NASA NACA-TN-759-1940 Pressure-distribution investigation of an N A C A 0009 airfoil with a 30-percent-chord plain flap and three tabs《带有30%弦的普通襟翼和三个标签的NACA 0009机翼的压力分布研究》.pdf

《NASA NACA-TN-759-1940 Pressure-distribution investigation of an N A C A 0009 airfoil with a 30-percent-chord plain flap and three tabs《带有30%弦的普通襟翼和三个标签的NACA 0009机翼的压力分布研究》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TN-759-1940 Pressure-distribution investigation of an N A C A 0009 airfoil with a 30-percent-chord plain flap and three tabs《带有30%弦的普通襟翼和三个标签的NACA 0009机翼的压力分布研究》.pdf（44页珍藏版）》请在麦多课文档分享上搜索。

1、. .1: 111111111111111111111111111111111111. 3 l76 0009 2813,-. +*. .4a71It.-. -.,. .:Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. -) ,.,/.1.-NATIONAL ADVISORY COMMITT_IIEFOR AERONAUTICS.7- - L. ,TECHNICAL NOTE NO. ?59v-.-.“PRESSURE-DISTRIBUTION

2、INVESTIGATION OF AN N,A,C,A. 0009t AIRFOIL WITH A 30-PERCENT-CHORD PLAIN FLAP“+,%.AND THREE TABS, .SUMMARY.-Pressure-distribution tests of an N.A.C.A. O?09 air-foil with a 3!3-percent-chord plain flap and three plaintabs, having chords 10, 20, and 30 percent of the flap chord, were made in the i?,A.

3、C.A.4- 3Y 6-foot verticaltunnel. The purpose of these tests was to continue an in-vestigation to supply structural and aerodynamic sectiondata that may be applied to the design of horizontal andvertical tail surfaces.The results are presented as diagrams of resultant“pressures and of resultant-yress

4、ure increments for theairfoil with the flap and the 25-percent-ckord tab, In-crements of normal-force and hine-moment coefficients for .the airfoil, the flap, and the three. tabs are S.lSO given.At all unstalled fiap and tah deflections, the exper- imental distributions agree well with those calcula

5、ted byan analytical method. The agreement is poor, however,when the stalled or the unstalled condition of the flapor the tab deflected alone was changed to an unstalled orstallqd condition by the simultaneous deflection of boththe flap and the tab. INTRODUCTION. The trailing-edge tab has proved to b

6、e an effectivea71 “ device in reducing the excessive control forces resltitig .from the recent increasesin size and speed of airplanes.a Although s,number of investigations have bepn conductedto determine the characteristics of the different factorsProvided by IHSNot for ResaleNo reproduction or net

7、working permitted without license from IHS-,-,-2 N,A. C.A. Technical Note No, 759affecting control surfaces (references 1, 2, and %), nodata giving the aerodynamic section characteristics of athin airfoil as affected by flaps and tabs seen.ed to heavailable that would he applicable to tail-eurface d

8、esi:;n.An investigation was therefore undertaken to supply in-formation applicable to the aerodynamic and the structuraldesign of tall surfaces with ta%s. The first pert of thisinvestigation comprised pressure-distribution tests madeof an N.A,C,A. 0009 airfoil with a 50-percent-chord plainflap and t

9、hree plain tabs; the, results o: these tests arcreported in reference 4.The results reported herein were obtained from pres-sure-distribution determinations over one section of anN.A,C,A. 0009 airfoil with a 3!3-percent-chord plain flapand with plain t”os 10, 20, and 39 percent of the flapchord. Fro

10、m the data obtained, normal-force and pitchin#-moment coefficients were calculated for the airfoil sec-tion complete with the.flap ,dudthe various tabs. Thenormal-force and the hinge-moment coefficients for theflap withthe different tabs and.for the tabs separatelywere also determined. “,.-.“ApAAfj

11、AND Ts.TcjModel and. Test Installation#The tests vvere made in the. N.A.C.A, vertical wind tun-nel. The test section of this tunnel has been convertedfrom the original open, circular, 5-foot-diameter jet (ref-erence 5) to a closed, rectangular, 4.- Yy -foot throatshown in fiure 1.The rectangular 3-f

12、oot-chord by 4-foot-s”pan model w,asmade of laminated mahogany to the N,AC.A. 0009 profile.It was equipped with a plain flap having a chord 30 per-cent of the airfoil chord, c, and with three seriallyhinyed plain tabs having chords 13, 2(I, and Zg percent ofthe flap chord, Cf, as shown in fiqure 2.

