Technical Seminar: Exploring Hypersonic Flow
NASA Aeronautics is developing a method for 2D and 3D imaging of hypersonic flows, called Nitric Oxide Planar Laser-Induced Fluorescence (NO-PLIF). NO-PLIF has been used to study basic transition f...
Multigrid for hypersonic inviscid flows
NASA Technical Reports Server (NTRS)
Decker, Naomi H.; Turkel, Eli
1990-01-01
The use of multigrid methods to solve the Euler equations for hypersonic flow is discussed. The steady state equations are considered with a Runge-Kutta smoother based on the time accurate equations together with local time stepping and residual smoothing. The effect of the Runge-Kutta coefficients on the convergence rate was examined considering both damping characteristics and convection properties. The importance of boundary conditions on the convergence rate for hypersonic flow is discussed. Also of importance are the switch between the second and fourth difference viscosity. Solutions are given for flow around the bump in a channel and flow around a biconic section.
Mathematical Models Of Turbulence In Hypersonic Flow
NASA Technical Reports Server (NTRS)
Marvin, J. G.; Coakley, T. J.
1991-01-01
Report discusses mathematical models of turbulence used in numerical simulations of complicated viscous, hypersonic flows. Includes survey of essential features of models and their statuses in applications.
Proximal bodies in hypersonic flow
Deiterding, Ralf; Laurence, Stuart J; Hornung, Hans G
2007-01-01
Hypersonic flows involving two or more bodies travelling in close proximity to one another are encountered in several important situations, both natural and man-made. The present work seeks to investigate one aspect of the resulting flow problem by exploring the forces experienced by a secondary body when it is within the domain of influence of a primary body travelling at hypersonic speeds. An analytical methodology based on the blast wave analogy is developed and used to predict the secondary force coefficients for simple geometries in both two and three dimensions. When the secondary body is entirely inside the primary shocked region, the nature of the lateral force coefficient is found to depend strongly on the relative size of the two bodies. For two spheres, the methodology predicts that the secondary body will experience an exclusively attractive lateral force if the secondary diameter is larger than one-sixth the primary diameter. The analytical results are compared with those from numerical simulations and reasonable agreement is observed if an appropriate normalization for the lateral displacement is used. Results from a series of experiments in the T5 hypervelocity shock tunnel are also presented and compared with perfect-gas numerical simulations, with good agreement. A new force-measurement technique for short-duration hypersonic facilities, enabling the experimental simulation of the proximal bodies problem, is described. This technique provides two independent means of measurement, and the agreement observed between the two gives a further degree of confidence in the results obtained.
Turbulence modeling for hypersonic flows
NASA Technical Reports Server (NTRS)
Marvin, J. G.; Coakley, T. J.
1989-01-01
Turbulence modeling for high speed compressible flows is described and discussed. Starting with the compressible Navier-Stokes equations, methods of statistical averaging are described by means of which the Reynolds-averaged Navier-Stokes equations are developed. Unknown averages in these equations are approximated using various closure concepts. Zero-, one-, and two-equation eddy viscosity models, algebraic stress models and Reynolds stress transport models are discussed. Computations of supersonic and hypersonic flows obtained using several of the models are discussed and compared with experimental results. Specific examples include attached boundary layer flows, shock wave boundary layer interactions and compressible shear layers. From these examples, conclusions regarding the status of modeling and recommendations for future studies are discussed.
Hypersonic flow past open cavities
NASA Technical Reports Server (NTRS)
Morgenstern, Alagacyr, Jr.; Chokani, Ndaona
1993-01-01
The hypersonic flow over a cavity is investigated. The time-dependent compressible Navier-Stokes equations, in terms of mass averaged variables, are numerically solved. An implicit algorithm, with a subiteration procedure to recover time-accuracy, is used to perform the time-accurate computations. The objective of the study is to investigate the effects of Reynolds number and cavity dimensions. The comparison of the computations with available experimental data, in terms of time mean static pressure, heat transfer, and Mach number show good agreement. In the computations large vortex structures, which adversely affect the cavity flow characteristics, are observed at the rear of the cavity. A self-sustained oscillatory motion occurs within the cavity over a range of Reynolds number and cavity dimensions. The frequency spectra of the oscillations show good agreement with a modified semi-empirical relation.
Hypersonic flow past open cavities
NASA Technical Reports Server (NTRS)
Morgenstern, Algacyr, Jr.; Chokani, Ndaona
1994-01-01
The hypersonic flow over a cavity is investigated. The time-dependent compressible Navier-Stokes equations are numerically solved. An implicit algorithm, with a subiteration procedure to recover time accuracy, is used to perform the time-accurate computations. The objective of the study is to investigate the effects of Reynolds number and cavity dimensions. The comparsion of the computations with available experimental data, in terms of time mean static pressure, heat transfer, and Mach number, show good agreement. In the computations large vortex structures, which adversely affect the cavity flow characteristics, are observed at the rear of the cavity. A self-sustained oscillatory motion occurs within the cavity over a range of Reynolds number and cavity dimensions. The frequency spectra of the oscillations show good agreement with a modified semiempirical relation.
Center for Hypersonic Combined Cycle Flow Physics
2015-03-24
AFRL-AFOSR-VA-TR-2015-0292 CENTER FOR HYPERSONIC COMBINED CYCLE FLOW PHYSICS James Mcdaniel UNIVERSITY OF VIRGINIA Final Report 03/24/2015...HYPERSONIC COMBINED CYCLE FLOW PHYSICS 5a. CONTRACT NUMBER 5b. GRANT NUMBER FA9550-09-1-0611 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) McDaniel, James C...DISTRIBUTION/AVAILABILITY STATEMENT Approved for Public Release 13. SUPPLEMENTARY NOTES 14. ABSTRACT Combined cycle flow physics were investigated using a
Conjugate Heat Transfer Study in Hypersonic Flows
NASA Astrophysics Data System (ADS)
Sahoo, Niranjan; Kulkarni, Vinayak; Peetala, Ravi Kumar
2017-05-01
Coupled and decoupled conjugate heat transfer (CHT) studies are carried out to imitate experimental studies for heat transfer measurement in hypersonic flow regime. The finite volume based solvers are used for analyzing the heat interaction between fluid and solid domains. Temperature and surface heat flux signals are predicted by both coupled and decoupled CHT analysis techniques for hypersonic Mach numbers. These two methodologies are also used to study the effect of different wall materials on surface parameters. Effectiveness of these CHT solvers has been verified for the inverse problem of wall heat flux recovery using various techniques reported in the literature. Both coupled and decoupled CHT techniques are seen to be equally useful for prediction of local temperature and heat flux signals prior to the experiments in hypersonic flows.
Advances in Computational Capabilities for Hypersonic Flows
NASA Technical Reports Server (NTRS)
Kumar, Ajay; Gnoffo, Peter A.; Moss, James N.; Drummond, J. Philip
1997-01-01
The paper reviews the growth and advances in computational capabilities for hypersonic applications over the period from the mid-1980's to the present day. The current status of the code development issues such as surface and field grid generation, algorithms, physical and chemical modeling, and validation is provided. A brief description of some of the major codes being used at NASA Langley Research Center for hypersonic continuum and rarefied flows is provided, along with their capabilities and deficiencies. A number of application examples are presented, and future areas of research to enhance accuracy, reliability, efficiency, and robustness of computational codes are discussed.
Vibrational relaxation in hypersonic flow fields
NASA Technical Reports Server (NTRS)
Meador, Willard E.; Miner, Gilda A.; Heinbockel, John H.
1993-01-01
Mathematical formulations of vibrational relaxation are derived from first principles for application to fluid dynamic computations of hypersonic flow fields. Relaxation within and immediately behind shock waves is shown to be substantially faster than that described in current numerical codes. The result should be a significant reduction in nonequilibrium radiation overshoot in shock layers and in radiative heating of hypersonic vehicles; these results are precisely the trends needed to bring theoretical predictions more in line with flight data. Errors in existing formulations are identified and qualitative comparisons are made.
Computational analysis of hypersonic airbreathing aircraft flow fields
NASA Technical Reports Server (NTRS)
Dwoyer, Douglas L.; Kumar, Ajay
1987-01-01
The general problem of calculating the flow fields associated with hypersonic airbreathing aircraft is presented. Unique aspects of hypersonic aircraft aerodynamics are introduced and their demands on computational fluid dynamics are outlined. Example calculations associated with inlet/forebody integration and hypersonic nozzle design are presented to illustrate the nature of the problems considered.
Aerodynamic heating in hypersonic flows
NASA Technical Reports Server (NTRS)
Reddy, C. Subba
1993-01-01
Aerodynamic heating in hypersonic space vehicles is an important factor to be considered in their design. Therefore the designers of such vehicles need reliable heat transfer data in this respect for a successful design. Such data is usually produced by testing the models of hypersonic surfaces in wind tunnels. Most of the hypersonic test facilities at present are conventional blow-down tunnels whose run times are of the order of several seconds. The surface temperatures on such models are obtained using standard techniques such as thin-film resistance gages, thin-skin transient calorimeter gages and coaxial thermocouple or video acquisition systems such as phosphor thermography and infrared thermography. The data are usually reduced assuming that the model behaves like a semi-infinite solid (SIS) with constant properties and that heat transfer is by one-dimensional conduction only. This simplifying assumption may be valid in cases where models are thick, run-times short, and thermal diffusivities small. In many instances, however, when these conditions are not met, the assumption may lead to significant errors in the heat transfer results. The purpose of the present paper is to investigate this aspect. Specifically, the objectives are as follows: (1) to determine the limiting conditions under which a model can be considered a semi-infinite body; (2) to estimate the extent of errors involved in the reduction of the data if the models violate the assumption; and (3) to come up with correlation factors which when multiplied by the results obtained under the SIS assumption will provide the results under the actual conditions.
CFD on hypersonic flow geometries with aeroheating
NASA Astrophysics Data System (ADS)
Sohail, Muhammad Amjad; Chao, Yan; Hui, Zhang Hui; Ullah, Rizwan
2012-11-01
The hypersonic flowfield around a blunted cone and cone-flare exhibits some of the major features of the flows around space vehicles, e.g. a detached bow shock in the stagnation region and the oblique shock wave/boundary layer interaction at the cone-flare junction. The shock wave/boundary layer interaction can produce a region of separated flow. This phenomenon may occur, for example, at the upstream-facing corner formed by a deflected control surface on a hypersonic entry vehicle, where the length of separation has implications for control effectiveness. Computational fluid-dynamics results are presented to show the flowfield around a blunted cone and cone-flare configurations in hypersonic flow with separation. This problem is of particular interest since it features most of the aspects of the hypersonic flow around planetary entry vehicles. The region between the cone and the flare is particularly critical with respect to the evaluation of the surface pressure and heat flux with aeroheating. Indeed, flow separation is induced by the shock wave boundary layer interaction, with subsequent flow reattachment, that can dramatically enhance the surface heat transfer. The exact determination of the extension of the recirculation zone is a particularly delicate task for numerical codes. Laminar flow and turbulent computations have been carried out using a full Navier-Stokes solver, with freestream conditions provided by the experimental data obtained at Mach 6, 8, and 16.34 wind tunnel. The numerical results are compared with the measured pressure and surface heat flux distributions in the wind tunnel and a good agreement is found, especially on the length of the recirculation region and location of shock waves. The critical physics of entropy layer, boundary layers, boundary layers and shock wave interaction and flow behind shock are properly captured and elaborated.. Hypersonic flows are characterized by high Mach number and high total enthalpy. An elevated
Tandem spheres in hypersonic flow
Laurence, Stuart J; Deiterding, Ralf; Hornung, Hans G
2009-01-01
The problem of determining the forces acting on a secondary body when it is travelling at some point within the shocked region created by a hypersonic primary body is of interest in such situations as store or stage separation, re-entry of multiple vehicles, and atmospheric meteoroid fragmentation. The current work is concerned with a special case of this problem, namely that in which both bodies are spheres and are stationary with respect to one another. We first present an approximate analytical model of the problem; subsequently, numerical simulations are described and results are compared with those from the analytical model. Finally, results are presented from a series of experiments in the T5 hypervelocity shock tunnel in which a newly-developed force-measurement technique was employed.
Surface electromagnetic actuator in rarefied hypersonic flow
NASA Astrophysics Data System (ADS)
Surzhikov, S. T.
2017-02-01
Hypersonic flow past the surface with sharp edge is investigated. This surface forms an obtuse angle therefore the shock wave generated by the leading edge interacts with the surface. The effect of influence of the surface direct current discharge and a transverse magnetic field on the gas dynamic characteristics is investigated. To solve this problem a numerical simulation is used. The calculation model includes the set of the Navier - Stokes and energy conservation equations, as well as equations of electrodynamics in the ambipolar approximation, and the Poisson equation. Results of numerical modelling the gas dynamics and electrodynamics of gas discharge with magnetic field show the change in the structure of the shock-wave interaction with surface far from location of the gas discharge. It is shown that the low-current glow discharge can be used as an electromagnetic actuator in hypersonic flows.
Hypersonic Flow Computations on Unstructured Meshes
NASA Technical Reports Server (NTRS)
Bibb, K. L.; Riley, C. J.; Peraire, J.
1997-01-01
A method for computing inviscid hypersonic flow over complex configurations using unstructured meshes is presented. The unstructured grid solver uses an edge{based finite{volume formulation. Fluxes are computed using a flux vector splitting scheme that is capable of representing constant enthalpy solutions. Second{order accuracy in smooth flow regions is obtained by linearly reconstructing the solution, and stability near discontinuities is maintained by locally forcing the scheme to reduce to first-order accuracy. The implementation of the algorithm to parallel computers is described. Computations using the proposed method are presented for a sphere-cone configuration at Mach numbers of 5.25 and 10.6, and a complex hypersonic re-entry vehicle at Mach numbers of 4.5 and 9.8. Results are compared to experimental data and computations made with established structured grid methods. The use of the solver as a screening tool for rapid aerodynamic assessment of proposed vehicles is described.
Algorithm For Hypersonic Flow In Chemical Equilibrium
NASA Technical Reports Server (NTRS)
Palmer, Grant
1989-01-01
Implicit, finite-difference, shock-capturing algorithm calculates inviscid, hypersonic flows in chemical equilibrium. Implicit formulation chosen because overcomes limitation on mathematical stability encountered in explicit formulations. For dynamical portion of problem, Euler equations written in conservation-law form in Cartesian coordinate system for two-dimensional or axisymmetric flow. For chemical portion of problem, equilibrium state of gas at each point in computational grid determined by minimizing local Gibbs free energy, subject to local conservation of molecules, atoms, ions, and total enthalpy. Major advantage: resulting algorithm naturally stable and captures strong shocks without help of artificial-dissipation terms to damp out spurious numerical oscillations.
Control of turbulent mixing in hypersonic flow
NASA Astrophysics Data System (ADS)
Nishioka, Michio
1990-10-01
The conventional engines for present supersonic aircrafts have a drawback in the subsonic flights: they generate strong shock waves, increase total pressure losses, dissociate gases due to increased temperature, and substantially decrease fuel burning efficiencies. When the gases are burned in the supersonic flow, duration times of gases in the combustion chamber become too short. The development of a new technology is required to mix rapidly fuel (hydrogen) and oxygen in the supersonic flow and burn them in time. Flow instability of the initial turbulent flow structure in the hypersonic shear layer is simulated as fuel injection, and the after flow (Mach number of 2.5) is analyzed using the linear stability theory and is studied on the amplified disturbances. The growth of this supersonic disturbance is observed using the Schlieren method. The method for additional mixing of gases is studied to accelerate disturbances.
Multi-Scale Modeling of Hypersonic Gas Flow
NASA Astrophysics Data System (ADS)
Boyd, Iain D.
On March 27, 2004, NASA successfully flew the X-43A hypersonic test flight vehicle at a velocity of 5000 mph to break the aeronautics speed record that had stood for over 35 years. The final flight of the X-43A on November 16, 2004 further increased the speed record to 6,600 mph which is almost ten times the speed of sound. The very high speed attainable by hypersonic airplanes could revolutionize air travel by dramatically reducing inter-continental flight times. For example, a hypersonic flight from New York to Sydney, Australia, a distance of 10,000 miles, would take less than 2 h. Reusable hypersonic vehicles are also being researched to significantly reduce the cost of access to space. Computer modeling of the gas flows around hypersonic vehicles will play a critical part in their development. This article discusses the conditions that can prevail in certain hypersonic gas flows that require a multi-scale modeling approach.
Nonequilibrium effects for hypersonic transitional flows
NASA Technical Reports Server (NTRS)
Moss, James N.; Simmonds, Ann L.; Cuda, Vincent, Jr.
1987-01-01
Presented are the results of numerical simulations of hypersonic flow about blunt cones and hemispherical nose configurations for reentry velocities of 7.5 and 10 km/s. Cone half angles 0, 5, and 10 deg are considered at zero angle of incidence; however, the focus is for the 5 deg cone. The body size and altitude ranges considered (70 to 110 km) are such that the flow is in the transitional regime. Translational, thermodynamic, and chemical nonequilibrium effects are considered in the numerical simulation by utilizing the direct simulation Monte Carlo (DSMC) method of Bird. The DSMC results are compared with those obtained with viscous shock-layer and Navier-Stokes methods. Comparisons between the DSMC and continuum calculations show the altitude range where differences in flowfield structure and surface quantities become significant. The current calculations show that the binary scaling similitude provides a means of correlating the blunt body surface quantities in the hypersonic, transitional regime. Furthermore, for the higher velocity entry conditions, the results highlight some of the concerns in the application of multitemperature continuum formulations, particularly the use of some proposed functional relations for the chemical rate constants under thermodynamic nonequilibrium conditions.
Development of an aerodynamic measurement system for hypersonic rarefied flows
NASA Astrophysics Data System (ADS)
Ozawa, T.; Fujita, K.; Suzuki, T.
2015-01-01
A hypersonic rarefied wind tunnel (HRWT) has lately been developed at Japan Aerospace Exploration Agency in order to improve the prediction of rarefied aerodynamics. Flow characteristics of hypersonic rarefied flows have been investigated experimentally and numerically. By conducting dynamic pressure measurements with pendulous models and pitot pressure measurements, we have probed flow characteristics in the test section. We have also improved understandings of hypersonic rarefied flows by integrating a numerical approach with the HRWT measurement. The development of the integration scheme between HRWT and numerical approach enables us to estimate the hypersonic rarefied flow characteristics as well as the direct measurement of rarefied aerodynamics. Consequently, this wind tunnel is capable of generating 25 mm-core flows with the free stream Mach number greater than 10 and Knudsen number greater than 0.1.
Development of an aerodynamic measurement system for hypersonic rarefied flows.
Ozawa, T; Fujita, K; Suzuki, T
2015-01-01
A hypersonic rarefied wind tunnel (HRWT) has lately been developed at Japan Aerospace Exploration Agency in order to improve the prediction of rarefied aerodynamics. Flow characteristics of hypersonic rarefied flows have been investigated experimentally and numerically. By conducting dynamic pressure measurements with pendulous models and pitot pressure measurements, we have probed flow characteristics in the test section. We have also improved understandings of hypersonic rarefied flows by integrating a numerical approach with the HRWT measurement. The development of the integration scheme between HRWT and numerical approach enables us to estimate the hypersonic rarefied flow characteristics as well as the direct measurement of rarefied aerodynamics. Consequently, this wind tunnel is capable of generating 25 mm-core flows with the free stream Mach number greater than 10 and Knudsen number greater than 0.1.
Assessment of nonequilibrium radiation computation methods for hypersonic flows
NASA Technical Reports Server (NTRS)
Sharma, Surendra
1993-01-01
The present understanding of shock-layer radiation in the low density regime, as appropriate to hypersonic vehicles, is surveyed. Based on the relative importance of electron excitation and radiation transport, the hypersonic flows are divided into three groups: weakly ionized, moderately ionized, and highly ionized flows. In the light of this division, the existing laboratory and flight data are scrutinized. Finally, an assessment of the nonequilibrium radiation computation methods for the three regimes in hypersonic flows is presented. The assessment is conducted by comparing experimental data against the values predicted by the physical model.
CFD Validation Studies for Hypersonic Flow Prediction
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
2001-01-01
A series of experiments to measure pressure and heating for code validation involving hypersonic, laminar, separated flows was conducted at the Calspan-University at Buffalo Research Center (CUBRC) in the Large Energy National Shock (LENS) tunnel. The experimental data serves as a focus for a code validation session but are not available to the authors until the conclusion of this session. The first set of experiments considered here involve Mach 9.5 and Mach 11.3 N2 flow over a hollow cylinder-flare with 30 degree flare angle at several Reynolds numbers sustaining laminar, separated flow. Truncated and extended flare configurations are considered. The second set of experiments, at similar conditions, involves flow over a sharp, double cone with fore-cone angle of 25 degrees and aft-cone angle of 55 degrees. Both sets of experiments involve 30 degree compressions. Location of the separation point in the numerical simulation is extremely sensitive to the level of grid refinement in the numerical predictions. The numerical simulations also show a significant influence of Reynolds number on extent of separation. Flow unsteadiness was easily introduced into the double cone simulations using aggressive relaxation parameters that normally promote convergence.
Pitot pressure analyses in CO2 condensing rarefied hypersonic flows
NASA Astrophysics Data System (ADS)
Ozawa, T.; Suzuki, T.; Fujita, K.
2016-11-01
In order to improve the accuracy of rarefied aerodynamic prediction, a hypersonic rarefied wind tunnel (HRWT) was developed at Japan Aerospace Exploration Agency. While this wind tunnel has been limited to inert gases, such as nitrogen or argon, we recently extended the capability of HRWT to CO2 hypersonic flows for several Mars missions. Compared to our previous N2 cases, the condensation effect may not be negligible for CO2 rarefied aerodynamic measurements. Thus, in this work, we have utilized both experimental and numerical approaches to investigate the condensation and rarefaction effects in CO2 hypersonic nozzle flows.
Hypersonic Viscous Flow Over Large Roughness Elements
NASA Technical Reports Server (NTRS)
Chang, Chau-Lyan; Choudhari, Meelan M.
2009-01-01
Viscous flow over discrete or distributed surface roughness has great implications for hypersonic flight due to aerothermodynamic considerations related to laminar-turbulent transition. Current prediction capability is greatly hampered by the limited knowledge base for such flows. To help fill that gap, numerical computations are used to investigate the intricate flow physics involved. An unstructured mesh, compressible Navier-Stokes code based on the space-time conservation element, solution element (CESE) method is used to perform time-accurate Navier-Stokes calculations for two roughness shapes investigated in wind tunnel experiments at NASA Langley Research Center. It was found through 2D parametric study that at subcritical Reynolds numbers, spontaneous absolute instability accompanying by sustained vortex shedding downstream of the roughness is likely to take place at subsonic free-stream conditions. On the other hand, convective instability may be the dominant mechanism for supersonic boundary layers. Three-dimensional calculations for both a rectangular and a cylindrical roughness element at post-shock Mach numbers of 4.1 and 6.5 also confirm that no self-sustained vortex generation from the top face of the roughness is observed, despite the presence of flow unsteadiness for the smaller post-shock Mach number case.
Hypersonic Viscous Flow Over Large Roughness Elements
NASA Technical Reports Server (NTRS)
Chang, Chau-Lyan; Choudhari, Meelan M.
2009-01-01
Viscous flow over discrete or distributed surface roughness has great implications for hypersonic flight due to aerothermodynamic considerations related to laminar-turbulent transition. Current prediction capability is greatly hampered by the limited knowledge base for such flows. To help fill that gap, numerical computations are used to investigate the intricate flow physics involved. An unstructured mesh, compressible Navier-Stokes code based on the space-time conservation element, solution element (CESE) method is used to perform time-accurate Navier-Stokes calculations for two roughness shapes investigated in wind tunnel experiments at NASA Langley Research Center. It was found through 2D parametric study that at subcritical Reynolds numbers of the boundary layers, absolute instability resulting in vortex shedding downstream, is likely to weaken at supersonic free-stream conditions. On the other hand, convective instability may be the dominant mechanism for supersonic boundary layers. Three-dimensional calculations for a rectangular or cylindrical roughness element at post-shock Mach numbers of 4.1 and 6.5 also confirm that no self-sustained vortex generation is present.
Scaled Rocket Testing in Hypersonic Flow
NASA Technical Reports Server (NTRS)
Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish
2015-01-01
NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.
Study on the numerical schemes for hypersonic flow simulation
NASA Astrophysics Data System (ADS)
Nagdewe, S. P.; Shevare, G. R.; Kim, Heuy-Dong
2009-10-01
Hypersonic flow is full of complex physical and chemical processes, hence its investigation needs careful analysis of existing schemes and choosing a suitable scheme or designing a brand new scheme. The present study deals with two numerical schemes Harten, Lax, and van Leer with Contact (HLLC) and advection upstream splitting method (AUSM) to effectively simulate hypersonic flow fields, and accurately predict shock waves with minimal diffusion. In present computations, hypersonic flows have been modeled as a system of hyperbolic equations with one additional equation for non-equilibrium energy and relaxing source terms. Real gas effects, which appear typically in hypersonic flows, have been simulated through energy relaxation method. HLLC and AUSM methods are modified to incorporate the conservation laws for non-equilibrium energy. Numerical implementation have shown that non-equilibrium energy convect with mass, and hence has no bearing on the basic numerical scheme. The numerical simulation carried out shows good comparison with experimental data available in literature. Both numerical schemes have shown identical results at equilibrium. Present study has demonstrated that real gas effects in hypersonic flows can be modeled through energy relaxation method along with either AUSM or HLLC numerical scheme.
Electron-Beam Diagnostic Methods for Hypersonic Flow Diagnostics
NASA Technical Reports Server (NTRS)
1994-01-01
The purpose of this work was the evaluation of the use of electron-bean fluorescence for flow measurements during hypersonic flight. Both analytical and numerical models were developed in this investigation to evaluate quantitatively flow field imaging concepts based upon the electron beam fluorescence technique for use in flight research and wind tunnel applications. Specific models were developed for: (1) fluorescence excitation/emission for nitrogen, (2) rotational fluorescence spectrum for nitrogen, (3) single and multiple scattering of electrons in a variable density medium, (4) spatial and spectral distribution of fluorescence, (5) measurement of rotational temperature and density, (6) optical filter design for fluorescence imaging, and (7) temperature accuracy and signal acquisition time requirements. Application of these models to a typical hypersonic wind tunnel flow is presented. In particular, the capability of simulating the fluorescence resulting from electron impact ionization in a variable density nitrogen or air flow provides the capability to evaluate the design of imaging instruments for flow field mapping. The result of this analysis is a recommendation that quantitative measurements of hypersonic flow fields using electron-bean fluorescence is a tractable method with electron beam energies of 100 keV. With lower electron energies, electron scattering increases with significant beam divergence which makes quantitative imaging difficult. The potential application of the analytical and numerical models developed in this work is in the design of a flow field imaging instrument for use in hypersonic wind tunnels or onboard a flight research vehicle.
Research on Aeroheating of Hypersonic Reentry Vehicle Base Flow Fields
NASA Astrophysics Data System (ADS)
Xuguo, Qin; Yongtao, Shui; Yonghai, Wang; Gang, Chen; Qiang, Li
2017-09-01
The structure of the base flow of a hypersonic reentry vehicle and the resulting base pressure and heat transfer have been studied by numerical study. The compressible Navier-Stokes equations are solved by the finite-volume method. SST k-ω turbulence model is used, and comparisons are made with flight test. Attention was focused on assessing the effects of angle of attack and Mach number. It was found that angle of attack can significantly alter the wake flow structure and reentry vehicle base pressure and heating distributions. The results of the simulation may provide a theoretical basis for the design of the thermal protection system of hypersonic reentry vehicles.
Experimental results for a hypersonic nozzle/afterbody flow field
NASA Technical Reports Server (NTRS)
Spaid, Frank W.; Keener, Earl R.; Hui, Frank C. L.
1995-01-01
This study was conducted to experimentally characterize the flow field created by the interaction of a single-expansion ramp-nozzle (SERN) flow with a hypersonic external stream. Data were obtained from a generic nozzle/afterbody model in the 3.5 Foot Hypersonic Wind Tunnel at the NASA Ames Research Center, in a cooperative experimental program involving Ames and McDonnell Douglas Aerospace. The model design and test planning were performed in close cooperation with members of the Ames computational fluid dynamics (CFD) team for the National Aerospace Plane (NASP) program. This paper presents experimental results consisting of oil-flow and shadow graph flow-visualization photographs, afterbody surface-pressure distributions, rake boundary-layer measurements, Preston-tube skin-friction measurements, and flow field surveys with five-hole and thermocouple probes. The probe data consist of impact pressure, flow direction, and total temperature profiles in the interaction flow field.
The Analysis of Underexpanded Jet Flows for Hypersonic Aerodynamic Experiments in Vacuum Chambers
NASA Astrophysics Data System (ADS)
Riabov, V. V.; Fedoseyev, A. I.
Underexpanded jets have become widely used in studies of rarefied-gas flows [1]- [3] and aerodynamics of hypersonic probes in wind tunnels [4]-[7]. The objective of the present study is to analyze shapes and flow parameters in internal regions of hypersonic underexpanded viscous jets, and to apply the jet theory to hypersonic studies
Analytical studies of hypersonic viscous dissociated flows
NASA Technical Reports Server (NTRS)
Inger, George R.
1995-01-01
This project primarily dealt with integral boundary-layer solution techniques that are directly applicable to the problem of determining aerodynamic heating rates of hypersonic vehicles like X-33 in the vicinity of stagnation points, windward centerlines, and swept-wing leading edges. The analyses include effects of finite-rate gas chemistry across the boundary layer and finite-rate catalysis of atom recombination at the surface. A new approach for combining the insight afforded by integral boundary-layer analysis with comprehensive (and expensive) computational fluid dynamic (CFD) flowfield solutions of the thin-layer Navier-Stokes equations was developed. The approach extracts CFD derived quantities at the wall and at the boundary layer edge for inclusion in a post-processing boundary-layer analysis. The post-processed data base allows a designer at a workstation to ask and answer the following questions: (1) How much does the heating change if one uses a thermal protection system (TPS) with different catalytic properties than was used in the original CFD solution? (2) How does the heating change when one moves the interface of two different TPS materials with different catalytic efficiencies for the purpose of reducing vehicle weight and expense? The answer to the second question is particularly critical, because abrupt changes from low catalytic efficiency to high catalytic efficiency can lead to localized increase in heating which exceeds the usually conservative estimate provided by a fully catalytic wall assumption. A secondary issue that was addressed involves the prediction of heating levels in the vicinity of sharp corners that are transverse to or aligned with the flow. An example of the first case is heating at the edge of the COMET reentry module. An example of the second case is heating along the side edge of a deflected body flap on an SSV. The difficulty of putting grids in the vicinity of such corners with continuously varying metric coefficients
Calculations of Supersonic and Hypersonic Flows using Compressible Wall Functions
NASA Technical Reports Server (NTRS)
Huang, P. G.; Coakley, T. J.
1993-01-01
The present paper presents a numerical procedure to calculate supersonic and hypersonic flows using the compressible law of the wall. The turbulence models under consideration include the Launder-Reece-Rodi-Gibson Reynolds-stress model and the k-epsilon model. The models coupled with the proposed wall function technique have been tested in both separated and unseparated flows. The flows include (1) an insulated flat plate flow over a range of Mach numbers, (2) a Mach 5 flat plate flow with cold wall conditions, (3) a two dimensional supersonic compression corner flow, (4) a hypersonic flow over an axisymmetric flare, and (5) a hypersonic flow over a 2-D compression corner. Results indicate that the wall function technique gives improved predictions of skin friction and heat transfer in separated flows compared with models using wall dampers. Predictions of the extent of separation are not improved over the wall damper models except with the Reynolds-stress model for the supersonic compression corner flow case.
Hot-wire anemometry in hypersonic helium flow
NASA Technical Reports Server (NTRS)
Wagner, R. D.; Weinstein, L. M.
1974-01-01
Hot-wire anemometry techniques are described that have been developed and used for hypersonic-helium-flow studies. The short run time available dictated certain innovations in applying conventional hot-wire techniques. Some examples are given to show the application of the techniques used. Modifications to conventional equipment are described, including probe modifications and probe heating controls.
Science and Technology Text Mining: Hypersonic and Supersonic Flow
2006-05-31
human analyst. DT was used to derive technical intelligence from a hypersonic/ supersonic flow ( HSF ) database derived from the Science Citation Index and...the Engineering Compendex. Phrase frequency analysis by the technical domain expert provided the pervasive technical themes of the HSF database, and...the phrase proximity analysis provided the relationships among the pervasive technical themes. Bibliometric analysis of the HSF literature
Investigation of Hypersonic Nozzle Flow Uniformity Using NO Fluorescence
NASA Technical Reports Server (NTRS)
O'Byrne, S.; Danehy, P. J.; Houwing, A. F. P.
2005-01-01
Planar laser-induced fluorescence visualisation is used to investigate nonuniformities in the flow of a hypersonic conical nozzle. Possible causes for the nonuniformity are outlined and investigated, and the problem is shown to be due to a small step at the nozzle throat. Entrainment of cold boundary layer gas is postulated as the cause of the signal nonuniformity.
An assessment of laser velocimetry in hypersonic flow
NASA Technical Reports Server (NTRS)
1992-01-01
Although extensive progress has been made in computational fluid mechanics, reliable flight vehicle designs and modifications still cannot be made without recourse to extensive wind tunnel testing. Future progress in the computation of hypersonic flow fields is restricted by the need for a reliable mean flow and turbulence modeling data base which could be used to aid in the development of improved empirical models for use in numerical codes. Currently, there are few compressible flow measurements which could be used for this purpose. In this report, the results of experiments designed to assess the potential for laser velocimeter measurements of mean flow and turbulent fluctuations in hypersonic flow fields are presented. Details of a new laser velocimeter system which was designed and built for this test program are described.
Review and assessment of turbulence models for hypersonic flows
NASA Astrophysics Data System (ADS)
Roy, Christopher J.; Blottner, Frederick G.
2006-10-01
Accurate aerodynamic prediction is critical for the design and optimization of hypersonic vehicles. Turbulence modeling remains a major source of uncertainty in the computational prediction of aerodynamic forces and heating for these systems. The first goal of this article is to update the previous comprehensive review of hypersonic shock/turbulent boundary-layer interaction experiments published in 1991 by Settles and Dodson (Hypersonic shock/boundary-layer interaction database. NASA CR 177577, 1991). In their review, Settles and Dodson developed a methodology for assessing experiments appropriate for turbulence model validation and critically surveyed the existing hypersonic experiments. We limit the scope of our current effort by considering only two-dimensional (2D)/axisymmetric flows in the hypersonic flow regime where calorically perfect gas models are appropriate. We extend the prior database of recommended hypersonic experiments (on four 2D and two 3D shock-interaction geometries) by adding three new geometries. The first two geometries, the flat plate/cylinder and the sharp cone, are canonical, zero-pressure gradient flows which are amenable to theory-based correlations, and these correlations are discussed in detail. The third geometry added is the 2D shock impinging on a turbulent flat plate boundary layer. The current 2D hypersonic database for shock-interaction flows thus consists of nine experiments on five different geometries. The second goal of this study is to review and assess the validation usage of various turbulence models on the existing experimental database. Here we limit the scope to one- and two-equation turbulence models where integration to the wall is used (i.e., we omit studies involving wall functions). A methodology for validating turbulence models is given, followed by an extensive evaluation of the turbulence models on the current hypersonic experimental database. A total of 18 one- and two-equation turbulence models are reviewed
Multigrid for hypersonic viscous two- and three-dimensional flows
NASA Technical Reports Server (NTRS)
Turkel, E.; Swanson, R. C.; Vatsa, V. N.; White, J. A.
1991-01-01
The use of a multigrid method with central differencing to solve the Navier-Stokes equations for hypersonic flows is considered. The time dependent form of the equations is integrated with an explicit Runge-Kutta scheme accelerated by local time stepping and implicit residual smoothing. Variable coefficients are developed for the implicit process that removes the diffusion limit on the time step, producing significant improvement in convergence. A numerical dissipation formulation that provides good shock capturing capability for hypersonic flows is presented. This formulation is shown to be a crucial aspect of the multigrid method. Solutions are given for two-dimensional viscous flow over a NACA 0012 airfoil and three-dimensional flow over a blunt biconic.
Investigation of flow fields within large scale hypersonic inlet models
NASA Technical Reports Server (NTRS)
Gnos, A. V.; Watson, E. C.; Seebaugh, W. R.; Sanator, R. J.; Decarlo, J. P.
1973-01-01
Analytical and experimental investigations were conducted to determine the internal flow characteristics in model passages representative of hypersonic inlets for use at Mach numbers to about 12. The passages were large enough to permit measurements to be made in both the core flow and boundary layers. The analytical techniques for designing the internal contours and predicting the internal flow-field development accounted for coupling between the boundary layers and inviscid flow fields by means of a displacement-thickness correction. Three large-scale inlet models, each having a different internal compression ratio, were designed to provide high internal performance with an approximately uniform static-pressure distribution at the throat station. The models were tested in the Ames 3.5-Foot Hypersonic Wind Tunnel at a nominal free-stream Mach number of 7.4 and a unit free-stream Reynolds number of 8.86 X one million per meter.
Three-Dimensional Aeroelastic and Aerothermoelastic Behavior in Hypersonic Flow
NASA Technical Reports Server (NTRS)
McNamara, Jack J.; Friedmann, Peretz P.; Powell, Kenneth G.; Thuruthimattam, Biju J.; Bartels, Robert E.
2005-01-01
The aeroelastic and aerothermoelastic behavior of three-dimensional configurations in hypersonic flow regime are studied. The aeroelastic behavior of a low aspect ratio wing, representative of a fin or control surface on a generic hypersonic vehicle, is examined using third order piston theory, Euler and Navier-Stokes aerodynamics. The sensitivity of the aeroelastic behavior generated using Euler and Navier-Stokes aerodynamics to parameters governing temporal accuracy is also examined. Also, a refined aerothermoelastic model, which incorporates the heat transfer between the fluid and structure using CFD generated aerodynamic heating, is used to examine the aerothermoelastic behavior of the low aspect ratio wing in the hypersonic regime. Finally, the hypersonic aeroelastic behavior of a generic hypersonic vehicle with a lifting-body type fuselage and canted fins is studied using piston theory and Euler aerodynamics for the range of 2.5 less than or equal to M less than or equal to 28, at altitudes ranging from 10,000 feet to 80,000 feet. This analysis includes a study on optimal mesh selection for use with Euler aerodynamics. In addition to the aeroelastic and aerothermoelastic results presented, three time domain flutter identification techniques are compared, namely the moving block approach, the least squares curve fitting method, and a system identification technique using an Auto-Regressive model of the aeroelastic system. In general, the three methods agree well. The system identification technique, however, provided quick damping and frequency estimations with minimal response record length, and therefore o ers significant reductions in computational cost. In the present case, the computational cost was reduced by 75%. The aeroelastic and aerothermoelastic results presented illustrate the applicability of the CFL3D code for the hypersonic flight regime.
Approximate convective heating equations for hypersonic flows
NASA Technical Reports Server (NTRS)
Zoby, E. V.; Moss, J. N.; Sutton, K.
1979-01-01
Laminar and turbulent heating-rate equations appropriate for engineering predictions of the convective heating rates about blunt reentry spacecraft at hypersonic conditions are developed. The approximate methods are applicable to both nonreacting and reacting gas mixtures for either constant or variable-entropy edge conditions. A procedure which accounts for variable-entropy effects and is not based on mass balancing is presented. Results of the approximate heating methods are in good agreement with existing experimental results as well as boundary-layer and viscous-shock-layer solutions.
Disturbances from Shock/Boundary-Layer Interactions Affecting Upstream Hypersonic Flow
2005-12-01
2180, NASA, 1983. 11. J. L. Stollery. Some Viscous Interactions Affecting the Design of Hypersonic Intakes and Nozzles. Advances in Hypersonics ...affecting upstream hypersonic flow F49620-03-1-0030 Craig Ryan Skoch Purdue University, School of Aeronautics and Astronautics none Air Force Office of...separations from propagating upstream. hypersonic laminar-turbulent transition, quiet wind tunnels, shock/boundary-layer interaction U U U Unlimited 132
Two-equation turbulence modeling for 3-D hypersonic flows
NASA Technical Reports Server (NTRS)
Bardina, J. E.; Coakley, T. J.; Marvin, J. G.
1992-01-01
An investigation to verify, incorporate and develop two-equation turbulence models for three-dimensional high speed flows is presented. The current design effort of hypersonic vehicles has led to an intensive study of turbulence models for compressible hypersonic flows. This research complements an extensive review of experimental data and the current development of 2D turbulence models. The review of experimental data on 2D and 3D flows includes complex hypersonic flows with pressure profiles, skin friction, wall heat transfer, and turbulence statistics data. In a parallel effort, turbulence models for high speed flows have been tested against flat plate boundary layers, and are being tested against the 2D database. In the present paper, we present the results of 3D Navier-Stokes numerical simulations with an improved k-omega two-equation turbulence model against experimental data and empirical correlations of an adiabatic flat plate boundary layer, a cold wall flat plate boundary layer, and a 3D database flow, the interaction of an oblique shock wave and a thick turbulent boundary layer with a free stream Mach number = 8.18 and Reynolds number = 5 x 10 to the 6th.
Progress with multigrid schemes for hypersonic flow problems
NASA Technical Reports Server (NTRS)
Radespiel, R.; Swanson, R. C.
1991-01-01
Several multigrid schemes are considered for the numerical computation of viscous hypersonic flows. For each scheme, the basic solution algorithm uses upwind spatial discretization with explicit multistage time stepping. Two level versions of the various multigrid algorithms are applied to the two dimensional advection equation, and Fourier analysis is used to determine their damping properties. The capabilities of the multigrid methods are assessed by solving three different hypersonic flow problems. Some new multigrid schemes based on semicoarsening strategies are shown to be quite effective in relieving the stiffness caused by the high aspect ratio cells required to resolve high Reynolds number flows. These schemes exhibit good convergence rates for Reynolds numbers up to 200 x 10(exp 6) and Mach numbers up to 25.
A technique for measuring hypersonic flow velocity profiles
NASA Technical Reports Server (NTRS)
Gartrell, L. R.
1973-01-01
A technique for measuring hypersonic flow velocity profiles is described. This technique utilizes an arc-discharge-electron-beam system to produce a luminous disturbance in the flow. The time of flight of this disturbance was measured. Experimental tests were conducted in the Langley pilot model expansion tube. The measured velocities were of the order of 6000 m/sec over a free-stream density range from 0.000196 to 0.00186 kg/cu m. The fractional error in the velocity measurements was less than 5 percent. Long arc discharge columns (0.356 m) were generated under hypersonic flow conditions in the expansion-tube modified to operate as an expansion tunnel.
Portable Fluorescence Imaging System for Hypersonic Flow Facilities
NASA Technical Reports Server (NTRS)
Wilkes, J. A.; Alderfer, D. W.; Jones, S. B.; Danehy, P. M.
2003-01-01
A portable fluorescence imaging system has been developed for use in NASA Langley s hypersonic wind tunnels. The system has been applied to a small-scale free jet flow. Two-dimensional images were taken of the flow out of a nozzle into a low-pressure test section using the portable planar laser-induced fluorescence system. Images were taken from the center of the jet at various test section pressures, showing the formation of a barrel shock at low pressures, transitioning to a turbulent jet at high pressures. A spanwise scan through the jet at constant pressure reveals the three-dimensional structure of the flow. Future capabilities of the system for making measurements in large-scale hypersonic wind tunnel facilities are discussed.
Perspectives on hypersonic viscous and nonequilibrium flow research
NASA Technical Reports Server (NTRS)
Cheng, H. K.
1992-01-01
An attempt is made to reflect on current focuses in certain areas of hypersonic flow research by examining recent works and their issues. Aspects of viscous interaction, flow instability, and nonequilibrium aerothermodynamics pertaining to theoretical interest are focused upon. The field is a diverse one, and many exciting works may have either escaped the writer's notice or been abandoned for the sake of space. Students of hypersonic viscous flow must face the transition problems towards the two opposite ends of the Reynolds or Knudsen number range, which represents two regimes where unresolved fluid/gas dynamic problems abound. Central to the hypersonic flow studies is high-temperature physical gas dynamics; here, a number of issues on modelling the intermolecular potentials and inelastic collisions remain the obstacles to quantitative predictions. Research in combustion and scramjet propulsion will certainly be benefitted by advances in turbulent mixing and new computational fluid dynamics (CFD) strategies on multi-scaled complex reactions. Even for the sake of theoretical development, the lack of pertinent experimental data in the right energy and density ranges is believed to be among the major obstacles to progress in aerothermodynamic research for hypersonic flight. To enable laboratory simulation of nonequilibrium effects anticipated for transatmospheric flight, facilities capable of generating high enthalpy flow at density levels higher than in existing laboratories are needed (Hornung 1988). A new free-piston shock tunnel capable of realizing a test-section stagnation temperature of 10(exp 5) at Reynolds number 50 x 10(exp 6)/cm is being completed and preliminary tests has begun (H. Hornung et al. 1992). Another laboratory study worthy of note as well as theoretical support is the nonequilibrium flow experiment of iodine vapor which has low activation energies for vibrational excitation and dissociation, and can be studied in a laboratory with modest
Optical Window Materials For Hypersonic Flow
NASA Astrophysics Data System (ADS)
Au, Robert H.
1989-09-01
Optical window materials were investigated for infrared sensor systems used in observing ground targets from a hypersonic-glide vehicle. The equilibrium temperature of the window in the glide region depends on the emissivity and varied between 1,370 and 2,250 K. The high temperatures showed that a protective cover over the window is required during the entire glide region of the trajectory. Ejection of the window cover at 70-kft altitude in the terminal region was assumed, resulting in maximum window temperatures of 565 K and 592 K for magnesium oxide and diamond windows, respectively, both 0.8-in thick. The window temperatures for germanium and sapphire were also calculated. Thermal shock, thermal expansion, the effects of the window radiation on the infrared detectors and methods to reduce the hot window problem were examined.
Computational study of generic hypersonic vehicle flow fields
NASA Technical Reports Server (NTRS)
Narayan, Johnny R.
1994-01-01
The geometric data of the generic hypersonic vehicle configuration included body definitions and preliminary grids for the forebody (nose cone excluded), midsection (propulsion system excluded), and afterbody sections. This data was to be augmented by the nose section geometry (blunt conical section mated with the noncircular cross section of the forebody initial plane) along with a grid and a detailed supersonic combustion ramjet (scramjet) geometry (inlet and combustor) which should be merged with the nozzle portion of the afterbody geometry. The solutions were to be obtained by using a Navier-Stokes (NS) code such as TUFF for the nose portion, a parabolized Navier-Stokes (PNS) solver such as the UPS and STUFF codes for the forebody, a NS solver with finite rate hydrogen-air chemistry capability such as TUFF and SPARK for the scramjet and a suitable solver (NS or PNS) for the afterbody and external nozzle flows. The numerical simulation of the hypersonic propulsion system for the generic hypersonic vehicle is the major focus of this entire work. Supersonic combustion ramjet is such a propulsion system, hence the main thrust of the present task has been to establish a solution procedure for the scramjet flow. The scramjet flow is compressible, turbulent, and reacting. The fuel used is hydrogen and the combustion process proceeds at a finite rate. As a result, the solution procedure must be capable of addressing such flows.
Computation of Hypersonic Flow about Maneuvering Vehicles with Changing Shapes
Ferencz, R M; Felker, F F; Castillo, V M
2004-02-23
Vehicles moving at hypersonic speeds have great importance to the National Security. Ballistic missile re-entry vehicles (RV's) travel at hypersonic speeds, as do missile defense intercept vehicles. Despite the importance of the problem, no computational analysis method is available to predict the aerodynamic environment of maneuvering hypersonic vehicles, and no analysis is available to predict the transient effects of their shape changes. The present state-of-the-art for hypersonic flow calculations typically still considers steady flow about fixed shapes. Additionally, with present computational methods, it is not possible to compute the entire transient structural and thermal loads for a re-entry vehicle. The objective of this research is to provide the required theoretical development and a computational analysis tool for calculating the hypersonic flow about maneuvering, deforming RV's. This key enabling technology will allow the development of a complete multi-mechanics simulation of the entire RV flight sequence, including important transient effects such as complex flight dynamics. This will allow the computation of the as-delivered state of the payload in both normal and unusual operational environments. This new analysis capability could also provide the ability to predict the nonlinear, transient behavior of endo-atmospheric missile interceptor vehicles to the input of advanced control systems. Due to the computational intensity of fluid dynamics for hypersonics, the usual approach for calculating the flow about a vehicle that is changing shape is to complete a series of steady calculations, each with a fixed shape. However, this quasi-steady approach is not adequate to resolve the frequencies characteristic of a vehicle's structural dynamics. Our approach is to include the effects of the unsteady body shape changes in the finite-volume method by allowing for arbitrary translation and deformation of the control volumes. Furthermore, because the Eulerian
Molecular-Based Optical Diagnostics for Hypersonic Nonequilibrium Flows
NASA Technical Reports Server (NTRS)
Danehy, Paul; Bathel, Brett; Johansen, Craig; Winter, Michael; O'Byrne, Sean; Cutler, Andrew
2015-01-01
This presentation package consists of seven different talks rolled up into one. These talks are all invited orals presentations in a special session at the Aviation 2015 conference and represent contributions that were made to a recent AIAA book that will be published entitled 'Hypersonic Nonequilibrium Flows: Fundamentals and Recent Advances'. Slide 5 lists the individual presentations that will be given during the special session.
Computational analysis of hypersonic flows past elliptic-cone waveriders
NASA Technical Reports Server (NTRS)
Yoon, Bok-Hyun; Rasmussen, Maurice L.
1991-01-01
A comprehensive study for the inviscid numerical calculation of the hypersonic flow past a class of elliptic-cone derived waveriders is presented. The theoretical background associated with hypersonic small-disturbance theory (HSDT) is reviewed. Several approximation formulas for the waverider compression surface are established. A CFD algorithm is used to calculate flow fields for the on-design case and a variety of off-design cases. The results are compared with HSDT, experiment, and other available CFD results. For the waverider shape used in previous investigations, the bow shock for the on-design condition stands off from the leading-edge tip of the waverider. It was found that this occurs because the tip was too thick according to the approximating shape formula that was used to describe the compression surface. When this was corrected, the bow shock became closer to attached as it should be. At Mach numbers greater than the design condition, a lambda-shock configuration develops near the tip of the compression surface. At negative angles of attack, other complicated shock patterns occur near the leading-edge tip. These heretofore unknown flow patterns show the power and utility of CFD for investigating novel hypersonic configurations such as waveriders.
Turbulence Models for Accurate Aerothermal Prediction in Hypersonic Flows
NASA Astrophysics Data System (ADS)
Zhang, Xiang-Hong; Wu, Yi-Zao; Wang, Jiang-Feng
Accurate description of the aerodynamic and aerothermal environment is crucial to the integrated design and optimization for high performance hypersonic vehicles. In the simulation of aerothermal environment, the effect of viscosity is crucial. The turbulence modeling remains a major source of uncertainty in the computational prediction of aerodynamic forces and heating. In this paper, three turbulent models were studied: the one-equation eddy viscosity transport model of Spalart-Allmaras, the Wilcox k-ω model and the Menter SST model. For the k-ω model and SST model, the compressibility correction, press dilatation and low Reynolds number correction were considered. The influence of these corrections for flow properties were discussed by comparing with the results without corrections. In this paper the emphasis is on the assessment and evaluation of the turbulence models in prediction of heat transfer as applied to a range of hypersonic flows with comparison to experimental data. This will enable establishing factor of safety for the design of thermal protection systems of hypersonic vehicle.
Surface pressure measurements for CFD code validation in hypersonic flow
Oberkampf, W.L.; Aeschliman, D.P.; Henfling, J.F.; Larson, D.E.
1995-07-01
Extensive surface pressure measurements were obtained on a hypersonic vehicle configuration at Mach 8. All of the experimental results were obtained in the Sandia National Laboratories Mach 8 hypersonic wind tunnel for laminar boundary layer conditions. The basic vehicle configuration is a spherically blunted 10{degrees} half-angle cone with a slice parallel with the axis of the vehicle. The bluntness ratio of the geometry is 10% and the slice begins at 70% of the length of the vehicle. Surface pressure measurements were obtained for angles of attack from {minus}10 to + 18{degrees}, for various roll angles, at 96 locations on the body surface. A new and innovative uncertainty analysis was devised to estimate the contributors to surface pressure measurement uncertainty. Quantitative estimates were computed for the uncertainty contributions due to the complete instrumentation system, nonuniformity of flow in the test section of the wind tunnel, and variations in the wind tunnel model. This extensive set of high-quality surface pressure measurements is recommended for use in the calibration and validation of computational fluid dynamics codes for hypersonic flow conditions.
