Document
., J. Y NATIONAL ADVISORY COMMITTEE FOR AERONAUTI .—
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ORIGINALLY ISSUED A@l 1945 8S L5C248 AdvanceConfidential Report BYmd$mmasm COMPARISON OF PITCHINGMOMENTSPORIXKZD BY SPOILERSAND SOME AERODYNAMIC CHARACTERISTICS OF AN llACA 23012 AJRFOILWITH VARIOUST13?ES OF AILERON By Paul E. Purser and Elizabeth G. MdUnney Lan@.eyMmm”ieJ.Aeronautical. Laboratory LangleyField,Va.
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N A CA LIBWY YZGUY “kEMORL4L AERONAUHCAL LAE301LlT()~y WASHINGTON ~W@ey .~ield, va.
NACA WARTHVIE REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- Some of these reports were not tech- viously held under a security status but are now unclassified.
nically edited. All have been reproduced without change in order to expedite general distribution.
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31176014033732 NACA ACR No, ~C2&!r” = _ I — —.—. . . .——” /“ NATIONAL ‘VISORY COMMiTTEE FOR AHWN AUTICS P .
- . . ..-. ,,, : ‘ ADvJWCE -C ONFIDENTI@ REPORT COMPARISON OF PITCHING MOMENTS PRODUCED 13YPLAIN FLAPS AND BY SPOILERS AND SOME AERODYNAMIC CHARACTERISTICS OF AN NACA 23012 AIRFOIL WITH VARIOUS TYPES OP AILERON By Paul E. Purser and Elizabeth G. MoKinney smmt An analysis and comparison has been mede of the pitching-moment characteristics of aihfolls with plain flaps end spoilers.
Aerodynamic section characteristics of en NACA a5012 airfoil having a retracted slotted flap with a plain, a slot-lip, and a retractable aileron are also presented for a lgrge range of eileron deflections.
The analysis indicated thet the pitchtig moments produced by spoilers were less positive than those pro- duced by plain flaps of equal effectiveness. The data from two isolated oeses indicsted that the pitching moments creeted by the spoiler incre~sed less with Mech number then the Ditching moments produced by the plain flap.
The positive. values of the pitching moments produced by both the plain fl~ps and the spoilers decraased as the devices were located nearer the sirfoil leadir,g edge.
INTRODUCTION The NACA has undertaken a brief investigation of the pitching-moment characteristics caused by verioue lateral- control devices for application to wing-twist problems in high-speed flight.
Pitahlng-momant data for plain-flap controls have been published previously (see references 1 to 5, for example) end some data for spoiler-type controls have been published in references 5 to 7.
The effects of trailing-edge modifications on the pitching-moment characteristics of eirfoils with plain fleps have been discussed in reference 8.
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NACA AOR No. L’jC2&a Tests in two-dimensional flOw of an NAcA 23012 ai.rfoll with a plain aileron and with two spoiler-type ailerons (a slot-lip and a retractable aileron) were reported in reference 9, but the pitching-m,oment dsta were not presented, The present report gives the section pitching-moment characteristics end other ssction data f’or these three arrengemen ts. A brief anslyais is also included of vartGus data on the pitching-moment characteristics of airfoils with plain flops and with snoilers.
COEFFICIENTS AND SYMBOIS The coefficients and the symbols used herein are defined as follows: Cz airfoil section lift coefficient (Z/qc) cd airfoil section profile-drag coefficient (do/qc) o airfoil section pitchin~-moment coefficient about cm quarter-chord point of airfoil (m/qc~) aileron section hinge-moment coaffi.cient (ha/qca2) cha where blrf’oil section lift do airfGil saction wGfile drag m airfoil section pitching moment about quarter- chmd noint ha aileron section hinge momsnt dynamic pressure f ?)
~P~ q [2 c chord of besic airfoil with flap neutrQl chord of eileron Ca I?ACA A(X No. LsC2ka ~ .
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velocity of-free stream . ., ..
v ‘ “’’’--”’ mass density of alr P .
