increasing the lift of airplane wings by flaps

8/2/2019 Increasing the Lift of Airplane Wings by Flaps

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2ECHN1 CAL MEIIOPLANDUXSADV1 SORY C01~MITTIH3 FOR AERONAUTICS

A S.IMPLE METHOD IOR INCREASING THE TIFT OF AIRPLANfi WINGS BY MEANS OF FLAPS

By Eugen Gruschwitz emd Os?.mr Schrenk

Z~itscnrift f& Flu&techqik un; Motorluftschiffahrtvol. 23, $~os 20, octo~er 28, 3.922.Verlag Won R, (pdenboure, Munc~en SDI~ Berlin

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~Illlllllllllllwmimillilllllllllllll31176014373790 !..NATIONAL ADVISORY COMMITTEE

TECHNICAL MEMORANDUM

FOR AERONAUTICS

NO, 714,.,

A SIMPLE METHOD FOR INCREASING THE LIFT OF AIRPLANEWINGS BY MEANS OF FLAPS*

By Eugen Gruschwitz and Oskar Schrenk

Aerodynamic considerations led us, not long ago, to in-vestigate a device which seemed to promise a contributionto the problem of reducing the landing speed of an airplane. ~-~ We have subsequently learned that similar devices had al-ready been proposed and investigated by others, but it seemsadvisable, nevertheless, to report our resclts, if only to .call attention to this hitherto hardly considered possibil-ity and to present it for discussion. Our results are bet-ter than those hitherto obtained, and the line of reasoningon which our investigation was based, along with the physi-cal explanation, also affords, under certain conditions, theopportunity for further improvement or for similar applica-tions. The problem is to create, in landing, a region ofturbulence on the lower side of the wing near the trailing *--edge by some obstacle to the air flow.The devices tested by us consitited of flaps of vary-Ing chord and position (fig. 1), the chord s being equalto the distance of the pivot from the trailing edge. The

three flaps tested had chords respectively equal to 5, 10and 20 percent of the wing chord, the flap angle ~ being ,set at 30, 60, 90 and 120 degrees.Figures 2 to 4 show the results of.the testsin the,.large wind tunnel. The 5-percent flap (figs. 2a to 2c) in-creased the maximum Ca $rom about 1.25 to 1.73 with flap

angles ~ = 60 and 90. The 10-percent flap yi,e:l.ded Ca =2 for @ = 60; Ca = 2.06 for P =,90?; and appiiac~ably lessfor P = 120. At the saine time the wing served as an airbrake. The additional profile drag for @ = 90 was about------------.-- -.*ueber. eine einfa.c.he M$glichkeit. z.ur Auftriebser,h~hp.n.g.vonTragfl&geln.n Zeitschrift f&r Ilugtechnik und Motorluft-schiffahrt, vol. 23, no. 20, October 28, 1932, pp. 597-601.Abstract of lecture by Oskar Schrenk at the twenty-firstannual meeting of the R.G.L. (Wissenschaftliche Gesellschaftf~r Luftfahrt), Berlin, 1932.

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2. N.,A.C.A. Technical Memorandum No. 714

*

5 times and ,for ~ = 60 about 2.5 times the profile dragof the smooth wing. The 20-percent flap yielded Ca = 2.18for P = 90 and Ca = 2.15 for ~ = 60. The single plot-ted point for Ca = 2.22 (fig. 2c) was obtained with thesame arrangement, but with the addition of small terminaldisks at both ends of the deflected flap, which created,fore and aft of the flap, pressure conditions less affectedby the edge effects.

It Pas to be expected that the moment lines would beshifted greatly toward the right ,(,figs.3a to 3c). Thisdisplacement is not so detrimental, however, as would ap-pear at first glance, because the. flap was used only forhigh Ca values, at which the c.p. falls in the region oc-cupied by it in normal flight. In designing an airplane,these conditions could also be influenced, within certainlimits, by the choice of a basic profile with suitable lo-cation of the cm line .

.The maximum lifts of all the devices tested occur atabout the same angle of attack as the maximum lift forP = 00 (figs. 4a to 4C). Hence the landing,.angl~ of at-tack with increas.ed ca remains within the normal limits.Moreover, flaps can be used on-the inner side of the aile-rons without causing any trouble with the angle of attack.

The reported c values were obtained ..withthe wingraised, (i.e. with i~creased al~gle of attack) while itstood ,in the wind. Starting with the condition of detachedflow,. we investigated the behavior of the wing in two cases,namely, for ,s/t = 0.05 and G.20 with the flap position$ = .90 and found that the flow already adhered on the up-per side of the wing at 0.50 below the value correspondi-ng to the C~ m ax of the raided wing.

~he considerable lift increase wa~ due in lesser de-gree to the dynamic effect on the lower side of the wing,but more to the. conditions on the upper side, and indeed through two concurrent circumstances:

1. A. negative pressure prevails in the turbulent re-tion behind the obstacle, which is simultaneously the ter-minal p~essure at the trailing e.dgeof the upper side. Thepressure level of the whole upper side sinks simultaneous-ly with the terminal pressure on the same side, just asreported by J. Ackeret years ago for wings with cutwwaytrailing edges (reference 1).


