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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 NTRS1981

Ion streams in the magnetotail

Ion mass spectrometer observations of low-temperature streaming plasmas in the earth's magnetotail are reported. Measurements in the energy per charge range 0 to 17 keV/e were made at geocentric radial distances less than 23 earth radii from the ISEE 1 spacecraft. Ion streams of solar wind origin in the magnetotail boundary layer and of ionospheric origin in the tail lobes and plasma sheet are described. A statistical study of some of the characteristics of the streams allows the inference that the central tail lobe plasmas are primarily constituted of ion streams of ionospheric origin and that the ionosphere is a significant contributor to the hot plasmas that form the plasma sheet.

Sharp, R. D.↗
NASA NTRS1982

Dilution jet mixing

Schematics are given for test section configurations, orifice configurations, and the dilution jet mixing test rig used to collect a data base on mixing a single sided and a two sided row of jets with a confined cross flow. Parameters investigated include momentum ratio; nonuniform cross stream temperature and velocity profiles; cold/hot injection, and cross stream flow area convergence. Graphs show measured theta momentum distributions for: (1) in line and staggered orifice configurations; and (2) the profiled mainstream; and (3) flow area convergence.

Srinivasan, S.↗
NASA NTRS1953

Analytical Investigation of Icing Limit for Diamond-Shaped Airfoil in Transonic and Supersonic Flow

Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.

Callaghan, Edmund E.↗
NASA NTRS1987

Stability relationship for water droplet crystallization with the NASA Lewis icing spray

In order to produce small droplets for icing cloud simulation, high pressure air atomizing nozzles are used. For certain icing testing applications, median drop sizes as small as 5 mm are needed, which require air atomizing pressures greater than 3000 kPa. Isentropic expansion of the ambient temperature atomizing air to atmospheric pressure can result in air stream temperatures of -160 C which results in ice crystals forming in the cloud. To avoid such low temperatures, it is necessary to heat the air and water to high initial temperatures. An icing spray research program was conducted to map the temperatures below which ice crystals form. A soot slide technique was used to determine the presence of crystals in the spray.

Marek, C. John↗
NASA NTRS1988

Stability relationship for water droplet crystallization with the NASA Lewis icing spray nozzle

In order to produce small droplets for icing cloud simulation, high pressure air atomizing nozzles are used. For certain icing testing applications, median drop sizes as small as 5 mm are needed, which require air atomizing pressures greater than 3000 kPa. Isentropic expansion of the ambient temperature atomizing air to atmospheric pressure can result in air stream temperatures of -160 C which results in ice crystals forming in the cloud. To avoid such low temperatures, it is necessary to heat the air and water to high initial temperatures. An icing spray research program was conducted to map the temperatures below which ice crystals form. A soot slide technique was used to determine the presence of crystals in the spray.

Marek, C. John↗
NASA NTRS2015

Numerical Modeling of the Transient Chilldown Process of a Cryogenic Propellant Transfer Line

Before cryogenic fuel depots can be fully realized, efficient methods with which to chill down the spacecraft transfer line and receiver tank are required. This paper presents numerical modeling of the chilldown of a liquid hydrogen tank-to-tank propellant transfer line using the Generalized Fluid System Simulation Program (GFSSP). To compare with data from recently concluded turbulent LH2 chill down experiments, seven different cases were run across a range of inlet liquid temperatures and mass flow rates. Both trickle and pulse chill down methods were simulated. The GFSSP model qualitatively matches external skin mounted temperature readings, but large differences are shown between measured and predicted internal stream temperatures. Discrepancies are attributed to the simplified model correlation used to compute two-phase flow boiling heat transfer. Flow visualization from testing shows that the initial bottoming out of skin mounted sensors corresponds to annular flow, but that considerable time is required for the stream sensor to achieve steady state as the system moves through annular, churn, and bubbly flow. The GFSSP model does adequately well in tracking trends in the data but further work is needed to refine the two-phase flow modeling to better match observed test data.

Chilldown↗
NASA NTRS2012

Wind-Tunnel and Flight Test Results for the Measurements of Flow Variables at Supersonic Speeds Using Improved Wedge and Conical Probes