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end a. angle of attack for airfoil .of Infinite as~ect ratio, degrees aileron deflection, degrees; positive when trailing edge moves down slotted-flap deflection, degrees; positive when trailing edge moves down chord of.spoiler or retractable aileron arojection from airfoil surfece of spoiler or retractable aileron x distance from airfoil leadlng edge to flap hinge line or to outer edge of spoiler
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rate of chenge of pitching-moment coefficient with
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control d~flection at ~onstsnt lift rete of change of angle of attack with control
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deflection et constsnt lift km ‘ rate of change of pitching-momer.t coefficient
(’a o Cz with effective engie ot’ attack at constant lift
M Mach number (v/a) a velocity of sound In free stream I NACA A(2RNo. I@~a TESTS APPARA??TS , MODEL, AND The apparatus, model, and tests are described In reference 9. h brief, the 3- try ?-foot model was built tO the NNA 23012 t3irf0il profile ~d, when mO~ted ti the Langley 7- by 10-foot tunnel (described in reference 3), completely spanned the test sectlon~ The tests were made at a d~mamic pressure of 16.37 pounds per square fcot, which corresponds to a velocity of about SO miles per hour and to e test Reynolds number of about 2.lg x 10°, based on the chord of the basic airfoil, The effective Reynold ?
number (for maximum lift coefficients) was about 3.5 x 10 , baaed on a turbulence factor of 1.6 for the Lengley 7- by 10-foot tunnel.
, The airfoil profile, t% slot and flsp dlmsnsions, end the arrangements of the plain-flap and spoiler-type The chords of the plain ailerons are given in ftgure 10 and slot-lip eilarons and the chord of the retractable aileron in its most extended position were 10 pe~cont of the basic airfoil chord. The slotted-flap inbtallsti.an.
was that designated 2-h In reference 3.
All tests reported herein were made with the slotted flap retracted (Gf =.OO).
?lETHOJ)S OF ANALYSIS The prlmsry eerodymcmic factor contributing to wing twist during rolling In high-speed flight is the p3tching For normal moment produced by the lstm~l-control &evice.
wings and ailerons, the pitching r.loment produced by ~ileron deflection twists the wing in such a way thet the lift induced by the twist o~~oses the lift induced by the aileron deflection end thus effectively reduces the lateral “ control evalleble. The pitchlnq moment (or the wing twist) produced by s given eilcron deflection incrasses approx- imately ES tie square of the speed and Pt some poi~,t tho lift h?ucad by the wing twist balcnces thq lift induced by the &il~ron deflection and tha ~i.r~ime does not roll when the Eil~rons m?e deflected. The spoad ~t which the lateral control becomes zero is known as tinereversal s~eed.
In order to .!udae the re~~ti.ve merits or various to wtig twist, the lateral-control d;vi.~es with res;~ect be compared on tha pltchhg-mo?nent cheracteristica must NACA AOR NO. L5C24a ‘~) basis of equal” effectiveness. The pitching-momen t L... parameter used should therefore be based on the ohanKe in rolling moment. iift;’’0ti-6ff”btitive enEle of attac~ ,.
produced Zy the a~leron”rather than on l%e all~ron.
a cm deflection or spoiler projection. The slope —
()
~=o cl was therefore used to oapare all the pla”in flaps and spoilers on en equal-effectiveness basis, since this parameter indicates the change in pitching-moment coef- ficient resultlng from a unit chenge in the effeotive angle of attack of the portion of the wing covered by the aileron.
The slope of the pitching-moment-coefficient curve was taken at constant lift (CZ = 0.1) because, when the airplane is rolled by the aileron”s, the aileron section of the wing operates at nearly constant lift. Althou@ spoiler-type ellerons, since they are used on anly one wing at s time, opereta farther from conditions of constent lift than the plain-flap ailerons, the perameter at constant lift is still believed to be more neerly repre- sentative of actual conditions than a pnrameter at constant .
angle of attack.
m f A logical abcissa a~einst which to plot — {::0 CL ) for plain flaps would be the flap chord but, when spoiler data must also be preseuted on a comparable basis, such en abcisss is no longer logical because spoiler chord (or projection) is snalogous to flap deflection rather than to flap chord. The data were therefore plotted against the chordwise location of the plain-flap hinge line or of the outer edge of the spoiler. For devices such as the slot-lip ailerons, which were considered to be spoilers, the location used was the average location of the aileron trailing edge over the deflection range considered.
All the finite-span data (references 5 and 6) were converted to section characteristics by use of refer- ences 10 and 11.
