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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Survey of Primary Flow Measurement Parameters at the NASA Langley Transonic Dynamics Tunnel

An assessment of the methods and locations used to measure the primary flow conditions in the NASA Langley Transonic Dynamics Tunnel was conducted during calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to R-134a. A survey of stagnation pressure, plenum static pressure, and stagnation temperature was undertaken at many pertinent locations in the settling chamber, plenum, and contraction section of the wind tunnel and these measurements were compared to those of the existing primary flow measurement systems. Local flow velocities were measured in the settling chamber using a pitot probe. Results illustrate that small discrepancies exist between measured primary tunnel flow conditions and the survey measurements. These discrepancies in tunnel stagnation pressure, plenum pressure, and stagnation temperature were found to be approximately +/- 1-3 psf and 2-3 degrees Fahrenheit. The propagation of known instrument errors in measured primary flow conditions and its impact on tunnel Mach number, dynamic pressure, flow velocity, and Reynolds number have been investigated analytically and shown to require careful attention when considering the uncertainty in measured test section conditions.

Piatak, David J.↗
NASA NTRS1974

Venting characteristics of gaseous helium and nitrogen discharging into a free stream at Mach numbers from 0.60 to 1.57

An experimental investigation was conducted in the Lewis Research Center 8- by 6-Foot Supersonic Wind Tunnel to compare the effects on discharge coefficient of venting gaseous helium and gaseous nitrogen into a free stream. The test was conducted from Mach 0.06 to 1.57 with the vent mounted in a flat plate. The plate was strut mounted to the tunnel ceiling and at a 0 angle of attack. The gases were discharged from a plenum chamber through a 2.54-centimeter (1.00-in.) diameter vent. The ratio of local static pressure to plenum pressure was varied from 0.51 to 0.975. The ratio of boundary layer thickness to vent diameter varied from a maximum of 1.34 at Mach 0.60 to a minimum of 0.55 at Mach 1.37.

Johns, A. L.↗
NASA NTRS1981

Advanced Low-Emissions Catalytic-Combustor Program, phase 1

Six catalytic combustor concepts were defined, analyzed, and evaluated. Major design considerations included low emissions, performance, safety, durability, installations, operations and development. On the basis of these considerations the two most promising concepts were selected. Refined analysis and preliminary design work was conducted on these two concepts. The selected concepts were required to fit within the combustor chamber dimensions of the reference engine. This is achieved by using a dump diffuser discharging into a plenum chamber between the compressor discharge and the turbine inlet, with the combustors overlaying the prediffuser and the rear of the compressor. To enhance maintainability, the outer combustor case for each concept is designed to translate forward for accessibility to the catalytic reactor, liners and high pressure turbine area. The catalytic reactor is self-contained with air-cooled canning on a resilient mounting. Both selected concepts employed integrated engine-starting approaches to raise the catalytic reactor up to operating conditions. Advanced liner schemes are used to minimize required cooling air. The two selected concepts respectively employ fuel-rich initial thermal reaction followed by rapid quench and subsequent fuel-lean catalytic reaction of carbon monoxide, and, fuel-lean thermal reaction of some fuel in a continuously operating pilot combustor with fuel-lean catalytic reaction of remaining fuel in a radially-staged main combustor.

Sturgess, G. J.↗
NASA NTRS1999

Numerical Investigation of the ARMSEF Arc-Jet

The NASA/JSC 10 MW Atmospheric Reentry Materials and Structures Evaluation Facility (ARMSEF) is used for the development and evaluation of thermal protection systems for manned reentry vehicles. In an effort at better understanding the flow characteristics of an arc-jet wind tunnel and hence the correlation between simulation and actual flight conditions, preliminary numerical analyses have been made. Difficulties in defining inlet boundary conditions to the plenum chamber, where there is very little usable experimental data, are discussed. Two different boundary conditions for the plenum, a straight profile and parabolic profile, are employed to describe the conditions in this very important flow defining region. Comparisons are made to experimental data with the results showing that some type of parabolic or step profile is needed to adequately resolve the strong core flow that is present in the nozzle and plume region. Recommendations are made regarding future efforts in analyzing this challenging flow field.

