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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Experimental evaluation of blockage ratio and plenum evacuation system flow effects on pressure distribution for bodies of revolution in 0.1 scale model test section of NASA Lewis Research Center's proposed altitude wind tunnel

An experimental investigation was conducted in the slotted test section of the 0.1-scale model of the proposed Altitude Wind Tunnel to evaluate wall interference effects at tunnel Mach numbers from 0.70 to 0.95 on bodies of revolution with blockage rates of 0.43, 3, 6, and 12 percent. The amount of flow that had to be removed from the plenum chamber (which surrounded the slotted test section) by the plenum evacuation system (PES) to eliminate wall interference effects was determined. The effectiveness of tunnel reentry flaps in removing flow from the plenum chamber was examined. The 0.43-percent blockage model was the only one free of wall interference effects with no PES flow. Surface pressures on the forward part of the other models were greater than interference-free results and were not influenced by PES flow. Interference-free results were achieved on the aft part of the 3- and 6-percent blockage models with the proper amount of PES flow. The required PES flow was substantially reduced by opening the reentry flaps.

Burley, Richard R.↗
NASA NTRS1986

Communications platform payload definition study, executive summary

Large geostationary communications platforms have been investigated in a number of studies since 1974 as a possible means to more effectively utilize the geostationary orbital arc and electromagnetic spectrum and to reduce overall satellite communications system costs. This NASA Lewis sponsored study addresses the commercial feasibility of various communications platform payload concepts circa 1998. It defines promising payload concepts, estimates recurring costs and identifies critical technologies needed to permit eventual commercialization. Ten communications service aggregation scenarios describing potential groupings of services were developed for a range of conditions. Payload concepts were defined for four of these scenarios: (1) Land Mobile Satellite Service (LMSS), meet 100% of CONUS plus Canada demand with a single platform; (2) Fixed Satellite Service (FSS) (Trunking + Customer Premises Service (CPS), meet 20% of CONUS demands; (3) FSS (Trunking + video distribution), 10 to 13% of CONUS demand; and (4) FSS (20% of demand) + Inter Satellite Links (ISL) + TDRSS/TDAS Data Distribution.

Clopp, H. W.↗
NASA NTRS1984

Why credible propeller noise measurements are possible in the acoustically untreated NASA Lewis 8 ft by 6 ft wind tunnel

An explanation is presented for the lack of acoustic reflections in noise studies of propfan models in the NASA-Lewis 8 x 6 ft wind tunnel, where trials were run at Mach numbers 0.5-0.85. The highly directional propeller noise, i.e., mainly in the plane of rotation, experiences a convective effect due to the high subsonic axial Mach number. Reflected sounds are carried downstream, out of range of the acoustic sensors in the tunnel. Furthermore, reflected noise is less audible, and therefore does not affect measurements near peak values. It is suggested that some data contamination may occur below Mach 0.6, and that measurements be performed on higher harmonics generated by low level reflected noise.

Dittmar, J. H.↗
NASA NTRS1989

Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take/off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller mode to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

Hughes, Christopher E.↗
NASA NTRS1957

Total-Pressure Distortion and Recovery of Supersonic Nose Inlet with Conical Centerbody in Subsonic Icing Conditions

Ice was formed on a full-scale unheated supersonic nose inlet in the NACA Lewis icing tunnel to determine its effect on compressor-face total-pressure distortion and recovery.Inlet angle of attack was varied from 0degrees to 12 degrees, free-stream Mach number from 0.17 to 0.28, and compressor-face Mach number from 0.10 to 0.47. Icing-cloud liquid-water content was varied from 0.65 to 1.8 grams per cubic meter at free-stream static air temperatures of 15 degrees and 0 degrees F. The addition of ice to the inlet components increased total-pressure-distortion levels and decreased recovery values compared withclear0air results, the losses increasing with time in ice. The combination of glaze ice, high corrected weight flow, and high angle of attack yielded the highest levels of distortion and lowest values of recovery. The general character of compressor-face distortion with an iced inlet was the same as that for the clean inlet, the total-pressure gradients being predominantly radial, with circumferential gradients occurring at angle of attack. At zero angle of attack, free-stream Mach number of 0.27, and a constant corrected weight flow of 150 pounds per second (compressor-face Mach number of 0.43), compressor-face total-pressure-distortion level increased from about 6 percent in clear air to 12 percent after 21 minutes of heavy glaze icing; concurrently, total-pressure recovery decreased from about 0.98 to 0.945. For the same operating conditions but with the inlet at 12 deg angle of attack, a change in distortion level occurred from about 9 percent in clear air to 14 percent after 2-1/4 minutes of icing, with a decrease in recovery from about 0.97 to 0.94.

