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.FJATIONAL ADVISORY .—COMMITTEE FOR AERONAUTICS -

ORIGINALLY ISSUED

September19k4‘“”bAdvanoeCotiidential

IIWE-431?EEDlWEYI!IWI’IONOF

J-=@w

.,, .

By NoxmenF.

ReportL4118

ImL’mAG M13?Glmlms

MemorialAeronautical~ey FieM, Va.

Idoratory

,. ,’, ,:“

INUKA ‘“WASHINGTON

. .,. .

NACA WARTIME REPORTS arereprintsofpapersori~inallvissuedtoprovideraDiddistributionofadvanceresearchresultstoanauth&izedgr&~ requirkgti-emforthewar effort:They werepre-viouslyheldundera securitystatusbutarenow unclassified.Some ofthesereportswere nottech-nicallyedited.Allhavebeen reproducedwithoutchangeinordertoexpeditegeneraldistribution.

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ITACAACR No, I&Y18

‘- ‘~ONAL ADVISORY COMMITTEE FOR AERONAUTICS

lllll~ll~lgyg~y]llllj —ADVANCE ”CONFIDENT~L REPORT—— --_——.—

H16H-SPEED INVESTIGATION OF lX?W-J3RAGWING INLETfj..,.

ByNorman F. Smith,.

SUNMARY

Tests of a low-drag wing-inlet section, which hadbeen developed in a previous low-speed investigation atLMAL, were conducted at high speeds in the NACA &foothigh-speed tunnel; Near the design angle of attack, theinlet section was found to have minimum profile dragcomparable to that of a similar low-drag plain airfoilsection and to have negligible inlet losses. These resultscorroborate those obtained in tlfielow-speed developmentprogram. The Inlet section was found to have a highercritical Mach number t%an a comparable airfoil section,as predicted in tb.eprevious low-speed tests of this inletsection. A gain in critical Mach number of about 0.02was ,measured, which is approximately one-half the gainindicated by the previous low-speed data and the dataobtained at low Mach numbers in the present investiga-tion. IToinordinate changes in section characteristicswith Mach number were found. In general, the variationswere quite similar to those variations found for thecomparable plain airfoil section tested simultaneously.

Satisfactory section characteristics could beobtained only for a small range of angle of attack andinlet-velocity ratio, as a result of internal separationand external pressure peaks. The maximum lift found forthe inlet section was considerably lower than that foundfor the similar low-drag plain airfoil section. Testsof the inlet section with two amounts of camber showedthat the introduction of a moderate amount of camberimproves the section characteristics and the usefulangle-of-attack range. Furtlnerdevelopment is shown tobe necessary to produce inlet shapes havingsatisfactorycharacteristics through the desired ranges of angle of. attack and air-flow quantity.,.

.

INTRODUCTION

A prQgram was initiated by the NACA that includedtests at low speeds in the NACA two-dimensional

2 ITACAACR NO. 4118

low-turbulence tunnel to develop air-inlet shapes forlow-drag airfoils and subsequent tests at high speedsin the NACA 8-foot high-speed tunnel to determinecompressibility effects on the characteristics of themost promising shapes. A development program in theNACA two-d.irnensionalIqw-turbulence tunnel producedwin,g-i.nletsections having a minimum section profiledrag comparable to that of similar low-drag plain airfoilsections (reference 1). The low-speed pressure data alsoindicated that the critical compressibility speed mightbe higher for an a;.rfoi.lof a given thickness ratio withalr flowing into an efficient air-inlet opening than fora similar sectionof the same thickness ratio with noinlet opening.

The present investigation wasrnade in the NACA8-foot high-speed tunnel with t!meeemodels designedfrom one of’the best inlet seotions (shape 9) presentedin reference 1. .4sa Dart of the general program,shape 9 was tested in {.tsoriginal symmetrical. form;in order to study the general application of the inletshape to cambered sections, two additional camberedmodels were designed and tested.

Section characteristics of the wing-inlet modelswere determined from pressure distributions, internal-flow measurements, and wake-survey measurements. Forpurposes. of comparison, the corresponding characteristicsof a comparable low-drag plain airfoil section weresimilarly determined.

SYMBOLS

a.

M.

Mcr

ao

Po

qo

speed of sound i.nfree-stream air, feet per second

free-stream Mach number (Vo/ao )

critical Mach number

section angle of attack, degrees

free-stream density, slugs per cubic foot

free-stream dynamic pressure, pounds per square

()

1foot ~Povo2

AH ...

AH/qo”

P

P*

P

Pc??

c,

d“’o

Cd. 3

dint

cdint

n...

,Cn,,.

total-pressure .lc~s.from,free s.ttie.am.to internalSim-ve~ ‘rake shown in”figure 1,,p“cundsper squarefoot

totii-pres szire-loss‘c-heffiiient”, ‘‘.’.,,’.

,,

local static pressure , pounds per square foot

free-stream static pressure, pounds per squarerOOt

(); PoD-pressure coefficient

q.critical pressure coefficient, corresponding to a

Mach number of 1.0 at that point .,

wing chord, feet

section profile drag, pounds per foot of spando

section profile-drag coef$ici.ent();-

internal drag, pounds per foot of span

()

‘intinternal drag coefficient -—qoc

section normal force, Pounds Per fCJOt Of Span.,’

section normal-force coefficient()+

.rn, -,section pitching moment ab’outquarter-chord .point,.,: foot-pounds per foot of span

{)section pitching-moment coefficient’ ~c%~q: , qoc2” .’.

R ;, Reynold?. number ., . ., ... .. .,,.,.

V. free-~tream velocity, 1%.et:Per.s,econ”d ,

VI ,velo,cit.yat air,inlet, feet p~r secqd.

yl/Vo’,:’inlet-:ve,locit~ra~,iO .. i ~~~1 .,

“area of!ai”r,inlet.,“square:feet ‘per““+dot,of”i“~i~,?.

,, .,.

‘4area”df a:$”’tiutlet,squ.aie ‘feet ~er-foQt”””ol?s“ipan

(station number designations follow those usedin reference ~)

I

4

x

3’9

CL

NACA ACR No, 4118

distance along chord from leading edge of airfoil(see fig, 1)

yd distance perpendicular to chord for respectivesurfaces (see fig. 1)

section lift coefficient

APPARATUS AND TESTS

Apparatus

The NACA 8-foot high-speed tunnel , !n.which thisinvestigation was conducted, is a closed-throat,circ~~lar.~-ecti~n,single-return tunnel with airspeedcontinuously controllable from about 75 to 600 milesper hour. The turbulence of the airstream. is low butis somewhat higher tb.anthe turbulence of free air.

Models

The three models testeclwere or 2-foot chord andwere constructed of wood. The general layout of themodels and a scale drawing of each section tested arepresented in figure 1. The inlet section of each modelextended over one-fourth of the span and was fa.ired,in an “end-closure !!length ap.~rox~natel.~ 2.75 tinl~s the

“basi,c airfoil section”inlet height, into a low--,dragthat made up the rest of the span (f’i&s.1 and 2). Generaland detail views of one of’the wing-inlet models arepresented in ?igures 2 and ~, respectively. It shouldbe noted that the basic airfoil section matches the inletsection only in maximum th:chmss, maximum-thicknesslocation, and camber, and is representative o.fmanysections that might be used in conjunction with theparticular inlet section. The model ordinates are givenin percent chord in table I.

