1、I:,1NATIONALADVISORY COMMITTEEFOR AERONAUTICSTECHNICAL NOTE 2741INVESTIGATION OF THE INFLUENCE OF FUSELAGE AND TAILSURFACES ON LOW-SPEED STATIC STABILITYAND ROLLING CHARACTERISTICSOF A SWEPT-WING MODELBy John D. Bird, Jacob H. Liechtenstein,and Byron M. JaquetLangley Aeronautical LaboratoryLangley F
2、ield, Va.WashingtonJtiy 1952AFM?cProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM -lZ.NATIONAL ADVISORY COMMITTEE.kIlllllllllllllllllllllillllllllll=ollb5707 =FOR AERONAUTICS “”- “-TECHNICAL NOTE 2741INVESTIGATION OF THE INFLUENC
3、E OF FUSEIXZ AND TAILSURFACES ON IX3W-SFEEDSTATIC STABILITYAND ROLLING CHARACTERISTICSOF A SWEPT-WING MODEL1By John D. Bird, Jacob H. Liechtenstein,and Byron M. JaquetSUMMARYA wind-tunnel investigationwas made in the Langley stability tunnelto determine the influence of the fuselage and tail surface
4、s on the staticstability and rotary derivatives in roll of a transonic airplane configu-ration which had 45 sweptback wing and tail surfaces8 The tests made in straight-flow showed that the wing alone has mar-ginal longitudinal stability characteristics near maximum lift. Thevariation of rolling-mom
5、ent coefficient with angle of yaw of-the complete .model is almost the same as for the wing alone.The results of the tests made in simulated rolling flight indicatethat for this model the effects of the fuselage and tail surfaces onthe rate of change of the rolling-moment, yawing-moment, and lateral
6、-force coefficients with wing-tip helix angle are small in comparisonwith the effect of the angle of attack on these rotary characteristics.The vertical tail produces larger incremmts of the rate of change oflateral-force and yawing-moment coefficients with wing-tip helix anglethan the fuselage or h
7、orizontal tail.INTRODUCTIONEstimation of the dynamic flight characteristics of aircraft requires “a knowledge of the component forces and nmments arising from the orientation%upersedes the recently declassified NACA RM L7H15, “Investigationofthe Influence of Fuselage and Tail Surfaces on Low-Speed S
8、tatic Stability.and Rolling Characteristics of a Swept-Wing Model” by John D. Bird,Jacob H. Liechtenstein,and Byron M. Jaquet, 1947.wProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 . . . . NACA TN 27+1of the model with .respectto the air stream (st
9、atic derivatives) andfrom the rate of angular displacementwith resyect to the air stream(rotary derivatives). The forces and moments arising from orienta-tion of the model are determinedby use of conventionalwind-ttieltests, and, until the recent use.of large amounts of wing sweep, therotary derivat
10、ives at other than very high angles of attack were satis-factorily estimated by theoretical means. Unpublished data and thecalculations of reference 1, however, show that for swept wings thederivatives in roll cannot be satisfactorilypredicted byexisting theo-retical means, particularly at moderate
11、and high lift coefficients. An ,investigationtherefore was conducted to determine the influence of thetail surfaces and fuselage of an airplane on the low-speed rotary deriva- -tives in roll of a transonic airplane configuration having 45 sweptbackwing and tail surfaces. The static stability charact
12、eristicsof variousconfigurations of the model were determined inThe results of this investigationare reportedSYMBOLSThe results of the tests are presented asthe course of the tests.herein.standard coefficients offorces and moments which are referred to the stability axes the originof which is assume
13、d to be at the projection on the plane of symmetry ofthe quarter-chordpoint”of the mean geometric chord of the wing of themodel tested. The stability axes system is shown in figure 1. The coef-ficients and symbols used herein are defined as follows:lift coefficient()LzCD ().1drag coefficient Q-qsCyC
14、mCnlateral-force coefficient ()qs()rolling-moment coefficient qs-rl()pitching-moment coefficient qsc()yawing-moment coefficient qSb,.L“.-,.-*A,.“Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN2741 3.LD“YL1MNPvs.b. ca*pb%Plift, negative of Z-f
15、orce in figure 1draglateral forcerolling momnt about X-axispitching moment about Y-axis,yawing moment about Z-axis()v2dynamic pressure amass density of airfree-stream velocitywing areaspan of wingchord of wing, measured parallel to axis of symmetryangle of attack, measured in plane of symmetry, degr
16、eesangle of yaw, degreeswing-tip helix angle,rate of roll, radiansradiansper secondProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 NACA TN 2741-.APPARATUS AND TESTS I.The tests described herein were conducted in the 6-foot-diameterrailing,-flowtest
17、 section of the Langley stability tunnel. This see-tion is equipped with a motor-driven rotor which imparts a twist to theair stream so that a model mounted rigidly in the tunnel is in a fieldof flow similar to that which exists about an airplane in rolling flight(reference 2). The test model is mou
18、nted on a single strut which isconnected to a conventional six-componentbalance system.The model used.for the subject tests was a transonic configurationhaving 45 sweptba.ckwing and tail surfaces, These surfaces hadNACA 0012 airfoil sections normal to the leading edge (thicknessratio 0.085 parallel”
19、to plane of symmetry) and a taper ratio of 1:The fuselage was a body of revolution which had a circular-arcprofileand a fineness ratio of 8.34. A view of the model mounted in the tun-nel”is shown as figure 2, and the geometric characteristicsof themodel are.given in figure 3.The test configurationsa
