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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

High pressure Mach 10 to 20 electrothermal hypersonic wind tunnel

A new technique is described for generating high enthalpy, long pulse flows for testing scramjet combustors at equivalent flight Mach numbers of 10 to 20. The approach derives from high pressure discharge technology developed for electrothermal guns and space thrusters. Cryogenic liquid air is heated with a high power electric discharge (500 MW for a 15 kg/sec test flow rate) in a capillary tube, generating a quasi-steady pressure of up to 6800 atm at 20,000 K. The heated air is mixed with additional cryogenic air and is expanded in a conventional supersonic nozzle to 0.5 to 1 atm at 3000 to 6000 m/sec. Combustion and nozzle expansion test time of approximately 300 m/sec are contemplated. Facility nozzle chemical kinetics calculations show that above Mach 16 increasing concentrations of argon and oxygen replenishment are required to obtain suitable test conditions. Initial experiments on a subscale test rig are described for 0.5 kg/sec and 300 atm pressure.

Burton, R. L.↗
NASA NTRS1991

Progress of simulations for reacting shear layers

An attempt was made to develop a high speed, chemically reactive shear layer test rig. The purpose of the experiment was to study the mixing of oxidizer and fuel streams in reacting shear layers for various density, velocity, and Mach number. The primary goal was to understand the effects of the compressibility upon mixing and combustion in a fundamental way. Therefore, a two-dimensional shear layer is highly desirable for its simplicity to quantify the compressibility effects. The RPLUS 2D code is used to calculate the flow fields of different sections of the test rig. The emphasis was on the supersonic nozzle design, the vitiation process for the hot air stream and the overall thermodynamic conditions of the test matrix. The k-epsilon turbulence model with wall function was successfully implemented in the RPLUS code. The k and epsilon equations are solved simultaneously and the LU scheme is used to make it compatible with the flow solver.

Yu, Sheng-Tao↗
NASA NTRS1991

Computational investigation of circular-to-rectangular transition ducts

A study demonstrating that flows inside transition ducts with unusual geometry can be analyzed by employing proper selections of a Navier-Stokes code, grid topology, and turbulence modeling is presented. Based on comparison between existing experimental data and the computed results for the same configurations, reasonable agreements were obtained for wall static pressure in the transition duct. Static pressure comparisons in the supersonic nozzle section were excellent, as well as agreement between computed and measured mass flow and thrust performance.

Pao, S. P.↗
NASA NTRS1992

Thermochemical nonequilibrium and radiative interactions in supersonic hydrogen-air combustion

The two-dimensional, elliptic Navier-Stokes equations are used to investigate supersonic flows with nonequilibrium chemistry and thermodynamics, coupled with radiation, for hydrogen-air systems. The chemistry source term in the species equation is treated implicitly to alleviate the stiffness associated with fast reactions. The explicit, unsplit MacCormack finite-difference scheme is used to advance the governing equations in time, until convergence is achieved. The specific problem considered is the premixed, expanding flow in a supersonic nozzle. The reacting flow consists of seven species, one of which is the inert N2 molecule. The thermal state of the gas is modeled with one translational-rotational temperature and five vibrational temperatures. The harmonic oscillator model is used in the formulation for vibrational relaxation. The tangent slab approximation is used in the radiative flux formulation. A pseudo-gray model is used to represent the absorption-emission characteristics of the participating species. Results obtained for specific conditions indicate the presence of nonequilibrium in the expansion region. This reduces the radiative interactions and can have a significant influence on the flowfield.

Chandrasekhar, R.↗
NASA NTRS1992

A wind tunnel application of large-field focusing schlieren

A large-field focusing schlieren apparatus was installed in the NASA Lewis Research Center 9 by 15 foot wind tunnel in an attempt to determine the density gradient flow field of a free jet issuing from a supersonic nozzle configuration. The nozzle exit geometry was designed to reduce acoustic emissions from the jet by enhancing plume mixing. Thus, the flow exhibited a complex three-dimensional structure which warranted utilizing the sharp focusing capability of this type of schlieren method. Design considerations concerning tunnel limitations, high-speed photography, and video tape recording are presented in the paper.

Ponton, Michael K.↗
NASA NTRS1992

The effects of temperature on supersonic jet noise emission

This paper examines the generation of sound produced by high temperature supersonic jets. In particular, the question of the importance of supersonic instability waves to noise emission is considered relative to the role of Kelvin-Helmholtz (K-H) instability waves. Here, these waves are taken to be synonymous with the Mach emission process. Jet total temperatures from 313 to 1534 K are investigated using an axisymmetric water cooled supersonic nozzle designed for Mach 2. The aerodynamic and acoustic results of this study indicate that the dominant noise contributors are the K-H waves over the entire temperature range. Good agreement between measured and numerically predicted plume properties are obtained and an elliptic nozzle is used to demonstrate reduction of the K-H waves.

