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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 847 records · Page 47
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Calculation of flow distribution in large radius ratio stages of axial flow turbines and comparison of theory and experiment

A method of calculating stage parameters and flow distribution of axial turbines is described. The governing equations apply to space between the blade rows and are based on the assumption of rotationally symmetrical, compressible, adiabatic flow conditions. Results are presented for stage design and flow analysis calculations. Theoretical results from the calculation system are compared with experimental data from low pressure steam turbine tests.

Herzog, J.↗
NASA NTRS1975

Composite engines for application to a single-stage-to-orbit vehicle

Seven composite engines were designed for application to a reusable single-stage-to-orbit vehicle. The engine designs were variations of the supercharged ejector ramjet engine. The resulting performance, weight, and drawings of each engine form a data base for establishing a potential of this class of composite engine to various missions, including the single-stage-to-orbit application. The impact of advanced technology in the design of the critical fan turbine was established.

Bendot, J. G.↗
NASA NTRS1976

Performance of low-pressure-ratio low-tip-speed fan stage with blade tip solidity of 0.65

The overall and blade-element performance of a low pressure ratio, low tip speed fan stage is presented over the stable operating range at rotative speeds from 90 to 120 percent of design speed. Stage peak efficiency of 0.927 was obtained at a weight flow of 32.4 kg/sec (190.31 kg/sec/sq m of annulus area) and a pressure ratio of 1.134. The stall margin at design speed and peak efficiency was 15.3 percent.

Kovich, G.↗
NASA NTRS1976

Effect of casing treatment on performance of an inlet stage for a transonic multistage compressor

An inlet stage of a transonic compressor was tested with three rotor tip casing treatment configurations: blade angle slots, circumferential grooves, and axial skewed slots. Significant increases in both rotor and stage total pressure ratio, total temperature ratio, efficiency, flow range, and very large improvements in stall margin were obtained with all three casing treatment configurations. The greatest improvement in performance was achieved with axial skewed slots.

Urasek, D. C.↗
NASA NTRS1976

Performance of a 1.57 pressure-ratio transonic fan stage with a screen-induced 90 deg circumferential inlet flow distortion

A transonic fan stage having a design pressure ratio of 1.57 was tested with a 90 degree circumferential distortion imposed on the inlet flow. The rotor diameter was approximately 50.8 cm, and the design pressure ratio was 1.60 at a tip speed of 425 m/sec. Overall performance at 70 and 100 percent of design speed showed a loss of stall pressure ratio and flow range at design speed and no significant loss in stall pressure ratio at 70 percent of design speed. Detailed flow measurements are presented to show the rotor-upstream flow interactions and the attenuation and amplification properties through the stage.

Sanger, N. L.↗
NASA NTRS1976

Aerodynamic performance of 0.4066-scale model to JT8D refan stage

The aerodynamic performance of a scale model of the split flow JT8D rafan stage is presented over a range of flows at speeds from 40 to 100 percent design. The bypass stage peak efficiency of 0.800 occurred at a total weight flow of 35.82 kilograms per second and a pressure ratio of 1.697. The stall margin was 15 percent based on pressure ratio and weight flow at stall and peak efficiency conditions. The data indicated that the hub region of the core stators was choked at design speed over the entire flow range tested.

Moore, R. D.↗
NASA NTRS1976

Single stage, low noise, advanced technology fan. Volume 5: Fan acoustics. Section 1: Results and analysis

The acoustic tests and data analysis for a 0.508-scale fan vehicle of a 111,300 newton (25,000 pound) thrust, full-size engine, which would have application on an advanced transport aircraft, is described. The single-stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec (1,650 ft/sec) to achieve the desired pressure ratio in a single-stage fan with low radius ratio (0.38), and to maintain adequate stall margin. The fan has 44 tip-shrouded rotor blades and 90 outlet guide vanes. The two basic approaches taken in the acoustic design were: (1) minimization of noise at the source, and (2) suppression of the generated noise in the inlet and bypass exhaust duct. Suppression of the generated noise was accomplished in the inlet through use of the hybrid concept (wall acoustic treatment plus airflow acceleration suppression) and in the exhaust duct with extensive acoustic treatment including a splitter. The goal of the design was attainment of twenty effective perceived noise decibels (20 EPNdB) below current Federal Air Regulation noise standards for a full-scale fan at the takeoff, cutback, and approach conditions. The suppression goal of FAR 36-20 was not reached, but improvements in the technology of both front and aft fan-noise suppression were realized. The suppressed fan noise was shown to be consistent with the proposed federal regulation on aircraft noise.

