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NASA NTRS1974

Tables and charts of equilibrium normal shock and shock-tube solutions for helium-hydrogen mixtures with velocities to 70 km/sec

Equilibrium thermodynamic and flow properties are presented in tabulated and graphical form for moving, standing, and reflected normal shock waves into helium-hydrogen mixtures representative of proposed outer planet atmospheres. The volumetric compositions of these mixtures are 0.35He-0.65H2, 0.20He-0.80H2, and 0.05He-0.95H2. Properties include pressure, temperature, density, enthalpy, speed of sound, entropy, molecular-weight ratio, isentropic exponent, velocity, and species mole fractions. Incident (moving) shock velocities are varied from 4 to 70 km/sec for a range of initial pressure of 5 N/sq m to 100 kN/sq m. The present results are applicable to shock-tube flows and to free-flight conditions for a blunt body at high velocities. A working chart illustrating idealized shock-tube performance with a 0.20He-0.80H2 test gas and heated helium driver gas is also presented.

Miller, C. G., III↗

NASA NTRS1995

Scaling in a scramjet combustor

Studies of scramjet combustor scaling have been extended to include testing of a 'small' combustor, thereby allowing comparison with results obtained in a 'large' combustor of nearly identical configuration. The scale of the two combustors varied by a factor of 5. It was anticipated that the combustion process would scale as the product of pressure with length if the temperature, the Mach number, the velocity, the free stream composition and the equivalence ratio were the same. To achieve this while allowing for the necessary difference in pressure the two models were tested with different shock tunnel nozzles. The large was tested in a M = 8 nozzle, with an intake consisting of a pair of opposing wedges preceding the model, which produced shocks to reduce the Mach number to approximately 3.4 before the flow entered the model inlet. It was then expanded to raise the Mach number again. The small model was tested in a M = 4 nozzle, with the nozzle flow passing directly to the inlet of the model. The wedge intake for the large model was adjusted to yield similar combustion chamber Mach number for both models. Then, by matching the stagnation enthalpy, it was possible also to match the temperature and the velocity. The free stream composition was matched by operating at the same nozzle supply pressure. Wall pressure distributions were measured along each of the ducts. The pressure is normalized with respect to the pressure measured at the point opposite to the injector exit, and the distance downstream of the exit is normalized with respect to the height of the duct. It can be seen that, when normalized in this manner, both ducts produce quantitatively similar pressure distributions, confirming the initial hypothesis that the combustion process would scale as the product of pressure with length.

Pulsonetti, M. V.↗

NASA NTRS1989

Thermophysical properties of Helium-4 from 0.8 to 1500 K with pressures to 2000 MPa

Tabular summary data of the thermophysical properties of fluid helium are given for temperatures from 0.8 to 1500 K, with pressures to 2000 MPa between 75 and 300 K, or to 100 MPa outside of this temperature band. Properties include density, specific heats, enthalpy, entropy, internal energy, sound velocity, expansivity, compressibility, thermal conductivity, and viscosity. The data are calculated from a computer program which is available from the National Institute of Standards and Technology. The computer program is based on carefully fitted state equations for both normal and superfluid helium.

Arp, Vincent D.↗

NASA NTRS1993

Summary of the GASP code application and evaluation effort for scramjet combustor flowfields

Numerical simulations of 3D turbulent mixing and reacting flows have been systematically evaluated by comparison with experimental data and other numerical solutions to provide confidence in the General Aerodynamic Simulation Program (GASP) code version 1.3. Specifically, the GASP flow solver has been used to model high speed flow through scram jet combustors. Unit injector problems evaluated include (1) UVA staged sonic normal injection of air into Mach 2 air stream, (2) the VPI Mach 1.7 low angled (15 deg) flush wall helium injection into an unconfined Mach 6 'cold air' stream, and (3) the HYPULSE angled (30 deg) flush wall hydrogen injection into a Mach 6, high enthalpy (flight Mach 17 simulation) nitrogen and air test gases. Results of the numerical simulation for the first two test cases have been compared with detailed in-stream measurements and with SPARK CFD solutions. The results for the last two cases have been compared with measured wall pressure and heat flux data and with SPARK CFD solutions. The present GASP solutions compare favorably with both the experimental data and the SPARK solutions.