Seiner, John M.↗
NASA NTRS1993

Raman scattering measurements of vibrational and rotational distributions in expanding nitrogen

Vibrational and rotational population distributions in expanding nitrogen in the supersonic nozzle of a reflected shock tunnel were investigated using spontaneous Raman scattering spectroscopy. The experimental data are compared with two multitemperature nonequilibrium calculations, one based on the Landau-Teller model for vibrational relaxation, the other on the solution of the vibrational master equations. Good agreement was found between both the Landau-Teller and the master equation solutions for the vibrational energy mode, but the experimentally inferred rotational temperatures were found to be systematically low, due to collisional narrowing. The dramatic increase in vibrational relaxation rates for a cooling flowfield reported in most previous studies was not reproduced by the present experiments.

Gillespie, Walter D.↗
NASA NTRS1950

Approximate Turbulent Boundary-Layer Development in Plane Compressible Flow Along Thermally Insulated Surfaces with Application to Supersonic-Tunnel Contour Correction

Numerical solutions of the differential equation obtained from the momentum theorem for the development of a turbulent boundary layer along a thermally insulated surface in two-dimensional and in radial shock-free flow are presented in tabular form for a range of Mach numbers from 0.100 to 10. The solution can be used in a step-wise procedure with any given distribution of favorable pressure gradients and for zero pressure gradients. Solutions are also given for use with moderate adverse pressure gradients. The mean velocity in the boundary layer is approximated by a power-law profile. In view of the stepwise integration methods to be used, the exponent designated the profile shape can be varied along the surface between the integral fraction limits 1/5 and 1/11 through interpolation. Agreement obtained between theoretical and experimental boundary-layer development in a supersonic nozzle at a nominal Mach number of 2 indicates the general validity of the approximations used in the analysis - in particular, the method of extrapolating low-speed skin-friction relations to high Mach number flows. The extrapolation method used assumes that the skin-friction coefficient depend primarily on Reynolds number, provided that the density and the kinematic viscosity are evaluated at surface conditions.

BOUNDARY LAYER - WIND TUNNEL WALLS↗
NASA NTRS1957

An Evaluation of Four Experimental Methods for Measuring Mean Properties of a Supersonic Turbulent Boundary Layer

Surveys were made through a turbulent boundary layer on a flat plate by means of a pitot probe, an x-ray densitometer, and hot-wire and cold-wire probes. Results from these surveys were analyzed to determine (a) the reliability of the basic data and hence the methods by which they were obtained, and (b) how well the actual distributions of properties in the boundary layer compare with those commonly assumed in semiempirical and theoretical analyses. All surveys were made at the same longitudinal station on the flat plate. The tests were conducted in a an 8- by 8-inch supersonic nozzle. The free-stream Mach number was 3.03 and the Reynolds number was approximately 210,000 based on boundary-layer thickness.

Nothwang, George J↗
NASA NTRS1994

HO2 detection by near infrared absorption using tunable diode lasers

The strongest absorption lines are in the v sub 2 vibrational band around 1400/cm; several near-coincident line pairs in this region provide max linestrengths of 1.2 x 10 exp -20 sq. cm/(molecule*cm). High Frequency wavelength modulation coupled with sweep integration were used to minimize noise and maintain spectral discrimination. Reduced pressure sampling was performed using supersonic nozzle and an aerodynamic flow cell to minimize residence time and wall losses of radicals. The astigmatic mirror design provides pathlengths up to 200 m in volume of 3 liters with minimal optical interference fringes; longer pathlengths are possible. The detection limit depends on pathlength and minimum fractional absorptions.

Nelson, David↗
NASA NTRS1996

Emission Spectral Measurements in the Plenum of an Arc Jet Wind Tunnel

Arc jet wind tunnel facilities are used to evaluate thermal protection system materials for re-entry vehicles. The high speed, high temperature flowfield generated by the arc jet can simulate the extreme aerodynamic heating environment experienced during re-entry so that the survivability of heat shield materials and performance of various designs options can be tested. Although the re-entry heating environment can be approximated in the arc jet facility, the flowfield only partially simulates the actual re-entry flight conditions. Reynolds numbers are not matched so that surface shear stress distributions and mass transfer rates due to ablation or other mechanisms are not modeled correctly. Unlike flight conditions the arc freestream air is in non-equilibrium because of the rapid expansion that occurs in the supersonic nozzle. To properly study the actual re-entry flow environment, computational fluid dynamics, computational chemistry and radiation models must be used. Arc jet tunnel tests serve to validate these models. To perform accurate simulations inlet and boundary-conditions are needed, which come from measurements of the flowfield. The present study is concerned with measurements in the plenum region of an arc heater. In the past, conditions in the arc heater flowfield have been predicted using simulations since conventional measurement techniques could not be used in the harsh extremely high temperature environment. The present study is part of a recent push to utilize optical techniques to help better characterize the arc jet flowfields. Emission measurements have been made in the shock layer and the constrictor section of the arc heater to determine temperatures and species number densities. LIF measurements have been made in the free stream to determine temperature and velocity.

