1、C A S _ _LEz. C0_NATIONAL ADVISORY COMMITTEEFOR AERONAUTICSREPORT No. 574PRESSURE DISTRIBUTION OVER AN AIRFOILSECTION WITH A FLAP AND TABBy CARL $. WENZINGERFiLE OiOPYIhefilesif theNa_nal1936For sale by the Superintendent of Documents, washington, D. C. Price 10 centsSubscription price, $3 per yearP
2、rovided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REPORT No. 574PRESSURE DISTRIBUTION OVER AN AIRFOILSECTION WITH A FLAP AND TABBy CARL J. WENZINGERLangley
3、 Memorial Aeronautical Laboratory78301-36 IProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICSHEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.LABORATORIES, LANGLEY FIELD, VA.Created by act of Congress apl,_roved
4、March 3, 1915, for the supervision and direction of the scientificstudy of the problems of flight (U. S. Code, Title 50, See. 151). Its membership was increased to 15 byact approved March 2, 1929. The members are appointed by the President, and serve as such withoutcompensation.,0,SEPI S. AMES, Ph.D
5、., Chairman, CHAIILE$ A. LINDBERGlt, LL.D., !Baltimore, Md. New York City.DAVID W. T both normal-deflection. In addition, the variation of increments oJ force and hinge,moment coefficients were computedtab normal-Jorce and hinge-moment coefficients with tab for the flap section with tab and for the
6、tab sectiondeflection for a given flap setting was practically inde- alone.pendent of flap deflection. Comparisons of the theoretical APPARATUS AND TESTSwith the experimental forces and moments for the airfoil The N. A. C. A. 7- by 10-foot wind tunnel, in whichsection withflap and tab show that the
7、theory agrees fairly the tests were made, is described in reference 2. A half-well with experiment for small flap deflections with the span Clark Y airfoil (fig. 1) that had originally beentab neutral, but that the theory indicates much greater built for pressure-distribution tests of high-lift devi
8、ceseffects than are actually obtained when the flap and tab was used. The model was altered by installing at theare simultaneously deflected, tip a flap having a chord 30 percent of the airfoil chordand a span 40 percent of the half-span model. An insetINTRODUCTION tab was mounted at the trailing ed
9、ge of the flap, theA considerable number of airplanes are fitted with a tab size and location being selected as representativesmall flap on one or more of the movable control sur- of the average. The tab chord was 20 percent of thefaces. Such an auxiliary flap is ordinarily referred to flap chord an
10、d its span was 50 percent of the flap span.as a “tab“ and is usually set into the trailing edge of the The gaps between the flap and the airfoil and thosecontrol surface. When the tab is used to reduce the between the tab and the flap were sealed with plasticinehinge moments of a control surface, it
11、 is known as a for all tests.“balancing tab“; when used to trim the airplane in The airfoil, flap, and tab were all constructed of lam-place of an adjustable stabilizer or fin, it is referred to inated mahogany to within _0.010 inch of the specifiedas a “trimming tab.“ ordinates. A row of small orif
12、ices was installed in theThe chief aerodynamic characteristics of tabs are upper and lower surfaces at one chord section located atcovered in reference 1, which describes an investigation the center of the span of the flap and tab. (See fig. 1.)of a wing with serveral arrangements of ailerons and Th
13、is location was 20 percent of the semispan of thetabs, alone and in coniunction with other types of model inboard of the rectangular tip so that satisfac-balancing arrangements. In reference 1 data are also tory section characteristics could be obtained which1Provided by IHSNot for ResaleNo reproduc
14、tion or networking permitted without license from IHS-,-,-2 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAIJTICSwould be outside the influence of the usually high local RESULTStip pressures. The half-span model was set up in con- The results of the investigation, in their originaljunction with _ refl
15、ection plane at its inboard end, the form, consisted of pressme diagrams for the section asplane extending from top to bottom of the air stream tested at different angles of attack and for differentand some distance ahead of and behind the model, tab and flap deflections. In order to facilitate the
16、A multiple-tube alcohol manometer photographically interpretation and-application of these results, therecorded tile pressures on the airfoil section, pressure diagrams are presented in the form of “incre-Pressures were measured for flap settings of 0 , ment“ diagrams, which represent the changes iX
17、. pres-_%15, and _:30 with the tab neutral. With the sure distribution due to changes in the significant iflap neutral, pressures were measured for tab settings of variables. The pressure diagrams for the basic section:t: 10, 20 , and 30 . The pressures were then (i. e., neutral tab and flap) are al
18、so given so that themeasured for various combinations Of flap up with tab resultant diagram for any ease may be obtained bydown and of flap down with tab up. The angles of addition of the increment and the basic-section dia-attack used in the tests (-5 , 0, 10, and 15) covered grams. The principal a
19、dvantage of the incrementOr-/C-lees located/n fh/_, )laneCO.O0“ I Or/- I f,-om_ flee I L.,A_g_Td-do_2.ood-_.o_o/4.5oo0 H /K_25J 18.875et=/.20, K 18.000“ . _2P_L/9.s2sPlan v/ew of hG/f-spGf7 model _- /2.00“_ ) /9.87524. O0“ H Ald PUAU 8U CU OU EU FU GU HU JU KU MU,“OUiOUSealed,i,:. . plas/c/ne- 1?i!
