REG NACA-RM-A51A16-1951 Low-speed investigation of a 0 16 scale model of the X-3 airplane lateral and directional characteristics.pdf

1、RESEARCH MEMORANDUM LOW-SPEED INVESTIGATION OF A 0.16SCALE MODEL OF THE X-3 -PLANE - LATERAL AND DIRECTIONAL CEURACTERISTICS By Noel K. Delany and Nora-Lee F. Hayter Ames Aeronautical Laboratory Moff ett Field, CaXf . “. fHlFlSMK 25. D. c - NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON Marc

2、h 16, 2951 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM A5lAl6 NATCONALADVISOR.YCOBQETTEEFOR .KERONAmCS LOW-TSPEED INVESTIGATION OF AO.164CAIZDELOF THFzX3 AI- - By NoelK. Debmy and Nora-Lee F. Hay-ter . A wind4xmnelinvestigation has been m

3、ade of the low-speed, static, lateral and directional characteristics of a model of an early desiep of the X3 airplane with the King flaps neutral and deflected. Measurements were also made of the fluctuations In rolling moment with time. . The model utilized a wing having an aspect ratio of 3.01, a

4、 4.5 percent4hick hexagonal section, and a taper ratio of 0.4. The wing was equipped with plain leating-edge flaps and split trailing-edge flaps. For all conditions investigated the data indicate that an airplane corresponding to the model tested will possess static lateral and direc- tional stabfli

5、ty and that the ailerons till produce satisfactory maximum values of pb/2V. Full rudder deflection will be sufficient to balance the airplane to 8 of sideslip, As indicated by the measured fluctuating rolling moments, the airplane may possess undesirable rolling*ment characterfstice near and after t

6、he stall with the flaps fully deflected. INTRODUC!JKCOW . -0 The X-3 sir-plane, designed as a supersonic research airplane incor- porating suchfeatures as athinlowespect-ratiowinganda large fuselage, might be expected to present stabflfty problems fn low-speed flight. TbJ3low-speedlongIWd5nal chexac

7、teristics as measured with a O. there waa no internal flow. Due to themanner inwhIchthemodelwas constructed, it wee Impossible to teat the complete model with the canopy (fig. 7(a) end the air scoopa (ffg. 7(b) tit the further addftion of the tail boom caused a large reduc- tion in the instability o

8、f the fuselage, contributinS a yawLng mnrrvrnt equivalent to approxe the wing stall, for the modelwith the flaps fully deflected (ELF= 30 , 6TR= -50) were quite erratic as indicated by the strain-gage equipment for measuring rolling moments. Recourse was therefore made to the use of an oscil.lograph

9、 which recorded the output from the roUing+mmLen t strain gage as a faction of time. An attempt was made to determine the elec- tronic and mechanical characteristics of the experimental setup and thus the relationship betwe- the indicated and the actual oscillating rolling moms however, it has been

10、included to permit a better understanding of the measurements of the oscillating rolling moments, Figures 23 to 28 are reproductions of typical portions of the oscillograh records for various model configurations but do not nec- easarily present the maximum rolling+mnent coefficients observed for a

11、given model configuration. Fuselage along -Figures 23 and 24 show the variation of rolling- moment coefficient at l-6 angle of attack a small- amplitude osciU.ation became apparent with a frequency varying between tz c.roles and 60 cycles per second. As the angle of attack was increased a larger osc

12、illation of the rolUng+nomen t coefficient developed that dd an amplitude of approximately iO.02 and a frequency of approx- imately 3 to 6 cycles per second. With the vertical tail on the fuselage (fig. 24), the development of the oscillating rolling moments followed the same pattern, However, the a

13、mplitudes became increasingly larger, reaching values as high a8 AC1 = fO.07. It Is believed that the low- frequency oscillating rolling moments were caused by intermIttent dis- charge of vortices from the sides of the fuselage, possibly in a manner similar to that for bodies of revolution noted in

14、reference 9. With the vertical tail an the fuselage, the vortices impinged on the tail, increasing the amplitude of the rolling moments. Visual studies at low wind-tunnel speeds were made using two filaments of smoke. These obsez- vations indicated two vortices to be forming on the forward portion o

15、f the fuselage and discharging alternately from the sides of the fuselage at approrimately the potit of maxImam fuselage breadth. The vortex that was not being discharged appeared to decay and intermingle with the turbulentfuselagewake. Complete model - flaps neutral.- The data for the complete mode

16、l, tail off and tail on, are presented in figures 25 and 26 for angles of attack of 8.3O to 20.6O. Above 20.60 angle of attack the oscillations did not increase in amplitude, nor did the amplitude of the rolllng- moment oscillation for the model become as large as for the fuselage alone or for the f

17、uselage with the tail m. It is possible that the reason the oscillations were smaller was that the large wake from the stalled wing caused a rapid decay or breaking up of the vortices being discharged by the fuselage. Complete model -flaps fully deflected.- With the wing leading- and trailing-edge f

18、laps deflected 30“ and 50, respectively (figs. 27 and 28), large rolllngaomen t oscillPB Span,feet 1.$7 Root chord, feet 0.752 Tip chord, feet . 0.293 Sweep of 5O+ercent-chord line, degrees . Incidence (variable), degrees. 23 10 to -19 Meeerodynamic chord of the eqosed area, 0.521 Provided by IHSNot

19、 for ResaleNo reproduction or networking permitted without license from IHS-,-,-EACA RMA5lA16 TABLF: I.- COlEWDED Horizontal tail (concluded) Exposed area, square feet 0.701 Hinge line, percent of M.A.C. of exposed area 25 Tail length (frm 15 percent ting M.A.C. to horizontal-tail hinge line), feet

20、. 3.375 Height above fuselage reference line, feet . o .587 Vertical tail Area, square feet. . 0.678 Aspect ratio 1.3 Taperratfo . 0.25 Span,feet . 0.947 Root chord, feet 1.147 Tip chord, feet . 0.287 Height of root chord above fuselage reference line, feet 0.633 Sweep of wpercen-kchord line, degree

21、s . 0 Meanaerodynamic chord, feet 0.803 Tail length (from 15 percent wing M.A.C. to 25 percent vertical tall M.A.C.), feet 3.411 Rudder(Hlngelinenormaltofuselagereferenceline) Span,feet . 0.705 Tip chord, feet . 0.l62 Root chord, feet 0.227 Deflection, degrees . k2C Jettisonable-nose fins Area of ea

22、ch fin, square feet 0.084 Aspect ratio 0.75 Taperratio . 0.25 Span,feet . 0.253 Root, pounds per square foot 100 c . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 . , . Note .- Coefficients, angles, and control suri%ce dsfhctions are shown positi

23、ve. Figure I: Diugmmmatic skefch indicating the s/gn conventions used. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Fuselage reference /ine 42 08 E, /4.85- / t - I -Tr- Sta. 0 00 All dimensions wd stations in in -. - -. -.- -.-.f (a) Main lanning

24、gear and doors. Cb) lhx landing gear and door. Figure6 .-Details of the landing gear and the land-ear doors. Provided 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

25、-,-,-RAC!A J3M A5m6 27 . . (b) Air scoops. Figure 7.- D&am of the canopy, the air scoops, and the jettisonable-noss ftns. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-L . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(c) Jettisonabose fFns. Pigure 7 .- Concltied. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-