1、SECURITY INFORMATION I8 RESEARCH MEMORANDUM 4 t I f I SPFED STATIC STABILITY AMD CONTROL CEAIUCTERISTICS OF i-SCAIX MODEL OF TE BELL X-1 AIRPLANE EQUIPPED 7 kt, WITH A 4-PERCENT-THICK, ASPECT-RATIO-4, ioh $C UNSWEPT WING I d i By William C. Moseley, Jr., and Robert T. Taylor Langley Aeronautical Lab
2、oratory Langley Field, Va. :NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS = WASHINGTON November 2, 1953 . . . .- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1s NMA RM L53H27 c NATIONAL ADVISORY Co“ITEZ FOR AERONAUTICS RESEARCH MEMORANDUM LOW-SPEED
3、STATIC STABILITY AND COFTROL CHARAI=IZRISTICS OF A 1- SCAL;E MODEL OF THE BXLG X-1 ALRpLANE EQUIPPED 4 WITE A k“PERCm-TffICK, ASPECT-RATIO-49 By William C . Moseley, Jr. , md Robert T. Taylor SUMMARY An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the low-speed sta
4、tic stability and control charac- teristics of a A - scale model of the Bell X-1 airphne equipped with a above CL = 0.6, there was Sf = oo , and S, = 60), it should be rehenibered that, although the tail length was the same, the mean aerodynamic chord was larger for the present model. flap-retracted
5、 condition. The data of figure 8(b) indicate that the elevator was sufficient to maintain a lift coefficient of 0.70 at an angle of attack of U0, a stabilizer setting of it = 3.8O, and 8 20 elevator deflection. The data of figure 8 show the aeroaynamic characteristics for the The data of figure 9 sh
6、ow the aeroaynamic characteristics for the landing condition. The data of figure g(b) indicate that the elevator was sufficient to maintain a lift coefficient of 1.04 at an angle of attack of 6 with a horizontal-tail setting of 2.4O and an elevator deflection of -5O. Figure 10 summrizes the longitud
7、inal-control characteristics of the model and indicates a moderate change in elevator deflection or sta- bilizer setting is required to balance the model through most of the lift- coefficient range. High Lift Devices The plain-wing tail-off lift-curve slope (fig. 5) was about 0.068 and campares well
8、 with the theoretical wing-alone lift-curve slope (0.64) as determined fran reference 3. It is felt that the contribution of the fuselage is significant in raising the lift-curve slope of the model and may account for stme of the discrepancy between experiment and theory. The lift data of figure ll
9、indicate that the lift-curve slope of the complete model varied between 0.074 at Sf = o to 0.066 at 6f = wo. The variation of lift coefficient with flap deflection (fig. 12) indicates tWt the opthum flap deflection is 35* for the slot tested. This flap deflection yields an increment in lift coeffici
10、ent of 0.64 at an me of attack of bo. Although the flap at ot = 0 and a = k0 yields 8, slightly higher lift coefficient for flap deflections above 35O the cor- responding increase in drag is excessively high. The values of C Lmax presented are peak values or were taken just beyond the point where an
11、 abrupt decrease in lift-curve slope occurred for flap deflections where no definite peak existed. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I_ TBe landing speed of the X-1 airplane can be determined if an empty weight of TOO0 pounds -is assume
12、d for the airplane. (All unexpended fuel t is jettisoned before landing. 1 From figure g(b) a balancedolift coeffi- cient of about“l.03 is obtained for an engle of attack of 4 . The wing loading would be 53.8 pounb per square foot and the landing apeed would be about 145 miles per hour. v Lateral an
13、d Directional Stability The data of figure 15 indicae that the model possessed both static lateral and static directional stability that was generally fairly con- stant with increase in lift coefficient. The stability parameters obtained fraan figures 13 and 14 agree very well with the values in fig
14、- ure 15 obtained at p = tj0. The effective dihedrd tended .to increase near the bigher lift coefficients for the flaps-retracted configuration, while for the flaps-deflected configuration the effective -dihedral decreased slightly near the highest lift coefficient. The data for the original X-1 mod
15、el with an aspect-ratio-6, 10-percent-thick wing yielded values generally the same at low lift coefficients but abrupt variations in effective dihedral occurred as the stall was approached. Deflecting the slotted flap had little effect on the static directionaJ- stability of the model. Lateral and D
16、frectional Control The data presented in figures 16 and 17 are for the right aileron deflected only. The tests were limited to approximate aileron travel on the actual airplane, Sa = +XF lo . For the flap-retracted configuration (fig. 16), the aileron retained its effectiveness up to about a = loo w
17、hich is near the stall. (See fig. 5. ) Above cc = loo, the aileron began to lose its effectiveness particularly for positive or dam deflections. The aileron rolling effectiveness paameter was about -0.00127, which is about what would be expected for 0.256 flap-type control with a blunt overhang of 0
18、.20. From the data of reference 4, an estimated value of Cz was -0.0012. The data for the flap-deflected configuration (fig. 17) show that the aileron effectiveness parameter of the original model was larger than that of the present model, = -0.00157 and % 6 % % = -O.OOU7, respectively; however, fig
