1、NATIONAL ADVISORY COMMITTEEFOR AERONAUTICSTECHNICAL NOTE 4174WIND-TUNNEL INVESTIGATION OF THE STATIC LATERALSTABILITY CHARACTERISTICS OF WING-FUSEIAGECOMBINATIONS AT HIGH SUBSONIC SPEEDSTAPER-RATIO SERIESBy James W. Wiggins and Paul G. FournierLangley Aeronautical LaboratoryLangley Field, Va.Washing
2、tonOctober 1957Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NMNATIONAL ADVISORY CCMMITTEE FOR AERONAUTICS Ilolllllllllilllllll!llllllllllrilwoliiTECHNICAL NOTEWIND-TUNNEL INVESTIGATION 0?STABILITY CHARACTERISTICS4174THESTATIC LA
3、TERALOF WING-FUSELAGECOMBINATIONS AT.HIGH SUESONIC SPEEDSTAPER-RATIO SEKES1By Jsmes W. Wiggins and Paul G. FourniersuMMARYAn investigationwas conducted in the Lsmgley high-speed 7- by10-foot tunnel to determine the effects of variations In taper ratiowithin the range of 0.3 to 1.0 on the static late
4、ral stability charac-teristics at high subsonic speeds of wing-fuselage cotiinations havingwings of 450 sweepback at the quarter-chord line and an aspect ratioof 4. As has been shown in previous experimental investigations ofother wing plan forms, the parameter c2#LY which expresses the rateof chang
5、e of effective dihedral with lift coefficient, was found toincrease at the high subsonic Mach rmuibersas the force-break Mach num-ber was approached. Above the force-bresk Mach nuttiber,CIP/CL decreasedin magnitude with the severity of the break increasing with a decrease intaper ratio. The experime
6、ntal variation of zp/% increases negativelywith taper ratio and agrees well with the prdicted trend; however, theexperimental values are shown to be appreciably larger than the predictedvalues. At low and moderate lift coefficients the derivative of yawingmoment due to sideslip smdla.teral force due
7、to sideslip CyP forthe wing-fuselage combinations are contributed almost entirely by thefuselage alone; however, at high lift coefficients the effects of thewing are quite large.INTRODUCTIONA systematic research program is being carried out in the Langleyhigh-speed 7- by 10-foct tumnel to determine
8、the aerodynamic character-istics of various arrangements of the component parts of research-type“lSupersedes recently declassified NACA Research Memorandum L53B25aby James W. Wiggins and Paul G. Fournier, 1953.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from
9、IHS-,-,-2 NACA TN 4174airplane models, including some completeare being obtained on characteristicsinsteady rolJ_at Mach nmibers from 0.4-0tomodel configurations. Datapitch, sideslipj and duringabout 0.95.This paper presents results which show the effect of taper ratioon the aerodynamic characterist
10、icsin sideslip of wing-fuselage combina-tions having wings with a sweep of “ at the quarter-chord line, anaspect ratio of 4, and an NACA 6ti.airfoil section. The three wingshave taper ratios of 0.3, 0.6, and 1.0. Investigationsof the effectsof sweep and aspect ratio on lateral stability characterist
11、icsare pre-sented in references 1 and 2, respectively. The characteristicsof thefuselage alone, which has been comumn to all configurationscovered inthe program, are included in reference 1. In order to expedite theissuance of the results, only a limited comparison of some of the nxxresignificant ch
12、aracteristicswith available theory is presented in thispaper. COEFFICIENTSAND SYMBOLSThe stability system of axes used for the presentation of the data,together with an indicationof the positive forces, moments, and angles,is presented in figure 1. All moments are referred to the quarter-chordpoint
13、of the mean aerodynamic chord.%Liftlift coefficient, qsRolling momentrolling-moment coefficient,qsbCn yawing-moment% lateral-forceYawing momentcoefficient,qsbLateral forcecoefficient,qs!I dyrmic pmmmj $ lb/sq ftP mass density of air, slugs/cu ftv free-stream velocity, fps .Provided by IHSNot for Res
14、aleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 4174 3Mw. sbcMxh numberpVEReynolds ntier, vabsolute viscosity of air, slugs/ft-secwing area, sq ftwing span, ftwing chord, ftb/2mean aerodynamic chord, JsoC%y, ftspanwise distance from plane of symmetry, ftangle of attac
15、k, degangle of sideslip, degdeflection, f%acnC% . per deg*= per degCYP bSubscript:WF-F wing-fuselage values minus fuselage valuesProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 NACA TN 4174MODEL AND ATPARATUSThe wing-fuselage combinationstested are
16、 shown in figure 2. Allwings had an NACA 65Ao06 airfoil section parallel to the plane of sym- M“metry and were attached to the fuselage in a midwing position. AUwings were constructed of solid al-m alloy except thewing which was of composite construction, consisting of aa bismuth-tin covering. The a