20、/ “ .)DL EL FL OL HL JL /fL /YfLOLtotNL PL t20. O0“Set/lanai v/ew Show/rTg OF/f/de /coo/ions on airfoil, flap, and lab.FIGUBE1.-Clark Y airfoil with tab and flap arranged for pressure-distribution tests.approximately the range hom zero lift to maximum diagrams is that they may, by the principle of s
21、uper-lift. position, be applied to pressure diagrams for any otherAngles of attack and flap deflections were measured basic airfoil section, including the symmetrical section,with respect to the airfoil chord; tab deflections were that does not depart too greatly from the Clark Y see-measured with r
22、espect to the flap chord. Positive flap tion on which the tests were made. The diagrams ofor tab angles indicate a downward deflection with resultant-pressure distribution for the basic airfoil see-respect to the airfoil or flap chord. The tests were tion are given in figure 2. The increments of res
23、ultantmade at a dynamic pressure of 16.37 pounds per pressure for various tab and flap deflections are pre-square foot, eorrespondir N to an air speed of 80 miles sented in figures 3 to 6. The figures give the results forper ham under standard sea-level conditions. The a low-angle-of-attack conditio
24、n, a=0 , and for a high-average Reynolds Number was 1,220,000, based on angle-of-attack condition, c_=t5 .the airfoil chord of 20 inches as the characteristic The important characteristics of the section as alength, whole and of the tab and flap, as functions of tab andProvided by IHSNot for ResaleN
25、o reproduction or networking permitted without license from IHS-,-,-PRESSURE DISTRIBUTION OVER AN AIRFOIL SECTION WITH A FLAP AND TAB 3I Flap section normal-force coefficient, -n-zCn_ fCyhr _ !_ _ Flap section hinge-moment coefficient, c_,s=qc/27_t I Tab section normal-force coefficient, c_ _-. _ -!
26、_ - Tab section hinge-moment coefficient, ch,Tqc,23.O -_.- in which/ “ t_ i -i nwistheresultantpressurefolcenormalto the!_ airfoil chord. ! row, the corresponding pitching moment about_-_Li nr, the resultant pressure force normal to theflap chord.2.0 _ ! -!-!. h, the corresponding moment about thefl
27、aPhinge.nt, resultant pressure force normal to the-P -_ _ -_-I tab chord.q hi, the corresponding moment about the tab hinge.The subscript w refers to the airfoil section with flap/.o and tab; the subscript j to the flap section with tab;the subscript t to the tab section alone._ The integrated coeff
28、icients for the basic airfoilsection are plotted in figure 7 against angle of attack.z_ Curves giving the increments for various tab and flapdeflections are presented in figures 8, 9, and 10.Figures 11 and 12 are plots of theoretical parameterso taken from reference 3 and modified so as to apply_ di
29、rectly to N. A. C. A. absolute coefficients. Com-parisons of theoretical with experimental values of theforces and moments for the Clark Y airfoil tested withseveral different deflections of tile tab and flap areshown in figures 3, 14, and 15.-/.0 J DISCUSSIONPressure distribution.-The effects on th
30、e distribu-! tion Of resultant-pressure increments due to tab or flapi deflection are shown by figures 3 and 4. Deflections- _ of the tab or of the ftap produce peak values of theFIG_JRE2.-Distribution of resultant pressure oii airfoil section with flap and tab pressure increments at the tab hinge o
31、r at the flapneutral, a=0and15. hinge, respectively. If the tab and flap are deflectedflap deflection, are also plotted as increments. These simuItaneously (tab deflection opposite to that of flap),increments were obtained by deducting the basic- then peak values of the pressure increments occur ats
32、ection characteristics from those for the section with both hinge axes but the resultant pressures act indeflected flaps, the characteristics being determined in opposite directions. (See figs. 5 and 6.)Section characteristics.-The characteristics of theeach case by integration of the original press
33、ure basic airfoil section given in figure 7 (tab and flapdiagrams. Calculations were made of the following neutral) exhibit no unusual tendencies. For a givenquantities in which lower-case letters are used to indi-cate section coefficients: setting of the tab, the flap and tab may be consideredas a
