NASA NACA-TN-734-1939 Pressure-distribution investigation of an N A C A 0009 airfoil with a 50-percent-chord plain flap and three tabs《带有50%弦普通襟翼和三个标签的NACA 0009机翼的压力分布研究》.pdf

1、. . - - . . - - . .: Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-h .r NATIONAL ADVISORY COXMITTEE I- FOR AERONAUTICS - TECHEICAL NOTE NO. 734 - - . PRESSURE-DISTRIBUTION INVESTIGATION OF A!? N.A.C.A. 0009 AIRFOIL WITH A BO-PERCENT-CHORD PLAIN FLA

2、P AND TXREE TABS By Willfam G. Street and Milton B. Ames, Jr. SUMJARY Pressure-distribution tests.of an N.A.C.A. 0009 air- s foil with a 50-percent-chord plain flap and three plain - tabs, having chords 10, 20, and 30 percent of the flap chord, were made in the N.A.C.A. #- by 6-foot vertical tunnel.

3、 The tests supplied aerodynamfc section data that may be applied to the design of horizontal and vertical tail surfaces. . The results are presented as resultant-pressure dia- grams for the airfofl with the flap and the 20-percent=- - - chord tab. Plots are also given of increments of nbr-mZl- - for

4、ce and hinge-moment coefficients for the airfoil, the flap, and the three tabs. The experimental results and values computed by analytical methods are in good agree- _. Y ment for small flap and tab deflectfons. The results of the tests indicated that the effectiveness of all three tab sizes in redu

5、cing flap hinge moments decreased wi;th increasing flap deflection. - .- _- l INTRODUCTION - ,. . The recent increases in speed and. size of airplanes have produced excessive control fo,rces. - - At the preraent time, one of the most effective devices for r8dUCing these high control forces is the tr

6、aglinq-edge tab. (Sea refer- -z ences 1 and 2.) Previous investigations have indicated -* the effects of such factors as plan form, cut-cuts, 8nd . I plates, elevator nose shape, gap, and balance on the aero- I -. _ dynamic characteristics of some tail surfaces of.finite span (references 3 and 4). P

7、O available data applicable to tail-surface design, however, give the section aerody- _ -A (“ - _ - .-. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 N.A.C.A. TechnicalNo-to lVo. 734 namic characteristics of a thin airfoil as affected by flaps

8、and tabs. The present investigation was therefore undertaken to supply information that would be applicable for use in aerodynamic and stmFnctura1 design of tail sur- faces w?th tabs. The tests consisted of pressure-distribution measure- ments over one section of an N.A.C.A. 0009 airfoil with a 50-p

9、ercent-chord plaih flap and with plain tabs 10, 20, and 30 percent of the flap chord. From the data obtained, normal-force and pitching-moment coefficients were calcu- lated for the airfoil section complete with the flap and the various tabs. In addition, normal-force and hfnge- moment coefficfents

10、were calculated for the flap nith the different tabs and for the tabs alone. APPARATUS AND TESTS Mod-e1 and Test Installation The tests were conducted in the N.A.C.A. 4- by 6- f%ut vertical wind tunnel. This tunnel, originally a 5- foot open-jet vertical wind tunnel (reference 51, has recently been

11、modifie- - L - i p*p-Po 9 .- - .I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 M.A.G.A. Technical Bate No. 734 P is static pressure at a point on the aFrfoi1, c PO static pressure in the free air stream. 99 dynamic pressure of the free sir strea

12、m. In order to enable the designer to obtain the origi- nal resultant-pressure diagram for any condition, resultant- pressure diagrams are qiven for the basic section (i.e., flap and tab neutral, fig. 4) which, when added to the fn- crement diagrams (figs, 5 to 10) oive the desired result. Because t

13、he large quantity of data prohibited the fn- elusion of the resultant-pressure-increment diagrams for all tab sizes, only the diagrams for the 0.2Ocf tab, which was consfdered to be an averase size, are presented. Val- ues of angle of attack were also selected to re resent the negative angle-of-atta