Gelder, Thomas F↗
NASA NTRS1988

Propulsion-airframe integration for commercial and military aircraft

A significant level of research is ongoing at NASA's Langley Research Center on integrating the propulsion system with the aircraft. This program has included nacelle/pylon/wing integration for turbofan transports, propeller/nacelle/wing integration for turboprop transports, and nozzle/afterbody/empennage integration for high performance aircraft. The studies included in this paper focus more specifically on pylon shaping and nacelle location studies for turbofan transports, nacelle and wing contouring and propeller location effects for turboprop transports, and nozzle shaping and empennage effects for high performance aircraft. The studies were primarily conducted in NASA Langley's 16-Foot Transonic Tunnel at Mach numbers up to 1.20. Some higher Mach number data obtained at NASA's Lewis Research Center is also included.

Henderson, William P.↗
NASA NTRS1991

A comparison of fiber effects on polymer matrix composite oxidation

A number of thermo-oxidative stability studies addressing the effects of fiber reinforcement on composite thermal stability and influence of geometry on the results of aging studies were performed at NASA-Lewis. The information presented herein, a compilation of some results from these studies, shows the influence of the reinforcement fibers on the oxidative degradation of various PMR-15 composites. Reinforcement of graphite and ceramics were studied and three composite oxidation mechanisms were observed. One was a dominant attack of the reinforcement fiber, the second was the aggressive oxidation of the matrix material, and the third was interfacial degradation.

Bowles, Kenneth J.↗
NASA NTRS1972

A finite difference method for predicting supersonic turbulent boundary layer flows with tangential slot injection

An implicit finite difference method has been applied to tangential slot injection into supersonic turbulent boundary layer flows. In addition, the effects induced by the interaction between the boundary layer displacement thickness and the external pressure field are considered. In the present method, three different eddy viscosity models have been used to specify the turbulent momentum exchange. One model depends on the species concentration profile and the species conservation equation has been included in the system of governing partial differential equations. Results are compared with experimental data at stream Mach numbers of 2.4 and 6.0 and with results of another finite difference method. Good agreement was generally obtained for the reduction of wall skin friction with slot injection and with experimental Mach number and pitot pressure profiles. Calculations with the effects of pressure interaction included showed these effects to be smaller than effects of changing eddy viscosity models.

Miner, E. W.↗
NASA NTRS1950

Effect of Compressor-outlet Bleedoff on Turbojet-engine Performance

An investigation was conducted in the NACA Lewis altitude wind tunnel to determine the effect of compressor-outlet bleedoff on the performance of an axial-flow turbojet engine equipped with a variable-area exhaust nozzle. Results presented indicate the effect of compressor-outlet bleedoff on performance at altitudes of 25,000 and 40,000 feet and a flight Mach number of 0.53. Variation of performance with bleedoff flow is indicated for operation with fixed- and variable-area exhaust nozzles. Temperature and pressure losses through the bleedoff ducting system are also discussed.

TURBINES, GAS - BLEED-OFF↗
NASA NTRS1994

Theoretical studies of the transport properties in compound semiconductors

This final report is an overview of the work done on Cooperative Agreement NCC 3-55 with the Solid State Technology Branch of the NASA-Lewis Research Center (LeRC). Over the period of time that the agreement was in effect, the principal investigator and, in the last three years, the co-principal investigator worked on a significant number of projects and interacted with members of the Solid State Technology (SST) branch in a number of different ways. For the purpose of this report, these efforts will be divided into five categories: 1) work directly with experimental electrical transport studies conducted by members of the SST branch; 2) theoretical work on electrical transport in compound semiconductors; 3) electronic structure calculations which are relevant to the electrical transport in polytypes of SiC and SiC-AlN alloys; 4) the electronic structure calculations of polar interfaces; and 5) consultative and supportive activities related to experiments and other studies carried out by SST branch members. Work in these categories is briefly discussed.

Segall, Benjamin↗
NASA NTRS1997

Modeling Cyclic Phase Change and Energy Storage in Solar Heat Receivers

Numerical results pertaining to cyclic melting and freezing of an encapsulated phase change material (PCM), integrated into a solar heat receiver, have been reported. The cyclic nature of the present melt/freeze problem is relevant to latent heat thermal energy storage (LHTES) systems used to power solar Brayton engines in microgravity environments. Specifically, a physical and numerical model of the solar heat receiver component of NASA Lewis Research Center's Ground Test Demonstration (GTD) project was developed and results compared with available experimental data. Multi-conjugate effects such as the convective fluid flow of a low-Prandtl-number fluid, coupled with thermal conduction in the phase change material, containment tube and working fluid conduit were accounted for in the model. A single-band thermal radiation model was also included to quantify reradiative energy exchange inside the receiver and losses through the aperture. The eutectic LiF-CaF2 was used as the phase change material (PCM) and a mixture of He/Xe was used as the working fluid coolant. A modified version of the computer code HOTTube was used to generate results for comparisons with GTD data for both the subcooled and two-phase regimes. While qualitative trends were in close agreement for the balanced orbit modes, excellent quantitative agreement was observed for steady-state modes.