13、During tests,all flap and tab gaps were sealed with plasticize and cel-lulose tape to prevent air leakaqe at the hic+es. Theradius of curvature at the hinge for both the flap and thet.bs was approximately one-half the airfoil thickness atthe respective hinge positions. ,A single chordwise row of pre

14、ssure orifices was built.bProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-N.A, C.A, Technical Note No. 759 3into the upper and the lower surfaces of the aiifoil, theflap, and the ta%s at the midspan, The orifice positionsare shown in fiqure 3,is the

15、model completely spanned the test section,two-dimensional flow was approximated, The model was at-tached to the balance frame by means OT toque tubes, whichextended through the sides of the tunnel and kere rotatedby a calibrated electric drive. to set the angle of attack,The flap and the ta% anqles

16、were set inside the tunnel byvarying the position of small lever arms on the movablesurfaces, .,The rubber tubes from the pressure or,ifices were%rought out of the model at one end through the torquetube and the tunnel wall to a photographically recordingb multiple-tu%”e manometer,.“.4TestsT,ess wer

17、e “conducted at an effective Reyolds Ntimberof approximately 7,4-10,000, (Effective Reynold5 “Number =test Rejnolds Number x turbulence factor. The turbulencefactor of the 4- by 6-foot vertical tunnelis 1.93.) Thetunnel was operated at,an average dynamic pressure of 10.8pounds per square foot, corre

18、sponding to an air speed ofabout 65 miles,fier hour at standard sea-level conditions.For direc”tcompaisonwith the results presented inreference 4, the tests were made a% angles of attack from-14*O to 10*0 at intervals of 5, The mctiel was testedwith the 30-percent-chord plain flap deflected 0, 5, 10

19、,200, 3(30, and 45:9Throughout the entire angle-of-attack ,range for each flap deflection, the three tabs were de-flected 0 10, *20, and +30,. , ,. . . . . , RESULTS $ , Presentation of D“ataThe results of the distribution of pressures aregiven in the form of diagrams of resultant pressures andresul

20、t:ant-pressure increments, which represent changes inresultant-pressure distribution caused by a change in an-gle of any one”part Qr any combination of the component.“Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 N,A. G.A, Technics.1 ,Note No; 75

21、9parts of the airfoil. The resultant normal pressure atany point along the chord. line of the airfoil was deter-mined by taking the algebraic difference of the pressuresnormal to the surface of the airfoil at that point. Alldiagrams of resultant pressures or resultaht-pressure in-crements of the air

22、foil, flap, and tab combination aroplotted as pressure coefficients, P, or as AP, whereP-POP= .- ., !Istatic _prqssure at a point on airfoil.static preesure in free air stream.dynamic pressure of free air stream.Resultant-pressure diagrams are given for the basicsection (i.ec, flap and tab neutral)

23、in fi%ure 4. Theresultant-pressure diagram for any other condition may %eobtained hy adding to the basic diaqram the resu.ltant-pressure-incremelnt diagram (figs. 5 to 10 for the partic-ular condition.The large quantity of data prohibited the inclusionof all the resultant-pressure-i ncrem.ent diaram

24、s, Onlythe diagrams for tab deflections of 0 and *30 for the0.20cf tab, which was considered to be an average size,are presented- Vdlues of angle of attack were selected torepresent the followinrProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-*-4N,A.

25、 C.A. Technical Note No. 759 13Two notable differences occur in this comparison be-tween data from tests in the 4- by 6-foot vertical tunnel”and in the full-scale tunnels First, the value of acn/aaof 0.095 obtained from the present tests and those reportedin reference 4, although not in accord with

26、the value of0.083 obtained in the tests reported in refere.ce 3, is inagreement with the value of 0,095 from other full-scale-tunnel tests reported in reference 7. The other differ-ence occurs in comparisons of values of 3chf/86f andch+. This difference is in accordance with expectedresults. As poin

27、ted out previously, the section charac- teristics of the flap and the tab herein reported were notcorrected for tunnel effects. In addition, the gaps be-tween the flap and the tab were sealed. Both of these factors would cause an increase in these slopes. .* .CONCLUDING REMARKS. .,.,Aerodynamic sect

