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COMMITTEE FOR AERONAUTICS ~ATIONAL ADVISORY
wi’urllm lMWOIU’
ORIGINALLY ISSUED June lg~ aa Ccmf~dent.lal Bulletin lNTCESTIGATION OF EXTR13fE LEADING-~E R07XX3NESS ON THICK LOhT-DRAG AIRFOIIS TO INDICATE THOSE CRITICAL TO SEPARATI(X By Easl;manN . Jaco%s, Ira H. Abbott, and Milton Davidson Langley Memorial AeronauiAcal Ltiboratory Langley Field, Va.
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NACA WARTWIE REPORTS arereprints ofpapers originally issued tvprovide rapid distribution of advance research results toan authorized grouprequiring them forthewar effort. They were pre- Some ofthese reports were nottech- viously held undera security status butarenow unclassified.
d tally edited. Allhave been reproduced without changeinordertoexpedite g:neral distribution.
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IIIPmS!CIGATIOIS OP IIKTRUME LEADIMG-EDG3 ROUGHNESS OH THIOK LOW-D~(3 AIES’OILS TO INDIOME l!E06El OEIZIUAL TO 8MPAHATIOE By Uastman M- Jaoob’e, Ira Es Abbott, and Milton Davidson SWeral airfo~le, Inaluding a conventional HAOA 23021 and come low-drag airfoils for whioh the thloknees had been Inoreaeed. to the point that they were ooneidered .
doubtfully coaservatlva with respeot to separation, were investigated as smooth airfoils and after the applloation of a standard roughneee.
The results show some of the air- .
foils to be oritloal to reparation resulting from suoh flow disturbances.
It is conoluded, pending the further .Investigatlon of eeparatlon dlffioultiem, that airfoil seotlons falling definitely within the aonservatlve range should be used.
INTRODUOTIOH The RAOA low-drag alrfolls first inveetlgated, the airfoils db”alt with in the earl~er ~plio”atiuns mainly to .
pursuit airplanes, and most of the airfoils for whlah”.data are preaiented in referenoe 1 were intended to be of oon- Bervative design.
Ho very serious reparation dlffioultleo should therefore arise in operation with theee airfoils, In other even though the leading edge beoomes very rough.
words, “moat of the airfoils were “SO ohoeen that the thlok- neae, the camber, and the poeltlon of. minimum preosure would lead to a conservative pressure recovery over the rearward part of the upper nurfaoe. Eor l uoh .a$rfoile the reoovery oould he made witlrout marked eeparatlOa, even In the ~resenoe of “a boundary layer sxueeoively thlokened by premature tr~neition ahd ““roughness near the leading edge of”the a3rfoil.” Thus, i-t wae.expeoted that . .
the new alrfoile would give drag ooeffioientm $n the came range as conventional airfoil-e when both were similarly roughened rather than give exoeesive drag ooeffloients aseooiated with turbulent separation.
On some more reoent applications to long-range bomb- ers, however, root seotione have been Increased In thlok- neos to the point that their relation to the conservative range has beoome, at least, dOubtful. The range of con- servative airfoil design as contrasted with the oritioal range as determined by the choioe of thlokness, oamber, and position of minimum pressure is disouseed in general terms in reference 1. Results are therein presented on at leaet one airfoil that vae eetlmated to fall in the doubtful range, or in the range wherein airf-oile may be oritical to separation resultlng from leading-edge rough- ness.
The present series of teets was undertaken to obtain quantitative data with regard to these limitm of ooneerv- ative airfoil design. The program contemplated an inves- tigation of a taerie.e of atrfolla emtlmated to lie olose to the doubtful range. It was thus thought that a com- parison of the test results ?or wings with and without a standard roughness applied to the leading edge of eaoh would give quantitative data tending to define the range of conservative design.
CHOICE 03’ STANDARD ROU3KEESS It was desired to ohoose an extreme rough aondition as a standard roughness to be applied to the leading edge of the various airfoile and at the came time one that would not alter the oontour of the seotion.
The etandard roughness might thus simulate an extremely rough oondi- tion that might result from mud or rough ioe on the lead- ing edge of the airfoil but, of course, aould not repre- sent thiok ic!e accumulations of the worst type, whioh would seriously alter the airfoil oontour.
