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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

High Temperature Mechanical Characterization and Analysis of Al2O3 /Al2O3 Composition

Sixteen ply unidirectional zirconia coated single crystal Al2O3 fiber reinforced polycrystalline Al2O3 was tested in uniaxial tension at temperatures to 1400 C in air. Fiber volume fractions ranged from 26 to 31%. The matrix has primarily open porosity of approximately 40%. Theories for predicting the Young's modulus, first matrix cracking stress, and ultimate strength were applied and evaluated for suitability in predicting the mechanical behavior of Al2O3/Al2O3 composites. The composite exhibited pseudo tough behavior (increased area under the stress/strain curve relative to monolithic alumina) from 22 to 1400 C. The rule-of-mixtures provides a good estimate of the Young's modulus of the composite using the constituent properties from room temperature to approximately 1200 C for short term static tensile tests in air. The ACK theory provides the best approximation of the first matrix cracking stress while accounting for residual stresses at room temperature. Difficulties in determining the fiber/matrix interfacial shear stress at high temperatures prevented the accurate prediction of the first matrix cracking stress above room temperature. The theory of Cao and Thouless, based on Weibull statistics, gave the best prediction for the composite ultimate tensile strength.

Gyekenyesi, John Z.↗
NASA NTRS1973

Subsonic and supersonic jets and supersonic suppressor characteristics.

The flow and acoustic characteristics over a Mach number range of 0.6 to 1.5 were determined for convergent and parallel flow nozzles with room temperature air. Mean velocity and fluctuating impact and static pressure distributions were investigated. Near- and far-field acoustic measurements were obtained, and overall sound power levels were compared with subsonic theory of Lighthill and supersonic theory of Nagamatsu and Horvay. Flow and acoustic characteristics of single and multiple shroud supersonic suppressors were studied. With 191 tubes and shrouds the overall sound power level for a Mach 1.4 jet was reduced 15.3 and 20.5 db respectively.

Nagamatsu, H. T.↗
NASA NTRS1991

Chemical vapor deposition coatings for oxidation protection of titanium alloys

Results of an experimental investigation of the oxidation protection afforded to Ti-14Al-21Nb and Ti-14Al-23Nb-2V titanium aluminides and Ti-17Mo-3Al-3Nb titanium alloy by aluminum-boron-silicon and boron-silicon coatings are presented. These coatings are applied by a combination of physical vapor deposition (PVD) and chemical vapor deposition (CVD) processes. The former is for the application of aluminum, and the latter is for codeposition of boron and silicon. Coating thickness is in the range of 2 to 7 microns, and coating weights are 0.6 to 2.0 mg/sq cm. Oxidation testing was performed in air at temperatures to 1255 K in both static and hypersonic flow environments. The degree of oxidation protection provided by the coatings is determined from weight change measurements made during the testing and post test compositional analyses. Temperature-dependent total normal emittance data are also presented for four coating/substrate combinations. Both types of coatings provided excellent oxidation protection for the exposure conditions of this investigation. Total normal emittances were greater than 0.80 in all cases.

Cunnington, G. R.↗
NASA NTRS1990

CARS temperature measurements in a hypersonic propulsion test facility

Static-temperature measurements performed in a reacting vitiated air-hydrogen Mach-2 flow in a duct in Test Cell 2 at NASA LaRC by using a coherent anti-Stokes Raman spectroscopy (CARS) system are discussed. The hypersonic propulsion Test Cell 2 hardware is outlined with emphasis on optical access ports and safety features in the design of the Test Cell. Such design considerations as vibration, noise, contamination from flow field or atmospheric-borne dust, unwanted laser- and electrically-induced combustion, and movement of the sampling volume in the flow are presented. The CARS system is described, and focus is placed on the principle and components of system-to-monochromator signal coupling. Contour plots of scramjet combustor static temperature in a reacting-flow region are presented for three stations, and it is noted that the measurements reveal such features in the flow as maximum temperature near the model wall in the region of the injector footprint.

Jarrett, O., Jr.↗
NASA NTRS1995

Characterization of a hot-film probe for hypersonic flow

The critical issues concerning the application of constant-temperature hot-film anemometry to hypersonic flow are reviewed and extended. Mass-flux static calibrations were conducted in a Mach 10 helium flow, while mass-flux and total-temperature static calibrations were made in a Mach 6 air flow. In addition, comparative hot-film/hot-wire turbulence measurements were made in a Mach 11 helium boundary layer to provide insight into the dynamic response of the hot film. The measurements indicate that substrate conduction 'losses' dominate the static response of the hot-film probe, thus resulting in poor sensitivity to mass-flux and total temperature. Furthermore, it has been found that it is not possible to isolate mass-flux fluctuations at high overheat ratios for the current hot-film design. Thus, the sapphire-substrate hot-film anemometer is a robust, high-bandwidth instrument limited to qualitative transition and turbulence measurements. Finally, the extension of this technique to providing quantitative information is dependent upon the development of lower thermal-conductivity substrate materials.

