Showing papers on "Wing root published in 1980"

Divergence of Forward Swept Composite Wings

Abstract: Forward swept wing aircraft may have superior aerodynamic performance for certain missions. Algebraic expressions to predict the static aeroelastic divergence characteristics of forward swept wings constructed of composite materials have been developed using a laminated box beam model to describe the wing structure and aerodynamic strip theory to predict the loads due to wing bending and torsional deformation. The expressions presented show that, because of elastic coupling between wing bending and torsion, wing divergence may be precluded for reasonably large forward sweep angles if the composite structure is properly tailored. The structural parameters that maximize divergence speed are readily identified. Two illustrative examples are presented. ao b

Aeroelastic Tailoring Studies in Fighter Aircraft Design

Abstract: Studies have been conducted on the use of the directional properties of composite material to provide design improvements for fighter aircraft. The TSO (Aeroelastic Tailoring and Structural Optimization) computer program, developed by the Air Force Flight Dynamics Laboratory, was used in these investigations. The configurations evaluated covered a wide spectrum of fighter aircraft aerodynamic surfaces, including the F-15 composite wing, a preliminary design horizontal tail, a prototype aircraft wing, and a future conceptual aircraft wing. Both drag reduction and increased roll effectiveness, with no weight cost, are predicted for the F-15 composite wing. A unique minimum weight design is shown for the preliminary design horizontal tail, in which the anisotropic characteristics of the composite material are used to perform the dual function of strength and flutter balance weight. A more conventional optimum flutter solution, based upon outer panel wing root pitch restraint increases, is shown for the prototype aircraft wing. Finally, significant wing twist, offering potential aerodynamic benefits, can be obtained on the conceptual aircraft wing.

Wing flapping with minimum energy

Abstract: For slow flapping motions it is found that the minimum energy loss occurs when the vortex wake moves as a rigid surface that rotates about the wing root - a condition analogous to that determined for a slow-turning propeller. The optimum circulation distribution determined by this condition differs from the elliptic distribution, showing a greater concentration of lift toward the tips. It appears that very high propulsive efficiencies are obtained by flapping.

Flight Measurements of a Wing Tip Vortex

Abstract: Detailed flight measurements of the wing tip vortex generated by a DeHavilland Beaver were made with instruments mounted on a sailplane, which was towed behind the aircraft. Data at three different downstream locations were obtained by using three different tow rope lengths. The measurement stations were 4, 8, and 16 span lengths behind the aircraft, and the vortex Reynolds number was slightly above T0/v=W6. Results of the tests differed considerably from those of previous authors. Although the customary similarity variables for vortex decay were verified, the velocities were a factor of 2.5 higher than previous results. The vortex core radius and turbulence levels also proved to be substantially smaller than earlier measurments. These results indicate that the final similarity state of the vortex may depend strongly upon the initial tip shape, initial turbulence levels of the vortex, and/or the ambient turbulence levels. a /Si b c(orC) CL k1,k2,k3 r ( or R )

Results of design studies and wind tunnel tests of high-aspect-ratio supercritical wings for an energy efficient transport

Abstract: These basic characteristics of critical wings included wing area, aspect ratio, average thickness, and sweep as well as practical constraints on the planform and thickness near the wing root to allow for the landing gear. Within these constraints, a large matrix of wing designs was studied with spanwise variations in the types of airfoils and distribution of lift as well as some small planform changes. The criteria by which the five candidate wings were chosen for testing were the cruise and buffet characteristics in the transonic regime and the compatibility of the design with low speed (high-lift) requirements. Five wing-wide-body configurations were tested in the NASA Ames 11-foot transonic wind tunnel. Nacelles and pylons, flap support fairings, tail surfaces, and an outboard aileron were also tested on selected configurations.

Active Controls for Flutter Suppression and Gust Alleviation in Supersonic Aircraft

Abstract: Application is made in the present paper of the recently developed relaxed aerodynamic energy concept and synthesis techniques to the definition of appropriate active control systems for the low-speed flutter model of the B-2707-300 supersonic cruise airplane. The effectiveness of the resulting activated systems is analytically tested for flutter suppression, wing root bending moment alleviation, and ride control (fuselage accelerations). The results obtained indicate that considerable increase in flutter speeds can be obtained by the various control systems, using a single trailing-edge control. In all cases, the flutter suppression control system led to a substantial reduction in both wing root bending moments and in fuselage and wing accelerations.

Theoretical study of nonadiabatic boundary-layer stabilization times in a cryogenic wind tunnel for typical stainless steel wing and fuselage models

Abstract: The time varying effect of nonadiabatic wall conditions on boundary layer properties was studied for a two dimensional wing section and an axisymmetric fuselage. The wing and fuselage sections are representative of the wing root chord and fuselage of a typical transport model for the National Transonic Facility. The analysis was made with a solid wing and three fuselage configurations (one solid and two hollow with varying skin thicknesses) all made from AISI type 310S stainless steel. The displacement thickness and local skin friction were investigated at a station on the model in terms of the time required for these two boundary layer properties to reach an adiabatic wall condition after a 50 K step change in total temperature. The analysis was made for a free stream Mach number of 0.85, a total temperature of 117 K, and stagnation pressures of 2, 6, and 9 atm.

Application of optimal control to structural load alleviation control systems

Abstract: Feedback laws based upon optimal control theory were derived. and these resulted in a reduction of the structural loads on the wing of a simulated aircraft. Various models of the aircraft dynamics were used. the most complete being of order 79. This model included rigid body motion, structural flexibility effects, unsteady aerodynamics, gust dynamics and actuator dynamics. The structural effects were characterised by the first fifteen bending modes. The subject aircraft studied, was considered to employ active ailerons and elevators and was subjected to manoeuvre commands and simulated atmospheric turbulence. Extensive numerical tests have shown that feedback laws derived from reduced dimension models performed comparably with the feedback law based on the most complete model. Tests were made on feedback laws ranging from order 79 to order 5. It was. however. not possible to reduce the number of feedback variables below five as this then affected the stability of the aircraft. The law based upon five state variable feedback was given the designation 'safety law'. One of the consequences of operating under the action of the 'safety law' was that the same level of load reduction could not be achieved as was obtained whenever a full state feedback law was employed. In addition 'safety law' operation was often marked by large transient oscillations of the wing root bending moment and it was considered that this would subsequently affect the fatigue life of the structure. An observer design was then investigated which reconstructed the complete state vector from a selection of measurements of the sensor signals appropriate to the 'safety law'. Results have shown that it is possible to achieve a practical implementation of such a scheme which will possess all the attendant advantages of full state feedback control. A consequence of reducing the strength of the wing of the aircraft as a result of employing an active load alleviation scheme is that a considerable degree of reliability of the control system, higher than that of both the basic airframe and its propulsive system, will be required. Because the use of hardware redundancy techniques as a means of providing the required degree of reliability would be expensive, software redundancy techniques suggest an attractive alternative. One example of how software redundancy may be employed is demonstrated in respect of , checking the analogue feedback gain controller used in the aircraft to implement linear feedback. It is shown how a-microprocessor may be effectively employed to introduce a surrogate gain should one or more of the channels of the controller fail.