Document
NATIONAL ADVISORYCOMMITTEE
FOR AERONAUTICS
TECHNICAL NOTE 2177 LOW-SPEED CHARACTERISTICS OF FOUR CAMBERED, 10-P&CENT-THICK NACA AIRFOIL SECTIONS By George B. McCullou@ and William M. Haire Ames Aeronautical Laboratory Moffett Fieldj Calif.
I I I I Washington August 1950 ...-=---- ~ 1> ,.
.“ ,., -.-. r--tr. A:.
i ~--- r--” - “ - , -, ill=--L- ,.- , —..
.-% . .. . ..-_ .. .. . -. . . . . . . .. . . . . -. ----- ...—. .
I tell lJ6Ff/WIY IUW3, NM
IllllllllImlllNllIIn
DOL51133 N4TIONAII ADVISQRYCOMMI’19?EE FCiR AERONAUTICS .
TECHNICAL NO~ 2177 LOW-SPEEDCHARACTERISTICS OF FOUR MMBERED, 10+ERCENT+EICK NACA AIRFOIL SEC!JZONS By GeorgeB. McCulloughand WilliamM. Haire SUMMARY A two+imensionellow-speedinvestigation was made of four thin, canibered airfoilsections. The airfoilsectionswere the NACA 64A31O, a= 1.0; the 64A81O, a = o.8 (modified); and the IIACA 0010 canberedto the same two mean lines. The data, obtainedfor Remolds nmibersof 3.7 x 106 and 5.2 x 10s, includemeasurements of lift, drag, pitching moment ~ and chordwise distribution of pressure. The effectof surface roughness was investigated as well as the effeet of a splitflap deflected 600.
,. It was found that the NACA four-digi~eries sectionsdeveloped greatermaximumlift than the corresponding NACA 64A+eries sections for all test conditions. The maximumlifts of both serieswithout “ flapswere reducedby surfaceroughness; the effectwas greaterfor the sectionswith the smalleramountof cauiber.The incrementof maxi– mum lift producedby the split flap deflected60° was greaterfor the NACA four-digitieriessections.
Visual observation of tufts attachedto the upper surfacesof the models indicated that the stall of the sectioris canberedfor an ideal lift coefficient of O.3 was the resultof separation of flow from the leadingedge almost immediately after the appearance of turbulentsepara- tion at the trailingedge;whereas,for the sectionscanibered for an ideal lift coefficient of 0.8, turbulentseparation from the trailingedge pro- gressedas far forwardas the ‘7&percent-chord stationbefore laminar separation appearednear the leadingedge.
INTRODUCTION o As part of a generalstudy of the stallingcharacteristics of thin . wings, a two+bnensionalinvestigation was made of the effeet of a large changeof airfoilthiclmessdistribution on the low~eed stallingchar– acteristics of cmiberedairfoilsections. Specifically, the basic thick- ness distributions compared were 10-percent+thick NACA four-digit and 6hA-series sections. C~culations based on the method describedin ref er– ence 1 hdicate that cliff erencesin the high+peed characteristics of sym- metricalconventional and low-dragairfoil sectionstend to disappear when . . . . . —----- . ..-. ———— . . ..— --—— .——— -—— -—--—-- —— --- -..---——.—— .
NACA TN ?277 these sectionsare canbered by conibining them with the S- NACA a-type mean line. If there is little or no cliff erencebetweentlw two types of caiberedairfoilsections from the standpoint of higkspeed” drag,the choiceof sectionfor aircraft employingcanbered wings probablywill dependon the low-speedstallingcbacteristics. For this reason>it seemd desfiable to comparethe characteristics of typicalsections at low speed. , The modelsemployedfor the investigations were canibered for ideal lift coefficients of 0.3 and 0.8, therebyenablingcomparisons of the two sect ionsto be made for two fairlywidely separated amountsof canber.
The mean lineswere the same for each pair of modelswith the same amount of caniber, but cliff ered for the two amountsof caniber.
The data obtahed includemeasurements of lift, drag,pitchingmoment, the chordwise distribution of pressure,and visual studiesof the character of the stallas indicated by tufts of thread attachedto the upper surfaces .
of the models. Data were obtained “forthe models in the smoothcondition and with roughnessappliedto the forward8 percentof the chord,both with and withouta simulated splitflap deflected60°.
The investigation was conductedin the -S 7-by 10-f oot wind tumnd , No. 1.