Uith suoh considerations in view, a standard rough- ness consisting of Carborundum partiolee thinly applled over the leading-edge part of the airfoil wae adopted. A .
mloroseopio examination of the partioles used showed them to be shaped like lumps of coal and to have orosswise dimensions near 0.010 Inch and seldom greater than 0.015 inoh. The partioleo were applied to one eurfaee of Sootch %.8RO; th.e. .ta-pe ~ae.,– i“n’rtu%tii’ attaohtid “to the. leadlng edge of the airfoil. The uee of Sootch tape in applylng the roughness permitted Its quick removal for the oomparatlve tests of the smooth airfoil. The Carborundum partloles were retained on the Sootoh tape by a thin coat of shellae allowed to heoome taoky before the application of the par- tloles. The tape and roughneso extended around the lead- ing edge of the airfoil sect~on for a total surfaoe length of 3% inohee, equally disposed above and below the lead- ing e~ge. The Carborundum was eo thinly. spread On thle surfaoe that 6 to 10 peroent of the area wag aotually oovered by Carborundum gralne.
The airfoil modelm were of ~-foot ohord and ~-foot span; the roughneso strap was extended aoross the entire spa from wall to wall in the tunnel.
Eor the full-eoale wing at a Reynolds number oorre- Eponding to that of these model tests, the corresponding roughness is geometrically similar to that on the model.
The roughnese m= thus be ocns-idered to be something like particles of sand somewhat less than & inoh aaross adher- .
Ing to the leading edge of a wing of 100-inoh ohord. Suoh roughness oondltions, of oourse, oannot b~ considered typloal but It W.aS hoped that the oomp~rative resulte of the same roughness On various wings would be of value as representing a standard roughness o,ondition, extreme, but Of a type not markedly altering the Original airfoil .oon- tour. .
.1 Ti9sYS AI?D RESULTS Tha”teete of the ~-foot-chord airfoils both vith and without roughness were of the routine type, app~oximately, as deso=lbed in reference 1. Meet of the results were obtained at a Reynolds number of about” 10 million. Som’e rOsults were also inoluded at a Reynolds number of approx- +matelyq.6 million in order td @ve some information on Q-.
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The following airfoil seotion~ were investigated smooth and with the standa@ roughneeo: ..
) EAOA 23021 ,Boeing 7-seri&s type,” 0920a thlokness “ EAOA 65,2=222 (&pprox.) - ‘ . .
NACA 66,2-422 (approx. ) “ l?ACA 65,3-418 ‘ The results of the airfoil tosta are presented In figures l.to 5 in standard chart form exoept that the drag aoe”ffibients h“ave been plotted to a emaller staale than that usually employed. The reduoed drag scale per- mits the high drag values associated with separations that oacur on some of the r“ough models to be shown.
DISCUSSION Comparison of airfolls.- The I?ACA 23021 aeotlon (fig.
1) is intended to represent E thick conventional airfoil.
It is evident that roughnese on suoh a seetion produoes a serious loss in the maximum lift eoeffioient. The mini- mum profile-drag coefficient is Inoreased from 0,0068 to nearly O.0100, Indicating the additional drag aseoolated with the premature transition and the roughness. The drag eo.efftcient appears to increase somewhat more rapidly with the lift coefficient than for the smooth alrfoll but the variation remaina normal, Increasing progressively with llft on approaching the reduoed maximum lift ooefflcient of the rough airfoil.
Thus , only the usual progressive separation effects are evident ae the maximum lift ooeffl- oient is approached.
The Boeing 7-eeries type airfoil in figure 2 is typ- Ioal of airfoils ehowing marked separation effeots due to roughness. The lift curve begins to chow a 10ss at small “ positive angles and the upper part of the ourve has a re- The maximum lift coefficient for the rough duoed slope.
airfoil is ap~roximately 1.1, a lower value than that of The mln$murn drag eoeffi- the oonventiomal rough airfoil.
oient for the rough airfoil is approximately 0.0116, a Talue only a l~ttle more than that of the conventional rough alrfoll; but the drag inoreases sharply above a lift coefficient of 0.6, indioatlng the onset of marked 1< .
- l “- J@Muwe-.-paawt %0=7d akiatt Alla G* l.t.al#.a .gw .@M)k . the 1~ft-
“. ~-o,urve slope ohangeam;”.-Tha. drag Ooeffioteniie awe seen to ...-.
.booomo rery ‘high at largsx. lift ooaffl.olente. : .- A oimilar, although lee-s ‘draetio, behavi& w~”ll’be .
obs~rved for the airfoils EAOA 6592-222 (approx.) and ,.
EAOA.65,2-4W ,(approx,) in figures 4 and 4. The airfOil vi$h ‘the hi,gh,or oambev ap.pearm to be eomewhat more ,unoon- 80.rVatlVOm . ‘ . , . .’-P .;.. .