Jutras, R. R.↗
NASA NTRS1976

High density propellant for single stage to orbit vehicles

Mixed mode propulsion concepts are studied for advanced, single stage earth orbital transportation systems (SSTO) for use in the post-1990 time period. These propulsion concepts are based on the sequential and/or parallel use of high density impulse and high specific impulse propellants in a single stage to increase vehicle performance and reduce dry weight. Specifically, the mixed mode concept utilizes two propulsion systems with two different fuels (mode 1 and mode 2) with liquid oxygen as a common oxidizer. Mode 1 engines would burn a high bulk density fuel for lift-off and early ascent to minimize performance penalties associated with carrying fuel tankage to orbit. Mode 2 engines will complete orbital injection utilizing liquid hydrogen as the fuel.

Notardonato, J. J.↗
NASA NTRS1976

Experimental performance of a small size two stage depressed collector for a 4.8-9.6 GHz high performance TWT

Three simple small size two stage depressed collectors were designed and experimentally evaluated in conjunction with a 330 to 520 watt CW, 4.8 to 9.6 GHz traveling wave tube (TWT). Each of the three designs produced a minimum collector efficiency of 80.0 percent considering saturated TWT operation at the maximum of output frequency and at band edges. The highest minimum collector efficiency produced was 80.5 percent with a two stage depressed collector of 4.8 cm diameter by 7 cm high internal dimensions.

Ramins, P.↗
NASA NTRS1976

Description and expected performance of flight-model, 12-gigahertz, output stage tube for the communications technology satellite

The flight model output stage tube for the Communications Technology Satellite is described. The output stage tube is a 12-GHz, 200-W, coupled cavity traveling wave tube. The tube has a multistage depressed collector for efficiency enhancement. Collector cooling is accomplished by direct radiation to space. Expected rf performance and factors affecting on orbit performance and life are discussed.

Chomos, G. J.↗
NASA NTRS1976

An 82 to 84 percent efficient, small size, 2 and 4 stage depressed collector for octave bandwidth high performance TWT's

Improvements were carried out on a TWT (traveling wave tube) for use in EMC (electronic countermeasure) systems by applying MDC (multi-stage depressed collector) and spent-beam refocusing techniques. The MDC geometric design consists of six collecting elements (electrodes) at ground and cathode potentials operating in 2 or 4 stage configuration. The refocusing system designed allows a controlled beam expansion to the point that space charge becomes negligible at the MDC entrance. Simultaneously, the standard deviation of the radial velocity (angular spread) is reduced. This system provides an analog real time readout of P(recovered). Data on all tests is obtained with an automated data acquisition system.

Kosmahl, H.↗
NASA NTRS1977

Effects of tip clearance on overall performance of transonic fan stage with and without casing treatment

The overall performance of a transonic fan stage is presented for various tip clearances, with and without casing treatment. The stage was tested with a solid casing, and with open skewed slots and closed skewed slots in the casing over the rotor blade tips. Four nominal nonrotating rotor blade tip clearances from 0.061 to 0.178 centimeter were used. For all three casings, the pressure ratio and efficiency decreased with increasing tip clearance. The stall margin for a given casing also decreased with increasing clearance. At design speed and a given tip clearance, the highest stall margin was obtained with the open-slot casing, and the lowest stall margin was obtained with the solid casing.