Im!wl!IoN The data &e presented W the form of standard NACA coefficients definedas follows:
‘A
sectia profile-drag coefficient c% qc
()
cl sectionlift coefficient~ qc
()
design sectionlift coefficient %i maximumsectionlift coefficient cl- sectionpitc~+noment coefficient, referredto the quarter+%ord % M point ~ () ‘4 a mwin-linedesignation,” fractionof chord over which’ designload is uniform .
c airfoilchord,feet .
.- —.- —— - .— - —.
NACA TM 21.77 profiledrag per unit span,poundsper “f cot “ free43tream total pressure, poundsper squarefoot lift per unit span,poundsper foot pitchingmomentrelativeto the qmte~hord petit per unit span, pound-feet per foot .
local staticpressure, poundsper squarefoot free-stream dynamicpressure, poundsper squarefoot ,.
Reynoldsnuniber, based on airfoilchord -P pressure coefficient — (’”q ) chordwisestation,feet sectionangle of attack,degrees sectionangle of attackcorre spendingto designlift coefficient, degrees deflection of trailing+dge splitflap from lower surface,degrees MmEzs Four models,each of &foot chord,were constructed of wood and cliff ered from on& anotheronly in airfoilsection. When mountedin the wind tunnel,each model spannedthe 7–footdimension. Attachedto the ends of the modelswere circularp~tes, 6 feet in di&meter,” which formd part of the tunnelfloor and ceildnge TO permit the“measurement of pressure distributions, flusbtype pressureorificeswere providedalongthe midspan sectionsof the models. The airfoil-sect ion designations were as follows: NACA .64Jw0, a=l.O1 NACA 0010, a=llO} czi = 0.3 NACA 64A81o,a=o.8 (modified)l r NACA 0010, a=o.8 (modified), czi=0.8 lCharacterist ics of the a=l. O mean line are given in reference2 and of the a=o.8 (modified) man line in reference3.
.. ..- ---- -- —.— -. ______ .-- . . .
—— ---- —- -——— — NACA TN 21.77 Coordinates of the four sectionsare given h table I and sketchesof the profilesin figure1.
Ths effectof a 20+ercent-chordsplitflap,hingedon the lower sur- face,was simulated by attaching a flat steelplate to the lower surface of the modelwith 60° woodenbrackets. The s- bracketswere used with all four models so that the 60° flap deflection was maintained with respect to the lower surfacein all cases.
Roughness “consist@ of Carborundum grainsof approximately O.011-inch diemter was appliedto the upper and lower surfacesfor a distanceof 8 percentof the chordrearwardfrom the leaddngedge of the model. TIM grainswere distributed so as to coverapproximately 15 to 20 percentof this area.
TESTS The test data were obtainedfor two values of Reynolds nunber: of 0.131 and 3.7 X 106 and 5.2 X 10s, which come spondedto Mach nunbers 0.187, respective=. Teas were made ~or both valuesof Re~lds n&ber - .
for the models in the smoothcondition, and for the higher value (R = 5.2x 10e) with roughnessappliedto the leading+dge regions.
$ a Force measurements of lift and pitchingmomentwere made with ths whd+mnnel balance system. Exceptwhere noted otherwise, the data were corrected ‘for the effectsof tunnel+mll constraint and compressibility .
by the methodsoutlinedin reference4.
RESUEl!S AND DISCUSSION Lift .
.
The lift and moment characteritiics of the four modelsas measured by the wiqd-tunnel balance systemare shown in figures2 and 3. For all test conditions, both ~thout and with the siqulated splitflap, greater maximumsectionlift coefficients were obtained, for the NACA four+ligit- seriesairfoilsthan for the corre spending IJACA 64&series airfofls. For the airfoilscaribered for a designlift coefficient of 0.3, ~- for the four-digit airfoilwas greaterby about O.2. The superiority y of the four4igit seriesdiminished when the camberwas increased to that for a designlift coefficient of O.8. For this amountof canberthe increment of maximum1~ coefficient was less than 0.1.
The lift characteristics of the four airfoilsare Wmmarized in the followhg table: .-. —. -—.. .-— — ——.-—- —--- —-- ---- .—. . —-.-.—— . . .._ . ..-= —.= —.— ,.
NACA TN 2177 .