On the instability of hypersonic flow past a flat plate
NASA Technical Reports Server (NTRS)
Blackaby, Nicholas D.; Cowley, Stephen J.; Hall, Philip
1993-01-01
Qualitative features of the inviscid instability characteristics of hypersonic boundary-layer flows over a flat plate are considered. The instability of a viscous hypersonic boundary layer which exists far downstream from the leading edge of the plate. It is shown that the vorticity mode of instability operates on a different lengthscale from that obtained using a Chapman viscosity law. The growth rate predicted by a linear viscosity law is found to overestimate the size of the growth rate. The inviscid instability of the boundary layer near the leading edge interaction zone is discussed focusing on the strong-interaction zone which occurs sufficiently close to the leading edge. The vorticity mode in this regime is found to be unstable.
On the instability of hypersonic flow past a wedge
NASA Technical Reports Server (NTRS)
Cowley, Stephen; Hall, Philip
1988-01-01
The instability of a compressible flow past a wedge is investigated in the hypersonic limit. Particular attention is given to the Tollmien-Schlichting waves governed by triple-deck theory though some discussion of inviscid modes is given. It is shown that the attached shock has a significant effect on the growth rates of Tollmien-Schlichting waves. Moreover, the presence of the shock allows for more than one unstable Tollmien-Schlichting wave. Indeed, an infinite discrete spectrum of unstable waves is induced by the shock, but these modes are unstable over relatively small but high frequency ranges. The shock is shown to have little effect on the inviscid modes considered by previous authors and an asymptotic description of inviscid modes in the hypersonic limit is given.
Unstructured Mesh Methods for the Simulation of Hypersonic Flows
NASA Technical Reports Server (NTRS)
Peraire, Jaime; Bibb, K. L. (Technical Monitor)
2001-01-01
This report describes the research work undertaken at the Massachusetts Institute of Technology. The aim of this research is to identify effective algorithms and methodologies for the efficient and routine solution of hypersonic viscous flows about re-entry vehicles. For over ten years we have received support from NASA to develop unstructured mesh methods for Computational Fluid Dynamics. As a result of this effort a methodology based on the use, of unstructured adapted meshes of tetrahedra and finite volume flow solvers has been developed. A number of gridding algorithms flow solvers, and adaptive strategies have been proposed. The most successful algorithms developed from the basis of the unstructured mesh system FELISA. The FELISA system has been extensively for the analysis of transonic and hypersonic flows about complete vehicle configurations. The system is highly automatic and allows for the routine aerodynamic analysis of complex configurations starting from CAD data. The code has been parallelized and utilizes efficient solution algorithms. For hypersonic flows, a version of the, code which incorporates real gas effects, has been produced. One of the latest developments before the start of this grant was to extend the system to include viscous effects. This required the development of viscous generators, capable of generating the anisotropic grids required to represent boundary layers, and viscous flow solvers. In figures I and 2, we show some sample hypersonic viscous computations using the developed viscous generators and solvers. Although these initial results were encouraging, it became apparent that in order to develop a fully functional capability for viscous flows, several advances in gridding, solution accuracy, robustness and efficiency were required. As part of this research we have developed: 1) automatic meshing techniques and the corresponding computer codes have been delivered to NASA and implemented into the GridEx system, 2) a finite
Nonlinear Instability of Hypersonic Flow past a Wedge
NASA Technical Reports Server (NTRS)
Seddougui, Sharon O.; Bassom, Andrew P.
1991-01-01
The nonlinear stability of a compressible flow past a wedge is investigated in the hypersonic limit. The analysis follows the ideas of a weakly nonlinear approach. Interest is focussed on Tollmien-Schlichting waves governed by a triple deck structure and it is found that the attached shock can profoundly affect the stability characteristics of the flow. In particular, it is shown that nonlinearity tends to have a stabilizing influence. The nonlinear evolution of the Tollmien-Schlichting mode is described in a number of asymptotic limits.
On the instability of hypersonic flow past a flat plate
NASA Technical Reports Server (NTRS)
Blackaby, Nicholas; Cowley, Stephen; Hall, Philip
1990-01-01
The instability of hypersonic boundary-layer flows over flat plates is considered. The viscosity of the fluid is taken to be governed by Sutherland's law, which gives a much more accurate representation of the temperature dependence of fluid viscosity at hypersonic speeds than Chapman's approximate linear law; although at lower speeds the temperature variation of the mean state is less pronounced so that the Chapman law can be used with some confidence. Attention is focussed on the so-called (vorticity) mode of instability of the viscous hypersonic boundary layer. This is thought to be the fastest growing inviscid disturbance at hypersonic speeds; it is also believed to have an asymptotically larger growth rate than any viscous or centrifugal instability. As a starting point the instability of the hypersonic boundary layer which exists far downstream from the leading edge of the plate is investigated. In this regime the shock that is attached to the leading edge of the plate plays no role, so that the basic boundary layer is non-interactive. It is shown that the vorticity mode of instability of this flow operates on a significantly different lengthscale than that obtained if a Chapman viscosity law is assumed. In particular, it is found that the growth rate predicted by a linear viscosity law overestimates the size of the growth rate by O(M(exp 2). Next, the development of the vorticity mode as the wavenumber decreases is described, and it is shown that acoustic modes emerge when the wavenumber has decreased from it's O(1) initial value to O(M (exp -3/2). Finally, the inviscid instability of the boundary layer near the leading edge in the interaction zone is discussed and particular attention is focussed on the strong interaction region which occurs sufficiently close to the leading edge. It is found that the vorticity mode in this regime is again unstable, and that it is concentrated in the transition layer at the edge of the boundary layer where the temperature
Internal corner flow fields. [calculations for super/hypersonic inlets
NASA Technical Reports Server (NTRS)
Marconi, F.
1979-01-01
A computational procedure has been developed to predict the inviscid super/hypersonic flow field of conical internal corners. The prediction of internal corner flow fields can be important in the design of supersonic 'box' type inlets. The computational procedure utilizes a second order finite difference marching technique to asymptote to the conical corner flow solution of Euler's equations. These flow fields are dominated by complex shock interactions. All discontinuities, shocks and slip surfaces are fitted with the appropriate jump conditions. The 'triple' points (the interaction of two shocks and a slip surface) are also computed exactly. Computed results are compared with experimental data and the computational results of other investigators. In addition, the sensitivity of these flow fields to a number of geometric parameters is studied, and the impact of these flows on inlet performance is assessed.
Euler and Navier-Stokes solutions for hypersonic flows
NASA Technical Reports Server (NTRS)
Thareja, Rajiv R.; Prabhu, Ramadas K.; Stewart, James R.; Morgan, Ken; Peraire, Jaime
1989-01-01
An upwind finite-element technique that uses cell-centered quantities and implicit and/or explicit time marching has been developed for computing hypersonic laminar viscous flows using adaptive unstructured grids in two and three dimensions. A perfect gas model as well as an equilibrium air model is implemented for solving high-speed flows. A first-order basic scheme and a higher-order flux-corrected transport (FCT) scheme have been implemented. This technique has been used to predict 'Type III and IV' shock interactions on a cylinder in two dimensions and a swept cylinder in three dimensions, with a view to determine the pressure and heating rate augmentation caused by an impinging shock on the leading edge of a cowl lip of an engine inlet. The predictions of wall pressure and heating rates compare very well with experimental data. The flow features are very distinctly captured with a sequence of adaptively-generated grids. Three-dimensional corner flow, typically encountered in engine inlets due to compression of the flow by ramps in the walls, is also modeled. This procedure is the first step in developing an integrated fluid, thermal, structural analysis capability for hypersonic flight vehicles like the National Aero-Space Plane.
Euler and Navier-Stokes solutions for hypersonic flows
NASA Technical Reports Server (NTRS)
Thareja, Rajiv R.; Prabhu, Ramadas K.; Stewart, James R.; Morgan, Ken; Peraire, Jaime
1989-01-01
An upwind finite-element technique that uses cell-centered quantities and implicit and/or explicit time marching has been developed for computing hypersonic laminar viscous flows using adaptive unstructured grids in two and three dimensions. A perfect gas model as well as an equilibrium air model is implemented for solving high-speed flows. A first-order basic scheme and a higher-order flux-corrected transport (FCT) scheme have been implemented. This technique has been used to predict 'Type III and IV' shock interactions on a cylinder in two dimensions and a swept cylinder in three dimensions, with a view to determine the pressure and heating rate augmentation caused by an impinging shock on the leading edge of a cowl lip of an engine inlet. The predictions of wall pressure and heating rates compare very well with experimental data. The flow features are very distinctly captured with a sequence of adaptively-generated grids. Three-dimensional corner flow, typically encountered in engine inlets due to compression of the flow by ramps in the walls, is also modeled. This procedure is the first step in developing an integrated fluid, thermal, structural analysis capability for hypersonic flight vehicles like the National Aero-Space Plane.
Flow-Tagging Velocimetry for Hypersonic Flows Using Fluorescence of Nitric Oxide
NASA Technical Reports Server (NTRS)
Danehy, P. M.; OByrne, S.; Houwing, A. F. P.
2001-01-01
We investigate a new type of flow-tagging velocimetry technique for hypersonic flows. The technique involves exciting a thin line of nitric oxide molecules with a laser beam and then, after some delay, acquiring an image of the displaced line. One component of velocity is determined from the time of flight. This method is applied to measure the velocity profile in a Mach 8.5 laminar, hypersonic boundary layer in the Australian National Universities T2 free-piston shock tunnel. The velocity is measured with an uncertainty of approximately 2%. Comparison with a CFD simulation of the flow shows reasonable agreement.
Sonic injection through diamond orifices into a hypersonic flow
NASA Astrophysics Data System (ADS)
Fan, Huaiguo
The objective for the present study was to experimentally characterize the performance of diamond shaped injectors for hypersonic flow applications. First, an extensive literature review was performed. Second, a small scale Mach 5.0 wind tunnel facility was installed. Third, a detailed experimental parametric investigation of sonic injection through a diamond orifice (five incidence angles and three momentum ratios) and a circular injector (three momentum ratios) into the Mach 5.0 freestream was performed. Also, the use of downstream plume vorticity control ramps was investigated. Fourth, a detailed analysis of the experimental data to characterize and model the flow for the present range of conditions was achieved. The experimental techniques include surface oil flow visualization, Mie-Scattering flow visualization, particle image velocimetry (PIV), shadowgraph photograph, and a five-hole mean flow probe. The results show that the diamond injectors have the potential to produce attached shock depending on the incidence angle and jet momentum ratio. For example, the incidence angles less than or equal to 45° at J = 0.43 generated attached interaction shocks. The attached shock produced reduced total pressure loss (drag for scramjet) and eliminated potential hot spots, associated with the upstream flow separation. The jet interaction shock angle increased with jet incidence angle and momentum ratio due to increased penetration and flow disturbances. The plume penetration and cross-sectional area increased with incidence angle and momentum ratio. The increased jet interaction shock angle and strength produced increased total pressure loss, jet interaction force and total normal force. The characteristic kidney bean shaped plume was not discernable from the diamond injectors indicating increased effectiveness for film cooling applications. A vorticity generation ramp increased the penetration of the plume and the plume shape was indicative of higher levels of
Nonintrusive Temperature and Velocity Measurements in a Hypersonic Nozzle Flow
NASA Technical Reports Server (NTRS)
OByrne, S.; Danehy, P. M.; Houwing, A. F. P.
2002-01-01
Distributions of nitric oxide vibrational temperature, rotational temperature and velocity have been measured in the hypersonic freestream at the exit of a conical nozzle, using planar laser-induced fluorescence. Particular attention has been devoted to reducing the major sources of systematic error that can affect fluorescence tempera- ture measurements, including beam attenuation, transition saturation effects, laser mode fluctuations and transition choice. Visualization experiments have been performed to improve the uniformity of the nozzle flow. Comparisons of measured quantities with a simple one-dimensional computation are made, showing good agreement between measurements and theory given the uncertainty of the nozzle reservoir conditions and the vibrational relaxation rate.
Supercomputer modeling of flow past hypersonic flight vehicles
NASA Astrophysics Data System (ADS)
Ermakov, M. K.; Kryukov, I. A.
2017-02-01
A software platform for MPI-based parallel solution of the Navier-Stokes (Euler) equations for viscous heat-conductive compressible perfect gas on 3-D unstructured meshes is developed. The discretization and solution of the Navier-Stokes equations are constructed on generalized S.K. Godunov’s method and the second order approximation in space and time. Developed software platform allows to carry out effectively flow past hypersonic flight vehicles simulations for the Mach numbers 6 and higher, and numerical meshes with up to 1 billion numerical cells and with up to 128 processors.
Hypersonic flow separation in shock wave boundary layer interactions
NASA Technical Reports Server (NTRS)
Hamed, A.; Kumar, Ajay
1992-01-01
An assessment is presented for the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists mainly of two dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the subsequent changes in the flow empirical correlations for incipient separation are reviewed. The effects of the Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed. The pertinent experimental data for the separated flow characteristics in separated turbulent boundary layer shock interaction are also presented and discussed.
Hypersonic flow separation in shock wave boundary layer interactions
NASA Technical Reports Server (NTRS)
Hamed, A.; Kumar, Ajay
1992-01-01
An assessment is presented for the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists mainly of two dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the subsequent changes in the flow empirical correlations for incipient separation are reviewed. The effects of the Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed. The pertinent experimental data for the separated flow characteristics in separated turbulent boundary layer shock interaction are also presented and discussed.
Flux-split algorithms for hypersonic flows
NASA Technical Reports Server (NTRS)
Cinnella, P.; Grossman, B.
1992-01-01
This study reviews numerical techniques for the simulation of fluid flows spanning the range of reactive regimes, from local chemical equilibrium to full thermochemical non-equilibrium. In particular, characteristic-based algorithms are considered, of the flux-vector and the flux-difference type, which are becoming ever more popular due to their accurate rendition of physically complicated, shock-wave dominated flows. Consideration is given to the problems associated with modeling the thermo-chemical behavior of reactive mixtures of thermally perfect gases, and a few topics of current research are outlined.
Laser ignition of hypersonic air-hydrogen flow
NASA Astrophysics Data System (ADS)
Brieschenk, S.; Kleine, H.; O'Byrne, S.
2013-09-01
An experimental investigation of the behaviour of laser-induced ignition in a hypersonic air-hydrogen flow is presented. A compression-ramp model with port-hole injection, fuelled with hydrogen gas, is used in the study. The experiments were conducted in the T-ADFA shock tunnel using a flow condition with a specific total enthalpy of 2.5 MJ/kg and a freestream velocity of 2 km/s. This study is the first comprehensive laser spark study in a hypersonic flow and demonstrates that laser-induced ignition at the fuel-injection site can be effective in terms of hydroxyl production. A semi-empirical method to estimate the conditions in the laser-heated gas kernel is presented in the paper. This method uses blast-wave theory together with an expansion-wave model to estimate the laser-heated gas conditions. The spatially averaged conditions found with this approach are matched to enthalpy curves generated using a standard chemical equilibrium code (NASA CEA). This allows us to account for differences that are introduced due to the idealised description of the blast wave, the isentropic expansion wave as well as thermochemical effects.
STAR FORMATION IN TURBULENT MOLECULAR CLOUDS WITH COLLIDING FLOW
Matsumoto, Tomoaki; Dobashi, Kazuhito; Shimoikura, Tomomi
2015-03-10
Using self-gravitational hydrodynamical numerical simulations, we investigated the evolution of high-density turbulent molecular clouds swept by a colliding flow. The interaction of shock waves due to turbulence produces networks of thin filamentary clouds with a sub-parsec width. The colliding flow accumulates the filamentary clouds into a sheet cloud and promotes active star formation for initially high-density clouds. Clouds with a colliding flow exhibit a finer filamentary network than clouds without a colliding flow. The probability distribution functions (PDFs) for the density and column density can be fitted by lognormal functions for clouds without colliding flow. When the initial turbulence is weak, the column density PDF has a power-law wing at high column densities. The colliding flow considerably deforms the PDF, such that the PDF exhibits a double peak. The stellar mass distributions reproduced here are consistent with the classical initial mass function with a power-law index of –1.35 when the initial clouds have a high density. The distribution of stellar velocities agrees with the gas velocity distribution, which can be fitted by Gaussian functions for clouds without colliding flow. For clouds with colliding flow, the velocity dispersion of gas tends to be larger than the stellar velocity dispersion. The signatures of colliding flows and turbulence appear in channel maps reconstructed from the simulation data. Clouds without colliding flow exhibit a cloud-scale velocity shear due to the turbulence. In contrast, clouds with colliding flow show a prominent anti-correlated distribution of thin filaments between the different velocity channels, suggesting collisions between the filamentary clouds.
A CFD validation roadmap for hypersonic flows
NASA Technical Reports Server (NTRS)
Marvin, Joseph G.
1992-01-01
A roadmap for computational fluid dynamics (CFD) code validation is developed. The elements of the roadmap are consistent with air-breathing vehicle design requirements and related to the important flow path components: forebody, inlet, combustor, and nozzle. Building block and benchmark validation experiments are identified along with their test conditions and measurements. Based on an evaluation criteria, recommendations for an initial CFD validation data base are given and gaps identified where future experiments would provide the needed validation data.
Description and Flow Characterization of Hypersonic Facilities
1994-08-01
a space marching algorithm which includes the induced pressure effects of boundary-layer growth from laminar to turbu- lent flow provided the...showing the onset of air liquifaction . decrease in stream Mach number and an increase in stream static pressure. As was the case with air liquefaction...tank, which then induces pressures to locally separate the tunnel boundary layer. Sequential shadowgraph pictures (Fig. 37) 15 AEDC-TR-94-8
High enthalpy hypersonic boundary layer flow
NASA Technical Reports Server (NTRS)
Yanow, G.
1972-01-01
A theoretical and experimental study of an ionizing laminar boundary layer formed by a very high enthalpy flow (in excess of 12 eV per atom or 7000 cal/gm) with allowance for the presence of helium driver gas is described. The theoretical investigation has shown that the use of variable transport properties and their respective derivatives is very important in the solution of equilibrium boundary layer equations of high enthalpy flow. The effect of low level helium contamination on the surface heat transfer rate is minimal. The variation of ionization is much smaller in a chemically frozen boundary layer solution than in an equilibrium boundary layer calculation and consequently, the variation of the transport properties in the case of the former was not essential in the integration. The experiments have been conducted in a free piston shock tunnel, and a detailed study of its nozzle operation, including the effects of low levels of helium driver gas contamination has been made. Neither the extreme solutions of an equilibrium nor of a frozen boundary layer will adequately predict surface heat transfer rate in very high enthalpy flows.
A hybrid particle/continuum approach for nonequilibrium hypersonic flows
NASA Astrophysics Data System (ADS)
Wang, Wen-Lan
A hybrid particle-continuum computational framework is developed and presented for simulating nonequilibrium hypersonic flows, aimed to be more accurate than conventional continuum methods and faster than particle methods. The frame work consists of the direct simulation Monte Carlo-Information Preservation (DSMC-IP) method coupled with a Navier-Stokes solver. Since the DSMC-IP method provides the macroscopic information at each time step, determination of the continuum fluxes across the interface between the particle and continuum domains becomes straightforward. Buffer and reservoir calls are introduced in the continuum domain and work as an extension of the particle domain. At the end of the particle movement phase, particles in either particle or buffer cells are retained. All simulated particles in the reservoir cells are first deleted for each time, step and re-generated based on the local cell values. The microscopic velocities for the newly generated particles are initialized to the Chapman-Enskog distribution using an acceptance-rejection scheme. Continuum breakdown in a flow is defined as when the continuum solution departs from the particle solution to at least 5%. Numerical investigations show that a Knudsen-number-like parameter can best predict the continuum breakdown in the flows of interest. Numerical experiments of hypersonic flows over a simple blunted cone and a much more complex hollow cylinder/flare are conducted. The solutions for the two geometries considered from the hybrid framework are compared with experimental data and pure particle solutions. Generally speaking, it is concluded that the hybrid approach works quite well. In the blunted cone flow, numerical accuracy is improved when 10 layers of buffer cells are employed and the continuum breakdown cut-off value is set to be 0.03. In the hollow cylinder/flare hybrid simulation, the size of the separation zone near the conjunction of the cylinder and flare is improved from the initial
NASA Astrophysics Data System (ADS)
Lafon, J.-P. J.; Acker, A.; Moffat, A. F. J.
The following topics were dealt with: hypersonic flows, applications in space industry, stellar winds, wind instabilities and variability on small and large scales, disk formation, interaction of winds in different stages and with their environment, colliding winds in binary systems, dust in stellar winds.
High-resolution shock-capturing schemes for inviscid and viscous hypersonic flows
NASA Technical Reports Server (NTRS)
Yee, H. C.; Klopfer, G. H.; Montagne, J.-L.
1988-01-01
The development of robust, accurate, and efficient implicit shock-capturing schemes for multidimensional compressible Navier-Stokes equations in the hypersonic and real gas flow regimes is presently undertaken by extending a class of implicit total variation-diminishing (TVD) schemes suitable for transonic and supersonic, multidimensional Euler and Navier-Stokes equations to hypersonic computations. Numerical aspects of TVD schemes are identified which affect the convergence rate for hypersonic Mach numbers and real gas flows, but which have a negligible effect on low Mach number or perfect gas flows.
Experimental investigation of hypersonic flow induced separation over double wedges
NASA Astrophysics Data System (ADS)
Hashimoto, Tokitada
2009-09-01
Flow separation occurs over the compression corners generated by deflected control surfaces on hypersonic re-entry vehicles and in the inlet of scram jet engines. Configurations like a double wedge and double cone model are useful for studying the separated flow features. Flow fields around concave corners are relatively complicated and produce several classical viscous flow features depending on the combination of the first and second wedge or cone half apex angles. Particularly characteristic phenomena are mainly shock/boundary layer, shock/shock interaction, unsteady shear layers and non-linear shock oscillations. Although most of these basic gas dynamics characteristics are well known, it is not clear what happens at high enthalpy conditions. This paper reports a result of flow fields over a double wedge at a stagnation enthalpy of 4.8 MJ/kg. The experiment was carried out in a free piston shock tunnel at a nominal Mach number of 6.99. Schlieren and double exposure holographic interferometry were applied to visualize the flow field over the double wedge.
High-resolution shock-capturing schemes for inviscid and viscous hypersonic flows
NASA Technical Reports Server (NTRS)
Yee, H. C.; Klopfer, G. H.; Montagne, J.-L.
1988-01-01
A class of implicit Total Variation Diminishing (TVD) type algorithms suitable for transonic and supersonic multidimensional Euler and Navier-Stokes equations was extended to hypersonic computations. The improved conservative shock-capturing schemes are spatially second- and third-order, and are fully implicit. They can be first- or second-order accurate in time and are suitable for either steady or unsteady calculations. Enhancement of stability and convergence rate for hypersonic flows is discussed. With the proper choice of the temporal discretization and suitable implicit linearization, these schemes are fairly efficient and accurate for very complex two-dimensional hypersonic inviscid and viscous shock interactions. This study is complimented by a variety of steady and unsteady viscous and inviscid hypersonic blunt-body flow computations. Due to the inherent stiffness of viscous flow problems, numerical experiments indicated that the convergence rate is in general slower for viscous flows than for inviscid steady flows.
Hypersonic Flows About a 25 degree Sharp Cone
NASA Technical Reports Server (NTRS)
Moss, James N.
2001-01-01
This paper presents the results of a numerical study that examines the surface heating discrepancies observed between computed and measured values along a sharp cone. With Mach numbers of an order of 10 and the freestream length Reynolds number of an order of 10 000, the present computations have been made with the direct simulation Monte Carlo (DSMC) method by using the G2 code of Bird. The flow conditions are those specified for two experiments conducted in the Veridian 48-inch Hypersonic Shock Tunnel. Axisymmetric simulations are made since the test model was assumed to be at zero incidence. Details of the current calculations are presented, along with comparisons between the experimental data, for surface heating and pressure distributions. Results of the comparisons show major differences in measured and calculated results for heating distributions, with differences in excess of 25 percent for the two cases examined.
The 2D and 3D hypersonic flows with unstructured meshes
NASA Technical Reports Server (NTRS)
Thareja, Rajiv
1993-01-01
Viewgraphs on 2D and 3D hypersonic flows with unstructured meshes are presented. Topics covered include: mesh generation, mesh refinement, shock-shock interaction, velocity contours, mesh movement, vehicle bottom surface, and adapted meshes.
NASA Technical Reports Server (NTRS)
Limanskiy, A. V.; Timoshenko, V. I.
1986-01-01
Numerical results on the hypersonic gas flow in viscous interaction regime past sharp circular cones with thermally destructible Teflon surface are presented. Characteristics of the mutual influence between the thermochemical decomposition of the surface and the viscous interaction are revealed.
Aerothermal characteristics of bleed slot in hypersonic flows
NASA Astrophysics Data System (ADS)
Yue, LianJie; Lu, HongBo; Xu, Xiao; Chang, XinYu
2015-10-01
Two types of flow configurations with bleed in two-dimensional hypersonic flows are numerically examined to investigate their aerodynamic thermal loads and related flow structures at choked conditions. One is a turbulent boundary layer flow without shock impingement where the effects of the slot angle are discussed, and the other is shock wave boundary layer interactions where the effects of slot angle and slot location relative to shock impingement point are surveyed. A key separation is induced by bleed barrier shock on the upstream slot wall, resulting in a localized maximum heat flux at the reattachment point. For slanted slots, the dominating flow patterns are not much affected by the change in slot angle, but vary dramatically with slot location relative to the shock impingement point. Different flow structures are found in the case of normal slot, such as a flow pattern similar to typical Laval nozzle flow, the largest separation bubble which is almost independent of the shock position. Its larger detached distance results in 20% lower stagnation heat flux on the downstream slot corner, but with much wider area suffering from severe thermal loads. In spite of the complexity of the flow patterns, it is clearly revealed that the heat flux generally rises with the slot location moving downstream, and an increase in slot angle from 20° to 40° reduces 50% the heat flux peak at the reattachment point in the slot passage. The results further indicate that the bleed does not raise the heat flux around the slot for all cases except for the area around the downstream slot corner. Among all bleed configurations, the slot angle of 40° located slightly upstream of the incident shock is regarded as the best.
Aerothermal characteristics of bleed slot in hypersonic flows
NASA Astrophysics Data System (ADS)
Yue, LianJie; Lu, HongBo; Xu, Xiao; Chang, XinYu
2015-10-01
Two types of flow configurations with bleed in two-dimensional hypersonic flows are numerically examined to investigate their aerodynamic thermal loads and related flow structures at choked conditions. One is a turbulent boundary layer flow without shock impingement where the effects of the slot angle are discussed, and the other is shock wave boundary layer interactions where the effects of slot angle and slot location relative to shock impingement point are surveyed. A key separation is induced by bleed barrier shock on the upstream slot wall, resulting in a localized maximum heat flux at the reattachment point. For slanted slots, the dominating flow patterns are not much affected by the change in slot angle, but vary dramatically with slot location relative to the shock impingement point. Different flow structures are found in the case of normal slot, such as a flow pattern similar to typical Laval nozzle flow, the largest separation bubble which is almost independent of the shock position. Its larger detached distance results in 20% lower stagnation heat flux on the downstream slot corner, but with much wider area suffering from severe thermal loads. In spite of the complexity of the flow patterns, it is clearly revealed that the heat flux generally rises with the slot location moving downstream, and an increase in slot angle from 20° to 40° reduces 50% the heat flux peak at the reattachment point in the slot passage. The results further indicate that the bleed does not raise the heat flux around the slot for all cases except for the area around the downstream slot corner. Among all bleed configurations, the slot angle of 40° located slightly upstream of the incident shock is regarded as the best.
Pressure Gradient Effects on Hypersonic Cavity Flow Heating
NASA Technical Reports Server (NTRS)
Everhart, Joel L.; Alter, Stephen J.; Merski, N. Ronald; Wood, William A.; Prabhu, Ramdas K.
2007-01-01
The effect of a pressure gradient on the local heating disturbance of rectangular cavities tested at hypersonic freestream conditions has been globally assessed using the two-color phosphor thermography method. These experiments were conducted in the Langley 31-Inch Mach 10 Tunnel and were initiated in support of the Space Shuttle Return-To-Flight Program. Two blunted-nose test surface geometries were developed, including an expansion plate test surface with nearly constant negative pressure gradient and a flat plate surface with nearly zero pressure gradient. The test surface designs and flow characterizations were performed using two-dimensional laminar computational methods, while the experimental boundary layer state conditions were inferred using the measured heating distributions. Three-dimensional computational predictions of the entire model geometry were used as a check on the design process. Both open-flow and closed-flow cavities were tested on each test surface. The cavity design parameters and the test condition matrix were established using the computational predictions. Preliminary conclusions based on an analysis of only the cavity centerline data indicate that the presence of the pressure gradient did not alter the open cavity heating for laminar-entry/laminar-exit flows, but did raise the average floor heating for closed cavities. The results of these risk-reduction studies will be used to formulate a heating assessment of potential damage scenarios occurring during future Space Shuttle flights.
Pressure Gradient Effects on Hypersonic Cavity Flow Heating
NASA Technical Reports Server (NTRS)
Everhart, Joel L.; Alter, Stephen J.; Merski, N. Ronald; Wood, William A.; Prabhu, Ramadas K.
2006-01-01
The effect of a pressure gradient on the local heating disturbance of rectangular cavities tested at hypersonic freestream conditions has been globally assessed using the two-color phosphor thermography method. These experiments were conducted in the Langley 31-Inch Mach 10 Tunnel and were initiated in support of the Space Shuttle Return-To-Flight Program. Two blunted-nose test surface geometries were developed, including an expansion plate test surface with nearly constant negative pressure gradient and a flat plate surface with nearly zero pressure gradient. The test surface designs and flow characterizations were performed using two-dimensional laminar computational methods, while the experimental boundary layer state conditions were inferred using the measured heating distributions. Three-dimensional computational predictions of the entire model geometry were used as a check on the design process. Both open-flow and closed-flow cavities were tested on each test surface. The cavity design parameters and the test condition matrix were established using the computational predictions. Preliminary conclusions based on an analysis of only the cavity centerline data indicate that the presence of the pressure gradient did not alter the open cavity heating for laminar-entry/laminar-exit flows, but did raise the average floor heating for closed cavities. The results of these risk-reduction studies will be used to formulate a heating assessment of potential damage scenarios occurring during future Space Shuttle flights.
Development and validation of CNS (compressible Navier-Stokes) for hypersonic external flows
NASA Technical Reports Server (NTRS)
Flores, Jolen; Chow, Chuen-Yen; Ryan, James S.
1989-01-01
CNS, a new computational fluid dynamics procedure, has been developed to aid in hypersonic vehicle design. The code can be used to model the entire external flow around hypersonic vehicle shapes, from the captured shock at the nose to the beginning of the wake. Unlike space-marching codes, the technique allows axially separated flow regions to be modeled. Validation trials using sphere-cone data reveal good solution accuracy for the surface pressure and flowfield temperature.
Characterization of a hot-film probe for hypersonic flow
NASA Technical Reports Server (NTRS)
Sheplak, M.; Spina, E.; Mcginley, C.
1995-01-01
The critical issues concerning the application of constant-temperature hot-film anemometry to hypersonic flow are reviewed and extended. Mass-flux static calibrations were conducted in a Mach 10 helium flow, while mass-flux and total-temperature static calibrations were made in a Mach 6 air flow. In addition, comparative hot-film/hot-wire turbulence measurements were made in a Mach 11 helium boundary layer to provide insight into the dynamic response of the hot film. The measurements indicate that substrate conduction 'losses' dominate the static response of the hot-film probe, thus resulting in poor sensitivity to mass-flux and total temperature. Furthermore, it has been found that it is not possible to isolate mass-flux fluctuations at high overheat ratios for the current hot-film design. Thus, the sapphire-substrate hot-film anemometer is a robust, high-bandwidth instrument limited to qualitative transition and turbulence measurements. Finally, the extension of this technique to providing quantitative information is dependent upon the development of lower thermal-conductivity substrate materials.
Computations of Axisymmetric Flows in Hypersonic Shock Tubes
NASA Technical Reports Server (NTRS)
Sharma, Surendra P.; Wilson, Gregory J.
1995-01-01
A time-accurate two-dimensional fluid code is used to compute test times in shock tubes operated at supersonic speeds. Unlike previous studies, this investigation resolves the finer temporal details of the shock-tube flow by making use of modern supercomputers and state-of-the-art computational fluid dynamic solution techniques. The code, besides solving the time-dependent fluid equations, also accounts for the finite rate chemistry in the hypersonic environment. The flowfield solutions are used to estimate relevant shock-tube parameters for laminar flow, such as test times, and to predict density and velocity profiles. Boundary-layer parameters such as bar-delta(sub u), bar-delta(sup *), and bar-tau(sub w), and test time parameters such as bar-tau and particle time of flight t(sub f), are computed and compared with those evaluated by using Mirels' correlations. This article then discusses in detail the effects of flow nonuniformities on particle time-of-flight behind the normal shock and, consequently, on the interpretation of shock-tube data. This article concludes that for accurate interpretation of shock-tube data, a detailed analysis of flowfield parameters, using a computer code such as used in this study, must be performed.
NASA Astrophysics Data System (ADS)
Gestrin, S. G.; Gorbatenko, B. B.; Mezhonnova, A. S.
2016-05-01
It is shown that the resonance effect of a magnetohydrodynamic hypersonic shear flow on an elastic plate placed in it causes the development of wind instability. Plate bending oscillations propagating along the flow are stabilized in the hypersonic flow regime, whereas waves running at an angle to the flow remain unstable. Expression derived for the instability increment allows conclusions about the effect of the magnetic field on the interaction of waves with the flow to be drawn as well as about the feasibility of its suppression in an unstable flow regime.
High-resolution shock-capturing schemes for inviscid and viscous hypersonic flows
NASA Technical Reports Server (NTRS)
Yee, H. C.; Klopfer, G. H.; Montagne, J.-L.
1990-01-01
Hypersonic computations are presently conducted with an extension of a class of high-resolution implicit TVD algorithms suited to transonic multidimensional Euler and Navier-Stokes equations. These conservative shock-capturing schemes, which are spatially second- and third-order, may be first- and second-order accurate in time and suitable for either steady or unsteady calculations. Attention is given to the enhancement of hypersonic flows' convergence rate and stability; accuracy and efficiency is achieved by these means for very complex two-dimensional hypersonic viscous and inviscid shock interactions.
Direct simulation of hypersonic flows over blunt slender bodies
NASA Technical Reports Server (NTRS)
Moss, J. N.; Cuda, V., Jr.
1986-01-01
Results of a numerical study of low-density hypersonic flow about cylindrically blunted wedges and spherically blunted cones with body half angles of 0, 5, and 10 deg are presented. Most of the transitional flow regime encountered during entry between the free molecule and continuum regimes is simulated for a reentry velocity of 7.5 km/s by including freestream conditions of 70 to 100 km. The bodies are at zero angle of incidence and have diffuse and finite catalytic surfaces. Translational, thermodynamic, and chemical nonequilibrium effects are considered in the numerical simulation by utilizing the direct simulation Monte Carlo (DSMC) method. The numerical simulations show that noncontinuum effects such as surface temperature jump, and velocity slip are evident for all cases considered. The onset of chemical dissociation occurs at a simulated altitude of 96 km for the two-dimensional configurations. Comparisons between the DSMC and continuum viscous shock-layer calculations highlight the significant difference in flowfield structure predicted by the two methods.
Numerical simulation of supersonic and hypersonic inlet flow fields
NASA Technical Reports Server (NTRS)
Mcrae, D. Scott; Kontinos, Dean A.
1995-01-01
This report summarizes the research performed by North Carolina State University and NASA Ames Research Center under Cooperative Agreement NCA2-719, 'Numerical Simulation of Supersonic and Hypersonic Inlet Flow Fields". Four distinct rotated upwind schemes were developed and investigated to determine accuracy and practicality. The scheme found to have the best combination of attributes, including reduction to grid alignment with no rotation, was the cell centered non-orthogonal (CCNO) scheme. In 2D, the CCNO scheme improved rotation when flux interpolation was extended to second order. In 3D, improvements were less dramatic in all cases, with second order flux interpolation showing the least improvement over grid aligned upwinding. The reduction in improvement is attributed to uncertainty in determining optimum rotation angle and difficulty in performing accurate and efficient interpolation of the angle in 3D. The CCNO rotational technique will prove very useful for increasing accuracy when second order interpolation is not appropriate and will materially improve inlet flow solutions.
Reattachment heating upstream of short compression ramps in hypersonic flow
NASA Astrophysics Data System (ADS)
Estruch-Samper, David
2016-05-01
Hypersonic shock-wave/boundary-layer interactions with separation induce unsteady thermal loads of particularly high intensity in flow reattachment regions. Building on earlier semi-empirical correlations, the maximum heat transfer rates upstream of short compression ramp obstacles of angles 15° ⩽ θ ⩽ 135° are here discretised based on time-dependent experimental measurements to develop insight into their transient nature (Me = 8.2-12.3, Re_h= 0.17× 105-0.47× 105). Interactions with an incoming laminar boundary layer experience transition at separation, with heat transfer oscillating between laminar and turbulent levels exceeding slightly those in fully turbulent interactions. Peak heat transfer rates are strongly influenced by the stagnation of the flow upon reattachment close ahead of obstacles and increase with ramp angle all the way up to θ =135°, whereby rates well over two orders of magnitude above the undisturbed laminar levels are intermittently measured (q'_max>10^2q_{u,L}). Bearing in mind the varying degrees of strength in the competing effect between the inviscid and viscous terms—namely the square of the hypersonic similarity parameter (Mθ )^2 for strong interactions and the viscous interaction parameter bar{χ } (primarily a function of Re and M)—the two physical factors that appear to most globally encompass the effects of peak heating for blunt ramps (θ ⩾ 45°) are deflection angle and stagnation heat transfer, so that this may be fundamentally expressed as q'_max∝ {q_{o,2D}} θ ^2 with further parameters in turn influencing the interaction to a lesser extent. The dominant effect of deflection angle is restricted to short obstacle heights, where the rapid expansion at the top edge of the obstacle influences the relaxation region just downstream of reattachment and leads to an upstream displacement of the separation front. The extreme heating rates result from the strengthening of the reattaching shear layer with the increase in
Computation of hypersonic flows with finite rate condensation and evaporation of water
NASA Technical Reports Server (NTRS)
Perrell, Eric R.; Candler, Graham V.; Erickson, Wayne D.; Wieting, Alan R.
1993-01-01
A computer program for modelling 2D hypersonic flows of gases containing water vapor and liquid water droplets is presented. The effects of interphase mass, momentum and energy transfer are studied. Computations are compared with existing quasi-1D calculations on the nozzle of the NASA Langley Eight Foot High Temperature Tunnel, a hypersonic wind tunnel driven by combustion of natural gas in oxygen enriched air.
Computation of hypersonic flows with finite rate condensation and evaporation of water
NASA Technical Reports Server (NTRS)
Perrell, Eric R.; Candler, Graham V.; Erickson, Wayne D.; Wieting, Alan R.
1993-01-01
A computer program for modelling 2D hypersonic flows of gases containing water vapor and liquid water droplets is presented. The effects of interphase mass, momentum and energy transfer are studied. Computations are compared with existing quasi-1D calculations on the nozzle of the NASA Langley Eight Foot High Temperature Tunnel, a hypersonic wind tunnel driven by combustion of natural gas in oxygen enriched air.
NASA Technical Reports Server (NTRS)
Dogra, V. K.; Moss, J. N.; Wilmoth, R. G.; Price, J. M.
1992-01-01
Results of a numerical study concerning flow past a 70-deg blunted cone in hypersonic low-density flow environments are presented using the direct simulation Monte-Carlo method. The flow conditions simulated are those that can be obtained in existing low-density hypersonic wind tunnels. Results indicate that a stable vortex forms in the near wake at and below a freestream Knudsen number (based on cone diameter) of 0.01 and the size of the vortex increases with decreasing Knudsen number. The base region of the flow remains in thermal nonequilibrium for all cases considered herein.
Adiabatic Shock Capturing in Perfect Gas Hypersonic Flows
NASA Technical Reports Server (NTRS)
Kirk, Benjamin S.
2009-01-01
This paper considers the streamline-upwind Petrov/Galerkin (SUPG) method applied to the compressible Euler and Navier-Stokes equations in conservation-variable form. The spatial discretization, including a modified approach for interpolating the inviscid flux terms in the SUPG finite element formulation, is briefly reviewed. Of particular interest is the behavior of the shock capturing operator, which is required to regularize the scheme in the presence of strong, shock-induced gradients. A standard shock capturing operator which has been widely used in previous studies by several authors is presented and discussed. Specific modifications are then made to this standard operator which are designed to produce a more physically consistent discretization in the presence of strong shock waves. The actual implementation of the term in a finite dimensional approximation is also discussed. The behavior of the standard and modified scheme is then compared for several supersonic/hypersonic flows. The modified shock capturing operator is found to preserve enthalpy in the inviscid portion of the flowfield substantially better than the standard operator.
Rotational and vibrational nonequilibrium effects in rarefied, hypersonic flow
NASA Technical Reports Server (NTRS)
Boyd, Iain D.
1989-01-01
Results are reported for an investigation into the methods by which energy transfer is calculated in the Direct Simulation Monte Carlo method. Description is made of a recently developed energy exchange model that deals with the translational and rotational modes. A new model for simulating the transfer of energy between the translational and vibrational modes is also explained. This model allows the vibrational relaxation time to follow the temperature dependence predicted by the Landau-Teller theory at moderate temperatures. For temperatures in excess of about 8000K the vibrational model is extended to include an empirical result for the relaxation time. The effect of introducing these temperature dependent collision numbers into the DSMC technique is assessed by making calculations representative of the stagnation streamline of a hypersonic space vehicle. Both thermal and chemical nonequilibrium effects are included while the flow conditions have been chosen such that ionization and radiation may be neglected. The introduction of these new models is found to significantly affect the degree of thermal nonequilibrium observed in the flowfield. Larger, and more widely ranging, differences in the results obtained with the different energy exchange probabilities are found when a significant amount of internal energy is included in the calculation of chemical nonequilibrium.
Blunt Body Aerodynamics for Hypersonic Low Density Flows
NASA Technical Reports Server (NTRS)
Moss, James N.; Glass, Christopher E.; Greene, Francis A.
2006-01-01
Numerical simulations are performed for the Apollo capsule from the hypersonic rarefied to the continuum regimes. The focus is on flow conditions similar to those experienced by the Apollo 6 Command Module during the high altitude portion of its reentry. The present focus is to highlight some of the current activities that serve as a precursor for computational tool assessments that will be used to support the development of aerodynamic data bases for future capsule flight environments, particularly those for the Crew Exploration Vehicle (CEV). Results for aerodynamic forces and moments are presented that demonstrate their sensitivity to rarefaction; that is, free molecular to continuum conditions. Also, aerodynamic data are presented that shows their sensitivity to a range of reentry velocities, encompassing conditions that include reentry from low Earth orbit, lunar return, and Mars return velocities (7.7 to 15 km/s). The rarefied results obtained with direct simulation Monte Carlo (DSMC) codes are anchored in the continuum regime with data from Navier-Stokes simulations.
Adiabatic Shock Capturing in Perfect Gas Hypersonic Flows
NASA Technical Reports Server (NTRS)
Kirk, Benjamin S.
2009-01-01
This paper considers the streamline-upwind Petrov/Galerkin (SUPG) method applied to the compressible Euler and Navier-Stokes equations in conservation-variable form. The spatial discretization, including a modified approach for interpolating the inviscid flux terms in the SUPG finite element formulation, is briefly reviewed. Of particular interest is the behavior of the shock capturing operator, which is required to regularize the scheme in the presence of strong, shock-induced gradients. A standard shock capturing operator which has been widely used in previous studies by several authors is presented and discussed. Specific modifications are then made to this standard operator which are designed to produce a more physically consistent discretization in the presence of strong shock waves. The actual implementation of the term in a finite dimensional approximation is also discussed. The behavior of the standard and modified scheme is then compared for several supersonic/hypersonic flows. The modified shock capturing operator is found to preserve enthalpy in the inviscid portion of the flowfield substantially better than the standard operator.
Application of Pressure Sensitive Paint in Hypersonic Flows
NASA Technical Reports Server (NTRS)
Jules, Kenol; Carbonaro, Mario; Zemsch, Stephan
1995-01-01
It is well known in the aerodynamic field that pressure distribution measurement over the surface of an aircraft model is a problem in experimental aerodynamics. For one thing, a continuous pressure map can not be obtained with the current experimental methods since they are discrete. Therefore, interpolation or CFD methods must be used for a more complete picture of the phenomenon under study. For this study, a new technique was investigated which would provide a continuous pressure distribution over the surface under consideration. The new method is pressure sensitive paint. When pressure sensitive paint is applied to an aerodynamic surface and placed in an operating wind-tunnel under appropriate lighting, the molecules luminesce as a function of the local pressure of oxygen over the surface of interest during aerodynamic flow. The resulting image will be brightest in the areas of low pressure (low oxygen concentration), and less intense in the areas of high pressure (where oxygen is most abundant on the surface). The objective of this investigation was to use pressure sensitive paint samples from McDonnell Douglas (MDD) for calibration purpose in order to assess the response of the paint under appropriate lighting and to use the samples over a flat plate/conical fin mounted at 75 degrees from the center of the plate in order to study the shock/boundary layer interaction at Mach 6 in the Von Karman wind-tunnel. From the result obtained it was concluded that temperature significantly affects the response of the paint and should be given the uppermost attention in the case of hypersonic flows. Also, it was found that past a certain temperature threshold, the paint intensity degradation became irreversible. The comparison between the pressure tap measurement and the pressure sensitive paint showed the right trend. However, there exists a shift when it comes to the actual value. Therefore, further investigation is under way to find the cause of the shift.
The computation of thermo-chemical nonequilibrium hypersonic flows
NASA Technical Reports Server (NTRS)
Candler, Graham
1989-01-01
Several conceptual designs for vehicles that would fly in the atmosphere at hypersonic speeds have been developed recently. For the proposed flight conditions the air in the shock layer that envelops the body is at a sufficiently high temperature to cause chemical reaction, vibrational excitation, and ionization. However, these processes occur at finite rates which, when coupled with large convection speeds, cause the gas to be removed from thermo-chemical equilibrium. This non-ideal behavior affects the aerothermal loading on the vehicle and has ramifications in its design. A numerical method to solve the equations that describe these types of flows in 2-D was developed. The state of the gas is represented with seven chemical species, a separate vibrational temperature for each diatomic species, an electron translational temperature, and a mass-average translational-rotational temperature for the heavy particles. The equations for this gas model are solved numerically in a fully coupled fashion using an implicit finite volume time-marching technique. Gauss-Seidel line-relaxation is used to reduce the cost of the solution and flux-dependent differencing is employed to maintain stability. The numerical method was tested against several experiments. The calculated bow shock wave detachment on a sphere and two cones was compared to those measured in ground testing facilities. The computed peak electron number density on a sphere-cone was compared to that measured in a flight test. In each case the results from the numerical method were in excellent agreement with experiment. The technique was used to predict the aerothermal loads on an Aeroassisted Orbital Transfer Vehicle including radiative heating. These results indicate that the current physical model of high temperature air is appropriate and that the numerical algorithm is capable of treating this class of flows.
Arc-heated gas flow experiments for hypersonic propulsion applications
NASA Astrophysics Data System (ADS)
Roseberry, Christopher Matthew
Although hydrogen is an attractive fuel for a hypersonic air-breathing vehicle in terms of reaction rate, flame temperature, and energy content per unit mass, the substantial tank volume required to store hydrogen imposes a drag penalty to performance that tends to offset these advantages. An alternative approach is to carry a hydrocarbon fuel and convert it on-board into a hydrogen-rich gas mixture to be injected into the engine combustors. To investigate this approach, the UTA Arc-Heated Wind Tunnel facility was modified to run on methane rather than the normally used nitrogen. Previously, this facility was extensively developed for the purpose of eventually performing experiments simulating scramjet engine flow along a single expansion ramp nozzle (SERN) in addition to more generalized applications. This formidable development process, which involved modifications to every existing subsystem along with the incorporation of new subsystems, is described in detail. Fortunately, only a minor plumbing reconfiguration was required to prepare the facility for the fuel reformation research. After a failure of the arc heater power supply, a 5.6 kW plasma-cutting torch was modified in order to continue the arc pyrolysis experiments. The outlet gas flow from the plasma torch was sampled and subsequently analyzed using gas chromatography. The experimental apparatus converted the methane feedstock almost completely into carbon, hydrogen and acetylene. A high yield of hydrogen, consisting of a product mole fraction of roughly 0.7, was consistently obtained. Unfortunately, the energy consumption of the apparatus was too excessive to be feasible for a flight vehicle. However, other researchers have pyrolyzed hydrocarbons using electric arcs with much less power input per unit mass.
N-S/DSMC hybrid simulation of hypersonic flow over blunt body including wakes
NASA Astrophysics Data System (ADS)
Li, Zhonghua; Li, Zhihui; Li, Haiyan; Yang, Yanguang; Jiang, Xinyu
2014-12-01
A hybrid N-S/DSMC method is presented and applied to solve the three-dimensional hypersonic transitional flows by employing the MPC (modular Particle-Continuum) technique based on the N-S and the DSMC method. A sub-relax technique is adopted to deal with information transfer between the N-S and the DSMC. The hypersonic flows over a 70-deg spherically blunted cone under different Kn numbers are simulated using the CFD, DSMC and hybrid N-S/DSMC method. The present computations are found in good agreement with DSMC and experimental results. The present method provides an efficient way to predict the hypersonic aerodynamics in near-continuum transitional flow regime.
NASA Technical Reports Server (NTRS)
Kapoor, Kamlesh; Anderson, Bernhard H.; Shaw, Robert J.
1994-01-01
A two-dimensional computational code, PRLUS2D, which was developed for the reactive propulsive flows of ramjets and scramjets, was validated for two-dimensional shock-wave/turbulent-boundary-layer interactions. The problem of compression corners at supersonic speeds was solved using the RPLUS2D code. To validate the RPLUS2D code for hypersonic speeds, it was applied to a realistic hypersonic inlet geometry. Both the Baldwin-Lomax and the Chien two-equation turbulence models were used. Computational results showed that the RPLUS2D code compared very well with experimentally obtained data for supersonic compression corner flows, except in the case of large separated flows resulting from the interactions between the shock wave and turbulent boundary layer. The computational results compared well with the experiment results in a hypersonic NASA P8 inlet case, with the Chien two-equation turbulence model performing better than the Baldwin-Lomax model.
Implementation of a hypersonic rarefied flow particle simulation on the Connection Machine
NASA Technical Reports Server (NTRS)
Dagum, Leonardo
1988-01-01
A very efficient direct particle simulation algorithm for hypersonic rarefied flows is presented and its implmentation on a Connection Machine is described. The implementation simulates ideal diatomic Maxwell molecules with three translational and two rotational degrees of freedom. Results for a 2-D simulation of supersonic flow over a 30 deg wedge are presented and used for validation.
Flow-Tagging Velocimetry for Hypersonic Flows Using Fluorescence of Nitric Oxide
NASA Technical Reports Server (NTRS)
Danehy, Paul M.; OByrne, Sean; Houwing, A. Frank P.; Fox, Jodie S.; Smith, Daniel R.
2003-01-01
We demonstrate a new variation of molecular-tagging velocimetry for hypersonic flows based on laser-induced fluorescence. A thin line of nitric-oxide molecules is excited with a laser beam and then, after a time delay, a fluorescence image of the displaced line is acquired. One component of velocity is determined from the time of flight. This method is applied to measure the velocity profile in a Mach 8.5 laminar, hypersonic boundary layer in the Australian National University s T2 free-piston shock tunnel. The single-shot velocity measurement uncertainty in the freestream was found to be 3.5%, based on 90% confidence. The method is also demonstrated in the separated flow region forward of a blunt fin attached to a flat plate in a Mach 7.4 flow produced by the Australian National University s T3 free-piston shock tunnel. The measurement uncertainty in the blunt fin experiment is approximately 30%, owing mainly to low fluorescence intensities, which could be improved significantly in future experiments. This velocimetry method is applicable to very high-speed flows that have low collisional quenching of the fluorescing species. It is particularly convenient in facilities where planar laser-induced fluorescence is already being performed.
Pressure and force measurements on models set in hypersonic flows: A review
NASA Technical Reports Server (NTRS)
Miller, Charles G.
