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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 181 records · Page 10
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NASA NTRS2016

Experimental Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Take-off and Landing Conditions

As part of the NASA Environmentally Responsible Aircraft project, an ultra high bypass ratio engine integration on a hybrid wing body demonstration was planned. The goal was to include engine and airframe integration concepts that reduced fuel consumption by at least 50% while still reducing noise 42 db cumulative on the ground. Since the engines would be mounted on the upper surface of the aft body of the aircraft, the inlets may be susceptible to vortex ingestion from the wing leading edge at high angles of attack and sideslip, and separated wing/body flow. Consequently, experimental and computational studies were conducted to collect flow surveys useful for characterizing engine operability. The wind tunnel tests were conducted at two NASA facilities, the 14- by 22-foot at NASA Langley and the 40- by 80-foot at NASA Ames Research Center. The test results included in this paper show that the distortion and pressure recovery levels were acceptable for engine operability. The CFD studies conducted to compare to experimental data showed excellent agreement for the angle of attacks examined, although failed to match the low speed experimental data at high sideslip angles.

Carter, Melissa B.↗
NASA NTRS1976

Preliminary Sizing and Performance Evaluation of Supersonic Cruise Aircraft

The basic processes of a method that performs sizing operations on a baseline aircraft and determines their subsequent effects on aerodynamics, propulsion, weights, and mission performance are described. The input requirements of the associated computer program are defined and its output listings explained. Results obtained by applying the method to an advanced supersonic technology concept are discussed. These results include the effects of wing loading, thrust-to-weight ratio, and technology improvements on range performance, and possible gains in both range and payload capability that become available through growth versions of the baseline aircraft. The results from an in depth contractual study that confirm the range gain predicted for a particular wing loading, thrust-to-weight ratio combination are also included.

Fetterman, D. E., Jr.↗
NASA NTRS1973

Analytical study of takeoff and landing performance for a jet STOL transport configuration with full-span, externally blown, triple-slotted flaps

Takeoff and landing performance characteristics and field length requirements were determined analytically for a jet STOL transport configuration with full-span, externally blown, tripleslotted flaps. The configuration had a high wing, high T-tail, and four pod-mounted high-bypass-ratio turbofan engines located under and forward of the wing. One takeoff and three approach and landing flap settings were evaluated. The effects of wing loading, thrust-to-weight ratio, weight, ambient temperature, altitude on takeoff and landing field length requirements are discussed.

Washington, H. P.↗
NASA NTRS1972

Prospects for low wing-loading STOL transports with ride smoothing.

Airplanes with low wing-loadings provide STOL capability without reliance on auxiliary propulsion or augmented lift, but require a ride smoothing control system to provide acceptable passenger comfort. A parametric study produced a configuration having a .35 thrust-to-weight ratio and a 50 psf wing loading, and which satisfied specified mission requirements and airworthiness standards. A ride-smoothing control system (RCS) synthesis was then performed which consisted of ride quality criteria definition, RCS concept trades, and analysis of RCS performance benefits at significant flight conditions. Within the limitations of the study it is concluded that this is a viable approach to STOL airplane design.

Holloway, R. B.↗
NASA NTRS1975

Tabulated pressure measurements on an executive-type jet transport model with a supercritical wing

A 1/9 scale model of an existing executive type jet transport refitted with a supercritical wing was tested on in the 8 foot transonic pressure tunnel. The supercritical wing had the same sweep as the original airplane wing but had maximum thickness chord ratios 33 percent larger at the mean geometric chord and almost 50 percent larger at the wing-fuselage juncture. Wing pressure distributions and fuselage pressure distributions in the vicinity of the left nacelle were measured at Mach numbers from 0.25 to 0.90 at angles of attack that generally varied from -2 deg to 10 deg. Results are presented in tabular form without analysis.