14、ck condition, a = -9 ; P the LOW angle-of-attack condition, cz = 1/2O; and the PO 8 itive angle-of-attack condition, cx = 5*“. The value of 59 may seem rather low; but, with such a large-chord flap, the stall occurs at low angles of attack. The section characteristics of the airfoil, the flap, and t

15、he tab, as functions of flap and tab deflection, are also plotted as increments, which were obtained by deduct- ing the basic section coefficients from those for the sec- tion with the tab, the flap, or the combination deflected. The characteristics were obtained in each case by mechan- ical integra

16、tion of the original plotted pressure diacrams. Computations were made to determine the section coeffi- cients, which are defined as follows: c I cn n = yy , airfoil section normal-force coefficient. m rn = -, 9 c= -airfoil section pitching-moment coefft- cient about quarter-chord point of airfofl.

17、nf Cnf = -I flap section normal-force coefficient. c 4. c-f ch hf f = -:, flap section hfnse-moment coefficient. q Cf * Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-N.A.C.A. Technical Note No. 734 5 Cnt nt = -* tab section normal-force coefficient

18、. q ct Cht = - tab section hinpe-moment coefficient, 9 ct= where the n, m, nf hf nt ht and Cl Cf* Ct 1 forces and moments per unit saan are: normal force of airfoil section. , pitching moment of airfoil section about the . quarter-chord point. I normal force of flap section. hinge moment of flap sec

19、tion. - normal force of tab section. : hinge moment of tab section. chord of basic airfoil with flap and tab neutral. flap chord. tab chord. The subscript f refers to the flap with the tab; and the subscript t, to the tab alone. In figure 11, the integrated coefficients for the basic airfoil are plo

20、tted against angle of attack. Curves Sfving the increments for various tab and flap deflections are presented fn figures 12 to 20. . The effect of tab size and deflection on flap section hinqe-moment coefficient is shown in figure 21. A compar- _ .- ison of theoretical and experimental airfoil secti

21、on normal- _ - force and pitching-moment coefficients is given in ff-Sure 22. Precision . Since no air-flow-alinement tests were made in the tunnel for the test set-up used in-the investigation, the absolute angle of attack may be slicghtly-in error. In the final data, however, all angles of attack

22、were corrected Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 N.A.C;A. Technical Note No-6 734 for a misalinement of l/2*, which appeared to be present in the air flow. The relative angles of attack are accu- rate ttrwithin fO.lO. Tabs and flaps m

23、ere set_ tv th-e spec- ified angles to-within *l.O“. Errors in orifice pressures were within f2 percent except-at the nose and the bingo lo- cations on the upper surfqce, where the variation may be about +5 percent a t high angles of attack. The dynamic- pressure readings were accurate to within *l.

24、O- percent. Inasmuch as two-dimensional flow was approximated, the integrated results may be taken as section character- istics. Corrections for tunnel-wall effects (re.ference 8) were made only to the airfoil section normal-force coeffi- cients Cn. The value of den/da for the airfoil with flap and

25、tab neutral Agrees well with values e axis and did not exceed 0;2q. o“ * For the flap deflected 13 and the ta;6 deflected -30, and 30 (fig. 7), the maximum varfation of the cornput- ed from the experimental results is Q.4q at the airfoil .- nose. The greatest variation at the hinges fsabout 0.2q. An

26、 excellent correlation was obtained between the .computed. and the experimental results for 30 and 0 tab. deflections. The comparisons made in figures 8 and 9 generally check to within O.lq. . . 4 - The comparison shown in figure 8 of computed and ex- perimental load distributions for an angle of at

27、tack d-f 1/2O o“, with the flap deflected 2Qo and.the tab deflected 30, and -30 fndicates a good asreement for values ahea-d . of the hinge of the flap. Behind the-hinge axis, the- c“om- puted values are about (3.3q lower than the test -re-sulfs In figure 9, the computed and the experLUmental result

