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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 1,657 records · Page 92
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

An empirical method for computing leeside centerline heating on the Space Shuttle Orbiter

An empirical method is presented for computing top centerline heating on the Space Shuttle Orbiter at simulated reentry conditions. It is shown that the Shuttle's top centerline can be thought of as being under the influence of a swept cylinder flow field. The effective geometry of the flow field, as well as top centerline heating, are directly related to oil-flow patterns on the upper surface of the fuselage. An empirical turbulent swept cylinder heating method was developed based on these considerations. The method takes into account the effects of the vortex-dominated leeside flow field without actually having to compute the detailed properties of such a complex flow. The heating method closely predicts experimental heat-transfer values on the top centerline of a Shuttle model at Mach numbers of 6 and 10 over a wide range in Reynolds number and angle of attack.

Helms, V. T., III↗
NASA NTRS1983

Flap effectiveness on subsonic longitudinal aerodynamic characteristics of a modified arrow wing

An investigation of the subsonic longitudinal aerodynamic characteristics of a modified arrow-wing model was conducted in the Langley 4- by 7-Meter Tunnel. This investigation addressed the effectiveness of the leading and trailing edge flap deflections of this model. The arrow wing was tested at a Mach number of 0.02 and at an angle-of-attack range from -4 deg to 24 deg. The results of the investigation showed that deflecting the leading edge and trailing edge in combination could promote an attached-flow condition at the wing leading edge. Also, the leading edge suction could be maximized over the complete lift-coefficient range by scheduling a combination of leading and trailing edge flap deflections.

Quinto, P. F.↗
NASA NTRS1983

Heating measurements on Space Shuttle Orbiter models with differentially deflected elevons

The phase change paint technique was used to make heat transfer measurements on the windward wing/elevons area of Space Shuttle Orbiter models with differentially deflected elevons. Outboard elevons were deflected windward at angles from 0 to 20 deg. The inboard elevons were deflected leeward. The models were tested in air at Mach 6 and 10 with two flow conditions and three angles of attack for both Mach numbers. Heat transfer coefficient contour maps of the wing/elevon area were obtained for all test conditions. The maximum heat transfer coefficient on the outboard elevon and on certain areas of the wing that were obtained when the elevons were deflected are compared to values obtained when the elevons were undeflected.

Wells, W. L.↗
NASA NTRS1984

Fuselage and nozzle pressure distributions of a 1/12-scale F-15 propulsion model at transonic speeds. Effect of fuselage modifications and nozzle variables

Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.

Pendergraft, O. C., Jr.↗
NASA NTRS1984

Prediction of vortex lift on interacting delta wings in incompressible flow

Vortex flow modeling is used to calculate the steady inviscid incompressible flow past one, two and three delta wing configurations. The wings are modeled with vortex lattices and the leading and trailing-edge sheets are modeled by segmented straight vortex filaments. Aerodynamic characteristics are obtained for a range of geometry and angle of attack.

Dodbele, S. S.↗
NASA NTRS1982

Effects of surface chemistry on hot corrosion life: Overview

This program concentrates on analyzing a limited number of hot corroded components from the field and the carrying out of a series of controlled laboratory experiments to establish the effects of oxide scale and coating chemistry on hot corrosion life. This is to be determined principally from the length of the incubation period, the investigation of the mechanisms of hot corrosion attack, and the fitting of the data generated from the test exposure experiments to an empirical life prediction model.

Merutka, J.↗
NASA NTRS1984

Experimental and predicted heating distributions for biconics at incidence in air at Mach 10

Heating distributions were measured on a 1.9-percent-scale model of a generic aeroassisted vehicle proposed for missions to a number of planets and for use as a moderate lift-drag ratio Earth orbital transfer vehicle. This vehicle is spherically blunted, 12.84 deg/7 deg biconic with the fore-cone bent upward 7 deg to provide self-trim capability. A straight biconic with the same nose radius and the same half-angles was also tested. The free-stream Reynolds numbers based on model length were equal to about 2 x 10(5) or 9 x 10 (5). The angle of attack, referenced to the aft-cone, was varied from 0 deg to 20 deg. Heating distributions predicted with a parabolized Navier-Stokes (PNS) code are compared with the measurements for the present Reynolds numbers and range of angles of attack. Leeward heating was greatly affected by Reynolds number, with the heating increasing with decreasing Reynolds number for attached flow (low incidence). The opposite was true for separated flow, which occurred when the fore-cone angle of attack exceeded 0.8 times the fore-cone half-angle. Windward heating distributions were predicted to within 10 percent with the PNS code. Leeward heating distributions were predicted qualitatively for both Reynolds numbers, but quantitative agreement was poorer than on the windward side.