Sheplak, M.↗
NASA NTRS1980

Ground-based measurements of DM-2 rocket exhaust effluents using fixed-flow samplers and electrets

Electret devices were used to collect samples of rocket exhaust products. The samples were analyzed by X-ray spectroscopy. The performance of the electrets was compared with that of Thiokol fixed flow samplers during the static test firing of a solid rocket motor (SRM) demonstration model. A vertical profile of wind speed, wind direction, and air temperature, used as inputs to the multilayer diffusion model for the SRM static test is presented.

Susko, M.↗
NASA NTRS1983

Numerical predictions of residence times behind a rearward facing step with transverse injection

A numerical investigation of the two-dimensional recirculation zone formed between a rear-ward facing step and a transverse hydrogen jet in a supersonic flow is presented. Calculations were made using an elliptic flow computer code with a time-relaxation finite difference integration algorithm. Cases were examined for jet static pressures between 2 and 5 times mainstream with jets located 3 to 7 step heights downstream of the step. Nominal mainstream flow conditions were Mach 2.4 air at 1 atm pressure and 1000K (1800R) static temperature. Results in the form of mean properties in the recirculation zone and residence time are presented. Calculations of the decay of hydrogen mass from the steady state are presented for selected cases. The combustion potential for each case is assessed by comparison of computed conditions in the recirculation zone with the results from a well-stirred reactor theory.

Rogers, R. C.↗
NASA NTRS2005

Airdata Measurement and Calibration

This Section provides a brief introduction to airdata measurement and calibration. Readers will learn about typical test objectives, quantities to measure, and flight maneuvers and operations for calibration. The Section informs readers about tower-flyby, trailing cone, pacer, radar-tracking, and dynamic airdata calibration maneuvers. Readers will also begin to understand how some data analysis considerations and special airdata cases, including high-angle-of-attack flight, high-speed flight, and nonobtrusive sensors are handled. This section is not intended to be all inclusive; readers should review AGARDograph 300, Volume 1, "Calibration of Airdata Systems and Flow Direction Sensors" for more detailed information. [11-1] References 11-2, 11-3, and 11-4 also supply pertinent information to better understand airdata measurement and calibration and related terminology. Airdata are vital to successfully complete an aircraft's mission and are derived from the air surrounding the aircraft. These airdata encompass indicated and true airspeed, pressure altitude, ambient air temperature, angles of attack and sideslip, Mach number, and rate of climb. Typically, pitot and static pressures are sensed and converted (by mechanical means in the instruments themselves) into indications on the altimeter, vertical speed indicator, airspeed indicator, and Machmeter. Similarly, measured local flow angles establish angles of attack and sideslip, and the outside air temperature is measured and indicated in the cockpit. (Instruments that can perform the conversion, such as airspeed indicators, altimeters, and Machmeters, do not correct for errors in the input values.) These measured parameters are commonly input to the airdata computer which, using appropriate algorithms and correction factors (or calibrations, as discussed later), can provide other parameters, such as true airspeed, required by the aircraft's avionics or flight control system. The presence of the aircraft in the airstream causes input errors to the measuring instruments - the aircraft disturbs the air that it flies through, thereby also disturbing the airdata measurements. Figure 11-1 shows the airflow around an airplane wing. The air above the wing has lower pressure than the ambient air, while the pressure below the wing is higher than the ambient air. Compressibility and shock waves also disturb the air and affect the measurements. Compressibility effects become important above approximately Mach number 0.3. As a result the static pressure around an airplane varies considerably with location. Local flow angles also differ from the free-stream flow direction. In straight-and-level flight the airflow rises to the wing leading edge and falls below the trailing edge, causing errors in flow direction measurements. To some extent these errors can be studied in wind tunnels, but wind-tunnel measurements cannot replace in-flight measurements.

Haering, Edward A., Jr.↗
NASA NTRS1973

Analytical and experimental study of supersonic combustion of hydrogen in a vitiated airstream

Detailed probe measurements of total temperature, pressure, and composition were taken in a two-dimensional test section 35.6 cm downstream of hydrogen injection. A high pressure gas generator supplied Mach 2.44 vitiated air or inert gas at elevated temperatures and at a static pressure equal to that of the hydrogen. Special water-cooled probes and sampling techniques were developed for the short test times required by heat-sink hardware. Independent methods of measuring stream total temperatures are compared. For the pure mixing case, the computed composition profile agreed well with the experimental profile. The analysis takes into account the wall boundary layer and the initial boundary layer in the main stream. Ignition of hydrogen, as determined from photographic exposures of the radiating gases, varied from 30 to 10 cm downstream from injection for a 45 K increase in local free-stream static temperature.

