SEARCH · Search NASA
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
This brief survey of the problems encountered in high-altitude flight deals in particular with the need for high lift coefficient in the wings, large aspect ratios in the wings, and also the problem of hermetically sealing the cabin.
The primary objective of this study was to determine whether Return To Launch Site (RTLS) abort at T seconds along the launch trajectory of the Personnel Launch System (PLS) is possible using optimal control theory. The secondary objective is to assess effects of bank angle constraint, lift coefficient constraint, free and fixed final boundary conditions, etc. of the vehicle. The PLS is a complementary system to the Space Shuttle.
Aims for improvement of fighter aircraft pursued by the unsteady flow community are high agility (the ability of the aircraft to make close turns in a low-speed regime) and super maneuverability (the ability of the aircraft to operate at high angles of attack in a post stall regime during quick maneuvers in a more extended speed range). High agility requires high lift coefficients at low speeds in a dynamic situation and this requirement can be met by dynamically forced separation or by quasistatic stall control. The competing methods will be assessed based on the known physics. Maneuvering into the post stall regime also involves dynamic separation but because even fast maneuvers involving the entire aircraft are 'aerodynamically slow' the resulting dynamic vortex structures should be considered 'elicited' rather than 'forced.' More work seems to be needed in this area of elicited dynamic separation.
Results are presented which define unsteady flow conditions associated with the high-dynamic structural response of a high-aspect-ratio, elastic, supercritical wing at transonic speeds. The wing was tested in the Langley Transonic Dynamics Tunnel with a heavy gas test medium. The supercritical wing, designed for a cruise lift coefficient of 0.53 at a Mach number of 0.80, experienced the high-dynamic structural response from Mach 0.90 to 0.94 with the maximum response occurring at about Mach 0.92. At the maximum response conditions of the wing, the forcing function appears to be the oscillatory chordwise movement of strong shocks located on the upper and lower surfaces of the wing in conjunction with the flow separation on the lower surface of the wing in the trailing-edge cove region.
The viability of applying a state-of-the-art Euler code to calculate the aerodynamic forces and moments through maximum lift coefficient for a generic sharp-edge configuration is assessed. The OVERFLOW code, a method employing overset (Chimera) grids, was used to conduct mesh refinement studies, a wind-tunnel wall sensitivity study, and a 22-run computational matrix of flow conditions, including sideslip runs and geometry variations. The subject configuration was a generic wing-body-tail geometry with chined forebody, swept wing leading-edge, and deflected part-span leading-edge flap. The analysis showed that the Euler method is adequate for capturing some of the non-linear aerodynamic effects resulting from leading-edge and forebody vortices produced at high angle-of-attack through C(sub Lmax). Computed forces and moments, as well as surface pressures, match well enough useful preliminary design information to be extracted. Vortex burst effects and vortex interactions with the configuration are also investigated.
VLMD computer program determines subsonic mean-camber surfaces of trimmed noncoplanar planforms with minimum vortex drag at specified lift coefficient. Up to two planforms designed together. Method used that of subsonic vortex lattice method of chord loading specification, ranging from rectangular to triangular, left specified by user. Program versatile and applied to isolated wings, wing/canard configurations, tandem wing, and wing/-winglet configuration. Written in FORTRAN.
To identify planform characteristics which have promise for a highly maneuverable vehicle, an investigation was conducted in the Langley Subsonic Basic Research Tunnel to determine the low-speed longitudinal aerodynamics of 21 planform geometries. Concepts studied included twin bodies, double wings, cutout wings, and serrated forebodies. The planform models tested were all 1/4-in.-thick flat plates with beveled edges on the lower surface to ensure uniform flow separation at angle of attack. A 1.0-in.-diameter cylindrical metric body with a hemispherical nose was used to house the six-component strain gauge balance for each configuration. Aerodynamic force and moment data were obtained across an angle-of-attack range of 0 to 70 deg with zero sideslip at a free-stream dynamic pressure of 30 psf. Surface flow visualization studies were also conducted on selected configurations using fluorescent minitufts. Results from the investigation indicate that a cutout wing planform can improve lift characteristics; however, cutout size, shape, and position and wing leading-edge sweep will all influence the effectiveness of the cutout configuration. Tests of serrated forebodies identified this concept as an extremely effective means of improving configuration lift characteristics; increases of up to 25 percent in the value of maximum lift coefficient were obtained.
