Real-Time Gas Analysis for the Carbothermal Demonstration Project’s Integrated Prototype Testing
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Joel A Olson↗
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Hypersonic non-equilibrium flow and its thermodynamic relations
Hypersonic nonequilibrium flow and its underlying thermodynamic relations, including real gas effects
NONEQUILIBRIUM FLOW↗
Experimental and analytical sonic nozzle discharge coefficients for Reynolds numbers up to 8 x 10 to the 6th power
Sonic discharge coefficients are presented for two different geometry flow nozzles using nitrogen gas at high pressures where real gas corrections are significant. Throat Reynolds number range extended up to 8 million. Experimentally obtained coefficients for a nozzle with a continuous and finite radius of curvature agreed with those obtained analytically to within 0.2 percent. Experimental coefficients for a long radius ASME nozzle agreed to within 1/4 percent to an empirical equation representing the most probable subsonic discharge coefficient.
Szaniszlo, A. J.↗
Enhancements to Program LAURA for computation of three-dimensional hypersonic flow
Changes to Program Laura (Langley Aerothermodynamic Upwind Relaxation Algorithm) are presented which enhance both stability and accuracy of the algorithm. A discussion of iteration/sweeping strategies and their relation to computer architectures is included to best exploit the capabilities of serial, vector, and parallel processor machines. Test cases for Mach 10 perfect gas flow and Mach 32 real gas flow in chemical nonequilibrium over a blunt, raked elliptic cone using the thin-layer Navier-Stokes equations are presented in order to demonstrate the current improved capabilities. Algorithm changes include the use of volume averaging, application of a symmetric total variation diminishing (TVD) scheme, and stronger interaction between the grid/shock alignment routine and the relaxation algorithm. Good comparisons with heat transfer and pitching moment data at three different angles of attack for the Mach 10 tests serve to further validate the present algorithm. Parameters are defined which control the coupling of the specie continuity equations with the solution of the mixture conservation equations. A discussion of the consequences involved in the choice of strong versus weak coupling is presented, and a sample nonequilibrium calculation on a fine grid over a full scale model of the Aeroassist Flight Experiment (AFE) demonstrates current capabilities.
Gnoffo, Peter A.↗
Supersonic minimum length nozzle design for dense gases
Recently, dense gases have been investigated for many engineering applications such as for turbomachinery and wind tunnels. Supersonic nozzle design for these gases is complicated by their nonclassical behavior in the transonic flow regime. In this paper a method of characteristics (MOC) is developed for two-dimensional (planar) and, primarily, axisymmetric flow of a van der Waals gas. Using a straight aortic line assumption, a centered expansion is used to generate an inviscid wall contour of minimum length. The van der Waals results are compared to previous perfect gas results to show the real gas effects on the flow properties and inviscid wall contours.
Aldo, Andrew C.↗
Experimental and analytical sonic nozzle discharge coefficients for Reynolds numbers up to 8 x 10 to the 6th power
Sonic discharge coefficients are obtained for two different geometry flow nozzles using high-pressure nitrogen gas (100 atm) with significant real-gas flow corrections. Throat Reynolds number range extended up to 8 million. Discharge coefficients for both nozzles monotonically increase in value at the high throat Reynolds numbers. The 95-percent confidence band for each nozzle is shown. Analytical discharge coefficients for the continuous and finite radius of curvature nozzle are presented. These analytical results for the laminar and turbulent boundary-layer cases are compared to experimental values for sonic flow. Experimental values are also compared to values calculated from the best empirical curve fit equation for subsonic flow.
Szaniszlo, A. J.↗
Experimental and analytical sonic nozzle discharge coefficients for Reynolds numbers up to 8,000,000
Sonic discharge coefficients are presented for two different geometry flow nozzles using nitrogen gas at high pressures where real-gas corrections are significant. Throat Reynolds number range extended up to 8,000,000. Experimentally obtained coefficients for a nozzle with a continuous and finite radius of curvature agreed with those obtained analytically to within 0.2 per cent. Experimental coefficients for a long-radius ASME nozzle agreed to within 0.25 per cent to an empirical equation representing the most probable subsonic discharge coefficient.
Szaniszlo, A. J.↗
Swirling flow of a dissociated gas
Most physical applications of the swirling flow, defined as a vortex superimposed on an axial flow in the nozzle, involve high temperatures and the possibility of real gas effects. The generalized one-dimensional swirling flow in a converging-diverging nozzle is analyzed for equilibrium and frozen dissociation using the ideal dissociating gas model. Numerical results are provided to illustrate the major effects and to compare with results obtained for a perfect gas with constant ratio of specific heats. It is found that, even in the case of real gases, perfect gas calculations can give a good estimate of the reduction in mass flow due to swirl.
