1、RE,SEARCH MEMORANDUM WIND-TUNNEL INVESTIGATION AT HIGH SUBSONIC SPEEDS OF SOME EFFECTS OF FUSELAGE CROSS-SECTION SHAPE AND WING HEIGHT ON THE STATIC LONGITUDINAL AND LATERAL STABILITY CRARACTERLSTICS OF A MODEL HAVING A 45O bWEPT WING, By Thomas J. King, Jr. . . . . ., . . ,- . TMS material contab M
2、ormatlonaffecting the National Defense of the United states within the meaning of the espionage laws, Title 18, U.S.C., Secs. 793 and 794, the trammission or revelation of which.in y: , manner to,an,u,rized.peTsonisprohibited plaw. . : . J .) . ., ,1 ,. ,., , . NATIONAL, ADVISORY COMMITTEE FOR AERON
3、AUTICS WASHINGTON February 3,1956 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L55J25 “ NATIONAL ADVISORY COMMITIEE FOR AERONAUTICS OF WIND-TUNNEL INVESTIGATSON AT HIGH SUBSONIC SPEEDS SOME EFFECTS OF FUSELAGE CROSS-SECTION SWE AND WING HE
4、IGHT ON THF: STATIC LONGITUDINAL AND LATERAT, STABILITY CHARACTERISTICS OF A MODEL HAVING A 45O SWEPT WING By Thomas J. King, Jr. SUMMARY An investigation was conducted in the Langley high-speed 7- by 10- foot tunnel at Mach nunibers from 0.80 to 0.92 to determine some effects of fuselage shape on t
5、he aerodynamic characteristics of a model having low and high wing arrangements. The results showed that when the cross section of a fuselage was changed from a circular to an essentially square shape, the location of the aerodynamic center for the wing-body combina- tion was moved forward. With the
6、 tail on, the high-wing model with the circular fuselage cross section had the most favorable variation of pitching moment over the lift-coefficient range. The directional stability was greatest for a low-wing configuration with a fuselage having a half-circular cross section on top and a half- squa
7、re cross section below. The square-fuselage configurations became directionally unstable at an angle of attack of about 12 with the wing in either high or low positipn; whereas the high-wing-circular-fuselage model became directionally unstable at an angle of attack of about 170 and the low-wing-cir
8、cular-fuselage model remained stable through the test angle-of-attack range. Fuselage cross section had little effect at low angles of attack on the effective dihedral derivative; but, at high angles of attack, the square fuselage provided considerably more effective dihedral than the circular fusel
9、age. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. I -.I - 2 INTRODUCTION The National Advisory Cormittee for Aeronautics is conducting wind- tunnel investigations to determine the aerodynamic characteristics of air- plane models with various arr
10、angements of the component parts. Some results of investigations at low speed have been reported in reference 1, at high subsonic speeds in reference 2, and at supersonic speeds in refer- ences 3 and 4. This paper presents results which show some effects of fuselage cross-section shape and wing heig
11、ht on the longitudinal aerodynamic characteristics and static lateral derivatives of a model having a 45 swept wing of aspect ratio 4, taper ratio 0.3, and with an NACA 65006 airfoil section in combination with a fuselage of fineness ratio 10.95. The test Mach nmiber range was from 0.80 to 0.92; the
12、 corresponding Reynolds numbers (based on wing mean aerodynamic chord) varied from 2.5 X 10 to 3.0 x 10 . 6 6 com1cms AND SYMBOLS The force and moment coefficients are presented about the stability axes system shown ir, figure 1. The pitching-moment and yawing-moment axes intersect on the fuselage c
13、enter line and are located 31.22 inches from the fuselage nose (longitudinal location of quarter-chordpoint of wing mean aerodynamic chord). CL lift coefficient, Lift 9s CD Cm pitching-moment coefficient, Pie ching moment sse side-force coefficient, Side force 9s Cn yawing-momen% coefficient, Yawing
14、 moment (2% rolling-moment coefficient, Rolling moment qsb 9 dynamic pressure, - p$, lb/sq ft Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I NACA RM L55J25 - V P S b F E v C free-stream velocity, ft/sec mass density of air, slugs/cu ft wing area,
15、2.25 sq ft wing span, 3 .OO ft wing mean aerodynamic chord, iLbI2 c2dy, 0.822 ft horizontal-tail mean aerodynamic chord, 0.388 ft vertical-tail mean aerodynamic chord, 0.757 ft local chord parallel to plane of symmetry, ft spanwise distance from plane of symmetry, ft Mach number angle of attack, deg
16、 angle of sideslip, deg 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I 1111 I 4 I I I Ill I I1 - MODELS AND APPARATUS A three-view drawing of the model is presented in figure 2 together with tables of the geometric characteristics of the wing an
17、d tail sur- faces. Coordinates of the fuselage profile and details of the fuselage cross-section shapes are given in figure 3. The corners of the rectangular-sided cross sections were rounded to a radius equal to 6.4 percent of the section width. The profiles of the fuselages were 1 identical for th