28、s are compared for the flap deflected 30 and the tab deflected 30, O”, ancnt/det and de momont became les8, as shown in figure 21. Increase -in tab chdriTa% the same - I. cn and 6t definitely decreased the flap hinge moments: the decrease in =hf was not proportional to bet decreased * -e with increa

29、se in tab chord. Comparisons with Theory In figure 22 is given a comparison of curves theoret- ically calculated, as outlfned in reference 1, with some representative curves from the present invest_ant pressures for various pressures for varipus angles qf attack wd various angles of attac r t i-i i

30、i i i -2 I I I I I t t I I If i i i . 0 I 11 -20 I I I I I i L ia PZnf CEd 8x3 . /W Figure 9.- Inorements of resultant pressures for various angles of attack and varioue deflections of a 0.2Ocf tab on a 0.50 plain flap deflected 300. Percsnf c 734 Fig. 12a,b,o . II I I I I r I 11 -1. * P.rn I I I I

31、I o&G7 -a 20 I.8 . E! - 9 d NPI 1 I t -+- -10- . - - -0 i I i -3y, -a i - I I i t I Cal I 1 I i : i 6 i * k. . i i i i i I .8 .4 0 i i i i i i I i i I ca, a =-14 l/i0 - (b) a =-9 l/20 (h) a =-A l./Z - Figure 12,a to f.- Inorents of airfoil aaotion normal-for08 and pitching-moment coefficient8 for va

32、rious deflection8 of a 0.500 plain flap and a 0.1Oof tab. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-N.A.C.A. Teohnioal Note No. 734 Fig. E!d,e,f .8 +- .I - .s 0 id) .8 VJ Figure 12 concluded. Provided by IHSNot for ResaleNo reproduction or netw

33、orking permitted without license from IHS-,-,-I . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . - L -. , . ! L . l Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo

34、 reproduction or networking permitted without license from IHS-,-,-t ii- / I / I.0 i? 8 .8 x F .9 ,8 .7 .6 Tab defkflbn, St, deg. a = -14 1/20 a=-9l/P rl.gtra 14, a ta f.- Inorunsnts of tab maotlon normal-foroe and hiw-moment aoatfloientafar rorioua daflrotiona of a 0.500 plain flapand a O.lOaf tab.

35、 0. I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-t.8 LO I I I I I I I I I I I4 .9 I I I -0 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, 1 , . I . Provided by IHSNot for ResaleNo reproduction

36、 or networking permitted without license from IHS-,-,-N.A.C.A. Technical Note No. 734 Fig. 15a,b,o . Figure l6, a to f.- Increments of airfoil se&ion normal-force and pitohing-mommt coefficient8 for . . various deflections of a 0.50 plain flap and a O.ZOOf tab. A Provided by IHSNot for ResaleNo repr

37、oduction or networking permitted without license from IHS-,-,-W.A.C.A. Technical Note Nd. 734 Fig. 15d,e,f . . 8 Y uu -I.2 o I _ F/i& dc?ncfn60s, a&g? 50 0 40 50 fg Gkfz&-?,*. (d) a = l./2O (0) a = 5 1/2O (f) a = 10 l/2O Figure 15 concluded, -. . - : 3 A- -=.z=zy .-. :- a- . . Provided by IHSNot for

38、 ResaleNo reproduction or networking permitted without license from IHS-,-,-, . * l I te2 xr.0 c .8 .s .6 E 84 Y 8 2 .2 % -0 f F -.2 .-A -.6 -.8 2.4 I I I I I 1 .6 I I I I I I I I I I I I I I I I I I I I I I ) I I I I I I 2.0 z 1 .41-7-rl-.- jb-H :6 -.s -,8 -1.0 0 ro 2o 30 40 50 (b) 0 IO 20 30 4o 50

39、 f-r ddkdi37, s, a+. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-f-.6 &.7 “.a -.I3 -3 -LO 0 ro 2u 30 40 50 Cc) to 20 .m 40 50 -LO 0 IO 20 30 40 50 Cd) 0 ro 20 30 40 50 -. c P Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-