, Smooth Rough Surface Smooth 3.7 x 106 R =5.2 XIOS R=5.2x10s- ondition R= 64A31O “ 0010 6kA310 64A31o Airfoil c@.3 c@l.3 c@=3 “ 1.33 1.32 ~ 1.53 1.36 1.53 1*E %= . io8 .108 l 109 l 109 .110 .109 dcz/dao 1.85 2.37 “ 1.92 2.36 1.98 2.20 Czm with af = 600 .
0010 0010 ~ 64A81o Airfoil 64A81o ~ ,@A810 Czt=o. c@O C21=0.8 1.68 1.76 ‘ 1.70 1.74 L 62 1.69 c2- dc~/dao .104 a104 ,107 .107 .104 l 103 .
with 2.39 2~62 2.43 2.66 2.41 2.53 C2= af = 600 1.
.
The valuesgiven for the lift-curve slopewere nwasuredat the designlift coefficient.
In orderto providea check on the wind-tuunel balancemeasurements, the sectionlift coefficients in the vicinityof the maximumlifi of tlm smoothmodelswere computed by integration of the pressuredistributions.
The values so determined were about0.02 greaterthan those determined from the balancesystemexcepteforthe four+igit-seriessectioncanibered for C21=0.8. For this section3 was 0.1 greaterthan the balance- c2~ systemvalue.
The increment of maximumlift producedby the simulated splitflap was greaterfor the four+ligit-series airfoilsthan for the 6kA series,and for both seriesthe increment of maximumlift producedby the flap was greater for the more highly canibered sections.
The additionof roughness to the unflappedairfoilsreducedthe maxi– mum lift coefficient about0.2 for the airfoilscambersdfOr C2i=O.3J~d less than 0.1 for the airfoilscanibered for cZi=0.8. For the airfoils with the simulated splitflap, the application of leading+dge roughness produced a different effecton each seriesof airfoils. For the @+A310airfoilthe maximumlift was ticreased, and for the @kA810airfoilthe maximumlift was decreased; whereas,.for the four+ligit serieswith eitheramountof csniber, .
the maximum lift coefficient was decreased more than 0.1.
—.— --- —.—— — .. —.- .-. — .—. . —. - . . ... — -------- NACA TN 2177 6 as indicated by tufts, sho%d . “ Visual studiesof the stallpatterns, that the flow over the modelswas unsteady near maximumlift. In general, both the models canibered for Cz~=0.3 stailed from flow separation at the lead~ edge almostimmediate~- &ter the appearance of turbulentsepara- .
tion at the trail- edge. The models caribered for czi=0.8 stalled primarily from turbulentseparation.The area of separated flow progressed forwardto aboutthe 7@percent-chordstationbefore the stallpattern became confused by smallareas of laminarseparation near the leadingedge.
These stalklngcharacteristics were also Indicated by the pressuredistri- butions(notpresented’) obtainedfor angles of attackgreaterthan those corresponding to maximumlift. For the sectionscanberedfor Cz~=0.3, the peak negativepressurescollapsed abruptlyat maximumlift;whereas, for the sectionscanibered for CZ1=0.8, the collapseof the peak negative pressurewas gradual. The pressuredistribution near the trailingedges of the latter sections were relatively flat, indicative of turbulentsepa- ration. Withinthe range of the investigation (R = 3.7 x 1~ to 5.2 x 1~), the stallingcharacteristics of all the four airfoilsections were rela- tivelyunaffected by Reynoldsnuuiber.
The reasonfor” the more deleterious effectsof roughnesson the maxi- l mum lift of the sectionsof lower caniber may be e@ained as follows: An uuptilished resultof an tive stigation of the stallingcharacteristics of the NACA 63+09 airfoilsectionwas that roughnesson the upper surface reducedmaximumlift,but roughuesson the lower surfaceincreased maximum lift. Hence, it may be inferredthat, for a regionwhere the flow is lamhar and where there is a severeadversepressuregradient,surface roughness tends to promoteearly separation of the boundarylayer. Since the sectionscanibered for czi=O. 3 stalledprimarilyfrom ~ sep=a- tion, it would be e~cted from the foregoingreasoning that roughness would have a more adverseeffecton these sections than on the sectias canibered for, Cz =0. 8J which staUed primarilyfrom turbulentseparation.
