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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.

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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

Effects of Forebody Deflection on the Stability and Control Characteristics of a Canard Airplane Configuration with a High Trapezoidal Wing at a Mach Number of 2.01

An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.01 to determine the effects of forebody deflection on the stability and control characteristics of a canard airplane configuration. The configuration had a high trapezoidal aspect-ratio-3 wing, a trapezoidal canard surface, and a single swept vertical tail. Forebody deflection angles of 0 deg, 2 deg and deg were investigated. The results indicated that nose-up deflections of the forebody provided positive increments of pitching moment with little increase in drag and hence would be useful in reducing the pitch-control requirements and the attendant losses in lift-drag ratio due to trimming. Deflection of the forebody, however, aggravated the decrease in directional stability with increasing angle of attack by causing a loss in tail contribution and by increasing the instability of the wing-body combination.

Spearman, M. Leroy↗
NASA NTRS1974

Effect of ground proximity on the longitudinal aerodynamic characteristics of an aspect-ratio-1 wing with and without wing-tip blowing

An investigation has been conducted to determine the effect of ground proximity on the aerodynamic characteristics of an aspect-ratio-1 wing with and without wing-tip blowing. This investigation was conducted in the Langley towing tank no. 1 with the model towed over the water to eliminate the effects of walls and of wind-tunnel ground-board boundary layers. The results indicate that the model is stable with height and pitch at positive angles of attack and that the lift-drag ratio is improved at small ground heights with some blowing at the wing tips.

Mineck, R. E.↗
NASA NTRS1977

Effect of a simulated engine jet blowing above an arrow wing at Mach 2.0

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.

Shrout, B. L.↗
NASA NTRS1988

Configuration trade and code validation study on a conical hypersonic vehicle

A study has been conducted on a generic wing-cone transatmospheric vehicle at Mach numbers form 2.5 to 4.5. The objectives of the study were to experimentally define the aerodynamic characteristics of the vehicle and evaluate several computational aerodynamic prediction methods through comparison with the experimental results. The baseline wing-cone configuration fuselage consisted of a 5 deg half-angle cone forebody, cylindrical midbody, and 9 deg truncated cone afterbody. The 4-percent-thick diamond airfoil wing had an aspect ratio of 1. Several configuration variables were investigated to provide trade information on canard, wing-position and incidence, vertical tail, and nose bluntness effects. Results of the study show that wing-position and wing-incidence effects on the longitudinal aerodynamic characteristics can be significantly influenced by wing-body interference. The use of positive wing incidence to provide favorable forebody orientation for possible inlet performance improvement is accompanied by trim drag and lift-drag ratio penalties. The lateral-directional stability characteristics were strongly influenced by the location of the vertical tails. The higher-order full-potential method provided better estimates of the aerodynamic characteristics than either the linearized supersonic potential method or the tangent-cone/tangent-wedge/shock-expansion on method.

Covell, Peter F.↗
NASA NTRS1989

Transonic aerodynamic characteristics of a proposed Assured Crew Return Capability (ACRC) lifting-body configuration

The investigation was conducted in the Calspan 8-Ft Transonic wind tunnel at Mach numbers from 0.6 to 1.2. The 0.07-scale model had a low aspect ratio body with a flat undersurface. A center fin and two tip fins were mounted on the aft upper body. The tip fins were rolled outboard 40 deg from the vertical. Elevon surfaces made up the trailing edges of the outboard fins and body flaps were located on the upper and lower aft fuselage. Results of the investigation indicated that the model was longitudinally and laterally stable about a center-of-gravity position of 0.54 body length. The maximum trimmed lift-drag ratio was about 3.1 at M = 0.6. The small center fin contributed only a small positive increment to lateral stability but was effective as a yaw control device. Protuberances on the forebody had little effect on the aerodynamic characteristics of the configuration. The model with pitch controls undeflected had desirable longitudinal trim characteristics.

Ware, George M.↗
NASA NTRS1958

Approximate Solutions for the Flow About Flat-Top Wing-Body Configurations at High Supersonic Airspeeds