Cox, Carey F.↗
NASA NTRS2002

Progress in the Development of a Nozzle Design Methodology for Pulsed Detonation Engines

The Johns Hopkins University Applied Physics Laboratory (JHU/APL), in support of the NASA Glenn Research Center (NASA GRC), is investigating performance methodologies and system integration issues related to Pulsed Detonation Engine (PDE) nozzles. The primary goal of this ongoing effort is to develop design and performance assessment methodologies applicable to PDE exit nozzle(s). APL is currently focusing its efforts on a common plenum chamber design that collects the exhaust products from multiple PDE tubes prior to expansion in a single converging-diverging exit nozzle. To accomplish this goal, a time-dependent, quasi-one-dimensional analysis for determining the flow properties in and through a single plenum and exhaust nozzle is underway. In support of these design activities, parallel modeling efforts using commercial Computational Fluid Dynamics (CFD) software are on-going. These efforts include both two and three-dimensional as well as steady and time-dependent computations to assess the flow in and through these devices. This paper discusses the progress in developing this nozzle design methodology.

Leary, B. A.↗
NASA NTRS2010

Emission Spectroscopic Measurements in the Plenum Region of the NASA IHF Arc Jet Facility

A newly designed segment with optical access was installed in the plenum chamber of the 60 MW Interaction Heating arcjet Facility at NASA Ames Research Center. This special segment has ports located off axis, and the optical fibers can be inserted into these ports. The special segment allows for optical examination of the arc-heated flow as it enters the plenum, and thus assists in determining estimates of the thermodynamic state of the inflow to the convergent section of the nozzle. In the present work, optical emission measurements have been made in VIS-NIR region (wavelengths between 500 nm to 900 nm) for two settings of the arc heater - a 6000 A condition (high condition) with the minimum amount of radial injection of cold air in the plenum, and a 3300 A condition (low condition) with significant amount of cold air injection to reduce the enthalpy of the arc-heated stream. The results presented here were obtained using an Acton SP300i spectrometer coupled to a Princeton Instruments PI-max intensified camera. In addition to the optical emission measurements, computations were performed for the flow in the plenum and radiation along lines of sight corresponding to the optical ports. Along the centerline, i.e., the longest line of sight across the plenum cross-section, there is good agreement between computations and measurements for the high enthalpy condition, although the off-axis radial profiles show some differences. For the low enthalpy condition, there are significant differences between computations and measurements. The current working hypothesis is that the computational model does not capture details of the mixing process in the plenum.

Winter, Michael W.↗
NASA NTRS1981

A theory for predicting boundary impedance and resonance frequencies of slotted-wall wind tunnels, including plenum effects

Wave-induced resonance associated with the geometry of wind-tunnel test sections can occur. A theory that uses acoustic impedance concepts to predict resonance modes in a two dimensional, slotted wall wind tunnel with a plenum chamber is described. The equation derived is consistent with known results for limiting conditions. The computed resonance modes compare well with appropriate experimental data. When the theory is applied to perforated wall test sections, it predicts the experimentally observed closely spaced modes that occur when the wavelength is not long compared with he plenum depth.

Barger, R. L.↗
NASA NTRS1977

Slotted test section dynamics of the NTF

A dynamic model of the National Transonic Facility (NTF) section and plenum chamber was developed. The model serves as a means for determining the dynamic performance of the NTF and the subsequent design of optimal control configurations, and is of one dimensional flow type as to facilitate computer simulation. The model allows for the simulation of the effects of test section blockage, boundary layer losses, slot flow required for supersonic flow generation, reentry flap position, and auxiliary plenum suction.

Gumas, G.↗
DOE OSTI2023

Performance Characteristics of a Rotating Detonation Combustor Exiting Into a Pressurized Plenum to Simulate Gas Turbine Inlet

Abstract The present study aims to experimentally characterize the performance of a rotating detonation combustion (RDC) system integrated with a pressurized downstream plenum to simulate the high-pressure inlet conditions of power-generating gas turbines. A thorough understanding of the operational behavior including wave mode behavior, static pressure profile along the combustor length, and dynamic features of pressure fluctuations is crucial for successful integration of RDC with the turbine. In this study, two RDC configurations are investigated, RDC with a constant area annulus and RDC with a converging nozzle. In both cases, the RDC flow exited into a plenum chamber kept at pressures varying from 155 kPa to 330 kPa. RDC was operated on methane and oxygen-enriched air to represent reactants used in land-based power generation. Experiments were conducted for the two RDCs configurations operated at three reactant mass flow rates (0.23, 0.32, and 0.46 kg/s). The RDC performance is characterized by time-averaged static pressure measurements, and wave velocity determined by ionization probes. In addition, dynamic pressure measurements were recorded both inside and near the exit of RDC channel to investigate wave interactions between RDC and downstream plenum. Results show that the RDC with the converging nozzle achieved superior performance while minimizing detrimental interactions with the reflected shock and/or acoustic waves from the downstream plenum.