Burrows, M. C.↗
NASA NTRS2008

Preliminary Test Results of a Non-Contacting Finger Seal on a Herringbone-Grooved Rotor

Low leakage, non-contacting finger seals have potential to reduce gas turbine engine specific fuel consumption by 2 to 3 percent and to reduce direct operating costs by increasing the time between engine overhauls. A non-contacting finger seal with concentric lift-pads operating adjacent to a test rotor with herringbone grooves was statically tested at 300, 533, and 700 K inlet air temperatures at pressure differentials up to 576 kPa. Leakage flow factors were approximately 70 percent less than state-of-the-art labyrinth seals. Leakage rates are compared to first order predictions. Initial spin tests at 5000 rpm, 300 K inlet air temperature and pressure differentials to 241 kPa produced no measurable wear.

Proctor, Margaret P.↗
NASA NTRS2008

Preliminary Test Results of a Non-Contacting Finger Seal on a Herringbone-Grooved Rotor

Low leakage, non-contacting finger seals have potential to reduce gas turbine engine specific fuel consumption by 2 to 3 percent and to reduce direct operating costs by increasing the time between engine overhauls. A non-contacting finger seal with concentric lift-pads operating adjacent to a test rotor with herringbone grooves was statically tested at 300, 533, and 700 K inlet air temperatures at pressure differentials up to 576 kPa. Leakage flow factors were approximately 70 percent less than state-of-the-art labyrinth seals. Leakage rates are compared to first order predictions. Initial spin tests at 5000 rpm, 300 K inlet air temperature and pressure differentials to 241 kPa produced no measurable wear.

Proctor, Margaret P.↗
NASA NTRS2017

Modeling of a Turbofan Engine with Ice Crystal Ingestion in the NASA Propulsion System Laboratory

The main focus of this study is to apply a computational tool for the flow analysis of the turbine engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The PSL has been used to test a highly instrumented Honeywell ALF502R-5A (LF11) turbofan engine at simulated altitude operating conditions. Test data analysis with an engine cycle code and a compressor flow code was conducted to determine the values of key icing parameters, that can indicate the risk of ice accretion, which can lead to engine rollback (un-commanded loss of engine thrust). The full engine aerothermodynamic performance was modeled with the Honeywell Customer Deck specifically created for the ALF502R-5A engine. The mean-line compressor flow analysis code, which includes a code that models the state of the ice crystal, was used to model the air flow through the fan-core and low pressure compressor. The results of the compressor flow analyses included calculations of the ice-water flow rate to air flow rate ratio (IWAR), the local static wet bulb temperature, and the particle melt ratio throughout the flow field. It was found that the assumed particle size had a large effect on the particle melt ratio, and on the local wet bulb temperature. In this study the particle size was varied parametrically to produce a non-zero calculated melt ratio in the exit guide vane (EGV) region of the low pressure compressor (LPC) for the data points that experienced a growth of blockage there, and a subsequent engine called rollback (CRB). At data points where the engine experienced a CRB having the lowest wet bulb temperature of 492 degrees Rankine at the EGV trailing edge, the smallest particle size that produced a non-zero melt ratio (between 3 percent - 4 percent) was on the order of 1 micron. This value of melt ratio was utilized as the target for all other subsequent data points analyzed, while the particle size was varied from 1 micron - 9.5 microns to achieve the target melt ratio. For data points that did not experience a CRB which had static wet bulb temperatures in the EGV region below 492 degrees Rankine, a non-zero melt ratio could not be achieved even with a 1 micron ice particle size. The highest value of static wet bulb temperature for data points that experienced engine CRB was 498 degrees Rankine with a particle size of 9.5 microns. Based on this study of the LF11 engine test data, the range of static wet bulb temperature at the EGV exit for engine CRB was in the narrow range of 492 degrees Rankine - 498 degrees Rankine , while the minimum value of IWAR was 0.002. The rate of blockage growth due to ice accretion and boundary layer growth was estimated by scaling from a known blockage growth rate that was determined in a previous study. These results obtained from the LF11 engine analysis formed the basis of a unique “icing wedge.”