The velues of aspect ratio used with the charts of reference 10 in computing section character- istics from the data of refarence 6 were corrected for Mach number effects by the method of reference 12. At each value or Mach number, the geometric aspect ratio was .— .
multiplied by the factor X:l - M2. This procedure gives an effective reduction in aspect ratio as the Mach number is increased.
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R3?S~ ‘!’S AND DISCUSSION Test deta. - The aerodynamic section characteristics of en NACA ~~2 airfoil heving a retrected slotted flap with e plain, a slob-llp, and a retractable aileron are preser”ted in figure 2. The lift and drag (or rolling- moment end yawing-inoment) ckaracteristlcs end the hinge- moment character~stics have bean amply discussed in refere- nce 9. The pitching-moment data presented in figure 2, tcgether with other mablisheci and unpublished data for Mach numbers up to ebout 0.3, have been s’ummarlzed and are presented in figure 3.
Pitching moments produced by pl~ln flaps.- The experim ental da~fi on.tk~epi tcning momenta produced by pl”ain flaps shown in flg~e 3 agree very wbll with values computed from Glauertl~ thin-airfoil tlieory (references.1 end 2) both in rmgnltude and in variation with x/c .
These dsts indicate that, for equal changes in effactive angle of attack (equal rolling mommts), wide-chord flaps than nerrcw-chord fl~ps prod’~ced smailer pitching momenta exd, consequently, that th8 usa of wtde-chord fl.!ws wouid ailow the sttairunent of higher values of the ravarsel suead. The use of wide-chord flaus, however, wiil be limited by wkethar their hinge moments can be well enough balanced to nroduce reasonable valuas of stick force.
It should be nated thet tha data of referenca ~ indicate th~t. the pitching momants m?oduced by plain flaflsmay be reduced by lncreesir!g the engle betwean the uppsr end lowar surfaces of the fiep et the trailing edge.
Pitchir!~ moments ~~oduced b spoilers.- The eXPe rTii3iiEaT +- tiate on l%e~cn fiHmomants produced by sqoilars form a reletivbly smooth curvu (fig. 3) and- Indicate th~t, for aqurl effectiveness, the spoiler located nearest ths eirf’oil leading edge aroduces the oc~’$ sm611est positive veluas of With a s~oiler ~,cz” () located ahe~d of about 0.45c, the wtig twist might augmnt rether than reduce tP.erolling effectiveness.
The use of spoilers located so near tha airfoil leading edge, however, Is not recommended since many prsvious wind-tunr,el and flight invastig&tions h~ve tidicated that the tendencies toward 1~.g ~.nderretic ection incrsme aE the snoilar is oleced naarar the ~irfoil leedlng edge.
The locations of the spoiler installations found acceptable .
- have ”variedfrom about 0.600-.to-*o.ut, 0.85c..
and s-;: ::::h+mt:z::%:d:L:;:2F
hat moments produced by plain flaps snd he pit thing spoilers have about equal variations with flap or spoiler M* chordwise location and thet the values of ?KCl ‘e
()
less positive for the spoilers than for the plain flaps.
When a ccmmarison 1s made at chordwise locations normally used - thet IS, 0.70c to 0.75c for spoilers and 0.80c to 0.85c for plain flans, the Pitching moments produced by spoilers are about one-half or less of the pitching moments produced by plain flaps of the same effectiveness.
Mach number effects.- Data on the effects of Mach number on the ~itchi~~ornents aroduced by control surfaces are relatively scarce. A compar~son is presented of the effects of Mach number on in figure J, however, the pitching moments produced by a spoiler aileron on the wing of reference 6 and on the pitching moments produced by a plain flap on an NACA 66,1-115 airfoil tested in the Langley 8-foot high-speed tunnel. The pitching moments produced by the plain flap, in addition to being larger than those produced by the spoilar, also increase witln Mach number at a rate greatey thar. that indicated by the Glauert-Prendtl factor -—. The Ditching moments 41: Y?
produced by the snoilar, howaver, increeseqat a rate slightly less thm that indicated by —-— Jl% over most of the Mach numbar range tested. Although the data shown In figure L are not strictly comparable and are for two isolated casas, there appears to be a possibility that Mach number effects may be smeller on pitching moments producad by spoilers than on pitching momants producad by Plain flaps.
CONCLUS1ONS An ~alysis of data on the pitching-moment character- istics of airfoils with plain flaos and snoilars indicated the following conclusions: iii— — — —— — -— . . -... .—— . ..— —.— 8 MACA’L ACR No. ~C&a 1.