The inlet section of the symmetrical wing isexternally an exact reproduction of shape 9 of refer.ence 1 (reduced to 2-foot chord). The inlet height isap,proxiinately32.5 percent of the imaximum thickness ofthe section. The basic airfoil section is the NACA66(218 )-018.9 airfoj.1 section (reference 3 ).

NACA “ACR ~TO. L!LT18 5

For the medium-camber wing, shape 9 of reference 1was fitted to an NACA 61j,3-018 airfoil section andcambered. The procedure was as fo~lows:

,, ...... .

(a) The le~~th of the origin~l ‘Inlet”saction ‘aheadof the maximum-thickness station (0.,45c)was reduced bythe ratio of .40,~5Jto make the location of maximumthickness coincide with that of the basic airfoil sec-tion (0,4.OC).

(b) The thickness was reduced to that of the basicairfoil sect%on by subtracting 0..44percent chord fromthe inlet-section ordinates.

(c) An arbitrary fairing to the exit was begun atthe 65-percent-chord station of the airfoil.

(d] The fairness of the resulting inlet section waschecked by computing the slope of the surface betweenconsecutive ordinates and modifying the ordinates wherenecessary to make the variation of slope along the chordsmooth.

(e) The final ordinates were combined with a camberline of.design c~ = 0.30, mean line a = 0.6 (refer-ence 3).

The inlet height Is approximately 29.5 percent of themaximum thickness of the section. The basic airfoilsectj.onis the NACA 65,3-31[3airfoil section.

In the high-camber wing, the inlet and the basicairfoil sections have the same thickness distributionas the symmetrical wing but are cambered to Cz = 0.50,mean line a = 0.6.

The three inlet sections very closely represent afamily of’cambered sections, and the test data can beanalyzed to establish the general effects of camber onthe inlet section. The duct for the three models hadthe same llthicknessdistribution’l and was designed togive a low value of internal 10ss. The camber line ofeach inlet section was applied to the duct for thatmode 1. (See fig. 1.) No simulated internal resistancewas employed, because resistance serves merely to reducethe inlet-velocity ratio that can be obtained with agiven exit area and does not appreciably affect theexternal conditions over the section or the internal

6 NACA ACR Noe ti+I18

conditions at the ‘entrance. Removable exit plates weremodified as shown in figure 1 to give the desired exitareas and inlet-velocity ratios. Streamline steelspacers were installed,both in entrances and exits toprovide additional strength. Duning the tests nomodifications to the sections were made except forseveral small changes in the internal-lip shape, whichwas designed to reduce entrance losses. (See fig. 4. )

Measurements

Each inlet section was equipped with surface static-pressure orifices to measure external pressure distribu-tions and with an internal survey rake to determineinternal-flow cond.iti.ons. (See fig. 1. ) The basic air-foil secticn of each wing was also provided with surfacestatic-pressure orifices at approximately the samedistance from the tunnel center line. The pressure tubingpassed from the model through a passage inthe wing andwas connected to a multiple-tube manometer in the testchamber.

Section normal-force and pitching-moment coeffi-cients were obtained by integration of pressure-distribution plots. Total-pressure loss and inlet-velocity ratio were computed from measui-ements obtainedwith. the internal survey rake. Section profile dragwas mess’ured by the wake-survey method behind sectionsnot influenced by the surface pressure-distributionorifices or the inlet end-closure,

For several tests, wool tufts were installed atappropriate points on the symmetrical inlet section topermit observation of flow conditions. The models weretested t’hrough the complete angle-of-attack range fromapproximately -4.0to an “angle higher than the angle formaximumlift at low speeds. The angle-of-attack rangeat higher speeds was reduced because of structurallimitations of the wing. The inlet-velocity ratio ofeach inlet section was varied from approximately 0.25to 0.85. The tests were run through a range of Machnumber from 0.20 to approximately 0.70 correspondingto a range of Reynolds number from 5,000,000to 7,700,000 (figi 5).

I

NACA .ACR~!O.~118 7

RESUIITSAND DISCUSSION

External-Flow Conditions-.. ,,,

Pressure distributions.- Pressure distributions areshow in rlGures t for the inlet section of’the threewings at a. = O? agcl V~/Vo = O.~; for coraparison, thepressure distributions over the basic tiirf’oilsectionsare also presented.

In figures 9 to 14 the pressures ove~ the inletsections are shown to be very sensitive to changesin a. and V1/Vo. The upper surfaces of the camberedinlet sections are less sensitive tb.anthe s~ymmetricalinlet section, because the introduction of camberresults in a favorable increase in curvature of theupper lip. The decrenso in curvature of’tilel~wer lipresults in a pressure peak on this surface at the designansle of attack, ~ternal-flow cond~tions impro-rewithincrease in inlet-velocity ratio.

The result~ show that only a small range of’angleof attack and inlet-velocity ratio exists wherein. afavoi-able pressure di~tribution can be maintained everthe inlet sectioris. When well established, the pres-sure peaks prodnced outside this small range willresult in ~r~clus~Lon of larninar flow cm one surface andin reduction of critical speed.

cl’itf.Ci21Mach numberc- Critical Mach number McrIS defined as the free-stream Mach number at which alocal Mach number of 1.0 is attained at some point onthe section. Figure lb shows the variatioilof peakpressures with Mach number for the symmetrical wing atvarious angles of attack. It is apparent that thevariations for the inlet section do not follow thenormally assumed variation. Prediction of’criticalspeeds by the usual methods from low-speed data ofthis kind would be greatly in error, because of thehigh peak pressures j.nvolved. Previous experienceIndicates that a very steep pressure gradient cancause the .f’orrnationof a local separation bubble whicheffectively changes the shape of the body and lowersthe pressure peak as 1;10 is increased. A separationbubble of this ktnd is evidently nroducedon the upperlip of’the symmetrical inlet sect~on at moderate angle’sof attack. Because the bubble is quite “small, no large

8 NACA ACR No. 411!3

increase in drag results. At the higher angles of attack,however, serious separation takes place over the lips.

,,

Figures 16 and 17 show the critical speeds at variousangles of’attack and inlet-velocity ratios for the inletand.basic airfoil sections of the symmetrical.wing. Thecritical speed of the inlet section’is sli@tly higherthan that of the basic airfoil s,~ctionnear the designangle of attack beta’u.seof the effective reduction inthickness caused hy the passage of air through the sec-tion. Gains in critical Mach numbers of approximately0.02 over those obtained for the basic airfoil sectionswere found for the symmetrical inlet section, Thesegains are somewhat less than the values of 0.035to O.O&O”obtained from extrapolating the results of thelow-speed tests of reference 1 and the low-speed datafrom the present investigation. When the angle of attackis reached ‘~eyond which the critical speed is governedby the pressure peaks on the inlet lips, the criticalspeed oi’the inlet section is reduced below that of thebasic airfoil section. The extent of’tke an~le-of-attackrange for high critical speed is a func;~ion of inlet-velocity ratio.

