1、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-rt it tt s National Aeronautics and Space Administration Scientific and Technical 1979 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-An investigation
2、was conducted in the Langley V/STOL tunnel to determine the static longitudinal and lateral-directional aerodynamic characteristics of an advanced high-aspect-ratio supercritical-wing transport model equipped with a full-span leading-edge slat and part-span double-slotted trailing-edge flaps. This w
3、ide-body transport model was also equipped with spoiler and aileron control surfaces, flow-through nacelles, landing gear, movable hori- zontal tails, and interchangeable wing tips with aspect ratios of 10 and 12. The model was tested with leading-edge slat and trailing-edge flap combina- tions repr
4、esentative of cruise, climb, take-offF and landing wing configu- rations. The tests were conducted at free-stream conditions corresponding to Reynolds numbers (based on mean geometric chord) of 0.97 to 1.63 x lo6 and corresponding Mach numbers of 0.12 to 0.20, through an angle-of-attack range of -2O
5、 to 24O and a sideslip-angle range of -loo to 5O. The test results show that, for the aspect-ratio-1 0 wing configurations, the cruise wing had a maximum trimmed lift coefficient of 1.23 and a lift-drag ratio of 16.24; the climb wing, a lift coefficient of 1.93 and a lift-drag ratio of 11.03; the ta
6、ke-off wing, a lift coefficient of 2.29 and a lift-drag ratio of 9.82; and the landing wing, a lift coefficient of 2.47 and a lift-drag ratio of 7.00. The aspect-ratio-12 take-off and landing wing configurations had only slightly higher maximum lift coefficient and lift-drag values than those for th
7、e corresponding aspect-ratio-10 wing configurations. Also, for the aspect-ratio-10 wing configurations, the climb wing had less lateral-directional stability than the cruise wing, and the take-off and landing wing configurations had almost identical stability at the same lift coefficient. The take-o
8、ff and landing wing configurations also had greater lateral-directional stability than the cruise wing configuration. INTRODUCTION In recent years, NASA has been actively involved in an aeronautical research project to improve the energy efficiency of modern wide-body jet transport aircraft. The Air
9、craft Energy Efficiency (ACEE) project was formu- lated to encourage industry participation and to coordinate the industry and NASA research efforts. One element of the ACEE project is the Energy Efficient Transport (Em) program which is concerned primarily with the developnent of advanced aerodynam
10、ics and active-controls technology for application to deriv- ative or next-generation transport aircraft. A part of the EET program has been the development, by NASA Langley Research Center personnel, of advanced supercritical wings with greater section thickness-cmrd ratios, higher aspect ratios, h
11、igher cruise lift coefficients, and lower sweepback than the conven- tional wings of current transports. These supercritical wings have been tested extensively in the Langley wind tunnels to determine their high-speed cruise performance characteristics (refs. 1 and 2). Because of their high cruise l
12、ift Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-coefficients and high aspect ratios, these wings could be smaller and more fuel efficient than currently used wings if take-off and landing requirements could be met The present investigation was co
13、nducted to determine the low-speed per- formance characteristics of a representative high-aspect-ratio supercritical wing equipped with a conventionally sized high-lift flap system. The model tested was a 3.66-m (12.0-ft) span model of an advanced long-range wide-body jet transport with cruise wing
14、and fuselage dimensions similar to those of the NASA SCW-2a supercritical wing tested in the Langley 8-foot transonic pres- sure tunnel and reported in references 1 and 2. This wing had an aspect ratio of 12r a 27O quarter-chord sweep, and streamwise supercritical airfoil sections that varied in max
15、imum thickness-chord ratio from approximately 0.15 at the wing root to 0.10 at the tip. The high-lift flap system consisted of a part-span double-slotted trailing-edge flap and a full-span leading-edge slat. The trailing-edge flap consisted of a large vane and small aft flap combination, as opposed
16、to the more conventionally used small vane and large flap combinations. Vane-flap combinations similar to the combination used on this model have recently been under development by several aircraft manufacturers and have achieved maximum two-dimensional lift coefficients approaching those of more co
17、mplex triple- slotted flap combinations. The model was also equipped with inboard high-speed ailerons, outboard low-speed ailerons, two wing-mounted flow-throiigh nacelles, landing gear, movable horizontal tails, and interchangeable wing tips with aspect ratios of 10 and 12. The model was tested in