18、e three cross-section shapes (see fig. 3) but the half- circular-half-square and square cross-section areas were greater than the circular cross-section area by about 13 percent and 27 percent, respec- tively. A photograph of the low-wing-square-fuselage model mounted on the sting in the Langley hig
19、h-speed 7- by 10-foot tunnel is shown in figure 4. The chord plane of the wing was located on the fuselage 2.00 inches from the plane of the fuselage center line (fig. 2). The fuselage nose and center sections could be rotated 180 about the fuselage longitudinal axis to place the wing in a low or hi
20、gh position. The complete model, consisting of wing and fuselage with or without tail surfaces, was attached to the supporting sting (fig. 4) by a six-component internal strain-gage balance. The model forces and moments were measured by the balance and recorded automatically. TESTS The sting-support
21、ed model was tested in the Langley high-speed 7- by 10-foot tunnel over a Mach nuniber range from 0.80 to 0.92. The Reynolds rimer (based on wing mean aerodynamic chord) varied from about 2.5 x 10 to 3 .O X 10 . The angle of attack varied from -3 to a maximum of 24O; but as the Mach nuniber was incr
22、eased, the mexhum angle of attack was limited by balance loads or available tunnel power. With the wing in the low position, tests were made with the circular, half- circular-half-square, and square fuselage shapes. Tests were made on the circular and square fuselage shapes with the wing in the high
23、 position. Static longitudinal characteristics were obtained through the angle-of- attack range at f3 = 0. During the longitudinal tests of the circular fuselage, only the horizontal tail was removed. In the rest of the tail- off tests, including the lateral parameter tests, the horizontal tail as w
24、ell as the vertical tail was removed. Static lateral characteristics were obtained through the angle-of-attack range at nominal sideslip angles of i4. The static lateral stability parameters were computed at each angle of attack by taking the algebraic differences between Cn, Cy, and C2 at the two a
25、ngles of sideslip (*bo). These values were then 6 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L55J25 - 5 divided by the difference in sideslip angle which varied slightly from j the nominal value of 8O because of corrections to p due to
26、 deflection 1 of the balance and sting under load. I CORRECTIONS 1 Blocking corrections applied to Mach nuniber and dynamic pressure were determined by the method of reference 5. Jet-boundary corrections determined from reference 6 were applied to the angle of attack and drag. Corrections due to lon
27、gitudinal pressure gradient were applied to the drag data. No model-support tares have been applied to the results. Drag data have been adjusted to correspond to a pressure at the base of the fuselage equal to free-stream static pressure. 1 The angles of attack and angles of sideslip have been corre
28、cted for deflection of the sting support and balance. No attempt has been made to correct the data for aeroelastic deformation of the model as the correc- tions are believed to be small. (See ref. 7. ) PRESENTATION OF RESULTS The results of this investigation are presented in figures listed as follo
29、ws: Longitudinal characteristics of: Figure Low-wing-circular-fuselage combination 5 Low-wing-square-fuselage conbination . 7 High-wing-circular-fuselage codination . 6 High-wing-square-fuselage combination 8 Variation of C with Mach nuniber . 9 % Summary of effects of body shape and wing height , I
30、 on variation of Cm against CL at M = 0.80 . 10 1 Static lateral stability parameters of: Low-wing-circular-fuselage combination 11 Low-wing-half-circular-half-square-fuselage combination . 13 High-wing.-circular-fuselage combination . 12 Low-wing-square-fuselage combination . 14 High-wing-square-fu
31、selage combination 15 - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA RM L55525 Figure Comparison of the variation of Cys, Cnp, and C with 2P a at M = 0.80 . . . . . . . . . . ,. . . . . . . . . . . . . 16 Increments of static lateral deriva
32、tives due to tail . . . . . 17 DISCUSSION Longitudinal Stability Characteristics Fuselage cross-section shape and wing position had little effect on the variation of lift coefficient with angle of attack (figs. 5(a) to 8(a). The drag of the square-fuselage configurations near zero lift was, in gener
33、al, slightly higher than the drag of the circular-fuselage configuration, probably because of the larger volume of the square fuselage. The slopes of the pitching-moment curves against CL for circular- and square-fuselage models have been measured at zero lift and are pre- sented in figure 9. In gen