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N.A. C.A. ~edhniceil M&inorandum No. 714 3

2. With the lowering of the tsrninal pressure, notonly the whole pressure level is lowered~ but the pressureincrease, which the flow can withstand with respect to theboundary layer, also becomes steeper and, on the whole, ,,greater. The following consideration led us to this con-elusion and consequently to the investigation. The- pres-sure increasu which can be overcome in a retarded flow, de-pends on the processes in the boundary layer. This is de-termined by the velocity and pressure variation along thesurface. In order to estimate the attainable pressure in-crease, it must be referred to a standard dyuamic pressure.There is little sense in using the flight dynamic pressureremote from the wing, since the flow of the boundary layeris not affected by it. It would. be better to choose, as thephysically sensible reference pressure, the pressure at anypoint of the surface, e.g., that at the trailing edge of thewing. A rough-i approximation can then be nade on the assump-tion that the ratio between the pressure increase and refer-ence pressure has a constant value. If, for example, thepressure at the trailing edge is doubled by any device, thepressure increase itself, or, which amounts to the samething, the dynamic Fressure at every -point of the. wing isdoubled. Any lowering of the 6tatic pressure at the trail- . ;.ing edge, as caused by the turbulent region, means, accord-ing to the law of Berngulli, an increase of the same amount ,,,;in the dynamic pressure. i{.. ,

This line of reasoning is justified hy figure 5, in iwhich two pressure distributions a.t.outwings (converted to \unit dynamic preseure) are plotted, which wererfecently Imeasured in connection with other pressure-distribution ,i.$.ests. The curve with hatched inclosure was obtained with-out , and the other with a d,e~le;cted flap. In both casesthe conditions, are thohe occurring just %efore . the Be~&~rA-tion. of the .f,low. On the whoie, they agree vefy well withthe atiove,-m,ent$oned consideration, though the suc,tlon pointat the leading edge issomewhat large~and the p:resBure inthe middle of the wing is souewhat.lo%ek than wa.isestinated. 1.,,..

A similar line of reasoqing was employed by ProfessorBetz ip 1922 for explaining the phenomena of slotted wings(reference 2) . The leading-edge slat of a slotted-wingsystem is in the negative-pressiire-regiont-theleadingedge of the main wing, so that the ne~ative-~.ressure levelof the leading-edge slat can be ~li~herand the pressurerise steeper. . .A(: Our reasoning shows that the Ca max values attainable

~ by like devices a r e approximately pro~ortional for differ-


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,4 N.A.C. A. Technical Memorandum No. 714

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ent profile shapes to the Ca Da= values of the basic Tro-files. This probably holds good also for the ease whenthe ca max of the basic ~;rofile kas already been raidedartificially by ~ther means. This Froyorti~nality makesthe application of our Frinciple, combined with slottedwings or the removal of the boundary lay,er by suction,seem very promising.

As already mentioned, we learned subsequently of sev-eral experiments with flaps which could be let down on thelower side of the wing. Two or three years ago experimentswere tried in Japan with a f~ap which could be let down bya small angle (reference 3). The Ca values reached aprox-imate$ly 1.8 Almost the same device was investigated inthe Gbttingen laboratory in 1.923,f,ora private firm, ithaving been suggested by experiments with wings with cut-away trailing edges. The question was not investigated.further at that time.

Moreover, Professor Betz, unbeknown to us, has sug-gested, in an article not yet published, another device forincreasing the lift according to the same principle. Thisconsisted in using the trailing edge of a wing attached tothe main part as a rotatable body after the manner of anaileron and capable of being deflected downward by an angleof 90 degrees. For this device also, we recently made afew lift determinations with the aid of pressure-distribu-tion meastirements. Figure 6 shows the result, along withthe results obtained by the same nethod with the smoothwing and with the t,,,en-percentflap deflected downward. Thetwo curves for the.wings with cutaway trailing edges differby the flap chord s. The lift ialues given here are allsomew-nat hiGher, as local values in a mean wing sectionsthan the previously ehown results corresp~nding to polarmaa.surementbp. If one is interested more in the aerodynamiceffect than in tile practical application to flight, he eaualso refer the Ca values of the cutaway wi=gs to the short-ened chord instead of to the original chord and th~s obtainthe ca max values indicated by the dash lines in figure 6,the upper one of which is e-specially high, owlm probablytti the very great flap chord.

Lastly we triecl a few check tests on negative-pressuredevelopment by obstacles with th[: following results.1. The shape of the obstacle is not very importarit asregards the effect. The important thing is the development

of a turbulent wake.


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N.A, C.A. Tec~nical Memorandum Nc. 714 5

2. It is likewise unimportant for the negative-pres-sure .d.eyqlop~ent,,as to whether there is an air space be-tweenthe main wing and ihb obstacle.3* The maximum negative pressures are not Inmediat+ly behind the obstacle, but from one to two widths of theobstacle downstream. Hence one has a rather free hand inits application and can, for exa~ple, apply it at the rearspar and, in any case, forward of the aileroa.It is obvious to us that a series of technical and.structural problems still requires elucidation. Among themost important are questions of stability and of the effi-cacy of the control surfaces. We considered it our task

first to establish a few aerodynamic facts and we thoughtit worth while to report our results and call attention tothese possibilities.

Translation by Dwight M. Miner,National AdvisGry Committeefor Aeronautics.

REFERENCES

1. Ackeret, J.: Vorl. Mi\}eilungen der A.V.A..~ noo 291924.2. Betz, A.: Die If.irlymgsweise von unterte51ten Flugel-profilen. Berichte und Abhlmdlungen der W.G.L.,no; 6, January 1922.3. Nods, Tetsuwo: On the, Variable Section AqrofoilllHira?ri-Baneltand its Merit in the Practical Applic-ation. Joflzr.S.MiE,, (foreign.edition)! VO1. 33,

no. 151, .Tokyo, 1930..,.,.

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N.A.C.A. Technic@ hiemoranihuio. 714

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Figure l.-Profile of wing testeiiwith flap.

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N.A.C.A. Technical Memorandum Ho. 714

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N.A.C.A. Technical Memorandum No. 714 Figs. 4c,5,6

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increasing the lift of airplane wings by flaps

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