18、the Langley V/STOL tunnel with wing leading-edge slat and trailing-edge flap combinations representative of cruise, climb, take- off, and landing wing configurations. The model was instrumented with a six- component strain-gage balance to measure aerodynamic forces and moments and with chordwise pre
19、ssure taps at three spanwise stations to determine represen- tative wing and flap loads. The pressure data obtained from this investigation are presented in graphic and tabular form in references 3 and 4. This report presents and discusses the static longitudinal and lateral-directional aero- dynami
20、c data obtained during this investigation. SYMBOLS AND ABBREVIATIONS The longitudinal forces and moments presented in this report are referenced to the stability-axis system and the lateral forces and moments to the body-axis system. The moment data are referred to a moment center located in the mod
21、el plane of symmetry at a point 50.14 cm (19.74 in.) longitudinally aft of the wing leading edge and 6.60 cm (2.60 in.) vertically below the wing reference plane. The longitudinal location of the moment center corresponds to the quarter-chord point location of the mean geometric chord of the aspect-
22、ratio-12 wing. The aerodynamic coefficient data for the aspect-ratio-1 2 wing configurations are based on a wing reference area of 1 .ll m2 (12 ft2) and a reference span of 3.66 m (12.0 ft), and the coefficient data for the aspect-ratio-10 wing configu- rations are based on a wing reference area of
23、1.04 m2 (11 .21 ft2) and a refer- 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-ence span of 3-23 m (70,59 ft). These reference wing areas are based on the trapezoidal planform which extends from the m el center line to the wing tip, The aspect-r
24、atio-12 wing had a mean geom hord of 33-02 cm (13,OO in.) and the aspect-ratio-10 wing, 34-90 cm (13,7 1 ; however I an approximate average value of 34-04 cm (13.40 in.) was used as the reference mean geometric chord for both the aspect-ratio-10 and aspect-ratio-12 wing configurations. This average
25、value also corresponded to the value of the local wing chord at the wing trailing-edge break station, All measurements and calculations were made in the U.S. Customary Units. Values presented herein are given in the International System of Units (SI), with the equivalent values in U.S. Customary Uni
26、ts given parenthetically. b2 A aspect ratio, - S b wing span, m (ft) C local streamise wing chord, cm (in.) E reference mean geometric chord, cm (in.) Drag drag coefficient, - ( CD qs CD Lift CIS lift coefficient, - ( CL CL CZ rolling-moment coefficient, tables effective dihedral parameter c% c, Cn
27、in computer-generated tables) in computer-generated tables) Rolling moment (CRM in computer-generated qSb based on increment of Cz between Pitching moment pitching-moment coefficient, (CPM in computer- qsc generated tables) Yawing moment yawing-moment coefficient, (CYM in computer-generated qSb tabl
28、es) 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-“6 C CY *m *L it - L/D L.E. M 9 B directional stability parameter based on increment of Cn between ac, Side force 9s side-force coefficient (CSF in computer-generated tables) side-force parameter
29、based on increment of Cy between = -loo 2, longitudinal stability parameter (CiiCL in computer-generated tables) incidence of horizontal tail, positive for leading edge up, deg (ISUB1 in computer-generated tables) lift-drag ratio (L/D in computer-generated tables) leading edge free-stream Mach numbe
30、r (MACH in computer-generated tables) free-stream dynamic pressure, kPa (lb/ft2) (Q in computer-generated tables) - free-stream Reynolds number based on c wing reference area, m2 (ft2) trailing edge wing th ickness-chord rat io angle of attack of model reference center line, positive nose up, deg (A
31、LPHA in computer-generated tables) angle of sideslip of model reference center line, positive nose left, deg (BETA in computer-generated tables) aileron deflection angle, positive for trailing edge down, deg flap deflection angle, positive for trailing edge down, deg slat deflection angle, positive
32、for trailing edge down, deg spoiler deflection angle, positive for trailing edge up, deg Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6, vane deflection anglep positive for trailing edge down, deg E downwash angle at horizontal tail, deg rl nondim
33、ensional wing semispan location Subscripts: cor r corrected R left max maximum r right Abbreviation: W.R.P. wing reference plane MODEL DESCRIPTION The model tested during this investigation had a 3.66-m (12.0-ft) span and was representative of an advanced long-range wide-body jet transport with crui