i , As can be seen in figures3(a) and 3(b), the lift curve of the IDICA 64A81o sectiondepartsfrom that of the four-digit-series sectionas the negativ+liftrange is approached from the positiveside. Coincident with the shiftin the lift curve is a strongpositivetrend of the pitching moment. The shLi%of the lift curve is thoughtto be producedby a local- ized region of 1~ paratedflow on the lower surfacefollowedby reattachment. .Lower-surface ~essure distributions shownin figure4 support this belief. For an angle of attackof -5.2°, immediately prior to the forcebreak,there is a strongnegativepressurepeak near khe leadingedge followed by a severepressuregradient. For an angle of attackof +. 7°, the negativepressurepeek has collapsedand there is a considerable chord- wise extentof S* stantially constant pressure. A shilar abruptredis- tributionof pressureoccurred-for the NACA 6hAO06airfoilsectionfor an angle of attackbetween4.5° and 5° (reference 5). It was shownthat the .
collapseof the peak pressures was accompanied by the appearance near the leadingedge of a regionof separated flow which subsequently reattached to ..._. .— _ .._ _~ ..
---. — —.——— —.. —.. -.—.- .--— .—. — ---- NACA TN 2177 the surfaceend progressed downstream as a thick,turbulent boundarylayer.
The collapseof the pressurepeak on the NACA 64AO06airfoilalso was accoqanied by discont inuit ies in the force and moment characteristics.
With roughnesson the surface(fig.3(c)),& lift- and moment char- acteristics of the two sectionscanibered for c =0.8 are in good agreement Ii throughout the enttie lift range. The pitchingmomentsof both sections show a positivetrend in the negative-lift range similarto that of the smoothNACA @A810 section(fig.3(b)). By analogy,it is evidentthat a localized regionof laminarseparated flow formedon the lower surfaceof the roughenedfour+ligit-series sectionprior to the attainment of the maximumnegative-Mft coefficient.However,for the two roughenedsect ions, the extentof the regionof separated flow increased more gradually with ticreasingly negativeangle of attackthan for the smoothNACA 64A81o section. These characteristics were borne out by inspection of the pressure distributions and the balance-system measurements of drag.
The slopesof the lift curvesat the designlift coefficient for the modelscanibered for Czi=o. 3 were not appreciably affected by Reynolds nuniber or by the additionof roughness. However,for the models caribered for CZL=O*8Sthe slopesof the lift curveswere reducedfrom o.1o7per de~ee to 0.104 per degreeby a reductionof the Reynoldsnunberfrom 5.2 x lF to 3.7 x l&. A similardecreasein lift+urve slopewas pro- ducedby the additionof surfaceroughness.
The greatersensitivity of the more highly canibered sectionsto Reynoldsnuniber and surfaceroughness may be e@ained by consideration of the factorswhich influence boundary-layer growth. For conditions corre- sponding to designlift,the adversepressuregradientsover the rear por- tions of the models canibered for C2~=o. 3 were relatively mild, and as a re’suit the boundarylayerswere thin and had no tendencyto separate. For the models canibered for czi=0.8, because of their greaterdesignlift coefficient, the adversepressuregradients were more severeand causeda more rapidboundary-layer growth. These boundarylayershad developed to a stagewhere they were aboutto thickenrapidlyas is shownby the no~ linearityof the lift curvesabovethe designlift coefficient, and as a consequence were more sensitive to any circumstance adverseto boundary- layer flow. For this reason,reductionof Reynoldsnunberor the addition of surfaceroughuessreducedthe lift-curveslopeof the models canbered for cZt=o.8e .
PitchingMoment , The magnitudesof the pitchhg+mxmnt coefficients are aboutwhat .
would be expectedfrom examination of data for airfoil. sections having camberlines similarto those tested. However,the variationof the pitchin&mment coefficient with lift coefficient (figs.2 and 3) showsa —— ——--—— --— -- . . - . .. . . —.---, ---—- —- ——-— -—-- ——— -- — .
llACA TIJ 2177 , positivetrendbeyondths stall;data for similarairfoilsections from other sourcesshowthe oppositetrend. The reasaufor this is not under- % stood. Integration of the pressuredistributions also showedpositive trendswhich were less pronounced than those shownby the forcemeasure- .
mnts.
Drag behind the four models The drag coefficients computedfrom wake surveys The minimumsectionprofile-drag coefficients and are shownin figure5.
the corresponding valuesof the sectionlift coefficient for a Reynolds nunberof 5.2 x 108 are tabulatedas follows: Surfacecondition I !
Airfoil Rc Smooth * sect ion Cz Cz
c%in C%nin
0.3 . 0.0082 0.2 64A31O 0.0042 .2 .2 0010 .0051 .0084 c@.3 .8 .4 64A8~ .0048 .0087 .4 .4 0010 .0055 .0088 czt=o.8 .