Donohue, Jim↗
NASA NTRS1997

Measurement of Momentum Transfer Coefficients for H2, N2, CO, and CO2 Incident Upon Spacecraft Surfaces

Measurements of momentum transfer coefficients were made for gas-surface interactions between the Space Shuttle reaction control jet plume gases and the solar panel array materials to be used on the International Space Station. Actual conditions were simulated using a supersonic nozzle source to produce beams of the gases with approximately the same average velocities as the gases have in the Shuttle plumes. Samples of the actual solar panel materials were mounted on a torsion balance that was used to measure the force exerted on the surfaces by the molecular beams. Measurements were made with H2, N2, CO, and CO2 incident upon the solar array material, Kapton, SiO2-coated Kapton, and Z93-coated Al. The measurements showed that molecules scatter from the surfaces more specularly as the angle of incidence increases and that the scattering behavior has a strong dependence upon both the incident gas and velocity. These results show that for some technical surfaces the simple assumption of diffuse scattering with complete thermal accommodation is entirely inadequate. It is clear that additional measurements are required to produce models that more accurately describe the gas-surface interactions encountered in rarefied flow regimes.

Cook, Steven R.↗
NASA NTRS1997

Measurements of momentum transfer coefficients for H2, N2, CO and CO2 incident upon spacecraft surfaces

Measurements of momentum transfer coefficients were made for gas-surface interactions between the Space Shuttle reaction control jet plume gases and the solar panel array materials to be used on the International Space Station. Actual conditions were simulated using a supersonic nozzle source to produce beams of the gases with approximately the same average velocities as the gases have in the Shuttle plumes. Samples of the actual solar panel materials were mounted on a torsion balance that was used to measure the force exerted on the surfaces by the molecular beams. Measurements were made with H2, N2, CO, and CO2 incident upon the solar array material, Kapton, SiO2-coated Kapton, and Z93-coated Al. The measurements showed that molecules scatter from the surfaces more specularly as the angle of incidence increases and that scattering behavior has a strong dependence upon both the incident gas and velocity. These results show that for some technical surfaces the simple assumption of diffuse scattering with complete thermal accommodation is entirely inadequate. It is clear that additional measurements are required to produce models that more accurately describe the gas-surface interactions encountered in rarefied flow regimes.

Cook, S. R.↗
NASA NTRS1999

Comparisons of Shock Noise Predictions with Flight Data

A flight test was performed at NASA Dryden Research Center in November 1991 utilizing both F18 and Fl6 aircraft. These flights were designed to provide (1) acoustic data that could be extrapolated to that of an HSCT at various points of its climb-to-cruise operation and (2) a data base for noise from a supersonic jet exhausting from an aircraft moving at high subsonic speeds. This presentation utilizes data obtained from these flyovers to evaluate predictions of broadband shock noise from supersonic jets in flight. The Fl8 is particularly suitable for flyovers of shock noise since it can be flown with one engine at flight idle. The second engine can then be operated at a pressure high enough to produce a supersonic nozzle exhaust and still maintain an unaccelerated, level flyover.

Norum, T. D.↗
NASA NTRS2001

MEMS-Based Satellite Micropropulsion Via Catalyzed Hydrogen Peroxide Decomposition

Micro-electromechanical systems (MEMS) techniques offer great potential in satisfying the mission requirements for the next generation of "micro-scale" satellites being designed by NASA and Department of Defense agencies. More commonly referred to as "nanosats", these miniature satellites feature masses in the range of 10-100 kg and therefore have unique propulsion requirements. The propulsion systems must be capable of providing extremely low levels of thrust and impulse while also satisfying stringent demands on size, mass, power consumption and cost. We begin with an overview of micropropulsion requirements and some current MEMS-based strategies being developed to meet these needs. The remainder of the article focuses the progress being made at NASA Goddard Space Flight Center towards the development of a prototype monopropellant MEMS thruster which uses the catalyzed chemical decomposition of high concentration hydrogen peroxide as a propulsion mechanism. The products of decomposition are delivered to a micro-scale converging/diverging supersonic nozzle which produces the thrust vector; the targeted thrust level approximately 500 N with a specific impulse of 140-180 seconds. Macro-scale hydrogen peroxide thrusters have been used for satellite propulsion for decades; however, the implementation of traditional thruster designs on a MEMS scale has uncovered new challenges in fabrication, materials compatibility, and combustion and hydrodynamic modeling. A summary of the achievements of the project to date is given, as is a discussion of remaining challenges and future prospects.