34、se wing and fuselage dimensions scaled from those of the NASA SCW-2a high-aspect- ratio supercritical model developed at the NASA Langley Research Center and reported in reference 1. The wing on this model was equipped with convention- ally sized low- and high-speed ailerons, full-span leading-edge
35、slat, and part- span double-slotted trailing-edge flaps. A drawing showing the overall model components is presented in figure l(a), and a detailed wing planform layout of the control and flap surfaces in figure l(b). Photographs of the model installed in the Langley V/STOL tunnel are shown in figur
36、e 2. The pertinent model geometric characteristics are summarized in table I. Detailed wing and flap component surface coordinates are given in reference 3. The model was fabricated with aluminum wings and glass fiber fuselage and empennage for minimal deflections at the design conditions of a maxim
37、um tunnel dynamic pressure of 2.87 kPa (60.0 lb/ft2) and a maximum wing lift coefficient of 3.0. The empennage consisted of movable horizontal tails without elevators and a fixed vertical fin without a rudder. The horizontal tails were mounted to the model with a geared, pivoting bracket that allowe
38、d for incidence angles from -15O to 15O in 5O increments. The model was also equipped with two wing- mounted, flaw-through nacelles with scaled external dimensions similar to those of a typical high-bypass-ratio (approximately 6).turbofan engine used on current wide-body jets. A third tail-mounted n
39、acelle was not simulated on this model because of the expected minimal influence this nacelle would have on the per- formance of the flap system. The model was also equipped with simulated land- ing gear and doors attached to the wing and fuselage underside near the nose. The wheel-well cavities wer
40、e not simulated on this model. 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The basic cruise was designed with an aspect-ratio-1 2 trapezoidal planform which extend om the model center line to the wing tip and had 27O quarter-chord sweep. starte
41、d at the r = 0,383 wing semispan station and increased the chord at the center line by 0 percent. The wing was fabricated with a removable tip sec- tion which could be easily replaced with a shorter tip to produce an aspect- ratio-10 wing planform. The aspect-ratio-12 wing had streamwise supercritic
42、al- wing sections with maximum thickness-chord ratios of 0.144 at the side-of- body semispan location (r = 0,096), 0.12 at the trailing-edge break station (17 = 0.3831, and 0.10 at the wing tip (Tl = 1.0). The wing was mounted to the fuselage with 5O dihedral and -lo of incidence at the wing root. T
43、he wing had an inboard trailing-edge extension that The wing was fabricated with removable leading- and trailing-edge segments. The cruise wing segments could be removed easily and replaced with a leading- edge slat and trailing-edge lap/aileron segments. The trailing-edge flap and low- and high-spe
44、ed aileron surface areas were sized and positioned spanwise for the aspect-ratio-10 wing based on a comparative analysis of several exist- ing designs for lower aspect-ratio-6 to aspect-ratio-8 transport wings. The aspect-ratio-12 flap system was obtained by simply extending the outboard leading-edg
45、e slat and the outboard low-speed aileron segments. Four wing con- figurations were possible with either the aspect-ratio-10 or aspect-ratio-12 tips. These configurations were: (1) cruise, with slat, vane, and flap nested; (2) climb, with slat deflected -50 and vane and flap nested; (3) take- off, w
46、ith slat deflected -50 and vane and flap deflected 15O; and (4) landing, with slat deflected -50 and vane and flap deflected 30. wing configurations were used during a majority of the tests because the flap and aileron surface areas were better proportioned than the corresponding areas for the aspec
47、t-ratio-1 2 wing The aspect-ratio-10 Control and Flap Systems The trailing-edge flap was a part-span inboard and outboard double-slotted flap that consisted of an advanced design large vane and small aft flap combi- nation as compared with the mre conventionally used small vane and large aft flap co
48、mbinations. Advanced designs similar to this combination have recently been under development by several aircraft manufacturers and have experimen- tally achieved maximum two-dimensional lift coefficients approaching those achieved by the more complex triple-slotted flap systems. The structural load
49、s produced by this flap combination are less severe than those of the conven- tional combinations because a greater percentage of the total vane/flap loads are generated by the more closely coupled large vane component. The flap segments were equipped with spoilers that could be deflected to either 45O or 60, which are primarily ground-spoiler and speed-brake deflec- tions. A simple hin