The pitchti~ mxnents produced by spoilers were less positive than those produced by plain flaps of equal effectiveness.
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The positive values of the pitching moments produced by both the plain flaps and the sfioilers decreased as the devices were located nearer the airfoil leading edge.
The date from two isolated cases Indicated that en increase in Mach number csused less increase In the pitching moments produced by the spoiler than in those produced by the plain flap.
Langley Memorial Aeronautic&l Laboratory National Advisory Committee for Aeronautics Langley Field, VP.
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NACA ACR MO. L5C24a ~~j REFERENCES .
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1. Ames, Milton B., Jr., and Sears, Richard I. : Deter- mination of Control-Surface Characteristics from NACA Plain-Flap end Tab Data. NACA Rep. No. 721, 1941.
2.
Glauert, H. : Theoretical Relationships for an Aerofoil with Hinged Flap. .RJ & M. No. 1095, British A.R.C., 1927.
Wenzinger, Carl J., and Harris, Thomas A. : Wind- Tunnel Investigation of en N. A.C. A. 23012 Airfoil with Various Arrangements of Slotted Flaps.
NACA Rep. No. 664, 1939.
4. purser, Paul E., end Riebe, John h!.: Wind-Tunnel Investigation of Control-Surface Characteristics.
xv - Various Contour Modifications of a 0.30- Airfoil-Chord Plein Flap on an NACA 66 (215 )-014 Airfoil. NACA ACR No.
3=0, 1943= Purser, Paul E., and Turner, Tnomes R.: Wind-Tunnel 5~ Investigation of Perforated Split Flaps for Use as Dive Brakes on a Rectenguler NACA 23012 Airfoil.
NACA ACR, July 1941.
6.
Leitone, Edmund V.: An Investiq~tion of the High- Speed Lateral-Control Characteristics of a S~oiler.
NACA ACR NO. 4c23, 1944.
Short al, J. A.: Effect of Retractable-Sooiler Location on Rolling- and Yawing-Moment Coefficients.
NACA TN No. 499, 1934.
8. Purser, Paul E., end Johnson, Harold S.: Effects of Trailing-Edge Modifications on Pitching-Moment Characteristics of Airfoils. NACA CB No. 4130, 1944.
a ,. Wenzinger, Carl J., and Bamber, W,illard J.: . Wind- - Tunnel Tests of Three Lateral-Control Devices in ,,.
Combination with a Full-Span Slotted Flap on an N.A.C.A. 23012 Arfoil. NLCA Tl?No. 659, ~938.
iii — ,,, . . . ..- . . ..— - I n l 1 II I NACA AC:iNo. L5C24a” 10. Weick, Fred E., and Jones, Robert T.: R~sume’ and Analysis of M. A.C.A. Lateral Control Research.
NNA Rep. NO. 605, 1937.
11. Pearson, Henry A., end Anderson, Raymond F.: Calculation of the Aerod~amic Characteristics of Tapered Wings with Partial-span Flaps. NACA Rep.
HQW 665, 1539w 12. Goldsteh, S., and Young, A. D,: The Linear Perturbation Theory of Compressible Flow, with Applications to ‘Wind-Tunnel Interference.
R. & M. No, 1909, British A.R.C., 1943.
, NACA ACR No. ‘L5C24a Fig. 1
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Plain aileron .
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COMMllTEE FORAERONAUTICS ~
fi’u~e /.- TheNAC~ 230/2 airfoil with var~ous
types Of ui/eron ond w)th u slotted fl~p.
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Airfoil Reference Air)toi/ Reference /
o WACA W09 q NACA 0009 ~~published
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v li~C/I 66(2/5)-2/6 U~yblfshed Q NACA 230/2 3 and presenz’ ~epo~~ ~ tKAC~ 66(2?L0-0~ + A A4ACA 66series A cjqf~ y ~ NACA 66(2M5)-#6 Unpub/is~~d z o .
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prduce d by pldn sedeu’ f/q~s md by s~oi~ers. M=W to Q3.
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fitiite-spdn di7ta 00 J .2 .3 .+ .5 .6 .7 .8 in reference 6 MQCb number, M NATIONAL ADVISORY COMMllTEE FORAERONAUTICS figure +.- Effec+ of Much nwnk on the ~I?chIh9 moments produced Q .-
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