Bartlett, D. W.↗
NASA NTRS1955

Aerodynamic Characteristics of a 0.04956-Scale Model of the Convair F-102A Airplane at Transonic Speeds

Tests have been conducted in the Langley 8-foot transonic tunnel on a 0.04956-scale model of the Convair F-102A airplane which employed an indented and extended fuselage, cambered wing leading edges, and deflected wing tips. Force and moment characteristics were obtained for Mach numbers from 0.60 to 1.135 at angles of attack up to 20 . In addition, tests were made over a limited angle-of-attack range to determine the effects of the cambered leading edges, deflected tips, and a nose section with a smooth area distribution. Fuselage modifications employed on the F-102A were responsible for a 25.percent reduction in the minimum drag-coefficient rise between the Mach numbers of 0.85 and 1.075 when compared with that for the earlier versions of the F-102. Although the wing modifications increased the F-102A subsonic minimum drag-coefficient level approximately 0.0020, they produced large decreases in drag at lifting conditions over that for the original (plane-wing) F-102. The F-102A had 15 to 25 percent higher maximum lift-drag ratios than did the original F-102. The F-102A had about 15 percent lower maximum lift-drag ratios at Mach numbers below 0.95 and slightly higher maximum lift-drag ratios at supersonic speeds when compared with those ratios for sn earlier modified-wing version of the F-102. Chordwise wing fences which provided suitable longitudinal stability for the original F-102 were not adequate for the cambered-wing F-102A The pitching-moment curves indicated a region of near neutral stability with possible pitch-up tendencies for the F-102A at high subsonic Mach numbers for lift coefficients between about 0.4 and 0.5.

Tempelmeyer, Kenneth E.↗
NASA NTRS1948

Transonic Drag Characteristics of a Wing-body Combination Showing the Effect of a Large Wing Fillet

Results of an investigation by the free-fall method are presented herein for a configuration having a body of revolution of fineness ratio 12 and 45 degrees sweptback wing mounted aft of the maximum diameter of the body. The fillets were designed to provide large increases in the sweep oof the leading edge and the line of maximum thickness as the wing root was approached. Comparison of these results with those for the same configuration without fillets shows that the addition of wing fillets increased the total drag of the configuration by about 35 percent at Mach numbers near 1.0 and about 15 percent at Mach numbers near 1.2. Results indicate that the fillets produced no appreciable change in the wing and tail drags but produced a large increase in body drag due to interference.

Cheatham, Donald C↗
NASA NTRS1951

Study of Effects of Sweep on the Flutter of Cantilever Wings

An experimental and analytical investigation of the flutter of sweptback cantilever wings is reported. The experiments employed groups of wings swept back by rotating and by shearing. The angle of sweep range from 0 degree to 60 degrees and Mach numbers extended to approximately 0.85. A theoretical analysis of the air forces on an oscillating swept wing of high length-chord ratio is developed, and the approximations inherent in the assumptions are discussed. Comparison with experiment indicates that the analysis developed in the present report is satisfactory for giving the main effects of sweep, at least for nearly uniform cantilever wings of high and moderate length-chord ratios.

Barmby, J G↗
NASA NTRS1977

The oblique wing - Aircraft design for transonic and low supersonic speeds

In tests at transonic speeds a conventional 'subsonic' wing turned at an oblique angle to the flow has demonstrated lift-drag ratios higher than conventional swept-back or delta wings. When returned to its straight position, perpendicular to the flow, such a wing would minimize the display of energy and noise in the airport environment. Following a brief review of theoretical and experimental results, the paper discusses various characteristics of the oblique wing as they relate to aircraft design. Topics covered are lift-drag ratio, flight control and trim and aeroelastic stability. The possibility of an all-wing aircraft for flight at Mach 2 is suggested.

Jones, R. T.↗
NASA NTRS1962

Pressure Distribution Induced on a Flat Plate at a Free-Stream Mach Number of 1.39 by Rockets Exhausting Upstream and Downstream

An experimental investigation was made of the pressures induced on a flat plate at a free-stream Mach number of 1.39 by a supersonic rocket jet exhausting upstream and downstream. Measurements of the pressure distribution on a flat plate were made at zero angle of attack for 11 different locations of the jet exhaust nozzle beneath the wing. Measurements were made at ratios of rocket-exit total pressure to freestream static pressure from 6 to 60 and at a Reynolds number per foot of approximately 10 x 10 6 . The rocket when exhausted upstream produced a strong shock that moved further upstream with increasing rocket-exit total-pressure ratio. Positive incremental normal-force coefficients were obtained at all test positions. Data at 11 test positions are tabulated for rocket-on and rocket-off pressure coefficients as well as for incremental pressure coefficients for the 48 orifices of the flat plate for the range of ratio of rocket-exit total pressure to free-stream static pressure of the investigation. Changing the location of the model with respect to the plate had a negligible effect when the rocket was varied in the chordwise direction, but the pressure coefficients were reduced as the rocket was lowered away from the flat-plate wing.