Turbo engine↗
NASA NTRS1998

High Temperature Mechanical Characterization of Ceramic Matrix Composites

A high temperature mechanical characterization laboratory has been assembled at NASA Lewis Research Center. One contribution of this work is to test ceramic matrix composite specimens in tension in environmental extremes. Two high temperature tensile testing systems were assembled. The systems were assembled based on the performance and experience of other laboratories and meeting projected service conditions for the materials in question. The systems use frames with an electric actuator and a center screw. A PC based data acquisition and analysis system is used to collect and analyze the data. Mechanical extensometers are used to measure specimen strain. Thermocouples, placed near the specimen, are used to measure the specimen gage section temperature. The system for testing in air has a resistance element furnace with molybdenum disilicide elements and pneumatic grips with water cooling attached to hydraulic alignment devices. The system for testing in an inert gas has a graphite resistance element furnace in a chamber with rigidly mounted, water cooled, hydraulically actuated grips. Unidirectional SiC fiber reinforced reaction bonded Si3N4 and triaxially woven, two dimensional, SiC fiber reinforced enhanced SiC composites were tested in unidirectional tension. Theories for predicting the Young's modulus, modulus near the ultimate strength, first matrix cracking stress, and ultimate strength were applied and evaluated for suitability in predicting the mechanical behavior of SiC/RBSN and enhanced SiC/SiC composites. The SiC/RBSN composite exhibited pseudo tough behavior (increased area under the stress/strain curve) from 22 C to 1500 C. The rule of mixtures provides a good estimate of the Young's modulus of the SiC/RBSN composite using the constituent properties from room temperature to 1440 C for short term static tensile tests in air or nitrogen. The rule of mixtures significantly overestimates the secondary modulus near the ultimate strength. The ACK theory provides the best approximation of the first matrix cracking stress when residual stresses are ignored. The theory of Cao and Thouless, based on Weibull statistics, gave the best prediction for the composite ultimate strength. The enhanced SiC/SiC composite exhibited nonlinear stress/strain behavior from 24 C to 1370 C in air with increased ultimate strain when compared to monolithic SiC. The theory of Yang and Chou with the assumption of a frictional fiber/matrix interface provided the best estimate of the Young's modulus. The theory of Cao and Thouless gave the best estimate for the ultimate strength.

Gyekenyesi, John Z.↗
NASA NTRS2008

Simultaneous Temperature and Velocity Measurements in a Large-Scale, Supersonic, Heated Jet

Two laser-based measurement techniques have been used to characterize an axisymmetric, combustion-heated supersonic jet issuing into static room air. The dual-pump coherent anti-Stokes Raman spectroscopy (CARS) measurement technique measured temperature and concentration while the interferometric Rayleigh scattering (IRS) method simultaneously measured two components of velocity. This paper reports a preliminary analysis of CARS-IRS temperature and velocity measurements from selected measurement locations. The temperature measurements show that the temperature along the jet axis remains constant while dropping off radially. The velocity measurements show that the nozzle exit velocity fluctuations are about 3% of the maximum velocity in the flow.

Danehy, P. M.↗
NASA NTRS1971

Analysis of heat-transfer tests of an impingement-convection- and film-cooled vane in a cascade

Experimental flow and heat transfer data obtained for an air-cooled turbine vane tested in a static cascade at gas temperatures and pressures to 1644 K (2500 F) and 31 N/cm2 (45 psia), respectively, are presented. Average and local vane temperatures were correlated in several ways. Calculated and measured coolant flows and vane temperatures are compared. Potential allowable increases in gas temperature are also discussed.

Gladden, H. J.↗
NASA NTRS1976

New aromatic silicon-containing polyamides

The experiments described were aimed at studying systematically the effect of silyl linkages on the thermogravimetric and thermomechanical properties of aromatic polyamides. Specifically, 18 aromatic silicon-containing polyamides were synthesized via interfacial polymerization of six silicon-containing diacid chlorides with 3,3-prime-diaminodiphenylmethane, 3,3-prime-diaminobenzophenone, and 1-(3-prime-aminobenzyl)-4-(3-double prime-aminobenzoyl)benzene. All polyamides were soluble in m-cresol and N,N-dimethylacetamide, and had glass transition temperatures between 178 and 254 C. Thermogravimetric analysis conducted in static air on film specimens showed 5 and 10 percent weight losses between 331 and 400 C and 354 and 440 C, respectively. The potential gain in processability engendered by incorporation of silane substituents was offset in most cases by substantial losses in thermo-oxidative stability.

Pratt, J. R.↗
NASA NTRS1983

Numerical predictions of residence times behind a rearward facing step with transverse injection

A numerical investigation of the two-dimensional recirculation zone formed between a rearward facing step and a transverse hydrogen jet in a supersonic flow is presented. Calculations were made using an elliptic flow computer code with a time-relaxation finite difference integration algorithm. Cases were examined for jet static pressures between 2 and 5 times mainstream with jets located 3 to 7 step heights downstream of the step. Nominal mainstream flow conditions were Mach 2.4 air at 1 atm pressure and 1000 K (1800 R) static temperature. Results in the form of mean properties in the recirculation zone and residence time are presented. Calculations of the decay of hydrogen mass from the steady state are presented for selected cases. The combustion potential for each case is assessed by comparison of computed conditions in the recirculation zone with the results from a well-stirred reactor theory.

Rogers, R. C.↗