34、flap unit. Then the effect of deflection of such a%wAirfoil section normal-force coefficient, c_-_ unit will be similar to that for an ordinary flap (e. g.,aileron, elevator, or rudder). Increments to the basicAirfoil section pitching-moment m. values of airfoil section normal-force and pitching-me-
35、coefficient, c_j4-qc,_2 ment coefficients are given in figure 8 for various flapProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for R
36、esaleNo reproduction or networking permitted without license from IHS-,-,-6 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICSdeflections with given tab settings. With the tab Lift coefficient of airfoil:deflected it will be noted that the curves are displaced_. dCJ , ba baparallel to the curve for
37、the undeflected tab. This U_=_La d-bs/a/d-_tta_) (1)parallel nature of the curves shows that the variationfof increments with flap deflection, considered withrespect to any given initial tab deflection, is independ- Pitching-moment coefficient:ent of tab deflection. At 30 deflection of the tab, _C,_
38、 bCm _.however, the effectiveness of the tab appears to have Umc/4:Umo- _- _f(_.f-d-_t8 t (2) .been considerably reduced so that tab deflections of _20 should not be exceeded with the arrangementstested. Hinge-moment coefficient of flap:Increments to the basic values of flap section normal- C C 5Chs
39、 C bChs bChfforce and hinge-moment coefficients are plotted in hs= hS0+_ L+_-1_s+_-t _ (3)figure 9 for various flap deflections with given tabsettings. The curves for the tab-deflected condition Hinge-moment coefficient of tab:are displaced parallel to the curve for the undeflectedtab, as was the ca
40、se for the airfoil section increments. C C bC_t bCht _C_,tThe variation of the flap increments with flap deflec- s_,= _,t0+_ CL+_-/_/+_ _t (4)-x-Omc/4 t _ - _ _- Chf -_ -_ Oht_.1_ I -.10 -.10LO -x/.80 / “enw / .40 I/ ./I0 . .40 -/i p :-t0 tx T_X_ _x-o 8 /6 -8 o -D D 8d, oeqree_ _, deglr-ee_ c_, oegr
41、-ee_FIGURE7.-Characteristics of the basic airfoil section. Tab and flap neutral.tion for a given tab deflection are likewise independent a is the angle of attack of the main portion of theof tab deflection, airfoil measured from zero lift of the un-Increments to the basic values of tab section norma
42、l- deformed section. (All angles are measuredforce and hinge-moment coeffcients are given in in radians.)figure 10 for various tab“ deflections with given flapsettings. The curves for the flap-deflected condition C, C_,lo, and Cht0 are moment coefficients at zerolift of the undeformed airfoil.are al
43、so displaced approximately parallel to the curvefor the undeflected flap, over the range of tab deflec- 5a ba bC_ bC_ bChf bCht bChl _Chttions from -20 to 20 . The curves show that the Parameters _, _t _-5-_t b_ 5_: _variation of increments with tab deflection for a given are given in figure 11.flap
44、 deflection is practically independent of flapdeflection. 5C_s bC_ tComparison with theory.-Theoretical expressions Parameters -_ and _ are given in figure 12.for the lift, pitching moment, and hinge moment for athin airfoil with any multiply hinged flap system have The curves given in figures 11 an
45、d 12 correspond tobeen derived by Perring (reference 3). The following those given in reference 3 except that the values haverelationships apply to a thin airfoil with a flap and been calculated and the curves redrawn on the basis ofa tab, N. A. C. A. absolute coefficients being used: N.A.C.A. absol
46、ute coefficients.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-PRESSURE DISTRIBUTION OVER AN AIRFOIL SECTION WITH A FLAP AND TAB 7i80 .20I- -:-_ i ! #,40 I _ _ - _ . / 0o_ ;-iL-O Che t -7_-I,40 Iflaphingemoment tab hingemomentFIGURE 2.-Iihge-moment
47、 parameters. Izlereil2ents in ratio of - ortal_ angle flap angleProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-PRESSURE DISTRIBUTION OVER AN AIRFOIL SECTION WITH A FLAP AND TAB 11?7 /o# 40 ,.I/._d/-.o:oo _ ,/ _: 20 _ _ -2o“_X ., 0o , _ -_ ,/ /o I Th
48、eo- -.10-.40 -; _ . “Feb- col i-.80. -.20II- 1.20 -30 -20 -I0 0 I0 20 30 -30 -20 -I0 0 I0 20 30 =80 t6_, degrees dy, degreesFIGURE13.-Comparison of theoretical, and experimental values of airfoil Beotionnormal-force and pitching-moment coefficients. Clark Y section with flap and tab. _ =0.30 - “, _ I I I I “ Theore/col- i .l I d_.2o _, “ =-_S:-_20 _ x =zx= 20 “-_ -_ _ _l _ ,., , ,15t-_ _ T/veorehcol ,_ “_ “, _ =-/5o“C ,- .,o