Wolfram, W. R., Jr.↗
Development and Validation of Physics-Based Sub-Models of High-Pressure Supercritical Fuel Injection at Diesel Conditions Insert
The project sought to develop and validate accurate, physics-based, numerical submodels and implement them in computational fluid dynamics (CFD) software codes to enable accurate predictions of diesel sprays at pressures higher than the current systems. The combined experimental-computational effort has four main objectives: (1) acquire spatially and temporally resolved scalar measurements of fuel-oxidizer mixing in the near field of the fuel jet, (2) develop software modules to calculate thermo-physical properties of diesel surrogates and their mixtures with oxidizer of interest in real-gas operating regime, (3) integrate the real-fluid model into a commercial and an open-source CFD code to simulate fuel-air mixing in high-pressure diesel sprays, and (4) assess the robustness, accuracy, and uncertainty of the integrated CFD solvers in computing diesel sprays. Project accomplishments are consistent with the original goals of the project. Although we acquired spatially and temporally resolved scalar measurements, they were limited to refractive index difference and not to thermodynamic properties such as temperature and equivalence ratio. However, we successfully developed the theoretical background needed to acquire such measurements at diesel conditions and demonstrated the capability of the RSD for scalar measurements in a canonical turbulent jet flow replicating features of the diesel jet. Additional work will be needed to implement these procedures at high-pressure, high-temperature engine-like conditions of the CPFR. Mainly, we did not anticipate the need to find the liquid boundary and its statistical features or the state relationships between refractive index and thermodynamic properties, assuming the widely known Dale-Gladstone equation would suffice. Still, efforts devoted to these tasks were essential to fulfil the goals of the project and serves as the foundation on which many of the conclusions and key contributions of this project are based. Further assessment and application of the developed CFD codes is recommended to fully realize the benefits of this research.
33 ADVANCED PROPULSION SYSTEMS↗
Atlas F entry aerothermic study
The feasibility of obtaining heat transfer data on an expended Atlas F booster launch vehicle was investigated in the altitude range of 300,000 to 200,000 feet during entry conditions, with a velocity in the range of 20,000 to 25,000 feet per second, and through a range of vehicle attitudes of plus or minus 90 degrees. These data are desired for correlation with turbulent heat transfer and boundary layer transition data obtained from wind tunnel test facilities. The data would also be valuable in assessing rarified gas and surface catalicity effects in a real gas environment.
Source record↗
Velocity measurements by laser resonance fluorescence
The photonburst correlation method was used to detect single atoms in a buffer gas. Real time flow velocity measurements with laser induced resonance fluorescence from single or multiple atoms was demonstrated and this method was investigated as a tool for wind tunnel flow measurement. Investigations show that single atoms and their real time diffusional motion on a buffer gas can be measured by resonance fluorescence. By averaging over many atoms, flow velocities up to 88 m/s were measured in a time of 0.5 sec. It is expected that higher flow speeds can be measured and that the measurement time can be reduced by a factor of 10 or more by careful experimental design. The method is clearly not ready for incorporation in high speed wind tunnels because it is not yet known whether the stray light level will be higher or lower, and it is not known what detection efficiency can be obtained in a wind tunnel situation.
She, C. Y.↗
A decade of computer simulations for space shuttle aerodynamics
Ten years ago computer simulations of the flow field around the Space Shuttle Orbiter were limited to inviscid calculations for the windward side of the forebody and viscous calculations for selected 2-D problems. Advances in computer hardware and numerical methods during the past ten years made it possible to calculate viscous flow over the complete orbiter configuration at angle of attack. The equations solved are the Reynolds-averaged, Navier-Stokes equation, simplified by either the thin-layer or parabolized approximation. An algebraic eddy viscosity model is used for turbulent flow. The free stream is assumed to be a perfect gas for wind tunnel conditions and a real gas in thermodynamic equilibrium for flight conditions. Four examples of recent computer simulations are presented. Flow field results include oil flow patterns on the surface and Mach number contours, isobars, and cross-flow velocity vectors in the shock layer.
Inouye, Mamoru↗
A decade of computer simulations for Space Shuttle aerodynamics
Ten years ago computer simulations of the flow field around the Space Shuttle Orbiter were limited to inviscid calculations for the windward side of the forebody and viscous calculations for selected 2-D problems. Advances in computer hardware and numerical methods during the past ten years made it possible to calculate viscous flow over the complete orbiter configuration at angle of attack. The equations solved are the Reynolds-averaged, Navier-Stokes equation, simplified by either the thin-layer or parabolized approximation. An algebraic eddy viscosity model is used for turbulent flow. The free stream is assumed to be a perfect gas for wind tunnel conditions and a real gas in thermodynamic equilibrium for flight conditions. Four examples of recent computer simulations are presented. Flow field results include oil flow patterns on the surface and Mach number contours, isobars, and cross-flow velocity vectors in the shock layer.