1993-01-01
A review of measurement techniques used to obtain aerodynamic forces and moments and surface/flow field pressures for models tested in impulse hypersonic-hypervelocity facilities and in conventional-type hypersonic wind tunnels is presented. Although force and moment measurement techniques presently used in hypersonic wind tunnels are relatively unchanged from the 1960's and 1970's, significant advances have recently been made for impulse facilities. For both hypersonic wind tunnels and impulse facilities, the state-of-the-art has advanced via refinements, improved test techniques, and advances in semiconductor technology, data acquisition systems, and computers. The introduction of electronically scanned pressure systems over a decade ago 'revolutionized' pressure measurements in hypersonic wind tunnels and a second 'revolution' is impending with the development and application of optical, two-dimensional, global pressure measurement techniques. The development and continued refinement of miniature piezoresistive transducers has provided the capability to perform detailed surface pressure measurements on relatively small, complex models in impulse facilities; these transducers also provided the capability for intrusive flow field pressure measurements with miniature survey rakes.
DSMC simulation of hypersonic flows using an improved SBT-TAS technique
NASA Astrophysics Data System (ADS)
Goshayeshi, Bijan; Roohi, Ehsan; Stefanov, Stefan
2015-12-01
The current paper examines a new DSMC approach to hypersonic flow simulation consisting of a combination between the Simplified Bernoulli Trials (SBT) collision algorithm and the transient adaptive subcell (TAS) selection procedure. The SBT collision algorithm has already been introduced as a scheme that provides accurate results with a quite small number of particles per cells and its combination with the transient adaptive subcell (TAS) technique will enable SBT to have coarser grid sizes as well. In the current research, the no-time-counter (NTC) collision algorithm and nearest neighbor (NN) pair selection procedure of Bird DS2V code are substituted by the SBT-TAS and comparisons between the new algorithm and NTC-NN are made considering appropriate test cases including hypersonic cylinder flow and axisymmetric biconic flow. Hypersonic cylinder flow is a well-known benchmark problem with a wide collision frequency range while the biconic flow exhibits laminar shock/shock and shock/boundary-layer interactions. Improvements implemented in the SBT-TAS technique, including subcell volume estimation, surface properties filter, and time controller, are discussed in detail. The simulations of these hypersonic test cases demonstrated that from the viewpoint of consumed sample-size, SBT-TAS is an efficient collision technique.
Atomistic Simulation of Non-Equilibrium Phenomena in Hypersonic Flows
NASA Astrophysics Data System (ADS)
Norman, Paul Erik
The goal of this work is to model the heterogeneous recombination of atomic oxygen on silica surfaces, which is of interest for accurately predicting the heating on vehicles traveling at hypersonic speeds. This is accomplished by creating a finite rate catalytic model, which describes recombination with a set of elementary gas-surface reactions. Fundamental to a description of surface catalytic reactions are the in situ chemical structures on the surface where recombination can occur. Using molecular dynamics simulations with the Reax GSISiO potential, we find that the chemical sites active in direct gas-phase reactions on silica surfaces consist of a small number of specific structures (or defects). The existence of these defects on real silica surfaces is supported by experimental results and the structure and energetics of these defects have been verified with quantum chemical calculations. The reactions in the finite rate catalytic model are based on the interaction of molecular and atomic oxygen with these defects. Trajectory calculations are used to find the parameters in the forward rate equations, while a combination of detailed balance and transition state theory are used to find the parameters in the reverse rate equations. The rate model predicts that the oxygen recombination coefficient is relatively constant at T (300-1000 K), in agreement with experimental results. At T > 1000 K the rate model predicts a drop off in the oxygen recombination coefficient, in disagreement with experimental results, which predict that the oxygen recombination coefficient increases with temperature. A discussion of the possible reasons for this disagreement, including non-adiabatic collision dynamics, variable surface site concentrations, and additional recombination mechanisms is presented. This thesis also describes atomistic simulations with Classical Trajectory Calculation Direction Simulation Monte Carlo (CTC-DSMC), a particle based method for modeling non
Aero-Heating of Shallow Cavities in Hypersonic Freestream Flow
NASA Technical Reports Server (NTRS)
Everhart, Joel L.; Berger, Karen T.; Merski, N. R., Jr.; Woods, William A.; Hollingsworth, Kevin E.; Hyatt, Andrew; Prabhu, Ramadas K.
2010-01-01
The purpose of these experiments and analysis was to augment the heating database and tools used for assessment of impact-induced shallow-cavity damage to the thermal protection system of the Space Shuttle Orbiter. The effect of length and depth on the local heating disturbance of rectangular cavities tested at hypersonic freestream conditions has been globally assessed using the two-color phosphor thermography method. These rapid-response experiments were conducted in the Langley 31-Inch Mach 10 Tunnel and were initiated immediately prior to the launch of STS-114, the initial flight in the Space Shuttle Return-To-Flight Program, and continued during the first week of the mission. Previously-designed and numerically-characterized blunted-nose baseline flat plates were used as the test surfaces. Three-dimensional computational predictions of the entire model geometry were used as a check on the design process and the two-dimensional flow assumptions used for the data analysis. The experimental boundary layer state conditions were inferred using the measured heating distributions on a no-cavity test article. Two test plates were developed, each containing 4 equally-spaced spanwise-distributed cavities. The first test plate contained cavities with a constant length-to-depth ratio of 8 with design point depth-to-boundary-layer-thickness ratios of 0.1, 0.2, 0.35, and 0.5. The second test plate contained cavities with a constant design point depth-to-boundary-layer-thickness ratio of 0.35 with length-to-depth ratios of 8, 12, 16, and 20. Cavity design parameters and the test condition matrix were established using the computational predictions. Preliminary results indicate that the floor-averaged Bump Factor (local heating rate nondimensionalized by upstream reference) at the tested conditions is approximately 0.3 with a standard deviation of 0.04 for laminar-in/laminar-out conditions when the cavity length-to-boundary-layer thickness is between 2.5 and 10 and for
A new computational method for reacting hypersonic flows
NASA Astrophysics Data System (ADS)
Niculescu, M. L.; Cojocaru, M. G.; Pricop, M. V.; Fadgyas, M. C.; Pepelea, D.; Stoican, M. G.
2017-07-01
Hypersonic gas dynamics computations are challenging due to the difficulties to have reliable and robust chemistry models that are usually added to Navier-Stokes equations. From the numerical point of view, it is very difficult to integrate together Navier-Stokes equations and chemistry model equations because these partial differential equations have different specific time scales. For these reasons, almost all known finite volume methods fail shortly to solve this second order partial differential system. Unfortunately, the heating of Earth reentry vehicles such as space shuttles and capsules is very close linked to endothermic chemical reactions. A better prediction of wall heat flux leads to smaller safety coefficient for thermal shield of space reentry vehicle; therefore, the size of thermal shield decreases and the payload increases. For these reasons, the present paper proposes a new computational method based on chemical equilibrium, which gives accurate prediction of hypersonic heating in order to support the Earth reentry capsule design.
Assessment of predictive capabilities for aerodynamic heating in hypersonic flow
NASA Astrophysics Data System (ADS)
Knight, Doyle; Chazot, Olivier; Austin, Joanna; Badr, Mohammad Ali; Candler, Graham; Celik, Bayram; Rosa, Donato de; Donelli, Raffaele; Komives, Jeffrey; Lani, Andrea; Levin, Deborah; Nompelis, Ioannis; Panesi, Marco; Pezzella, Giuseppe; Reimann, Bodo; Tumuklu, Ozgur; Yuceil, Kemal
2017-04-01
The capability for CFD prediction of hypersonic shock wave laminar boundary layer interaction was assessed for a double wedge model at Mach 7.1 in air and nitrogen at 2.1 MJ/kg and 8 MJ/kg. Simulations were performed by seven research organizations encompassing both Navier-Stokes and Direct Simulation Monte Carlo (DSMC) methods as part of the NATO STO AVT Task Group 205 activity. Comparison of the CFD simulations with experimental heat transfer and schlieren visualization suggest the need for accurate modeling of the tunnel startup process in short-duration hypersonic test facilities, and the importance of fully 3-D simulations of nominally 2-D (i.e., non-axisymmmetric) experimental geometries.
Numerical simulations of heat and mass transfer at ablating surface in hypersonic flow
NASA Astrophysics Data System (ADS)
Bocharov, A. N.; Golovin, N. N.; Petrovskiy, V. P.; Teplyakov, I. O.
2015-11-01
The numerical technique was developed to solve heat and mass transfer problem in 3D hypersonic flow taking into account destruction of thermal protection system. Described technique was applied for calculation of heat and mass transfer in sphere-cone shaped body. The data on temperature, heat flux and mass flux were obtained.
Hypersonic Flow over a Cylinder with a Nanosecond Pulse Electrical Discharge
2014-03-01
Hypersonic Flow over a Cylinder with a Nanosecond Pulse Electrical Discharge Nicholas J. Bisek∗ and Jonathan Poggie† U.S. Air Force Research... pulsed at nanosecond time scales and it rapidly added thermal energy to the flow, creating a shock wave that traveled away from the pulse source. As the...control studies [5–7] using LAFPA actuators in atmospheric pressure jet flows for Mach 0.9 to Mach 2 demonstrated significant localized heating and
Multi Laser Pulse Investigation of the DEAS Concept in Hypersonic Flow
Minucci, M.A.S.; Toro, P.G.P.; Oliveira, A.C.; Chanes, J.B. Jr.; Ramos, A.G.; Nagamatsu, H.T.; Myrabo, L.N.
2004-03-30
The present paper presents recent experimental results on the Laser-Supported Directed Energy 'Air Spike' - DEAS in hypersonic flow achieved by the Laboratory of Aerothermodynamics and Hypersonics - LAH, Brazil. Two CO2 TEA lasers, sharing the same optical cavity, have been used in conjunction with the IEAv 0.3m Hypersonic Shock Tunnel - HST to demonstrate the Laser-Supported DEAS concept. A single and double laser pulse, generated during the tunnel useful test time, were focused through a NaCl lens upstream of a Double Apollo Disc model fitted with seven piezoelectric pressure transducers and six platinum thin film heat transfer gauges. The objective being to corroborate previous results as well as to obtain additional pressure and heat flux distributions information when two laser pulses are used.
NASA Technical Reports Server (NTRS)
Scott, Carl D.
1992-01-01
The meaning of catalysis and its relation to aerodynamic heating in nonequilibrium hypersonic flows are discussed. The species equations are described and boundary conditions for them are derived for a multicomponent gas and for a binary gas. Slip effects are included for application of continuum methods to low-density flows. Measurement techniques for determining catalytic wall recombination rates are discussed. Among them are experiments carried out in arc jets as well as flow reactors. Diagnostic methods for determining the atom or molecule concentrations in the flow are included. Results are given for a number of materials of interest to the aerospace community, including glassy coatings such as the RCG coating of the Space Shuttle and for high temperature refractory metals such as coated niobium. Methods of calculating the heat flux to space vehicles in nonequilibrium flows are described. These methods are applied to the Space Shuttle, the planned Aeroassist Flight Experiment, and a hypersonic slender vehicle such as a transatmospheric vehicle.
Molecular cloud formation in high-shear, magnetized colliding flows
NASA Astrophysics Data System (ADS)
Fogerty, E.; Frank, A.; Heitsch, F.; Carroll-Nellenback, J.; Haig, C.; Adams, M.
2016-08-01
The colliding flows (CF) model is a well-supported mechanism for generating molecular clouds. However, to-date most CF simulations have focused on the formation of clouds in the normal-shock layer between head-on colliding flows. We performed simulations of magnetized colliding flows that instead meet at an oblique-shock layer. Oblique shocks generate shear in the post-shock environment, and this shear creates inhospitable environments for star formation. As the degree of shear increases (i.e. the obliquity of the shock increases), we find that it takes longer for sink particles to form, they form in lower numbers, and they tend to be less massive. With regard to magnetic fields, we find that even a weak field stalls gravitational collapse within forming clouds. Additionally, an initially oblique collision interface tends to reorient over time in the presence of a magnetic field, so that it becomes normal to the oncoming flows. This was demonstrated by our most oblique shock interface, which became fully normal by the end of the simulation.
NASA Astrophysics Data System (ADS)
Bender, Jason D.
Understanding hypersonic aerodynamics is important for the design of next-generation aerospace vehicles for space exploration, national security, and other applications. Ground-level experimental studies of hypersonic flows are difficult and expensive; thus, computational science plays a crucial role in this field. Computational fluid dynamics (CFD) simulations of extremely high-speed flows require models of chemical and thermal nonequilibrium processes, such as dissociation of diatomic molecules and vibrational energy relaxation. Current models are outdated and inadequate for advanced applications. We describe a multiscale computational study of gas-phase thermochemical processes in hypersonic flows, starting at the atomic scale and building systematically up to the continuum scale. The project was part of a larger effort centered on collaborations between aerospace scientists and computational chemists. We discuss the construction of potential energy surfaces for the N4, N2O2, and O4 systems, focusing especially on the multi-dimensional fitting problem. A new local fitting method named L-IMLS-G2 is presented and compared with a global fitting method. Then, we describe the theory of the quasiclassical trajectory (QCT) approach for modeling molecular collisions. We explain how we implemented the approach in a new parallel code for high-performance computing platforms. Results from billions of QCT simulations of high-energy N2 + N2, N2 + N, and N2 + O2 collisions are reported and analyzed. Reaction rate constants are calculated and sets of reactive trajectories are characterized at both thermal equilibrium and nonequilibrium conditions. The data shed light on fundamental mechanisms of dissociation and exchange reactions -- and their coupling to internal energy transfer processes -- in thermal environments typical of hypersonic flows. We discuss how the outcomes of this investigation and other related studies lay a rigorous foundation for new macroscopic models for
Use of arc-jet hypersonic blunted wedge flows for evaluating performance of Orbiter TPS
NASA Technical Reports Server (NTRS)
Rochelle, W. C.; Battley, H. H.; Gallegos, J. J.
1979-01-01
Arc-jet tests at NASA/JSC have been conducted recently to evaluate the performance of the Orbiter Thermal Protection System (TPS) on three critical areas of the side and top of the Orbiter fuselage: (1) cargo bay door, (2) crew access door, and (3) LRSI/FRSI joint regions. Test articles corresponding to these three areas on the Orbiter were mounted in an arc-jet test chamber in a blunted-wedge holder and exposed to hypersonic flow at various angles of attack. The effects of flow direction, heating load, and overtemperature were investigated. In addition, the reuse capability of the TPS materials was evaluated, along with the protection of the pressure seals within the test articles. Thermal match model predictions correlated well with primary structure thermocouple data. Heating rate and pressure predictions based on a nonequilibrium flow field computer program showed good agreement with arc-jet test data and existing hypersonic flow theories.
High-speed flow visualization in hypersonic, transonic, and shock tube flows
NASA Astrophysics Data System (ADS)
Kleine, H.; Olivier, H.
2017-02-01
High-speed flow visualisation has played an important role in the investigations conducted at the Stoßwellenlabor of the RWTH Aachen University for many decades. In addition to applying the techniques of high-speed imaging, this laboratory has been actively developing new or enhanced visualisation techniques and approaches such as various schlieren methods or time-resolved Mach-Zehnder interferometry. The investigated high-speed flows are inherently highly transient, with flow Mach numbers ranging from about M = 0.7 to M = 8. The availability of modern high-speed cameras has allowed us to expand the investigations into problems where reduced reproducibility had so far limited the amount of information that could be extracted from a limited number of flow visualisation records. Following a brief historical overview, some examples of recent studies are given, which represent the breadth of applications in which high-speed imaging has been an essential diagnostic tool to uncover the physics of high-speed flows. Applications include the stability of hypersonic corner flows, the establishment of shock wave systems in transonic airfoil flow, and the complexities of the interactions of shock waves with obstacles of various shapes.
Hypersonic Flow over a Cylinder with a Nanosecond-Pulse Electrical Discharge
2013-01-01
Hypersonic Flow over a Cylinder with a Nanosecond- Pulse Electrical Discharge Nicholas J. Bisek,∗ Jonathan Poggie† Air Force Research Laboratory...study of Mach 5 air over a cylinder with a dielectric bar- rier discharge actuator was performed. The actuator was pulsed at nanosec- ond time scales, and...it rapidly added thermal energy to the flow, creating a shock wave that traveled away from the pulse source. As the shock wave traveled upstream, it
NASA Technical Reports Server (NTRS)
Chalot, F.; Hughes, T. J. R.; Johan, Z.; Shakib, F.
1991-01-01
A finite element method for the compressible Navier-Stokes equations is introduced. The discretization is based on entropy variables. The methodology is developed within the framework of a Galerkin/least-squares formulation to which a discontinuity-capturing operator is added. Results for four test cases selected among those of the Workshop on Hypersonic Flows for Reentry Problems are presented.
Parametric Study of Cantilever Plates Exposed to Supersonic and Hypersonic Flows
NASA Astrophysics Data System (ADS)
Sri Harsha, A.; Rizwan, M.; Kuldeep, S.; Giridhara Prasad, A.; Akhil, J.; Nagaraja, S. R.
2017-08-01
Analysis of hypersonic flows associated with re-entry vehicles has gained a lot of significance due to the advancements in Aerospace Engineering. An area that is studied extensively by researchers is the simultaneous reduction aerodynamic drag and aero heating in re-entry vehicles. Out of the many strategies being studied, the use of aerospikes at the stagnation point of the vehicle is found to give favourable results. The structural stability of the aerospike becomes important as it is exposed to very high pressures and temperatures. Keeping this in view, the deflection and vibration of an inclined cantilever plate in hypersonic flow is carried out using ANSYS. Steady state pressure distribution obtained from Fluent is applied as load to the transient structural module for analysis. After due validation of the methods, the effects of parameters like flow Mach number, plate inclination and plate thickness on the deflection and vibration are studied.
Evaluation of thermochemical models for particle and continuum simulations of hypersonic flow
NASA Technical Reports Server (NTRS)
Boyd, Iain D.; Gokcen, Tahir
1992-01-01
Computations are presented for one-dimensional, strong shock waves that are typical of those that form in front of a reentering spacecraft. The fluid mechanics and thermochemistry are modeled using two different approaches. The first employs traditional continuum techniques in solving the Navier-Stokes equations. The second approach employs a particle simulation technique (the direct simulation Monte Carlo method, DSMC). The thermochemical models employed in these two techniques are quite different. The present investigation presents an evaluation of thermochemical models for nitrogen under hypersonic flow conditions. Four separate cases are considered that are dominated in turn by vibrational relaxation, weak dissociation, strong dissociation and weak ionization. In near-continuum, hypersonic flow, the nonequilibrium thermochemical models employed in continuum and particle simulations produce nearly identical solutions. Further, the two approaches are evaluated successfully against available experimental data for weakly and strongly dissociating flows.
Evaluation of thermochemical models for particle and continuum simulations of hypersonic flow
NASA Technical Reports Server (NTRS)
Boyd, Iain D.; Gokcen, Tahir
1992-01-01
Computations are presented for one-dimensional, strong shock waves that are typical of those that form in front of a reentering spacecraft. The fluid mechanics and thermochemistry are modeled using two different approaches. The first employs traditional continuum techniques in solving the Navier-Stokes equations. The second approach employs a particle simulation technique (the direct simulation Monte Carlo method, DSMC). The thermochemical models employed in these two techniques are quite different. The present investigation presents an evaluation of thermochemical models for nitrogen under hypersonic flow conditions. Four separate cases are considered that are dominated in turn by vibrational relaxation, weak dissociation, strong dissociation and weak ionization. In near-continuum, hypersonic flow, the nonequilibrium thermochemical models employed in continuum and particle simulations produce nearly identical solutions. Further, the two approaches are evaluated successfully against available experimental data for weakly and strongly dissociating flows.
2010-05-01
materials under hypersonic flow conditions was developed and evaluated. A direct-connect scramjet combustor rig, de- signed based on the needs for...hypersonic flight and de- scribed elsewhere,10 was explored for use as a wind tunnel. A methodology to introduce samples into the flow path of the combustor ...Scramjet Direct-Connect Rig Gruber et al.10 have designed and fabricated a di- rect-connect full-scale scramjet combustor test facility for studies on
Simultaneous PSP and TSP measurements of transient flow in a long-duration hypersonic tunnel
NASA Astrophysics Data System (ADS)
Peng, Di; Jiao, Lingrui; Sun, Zhijun; Gu, Yunsong; Liu, Yingzheng
2016-12-01
The current work presents simultaneous measurements of transient flow using fast-responding pressure- and temperature-sensitive paints in a long-duration hypersonic tunnel; the pressure, temperature and heat flux fields were obtained on a standard model (HB-2) at Ma = 5. Fast PSP and TSP were applied symmetrically on the model with low thermal conductivity. Both coatings were illuminated by a UV-LED, and unsteady pressure and temperature data were recorded at 500 Hz using a high-speed camera. Time-dependent temperature correction was applied on the PSP data based on the TSP results, while the heat flux was calculated from the time-resolved temperature fields using a 1D semi-finite heat conduction model. The temperature-induced errors in PSP data were effectively removed by the current compensation method. The pressure and heat flux results showed good agreement with the reference data from previous studies. The key events throughout the hypersonic tunnel run were captured by the unsteady PSP/TSP data, including the tunnel start-up, the flow build-up, the steady flow period and the tunnel shutdown. The differences caused by the change of attack angle were also clearly recognized. The current PSP/TSP system has shown great potential for unsteady flow diagnostics in hypersonic flows.
Comparative study of turbulence models in predicting hypersonic inlet flows
NASA Technical Reports Server (NTRS)
Kapoor, Kamlesh; Anderson, Bernhard H.; Shaw, Robert J.
1992-01-01
A numerical study was conducted to analyze the performance of different turbulence models when applied to the hypersonic NASA P8 inlet. Computational results from the PARC2D code, which solves the full two-dimensional Reynolds-averaged Navier-Stokes equation, were compared with experimental data. The zero-equation models considered for the study were the Baldwin-Lomax model, the Thomas model, and a combination of the Baldwin-Lomax and Thomas models; the two-equation models considered were the Chien model, the Speziale model (both low Reynolds number), and the Launder and Spalding model (high Reynolds number). The Thomas model performed best among the zero-equation models, and predicted good pressure distributions. The Chien and Speziale models compared very well with the experimental data, and performed better than the Thomas model near the walls.
Comparative study of turbulence models in predicting hypersonic inlet flows
NASA Technical Reports Server (NTRS)
Kapoor, Kamlesh; Anderson, Bernhard H.; Shaw, Robert J.
1992-01-01
A numerical study was conducted to analyze the performance of different turbulence models when applied to the hypersonic NASA P8 inlet. Computational results from the PARC2D code, which solves the full two-dimensional Reynolds-averaged Navier-Stokes equation, were compared with experimental data. The zero-equation models considered for the study were the Baldwin-Lomax model, the Thomas model, and a combination of the Baldwin-Lomax and Thomas models; the two-equation models considered were the Chien model, the Speziale model (both low Reynolds number), and the Launder and Spalding model (high Reynolds number). The Thomas model performed best among the zero-equation models, and predicted good pressure distributions. The Chien and Speziale models compared wery well with the experimental data, and performed better than the Thomas model near the walls.
Three-dimensional hypersonic rarefied flow calculations using direct simulation Monte Carlo method
NASA Technical Reports Server (NTRS)
Celenligil, M. Cevdet; Moss, James N.
1993-01-01
A summary of three-dimensional simulations on the hypersonic rarefied flows in an effort to understand the highly nonequilibrium flows about space vehicles entering the Earth's atmosphere for a realistic estimation of the aerothermal loads is presented. Calculations are performed using the direct simulation Monte Carlo method with a five-species reacting gas model, which accounts for rotational and vibrational internal energies. Results are obtained for the external flows about various bodies in the transitional flow regime. For the cases considered, convective heating, flowfield structure and overall aerodynamic coefficients are presented and comparisons are made with the available experimental data. The agreement between the calculated and measured results are very good.
Drag Reduction by Laser-Plasma Energy Addition in Hypersonic Flow
Oliveira, A. C.; Minucci, M. A. S.; Toro, P. G. P.; Chanes, J. B. Jr; Myrabo, L. N.
2008-04-28
An experimental study was conducted to investigate the drag reduction by laser-plasma energy addition in a low density Mach 7 hypersonic flow. The experiments were conducted in a shock tunnel and the optical beam of a high power pulsed CO{sub 2} TEA laser operating with 7 J of energy and 30 MW peak power was focused to generate the plasma upstream of a hemispherical model installed in the tunnel test section. The non-intrusive schlieren optical technique was used to visualize the effects of the energy addition to hypersonic flow, from the plasma generation until the mitigation of the shock wave profile over the model surface. Aside the optical technique, a piezoelectric pressure transducer was used to measure the impact pressure at stagnation point of the hemispherical model and the pressure reduction could be observed.
Review of blunt body wake flows at hypersonic low density conditions
NASA Technical Reports Server (NTRS)
Moss, J. N.; Price, J. M.
1996-01-01
Recent results of experimental and computational studies concerning hypersonic flows about blunted cones including their near wake are reviewed. Attention is focused on conditions where rarefaction effects are present, particularly in the wake. The experiments have been performed for a common model configuration (70 deg spherically-blunted cone) in five hypersonic facilities that encompass a significant range of rarefaction and nonequilibrium effects. Computational studies using direct simulation Monte Carlo (DSMC) and Navier-Stokes solvers have been applied to selected experiments performed in each of the facilities. In addition, computations have been made for typical flight conditions in both Earth and Mars atmospheres, hence more energetic flows than produced in the ground-based tests. Also, comparisons of DSMC calculations and forebody measurements made for the Japanese Orbital Reentry Experiment (OREX) vehicle (a 50 deg spherically-blunted cone) are presented to bridge the spectrum of ground to flight conditions.
Spatially resolved excitation temperature measurements in a hypersonic flow using the hook method.
Sandeman, R J; Ebrahim, N A
1977-05-01
The extension of the hook method to include spatial resolution of nonuniformities in the test plane as suggested by Huber (1971) and Sandeman (1971) is demonstrated experimentally by measurements of the variation of the integrated line density of ground state sodium in a flame. Experiments are also described in which the variations in the flow of CO(2) in a hypersonic shock tunnel are spatially resolved along the spectrometer slit. The variations in the hook separations of the 425.4-nm Cr1 resonance and the 434.4-nm CrI 1-eV lower state line are simultaneously measured. The chromium exists as an impurity in the hypersonic flow of CO(2) over a cylinder in a shock tunnel. The populations of the levels so obtained have enabled the comparison of the excitation temperature of the Cr 1-eV level with the calculated gas temperature.
Real-Gas Correction Factors for Hypersonic Flow Parameters in Helium
NASA Technical Reports Server (NTRS)
Erickson, Wayne D.
1960-01-01
The real-gas hypersonic flow parameters for helium have been calculated for stagnation temperatures from 0 F to 600 F and stagnation pressures up to 6,000 pounds per square inch absolute. The results of these calculations are presented in the form of simple correction factors which must be applied to the tabulated ideal-gas parameters. It has been shown that the deviations from the ideal-gas law which exist at high pressures may cause a corresponding significant error in the hypersonic flow parameters when calculated as an ideal gas. For example the ratio of the free-stream static to stagnation pressure as calculated from the thermodynamic properties of helium for a stagnation temperature of 80 F and pressure of 4,000 pounds per square inch absolute was found to be approximately 13 percent greater than that determined from the ideal-gas tabulation with a specific heat ratio of 5/3.
Approximate Analytical Solutions for Hypersonic Flow Over Slender Power Law Bodies
NASA Technical Reports Server (NTRS)
Mirels, Harold
1959-01-01
Approximate analytical solutions are presented for two-dimensional and axisymmetric hypersonic flow over slender power law bodies. Both zero order (M approaches infinity) and first order (small but nonvanishing values of 1/(M(Delta)(sup 2) solutions are presented, where M is free-stream Mach number and Delta is a characteristic slope. These solutions are compared with exact numerical integration of the equations of motion and appear to be accurate particularly when the shock is relatively close to the body.
Surface Characterization of LMMS Molybdenum Disilicide Coated HTP-8 Using Arc- Jet Hypersonic Flow
NASA Technical Reports Server (NTRS)
Stewart, David A.
2000-01-01
Surface properties for an advanced Lockheed Martin Missile and Space (LMMS) molybdenum disilicide coated insulation (HTP-8) were determined using arc-jet flow to simulate Earth entry at hypersonic speeds. The catalytic efficiency (atom recombination coefficients) for this advanced thermal protection system was determined from arc-jet data taken in both oxygen and nitrogen streams at temperatures ranging from 1255 K to roughly 1600 K. In addition, optical and chemical stability data were obtained from these test samples.
Simulation of Stagnation Region Heating in Hypersonic Flow on Tetrahedral Grids
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
2007-01-01
Hypersonic flow simulations using the node based, unstructured grid code FUN3D are presented. Applications include simple (cylinder) and complex (towed ballute) configurations. Emphasis throughout is on computation of stagnation region heating in hypersonic flow on tetrahedral grids. Hypersonic flow over a cylinder provides a simple test problem for exposing any flaws in a simulation algorithm with regard to its ability to compute accurate heating on such grids. Such flaws predominantly derive from the quality of the captured shock. The importance of pure tetrahedral formulations are discussed. Algorithm adjustments for the baseline Roe / Symmetric, Total-Variation-Diminishing (STVD) formulation to deal with simulation accuracy are presented. Formulations of surface normal gradients to compute heating and diffusion to the surface as needed for a radiative equilibrium wall boundary condition and finite catalytic wall boundary in the node-based unstructured environment are developed. A satisfactory resolution of the heating problem on tetrahedral grids is not realized here; however, a definition of a test problem, and discussion of observed algorithm behaviors to date are presented in order to promote further research on this important problem.
DSMC Simulation and Experimental Validation of Shock Interaction in Hypersonic Low Density Flow
2014-01-01
Direct simulation Monte Carlo (DSMC) of shock interaction in hypersonic low density flow is developed. Three collision molecular models, including hard sphere (HS), variable hard sphere (VHS), and variable soft sphere (VSS), are employed in the DSMC study. The simulations of double-cone and Edney's type IV hypersonic shock interactions in low density flow are performed. Comparisons between DSMC and experimental data are conducted. Investigation of the double-cone hypersonic flow shows that three collision molecular models can predict the trend of pressure coefficient and the Stanton number. HS model shows the best agreement between DSMC simulation and experiment among three collision molecular models. Also, it shows that the agreement between DSMC and experiment is generally good for HS and VHS models in Edney's type IV shock interaction. However, it fails in the VSS model. Both double-cone and Edney's type IV shock interaction simulations show that the DSMC errors depend on the Knudsen number and the models employed for intermolecular interaction. With the increase in the Knudsen number, the DSMC error is decreased. The error is the smallest in HS compared with those in the VHS and VSS models. When the Knudsen number is in the level of 10−4, the DSMC errors, for pressure coefficient, the Stanton number, and the scale of interaction region, are controlled within 10%. PMID:24672360
DSMC simulation and experimental validation of shock interaction in hypersonic low density flow.
Xiao, Hong; Shang, Yuhe; Wu, Di
2014-01-01
Direct simulation Monte Carlo (DSMC) of shock interaction in hypersonic low density flow is developed. Three collision molecular models, including hard sphere (HS), variable hard sphere (VHS), and variable soft sphere (VSS), are employed in the DSMC study. The simulations of double-cone and Edney's type IV hypersonic shock interactions in low density flow are performed. Comparisons between DSMC and experimental data are conducted. Investigation of the double-cone hypersonic flow shows that three collision molecular models can predict the trend of pressure coefficient and the Stanton number. HS model shows the best agreement between DSMC simulation and experiment among three collision molecular models. Also, it shows that the agreement between DSMC and experiment is generally good for HS and VHS models in Edney's type IV shock interaction. However, it fails in the VSS model. Both double-cone and Edney's type IV shock interaction simulations show that the DSMC errors depend on the Knudsen number and the models employed for intermolecular interaction. With the increase in the Knudsen number, the DSMC error is decreased. The error is the smallest in HS compared with those in the VHS and VSS models. When the Knudsen number is in the level of 10(-4), the DSMC errors, for pressure coefficient, the Stanton number, and the scale of interaction region, are controlled within 10%.
Hypersonic Laminar Viscous Flow Past Spinning Cones at Angle of Attack
NASA Technical Reports Server (NTRS)
Agarwal, Ramesh; Rakich, John V.
1982-01-01
Computational results are presented for hypersonic viscous flow past spinning sharp and blunt cones of angle of attack, obtained with a parabolic Navier-Stokes marching code. The code takes into account the asymmetries in the flowfield resulting from spinning motion and computes the asymmetric shock shape, cross-flow and streamwise shear, heat transfer, cross-flow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results.
Line-shape flattening resulting from hypersonic nozzle wedge flow in low-pressure chemical lasers.
Livingston, P M; Bullock, D L
1980-07-01
The new hypersonic wedge nozzle (HYWN) supersonic wedge nozzle design produces a significant component of directed gas flow along the optical axis of a laser cavity comparable to thermal speeds. The gain-line-shape function is broadened and the refractive-index line shape is also spread as a function of wedge-flow half-angle. An analytical treatment as well as a numerical study is presented that evaluates the Doppler-directed-flow impact on the number of longitudinal modes and their frequencies as well as on gain and refractive-index saturation of those that lase in a Fabry-Perot cavity.
New method of asymmetric flow field measurement in hypersonic shock tunnel.
Yan, D P; He, A Z; Ni, X W
1991-03-01
In this paper a method of large aperture (?500 mm) high sensitivity moire deflectometry is used to obtain multidirectional deflectograms of the asymmetric flow field in hypersonic (M = 10.29) shock tunnel. At the same time, a 3-D reconstructive method of the asymmetric flow field is presented which is based on the integration of the moire deflective angle and the double-cubic many-knot interpolating splines; it is used to calculate the 3-D density distribution of the asymmetric flow field.
Numerical study of unsteady viscous hypersonic blunt body flows with an impinging shock
NASA Technical Reports Server (NTRS)
Klopfer, G. H.; Yee, H. C.; Kutler, P.
1988-01-01
A complex two-dimensional, unsteady, viscous hypersonic shock wave interaction is numerically simulated by a high-resolution, second-order fully implicit shock-capturing scheme. The physical model consists of a nonstationary oblique shock impinging on the bow shock of a blunt body. Studies indicated that the unsteady flow patterns are slightly different from their steady counterparts. However, for the sample cases investigated the peak surface pressures for the unsteady flows seem to occur at very different impingement locations than for the steady flow cases.
Modeling hypersonic boundary-layer flows with second-moment closure
NASA Technical Reports Server (NTRS)
Huang, P. George
1991-01-01
An ongoing research effort designed to apply the best possible second-moment-closure model to simulate complex hypersonic flows is presented. The baseline model under consideration is the Launder-Reece-Rodi Reynolds stress transport turbulence model. Two add-ons accounting for wall effects, namely, the Launder-Shima low-Reynolds-number model and the compressible wall-function technique, are tested. Results are reported for flow over a flat plate, both adiabatic-wall and cooled-wall cases. It has been found that further improvements of the existing models are necessary to achieve accurate prediction in high Mach number flow range.
Hypersonic Laminar Viscous Flow Past Spinning Cones at Angle of Attack
NASA Technical Reports Server (NTRS)
Agarwal, Ramesh; Rakich, John V.
1982-01-01
Computational results are presented for hypersonic viscous flow past spinning sharp and blunt cones of angle of attack, obtained with a parabolic Navier-Stokes marching code. The code takes into account the asymmetries in the flowfield resulting from spinning motion and computes the asymmetric shock shape, cross-flow and streamwise shear, heat transfer, cross-flow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results.
Flow characteristics of hypersonic inlets with different cowl-lip blunting methods
NASA Astrophysics Data System (ADS)
Lu, HongBo; Yue, LianJie; Chang, XinYu
2014-04-01
Under hypersonic flight conditions, the sharp cowl-lip leading edges have to be blunted because of the severe aerodynamic heating. This paper proposes four cowl-lip blunting methods and studies the corresponding flow characteristics and performances of the generic hypersonic inlets by numerical simulation under the design conditions of a flight Mach number of 6 and an altitude of 26 km. The results show that the local shock interference patterns in the vicinity of the blunted cowl-lips have a substantial influence on the flow characteristics of the hypersonic inlets even though the blunting radius is very small, which contribute to a pronounced degradation of the inlet performance. The Equal Length blunting Manner (ELM) is the most optimal in that a nearly even reflection of the ramp shock produces an approximately straight and weak cowl reflection shock. The minimal total pressure loss, the lowest cowl drag, maximum mass-capture and the minimal aeroheating are achieved for the hypersonic inlet. For the other blunting manners, the ramp shock cannot reflect evenly and produces more curved cowl reflection shock. The Type V shock interference pattern occurs for the Cross Section Cutting blunting Manner (CSCM) and the strongest cowl reflection shock gives rise to the largest flow loss and drag. The cowl-lip blunted by the other two blunting manners is subjected to the shock interference pattern that transits with an increase in the blunting radius. Accordingly, the peak heat flux does not fall monotonously with the blunting radius increasing. Moreover, the cowl-lip surface suffers from severe aerothermal load when the shear layer or the supersonic jet impinges on the wall.
Computational Study of Flow Establishment in Hypersonic Pulse Facilities
NASA Technical Reports Server (NTRS)
Yungster, S.; Radhakrishnan, K.
1995-01-01
This paper presents a study of the temporal evolution of the combustion flowfield established by the interaction of ram-accelerator-type projectiles with an explosive gas mixture accelerated to hypersonic speeds in an expansion tube. The Navier-Stokes equations for a chemically reacting gas are solved in a fully coupled manner using an implicit, time accurate algorithm. The solution procedure is based on a spatially second order, total variation diminishing (TVD) scheme and a temporally second order, variable-step, backward differentiation formula method. The hydrogen-oxygen chemistry is modeled with a 9-species, 19-step mechanism. The accuracy of the solution method is first demonstrated by several benchmark calculations. Numerical simulations of expansion tube flowfields are then presented for two different configurations. In particular, the development of the shock-induced combustion process is followed. In one case, designed to ensure ignition only in the boundary layer, the lateral extent of the combustion front during the initial transient phase was surprisingly large. The time histories of the calculated thrust and drag forces on the ram accelerator projectile are also presented.
An efficient collision limiter Monte Carlo simulation for hypersonic near-continuum flows
NASA Astrophysics Data System (ADS)
Liang, Jie; Li, Zhihui; Li, Xuguo; Fang, Boqiang Du Ming
2016-11-01
The implementation of a collision limiter DSMC-based hybrid approach is presented to simulate hypersonic near-continuum flow. The continuum breakdown parameters based on gradient-length local Knudsen number are characterized different regions of the flowfield. The collision limiter is used in continuum inviscid regions with large time step and cell size. Local density gradient-based dynamic adaptation of collision and sampling cells refinement is employed in high gradient regions including strong shocks and boundary layer near surface. A variable time step scheme is adopted to make sure a more uniform distribution of model particles per collision cell throughout the computational domain, with a constant ratio of local time step interval to particle weights to avoid particles cloned or destroyed when crossing interface from cell to cell. The surface pressure and friction coefficients of hypersonic reentry flow for a blunt capsule are computed in different conditions and compared with benchmark case in transitional regime to examine the efficiency and accuracy. The aerodynamic characteristics of a wave rider shape with sharp leading edge are simulated in the test state for hypersonic near-continuum. The computed aerodynamic coefficients have good agreements with experimental data in low density wind tunnel of CARDC and have less computational expense.
Interference effects on the hypersonic, rarefied flow about a flat plate
NASA Technical Reports Server (NTRS)
Wilmoth, Richard G.
1989-01-01
The Direct Simulatiaon Monte Carlo method is used to study the hypersonic, rarefied flow interference effects on a flat plate caused by nearby surfaces. Calculations focus on shock-boundary-layer and shock-lip interactions in hypersonic inlets. Results are presented for geometries consisting of a flat plate with different leading-edge shapes over a flat lower wall and a blunt-edge flat plate over a 5-degree wedge. The problems simulated correspond to a typical entry flight condition of 7.5 km/s at altitudes of 75 to 90 km. The results show increases in predicted local heating rates for shock-boundary-layer and shock-lip interactions that are quantitatively similar to those observed experimentally at much higher densities.
The application of laser Rayleigh scattering to gas density measurements in hypersonic helium flows
NASA Technical Reports Server (NTRS)
Hoppe, J. C.; Honaker, W. C.
1979-01-01
Measurements of the mean static free-stream gas density have been made in two Langley Research Center helium facilities, the 3-inch leg of the high-Reynolds-number helium complex and the 22-inch hypersonic helium tunnel. Rayleigh scattering of a CW argon ion laser beam at 514.5 nm provided the basic physical mechanism. The behavior of the scattered signal was linear, confirmed by a preliminary laboratory study. That study also revealed the need to introduce baffles to reduce stray light. A relatively simple optical system and associated photon-counting electronics were utilized to obtain data for densities from 10 to the 23rd to 10 to the 25th per cu m. The major purpose, to confirm the applicability of this technique in the hypersonic helium flow, was accomplished.
Interference effects on the hypersonic, rarefied flow about a flat plate
NASA Technical Reports Server (NTRS)
Wilmoth, Richard G.
1988-01-01
The Direct Simulation Monte Carlo method is used to study the hypersonic, rarified flow interference effects on a flat plate caused by nearby surfaces. Calculations focus on shock-boundary-layer and shock-lip interactions in hypersonic inlets. Results are presented for geometries consisting of a flat plate with different leading-edge shapes over a flat lower wall and a blunt-edge flat plate over a 5-degree wedge. The problems simulated correspond to a typical entry flight condition of 7.5 km/s at altitudes of 75 to 90 km. The results show increases in predicted local heating rates for shock-boundary-layer and shock-lip interactions that are quantitatively similar to those observed experimentally at much higher densities.
The application of laser Rayleigh scattering to gas density measurements in hypersonic helium flows
NASA Technical Reports Server (NTRS)
Hoppe, J. C.; Honaker, W. C.
1979-01-01
Measurements of the mean static free-stream gas density have been made in two Langley Research Center helium facilities, the 3-inch leg of the high-Reynolds-number helium complex and the 22-inch hypersonic helium tunnel. Rayleigh scattering of a CW argon ion laser beam at 514.5 nm provided the basic physical mechanism. The behavior of the scattered signal was linear, confirmed by a preliminary laboratory study. That study also revealed the need to introduce baffles to reduce stray light. A relatively simple optical system and associated photon-counting electronics were utilized to obtain data for densities from 10 to the 23rd to 10 to the 25th per cu m. The major purpose, to confirm the applicability of this technique in the hypersonic helium flow, was accomplished.
DSMC Simulation of Separated Flows About Flared Bodies at Hypersonic Conditions
NASA Technical Reports Server (NTRS)
Moss, James N.
2000-01-01
This paper describes the results of a numerical study of interacting hypersonic flows at conditions that can be produced in ground-based test facilities. The computations are made with the direct simulation Monte Carlo (DSMC) method of Bird. The focus is on Mach 10 flows about flared axisymmetric configurations, both hollow cylinder flares and double cones. The flow conditions are those for which experiments have been or will be performed in the ONERA R5Ch low-density wind tunnel and the Calspan-University of Buffalo Research Center (CUBRC) Large Energy National Shock (LENS) tunnel. The range of flow conditions, model configurations, and model sizes provides a significant range of shock/shock and shock/boundary layer interactions at low Reynolds number conditions. Results presented will highlight the sensitivity of the calculations to grid resolution, contrast the differences in flow structure for hypersonic cold flows and those of more energetic but still low enthalpy flows, and compare the present results with experimental measurements for surface heating, pressure, and extent of separation.
Effects of nose bluntness and shock-shock interactions on blunt bodies in viscous hypersonic flows
NASA Technical Reports Server (NTRS)
Singh, D. J.; Tiwari, S. N.
1990-01-01
A numerical study was conducted to investigate the effects of blunt leading edges on the viscous flow field around a hypersonic vehicle such as the proposed National Aero-Space Plane. Attention is focused on two specific regions of the flow field. In the first region, effects of nose bluntness on the forebody flow field are investigated. The second region of the flow considered is around the leading edges of the scramjet inlet. In this region, the interaction of the forebody shock with the shock produced by the blunt leading edges of the inlet compression surfaces is analyzed. Analysis of these flow regions is required to accurately predict the overall flow field as well as to get necessary information on localized zones of high pressure and intense heating. The results for the forebody flow field are discussed first, followed by the results for the shock interaction in the inlet leading edge region.
Collision partner selection schemes in DSMC: From micro/nano flows to hypersonic flows
NASA Astrophysics Data System (ADS)
Roohi, Ehsan; Stefanov, Stefan
2016-10-01
The motivation of this review paper is to present a detailed summary of different collision models developed in the framework of the direct simulation Monte Carlo (DSMC) method. The emphasis is put on a newly developed collision model, i.e., the Simplified Bernoulli trial (SBT), which permits efficient low-memory simulation of rarefied gas flows. The paper starts with a brief review of the governing equations of the rarefied gas dynamics including Boltzmann and Kac master equations and reiterates that the linear Kac equation reduces to a non-linear Boltzmann equation under the assumption of molecular chaos. An introduction to the DSMC method is provided, and principles of collision algorithms in the DSMC are discussed. A distinction is made between those collision models that are based on classical kinetic theory (time counter, no time counter (NTC), and nearest neighbor (NN)) and the other class that could be derived mathematically from the Kac master equation (pseudo-Poisson process, ballot box, majorant frequency, null collision, Bernoulli trials scheme and its variants). To provide a deeper insight, the derivation of both collision models, either from the principles of the kinetic theory or the Kac master equation, is provided with sufficient details. Some discussions on the importance of subcells in the DSMC collision procedure are also provided and different types of subcells are presented. The paper then focuses on the simplified version of the Bernoulli trials algorithm (SBT) and presents a detailed summary of validation of the SBT family collision schemes (SBT on transient adaptive subcells: SBT-TAS, and intelligent SBT: ISBT) in a broad spectrum of rarefied gas-flow test cases, ranging from low speed, internal micro and nano flows to external hypersonic flow, emphasizing first the accuracy of these new collision models and second, demonstrating that the SBT family scheme, if compared to other conventional and recent collision models, requires smaller
Towards High-Reynolds Number Quiet Flow in Hypersonic Tunnels
2009-02-23
MOC procedure is used in an inverse design mode to determine an inviscid nozzle wall contour that produces the desired uniform exit flow. This...determine the optimal shape of the supersonic nozzle to achieve laminar flow on the nozzle walls and hence quiet flow in the test section. In the...methodology was developed to determine the optimal shape of the supersonic nozzle to achieve laminar flow on the nozzle walls and maximize the length of
A Numerical Simulation of a Normal Sonic Jet into a Hypersonic Cross-Flow
NASA Technical Reports Server (NTRS)
Jeffries, Damon K.; Krishnamurthy, Ramesh; Chandra, Suresh
1997-01-01
This study involves numerical modeling of a normal sonic jet injection into a hypersonic cross-flow. The numerical code used for simulation is GASP (General Aerodynamic Simulation Program.) First the numerical predictions are compared with well established solutions for compressible laminar flow. Then comparisons are made with non-injection test case measurements of surface pressure distributions. Good agreement with the measurements is observed. Currently comparisons are underway with the injection case. All the experimental data were generated at the Southampton University Light Piston Isentropic Compression Tube.
Hypersonic rarefied-flow aerodynamics inferred from Shuttle Orbiter acceleration measurements
NASA Technical Reports Server (NTRS)
Blanchard, R. C.; Hinson, E. W.
1989-01-01
Data obtained from multiple flights of sensitive accelerometers on the Space Shuttle Orbiter during reentry have been used to develop an improved aerodynamic model for the Orbiter normal- and axial-force coefficients in hypersonic rarefied flow. The lack of simultaneous atmospheric density measurements was overcome in part by using the ratio of normal-to-axial acceleration, in which density cancels, as a constraint. Differences between the preflight model and the flight-acceleration-derived model in the continuum regime are attributed primarily to real gas effects. New insights are gained into the variation of the force coefficients in the transition between the continuum regime and free molecule flow.
Towards High Reynolds Number Quiet Flow in Hypersonic Tunnels
2009-02-23
design methodology was developed to determine the optimal shape of the supersonic nozzle to achieve laminar flow on the nozzle walls and hence quiet...fully automated optimal design methodology was developed to determine the optimal shape of the supersonic nozzle to achieve laminar flow on the... nozzle walls and maximize the length of the region of quiet flow in the test section. In the latter case, detailed time-accurate numerical simulations
Fluid flow analysis of a hot-core hypersonic wind-tunnel nozzle concept
NASA Technical Reports Server (NTRS)
Anders, J. B.; Sebacher, D. I.; Boatright, W. B.
1972-01-01
A hypersonic-wind-tunnel nozzle concept which incorporates a hot-core flow surrounded by an annular flow of cold air offers a promising technique for maximizing the model size while minimizing the power required to heat the test core. This capability becomes especially important when providing the true-temperature duplication needed for hypersonic propulsion testing. Several two-dimensional wind-tunnel nozzle configurations that are designed according to this concept are analyzed by using recently developed analytical techniques for prediction of the boundary-layer growth and the mixing between the hot and cold coaxial supersonic airflows. The analyses indicate that introduction of the cold annular flow near the throat results in an unacceptable test core for the nozzle size and stagnation conditions considered because of both mixing and condensation effects. Use of a half-nozzle with a ramp on the flat portion does not appear promising because of the thick boundary layer associated with the extra length. However, the analyses indicate that if the cold annular flow is introduced at the exit of a full two-dimensional nozzle, an acceptable test core will be produced. Predictions of the mixing between the hot and cold supersonic streams for this configuration show that mixing effects from the cold flow do not appreciably penetrate into the hot core for the large downstream distances of interest.
The hypersonic Mach number independence principle in the case of viscous flow
NASA Astrophysics Data System (ADS)
Kliche, D.; Mundt, Ch.; Hirschel, E. H.
2011-08-01
The hypersonic Mach number independence principle of Oswatitsch is important for hypersonic vehicle design. It explains why, above a certain flight Mach number ( M ∞ ≈ 4-6, depending on the body shape), some aerodynamic properties become independent of the flight Mach number. For ground test facilities this means that it is sufficient for the Mach number in the test section to be high enough, that Mach number independence exists. However, the principle was derived for calorically perfect gas and inviscid flow only. In this paper a theoretical study for blunt bodies in the case of viscous flow is presented. We provide numerical results which give insight into how attached viscous flow behaves at high Mach numbers. The flow past an axisymmetric configuration is analysed by applying a coupled Euler/second-order boundary-layer method. Wall boundaries are treated by assuming an adiabatic or radiation-adiabatic wall for laminar flow. Calorically perfect or equilibrium air is accounted for. Lift, drag, and moment coefficients, and lift-to-drag ratios are given for several combinations of flight Mach number and altitude, i.e. Reynolds number. For blunt bodies considered here, which are pressure dominated, Mach number independence occurs for the adiabatic wall, but not for the radiation-adiabatic wall assumption.
NASA Technical Reports Server (NTRS)
Holland, Scott Douglas
1991-01-01
A combined computational and experimental parametric study of the internal aerodynamics of a generic three dimensional sidewall compression scramjet inlet configuration was performed. The study was designed to demonstrate the utility of computational fluid dynamics as a design tool in hypersonic inlet flow fields, to provide a detailed account of the nature and structure of the internal flow interactions, and to provide a comprehensive surface property and flow field database to determine the effects of contraction ratio, cowl position, and Reynolds number on the performance of a hypersonic scramjet inlet configuration.
Flow analysis and design optimization methods for nozzle-afterbody of a hypersonic vehicle
NASA Technical Reports Server (NTRS)
Baysal, O.
1992-01-01
This report summarizes the methods developed for the aerodynamic analysis and the shape optimization of the nozzle-afterbody section of a hypersonic vehicle. Initially, exhaust gases were assumed to be air. Internal-external flows around a single scramjet module were analyzed by solving the 3D Navier-Stokes equations. Then, exhaust gases were simulated by a cold mixture of Freon and Ar. Two different models were used to compute these multispecies flows as they mixed with the hypersonic airflow. Surface and off-surface properties were successfully compared with the experimental data. The Aerodynamic Design Optimization with Sensitivity analysis was then developed. Pre- and postoptimization sensitivity coefficients were derived and used in this quasi-analytical method. These coefficients were also used to predict inexpensively the flow field around a changed shape when the flow field of an unchanged shape was given. Starting with totally arbitrary initial afterbody shapes, independent computations were converged to the same optimum shape, which rendered the maximum axial thrust.
