1、l I . 2 - TECHNICAL NOTES -A.- i-e IS . LT. NATIONAL ADVISORY COMMITTEE FOR AEROltAUTIcS ; i I* - No. 663 ._- _ _._._ _._. THE EFFEcTs OF FARTIAL-SPAa PtiIll FL. -A section and with tip-section flaps vere tested. - The results showed that the aerodynamic characteris- tics of partial-span plain flaps
2、 nere, in general, sim- ilar to those of split flaps of the sane span, bit thank the lift and the drag nere loss for the ning with plaiti .- flaps than for the wing vith split flaps of conparable size. For the rectangular wing with center-section plain - flaps, the maximum lift and the lift-drag rat
3、io at %5i- null lift vere greater,and the drag at naxfnun lift nas less than for the wing vith tip-section plain flaps 5f -.- the same size. The maximum lift of the tapered nfng var- i- -. ied in the same nanner as that of the rectangular wing but the drag and the lfft-drag-ratfo relationships mere
4、opposite. IBTRODUCTIOU - - Ilany arrangements of ning flaps have been tested in wina tunnels and in flight. Partial-span flaps are em- ployed in nearly all cases so that part of the trailing -A edge cali be used for lateral control, and frequently a section must be cut out at the center to allow for
5、hthe s fuselage. L Bind-tunnel tests of partial-span split flaps on rec- tangular and tapered rvings ha-re been reported in refer- ences 1 and 2 The present investigation deais with simi- lar arrangements of plain flaps. ._ _- _._. .-A -. - I: _ .- - - =:A -w Provided by IHSNot for ResaleNo reproduc
6、tion or networking permitted without license from IHS-,-,-2 K.A.C.A. Technical ITote Uo. 663 . APPARATUS AWD TESTS I frodel8. - The models used in these tests are Clark Y mings of laminated mahogany, each model having a span of 60 inches and an aspect ratio of 6. One airfoil is rec- tangular in plan
7、 form (fig. 1) and the other is tapered 5:l (fig. 2). For the tapered wing, the Clark Y profile was used at all sections along the span and the maximum ordinates of all sections were in a horizontal plane on the upper surface, The chord of the flaps is 20 percent of the wing chord at any longitudina
8、l section for both wrings. The flaps mere deflected about the axes shown, the anglos bekng measured in a.plane normal to the axis of dofloc- tion. The gaps between flap and wing nero sealed for all tests. The flaps t-rere cut into sections to form ton flaps of equal span. Rind tunnel.- II- The model
9、s fvere mounted on the stand- ard force-test tripod in the B.A.C.A. 7-by lo-foot closed- throat mind tunnel, which is described in detail in ref- erence 3. 1 L . Tests.- The.tests were made at a dynamic pressure of 16.37 pounds per square foot, corresponding to an air speed of about 80 miles per hou
10、r at standard sea-level conditions. Tests mere nade with center-section and tip-section flaps 20, 40, 60, and 80 percent of the span deflected 60 - and with full-span flaps neutral and deflo_cted_60. The angles of attack covered a range from approximately -14 _ to 2o“, which included zero and maximu
11、m lifts. r RJZSULTS AITD DISCUSSIOB . Coefficients : .I, The results are given in the.forn of absolute cbeffi- cients of lift, drag, and p,itchfng moment. F Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-3i;A.C.A. Technical ZIote 20. 663 -3 where Cm
12、( M . a.c. lo = s;crs L is wing lift. D, wing drag. . - - - M, pitching moment about aerodynamic center of plain wing. 9, dynamic pressure. - . - - - s, wing area. C9 mean mfng chord. The data have been corrected for the effects of the wind-tunnel j.et boundaries to aspect ratio 6 in free air. Iiect
13、angular IPing Curves of lift, drag, and-center-of-pressufelocation for the rectangular wing with center-section flaps are given in figure 3 and curves of pitching moment in fi CLmax decrease with. increase of flap span except for the wing w$th a center.-section flap of span less than 0.20b. Lower va