Supersonic jet↗
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.↗
NASA NTRS1956

Theoretical and Experimental Investigation of the Effect of Tunnel Walls on the Forces on an Oscillating Airfoil in Two-Dimensional Subsonic Compressible Flow

This report presents a theoretical and experimental investigation of the effect of wind-tunnel walls on the air forces on an oscillating wing in two-dimensional subsonic compressible flow. A method of solving an integral equation which relates the downwash on a wing to the unknown loading is given, and some comparisons are made between the theoretical results and the experimental results. A resonance condition, which was predicted by theory in a previous report (NACA report 1150), is shown experimentally to exist. In addition, application of the analysis is made to a number of examples in order to illustrate the influence of walls due to variations in frequency of oscillation, Mach number , and ratio of tunnel height to wing semichord.

Runyan, Harry L↗
NASA NTRS1946

High-Speed Longitudinal-Stability and Control Characteristics of the Consolidated Vultee Lark Missile as Predicted from Wind-Tunnel Tests (TED No. NACA 2391)

A high-speed wind-tunnel investigation of the aerodynamic characteristics of a full-scale model of the Consolidated Vultee Lark indicates that the missile possesses satisfactory longitudinal-stability and-control characteristics throughout the Mach number range from 0.2 to 0.85, but that the maximum lift coefficients developed are not high enough to insure interception of the target at high altitudes. A reduction in wing loading appears advisable. Although the static longitudinal stability at zero angle of attack changes with Mach number and with lift coefficient, satisfactory control should be possible at all times as the tails retain their relatively large effectiveness throughout the range of Mach numbers and lift coefficients tested. Minimum stability and maximum maneuverability occur around 0.80 Mach number and 0.2 lift coefficient, which corresponds to level flight conditions of the missile. The optimum ratio of tail-to-wing deflection is 0.4.

Axelson, John A.↗
NASA NTRS2014

Core and Wing Densities of Asymmetric Coronal Spectral Profiles: Implications for the Mass Supply of the Solar Corona

Recent solar spectroscopic observations have shown that coronal spectral lines can exhibit asymmetric profiles, with enhanced emissions at their blue wings. These asymmetries correspond to rapidly upflowing plasmas at speeds exceeding approximately equal to 50 km per sec. Here, we perform a study of the density of the rapidly upflowing material and compare it with that of the line core that corresponds to the bulk of the plasma. For this task, we use spectroscopic observations of several active regions taken by the Extreme Ultraviolet Imaging Spectrometer of the Hinode mission. The density sensitive ratio of the Fe(sub XIV) lines at 264.78 and 274.20 Angstroms is used to determine wing and core densities.We compute the ratio of the blue wing density to the core density and find that most values are of order unity. This is consistent with the predictions for coronal nanoflares if most of the observed coronal mass is supplied by chromospheric evaporation driven by the nanoflares. However, much larger blue wing-to-core density ratios are predicted if most of the coronal mass is supplied by heated material ejected with type II spicules. Our measurements do not rule out a spicule origin for the blue wing emission, but they argue against spicules being a primary source of the hot plasma in the corona. We note that only about 40% of the pixels where line blends could be safely ignored have blue wing asymmetries in both Fe(sub XIV) lines. Anticipated sub-arcsecond spatial resolution spectroscopic observations in future missions could shed more light on the origin of blue, red, and mixed asymmetries.

coronal spectral profiles↗
NASA NTRS1948

Damping in Pitch and Roll of Triangular Wings at Supersonic Speeds

A method is derived for calculating the damping coefficients in pitch and roll for a series of triangular wings and a restricted series of sweptback wings at supersonic speeds. The elementary "supersonic source" solution of the linearized equation of motion is used to find the potential function of a line of doublets, and the flows are obtained by surface distributions of these doublet lines. The damping derivatives for triangular wings are found to be a function of the ratio of the tangent of the apex angle to the tangent of the Mach angle. As this ratio becomes equal to and greater than 1.0 for triangular wings, the damping derivatives, in pitch and in roll, become constant. The damping derivative in roll becomes equal to one-half the value calculated for an infinite rectangular wing, and the damping derivative in pitch for pitching about the apex becomes equal to 3.375 times that of an infinite rectangular wing.

Brown, Clinton E↗