NASA NACA-RM-L52G07-1952 Wind-tunnel investigation of the low-speed static and rotary stability derivatives of a 0 13-scale model of the Douglas D-558-II airplane in the landing co.pdf

1、RESEARCH MEMORANDUM WIND-TUNNEL INVESTIGATION OF “HE LOW-SPEED STATIC AND ROTARY STABILITY DERIVATIVES OF A 0.13-SCALE MODEL OF Tm D3UGLAS D-558-IC ALRPLANh IN THE LANDING CONFIGURATION Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1v . RESFARCH ME

2、MORANDUM WIND-TUNNEL INYESTIGATION OF THE LOW-SPEED STATIC AM3 ROTARY STABILITY DEUVA!IIVES OF A 0. U-SCATX MODEL OF THE DOUGLAS D-558-11 AIRPIJUG3 IN THE LANDIE CONFIGURATION By M. J. QueiJo and Evalyn G. Welb SUMMARY A wind-tunnel investigation has been made to determine the low- speed static and

3、rotary stability derivatives of a 0.13-scale model of the Douglas D-558-11 airplane in the lending configuration. The Eft coefficient of the mdel varied Unearly with angle of attack up to a maximum lift coefficient of 1.24 which occurred at an angle of attack of U0. The Ufi-curve slope was about 0.0

4、6 per degree in this range. The model was longitudinallystable in the angle-of-attack range from Oo to 16O, with a static margin of about 16 percent of the wing mean aerodynamic chord over most of this range. The model was approximately neutrally stable near an angle of attack of Il0. The directiona

5、l stability of the model decreased slowly with increase in angle of attack up to an angle of attack of about l3O. At higher angles, the stability deteriorated more rapidly. The yawing moment due to rolling velocity was negative throughout the angle-of- attack range, and the magnitude of the tail co-

6、ntribution to this moment near zero angle of attack indicated a stronger sidewash effect for the flapped wing than generally has been obtained for plain wings. The derivatives associated with yawing flow were nearly constant for angles of attack from Oo to about uO, but varied considerably at higher

7、 angles. INTRODUCTION Various investigations have shown that the dynamic lateral stability characteristics of high-speed aircraft are critically dependent on Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM L52G07 certain mass and aerodynami

8、c parameters. and, hence, that reliable estimates ofthe dynamic stability of such aircraft catl be made only if these parameters are determined accurately. The static derivatives of an airplane can be determined accurately by 113811s of conventional wind-tunnel tests of a model; however, only a few

9、facilities are avail- able for measuring rotary (rolling and ming) derivatives. The Langley stability tunnel, which is equipped with facilities for simulating rolling and yawing flaw, was qtilized to make available measured low- speed static and-rotary derivatives of a mdel of the Douglas D-558-11 a

10、irplane in the landing configuration (slats, flaps, and landing gear extended). The measured low-speed parameters of the same model with slats, flaps, and landing gear retracted are given in reference 1. SYMBOLS AND COEFFICIENTS The data presented herein are in the form of standard IUCA coef- ficien

11、ts of forces and moments which are referred to the system of stability axes (fig. 1) with the origin.at the projectiun of the quarterrchardpoint of the wing mean aerodysamlc cbord on the plane of symmetry. This syEltem of axes is defined as an orthogonal system having the origin at the center of gra

12、vity and in which the Z-axis is in the plane of symmetry and perpendicular to the relative wind, the X-axis i8 in the plane of symmetry and perpendicular to the Z-axis, and the Y-axis is perpendicular to the phne of symetry. Positive directions of forces, moments, and displacements-are sham in figur

13、e 1. b wing span, ft C local wing chord, parallel to plane of symnetry, ft C mean aerodynamic chord, ft - P rolling angular velocity, radians/sec 9 dynamic pressure, $, lb/sq ft 2 r yawfng angular velocity, raaans/aec S wing area, sq ft V free-stream velocity, ft/sec a angle of attack, deg P sides l

14、ip angle, radians 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. D. Y angle of clirib, deg 4f angle of yawy deg # angle of roll, deg P mass density of air, slugs/cu ft D drag, lb L lift, lb Y side force, lb M pitching moment, ft-lb N yawing mome