The results of supersonic wind-tunnel tests on three probes at nominal Mach numbers of 1.6, 1.8 and 2.0 and flight tests on two of these probes up to a Mach number of 1.9 are described. One probe is an 8 deg. half-angle wedge with two total-pressure measurements and one static. The second, a conical probe, is a cylinder that has a 15 deg., semi-angle cone tip with one total-pressure orifice at the apex and four static-pressure orifices on the surface of the cone, 90 deg. apart, and about two-thirds of the distance from the cone apex to the base of the cone. The third is a 2 deg. semi-angle cone that has two static ports located 180 deg. apart about 1.5 inches behind the apex of the cone. The latter probe was included since it has been the "probe of choice" for wind-tunnel flow-field pressure measurements (or one similar to it) for the past half-century. The wedge and 15 deg. conical probes used in these tests were designed for flight diagnostic measurements for flight Mach numbers down to 1.35 and 1.15 respectively, and have improved capabilities over earlier probes of similar shape. The 15. conical probe also has a temperature sensor that is located inside the cylindrical part of the probe that is exposed to free-stream flow through an annulus at the apex of the cone. It enables the determination of free-stream temperature, density, speed of sound, and velocity, in addition to free-stream pressure, Mach number, angle of attack and angle of sideslip. With the time-varying velocity, acceleration can be calculated. Wind-tunnel tests of the two probes were made in NASA Langley Research Center fs Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8, and 2.0. Flight tests were carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15B aircraft up to Mach numbers of 1.9. The probes were attached to a fixture, referred to as the Centerline Instrumented Pylon (CLIP), under the fuselage of the aircraft. Problems controlling the velocity of the flow through the conical probe required for accurate temperature measurements are noted, as well as some calibration problems of the miniature pressure sensors that required a re-calculation of the flow variables. Data are presented for angle of attack, pressure and Mach number obtained in the wind tunnel and in flight. In the wind tunnel some transient data were obtained by translating the probes through the shock flow field created by a bump on the wind-tunnel wall.

Bobbitt, Percy J.↗
NASA NTRS1973

Fluidic device for measuring constituent masses of a flowing binary gas mixture

Device consists of fluidic humidity sensor and specially designed flow calorimeter. Calorimeter provides readings of gas stream temperature rise produced by measured amount of heat that is dissipated into gas stream, and humidity sensor is used to obtain continuous calculation of specific heat capacity of gas mixture.

Prokopius, P. R.↗
NASA NTRS1973

Remote-sensing studies of hydrologic environments in the lower Raritan River System, New Jersey

A series of remote sensing experiments were conducted in January and November 1968. The airborne multisensor missions included photographic and thermal-infrared sensors operated over a single site at a time when streamflow and water temperature observations were being made on the ground. Remote sensing data show: (1) effect of thermal waste water discharges on stream temperatures, (2) cross-channel variations in thermal characteristics due to waste water discharge, channel characteristics, and tidal currents, (3) influence of flow rates on dispersion, (4) patterns and distribution of ice cover, and (5) movement of sediment loads.

Anderson, P. W.↗
NASA NTRS1985

Transient technique for measuring heat transfer coefficients on stator airfoils in a jet engine environment

A transient technique was used to measure heat transfer coefficients on stator airfoils in a high-temperature annular cascade at real engine conditions. The transient response of thin film thermocouples on the airfoil surface to step changes in the gas stream temperature was used to determine these coefficients. In addition, gardon gages and paired thermocouples were also utilized to measure heat flux on the airfoil pressure surface at steady state conditions. The tests were conducted at exit gas stream Reynolds numbers of one-half to 1.9 million based on true chord. The results from the transient technique show good comparison with the steady-state results in both trend and magnitude. In addition, comparison is made with the STAN5 boundary layer code and shows good comparison with the trends. However, the magnitude of the experimental data is consistently higher than the analysis.

Gladden, H. J.↗
NASA NTRS1985

Transient technique for measuring heat transfer coefficients on stator airfoils in a jet engine environment

A transient technique was used to measure heat transfer coefficients on stator airfoils in a high-temperature annular cascade at real engine conditions. The transient response of thin film thermocouples on the airfoil surface to step changes in the gas stream temperature was used to determine these coefficients. In addition, gardon gages and paired thermocouples were also utilized to measure heat flux on the airfoil pressure surface at steady state conditions. The tests were conducted at exit gas stream Reynolds numbers of one-half to 1.9 million based on true chord. The results from the transient technique show good comparison with the steady-state results in both trend and magnitude. In addition, comparison is made with the STAN5 boundary layer code and shows good comparison with the trends. However, the magnitude of the experimental data is consistently higher than the analysis.

Gladden, H. J.↗
NASA NTRS1961

Recovery Temperature, Transition, and Heat Transfer Measurements at Mach 5

Schlieren, recovery temperature, and heat-transfer measurements were made on a hollow cylinder and a cone with axes alined parallel to the stream. Both the cone and cylinder were equipped with various bluntnesses, and the tests covered a Reynolds number range up to 20 x 10(exp 6) at a free-stream Mach number of 4.95 and wall to free-stream temperature ratios from 1.8 to 5.2 (adiabatic). A substantial transition delay due to bluntness was found for both the cylinder and the cone. For the present tests (Mach 4.95), transition was delayed by a factor of 3 on the cylinder and about 2 on the cone, these delays being somewhat larger than those observed in earlier tests at Mach 3.1. Heat-transfer tests on the cylinder showed only slight effects of wall temperature level on transition location; this is to be contrasted to the large transition delays observed on conical-type bodies at low surface temperatures at Mach 3.1. The schlieren and the peak-recovery-temperature methods of detecting transition were compared with the heat-transfer results. The comparison showed that the first two methods identified a transition point which occurred just beyond the end of the laminar run as seen in the heat-transfer data.

