NASA NACA-TN-2819-1952 Effect of high-lift devices on the static-lateral-stability derivatives of a 45 degree sweptback wing of aspect ratio 4 0 and taper ratio 0 6 in combination .pdf

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1、, INACATN2819measured parallel.S. measured parallel to plane of symmetry, ftchord of slat, measured parallel to plane of symmetry, ftchord of flap, measured parallel to plane of symmetry, ftlongitudinal distance rearward from airplane center of gravityto wing aerodynamic center, ft -longitudinal dis

2、tance forward from”wing aerodynamiccenter to-center of pressure of l“iftload dfieto flap deflection, ftlongitudinal distance forward from fiingaerodynamic center tocenter of pressure of drag load due to flap deflection, ft “2aspect ratio, %Yfeffective aspect ratio of flapped part of wing, Ataper rat

3、io, ratio of tip chord to root chordangle of sweep, positive for sweepback,degangle of sweep of flap hinge line, positive”for sweepback,degangle of atta”ck,.measured in plane of symmetry, deginduced angle of attackangle of sideslip, degflapisdeflection relative to wing, positive when trailing edge d

4、own, measured in plane normal to hinge line, degProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.NACA TN 2819 5“a. section lift-curve slope whenc1 section lift coefficientab flap-effectivenessparameter,hinge lineplaced normal to air streammeasured in

5、 plane normal toacyc%TiiAclfj Cn9 ACYP increment in CZP, C%, CyB due to flaP deflec-(tion at constant a or CL for example,.)Czp ()- c1wing with flaps ) wing without flaps. ACLf increment in lift coefficient due to flap deflection at aspecific angle of attackACDO increment in profile drag coefficient

6、 due to flap deflection( )CDO ()- CD.with flaps without flaps)Subscripts:L left semispan of wing, retreating semispan for positivesideslipR right sespan of wing, advancing semispan for positivesidesl.ipMODEL-COMPONENT DESIGNATIONSThe components for the various configurationsused in the presentinvest

7、igation are identified herein by the following letter designations:a71 w wing aloneWB wing-body configuration*Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA TN 2819s slatF1 plain flap with uutboard end at 0.”kb/2F2 plain flap with outboard en

8、d-at 0./2F3 plain flay with outboard end at 1.6b/2F4 split flap with outboard end at 0.i%/2 “-F5 split flap with outboard end at 0./2F6 split flap with outboard end at 1.Ob/2.-._.-.-MODEL, APPARATUS, AND k!ESTSThe gener-alresearch model used for the-present investigationwas “- “designed to permit te

9、sts of the wing-body cofifigurationalone or with alZYof various combinationsof slats and trailing-edgeflaps. A sketch of .the completemodel Is presented in figure 2, and a list of pertinentgeometric characteristicsof the various componentparts is given intable I. .The wing had 45 sweepback of.the qu

10、arter-chordline, an aspectratio of 4.0, a taper ratio of 0.6, and NACA 65AO08 airfoil sectionsparallel to the plane of symmetry. The ordinates for the NACA 65AO08airfoil section are given in table II. The wing was mounted along the _ ;body center line. The body was a body of r“eflutionwith a finenes

11、sratio of 6.67. The body profile followed a whereas, with the slat extended, the initialbreak is delayed untilabout 14. For the wing-body configurationwithout slats, correspondingbreaks were found in the CL) Cm and Cl curves.P No such breakswere found for the configurationswith slat extended. Inasmu

12、ch astares were not taken into account, the absolute values of the drag coef-ficients should not be consideredas representativeof free-air values.The increments in drag coefficient due to flap deflection and the varia-tion of drag with lift, however, should be re”=sonablyacctiate.Although the increm

13、entsin lift due to flap deflection for theplain flap were equal to or greater than those for the sPIJt flaP, the .- _ _.increments in drag were somewhatless for the plain flap than for thesplit flap. The lift-dragratio, therefore, for a given lift coefficientwas higher for the plain flap than for th