The minimum” drags of the @U&series airfoils, as wouldbe expected, were less than those-ofthe corresponding fOur+ligit-series airfoilsfor all test conditions.The minimumdrags and the centersof the low+rag ranges of the mooth 6kA-series airfoilsoccurredfor lif% coefficients closeto the designvalues. The minimumdrags of the four-digit-series airfoilsoccurredfor l+ift coefficients considerably lawer than the design values.
?!he effectof increased carherwas to increasethe drag corresponding to the designlift of both the &A- and the four-digit-series airfoils.
The low+ag range of the @kA+eries airfoilbecam less well definedend — — of smallerexknt with increased caniber.
The effectof Reynbldsnunber on minimum drag was inappreciable for all the airfoils, but increasing the Reynolds nunberreducedthe drag out- .
sidetha law+lragrange.
..- -. ___.-. _ —— .——.. -- -.. --— . . ..- .
.
I!ACA ‘Ill 2177 The additionof surfacero@ness, of course,increased the drag in all casesand completely removedthe low-drag“bucket”of the 64&series.
The minimumdrag of the rough 6kA-seriessectionsoccurredfor the sam valuesof lift coefficient as for the four-digitseries.
l&essureDistribution Experimental pressuredistributions corresponding to three conditions, one near zero lift, one nesr designlift, and one near msxinmmlift, are shownfor the four airfoilsin figure6. These data have not been corrected for the effectsof tunnel+alJ. constraint or compressibility. The pressure distributions of the two atifoilscanberedfor c2i=0.3 (fig.6(a))do not indicateany flow separation near maximumlift;whereasthose of the airfoils caribered for c2i=0.8 indicateturbulentseparation over the rear 20 percent of the chord (fig.6(b)). In alZ respectsthe pressuredistributions are typicalof the two types of airfoilsectionsandno une~ected abnormalities are revealed.
.
Comparisons of the theoretical &nd eqerhental pressuredistributions corresponding to the designvaluesof lift are shuwnin figure70 The theo- reticalpressuredistributions were derivedby the mthod of velocitysuper- positiondescribedin reference2, and the e~erimental pressuredistributions .
were obtainedfrom data measuredon either side of the designlift coeff i- cient. The experimental data are for a Reynoldsnunber of 5.2 x 106 and are corrected for the effeets of tunnel+mll constraint and compressibility by the methodof reference4.
As is pointedout in reference2, the method of velocitysuperposition givespressuredistributions which corre spendto lift coefficients greater than the designvaluesby an amountdependenton the thickness ratio of the basic thiclmessform. (In the presentcase the amountwas about 10 percent. ) Both the theoretical and experimental pressuredistributions were adjusted by interpolation until the valuesof lift coefficient obtained by integration of the pressurediagramsagreedcloselywith the designvalues. The sect ion anglesof attack ~ shown in figure7 are thbse corre@ondimgto the adjustedpressurediagrams. (As is also mentionedin reference2, the theo- reticalanglesof attackare only approximately correctand shouldnot be used where great accuracyis requiredwithoute~rimental verification. ) In general,the agreemmt betweenthe e~rimental and computed pres- sure distributions is good. Such discrepancies as do exist can be charged to the limitations of the method of computation (which,for one thing, ignoressny viscouseffects)and possiblyto smallconstruction errors in the profilesof the experimental models.
.—-—.. ———. . -. -..-—— —-- —.—— _—. . ....— — . . . . ..—7 -—. —. —-. -— .
IIACA TIt 2177 COWLUSIOIiS The stallingcharacteristics of canibered airfoilsectionsas affected by the ItACA 0010 and @lAOIObasic thicknessdistribution were investigated at low speeds. The amountsof canherwere those for designlift coeffi- cientsof 0.3 and 0.8. From data obtainedat Reynoldsnunibers of 3.7 x 108 and 5.2X106 the followingconclusions canbe drawn: 1. The maximum sectionlift coefficients of the four+igit sections were greaterthan those of the 64&series sectionsfor u test conditions.
The superiority of the fotiigit sectionsdiminished with increasedamount of cmiber.