28、ion characteri,stics,and resultant-pressure distibutions have %een presented for the N.A.C,A.0009 airfoil ,with a Go-percent-chord flap and three plaintabs having chords 10, 20, and 30 percent of the”,flapchord.For all unstalled flap and ta% deflections, the ex-primental,a,n,d the calculated distrib

29、utions of resultant-pressure increments are in good agreement. .The results of the analytical method of calculatinthe resultant-pressure distribution apd the experimentalresults are not in agreement for the cases riwhich thestalled or the unstaled condition of the flap or the tabdeflected. alone was

30、 changed by the simultaneous deflec-tion of the flap and the tat. This poor agreement betweenthe experi,mential an;the calculated results,is attributedto the fact that the coefficient increments for thesecritical conditions are not additive, ,as they must be toobtain good agreement. In the applicati

31、on of these data for design purposes,it should he remembered that,for all cases, gaps werecompletely sealed, resulting in higher peak pressures atthe hinge axes and in higher hinge-moment and normal-forcecoefficients than WOUlti have %een obtained with unsealedProvided by IHSNot for ResaleNo reprodu

32、ction or networking permitted without license from IHS-,-,-.14 N.AC.A. Technical Note No. 759a71gaps. It should also be noted that only the values of thenormal-force coefficients were corrected for tunnel effects, F-Langley Memorial Aeronautical Laboratory,National Advisory Committee for Aeronautics

33、,Langley Field, Vs., March 27, 1940.REFERENCES1.2.3.4.,5.6,Wenzinger, Carl J,: Pressure Distrilmtion over anAirfoil Section with a Flap and Tab. TR, NO. 574,N.A,C.A. , 1936. .Earris, Thomas A.: Reduction of Hinge Moments of Air- *plane Control Surfaces bayTabs. T.R, No, 528,N,A.C.A. , 193!5, “AGoett

34、, Harry J., and Reeder, J. P, : Effects of Eleva-tor Nose Shape, Gap, Balance, and Ta.%s on the Aero-dynamic Characteristics of a Horizontal. Tail Sur-face. T.R. No. 675, N.A.C.A. , 1939.Street, William G., and Ames, Milton B., Jr.: Pressure-Distribution Investigation of an N.A.C.A. 0009 Air-foil wi

35、th a 50-Percent-Chord Plain Flap and ThreeTabs, T.N. No. 734, N.A.C.A., 1939.Wenzinger, Carl J., and Harris, Thomas A.: The Verti- .cal Wind Tunnel of the National Advisory Committeefor Aeronautics. T.R. No, 587, K.A.C,A., 1931, .,Allen, H. Julian: Calculation of the Chordwise LoadDistribution over

36、Airfoil Sections with Plain,Split, or Serially Hinged Trailing-Edge Flaps.T*R* No. 634, N.A.C.A. , 1938.Croett, Harry J., and Bullivant, W. Kenneth: Tests ofN.A,C.A, 0009, 0012, and 0018 Airfoils in the Full-Scale Tunnel. T.R. No, 647, N,A.C.A. , 1938. . .,.Provided by IHSNot for ResaleNo reproducti

37、on or networking permitted without license from IHS-,-,-B. A.C.A. ?oohic81 SOte ik. 759 FB.1.2,3.,w.1 (1-(.SqIsb.-”,.,1.1 ,:1.8-1.6 J :mll%;!efi1,/.4 t 1./.o :.3/ /-;-l;Eg -.8 QJ-.6$ -Lo -.7-1.2 :8-/.4 -.9-30 -20 -10 0 /0 20 30 W5v -20 -/0 o 10 20 30Tab deflection, 6, , dega= -4 1/20Figure 14 . .Tab

38、 deflection, 6, , dega = 1/20 /*-i. .-i. -.55 $EQJ-, c -.6&-1&-.7-1. -.8-1. -.9-30 -20 -10 0 /0 20 30 (f )-30 -20 -lo 0 10 20 30Tffb deflection, C$t, deg . 10 1/20 -Yu.1 1:!Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.I. 4z“*.0.1-Cn.,%-$aIncremen

39、f of airfoil Se&I-On normal-force coefficient, A cmProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. .4*.c“.N.A.C.A. Technical Note No. 759 Fig. 15d,e,fFlqodef/ecfion, G,deg(d) u = 1/2 (e) U= 5 1/2 (f) u = 10 1/20Figure 15 concluded.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-