Srinivasan, Shivakumar↗

NASA NTRS1964

Experimental Data on Stagnation-Point Gas Injection Cooling on a Hemisphere-Cone in a Hypersonic Arc Tunnel

Exploratory tests were made to determine the reduction of heat transfer resulting from ejection of gases at the stagnation point on a hemisphere-cone at a nominal Mach number of about 9 and at stagnation enthalpies up to 1600 BTU per pound. Helium, nitrogen, and argon gases were used as coolants. The ratio of mass flow of coolant to the mass flow of air swept out by the model projected area was less than 0.20 in all cases. The experimental heat-transfer data were normalized by the calculated stagnation-point values based on test conditions in the tunnel and are presented as a function of the surface distance from the stagnation point. The variation of shock standoff distances with mass flow and volumetric flow of the gaseous ejectants is presented along with photographs of the luminous flow field of the model during tests.

STAGNATION FLOW↗

NASA NTRS2012

Photogrammetric recession measurements of an ablating surface

An instrument and method for measuring the time history of recession of an ablating surface of a test article during testing in a high enthalpy thermal test facility, such as an arcjet. The method advances prior art by providing time-history data over the full ablating surface without targets and without any modifications to the test article. The method is non-intrusive, simple to implement, requires no external light source, and does not interfere with normal operations of the arcjet facility.

Schairer, Edward T.↗

NASA NTRS1946

An Investigation of the Icing and Heated-air De-icing Characteristics of the R-2600-13 Induction System

A laboratory investigation was made on a Holley 1685-HB carburetor mounted on an R-2600-13 supercharger assembly to determine the icing characteristics and the heated-air de-icing requirements of this portion of the B-25D airplane induction system. Icing has been found to be most prevalent at relatively small throttle openings and, consequently, all runs were made at simulated 60-percent normal rated power condition. Icing characteristics were determined during a series of 15-minute runs over a range of inlet-air conditions. For the de-icing investigation severe impact ice was allowed to form in the induction system and the time required for the recovery of 95 percent of the maximum possible air flow at the original throttle setting was then determined for a range of wet-bulb temperatures. Results of these runs showed that ice on the walls of the carburetor adapter and on the rim of the impeller-shroud portion of the supercharger diffuser plate did not affect engine operation at 60-percent normal rated power. Ice that adversely affected the air flow and the fuel-air ratio was formed only on the central web of the carburetor and then only when the inlet air was saturated or contained free moisture in excess of saturation. No serious ice formations were observed at inlet-air temperatures above 66 0 F or with an inlet-air enthalpy greater than 34 Btu per pound. The maximum temperature at. which any trace of icing could be detected was 1110 F with a relative humidity of approximately 28 percent, The air-flow recovery time for emergency de-icing was 0.3 minute for.an enthalpy of 35 Btu per pound or wet-bulb temperature of 68 0 F. Further increase in enthalpy and wet-bulb temperature above these values resulted in very slight improvement in recovery time. The fuel-air ratio restored by a 5-Minute application of heated air was approximately 7 percent less than the initial value for cold-air conditions.

Chapman, Gilbert E.↗

NASA NTRS1975

Normal- and oblique-shock flow parameters in equilibrium air including attached-shock solutions for surfaces at angles of attack, sweep, and dihedral

Normal- and oblique-shock flow parameters for air in thermochemical equilibrium are tabulated as a function of shock angle for altitudes ranging from 15.24 km to 91.44 km in increments of 7.62 km at selected hypersonic speeds. Post-shock parameters tabulated include flow-deflection angle, velocity, Mach number, compressibility factor, isentropic exponent, viscosity, Reynolds number, entropy difference, and static pressure, temperature, density, and enthalpy ratios across the shock. A procedure is presented for obtaining oblique-shock flow properties in equilibrium air on surfaces at various angles of attack, sweep, and dihedral by use of the two-dimensional tabulations. Plots of the flow parameters against flow-deflection angle are presented at altitudes of 30.48, 60.96, and 91.44 km for various stream velocities.

Hunt, J. L.↗

NASA NTRS1974

Calculation of supersonic stream parameters of a real gas from measurable quantities using FORTRAN 4 routines

Sets of routines are presented that calculate such real-gas supersonic stream properties as velocity, density, enthalpy, entropy, and isentropic exponent. Two alternative sets of measurements are assumed: (1) stagnation pressure, stagnation temperature, and the pressure on the surface of a static-pressure wedge; or (2) pressure and temperature in a plenum upstream of a supersonic nozzle and the stagnation pressure at the exit of this nozzle. The routines are applicable to any gas whose properties are known. Explicit routines are presented for determining these properties for air, nitrogen, oxygen, normal hydrogen, para-hydrogen, helium, argon, steam, methane, and natural gas.