Moore, R. D.↗
NASA NTRS1977

Acoustic and aerodynamic performance of a variable-pitch 1.83-meter-(6-ft) diameter 1.20-pressure-ratio fan stage (QF-9)

Far field noise data and related aerodynamic performance are presented for a variable pitch fan stage having characteristics suitable for low noise, STOL engine application. However, no acoustic suppression material was used in the flow passages. The fan was externally driven by an electric motor. Tests were made at several forward thrust rotor blade pitch angles and one for reverse thrust. Fan speed was varied from 60 to 120 percent of takeoff (design) speed, and exhaust nozzles having areas 92 to 105 percent of design were tested. The fan noise level was at a minimum at the design rotor blade pitch angles of 64 deg for takeoff thrust and at 57 deg for approach (50 percent takeoff thrust). Perceived noise along a 152.4-m sideline reached 100.1 PNdb for the takeoff (design) configuration for a stage pressure ratio of 1.17 and thrust of 57,600 N. For reverse thrust the PNL values were 4 to 5 PNdb above the takeoff values at comparable fan speeds.

Glaser, F. W.↗
NASA NTRS1977

Ultra high tip speed (670.6 m/sec) fan stage with composite rotor: Aerodynamic and mechanical design

A highly loaded, single-stage compressor having a tip speed of 670.6 m/sec was designed for the purpose of investigating very high tip speeds and high aerodynamic loadings to obtain high stage pressure ratios at acceptable levels of efficiency. The design pressure ratio is 2.8 at an adiabatic efficiency of 84.4%. Corrected design flow is 83.4 kg/sec; corrected design speed is 15,200 rpm; and rotor inlet tip diameter is 0.853 m. The rotor uses multiple-circular-arc airfoils from 0 to 15% span, precompression airfoils assuming single, strong oblique shocks from 21 to 43% span, and precompression airfoils assuming multiple oblique shocks from 52% span to the tip. Because of the high tip speeds, the rotor blades are designed to be fabricated of composite materials. Two composite materials were investigated: Courtaulds HTS graphite fiber in a Kerimid 601 polyimide matrix and the same fibers in a PMR polyimide matrix. In addition to providing a description of the aerodynamic and mechanical design of the 670.0 m/sec fan, discussion is presented of the results of structural tests of blades fabricated with both types of matrices.

Halle, J. E.↗
NASA NTRS1976

Single stage, low noise, advanced technology fan. Volume 5: Fan acoustics. Section 2: One-third octave data tabulations and selected narrowband traces

The raw-acoustic data corrected to standard day, from acoustic tests performed on a 0.508-scale fan vehicle of a 111,300 newton thrust, full-size engine, which has application on an advanced transport aircraft, are presented. The single-stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec to achieve the desired pressure ratio in a single-stage fan with low radius ratio, and to maintain adequate stall margin. The two basic approaches taken in the acoustic design were: (1) minimization of noise at the source, and (2) suppression of the generated noise in the inlet and bypass exhaust duct. Suppression of the generated noise was accomplished in the inlet through use of the hybrid concept (wall acoustic treatment plus airflow acceleration suppression) and in the exhaust duct with extensive acoustic treatment including a splitter. The goal of the design was attainment of twenty effective perceived noise decibels. The suppression goal of FAR 36-20 was not reached, but improvements in the technology of both front and aft fan-noise suppression were realized.

Jutras, R. R.↗
NASA NTRS1976

Stage efficiency in the analysis of thermochemical water decomposition processes

The procedure for analyzing thermochemical water-splitting processes using the figure of merit is expanded to include individual stage efficiencies and loss coefficients. The use of these quantities to establish the thermodynamic insufficiencies of each stage is shown. A number of processes are used to illustrate these concepts and procedures and to demonstrate the facility with which process steps contributing most to the cycle efficiency are found. The procedure allows attention to be directed to those steps of the process where the greatest increase in total cycle efficiency can be obtained.

Conger, W. L.↗
NASA NTRS1977

An evaluation of composite propulsion for single-stage-to-orbit vehicles designed for horizontal take-off

Composite propulsion was analyzed for single-stage-to-orbit vehicles designed for horizontal take-off. Trajectory, geometric, and mass analyses were performed to establish the orbital payload capability of six engines. The results indicated that none of the engines performed adequately to deliver payloads to orbit as analyzed. The single-stage turbine and oxidizer-rich gas generator resulted in a low engine specific impulse, and the performance increment of the ejector subsystem was less than that of a separate rocket system with a high combustion pressure. There was a benefit from incorporating a fan into the engine, and removal of the fan from the airstream during the ramjet mode increased the orbital payload capability.

Martin, J. A.↗