Study of flow control by localized volume heating in hypersonic boundary layers
NASA Astrophysics Data System (ADS)
Keller, M. A.; Kloker, M. J.; Kirilovskiy, S. V.; Polivanov, P. A.; Sidorenko, A. A.; Maslov, A. A.
2014-12-01
Boundary-layer flow control is a prerequisite for a safe and efficient operation of future hypersonic transport systems. Here, the influence of an electric discharge—modeled by a heat-source term in the energy equation—on laminar boundary-layer flows over a flat plate with zero pressure gradient at Mach 3, 5, and 7 is investigated numerically. The aim was to appraise the potential of electro-gasdynamic devices for an application as turbulence generators in the super- and hypersonic flow regime. The results with localized heat-source elements in boundary layers are compared to cases with roughness elements serving as classical passive trips. The numerical simulations are performed using the commercial code ANSYS FLUENT (by ITAM) and the high-order finite-difference DNS code NS3D (by IAG), the latter allowing for the detailed analysis of laminar flow instability. For the investigated setups with steady heating, transition to turbulence is not observed, due to the Reynolds-number lowering effect of heating.
Fluid dynamic modeling and numerical simulation of low-density hypersonic flow
NASA Technical Reports Server (NTRS)
Cheng, H. K.; Wong, Eric Y.
1988-01-01
The concept of a viscous shock-layer and several related versions of continuum theories/methods are examined for their adequacy as a viable framework to study flow physics and aerothermodynamics of relevance to sustained hypersonic flights. Considering the flat plate at angle of attack, or the wedge, as a generic example for the major aerodynamic component of a hypersonic vehicle, the relative importance of the molecular-transport effects behind the shock (in the form of the 'shock slip') and the wall-slip effects are studied. In the flow regime where the shock-transition-zone thickness remains small compared to the shock radius of curvature, a quasi-one-dimensional shock structure under the Burnett/thirteen-moment approximation, as well as particulate/collisional models, can be consistently developed. The fully viscous version of the shock-layer model is shown to provide the crucial boundary condition downstream the shock in this case. The gas-kinetic basis of the continuum description for the flow behind the bow shock, and certain features affecting the non-equilibrium flow chemistry, are also discussed.
NASA Technical Reports Server (NTRS)
Hartill, W. R.
1977-01-01
A hypersonic wind tunnel test method for obtaining credible aerodynamic data on a complete hypersonic vehicle (generic X-24c) with scramjet exhaust flow simulation is described. The general problems of simulating the scramjet exhaust as well as accounting for scramjet inlet flow and vehicle forces are analyzed, and candidate test methods are described and compared. The method selected as most useful makes use of a thrust-minus-drag flow-through balance with a completely metric model. Inlet flow is diverted by a fairing. The incremental effect of the fairing is determined in the testing of two reference models. The net thrust of the scramjet module is an input to be determined in large-scale module tests with scramjet combustion. Force accounting is described, and examples of force component levels are predicted. Compatibility of the test method with candidate wind tunnel facilities is described, and a preliminary model mechanical arrangement drawing is presented. The balance design and performance requirements are described in a detailed specification. Calibration procedures, model instrumentation, and a test plan for the model are outlined.
Flow analysis and design optimization methods for nozzle afterbody of a hypersonic vehicle
NASA Technical Reports Server (NTRS)
Baysal, Oktay
1991-01-01
This report summarizes the methods developed for the aerodynamic analysis and the shape optimization of the nozzle-afterbody section of a hypersonic vehicle. Initially, exhaust gases were assumed to be air. Internal-external flows around a single scramjet module were analyzed by solving the three dimensional Navier-Stokes equations. Then, exhaust gases were simulated by a cold mixture of Freon and Argon. Two different models were used to compute these multispecies flows as they mixed with the hypersonic airflow. Surface and off-surface properties were successfully compared with the experimental data. In the second phase of this project, the Aerodynamic Design Optimization with Sensitivity analysis (ADOS) was developed. Pre and post optimization sensitivity coefficients were derived and used in this quasi-analytical method. These coefficients were also used to predict inexpensively the flow field around a changed shape when the flow field of an unchanged shape was given. Starting with totally arbitrary initial afterbody shapes, independent computations were converged to the same optimum shape, which rendered the maximum axial thrust.
Quantifying Non-Equilibrium in Hypersonic Flows Using Entropy Generation
2007-03-01
Lennard - Jones model is offered as an alternative to Sutherland?s Law for calculating viscosity and thermal conductivity. The two are compared, and parameters offering a good fit for these flows are suggested for the Lennard - Jones
Oliveira, A. C.; Minucci, M. A. S.; Toro, P. G. P.; Chanes, J. B. Jr; Salvador, I. I.; Myrabo, L. N.; Nagamatsu, H. T.
2006-05-02
Experimental results on the visualization of the time evolution of the laser-plasma induced breakdown produced in low density hypersonic flow using the Schlieren technique are presented. The plasma was generated by focusing the high power laser pulse of a CO2 TEA laser in the test section of the IEAv 0.3m Hypersonic Shock Tunnel. An ultra-high speed electronic tube camera was used to register the event. The photographs reveal the expansion of the shock wave produced by the laser generated hot plasma and the convection of the plasma kernel by the hypersonic flow. It is also observed the interaction between the plasma disturbed region and the shock established by the flow around an hemisphere-cylinder model. A strong change in the shock wave structure near the model was observed, corroborating the DEAS concept.
Li, Zhihui; Ma, Qiang; Wu, Junlin; Jiang, Xinyu; Zhang, Hanxin
2014-12-09
Based on the Gas-Kinetic Unified Algorithm (GKUA) directly solving the Boltzmann model equation, the effect of rotational non-equilibrium is investigated recurring to the kinetic Rykov model with relaxation property of rotational degrees of freedom. The spin movement of diatomic molecule is described by moment of inertia, and the conservation of total angle momentum is taken as a new Boltzmann collision invariant. The molecular velocity distribution function is integrated by the weight factor on the internal energy, and the closed system of two kinetic controlling equations is obtained with inelastic and elastic collisions. The optimization selection technique of discrete velocity ordinate points and numerical quadrature rules for macroscopic flow variables with dynamic updating evolvement are developed to simulate hypersonic flows, and the gas-kinetic numerical scheme is constructed to capture the time evolution of the discretized velocity distribution functions. The gas-kinetic boundary conditions in thermodynamic non-equilibrium and numerical procedures are studied and implemented by directly acting on the velocity distribution function, and then the unified algorithm of Boltzmann model equation involving non-equilibrium effect is presented for the whole range of flow regimes. The hypersonic flows involving non-equilibrium effect are numerically simulated including the inner flows of shock wave structures in nitrogen with different Mach numbers of 1.5-Ma-25, the planar ramp flow with the whole range of Knudsen numbers of 0.0009-Kn-10 and the three-dimensional re-entering flows around tine double-cone body.
NASA Technical Reports Server (NTRS)
Anders, J. B., Jr.
1975-01-01
An axisymmetric, hypersonic nozzle for arc-heated air is described. The method of characteristics is used to compute an inviscid nozzle contour in which vibrational nonequilibrium is approximated by the sudden-freeze technique. Chemical reactions are shown to freeze early in the nozzle expansion, and the result of vibrational and chemical freezing on the nozzle contour is demonstrated. The approximate nozzle design is analyzed by an exact calculation based on the method of characteristics for flow with vibrational nonequilibrium. Exit profiles are computed, and the usefulness of the approximate design is discussed. An analysis of the nozzle performance at off-design conditions is presented.
Temperature measurements in hypersonic air flows using laser-induced O2 fluorescence
NASA Technical Reports Server (NTRS)
Laufer, Gabriel; Mckenzie, Robert L.
1988-01-01
An investigation is reported of the use of laser-induced fluorescence on oxygen for the measurement of air temperature and its fluctuations owing to turbulence in hypersonic wind tunnel flows. The results show that for temperatures higher than 60 K and densities higher than 0.01 amagat, the uncertainty in the temperature measurement can be less than 2 percent if it is limited by photon-statistical noise. The measurement is unaffected by collisional quenching and, if the laser fluence is kept below 1.5 J/sq cm, it is also unaffected by nonlinear effects which are associated with depletion of the absorbing states.
An iodine hypersonic wind tunnel for the study of nonequilibrium reacting flows
NASA Technical Reports Server (NTRS)
Pham-Van-diep, G. C.; Muntz, E. P.; Weaver, D. P.; Dewitt, T. G.; Bradley, M. K.; Erwin, D. A.; Kunc, J. A.
1992-01-01
A pilot scale hypersonic wind tunnel operating on pure iodine vapor has been designed and tested. The wind tunnel operates intermittently with a run phase lasting approximately 20 minutes. Successful recirculation of the iodine used during the run phase has been achieved but can be improved. Relevant issues regarding the full scale facility's design and operation, and the use of iodine as a working gas are discussed. Continuous wave laser induced fluorescence was used to monitor number densities within the plume flowfield, while pulsed laser induced fluorescence was used in an initial attempt to measure vibrational energy state population distributions. Preliminary nozzle flow calculations based on finite rate chemistry are presented.
Aerodynamic Modeling of Oscillating Wing in Hypersonic Flow: a Numerical Study
NASA Astrophysics Data System (ADS)
Zhu, Jian; Hou, Ying-Yu; Ji, Chen; Liu, Zi-Qiang
2016-06-01
Various approximations to unsteady aerodynamics are examined for the unsteady aerodynamic force of a pitching thin double wedge airfoil in hypersonic flow. Results of piston theory, Van Dyke’s second-order theory, Newtonian impact theory, and CFD method are compared in the same motion and Mach number effects. The results indicate that, for this thin double wedge airfoil, Newtonian impact theory is not suitable for these Mach number, while piston theory and Van Dyke’s second-order theory are in good agreement with CFD method for Ma<7.
NASA Technical Reports Server (NTRS)
Grantz, A. C.; Dejarnette, F. R.; Thompson, R. A.
1989-01-01
The approximate axisymmetric method presented for accurately calculating the surface and flowfield properties of fully viscous hypersonic flow over blunt-nosed bodies incorporates the turbulence model of Cebeci-Smith (1970) and the equilibrium air tables of Hansen (1959). The method is faster than the parabolized Navier-Stokes or viscous shock layer solvers that it could replace for preliminary design determinations. Surface heat transfer and pressure predictions for the present method are comparable with the more accurate viscous shock layer method as well as flight test and wind tunnel data. A starting solution is not required.
Dissociation-recombination models in hypersonic boundary layer O2/O flows
NASA Astrophysics Data System (ADS)
Armenise, I.; Esposito, F.
2012-04-01
A recent complete set of oxygen atom-molecule collision rate coefficients, calculated by means of a quasiclassical trajectory (QCT) method, has been used to evaluate the vibrational non-equilibrium in hypersonic boundary layer flows. The importance of multiquanta transitions has been demonstrated. Moreover a new 'direct dissociation-recombination' (DDR) model has been adopted and the corresponding results differ from the ones obtained with the ladder-climbing (LC) model, characterized by the extrapolation of bound-to-bound transitions to the continuum. The heat flux through the boundary layer and at the surface has been calculated too.
Hypersonic stagnation line merged layer flow on blunt axisymmetric bodies of arbitrary shape
NASA Technical Reports Server (NTRS)
Jain, Amolak S.
1993-01-01
The problem of hypersonic stagnation line merged-layer flow of variously shaped blunt asisymmetric bodies is here formulated in such a way as to allow analytical calculations for bodies generated by a conic section. The governing equations encompass, apart from the usual parameters, the eccentricity of the conic section that generates the body-of-revolution for the effect of body shape on the solution obtained. The stagnation-point surface pressure increases as the favorable pressure gradient decreases, in the course of a change of body shape from spherical to hyperboloid.
An iodine hypersonic wind tunnel for the study of nonequilibrium reacting flows
NASA Technical Reports Server (NTRS)
Pham-Van-diep, G. C.; Muntz, E. P.; Weaver, D. P.; Dewitt, T. G.; Bradley, M. K.; Erwin, D. A.; Kunc, J. A.
1992-01-01
A pilot scale hypersonic wind tunnel operating on pure iodine vapor has been designed and tested. The wind tunnel operates intermittently with a run phase lasting approximately 20 minutes. Successful recirculation of the iodine used during the run phase has been achieved but can be improved. Relevant issues regarding the full scale facility's design and operation, and the use of iodine as a working gas are discussed. Continuous wave laser induced fluorescence was used to monitor number densities within the plume flowfield, while pulsed laser induced fluorescence was used in an initial attempt to measure vibrational energy state population distributions. Preliminary nozzle flow calculations based on finite rate chemistry are presented.
Development of braided rope seals for hypersonic engine applications. Part 2: Flow modeling
NASA Technical Reports Server (NTRS)
Mutharasan, Rajakkannu; Steinetz, Bruce M.; Tao, Xiaoming; Ko, Frank
1991-01-01
Two models based on the Kozeny-Carmen equation were developed to analyze the fluid flow through a new class of braided rope seals under development for advanced hypersonic engines. A hybrid seal geometry consisting of a braided sleeve and a substantial amount of longitudinal fibers with high packing density was selected for development based on its low leakage rates. The models developed allow prediction of the gas leakage rate as a function of fiber diameter, fiber packing density, gas properties, and pressure drop across the seal.
Influence of Energy Input on the Flow Past Hypersonic Aircraft X-43
NASA Astrophysics Data System (ADS)
Khankhasaeva, Ya V.; E Borisov, V.; E Lutsky, A.
2017-02-01
This paper deals with a numerical study of the influence of energy sources on the flow past hypersonic aircraft X-43. Flight mode with M = 6 and angle of attack α = 0°, 4° with energy deposition in areas around various parts of HA was considered. It is shown that energy input in front of the bow of the HA leads to a significant weakening of the bow shock wave and an increase in aerodynamic efficiency of the vehicle. The results of studies on the impact of energy input in the scramjet intake are also presented.
Temperature measurements in hypersonic air flows using laser-induced O2 fluorescence
NASA Technical Reports Server (NTRS)
Laufer, Gabriel; Mckenzie, Robert L.
1988-01-01
An investigation is reported of the use of laser-induced fluorescence on oxygen for the measurement of air temperature and its fluctuations owing to turbulence in hypersonic wind tunnel flows. The results show that for temperatures higher than 60 K and densities higher than 0.01 amagat, the uncertainty in the temperature measurement can be less than 2 percent if it is limited by photon-statistical noise. The measurement is unaffected by collisional quenching and, if the laser fluence is kept below 1.5 J/sq cm, it is also unaffected by nonlinear effects which are associated with depletion of the absorbing states.
An assessment and application of turbulence models for hypersonic flows
NASA Technical Reports Server (NTRS)
Coakley, T. J.; Viegas, J. R.; Huang, P. G.; Rubesin, M. W.
1990-01-01
The current approach to the Accurate Computation of Complex high-speed flows is to solve the Reynolds averaged Navier-Stokes equations using finite difference methods. An integral part of this approach consists of development and applications of mathematical turbulence models which are necessary in predicting the aerothermodynamic loads on the vehicle and the performance of the propulsion plant. Computations of several high speed turbulent flows using various turbulence models are described and the models are evaluated by comparing computations with the results of experimental measurements. The cases investigated include flows over insulated and cooled flat plates with Mach numbers ranging from 2 to 8 and wall temperature ratios ranging from 0.2 to 1.0. The turbulence models investigated include zero-equation, two-equation, and Reynolds-stress transport models.
Translation-vibration-dissociation coupling in nonequilibrium hypersonic flows
NASA Technical Reports Server (NTRS)
Candler, Graham
1989-01-01
A new simple and computationally efficient model was developed, describing the evolution of vibrational states during relaxation and dissociation. The model is based on dividing the nitrogen molecules into two types, those in the vibrational states at a lower level, whose vibrational energy is below a cutoff energy, and those in an upper level, with vibrational energy above the cutoff. Dissociation occurs at the upper level, and recombination returns molecules to the lower level. The model was applied to two flows of engineering interest, the flow through a normal Mach 15 shock wave at 60 km, and a supersonic quasi-one-dimensional flow in a nozzle. Results are compared to those obtained by existing translation-vibration-dissociation coupling models, with results indicating significant differences between the models.
Translation-vibration-dissociation coupling in nonequilibrium hypersonic flows
NASA Technical Reports Server (NTRS)
Candler, Graham
1989-01-01
A new simple and computationally efficient model was developed, describing the evolution of vibrational states during relaxation and dissociation. The model is based on dividing the nitrogen molecules into two types, those in the vibrational states at a lower level, whose vibrational energy is below a cutoff energy, and those in an upper level, with vibrational energy above the cutoff. Dissociation occurs at the upper level, and recombination returns molecules to the lower level. The model was applied to two flows of engineering interest, the flow through a normal Mach 15 shock wave at 60 km, and a supersonic quasi-one-dimensional flow in a nozzle. Results are compared to those obtained by existing translation-vibration-dissociation coupling models, with results indicating significant differences between the models.
NASA Technical Reports Server (NTRS)
Sehgal, A. K.; Tiwari, S. N.; Singh, D. J.
1991-01-01
Hypersonic flows over cones and straight biconic configurations are calculated for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. Effect of angle of attack and nose bluntness on these slender cones in air is studied extensively. The numerical procedures are based on the solution of complete Navier-Stokes equations at the nose section and parabolized Navier-Stokes equations further downstream. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The complete flow field is thoroughly analyzed with respect to velocity, temperature, pressure, and entropy profiles. The post shock flow field is studied in detail from the contour plots of Mach number, density, pressure, and temperature. The effect of nose bluntness for slender cones persists as far as 200 nose radii downstream.
Flow separation in shock wave boundary layer interactions at hypersonic speeds
NASA Technical Reports Server (NTRS)
Hamed, A.
1990-01-01
An assessment is presented for the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists mainly of two dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the subsequent changes in the flow empirical correlations for incipient separation are reviewed. The effects of the Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed. The pertinent experimental data for the separated flow characteristics in separated turbulent boundary layer shock interaction are also presented and discussed.
A Structured-Grid Quality Measure for Simulated Hypersonic Flows
NASA Technical Reports Server (NTRS)
Alter, Stephen J.
2004-01-01
A structured-grid quality measure is proposed, combining three traditional measurements: intersection angles, stretching, and curvature. Quality assesses whether the grid generated provides the best possible tradeoffs in grid stretching and skewness that enable accurate flow predictions, whereas the grid density is assumed to be a constraint imposed by the available computational resources and the desired resolution of the flow field. The usefulness of this quality measure is assessed by comparing heat transfer predictions from grid convergence studies for grids of varying quality in the range of [0.6-0.8] on an 8'half-angle sphere-cone, at laminar, perfect gas, Mach 10 wind tunnel conditions.
Molecule-based approach for computing chemical-reaction rates in upper atmosphere hypersonic flows.
Gallis, Michail A.; Bond, Ryan Bomar; Torczynski, John Robert
2009-08-01
This report summarizes the work completed during FY2009 for the LDRD project 09-1332 'Molecule-Based Approach for Computing Chemical-Reaction Rates in Upper-Atmosphere Hypersonic Flows'. The goal of this project was to apply a recently proposed approach for the Direct Simulation Monte Carlo (DSMC) method to calculate chemical-reaction rates for high-temperature atmospheric species. The new DSMC model reproduces measured equilibrium reaction rates without using any macroscopic reaction-rate information. Since it uses only molecular properties, the new model is inherently able to predict reaction rates for arbitrary nonequilibrium conditions. DSMC non-equilibrium reaction rates are compared to Park's phenomenological non-equilibrium reaction-rate model, the predominant model for hypersonic-flow-field calculations. For near-equilibrium conditions, Park's model is in good agreement with the DSMC-calculated reaction rates. For far-from-equilibrium conditions, corresponding to a typical shock layer, the difference between the two models can exceed 10 orders of magnitude. The DSMC predictions are also found to be in very good agreement with measured and calculated non-equilibrium reaction rates. Extensions of the model to reactions typically found in combustion flows and ionizing reactions are also found to be in very good agreement with available measurements, offering strong evidence that this is a viable and reliable technique to predict chemical reaction rates.
An approximate method for calculating three-dimensional inviscid hypersonic flow fields
NASA Technical Reports Server (NTRS)
Riley, Christopher J.; Dejarnette, Fred R.
1990-01-01
An approximate solution technique was developed for 3-D inviscid, hypersonic flows. The method employs Maslen's explicit pressure equation in addition to the assumption of approximate stream surfaces in the shock layer. This approximation represents a simplification to Maslen's asymmetric method. The present method presents a tractable procedure for computing the inviscid flow over 3-D surfaces at angle of attack. The solution procedure involves iteratively changing the shock shape in the subsonic-transonic region until the correct body shape is obtained. Beyond this region, the shock surface is determined using a marching procedure. Results are presented for a spherically blunted cone, paraboloid, and elliptic cone at angle of attack. The calculated surface pressures are compared with experimental data and finite difference solutions of the Euler equations. Shock shapes and profiles of pressure are also examined. Comparisons indicate the method adequately predicts shock layer properties on blunt bodies in hypersonic flow. The speed of the calculations makes the procedure attractive for engineering design applications.
NASA Astrophysics Data System (ADS)
Gao, WenZhi; Li, ZhuFei; Yang, JiMing
2015-10-01
A hybrid CFD/characteristic method (CCM) was proposed for fast design and evaluation of hypersonic inlet flow with nose bluntness, which targets the combined advantages of CFD and method of characteristics. Both the accuracy and efficiency of the developed CCM were verified reliably, and it was well demonstrated for the external surfaces design of a hypersonic forebody/inlet with nose bluntness. With the help of CCM method, effects of nose bluntness on forebody shock shapes and the flowfield qualities which dominate inlet performance were examined and analyzed on the two-dimensional and axisymmetric configurations. The results showed that blunt effects of a wedge forebody are more substantial than that of related cone cases. For a conical forebody with a properly blunted nose, a recovery of the shock front back to that of corresponding sharp nose is exhibited, accompanied with a gradually fading out of entropy layer effects. Consequently a simplification is thought to be reasonable for an axisymmetric inlet with a proper compression angle, and a blunt nose of limited radius can be idealized as a sharp nose, as the spillage and flow variations at the entrance are negligible, even though the nose scale increases to 10% cowl lip radius. Whereas for two-dimensional inlets, the blunt effects are substantial since not only the inlet capturing/starting capabilities, but also the flow uniformities are obviously degraded.
NASA Technical Reports Server (NTRS)
Cockrell, Charles E., Jr.; Huebner, Lawrence D.; Finley, Dennis B.
1995-01-01
The component integration of a class of hypersonic high-lift configurations known as waveriders into hypersonic cruise vehicles was evaluated. A wind-tunnel model was developed which integrates realistic vehicle components with two waverider shapes, referred to as the straight-wing and cranked-wing shapes. Both shapes were conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) predictions were obtained over a Mach number range of 1.6 to 4.63 at a Reynolds number of 2.0 x 10(exp 6) per foot. The CFD predictions and flow visualization data confirmed the shock attachment characteristics of the baseline waverider shapes and illustrated the waverider flow-field properties. Experimental data showed that no significant performance degradations, in terms of maximum lift-to-drag ratios, occur at off-design Mach numbers for the waverider shapes and the integrated configurations. A comparison of the fully-integrated waverider vehicles to the baseline shapes showed that the performance was significantly degraded when all of the components were added to the waveriders, with the most significant degradation resulting from aftbody closure and the addition of control surfaces. Both fully-integrated configurations were longitudinally unstable over the Mach number range studied with the selected center of gravity location and for unpowered conditions. The cranked-wing configuration provided better lateral-directional stability characteristics than the straight-wing configuration.
Effect of dielectric barrier discharge plasma actuators on non-equilibrium hypersonic flows
NASA Astrophysics Data System (ADS)
Bhatia, Ankush; Roy, Subrata; Gosse, Ryan
2014-10-01
A numerical study employing discontinuous Galerkin method demonstrating net surface heat reduction for a cylindrical body in Mach 17 hypersonic flow is presented. This application focuses on using sinusoidal dielectric barrier discharge plasma actuators to inject momentum near the stagnation point. A 5 species finite rate air chemistry model completes the picture by analyzing the effect of the actuator on the flow chemistry. With low velocity near the stagnation point, the plasma actuator sufficiently modifies the fluid momentum. This results in redistribution of the integrated surface heating load on the body. Specifically, a particular configuration of normally pinching plasma actuation is predicted to reduce the surface heat flux at the stagnation point. An average reduction of 0.246% for the integrated and a maximum reduction of 7.68% are reported for the surface heat flux. The temperature contours in the fluid flow (with maximum temperature over 12 000 K) are pinched away from the stagnation point, thus resulting in reduced thermal load. Plasma actuation in this configuration also affects the species concentration distribution near the wall, in addition to the temperature gradient. The combined effect of both, thus results in an average reduction of 0.0986% and a maximum reduction of 4.04% for non-equilibrium calculations. Thus, this study successfully demonstrates the impact of sinusoidal dielectric barrier discharge plasma actuation on the reduction of thermal load on a hypersonic body.
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
1989-01-01
The code development and application program for the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA), with emphasis directed toward support of the Aeroassist Flight Experiment (AFE) in the near term and Aeroassisted Space Transfer Vehicle (ASTV) design in the long term is reviewed. LAURA is an upwind-biased, point-implicit relaxation algorithm for obtaining the numerical solution to the governing equations for 3-D, viscous, hypersonic flows in chemical and thermal nonequilibrium. The algorithm is derived using a finite volume formulation in which the inviscid components of flux across cell walls are described with Roe's averaging and Harten's entropy fix with second-order corrections based on Yee's Symmetric Total Variation Diminishing scheme. Because of the point-implicit relaxation strategy, the algorithm remains stable at large Courant numbers without the necessity of solving large, block tri-diagonal systems. A single relaxation step depends only on information from nearest neighbors. Predictions for pressure distributions, surface heating, and aerodynamic coefficients compare well with experimental data for Mach 10 flow over an AFE wind tunnel model. Predictions for the hypersonic flow of air in chemical and thermal nonequilibrium over the full scale AFE configuration obtained on a multi-domain grid are discussed.
SUPERSONIC AND HYPERSONIC INTERFERENCE FLOW FIELDS AND HEATING
NASA Technical Reports Server (NTRS)
Morris, D. J.
1994-01-01
Small areas of high heat transfer and pressure can occur on a vehicle surface due to the influence of an impinging shock on the local flow. A method was needed to determine peak pressure and heating of these areas. This package is a system of computer programs designed to calculate two-dimensional shock interference patterns for six types of interference flows. Results also include properties of the inviscid flow field and the inviscid-viscous interaction at the surface along with peak pressure and peak heating at the impingement point. The six types of interference flow patterns considered are: 1) Type I interference patterns, occurring when two weak shocks of opposite families, BS (bow shock) and IS (impingment shock), intersect when the flow upstream of the impingement point is supersonic, or in the case of a blunt body, takes place well below the sonic point. 2) Type II interference pattern occurs when two shocks of opposite families (bow shock and impinging shock) intersect. Both shocks are weak as in type I, but are of such strength that in order to turn the flow, a Mach reflection must exist in the center of the flow field with an embedded subsonic region occurring between the intersection points (A & B) and the accompanying shear layers. Type II interference occurs on a blunt body when the impinging shock intersects the bow shock near the sonic point. 3) Type III shock interference pattern occurs when a weak impinging shock intersects a strong detached bow shock. On a blunt body the shock intersection occurs near or above the lower sonic point. 4) Type IV interference can occur when the impinging shock intersects a strong bow shock ahead of a subsonic flow region. On a blunt body this shock intersection is located between the lower sonic point and just above the body axis. The impinging shock causes a displacement of the bow shock and the formation of a supersonic jet that is embedded in the subsonic region. A jet bow shock is produced when the jet impinges
DSMC Simulations of Hypersonic Flows and Comparison With Experiments
NASA Technical Reports Server (NTRS)
Moss, James N.; Bird, Graeme A.; Markelov, Gennady N.
2004-01-01
This paper presents computational results obtained with the direct simulation Monte Carlo (DSMC) method for several biconic test cases in which shock interactions and flow separation-reattachment are key features of the flow. Recent ground-based experiments have been performed for several biconic configurations, and surface heating rate and pressure measurements have been proposed for code validation studies. The present focus is to expand on the current validating activities for a relatively new DSMC code called DS2V that Bird (second author) has developed. Comparisons with experiments and other computations help clarify the agreement currently being achieved between computations and experiments and to identify the range of measurement variability of the proposed validation data when benchmarked with respect to the current computations. For the test cases with significant vibrational nonequilibrium, the effect of the vibrational energy surface accommodation on heating and other quantities is demonstrated.
Downstream influence of swept slot injection in hypersonic turbulent flow
NASA Technical Reports Server (NTRS)
Hefner, J. N.; Cary, A. M., Jr.; Bushnell, D. B.
1977-01-01
Results of an experimental and numerical investigation of tangential swept slot injection into a thick turbulent boundary layer at Mach 6 are presented. Film cooling effectiveness, skin friction, and flow structure downstream of the swept slot injection were investigated. The data were compared with that for unswept slots, and it was found that cooling effectiveness and skin friction reductions are not significantly affected by sweeping the slot.
Ultrasonically Absorptive Coatings for Hypersonic Laminar Flow Control
2007-12-01
20503. 1. AGENCY USE ONLY ( Leave Blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED February 2008 5. FUNDING NUMBERS 4. TITLE AND SUBTITLE...efforl nclude theoretical analysis, direct numerical simulation (DNS), wind-tunnel experiments, as well as fabrication of ceramic materials that...increase of the laminar run. First samples of a ceramic UAC integrated into TPS tile were fabricated using a stampinj echnique. Benchmark (no flow
NASA Astrophysics Data System (ADS)
Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B.; Pereira, A. L.; Nagamatsu, H. T.
2006-05-01
A new 0.6-m. diameter Hypersonic Shock Tunnel is been designed, fabricated and will be installed at the Laboratory of Aerothermodynamics and Hypersonics IEAv-CTA, Brazil. The brand new hypersonic facility, designated as T3, is primarily intended to be used as an important tool in the investigation of supersonic combustion management and of electromagnetic energy addition for flow control. The design of the runnel enables relatively long test times, 2-10 milliseconds, suitable for basic supersonic combustion and energy addition by laser experiments. Free stream Mach numbers ranging from 6 to 25 can be produced and stagnation pressures and temperatures of 200 atm. and 5,500 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization and the new facility is expected to be commissioned by the end of 2006.
Pulsed Electron Beam Spectroscopy for Temperature Measurements in Hypersonic Flows
2010-01-01
atmospheric pr essures wit hin the fligh t envelope of scramjet-powered flight vehicles. Because of the pressure disparity between measured flow and me...represents what might be o btained from the pulse d e-beam s ystem if it were used in the high-te mperature (but high-pr essure ) st agnation cha...di fferential pressure pump has been developed for pressure separations up to approximately 1 torr. F or higher pr essures , a f ast act ion r otary
Optimum shape of a blunt forebody in hypersonic flow
NASA Technical Reports Server (NTRS)
Maestrello, L.; Ting, L.
1989-01-01
The optimum shape of a blunt forebody attached to a symmetric wedge or cone is determined. The length of the forebody, its semi-thickness or base radius, the nose radius and the radius of the fillet joining the forebody to the wedge or cone are specified. The optimum shape is composed of simple curves. Thus experimental models can be built readily to investigate the utilization of aerodynamic heating for boundary layer control. The optimum shape based on the modified Newtonian theory can also serve as the preliminary shape for the numerical solution of the optimum shape using the governing equations for a compressible inviscid or viscous flow.
Parabolized Navier-Stokes methods for hypersonic flows
NASA Technical Reports Server (NTRS)
Lawrence, Scott L.
1991-01-01
A representative sampling of the techniques used in the integration of the Parabolized Navier-Stokes (PNS) equations is presented. Special atention is given to recent algorithms developed specifically for application to high speed flows, characterized by the presence of strong embedded shock waves and real gas effects. It is shown that PNS solvers are being used in the analysis of sonic boom signatures. Methods for modeling physical effects are discussed, including an overview of commonly used turbulence models and a more detailed discussion of techniques for including equilibrium and finite rate real gas effects.
NASA Technical Reports Server (NTRS)
Steinetz, Bruce M.; Mutharasan, Rajakkannu; Du, Guang-Wu; Miller, Jeffrey H.; Ko, Frank
1992-01-01
A critical mechanical system in advanced hypersonic engines is the panel-edge seal system that seals gaps between the articulating horizontal engine panels and the adjacent engine splitter walls. Significant advancements in seal technology are required to meet the extreme demands placed on the seals, including the simultaneous requirements of low leakage, conformable, high temperature, high pressure, sliding operation. In this investigation, the seal concept design and development of two new seal classes that show promise of meeting these demands will be presented. These seals include the ceramic wafer seal and the braided ceramic rope seal. Presented are key elements of leakage flow models for each of these seal types. Flow models such as these help designers to predict performance-robbing parasitic losses past the seals, and estimate purge coolant flow rates. Comparisons are made between measured and predicted leakage rates over a wide range of engine simulated temperatures and pressures, showing good agreement.
Aerothermodynamics of nozzle flows for advanced hypersonic propulsion systems
NASA Astrophysics Data System (ADS)
Weiland, C.; Hartmann, G.; Menne, S.
1992-02-01
One of the major tasks for the development of novel airbreathing space transportation systems, operating from usual airports by horizontal take off and landing, is the integration of an advanced propulsion system in the cell of that spacecraft. The air intake and in particular the free expansion nozzle affect not only the efficiency of the engine but also the forces and moments, and with that, the control of the complete spacecraft. Therefore, it is necessary to know in detail the flow fields through such nozzles and its interaction with the external airflow. Another project deals with conventional rocket motor nozzles whereby injection of turbine exhaust gases in the expansion part of the nozzle the wall of the nozzle is cooled (filmcooling concept) and the thrust is slightly increased. Theoretical investigation of these and other nozzles is the objective of this paper. Euler and boundary layer methods will be applied to predict the flow fields of the nozzles where special emphasis is laid on the consideration of real gas effects. The theory of the Euler method will be described in detail while for the second order boundary layer method the governing equations are presented and the range of its applicability is shortly discussed. Finally results for a variety of nozzles will be given.
Planar Laser-Induced Iodine Fluorescence Measurements in Rarefied Hypersonic Flow
NASA Technical Reports Server (NTRS)
Cecil, Eric; McDaniel, James C.
2005-01-01
A planar laser-induced fluorescence (PLIF) technique is discussed and applied to measurement of time-averaged values of velocity and temperature in an I(sub 2)-seeded N(sub 2) hypersonic free jet facility. Using this technique, a low temperature, non-reacting, hypersonic flow over a simplified model of a reaction control system (RCS) was investigated. Data are presented of rarefied Mach 12 flow over a sharp leading edge flat plate at zero incidence, both with and without an interacting jet issuing from a nozzle built into the plate. The velocity profile in the boundary layer on the plate was resolved. The slip velocity along the plate, extrapolated from the velocity profile data, varied from nearly 100% down to 10% of the freestream value. These measurements are compared with results of a DSMC solution. The velocity variation along the centerline of a jet issuing from the plate was measured and found to match closely with the correlation of Ashkenas and Sherman. The velocity variation in the oblique shock terminating the jet was resolved sufficiently to measure the shock wave thickness.
Schlieren Visualization of the Energy Addition by Multi Laser Pulse in Hypersonic Flow
Oliveira, A. C.; Minucci, M. A. S.; Toro, P. G. P.; Chanes, J. B. Jr; Myrabo, L. N.
2008-04-28
The experimental results of the energy addition by multi laser pulse in Mach 7 hypersonic flow are presented. Two high power pulsed CO{sub 2} TEA lasers (TEA1 5.5 J, TEA2 3.9 J) were assembled sharing the same optical cavity to generate the plasma upstream of a hemispherical model installed in the tunnel test section. The lasers can be triggered with a selectable time delay and in the present report the results obtained with delay between 30 {mu}s and 80 {mu}s are shown. The schlieren technique associated with a high speed camera was used to accomplish the influence of the energy addition in the mitigation of the shock wave formed on the model surface by the hypersonic flow. A piezoelectric pressure transducer was used to obtain the time history of the impact pressure at stagnation point of the model and the pressure reduction could be measured. The total recovery of the shock wave between pulses as well as the prolonged effect of the mitigation without recovery was observed by changing the delay.
NASA Technical Reports Server (NTRS)
Bathel, Brett F.; Danehy, Paul M.; Inman, Jennifer A.; Jones, Stephen B.; Ivey,Christopher b.; Goyne, Christopher P.
2010-01-01
Nitric-oxide planar laser-induced fluorescence (NO PLIF) was used to perform velocity measurements in hypersonic flows by generating multiple tagged lines which fluoresce as they convect downstream. For each laser pulse, a single interline, progressive scan intensified CCD (charge-coupled device) camera was used to obtain two sequential images of the NO molecules that had been tagged by the laser. The CCD configuration allowed for sub-microsecond acquisition of both images, resulting in sub-microsecond temporal resolution as well as sub-mm spatial resolution (0.5-mm horizontal, 0.7-mm vertical). Determination of axial velocity was made by application of a cross-correlation analysis of the horizontal shift of individual tagged lines. A numerical study of measured velocity error due to a uniform and linearly-varying collisional rate distribution was performed. Quantification of systematic errors, the contribution of gating/exposure duration errors, and the influence of collision rate on temporal uncertainty were made. Quantification of the spatial uncertainty depended upon the signal-to-noise ratio of the acquired profiles. This velocity measurement technique has been demonstrated for two hypersonic flow experiments: (1) a reaction control system (RCS) jet on an Orion Crew Exploration Vehicle (CEV) wind tunnel model and (2) a 10-degree half-angle wedge containing a 2-mm tall, 4-mm wide cylindrical boundary layer trip. The experiments were performed at the NASA Langley Research Center's 31-Inch Mach 10 Air Tunnel.
NASA Technical Reports Server (NTRS)
Bathel, Brett F.; Danehy, Paul M.; Inmian, Jennifer A.; Jones, Stephen B.; Ivey, Christopher B.; Goyne, Christopher P.
2010-01-01
Nitric-oxide planar laser-induced fluorescence (NO PLIF) was used to perform velocity measurements in hypersonic flows by generating multiple tagged lines which fluoresce as they convect downstream. For each laser pulse, a single interline, progressive scan intensified CCD camera was used to obtain separate images of the initial undelayed and delayed NO molecules that had been tagged by the laser. The CCD configuration allowed for sub-microsecond acquisition of both images, resulting in sub-microsecond temporal resolution as well as sub-mm spatial resolution (0.5-mm x 0.7-mm). Determination of axial velocity was made by application of a cross-correlation analysis of the horizontal shift of individual tagged lines. Quantification of systematic errors, the contribution of gating/exposure duration errors, and influence of collision rate on fluorescence to temporal uncertainty were made. Quantification of the spatial uncertainty depended upon the analysis technique and signal-to-noise of the acquired profiles. This investigation focused on two hypersonic flow experiments: (1) a reaction control system (RCS) jet on an Orion Crew Exploration Vehicle (CEV) wind tunnel model and (2) a 10-degree half-angle wedge containing a 2-mm tall, 4-mm wide cylindrical boundary layer trip. The experiments were performed at the NASA Langley Research Center's 31-inch Mach 10 wind tunnel.
Coupled computational fluid-thermal investigation of hypersonic flow over a quilted dome surface
NASA Astrophysics Data System (ADS)
Ostoich, Christopher; Bodony, Daniel; Geubelle, Philippe
2009-11-01
The hypersonic environment is characterized by the high temperatures that are generated in the fluid at a vehicle surface. In the effort to enable the operation of lightweight, reusable hypersonic vehicles, flexible, thin thermal protection panels have been considered to mitigate thermal loads. High surface temperatures create through-the-thickness thermal gradients which cause the panels to bow, resulting in changes to the external flow field and leading to a fully coupled fluid-thermal-structural problem. Certain aspects of the fluid-thermal (no structural) coupling were examined in a 1980s NASA Langley experiment of a Mach 5.74 laminar boundary past an array of spherical domes. We reexamine this case computationally using a high-fidelity Navier-Stokes solver coupled with a thermal solver to investigate the effects on the flow and resulting heat load on the structure due to the bowed panels. Specifically the surface temperature, surface heat flux, and downstream boundary developments are reported, and compared with experiment.
Sensitivity analysis of DSMC parameters for an 11-species air hypersonic flow
NASA Astrophysics Data System (ADS)
Higdon, Kyle J.; Goldstein, David B.; Varghese, Philip L.
2016-11-01
This research investigates the influence of input parameters in the direct simulation Monte Carlo (DSMC) method for the simulation of a hypersonic flow scenario. Simulations are performed using the Computation of Hypersonic Ionizing Particles in Shocks (CHIPS) code to reproduce NASA Ames Electric Arc Shock Tube (EAST) experimental results for a 10.26 km/s, 0.2 Torr scenario. Since the chosen nominal simulation involves an energetic flow, an electronic excitation model is introduced into CHIPS to complement the pre-existing 11-species air models. A global Monte Carlo sensitivity analysis was completed for this chosen scenario and three quantities of interest (QoIs) were investigated: translational temperature, electronic temperature, and electron number density. The electron impact ionization reaction, N + e- ⇌ N+ + e- + e-, was determined to have the greatest effect on all three QoIs as it defines the electron cascade that occurs post-shock. In addition, molecular nitrogen dissociation, associative ionization, and the N + NO+ ⇌ N+ + NO charge exchange reaction were all found to be important for these QoIs.
Planar Laser-Induced Iodine Fluorescence Measurements in Rarefied Hypersonic Flow
NASA Astrophysics Data System (ADS)
Cecil, Eric; McDaniel, James C.
2005-05-01
A planar laser-induced fluorescence (PLIF) technique is discussed and applied to measurement of time-averaged values of velocity and temperature in an I2-seeded N2 hypersonic free jet facility. Using this technique, a low temperature, non-reacting, hypersonic flow over a simplified model of a reaction control system (RCS) was investigated. Data are presented of rarefied Mach 12 flow over a sharp leading edge flat plate at zero incidence, both with and without an interacting jet issuing from a nozzle built into the plate. The velocity profile in the boundary layer on the plate was resolved. The slip velocity along the plate, extrapolated from the velocity profile data, varied from nearly 100% down to 10% of the freestream value. These measurements are compared with results of a DSMC solution. The velocity variation along the centerline of a jet issuing from the plate was measured and found to match closely with the correlation of Ashkenas and Sherman. The velocity variation in the oblique shock terminating the jet was resolved sufficiently to measure the shock wave thickness.
Hypersonic Separated Flows About "Tick" Configurations With Sensitivity to Model Design
NASA Technical Reports Server (NTRS)
Moss, J. N.; O'Byrne, S.; Gai, S. L.
2014-01-01
This paper presents computational results obtained by applying the direct simulation Monte Carlo (DSMC) method for hypersonic nonequilibrium flow about "tick-shaped" model configurations. These test models produces a complex flow where the nonequilibrium and rarefied aspects of the flow are initially enhanced as the flow passes over an expansion surface, and then the flow encounters a compression surface that can induce flow separation. The resulting flow is such that meaningful numerical simulations must have the capability to account for a significant range of rarefaction effects; hence the application of the DSMC method in the current study as the flow spans several flow regimes, including transitional, slip, and continuum. The current focus is to examine the sensitivity of both the model surface response (heating, friction and pressure) and flowfield structure to assumptions regarding surface boundary conditions and more extensively the impact of model design as influenced by leading edge configuration as well as the geometrical features of the expansion and compression surfaces. Numerical results indicate a strong sensitivity to both the extent of the leading edge sharpness and the magnitude of the leading edge bevel angle. Also, the length of the expansion surface for a fixed compression surface has a significant impact on the extent of separated flow.
Hypersonic separated flows about "tick" configurations with sensitivity to model design
NASA Astrophysics Data System (ADS)
Moss, J. N.; O'Byrne, S.; Gai, S. L.
2014-12-01
This paper presents computational results obtained by applying the direct simulation Monte Carlo (DSMC) method for hypersonic nonequilibrium flow about "tick-shaped" model configurations. These test models produces a complex flow where the nonequilibrium and rarefied aspects of the flow are initially enhanced as the flow passes over an expansion surface, and then the flow encounters a compression surface that can induce flow separation. The resulting flow is such that meaningful numerical simulations must have the capability to account for a significant range of rarefaction effects; hence the application of the DSMC method in the current study as the flow spans several flow regimes, including transitional, slip, and continuum. The current focus is to examine the sensitivity of both the model surface response (heating, friction and pressure) and flowfield structure to assumptions regarding surface boundary conditions and more extensively the impact of model design as influenced by leading edge configuration as well as the geometrical features of the expansion and compression surfaces. Numerical results indicate a strong sensitivity to both the extent of the leading edge sharpness and the magnitude of the leading edge bevel angle. Also, the length of the expansion surface for a fixed compression surface has a significant impact on the extent of separated flow.
An engineering aerodynamic heating method for hypersonic flow
NASA Technical Reports Server (NTRS)
Riley, Christopher J.; Dejarnette, Fred R.
1992-01-01
A capability to calculate surface heating rates has been incorporated in an approximate three-dimensional inviscid technique. Surface streamlines are calculated from the inviscid solution, and the axisymmetric analog is then used along with a set of approximate convective-heating equations to compute the surface heat transfer. The method is applied to blunted axisymmetric and three-dimensional ellipsoidal cones at angle of attack for the laminar flow of a perfect gas. The method is also applicable to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes (NS) and viscous shock-layer (VSL) equations. The new technique represents a significant improvement over current engineering aerothermal methods with only a modest increase in computational effort.
An Engineering Aerodynamic Heating Method for Hypersonic Flow
NASA Technical Reports Server (NTRS)
Riley, Christopher J.; DeJarnette, Fred R.
1992-01-01
A capability to calculate surface heating rates has been incorporated in an approximate three-dimensional inviscid technique. Surface streamlines are calculated from the inviscid solution, and the axisymmetric analog is then used along with a set of approximate convective-heating equations to compute the surface heat transfer. The method is applied to blunted axisymmetric and three-dimensional ellipsoidal cones at angle of attack for the laminar flow of a perfect gas. The method is also applicable to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes (NS) and viscous shock-layer (VSL) equations. The new technique represents a significant improvement over current engineering aerothermal methods with only a modest increase in computational effort.
An engineering aerodynamic heating method for hypersonic flow
NASA Technical Reports Server (NTRS)
Riley, Christopher J.; Dejarnette, Fred R.
1992-01-01
A capability to calculate surface heating rates has been incorporated in an approximate three-dimensional inviscid technique. Surface streamlines are calculated from the inviscid solution, and the axisymmetric analog is then used along with a set of approximate convective-heating equations to compute the surface heat transfer. The method is applied to blunted axisymmetric and three-dimensional ellipsoidal cones at angle of attack for the laminar flow of a perfect gas. The method is also applicable to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes (NS) and viscous shock-layer (VSL) equations. The new technique represents a significant improvement over current engineering aerothermal methods with only a modest increase in computational effort.
NASA Technical Reports Server (NTRS)
Huebner, Lawrence D.; Kniskern, Marc W.; Monta, William J.
1993-01-01
The purpose of this investigation were twofold: first, to determine whether accurate force and moment data could be obtained during hypersonic wind tunnel tests of a model with a scramjet exhaust flow simulation that uses a representative nonwatercooled, flow-through balance; second, to analyze temperature time histories on various parts of the balance to address thermal effects on force and moment data. The tests were conducted in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel at free-stream Reynolds numbers ranging from 0.5 to 7.4 x 10(exp 6)/ft and nominal angles of attack of -3.5 deg, 0 deg, and 5 deg. The simulant exhaust gases were cold air, hot air, and a mixture of 50 percent Argon and 50 percent Freon by volume, which reached stagnation temperatures within the balance of 111, 214, and 283 F, respectively. All force and moment values were unaffected by the balance thermal response from exhaust gas simulation and external aerodynamic heating except for axial-force measurements, which were significantly affected by balance heating. This investigation showed that for this model at the conditions tested, a nonwatercooled, flow-through balance is not suitable for axial-force measurements during scramjet exhaust flow simulation tests at hypersonic speeds. In general, heated exhaust gas may produce unacceptable force and moment uncertainties when used with thermally sensitive balances.
Hypersonic turbulent expansion-corner flow with shock impingement
NASA Technical Reports Server (NTRS)
Chung, Kung-Ming; Lu, Frank K.
1992-01-01
Mean and fluctuating surface pressure data were obtained in a Mach 8, turbulent, cold flow past an expansion corner subjected to shock impingement. The expansion corner of 2.5 or 4.25 deg was located at 0.77 m (30.25 in.) from the leading edge of a shape-edged flat plate while an external shock, generated by either a 2- or 4-deg sharp wedge, impinged at the corner, or at one boundary layer thickness ahead or behind the corner. The mean pressure distribution was strongly influenced by the mutual interaction between the shock and the expansion. For example, the upstream influence decreased when the shock impinged downstream of the corner. Also, the unsteadiness of the interactions was characterized by an intermittent region and a local rms pressure peak near the upstream influence line. The peak rms pressure fluctuations increased with a larger overall interaction strength. Shock impingement downstream of the corner resulted in lower peaks and also in a shorter region of reduced fluctuation levels. These features may be exploited in inlet design by impinging the cowl shock downstream of an expansion corner instead of at the corner. In addition, a limited Pitot pressure survey showed a thinning of the boundary layer downstream of the corner.
NASA Technical Reports Server (NTRS)
Ashby, George C.
1988-01-01
An experimental investigation of the design of pitot probes for flowfield surveys in hypersonic wind tunnels is reported. The results show that a pitot-pressure probe can be miniaturized for minimum interference effects by locating the transducer in the probe support body and water-cooling it so that the pressure-settling time and transducer temperature are compatible with hypersonic tunnel operation and flow conditions. Flowfield surveys around a two-to-one elliptical cone model in a 20-inch Mach 6 wind tunnel using such a probe show that probe interference effects are essentially eliminated.
NASA Technical Reports Server (NTRS)
Hamaker, Frank M; Neice, Stanford E; Wong, Thomas J
1953-01-01
The similarity law for nonsteady, inviscid, hypersonic flow about slender three-dimensional shapes is derived. Conclusions drawn are shown to be valid for rotational flow. Requirements for dynamic similarity of related shapes in free flight are obtained. The law is examined for steady flow about related three-dimensional shapes. Results of an experimental investigation of the pressures acting on two inclined cones are found to check the law as it applies to bodies of revolution.
NASA Astrophysics Data System (ADS)
Hu, Z. M.; Myong, R. S.; Yang, Y. R.; Cho, T. H.
2010-12-01
Shock polar analysis as well as 2-D numerical computation technique are used to illustrate a diverse flow topology induced by shock/shock interaction in a M ∞ = 9 hypersonic flow. New flow features associated with inviscid shock wave interaction on double-wedge-like geometries are reported in this study. Transition of shock interaction, unsteady oscillation, and hysteresis phenomena in the RR ↔ MR transition, and the physical mechanisms behind these phenomena are numerically studied and analyzed.
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.; Johnston, Christopher O.; Thompson, Richard A.
2009-01-01
A description of models and boundary conditions required for coupling radiation and ablation physics to a hypersonic flow simulation is provided. Chemical equilibrium routines for varying elemental mass fraction are required in the flow solver to integrate with the equilibrium chemistry assumption employed in the ablation models. The capability also enables an equilibrium catalytic wall boundary condition in the non-ablating case. The paper focuses on numerical implementation issues using FIRE II, Mars return, and Apollo 4 applications to provide context for discussion. Variable relaxation factors applied to the Jacobian elements of partial equilibrium relations required for convergence are defined. Challenges of strong radiation coupling in a shock capturing algorithm are addressed. Results are presented to show how the current suite of models responds to a wide variety of conditions involving coupled radiation and ablation.
DSMC method on aerodynamic heating and temperature characteristic of hypersonic rarefied flows
NASA Astrophysics Data System (ADS)
Ma, Jing; Bao, Xingdong; Mao, Hongxia; Dong, Yanbing
2016-10-01
Aerodynamic heating is one of important factors affecting hypersonic aircraft design. The Direct Simulation Monte Carlo method (DSMC) has evolved years into a powerful numerical technique for the computation of complex, non-equilibrium gas flows. In atmospheric target, non-equilibrium conditions occur at high altitude and in regions of flow fields with small length scales. In this paper, the theoretical basis of the DSMC technique is discussed. In addition, the methods used in DSMC are described for simulation of high temperature, real gas effects and gas-surface interactions. Combined with the solution of heat transfer in material, heat-flux distribution and temperature distribution of the different shape structures was calculated in rarefied conditions.
Mean flow field and surface heating produced by unequal shock interactions at hypersonic speeds
NASA Technical Reports Server (NTRS)
Birch, S. F.; Rudy, D. H.