A similar gain in c~it!.calMach number was foundfor the medium-camber Irlletsection. A comparison offigures 18 and 1.9with figures 16 and 17 indicates thata moderate amount of camber improves the c.riti.cal.-speedcharacteristics in the rarlgebeyoml the design a~~gleofattack without agnreciably affecting conditions at thedesign angle. Figures 20 and 21 show, however, that alarge amount of camber changes the shape of the lowerinlet lip to such an extent that the critical speed nearthe design angle of attack is sertously reduced.

Internal--FlowConditions

The variation of vl/vo with angle of attack forthe three inlet sections witln various exit areas areshown in figures 22 to 24:. The inlet-velocity ratio,in most cases, decreased slightly as ‘theangl-eof’attackwas changed from the design ar.gleof’attack, mainlybecause of internal-flow separation at the entrance.

Total-;oressure-loss-coefficient data for the inletsection of the s-ymmetrical wing is presented in.figure 25as a function of angle of attack. The angle-of’-attack

NACA AtiR]TO.L!I_I18 9

range for low total-pressure 10ss was””smal”l,particularlyfor high inlet-velocity ratio-s. Tuft tests showed that,separation occurred on.,the’ins~de of.the inlet lower lip‘at ao = 20,. ,Vljvo =0.5~;-’this separation.,led to,animmediate sliargrise in @q ~ as the angle of’attack

was:-increased,further. ,Because of increased entrancelosses, AH/q. rises with increase I.ilinlet-ve~o.cft~

ratio. Simil~r lnt~rnal-flow cfiar’acteristicswere foundfor the two cambered”inlet sections’ (figs. 26 and 27),

Chang:s in irile.t-velocityratio a@” total-pressure-10SS coefficient with Mach,”number were si~all. The angleof attack at which internal separation be@n did,n,otchange.appreciably with”Mach number.

,,A tuft test wai.,made tO invas.tfgat~ the flOW in the .

vicinity of’the inlet end-closure” of the sym.meti-icalwing,The test showed that, as tkleangle of attack’wasincreased, both internal”and external. separation occurredfirst in the end-closure sect~on,of the inlkt. .Furtherdevelopment of inlet en”d-closure shapes “isshown to benecessary.

Section Characteristics ““

Profile drag.- The results of the profile-dragmeasurements for the three wings are given in figures 28to 30. The values of’the profile-drag coefficientspresented for the inlet sections include both internaland external drag, and the internal drag for each con-figuration is also shown. Inasmuch as the change ininternal drag with Mach number was small, data foronly one Mach number are presented.

The mfnimum VLilUf?sof’ Cd. for the inlet”and the..

basic airfoil se”ctioiisof the symmetrical,wing areapproximately equal at medium inlet-velocity ratios;At the hi@lest v@To tested,’no low-drag range exists(fig. ~8(e)). Because the internal drag for”thtk condf-tion is very low, the high dra~ is.believed due””primarilyto the flared exit that .was.required to.jjroduce this highinlet-velocity ratio. (See ex}t details, fig. 1;) Inas-much as flaps are usually used in connection with trailing-

‘,‘edge air exits to obtain high.f’low,rates, high drag forthis condition is usually encountered. Some reduction

10 NACA ACR No. L4.118

in this drag may reobtained through improvements indesign. At the lowest test value of vl/vo no low-

drag’range exists (fig. 28(a)). Because ’of.the high.local angle of attack of both inlet lips, Pressure,peaksoccur and preclude the existence of laminar flowover both external surfaces. The drag coefficient ata. = 1° is less than at a. = Oo because the pressurepeak on the lower surface has disappeared and somelaminar flow exists on that surface.

The low-drag range is smaller for the Inlet sectionthan for the basic airfoil section, The extent of.thelow-drag ran.gedecrea:es as V1/Vo decreases becauseexternal-flow conditions become more critical at lowvalues Qf lJ~/Vo. The drag of tlieinlet section beyondthe low-drag range .(above approximately 4°) increasesat a much greater rate with angle of attack at lowinlet-velocity ratios than does the drag of the ‘basicairfoil section. Examination o.fthe internal-drag datashows that this steeper slope is due principally tounfavorable external-flow conditions.

The medium-camber wing shows the same general dragcharacteristics found for the symmetrical wing, exceptthat the low-drag range is somewhat greater and,thecenter of the range is shifted in the positive angle-of-attack direction. The shape of the upper lip has beenimproved “oycambering and the lower lip has been impairedonly slightly.

~Or tjhe h_i,gh-camber wing, the minimum profile-drag

coefficient of the inlet section at all values of v~/vo

is higher than that of the basic airfoil section. The

drag ~’ise is rapid at angles of’ a~taclt below the angle for

minimum drag, largely because of separation over thelower lip of’the inlet. At positive angles of attack,separation over the upper surface causes a rapid risein.,drag. From observations of the wake profile and

pressure distributions, this separation was found to

OCCUZW back cf the maximum-thickness station. Theprofi,ledrag could not be accurately measured because.of the extreme width and rapid fluctuations of the wingwake . At high Mach numbers, the separation becamesevere ever both the inlet and. the basic airfoil sQc-tions at all angles o.f attack. The camber for cl = o.~o,used with a thickness ratio,of 18.9 percent, apparentlyresults in sections with serious flow-separation tendencbs.

NACA ACR NO. ~118 “11

The minimum values of Cdo and the width of the

low-drag range for the batiic a~r~oil sections of the“ three wings tiestedare in agreement with values obtainedf“rom”ttio~dimensionalteistsof similar sections (refer-ence 3).

Pitching moment.- The variations of sectionp“itch~ng-moment coefficient”with Mach number are”of thesame order for the inlet and the basic airfoil sections

““.of the symmetrical wing (fig. 31). At the low inlet-veloci.tyratio a larger variation of pitching momentwith angle of attack occurs, ,because of alteration ofthe”chordwise lift loading by pressure peaks on the inletl“ips.

The same general trends are indicated for themedium- and high-camber wings (figs. 52 and 33). Largerchanges in pitching-moment coefficient with angle ofattack and.Mach nuiiber toolkplace with the high-camberwf.ng,probably because of the separation effects pre-viously noted.

Although not conclusive, these tests indicate thatthe addition ofa properly designed air inlet to a low-drag airfoil section nesd’not appreciably change thepitching-moment characteristics of the original section.

Lift ● - Section normal-force coefficients are~rese= instead of section lift coefficients. Analysisshowsthat the two are approximately equal; the differ-ence is less than 5 percent at maximum lift.

Normal-force-coefficient curves for the inlet andthe basic airfoil sections of the symmetrical wing areshovn in figure,34. A considerable deficiency inmaximm lift for the inlet section is evident. Maximumlift increases with increase in inlet-velocity ratiobecause. of the improvement of external-flow conditions.Tuft tests indicated tihabearly separation over theupper inlet lip Is responsible for the low value ofmaximum lift.