14、lues of L/D are obtained with the tip-section flaps than with the cent-er- section flap. The foregoing discussion indicates that, for landing, the advantage of the higher maximum lift obtained with fho canter-section flap might bo somomhat offset by the loner drag, and that a slower but flatter land
15、ing nould bo at- tainod with the centor-section flap. The curves of 11ft, drag, and center-of-prossure lo- CatLon for the wings with partial-span plain flap8 show the same general characteristics as those for a wing with partial-span split flaps (reference 1). The maximum lift and the drag at maximu
16、m lift for the ning with the plain flaps are less and the stall occurs at a somewhat lorrer angle of attack than for the wing rralth split flaps of the same size. (The data of reference 1 were corrected to aspect ratio 6 in free air for these comparisons.) r The difference in maximum lift for the wi
17、ngs with the center-sectfon flap or with tip-section flaps of equal span is less for plain flaps than for split flaps. It should also be noted that, whereas the drag is the same for the wings fitted with either tip-section split flaps or the center-section split flap, the wing with tip- section plai
18、n flaps,hns higher drag than the one with,thQ cantor-section flap;, Tapered tiing Plots of lift, drag, and center-of-pressure location * for the 5:1 tapered wing mith a center-section plain flap are given in ,figure 8, and pitching mononts are plotted in figure- 9. Plots of the same coofficionts for
19、 the ta- pored wing with tip-section flaps are,givQn in figurQs 10 .and 11. It mill be noticed that the values of lift, drag, and pitching-moment coefficients increase as the flap span is inCrQ the reverse was true Tor the rectan- gular ning. 1. The changes in lift and drug vith flap span for both r
20、ectangular and tapered nings having partial-span _ plain flaps vere similar to-those for corresponding wings -, with partial-span split flaps. 2. The maximum lift and the drag at maximum lift of a rectangular and of a tapered Clark Y wing havfng partial- - span flaps vere less for the sing nith _ula
21、 Yo. 505, iT.A.C.A., 193,4. 3. Venzinger, Carl J., and Harris, Thomas A.: Tests of an N.A.C.A. 23012 Airfoil with Various Arrangements of Slotted Flaps in the Closed-Throat 7- by IO-Foot Vfnd Tunnel. T.R. iTo. 1938. (to bo published), IT,A.C.A., . . . . Provided by IHSNot for ResaleNo reproduction o
22、r networking permitted without license from IHS-,-,-N.A.C.A. Technical Note No. 663 Fig.i - - .$ _. - _ . B b - 2 cd u” - - , Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-X.A.C.A. Technical Not2 No. 663 R Fig. 2 I . Provided by IHSNot for ResaleNo
23、 reproduction or networking permitted without license from IHS-,-,-Ii A.O.A. TdlIIiC81 Mote HO. 883 Flsp aprrr Flap -1.00 b -.40 wy -.60 X- -a0 PGPy nap apan :eo - - -.40 A- - x- .a0 w-b- .a +-0 c . ” I! ! ! I I I 1. ! I ! ! (! ! I I I I ! I I 1 2 I v I .I I I I I I I I id?! I I #j 7a bgle of attsok
24、,a,dag. Figure 6.- vuirtion of lift,drag,rod center of seeurr rith -19 of &tack. Figure lo.- Aqle of rttack,X,deg. Vuirtion of lift.dmg,snd center of preeourm with angle of rttrak. Cemter-eeotion plain flap on the tegerad OlukY Tip seotlon pip flapa on the tapered Cluk Y wtw. 6 6Q”. dng. 6 60 . Prov
25、ided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-H.A.C.A. Teohnloal Mote Ho. 663 Rgr.S,ll -_ b . -P l w o-l.00 8 F1r rpan - .po v - -.00 - .a a-.a +-em- 0 .l 0 +. . 5 x. . -.l -A- b -+ - b-c- + - -I-+-+-& _ ilk a.+- -+-+-l-x+ _c + -+ . -.3 -16 -12 -8 4 0
26、4 8 1% 16 a0 Angle of abtaok,a,deg. Figure S.- Variation of pitching-mment ooef:icient rlth le of rttaok. ontsr-aeotlon plain flape on the taperd Olrrk Y wing. 6f - 3. - - .a0 x-.a0 - - - -60 +-m-e- 0 .l 2 6 . % 0. ? . -a 4 Angle :f .ttaok,:deg. 8 ia- 18 - x- - - Figuxo ll.- vuirt1oB Of pitoblcmt 000ffc1mlt +:tb plrin flaps on the twmred Okrk I wing. 6f = 8 %. 10 of l ttrok Tlp-seotlcm .- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