2. The stall. of the sections was little sffectedby the changeof thicknessdistribution, but was significantly affectedly cdber. Visual observation of tufts attachedto the upper surfacesof the models indicated that the stall of the sectionscenberedfor an ideal lift coefficient of 0.3 was the result of separation of flow from the leadingedge almostimme- diatelyafter the appearance of turbulentseparation at the traillugedge; whereas,for the sectionscanberedfor an ideal lift coefficient of 0.8, turbulentseparation from the trailhg edge progressedas far forwsrdas .
the 7~ercent+hord stationbefore laminar separation appearednear the leadingedge.
3. The effect of Retilds nuniber on mazimumlift was smallfor the “ range invest igated.
k. Surface roughnessdecreasedthe ~ lift of all the unflappedairfoilsections. This reduction was greaterfor the sections cam.eredfor a designlift coefficient of 0.3 than for those csmbered for a designlift coefficient of 0.8. Roughnessreducedthe maximum lift of the flamed fourQligit-series sections, but showedno consistent effecton the flapped64~eries sections.
split 5. The incrementof maximumlift producedby a s~ated flap deflected600 was greaterfor the four-digit-series airfoilsthan for the 64A series. For eitherseries,the incrementof msximumlift producedby the flap was greater for the greateramountof csriber. ” Ames Aeronautical. W%oratory, for Aeronautics, I?ational AdvisoryCommittee May 12, 1950. .
MoffettField, Calif. , -—— ______ . ..— —— -. -- -.. .-—---- .— - ..—— ..—. —-—- u NACA ~ 21.77 RElmHENcEs 1. IUtzberg,GeraldE., ahd Crandall,Stewart: A Study of Flow Changes Associated With AirfoilSectionDrag Rise at Supercritical Speeds.
llACA TN 18u, 1949.
.
2. Abbott$Ira H.~ von Doenhoff~AlbertE.~ and Stivers~Louis S.~ Jr.: Sumary of AirfoilData. IIACA Rep. 824, 1945.
Theoretical and Experimental Data for a 3. Loftin,LaurenceK., Jr.: Rep. 903, 1948.
Nuniber of lWU2A 6A+eries AirfoilSections. lIACA 4. Allen,,H.Juliu, - Vincenti,Walter G.: Wall Interference b a Two+Oimnsional+lowWind Tunnel,with-Consideration of the Effect MACA Rep. 782, 1944.
of Compressibility.
5. McCullough, GeorgeB., and Gault,DonaldE.: Bo_y-myer and StallingCharacteristics of the liACA ~AO06 AirfoilSection. NACA TN1923, 1949.
.
.
.
L... .— --— —— —— ---—-— --— . . ..—. .----- .—-. ——..— --— —.. -.— -.—.-—.-— IWICA I-2 m 2177 TABLE I:- COORDIUTES OF CAMBEREDAIRFOIL SECTI(X’W l?AcAoolo CU*.3, ado [Stdm19 ad OrMnates Riven in Mrcent of rstatlann alaord&Tk&i%-8giv8 in pal’cent of .d150il C&ml] - Umermlrraca Louermlrface St8timl I ardlaata Stntilm [ cTdia8t4 ordinate
==1-==-
0 0 0 1.414 1.726 -:.m3 -1.260 -1.45 %% ;:~ -1.4J 7.W w:635 -lah&
$% 2%
X3.565 20.512 1 .&w 20.x58 :%$ t 25.4’29 -.7=7 ~ .8’70 30.34 pg -.3* 2!J.8g %S4 50:OW 50:W0 l:E 59.&Q ~.w 60.037 5.409 59.953 70.o&2 4.512 69.938 g%j l;% 80.q2 79.s28 3.380 90.CX53 l.gao l.ti #.87& 95.* $.g . -:% loom .
.KJ5 L.L13.Etionn&l.MMA opeofl’adinsthrcm@ ;. K a’amls: 1.1oo. SlOga of radlna tInm@l L. E.:0.380.
. .:. .
.
H&oA64A310 ?rAcA6bA810 a-lo a-o.8 (Idifiea ) .
.