Miller, C. G., III↗
NASA NTRS1985

A sensitivity analysis of the Shuttle Orbiter heating

A procedure for estimating relative sensitivities in Space Shuttle Orbiter entry heating due to variations from nominal flight parameters is presented. Ballpark values are provided for assessing heating changes caused by off-nominal variations in anglle of attack and Reynolds number for laminar flow conditions. Also, a procedure for scaling heating data obtained on subscale models tested at low-enthalpy wind-tunnel test conditions to high-enthalpy flight conditions using simplified flow analysis techniques is presented and verified with flight and wind-tunnel data. In addition, data are presented illustrating that a significant increase in heating was experienced on the fifth Orbiter flight. The increase was possibly caused by changes in the catalytic properties of the Orbiter thermal protection system surface.

Williams, S. D.↗
NASA NTRS1987

Effect of a trade between boattail angle and wedge size on the performance of a nonaxisymmetric wedge nozzle

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effect of a boattail angle and wedge-size trade on the performance of nonaxisymmetric wedge nozzles installed on a generic twin-engine fighter aircraft model. Test data were obtained at static conditions and at Mach numbers from 0.60 to 1.25. Angle of attack was held constant at 0 deg. High-pressure air was used to simulate jet exhaust, and the nozzle pressure ratio was varied from 1.0 (jet off) to slightly over 15.0. For the configurations studied, the results indicate that wedge size can be reduced without affecting aeropropulsive performance.

Carson, George T., Jr.↗
NASA NTRS1990

Unsteady potential flow past a propeller blade section

An analytical study was conducted to predict the effect of an oscillating stream on the time dependent sectional pressure and lift coefficients of a model propeller blade. The assumption is that as the blade sections encounter a wake, the actual angles of attack vary in a sinusoidal manner through the wake, thus each blade is exposed to an unsteady stream oscillating about a mean value at a certain reduced frequency. On the other hand, an isolated propeller at some angle of attack can experience periodic changes in the value of the flow angle causing unsteady loads on the blades. Such a flow condition requires the inclusion of new expressions in the formulation of the unsteady potential flow around the blade sections. These expressions account for time variation of angle of attack and total shed vortices in the wake of each airfoil section. It was found that the final expressions for the unsteady pressure distribution on each blade section are periodic and that the unsteady circulation and lift coefficients exhibit a hysteresis loop.

Takallu, M. A.↗
NASA NTRS1991

Evaluation of panel code predictions with experimental results of inlet performance for a 17-inch ducted prop/fab simulator operating at Mach 0.2

An axisymmetric panel code was used to evaluate a series of ducted propeller inlets. The inlets were tested in the Lewis 9 by 15 Foot Low Speed Wind Tunnel. Three basic inlets having ratios of shroud length to propeller diameter of 0.2, 0.4, and 0.5 were tested with the Pratt and Whitney ducted prop/fan simulator. A fourth hybrid inlet consisting of the shroud from the shortest basic inlet coupled with the spinner from the largest basic inlet was also tested. This later configuration represented the shortest overall inlet. The simulator duct diameter at the propeller face was 17.25 inches. The short and long spinners provided hub-to-tip ratios of 0.44 at the propeller face. The four inlets were tested at a nominal free stream Mach number of 0.2 and at angles of attack from 0 degrees to 35 degrees. The panel code method incorporated a simple two-part separation model which yielded conservative estimates of inlet separation.

Boldman, D. R.↗
NASA NTRS1991

Wind tunnel blockage study of a generic three-dimensional sidewall compression scramjet inlet at Mach 10

A large scale model of a generic three-dimensional sidewall compression scramjet inlet has been designed based on the results of a computational parametric study for testing in the 31-inch Mach 10 Hypersonic Wind Tunnel at the NASA Langley Research Center. In order to increase the instrumentation density in interaction regions for a highly instrumented model, it is desirable to make the model as large as possible. When the cross-sectional area of a model becomes large relative to the inviscid core size of the tunnel, the effects of blockage must be considered. In order to assess these effects, a blockage model (an inexpensive, much less densely instrumented version of the configuration) was fabricated for preliminary testing. Since it was desired to determine both the effect of the model on the performance of the wind tunnel and also to determine if the inlet would start, the model possessed a total of 32 static pressure orifices distributed on the forebody plane and sidewalls; seventeen static pressure orifices on the tunnel wall and 3 pitot probes on the model monitored the tunnel performance. This paper presents the design considerations in the development of the wind tunnel model and the blockage aspects of the effects of contraction ratio, cowl location, Reynolds number, and angle of attack.

Holland, Scott D.↗
NASA NTRS1991

Incipient torsional stall flutter aerodynamic experiments on a swept three-dimensional wing

The aerodynamics of small amplitude pitching motions near stall have been studied experimentally in order to improve understanding of the torsional stall flutter problem for propeller blades. A model wing was oscillated in pitch at several small amplitudes over a wide and representative range of conditions. Unsteady surface pressures were measured and integrated to determine the aerodynamic damping at five spanwise stations. Strong negative damping was found for motions centered near static stall for all studied reduced frequencies, Mach numbers, and sweep angles. The 30-deg sweptback configuration was found to become negatively damped over the entire span nearly simultaneously, while the unswept model exhibited local regions of negative damping that moved toward the wing tip as the mean angle of attack was increased.

