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Application of low drag-due-to-lift concept of linearized theory to the studies of twist and camber of isolated wing and wing-body configurations of supersonic aircraft
A method is presented for estimating the lift due to flap deflection at low angles of attack in incompressible flow. In this method provision is made for the use of incremental section-lift data for estimating the effectiveness of high-lift flaps. The method is applicable to swept wings of any aspect ratio or taper ratio. The present method differs from other current methods mainly in its ease of application and its more general application. Also included is a simplified method of estimating the lift-curve slope throughout the subsonic speed range.
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An iteration method is presented by which the detailed aerodynamic loading and twist characteristics of a flexible wing with known elastic properties may be calculated. The method is applicable at Mach numbers approaching 1.0 as well as at subsonic Mach numbers. Calculations were made for a wing-body combination; the wing was swept back 45 deg and had an aspect ratio of 4. Comparisons were made with experimental results at Mach numbers from.0.80 to 0.98.
Boeing and NASA are conducting a joint study program to design a wing flap system that will provide mission-adaptive lift and drag performance for future transport aircraft having light-weight, flexible wings. This Variable Camber Continuous Trailing Edge Flap (VCCTEF) system offers a lighter-weight lift control system having two performance objectives: (1) an efficient high lift capability for take-off and landing, and (2) reduction in cruise drag through control of the twist shape of the flexible wing. This control system during cruise will command varying flap settings along the span of the wing in order to establish an optimum wing twist for the current gross weight and cruise flight condition, and continue to change the wing twist as the aircraft changes gross weight and cruise conditions for each mission segment. Design weight of the flap control system is being minimized through use of light-weight shape memory alloy (SMA) actuation augmented with electric actuators. The VCCTEF program is developing better lift and drag performance of flexible wing transports with the further benefits of lighter-weight actuation and less drag using the variable camber shape of the flap.
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A ground loads test of a full-scale wing (175-ft span) was conducted using a fiber optic strain-sensing system to obtain distributed surface strain data. These data were input into previously developed deformed shape equations to calculate the wing s bending and twist deformation. A photogrammetry system measured actual shape deformation. The wing deflections reached 100 percent of the positive design limit load (equivalent to 3 g) and 97 percent of the negative design limit load (equivalent to -1 g). The calculated wing bending results were in excellent agreement with the actual bending; tip deflections were within +/- 2.7 in. (out of 155-in. max deflection) for 91 percent of the load steps. Experimental testing revealed valuable opportunities for improving the deformed shape equations robustness to real world (not perfect) strain data, which previous analytical testing did not detect. These improvements, which include filtering methods developed in this work, minimize errors due to numerical anomalies discovered in the remaining 9 percent of the load steps. As a result, all load steps attained +/- 2.7 in. accuracy. Wing twist results were very sensitive to errors in bending and require further development. A sensitivity analysis and recommendations for fiber implementation practices, along with, effective filtering methods are included
Tests of two wing-body combinations have been conducted in the Langley 19-foot pressure tunnel at a Reynolds number of 4 x 10(exp 6) and a Mach number of 0.19 to determine the effects of the bodies on the wing span load distributions. The wings had 45 degrees sweepback of the quarter-chord line, aspect ratio 8.02, taper ratio 0.45, and incorporated 12-percent-thick airfoil sections streamwise. One wing was untwisted and uncambered whereas the second wing incorporated both twist and camber. Identical bodies of revolution, of 10:1 fineness ratio, having diameter-to-span ratios of 0.10, were mounted in mid-high-wing arrangements. The effects of wind incidence, wing fences, and flap deflection were determined for the plane uncambered wing. The addition of the body to the plane wing increased the exposed wing loading at a given lift coefficient as much as 10 percent with the body at 0 degrees incidence and 4 percent at 4 degrees incidence. The bending-moment coefficients at the wing-body juncture were increased about 2 percent with the body at 0 degrees incidence, whereas the increases were as much as 10 percent with the body at 4 degrees incidence. The spanwise load distributions due to the body on the plane wing as calculated by using a swept-wing method employing 19 spanwise lifting elements and control points generally showed satisfactory agreement with experiment. The spanwise load distributions due to body on the flapped plane wing and on the twisted and cambered wing were dissimilar to those obtained on the plane wing. Neither of the methods of calculation which were employed yielded distributions that agreed consistently with experiment for either the flapped plane wing or the twisted and cambered wing.
The results presented in this note show the effect of twist and sweepback on the span load distribution over two monoplane wing models. The tests were made in the Atmospheric Wind Tunnel of the Langley Memorial Aeronautical Laboratory. The data are taken from the results of an investigation dealing primarily with lateral stability. As presented, they are suitable as an aid in the structural design of certain monoplane wings.