14、e split flap, either with orwithout the slat.hProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2A NACA TN 2819 9. Static Lateral Stability CharacteristicsThe static lateral stability characteristics for the wing-binationwith plain and split flaps both

15、 with and without leading-edge slats are presented in figure 8.For the configurationswithout slats, the effect of flap deflectionon C2P, CnB, and Cyp is generally similar to the effects found pre-.-viously in reference 4. As discussed in reference 4, the short-spanflap shifts the center,of pressure

16、inward from its position withoutflaps; consequently the Czp curve is shifted in a positive direction.Increasing the flap span generally shifts the curves in a negativedirection because the center of pressure is moved outward from its posi-tion with short-span flaps. n addition, the flaps delay the p

17、ositivebreak in the Clp curve until higher lift coefficients are attained sothat at high lift coefficients the value of CZP becomes more negativefor all the configurationswith flaps than for the configurationwithoutflaps.k due f np = -0.001 for the wing-body configuration is ingood agreement with th

18、e results presented in reference 6 for this con-.figuration, and this instability is entirely due to”the unstable momentof the body. Increasing the flap span generally tended to make c%less negative (decreasingthe directional instability) particularlyatthe higher lift coefficients. As a matter of fa

19、ct, at about 0.9 maximumlift coefficient, the instability-introducedby the body was nearlyremoved by the largest-span plain flap and fully removed by the largest-span split flap.Addition of full-span leading-edge slats to the vsrious configura-tions with and without trailing-edge flaps (figs.:8(c) a

20、nd 8(d) gen-erally extended the trends of the ClP and CyP curves obtained at lowlift coefficients to higher lift coefficients. However, the slats gen-erally introduce a slightly stable vsriation of CnP with increasinglift coefficients until the final break occurs just before maximum lift.The shifts

21、in the values of Czp due to trailing-edge-flapdeflectionwere similar in nature but of different magnitude with slats added tothe wing as compared to the wing without slats (compare fis. 8(c)and 8(d with 8(a) and 8(b). lthough the slatsthe slope of Cz against CL, they extended theBcurve to nearly max

22、imum lift, and, therefore, thea71were greater negatively (greater dihedral effect)Ag The effecttaken into account in the expression for the lift and dragpressure which in this analysis depend only upon flap spanJWICJITN 2819oftaperwas *centers ofand taper.The span-load distributionof a wing with fla

23、ps indicates that, for thepurpose of determiningthe aerodynamic induction,the effective aspectratio of the flapped part ofthe wing A should be used rather thanthe aspect ratio of the wing.Expressions for the centers of pressure of the flap loads are. .-Y-f ()yf 2()yf 31+X+ (7+A)YLf 1 :L(9-5k) +3(1 -

24、k)=z.( )8+h(7-4)+2(i”- 2b/2Y% 11+2A” “31+A “- .Lf ().5(x - I) + z - _XLf AtanAl+2h.+- l+k 1A tanA 0.5-tc 6 l+AXD ( )Df Atari A-1e3 -_r AtanAl+2A- 1.3=-tc 6 l+A +2 _l+A l+A.SideslippingFlight In sideslippingflight for a constant-chgrdswept wing, the span-load distribution is considered,for this analy

25、sis, to be the same onboth wing semispansalthough the magnitudes are different. .The loadsare affected by sidesIipbecause of the mannefiin which sideslip a-ffects“” 1 -.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-15.NACA TN2819the velocities norm

26、al to the quarter-chordline. In sideslipping flight,the leading semispan (right semispan in this analysis) has less effectivesweepback,whereas the trailing semispan (left semispan) has greatereffective Bweepback. The velocity component on the left semispan isaltered, therefore,by the factor COS(A +

27、). For a flapped wing ofcos Ainfinite span, the increment in lift due to sideslip on the left semi-span can, therefore,be expressed asACZL = AC2P40 1cOS(A + )cOs(Ah”+ P - COS A COs AhInasmuch asand for= aoCLb coB A cOS Ah = Acl“lp+o COS A COS A.oosmall angles of sideslip sin p = P, cos = 1, and sin2

