1、i LONGITUDINAL STABILITY AND CONTROL CHARACTERISTICS OF A POWERED MODEL OF STOL AIRPLANE CONFIGURATION A T WIN-PROPELLER DEFLECTED-SLIPSTREAM by Richard J. Margason, Alexander D. Hammond, and Garl L. Gentry Ldngley Research Center LangZey Station, Hampton, Va. , $ I 1 ,: ,1 z. 1 i NATIONAL AERONAUTI
2、CS AND SPACE ADMINISTRATION WASHINGTON, D. C. JULY 1966 1 F Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- TECH LIBRARY KAFB, NM LONGITUDINAL STABILITY AND CONTROL CHARACTERISTICS OF A POWERED MODEL OF A TWIN-PROPELLER DEFLECTED-SLIPSTREAM STOL AIR
3、PLANE CONFIGURATION By Richard J. Margason, Alexander D. Hammond, and Gar1 L. Gentry Langley Research Center Langley Station, Hampton, Va. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sole by the Clearinghouse for Federol Scientific and Technical Information Springfield, Virginia 22151 - Price
4、$4.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LONGITUDINAL STABILITY AND CONTROL CHARACTERISTICS OF A POWERED MODEL OF A TWIN-PROPELLEE DEFLECTED-SLIPSTREAM STOL AIRPLANE CONFIGURATION By Richard J. Margason, Alexander D. Hmond, and Gar1 L. G
5、entry Langley Research Center SUMMARY Results are presented of a wind-tunnel investigation of the static longi- tudinal stability and control capabilities of a twin-propeller deflected- slipstream STOL airplane in the take-off and landing speed range through the post-stall region at angles of attack
6、 up to 44O. The results of this investigation show that the magnitudes of the pitching moments of the wing-body combination for the flaps-retracted (0 flap deflection) configuration were small. The tail-on data show that any of the tail configu- rations with the flaps-retracted configuration provide
7、 an adequate stability contribution and are capable of trimming the airplane. The wing-body combina- tion for the flaps-deflected (451 flap deflection) configuration had a large tail lift requirement for longitudinal trim, particularly for the highest power setting of the investigation. The small ta
8、il in the high position had the capability of trimming the airplane for the low power conditions (thrust coeffi- cients of 0 and 0.70). At higher power conditions the tail stalled before trim was achieved. The large tail in either position had the capability of trimming the airplane up to the angle
9、of attack corresponding to the maximum lift coeffi- cient for all but the highest power condition (thrust coefficient of 2.42). This tail stalled before trim was achieved for the highest power condition. INTRODUCTION Recent interest in developing a small deflected-slipstream short take-off and landi
10、ng (STOL) airplane has led to a need for stability and control data on this type of configuration. A static wind-tunnel investigation of a powered model of a twin-propeller deflected-slipstream STOL aircraft configuration was conducted to provide some of this information. The lateral control charact
11、er- istics of this model have been presented in reference 1. The longitudinal stability and control characteristics are presented in the present report. This investigation was undertaken to determine the longitudinal stability and control characteristics through the angle-of-attack range from -bo in
12、to the post-stall region (to 44O). Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The investigation was conducted in the 17-foot (5.18 meter) test section of the Langley 3OO-MPH 7- by 10-foot tunnel and covered two flap deflections and several power
13、 conditions. SYMBOLS The units used for the physical quantities defined in this paper are given both in the U.S. Customary Units and in the International System of Units (SI). Factors relating these two systems of units are presented in reference 2. The symbols used are defined as follows: wing chor
14、d, 1.29 feet (0.39 meter) drag coefficient, Drag qs Lift lift coefficient, - qs maximum trimmed lift coefficient lift-curve slope, per degree pitching-moment coefficient referred to model moment, center at wing quarter-chord (c/4), Pitching moment ,(see fig. 1) qsc propeller thrust coefficient based
15、 on free-stream velocity and wing T area, - (often designated in literature as Tc) qs propeller thrust coefficient based on slipstream velocity and T propeller disk area, - WJSp propeller diameter, feet (meters) height of the horizontal-tail chord above the wing chord, feet (meters ) tail incidence,
16、 degrees tail length measured horizontally from the wing quarter-chord to the horizontal-tail quarter-chord, feet (meters) number of propellers Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-dynamic pressure, e?, pounds/foot2 (newtons/meter 2 ) 2 sl
17、ipstream dynamic pressure, q + -, T pounds/foot2 (newtons/mete$) Nsp propeller disk area, tail area, foot (meter21 d, foot2 (mete2) 4 wing area, 9.04 foot2 (0.84 meter2) total propeller thrus5, pounds (newtons) free-stream velocity, feet/second (meters/second) nondimensional horizontal-tail volume,