The flow about slender flat-top wing-body configurations traveling at high supersonic speeds and small angles of attack is investigated analytically. In the case of conical configurations, approximate algebraic solutions to the flow field are obtained. In the case of configurations which are conical at the vertex but curved in the stream direction, these solutions are combined with a slender-body approximation to the generalized shock-expansion method to obtain the flow downstream of the vertex. Surface pressures were obtained experimentally at Mach numbers from 3.0 to 6.0 and angles of attack up to 6 deg for several flat-top wing-body configurations. These configurations consisted of half-bodies of revolution mounted beneath thin highly swept wings. Three different bodies were employed. The two conical bodies consisted of one-half of a fineness-ratio-5 cone and one-half of a fineness-ratio-2-1/2 cone. The body of the third configuration consisted of one-half of a fineness-ratio-5 ogive. For the ogive configuration, the leading edges of the wing were curved and designed to just maintain the theoretically determined bow shock along the leading edge at a Mach number of 5.0 and an angle of attack of 3 deg. The predictions of the conical flow theory of this paper for the surface pressures are found to be in good agreement with experiment at Mach numbers of 5.0 and 6.0 up to angles of attack of approximately 3 deg. Estimated lift, drag, and pitching-moment coefficients, as well as maximum lift-drag ratio, are also in good agreement with existing experimental data at a Mach number of 5.0 for a conical configuration having an arrow plan-form wing. It is also found that the generalized shock-expansion method yields reasonable good agreement with experiment for the surface pressures on the half-ogive configuration at a Mach number of 5.0 and an angle of attack of 3 deg.

Savin, Raymond C.↗
NASA NTRS1960

Investigation at Mach Numbers of 0.20 to 3.50 of Blended Wing-Body Combinations of Sonic Design with Diamond, Delta, and Arrow Plan Forms

The models had aspect-ratio-2 diamond, delta, and arrow wings with the leading edges swept 45.00 deg, 59.04 deg, and 70.82 deg, respectively. The wing sections were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local chords in a streamwise direction. The models were tested with transition fixed at Reynolds numbers of about 4,000,000 to 9,000,0000, based on the mean aerodynamic chord of the wings. The effect of varying Reynolds number was checked at both subsonic and supersonic speeds. The diamond model was superior to the other plan forms at transonic speeds ((L/D)max = 11.00 to 9.52) because of its higher lift-curve slope and near optimum wave drag due to the blending process. For the wing thickness tested with the diamond model, the marked body and wing contouring required for transonic conditions resulted in a large wave-drag penalty at the higher supersonic Mach numbers where the leading and trailing edges of the wing were supersonic. Because of the low sweep of the trailing edge of the delta model, this configuration was less adaptable to the blending process. Removing a body bump prescribed by the Mach number 1.00 design resulted in a good supersonic design. This delta model with 10 percent less volume was superior to the other plan forms at Mach numbers of 1.55 to 2.35 ((L/D)max = 8.65 to 7.24), but it and the arrow model were equally good at Mach numbers of 2.50 to 3.50 ((L/D)max - 6.85 to O.39). At transonic speeds the arrow model was inferior because of the reduced lift-curve slope associated with its increased sweep and also because of the wing base drag. The wing base-drag coefficients of the arrow model based on the wing planform area decreased from a peak value of 0.0029 at Mach number 1.55 to 0.0003 at Mach number 3.50. Linear supersonic theory was satisfactory for predicting the aerodynamic trends at Mach numbers from 1.55 to 3.50 of lift-curve slope, wave drag, drag due to lift, aerodynamic-center location, and maximum lift-drag ratios for each of the models.

Holdaway, George H.↗
NASA NTRS1958

Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One with Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One with a Flared Afterbody and All-Movable Controls

An investigation has been made to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and with all-movable control surfaces. The flaps were tested on a configuration having low-aspect-ratio cruciform fins with an apex angle of 5 deg the all-movable controls were mounted at the 46.7-percent body station on a configuration having a 10 deg flared afterbody. The tests were made through an angle-of-attack range of -2 deg to 20 deg at zero sideslip in the Langley 11-inch hypersonic tunnel. The results indicated that the all-movable controls on the flared afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge -flap configuration. The flared -after body configuration had considerably higher drag than the cruciform-fin model but only slightly lower values of lift drag ratio.

Robinson, Ross B↗
NASA NTRS1958

Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One with Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One with a Flared Afterbody and All-Movable Controls

An investigation has been made to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and with all-movable control surfaces. The flaps were tested on a configuration having low-aspect-ratio cruciform fins with an apex angle of 5 degrees; the all-movable controls were mounted at the 46.7-percent body station on a configuration having a 10 degrees flared afterbody. The tests were made through an angle-of-attack range of -2 degrees to 20 degrees at zero sideslip in the Langley 11-inch hypersonic tunnel. The results indicated that the all-movable controls on the flared-afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge-flap configuration. The flared-afterbody configuration had considerably higher drag than the cruciform-fin model but only slightly lower values of lift-drag ratio.