19、ure 18 shows the present model to have a rolling velocity per degree of aileron deflection approximately 20 percent higher than that of the original X-1 model. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2s NACA RM L53H27 “.T, The aileron binge m
20、oments indicate that for the flap-retracted con- figuration the hinge-moment parameters were cha = -o.oo, c = -0.- a 4% and C = -0.0083. From reference 5, estimated dues (flap retracted), ha, U with 0.20 blunt overhang at law Mach nunbers. In order to compare estimated flight conditions with those o
21、f the original X-lmodel, two flight conditions were selected and hta for the conditions are given in table I. Condition I is for a heavily loaded, flap-retracted configuration, operating at a high Uft coefficient. Con- dition I1 simulates an empty, flap-deflected, landing configuration. Figure 18 sh
22、ow6 the rolling velocity, *el force, and wing-tip helix angle plotted against aileron deflection at steady roll for conditions I and I1 for the model of the present investigatiop and condition I for the original X-1 model. The data indicated that, for a given aileron deflection, the present model ha
23、s a stick force approximately 70 percent higher than that of the original X-1 model. However, for a given value. of the rolling velocity, the present model has a wheel force about 4.0 per- cent higher than that of the original X-1 model. No attempt was made to evaluate the effect of deflecting the s
24、lotted flap on the damping-in-roll coefficients used in these calculations. . The data of figures 19 and 20 indicate that the rolline;-mment- coefficient, yawing-moment-coefficient, and lateral-force-coefficient curves were generally linear for angles of sideslip of about Oo. The data of figure 21 s
25、how that the rudder and ailerons were capable of trbuing the model through sideslip yles of this deflection caused an increment in lift coefficient of 0.64 for an aagle of attack of ko . 3. The model possessed static directional stability through the range investigated and this stability was general
26、ly unaffected by deflecting the slotted flap. The effective dihedral.ms positive and generally con- stant throughout the angle-of-attack range and was only slightly affected by deflecting the slotted flaps. 4. The effectiveness of the rudder through the deflection range investigated (CLO) was adequa
27、te to trim the model through a sides.p range of +loo flaps retracted and *lo for flaps deflected 35*. 5. The aileron effectiveness was satisfactory through the stall for both the fleps-retracted and flaps-deflected conditions although a loss in effectiveness was present near the staU. I+ Langley Aer
28、onautical Laboratory, National Advisory Connnittee for Aeronautics, Langley Field, Va., August 18, 1953. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-XACA RM L53H27 11 1. Gillis, Clarence L., Polhamus, Edwud C., and Gray, Joseph L., Jr.,: Charts
29、 for Determining Jet-Boundaxy Corrections for Camplete Models in 7- by 10-Foot Closed Rectesgularr Wind Tunnels. NACA WR L-123, 1945. (Formerly NACA ARR L5G31) 2. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flaw Closed Throat Wind Tunnels, With Consideration of the Effect of Compres
30、sibility. WA Rep. 995, 1-950. (Supersedes NACA FM A7M. ) 3. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings With Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN 1491, 1947. 4. Lowry, John G., and Schneiter, Ledlie E.: Estimation of Effectiveness of Flap-Type Contr
31、ols on Sweptback Wings. NACA TN 1674, 1948. 5. Langley Reseazch Staff (Cmgiled by Thomas A. Toll).: SummazY of Lateral-Control Research. NACA Rep. 868, 1947. (Supersedes XACA TN 1245.) Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- NACA RM L53H27 T
32、ABLE: I.- FLIGKT CONDITIONS INVESTIGATED Condition I Flap deflection, Ef, deg . 0 Lift coefficient, CL . 0.70 Velocity, mph 240 Dynamic pressure, lb/ft2 . 148 Mach n-er . 0.32 Weight, lb 13,488 Altitude, f+ . 0 Condition I1 35 7000 1.03 0 143 52-5 0.19 =-%7 Provided by IHSNot for ResaleNo reproducti
33、on or networking permitted without license from IHS-,-,-NACA RM L53H27 Figure 1.- System of axe6 showing forces, moments, and angles. Positive values indicated by arrow heads. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-“-3420“-1 TABULATED DATA W
34、iiy Area,lbtal 8.125sqff Area, aileron 0.476sqft Area, slofhd flap L 104 sq ft span 5.69 ff 6a = 0; 6r = 0; p = 00. n_l. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L53H27 - 48 44 t5ro -36 .32 28 CD .24 20 ./6 . /2 .08 .04 0 -4 72 0 .2 4
35、.6 .8 LO Figure 5. - Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 .3 MACA RM L53E27 Figure 6.- Effect of-stabilizer setting on the aerodynamic characteristics - ir pitch of a l/ Sa = Oo; 6r = Oo; s, = 00; p = 00. “L Provided by IHSNo
36、t for ResaleNo reproduction or networking permitted without license from IHS-,-,-48 NACA RM L53H27 Figure 6.- Concluded. ” .2 4 .6 .8 LO 12 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 -2 0 .2 B .6 .8 LO 1.2 . I Figure 7.- Variation of static ma
37、rgin with Uft coefficied for the l/LSC8 6, = Oo; 8, = Oo; p = Oo. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 _I NACA RM L53H27 (a) Concluded. Figure 8.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA FM L53H27 - .3 20 . /2 4 -4 -8 -.6 -4 -2 0 :z 4 .6 .8 LU CL (b) it = 38. Figure 8. - Continued. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