Despite the 80-year history of classical wing theory, considerable research has recently been directed toward planform and wake effects on induced drag. Nonlinear interactions between the trailing wake and the wing offer the possibility of reducing drag. The nonlinear effect of compressibility on induced drag characteristics may also influence wing design. This thesis deals with the prediction of these nonlinear aspects of induced drag and ways to exploit them. A potential benefit of only a few percent of the drag represents a large fuel savings for the world's commercial transport fleet. Computational methods must be applied carefully to obtain accurate induced drag predictions. Trefftz-plane drag integration is far more reliable than surface pressure integration, but is very sensitive to the accuracy of the force-free wake model. The practical use of Trefftz plane drag integration was extended to transonic flow with the Tranair full-potential code. The induced drag characteristics of a typical transport wing were studied with Tranair, a full-potential method, and A502, a high-order linear panel method to investigate changes in lift distribution and span efficiency due to compressibility. Modeling the force-free wake is a nonlinear problem, even when the flow governing equation is linear. A novel method was developed for computing the force-free wake shape. This hybrid wake-relaxation scheme couples the well-behaved nature of the discrete vortex wake with viscous-core modeling and the high-accuracy velocity prediction of the high-order panel method. The hybrid scheme produced converged wake shapes that allowed accurate Trefftz-plane integration. An unusual split-tip wing concept was studied for exploiting nonlinear wake interaction to reduced induced drag. This design exhibits significant nonlinear interactions between the wing and wake that produced a 12% reduction in induced drag compared to an equivalent elliptical wing at a lift coefficient of 0.7. The performance of the split-tip wing was also investigated by wing tunnel experiments. Induced drag was determined from force measurements by subtracting the estimated viscous drag, and from an analytical drag-decomposition method using a wake survey. The experimental results confirm the computational prediction.
An investigation has been conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel in order to further the development of semi-span testing capabilities. A twin engine, energy efficient transport (EET) model with a four-element wing in a takeoff configuration was used for this investigation. Initially a full span configuration was tested and force and moment data, wing and fuselage surface pressure data, and fuselage boundary layer measurements were obtained as a baseline data set. The semi-span configurations were then mounted on the wind tunnel floor, and the effects of fuselage standoff height and shape as well as the effects of the tunnel floor boundary layer height were investigated. The effectiveness of tangential blowing at the standoff/floor juncture as an active boundary-layer control technique was also studied. Results indicate that the semi-span configuration was more sensitive to variations in standoff height than to variations in floor boundary layer height. A standoff height equivalent to 30 percent of the fuselage radius resulted in better correlation with full span data than no standoff or the larger standoff configurations investigated. Undercut standoff leading edges or the use of tangential blowing in the standoff/ floor juncture improved correlation of semi-span data with full span data in the region of maximum lift coefficient.
The problem of efficient wind tunnel testing for multi-element airfoils was first addressed by the author during a previous ASEE fellowship. A modern three element model with internal actuators to position a flap in two degrees of freedom was designed and later built. Some preliminary testing proved that the approach was viable. The purpose of this summer's work was to fully develop experimental methods including efficient data acquisition. The final goal is to develop dense data sets for both lift and drag measurements as a function of flap position for both take-off and landing configurations. The model has a span of 36 in. and chord of 18 in. and is currently being fitted for a 3 ft. x 4 ft. low speed wind tunnel. The flap was reworked to allow all pressure taps to function after initial tests showed two blocked ports. The serial method of obtaining pressures from the surface taps was found to be exceedingly slow so a new method using 12 pressure transducers and a 12 port parallel scanning valve were developed. A new automated data acquisition and control algorithm was developed using LabView software and a PC platform. Flow two-dimensionality is currently under investigation with boundary layer control by blowing; this was previously omitted for initial testing. By the end of the summer a detailed data set (uncorrected) consisting of lift coefficient versus flap position for the landing configuration should be available.
A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integml turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the larninar flow toward the desired amounl An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.
This report presents results from the first icing tests performed in the Modem Airfoils program. Two airfoils have been subjected to icing tests in the NASA Lewis Icing Research Tunnel (IRT). Both airfoils were two dimensional airfoils; one was representative of a commercial transport airfoil while the other was representative of a business jet airfoil. The icing test conditions were selected from the FAR Appendix C envelopes. Effects on aerodynamic performance are presented including the effects of varying amounts of glaze ice as well as the effects of approximately the same amounts of glaze, mixed, and rime ice. Actual ice shapes obtained in these tests are also presented for these cases. In addition, comparisons are shown between ice shapes from the tests and ice shapes predicted by the computer code, LEWICE for similar conditions. Significant results from the tests are that relatively small amounts of ice can have nearly as much effect on airfoil lift coefficient as much greater amounts of ice and that glaze ice usually has a more detrimental effect than either rime or mixed ice. LEWICE predictions of ice shapes, in general, compared reasonably well with ice shapes obtained in the IRT, although differences in details of the ice shapes were observed.