17、luminum fuselage was commonrations and it+ ordinates sre presented in reference 3.The three wings used in this investigationrepresenttaper-ratio-().6 *-steel core andto all configu-onl.ya partof a family of wings being studied in a nmre eensive program; there-fore, a common wing designation system i
18、s bedng used for the entireprogram. For example, the wing designatedby 45-4-o.6-006 has thequarter-chord line swept back 45, an aspect ratio of 4, and a taperratio of 0.6. The nuniber006 refers to the section designation; inthis case the design lift coefficient is zero and the thickness is6 percent
19、of the chord.The models were tested on the sting support system shown in fig-ures 3 and 4. With this support system the mcdel canbe remotely eoperated through a 28angle range in the plane of the vertical strut.By means of couplings in the sting, the model can be rolled through $so that either angle
20、of attack (fig. 3) or angle of sideslip (fig. 4) *can be the remotely controlled variable. With the mdel horizontal(fig. 3) couplings can be used to support the model at angles of side-slip of approximately -40 or 4 while the model is tested through theangle-of-attackrange.TEST AND CORRECTIONSThe te
21、sts were conducted in the Langley high-speed 7- by 10-foottunnel through a Mach number range from O.40 to 0.95. The size of themdels caused the tunnel to choke at correctedMach numbers of 0.95to 0.96, depending on the wing being tested. Th.e.block correctionswhich were applied to the data were deter
22、minedby the velocity-ratiomethod of reference 4.The present investigationconsisted of two groups of tests. Thefirst, from which most of the data were obtained, involved runs at anglesof sideslip of -4 and 40 through an angle-of-attackrange from -3to 24. The second series of tests were made at severa
23、lpredeterminedangles of attack through a sideslip-anglerange from 4 to -1OO. MdProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 4174 5The jet-boundary corrections applied to the aze of attack weredetermined by the method of reference 5. The co
24、rrections to lateralforce, yawing moment, and rolling moment were considered negligible.!lkrecorrectionswere obtained but were found to be negligible for awing-fuselage configuration and therefore were not applied. The angleof attack and the angle of sideslip have been corrected for the deflec-tion
25、of the sting support systa,and balance under load.Corrections for the spanwise dihedral distribution due to wing dis-tortion while under aemdynsmic load (figs. 5 and 6) have been appliedto these data sndwere determintiby the method discussed in detail inreference 1.The Reynolds numibervariations wit
26、h test Mhchnuniber ere presentedin figure 7 for the three wings. Reynolds nunibersrange froml.75x ldto 3.20 x lhowever, results from the 0.3-taper-ratiowingindicate a recovery from this condition at a lower Mach nuniberthan for -the wings of taper ratio 0.6 and 1.0 (figs. 8, 9, and lO). The range of
27、lift coefficientover which CIB is linear increaseswith Mach numberup to the force break (M %0.93) for the wings of taper ratios of 0.6and 1.0 (fig. 9 and ref. 1), but the results from the 0.3-taper-ratiowing (fig. 8) tidicate essentially no effect of Mach number cm the linearrsnge of CZP at these Ma
28、ch numbers. The variation of the slope c%/cL =wedne= erouft ithMach nuniberis presented in figure 12 and the variation with taper ratiois shown in figure 13 along with a comparisonwith wing-alone theory andexperimentalvalues corrected for aeroelastic distortion. The theoretical “predictions were det
29、erminedby applying the compressibilitycorrectionscalculatedby the method of reference 6 to the incompressible-flowvaluescalculatedby the method of reference 7. bAs was generally found for thewings investigated in references 1 and 2, values of /CZ increasedat the higher subsonicMach tiers as the forc
30、e-breskalthough thepresent available theory (ref. 6) predicts a slight decrease inwithin this Mach nuniberrange. . c%/All wing-fuselage combinations exhibited a reduction in Cl PIabove the force-breakBkch nuniber,with the severity of the breskincreasingwith a decrease in taper ratio.The experimental
31、variation of cZp/cL with taper ratio (fig. 13)agrees well with the predicted variation; however, as also shown in fig-ure 12, the experimentalvalues are appreciably greater than the pre-dicted values.Lateral-Force and Yawing-Moment CharacteristicsComparisons of the variations of the lateral-stabilit
32、yparametersCn9 and Cy$ with lift coefficientare presented in figures 14 and 15Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 4174 7for the wing-fuselage configurations. The wing-plus-wing-fuselage-sinterference data for the same conditions,w