1975-01-01
Mean velocity profiles were measured in a free shear layer produced by the interaction of two unequal strength shock waves at hypersonic free-stream Mach numbers. Measurements were made over a unit Reynolds number range of 3,770,000 per meter to 17,400,000 per meter based on the flow on the high velocity side of the shear layer. The variation in measured spreading parameters with Mach number for the fully developed flows is consistent with the trend of the available zero velocity ratio data when the Mach numbers for the data given in this study are taken to be characteristic Mach numbers based on the velocity difference across the mixing layer. Surface measurements in the shear-layer attachment region of the blunt-body model indicate peak local heating and static pressure consistent with other published data. Transition Reynolds numbers were found to be significantly lower than those found in previous data.
NASA Technical Reports Server (NTRS)
Jones, R. A.; Hunt, J. L.
1973-01-01
An experimental study of surface pressure distributions on a family of blunt and sharp large angle cones was made in hypersonic flows of helium, air, and tetrafluoromethane. The effective isentropic exponents of these flows were 1.67, 1.40, and 1.12. Thus, the effect of large shock density ratios such as might be encountered during planetary entry because of real-gas effects could be studied by comparing results in tetrafluoromethane with those in air and helium. It was found that shock density ratio had a large effect on both shock shape and pressure distribution. The differences in pressure distribution indicate that for atmospheric flight at high speed where real-gas effects produce large shock density ratios, large-angle cone vehicles can be expected to experience different trim angles of attack, drag coefficient, and lift-drag ratios than those for ground tests in air wind tunnels.
Hypersonic rarefied flow about a delta wing - Direct simulation and comparison with experiment
NASA Technical Reports Server (NTRS)
Celenligil, M. C.; Moss, James N.
1991-01-01
Three-dimensional simulations of hypersonic rarefied flow about a delta wing are made using the direct simulation Monte Carlo (DSMC) method of Bird, and the results of the computations are compared with recent experimental data obtained in a vacuum wind tunnel at the DLR in Gottingen, Germany. The present study considers Mach 8.89 nitrogen flow for a range of conditions that include Knudsen numbers of 0.016 to 3.505 for an incidence angle of 30 deg, and angles of incidence of 15 to 60 deg for a constant Knudsen number of 0.389. The calculations provide details concerning the flowfield structure and surface quantities. Comparisons between the calculations and the available experimental measurements are made for aerodynamic and overall heat-transfer coefficients and recovery temperature. The agreement between the measured and calculated data are very good, well within the estimated measurement uncertainty. Comparisons are also made with modified Newtonian and free-molecule theories.
NASA Technical Reports Server (NTRS)
Cheatwood, F. M.; Dejarnette, F. R.
1992-01-01
An approximate axisymmetric method has been developed which can reliably calculate nonequilibrium fully viscous hypersonic flows over blunt-nosed bodies. By substituting Maslen's second-order pressure expression for the normal momentum equation, a simplified form of the viscous shock layer (VSL) equations is obtained. This approach can solve both the subsonic and supersonic regions of the shock layer without a starting solution for the shock shape. This procedure is significantly faster than the parabolized Navier-Stokes and VSL solvers and would be useful in a preliminary design environment. Solutions have been generated for air flows over several analytic body shapes. Surface heat transfer and pressure predictions are comparable to VSL results. Computed heating rates are in good agreement with experimental data. The present technique generates its own shock shape as part of its solution, and therefore could be used to provide more accurate initial shock shapes for higher-order procedures which require starting solutions.
An improved flux-split algorithm applied to hypersonic flows in chemical equilibrium
NASA Technical Reports Server (NTRS)
Palmer, Grant
1988-01-01
An explicit, finite-difference, shock-capturing numerical algorithm is presented and applied to hypersonic flows assumed to be in thermochemical equilibrium. Real-gas chemistry is either loosely coupled to the gasdynamics by way of a Gibbs free energy minimization package or fully coupled using species mass conservation equations with finite-rate chemical reactions. A scheme is developed that maintains stability in the explicit, finite-rate formulation while allowing relatively high time steps. The codes use flux vector splitting to difference the inviscid fluxes and employ real-gas corrections to viscosity and thermal conductivity. Numerical results are compared against existing ballistic range and flight data. Flows about complex geometries are also computed.
Interaction between Shock Wave and Boundary Layer in Nonequilibrium Hypersonic Rarefied Flow
NASA Astrophysics Data System (ADS)
Tsuboi, Nobuyuki; Matsumoto, Yoichiro
An experimental study of the interaction between a shock wave and a boundary layer over a flat plate with a sharp leading edge in hypersonic rarefied gas flow is presented. Experiments in a low-density wind tunnel using an electron beam probe were conducted at the Shock Wave Laboratory, RWTH Aachen, Germany. Rotational temperatures for stagnation temperatures of T0=670˜1000 K and Kn=0.024˜0.028 based on a reference length of 0.05m were calculated using Muntz’s method and Robben and Talbot’s method. The domain of quasi two-dimensional flow over the plate was determined from three-dimensional rotational temperature measurements. Nonequilibrium between translational and rotational temperatures was observed near the leading edge, and the experimental rotational relaxation length explains the rotational collision number of 2˜4.
Shock induced detonation on projectiles in hypersonic flows of detonable gas mixtures
NASA Astrophysics Data System (ADS)
Rom, Josef
1995-06-01
The following subjects were investigated: (1) CFD results on the External Propulsion Accelerator (EPA) projectile configurations; (2) an analytical study on the stability of hypersonic reacting flow at the stagnation region of a blunt body using dynamical system analysis; (3) the use of the EPA for scramjet combustion research; (4) the use of the EPA for hypersonic aerodynamic test facility; (5) analysis of the initiation of detonation on a hypervelocity projectile and it's maximum velocity in the EPA; and (6) preparations for testing at the Army Research Laboratory's ram accelerator facility at Aberdeen, MD. The CFD calculations on the projectile configurations indicated a well established external combustion zone and reasonably large thrust. Analysis using energy balance considerations indicated that the maximum projectile velocity in the EPA is about 6 times the detonation speed while that for the ram accelerator is about 1.3 times the detonation speed of the mixture. Therefore, the EPA is capable of accelerating missile-projectiles to beyond the escape velocity and can be considered also for single stage to orbit missions.
Hypersonic Engine Leading Edge Experiments in a High Heat Flux, Supersonic Flow Environment
NASA Technical Reports Server (NTRS)
Gladden, Herbert J.; Melis, Matthew E.
1994-01-01
A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Three aerothermal load related concerns are the boundary layer transition from laminar to turbulent flow, articulating panel seals in high temperature environments, and strut (or cowl) leading edges with shock-on-shock interactions. A multidisciplinary approach is required to address these technical concerns. A hydrogen/oxygen rocket engine heat source has been developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to experimentally evaluate the heat transfer and structural response of the strut (or cowl) leading edge. A recent experimental program conducted in this facility is discussed and related to cooling technology capability. The specific objective of the experiment discussed is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Heat transfer analyses of a similar leading edge concept cooled with gaseous hydrogen is included to demonstrate the complexity of the problem resulting from plastic deformation of the structures. Macro-photographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight.
Computational analysis of a rarefied hypersonic flow over combined gap/step geometries
NASA Astrophysics Data System (ADS)
Leite, P. H. M.; Santos, W. F. N.
2015-06-01
This work describes a computational analysis of a hypersonic flow over a combined gap/step configuration at zero degree angle of attack, in chemical equilibrium and thermal nonequilibrium. Effects on the flowfield structure due to changes on the step frontal-face height have been investigated by employing the Direct Simulation Monte Carlo (DSMC) method. The work focuses the attention of designers of hypersonic configurations on the fundamental parameter of surface discontinuity, which can have an important impact on even initial designs. The results highlight the sensitivity of the primary flowfield properties, velocity, density, pressure, and temperature due to changes on the step frontal-face height. The analysis showed that the upstream disturbance in the gap/step configuration increased with increasing the frontal-face height. In addition, it was observed that the separation region for the gap/step configuration increased with increasing the step frontal-face height. It was found that density and pressure for the gap/step configuration dramatically increased inside the gap as compared to those observed for the gap configuration, i. e., a gap without a step.
NASA Astrophysics Data System (ADS)
Xu, Kun; He, Xin; Cai, Chunpei
2008-07-01
It is well known that for increasingly rarefied flowfields, the predictions from continuum formulation, such as the Navier-Stokes equations lose accuracy. For the high speed diatomic molecular flow in the transitional regime, the inaccuracies are partially attributed to the single temperature approximations in the Navier-Stokes equations. Here, we propose a continuum multiple temperature model based on the Bhatnagar-Gross-Krook (BGK) equation for the non-equilibrium flow computation. In the current model, the Landau-Teller-Jeans relaxation model for the rotational energy is used to evaluate the energy exchange between the translational and rotational modes. Due to the multiple temperature approximation, the second viscosity coefficient in the Navier-Stokes equations is replaced by the temperature relaxation term. In order to solve the multiple temperature kinetic model, a multiscale gas-kinetic finite volume scheme is proposed, where the gas-kinetic equation is numerically solved for the fluxes to update the macroscopic flow variables inside each control volume. Since the gas-kinetic scheme uses a continuous gas distribution function at a cell interface for the fluxes evaluation, the moments of a gas distribution function can be explicitly obtained for the multiple temperature model. Therefore, the kinetic scheme is much more efficient than the DSMC method, especially in the near continuum flow regime. For the non-equilibrium flow computations, i.e., the nozzle flow and hypersonic rarefied flow over flat plate, the computational results are validated in comparison with experimental measurements and DSMC solutions.
Assessment of gas-surface interaction models for computation of rarefied hypersonic flows
NASA Astrophysics Data System (ADS)
Padilla, Jose Fernando
Over the next few decades, spaceflight is expected to become more common through the resurgence of manned space exploration and the rise of commercial manned spaceflight. An essential role for the efficient research and development of suborbital spaceflight is played by computational simulation of rarefied hypersonic flows. Among the few classes of computational approaches for examining rarefied gas dynamics, the most widely used approach, for spatial scales relevant to suborbital spaceflight, is the direct simulation Monte Carlo (DSMC) method. Although the DSMC method has been under development for over forty years, there are still many areas where improvements can be made. One particular area is the associated numerical modeling of interactions between gas molecules and solid surfaces. Gas-surface interactions are not well understood for rarefied hypersonic conditions, although various models have been developed. This thesis ultimately focuses on assessing two common gas-surface interaction models in use with the DSMC method, the Maxwell model and the Cercignani, Lampis and Lord (CLL) model. In the search for a definitive thesis goal and as a consequence of the analysis tools developed for achieving the definitive thesis goal, several aspects of DSMC analysis are examined. Initially, procedures to determine aerodynamic coefficients from DSMC simulations are validated against certain windtunnel test data and an independent DSMC code. Then, sensitivity studies are performed involving aerothermodynamics predictions for the Apollo 6, at the 110 km altitude return trajectory point. This reveals the significance of gas-surface surface interaction models in rarefied hypersonic flows. A review of existing gas-surface interaction models motivates the assessment of the Maxwell and CLL models. The two models are scrutinized with the help of relatively recent windtunnel test measurements and procedures to extract surface scattering distributions. Both models yield similar
Direct simulation Monte Carlo of rarefied hypersonic flow on power law shaped leading edges
NASA Astrophysics Data System (ADS)
Santos, Wilson Fernando Nogueira Dos
A numerical study of several parameters that influence the flowfield structure, aerodynamic surface quantities and shock wave structure at rarefied hypersonic flow conditions is conducted on power law shaped leading edges. The calculations are performed with a detailed computer code that properly accounts for nonequilibrium effects and that has been demonstrated to yield excellent comparisons with flight- and ground-test data. The flowfield structure, aerodynamic surface quantities and shock wave structure of power law shaped leading edges are examined in order to provide information on how well these shapes could stand as possible candidates for blunting geometries of hypersonic leading edges. Newtonian flow analysis has shown that these shapes exhibit both blunt and sharp aerodynamic properties. Moreover, computational investigation of minimum-drag bodies at supersonic and moderate hypersonic speeds has indicated that power law shapes for certain exponents yield the lowest wave drag. These qualities make power law shapes strong candidates for leading edge design. A very detailed description of the impact on the flow properties, such as velocity, density, temperature and pressure, has been presented separately in the vicinity of the nose of the leading edges due to changes in their shapes. Numerical solutions show that the shape of the leading edge disturbed the flowfield far upstream, where the domain of influence decreased as the leading edge became aerodynamically sharp. A detailed procedure is presented to predict the pressure gradient along the body surface in a rarefied environment. Numerical solutions show that the pressure gradient behavior follows that predicted by Newtonian theory. It is found that the pressure gradient along the body surface goes to zero at the nose of the leading edge for power law exponents less than 2/3, a characteristic of a blunt body. It is finite for power law exponent of 2/3 and goes to minus infinite for power law exponents
Zonally-decoupled DSMC solutions of hypersonic blunt body wake flows
NASA Technical Reports Server (NTRS)
Wilmoth, Richard G.; Mitcheltree, Robert A.; Moss, James N.; Dogra, Virendra K.
1993-01-01
Direct simulation Monte Carlo (DSMC) solutions are presented for the hypersonic flow behind a blunt body in which the wake region is solved in a zonally-decoupled manner. The forebody flow is solved separately using either a DSMC or a Navier-Stokes method, and the forebody exit plane solution is specified as the inflow condition to the decoupled DSMC solution of the wake region. Results are presented for a 70-deg, blunted cone at flow conditions that can be accommodated in existing low-density wind tunnels with the Knudsen number based on base diameter ranging from 0.03 to 0.001. The zonally-decoupled solutions show good agreement with fully-coupled DSMC solutions of the wake flow densities and velocities. The wake closure predicted by the zonally-decoupled solutions is in better agreement with fully-coupled results than that predicted by a fully-coupled Navier-Stokes method indicating the need to account for rarefaction in the wake for the cases considered. The combined use of Navier-Stokes for the forebody with a decoupled DSMC solution for the wake provides an efficient method for solving transitional blunt-body flows where the forebody flow is continuum and the wake is rarefied.
Simulation of hypersonic rarefied flows with the immersed-boundary method
NASA Astrophysics Data System (ADS)
Bruno, D.; De Palma, P.; de Tullio, M. D.
2011-05-01
This paper provides a validation of an immersed boundary method for computing hypersonic rarefied gas flows. The method is based on the solution of the Navier-Stokes equation and is validated versus numerical results obtained by the DSMC approach. The Navier-Stokes solver employs a flexible local grid refinement technique and is implemented on parallel machines using a domain-decomposition approach. Thanks to the efficient grid generation process, based on the ray-tracing technique, and the use of the METIS software, it is possible to obtain the partitioned grids to be assigned to each processor with a minimal effort by the user. This allows one to by-pass the expensive (in terms of time and human resources) classical generation process of a body fitted grid. First-order slip-velocity boundary conditions are employed and tested for taking into account rarefied gas effects.
NASA Technical Reports Server (NTRS)
Chen, Y. K.; Henline, W. D.
1993-01-01
The general boundary conditions including mass and energy balances of chemically equilibrated or nonequilibrated gas adjacent to ablating surfaces have been derived. A computer procedure based on these conditions was developed and interfaced with the Navier-Stokes solver for predictions of the flow field, surface temperature, and surface ablation rates over re-entry space vehicles with ablating Thermal Protection Systems (TPS). The Navier-Stokes solver with general surface thermochemistry boundary conditions can predict more realistic solutions and provide useful information for the design of TPS. A test case with a proposed hypersonic test vehicle configuration and associated free stream conditions was developed. Solutions with various surface boundary conditions were obtained, and the effect of nonequilibrium gas as well as surface chemistry on surface heating and ablation rate were examined. The solutions of the GASP code with complete ablating surface conditions were compared with those of the ASC code. The direction of future work is also discussed.
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.; Gupta, Roop N.; Shinn, Judy L.
1989-01-01
The conservation equations for simulating hypersonic flows in thermal and chemical nonequilibrium and details of the associated physical models are presented. These details include the curve fits used for defining thermodynamic properties of the 11 species air model, curve fits for collision cross sections, expressions for transport properties, the chemical kinetics models, and the vibrational and electronic energy relaxation models. The expressions are formulated in the context of either a two or three temperature model. Greater emphasis is placed on the two temperature model in which it is assumed that the translational and rotational energy models are in equilibrium at the translational temperature, T, and the vibrational, electronic, and electron translational energy modes are in equilibrium at the vibrational temperature, T sub v. The eigenvalues and eigenvectors associated with the Jacobian of the flux vector are also presented in order to accommodate the upwind based numerical solutions of the complete equation set.
NASA Technical Reports Server (NTRS)
Dagum, Leonardo
1989-01-01
The data parallel implementation of a particle simulation for hypersonic rarefied flow described by Dagum associates a single parallel data element with each particle in the simulation. The simulated space is divided into discrete regions called cells containing a variable and constantly changing number of particles. The implementation requires a global sort of the parallel data elements so as to arrange them in an order that allows immediate access to the information associated with cells in the simulation. Described here is a very fast algorithm for performing the necessary ranking of the parallel data elements. The performance of the new algorithm is compared with that of the microcoded instruction for ranking on the Connection Machine.
DSMC Grid Methodologies for Computing Low-Density, Hypersonic Flows About Reusable Launch Vehicles
NASA Technical Reports Server (NTRS)
Wilmoth, Richard G.; LeBeau, Gerald J.; Carlson, Ann B.
1996-01-01
Two different grid methodologies are studied for application to DSMC simulations about reusable launch vehicles. One method uses an unstructured, tetrahedral grid while the other uses a structured, variable-resolution Cartesian grid. The relative merits of each method are discussed in terms of accuracy, computational efficiency, and overall ease of use. Both methods are applied to the computation of a low-density, hypersonic flow about a winged single-stage-to-orbit reusable launch vehicle concept at conditions corresponding to an altitude of 120 km. Both methods are shown to give comparable results for both surface and flowfield quantities as well as for the overall aerodynamic behavior. For the conditions simulated, the flowfield about the vehicle is very rarefied but the DSMC simulations show significant departure from free-molecular predictions for the surface friction and heat transfer as well as certain aerodynamic quantities.
Similar solutions for viscous hypersonic flow over a slender three-fourths-power body of revolution
NASA Technical Reports Server (NTRS)
Lin, Chin-Shun
1987-01-01
For hypersonic flow with a shock wave, there is a similar solution consistent throughout the viscous and inviscid layers along a very slender three-fourths-power body of revolution The strong pressure interaction problem can then be treated by the method of similarity. Numerical calculations are performed in the viscous region with the edge pressure distribution known from the inviscid similar solutions. The compressible laminar boundary-layer equations are transformed into a system of ordinary differential equations. The resulting two-point boundary value problem is then solved by the Runge-Kutta method with a modified Newton's method for the corresponding boundary conditions. The effects of wall temperature, mass bleeding, and body transverse curvature are investigated. The induced pressure, displacement thickness, skin friction, and heat transfer due to the previously mentioned parameters are estimated and analyzed.
NASA Technical Reports Server (NTRS)
Hung, C. M.; Maccormack, R. W.
1975-01-01
An efficient time-splitting, second-order accurate, numerical scheme is used to solve the complete Navier-Stokes equations for supersonic and hypersonic laminar flow over a two-dimensional compression corner. A fine, exponentially stretched mesh spacing is used in the region near the wall for resolving the viscous layer. Good agreement is obtained between the present computed results and experimental measurement for a Mach number of 14.1, a Reynolds number of 104,000, and wedge angles of 15, 18, and 24 deg. The details of the pressure variation across the boundary layer are given, and a correlation between the leading edge shock and the peaks in surface pressure and heat transfer is observed.
Asynchronous, macrotasked relaxation strategies for the solution of viscous, hypersonic flows
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
1991-01-01
A point-implicit, asynchronous macrotasked relaxation of the steady, thin-layer, Navier-Stokes equations is presented. The method employs multidirectional, single-level storage Gauss-Seidel relaxation sweeps, which effectively communicate perturbations across the entire domain in 2n sweeps, where n is the dimension of the domain. In order to enhance convergence the application of relaxation factors to specific components of the Jacobian is examined using a stability analysis of the advection and diffusion equations. Attention is also given to the complications associated with asynchronous multitasking. Solutions are generated for hypersonic flows over blunt bodies in two and three dimensions with chemical reactions, utilizing single-tasked and multitasked relaxation strategies.
Applications of Quantum Theory of Atomic and Molecular Scattering to Problems in Hypersonic Flow
NASA Technical Reports Server (NTRS)
Malik, F. Bary
1995-01-01
The general status of a grant to investigate the applications of quantum theory in atomic and molecular scattering problems in hypersonic flow is summarized. Abstracts of five articles and eleven full-length articles published or submitted for publication are included as attachments. The following topics are addressed in these articles: fragmentation of heavy ions (HZE particles); parameterization of absorption cross sections; light ion transport; emission of light fragments as an indicator of equilibrated populations; quantum mechanical, optical model methods for calculating cross sections for particle fragmentation by hydrogen; evaluation of NUCFRG2, the semi-empirical nuclear fragmentation database; investigation of the single- and double-ionization of He by proton and anti-proton collisions; Bose-Einstein condensation of nuclei; and a liquid drop model in HZE particle fragmentation by hydrogen.
NASA Technical Reports Server (NTRS)
Chen, Y. K.; Henline, W. D.
1993-01-01
The general boundary conditions including mass and energy balances of chemically equilibrated or nonequilibrated gas adjacent to ablating surfaces have been derived. A computer procedure based on these conditions was developed and interfaced with the Navier-Stokes solver for predictions of the flow field, surface temperature, and surface ablation rates over re-entry space vehicles with ablating Thermal Protection Systems (TPS). The Navier-Stokes solver with general surface thermochemistry boundary conditions can predict more realistic solutions and provide useful information for the design of TPS. A test case with a proposed hypersonic test vehicle configuration and associated free stream conditions was developed. Solutions with various surface boundary conditions were obtained, and the effect of nonequilibrium gas as well as surface chemistry on surface heating and ablation rate were examined. The solutions of the GASP code with complete ablating surface conditions were compared with those of the ASC code. The direction of future work is also discussed.
Asynchronous, macrotasked relaxation strategies for the solution of viscous, hypersonic flows
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
1991-01-01
A point-implicit, asynchronous macrotasked relaxation of the steady, thin-layer, Navier-Stokes equations is presented. The method employs multidirectional, single-level storage Gauss-Seidel relaxation sweeps, which effectively communicate perturbations across the entire domain in 2n sweeps, where n is the dimension of the domain. In order to enhance convergence the application of relaxation factors to specific components of the Jacobian is examined using a stability analysis of the advection and diffusion equations. Attention is also given to the complications associated with asynchronous multitasking. Solutions are generated for hypersonic flows over blunt bodies in two and three dimensions with chemical reactions, utilizing single-tasked and multitasked relaxation strategies.
Solution of the Burnett equations for hypersonic flows near the continuum limit
NASA Technical Reports Server (NTRS)
Imlay, Scott T.
1992-01-01
The INCA code, a three-dimensional Navier-Stokes code for analysis of hypersonic flowfields, was modified to analyze the lower reaches of the continuum transition regime, where the Navier-Stokes equations become inaccurate and Monte Carlo methods become too computationally expensive. The two-dimensional Burnett equations and the three-dimensional rotational energy transport equation were added to the code and one- and two-dimensional calculations were performed. For the structure of normal shock waves, the Burnett equations give consistently better results than Navier-Stokes equations and compare reasonably well with Monte Carlo methods. For two-dimensional flow of Nitrogen past a circular cylinder the Burnett equations predict the total drag reasonably well. Care must be taken, however, not to exceed the range of validity of the Burnett equations.
Analysis of hypersonic arcjet flow fields and surface heating of blunt bodies
NASA Technical Reports Server (NTRS)
Chen, Y.-K.; Henline, W. D.
1993-01-01
A Gauss-Seidel implicit aerothermodynamic Navier-Stokes computational CFD code with thermochemical surface conditions (GIANTS) and with appropriate boundary conditions was developed for the analysis of hypersonic arcjet flows over large angle blunt bodies. Two-dimensional axisymmetric multicomponent full Navier-Stokes equations were solved, together with boundary conditions, to account for the surface slip, surface catalysis, and energy and mass conservation at the body surface. It is shown that the predictions for isothermal, noncatalytic, and no-slip surface conditions which are made using the GIANTS code with simple boundary conditions agree with those of the LAURA code, and the predicted normalized heating distributions over the forebody surface with boundary conditions for various cone angles are consistent with the trends of data from arcjet experiments.
Hypersonic ionizing air viscous shock-layer flows over nonanalytic blunt bodies
NASA Technical Reports Server (NTRS)
Miner, E. W.; Lewis, C. H.
1975-01-01
The equations which govern the viscous shock-layer flow are presented and the method by which the equations are solved is discussed. The predictions of the present finite-difference method are compared with other numerical predictions as well as with experimental data. The principal emphasis is placed on predictions of the viscous flowfield for the windward plane of symmetry of the space shuttle orbiter and other axisymmetric bodies which approximate the shuttle orbiter geometry. Experimental data on two slender sphere-cones at hypersonic conditions are also considered. The present predictions agreed well with experimental data and with the past predictions. Substantial differences were found between present predictions and more approximate methods.
Solution of the Burnett equations for hypersonic flows near the continuum limit
NASA Technical Reports Server (NTRS)
Imlay, Scott T.
1992-01-01
The INCA code, a three-dimensional Navier-Stokes code for analysis of hypersonic flowfields, was modified to analyze the lower reaches of the continuum transition regime, where the Navier-Stokes equations become inaccurate and Monte Carlo methods become too computationally expensive. The two-dimensional Burnett equations and the three-dimensional rotational energy transport equation were added to the code and one- and two-dimensional calculations were performed. For the structure of normal shock waves, the Burnett equations give consistently better results than Navier-Stokes equations and compare reasonably well with Monte Carlo methods. For two-dimensional flow of Nitrogen past a circular cylinder the Burnett equations predict the total drag reasonably well. Care must be taken, however, not to exceed the range of validity of the Burnett equations.
Experimental investigation on drag and heat flux reduction in supersonic/hypersonic flows: A survey
NASA Astrophysics Data System (ADS)
Wang, Zhen-guo; Sun, Xi-wan; Huang, Wei; Li, Shi-bin; Yan, Li
2016-12-01
The drag and heat reduction problem of hypersonic vehicles has always attracted the attention worldwide, and the experimental test approach is the basis of theoretical analysis and numerical simulation. In the current study, research progress of experimental investigations on drag and heat reduction are summarized by several kinds of mechanism, namely the forward-facing cavity, the opposing jet, the aerospike, the energy deposition and their combinational configurations, and the combinational configurations include the combinational opposing jet and forward-facing cavity concept and the combinational opposing jet and aerospike concept. The geometric models and flow conditions are emphasized, especially for the basic principle for the drag and heat flux reduction of each device. The measurement results of aerodynamic and aerothermodynamic are compared and analyzed as well, which can be a reference for assessing the accuracy of numerical results.
Influence of Nose Radius of Blunt Cones on Drag in Supersonic and Hypersonic Flows
NASA Astrophysics Data System (ADS)
Hemateja, A.; Teja, B. Ravi; Dileep Kumar, A.; Rakesh, S. G.
2017-08-01
The objects moving at high speeds encounter forces which tend to decelerate the objects. This resistance in the medium is termed as drag which is one of the major concerns while designing high speed aircrafts. Another key factor which influences the design is the heat transfer. The main challenge faced by aerospace industries is to design the shape of the flying object that travels at high speeds with optimum values of heat generation and drag. This study deals with computational analysis of sharp and blunt cones with varying cone angles and nose radii. The effect of nose radius on the drag is studied at supersonic and hypersonic flows and at various angles of attack. It is observed that as the nose radius is increased, the heat transfer reduces & the drag increases and vice-versa. Looking at the results, the optimum value of nose radius can be chosen depending on the need of the problem.
Prediction of Drag Reduction in Supersonic and Hypersonic Flows with Counterflow Jets
NASA Technical Reports Server (NTRS)
Daso, Endwell O.; Beaulieu, Warren; Hager, James O.; Turner, James E. (Technical Monitor)
2002-01-01
Computational fluid dynamics solutions of the flowfield of a truncated cone-cylinder with and without counterflow jets have been obtained for the short penetration mode (SPM) and long penetration mode (LPM) of the freestream-counterflow jet interaction flowfield. For the case without the counterflow jet, the comparison of the normalized surface pressures showed very good agreement with experimental data. For the case with the SPM jet, the predicted surface pressures did not compare as well with the experimental data upstream of the expansion corner, while aft of the expansion corner, the comparison of the solution and the data is seen to give much better agreement. The difference in the prediction and the data could be due to the transient character of the jet penetration modes, possible effects of the plasma physics that are not accounted for here, or even the less likely effect of flow turbulence, etc. For the LPM jet computations, one-dimensional isentropic relations were used to derived the jet exit conditions in order to obtain the LPM solutions. The solution for the jet exit Mach number of 3 shows a jet penetration several times longer than that of the SPM, and therefore much weaker bow shock, with an attendant reduction in wave drag. The LPM jet is, in essence, seen to be a "pencil" of fluid, with much higher dynamic pressure, embedded in the oncoming supersonic or hypersonic freestream. The methodology for determining the conditions for the LPM jet could enable a practical approach for the design and application of counterflow LPM jets for the reduction of wave drag and heat flux, thus significantly enhancing the aerodynamic characteristics and aerothermal performance of supersonic and hypersonic vehicles. The solutions show that the qualitative flow structure is very well captured. The obtained results, therefore, suggest that counterflowing jets are viable candidate technology concepts that can be employed to give significant reductions in wave drag, heat
Effects of nose bluntness and angle of attack on slender bodies in hypersonic flows
NASA Technical Reports Server (NTRS)
Tiwari, S. N.; Sehgal, A. K.; Singh, D. J.
1992-01-01
The effects of angle of attack and nose bluntness on the flow field and wall quantities are investigated for hypersonic flows of air over slender bodies. The bodies considered are slender cones and straight biconic configurations. The numerical procedures used are based on the solution of complete Navier-Stokes equations in the nose region and parabolized Navier-Stokes equations in the downstream region. Results are obtained for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The postshock flow field is studied in detail from the contour plots of Mach number, density, and temperature. Flow separation is observed on the leeward plane for an on-axis, 12.84 deg/7 deg (fore-cone and aft-cone angles) biconic geometry at 12 deg angle of attack. Also, the windward and leeward heating rates for the fore-cone section decrease by a factor of four and five, respectively, when the nose bluntness is increased by an order of magnitude. The effect of nose bluntness for slender cone persists as far as 200 nose radii downstream.
Development of braided rope seals for hypersonic engine applications: Flow modeling
NASA Technical Reports Server (NTRS)
Mutharasan, Rajakkannu; Steinetz, Bruce M.; Tao, Xiaoming; Du, Guang-Wu; Ko, Frank
1992-01-01
A new type of engine seal is being developed to meet the needs of advanced hypersonic engines. A seal braided of emerging high temperature ceramic fibers comprised of a sheath-core construction was selected for study based on its low leakage rates. Flexible, low-leakage, high temperature seals are required to seal the movable engine panels of advanced ramjet-scramjet engines either preventing potentially dangerous leakage into backside engine cavities or limiting the purge coolant flow rates through the seals. To predict the leakage through these flexible, porous seal structures new analytical flow models are required. Two such models based on the Kozeny-Carman equations are developed herein and are compared to experimental leakage measurements for simulated pressure and seal gap conditions. The models developed allow prediction of the gas leakage rate as a function of fiber diameter, fiber packing density, gas properties, and pressure drop across the seal. The first model treats the seal as a homogeneous fiber bed. The second model divides the seal into two homogeneous fiber beds identified as the core and the sheath of the seal. Flow resistances of each of the main seal elements are combined to determine the total flow resistance. Comparisons between measured leakage rates and model predictions for seal structures covering a wide range of braid architectures show good agreement. Within the experimental range, the second model provides a prediction within 6 to 13 percent of the flow for many of the cases examined. Areas where future model refinements are required are identified.
Radial Gas Flows in Colliding Galaxies: Connecting Simulations and Observations
NASA Astrophysics Data System (ADS)
Iono, Daisuke; Yun, Min S.; Mihos, J. Christopher
2004-11-01
We investigate the detailed response of gas to the formation of transient and long-lived dynamical structures induced in the early stages of a disk-disk collision and identify observational signatures of radial gas inflow through a detailed examination of the collision simulation of an equal-mass bulge-dominated galaxy. Our analysis and discussion mainly focuses on the evolution of the diffuse and dense gas in the early stages of the collision, when the two disks are interacting but have not yet merged. Stars respond to the tidal interaction by forming both transient arms and long-lived m=2 bars, but the gas response is more transient, flowing directly toward the central regions within about 108 yr after the initial collision. The rate of inflow declines when more than half of the total gas supply reaches the inner few kiloparsecs, where the gas forms a dense nuclear ring inside the stellar bar. The average gas inflow rate to the central 1.8 kpc is ~7 Msolar yr-1 with a peak rate of 17 Msolar yr-1. Gas with high volume density is found in the inner parts of the postcollision disks at size scales close to the spatial resolution of the simulations, and this may be a direct result of shocks traced by the discontinuity in the gas velocity field. The evolution of gas in a bulgeless progenitor galaxy is also discussed, and a possible link to the ``chain galaxy'' population observed at high redshifts is inferred. The evolution of the structural parameters such as asymmetry and concentration of both stars and gas are studied in detail. Further, a new structure parameter (the compactness parameter K) that traces the evolution of the size scale of the gas relative to the stellar disk is introduced, and this may be a useful tracer to determine the merger chronology of colliding systems. Noncircular gas kinematics driven by the perturbation of the nonaxisymmetric structure can produce distinct emission features in the ``forbidden velocity quadrants'' of the position
NASA Astrophysics Data System (ADS)
Luo, Xiaobing
Relative-moving boundary problems have a wide variety of applications. They appear in staging during a launch process, store separation from a military aircraft, rotor-stator interaction in turbomachinery, and dynamic aeroelasticity. The dynamic unstructured technology (DUT) is potentially a strong approach to simulate unsteady flows around relative-moving bodies, by solving time-dependent governing equations. The dual-time stepping scheme is implemented to improve its efficiency while not compromising the accuracy of solutions. The validation of the implicit scheme is performed on a pitching NACA0012 airfoil and a rectangular wing with low reduced frequencies in transonic flows. All the matured accelerating techniques, including the implicit residual smoothing, the local time stepping, and the Full- Approximate-Scheme (FAS) multigrid method, are resorted once a dynamic problem is transformed into a series of ``static'' problems. Even with rather coarse Euler-type meshes, one order of CPU time savings is achieved without losing the accuracy of solutions in comparison to the popular Runge-Kutta scheme. More orders of CPU time savings are expected in real engineering applications where highly stretched viscous-type meshes are needed. The applicability of DUT is also extended from transonic/supersonic flows to hypersonic flows through special measures in spatial discretization to simulate the staging of a hypersonic vehicle. First, the simulations in Mach 5 and Mach 10 flights are performed on the longitudinal symmetry plane. A network of strong shocks and expansion waves are captured. A prescribed two-degrees-of-freedom motion is imposed on the booster and the adapter to mimic the staging. Then, a 3-D static Euler solver with an efficient edge- based data structure is modified for time-accurate flows. The overall history of aerodynamic interference during the staging in Mach 5 flight is obtained by an animation method, consisting of six static solutions along the
Simulating flow around scaled model of a hypersonic vehicle in wind tunnel
NASA Astrophysics Data System (ADS)
Markova, T. V.; Aksenov, A. A.; Zhluktov, S. V.; Savitsky, D. V.; Gavrilov, A. D.; Son, E. E.; Prokhorov, A. N.
2016-11-01
A prospective hypersonic HEXAFLY aircraft is considered in the given paper. In order to obtain the aerodynamic characteristics of a new construction design of the aircraft, experiments with a scaled model have been carried out in a wind tunnel under different conditions. The runs have been performed at different angles of attack with and without hydrogen combustion in the scaled propulsion engine. However, the measured physical quantities do not provide all the information about the flowfield. Numerical simulation can complete the experimental data as well as to reduce the number of wind tunnel experiments. Besides that, reliable CFD software can be used for calculations of the aerodynamic characteristics for any possible design of the full-scale aircraft under different operation conditions. The reliability of the numerical predictions must be confirmed in verification study of the software. The given work is aimed at numerical investigation of the flowfield around and inside the scaled model of the HEXAFLY-CIAM module under wind tunnel conditions. A cold run (without combustion) was selected for this study. The calculations are performed in the FlowVision CFD software. The flow characteristics are compared against the available experimental data. The carried out verification study confirms the capability of the FlowVision CFD software to calculate the flows discussed.
A database of aerothermal measurements in hypersonic flow for CFD validation
NASA Technical Reports Server (NTRS)
Holden, M. S.; Moselle, J. R.
1992-01-01
This paper presents an experimental database selected and compiled from aerothermal measurements obtained on basic model configurations on which fundamental flow phenomena could be most easily examined. The experimental studies were conducted in hypersonic flows in 48-inch, 96-inch, and 6-foot shock tunnels. A special computer program was constructed to provide easy access to the measurements in the database as well as the means to plot the measurements and compare them with imported data. The database contains tabulations of model configurations, freestream conditions, and measurements of heat transfer, pressure, and skin friction for each of the studies selected for inclusion. The first segment contains measurements in laminar flow emphasizing shock-wave boundary-layer interaction. In the second segment, measurements in transitional flows over flat plates and cones are given. The third segment comprises measurements in regions of shock-wave/turbulent-boundary-layer interactions. Studies of the effects of surface roughness of nosetips and conical afterbodies are presented in the fourth segment of the database. Detailed measurements in regions of shock/shock boundary layer interaction are contained in the fifth segment. Measurements in regions of wall jet and transpiration cooling are presented in the final two segments.
Characterization of CO2 flow in a hypersonic impulse facility using DLAS
NASA Astrophysics Data System (ADS)
Meyers, J. M.; Paris, S.; Fletcher, D. G.
2016-02-01
This work documents diode laser absorption measurements of CO2 flow in the free stream of the Longshot hypersonic impulse facility at Mach numbers ranging from 10 to 12. The diode laser sensor was designed to measure absorption of the P12 (30013) ← (00001) transition near 1.6 \\upmum, which yields relatively weak direct absorption levels (3.5 % per meter at peak Longshot free-stream conditions). Despite this weak absorption, measurements yielded valuable flow property information during the first 20 ms of facility runs. Simultaneous measurements of static temperature, pressure, and velocity were acquired in the inviscid core flow region using a laser wavelength scanning frequency of 600 Hz. The free-stream values obtained from DLAS measurements were then compared to Longshot probe-derived values determined from settling chamber and probe measurements. This comparison enabled an assessment of the traditional method of flow characterization in the facility, which indicated negligible influence from possible vibrational freezing of reservoir gases.
A database of aerothermal measurements in hypersonic flow for CFD validation
NASA Technical Reports Server (NTRS)
Holden, M. S.; Moselle, J. R.
1992-01-01
This paper presents an experimental database selected and compiled from aerothermal measurements obtained on basic model configurations on which fundamental flow phenomena could be most easily examined. The experimental studies were conducted in hypersonic flows in 48-inch, 96-inch, and 6-foot shock tunnels. A special computer program was constructed to provide easy access to the measurements in the database as well as the means to plot the measurements and compare them with imported data. The database contains tabulations of model configurations, freestream conditions, and measurements of heat transfer, pressure, and skin friction for each of the studies selected for inclusion. The first segment contains measurements in laminar flow emphasizing shock-wave boundary-layer interaction. In the second segment, measurements in transitional flows over flat plates and cones are given. The third segment comprises measurements in regions of shock-wave/turbulent-boundary-layer interactions. Studies of the effects of surface roughness of nosetips and conical afterbodies are presented in the fourth segment of the database. Detailed measurements in regions of shock/shock boundary layer interaction are contained in the fifth segment. Measurements in regions of wall jet and transpiration cooling are presented in the final two segments.
Application of a Modular Particle-Continuum Method to Partially Rarefied, Hypersonic Flow
NASA Astrophysics Data System (ADS)
Deschenes, Timothy R.; Boyd, Iain D.
2011-05-01
The Modular Particle-Continuum (MPC) method is used to simulate partially-rarefied, hypersonic flow over a sting-mounted planetary probe configuration. This hybrid method uses computational fluid dynamics (CFD) to solve the Navier-Stokes equations in regions that are continuum, while using direct simulation Monte Carlo (DSMC) in portions of the flow that are rarefied. The MPC method uses state-based coupling to pass information between the two flow solvers and decouples both time-step and mesh densities required by each solver. It is parallelized for distributed memory systems using dynamic domain decomposition and internal energy modes can be consistently modeled to be out of equilibrium with the translational mode in both solvers. The MPC results are compared to both full DSMC and CFD predictions and available experimental measurements. By using DSMC in only regions where the flow is nonequilibrium, the MPC method is able to reproduce full DSMC results down to the level of velocity and rotational energy probability density functions while requiring a fraction of the computational time.
DSMC Simulations of Hypersonic Flows With Shock Interactions and Validation With Experiments
NASA Technical Reports Server (NTRS)
Moss, James N.; Bird, Graeme A.
2004-01-01
The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolution, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.
DSMC Simulations of Hypersonic Flows With Shock Interactions and Validation With Experiments
NASA Technical Reports Server (NTRS)
Moss, James N.; Bird, Graeme A.
2004-01-01
The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolutions, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.
NASA Astrophysics Data System (ADS)
Erwin, Daniel A.; Kunc, Joseph A.; Muntz, E. P.
1991-08-01
The goal was to develop an experimental diagnostic technique suitable for gas flows of densities intermediate between atmospheric and rarefied. A laser assisted electron beam fluorescense technique which we call electron photon fluorescense was developed. Theoretical work was done to predict the time dependence of the excitation/deexcitation processes. As described in the original proposal, our goal in this work was the attainment of an experimental diagnostic technique suitable for gas flows of densities intermediate between atmospheric and rarefied. Measurements in such intermediate density flows, typical of hypersonic flight at altitudes above about 50 km, present difficulties in that traditional wind-tunnel techniques (shadow and schlieren, as well as laser based scattering techniques) provide insufficient signal. Moreover, the resonant scattering techniques may require an absorptive species as a tracer to be seeded into the flow, a requirement inconsistent with the realities of existing large facilities. On the other hand, the densities are not enough for continuous electron-beam fluorescence (EBF) to be used due to beam spreading and collisional quenching.
Fluorescence Visualization of Hypersonic Flow Past Triangular and Rectangular Boundary-layer Trips
NASA Technical Reports Server (NTRS)
Danehy, Paul M.; Garcia, A. P.; Borg, Stephen E.; Dyakonov, Artem A.; Berry, Scott A.; Inman, Jennifer A.; Alderfer, David W.
2007-01-01
Planar laser-induced fluorescence (PLIF) flow visualization has been used to investigate the hypersonic flow of air over surface protrusions that are sized to force laminar-to-turbulent boundary layer transition. These trips were selected to simulate protruding Space Shuttle Orbiter heat shield gap-filler material. Experiments were performed in the NASA Langley Research Center 31-Inch Mach 10 Air Wind Tunnel, which is an electrically-heated, blowdown facility. Two-mm high by 8-mm wide triangular and rectangular trips were attached to a flat plate and were oriented at an angle of 45 degrees with respect to the oncoming flow. Upstream of these trips, nitric oxide (NO) was seeded into the boundary layer. PLIF visualization of this NO allowed observation of both laminar and turbulent boundary layer flow downstream of the trips for varying flow conditions as the flat plate angle of attack was varied. By varying the angle of attack, the Mach number above the boundary layer was varied between 4.2 and 9.8, according to analytical oblique-shock calculations. Computational Fluid Dynamics (CFD) simulations of the flowfield with a laminar boundary layer were also performed to better understand the flow environment. The PLIF images of the tripped boundary layer flow were compared to a case with no trip for which the flow remained laminar over the entire angle-of-attack range studied. Qualitative agreement is found between the present observed transition measurements and a previous experimental roughness-induced transition database determined by other means, which is used by the shuttle return-to-flight program.
NASA Technical Reports Server (NTRS)
Chalot, F.; Hughes, T. J. R.; Johan, Z.; Shakib, F.
1991-01-01
An FEM for the compressible Navier-Stokes equations is introduced. The discretization is based on entropy variables. The methodology is developed within the framework of a Galerkin/least-squares formulation to which a discontinuity-capturing operator is added. Results for three test cases selected among those of the Workshop on Hypersonic Flows for Reentry Problems are presented.
Viscous-shock-layer analysis of hypersonic flows over long slender vehicles. Ph.D. Thesis, 1988
NASA Technical Reports Server (NTRS)
Lee, Kam-Pui; Gupta, Roop N.
1992-01-01
An efficient and accurate method for solving the viscous shock layer equations for hypersonic flows over long slender bodies is presented. The two first order equations, continuity and normal momentum, are solved simultaneously as a coupled set. The flow conditions included are from high Reynolds numbers at low altitudes to low Reynolds numbers at high altitudes. For high Reynolds number flows, both chemical nonequilibrium and perfect gas cases are analyzed with surface catalytic effects and different turbulence models, respectively. At low Reynolds number flow conditions, corrected slip models are implemented with perfect gas case. Detailed comparisons are included with other predictions and experimental data.
NASA Astrophysics Data System (ADS)
Roohi, Ehsan; Stefanov, Stefan
2016-11-01
This paper reviews the accuracy of the Simplified Bernoulli Trial (SBT) algorithm and its variants, i.e., SBT-TAS (SBT on transient adaptive subcells) and ISBT (intelligence SBT) in the simulation of a wide spectrum of rarefied flow problems, including collision frequency ratio evaluation in the equilibrium condition, comparison of the Sonine-polynomial coefficients prediction in the Fourier flow with the theoretical prediction of the Chapman-Enskog expansion, accurate wall heat flux solution for the Fourier flow in the early slip regime, and hypersonic flows over cylinder and biconic geometries. We summarize advantages and requirements that utilization of the SBT collision families brings to a typical DSMC solver.
DSMC simulation for effects of angles of attack on rarefied hypersonic cavity flows
NASA Astrophysics Data System (ADS)
Jin, Xuhon; Huang, Fei; Shi, Jiatong; Cheng, Xiaoli
2016-11-01
The present work investigates rarefied hypersonic flows over a flat plate with two-dimensional and three-dimensional cavities by employing the direct simulation Monte Carlo (DSMC) method, focusing on the effect of angles of attack (AOAs) on flow structure inside the cavity and aerodynamic surface quantities. It was found that only one primary recirculation structure was formed inside the cavity at the angle of attack (AOA) of 0°, while a second vortex system was produced just beneath the primary one with the angle of attack increased to 30°. As AOAs grow, the freestream flow is able to penetrate deeper into the cavity and attach itself to the cavity base, making the "dead-water" region shrink. Meantime, with the increment in the AOA, both heat transfer and pressure coefficients show a similar and gradual quantitative behavior, and along centerlines of the two side surfaces of the cavity, both pressure and heat transfer coefficients become growing, indicating that the increase in the AOA does enhance momentum and energy transfer to both the two aforementioned surfaces. However, the heat flux over the cavity floor does not keep increasing with the growth of the AOA, while the pressure does, indicating that augmenting AOAs does not enhance momentum to the cavity floor, but does make compressibility stronger and stronger near the cavity base.
Numerical simulation of hypersonic inlet flows with equilibrium or finite rate chemistry
NASA Technical Reports Server (NTRS)
Yu, Sheng-Tao; Hsieh, Kwang-Chung; Shuen, Jian-Shun; Mcbride, Bonnie J.
1988-01-01
An efficient numerical program incorporated with comprehensive high temperature gas property models has been developed to simulate hypersonic inlet flows. The computer program employs an implicit lower-upper time marching scheme to solve the two-dimensional Navier-Stokes equations with variable thermodynamic and transport properties. Both finite-rate and local-equilibrium approaches are adopted in the chemical reaction model for dissociation and ionization of the inlet air. In the finite rate approach, eleven species equations coupled with fluid dynamic equations are solved simultaneously. In the local-equilibrium approach, instead of solving species equations, an efficient chemical equilibrium package has been developed and incorporated into the flow code to obtain chemical compositions directly. Gas properties for the reaction products species are calculated by methods of statistical mechanics and fit to a polynomial form for C(p). In the present study, since the chemical reaction time is comparable to the flow residence time, the local-equilibrium model underpredicts the temperature in the shock layer. Significant differences of predicted chemical compositions in shock layer between finite rate and local-equilibrium approaches have been observed.
Investigation of hypersonic rarefied flow on a spherical nose of the AOTV
NASA Technical Reports Server (NTRS)
Jain, Amolak C.; Woods, G. Hamilton
1987-01-01
The Navier-Stokes (NS) equations were integrated numerically for investigating the flow characteristics on the forepart of the spherical nose of a space vehicle such as the AOTV or AFE by a modified Accelerated Successive Replacement (ASR) scheme under hypersonic rarefied conditions. Technical feasibility of the mathematical approach was demonstrated by computing the flowfield on a spherical nose under conditions that the AFE encounters at times t = 15 and 20 seconds after its reentry into the atmosphere. Local similar solutions for the merged layer flow along the stagnation line of the sphere were developed. These are correct to the same degree of accuracy as the NS equations. These solutions provided stagnation line boundary conditions for the domain of integration on the spherical noise. Also, a parametric study of the stagnation line solution was made with a view to understand the flow characteristics in tunnels with different ambient fluids. Analytical expressions for surface slip temperature, jump conditions, and concentration level in the presence of the real gas effects at the top of the Knudsen layer were derived and used to calculate the stagnation line flowfield with nonequilibrium dissociation and ionization. A number of graphics were drawn to illustrate the basic physics of the flowfields. The present analysis can be extended to include real gas effects and to bodies of arbitrary shapes. It can further provide boundary conditions for integrating the NS equations in the near wake region.
Comparative study on aerodynamic heating under perfect and nonequilibrium hypersonic flows
NASA Astrophysics Data System (ADS)
Wang, Qiu; Li, JinPing; Zhao, Wei; Jiang, ZongLin
2016-02-01
In this study, comparative heat flux measurements for a sharp cone model were conducted by utilizing a high enthalpy shock tunnel JF-10 and a large-scale shock tunnel JF-12, responsible for providing nonequilibrium and perfect gas flows, respectively. Experiments were performed at the Key Laboratory of High Temperature Gas Dynamics (LHD), Institute of Mechanics, Chinese Academy of Sciences. Corresponding numerical simulations were also conducted in effort to better understand the phenomena accompanying in these experiments. By assessing the consistency and accuracy of all the data gathered during this study, a detailed comparison of sharp cone heat transfer under a totally different kind of freestream conditions was build and analyzed. One specific parameter, defined as the product of the Stanton number and the square root of the Reynold number, was found to be more characteristic for the aerodynamic heating phenomena encountered in hypersonic flight. Adequate use of said parameter practically eliminates the variability caused by the deferent flow conditions, regardless of whether the flow is in dissociation or the boundary condition is catalytic. Essentially, the parameter identified in this study reduces the amount of ground experimental data necessary and eases data extrapolation to flight.
Computation of axisymmetric and ionized hypersonic flows using particle and continuum methods
NASA Technical Reports Server (NTRS)
Boyd, Iain D.; Gokcen, Tahir
1994-01-01
Comparisons between particle and continuum simulations of hypersonic near-continuum flows are presented. The particle approach employs the direct simulation Monte Carlo (DSMC) method, and the continuum approach solves the appropriate equations of fluid flow. Both simulations have thermochemistry models for air implemented including ionization. A new axisymmetric DSMC code that is efficiently vectorized is developed for this study. In this DSMC code, particular attention is paid to matching the relaxation rates employed in the continuum approach. This investigation represents a continuum of a previous study that considered thermochemical relaxation in one-dimensional shock waves of nitrogen. Comparison of the particle and continuum methods is first made for an axisymmetric blunt-body flow of air at 7 km/s. Very good agreement is obtained for the two solutions. The two techniques also compare well for a one-dimensional shock wave in air at 10 km/s. In both applications, the results are found to be sensitive to various aspects of the chemistry model employed.
Computation of axisymmetric and ionized hypersonic flows using particle and continuum methods
NASA Technical Reports Server (NTRS)
Boyd, Iain D.; Gokcen, Tahir
1994-01-01
Comparisons between particle and continuum simulations of hypersonic near-continuum flows are presented. The particle approach employs the direct simulation Monte Carlo (DSMC) method, and the continuum approach solves the appropriate equations of fluid flow. Both simulations have thermochemistry models for air implemented including ionization. A new axisymmetric DSMC code that is efficiently vectorized is developed for this study. In this DSMC code, particular attention is paid to matching the relaxation rates employed in the continuum approach. This investigation represents a continuum of a previous study that considered thermochemical relaxation in one-dimensional shock waves of nitrogen. Comparison of the particle and continuum methods is first made for an axisymmetric blunt-body flow of air at 7 km/s. Very good agreement is obtained for the two solutions. The two techniques also compare well for a one-dimensional shock wave in air at 10 km/s. In both applications, the results are found to be sensitive to various aspects of the chemistry model employed.