[Watllm Ed Ord.h8tas Elven in ~.or [Staticmn and=*&m la peroalt of airronciardl - 1 Uppsrmrfaue 1 Imf8raurfMa I Stitlm I Ora.lmata station I &MIUti matlm I Oraimta
==I==
o 0 0 0 o 0 0 0 .i?lk .976 -.723 .785 -.526 .39 .601 -.858 .638 1:% .852 1.231 1.123 l.~
:E -Um7 ?% :32 1.650
2.0S4 -1.403 2.4T5 l.%1 i?.ti -.787
-1.w ;“%$ -.832 2.759 4.8~
$g -2.164 8:016 -All %$ 3.436 7.332 +.420 p& I&al -.’m 3.970 9.832 14:500 IS.500 -.658 :% 20.457 19.543 %% -.526 7&& w% 24.6CU 3.264? 24.8w -.383 . 5.9M ~;g 29.668 2J.S& 34.742 %% M% 2% ::ZJ 35:258 I&# 40.MO .l.il.li)i &r& -3:252 4s.loo 1o.150 :3 ??;% .637 lo.2i3 %ff7 50.m 3.0% 6:334 -s.p %.$1 .s7 m.a?l 9.693 S4.979 2% 1.W( 60.039 *MS 59.885 ;:%J !&& f&69 M26 +X52 i?:% 64.8s 1.63.0 70.= 4:584 .0? %=3 4.668 7x50 -1.280 %!?
75.2s2 %% 75.U59 80.W ::% C&.g 79.W 7&930 % ;:?? -.* .
84.’@ 4.441 85.292 2.5e2 -.5& go.204 -.28s 3.004 1.836 m:ti 89:X g5.lob 1% i.5u 1.01.4 g5.038 MI% -.* &.%& loo.mo -.021 I.oo.ow --.a?l .Cel 1oo.000 . kel 0.6&’. T. E. radius: O.~.
, H.&&: L.E rdlns:O.@. T. E. radius: 0.C@3.
radius tlUol@l L. x.:0.380.
m of radios tbm@I L. E.:0.326.
I
_.. ----- .—. — ——.- . . . . . . . ... —..- .. ---- ... . .
. . . ..— —.. . ..
E
*
--.-—---
s E -------- C[l = 0.3, NAGA a”/. O mean he ---___ --_ —— ---- --- —--- - ql = 0.8, NAGA a =0.8 (motfh%edj mean fine — NAGA four- dig/t series . ..- IVAGA 64A - series I ~, Figun? l–Comparison of tie four aihil sections. - . ..-_ .. . . . . . . . --- ~ 14 ITACA ~ 2177 2.4 2.2 2.0 /.8 , ,
! !HIIII 1411111
1.6 1.4 /.2 c’ -.2 m , , , - -.8 .
I I 1 1 1 1 1 1 1 1 1 I I - -/.0
~y$qy7’-
.1 0 d -2 73 -16 -12 -8 -4 0 4 8 12 Section angle of attack, CO, deg Section pitching-moment coefficient, Cm (’d Smooth airfoil’s; Reynolds number, .37x IO=.
Figure 2.— Section liiff and pifchhg-momenf characteristics of the two models cambered for o design I’M coefficient of 0.3.
—.— _________ __ -——— —..
.— —— .. _. _.. . . ... . .. ----- .
NACA TN 2177 l /.8 1.6 /.4 .
.
-.2 -.6 -.8 I I , K-: , , , I I , I
I 11.1111111111181111
-/2-8-404 8 /2 ./ o -./ -.2 -.3 Secfionongle of affack, tqo,o’eg Secfibn pifchhg-mommt coefficien~ Cm (b) Smooth oi?foilsj Reynolds number, 52x IO=.
i .
.
Figure 2.– Continued .. —- - .— -.. — .— .—. ..— ----- .— — —- — — —- —--— ---- NACA TN 2177 .
2.2 .
/.8 /.6 /.4 /.2 Q- .
.
-.2 -.4 -.8 -/2-8 ”-404 8 /2.16 o + :2 -.3 Section angle of attick, G, o’eg Section pt)’thing-moment coefficfbn~ cm (c)Ai~oilS wjfhfeo~jng-@e roughness; ~eyno~dsnumber,S.ZX106.
.
* Figure 2.– Concluded.
——-. . . . .
_..-..-_ _____ ._-. ---- ___ _ ..— .
NACA TN 2177 2.6 2.4 2.2 2.0 /.8 .4 .2 -.2 81216 .JO+ :2 73 -:4 -/2-8-404 Section pitching-moment Section angle of attack, CO, o’eg coefthen~ C* (h) Smooth airfoils; l?eyno/ds number 3.7x /0’.
pitching-moment characteristics of the two Figure 3.—Section tiff and models cambered for a design /ift coefficient of 0.8.
.—. — ----- —-- -—— -—---- ------- ---- , .— — .,— —.. ——----- -—— 18 NACA TN 21.77 . .
2.6 2.4 /.8 /.6 & $ .8 l $! .6 , .4 .
.2 -.2 -.4 . .