33、hich were obtained by sti-tracting the fuselage-alone data of reference 1 from the data of fig-ures 14 and 15 are presented in figures 16 and 17.JThe results indicate that, at lift coefficientsbelow about 0.8,the derivatives %P - %p are contributed almost entirely by the.fuselage alone. The breaks i
34、n the curves at the higher lift coeffi-cients are probably due to wing stalling which changes the magnitudeand orientation of the resultant force on the two wing semispans.CONCLUSIONSThe results of the present investigation of the aerodynamic char-acteristics in sideslip at high subsonic speeds of w
35、ings having varioustaper ratios, witha sweep angle of 4Y, an aspect ratio 6, and anNACA 65ACK% airfoil section indicate the foldowing conclusions:1. The experimental variation of clpl (raw Of change of effec-tive”dihedral with lift inefficient) indi however, the eeri-mental values are shown ta be ap
36、preciably greater tk the predictedvalues.3. At low and moderate lift coefficients the derivative of yawingmoment due ta sideslip % and lateral force due to sideslip P forthe wing-fusekge conibinationsare contributed almost entirely by thefuselage alone; however, at high lift coefficients the effects
37、 of thewing are quite large.Langley Aeronautical Laboratory,National Advisory C!csmdtteefor Aeronautics,LangleyField, Vs., February , 1953.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACATN4174IWTERENCES1. Kuhn, Richard E., and Fournier, Paul G
38、.: Wind-Tunnel Investigationof the Static Lateral Stability Characteristicsof Wing-FusebgeCombinations at High Subsonic Speeds - Sweep Series. NACARML52Glla, 1952.2. Fournier, Paul G., and Byrnes, Andrew L., Jr.: Wing-Tunnel Investiga-tion of the Static Lateral Stability Characteristics of Wing-Fuse
39、lage Conibinationsat High Subsonic Speeds - Aspect-Ratio Series.NACARML52L18, 1953.3.Kuhn, RichardE., and Wiggins, Jsmes W.: Wind-Tunnel Investigationof the Aerodynamic Characteristics in Pitch of Wing-Fuselage Com-binations at High Subsonic Speeds - Aspect-Ratio Series. N.ACARML52A29, 1952.4.Eensel
40、, Rudolpf W.: Rectsm.ar-Wind-Tunnel Blocking CorrectionsUsing the Velocity-RatioMethod. NACA TN 2372,1951.5. Gillis, Clarence L., Polhsmus, Edward C., and Gray, Joseph L., Jr.:.-,Charts for Determining Jet-Boundary Corrections for Complete Modelsin 7-by10-Foot Closed Rect Wind Tunnels. NACA WR L-123
41、,194.5. (FormerlyNACAARRL5G31. )“6.Fisher, Lewis R.: Approximate Corrections for the Effects of Com-pressibility on the Subsonic Stability Derivatives of Swept Wings.NACATN 1854, 1949.7. Toll, Thomas A., and Queijo, M. J.: Approxi.nxxteRelations and Chartsfor Low-Speed StabilityDerivatives of Swept
42、Wings. NACA TN 1581,1948.K.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2C NACA TN 4174 9Lofefol forceRelutive windYuwi momentJ-(3Rolling momentLiftYowing momentRolling momentz !URehfive windFigure 1.- System of axes used showing the positive dire
43、ction of forces,moments, and angles.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Wi geomehyAR?O 2.25 Sqffspa? JoOftSwwep at 025 cinvd line 45Aqwct mtioIm?idenceDitwdmfAiI1. *F1.gure4.- A typical. model installed for wiable angle-of -sidediptests .
44、 Shown at 0 e of attack.* !, .1 ,1,.: ,i, 8,L-67495,.i.,.!,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.(KWO.mo5 “o0 .2 “4 ybh .6 .8 /.0Figure 5.- Deflection curves for the test wbgs.45-4-/.0-00645-4-,6-00645-4-,3-006Provided by IHSNot for Resale
45、No reproduction or networking permitted without license from IHS-,-,-ACZPACL.(702.00/o4 .5 .7 .8 .9 LO45-4-10-00645-4-.6-00645-4-.3-006Much number, M wl?igure6.- CorrectIon factors used to correct for the effects of aero-elastic distortion. 8Provided by IHSNot for ResaleNo reproduction or networking
46、 permitted without license from IHS-,-,- al45-4-.3- (W645-4-.6-00645-4-10-0064 ,5 .6 .7 .8 .9 LoMoth number, MFigure 7.- Variation of mean Reynolds nuuker with test Mach nmnber based Gon the wing mean aerodynend.c chord.Provided by IHSNot for ResaleNo reproduction or networking permitted without lic
47、ense from IHS-,-,-. . .- -,Al0 J?5b0 .94V0 93 ,0 92 bo 91 40 E5 .,.0 80 .0 .70 Ao 600Om m.G o .40m.zo2#6 E&7Lift coeffkient, CLo00000000M95 94,93 v9.? a.91 tB5 .m.70 A.6000 ?J.X72 .%# o .400#-J?024.6.8iDL/ft awffichnto00000“oodM95.SF93 92 491 v.85 0m.m A600O&w uc% o .400m-J2024.IS.81DLift cwffiiit,
48、CLFigure 8.- Lateral stability chm-acteristics of the 4.5-4-0.3-(K%wing-fuselage combination. Not corrected for aeroelastic M.s-kotiion.Flagged symbols represent tests In which the angle of sided.ip wasvaried., , 1 II,1!,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, F.o00000000+? O ?46.8,WIJft confflclent ,CLo00000000, ,720.2 f?6.8LoLtft mefficfant, CLo00