Heat Transfer to Surfaces of Finite Catalytic Activity in Frozen Dissociated Hypersonic Flow
NASA Technical Reports Server (NTRS)
Chung, Paul M.; Anderson, Aemer D.
1961-01-01
The heat transfer due to catalytic recombination of a partially dissociated diatomic gas along the surfaces of two-dimensional and axisymmetric bodies with finite catalytic efficiencies is studied analytically. An integral method is employed resulting in simple yet relatively complete solutions for the particular configurations considered. A closed form solution is derived which enables one to calculate atom mass-fraction distribution, therefore catalytic heat transfer distribution, along the surface of a flat plate in frozen compressible flow with and without transpiration. Numerical calculations are made to determine the atom mass-fraction distribution along an axisymmetric conical body with spherical nose in frozen hypersonic compressible flow. A simple solution based on a local similarity concept is found to be in good agreement with these numerical calculations. The conditions are given for which the local similarity solution is expected to be satisfactory. The limitations on the practical application of the analysis to the flight of the blunt bodies in the atmosphere are discussed. The use of boundary-layer theory and the assumption of frozen flow restrict application of the analysis to altitudes between about 150,000 and 250,000 feet.
Estimation of Stability Derivatives in Pitch for an Oscillating Wedge in Hypersonic Flow
NASA Astrophysics Data System (ADS)
Afghan Khan, Sher; Asadullah, Mohammed
2017-03-01
A similitude has been obtained for an oscillating two-dimensional wedge in pitch with attached shock at high angle of incidence in hypersonic flow. A strip theory, in which flow at infinite span wise location is two dimensional and independent of each other is being used. Further this unites the similitude with piston theory to give one dimensional piston theory with large flow deflection. Closed form solution is obtained for stiffness and damping derivatives in pitch. The present theory is valid when the shock wave is attached with the nose of the wedge. With the increase in semi vertex angle of the wedge the stiffness and the damping derivatives are found to increase progressively, and with the increase in the Mach number both stiffness and damping derivatives are found to decrease then with increase in Mach number it becomes independent of Mach number. From the amalgamation of theory developed, we can evaluate the stiffness and damping derivatives for in pitch for a wide range of Mach numbers, for various pivot positions, and angle of attack. Significantly the same results are being validated with the analytical results of Liu and Hui [4] and Lighthill’s [3] theory for some cases.
A survey of simulation and diagnostic techniques for hypersonic nonequilibrium flows
NASA Technical Reports Server (NTRS)
Sharma, Surendra P.; Park, Chul
1987-01-01
The possible means of simulating nonequilibrium reacting flows in hypersonic environments, and the required diagnostic techniques, are surveyed in two categories: bulk flow behavior and determination of chemical rate parameters. Flow visualization of shock shapes for validation of computational-fluid dynamic calculations is proposed. The facilities and the operating conditions necessary to produce the required nonequilibrium conditions, the suitable optical techniques, and their sensitivity requirements, are surveyed. Shock-tubes, shock-tunnels, and ballistic ranges in a wide range of sizes and strengths are found to be useful for this purpose, but severe sensitivity requirements are indicated for the optical instruments, which can be met only by using highly-collimated laser sources. Likewise, for the determination of chemical parameters, this paper summarizes the quantities that need to be determined, required facilities and their operating conditions, and the suitable diagnostic techniques and their performance requirements. Shock tubes of various strengths are found to be useful for this purpose. Vacuum ultraviolet absorption and fluorescence spectroscopy and coherent anti-Stokes Raman spectroscopy are found to be the techniques best suited for the measurements of the chemical data.
NASA Astrophysics Data System (ADS)
Fogerty, Erica; Carroll-Nellenback, Jonathan; Frank, Adam; Heitsch, Fabian; Pon, Andy
2017-09-01
We present numerical simulations of reorienting oblique shocks that form in the collision layer between magnetized colliding flows. Reorientation aligns post-shock filaments normal to the background magnetic field. We find that reorientation begins with pressure gradients between the collision region and the ambient medium. This drives a lateral expansion of post-shock gas, which reorients the growing filament from the outside-in (i.e. from the flow/ambient boundary, towards the colliding flows axis). The final structures of our simulations resemble polarization observations of filaments in Taurus and Serpens South, as well as the integral-shaped filament in Orion A. Given the ubiquity of colliding flows in the interstellar medium, shock reorientation may be relevant to the formation of filaments normal to magnetic fields.
NASA Technical Reports Server (NTRS)
Bertin, J. J.; Lamb, J. P.; Center, K. R.; Graumann, B. W.
1971-01-01
Windward and leeward measurements were made for a variety of simulated infinite cylinders exposed to hypersonic streams over an angle of attack from 30 deg to 90 deg. For the range of conditions included in the study, the following conclusions are made: (1) Swept cylinder theory provides a reasonable correlation of the measured laminar heat transfer rates from the plane of symmetry. (2) The boundary layer transition criteria in the plane of symmetry are a function of the transverse curvature. (3) Relaminarization of the circumferential boundary layer for a right circular cylinder was observed at the highest Reynolds number tested. (4) The effect of leeside geometry on the average heat transfer rate can be correlated with a single geometric parameter which is dependent on the location of separation. (5) The relationship of leeward heating to angle of attack is virtually linear for each cross section. (6) No systematic effect of free stream Reynolds number was observed.
Guarendi, Andrew N.; Chandy, Abhilash J.
2013-01-01
Numerical simulations of magnetohydrodynamic (MHD) hypersonic flow over a cylinder are presented for axial- and transverse-oriented dipoles with different strengths. ANSYS CFX is used to carry out calculations for steady, laminar flows at a Mach number of 6.1, with a model for electrical conductivity as a function of temperature and pressure. The low magnetic Reynolds number (≪1) calculated based on the velocity and length scales in this problem justifies the quasistatic approximation, which assumes negligible effect of velocity on magnetic fields. Therefore, the governing equations employed in the simulations are the compressible Navier-Stokes and the energy equations with MHD-related source terms such as Lorentz force and Joule dissipation. The results demonstrate the ability of the magnetic field to affect the flowfield around the cylinder, which results in an increase in shock stand-off distance and reduction in overall temperature. Also, it is observed that there is a noticeable decrease in drag with the addition of the magnetic field. PMID:24307870
Aerodynamic performance investigation on waverider with variable blunt radius in hypersonic flows
NASA Astrophysics Data System (ADS)
Li, Shibin; Wang, Zhenguo; Huang, Wei; Xu, Shenren; Yan, Li
2017-08-01
Waverider is an important candidate for the design of hypersonic vehicles. However, the ideal waverider cannot be manufactured because of its sharp leading edge, so the leading edge should be blunted. In the paper, the HMB solver and laminar flow model have been utilized to obtain the flow field properties around the blunt waverider with the freestream Mach number being 8.0, and several novel strategies have been suggested to improve the aerodynamic performance of blunt waverider. The numerical method has been validated against experimental data, and the Stanton number(St) of the predicted result has been analyzed. The obtained results show good agreement with the experimental data. Stmax decreases by 58% and L/D decreases by 8.2% when the blunt radius increases from 0.0002 m to 0.001 m. ;Variable blunt waverider; is a good compromise for aerodynamic performance and thermal insulation. The aero-heating characteristics are very sensitive to Rmax. The position of the smallest blunt radius has a great effect on the aerodynamic performance. In addition, the type of blunt leading edge has a great effect on the aero-heating characteristics when Rmax is fixed. Therefore, out of several designs, Type 4is the best way to achieve the good overall performance. The ;Variable blunt waverider; not only improves the aerodynamic performance, but also makes the aero-heating become evenly-distributed, leading to better aero-heating characteristics.
NASA Astrophysics Data System (ADS)
Egorov, I. V.; Novikov, A. V.; Fedorov, A. V.
2017-08-01
A method for direct numerical simulation of three-dimensional unsteady disturbances leading to a laminar-turbulent transition at hypersonic flow speeds is proposed. The simulation relies on solving the full three-dimensional unsteady Navier-Stokes equations. The computational technique is intended for multiprocessor supercomputers and is based on a fully implicit monotone approximation scheme and the Newton-Raphson method for solving systems of nonlinear difference equations. This approach is used to study the development of three-dimensional unstable disturbances in a flat-plate and compression-corner boundary layers in early laminar-turbulent transition stages at the free-stream Mach number M = 5.37. The three-dimensional disturbance field is visualized in order to reveal and discuss features of the instability development at the linear and nonlinear stages. The distribution of the skin friction coefficient is used to detect laminar and transient flow regimes and determine the onset of the laminar-turbulent transition.
Guarendi, Andrew N; Chandy, Abhilash J
2013-01-01
Numerical simulations of magnetohydrodynamic (MHD) hypersonic flow over a cylinder are presented for axial- and transverse-oriented dipoles with different strengths. ANSYS CFX is used to carry out calculations for steady, laminar flows at a Mach number of 6.1, with a model for electrical conductivity as a function of temperature and pressure. The low magnetic Reynolds number (<1) calculated based on the velocity and length scales in this problem justifies the quasistatic approximation, which assumes negligible effect of velocity on magnetic fields. Therefore, the governing equations employed in the simulations are the compressible Navier-Stokes and the energy equations with MHD-related source terms such as Lorentz force and Joule dissipation. The results demonstrate the ability of the magnetic field to affect the flowfield around the cylinder, which results in an increase in shock stand-off distance and reduction in overall temperature. Also, it is observed that there is a noticeable decrease in drag with the addition of the magnetic field.
2010-04-01
complete supersonic combustion ramjet configurations. The complexity of these flows requires that experiments in ground based facilities are strongly...flight testing. Therefore, ground based testing facilities were developed and they played an important role since the early era of hypersonic flight...namely hypersonic ground based testing, CFD and flight testing, are becoming economically achievable (Figure 1). The means to arrive at this goal is to
NASA Astrophysics Data System (ADS)
Saile, D.; Gülhan, A.; Henckels, A.; Glatzer, C.; Statnikov, V.; Meinke, M.
2013-06-01
The turbulent wake flow of generic rocket configurations is investigated experimentally and numerically at a freestream Mach number of 6.0 and a unit Reynolds number of 10·106 m-1. The flow condition is based on the trajectory of Ariane V-like launcher at an altitude of 50 km, which is used as the baseline to address the overarching tasks of wake flows in the hypersonic regime like fluid-structural coupling, reverse hot jets and base heating. Experimental results using pressure transducers and the high-speed Schlieren measurement technique are shown to gain insight into the local pressure fluctuations on the base and the oscillations of the recompression shock. This experimental configuration features a wedgeprofiled strut orthogonally mounted to the main body. Additionally, the influence of cylindrical dummy nozzles attached to the base of the rocket is investigated, which is the link to the numerical investigations. Here, the axisymmetric model possesses a cylindrical sting support of the same diameter as the dummy nozzles. The sting support allows investigations for an undisturbed wake flow. A time-accurate zonal Reynolds-Averaged Navier-Stokes/Large Eddy Simulation (RANS/LES) approach is applied to identify shocks, expansion waves, and the highly unsteady recompression region numerically. Subsequently, experimental and numerical results in the strut-averted region are compared with regard to the wall pressure and recompression shock frequency spectra. For the compared configurations, experimental pressure spectra exhibit dominant Strouhal numbers at about SrD = 0.03 and 0.27, and the recompression shock oscillates at 0.2. In general, the pressure and recompression shock fluctuations numerically calculated agree reasonably with the experimental results. The experiments with a blunt base reveal base-pressure spectra with dominant Strouhal numbers at 0.08 at the center position and 0.145, 0.21-0.22, and 0.31-0.33 at the outskirts of the base.
NASA Astrophysics Data System (ADS)
Salvador, I. I.; Minucci, M. A. S.; Toro, P. G. P.; Oliveira, A. C.; Channes, J. B.; Myrabo, L. N.; Nagamatsu, H. T.
2006-05-01
Due to high heat transfer rates in hypersonic flight and its consequent necessity of prohibitively massive thermal protection system, new methods of flow control are required to enable flight in such regimes. Here arises the Direct Energy Air Spike concept, where electromagnetic energy (laser/microwaves) is focalized upstream of the model causing the breakdown of the air and the generation of a Laser Supported Detonation wave which diverts the incoming stream parabolically. In this preliminary work, the heat transfer rates to the surface of a blunt body, downstream the laser induced shock wave, were qualitatively measured and compared with the results without the DEAS. These measurements were conducted with the use of fast response coaxial thermocouples and piezoelectric pressure transducers installed on the surface of the model in the 0.30m IEAv's T2 Hypersonic Shock Tunnel. The laser energy was supplied by a CO2 TEA Laser.
NASA Astrophysics Data System (ADS)
Avallone, F.; Greco, C. S.; Schrijer, F. F. J.; Cardone, G.
2015-04-01
The measurement of the convective wall heat flux in hypersonic flows may be particularly challenging in the presence of high-temperature gradients and when using high-thermal-conductivity materials. In this case, the solution of multidimensional problems is necessary, but it considerably increases the computational cost. In this paper, a low-computational-cost inverse data reduction technique is presented. It uses a recursive least-squares approach in combination with the trust-region-reflective algorithm as optimization procedure. The computational cost is reduced by performing the discrete Fourier transform on the discrete convective heat flux function and by identifying the most relevant coefficients as objects of the optimization algorithm. In the paper, the technique is validated by means of both synthetic data, built in order to reproduce physical conditions, and experimental data, carried out in the Hypersonic Test Facility Delft at Mach 7.5 on two wind tunnel models having different thermal properties.
NASA Technical Reports Server (NTRS)
Sanders, Bobby W.; Weir, Lois J.
2008-01-01
A new hypersonic inlet for a turbine-based combined-cycle (TBCC) engine has been designed. This split-flow inlet is designed to provide flow to an over-under propulsion system with turbofan and dual-mode scramjet engines for flight from takeoff to Mach 7. It utilizes a variable-geometry ramp, high-speed cowl lip rotation, and a rotating low-speed cowl that serves as a splitter to divide the flow between the low-speed turbofan and the high-speed scramjet and to isolate the turbofan at high Mach numbers. The low-speed inlet was designed for Mach 4, the maximum mode transition Mach number. Integration of the Mach 4 inlet into the Mach 7 inlet imposed significant constraints on the low-speed inlet design, including a large amount of internal compression. The inlet design was used to develop mechanical designs for two inlet mode transition test models: small-scale (IMX) and large-scale (LIMX) research models. The large-scale model is designed to facilitate multi-phase testing including inlet mode transition and inlet performance assessment, controls development, and integrated systems testing with turbofan and scramjet engines.
An approximate viscous shock layer approach to calculating hypersonic flows about blunt-nosed bodies
NASA Technical Reports Server (NTRS)
Cheatwood, F. MCN.; Dejarnette, F. R.
1991-01-01
An approximate axisymmetric method has been developed which can reliably calculate fully viscous hypersonic flows over blunt-nosed bodies. By substituting Maslen's second order pressure expression for the normal momentum equation, a simplified form of the viscous shock layer (VSL) equations is obtained. This approach can solve both the subsonic and supersonic regions of the shock layer without a starting solution for the shock shape. Since the method is fully viscous, the problems associated with coupling a boundary-layer solution with an inviscid-layer solution are avoided. This procedure is significantly faster than the parabolized Navier-Stokes (PNS) or VSL solvers and would be useful in a preliminary design environment. Problems associated with a previously developed approximate VSL technique are addressed. Surface heat transfer and pressure predictions are comparable to both VSL results and experimental data. The present technique generates its own shock shape as part of its solution, and therefore could be used to provide more accurate initial shock shapes for higher-order procedures which require starting solutions.
Wavefront sensor testing in hypersonic flows using a laser-spark guide star
NASA Astrophysics Data System (ADS)
Neal, Daniel R.; Armstrong, Darrell J.; Hedlund, Eric; Lederer, Melissa; Collier, Arnold S.; Spring, Charles; Gruetzner, James K.; Hebner, Gregory A.; Mansell, Justin D.
1997-11-01
The flight environment of next-generation theater missile defense interceptors involves hypersonic speeds that place severe aero-thermodynamic loads on missile components including the windows used for optical seekers. These heating effects can lead to significant boresight error and aberration. Ground-based tests are required to characterize these effects. We have developed methods to measure aberrations in seeker windows using a Shack-Hartmann wavefront sensor. Light from a laser or other source with a well known wavefront is passed through the window and falls on the sensor. The sensor uses an array of micro-lenses to generate a grid of focal spots on a CCD detector. The positions of the focal spots provide a measure of the wavefront slope over each micro-lens. The wavefront is reconstructed by integrating the slopes, and analyzed to characterize aberrations. During flight, optical seekers look upstream through a window at 'look angles' angles near 0 degrees relative to the free stream flow. A 0 degree angle corresponds to large angles approaching 90 degrees when measured relative to the normal of the window, and is difficult to simulate using conventional techniques to illuminate the wavefront sensor during wind tunnel tests. For this reason, we developed a technique using laser- induced optical breakdown that allows arbitrary look angles down to 0 degrees.
NASA Astrophysics Data System (ADS)
Hao, Jiaao; Wang, Jingying; Lee, Chunhian
2016-09-01
Effects of two different 11-species chemical reaction models on hypersonic reentry flow simulations are numerically investigated. These two models were proposed by Gupta (1990) and Park (1990) [12,15], respectively. In this study, two typical configurations, the RAM-C II vehicle and FIRE II capsule, are selected as test cases, whose thermo-chemical nonequilibrium flowfields are computed by a multi-block finite volume code using a two-temperature model (a translational-rotational temperature and a vibrational-electron-electronic temperature). In the RAM-C II case, it is indicated that although electron number density distributions of the two reaction models appear in a similar trend, their values are distinctively different. Results of the Gupta's model show a better agreement with the electrostatic probe data, while those of the Park's model are more consistent with the reflectometers data. Both models give similar temperature distributions. In the FIRE II case, the two models yield significantly different distribution profiles of ions and electrons, whose differences could reach an order of magnitude. In addition, an abnormal nonequilibrium relaxation process in the shock layer is found in the FIRE II flowfield simulated by the Gupta's model, which proves to be a consequence of electron impact ionization reactions.
Nonequilibrium Rotational Temperature Measurements over Flat Plates in Hypersonic Rarefied Gas Flow
NASA Astrophysics Data System (ADS)
Tsuboi, Nobuyuki; Matsumoto, Yoichiro
2008-12-01
An experimental study of the interaction between a shock wave and a boundary layer over a flat plate with a sharp leading edge in hypersonic rarefied gas flow is presented. Experiments in a low-density wind tunnel using an electron beam probe were conducted at the Shock Wave Laboratory, RWTH Aachen, Germany. Rotational temperatures for stagnation temperatures of T0 = 1000 K and Kn = 0.028 based on a reference length of 0.05 m were calculated using Robben and Talbot's method. The rotational temperature profiles at X = 3 mm for LE = 45 and 90 are 100 K larger than those for LE = 30. This means that a bow shock wave in front of the leading edge affects the rotational temperature profiles over the plate. The rotational energy distributions differ from the Maxwell-Boltzmann distributions and they are non-equilibrium distributions. The rotational temperature profiles for α = 12 deg are also 100 K larger than those for α = 0 deg. The feature of the rotational temperature over the plate for α = 12 deg is similar to that for LE = 45 and 90 deg.
NASA Technical Reports Server (NTRS)
Gonor, A. L. (Editor)
1982-01-01
The results of flow around wings, the determination of the optimal form, and the interaction of the wake with the accompanying flow at supersonic and hypersonic speeds of the free-stream flow are given. Methods of numerical and analytical calculation of one dimensional unsteady and two dimensional steady motions of fuel-gas mixtures with exothermic reactions are also considered.
High Enthalpy Effects on Two Boundary Layer Disturbances in Supersonic and Hypersonic Flow
NASA Astrophysics Data System (ADS)
Wagnild, Ross Martin
The fluid flow phenomenon of boundary layer transition is a complicated and difficult process to model and predict. The importance of the state of the boundary layer with regard to vehicle design cannot be understated. The high enthalpy environment in which high speed vehicles operate in further complicates the transition process by adding several more degrees of freedom. In this environment, the internal properties of the gas can stabilize or destabilize the boundary layer as well as modify the disturbances that cause transition. In the current work, the interaction of two types of disturbances with the high enthalpy flow environment are analyzed. The first is known as a second mode disturbance, which is acoustic in nature. The second type is known as a transient growth disturbance and is associated with flows behind roughness elements. Theoretical analyses, linear stability analyses, and computation fluid dynamics (CFD) are used to determine the ways in which these disturbances interact with the high enthalpy environment as well as the consequences of these interactions. First, acoustic wave are directly studied in order to gain a basic understanding of the response of second mode disturbances in the high enthalpy boundary layer. Next, this understanding is used in interpreting the results of several computations attempting to simulate the flow through a high enthalpy flow facility as well as experiments attempting to take advantage of the acoustic interaction with the high enthalpy environment. Because of the difficulty in modeling these experiments, direct simulations of acoustic waves in a hypersonic flow of a gas with molecular vibration are performed. Lastly, compressible transient growth disturbances are simulated using a linear optimal disturbance solver as well as a CFD solver. The effect of an internal molecular process on this type of disturbance is tested through the use of a vibrational mode. It is the goal of the current work to reinforce the
NASA Technical Reports Server (NTRS)
Warsi, Z. U. A.; Weed, R. A.; Thompson, J. F.
1980-01-01
A formulation of the complete Navier-Stokes problem for a viscous hypersonic flow in general curvilinear coordinates is presented. This formulation is applicable to both the axially symmetric and three dimensional flows past bodies of revolution. The equations for the case of zero angle of attack were solved past a circular cylinder with hemispherical caps by point SOR finite difference approximation. The free stream Mach number and the Reynolds number for the test case are respectively 22.04 and 168883. The whole algorithm is presented in detail along with the preliminary results for pressure, temperature, density and velocity distributions along the stagnation line.
NASA Technical Reports Server (NTRS)
Bardina, Jorge; Lombard, C. K.
1987-01-01
The Bardina and Lombard (1985) three-dimensional CSCM Navier-Stokes method is presently extended to the simulation of complex hypersonic reentry vehicle external flows at angle of attack. The robust stability of the method derives from the combination of conservative implicit upwind flux difference splitting with a three-dimensional diagonally-dominant approximate factorization and relaxation scheme and characteristic-based implicit boundary approximations. The method's efficiency derives from an implicit symmetric Gauss-Seidel 'method of planes' relaxation scheme with alternating directional space marching sweeps along the flow coordinate direction.
NASA Astrophysics Data System (ADS)
Golikov, E. A.; Izmodenov, V. V.; Alexashov, D. B.
2017-02-01
In the present paper we consider the steady flow produced by the hypersonic magnetized spherical source into the steady unmagnetized medium. The magnetic field of the source is considered purely azimuthal. The basic dimensionless parameter of the problem is Alfvén number at the inflow boundary at the source equatorial plane, which we denote as ε. We first review the numerical results, that was already presented in our recent paper (Golikov et al, 2017). According to our numerical results there should be a tangential discontinuity dividing the flow and the ambient medium. In the present paper we analytically derive the asymptotic behaviour of the tangential discontinuity radius as epsilon approaches zero.
Modeling and simulation of radiation from hypersonic flows with Monte Carlo methods
NASA Astrophysics Data System (ADS)
Sohn, Ilyoup
During extreme-Mach number reentry into Earth's atmosphere, spacecraft experience hypersonic non-equilibrium flow conditions that dissociate molecules and ionize atoms. Such situations occur behind a shock wave leading to high temperatures, which have an adverse effect on the thermal protection system and radar communications. Since the electronic energy levels of gaseous species are strongly excited for high Mach number conditions, the radiative contribution to the total heat load can be significant. In addition, radiative heat source within the shock layer may affect the internal energy distribution of dissociated and weakly ionized gas species and the number density of ablative species released from the surface of vehicles. Due to the radiation total heat load to the heat shield surface of the vehicle may be altered beyond mission tolerances. Therefore, in the design process of spacecrafts the effect of radiation must be considered and radiation analyses coupled with flow solvers have to be implemented to improve the reliability during the vehicle design stage. To perform the first stage for radiation analyses coupled with gas-dynamics, efficient databasing schemes for emission and absorption coefficients were developed to model radiation from hypersonic, non-equilibrium flows. For bound-bound transitions, spectral information including the line-center wavelength and assembled parameters for efficient calculations of emission and absorption coefficients are stored for typical air plasma species. Since the flow is non-equilibrium, a rate equation approach including both collisional and radiatively induced transitions was used to calculate the electronic state populations, assuming quasi-steady-state (QSS). The Voigt line shape function was assumed for modeling the line broadening effect. The accuracy and efficiency of the databasing scheme was examined by comparing results of the databasing scheme with those of NEQAIR for the Stardust flowfield. An accuracy of
NASA Technical Reports Server (NTRS)
Blanchard, R. C.; Walberg, G. D.
1980-01-01
Results of an investigation to determine the full scale drag coefficient in the high speed, low density regime of the Viking lander capsule 1 entry vehicle are presented. The principal flight data used in the study were from onboard pressure, mass spectrometer, and accelerometer instrumentation. The hypersonic continuum flow drag coefficient was unambiguously obtained from pressure and accelerometer data; the free molecule flow drag coefficient was indirectly estimated from accelerometer and mass spectrometer data; the slip flow drag coefficient variation was obtained from an appropriate scaling of existing experimental sphere data. Comparison of the flight derived drag hypersonic continuum flow regime except for Reynolds numbers from 1000 to 100,000, for which an unaccountable difference between flight and ground test data of about 8% existed. The flight derived drag coefficients in the free molecule flow regime were considerably larger than those previously calculated with classical theory. The general character of the previously determined temperature profile was not changed appreciably by the results of this investigation; however, a slightly more symmetrical temperature variation at the highest altitudes was obtained.
Flow field around a sphere colliding against a wall.
NASA Astrophysics Data System (ADS)
Zenit, R.; Hunt, M. L.
1998-11-01
This study investigates the flow field and the fluid agitation generated by particle collisions. The motion of a particle towards a wall, or towards another particle, will result in a collision if the Reynolds number of the flow is large. As the particle approaches the wall, the fluid in the gap between the particle and the wall will be displaced. When the particle touches the wall and rebounds, the direction of the flow will reverse. This process produces a considerable agitation in the fluid phase. To study this process an immersed pendulum experiment was built to produce controlled collisions of particles. A fine string is attached to a particle, which is positioned at rest from some initial angle. Once released, the particle accelerates towards a wall, or to another suspended particle, resulting in a collision. The fluid is seeded with neutrally buoyant micro-spheres, which illuminated by a laser sheet serve as flow tracers. The motion of the particles and tracers is recorded using a high speed digital camera. The images are digitally processed to calculate displacements and velocities for different times before and after the collision. Flow fields are obtained for different impact velocities, particle diameters and solid-fluid density ratios, as well as for particle-wall and particle-particle collisions. Preliminary results show that for the flow conditions tested, the rebound of the particle is dependent on the shape of the wake behind the particle at the moment of collision, and not only on the flow in the gap between the particle and the wall. The amount of collision-generated agitation appears to increase with impact velocity and density ratio.
NASA Astrophysics Data System (ADS)
Henckels, A.; Kreins, A. F.; Maurer, F.
For the investigation of hypersonic heat transfer phenomena experimental studies were performed in the blowdown facility 112K of DI.R Cologne. This facilily with 60 cm nozzle exit diameter provides test run times up to 40 seconds at test conditions of Mach number 5.3 or 11.2 and Reynolds numbers between 0.3 and 19 million per meter. An infrared camera system (Inframetrics 600) installed inside the plenum chamber of this tunnel recorded the establishing temperature distribution on the model surface during the whole run time. To get a comprehensive information about the character of the flow field the infrared results were supported by additional diagnostic techniques like coincidence Schlieren optics, Pitot pressure measurements and oil flow visualization. The paper will demonstrate the capability of the applied measurement techniques by presenting infrared measurement results as for instance from a study of an oblique shock front impinging on the laminar boundary layer of a flat plate model. Also results from studies concerning a longitudinal corner flow field and the delta wing configuration are included. All these activities will be accompanied by efforts to improve the present infrared and oil flow visualition techniques, driven by the needs of increased precision for numerical validation in hypersonic aerothermodynamics.
Interaction theory of hypersonic laminar near-wake flow behind an adiabatic circular cylinder
NASA Astrophysics Data System (ADS)
Hinman, W. Schuyler; Johansen, C. T.
2016-11-01
The separation and shock wave formation on the aft-body of a hypersonic adiabatic circular cylinder were studied numerically using the open source software OpenFOAM. The simulations of laminar flow were performed over a range of Reynolds numbers (8× 10^3 < Re < 8× 10^4) at a free-stream Mach number of 5.9. Off-body viscous forces were isolated by controlling the wall boundary condition. It was observed that the off-body viscous forces play a dominant role compared to the boundary layer in displacement of the interaction onset in response to a change in Reynolds number. A modified free-interaction equation and correlation parameter has been presented which accounts for wall curvature effects on the interaction. The free-interaction equation was manipulated to isolate the contribution of the viscous-inviscid interaction to the overall pressure rise and shock formation. Using these equations coupled with high-quality simulation data, the underlying mechanisms resulting in Reynolds number dependence of the lip-shock formation were investigated. A constant value for the interaction parameter representing the part of the pressure rise due to viscous-inviscid interaction has been observed at separation over a wide range of Reynolds numbers. The effect of curvature has been shown to be the primary contributor to the Reynolds number dependence of the free-interaction mechanism at separation. The observations in this work have been discussed here to create a thorough analysis of the Reynolds number-dependent nature of the lip-shock.
The application and analysis of liquid crystal thermographs in short duration hypersonic flow
NASA Astrophysics Data System (ADS)
Babinsky, H.; Edwards, J. A.
1993-01-01
Liquid crystal thermography is applied here in two different hypersonic wind tunnels, a Mach 8 gun tunnel and a Mach 5 blow-down wind tunnel. A technique to extract surface heat transfer levels is introduced. It is shown how the method can be adopted to the specific difficulties encountered in hypersonic short duration facilities. Digital image processing is used to automatically obtain quantitative information from the liquid crystal experiments, the main features of this process are discussed. Results are shown for a variety of models and the main experimental errors are discussed.
NASA Technical Reports Server (NTRS)
Narain, J. P.; Muramoto, K. K.; Lawrence, S. L.
1991-01-01
A three-dimensional parabolized Navier-Stokes computer code which employs an upwind algorithm is used to conduct a numerical study of an advanced maneuvering reentry vehicle configuration. Comparisons between numerical solutions and experimental data are presented for surface pressure, wall heat flux, and overall forces and moments. The effects of angle of attack, angle of yaw, and surface mass injection are investigated. Good agreement is observed between the calculated and measured data. The results of this investigation demonstrate the accuracy and efficiency of an upwind scheme in predicting the hypersonic flow field characteristics about a complex configuration.
NASA Technical Reports Server (NTRS)
Codding, William H.; Lombard, C. K.; Yang, J. Y.
1988-01-01
The Conservative Supra-Characteristic Method (CSCM) Navier-Stokes solver is applied to ascertain the problems inherent in the design of a nominal Mach 14 nozzle for NASA-Ames' 3.5-ft Hypersonic Wind Tunnel; attention is given to the effects of boundary layer cooling systems on the aerodynamic redesign of the nozzle throat region. Complete nozzle flowfields are calculated with and without slot injection of either hot or cold fluid into the boundary layer just upstream of the throat, as well as with alternatively adiabatic and cold walls. The CSCM method is capable of resolving subtle differences in the flows.
NASA Technical Reports Server (NTRS)
Sharma, Surendra P.
1992-01-01
Basic requirements for a ground test facility simulating low density hypersonic flows are discussed. Such facilities should be able to produce shock velocities in the range of 10-17 km/sec in an initial pressure of 0.010 to 0.050 Torr. The facility should be equipped with diagnostics systems to be able to measure the emitted radiation, characteristic temperatures and populations in various energy levels. In the light of these requirements, NASA Ames's electric arc-driven low density shock tube facility is described and available experimental diagnostics systems and computational tools are discussed.
NASA Astrophysics Data System (ADS)
Regert, T.; Horvath, I.; Buchlin, J.-M.; Masutti, D.; Chazot, O.; Vetrano, M. R.; Lapebie, C.; Le Gallic, C.
2017-06-01
This paper presents and discusses the results of tests of breakup phenomenon of liquid water into a hypersonic cross §ow from the surface of a 7 degree half-angle cone model at zero degree angle of incidence. The present work shows the dependence of the liquid phase characteristics on the cross-section area of the injection hole in a Mach 6 cross flow. The results are analyzed qualitatively by imaging, by Interferometric Laser Imaging for Droplet Sizing (ILIDS), and by InfraRed Light Extinction Spectroscopy (IR-LES). Conclusions are drawn concerning the droplet size distribution and the liquid §ow ¦eld characteristics.
NASA Technical Reports Server (NTRS)
Codding, William H.; Lombard, C. K.; Yang, J. Y.
1988-01-01
The Conservative Supra-Characteristic Method (CSCM) Navier-Stokes solver is applied to ascertain the problems inherent in the design of a nominal Mach 14 nozzle for NASA-Ames' 3.5-ft Hypersonic Wind Tunnel; attention is given to the effects of boundary layer cooling systems on the aerodynamic redesign of the nozzle throat region. Complete nozzle flowfields are calculated with and without slot injection of either hot or cold fluid into the boundary layer just upstream of the throat, as well as with alternatively adiabatic and cold walls. The CSCM method is capable of resolving subtle differences in the flows.
Second-order small disturbance theory for hypersonic flow over power-law bodies. Ph.D. Thesis
NASA Technical Reports Server (NTRS)
Townsend, J. C.
1974-01-01
A mathematical method for determining the flow field about power-law bodies in hypersonic flow conditions is developed. The second-order solutions, which reflect the effects of the second-order terms in the equations, are obtained by applying the method of small perturbations in terms of body slenderness parameter to the zeroth-order solutions. The method is applied by writing each flow variable as the sum of a zeroth-order and a perturbation function, each multiplied by the axial variable raised to a power. The similarity solutions are developed for infinite Mach number. All results obtained are for no flow through the body surface (as a boundary condition), but the derivation indicates that small amounts of blowing or suction through the wall can be accommodated.
NASA Technical Reports Server (NTRS)
Marconi, F.; Salas, M.; Yaeger, L.
1976-01-01
A numerical procedure has been developed to compute the inviscid super/hypersonic flow field about complex vehicle geometries accurately and efficiently. A second order accurate finite difference scheme is used to integrate the three dimensional Euler equations in regions of continuous flow, while all shock waves are computed as discontinuities via the Rankine Hugoniot jump conditions. Conformal mappings are used to develop a computational grid. The effects of blunt nose entropy layers are computed in detail. Real gas effects for equilibrium air are included using curve fits of Mollier charts. Typical calculated results for shuttle orbiter, hypersonic transport, and supersonic aircraft configurations are included to demonstrate the usefulness of this tool.
NASA Astrophysics Data System (ADS)
Riley, Zachary Bryce
The use of thin-gauge, light-weight structures in combination with the severe aero-thermodynamic loading makes reusable hypersonic cruise vehicles prone to fluid-thermal-structural interactions. These interactions result in surface perturbations in the form of temperature changes and deformations that alter the stability and eventual transition of the boundary layer. The state of the boundary layer has a significant effect on the aerothermodynamic loads acting on a hypersonic vehicle. The inherent relationship between boundary-layer stability, aerothermodynamic loading, and surface conditions make the interaction between the structural response and boundary-layer transition an important area of study in high-speed flows. The goal of this dissertation is to examine the interaction between boundary layer transition and the response of aerothermally compliant structures. This is carried out by first examining the uncoupled problems of: (1) structural deformation and temperature changes altering boundary-layer stability and (2) the boundary layer state affecting structural response. For the former, the stability of boundary layers developing over geometries that typify the response of surface panels subject to combined aerodynamic and thermal loading is numerically assessed using linear stability theory and the linear parabolized stability equations. Numerous parameters are examined including: deformation direction, deformation location, multiple deformations in series, structural boundary condition, surface temperature, the combined effect of Mach number and altitude, and deformation mode shape. The deformation-induced pressure gradient alters the boundary-layer thickness, which changes the frequency of the most-unstable disturbance. In regions of small boundary-layer growth, the disturbance frequency modulation resulting from a single or multiple panels deformed into the flowfield is found to improve boundary-layer stability and potentially delay transition. For the
NASA Technical Reports Server (NTRS)
Agarwal, R.; Rakich, J. V.
1978-01-01
Computational results, obtained with a parabolic Navier-Stokes marching code, are presented for hypersonic viscous flow past spinning sharp and blunt cones at angle of attack. The code takes into account the asymmetries in the flow field resulting from spinning motion and computes the asymmetric shock shape, crossflow and streamwise shear, heat transfer, crossflow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results. In addition, a new criterion for defining crossflow separation behind spinning bodies is introduced which generalizes the Moore-Rott-Sears criterion for two-dimensional unsteady separation. A condition which characterizes the onset of separation in the flow field is defined.
NASA Technical Reports Server (NTRS)
Agarwal, R.; Rakich, J. V.
1978-01-01
Computational results, obtained with a parabolic Navier-Stokes marching code, are presented for hypersonic viscous flow past spinning sharp and blunt cones at angle of attack. The code takes into account the asymmetries in the flow field resulting from spinning motion and computes the asymmetric shock shape, crossflow and streamwise shear, heat transfer, crossflow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results. In addition, a new criterion for defining crossflow separation behind spinning bodies is introduced which generalizes the Moore-Rott-Sears criterion for two-dimensional unsteady separation. A condition which characterizes the onset of separation in the flow field is defined.
NASA Astrophysics Data System (ADS)
Xingxing, Chen; Zhihui, Wang; Yongliang, Yu
2016-11-01
Hypersonic chemical non-equilibrium gas flows around blunt nosed bodies are studied in the present paper to investigate the Reynolds analogy relation on curved surfaces. With a momentum and energy transfer model being applied through boundary layers, influences of molecular dissociations and recombinations on skin frictions and heat fluxes are separately modeled. Expressions on the ratio of Cf / Ch (skin friction coefficient to heat flux) are presented along the surface of circular cylinders under the ideal dissociation gas model. The analysis indicates that molecular dissociations increase the linear distribution of Cf / Ch, but the nonlinear Reynolds analogy relation could ultimately be obtained in flows with larger Reynolds numbers and Mach numbers, where the decrease of wall heat flux by molecular recombinations signifies. The present modeling and analyses are also verified by the DSMC calculations on nitrogen gas flows.
Flow resolution and domain of influence in rarefied hypersonic blunt-body flows
NASA Technical Reports Server (NTRS)
Haas, Brian L.
1993-01-01
The study assesses the effects of the upstream domain size and grid resolution upon flow properties and body aerodynamics computed for rarefied flows over cold blunt bodies with a direct simulation Monte Carlo (DSMC) particle method. Empirical correlations are suggested for aerodynamic coefficients for two-dimensional flows past a perpendicular flat plate. Free-stream parameters which were varied in the study include the Mach number, Knudsen number, surface temperature, and intermolecular potential. Insufficient grid resolution leads to overprediction of aerodynamic heating and forces in the DSMC method. Solution accuracy correlates well with the Reynolds number defined at the wall temperature and the stagnation mean free path relative to the cell dimension. Insufficient upstream domain size in the DSMC method leads to overprediction of heating and drag. Errors in aerodynamic coefficients correlate well with the distance ahead of the body where flow temperature reaches half of its peak value. Simulation of a hard-sphere gas is more sensitive to grid resolution, while simulation of a Maxwell gas is more sensitive to upstream domain size.
Experimental studies of shock-wave/wall-jet interaction in hypersonic flow, part A
NASA Technical Reports Server (NTRS)
Holden, Michael S.; Rodriguez, Kathleen
1994-01-01
Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident-shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not
Experimental studies of shock-wave/wall-jet interaction in hypersonic flow
NASA Technical Reports Server (NTRS)
Holden, Michael S.; Rodriguez, Kathleen
1994-01-01
Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not
Sensitivity of heat fluxes in hypersonic CO2 flows to the state-to-state kinetic schemes
NASA Astrophysics Data System (ADS)
Armenise, I.; Kustova, E.
2016-11-01
Kinetics and heat transfer in a CO2/CO/O2/O/C mixture in a hypersonic boundary layer is studied using a state-to-state vibrational-chemical kinetic model. The CO2 molecule is detailed in its symmetric stretching, bending and asymmetric stretching modes, which are strongly coupled through inter-mode vibrational energy transfers. Two sets of rate coefficients for the vibrational energy transitions are used. Different kinetic schemes including various physical and chemical processes are assessed. The heat flux is calculated, in the framework of the modified Chapman-Enskog theory, accounting for the vibrational states of involved molecules. Comparisons with results obtained using a simplified model, including mainly vibrational levels of the asymmetric stretching mode, are carried out. It is shown that VT transitions in the symmetric and asymmetric modes do not alter the flow and can be neglected. The heat flux is not sensitive to the rates of vibrational energy transitions but depends noticeably on the processes implemented to the kinetic scheme. Using the simplified model yields under-predicted surface heat fluxes; nevertheless we can recommend it for fast estimates of the fluid dynamic variables and heat transfer in hypersonic flows since its implementation essentially reduces computational costs.
Cristofolini, Andrea; Neretti, Gabriele; Borghi, Carlo A.
2012-08-01
This work proposes an experimental analysis on the magneto hydro dynamic (MHD) interaction induced by a magnetic test body immersed into a hypersonic argon flow. The characteristic plasma parameters are measured. They are related to the voltages arising in the Hall direction and to the variation of the fluid dynamic properties induced by the interaction. The tests have been performed in a hypersonic wind tunnel at Mach 6 and Mach 15. The plasma parameters are measured in the stagnation region in front of the nozzle of the wind tunnel and in the free stream region at the nozzle exit. The test body has a conical shape with the cone axis in the gas flow direction and the cone vertex against the flow. It is placed at the nozzle exit and is equipped with three permanent magnets. In the configuration adopted, the Faraday current flows in a closed loop completely immersed into the plasma of the shock layer. The electric field and the pressure variation due to MHD interaction have been measured on the test body walls. Microwave adsorption measurements have been used for the determination of the electron number density and the electron collision frequency. Continuum recombination radiation and line radiation emissions have been detected. The electron temperature has been determined by means of the spectroscopic data by using different methods. The electron number density has been also determined by means of the Stark broadening of H{sub {alpha}} and the H{sub {beta}} lines. Optical imaging has been utilized to visualize the pattern of the electric current distribution in the shock layer around the test body. The experiments show a considerable effect of the electromagnetic forces produced by the MHD interaction acting on the plasma flow around the test body. A comparison of the experimental data with simulation results shows a good agreement.
NASA Astrophysics Data System (ADS)
Cristofolini, Andrea; Neretti, Gabriele; Borghi, Carlo A.
2012-08-01
This work proposes an experimental analysis on the magneto hydro dynamic (MHD) interaction induced by a magnetic test body immersed into a hypersonic argon flow. The characteristic plasma parameters are measured. They are related to the voltages arising in the Hall direction and to the variation of the fluid dynamic properties induced by the interaction. The tests have been performed in a hypersonic wind tunnel at Mach 6 and Mach 15. The plasma parameters are measured in the stagnation region in front of the nozzle of the wind tunnel and in the free stream region at the nozzle exit. The test body has a conical shape with the cone axis in the gas flow direction and the cone vertex against the flow. It is placed at the nozzle exit and is equipped with three permanent magnets. In the configuration adopted, the Faraday current flows in a closed loop completely immersed into the plasma of the shock layer. The electric field and the pressure variation due to MHD interaction have been measured on the test body walls. Microwave adsorption measurements have been used for the determination of the electron number density and the electron collision frequency. Continuum recombination radiation and line radiation emissions have been detected. The electron temperature has been determined by means of the spectroscopic data by using different methods. The electron number density has been also determined by means of the Stark broadening of Hα and the Hβ lines. Optical imaging has been utilized to visualize the pattern of the electric current distribution in the shock layer around the test body. The experiments show a considerable effect of the electromagnetic forces produced by the MHD interaction acting on the plasma flow around the test body. A comparison of the experimental data with simulation results shows a good agreement.
Schweigert, I. V.
2012-08-15
The plasma sheath near the surface of a hypersonic aircraft formed under associative ionization behind the shock front shields the transmission and reception of radio signals. Using two-dimensional kinetic particle-in-cell simulations, we consider the change in plasma-sheath parameters near a flat surface in a hypersonic flow under the action of electrical and magnetic fields. The combined action of a high-frequency 2-MHz capacitive discharge, a constant voltage, and a magnetic field on the plasma sheath allows the local electron density to be reduced manyfold.
Aero-thermal analysis of lifting body configurations in hypersonic flow
NASA Astrophysics Data System (ADS)
Kumar, Sachin; Mahulikar, Shripad P.
2016-09-01
The aero-thermal analysis of a hypersonic vehicle is of fundamental interest for designing its thermal protection system. The aero-thermal environment predictions over several critical regions of the hypothesized lifting body vehicle, including the stagnation region of the nose-cap, cylindrically swept leading edges, fuselage-upper, and fuselage-lower surfaces, are discussed. The drag (Λ=70°) and temperature (Λ=80°) minimized sweepback angles are considered in the configuration design of the two hypothesized lifting body shape hypersonic vehicles. The main aim of the present study is to analyze and compare the aero-thermal characteristics of these two lifting body configurations at same heat capacity. Accordingly, a Computational Fluid Dynamics simulation has been carried out at Mach number (M∞=7), H=35 km altitude with zero Angle of Attack. Finally, the material selection for thermal protection system based on these predictions and current methodology is described.
Experimental studies of transpiration cooling with shock interaction in hypersonic flow, part B
NASA Technical Reports Server (NTRS)
Holden, Michael S.
1994-01-01
This report describes the result of experimental studies conducted to examine the effects of the impingement of an oblique shock on the flowfield and surface characteristics of a transpiration-cooled wall in turbulent hypersonic flow. The principal objective of this work was to determine whether the interaction between the oblique shock and the low-momentum region of the transpiration-cooled boundary layer created a highly distorted flowfield and resulted in a significant reduction in the cooling effectiveness of the transpiration-cooled surface. As a part of this program, we also sought to determine the effectiveness of transpiration cooling with nitrogen and helium injectants for a wide range of blowing rates under constant-pressure conditions in the absence of shock interaction. This experimental program was conducted in the Calspan 48-Inch Shock Tunnel at nominal Mach numbers of 6 and 8, for a Reynolds number of 7.5 x 10(exp 6). For these test conditions, we obtained fully turbulent boundary layers upstream of the interaction regions over the transpiration-cooled segment of the flat plate. The experimental program was conducted in two phases. In the first phase, we examined the effects of mass-addition level and coolant properties on the cooling effectiveness of transpiration-cooled surfaces in the absence of shock interaction. In the second phase of the program, we examined the effects of oblique shock impingement on the flowfield and surface characteristics of a transpiration-cooled surface. The studies were conducted for a range of shock strengths with nitrogen and helium coolants to examine how the distribution of heat transfer and pressure and the characteristics of the flowfield in the interaction region varied with shock strength and the level of mass addition from the transpiration-cooled section of the model. The effects of the distribution of the blowing rate along the interaction regions were also examined for a range of blowing rates through the
Modeling and Evaluating the Environmental Degradation of UHTCs under Hypersonic Flow (Preprint)
2014-02-01
34Oxidation behavior of zirconium diboride silicon carbide produced by the spark plasma sintering method," J Amer. Ceram. Soc., 92 [9] 2046-2052...vehicles. It uses a plasma generated gas mixture which contains a high fraction of ionized atoms of oxygen and nitrogen. Recombination of the charged...species is catalyzed by the sample surface to release heat. The plasma is accelerated to high velocities (supersonic but not hypersonic) using an
A system to measure flow moisture content in hypersonic wind tunnels
NASA Technical Reports Server (NTRS)
West, James W.
1992-01-01
The technique and equipment is described which is used for obtaining data on the moisture content in two NASA Langley Hypersonic Wind Tunnels. A detailed description of the sampling system and its operation is presented along with the moisture analyzer used. The procedure used for converting dew point to parts of water per million by volume (ppmv) is included with graphs that show tunnel moisture content at various pressures.
A Numerical Study of Novel Drag Reduction Techniques for Blunt Bodies in Hypersonic Flows
2011-01-01
Injection System and Operation. a) Sketch of Injector Nozzle and Reservoir. b-d) Distinct Modes of Injection. 9 Experimental investigations of...Paull, A., “ Numerical Investigation of a Spiked Blunt Nose Cone at Hypersonic Speeds,” Journal of Spacecraft and Rockets , Vol. 45, No. 3, 2008, pp. 449...both experimental and numerical ) have also shown that forward facing mass injection has the ability to significantly reduce drag on blunt bodies by
NASA Technical Reports Server (NTRS)
Stewart, John E.; Smith, Robert E.; Ashby, George C., Jr.
1992-01-01
A class of vehicles for a mission to Mars are analyzed for aerodynamic characteristics using advanced Computational Fluid Dynamics (CFD). The general configuration is a modified cone-conical-frustum geometry where the nose radius has a large influence on the flowfield. Inviscid-compressible flow using the Euler equations and viscous-compressible flow using the thin-layer Navier-Stokes equations is applied to the configuration. The surface modeling, grid generation and application of state-of-the-art CFD software are described. The effects of nose radius, angle of attack, and hypersonic velocity on the flight characteristics of the vehicle are discussed. The numerical simulations demonstrate the merits of the inviscid and viscous software. Results are compared with wind tunnel experiments.
NASA Astrophysics Data System (ADS)
Cristofolini, Andrea; Borghi, Carlo A.; Neretti, Gabriele; Schettino, Antonio; Trifoni, Eduardo; Battista, Francesco; Passaro, Andrea; Baccarella, Damiano
2012-11-01
This paper deals with the experimental investigation on the MHD (magneto-hydro-dynamic or magneto-fluid-dynamic) interaction around a test body immersed into a hypersonic unseeded air flow. The experiments have been carried out in the CIRA plasma wind tunnel SCIROCCO. Two test conditions have been utilized for the experiments with a total pressure of 2.5 and 2.3 bar respectively, a total specific enthalpy of 16 and 12.1 MJ/kg respectively. The air flow was accelerated in the nozzle up to Mach 10. The magnetic induction field is generated by an electromagnet enclosed in the test body and reaches a 0.8 T maximum value in the interaction region.
NASA Technical Reports Server (NTRS)
Hefner, J. N.; Cary, A. M., Jr.; Bushnell, D. M.
1974-01-01
Results of an experimental and numerical investigation of tangential swept slot injection (sweep angles of 22.5 and 45 deg) into a thick turbulent boundary layer at Mach 6 are presented. Film cooling effectiveness, skin friction, and flow structure downstream of the swept slot injection are investigated. The data are compared to that for unswept slots, and it is found that cooling effectiveness and skin-friction reductions are not significantly affected by sweeping the slot. Predictions of cooling effectiveness and skin friction obtained by a numerical finite-difference technique agree reasonably well with experimental surface variables. As in previous supersonic two-dimensional slot research, reduced mixing was found downstream of the slot lip in the present three-dimensional case.
System-size independence of directed flow measured at the BNL relativistic heavy-ion collider.