Inouye, Mamoru↗
An experimental parametric study of geometric, Reynolds number, and ratio of specific heats effects in three-dimensional sidewall compression scramjet inlets at Mach 6
The effects of the decrease in the ratio of specific heats have been investigated in generic 3D sidewall compression scramjet inlets with leading-edge sweep angles of 30 and 70 degrees. The effects of a decrease in ratio of specific heats were seen by comparing data from two facilities in two test gases: in the Langley Mach 6 CF4 Tunnel in tetrafluoromethane and in the Langley 15-Inch Mach 6 Air Tunnel in perfect gas air. In addition to the simulated real gas effects, the parametric effects of cowl position, contraction ratio, leading-edge sweep, and Reynolds number were investigated in the 15-Inch Mach 6 Air Tunnel. The models were instrumented with a total of 45 static pressure orifices distributed on the sidewalls and baseplate. Surface streamline patterns were examined via oil flow, and schlieren videos were made of the external flow field. The results of these tests have significant implications to ground based testing of inlets in facilities which do not operate at flight enthalpies.
Holland, Scott D.↗
Role of Inlet Boundary Conditions on Fuel-Air Mixing at Supercritical Conditions
Advanced engine design and alternative fuels present the possibility of fuel injection at purely supercritical conditions in diesel engines and gas turbines. The complex interactions that govern this phenomenon still need significant research, particularly the boundary conditions for fuel injection are critical for accurate simulation. However, the flow inside the injector itself is often omitted to reduce the computational efforts, and thus, velocity, mass flux, or total pressure is specified at the injector exit (or domain inlet), often with simplified velocity profiles and turbulence levels. This simplified inlet boundary treatment has minimal effects on results for conventional fuel injection conditions, however, the validity of this approach at supercritical conditions has not been assessed. Comprehensive real-gas and binary fluid mixing models have been implemented for computational fluid dynamics (CFD) analysis of fuel-air mixing at supercritical conditions. The model is verified using prior CFD results from the literature. The model is used to investigate the effects of the shape of axial velocity and mass fraction profiles at the inlet boundary with the goal to improve the comparison of predictions to experimental data. Results show that the boundary conditions have a significant effect on the predictions, and none of the cases match precisely with experimental data. Furthermore, the study reveals that the physical location of the inlet boundary might be difficult to infer correctly from the experiments and highlights the need for high-quality, repeatable measurements at supercritical conditions to support the development of relevant high-fidelity models for fuel-air mixing.
02 PETROLEUM↗
X-38 Experimental Aerothermodynamics
The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic environment is being defined through a synergistic combination of ground based testing and computational fluid dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp 6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were made and generally compared to within 50 deg F of flight prediction.
Horvath, Thomas J.↗
Simulation of ideal-gas flow by nitrogen and other selected gases at cryogenic temperatures
The real gas behavior of nitrogen, the gas normally used in transonic cryogenic tunnels, is reported for the following flow processes: isentropic expansion, normal shocks, boundary layers, and interactions between shock waves and boundary layers. The only difference in predicted pressure ratio between nitrogen and an ideal gas which may limit the minimum operating temperature of transonic cryogenic wind tunnels occur at total pressures approaching 9 atm and total temperatures 10 K below the corresponding saturation temperature. These pressure differences approach 1 percent for both isentropic expansions and normal shocks. Alternative cryogenic test gases were also analyzed. Differences between air and an ideal diatomic gas are similar in magnitude to those for nitrogen and should present no difficulty. However, differences for helium and hydrogen are over an order of magnitude greater than those for nitrogen or air. It is concluded that helium and cryogenic hydrogen would not approximate the compressible flow of an ideal diatomic gas.
Hall, R. M.↗
Experimental perfect-gas study of expansion-tube flow characteristics
Results of an experimental investigation of expansion tube flow characteristics performed with helium test gas and acceleration gas are presented. The use of helium, eliminates complex real gas chemistry in the comparison of measured and predicted flow quantities. The driver gas was unheated helium at a nominal pressure of 33 MN sq m. The quiescent test gas pressure and quiescent acceleration gas pressure were varied from 0.7 to 50 kN/sq m and from 2.5 to 53 N/sq m, respectively. The effects of tube-wall boundary layer growth and finite secondary diaphragm opening time were examined through the variation of the quiescent gas pressures and secondary diaphragm thickness. Optimum operating conditions for helium test gas were also defined.
Shinn, J. L.↗