Brinich, Paul F.↗
NASA NTRS1978

Tabulation and summary of thermodynamic effects data for developed cavitation on ogive-nosed bodies

Thermodynamic effects data for developed cavitation on zero and quarter caliber ogives in Freon 113 and water are tabulated and summarized. These data include temperature depression (delta T), flow coefficient (C sub Q), and various geometrical characteristics of the cavity. For the delta T tests, the free-stream temperature varied from 35 C to 95 C in Freon 113 and from 60 C to 125 C in water for a velocity range of 19.5 m/sec to 36.6 m/sec. Two correlations of the delta T data by the entrainment method are presented. These correlations involve different combinations of the Nusselt, Reynolds, Froude, Weber, and Peclet numbers and dimensionless cavity length.

Holl, J. W.↗
NASA NTRS1991

Heat transfer, velocity-temperature correlation, and turbulent shear stress from Navier-Stokes computations of shock wave/turbulent boundary layer interaction flows

The properties of 2-D shock wave/turbulent boundary layer interaction flows were calculated by using a compressible turbulent Navier-Stokes numerical computational code. Interaction flows caused by oblique shock wave impingement on the turbulent boundary layer flow were considered. The oblique shock waves were induced with shock generators at angles of attack less than 10 degs in supersonic flows. The surface temperatures were kept at near-adiabatic (ratio of wall static temperature to free stream total temperature) and cold wall (ratio of wall static temperature to free stream total temperature) conditions. The computational results were studied for the surface heat transfer, velocity temperature correlation, and turbulent shear stress in the interaction flow fields. Comparisons of the computational results with existing measurements indicated that (1) the surface heat transfer rates and surface pressures could be correlated with Holden's relationship, (2) the mean flow streamwise velocity components and static temperatures could be correlated with Crocco's relationship if flow separation did not occur, and (3) the Baldwin-Lomax turbulence model should be modified for turbulent shear stress computations in the interaction flows.

Wang, C. R.↗
NASA NTRS1992

Full system engineering design and operation of an oxygen plant

The production of oxygen from the indigenous resources on Mars is described. After discussing briefly the project's background and the experimental system design, specific experimental results of the electrolytic cell are presented. At the heart of the oxygen production system is a tubular solid zirconia electrolyte cell that will electrochemically separate oxygen from a high-temperature stream of Coleman grade carbon dioxide. Experimental results are discussed and certain system efficiencies are defined. The parameters varied include (1) the cell operating temperature; (2) the carbon dioxide flow rate; and (3) the voltage applied across the cell. The results confirm our theoretical expectations.

Colvin, James↗
NASA NTRS1959

Turbulent Boundary Layer on a Yawed Cone in a Supersonic Stream

The momentum integral equations are derived for the boundary layer on an arbitrary curved surface, using a streamline coordinate system. Computations of the turbulent boundary layer on a slightly yawed cone are made for a Prandtl number of 0.729, wall to free-stream temperature ratios of 1/2, 1, and 2, and Mach numbers from 1 to 4. Deflection of the fluid in the boundary layer from outer stream direction, local friction coefficient, displacement surface, lift coefficient, and pitching-moment coefficient are presented.

Braun, Willis H.↗
NASA NTRS1991

Propellant production on Mars - Single cell oxygen production test bed

A study focusing on oxygen production using resources indigenous to Mars is presented. A bank of solid zirconia electrolytic cells that will electrochemically separate oxygen from a high temperature stream of carbon dioxide is at the center of the oxygen production system. The experimental data are discussed with attention given to the cell operating temperature, the carbon dioxide flow rate, and the voltage applied across the cell.

Colvin, James↗
NASA NTRS1961

Measurement and Empirical Correlation of Transpiration-Cooling Parameters on a 25 degree Cone in a Turbulent Boundary Layer in Both Free Flight and a Hot-Gas Jet

Transpiration-cooling parameters are presented for a turbulent boundary layer on a cone configuration with a total angle of 250 which was tested in both free flight and in an ethylene-heated high-temperature jet at a Mach number of 2.0. The flight-tested cone was flown to a maximum Mach number of 4.08 and the jet tests were conducted at stagnation temperatures ranging from 937 R to 1,850 R. In general, the experimental heat transfer was in good agreement with the theoretical values. Inclusion of the ratio of local stream temperature to wall temperature in the nondimensional flow rate parameter enabled good correlation of both sets of transpiration data. The measured pressure at the forward station coincided with the theoretical pressure over a sharp cone; however, the measured pressure increased with distance from the nose tip.

Walton, Thomas E., Jr.↗