Johnson, R. C.↗

NASA NTRS1976

Full-coverage film cooling heat transfer study: Summary of data for normal-hole injection and 30 deg slant-hole injection

Heat transfer to a full coverage film cooled turbulent boundary layer over a flat surface was studied. The surface consisted of a discrete hole test section containing 11 rows of holes spaced 5 diameters apart in a staggered array and an instrumented recovery region. Ten diameter spacing was also studied by plugging appropriate holes. Two test sections were used, one having holes normal to the surface and the other having holes angled 30 deg to the surface in the downstream direction. Stanton number data were obtained both in the full coverage region and in the downstream recovery region for a range of blowing ratios, or mass flux ratios, from 0 to 1.3. Initial conditions at the upstream edge of the blowing region were varied from 500 to 5000 for momentum thickness Reynolds number and from 100 to 1800 for enthalpy thickness Reynolds number. The range of Reynolds numbers based on hole diameter and mainstream velocity was 6000 to 22000. Initial boundary layer thicknesses range from 0.5 to 2.0 hole diameters. Air was used as the working fluid. The data were taken for the secondary injection temperature equal to the wall temperature and also equal to the mainstream temperature. Superposition was then used to obtain Stanton number as a continuous function of the injectant temperature. The heat transfer coefficient was defined on the basis of a mainstream-to-wall temperature difference. This definition permits direct comparison of performance between film cooling and transpiration cooling.

Crawford, M. E.↗

NASA NTRS1972

Anode heat transfer for a flowing argon plasma at elevated electron temperature.

Theoretical investigation of heat transfer from a preionized gaseous plasma flowing over an anode surface at an elevated electron temperature in the presence of an electric field normal to the surface. A laminar boundary layer is considered in which only the velocity profile is locally similar and fluid properties are assumed to change uniformly in the gas flow direction. Results obtained by an approximation method show that for moderate current densities, the velocity and temperature distributions are insensitive to current. In addition, the effect of elevated electron temperature is negligible on convective heat transfer, but is significant for the overall heat transfer due to the enthalpy transport by current. Total heat flux to the anode is obtained by evaluating the Nusselt number and adding terms due to the potential drop in the sheath and the surface work function.

Bose, T. K.↗

NASA NTRS2013

Jet Surface Interaction-Scrubbing Noise

Generation of sound due to scrubbing of a jet flow past a nearby solid surface is investigated within the framework of the generalized acoustic analogy theory. The analysis applies to the boundary layer noise generated at and near a wall, and excludes the scattered noise component that is produced at the leading or the trailing edge. While compressibility effects are relatively unimportant at very low Mach numbers, frictional heat generation and thermal gradient normal to the surface could play important roles in generation and propagation of sound in high speed jets of practical interest. A general expression is given for the spectral density of the far-field sound as governed by the variable density Pridmore- Brown equation. The propagation Green's function should be solved numerically starting with the boundary conditions on the surface and subject to specified mean velocity and temperature profiles between the surface and the observer. The equivalent sources of aerodynamic sound are associated with non-linear momentum flux and enthalpy flux terms that appear in the linearized Navier-Stokes equations. These multi-pole sources should be modeled and evaluated with input from a Reynolds-Averaged Navier-Stokes (RANS) solver with an appropriate turbulence model.

Khavaran, Abbas↗

NASA NTRS1980

Full-coverage film cooling. I - Comparison of heat transfer data for three injection angles

Wind tunnel experiments were carried out at Stanford between 1971 and 1977 to study the heat transfer characteristics of full-coverage film cooled surfaces with three geometries; normal-, 30 deg slant-, and 30 deg x 45 deg compound-angled injection. A flat full-coverage section and downstream recovery section comprised the heat transfer system. The experimental objectives were to determine, for each geometry, the effects on surface heat flux of injection blowing ratio, injection temperature ratio, and upstream initial conditions. Spanwise-averaged Stanton numbers were measured for blowing ratios from 0 to 1.3, and for two values of injection temperature at each blowing ratio. The heat transfer coefficient was defined on the basis of a mainstream-to-wall temperature difference. Initial momentum and enthalpy thickness Reynolds numbers were varied from 500 to about 3000.

Crawford, M. E.↗

NASA NTRS1992

Estimation of excitation energy of diatomic molecules in expanding nonequilibrium flows

The energy contained in the highly excited vibrational and rotational states in a diatomic gas in a thermochemical nonequilibrium state during expansion is estimated. The estimation is made on the assumption that the populations of the vibrational and rotational states, when normalized by their respective equilibrium values, are describable by simple functions containing no more than four arbitrary parameters. A cubic polynomial, a logarithmic-cubic polynomial, and a bimodal step function are used for this purpose. The four parameters are determined by imposing conditions known at the ground state and the dissociation limit and the mass conservation law. The energy in excess of that accounted for by assuming a Boltzmann distribution of these states, defined here as excess excitation energy, is calculated for N2, O2, NO, CO, OH, and H2. A calculation made for a typical nozzle flow shows that the excess energy may reach 6 percent of the total enthalpy of the flow, and that the flow velocity may decrease by as much as 4 percent due to the nonequilibrium excitation phenomenon.