The unsteady, three-dimensional Navier-Stokes equations coupled with the Euler equations of rigid-body dynamics are sequentially solved to simulate and analyze the aerodynamic response of a high angle of attack delta wing undergoing oscillatory motion. The governing equations of fluid flow and dynamics of the multidisciplinary problem are solved using a time-accurate solution of the laminar, unsteady, compressible, full Navier- Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The primary model under consideration consists of a 65 deg swept, sharp-edged, cropped delta wing of zero thickness at 20 deg angle of attack. In a freestream of Mach 0.85 and Reynolds number of 3.23 x 10(exp 6), the flow over the upper surface of the wing develops a complex shock system which interacts with the leading-edge primary vortices producing vortex breakdown. The effect of the oscillatory motion of the wing on the vortex breakdown and overall aerodynamic response is detailed to provide insight to the complicated physics associated with unsteady flows and the phenomenon of wing rock. Forced sinusoidal single and coupled mode rolling and pitching motion is presented for the wing in a transonic freestream. The Reynolds number, frequency of oscillation, and the phase angle are varied. Comparison between the single and coupled mode forced rolling and pitching oscillation cases illustrate the effects of coupling the motion. This investigation shows that even when coupled, forced rolling oscillation at a reduced frequency of 2(pi) eliminates the vortex breakdown which results in an increase in lift. The coupling effect for in phase forced oscillations show that the lift coefficient of the pitching-alone case and the rolling-moment coefficient of the rolling-alone case dominate the resulting response. However, with a phase lead in the pitching motion, the coupled motion results in a non-periodic response of the rolling moment. The second class of problems involve releasing the wing in roll to respond to the flowfield. Two models of sharp-edged delta wings, the previous 65 deg swept model and an 80 deg swept, sharp-edged delta wing, are used to observe the aerodynamic response of a wing free to roll in a transonic and subsonic freestream, respectively. These cases demonstrate damped oscillations, self-sustained limit cycle oscillations, and divergent rolling oscillations. Ultimately, an active control model using a mass injection system was applied on the surface of the wing to suppress the self-sustained limit cycle oscillation known as wing rock. Comparisons with experimental investigations complete this study, validating the analysis and illustrating the complex details afforded by computational investigations.
A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integral turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the laminar flow toward the desired amount. An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.
Studies have been made on several wing leading-edge modifications applicable at present to single-engine light aircraft, which produce stabilizing vortices at stall and beyond. These vortices have the effect of fixing the stall pattern of the wing such that the various portions of the wing upper surface stall nearly symmetrically. The lift coefficient produced is maintained at a high level to angles of attack significantly above the stall angle of the unmodified wing, and the divergence in roll usually is reduced to a controllable level. It is hypothesized that these characteristics will help prevent inadvertent spin entry after a stall. Results are presented from recent large-scale wind-tunnel tests of a typical light aircraft, both with and without the modifications. The data indicate (hot the static stall and poststall characteristics of this aircraft, in a typical landing-approach condition, are noticeably improved when it suitable leading-edge modification is employed; and also that no appreciable aerodynamic penalties are evident in the normal flight envelope.
A flight investigation has been conducted to study how pilots use the high lift available with blowing-type boundary-layer control applied to the leading- and trailing-edge flaps of a 45 deg. swept-wing airplane. The study includes documentation of the low-speed handling qualities as well as the pilots' evaluations of the landing-approach characteristics. All the pilots who flew the airplane considered it more comfortable to fly at low speeds than any other F-100 configuration they had flown. The major improvements noted were the reduced stall speed, the improved longitudinal stability at high lift, and the reduction in low-speed buffet. The study has shown the minimum comfortable landing-approach speeds are between 120.5 and 126.5 knots compared to 134 for the airplane with a slatted leading edge and the same trailing-edge flap. The limiting factors in the pilots' choices of landing-approach speeds were the limits of ability to control flight-path angle, lack of visibility, trim change with thrust, low static directional stability, and sluggish longitudinal control. Several of these factors were found to be associated with the high angles of attack, between 13 deg. and 15 deg., required for the low approach speeds. The angle of attack for maximum lift coefficient was 28 deg.
A two-dimensional wind-tunnel investigation of the pressure distributions over several NACA 16-series airfoils with thicknesses of 4, 6, 9, and 12 percent of the chord and design lift coefficients of 0, 0.2, 1 and 0.5 has been conducted in the Langley airfoil test apparatus at transonic Mach numbers from 0.7 to 1.25. The tests ranged in Reynolds number from 2.4 x 10 (exp 6) to 2.8 x 10 (exp 6) and in angle of attack from -10 to 12 degrees. Chordwise pressure distributions and schlieren flow photographs are presented without analysis.
An experimental investigation has been made in the Langley highspeed hydrodynamics facility to determine the force and moment characteristics of two hydrofoils (one having an aspect ratio of 1 and the other having an aspect ratio of 3) designed to have improved lift-drag ratios when operating under either supercavitating or ventilated conditions. Measurements were made of lift, drag, and pitching moment over a range of angles of attack from 40 to 200 for depths of submersion varying from 0 to approximately 1 chord. The range of speed for the investigation was from 110 to 200 feet per second. When the upper surface of the hydrofoils was completely unwetted, the experimental values of lift and drag forces were in good agreement with the theoretical values obtained from the zero-cavitation-number theory. The theoretical values for minimum angle of attack for operation with the upper surface of the hydrofoil unwetted define the lower limits of angle of attack for which the experimental values of lift coefficient are either in agreement with or slightly greater than those predicted by theory.