500-~eriesairfolla.

1

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Figure

# -- ——+---— :...- ., --- ---- :. :..---:

._ --- .. .----- ,----- ---.- .. ...-”’.

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Fig- 25

NATIONALADVISORY00MMITTEEFOR AERONAUTICS .

Much numbec M

a5.- Comparison of airfoil minimum

.

drag

. . ... . ::..-. .-..;

—

.—,.,

.,.

. ..—..s

..—.....” ---

.-.-.._

,.. - . . . .

coefficients.

.