Park, Chul↗

NASA NTRS2002

Fundamental Mixing and Combustion Experiments for Propelled Hypersonic Flight

The first experiment is a study of a coaxial jet discharging into stagnant laboratory air, with center jet of a mixture of 5% oxygen and 95% helium by volume and coflow jet of air. The exit flow pressure of both center-jet and coflow nozzles is 1 atmosphere. The presence of oxygen in the center jet is to allow the use of an oxygen flow-tagging technique (RELIEF4) to obtain non-intrusive velocity measurements. Both jets are nominally Mach 1.8, but, because of the greater speed of sound, the center jet velocity is more than twice that of the coflow. The mixing layer which forms between the center jet and the coflow near the nozzle exit is compressible, with a calculated convective Mach number of approximately 0.7. This geometry has several advantages: The streamwise development of the flow is generally dominated by turbulent stresses (rather than pressure forces), and thus calculations are sensitive to turbulence modeling. It includes features present in supersonic combustors, including a compressible mixing layer near the nozzle exit and a light-gas/air plume downstream. Since it is a free jet, it provides easy access for both optical instrumentation and probes. Since it is axisymmetric, it requires fewer experimental measurements to fully characterize, and calculations can be performed with more modest computer resources. However, weak shock waves formed at the nozzle exit strengthen and turn normal as they approach the axis, complicating the flow. Care is thus taken in the design of the facility to provide as near as possible to 1-D flow at the exit of both center and coflow nozzles, and to minimize the strength of waves generated at the nozzle exit. Results from this experiment are compared to CFD solutions obtained by VULCAN, a previously developed code used in engine analysis. The second experiment is a study of a supersonic combustor consisting of a diverging duct with single downstream-angled wall injector. Thus, the geometry is relatively simple and large regions of subsonic recirculating flow are avoided. The nominal entrance Mach number is 2 and the enthalpy of the test gas (hot air "simulant") is nominally that of Mach 7 flight. It was believed, on the basis of calculations performed that this would produce mixing-limited flow, that is to say, one for which chemical reaction to equilibrium proceeds at a much greater rate than mixing. It later proved that this was not the case. The primary experimental technique employed is coherent anti-Stokes Raman spectroscopy, known by its acronym CARS. The species probed is molecular nitrogen and the quantity measured is temperature. Intrusive probes, such as Pitot, total temperature, hot-wire, etc., are not used due to access difficulty and high heat flux in the combustor, and because they may alter the flow. CARS has several advantages over other optical methods. It is a relatively mature and well-understood technique. Signal levels are relatively high and the signal is in the form of a coherent (laser) beam that can be collected through small windows. Incoherent (non-CARS) interferences are rejected by spatial filtering.