15、nt, ft-lb 2 rolling moment, ft-lb CD drag coefficient, D/qS CL lift coefficient, L/qs CY side-force coefficient, Y/QS cm pitching-mment coefficient, M/qS Cl rolling-moment coefficient, Z/qSb Cn yawing-mment coefficient, N/q% 3 - c- however, the degree of stability generally decreased with increase i

16、n angle of attack. The effective-dihedral parameter C.lp was approximately the same for the wing-fuselage combination as it was for the complete mdel, and generally increased negatively with an increase in angle of attack. Characteristics in Rolling Flow The aerodynamic derivatives of the model in s

17、imulated roll are shown in figure 7 a5 curves of Cyp, Cnp, and Czp plotted against . angle of attack. Addition of the tail surfaces to the wing-fuselage combination produced a negative increment of Cnp. From geometric considerations, such as those of reference 6, a positive increment to would have b

18、een expected near a = Oo from addition of the tail cnp Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 surfaces. The negative increment actually obtained probably is caused by changes in flow angularity at the tall. plane, associated wfth wing wake

19、 characteristics (ref. 7). It appears that deflection of the flaps has a rather powerful effect. on the flow angularity at the. tazl since, in investigations with models having plain wings, the positive tail contribution to Cnp at a = Oo, although reduced by wing wake angu- larity, has not been made

20、 negative (for example, 6ee refs. 7, 8, or 9). Characteristics in Yawing Flow The yawing-flow parameters Cyr, Cnr, and Cz are plotted against r angle of atkack in figure 8. These parameters remained approximately constant for angles of attack up to about. U0 for the model with the tail surfaces on o

21、r off. At angles of attack greater than 13, the pkrameters varied over a rather large range with increase in angle of attack. CONC WSIONS An investigation was made in the Langley stability tunnel to deter- mine the low-speed static and rotary stability derivative8 of a 0.13-scale model of the Dougla

22、s D-558-11 airplane in the-laiding configuration. The results of the Fnveatigation have led to the following conclusions: 1. The lift coefficient varied linearly with angle of attack up to a maximum lift coefficient of 1.24, which occurred at an angle of 8ttaCk of 13. The lift-curve slope was about

23、0.06 per degree in this range. 2. The model had static lo itudinal stability in the angle-of- attack range from Oo to about 16 3 with a static margin of about 16 per- cent of the wing mean aerodynamic chord over most of the range. The model was approximately neutrally stable near an angle of attack

24、of 1l0. 3. The directional stability of the mdel decreased slowly with increase in aagle of attack up to about 13. At higher angles of attack, the directional stability generally decreased more rapidly. 4. The yawing moment due to roll C was negative throughout the nP angle-of-attack range. The magn

25、itude of the tail contribution to Cn at low angles of attack indicated a stronger sidewash effect in roll for the flapped wing than generally hag beep obtained for plain wings. P Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L5xO7 7 5. The

26、.derivatives associated with- yawing flow were about constant in the angle-of-attack range from Oo to U0, and varied considerably with angle of attack above Uo. Langley Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. Provided by IHSNot for ResaleNo reproduction

27、 or networking permitted without license from IHS-,-,-8 NACA RM L52GO7 REFERENCES 1. QueiJo, M. J., and Goodman, Alex: Calculations of the DynWc Lateral Stability Characteristics of the Douglas D-558-11 Airplane in High-Speed-Flight. for .Various Wing had. Reynolds number, 1,100,000. Provided by IHS

28、Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L5xO7 - 15 0 L .I 0 4 8 /2 /6 20 24 Angle of afhck, a3, deg Figure 6. - Variation of Cyp, Cnp, and C with a Tor a 0.13-scale model of the D-3S-11 airplane in the landing configuration. Mach number, 0.17; Reyno

29、lds number, 1,100,000. 28 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-11 c G 2P - NACA RM L52GO7 0 4 8 12 16 20 24 Angle of aftuch, a;, de9 Figure 7.- Variation of Cyp, C and C2 with a for a O.13-sale np P model of the D-598-11 airplane in the la

30、nding canffguration. Mach number, 0.17; Reynold6 number, 1,100,000. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-3v a -.4 fr -“ -3 - .L 0 4 8 /2 16 20 24 Angle of attack, zl deg Figure 8. - Variation of Cy Cnr and C 2, with a for a 0.13-scale mode

31、l of the D-558-11 airplane In the landing configuration. Mach number, 0.13; Reynolds nuniber, 863,000. rJ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-