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. . The ohar&ot,er.istloo o.f.a low-drsg,air.foil of the oon- .s.or~a.tive tyyo... BA0466t:~P418181. aps+shoyn In f~~re 6. A rQ.du@iw ia mazlmqm..l,i:$t ~to L.S9 Is observed ta he Of tha. same -type as that.~~.owp for. the oonventlonal airfoil .
due to rougbneea..
$hb.rninlmum drag eoef.fioient .iB.in- oreao@ to approximately ,t,he same value as that sf the ..
oanveati.onal airfoil alth~ugh. the inorement due to rough- nops SE, of oourse, larger on aooouat of the”lower initial value for the low-drag airfoil.. The drag ooeffioleats show a:prqgressive Inoreaee on approaohlng the maximum lift ooeffioieEt, as di~ the oonventiona} ai%foil., %hua ohoving drag ooeffioients in the same range- a~ tlibse’”of the.roug~ aonyentional air~atl for lzf,t. ~oefflo~onte be- I 1.OW I,L.*”: It is .therefose aonoludod that. for”a oone.erva- tive airfoil .of the low-drag typo, roughnetio eh-ould pro- dupe ao mqrked separation effects apast from the effeots that normally ottourwhen the maximum-lift attitude la approaeh,edw “ ‘.S~~nifioaaoe of wqke-surve~ measurements made ia the ..
preseaoe of separattoa.- Ia toots made .to determiae drag by m.eaas of wake ’surveye .ia the presense of eeparatl’oa.
the dead air la tho rdgloas of 100EU separation may. tead ‘- .t.o ’deV.iate ‘spaawi-ee” iti duoh a wa# as to paoe pff at ~the ‘. ~., -‘ l u%voy ’-pl~-e,:.indtd”a%~ng axd.essi~.e dra-g;. or 40 devia%.
Outward l O ao to pass off ~a some ot~am,:pl~nes...iad~oating a deflolent drag ia the l urvey plane. ~or that reaoon, .Zi niokti”’’ixietaad~g wlM*e-me~aratloti wag “liEely to ooour ,.
apaawise” daag. marveym were made.
Some of ‘the reeulte of “-such, eurv”eys as?d.mhowa tn,.tl~re.m .6 aad 7. ~or.the: aaooa- .“ serwat$ve u~sfoll (“fig,. “6)s.it will.be noted that..separa- ..+ . .tloa doeti’%~”d :to-’be. XocaXIWdilWn s raglom.sear tiidapan.
:. .’ .Zt v.as. fmmd+th’at %his. ~.epasat.l.oa regtoa-migkt tend .to ,.
“shift .opanwl-se;. thi’g tehdezioy .Zaads to.’laboasistsnt re- ..
l skts -at the ourv.ey.”plaad, auah as.tiro~e ohowa ia figure “:a. ..The results show,, thereforo, that the .sep.arated .seglon~ may .8s 10Qal a.@.that the d~ags measured behind . ..- . -, . . .. . . . ,.
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these re@ons aay ~S excssstve rather than quantitat~vely correcte It seems c~~ar, nevertheless, that such seatio~s shotving even local Separation ca~not he considered conserv- ative and their use should be avoided, “ On the other hand the spanwise. surveys shown for a conservative low-drag airfoil in figure 7 indicate a con- sistent spanivise vayiation of drag, hence, an airfoil that .“is not prone to local-flow breakdown. Spanyise surveys are usually made as part of the testing procedure in the NACA two-dimensional *unnela The results shows in figure 7 for the smooth condition are typical, of the results usually obtained when no pronounced separation is preseqt.
Spanwise surveys made on 2.arge-.chord wing sections repre- senting practical c.onptruction sometimes show moderately , large variations of drag along the span even when the sections are considered. to be well within the’ conservative range, SuGh variations ar+ attributed to changes in skin friction resulting from l?cal. accidentally distributed surfac~ roughness and these variations tend to dieappear as the surfaces are improved.
The airfoil boundary layers near the tunnel walls are~ “of course, aff+cted by the presence of the walls with the possibility of’resultirig spanwise flows that might af- .
fe~t tl..e resistance to separation of the flows near the center of the airfoil as well as the drag measurements Tests with different chord-length smooth models of” air- “’ foils within the conservative range have failed to show any significant spanwise drag variation or “va??iation Of airf::il cha~ac$eristics wi~h chord, that would be expected if 3zL*h effects were presente it is planned to extend sue”l tk>sts$ however., to” include rough as well as smooth mo(ieks of airfoils in or near the critical range.because Of the “possibility that such effects ’may be present under these conditions, “ Avplicatiion of resultsa- “The present ”results strongly su~g”o~=t that the use of airfoil-s ‘which. do qot fall within the c?nservati’ve range should”be avoided.