Abelev, B I; Aggarwal, M M; Ahammed, Z; Anderson, B D; Arkhipkin, D; Averichev, G S; Bai, Y; Balewski, J; Barannikova, O; Barnby, L S; Baudot, J; Baumgart, S; Beavis, D R; Bellwied, R; Benedosso, F; Betts, R R; Bhardwaj, S; Bhasin, A; Bhati, A K; Bichsel, H; Bielcik, J; Bielcikova, J; Biritz, B; Bland, L C; Bombara, M; Bonner, B E; Botje, M; Bouchet, J; Braidot, E; Brandin, A V; Bueltmann, S; Burton, T P; Bystersky, M; Cai, X Z; Caines, H; Calderón de la Barca Sánchez, M; Callner, J; Catu, O; Cebra, D; Cendejas, R; Cervantes, M C; Chajecki, Z; Chaloupka, P; Chattopadhyay, S; Chen, H F; Chen, J H; Chen, J Y; Cheng, J; Cherney, M; Chikanian, A; Choi, K E; Christie, W; Chung, S U; Clarke, R F; Codrington, M J M; Coffin, J P; Cormier, T M; Cosentino, M R; Cramer, J G; Crawford, H J; Das, D; Dash, S; Daugherity, M; de Moura, M M; Dedovich, T G; Dephillips, M; Derevschikov, A A; Derradi de Souza, R; Didenko, L; Dietel, T; Djawotho, P; Dogra, S M; Dong, X; Drachenberg, J L; Draper, J E; Du, F; Dunlop, J C; Dutta Mazumdar, M R; Edwards, W R; Efimov, L G; Elhalhuli, E; Elnimr, M; Emelianov, V; Engelage, J; Eppley, G; Erazmus, B; Estienne, M; Eun, L; Fachini, P; Fatemi, R; Fedorisin, J; Feng, A; Filip, P; Finch, E; Fine, V; Fisyak, Y; Gagliardi, C A; Gaillard, L; Gangadharan, D R; Ganti, M S; Garcia-Solis, E; Ghazikhanian, V; Ghosh, P; Gorbunov, Y N; Gordon, A; Grebenyuk, O; Grosnick, D; Grube, B; Guertin, S M; Guimaraes, K S F F; Gupta, A; Gupta, N; Guryn, W; Haag, B; Hallman, T J; Hamed, A; Harris, J W; He, W; Heinz, M; Heppelmann, S; Hippolyte, B; Hirsch, A; Hoffman, A M; Hoffmann, G W; Hofman, D J; Hollis, R S; Huang, H Z; Hughes, E W; Humanic, T J; Igo, G; Iordanova, A; Jacobs, P; Jacobs, W W; Jakl, P; Jin, F; Jones, P G; Judd, E G; Kabana, S; Kajimoto, K; Kang, K; Kapitan, J; Kaplan, M; Keane, D; Kechechyan, A; Kettler, D; Khodyrev, V Yu; Kiryluk, J; Kisiel, A; Klein, S R; Knospe, A G; Kocoloski, A; Koetke, D D; Kollegger, T; Kopytine, M; Kotchenda, L; Kouchpil, V; Kravtsov, P; Kravtsov, V I; Krueger, K; Kuhn, C; Kumar, A; Kumar, L; Kurnadi, P; Lamont, M A C; Landgraf, J M; Lange, S; Lapointe, S; Laue, F; Lauret, J; Lebedev, A; Lednicky, R; Lee, C-H; Levine, M J; Li, C; Li, Y; Lin, G; Lin, X; Lindenbaum, S J; Lisa, M A; Liu, F; Liu, J; Liu, L; Ljubicic, T; Llope, W J; Longacre, R S; Love, W A; Lu, Y; Ludlam, T; Lynn, D; Ma, G L; Ma, J G; Ma, Y G; Mahapatra, D P; Majka, R; Mangotra, L K; Manweiler, R; Margetis, S; Markert, C; Matis, H S; Matulenko, Yu A; McShane, T S; Meschanin, A; Millane, J; Miller, M L; Minaev, N G; Mioduszewski, S; Mischke, A; Mitchell, J; Mohanty, B; Morozov, D A; Munhoz, M G; Nandi, B K; Nattrass, C; Nayak, T K; Nelson, J M; Nepali, C; Netrakanti, P K; Ng, M J; Nogach, L V; Nurushev, S B; Odyniec, G; Ogawa, A; Okada, H; Okorokov, V; Olson, D; Pachr, M; Pal, S K; Panebratsev, Y; Pawlak, T; Peitzmann, T; Perevoztchikov, V; Perkins, C; Peryt, W; Phatak, S C; Planinic, M; Pluta, J; Poljak, N; Porile, N; Poskanzer, A M; Potekhin, M; Potukuchi, B V K S; Prindle, D; Pruneau, C; Pruthi, N K; Putschke, J; Qattan, I A; Raniwala, R; Raniwala, S; Ray, R L; Ridiger, A; Ritter, H G; Roberts, J B; Rogachevskiy, O V; Romero, J L; Rose, A; Roy, C; Ruan, L; Russcher, M J; Rykov, V; Sahoo, R; Sakrejda, I; Sakuma, T; Salur, S; Sandweiss, J; Sarsour, M; Schambach, J; Scharenberg, R P; Schmitz, N; Seger, J; Selyuzhenkov, I; Seyboth, P; Shabetai, A; Shahaliev, E; Shao, M; Sharma, M; Shi, S S; Shi, X-H; Sichtermann, E P; Simon, F; Singaraju, R N; Skoby, M J; Smirnov, N; Snellings, R; Sorensen, P; Sowinski, J; Spinka, H M; Srivastava, B; Stadnik, A; Stanislaus, T D S; Staszak, D; Stock, R; Strikhanov, M; Stringfellow, B; Suaide, A A P; Suarez, M C; Subba, N L; Sumbera, M; Sun, X M; Sun, Y; Sun, Z; Surrow, B; Symons, T J M; Szanto de Toledo, A; Takahashi, J; Tang, A H; Tang, Z; Tarnowsky, T; Thein, D; Thomas, J H; Tian, J; Timmins, A R; Timoshenko, S; Tokarev, M; Trainor, T A; Tram, V N; Trattner, A L; Trentalange, S; Tribble, R E; Tsai, O D; Ulery, J; Ullrich, T; Underwood, D G; Van Buren, G; van der Kolk, N; van Leeuwen, M; Vander Molen, A M; Varma, R; Vasconcelos, G M S; Vasilevski, I M; Vasiliev, A N; Videbaek, F; Vigdor, S E; Viyogi, Y P; Vokal, S; Voloshin, S A; Wada, M; Waggoner, W T; Wang, F; Wang, G; Wang, J S; Wang, Q; Wang, X; Wang, X L; Wang, Y; Webb, J C; Westfall, G D; Whitten, C; Wieman, H; Wissink, S W; Witt, R; Wu, J; Wu, Y; Xu, N; Xu, Q H; Xu, Y; Xu, Z; Yang, Y Y; Yepes, P; Yoo, I-K; Yue, Q; Zawisza, M; Zbroszczyk, H; Zhan, W; Zhang, H; Zhang, S; Zhang, W M; Zhang, Y; Zhang, Z P; Zhao, Y; Zhong, C; Zhou, J; Zoulkarneev, R; Zoulkarneeva, Y; Zuo, J X
2008-12-19
We measure directed flow (v_{1}) for charged particles in Au+Au and Cu+Cu collisions at sqrt[s_{NN}]=200 and 62.4 GeV, as a function of pseudorapidity (eta), transverse momentum (p_{t}), and collision centrality, based on data from the STAR experiment. We find that the directed flow depends on the incident energy but, contrary to all available model implementations, not on the size of the colliding system at a given centrality. We extend the validity of the limiting fragmentation concept to v_{1} in different collision systems, and investigate possible explanations for the observed sign change in v_{1}(p_{t}).
NASA Astrophysics Data System (ADS)
Leite, Paulo H. M.; Santos, Wilson F. N.
2014-12-01
A computational analysis of a hypersonic flow over a combined gap/step configuration at zero degree angle of attack, in chemical equilibrium and thermal non-equilibrium is presented in this work. Effects on pressure and heating loads due to changes on the freestream Mach number and on the step frontal-face height have been investigated by employing the Direct Simulation Monte Carlo (DSMC) method. The work focuses the attention of designers of hypersonic configurations on the fundamental parameter of surface discontinuity, which can have an important impact on even initial design. The analysis showed that heating and pressure loads increased with increasing not only the step height but also with the freestream Mach number. In addition, peak values for both loads took place at the vicinity of the step convex corner, a similar behavior observed for a forward-facing step configuration. It was also found that these loads for the gap/step configuration are slightly smaller than those for a forward-facing step.
Vallon, Raphäel; Soutadé, Jacques; Vérant, Jean-Luc; Meyers, Jason; Paris, Sébastien; Mohamed, Ajmal
2010-01-01
Since the beginning of the Mars planet exploration, the characterization of carbon dioxide hypersonic flows to simulate a spaceship's Mars atmosphere entry conditions has been an important issue. We have developed a Tunable Diode Laser Absorption Spectrometer with a new room-temperature operating antimony-based distributed feedback laser (DFB) diode laser to characterize the velocity, the temperature and the density of such flows. This instrument has been tested during two measurement campaigns in a free piston tunnel cold hypersonic facility and in a high enthalpy arc jet wind tunnel. These tests also demonstrate the feasibility of mid-infrared fiber optics coupling of the spectrometer to a wind tunnel for integrated or local flow characterization with an optical probe placed in the flow.
Vallon, Raphäel; Soutadé, Jacques; Vérant, Jean-Luc; Meyers, Jason; Paris, Sébastien; Mohamed, Ajmal
2010-01-01
Since the beginning of the Mars planet exploration, the characterization of carbon dioxide hypersonic flows to simulate a spaceship’s Mars atmosphere entry conditions has been an important issue. We have developed a Tunable Diode Laser Absorption Spectrometer with a new room-temperature operating antimony-based distributed feedback laser (DFB) diode laser to characterize the velocity, the temperature and the density of such flows. This instrument has been tested during two measurement campaigns in a free piston tunnel cold hypersonic facility and in a high enthalpy arc jet wind tunnel. These tests also demonstrate the feasibility of mid-infrared fiber optics coupling of the spectrometer to a wind tunnel for integrated or local flow characterization with an optical probe placed in the flow. PMID:22219703
Effect of Body Perturbations on Hypersonic Flow Over Slender Power Law Bodies
NASA Technical Reports Server (NTRS)
Mirels, Harold; Thornton, Philip R.
1959-01-01
Hypersonic-slender-body theory, in the limit as the free-stream Mach number becomes infinite, is used to find the effect of slightly perturbing the surface of slender two-dimensional and axisymmetric power law bodies, The body perturbations are assumed to have a power law variation (with streamwise distance downstream of the nose of the body). Numerical results are presented for (1) the effect of boundary-layer development on two dimensional and axisymmetric bodies, (2) the effect of very small angles of attack (on tow[dimensional bodies), and (3) the effect of blunting the nose of very slender wedges and cones.
Flat plate at incidence as a waverider in rarefied hypersonic flow
NASA Technical Reports Server (NTRS)
Cheng, H. K.; Wong, E. Y.; Hoover, L. N.; Dogra, V. K.
1990-01-01
The physical validity of continuum models and their ability to predict the critical aerothermodynamic properties of a waverider at high altitudes are examined using a flat plate at angle of attack as a generic hypersonic lifting vehicle. For a shock layer far from local translational equilibrium, a theoretical study based on Grad's thirteen-moment equations shows that the Navier-Stokes based solutions can correctly predict the drag, lift, and surface heat transfer rate, with the prediction error comparable to that of the standard shock-layer theory. The conclusion is supported by a comparison with direct simulation Monte Carlo calculations.
Measured and calculated mean flow properties of a two-dimensional, hypersonic, turbulent wake
NASA Technical Reports Server (NTRS)
Wagner, R. D.
1972-01-01
The hypersonic turbulent wake produced by a wedge was studied experimentally and its properties were compared with predictions obtained from a numerical computation procedure. In the computation procedure several models for the eddy viscosity formulation of the turbulent transport were examined. Conventional-defect models and a modified mixing-length model were found to yield good predictions of the experimental data. The classical mixing-length model gave unrealistic results. The experimental data displayed similarity when velocity and temperature defects were scaled by the maximum defects and the transverse coordinate was scaled by the velocity-defect half-width.
The computation of hypersonic ionized flows in chemical and thermal nonequlibrium
NASA Technical Reports Server (NTRS)
Maccormack, Robert W.; Candler, Graham V.
1988-01-01
A numerical method to compute a two-dimensional hypersonic flowfield that is ionized and in thermochemical nonequilibrium has been developed. Such a flowfield is described by coupled time-dependent partial differential equations for the conservation of species mass, mass-average momentum, vibrational energy of each diatomic species, electron energy, and total mass-averaged energy. The steady-state solution to these fully coupled equations is obtained using an implicit Gauss-Seidel line relaxation technique. The computed electron densities in the flowfield compare well with experimental results.
PLIF Temperature and Velocity Distributions in Laminar Hypersonic Flat-plate Flow
NASA Technical Reports Server (NTRS)
OByrne, S.; Danehy, P. M.; Houwing, A. F. P.
2003-01-01
Rotational temperature and velocity distributions have been measured across a hypersonic laminar flat-plate boundary layer, using planar laser-induced fluorescence. The measurements are compared to a finite-volume computation and a first-order boundary layer computation, assuming local similarity. Both computations produced similar temperature distributions and nearly identical velocity distributions. The disagreement between calculations is ascribed to the similarity solution not accounting for leading-edge displacement effects. The velocity measurements agreed to within the measurement uncertainty of 2 % with both calculated distributions. The peak measured temperature was 200 K lower than the computed values. This discrepancy is tentatively ascribed to vibrational relaxation in the boundary layer.
NASA Technical Reports Server (NTRS)
Miller, C. G.; Micol, J. R.; Gnoffo, P. A.; Wilder, S. E.
1983-01-01
Laminar heat transfer rates were measured on spherically blunted, 13 deg/7 deg on axis and bent biconics (fore cone bent 7 deg upward relative to aft cone) at hypersonic hypervelocity flow conditions in the Langley Expansion Tube. Freestream velocities from 4.5 to 6.9 km/sec and Mach numbers from 6 to 9 were generated using helium, nitrogen, air, and carbon dioxide test gases, resulting in normal shock density ratios from 4 to 19. Angle of attack, referenced to the axis of the aft cone, was varied from 0 to 20 deg in 4 deg increments. The effect of nose bend, angle of attack, and real gas phenomena on heating distributions are presented along with comparisons of measurement to prediction from a code which solves the three dimensional parabolized Navier-Stokes equations.
NASA Technical Reports Server (NTRS)
Hollis, Brian R.; Griffith, Wayland C.; Yanta, William J.
1991-01-01
A fine-wire thermocouple probe was used to determine freestream stagnation temperatures in hypersonic flows. Data were gathered in a N2 blowdown wind tunnel with runtimes of 1-5 s. Tests were made at supply pressures between 30 and 1400 atm and supply temperatures between 700 and 1900 K, with Mach numbers of 14 to 16. An iterative procedure requiring thermocouple data, pilot pressure measurements, and supply conditions was used to determine test cell stagnation temperatures. Probe conduction and radiation losses, as well as real gas behavior of N2, were accounted for during analysis. Temperature measurement error was found to be 5 to 10 percent. A correlation was drawn between thermocouple diameter Reynolds number and temperature recovery ratio. Transient probe behavior was studied and was found to be adequate in temperature gradients up to 1000 K/s.
NASA Technical Reports Server (NTRS)
Grose, W. L.
1971-01-01
An approximate inverse solution is presented for the nonequilibrium flow in the inviscid shock layer about a vehicle in hypersonic flight. The method is based upon a thin-shock-layer approximation and has the advantage of being applicable to both subsonic and supersonic regions of the shock layer. The relative simplicity of the method makes it ideally suited for programming on a digital computer with a significant reduction in storage capacity and computing time required by other more exact methods. Comparison of nonequilibrium solutions for an air mixture obtained by the present method is made with solutions obtained by two other methods. Additional cases are presented for entry of spherical nose cones into representative Venusian and Martian atmospheres. A digital computer program written in FORTRAN language is presented that permits an arbitrary gas mixture to be employed in the solution. The effects of vibration, dissociation, recombination, electronic excitation, and ionization are included in the program.
NASA Technical Reports Server (NTRS)
Hollis, Brian R.; Griffith, Wayland C.; Yanta, William J.
1991-01-01
A fine-wire thermocouple probe was used to determine freestream stagnation temperatures in hypersonic flows. Data were gathered in a N2 blowdown wind tunnel with runtimes of 1-5 s. Tests were made at supply pressures between 30 and 1400 atm and supply temperatures between 700 and 1900 K, with Mach numbers of 14 to 16. An iterative procedure requiring thermocouple data, pilot pressure measurements, and supply conditions was used to determine test cell stagnation temperatures. Probe conduction and radiation losses, as well as real gas behavior of N2, were accounted for during analysis. Temperature measurement error was found to be 5 to 10 percent. A correlation was drawn between thermocouple diameter Reynolds number and temperature recovery ratio. Transient probe behavior was studied and was found to be adequate in temperature gradients up to 1000 K/s.
Laufer, G; McKenzie, R L; Fletcher, D G
1990-11-20
Laser-induced fluorescence in oxygen, in combination with Raman scattering, is shown to be an accurate means by which temperature, density, and their fluctuations owing to turbulence can be measured in air flows associated with high speed wind tunnels. For temperatures above 60 K and densities above 0.01 amagat, the uncertainties in the temperature and density measurements can be <2%, if the signal uncertainties are dominated by photon statistical noise. The measurements are unaffected by collisional quenching and can be achieved with laser fluences for which nonlinear effects are insignificant. Temperature measurements using laser-induced fluorescence alone have been demonstrated at known densities in the range of low temperatures and densities which are expected in a hypersonic wind tunnel.
Gallis, Michael A; Bond, Ryan B; Torczynski, John R
2009-09-28
Recently proposed molecular-level chemistry models that predict equilibrium and nonequilibrium reaction rates using only kinetic theory and fundamental molecular properties (i.e., no macroscopic reaction-rate information) are investigated for chemical reactions occurring in upper-atmosphere hypersonic flows. The new models are in good agreement with the measured Arrhenius rates for near-equilibrium conditions and with both measured rates and other theoretical models for far-from-equilibrium conditions. Additionally, the new models are applied to representative combustion and ionization reactions and are in good agreement with available measurements and theoretical models. Thus, molecular-level chemistry modeling provides an accurate method for predicting equilibrium and nonequilibrium chemical-reaction rates in gases.
NASA Astrophysics Data System (ADS)
Hollis, Brian R.; Griffith, Wayland C.; Yanta, William J.
A fine-wire thermocouple probe was used to determine freestream stagnation temperatures in hypersonic flows. Data were gathered in a N2 blowdown wind tunnel with runtimes of 1-5 s. Tests were made at supply pressures between 30 and 1400 atm and supply temperatures between 700 and 1900 K, with Mach numbers of 14 to 16. An iterative procedure requiring thermocouple data, pilot pressure measurements, and supply conditions was used to determine test cell stagnation temperatures. Probe conduction and radiation losses, as well as real gas behavior of N2, were accounted for during analysis. Temperature measurement error was found to be 5 to 10 percent. A correlation was drawn between thermocouple diameter Reynolds number and temperature recovery ratio. Transient probe behavior was studied and was found to be adequate in temperature gradients up to 1000 K/s.
Transition at hypersonic speeds
NASA Technical Reports Server (NTRS)
Morkovin, Mark V.
1987-01-01
Certain conjectures on the physics of instabilities in high-speed flows are discussed and the state of knowledge of hypersonic transition summarized. The case is made for an unpressured systematic research program in this area consisting of controlled microscopic experiments, theory, and numerical simulations.
NASA Technical Reports Server (NTRS)
Bardina, J. E.
1994-01-01
A new computational efficient 3-D compressible Reynolds-averaged implicit Navier-Stokes method with advanced two equation turbulence models for high speed flows is presented. All convective terms are modeled using an entropy satisfying higher-order Total Variation Diminishing (TVD) scheme based on implicit upwind flux-difference split approximations and arithmetic averaging procedure of primitive variables. This method combines the best features of data management and computational efficiency of space marching procedures with the generality and stability of time dependent Navier-Stokes procedures to solve flows with mixed supersonic and subsonic zones, including streamwise separated flows. Its robust stability derives from a combination of conservative implicit upwind flux-difference splitting with Roe's property U to provide accurate shock capturing capability that non-conservative schemes do not guarantee, alternating symmetric Gauss-Seidel 'method of planes' relaxation procedure coupled with a three-dimensional two-factor diagonal-dominant approximate factorization scheme, TVD flux limiters of higher-order flux differences satisfying realizability, and well-posed characteristic-based implicit boundary-point a'pproximations consistent with the local characteristics domain of dependence. The efficiency of the method is highly increased with Newton Raphson acceleration which allows convergence in essentially one forward sweep for supersonic flows. The method is verified by comparing with experiment and other Navier-Stokes methods. Here, results of adiabatic and cooled flat plate flows, compression corner flow, and 3-D hypersonic shock-wave/turbulent boundary layer interaction flows are presented. The robust 3-D method achieves a better computational efficiency of at least one order of magnitude over the CNS Navier-Stokes code. It provides cost-effective aerodynamic predictions in agreement with experiment, and the capability of predicting complex flow structures in
NASA Astrophysics Data System (ADS)
Risius, Steffen; Beck, Walter H.; Klein, Christian; Henne, Ulrich; Wagner, Alexander
2017-09-01
Heat loads on spacecraft traveling at hypersonic speed are of major interest for their designers. Several tests using temperature-sensitive paints (TSP) have been carried out in long duration shock tunnels to determine these heat loads; generally paint layers were thin, so that certain assumptions could be invoked to enable a good estimate of the thermal parameter ρck (a material property) to be obtained—the value of this parameter is needed to determine heat loads from the TSP. Very few measurements have been carried out in impulse facilities [viz. shock tunnels such as the High Enthalpy Shock Tunnel Göttingen (HEG)], where test times are much shorter. Presented here are TSP temperature measurements and subsequently derived heat loads on a ramp model placed in a hypersonic flow in HEG (specific enthalpy h 0 = 3.3 MJ kg-1, Mach number M = 7.4, temperature T ∞ = 277 K, density ρ ∞ = 11 g m-3). A number of fluorescence intensity images were acquired, from which, with the help of calibration data, temperature field data on the model surface were determined. From these the heat load into the surface was calculated, using an assumption of a 1D, semi-infinite heat transfer model. ρck for the paint was determined using an insitu calibration with a Medtherm coaxial thermocouple mounted on the model; Medtherm ρck is known. Finally presented are sources of various measurement uncertainties, arising from: (1) estimation of ρck; (2) intensity measurement in the chosen interrogation area; (3) paint time response.
A Hot Dynamic Seal Rig for Measuring Hypersonic Engine Seal Durability and Flow Performance
NASA Technical Reports Server (NTRS)
Miller, Jeffrey H.; Steinetz, Bruce M.; Sirocky, Paul J.; Kren, Lawrence A.
1993-01-01
A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts was installed at NASA Lewis Research Center. The test fixture was designed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C (1550 F) and air pressure differentials up to 690 kPa (100 psi). Performance of the seals can be measured while sealing against flat or distorted walls. In the fixture two seals are preloaded against the sides of a 30 cm (1 ft) long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. The capabilities of this test fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling are addressed.
Evaluation of aero-optical effects in hypersonic flow using holographic interferometry
NASA Astrophysics Data System (ADS)
Azzazy, M.
1991-01-01
The aerooptical performance of a generic high endoatmospheric defense interceptor (HEDI) in a hypersonic tunnel is evaluated. The problem of a plane monochromatic wave incident on volume V of a turbulent medium is studied, and it is concluded that in order to evaluate the aerooptical effects on HEDI performance, it is imperative to measure phase fluctuations. A holographic optical system employed in the experiment is outlined, along with the test setup where the object beam travels through the wind-tunnel Schlieren window on the top of the tunnel, the splitter window on the bottom of the pylon, and the HEDI window before it is reflected back on itself through the same path to a steering mirror and on a holographic plate. The interferogram images are analyzed digitally to yield the fringe RMS and the phase correction function. It is noted that the fringe RMS influences the attenuation of the coherent beam, while the correlation length scale affects the circle of blur.
A hot dynamic seal rig for measuring hypersonic engine seal durability and flow performance
NASA Technical Reports Server (NTRS)
Miller, Jeffrey H.; Steinetz, Bruce M.; Sirocky, Paul J.; Kren, Lawrence A.
1993-01-01
A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts has been installed at NASA Lewis Research Center. The test fixture has been designed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C (1550 F) and air pressure differentials up to 690 kPa (100 psi). Performance of the seals can be measured while sealing against flat or distorted walls. In the fixture two seals are preloaded against the sides of a 30 cm (1 ft) long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. This report covers the capabilities of this test fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling.
Hot dynamic test rig for measuring hypersonic engine seal flow and durability
NASA Technical Reports Server (NTRS)
Miller, Jeffrey H.; Steinetz, Bruce M.; Sirocky, Paul J.; Kren, Lawrence A.
1994-01-01
A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts was developed. The test fixture was developed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C and air pressure differentials of to 0.7 MPa. Performance of the seals can be measured while sealing against flat or engine-simulated distorted walls. In the fixture, two seals are preloaded against the sides of a 0.3 m long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. The capabilities of this text fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling are covered.
System-size independence of directed flow at the RelativisticHeavy-Ion Collider
STAR Coll
2008-09-20
We measure directed flow (v{sub 1}) for charged particles in Au + Au and Cu + Cu collisions at {radical}s{sub NN} = 200 GeV and 62.4 GeV, as a function of pseudorapidity ({eta}), transverse momentum (p{sub t}) and collision centrality, based on data from the STAR experiment. We find that the directed flow depends on the incident energy but, contrary to all available model implementations, not on the size of the colliding system at a given centrality. We extend the validity of the limiting fragmentation concept to v{sub 1} in different collision systems, and investigate possible explanations for the observed sign change in v{sub 1}(p{sub t}).
System-size independence of directed flow measured at the BNL relativistic heavy-ion collider.
Abelev, B. I.; Aggarwal, M. M.; Ahammed, Z.; Anderson, B. D.; Arkhipkin, D.; Krueger, K.; Spinka, H. M.; Underwood, D. G.; High Energy Physics; Univ. of Illinois; Panjab Univ.; Variable Energy Cyclotron Centre; Kent State Univ.; Particle Physic Lab.; STAR Collaboration
2008-01-01
We measure directed flow (v{sub 1}) for charged particles in Au+Au and Cu+Cu collisions at {radical}s{sub NN} = 200 and 62.4 GeV, as a function of pseudorapidity ({eta}), transverse momentum (p{sub t}), and collision centrality, based on data from the STAR experiment. We find that the directed flow depends on the incident energy but, contrary to all available model implementations, not on the size of the colliding system at a given centrality. We extend the validity of the limiting fragmentation concept to v{sub 1} in different collision systems, and investigate possible explanations for the observed sign change in v{sub 1}(p{sub t}).
NASA Technical Reports Server (NTRS)
Miller, C. G., III; Wilder, S. E.
1972-01-01
Data-reduction procedures for determining free stream and post-normal shock kinetic and thermodynamic quantities are derived. These procedures are applicable to imperfect real air flows in thermochemical equilibrium for temperatures to 15 000 K and a range of pressures from 0.25 N/sq m to 1 GN/sq m. Although derived primarily to meet the immediate needs of the 6-inch expansion tube, these procedures are applicable to any supersonic or hypersonic test facility where combinations of three of the following flow parameters are measured in the test section: (1) Stagnation pressure behind normal shock; (2) freestream static pressure; (3) stagnation point heat transfer rate; (4) free stream velocity; (5) stagnation density behind normal shock; and (6) free stream density. Limitations of the nine procedures and uncertainties in calculated flow quantities corresponding to uncertainties in measured input data are discussed. A listing of the computer program is presented, along with a description of the inputs required and a sample of the data printout.
NASA Technical Reports Server (NTRS)
Marconi, F.; Yaeger, L.
1976-01-01
A numerical procedure was developed to compute the inviscid super/hypersonic flow field about complex vehicle geometries accurately and efficiently. A second-order accurate finite difference scheme is used to integrate the three-dimensional Euler equations in regions of continuous flow, while all shock waves are computed as discontinuities via the Rankine-Hugoniot jump conditions. Conformal mappings are used to develop a computational grid. The effects of blunt nose entropy layers are computed in detail. Real gas effects for equilibrium air are included using curve fits of Mollier charts. Typical calculated results for shuttle orbiter, hypersonic transport, and supersonic aircraft configurations are included to demonstrate the usefulness of this tool.
NASA Astrophysics Data System (ADS)
Guo, Guang-ming; Liu, Hong; Zhang, Bin
2017-03-01
Near space has been paid more and more attentions in recent years due to its militarily application value. Direct simulation Monte Carlo (DSMC) which is one of the most successful particle simulation methods in treating rarefied gas dynamics is employed to investigate the flow characteristics of a hypersonic backward-facing step (BFS) under active flow control using supersonic jet in near space. The numerical tool is validated by an experimental flow of dual cusped-plate model, shock wave structures from the numerical simulation are shown in quite good agreement with the experimental result. The influence of altitude and active flow control on BFS flow are then studied in detail. Three parameters, i.e. boundary layer thickness, recirculation region length, and lean angle of the primary recirculation region that is first defined to describe recirculation region shape, are used to evaluate the flow characteristics of every case computed. The numerical results indicate that the main effect of vertical jet upstream of the step is the enhancement of boundary layer thickness downstream of the jet slot, then, it shows a weak influence on recirculation region length and a negligible effect on lean angle. Conversely, the horizontal jet near the step edge can greatly change the recirculation region length by adjusting jetting angle, but it only has a weak influence both on boundary layer thickness and on lean angle for every jetting angle considered. A significant finding is that the recirculation region length is decreased severely in near space compared with experimental and numerical results presented in the open literature.
Hypersonic rarefied wake characterization
NASA Technical Reports Server (NTRS)
Brewer, E. B.
1993-01-01
Results of a numerical study using the direct simulation Monte Carlo (DSMC) method are presented for hypersonic rarefied flow over an aeroassisted space transfer vehicle (ASTV). The emphasis of the study is the characterization of the near wake region which includes the ASTV payload. The study covered the transitional flow regime from near continuum to free molecular. Calculations show that the character of the near wake is significantly affected by the presence of the payload. Flow separation occurs when an afterbody is present throughout the transitional flow regime. In contrast, when no afterbody is present, no separation is observed until the flow approaches continuum.
NASA Astrophysics Data System (ADS)
Bonelli, Francesco; Tuttafesta, Michele; Colonna, Gianpiero; Cutrone, Luigi; Pascazio, Giuseppe
2017-10-01
This paper describes the most advanced results obtained in the context of fluid dynamic simulations of high-enthalpy flows using detailed state-to-state air kinetics. Thermochemical non-equilibrium, typical of supersonic and hypersonic flows, was modeled by using both the accurate state-to-state approach and the multi-temperature model proposed by Park. The accuracy of the two thermochemical non-equilibrium models was assessed by comparing the results with experimental findings, showing better predictions provided by the state-to-state approach. To overcome the huge computational cost of the state-to-state model, a multiple-nodes GPU implementation, based on an MPI-CUDA approach, was employed and a comprehensive code performance analysis is presented. Both the pure MPI-CPU and the MPI-CUDA implementations exhibit excellent scalability performance. GPUs outperform CPUs computing especially when the state-to-state approach is employed, showing speed-ups, of the single GPU with respect to the single-core CPU, larger than 100 in both the case of one MPI process and multiple MPI process.
NASA Technical Reports Server (NTRS)
Hackett, Charles M.
1993-01-01
The interaction between a swept shock wave and a laminar boundary layer was investigated experimentally in high-enthalpy hypersonic flow. The effect of high-temperature, real gas physics on the interaction was examined by conducting tests in air and helium. Heat transfer measurements were made on the surface of a flat plate and a shock-generating fin using thin-film resistance sensors for fin incidence angles of 0, 5, and 10 deg at Mach numbers of 6.9 in air and 7.2 in helium. The experiments were conducted in the NASA HYPULSE expansion tube, an impulse-type facility capable of generating high-enthalpy, high-velocity flow with freestream levels of dissociated species that are particularly low. The measurements indicate that the swept shock wave creates high local heat transfer levels in the interaction region, with the highest heating found in the strongest interaction. The maximum measured heating rates in the interaction are order of magnitude greater than laminar flat plate boundary layer heating levels at the same location.
Nonlinear development and secondary instability of Gortler vortices in hypersonic flows
NASA Technical Reports Server (NTRS)
Fu, Yibin B.; Hall, Philip
1991-01-01
In a hypersonic boundary layer over a wall of variable curvature, the region most susceptible to Goertler vortices is the temperature adjustment layer over which the basic state temperature decreases monotonically to its free stream value. Except for a special wall curvature distribution, the evolution of Goertler vortices trapped in the temperature adjustment layer will in general be strongly affected by the boundary layer growth through the O(M sup 3/2) curvature of the basic state, where M is the free stream Mach number. Only when the local wavenumber becomes as large as of order M sup 3/8, do nonparallel effects become negligible in the determination of stability properties. In the latter case, Goertler vortices will be trapped in a thin layer of O(epsilon sup 1/2) thickness which is embedded in the temperature adjustment layer; here epsilon is the inverse of the local wavenumber. A weakly nonlinear theory is presented in which the initial nonlinear development of Goertler vortices in the neighborhood of the neutral position is studied and two coupled evolution equations are derived. From these, it can be determined whether the vortices are decaying or growing depending on the sign of a constant which is related to wall curvature and the basic state temperature.
NASA Technical Reports Server (NTRS)
Danehy, Paul M.; Alderfer, David W.; Inman, Jennifer A.; Berger, Karen T.; Buck, Gregory M.; Schwartz, Richard J.
2008-01-01
Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Only a few of the models survived repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2- inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various model configurations and NO seeding methods were used, including a new streamwise visualization method based on PLIF. Virtual Diagnostics Interface (ViDI) technology, developed at NASA Langley Research Center, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images.
NASA Astrophysics Data System (ADS)
Birrer, Marcel; Stemmer, Christian; Adams, Nikolaus N.
2011-05-01
Investigations of hypersonic boundary-layer flows around a cubical obstacle with a height in the order of half the boundary layer thickness were carried out in this work. Special interest was laid on the influence of chemical non-equilibrium effects on the wake flow of the obstacle. Direct numerical simulations were conducted using three different gas models, a caloric perfect, an equilibrium and a chemical non-equilibrium gas model. The geometry was chosen as a wedge with a six degree half angle, according to the aborted NASA HyBoLT free flight experiment. At 0.5 m downstream of the leading edge, a surface trip was positioned. The free-stream flow was set to Mach 8.5 with air conditions taken from the 1976 standard atmosphere at an altitude of 42 km according to the predicted flight path. The simulations were done in three steps for all models. First, two-dimensional calculations of the whole configuration including the leading edge and the obstacle were conducted. These provide constant span-wise profiles for detailed, steady three-dimensional simulations around the close vicinity of the obstacle. A free-stream Mach number of about 6.3 occurs behind the shock. A cross-section in the wake of the object then delivers the steady inflow for detailed unsteady simulations of the wake. Perturbations at unstable frequencies, obtained from a bi-global secondary stability analysis, were added to these profiles. The solutions are time-Fourier transformed to investigate the unsteady downstream development of the different modes due to the interaction with the base-flow containing two counter-rotating vortices. Results will be presented that show the influence of the presence of chemical non-equilibrium on the instability in the wake of the object leading to a laminar or a turbulent wake.
Rossmann, Tobias; Mungal, M Godfrey; Hanson, Ronald K
2003-11-20
The scalar-field imaging of a hypersonic mixing flow is performed in a mixing facility that is shock tunnel driven. The instantaneous mixture-fraction field of a hypersonic two-dimensional mixing layer (M1 = 5.1, M2 = 0.3) is determined with a temperature-insensitive planar laser-induced fluorescence technique with nitric oxide (NO) as the tracer species. Single-shot images are obtained with the broadband excitation of a reduced temperature-sensitivity transition in the A2 sigma+ <-- X2 II(1/2) (0, 0) band of NO near 226 nm. The instantaneous mixture-fraction field at a convective Mach number of 2.64 is shown to be nearly identical to a typical diffusive process, supporting the notion of gradient-transport mixing models for highly compressible mixing layers.
Production of high-beta magnetised plasmas by colliding supersonic flows from inverse wire arrays
NASA Astrophysics Data System (ADS)
Hare, Jack; Suttle, Lee; Lebedev, Sergey; Bennett, Matthew; Burdiak, Guy; Clayson, Thomas; Suzuki-Vidal, Francisco; Swadling, George; Patankar, Siddharth; Robinson, Timothy; Stuart, Nicholas; Smith, Roland; Yang, Qingguo; Wu, Jian; Rozmus, Wojciech
2015-11-01
HEDP often exhibit a high plasma β and an electron Hall parameter greater than one. This results in a complex interplay between the transport of heat and magnetic fields, relevant to the Magnetised Liner Inertial Fusion (MagLIF) concept. We can produce such plasmas by colliding two supersonic quasi-planar flows from two adjacent inverse wire arrays made from carbon. The standing shock formed by the collision heats and compresses the plasma. The plasma flows advect magnetic fields which are perpendicular to the flow direction. Depending on the experimental set up, this can result in either flux compression or reconnection in the interaction region. The experiments are conducted on MAGPIE (1.4 MA, 250 ns current pulse). The formed shock is stable over long timescales (~100 ns), and the electron temperature (100 eV) is close to the ion temperature (500 eV), measured by spatially resolved Thomson scattering. Magnetic fields above 5 T is observed using a Faraday rotation diagnostic, and an electron density of around 5x1017 cm-3 is measured by interferometry.
NASA Technical Reports Server (NTRS)
Cheatwood, F. Mcneil; Dejarnette, Fred R.
1991-01-01
An approximate axisymmetric method was developed which can reliably calculate fully viscous hypersonic flows over blunt nosed bodies. By substituting Maslen's second order pressure expression for the normal momentum equation, a simplified form of the viscous shock layer (VSL) equations is obtained. This approach can solve both the subsonic and supersonic regions of the shock layer without a starting solution for the shock shape. The approach is applicable to perfect gas, equilibrium, and nonequilibrium flowfields. Since the method is fully viscous, the problems associated with a boundary layer solution with an inviscid layer solution are avoided. This procedure is significantly faster than the parabolized Navier-Stokes (PNS) or VSL solvers and would be useful in a preliminary design environment. Problems associated with a previously developed approximate VSL technique are addressed before extending the method to nonequilibrium calculations. Perfect gas (laminar and turbulent), equilibrium, and nonequilibrium solutions were generated for airflows over several analytic body shapes. Surface heat transfer, skin friction, and pressure predictions are comparable to VSL results. In addition, computed heating rates are in good agreement with experimental data. The present technique generates its own shock shape as part of its solution, and therefore could be used to provide more accurate initial shock shapes for higher order procedures which require starting solutions.
Song, Yang; Zhang, Bin; He, Anzhi
2006-11-01
A novel algebraic iterative algorithm based on deflection tomography is presented. This algorithm is derived from the essentials of deflection tomography with a linear expansion of the local basis functions. By use of this algorithm the tomographic problem is finally reduced to the solution of a set of linear equations. The algorithm is demonstrated by mapping a three-peak Gaussian simulative temperature field. Compared with reconstruction results obtained by other traditional deflection algorithms, its reconstruction results provide a significant improvement in reconstruction accuracy, especially in cases with noisy data added. In the density diagnosis of a hypersonic wind tunnel, this algorithm is adopted to reconstruct density distributions of an axial symmetry flow field. One cross section of the reconstruction results is selected to be compared with the inverse Abel transform algorithm. Results show that the novel algorithm can achieve an accuracy equivalent to the inverse Abel transform algorithm. However, the novel algorithm is more versatile because it is applicable to arbitrary kinds of distribution.
NASA Astrophysics Data System (ADS)
Sun, Xi-wan; Guo, Zhen-yun; Huang, Wei; Li, Shi-bin; Yan, Li
2016-09-01
The drag and heat reduction problem of hypersonic reentry vehicles has always attracted the attention worldwide, and many novel schemes have been proposed recently. In the current study, the research progress of the combinational configuration of the forward-facing cavity and the counterflowing jet has been reviewed, and the conventional cavity configuration has been substituted by an approximate maximum thrust nozzle contour for better heat and surface pressure reduction efficiency. The Reynolds-average of Navier-Stokes (RANS) equations coupled with the SST k-ω turbulence model have been employed to calculate its surrounding flow fields. A validation metric and the grid convergence index (GCI) have been employed to conduct the turbulence model assessment and the grid independence analysis respectively. The axisymmetric assumption has been verified by three-dimensional computational results as well. The obtained results show that the SST k-ω model is more suitable for the novel drag and heat flux reduction scheme proposed in this article, and the axisymmetric assumption is approximately reasonable. After investigating the influence of jet pressure ratio, the novel combinational configuration has been verified to be more effective in heat and surface pressure reduction, and this is because the approximate maximum thrust nozzle contour contributes to better expansion and avoids total pressure loss of the jet.
Fluorescence Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models
NASA Technical Reports Server (NTRS)
Alderfer, D. W.; Danehy, P. M.; Inma, J. A.; Berger, K. T.; Buck, G. M.; Schwartz, R J.
2007-01-01
Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Most of the models did not survive repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2-inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various configurations were studied including different sting placements relative to the models, different model orientations and attachment angles, and different NO seeding methods. The angle of attack of the models was also varied and the location of the laser sheet was scanned to provide three-dimensional flowfield information. Virtual Diagnostics Interface technology, developed at NASA Langley, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images. Lessons learned and recommendations for future experiments are discussed.
Non-thermal radio emission from colliding flows in classical nova V1723 Aql
NASA Astrophysics Data System (ADS)
Weston, Jennifer H. S.; Sokoloski, J. L.; Metzger, Brian D.; Zheng, Yong; Chomiuk, Laura; Krauss, Miriam I.; Linford, Justin D.; Nelson, Thomas; Mioduszewski, Amy J.; Rupen, Michael P.; Finzell, Tom; Mukai, Koji
2016-03-01
The importance of shocks in nova explosions has been highlighted by Fermi's discovery of γ-ray-producing novae. Over three years of multiband Very Large Array radio observations of the 2010 nova V1723 Aql show that shocks between fast and slow flows within the ejecta led to the acceleration of particles and the production of synchrotron radiation. Soon after the start of the eruption, shocks in the ejecta produced an unexpected radio flare, resulting in a multipeaked radio light curve. The emission eventually became consistent with an expanding thermal remnant with mass 2 × 10-4 M⊙ and temperature 104 K. However, during the first two months, the ≳106 K brightness temperature at low frequencies was too high to be due to thermal emission from the small amount of X-ray-producing shock-heated gas. Radio imaging showed structures with velocities of 400 km s-1 (d/6 kpc) in the plane of the sky, perpendicular to a more elongated 1500 km s-1 (d/6 kpc) flow. The morpho-kinematic structure of the ejecta from V1723 Aql appears similar to nova V959 Mon, where collisions between a slow torus and a faster flow collimated the fast flow and gave rise to γ-ray-producing shocks. Optical spectroscopy and X-ray observations of V1723 Aql during the radio flare are consistent with this picture. Our observations support the idea that shocks in novae occur when a fast flow collides with a slow collimating torus. Such shocks could be responsible for hard X-ray emission, γ-ray production, and double-peaked radio light curves from some classical novae.
NASA Technical Reports Server (NTRS)
Anderson, E. C.; Moss, J. N.
1975-01-01
The viscous shock layer equations applicable to hypersonic laminar, transitional, and turbulent flows of a perfect gas over two-dimensional plane or axially symmetric blunt bodies are presented. The equations are solved by means of an implicit finite difference scheme, and the results are compared with a turbulent boundary layer analysis. The agreement between the two solution procedures is satisfactory for the region of flow where streamline swallowing effects are negligible. For the downstream regions, where streamline swallowing effects are present, the expected differences in the two solution procedures are evident.
On the Goertler instability in hypersonic flows: Sutherland law fluids and real gas effects
NASA Technical Reports Server (NTRS)
Fu, Yibin B.; Hall, Philip; Blackaby, Nicholas D.
1990-01-01
The Goertler vortex instability mechanism in a hypersonic boundary layer on a curved wall is investigated. The precise roles of the effects of boundary layer growth, wall cooling, and gas dissociation is clarified in the determination of stability properties. It is first assumed that the fluid is an ideal gas with viscosity given by Sutherland's law. It is shown that when the free stream Mach number M is large, the boundary layer divides into two sublayers: a wall layer of O(M sup 3/2) thickness over which the basic state temperature is O(M squared) and a temperature adjustment layer of O(1) thickness over which the basic state temperature decreases monotonically to its free stream value. Goertler vortices which have wavelengths comparable with the boundary layer thickness are referred to as wall modes. It is shown that their downstream evolution is governed by a set of parabolic partial differential equations and that they have the usual features of Goertler vortices in incompressible boundary layers. As the local wavenumber increases, the neutral Goertler number decreases and the center of vortex activity moves towards the temperature adjustment layer. Goertler vortices with wavenumbers of order one or larger must necessarily be trapped in the temperature adjustment layer and it is this mode which is most dangerous. For this mode, it was found that the leading order term in the Goertler number expansion is independent of the wavenumber and is due to the curvature of the basic state. This term is also the asymptotic limit of the neutral Goertler numbers of the wall mode. To determine the higher order corrections terms in the Goertler number expansion, two wall curvature cases are distinguished. Real gas effects were investigated by assuming that the fluid is an ideal dissociating gas. It was found that both gas dissociation and wall cooling are destabilizing for the mode trapped in the temperature adjustment layer, but for the wall mode trapped near the wall the
Experimental And Numerical Study Of CMC Leading Edges In Hypersonic Flows
NASA Astrophysics Data System (ADS)
Kuhn, Markus; Esser, Burkard; Gulhan, Ali; Dalenbring, Mats; Cavagna, Luca
2011-05-01
Future transportation concepts aim at high supersonic or hypersonic speeds, where the formerly sharp boundaries between aeronautic and aerospace applications become blurred. One of the major issues involved to high speed flight are extremely high aerothermal loads, which especially appear at the leading edges of the plane’s wings and at sharp edged air intake components of the propulsion system. As classical materials like metals or simple ceramics would thermally and structurally fail here, new materials have to be applied. In this context, lightweight ceramic matrix composites (CMC) seem to be prospective candidates as they are high-temperature resistant and offer low thermal expansion along with high specific strength at elevated temperature levels. A generic leading edge model with a ceramic wing assembly with a sweep back angle of 53° was designed, which allowed for easy leading edge sample integration of different CMC materials. The samples consisted of the materials C/C-SiC (non-oxide), OXIPOL and WHIPOX (both oxide) with a nose radius of 2 mm. In addition, a sharp edged C/C-SiC sample was prepared to investigate the nose radius influence. Overall, 13 thermocouples were installed inside the entire model to measure the temperature evolution at specific locations, whereby 5 thermocouples were placed inside the leading edge sample itself. In addition, non-intrusive techniques were applied for surface temperature measurements: An infrared camera was used to measure the surface temperature distribution and at specific spots, the surface temperature was also measured by pyrometers. Following, the model was investigated in DLR’s arc-heated facility L3K at a total enthalpy of 8.5 MJ/kg, Mach number of 7.8, different angles of attack and varying wing inclination angles. These experiments provide a sound basis for the simulation of aerothermally loaded CMC leading edge structures. Such fluid-structure coupled approaches have been performed by FOI, basing on a
Shear modulation of intercellular contact area between two deformable cells colliding under flow
Jadhav, Sameer; Chan, Kit Yan; Konstantopoulos, Konstantinos
2007-01-01
Shear rate has been shown to critically affect the kinetics and receptor specificity of cell-cell interactions. In this study, the collision process between two modeled cells interacting in a linear shear flow is numerically investigated. The two identical biological or artificial cells are modeled as deformable capsules composed of an elastic membrane. The cell deformation and trajectories are computed using the Immersed Boundary Method for shear rates of 100–400 s−1. As the two cells collide under hydrodynamic shear, large local cell deformations develop. The effective contact area between the two cells is modulated by the shear rate, and reaches a maximum value at intermediate levels of shear. At relatively low shear rate, the contact area is an enclosed region. As the shear rate increases, dimples form on the membrane surface, and the contact region becomes annular. The non-monotonic increase of the contact area with the increase of shear rate from computational results implies that there is a maximum effective receptor-ligand binding area for cell adhesion. This finding suggests the existence of possible hydrodynamic mechanism that could be used to interpret the observed maximum leukocyte aggregation in shear flow. The critical shear rate for maximum intercellular contact area is shown to vary with cell properties such as radius and membrane elastic modulus. PMID:17467716
NASA Technical Reports Server (NTRS)
Gai, S. L.; Cain, T.; Joe, W. S.; Sandeman, R. J.; Miller, C. G.
1988-01-01
Heat transfer rate measurements have been obtained at 0, 5, 15, and 21 deg angles-of-attack for a straight biconic scale model of an aeroassisted orbital vehicle proposed for planetary probe missions. Heat-transfer distributions were measured using palladium thin-film resistance gauges deposited on a glass-ceramic substrate. The windward heat transfer correlations were based on equilibrium flow in the shock layer of the model, although the flow may depart from equilibrium in the flow-field.
NASA Technical Reports Server (NTRS)
Gai, S. L.; Cain, T.; Joe, W. S.; Sandeman, R. J.; Miller, C. G.
1988-01-01
Heat transfer rate measurements have been obtained at 0, 5, 15, and 21 deg angles-of-attack for a straight biconic scale model of an aeroassisted orbital vehicle proposed for planetary probe missions. Heat-transfer distributions were measured using palladium thin-film resistance gauges deposited on a glass-ceramic substrate. The windward heat transfer correlations were based on equilibrium flow in the shock layer of the model, although the flow may depart from equilibrium in the flow-field.
Laser-spectroscopic measurement techniques for hypersonic, turbulent wind tunnel flows
NASA Technical Reports Server (NTRS)
Mckenzie, Robert L.; Fletcher, Douglas G.
1992-01-01
A review is given of the nature, present status, and capabilities of two laser spectroscopic methods for the simultaneous measurement of temperature, density, and their fluctuations owing to turbulence in high speed wind tunnel flows. One method is based on the two frequency excitation of nitric oxide seeded into a nitrogen flow, using tunable dye lasers. The second, more recent method relies on the excitation of oxygen in air flows using a tunable, ArF excimer laser. Signal are obtained from both the laser induced fluorescence and from Raman scattering of the same laser pulse. Measurements are demonstrated in the turbulent boundary layer of a Mach-2 channel flow.
NASA Astrophysics Data System (ADS)
Kim, Young-Min; Lee, Chang-Hwan; Teaney, Derek; Zahed, Ismail
2017-07-01
We use an event-by-event hydrodynamical description of the heavy-ion collision process with Glauber initial conditions to calculate the thermal emission of photons. The photon rates in the hadronic phase follow from a spectral function approach and a density expansion, while in the partonic phase they follow from the Arnold-Moore-Yaffe (AMY) perturbative rates. The calculated photon elliptic flows are lower than those reported recently by both the ALICE and PHENIX collaborations.
Computer programs for predicting supersonic and hypersonic interference flow fields and heating
NASA Technical Reports Server (NTRS)
Morris, D. J.; Keyes, J. W.
1973-01-01
This report describes computer codes which calculate two-dimensional shock interference patterns. These codes compute the six types of interference flows as defined by Edney (Aeronaut. Res. Inst. of Sweden FAA Rep. 115). Results include properties of the inviscid flow field and the inviscid-viscous interaction at the surface along with peak pressure and peak heating at the impingement point.
NASA Astrophysics Data System (ADS)
Prakash, Ram; Gai, Sudhir L.; O'Byrne, Sean; Brown, Melrose
2016-11-01
The flow over a `tick' shaped configuration is performed using two Direct Simulation Monte Carlo codes: the DS2V code of Bird and the code from Sandia National Laboratory, called SPARTA. The configuration creates a flow field, where the flow is expanded initially but then is affected by the adverse pressure gradient induced by a compression surface. The flow field is challenging in the sense that the full flow domain is comprised of localized areas spanning continuum and transitional regimes. The present work focuses on the capability of SPARTA to model such flow conditions and also towards a comparative evaluation with results from DS2V. An extensive grid adaptation study is performed using both the codes on a model with a sharp leading edge and the converged results are then compared. The computational predictions are evaluated in terms of surface parameters such as heat flux, shear stress, pressure and velocity slip. SPARTA consistently predicts higher values for these surface properties. The skin friction predictions of both the codes don't give any indication of separation but the velocity slip plots indicate an incipient separation behavior at the corner. The differences in the results are attributed towards the flow resolution at the leading edge that dictates the downstream flow characteristics.
NASA Technical Reports Server (NTRS)
Kumar, A.; Graves, R. A., Jr.
1980-01-01
A user's guide is provided for a computer code which calculates the laminar and turbulent hypersonic flows about blunt axisymmetric bodies, such as spherically blunted cones, hyperboloids, etc., at zero and small angles of attack. The code is written in STAR FORTRAN language for the CDC-STAR-100 computer. Time-dependent, viscous-shock-layer-type equations are used to describe the flow field. These equations are solved by an explicit, two-step, time asymptotic, finite-difference method. For the turbulent flow, a two-layer, eddy-viscosity model is used. The code provides complete flow-field properties including shock location, surface pressure distribution, surface heating rates, and skin-friction coefficients. This report contains descriptions of the input and output, the listing of the program, and a sample flow-field solution.