20、nd thedata in the figures,are given in thedata were obtained from reference 3.wing. . . a71 . . a71 . . . . . . . a71 . .Fuselage. . . . . . . . . . . . .Wing and fuselage . . . . . “ . .Wing, fuselage, and vertical tail .Wing, fuselage, vertical tail, andhorizontal tail . . . . . . . . .Six-compone
21、nt asurements weresymbols used in iults i;icated that,although there were large tare corrections to the drag coefficient, the. corrections to the derivatives of the forces and moments with respectto yaw angle and wing-tip helix angle were in most cases negligible.Provided by IHSNot for ResaleNo repr
22、oduction or networking permitted without license from IHS-,-,-6 NACA TN 2741 5Although reference 6 presents a more exact method of detetini ,the method used herein, as outlined in reference 4, is believed to give .sufficiently accurate results for the model aridtunnel used in this ainvestigation.RES
23、UITS AND DISCUSSION .Wesentation of Data .-The results of-this investigationare presented in figures 4 to-9.Curves are given in each plot.for all configurationstested in order tofacilitate comparison. Figure 4 presents the lift, drag, and pitching-moment characteristicsof the test configurationsfor
24、the angle-of-attackrange at $ = 0, together with a cross plot of the.pitching-moment coef- .ficient against lift-coefficient. Figures 5, 6, and 7 present the varia- tion of the rolling-moment,yawing-moment, and lateral-forcecoefficientswith angle of yaw for angles of attack of 0,6.20,and 12.o. The .
25、derivatives cz) c and CY are presented for the angle-of-attackrange in figure 8. Figure 9 presents the derivatives Czp) Cnp.and Cyp for the angle-of-attackrange. .Characteristics in Straight Flow*The longitudinal stability characteristicsof all model configure-.tions other than the complete model an
26、d the fuselage alonewere marginal .-in the critical region near maximum lift. The longitudinal stabilitycharacteristics of the complete model are satisfactory for the entire lift range (fig. 4). Marginal characteristicsfor the wing alone arepredicted by-the correlation of longitudinal stability char
27、acteristicsof swept wings presented in reference 7.The curves of figures 5, 6, and 7 indicate approximately a linearvariation of yawing-moment, rolling-moment,and pitching-moment coef- .:ficients with angle of yaw for afiglesof attack up to 12.5.The curves of figure 8 indicate that, up to maximum li
28、ft, %*is primarily a function of the characteristicsof the wing alone. Thisfact is evidenced by the proximity of the curves of Cz$ plotted againstangle of attack for the various test configurations. With regard tocn the vertical tail produces a stabilizingeffect which, except at “-. ,-very high angl
29、es of attack, is larger than the destabilizing effect(positive increment of CnV) produced by the selage (fig, 8). The .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 2741.influence of the vertical tail and the fuselage”on Cy is of the samesi
30、gn except at high angles of attack (fig. 8).Characteristics in Rolling FlowFrom calibration tests it was determined that the lift, drag, andpitching-moment coefficients of the model were almost independent ofthe rate of rotation; whereas the lateral-force,rolling-moment, andyawing-moment coefficient
31、s varied linearly with rate of rotation. Thederivatives, however, presented herein were obtained from tests made-values of the derivatives- Clp, C ,and-%in comparison with the effects of angle ofCONCLUSIONSWind-tunnel tests for determining theCYP- of the wing are small ._“_:- - -.:-. .-attack on the
32、se derivatives. .:“ -static stability character-” -” istics and the rotary derivatives in roll of a transonic mide configu- “- “”-”ration having 45 sweptback wing and tail surfaces indicate the following “:conclusions:1. The longitudinalstability characteristicsof the wing alone andthe”mode.1without
33、 the horizontal tail surfaces are marginal in thecritical region near maximum lift. The characteristicsof the completemodel are satisfactory.2. The variation of the lateral-stabilityparameter CZQ isprimarily a function of the characteristics of the wing alone up tomaimum lift. 3.Theaddition of”the f
34、uselage aridhorizontal tail surfaces tothe wing has little effect on the rate of change of the rolling-moment”,yawing-moment, and lateral-forcecoefficient%with wing-tip helix angl.4. The addition of-the vertical tail to the model produces appre-ciable increments in the rate-of change of te rolling-m
35、oment, yawing-moment, and lateral-force coefficients with wing-tip helix angle, butthese variations are small in comparisonwith the effects of angleof attack on these rotary characteristics.Langley Aeronautical LaboratoryNational Advisory Committee for AeronauticsLangley Field, Va.,.August 21, 1947.
36、,.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2Z NACA TN 274141.2.3.4.5.6. 7.REFERENCESWeissinger, J.: The Lift Distribution of Swept-Back Wings. NACATM 1120, 1947.Mackchlan, Robert, and Letko, William: Correl drag, and pitching+noment coefficien
37、tswith angle of attack for all model configurations. $ = 00.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.-. -. -,NACA TN 27417.*-.-.-,“.,-.#“ .-.:.-.,:=. ,. -:. - -.*M-NACA-LIey -7-22-62-1000Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