The inlet sectionof the medium-camber wtng, whencompared with the ‘oasi.cairfoil ,section (fig, 55), showsonly a small loss -inmaximum lift becapse the inletupper lip has been improved by cambering. 130ththetnlet and the basic airfoil sections of the .high-camberwing show a decrease”in lift-cturve‘slope at angles of

.“

12 NACA.ACR NO. L4118

attack”greater than approximately 4° as a result ofseparation over .the rear portion of the section (fig. 36).The angle”of maximum lift was “not reached in the tests ofthis wing.

The angle of zero lift for the two cambered inletsections shifts somewhat with inlet-velocity ratio(figs. 35 and 36). ” This effect is due larg;ly to changesin exit fairing. The symmetvfcal inlet section exhibitsvery little shift in angle of zero lift because, accuratelysymmetrical exit fairings are easily produced.

The variations of normal=force coefficient with Machnumber are of the same order for the inlet section andthe corresponding basic airfoil section of each wing(figs. 37,tJo39). Data for only one inlet-velocityratio are presented~because the effect of inlet-velocityratio was very small for moderate values of lift.

Modifications

Two internal inlet-lip modifications, which weredesigned to improve entrance conditions, were tested onthe inlet section of the symmetrical wing. Modifica-tion A was an arbitrary fairing involving no change inlip radius; modification B was the same as modifica-tion A with a ~0-percent increase in lip radius (fig. 4).

. .

The results show (fig. 40) that an addition to thelower lip of a simple fail”ing such as modification Aincreases the angle-of-attack range for low total-pressure “loss in the inlet from 40 to 80. Tuft testsshowed,that the .fairing increased the angle-of-attackrange by delaying internal-flow separation off the inletlower lip. The results obtained with modification B shaw “little improvement over results obtained with the originalinlet. The larger lip radius apparently nullifies theeffect of the fairing and produces, in addition, anunfavorable effect on the external flow.

.DesignConsiderations

The angle-of-atta’ck range through which low inletlosses and low section drag are desired is approxi-mately 7°, or from high-speed attitude to climb attitude.The data indicate that the original inlet shape (shape 9of reference 1) does not have the desired range.

“NJiCAACRNO. .“~I18 13

Tests of the lnle%:se””ctidri~:o~the symmetrical wingwith m.odificatlons A and B indicate that the angle-of-attack range for low inlet losses can be easily increasedtio’’a”sa~isfac’tor”yextent’;: Notes.ts of external~,rnodifica-tionsj”assuch;were’ included hi’’thepresent investiga-tio~i. Tests of’the Ca’m.bebed-:wings’indicate.,hoyever,’that ‘tncl’qaSingthe curvatuhe of the,inlet upper lip”,:

‘-re”~ults.in a~imppovem~nt over ‘the,o.riginalsection....““Unpublished,data from wi.nd-t~fiel programs :inwhich ,,wing,’inletg w“eredeveloped for specific airplanes,,corroborate’ this finding and show ,t.hatan appreci,~b~e’gain in maximum lift can be realized by ~mproving t,he..,flow o,vert~}e,inlet upFer lip. These development programsindicate a~s”oth:atjuMcious use’’.ofIlp s.tagger,;beyondtheamouritptiodu~cbdby cambe~ ‘can improve both intern,al-llowconditions ‘and‘rriaxikinnIi”ft.~~ -.

,.,““ T~e te”stsindicate that cambering a symmetrical

~‘5nle”tsection by normal methods (reference .3) .i.sunsatisfactory ?.nthe vicinity of the inlet lips.

,,~h,e,<nle,tlower lip, because of decreased curvature“-,’duqto ’”camber”,produced an adverse pressure ~is~ribution.’”a.t”“the des~,gn angle of ,att”ack on the mediurti-and high-c’~be,rwikgs~ - ,. ,,

.. ..,.., .“t .I:method for fitting”~n:inlet s6ction to a..given

‘“~-irfoi~secti,on‘has been described. untier the’ design of

‘.t’he,he”dium-c:~bei Wing, The chara.cterlstics of the’‘i”rilets“e;tionproduced -by this method depend, o.f.coufise,upon the characteristics of the inlet sectiGn from whichthis,section is designed. In addition, the procedureme~ely ut}lizes normal” camberihg methods and.dotisnotg}”v~,r~eededs~6cfalconsideration to the inlet lips.;The fiediiim-camberifilet,-therefore;evinces the:same.:lim>,tationsfetid “for ‘the”ori&inal inlet section.e“xc’eptfOr the slight improvement due.to camber. “.AS a generalmethod’ o“fap’pl,ication,lthe procedure is Indicated by thet~$t.sto,b’esatisfactory from cor~silfle.ratio~lS”Df pressuredistribution drag; andcritical speed near the deslgriangle of attack. * ...,,’. ,,.

,-”! :.,: .“. .:.: ...:F~,>her”development is a~~a”rently.need~”dto produce

“u”spful.,ef~ic~~qt’inlet shapes. .Sat~isf’actorysectiom”,.characteristics mus-tbe”available for sufficiently wideranges o.finlet-velocity ratio and angle of attack+

14’ “~TA~A~cR”N~.tiIj-8

., .,, ,.GO~lC”LU;IONS‘“.,

,’

1 Tests at high speed of three wing-inlet models.desig~~d from on? of the best wing-inlet sectionsdeveloped in”a previous investigation at low speedsshowed that the inlet section b.asininimum section dragcomparable with that .of a similar’low-drag plain airfoilsection and.~-.as~JeIigible’ inlet losses near the design6angle of attac~. A properly designed air inlet can beinstalled in a ‘low-dragwing at virtually no cost inexternal drag.

2. Cr,itical Mach numbers approximately 0,02fiigher than those of the basic airfoil sections werefound for the symmetrical and medium-camberinlets.”These values are approximately one-half the gainsindicated by extrapolation o,flow-s~eed data f’rom ‘theprevious development’ ~rogram. or fr”o~.the present investi-.gatio.no

3. The inlet section i:,quite sensitive ‘to“changeiin angle of attack. Adverse effects are produced on th~inlet lips that result in small angle-cf-attack rangesfor,low drag, high critical speed, and low entrancelosses. A considerable deficiency in in.ax”imumlift, ascompared to the maximum lift of the basic airfoil sec-tion, results from unfavorable’ flow conditions over the“inlet upper lip for the symmetrical and low-camber inletsections.

~i Introduction of a moderate amo~t of camberimpro”vesmost of the section c,har’acteristicsand theuseful ”angle-of-attack range, The improvement is dueprimarily to the increased curvature of the l~pper lip,which reduces dr delays the adverse e.f.~,ec-tsIncurred”13”:TL::e original shave . m7,.~~!e decrsas’ed c“vrvat~lre

of’ the lower lip, howe,ver, groduces adverse effectsl

indicating t.i-~atspecial metlaods must be devised for

cambering inlet sections.

5.’The variations in inlet section characteristicsdue to compressibility were, in general, quite similarto the variations foun~ for the comparable plain afrfoilsection.

6. The method devised lor fitting an inlet sectionto a basic airfoil section (used in the design of the

NACA ACR NO c I@18 15

medium-oamber wing) is shown by the data to be satis-factory, as compared with the original symmetirlcalsection, with regard to pressure distribution, criticalspeed, and””’drag.