+2 -8 -4 0 4 8 /2 ./ o -./ -.2 -.3 -.4 Section angle of affack, Co, o’eg Secfion pifching-momenf coefffcien( Cm “ (b) Smoofh airfoils; ffeynolds number, 5.2x 10’.
.
l .
Figure 3.— Gonfinued.
.
...— -.. —-.. —- --—— . -. —-- -—--. — ---—--—- NACA TM 217’7 2.6 2.4 2.2 2.0 /.8 /.6
Ii/i iiii J?iiiiiiiiiiiiifll I [I 1
I I 1 1 I 1 1 $ .8 .
I
I-m-t 1 -t -m-H--t
T m
$ .6 .
.4 .2 :2 :4 o -./ 72 -.3 -.4 -/2-8-404 8 /2 /6 Section pitching-moment Section angle of affack, uO, deg coefficient cm (c) Airfoils with leading-edge roughness; Reynolds number; 5.2 x /06.
Figure 3.— Concluded.
.—-. .— - —----- -— --- ——. .— - .—— —— -— .. -. . — —-- - —--- ..— -.-.-— -— -. -- NACA TN 2177 .
a .
.
.
.
.
4B 4.4 J 4.0 .
f 3.6 - A!ACA 64A8/O , . .
- Q%, ~e9 Cj d - S2 0.08 -5.7 0.02 d * 2.8 (“ 2.4 \ % l 1.6 – “, /.2 I . .
. .
.8 — — — — — — ~ *’ .4 =E= , ‘O J .2 .3 .4 .5 .6 .7 -8 B lo G%ordwise sfofion, x/c .
.
Figure 4–Lower-suribce pressure a’istr&tion for the NAGA 64A810 airfoil.
Reynolds number, 5.2 x /0!
—.——. . .-— ..-— -— .—— —-.
.
I . .
I Figure 5,- Section profile -drag characteristics of the four airfoils.
!
\ I 1, I .
NACA TN 21-77 . .
/ Flagged symbols ino’icofe “ lower surface.
v NACA OOl~C@3 Symbol ao,deg q o -2./ -00/ /.0 0.34 q A /24 /.48 \ , , > ‘O .2 .4 .6 .8 LO 0-2.4.6.810 Chordwise station, x/c (u) Ai~oils cambered for a design I..t coefficient of 0.3.
, Figure 6.— Representofive pressure distributions of the four .
Reynolds number, 52 x 10’.
cambered airfoils.
n . .
--———— --—. .— -.. _— __ --- .. . —..
NACA TN 2177 .
Flogged symbols indicate lower surface.
/2 , // .
/0 NACA 64A8/O NA6’A OLV~ c@8 Symbol ao,deg Ct 4, 0 -62-003 q 20 C186 A /24 1.7V .4 ‘ .6 .8 LO Ghortfwise station, xfc (b) Airfoils combered for Q design lift coefficient of 0.8.
e Figure 6.— Concluded. “ .
.
——-.—- —--- .—. — ——... . .——. ———-— — .
NACAm 2177 .
‘o -./ .2 .3 .4 .5 .6 .7 .8 .9
‘lb Ghordwise station, x/c .
Flagged symbols indicofe lower surface.
w l /.2 .8‘ iVAGA 001~ C&=~ I ‘O J .2 .3 .4 .5 .6 .7 .8 .9 Lo Ghoro’wisestation, x/c (u) Airfoils cambered for’ u design lift coefficient of 0.3.
.
Figure Z— Comparison of the theoretical and experimental pressure v distribution corresponding to the design lift coefficient.
.
.
.
— ——. ..—._ _ -—-—. .— —— -.— .-. .- -— - 2.0 /.6 /.2 .8 .4 Chordwise statio~ /Vaggeo’ symbols Micafe lower suhce.
2.0 .
/.6 I I NAGA 00/(2 C. =0.8 —Theory - a>0.5° O ExDen”ment a,=L40
A I I I T
.
I 1 I I I I 1 I I I
I ‘=Q9+-
I I I I I I I I
“O ./ .2 .3 .4 .5 .6 .7 .8 .9 40 l Chordwise station, x/c (’b) Airfoiis cumbered for a design M coefficient of 0.8.
.
Figure Z— Concluded.
.- .
NhcA-Lmgley - 8-s1-s0 - 985 —---.——— —— .-— .------— —— . .. . —.. .-— -—- --———— --— — --—. — --—- -— -