Sections of this ~ype that have shown a tendency to I)reak down locally in the pr”esence of a leading-edg~ dist~rl)ance may also break &-own .iri the presence .of other :distrubances, such as those dtie to fus”elage or na,celle interferences, construc- The p~eci,se limit.~ of the c~n- tion irregularities, e$c= servativerange, however, are at present not definitely establisheil. “In fact, the only quantitative data tending I’ortu- to define the limits are iQose herein presented.
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nat ely, the low-d~ag airfoils, with the exception of a few above 20-percent thickness, in use or under consideration for practical applications may be judged to be conserva- ‘ tive by comparison with the NACA 65,3-418 shotin to _oe m’ satisfactory by the present data. Pending further inves- W$ tigations, sections that cannot be judged satisfactory by !3 such comparisons should be specially investigated by tests such as the present ones in the two-dimensional tunnel.
Difficulties have usually arisen ~nrough the use of excessively thick sections.
A suitable remedy is o%vious: While keeping the same spar depth, the wing chO?d may be increased to reduce the thickness ratio until the section falls within the conservative range. During the tests of a bomber model in the 8-foot high-speed tunnel a leading- edge glove was used to reduce the section thickness ratio.
A _p@tter plan would have been to increase the chord of the entire section in ordep to obtain the same reduction in thickness ratio.
Finally, two other possible methods that may eventu- ally lead to obviation of the difficulties herein consid- ered may be mentioned. With relatively large nose-opening air intakes it appears to be possible to employ thick sec- tions without excessively low minimum pressures and the attendant unconservative press”ure recoveries. The use of suitable lift-control flaps with slots or other forms of boundary-layer control should be advantageous in obviat- ing the separation difficulties.
CONCLUSIONS Pending the further investigation of separation dif- ficulties, airfoil sections falling definitely within the conservative range should be used.
Langley Memorial Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va.
IHFER3NCES 1. Jacobs , Eastman N., Abbott, Ira H. , and l?avidson, Milton: Prelirnin.ary Low-Drag-Airfoil and l?lap Data from Tests at Large Reynolds Numbers and Low Turbulence.
NACA -4.C.R. $ March 1942.
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i I The values of eection ltft coeff~cleat (ftgs. 1 to 6) ahoald be corrected ?Y the following ’equation = 0.965cZ + 0.015
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: TDT 172 -1.6 -16 -a o 8 18 24 aeOtiOn angle of attank, a, deg Seotion lift ooeffioient, ‘t airfoil.
Figure 1.- MACA XMMl >, v & I -La -.8 -.4 0 1.8 M 8.0 8.4 awtion”Lft m:fioient,OJ -.
s Boeing 7-eerie8 type, O.* tbiobess, airfoil.
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-. 8 Airfoil” EAOA & ,aaaa -- m m 1.0 fmmlrox. ) —.- ._r.-—., -l”a~ - i R Ohord 84 in.
w TDT 108 and 17s TOmt I I I -1.6 K .0 .
.3.8 -a o a la 84 -1.8 -.8 -.4 . 1.6 8.0 8.4 ‘-6wtiun a@Le of attmk, u, de w fi~. &- IAO, 66,8-88, (~1’ox.) d,,Oi!yim ‘;fi ‘;;:ioid” 80’ “ I I
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o moth a.8— I RY1;1306 8.4 –x R= 8x108 +R=lOXIOe a.o 1.6 w o - l.a.
j ~ .8 c -1 d .4 # J :0 -. 4 -.8 AIrfoifi IAOA &,== s - 1.0 (qlprox. L -loa R: ytiti:a4 in.
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‘1 -1.0 la a4 -a4 -la lb lMOtiOf UVJ1O :f attao:, 6, d~ b Figure 4.- mcA 65,a-43a (qWOX. ) ~rfoil.
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ho 8.8—0 R?:;:;nx @ — Lewllq 8,4-x R-S._ _ +R=lO.lx lOal l!ll! 111 en Figure \ \ Pig. 6 \ + \ boa’ mndltion .
\ e E-1CM06 : (%=00 .09 Leading edge ro’@h ..*o.. ~ . .*-, .
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d Mgure 6.. 8panwlse drag eurv~ for the BAOA 65,2422, a=l.O(a~rox. ) airfoil.
Fig. 7 N’ACM.
f Smoo4h condition O &&@; —
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d + E k VI + Figure 7.- Spenwise drag survey for the NM2A 65,3-418, a=l.0, airfoil.
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