18、- St - It sw distance measured along airfoil chord line from the leading edge, feet (meters ) distance measured perpendicular from airfoil chord line to airfoil lower surface, feet (meters) distance measured perpendicular from airfoil chord line to airfoil upper surface, feet (meters) angle of attac
19、k, degrees deflection of movable surface (with subscript to denote surface deflected) , degrees downwash angle at the horizontal tail, degrees downwash angle at the horizontal tail when wing angle of attack is zero, degrees air density, slugs/foot3 (kilogram/meter3) Subscripts: f flap (see fig. 3) t
20、 tail V vane (see fig. 3) 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MODEL AND APPARATUS A three-view drawing of the model is presented in figure 1 and photo- graphs are presented in figure 2. a l5.3O-inch (0.39 meter) chord, was unswept, and
21、had a span of 7.00 feet (2.13 meters) with an aspect ratio of 5.42. faired wooden blocks fastened to a metal spar which supported the two motor nacelles and the fuselage strongback as well as the brackets which held the flap system. The wing had an NACA 44-13 airfoil section, The wing contour was fo
22、rmed with The double-slotted high-lift flap system consisted of a 20-percent-wing- chord vane with a St. Cyr 156 airfoil section and a 40-percent-wing-chord flap with a modified Rhode St. Genese 35 airfoil section over the forward 30 percent of its chord faired into the wing airfoil section over the
23、 rear 70 percent of its chord. The flap and vane ordinates, as well as the flap and vane positions when deflected, are given in figure 3. Two different horizontal tails were tested. Both had an aspect ratio of 3-15 and an NACA 4415 airfoil section whose profile was modified to give a 9 percent maxim
24、um thickness and were mounted inverted to provide an inverse camber. The two tails had different areas, spans, and chords. (See fig. 1.) The small tail was tested in a high position (ht = 0.94) only; the large tail was tested in both the high and the low position (ht = 0.15). Both tail posi- tions w
25、ere above the wing-chord plane. For the tail configurations tested, the nondimensional horizontal-tail volumes V, are the following: small tail in the high position, 0.85; large tail in the high position, 1.15; large tail in the low position, 1.04. Additional data on the geometric characteristics ar
26、e also presented in figure 1. Since no directional stability tests were included in the investigation, the vertical tail served only as a support for the horizontal tail. cal surface consisted of a sheet of 1/2-inch (1.27 centimeters) aluminum with a rounded leading edge and a beveled trailing edge.
27、 The verti- The three-blade propellers were made of balsa covered with glass-fiber cloth and were driven by water-cooled variable-frequency electric motors oper- ated in parallel from a variable-frequency power supply, which kept the motor speeds matched within 20 revolutions per minute. propeller w
28、as determined by a stroboscopic indicator which received the output frequency of small alternators connected to each motor shaft. tests the right propeller rotated in a clockwise direction and the left pro- peller rotated in a counterclockwise direction when viewed from the rear of the model. The th
29、rust coefficient was varied by changing the wind-tunnel speed. The speed of rotation of each For all the During the tests the speed of rotation was maintained at 6000 rpm. The motors were mounted inside aluminum-alloy nacelles by means of strain- gage beams so that the propeller thrust could be meas
30、ured. The total lift, longitudinal force, pitching moment, rolling moment, yawing moment, and side force were measured by a strain-gage balance mounted to the fuselage at the wing quarter-chord. Only longitudinal components of the data are presented in 4 Provided by IHSNot for ResaleNo reproduction
31、or networking permitted without license from IHS-,-,-this report. in reference 1. The results of lateral control tests of this model are presented TESTS AND CORRECTIONS The investigation was made in the 17-foot (5.18 meter) test section of the Langley 3OO-MPH 7- by 10-foot tunnel. For the powered te
32、sts the free- stream dynamic pressure was varied from about 1.3 to 5.3 pounds/foot2 (72 to 254 newtons/meter2), depending on the desired thrust coefficient. The slip- stream dynamic pressure was relatively constant at about 7.5 pounds/foot2 (359 newt;ons/meter2) for all thrust coefficients. A free-s
33、tream dynamic pres- sure of about 6.0 pounds/foot2 (287 newtons/meterZ) was used for the propeller- off tests. For the powered tests the Reynolds number, based on wing chord, of the flow in the slipstream averaged about 0.65 X lo6; for the propeller-off tests the Reynolds number in the free stream a