34、eral, the aerodynamic center moved rearward with increasing Mach nurriber for all configurations. The aerodynamic- center location of the circular-fuselage configuration (tail off) was from 1.0 to 2.5 percent of the mean aerodynamic chord more rearward than that of the square-fuselage configuration
35、except at the highest Mach number. The aerodynamic-center location of the circular-fuselage con- figuration with the tail on was about 2.0 to 3.0 percent of the mean aerodynamic chord more rearward than that of the square-fuselage con- figuration at all Mach nmibers. In reference 8, it is shown that
36、 the shape of the static pitching- moment curve is a primary factor affecting the dynamic pitch-up motions of an airplane. Examination of the pitching-moment curves of figures 5 to 8 indicates that at moderate lift coefficients, regions of decreased stability were present for all configurations inve
37、stigated. The pitching- moment curves of the circular-fuselage configurations (high and low wing positions) had less severe breaks than those of comparable square-fuselage configurations (fig. 10). The addition of the horizontal tail compensated a large part of the unstable breaks for both fuselage
38、shapes with the wing in the high position; the stabilizing effect of the horizontal tail was not as strong on the low-wing configurations. In general, the complete model with the high wing and circular fuselage had the most favorable variation of pitching moment with lift over the Mach number range
39、investigated. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Lateral Stability Characteristics Low-wing configurations.- Fuselage cross-section shape had large effects on the lateral stability characteristics of,the low-wing models, particularly at
40、angles of attack above about 4. Comparison curves showing the variation of CyP, Cnp, and C with angle of attack at M = 0.80 are presented in figure 16. A decrease in the directional stability of the square-fuselage configuration resulted from the decrease in the increment of Cy due to the tail. From
41、 figure 16 it is seen that the square-fuselage configuration (tail on) became directionally unstable at a = 12. The value of Cn and the increment in due to the tail at M = 0.80 (figs. 16 and 17) were larger for the half-circular- half-square-fuselage than for either the circular- or the square-fusel
42、age configuration. In general, for the three low-wing configurations tested, variation in Mach number from 0.80 to 0.92 produced slight improvements in directional stability characteristics. 2P P P CnP In the low-angle-of-attack range, fuselage cross-section shape had little effect on C . For all co
43、nfigurations the variation of C with low and moderate angles of attack increased slightly with increase in Mach number. At angles of attack above approximately bo, the variation of C with a became markedly nonlinear and behaved in the .manner 2P 28 2P described in reference 9 relating to swept wings
44、. At angles of attack above loo, AC (fig. 17) became positive for the circular- and half- circular-half-square-fuselage configurations but remained negative for the square-fuselage model. 2P High-wing configurations.- The change in wing position from low to high had little effect on the angle of att
45、ack at which the square- fuselage configuration (tail on) became directionally unstable; although, as has been shown in other inve-stigations, changing the wing position from low to high on the circular-fuselage configuration (tail on) re- sulted in a significant deterioriation in directional stabil
46、ity, partic- ularly at high anglesof attack (fig. 16). At low angles of attack, raising the wing produced the usual reduction in AC for all config- urations. For the high-wing configurations there was little effect of fuselage cross-section shape on the increment in due to the tail. c% At low angles
47、 of attack, about the same increase in effective dihedral position for either the circular-.or square-fuselage configurations. At high angles of attack, the square-fuselage model had considerably higher effective dihedral than the circular-fuselage model. (-%) resulted from raising the wing from a l
48、ow to a high Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 CONCLUSIONS, An investigation was made to determine the aerodynamic characteris- tics at high subsonic speeds of a wing of aspect ratio 4, taper ratio 0.3, sweep of 45O, and with an NACA
49、65006 airfoil section mounted in a law and a high position on fuselages of fineness ratio 10.95 with circular, half- circular-half-square, and square cross-section shapes. The results of this investigation indicate the following conclusions: 1. The configurations with the circular-fuselage cross sections generally had the more rearwa