7. Further development is apparently needed toproduce efficient inlet shapes from which satisfactorywing inlets for any desired ran’gesof inlet-velocityratio and angle of attack can be designed and adaptedto a wing section having any design camber.

Langley Memorial Aeronautical LaboratoryNational Advisor Committee for Aeronautics

TLangZey Fle d, Va,

REFERENCES

1. von Doenhoff, Albert E., and Horton, Elmer A.:Preliminary Investigation in the NACA Low-Turbulence Tunnel of Low-Drag-Airfoil SectionsSuitable for Admitting Air at the Leading Edge.NACA ACR, July 19420

2. Becker, John V., and BaaIs, Donald D.: Analysis ofHeat and Compressibility Effects in Internal FlowSyStems and High-Speed Tests of a Ram-Jet System.NACA Rep. No. 773, 1943.

3. Abbott, Ira H., von Doenhoff, Albert E., and Stivers,Louis S., Jr.: Summary of Airfoil Data. NAC.AACR ~Oo L5C05, 19450

“-’’%$s%

I Smetrical wingI\Basic airfoil[ section

bper end lowersurfaces

1

ML radius=2.1+17

mlot aootion

w

3.033

ttralgh’1lne

I5.2085.2085.2085.208;;:;:

l?:%3.583

3:%-----1.921j

Fairlng point, = 0.408, y = 3.067L.E. radlm center

:= 0.250,Y = 3.342,.E.radhu = 0.250

TABLS I.- MODEL OF(DINATES

Ordinate8 in Permnt chord

Uedlm-camber wing

saaic airfoil ooction Inlet section

rppormrfaoo IMm surface ] UPP13r surface II Lower surfaceI I I

r , r ,

x 171X IYIXIX31XIYIXI 7d

x = %%? ~%.671ralr~ngPoiat

X = 0.821,Y = 2.554L.K radius = 1.921 L.E. mlluo center L.E. radius cont.?

x = .0.296,Y=2.9@ x= 0.800,Y= 2.792L.B.rodluo = 0.250 L.B.rullu-= 0.250

1lfATIOMALADVISORY

COW XTTEE FOR AKROIIAGTIOS

TABLE I - MODEL ORDINATES - Continued

Basicairfoilaectlon Inlet aecti.on

upper surface Lower surface Upper surface Lowersurface

x Y x Y x Y yd x Y Y~

o 0 0 0 .16-1.021 ●8

!$

----- 1.02133.16

.100 1.617-----

1.292 -.512 ;.$ 1.51.2 2.833IX& 1.779 1:?;!

8L 29 -.22

d $;$1.729 3:?25 ~2. 50

2.5713

;.;; 3::~; 2.192i:4;? 1:?00 . ?

.7791.917 ;*:23

t k3. 0; ::417”

; Zoz 4~:~1 ~!g~ :;;?k

.354z

:6.829 6:9 2%

3.083 ~p.6 2.167

&3:~8~ :219 : 87 :::;

J4,321 7* 5

i

10.78 ●

● 52

9:g; 2::{;1;:;{

:$g $:% $:$ ;;:zG ::$5; {d ~:gi %!10.746 % 17

%4;1 11.342i

5.30 25.212 .592?

12.100 2:53 30.13 .69

iz :$’% & i% 4.808 Wt $@ f:$ !:% :$!

3 .692.80

13:11750::2; 184~:933 $:52;

13.117 5.75013.00

t50%; 8: 17 4:93 ::~3z

;::W 8.~33 ?35. 08

12.76 55.217 L%tz

~t!g z

5*73 5* 912.150 .267 60.425

k12*00’ 8.625 5.7

d?+5.517 *

65.58z

1:6711.lK .837 64.71 11. 5

k8. 33

70.538.708

i4.200 7;. ~ 10.533 z

J&

● 33;. 25 7 :475

}4z %@ f~ :!i~

86:~!$ k

3. 50 9.167 .833

z.571 2. 37 0: 8

t.637

85:30 4.511

.672 1 2● 333

c~1*53

1.800 85.~4:;;$

.56 : 1.62590.19

12.74

18 .821 .987 90.333 3::;; 8.67

1z.2:; 2.042

95.02 1011.2 &.937 ~9287 ~.;;; ----- 9 .787 2.175.00

-----0 1%:~~6 . 1.925 99.94 1.979 1.925I

Fairingpoint Falrlngpointx= -0.717,y=3.000 x = 1.237,y = 2.896

L.E.radius= 1.808 L.E.radiuacenter L.E.radiuscenterx= -0.771,y = 3.23? x = 1.292,Y = 3.125L.3.radiua= 0.250 L.E.radiua= 0.250

I

~ATIONAL ADVISORYCOHHITl%B F’ORAERONA17MCS

Ii

,;,,<,.<,,+c/L.5Jre-! e~-::- r- —“,.lqa/ 1,,/~//

L, mI

IL 14~-~ 42- 2A~ 4“ ~6~ “ J

Bos IC O(rfoil Inlet secf, on G5ect/6f)

BOSIC aIr/oIl section oz~ri

Model orronu~menf-

Front y,ew

k.~>.. . — -

Bosic otrfotl section B031C OIrfo! sect,on

/y ...... .. ,. ,,

?_, ____ _— ——— —

/n Ie t section

The 5~/Wi7#riC0/ u/hq

20s Ic oirfoll secfion

,//—’/‘;;>:“,;

.—— — — .— —.

///,c ,/. ,, —/

inlet section

i%e hlqh-camber wIhq

f@re

Inlet section ; ~ , , .,, ;/,,

The medium-comber wihq

IncncxDetails of exit fairinqs for fhe inlet

section of the s ymm efrical win q

mm.,,,,,Ml,Co”uI,,,,t,Ilm”,m

[- Mode/ orrunqement and section detui/s.

zo.

%

.

Figure 2.- Installation of wing-inlet model in NACA 8-foot high-speed tunnel.

zo.

co

NACA ACR No. L4118 Fig. 3a,b

(a) Air inlet.

!’, ..i-. .-} o = .-.: “

—._ —. .._. -.

(b) Air outlet.

Figure 3.- Details of inlet section of symmetrical wing.

NATJONALADVISORYCOMMIJEEFORAERONAUTICS

duct Outline

Zhk?rnaf+’”F/qure g -modl$ko’l)ons

Secf/bn of the symme 7’7ic L7/

o //2, I IAches

●

winy●

zo●

lb

/ox /06

8

o

.

‘=sr’m

●

m

o ./ .2 ‘3M. “4 “5 ●6 “7 bZJ

~igure5.- Variationof Reynoldsnumberwith free-streamWoh numberz

for 2-feet-ohordairfoiltn EACA 8-foot hlgb-sp.od tunn.1..

FlbHP03

.

NACA ACR No. L4118 Fig. 6a

-.8

-9P

oli7/e f aec from

NATIONALADVISORYCOMMlllEEFORAEI?ONAUTICS

-dP

o+I

1-

0Pexenf choPd

(a) ~ = 0.200Fl~e 6.- Preesuredistributionsover syrnmetrioaluin~

a. = OO; Inlet eectlon, vl/vo = 0.53.