Off-design aerodynamic characteristics of swept arrow wings with twist and camber at supersonic velocities
This paper summarizes the results of a force, moment, and pressure experiment involving six thick, cambered and twisted, delta wings with 58 deg leading-edge sweep. This experiment was conducted in the NASA Langley 7- by 10-foot High-Speed Tunnel at Mach numbers of 0.75, 0.80, and 0.83. The design goal was a configuration which was self-trimming at a lift coefficient of 0.25 and Mach number of 0.80. Although the design goal was not met, the configuration which came closest and which had the best overall performance was selected for further study. Wing surface pressure data and limited surface oil flow data for this configuration are presented to show the extent of attached flow at the design point. For selected cases, inviscid solutions from vortex lattice method/suction analogy, PAN AIR, FLO-28, and FLO-57 are compared with the experimental force, moment, and pressure data.
A cambered and twisted triangular wing of aspect ratio 2 in combination with a cambered body was investigated experimentally to determine the effectiveness of the camber in reducing the drag due to lift at trim at supersonic speeds. Four arrangements were tested comprising all combinations of a symmetrical and a cambered wing with a symmetrical and a cambered body. The camber shape investigated was derived by linearized lifting surface theory for triangular wings with sonic leading edges and satisfied the requirement that the wing be trimmed at the design Mach number and lift coefficient. The experimental results for the cambered wing and cambered body showed that the drag coefficient at trim was always greater, at the same lift coefficient, than that for the untrimmed symmetrical wing and body. The trim lift coefficient was positive and decreased with increasing Mach number. At the design Mach number of 2.24, the trim lift coefficient was somewhat lower and the drag coefficient was higher than values predicted by linearized lifting surface theory for the wing alone. A comparison of the trim lift-drag ratio of the cambered wing and cambered body with values obtained by trimming the symmetrical wing and symmetrical body either with a canard or a trailing-edge flap showed that, at approximately the design Mach number the cambered configuration developed a somewhat higher value than the trailing-edge flap configuration but a lower value than the canard configuration.
The effects of a gas jet simulating a turbojet engine exhaust blowing above a cambered and twisted arrow wing were investigated. Tests were conducted in the Langley 4-foot supersonic pressure tunnel at a Mach number of 2.0. Nozzle pressure ratios from 1 to 64 were tested with both helium and air used as jet gases. The tests were conducted at angles of attack from -2 deg to 8 deg at a Reynolds number of 9,840,000 per meter. Only the forces and moments on the wing were measured. Results of the investigation indicated that the jet blowing over the wing caused reductions in maximum lift-drag ratio of about 4 percent for helium and 6 percent for air at their respective design nozzle pressure ratios, relative to jet-off data. Moderate changes in the longitudinal, vertical, or angular positions of the jet relative to the wing had little effect on the wing aerodynamic characteristics.
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This paper presents a jig twist optimization study of Mach 0.8 Transonic Truss-Braced Wing (TTBW) aircraft using FUN3D. The solver has been validated against wind tunnel test data. An aeroelastic model deformation correction method is developed for the wind tunnel model to improve the validation. The flight condition corresponding to Mach 0.8 and design lift coefficient 0.695 is selected for the TTBW aircraft jig twist optimization to reduce the drag coefficient. The static aeroelastic simulation module in FUN3D is used for the aeroelastic simulation. Three different orders of the Chebyshev polynomials are used as shape function to represent the variations of the jig twist distribution along the wing span. The preliminary optimization results show that the Mach 0.8 TTBW aircraft with the optimized jig twist achieves a drag reduction of about 5.7 counts at the design condition when 5 𝑡ℎ order Chebyshev polynomials are used as shape function.
A rapid shape parameterization tool called PROTEUS is developed for aircraft shape optimization. This tool can be applied directly to any aircraft geometry that has been defined in PLOT3D format, with the restriction that each aircraft component must be defined by only one data block. PROTEUS has eight types of parameterization schemes: planform, wing surface, twist, body surface, body scaling, body camber line, shifting/scaling, and linear morphing. These parametric schemes can be applied to two types of components: wing-type surfaces (e.g., wing, canard, horizontal tail, vertical tail, and pylon) and body-type surfaces (e.g., fuselage, pod, and nacelle). These schemes permit the easy setup of commonly used shape modification methods, and each customized parametric scheme can be applied to the same type of component for any configuration. This paper explains the mathematics for these parametric schemes and uses two supersonic configurations to demonstrate the application of these schemes.
A generic supersonic arrow-wing configuration has been extensively studied to develop a consistent experimental data base to aid in verifying analytical prediction methods. One of the wing configurations studied was a twisted and cambered wing with wing-mounted vertical fins and trailing-edge flaps. The pressure data obtained on that configuration was limited in the wing-fin juncture region, and no pressures were obtained on the fin itself. However, detailed data in the juncture region and on the fin are essential to check out the analytical prediction methods. Therefore, the model was modified to provide details of the juncture pressure field as well as pressures on the fin. The results of this experimental investigation showed that the addition of a fin helped improve the wing pressure distribution outboard of the fin. Deflection of the trailing-edge flap did not adversely affect this trend. The pressure distributions on the vertical fin were generally affected more by change in angle of attack than by flap deflection. The spanwise pressure distribution on the fin itself indicated a load reversal with angle of attack.
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