Bernot, P. T.↗
NASA NTRS1983

A comparison of trim drag for conventional and supercritical wings

The magnitude of the trim drag incurred by advanced supercritical wings and conventional wide-body wings at cruise conditions are investigated. Experiments were performed in a transonic pressure tunnel at a Mach number of 0.82 with a conventional wide-body wing and supercritical wing with high aspect ratio (9.80) using one of three low-tail configurations or two T-tail configurations in order to determine the effects of horizontal tail size, location, camber and static margin. Drag measurements indicate trim drag for the supercritical wing not to be significantly higher than for the conventional wing, although the minimum drag values for the supercritical wing occurred at lower static margins than for conventional wings. Both wings exhibited a reduction in trim drag with reduced cambered tail size, and greater minimum drag increments for cambered tails than for symmetrical tails. Lower trim-drag increments were also observed for the T-tail configuration than the low tails. The increase in lift-drag ratio for the supercritical wing over the conventional wing amounted to 11% for the best tail configurations.

Jacobs, P. F.↗
DOE OSTI2025

Leveraging interpolation models and error bounds for verifiable scientific machine learning

Effective verification and validation techniques for modern scientific machine learning workflows are challenging to devise. Statistical methods are abundant and easily deployed, but often rely on speculative assumptions about the data and methods involved. Error bounds for classical interpolation techniques can provide mathematically rigorous estimates of accuracy, but often are difficult or impractical to determine computationally. Here, in this work, we present a best-of-both-worlds approach to verifiable scientific machine learning by demonstrating that (1) multiple standard interpolation techniques have informative error bounds that can be computed or estimated efficiently; (2) comparative performance among distinct interpolants can aid in validation goals; (3) deploying interpolation methods on latent spaces generated by deep learning techniques enables some interpretability for black-box models. We present a detailed case study of our approach for predicting lift-drag ratios from airfoil images. Code developed for this work is available in a public Github repository.

97 MATHEMATICS AND COMPUTING↗
NASA NTRS1962

PROGRESS REPORT ON THE DEVELOPMENT OF PROTECTED CONSTRUCTION FOR HYPERSONIC VEHICLES

The structural problems of re-entry occasioned by aerodynamic heating are now generally well known. Typical of these heating problems is the environment experienced by the manned lifting vehicle re-entering from a low altitude orbit. Figure 1 shows typical curves of lower surface temperature as a function of time for a wing loading of 25 Ibs/ft(exp 2), and a lift-drag ratio of 2.5. The two curves cover a practical range of re-entry angles, the lower curve representing an ideal re-entry path with zero dive angle and the upper curve assuming a 2° error in the re-entry angle. Maneuvers for course change or correction are also added to the upper curve. Significant factors from this curve are the relationship between the maximum temperatures and the capabilities of available metallic material, and also the long re-entry time and its effect on total heat load. If a very shallow re-entry is made, maximum lower surface temperatures reach 2000°F which is just within the range of conventional superalloys. To accommodate practical re-entry angles and maneuvers, however, the temperature capability must reach about 2500°F which requires refractory metals. The re-entry time may be as high as 100 minutes, which gives total heat loads of approximately 40,000 BTU/ft(exp 2). A heat load of this magnitude, with equilibrium surface temperatures of the values shown, suggests that a lighter airframe can be constructed by dissipating the heat by radiation from the surface, rather than by absorbing it with a heat sink, or surface cooling, or ablation. Air Force programs to provide airframes for this type of environment have involved the parallel development of a number of different structural concepts. The development to be discussed here was carried out by Bell Aerosystems Company for the Fabrication Branch, Manufacturing Technology Laboratory, Directorate of Materials and Processes, Aeronautical Systems' Division, Wright-Patterson Air Force Base, Ohio, under Contract AF33(600)-40100 (Double-Wall Cooled Structure).

Reentry vehicle↗
NASA NTRS1968

Aerodynamic Characteristics of Several Hypersonic Boost-Glide-Type Configurations at Mach Numbers from 2.30 to 4.63

An investigation has been conducted at Mach numbers from 2.30 to 4.63 to determine the static aerodynamic characteristics of several configurations designed for flight at hypersonic Mach numbers. Two all-wing and three wing-body configurations were tested through an angle-of-attack range from about -4 degrees to 33 degrees and an angle-of-sideslip range from about -4 degrees to 8 degrees at a Reynolds number of 3 times 10 (sup 6) per foot (9.84 times 10 (sup 6) per meter). The results of the investigation indicated that the wing-body configurations produced higher values of maximum lift-drag ratio than those produced by the all-wing models. The high wing-body configurations tend to have a self-trimming capability as opposed to that for the low wing-body configurations. Each of the configurations produced a positive dihedral effect that increased with increasing angle of attack and decreased with increasing Mach number. The high wing-body models produced decreasing values of directional stability with increase in angle of attack, whereas the low wing-body models provided increasing values of directional stability with increase in angle of attack.

Graves, Ernald B.↗