,,, . ,.———.. .- , ,.. .,,,,,.,! . . . . . . ..-l.!. .

Fig. 6b

-.+

P

o

●4

NACA ACR Nc. L4118

MTIONAlADWSORYCOWllEE FORAERONAUTICS .

(b) ~ = 0.40.

Flguro6.- Oontdnud.

. . . . .

NACA ACR No. L4118 Fig. 6C

o

48

-+

P

o

o 20 @ 60 80 /00

I I I I I I I I..0 20 40 &l %0 /00

Pencent chord

(o) & = O.65.

Figure 6.- Conoludod.

Fig. 7a NACA ACR No. L4118

P

—. 8

4

0

+

.8

-.4P

o

.9

o 20 40 60P’rceof chord

80 /00

20 40R7cenf chord

(a) U. = 0.20.

~lgure7.-Premwre distrlhtions over ❑edium-oamber wing.

=0 - 0°; inlet oectloq, v~/vo = 0053.

——. -. - ...—. — -! .,. -. , ,, ,,, ,,- ,-,,, -,-,- -. ,,., I .! -1.---ml ..! ,11 ,, , , , ,

NACA ACR No. L4118 Fig. 7b

-.4P

o

.4

–23

-4-P

0

.4

0 20 & 60 &IO /mPercen+ CAOF9’

(b) ~ = 0.40.

Figure 7.- Continued.

............ .. .. . ,,, , , , , , ,,,. . ...—----- ,, ..—

Fig. 7C

-.

NACA ACR No. L4118

‘L 6

-L?

-z8

P

-.4

0

,+

‘i,6

-AZ

:8

P

-.4

0

o 20 m /00/?erci2?cho S \

L )NATIOW ADIMY

MMITTE~ FORAlpw w

–g1 ‘

bpper surfbce

/* .Lower surface

I

I \

41z

o 20 40 60 80 /00Ferrenf Chad

(0) ~ = 0.650

Figure 7.- Concluded.

NACA ACR No. L4118 Fig. 8a

P

.4

-.8

.4

NATIONALADVISORY\

A.

+

/ /

—— — +>

\, .\

z5!uwc air fiil seci%nA

(2a#6a&)/n9/+PCC# chord

(a) MO = 0.200

~iz. &- Pressuxw dlntribut ions ovor high-oamber wing.

=0 = 0°; inlet sootion,V1/Vo= 0.55.

NACA ACR No. L4118

A — NATIONALADVMRY}’

/’‘1 +

/

/ I I

/

Low St?wfi?ce

+ —y—+ ,_A-t. -\

\ l\

~J~_‘\+

OCzrjc u/kW/ Getsfim

(b) Ho = 0.40,

Figure 8.- Contlnuod.

#J

NACA ACR No. L4118 Fig. 8C

,.,

-.4P

o

+3

-.-’+P

o

Perce#t ho~d/+-%-,

t

\-f- .— ‘5.I W ML AlvLSORY+ (DMMHl@FOR#@iAu7Cs

I Zk7aic t2if#ioi/ secf/ao

027w60m/a9Rmamf &o&

(C) ~ = 0.600

Figure 8.- Conoludod.

Fig. 9a . NACA ACR No. L4118

-36

?

-32“

-28“‘

jgi) —

-2+‘ 00

R—Uppsuduce

c

b Z7?kt aecfionI/u .0.53 I

o ~epcen+:hpdJo10 0 ;:pcen$oJ30

(a) ~ = 0.20.

~i~ra 9.- Pressure diatributlona over forward portion.of aymetrlcal wing for varloua angles of attaok.

NACA ACR No. L4118 “ Fig. 9b

-32

-26

-24

-Z@

k-/.6

o

#

.8

o ~eycen+:i?opdm

(b) Ho = 0.40.

Figure 9.- Oontlnua&

Fig. 9C NACA ACR No. L4118

-3.2

do-26

(h?v –z

‘o- a “/

‘2

-20‘O

—Upper surfme1?-- -–42WW W7FQCG

4

JIn/ f Sec7+00

I./N=0.53

(o) Q = 0.65.

Figure 9.- OonolucleL

,

— I

NACA ACR No. L4118 Fig. 10a

““8mT=?l

o /0 30?+rcen~ocho~

o 30Lmn+2:ho&

(a)M. = 0.20.

Figure 10. - Prestwre dl.strilmtlons over forwe-’d portionof medium-cambek wing for varlouo angle B of attaok.

..———,—-..—, —..-..--.,-..- ..—--..—. .,—--..-—

Fig. 10b

o 30Pe%+ ‘2dd

NACA ACR No. L4118

(b) w = 0.40.

rigura 10.- COnttiu*

p-?

NACA ACR No. L4118 Fig. 10c

-8%3

-44

-40k

o mR&of$%7’

o

.+

.8

o #F?!?%%+%&d(0) ~ = 0.60.

Figt?ra 10.- Conolud@d.

Fig. lla NACA ACR No. L4118

—

‘&y *

II

(

o 30Pe?ce ntz~hord

(a) ~ =S0.20.

Figure 11.- Pressure dlotributlona over forward portionof high-camber wing for varioue angle B of att.aok.

NACA ACR No. L4118 Fig. llb

-32

vI

&

-2.8 (de;) _I “ -2II ‘e o

-2.4 I ❑ II ‘2

I 04

-2.0 b 6II UpDer surface

I——. Lower surtie

II

.1

0\

<

D

o 30Pe r;ent ~ord

(b) Ii.= 0.40. Fagure 11.- contifl~~do

11 — —

Fig. llc NACA ACR No. L4118

M ‘IOML1CwwfWMMI1EEFOR hERONILm

-gr

A

l\v ~ i

I 5’~,1 \

/

I

Busic ui-?’fo17 St?a’h?

o /0 20 30 ; /0 20 30Percenf chord Percen+ chord

(c) ~ = 0.60.

Figure 11.- Concluded.

,,, , ,,, , ,-, ,--,P I & .83

I I I.H I

—Upper wrfuce

Buaic airfoil aeotim–— upper surface

20 .30 40*rcent chrd

(a) a. = OO.

Fi~e 12.- Pressure dlstributione over forward portion ofwing for various inlet-velooity ratio0.

?%rcenf ehob-d

(b) a. = 1°.

inlet ●eotion of mymmetrioal1$.0.20.

z0.

rlb

PCD

l-lkP

I

-/.0

P

-.8

-,6

-.4 LK- 1 [ I I I I I I I

b /’0 Ao ;0 4’0P8rcen P chord

(c) a. = 2°.

-L6

-L4

P

-L2

-Lo

-.8

NATIONU AOkW~

WMMITIEE FOR ~

t

—

+

+ 0.23x .33

r A3

v .s3L ,8S

Inlet secfion—— Upper wrface

Basic w“rfoil mction

1 P 1 1 1 I

-.6

(d) a. = 4°.

nwface

zo.

Figure 12.- Conoluded.

------:>_

-.8

-.6

F

-,4

-.2

0

-,8

-.6

P

-.4

- ,&

o—

Atrcenf chord

(a) a. = OO.