Diskin, G. S.↗

NASA NTRS1995

Numerical Simulations of High Enthalpy Pulse Facilities

Axisymmetric flows within shock tubes and expansion tubes are simulated including the effects of finite rate chemistry and both laminar and turbulent boundary layers. The simulations demonstrate the usefulness of computational fluid dynamics for characterizing the flows in high enthalpy pulse facilities. The modeling and numerical requirements necessary to simulate these flows accurately are also discussed. Although there is a large body of analysis which explains and quantifies the boundary layer growth between the shock and the interface in a shock tube, there is a need for more detailed solutions. Phenomena such as thermochemical nonequilibrium. or turbulent transition behind the shock are excluded in the assumptions of Mirels' analysis. Additionally there is inadequate capability to predict the influence of the boundary layer on the expanded gas behind the interface. Quantifying the gas in this region is particularly important in expansion tubes because it is the location of the test gas. Unsteady simulations of the viscous flow in shock tubes are computationally expensive because they must follow features such as a shock wave over the length of the facility and simultaneously resolve the small length scales within the boundary layer. As a result, efficient numerical algorithms are required. The numerical approach of the present work is to solve the axisymmetric gas dynamic equations using an finite-volume formulation where the inviscid fluxes are computed with a upwind TVD scheme. Multiple species equations are included in the formulation so that finite-rate chemistry can be modeled. The simulations cluster grid points at the shock and interface and translate this clustered grid with these features to minimize numerical errors. The solutions are advanced at a CFL number of less than one based on the inviscid gas dynamics. To avoid limitations on the time step due to the viscous terms, these terms are treated implicitly. This requires a block tri-diagonal matrix inversion along each line of cells normal to the wall. The cost of this inversion is more than offset by the larger allowable time step. The source terms representing the finite-rate chemical kinetics are also treated implicitly. An algebraic turbulence model for compressible flow is used. The flow in a low pressure shock tube is computed and the results are compared with Mirels'analysis. The driven gas is nitrogen at 70 Pa, and the incident shock speed is approximately 2.9 km/sec so that there is little dissociation. The simulations include a laminar boundary layer and are run until the limiting flow regime is achieved. At this limit, the shock and interface travel at the same velocity because the amount of driven gas between these two features remains the same: the mass flow across the shock is equal to the mass of gas being entrained at the interface by the boundary layer. Simulations with several grids are presented to establish the grid independence of the solution, Good agreement is achieved between Mirels' correlations and the computations. This is expected since the flow conditions are chosen to be consistent with the assumptions used in Mirels' analysis. This comparison adds credibility to the numerical approach and highlights some of the differences between the theory and the detailed simulations. In addition, simulations of the HYPULSE expansion tube are presented for two operating conditions and the computations are compared to experimental data. The operating gas for both cases is nitrogen. One test condition is at a total enthalpy of 15.2 MJ/Kg and a relatively low pressure of 2 kPa. This case is characterized by a laminar boundary layer and significant chemical nonequilibrium. in the acceleration gas. The second test condition is at a total enthalpy of 10.2 MJ/Kg and a pressure of 38 kPa and is characterized by a turbulent boundary layer. The simulations compare well with experiment and reveal that the nonuniformity in pressure observed during the test time is related to variations in the boundary layer displacement thickness.

Wilson, Gregory J.↗

NASA NTRS1976

Equilibrium properties of chemically reacting gases

The equilibrium energy, enthalpy, entropy, specific heat at constant volume and constant pressure, and the equation of state of the gas are all derived for chemically reacting gas mixtures in terms of the compressibility, the mol fractions, the thermodynamic properties of the pure gas components, and the change in zero point energy due to reaction. Results are illustrated for a simple diatomic dissociation reaction and nitrogen is used as an example. Next, a gas mixture resulting from combined diatomic dissociation and atomic ionization reactions is treated and, again, nitrogen is used as an example. A short discussion is given of the additional complexities involved when precise solutions for high-temperature air are desired, including effects caused by NO produced in shuffle reactions and by other trace species formed from CO2, H2O and Ar found in normal air.

Source record↗

NASA NTRS1994

Observation of anomalous slow-mode shock and reconnection layer in the dayside magnetospause

Plasma and magnetic field data from the International Sun-Earth Explorer ISEE 2 spacecraft recorded during an outbound crossing of the dayside, northern hemisphere magnetopause in October 29, 1979, provide evidence for a slow shock (SS) in the observed reconnection layer. This layer is found to be bounded on the magnetosheath side by the SS; near the magnetospheric side of the layer, a second current sheet is found that may have been rotational discontinuity (RD). The direction of the accelerated plasma flow, the earthward sense of the normal magnetic filed across the SS and RD, and the relative orientation of the SS and the RD all indicate that the reconection site was located south of the spacecraft. Quantitative tests show that, allowing for experiemental uncertanties, data taken upstream and downstream of the SS are consistent with coplanarity and other Rankine-Hugoniot (RH) conditions. Examiniations of the flow parameters indicates two anomalous properties of the SS: the upstream flow, viewed in the deHoffmann-Teller frame, is superalfvenic and the downstream plasma is firehose unstable. In comparison to the long-wavelength slow-mode phase speed, however, the flow in the upstream region is super slow, while in the downstream region it is subslow, as requeired for a slow-mode shock. Further properties of the shock include a large decrease in total enthalpy across it, indicating the escape of a sizable heat flux from the shock structure, and the occurrence of a polarization reversal of the tangential magnetic field within the shock layer, a feature that is predicted by linear double-polytropic Hall-MHD and results from a large increase in pressure anistropy from the upstream region, where p(sub parallel) approximately equal to p(sub perp), to the downstream region, where p(sub parrallel) greater than p(sub perp). Quantitative tests of the RD-like discontinuity show that it satisfies the necessary RH conditions within experimental uncertainties and indicate that the flow across it is earthward, that is, it is the same as for the SS. Thus the RD is propagating behind the SS in the direction away from the Earth. The reversal in the order of the discontinuities is attributed to the reversal of the slow and intermediate phase speeds produced by the pressure anistropy in the region between the RD and the SS. The linear dispersion relation and polarization properties of dispersive MHD waves in an anisotropic plasma are examined in the appendix.

Walthour, D. W↗