Hall, Carsie A., III↗
NASA NTRS1995

Ice Accretion with Varying Surface Tension

During an icing encounter of an aircraft in flight, super-cooled water droplets impinging on an airfoil may splash before freezing. This paper reports tests performed to determine if this effect is significant and uses the results to develop an improved scaling method for use in icing test facilities. Simple laboratory tests showed that drops splash on impact at the Reynolds and Weber numbers typical of icing encounters. Further confirmation of droplet splash came from icing tests performed in the NaSA Lewis Icing Research Tunnel (IRT) with a surfactant added to the spray water to reduce the surface tension. The resulting ice shapes were significantly different from those formed when no surfactant was added to the water. These results suggested that the droplet Weber number must be kept constant to properly scale icing test conditions. Finally, the paper presents a Weber-number-based scaling method and reports results from scaling tests in the IRT in which model size was reduced up to a factor of 3. Scale and reference ice shapes are shown which confirm the effectiveness of this new scaling method.

Bilanin, Alan J.↗
NASA NTRS1972

Hall current effects in the Lewis magnetohydrodynamic generator

Data obtained in a magnetohydrodynamic generator are compared with theoretical values calculated by using the Dzung theory. The generator was operated with cesium-seeded argon as the working fluid. The gas temperature varied from 1800 to 2100 K, the gas pressure from 19 to 22 N/sq cm, the Mach number from 0.3 to 0.5, and the magnetic field strength from 0.2 to 1.6 T. The analysis indicates that there is incomplete seed vaporization and that Hall current shorting paths (through the working fluid to ground at both the entrance and exit of the channel) limit generator performance.

Nichols, L. D.↗
NASA NTRS1974

Gust Generator for a Supersonic Wind Tunnel

The effectiveness was investigated of a flat plate gust generator that was located in the nozzle throat of the Lewis 10- by 10-foot supersonic wind tunnel. Gust plates were tested at nozzle wall Mach numbers of 3.1, 2.4, and 2.0. Test results show that the flat plate concept may be used as a gust generator for a wind tunnel; however, more extensive investigation is required to completely define its capabilities and limitations. For the single transient data point recorded, a gust amplitude (decrement) of 0.15 in Mach number was obtained. Analysis of these transient data indicates a response with a corner frequency of at least 8 hertz.

Sanders, Bobby W.↗
NASA NTRS1989

Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take-off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller model to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

Hughes, Christopher E.↗
NASA NTRS1994

Numerical simulation of incidence and sweep effects on delta wing vortex breakdown

The structure of the vortical flowfield over delta wings at high angles of attack was investigated. Three-dimensional Navier-Stokes numerical simulations were carried out to predict the complex leeward-side flowfield characteristics, including leading-edge separation, secondary separation, and vortex breakdown. Flows over a 75- and a 63-deg sweep delta wing with sharp leading edges were investigated and compared with available experimental data. The effect of variation of circumferential grid resolution grid resolution in the vicinity of the wing leading edge on the accuracy of the solutions was addressed. Furthermore, the effect of turbulence modeling on the solutions was investigated. The effects of variation of angle of attack on the computed vortical flow structure for the 75-deg sweep delta wing were examined. At moderate angles of attack no vortex breakdown was observed. When a critical angle of attack was reached, bubble-type vortex breakdown was found. With further increase in angle of attack, a change from bubble-type breakdown to spiral-type vortex breakdown was predicted by the numerical solution. The effects of variation of sweep angle and freestream Mach number were addressed with the solutions on a 63-deg sweep delta wing.

INCIDENCE↗
NASA NTRS1955

Effect of centerbody boundary-layer removal near the throat of three coniccal nose inlets at Mach 1.6 to 2.0

A zero angle-of-attack investigation of the effect of compression-surface boundary-layer bleed through perforations near the throat of three full-scale conical nose inlets was conducted in the Lewis 8- by 6- foot supersonic wind tunnel for a Mach number range from 1.6 to 2.0. The bleed system increased pressure recovery, shifted the peak of the diffuser-discharge total-pressure profile toward the center-body, and decreased the range of stable inlet operation. A propulsion-system thrust minus drag analysis indicated that the increases in inlet pressure recovery were too small to compensate for the esimated bleed system drags.

Kremzier, Emil J↗
NASA NTRS1996

Evaluation of constant-Weber-number scaling for icing tests

Previous studies showed that for conditions simulating an aircraft encountering super-cooled water droplets the droplets may splash before freezing. Other surface effects dependent on the water surface tension may also influence the ice accretion process. Consequently, the Weber number appears to be important in accurately scaling ice accretion. A scaling method which uses a constant-Weber-number approach has been described previously; this study provides an evaluation of this scaling method. Tests are reported on cylinders of 2.5 to 15-cm diameter and NACA 0012 airfoils with chords of 18 to 53 cm in the NASA Lewis Icing Research Tunnel (IRT). The larger models were used to establish reference ice shapes, the scaling method was applied to determine appropriate scaled test conditions using the smaller models, and the ice shapes were compared. Icing conditions included warm glaze, horn glaze and mixed. The smallest size scaling attempted was 1/3, and scale and reference ice shapes for both cylinders and airfoils indicated that the constant-Weber-number scaling method was effective for the conditions tested.

Anderson, David N.↗