Lorber, Peter F.↗
NASA NTRS1991

Evaluation of panel code predictions with experimental results of inlet performance for a 17-inch ducted prop/fan simulator operating at Mach 0.2

An axisymmetric panel code was used to evaluate a series of ducted propeller inlets. The inlets were tested in the Lewis 9 by 15 Foot Low Speed Wind Tunnel. Three basic inlets having ratios of shroud length to propeller diameter of 0.2, 0.4, and 0.5 were tested with the Pratt and Whitney ducted prop/fan simulator. A fourth hybrid inlet consisting of the shroud from the shortest basic inlet coupled with the spinner from the largest basic inlet was also tested. This later configuration represented the shortest overall inlet. The simulator duct diameter at the propeller face was 17.25 inches. The short and long spinners provided hub-to-tip retios of 0.44 at the propeller face. The four inlets were tested at a nominal free stream Mach number of 0.2 and at angles of attack from 0 degrees to 35 degrees. The panel code method incorporated a simple two-part separation model which yielded conservative estimates of inlet separation.

Boldman, D. R.↗
NASA NTRS1992

A high-fidelity batch simulation environment for integrated batch and piloted air combat simulation analysis

A batch air combat simulation environment known as the Tactical Maneuvering Simulator (TMS) is presented. The TMS serves as a tool for developing and evaluating tactical maneuvering logics and to evaluate the tactical implications of perturbations to aircraft performance or supporting systems. The TMS is capable of simulating air combat between any number of engagement participants, with practical limits imposed by computer memory and processing power. Aircraft are modeled using equations of motion, control laws, aerodynamics and propulsive characteristics, and databases representative of a modern high-performance aircraft with and without thrust-vectoring capability are included. A Tactical Autopilot is implemented in the aircraft simulation model to convert guidance commands issued by computerized maneuvering logics in the form of desired angle-of-attack and wind axis-bank angle into inputs to the inner-loop control augmentation system of the aircraft.

Goodrich, Kenneth H.↗
NASA NTRS1993

Ice Accretion Prediction for a Typical Commercial Transport Aircraft

Ice accretion calculations were made for a modern commercial transport using the NASA Lewis LEWICE3D ice accretion code. The ice accretion calculations were made for the wing and horizontal tail using both isolated flow models and flow models incorporating the entire airplane. The isolated flow model calculations were made to assess the validity of using these simplified models in lieu of the entire model in the ice accretion analysis of full aircraft. Ice shapes typifying a rime and a mixed ice shape were generated for a 30 minute hold condition. In general, the calculated ice shapes looked reasonable and appeared representative of a rime and a mixed ice conditions. The isolated flow model simplification was good for the main wing except at the root where it overpredicted the amount of accreted ice relative to the full aircraft flow model. For the horizontal tail the size and amount of predicted ice compared well for the two flow models, but the position of the accretions were more towards the upper surface for the aircraft flow model relative to the isolated flow model. This was attributed to downwash from the main wing which resulted in a lower effective angle of attack for the aircraft tail.

Bidwell, C. S.↗
NASA NTRS1993

Ice accretion prediction for a typical commercial transport aircraft

Ice accretion calculations were made for a modern commercial transport using the NASA Lewis LEWICE3D ice accretion code. The ice accretion calculations were made for the wing and horizonal tail using both isolated flow models and flow models incorporating the entire airplane. The isolated flow model calculations were made to assess the validity of using these simplified models in lieu of the entire model in the ice accretion analysis for full aircraft. Ice shapes typifying a rime and a mixed ice shape were generated for a 30 minute hold condition. In general, the calculated ice shapes looked reasonable and appeared representative of a rime and a mixed ice conditions. The isolated flow model simplification was good for the main wing except at the root where it overpredicted the amount of accreted ice relative to the full aircraft flow model. For the horizontal tail the size and amount of predicted ice compared well for the two flow models, but the position of the accretions were more towards the upper surface for the aircraft flow model relative to the isolated flow model. This was attributed to downwash from the main wing which resulted in a lower effective angle of attack for the aircraft tail.

Bidwell, C. S.↗
NASA NTRS1993

Base drag prediction on missile configurations

New wind tunnel data have been taken, and a new empirical model has been developed for predicting base drag on missile configurations. The new wind tunnel data were taken at NASA-Langley in the Unitary Wind Tunnel at Mach numbers from 2.0 to 4.5, angles of attack to 16 deg, fin control deflections up to 20 deg, fin thickness/chord of 0.05 to 0.15, and fin locations from 'flush with the base' to two chord-lengths upstream of the base. The empirical model uses these data along with previous wind tunnel data, estimating base drag as a function of all these variables as well as boat-tail and power-on/power-off effects. The new model yields improved accuracy, compared to wind tunnel data. The new model also is more robust due to inclusion of additional variables. On the other hand, additional wind tunnel data are needed to validate or modify the current empirical model in areas where data are not available.

Moore, F. G.↗