Figure 13.- PreBsure dletributiona over forward portion ofwing rOr various Wet-reloclty ratios.

llATIONAL-

caMMmf.E Fw.~~~

o.

r+Huin

A9rcenf ohdrd

(b) a. = 1°.

inlet section of medl~mberM. = 0.200

PCNP

-.8

-.6

P

“.4

“ .2f

- —UPPaP .~urfuoe

I 10 /0 20 so 40

Perced chord

(c) a. = 2°.

= NAliwlmsmCOMMllTEEFORMONUM6

(d) a. = 4°.zoFigure 13.- Concluded.

r’IPl-lwco

H llmn!L-

CowrlEEml AEnwlJm

-,2

0

J%rcenf chord

(d a. = OO.

lUguro 14.- Pressure difitributlonewing for Tarlous

orer forward portion Orinlet-~3100ity reties.

?%reen? chord

(b) a. = 1°.

Met caotion of Iiigh-*her~ -0.20.

- ,&

F’-.6

-,4

-,2

&--

VK1 I I I , ! 1 1

Percent chord

(o) a. = 20.

WINAL AOVISMyCoMMl~EE FOR AEROWTl~

P Inlet Afien

H—H—t—H““s’=‘i”’”“’’i””-,6 – — -UpPeF au*04

o )0 20 30 40Percenf chord

(d)a.= 4°.

Fl~re 1A- Concluded.

NACA ACR No. L4118 CONFIDENTIAL Fig. 15a

o ./ .2 .3do “5 “6 “7 “8

(a) Inlet section, v~ /’v.= 0.53.

Figure 15.- Varlatlon with Mach number of peak ns&atlvepressure coefficients for symetri=d wing.

.?i.- $. 15b NACA ACR No. L4118

(b) Baslc airfoil section.

Figure ~.- con~l~ded,

NACA ACR No. L4118 Figs . 16,17

-2 -/ o / 2 3 4

oqo,deg

Figure 16.- Varlatlon of orltlcal Haoh number withangle of attaok for symmetrloal mln&

HATIONAL AOklWV

COMMlllEE FOR AE~

.7

,6k

$

.5

0 –– 7-, , ~ F: y_- .-

/ /

/ / - --

//

/ //

z / /

‘3

Figure 17.- Variation of orltical &oh number withinlet-velool.ty ratio for inlet section ofsymmetrical wing.

NACA ACR No. L4118

–2 -/ o / 2 3

0$0,deg

Figure 18.- Variation of critical Mach number withangle of attack for ❑edium-camber wing.

NATIONAL AOVISORY

CoMMITIEE FOR AERONAUTICS

7

050

(dq)

,/ -//

, .—0/ / -/

1//jr

z-. -—/ /

-Z~.— -— —/

._-. -

~ -z---- ‘

Flg.ire 19. - Variation of critical :,!achnumber withln16t-v510city ratio for inlet sectjon ofmedium-camber wing.

NACA ACR ~0. L4118 Figs . 20,21

.7

~

$.6

.5

‘ ““0. ‘yf ~’\.4.5 – - Bus/c (7I)tiilsecf%n——.33 – /

A /~ \

‘ -- -

/./ ,/-’”

,/ / ‘ \ / /// ./”.

/ .

-z -/ o I z 3 4

wd,u’eg

Fignre 20. - Varlation of oritioal Mach nnmber wlt.oangle of attaok for high-camber wing.

NATIONM ADVIWW

COMMli_FEEFOR AERON41JTK$

0 .2 .4 .6 .8 AO

.In/et Velo c.;fy roftb, I/”<

Figure 21.- Variation of crltioal MBOh number withl.nl.et-velocltyratio for inlet aectlon ofhigh-oamber nin&

Figs. 22,23 NACA ACR No. L4118

\<T —. ‘ a.b9—

. .\r 7

[ ~ — . _

/— 7—

!

1J 4 ‘3

w- [ Pfi

1 1 I 1 —.— ~ II

1 I

j+

+ A

‘4 ‘Z 0 2 4 6 8OC,deg

Figure 22.- Val%atlon of Inlet-velooity ratio wit~~~e of attack for varioua exit arean for Metaeotion of aymmetrioal wing. M. = 0.20.

I I I I 1 I7-+--+-’++___ J I II I +—— r 1

—L-L--l-&4-l I 1 r-% .42 ‘11v

IIx‘ -$-i1“7

i 1 , 1 , , 1 1 1 1 k 1 1 1 1. Iz-—-+-+ ; 1

+ + .k

NATIOtiALAOVWRY

C(MlhE FOR AE Iomlm s

Figure 230- Variation Or tilet-velocity ratio withangle of attack for veriobe exit ereaa for inletsection of medium-camber win~ 1% .0.20.

NACA ACR No. L4118

—

.6

.4

.2

0

-4 -z o 2 4Ko,&g

Figure 24.- Variatlon of Irdet-velooity ratioawe Of attick for variouO exit arean foraection of high-camber win~ U. . 0.20.

6 8

withInlet

F@ure 25. - Varlatibn Or internal 10BS wl~ an~e Of attackfor various inlet-veloclty ~tioa for inlet section of “BWKOetrlcal wln~

Figs. 26,27

—.—...——

I

NACA ACR No. L4118

-9 -2 02 -+ 6 8 /0 /2L=%,tieg

Figure 26.- Variation of internal lees with angle of attaokfor various inlet-velocity ratios for inlet section Ormedium-oamber win~

,6

,2

0

II l/1. ..55l I I I—— —.4O I I I I I I I

t

? / 1/..-

NATIONim IRY

–=+ -2 0 + 6 8 /0 /2do)CZg

Figure 27. - Variatlon of internal 10.9Bwith angle of attackfor vzrlous inlet-velocity ratlOB for inlet section ofhigh-camber wing.

NACA ACR No. L4118 Fig. 28a,b

—

.03

●✟

cdo

.0/

o+ -z o 4

G%, Ck.gz(a) ItiOt section, vl/Vo= 0.23.

1

()

/

<>

.03

J2Zcd

o

c b

o -4 -‘-~o,deg

—NATIONALADVISORY

CWWTEE FORAEmAm

{b)‘et ‘eottin’‘lFO = 0“33”

Figure 28. - Varlatlon of profile drag with angle of attackrOr t3ynunetrloalwing,

Fig. 28c, d NACA ACR No. L4118

.03

.02

.0/

o

.03

.Ozcd

o

.0/

o

-4 -2 0 z 4 6&O,deg

(c) Inletsection,VJVo = 0.4%

-4 –2 o 2 6~O,dq NATIONAL ADVISORY

CCMtiiHEEioRAERoNAuT~

(d) Inlet seotion, V1/Vo = 0.S3.

Figure 28.- Continued,

. . ------- ..- ., . . ,,. .

NACA ACR No. L4118 Fig. 28e, f

.03

.02

cdo

.0/

o

,03

.02

Cdo

.0/

o

1 I 1 // 1 1

/,v- I0 -#o

-4 -z 024 6~njdeq

(e) Inlet eectlon, V1/Vo= 0.85.