28、P isnegligible, the increment in lift can be rewritten asThe incrementis expressed asACZL = -AcZA40 sin(A + Ah) (Al)in primary force coefficient for a finite span wingandACLL = AcZL - Induced liftThe lift distributionresulting from sideslip isrespect to the plane of symmetry; therefore, thedetermine

29、s the magnitude of the induced angle of,sion ai = is one-h of the wing geometricanisymmetricalwithasect ratio thatattack in the expres-aspect ratio. Since8Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-916 NACA TN 2819the aspect ratio is consideredp

30、erpendicular tothe relative wind, the 4A COS2(A+ )effective aspect ratio becomes .- Inasmuch as the inducedCOS2A .angle of attack in a plane parallel to the pltherefore, the expression for a swept wing becomesA, Sf COS A COS % (A2)ACL = Acl_oA+2COSA,Provided by IHSNot for ResaleNo reproduction or ne

31、tworking permitted without license from IHS-,-,-3A.NACA TN 2819With the aid of equations (Al) and (A2), %L can be reduced toACLf sin(A+ Ah) A + 2 COSAACIL = - COS A COS AhA + cos A17,The total of the symmetrical and unsymmetrical values of the primary force coefficient on the left semispan isACLf 44

32、sin(A +Ah) A! + 2 =Os AclL=l- PCOSACOSAhA + Cos (A4)and by similar analysis for the right semispanLf, 1sin(A + Ah) A + 2 cos ACIR=l+ cosA”cosAhA, +4 cos A (A5).Continuing in a manner parallel with that of reference 8 permitsthe primary force coefficient C2L to be expressed aBandSince the profileof t

33、he flap hinge lineshown, therefore, that,L = “Ly $ COS2AY(IL COS3(A+ P):()2C1R Swc COS2A2R = 7A COS3(A - P)(A6)(A7)drag acts parallel to the air stream, the sweepdoes not enter the consideration. It can beACDo COS2(A + B)ACDOL = COS2A(A8)Provided by IHSNot for ResaleNo reproduction or networking per

34、mitted without license from IHS-,-,-mandNWA TN 2819.Rolling moment.-to flap dflection inWDo COS.2.A-“ J3)ACD = %2 co2AThe increment in rolling-momentsideslipACZ =Equations (A4) and (A5) can becan be expressed as( YLf)CIL - .CIR ycombined with equation(A9) “coefficient due(A1O)(A1O) to obtainYLf“LfA

35、+2 (tarii + tanAh)AClfl= - 2A+ (All)Lateral force.- Theforce coefficient due to,theory indicates that the increment in lateral.flap deflection in sideslip should be.A% = (2L- C?sinA - (ACDOL +AcD%)sin (A12)Substituting equations (A4) to (A9) into equation (AI-2)gives the fol-lowingAcyp =expressions

36、for the derivative:“ % (A13)Yawing moment.- The increment in yawing-moment coefficient dueto flap deflection in sideslip can be expressed as(YLf %- LfC2R)cosA -( C2R)sinA b +ACn=- C2L - C2L - .yDf iZ-(Df) ( )ACDOL - ACD% Cos + DoL.+ ACDOR in b (A14) “-“.Provided by IHSNot for ResaleNo reproduction o

37、r networking permitted without license from IHS-,-,-NACA TN 2819 19With the aid of equations (A4) to (A9), the incremental derivativebecomes()2 AcLf 2ACn =B ;: l(tanA+tanAh) . .ti cos ASw1( z- )3 tan A sin A Lf + cosAyLf2 A -!c m +(l?- XD )Y-DfACDO AE r+tan Ahv (A15)Provided by IHSNot for ResaleNo r

38、eproduction or networking permitted without license from IHS-,-,-820 NACA TN 2819 “REFERENCES1. Queijo, M. J., and Wolhsrt, Walter D.: Experimental Investigationof the Effect of Vertical-Tail.Sizeand Length and of FuselageShape and Length on the Static Lateral”StabilityCharacteristicsof a Model With