Engineering↗
NASA NTRS1968

Device damps fluid pressure oscillations in vent valve

Device, containing a tuned series arrangement of two plenum chambers and two orifices, damps high pressure fluid oscillations in a vent valve. Used in conjunction with vent valves, it relieves gas pressure that develops in liquid hydrogen and liquid oxygen tanks used on a space vehicle.

Nein, H. J.↗
NASA NTRS1969

Liquid laser cavities

Liquid laser cavities have plenum chambers at the ends of the capillary cell which are terminated in transparent optical flats. By use of these cavities, several new europium chelates and a terbium chelate can provide laser action in solution at room temperature.

Bjorklund, S.↗
NASA NTRS1971

Combustion products generating and metering device

An apparatus for generating combustion products at a predetermined fixed rate, mixing the combustion products with air to achieve a given concentration, and distributing the resultant mixture to an area or device to be tested is described. The apparatus is comprised of blowers, a holder for the combustion product generating materials (which burn at a predictable and controlled rate), a mixing plenum chamber, and a means for distributing the air combustion product mixture.

Wiberg, R. E.↗
NASA NTRS1973

Molecular flux measurements in the back flow region of a nozzle plume

A series of tests were conducted to measure the mass flux in the far field of a nozzle plume in a high vacuum with emphasis on the back flow region. The measurements presented provided fairly accurate data for off-axis angles as large as 140 deg (i.e., in the back flow region). This region, since it is well behind the exit plane, is of paticular interest to those concerned with instrument contamination. Usually sensitive spacecraft surfaces are located in the region affected by the back flow. Parameters such as expansion ratio, throat diameter, nozzle lip shape, and plenum (chamber) pressure were varied, carbon dioxide and nitrogen gases were flowed and mass flux measurements were taken using quartz crystal microbalances in as many as nine different locations relative to the tests nozzle. Several conclusions with respect to the effect of nozzle and gas parameters on the amount of back flow mass flux are offered, and it was demonstrated that gaseous mass fluxes, which are not predictable by present theories, are encountered in the region behind the nozzle exit plane. This knowledge is significant if materials incompatible with the gaseous exhaust products are used in this region.

Chirivella, J. E.↗
NASA NTRS1973

Mass flux measurements and correlations in the back flow region of a nozzle plume.

A series of tests were conducted to measure the mass flux in the far field of a nozzle plume in a high vacuum with emphasis on the back flow region. The measurements presented in this report provide fairly accurate data for off-axis angles as large as 140 deg (i.e., the back flow region). The tests utilized five different nozzles. Parameters such as expansion ratio, throat diameter, nozzle lip shape, and plenum (chamber) pressure were varied. Carbon dioxide and nitrogen gases were passed through, and mass flux measurements were taken using quartz crystal microbalances in as many as nine different locations relative to the test nozzle. The tests resulted in a large matrix of data that were correlated and compared to existing prediction theories. Several conclusions with respect to the effect of nozzle and gas parameters on the amount of back flow mass flux are offered, and it was demonstrated that gaseous mass fluxes, which are not predictable by present theories, are encountered in the region behind the nozzle exit plane.

Chirivella, J. E.↗
NASA NTRS1974

Wind tunnel flow generation section

A flow generation section for a wind tunnel test facility is described which provides a uniform flow for the wind tunnel test section over a range of different flow velocities. The throat of the flow generation section includes a pair of opposed boundary walls which are porous to the flowing medium in order to provide an increase of velocity by expansion. A plenum chamber is associated with the exterior side of each of such porous walls to separate the same from ambient pressure. A suction manifold is connected by suction lines with each one of the chambers. Valves are positioned in each of the lines to enable the suction manifold to be independently varied.

Sorensen, N. E.↗
NASA NTRS1975

Wind tunnel investigation of a Centaur standard shroud compartment vent from Mach number of 0.70 to 1.96

An experimental investigation was conducted in the Lewis Research Center 8- by 6-foot supersonic wind tunnel to determine the vent discharge coefficient for the Centaur standard shroud/liquid hydrogen tank compartment vent. The test was conducted from Mach 0.70 to 1.96 with the vent mounted in a flat plate. Full scale simulated flight hardware, such as the vent, corrugations, aft field joint ring and ice bag clip was used. Air was discharged from a plenum chamber, located on the tunnel sidewall behind the plate, through five 6.35 cm diameter vent orifices into the free stream. Boundary layer thickeners, analytically predicted displacement thickness for the vehicle nominal flight trajectory could be simulated over the Mach number range. The highest vent discharge coefficient for any given Mach number and vent pressure ratio generally occurred at the maximum displacement thickness.

Johns, A. L.↗