–4 -z OZ4’6%,~eg

(r) Basicairfoil section.

Figure 28.- Concluded.

I

Fig. 29a,b NACA ACR No. L4118

,03

o

.03

,02

c~o

sol

o

-4 ‘2 o 2 4 6M.] dey

(a) Inlet oeotlon, V1/Vo = 0.26.

-4 “2 024 (50COJc2’ey

(b) Inlet section,V1/VO = 0.S8.

Figure 29.- Varlatlon of profile drag with angle of attackfor medium-oamber wing.

NACA ACR No. L4118 Fig. 29c,d

.-...03

L!!

cdO.0/

o

.03

,02

cdO.0/

o

mlIMlrIELIQBm mmu

/

–4 -2 02.4 6OCO, deq

(o) Inletcootlon,V1/Vo = 0.45.

(d ,Inlet aeotion, vJvo = 0.530

~lgure29.-Contlnu@&

‘ig. 29e, f NACA ACR No. L4118

,03

●O2

cdo

.0/

o–4 -2

(o) Inlet

02OC#e 9

‘eotion’‘l/~o

4 6

= 0.80.

n!Y●UJ

.02cd

o

.0/

o–4 -z

(r)

0246~o)deq

NATIONALADVISJRYBnsiO airfoil aectlon. COMMllTEEFORAERONAUTICS

Figure 29.. ConOl~fio

r,

NACA ACR No. L4118

.03

.0/

o

.03

.02

Cdo.0/

o

f

(

\ \-. -.. --- .-—

>

(a) Inlet aectlon, V1/Vo= 0.33.

/ I 1.......I *

-c$/ “~: “=”

Q

~

MAOM .

//’)?+

\ 21 LCOMMIEE FOR------—. -—---

-—. _ —.

(b) Inlet section, V1/Vo = 0.45.

Figure 30.- Variation of profile drag with angle of :for high-oamber wing.

Fit?.30c,d NACA AC”RNo. L4118

.0/5

cdO.0/

o

,03

.02

cdo

.0/

o

-# –z o

&01de2q(c) Inlet seotion, V1/Vo

NATiONALAOVISORYCOW;HTEE FOiiAEROMUTKI$

= 0.55.

-4 -. 0

(d) Inlet section, V1/Vo =

Fl~re 30.- continued.

46

0.800

.-. . ... . . .

NACA ACR No. L4118

.,,

.03

.02

cdo

.0/

o

Fig. 30e

o 40

NATIO4ALADVWRY.......—-- -——.—.— —.

-4 -2 0 z 4 6Ko) deq

(e) Basic airfoil section.

Figure 30.- Concluded.

oco–

(deg)./ . .2 v

c/yJ=/4

::>‘? -...4‘1 40

A w— ; 6 D

I I

,

In/et sec+io~,~/~= 0.23

–./.2. .3

“4M0 “5 “6 “7

./

cm%o-- — ~ ~ —“

Zf7/etsecf/ on, q/~=0.53.2 .3 .~ti 5 .6 .7

./ o

cm% Ao $ A Y

J~I

A

BusIc ulrfoj~ secfion w

-./ 1 1 i 1.2 3 .4 .5 .6 .7

No NATIONALAOVISORY

COMMITTEEFORAERONAUTICS

Flgtuw 31. - Variation d aeotionpitohlng-moment coefficlentawith L&oh number for symmetricalwin~

(d;;)-2 v00

0 2A

cm40

/ —bx F

=4 . 6 D

–./

In/GGf .secfl’oh, y/bj= 0.26.Z .%~o .5 ,6 .Y

ocm

C/+

-,/-

—~n/et s~c7’/’0/),y/bg=a53 I I I 1

.2 .3 .4 ~ .5 ,6 .70

0

cmY4 ‘ iL —“ “_ + &

-./@QsIc Qirfoi/ secfion

.2 .3“4 M. ‘s “g ‘7NATIONALAOVISORy

coMMl~EE FOR AIRONALITIN

Figure 32.- Variation of sectionpitching-moment coefficientswith Mach number for medium-camber win&

zo●

1-*H1-

(3

—

z

*c1s-

I

I

#z .3 .4 ;;” :5 .6 .7

,2 .3 .5 .6 ,74 M*

~iguro 33.- Variation of ●eotionpitohlng+oment coeffioiantswith Mach nomber for higb-oember win%

,. +./

.8

.6

Cfl

.4

0

~igure 3L - Seotlon nomal-f ore.coefficients for ●yrnetriomlwin~ ~ . 0.20.

zo.

rlb

1-Cn

❑

.8

0

.2

i

/

/’

//

/o

/

/

Figure 35.- Seotlon normal-foroecoefflclonto for medium-camber -in& Ho = 0.20.

.8

.6

c“

.4

.2

0

.2

P//(f1

,(/

\ .

/’/‘ 1’ 2Ze7’.sechon

/ / V%( ?

/

/’ ● 0.$0f /

v .55/ v .45I

4// v

x .33

. .1 O/dQ V

/+t%on

f

/~ !

1I

/

/

/

,

/‘

NATIONIL AOVI:3Ry

cm MlllEEFOR AmT wru

-4 0 4 /2 /6 20a-:+=

Mgura36.. Sootion normal-foroeooeffioionts for high-oamberWa& Ho = 0.20.

wCn-wCn

!2

zo●

.

1-

OI

I

,8

.6

,4

c~,2

0

72.

5+

&?.

(a) InletHeotlon, V1/Vo s 0.530

.6

.6

.+

c~.2

0

72

+.2 .3 # .5 .6 .7

Ma

(b) rnelo airfoil eeotlo~

Figure 37. - Variation of aeotlon normal-foroo ooeffloiento with ~oh numberfor aymmetrioal wln~

zo.

1.2 .3 & 5 i6 .7

M.

.8

.6

.4

Cri.2

0

-. 2

–.4A .3 .4 .5 .6 .7

&

%1-m

●

z0.

II(A) Inlet ●ootlon, Ylpo - 0.53. (b) mBIOaltioll100tiO& P

*H

rlmro38.- Variationor●eotionnormal-fore. oooffioionts with udi numberfti ~dium-oombgr TjJI&

FCD !

I

I

io(

.8

,6

.4Ch

.2

0

-.2

-.4

Lo

.8

.6

4c~

.2

0

-./ .

-.4.2 .3 .4 .5 .6 -7

m.

(a) Inlet aeotion, VI PO - 0.59.

._—_+..-a ._$----- ..—.-

z>c-l

P

~igure 39. - Variation of seotion nomal-f oroe oocffloientn with l@ch numbwfor high-camber win~ .

Fig. 40a,b NACA ACR No. L4118

A* o

/AH +0

“4 -z o 2 4 6OCOa deg

(a) MO= 0.20.

NATIONALADVISORY

.4 ‘COMMITTEEFORAERONAUTICS

.2 ‘

o I

-4OCo,deg

(b) ~ = O.SOo

Pigur@ 40.- ~feots of s~eral lnte~l-lip falrlngs on internal100s08 for Inlet Seotion of mymmetrleal WI= vl/vo = 0.5s0

I

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