34、veraged about 0.58 X lo6. Since errors due to blockage, slipstream contraction, and tunnel-wall effects have been found to be small for models of this size in the l7-foot test section (ref. 3), no corrections for these types of error have been applied to the data. The propeller thrust data have been
35、 presented as the conventional thrust coefficient, that is, thrust nondimensionalized by free-stream dynamic pressure times wing area obtained by removing the propellers from the model. the thrust was measured by strain-gage beams at the motors. cients based on these measurements are presented with
36、the basic data. the motor rotation speed was held constant, the thrust increased as the angle of attack of the model increased; as a result, the thrust coefficients are not constant for a particular test. For convenience the average values of thrust coefficients near zero angle of attack for the dat
37、a presented in this report (used as reference values throughout the report) are listed in the following table : (CT = T/qS). In all cases a thrust coefficient of zero was For the propeller-on data The thrust coeffi- Although 45 Reference value of - I T,s 0 .20 -39 0 .50 .64 .78 I 0 17 .44 :is 1 2.42
38、 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-It is often desirable to use the propeller thrust coefficient based on slip- stream velocity and propeller disk area. Figure 4 is a plot of the relation between these two thrust coefficients for the
39、model tested. EULTS AND DISCUSSION The results of a wind-tunnel investigation of the longitudinal control and stability characteristics of a model of a twin-propeller deflected-slipstream STOL airplane are presented in the following figures: Figure Basic data: Flaps retracted, 6f = 0: Tailoff . 5 La
40、rge tail, high position 9 to 11 Large tail, low position . 12 to 14 Tailoff 15 Small tail, high position 16 to 1.9 Large tail, high position 20 to 23 Large tail, low position . 24 to 27 Small tail, high position 6 to 8 Flaps deflected, 6f = 45: . Comparisons : Effect of tail area 28 to 30 Effect of
41、tail height . 3lto 32 Basic Data The basic data figures present the variation of the lift and pitching- moment coefficients with angle of attack and the variation of drag and pitching- moment coefficients with lift coefficient. In addition, the variation of thrust coefficient with angle of attack is
42、 presented for the propeller-on tests. pitching-moment coefficients for all the data are presented about the wing quarter-chord line. The angle of attack used in this investigation ranged from The -4O to 44O. The basic data for the configuration with theohorizontal tail off are pre- sented in figure
43、 5 for the flaps-retracted (tif = 0 ) configuration and in fig- ure 15 for the flaps-deflected (6f = 45) configuration. The lift-curve slope Cs, as well as the maximum lift increases with increasing power. The following table gives a summary of the lift-curve slope and of the maximum lift coeffi- ci
44、ents (model with the horizontal tail off) for both of the flap deflections: 6 . I,.,.- . -_ . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.- CL ,max 1.12 1.83 2.11 2.47 4.40 5.35 6.48 I; 0.067 075 0.068 .l50 tion tral The tail-off pitching moment
45、s for the flaps-retracted (6f = 00) configura- are unstabl-e for angles of attack up to wing stall and are generally neu- beyond that angle of attack (fig. 5). The small magnitude of the pitching moments for this configuration shows that the horizontal tail is required mainly to provide stability fo
46、r the flaps-retracted (6f = 00 The tail-on data (figs. 6 to 14) show that any of the tail configurations tested with the flaps-retracted configuration provide an adequate stability contribution and, in addition, are capable of trimming the airplane throughout the lift-coefficient range of the invest
47、igation. configurations. ) In contrast to the flaps-retracted (6f = Oo) configuration, the flaps- deflected (6f = 45O) configuration requires a large increment of pitching moment for trim. The tail-off data of figure 15 show that nose-up increments of pitching-mom2nt coefficient ranging from approxi
48、mately 0.5 to 1.1 are needed, depending on the power condition; therefore, a large down load is required at the tail position. force coefficient as large as -0.95 must be developed by the large horizontal tail in the high position, assuming no loss in dynamic pressure at the tail. This value approac
49、hes the maximum normal-force coefficient for this tail. requirement for tail normal-force coefficient is even more severe for the other two tail configurations tested because of their smaller tail volumes. In order to satisfy this requirement, a normal- The The data for the flaps-deflected (6f = 45O) configuration with each of the several horizontal
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