39、 4.5Sweptback Wing and Tail Sur,faces.NACARep. 1049, 1951. (SupersedesNACATN 2168.)2. Letko, William: Effect of Vertical-Tailkrea and Length on the YawingStability Characteristicsof a Model Having a OSweptback Wing.NACA TN 2358, 1951.”3. Herriot, JohnG.: Blockage Corrections for Three-Dimensional-Fl

40、owClosed-ThroatWind Tunnels, With Considerationof the Effect ofCompressibility. NACA Rep. 995, 1950. (SupersedesNACA RM A7B28.)4. Liechtenstein,Jacob H.: Effect of High-LHt Devices on the Low-SpeedStatic Lateral and Yawing Stability Characteristicsof an Untapered47 Sweptback Wing. NACA TN 2689, 1952

41、, (SupersedesNACA RM L8G20.)5. Thiel, G., and Weissinger, F.: Six-ComponentMeasurements on aStraight and a 35 Swept-BackTrapezoidalWing With and WithoutSplit Flap. NACA TM”1107, 1947.6. Brewer, Jack D., and Liechtenstein,Jacob H.: Effectof HorizontalTail on Low-Speed Stitic Lateral Stability Charact

42、eristicsof aModel Having 45 SweptbackWing and Tail Surfaces. NACA TN 201O,1950.7. Goodman, Alex, and Fisher, Lewis R.: Investigationat Low Speeds ofthe Effect of Aspect Ratio and Sweep on Rolling Stability Deriva-tives of Untapered Wings. NACA Rep. 968, 1950. (SupersedesNACATN 1835.)8. Toll, Thomas

43、A., and Queijo, M. J.: Appr324a71 369.190.830.100.20607.2012.6018.000.20407.2012.6018.00Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-22 NACA TN 2819TABLE IIoAs FOR NACA 65Aoo8AIRFOILtatin andorunates in percent airfoilStation Ordinateo 0.YO .615a7

44、1 75 .7461.25 .9512.5-0 1.3035.00 1a71 7497.50 2.32010.00 2.43215.00 2.92620.00 3.30125.00 3.58530.00 3a71 79135.00 3.92840.00 3a71 99545.00 3.98850.00 3.89555.00 3.71460.00 3;45665.00 3.13570,00 2.76373.00 2.34880.00 1.89885.00 1.43090.00 .96Q95.00 .489100.00 .018L. E. radius: 0.408.1chord.a71Provi

45、ded by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 2819 23,.TABLE IIIBODY ORDINATEStation and ordinates in percent bodyStation Ordinateo 02.5 .71.4?; 2.110.0 2.712.5 3.315.020.0 ;:25.0 5*730.0 6.335.0 6.840.045.0 ;:50.055.0 ;:;60.0 7.365.0 7.270.0 6.

46、975.080.0 2:85.o 5.790.0 5.195.0 4.5100.0 3.8lengtiij.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 NACA TN 2819TABLE IVORDINATES FOR SLAT AND MAIN AIRFOIL SECTIONtations and ordinates in percent airfoil chorSlat Main airfoillUpper surface Lower

47、 surface Upper surfaceStation Ordinate Station Ordinate Station Ordinateo 0 0 0.498 .622 .498 -.622.747 .747 .747 -.7471.253 .951 1.253 -.%1 “-2.498 1.298 2.000 -1.173 2.000 -1.1734.996 1.751 2.667 -.098 2.667 -.0987.502 2.116 3a71 333 .338 3*333 .33810.000 2.427 4.000 .658 4.000 .65815.004 2.933 4.

48、667 .924 4.667 .9243*333 1.173 5.333 1.1736.222 1.449 6.222 1.4497.111 1.689 7.111 1.6898.000 1.902 8.000 I.9028.889 2.116 8.889 2.1169.778 2.311 9.778 2.31110.000 2.338 10.000 2.33815.004 2.933 10.667 2.480II556 2.61315.004 2.933%ehind the 15-percent station, the upper surface of the airfoilis the same as the basic NACA 65AO08 airfoil. -:Behind the 2-perc