Hitt, Darren L.↗
NASA NTRS1998

Application of a Planar Doppler Velocimetry System to a High Reynolds Number Compressible Jet

A Planar Doppler Velocimetry (PDV) system has been constructed and used to investigate the instantaneous turbulent velocity structure of a round high-speed compressible air jet with a low-speed co-flow. The exit condition was Mach=0.85 at ambient pressure, yielding a Reynolds number of about 650,000 on diameter. The PDV system was installed at NASA Langley Research Center in the Small Anechoic Jet Facility (SAJF), a chamber in which both the acoustic and aerodynamic properties of jets can be studied. For this test, the goal was to gather data which can be used to relate the turbulence structure of the jet to the levels and character of the acoustic noise produced by the jet. The current PDV system can acquire single-velocity-component, single-shot, planar images (15ns exposures) at 30 Hz. For this paper, the primary data set consists of 240 frames of velocity data acquired with both the jet and the low-speed co-flow seeded with light-scattering articles. Thus, velocities could be measured everywhere in the jet shear layer, both in the jet fluid and in the entrained co-flow. Some data were also taken with only the jet flow seeded. These provided mixing concentration images along with the reduced velocity fields. Other images were taken with only the co-flow seeded. These produced unique quantitative images of high speed entrainment. Optical "laser speckle" noise is the largest source of random noise in pulsed PDV systems. Components for the PDV imaging system were specifically selected to minimize speckle noise. To reduce systematic velocity errors due to laser drift, a frequency monitoring reference leg with a temperature-tuned reference iodine cell, was employed. In the course of this study, a novel flow seeder was developed. It enabled continuously variable seeding of the flow with particles of Sheared Pyrogenic Amorphous Hydrophobic Silica (SPAHS). The seeder comprised a dry fluidized bed hopper and a supersonic nozzle "pickup." Shearing action in the pickup dispersed the seed material in an exceptionally fine cloud (approximately 0.3 micron). These particles followed the flow well, did not clump or cake on screens or model surfaces, and were not susceptible to evaporation. Because of the refractory nature of the particles, SPAHS seeding should also be applicable to anticipated future testing at high temperatures.

Smith, Michael W.↗
NASA NTRS2004

Pulsed Ejector Thrust Amplification Tested and Modeled

There is currently much interest in pulsed detonation engines for aeronautical propulsion. This, in turn, has sparked renewed interest in pulsed ejectors to increase the thrust of such engines, since previous, though limited, research had indicated that pulsed ejectors could double the thrust in a short device. An experiment has been run at the NASA Glenn Research Center, using a shrouded Hartmann-Sprenger tube as a source of pulsed flow, to measure the thrust augmentation of a statistically designed set of ejectors. A Hartmann- Sprenger tube directs the flow from a supersonic nozzle (Mach 2 in the present experiment) into a closed tube. Under appropriate conditions, an oscillation is set up in which the jet flow alternately fills the tube and then spills around flow emerging from the tube. The tube length determines the frequency of oscillation. By shrouding the tube, the flow was directed out of the shroud as an axial stream. The set of ejectors comprised three different ejector lengths, three ejector diameters, and three nose radii. The thrust of the jet alone, and then of the jet plus ejector, was measured using a thrust plate. The arrangement is shown in this photograph. Thrust augmentation is defined as the thrust of the jet with an ejector divided by the thrust of the jet alone. The experiments exhibited an optimum ejector diameter and length for maximizing the thrust augmentation, but little dependence on nose radius. Different frequencies were produced by changing the length of the Hartmann-Sprenger tube, and the experiment was run at a total of four frequencies. Additional measurements showed that the major feature of the pulsed jet was a starting vortex ring. The size of the vortex ring depended on the frequency, as did the optimum ejector diameter.

Wilson, Jack↗
NASA NTRS2010

Turbulent Output-Based Anisotropic Adaptation

Controlling discretization error is a remaining challenge for computational fluid dynamics simulation. Grid adaptation is applied to reduce estimated discretization error in drag or pressure integral output functions. To enable application to high O(10(exp 7)) Reynolds number turbulent flows, a hybrid approach is utilized that freezes the near-wall boundary layer grids and adapts the grid away from the no slip boundaries. The hybrid approach is not applicable to problems with under resolved initial boundary layer grids, but is a powerful technique for problems with important off-body anisotropic features. Supersonic nozzle plume, turbulent flat plate, and shock-boundary layer interaction examples are presented with comparisons to experimental measurements of pressure and velocity. Adapted grids are produced that resolve off-body features in locations that are not known a priori.

Park, Michael A.↗