Issues and approach to develop validated analysis tools for hypersonic flows: One perspective
NASA Technical Reports Server (NTRS)
Deiwert, George S.
1992-01-01
Critical issues concerning the modeling of low-density hypervelocity flows where thermochemical nonequilibrium effects are pronounced are discussed. Emphasis is on the development of validated analysis tools. A description of the activity in the Ames Research Center's Aerothermodynamics Branch is also given. Inherent in the process is a strong synergism between ground test and real-gas computational fluid dynamics (CFD). Approaches to develop and/or enhance phenomenological models and incorporate them into computational flow-field simulation codes are discussed. These models have been partially validated with experimental data for flows where the gas temperature is raised (compressive flows). Expanding flows, where temperatures drop, however, exhibit somewhat different behavior. Experimental data for these expanding flow conditions are sparse; reliance must be made on intuition and guidance from computational chemistry to model transport processes under these conditions. Ground-based experimental studies used to provide necessary data for model development and validation are described. Included are the performance characteristics of high-enthalpy flow facilities, such as shock tubes and ballistic ranges.
CFD Tools for Design and Simulation of Transient Flows in Hypersonic Facilities
2010-03-24
slugs) and CFL is the specified Courant - Friedrichs-Lewy number. It is normally restricted to CFL ≤ 0.5 in the simulations discussed later. For each cell...here. In an unsteady, time-accurate flow simulation, the allowable timestep is constrained by the Courant -Friedrichs-Lewy (CFL) criterion. In a...flow simulation codes. Beyond the list of authors, material in this paper has come from the efforts of Richard Morgan, David Mee, Tim McIntyre, Paul
Freon-14 as a working medium in wind tunnels for modeling hypersonic air flows
NASA Astrophysics Data System (ADS)
Komarov, V. N.; Polianskii, O. Iu.; Chirikhin, A. V.
The thermodynamic and kinetic properties of Freon-14, a gas characterized by a low adiabat and well suited to the modelling of real gas effects in wind tunnels, are examined. It is shown that flows that are free from condensation and closely resemble equilibrium flows can be obtained over a sufficiently wide range of wind tunnel regimes and model sizes. Viscosity coefficients of Freon-14 are presented.
NASA Technical Reports Server (NTRS)
Miller, C. G., III
1975-01-01
Measured shock shapes are presented for sphere and hemisphere models in helium, air, CF4, C2F6, and CO2 test gases, corresponding to normal-shock density ratios (primary factor governing shock detachment distance of blunt bodies at hypersonic speeds) from 4 to 19. These shock shapes were obtained in three facilities capable of generating the high density ratios experienced during planetary entry at hypersonic conditions; namely, the 6-inch expansion tube, with hypersonic CF4 tunnel, and pilot CF4 Mach 6 tunnel (with CF4 replaced by C2F6). Measured results are compared with several inviscid perfect-gas shock shape predictions, in which an effective ratio of specific heats is used as input, and with real-gas predictions which include effects of a laminar viscous layer and thermochemical nonequilibrium.
Shock-Wave/Boundary-Layer Interactions in Hypersonic Low Density Flows
NASA Technical Reports Server (NTRS)
Moss, James N.; Olejniczak, Joseph
2004-01-01
Results of numerical simulations of Mach 10 air flow over a hollow cylinder-flare and a double-cone are presented where viscous effects are significant. The flow phenomena include shock-shock and shock- boundary-layer interactions with accompanying flow separation, recirculation, and reattachment. The purpose of this study is to promote an understanding of the fundamental gas dynamics resulting from such complex interactions and to clarify the requirements for meaningful simulations of such flows when using the direct simulation Monte Carlo (DSMC) method. Particular emphasis is placed on the sensitivity of computed results to grid resolution. Comparisons of the DSMC results for the hollow cylinder-flare (30 deg.) configuration are made with the results of experimental measurements conducted in the ONERA RSCh wind tunnel for heating, pressure, and the extent of separation. Agreement between computations and measurements for various quantities is good except that for pressure. For the same flow conditions, the double- cone geometry (25 deg.- 65 deg.) produces much stronger interactions, and these interactions are investigated numerically using both DSMC and Navier-Stokes codes. For the double-cone computations, a two orders of magnitude variation in free-stream density (with Reynolds numbers from 247 to 24,7 19) is investigated using both computational methods. For this range of flow conditions, the computational results are in qualitative agreement for the extent of separation with the DSMC method always predicting a smaller separation region. Results from the Navier-Stokes calculations suggest that the flow for the highest density double-cone case may be unsteady; however, the DSMC solution does not show evidence of unsteadiness.
NASA Astrophysics Data System (ADS)
Tarnavskii, G. A.
2006-07-01
The physical aspects of the effective-adiabatic-exponent model making it possible to decompose the total problem on modeling of high-velocity gas flows into individual subproblems (“physicochemical processes” and “ aeromechanics”), which ensures the creation of a universal and efficient computer complex divided into a number of independent units, have been analyzed. Shock-wave structures appearing at entry into the duct of a hypersonic aircraft have been investigated based on this methodology, and the influence of the physical properties of the gas medium in a wide range of variations of the effective adiabatic exponent has been studied.
Experimental aerothermodynamic research of hypersonic aircraft
NASA Technical Reports Server (NTRS)
Cleary, Joseph W.
1987-01-01
The 2-D and 3-D advance computer codes being developed for use in the design of such hypersonic aircraft as the National Aero-Space Plane require comparison of the computational results with a broad spectrum of experimental data to fully assess the validity of the codes. This is particularly true for complex flow fields with control surfaces present and for flows with separation, such as leeside flow. Therefore, the objective is to provide a hypersonic experimental data base required for validation of advanced computational fluid dynamics (CFD) computer codes and for development of more thorough understanding of the flow physics necessary for these codes. This is being done by implementing a comprehensive test program for a generic all-body hypersonic aircraft model in the NASA/Ames 3.5 foot Hypersonic Wind Tunnel over a broad range of test conditions to obtain pertinent surface and flowfield data. Results from the flow visualization portion of the investigation are presented.
Influence of leading edge bluntness on hypersonic flow in a generic internal-compression inlet
NASA Astrophysics Data System (ADS)
Borovoy, V.; Egorov, I.; Mosharov, V.; Radchenko, V.; Skuratov, A.; Struminskaya, I.
2015-06-01
Flow and heat transfer inside a generic inlet are investigated experimentally. The cross section of the inlet is rectangular. The inlet is installed on a flat plat at a significant distance from the leading edge. The experiments are performed in TsAGI wind tunnel UT-1M working in the Ludwieg tube mode at Mach number M∞ = 5 and Reynolds numbers (based on the plate length L = 320 mm) Re∞L = 23 · 106 and 13 · 106. Steady flow duration is 40 ms. Optical panoramic methods are used for investigation of flow outside and inside the inlet as well. For this purpose, the cowl and one of two compressing wedges are made of a transparent material. Heat flux distribution is measured by thin luminescent Temperature Sensitive Paint (TSP). Surface flow and shear stress visualization is performed by viscous oil containing luminophor particles. The investigation shows that at high contraction ratio of the inlet, an increase of plate or cowl bluntness to some critical value leads to sudden change of the flow structure.
Grid-refinement study of hypersonic laminar flow over a 2-D ramp
NASA Technical Reports Server (NTRS)
Thomas, James L.; Rudy, David H.; Kumar, Ajay; Van Leer, Bram
1991-01-01
Computations were made for those test cases of Problem 3 which were designated as laminar flows, viz., test cases 3.1, 3.2, 3.4, and 3.5. These test cases corresponded to flows over a flat plate and a compression ramp at high Mach number and at high Reynolds number. The computations over the compression ramps indicate a substantial streamwise extent of separation. Based on previous experience with separated laminar flows at high Mach numbers which indicated a substantial effect with spatial grid refinement, a series of computations with different grid sizes were performed. Also, for the flat plate, comparisons of the results for two different algorithms were made.
Estimation of propulsion-induced effects on transonic flows over a hypersonic configuration
NASA Technical Reports Server (NTRS)
Hartwich, Peter M.; Frink, Neal T.
1992-01-01
Boundary conditions are formulated for treating selected patches of a subject configuration as inlet or nozzle areas. For subsonic inflow, the mass flow through the inlet is controlled by the exhaust conditions and the effects of mass and heat addition. For supersonic inflow, the exhaust conditions are based on the inlet conditions and on combustion data. These formulations were included into an existing Euler/Navier-Stokes solver. Comparisons with experimental data demonstrate that the resulting software package efficiently permits the assessment of propulsion-induced effects on external flow fields, particularly around highly blended configurations.
NASA Astrophysics Data System (ADS)
Menezes, V.; Sun, M.; Jagadeesh, G.; Reddy, K. P. J.; Takayama, K.
The problem of wake flow at high speeds and the drag associated with it are a significant source of observation in the design of missiles, projectiles and other typical high speed vehicles. A large separated wake at the base of the body in flight would cause an increase in the overall drag due to reduced base pressure force, which otherwise would oppose the axial force on the body. The wake studies of high speed bodies also gain importance due to the severe aerodynamic heating problem and a high rise in the temperature of the base flow.
Numerical simulation of unsteady flow in a hypersonic shock tunnel facility
NASA Technical Reports Server (NTRS)
Cambier, Jean-Luc; Tokarcik, Susan; Prabhu, Dinesh K.
1992-01-01
This paper describes the computational work performed on the simulation of a 16-in shock-tunnel facility. The numerical problems encountered during the computation of these flows are discussed along with the validity of some approximations used, notably concerning the reduction of the problem into problems of smaller dimensionality. Quasi-1D simulations can be used to help design experiments, or to better understanding the characteristics of the facility. An application to the design of a nonintrusive diagnostic is shown. The multidimensional flow transients computed include the shock reflection at the end of the driven tube, the shock propagation down the nozzle, and the breaking of the main diaphragm.
Experimental Investigation of a Hypersonic Inlet with Variable Sidewall for Flow Control
NASA Astrophysics Data System (ADS)
Rolim, T. C.; Lu, F. K.
The main function of a scramjet inlet is to decelerate and compress the air for subsequent reaction with the fuel inside the combustor and, of course, contribute toward meeting the thrust requirement for the entire mission by providing adequate mass flow. It is desirable that the inlet be lightweight and that its geometry be capable of producing a uniform flow in an appropriate state to permit efficient mixing and subsequent combustion. Engine cycle analysis indicates that high contraction ratios CR are desirable for achieving high overall engine efficiency.
Second-order small-disturbance solutions for hypersonic flow over power-law bodies
NASA Technical Reports Server (NTRS)
Townsend, J. C.
1975-01-01
Similarity solutions were found which give the adiabatic flow of an ideal gas about two-dimensional and axisymmetric power-law bodies at infinite Mach number to second order in the body slenderness parameter. The flow variables were expressed as a sum of zero-order and perturbation similarity functions for which the axial variations in the flow equations separated out. The resulting similarity equations were integrated numerically. The solutions, which are universal functions, are presented in graphic and tabular form. To avoid a singularity in the calculations, the results are limited to body power-law exponents greater than about 0.85 for the two-dimensional case and 0.75 for the axisymmetric case. Because of the entropy layer induced by the nose bluntness (for power-law bodies other than cones and wedges), only the pressure function is valid at the body surface. The similarity results give excellent agreement with the exact solutions for inviscid flow over wedges and cones having half-angles up to about 20 deg. They give good agreement with experimental shock-wave shapes and surface-pressure distributions for 3/4-power axisymmetric bodies, considering that Mach number and boundary-layer displacement effects are not included in the theory.
Assessment of One- and Two-Equation Turbulence Models for Hypersonic Transitional Flows
ROY,CHRISTOPHER J.; BLOTTNER,FREDERICK G.
2000-01-14
Many Navier-Stokes codes require that the governing equations be written in conservation form with a source term. The Spalart-Allmaras one-equation model was originally developed in substantial derivative form and when rewritten in conservation form, a density gradient term appears in the source term. This density gradient term causes numerical problems and has a small influence on the numerical predictions. Further work has been performed to understand and to justify the neglect of this term. The transition trip term has been included in the one-equation eddy viscosity model of Spalart-Allmaras. Several problems with this model have been discovered when applied to high-speed flows. For the Mach 8 flat plate boundary layer flow with the standard transition method, the Baldwin-Barth and both k-{omega} models gave transition at the specified location. The Spalart-Allmaras and low Reynolds number k-{var_epsilon} models required an increase in the freestream turbulence levels in order to give transition at the desired location. All models predicted the correct skin friction levels in both the laminar and turbulent flow regions. For Mach 8 flat plate case, the transition location could not be controlled with the trip terms as given in the Spalart-Allmaras model. Several other approaches have been investigated to allow the specification of the transition location. The approach that appears most appropriate is to vary the coefficient that multiplies the turbulent production term in the governing partial differential equation for the eddy viscosity (Method 2). When this coefficient is zero, the flow remains laminar. The coefficient is increased to its normal value over a specified distance to crudely model the transition region and obtain fully turbulent flow. While this approach provides a reasonable interim solution, a separate effort should be initiated to address the proper transition procedure associated with the turbulent production term. Also, the transition process
Numerical Investigation of PLIF Gas Seeding for Hypersonic Boundary Layer Flows
NASA Technical Reports Server (NTRS)
Johanson, Craig T.; Danehy, Paul M.
2012-01-01
Numerical simulations of gas-seeding strategies required for planar laser-induced fluorescence (PLIF) in a Mach 10 air flow were performed. The work was performed to understand and quantify adverse effects associated with gas seeding and to compare different flow rates and different types of seed gas. The gas was injected through a slot near the leading edge of a flat plate wedge model used in NASA Langley Research Center's 31- Inch Mach 10 Air Tunnel facility. Nitric oxide, krypton, and iodine gases were simulated at various injection rates. Simulation results showing the deflection of the velocity field for each of the cases are presented. Streamwise distributions of velocity and concentration boundary layer thicknesses as well as vertical distributions of velocity, temperature, and mass distributions are presented for each of the cases. Relative merits of the different seeding strategies are discussed.
A multi-temperature TVD algorithm for relaxing hypersonic flows. [Total Variation Diminishing
NASA Technical Reports Server (NTRS)
Cambier, Jean-Luc; Menees, Gene P.
1989-01-01
In this paper, the extension of a multispecies TVD algorithm, second-order accurate for real-gas flows to a multitemperature formulation is described. The convection algorithm is coupled to internal relaxation processes, and the features of the coupling are examined. The first version consists of a three-temperature model, where translational-rotational, vibrational, and electronic energy modes are separately convected. Although several species are present, there is only one vibrational temperature in this model. The second version generalizes to a vibrational temperature for each molecular specie, with additional couplings between species. The algorithms are applied to a generic two-dimensional flow field, and results are compared with experimental observations.
Olstad, S.J.
1995-08-01
The application of a method for determining the temperature of an oxygen-replenished air stream heated to 2600 K by a hydrogen burner is reviewed and discussed. The purpose of the measurements is to determine the spatial uniformity of the temperature in the core flow of a ramjet test facility. The technique involves sampling the product gases at the exit of the test section nozzle to infer the makeup of the reactant gases entering the burner. Knowing also the temperature of the inlet gases and assuming the flow is at chemical equilibrium, the adiabatic flame temperature is determined using an industry accepted chemical equilibrium computer code. Local temperature depressions are estimated from heat loss calculations. A description of the method, hardware and procedures is presented, along with local heat loss estimates and uncertainty assessments. The uncertainty of the method is estimated at {+-}31 K, and the spatial uniformity was measured within {+-}35 K.
NASA Technical Reports Server (NTRS)
Harloff, G. J.; Lai, H. T.; Nelson, E. S.
1988-01-01
The PARC2D code has been selected to analyze the flowfields of a representative hypersonic scramjet nozzle over a range of flight conditions from Mach 3 to 20. The flowfields, wall pressures, wall skin friction values, heat transfer values and overall nozzle performance are presented.
Development of Combined Asymptotic and Numerical Procedures for Transonic and Hypersonic Flows.
1996-04-01
transparent with respect to these disturbances. It is this property that provides an opportunity to use absorption to maintain or increase the laminar flow... properties . For thin shock layers, we have extended steady Newtonian asymptotic theory to the unsteady case. We have used the theory to demonstrate...Norman D. Malmuth a Professor Julian D. Cole 0 Dr. Alexander V. Fedorov a Dr. Andrd lhokhlov * Professor Vladimir Ya Neiland * Dr. Vera M. Neyland
High-fidelity simulation of compressible flows for hypersonic propulsion applications
NASA Astrophysics Data System (ADS)
Otis, Collin C.
In the first part of this dissertation, the scalar filtered mass density function (SFMDF) methodology is implemented into the computer code US3D. The SFMDF is a sub-grid scale closure and is simulated via a Lagrangian Monte Carlo solver. US3D is an Eulerian finite volume code and has proven very effective for compressible flow simulations. The resulting SFMDF-US3D code is employed for large eddy simulation (LES) of compressible turbulent flows on unstructured meshes. Simulations are conducted of subsonic and supersonic flows. The consistency and accuracy of the simulated results are assessed along with appraisal of the overall performance of the methodology. In the second part of this dissertation, a new methodology is developed for accurate capturing of discontinuities in multi-block finite difference simulations of hyperbolic partial differential equations. The fourth-order energy-stable weighted essentially non-oscillatory (ESWENO) scheme on closed domains is combined with simultaneous approximation term (SAT) weak interface and boundary conditions. The capability of the methodology is demonstrated for accurate simulations in the presence of significant and abrupt changes in grid resolution between neighboring subdomains. Results are presented for the solutions of linear scalar hyperbolic wave equations and the Euler equations in one and two dimensions. Strong discontinuities are passed across subdomain interfaces without significant distortions. It is demonstrated that the methodology provides stable and accurate solutions even when large differences in the grid-spacing exist, whereas strong imposition of the interface conditions causes noticeable oscillations. Keywords: Large eddy simulation, filtered density function, turbulent reacting flows, multi-block finite difference schemes, high-order numerical methods, WENO shock-capturing, computational fluid dynamics.
Effect of Dielectric Barrier Discharge Plasma Actuators on Non-equilibrium Hypersonic Flows
2014-10-28
alternative, we propose the use of micro- second pulsed Dielectric Barrier Discharge (DBD) plasma actuator7,20–23 for the net reduction of thermal load...include counter-flow plasma jets , energy deposition methods using high energy beams like electron and micro-wave, and gas heating using arc or electric...actuators have been implemented in nanosecond and microsecond pulse widths. Nanosecond pulsed DBD plasma actuators have also been successful in high
Munafò, A; Panesi, M; Magin, T E
2014-02-01
A Boltzmann rovibrational collisional coarse-grained model is proposed to reduce a detailed kinetic mechanism database developed at NASA Ames Research Center for internal energy transfer and dissociation in N(2)-N interactions. The coarse-grained model is constructed by lumping the rovibrational energy levels of the N(2) molecule into energy bins. The population of the levels within each bin is assumed to follow a Boltzmann distribution at the local translational temperature. Excitation and dissociation rate coefficients for the energy bins are obtained by averaging the elementary rate coefficients. The energy bins are treated as separate species, thus allowing for non-Boltzmann distributions of their populations. The proposed coarse-grained model is applied to the study of nonequilibrium flows behind normal shock waves and within converging-diverging nozzles. In both cases, the flow is assumed inviscid and steady. Computational results are compared with those obtained by direct solution of the master equation for the rovibrational collisional model and a more conventional multitemperature model. It is found that the proposed coarse-grained model is able to accurately resolve the nonequilibrium dynamics of internal energy excitation and dissociation-recombination processes with only 20 energy bins. Furthermore, the proposed coarse-grained model provides a superior description of the nonequilibrium phenomena occurring in shock heated and nozzle flows when compared with the conventional multitemperature models.
Viscous shock-layer solutions for the low-density hypersonic flow past long slender bodies
NASA Technical Reports Server (NTRS)
Gupta, R. N.; Moss, J. N.; Zoby, E. V.; Tiwari, S. N.; Lee, K. P.
1988-01-01
Results are obtained for the surface pressure, drag, heat-transfer, and skin-friction coefficients for hyperboloids and sphere cones. Body half angles from 5 to 22.5 degrees are considered for various low-density flow conditions. Recently obtained surface-slip and shock-slip equations are employed to account for the low-density effects. The method of solution employed for the viscous shock-layer (VSL) equations is a partially coupled spatial-marching implicit finite-difference technique. The flow cases analyzed include highly cooled long slender bodies in high Mach number flows. The present perfect-gas VSL calculations compare quite well with available experimental data. Results have also been obtained from the steady-state Navier-Stokes (NS) equations by successive approximations. Comparison between the NS and VSL results indicates that VSL equations even with body and shock-slip boundary conditions may not be adequate in the stagnation region at altitudes greater than about 75 km for the cases analyzed here.
Dual-Code Solution Strategy for Chemically-Reacting Hypersonic Flows
NASA Technical Reports Server (NTRS)
Wood, William A.; Eberhardt, Scott
1995-01-01
A new procedure seeks to combine the thin-layer Navier-Stokes solver LAURA with the parabolized Navier-Stokes solver UPS for the aerothermodynamic solution of chemically-reacting air flow fields. The interface protocol is presented and the method is applied to two slender, blunted shapes. Both axisymmetric and three-dimensional solutions are included with surface pressure and heat transfer comparisons between the present method and previously published results. The case of Mach 25 flow over an axisymmetric six degree sphere-cone with a non-catalytic wall is considered to 100 nose radii. A stability bound on the marching step size was observed with this case and is attributed to chemistry effects resulting from the non-catalytic wall boundary condition. A second case with Mach 28 flow over a sphere-cone-cylinder-flare configuration is computed at both two and five degree angles of attack with a fully-catalytic wall. Surface pressures are seen to be within five percent with the present method compared to the baseline LAURA solution and heat transfers are within 10 percent. The effect of grid resolution is investigated in both the radial and streamwise directions. The procedure demonstrates significant, order of magnitude reductions in solution time and required memory for the three-dimensional case in comparison to an all thin-layer Navier-Stokes solution.
Apparatus and method for generating large mass flow of high temperature air at hypersonic speeds
NASA Technical Reports Server (NTRS)
Sabol, A. P.; Stewart, R. B. (Inventor)
1973-01-01
High temperature, high mass air flow and a high Reynolds number test air flow in the Mach number 8-10 regime of adequate test flow duration is attained by pressurizing a ceramic-lined storage tank with air to a pressure of about 100 to 200 atmospheres. The air is heated to temperatures of 7,000 to 8,000 R prior to introduction into the tank by passing the air over an electric arc heater means. The air cools to 5,500 to 6,000 R while in the tank. A decomposable gas such as nitrous oxide or a combustible gas such as propane is injected into the tank after pressurization and the heated pressurized air in the tank is rapidly released through a Mach number 8-10 nozzle. The injected gas medium upon contact with the heated pressurized air effects an exothermic reaction which maintains the pressure and temperature of the pressurized air during the rapid release.
NASA Technical Reports Server (NTRS)
Cook, W. J.
1975-01-01
The laminar boundary layer has been theoretically studied for six gases for flows over cold walls with zero pressure gradient at Mach numbers between 5.5 and 12.5 to correlate boundary layer quantities for the various gases. The flow conditions considered correspond to those that can be generated in test facilities such as the shock tunnel and the expansion tube. Computed results obtained using real gas properties indicate that the Eckert number based on edge conditions serves to correlate the results in terms of the wall shear stress and enthalpy gradient, the Stanton number, and the momentum thickness for the various gases within plus or minus 10 per cent for Te = Tw and Te approximately 3Tw. Computed Reynolds analogy factors exhibit very good agreement with those predicted by the Colburn analogy. Velocity and displacement thicknesses correlate well with Eckert number for Te = Tw, but fail to correlate for Te approximately 3Tw. Differences in results are traced to property variations. Results show that the Eckert number is a significant correlating variable for the flows considered.
Plume effects on the flow around a blunted cone at hypersonic speeds
NASA Technical Reports Server (NTRS)
Atcliffe, P.; Kumar, D.; Stollery, J. L.
1992-01-01
Tests at M = 8.2 show that a simulated rocket plume at the base of a blunted cone can cause large areas of separated flow, with dramatic effects on the heat transfer rate distribution. The plume was simulated by solid discs of varying sizes or by an annular jet of gas. Flow over the cone without a plume is fully laminar and attached. Using a large disc, the boundary layer is laminar at separation at the test Reynolds number. Transition occurs along the separated shear layer and the boundary layer quickly becomes turbulent. The reduction in heat transfer associated with a laminar separated region is followed by rising values as transition occurs and the heat transfer rates towards the rear of the cone substantially exceed the values obtained without a plume. With the annular jet or a small disc, separation occurs much further aft, so that heat transfer rates at the front of the cone are comparable with those found without a plume. Downstream of separation the shear layer now remains laminar and the heat transfer rates to the surface are significantly lower than the attached flow values.
Simulations of hypersonic, high-enthalpy separated flow over a 'tick' configuration
NASA Astrophysics Data System (ADS)
Moss, J. N.; O'Byrne, S.; Deepak, N. R.; Gai, S. L.
2012-11-01
The effect of slip is investigated in direct simulation Monte Carlo and Navier-Stokes-based computations of the separated flow between an expansion and a following compression surface, a geometry we call the 'tick' configuration. This configuration has been chosen as a test of separated flow with zero initial boundary layer thickness, a flowfield well suited to Chapman's analytical separated flow theories. The predicted size of the separated region is different for the two codes, although both codes meet their respective particle or grid resolution requirements. Unlike previous comparisons involving cylinder flares or double cones, the separation does not occur in a region of elevated density, and is therefore well suited to the direct simulation Monte Carlo method because the effect of slip at the surface is significant. The reasons for the difference between the two calculations are hypothesized to be a combination of significant rarefaction effects near the expansion surface and the non-zero radius of the leading edge. When the leading edge radius is accounted for, the rarefaction effect at the leading edge is less significant and the behavior of the flowfields predicted by the two methods becomes more similar.
NASA Astrophysics Data System (ADS)
Lorzel, Heath
The time-dependent, 2½-dimensional, axisymmetric, magnetohydrodynamics (MHD) solver MACH2 has been upgraded to include the effects of non-equilibrium air chemistry using the well-established reaction model developed by Park. Several validation cases are presented based on comparisons to the experimentally deduced shock stand-off distance of nitrogen flow over spheres, the shock stand-off distance of spheres fired into air in a ballistic test facility, and the electron number density on the surface of the Ram-C re-entry experiment. In addition, the magnetic induction equation has been upgraded with new verified models that compute the effects of the Hall and ion slip terms. The upgraded code is utilized to model an annular, Hall-type MHD generator that can be employed upstream of a turbojet engine for freestream conditions corresponding to Mach 5 flight at an altitude of 20km. The simulations demonstrate the feasibility of convening inlet kinetic power to storable electric power. Using ionization provided by electron-beam guns and a radial magnetic field B=3T, the generator is shown to produce a maximum of 4.8MW of electric power while reducing the total kinetic power of the flow by 31%. Optimizing the loading parameter, K*Load, across the electrodes demonstrates that the generator could produce 1.54MW of excess electric power that can be stored and used for on-board power requirements. Further, the reduction in flow kinetic power results in an increase in static pressure of 30% and a reduction in stagnation temperature of 3% at the turbojet's compressor inlet that aids the subsequent process of combustion.
Laser absorption flow diagnostics for use in hypersonic ground-based and flight experiments
NASA Astrophysics Data System (ADS)
Cavolowsky, John A.; Newfield, Mark E.
1992-12-01
Consideration is given to laboratory ground-based scale systems for measuring OH and NO and diode laser systems for measuring O2 and H2O. Several diagnostic instruments which are currently being used in a scramjet combustor test rig of the Ames 16-inch shock tunnel are described. Data on mole fraction, temperature, density, pressure, and velocity is obtained using specific data reduction techniques. These measurements provide critical data for evaluation of the propulsion system performance which include mass capture for quantification of fuel equivalence ratio (O2), flow contamination and emission levels (NO), combustion progress (OH), and injector mixing and combustion efficiency (H2O).
Analytical and experimental studies of shock interference heating in hypersonic flows
NASA Technical Reports Server (NTRS)
Keyes, J. W.; Hains, F. D.
1973-01-01
An analytical and experimental study is presented of the aerodynamic heating resulting from six types of shock interference patterns encountered in high speed flow. Centerline measurements of pressure and heat transfer distributions on basic bodies were obtained in four wind tunnels for Mach numbers from 6 to 20, specific heat ratios from 1.27 to 1.67, and free stream Reynolds numbers from 3 million to 25.6 million per meter. Peak heating and peak pressures up to 17 and 7.5 times stagnation values, respectively, were measured. In general, results obtained from semiempirical methods developed for each of the six types of interference agreed with the experimental peaks.
Shock Wave/Turbulent Boundary Layer Interaction in High-Reynolds-Number Hypersonic Flows
1987-07-01
XtSTART TRW WEDGE SURFACERe4/" WEDGE SHOCK 0l 103, TRIPLE PLATE SHOCK UPSTREAM,- POINT I• / tFLUENCE SHOCK "JET" -PLATE BOUNDARY.,.) • -:3•< ......LAYER...particularly for turbulent interacting flows, an analysis of the characteristic scale lengths, like that employed in triple deck theory, should be performed...constant A wavelength of light •= extent of 2-D field traversed by light waves , 0 tref = relative change in density between the reference point and the
Structure formation in a colliding flow: The Herschel view of the Draco nebula
NASA Astrophysics Data System (ADS)
Miville-Deschênes, M.-A.; Salomé, Q.; Martin, P. G.; Joncas, G.; Blagrave, K.; Dassas, K.; Abergel, A.; Beelen, A.; Boulanger, F.; Lagache, G.; Lockman, F. J.; Marshall, D. J.
2017-03-01
Context. The Draco nebula is a high Galactic latitude interstellar cloud observed at velocities corresponding to the intermediate velocity cloud regime. This nebula shows unusually strong CO emission and remarkably high-contrast small-scale structures for such a diffuse high Galactic latitude cloud. The 21 cm emission of the Draco nebula reveals that it is likely to have been formed by the collision of a cloud entering the disk of the Milky Way. Such physical conditions are ideal to study the formation of cold and dense gas in colliding flows of diffuse and warm gas. Aims: The objective of this study is to better understand the process of structure formation in a colliding flow and to describe the effects of matter entering the disk on the interstellar medium. Methods: We conducted Herschel-SPIRE observations of the Draco nebula. The clumpfind algorithm was used to identify and characterize the small-scale structures of the cloud. Results: The high-resolution SPIRE map reveals the fragmented structure of the interface between the infalling cloud and the Galactic layer. This front is characterized by a Rayleigh-Taylor (RT) instability structure. From the determination of the typical length of the periodic structure (2.2 pc) we estimated the gas kinematic viscosity. This allowed us to estimate the dissipation scale of the warm neutral medium (0.1 pc), which was found to be compatible with that expected if ambipolar diffusion were the main mechanism of turbulent energy dissipation. The statistical properties of the small-scale structures identified with clumpfind are found to be typical of that seen in molecular clouds and hydrodynamical turbulence in general. The density of the gas has a log-normal distribution with an average value of 103 cm-3. The typical size of the structures is 0.1-0.2 pc, but this estimate is limited by the resolution of the observations. The mass of these structures ranges from 0.2 to 20 M⊙ and the distribution of the more massive structures
NASA Technical Reports Server (NTRS)
Miller, C. G., III
1975-01-01
Shock shape results for flat-faced cylinders, spheres, and spherically blunted cones in various test gases, along with preliminary results from a calibration study performed in the Langley 6-inch expansion tube are presented. Free-stream velocities from 5 to 7 km/sec are generated at hypersonic conditions with helium, air, and CO2, resulting in normal shock density ratios from 4 to 19. Ideal-gas shock shape predictions, in which an effective ratio of specific heats is used as input, are compared with the measured results. The effect of model diameter is examined to provide insight to the thermochemical state of the flow in the shock layer. The regime for which equilibrium exists in the shock layer for the present air and CO2 test conditions is defined. Test core flow quality, test repeatability, and comparison of measured and predicted expansion-tube flow quantities are discussed.
Comparison of Nonequilibrium Solution Algorithms Applied to Chemically Stiff Hypersonic Flows
NASA Technical Reports Server (NTRS)
Palmer, Grant; Venkatapathy, Ethiraj
1995-01-01
Three solution algorithms, explicit under-relaxation, point implicit, and lower-upper symmetric Gauss-Seidel, are used to compute nonequilibrium flow around the Apollo 4 return capsule at the 62-km altitude point in its descent trajectory. By varying the Mach number, the efficiency and robustness of the solution algorithms were tested for different levels of chemical stiffness.The performance of the solution algorithms degraded as the Mach number and stiffness of the flow increased. At Mach 15 and 30, the lower-upper symmetric Gauss-Seidel method produces an eight order of magnitude drop in the energy residual in one-third to one-half the Cray C-90 computer time as compared to the point implicit and explicit under-relaxation methods. The explicit under-relaxation algorithm experienced convergence difficulties at Mach 30 and above. At Mach 40 the performance of the lower-upper symmetric Gauss-Seidel algorithm deteriorates to the point that it is out performed by the point implicit method. The effects of the viscous terms are investigated. Grid dependency questions are explored.
Atomic-level simulation of gas-surface interactions and hypersonic flow features
NASA Astrophysics Data System (ADS)
Valentini, Paolo
First-principles based atomic-level simulations can be used to develop realistic models of gas-surface chemistry and gas-phase chemistry, or to inform or validate the existing ones. A mechanism-based finite rate catalytic wall boundary condition for use in reacting flow simulation is presented. The input parameters of this model are the reaction rates kr of each elementary one-step process, including molecular and atomic adsorption. Therefore, ReaxFF Molecular Dynamics (MD) simulations of O2 adsorption on a platinum (111) surface are presented and compared to the existing experimental and computational results in the literature. Furthermore, the surface state of a catalyst strongly affects its performance. Grand Canonical Monte Carlo simulations are used to study oxygen coverage on Pt(111) exposed to molecular oxygen at certain pressure and temperature conditions. The results compare well with the available experimental and first-principles data. Finally, full-scale MD simulations of normal shock waves in dilute and rarefied argon are presented, and are critically compared to similar Direct Simulation Monte Carlo (DSMC) calculations as well as experiments. For very dilute gases, a novel Event-Driven/Time-Driven algorithm is developed to speed up the MD simulation of rarefied gases using realistic spherically symmetric soft potentials. This technique could pave the way for the application of much more refined and expensive interatomic potentials (potentially ab initio) to MD simulations of nonequilibrium flow features in rarefied gases, involving vibrational excitation and chemical reactivity.
Infrared thermography of transition due to isolated roughness elements in hypersonic flows
NASA Astrophysics Data System (ADS)
Avallone, F.; Schrijer, F. F. J.; Cardone, G.
2016-02-01
Boundary layer transition in high-speed flows is a phenomenon that despite extensive research over the years is still extremely hard to predict. The presence of protrusions or gaps can lead to an accelerated laminar-to-turbulent transition enhancing the thermal loads and the skin friction coefficient. In the current investigation, inverse heat transfer measurements using infrared thermography are performed on the flow past different roughness geometries in the form of cylinders and diamond at free stream Mach number equal to 7.5, h/δ ranging between 0.5 and 0.9 (where h is the roughness height and δ is the boundary layer thickness), and Reθ ranging between 1305 and 2450. The roughness elements are positioned on a 5° ramp placed at zero angle of attack. The measurements indicate that the roughness geometry influences the transitional pattern while the frontal area influences both the transition location and the maximum value of the Stanton number along the centreline. Moreover, there is a strong connection between the streamwise centreline Stanton number and the spreading of the wake width. In particular, the transition process is characterized by an approximately constant wake width. Differently, the wake width spreads at the location where the streamwise centreline Stanton number reaches the turbulent level. This point corresponds to a local maximum of the wake amplitude defined as one half of the maximum spanwise variation of the Stanton number.
NASA Technical Reports Server (NTRS)
2008-01-01
[figure removed for brevity, see original site] Click on the image for the animation
NASA's Phoenix Mars Lander will enter the Martian atmosphere at hypersonic speeds. Friction will heat the forward-facing surface of the heat shield to a peak of about 1,420 degrees Celsius (2,600 degrees Fahrenheit) at an altitude of 41 kilometers (25.5 miles).
This illustration is part of the animation featured above.
The Phoenix Mission is led by the University of Arizona, Tucson, on behalf of NASA. Project management of the mission is by NASA's Jet Propulsion Laboratory, Pasadena, Calif. Spacecraft development is by Lockheed Martin Space Systems, Denver.
Coupled nonequilibrium flow, energy and radiation transport for hypersonic planetary entry
NASA Astrophysics Data System (ADS)
Frederick, Donald Jerome
An ever increasing demand for energy coupled with a need to mitigate climate change necessitates technology (and lifestyle) changes globally. An aspect of the needed change is a decrease in the amount of anthropogenically generated CO2 emitted to the atmosphere. The decrease needed cannot be expected to be achieved through only one source of change or technology, but rather a portfolio of solutions are needed. One possible technology is Carbon Capture and Storage (CCS), which is likely to play some role due to its combination of mature and promising emerging technologies, such as the burning of hydrogen in gas turbines created by pre-combustion CCS separation processes. Thus research on effective methods of burning turbulent hydrogen jet flames (mimicking gas turbine environments) are needed, both in terms of experimental investigation and model development. The challenge in burning (and modeling the burning of) hydrogen lies in its wide range of flammable conditions, its high diffusivity (often requiring a diluent such as nitrogen to produce a lifted turbulent jet flame), and its behavior under a wide range of pressures. In this work, numerical models are used to simulate the environment of a gas turbine combustion chamber. Concurrent experimental investigations are separately conducted using a vitiated coflow burner (which mimics the gas turbine environment) to guide the numerical work in this dissertation. A variety of models are used to simulate, and occasionally guide, the experiment. On the fundamental side, mixing and chemistry interactions motivated by a H2/N2 jet flame in a vitiated coflow are investigated using a 1-D numerical model for laminar flows and the Linear Eddy Model for turbulent flows. A radial profile of the jet in coflow can be modeled as fuel and oxidizer separated by an initial mixing width. The effects of species diffusion model, pressure, coflow composition, and turbulent mixing on the predicted autoignition delay times and mixture
Guy, Aurélien Bourdon, Anne Perrin, Marie-Yvonne
2015-04-15
In this work, a state-to-state vibrational and electronic collisional model is developed to investigate nonequilibrium phenomena behind a shock wave in an ionized nitrogen flow. In the ionization dynamics behind the shock wave, the electron energy budget is of key importance and it is found that the main depletion term corresponds to the electronic excitation of N atoms, and conversely the major creation terms are the electron-vibration term at the beginning, then replaced by the electron ions elastic exchange term. Based on these results, a macroscopic multi-internal-temperature model for the vibration of N{sub 2} and the electronic levels of N atoms is derived with several groups of vibrational levels of N{sub 2} and electronic levels of N with their own internal temperatures to model the shape of the vibrational distribution of N{sub 2} and of the electronic excitation of N, respectively. In this model, energy and chemistry source terms are calculated self-consistently from the rate coefficients of the state-to-state database. For the shock wave condition studied, a good agreement is observed on the ionization dynamics as well as on the atomic bound-bound radiation between the state-to-state model and the macroscopic multi-internal temperature model with only one group of vibrational levels of N{sub 2} and two groups of electronic levels of N.
Self-diffusion of vibrational states: Impact on the heat transfer in hypersonic flows
NASA Astrophysics Data System (ADS)
Josyula, E.; Kustova, E. V.; Vedula, P.
2014-12-01
In the present paper, the influence of self-diffusion of vibrationally excited states on the fluid dynamics and surface heat transfer in an axisymmetric Mach 7.2 air flow past a sphere-cone is discussed. Two models for state-to-state transport properties are considered: a simplified model using the Eucken's relation for thermal conductivity and Fick's law for diffusion velocities with the constant Lewis number, and a rigorous kinetic theory based model for the calculation of state-specific thermal conductivity, diffusion and thermal diffusion coefficients. The simplified model is applied for the flowfield simulation to avoid high computational costs. For the application of the accurate kinetic theory approach, a post-processing procedure is used. Inclusion of self-diffusion results in an increase in the surface heat flux of up to 6.5% upstream of a shoulder region. Thermal conductivity is found to be the primary contributor to surface heat flux; the influence of mass and thermal diffusion is found to be negligible. Self-diffusion has a considerably greater influence in decreasing heat flux in the downstream regions far from stagnation point.
1993-04-01
AD-A267 032 AGARD-51 4 ADVISORY GROUP FOR AEROSPACE RESEARCH & DEVELOPMENT 7 RUE ANCELLE 92200 NEUILLY SUR SEINE FRANCE DTIC AGARD CONFERENCE...the authors. Published April 1993 Copyright 0 AGARD 1993 All Rights Reserved ISBN 92-835-0694- 4 Printed by Speclalised Printing r.?rvices Limited 40...oif Hypersonic Vehicles I 7 papers) - Facilitics ( 4 papers) - Instrumentation (5 papers) - (TI) Validation and Data Accuracy (7 papers j - Rarefied Gas
NASA Astrophysics Data System (ADS)
Yatsukhno, D. S.
2017-02-01
The aim of the study is to obtain hypersonic waverider aerodynamic characteristics over the angle-of-attack range from 0° to 10° at Mach 4. The simplified virtual model of the waverider was designed. The Navier-Stokes equations solution was performed using computer code into which the method for splitting into physical processes was implemented. The unstructured grid was used in the computational process.
NASA Astrophysics Data System (ADS)
Sun, Xi-wan; Guo, Zhen-yun; Huang, Wei; Li, Shi-bin; Yan, Li
2017-02-01
The drag reduction and thermal protection system applied to hypersonic re-entry vehicles have attracted an increasing attention, and several novel concepts have been proposed by researchers. In the current study, the influences of performance parameters on drag and heat reduction efficiency of combinational novel cavity and opposing jet concept has been investigated numerically. The Reynolds-average Navier-Stokes (RANS) equations coupled with the SST k-ω turbulence model have been employed to calculate its surrounding flowfields, and the first-order spatially accurate upwind scheme appears to be more suitable for three-dimensional flowfields after grid independent analysis. Different cases of performance parameters, namely jet operating conditions, freestream angle of attack and physical dimensions, are simulated based on the verification of numerical method, and the effects on shock stand-off distance, drag force coefficient, surface pressure and heat flux distributions have been analyzed. This is the basic study for drag reduction and thermal protection by multi-objective optimization of the combinational novel cavity and opposing jet concept in hypersonic flows in the future.
NASA Technical Reports Server (NTRS)
Bertram, Mitchel H.; Feller, William V.
1959-01-01
A procedure based on the method of similar solutions is presented by which the skin friction, heat transfer, and boundary-layer thickness in a laminar hypersonic flow with pressure gradient may be rapidly evaluated if the pressure distribution is known. This solution, which at present is. restricted to power-law variations of pressure with surface distance, is presented for a wide range of exponents in the power law corresponding to both favorable and adverse pressure gradients. This theory has been compared to results from heat-transfer experiments on blunt-nose flat plates and a hemisphere cylinder at free-stream Mach numbers of 4 and 6.8. The flat-plate experiments included tests made at a Mach number of 6.8 over a range of angle of attack of +/- 10 deg. Reasonable agreement of the experimental and theoretical heat-transfer coefficients has been obtained as well as good correlation of the experimental results over the entire range of angle of attack studied. A similar comparison of theory with experiment was not feasible for boundary-layer-thickness data; however, the hypersonic similarity theory was found to account satisfactorily for the variation in boundary-layer thickness due to local pressure distribution for several sets of measurements.
Condensation in hypersonic nitrogen wind tunnels
NASA Technical Reports Server (NTRS)
Lederer, Melissa A.; Yanta, William J.; Ragsdale, William C.; Hudson, Susan T.; Griffith, Wayland C.
1990-01-01
Experimental observations and a theoretical model for the onset and disappearance of condensation are given for hypersonic flows of pure nitrogen at M = 10, 14 and 18. Measurements include Pitot pressures, static pressures and laser light scattering experiments. These measurements coupled with a theoretical model indicate a substantial non-equilibrium supercooling of the vapor phase beyond the saturation line. Typical results are presented with implications for the design of hypersonic wind tunnel nozzles.
1957-04-15
Hypersonic Boost Glider in 11 Inch Hypersonic Tunnel L57-1681 In 1957 Langley tested its HYWARDS design in the 11 Inch Hypersonic Tunnel. Photograph published in Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 by James R. Hansen. Page 369.
Theory of Molecular Cloud Formation through Colliding Flows: Successes and Limits
NASA Astrophysics Data System (ADS)
Hennebelle, P.
2013-10-01
We discuss the recent efforts which have been made to understand the formation of molecular clouds through the accumulation of diffuse material, a scenario sometimes called “colliding flows”. We present a set of statistics which have been inferred from these simulations and which seem to agree reasonably with observations seemingly suggesting that this scenario could indeed be applied to understand molecular cloud formation. We also emphasize the limits of this highly idealized model.
Morris, N.; Buttsworth, D.; Jones, T.; Brescianini, C. |
1995-09-01
Rocket plume exhaust structures are aerodynamically and thermochemically very complex and the prediction of plume properties such as temperature, velocity, pressure, chemical species concentrations and turbulence properties is a formidable task as there are no definitive models for viscous and chemical effects. Contemporary computational techniques are still in their infancy and cannot yet reliably predict plume properties. Only through validation of computer codes using experimental data, can computational models be developed to the point where they can be confidently used as design and predictive tools. The motivation for this study was to acquire well defined data for rocket plumes at low altitude hypersonic flight conditions so that the above issues could be investigated.
NASA Astrophysics Data System (ADS)
Morris, N.; Buttsworth, D.; Jones, T.; Brescianini, C.
Rocket plume exhaust structures are aerodynamically and thermochemically very complex and the prediction of plume properties such as temperature, velocity, pressure, chemical species concentrations and turbulence properties is a formidable task as there are no definitive models for viscous and chemical effects. Contemporary computational techniques are still in their infancy and cannot yet reliably predict plume properties. Only through validation of computer codes using experimental data, can computational models be developed to the point where they can be confidently used as design and predictive tools. The motivation for this study was to acquire well defined data for rocket plumes at low altitude hypersonic flight conditions so that the above issues could be investigated.
NASA Technical Reports Server (NTRS)
Grose, W. L.
1994-01-01
An approximate inverse solution is presented for the nonequilibrium flow in the inviscid shock layer about a vehicle in hypersonic flight. The method is based upon a thin-shock-layer approximation and has the advantage of being applicable to both subsonic and supersonic regions of the shock layer. The relative simplicity of the method makes it ideally suited for programming on a digital computer with a significant reduction in storage capacity and computing time required by other more exact methods. Comparison of nonequilibrium solutions for an air mixture obtained by the present method is made with solutions obtained by two other methods. Additional cases are presented for entry of spherical nose cones into representative Venusian and Matrian atmospheres. A digital computer program written in FORTRAN language is presented that permits an arbitrary gas mixture to be employed in the solution. The effects of vibration, dissociation, recombination, electronic excitation, and ionization are included in the program.
NASA Astrophysics Data System (ADS)
Alexandrov, S. V.; Vaganov, A. V.; Shalaev, V. I.
2016-10-01
Processes of vortex structures formation and they interactions with the boundary layer in the hypersonic flow over delta wing with blunted leading edges are analyzed on the base of experimental investigations and numerical solutions of Navier-Stokes equations. Physical mechanisms of longitudinal vortexes formation, appearance of abnormal zones with high heat fluxes and early laminar turbulent transition are studied. These phenomena were observed in many high-speed wind tunnel experiments; however they were understood only using the detailed analysis of numerical modeling results with the high resolution. Presented results allowed explaining experimental phenomena. ANSYS CFX code (the DAFE MIPT license) on the grid with 50 million nodes was used for the numerical modeling. The numerical method was verified by comparison calculated heat flux distributions on the wing surface with experimental data.
NASA Technical Reports Server (NTRS)
Seiff, Alvin; Whiting, Ellis E.
1961-01-01
A method by which known bow-wave profiles may be analyzed to give the flow fields around blunt-nosed cylinders in axial hypersonic flow is presented. In the method, the assumption is made that the pressure distribution curve in a transverse plane is similar to that given by blast- wave theory. Numerical analysis based on the one-dimensional energy and continuity equations then leads to distributions of all the flow variables in the cross section, for either a perfect gas or a real gas. The entire flow field need not be solved. Attention can be confined to any desired station. The critical question is the validity of the above assumption. It is tested for the case of a hemisphere cylinder in flight at 20,000 ft/sec. The flow is analyzed for three stations along the cylindrical afterbody, and found to compare very closely with the results of an exact (inviscid) solution. The assumed form of the pressure distribution occurs at stations as close as 1.2 diameters to the body nose. However, it is suggested that the assumption may not apply this far forward in general, particularly when bodies of nonsmooth contour are considered.
Experimental aerothermodynamic research of hypersonic aircraft
NASA Technical Reports Server (NTRS)
Cleary, Joseph W.
1990-01-01
Wind tunnel tests were conducted to establish a benchmark experimental data base for a genetic hypersonic aircraft vehicle. Comprehensive measurements were made at Mach 7 to give flow visualization, surface pressure, surface convective heat transfer, and flow field Pitot pressure for a delta platform all-body vehicle. The tests were conducted in the NASA/Ames 3.5-Foot Hypersonic Wind Tunnel at Reynolds numbers sufficient to give turbulent flow. Comparisons are made of the experimental results with computational solutions of the flow by an upwind parabolized Navier-Stokes code developed at Ames. Good agreement of experiment with solutions by the code is demonstrated.
Review of Rarefied Gas Effects in Hypersonic Applications
2011-01-01
nonequilibrium gas flows are discussed, and several numerical methods for rarefied flow simulation are outlined. A variety of hypersonic flow ...path of the gas . If the Knudsen number is used to characterize the degree of rarefaction effects, then we can expect rarefied flow to result from...194 14. ABSTRACT Rarefied gas phenomena are found in a wide variety of hypersonic flow applications, and accurate characterization of such phenomena
NASA Astrophysics Data System (ADS)
Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B.; Oliveira, A. C.; Gomes, F. A. A.; Myrabo, L. N.; Nagamatsu, Henry T.
2008-04-01
The new 0.60-m. nozzle exit diameter hypersonic shock tunnel was designed to study advanced air-breathing propulsion system such as supersonic combustion and/or laser technologies. In addition, it may be used for hypersonic flow studies and investigations of the electromagnetic (laser) energy addition for flow control. This new hypersonic shock tunnel was designed and installed at the Laboratory for of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu, IEAv-CTA, Brazil. The design of the tunnel enables relatively long test times, 2-10 milliseconds, suitable for the experiments performed at the laboratory. Free stream Mach numbers ranging from 6 to 25 can be produced and stagnation pressures and temperatures up to 360 atm. and up to 9,000 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization.
Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B. Jr; Oliveira, A. C.; Gomes, F. A. A.; Myrabo, L. N.; Nagamatsu, Henry T.
2008-04-28
The new 0.60-m. nozzle exit diameter hypersonic shock tunnel was designed to study advanced air-breathing propulsion system such as supersonic combustion and/or laser technologies. In addition, it may be used for hypersonic flow studies and investigations of the electromagnetic (laser) energy addition for flow control. This new hypersonic shock tunnel was designed and installed at the Laboratory for of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu, IEAv-CTA, Brazil. The design of the tunnel enables relatively long test times, 2-10 milliseconds, suitable for the experiments performed at the laboratory. Free stream Mach numbers ranging from 6 to 25 can be produced and stagnation pressures and temperatures up to 360 atm. and up to 9,000 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization.
Internal convective cooling systems for hypersonic aircraft
NASA Technical Reports Server (NTRS)
Anthony, F. M.; Dukes, W. H.; Helenbrook, R. G.
1975-01-01
Parametric studies were conducted to investigate the relative merits of construction materials, coolants, and cooled panel concepts for internal convective cooling systems applied to airframe structures of hydrogen-fueled hypersonic aircraft. These parametric studies were then used as a means of comparing various cooled structural arrangements for a hypersonic transport and a hypersonic research airplane. The cooled airplane studies emphasized weight aspects as related to the choice of materials, structural arrangements, structural temperatures, and matching of the cooling system heat load to the available hydrogen fuel-flow heat sink. Consideration was given to reliability and to fatigue and fracture aspects, as well. Even when auxiliary thermal protection system items such as heat shielding, insulation, and excess hydrogen for cooling are considered the more attractive actively cooled airframe concepts indicated potential payload increases of from 40 percent to over 100 percent for the hypersonic transport as compared to the results of previous studies of the same vehicle configuration with an uncooled airframe.
Csernai, L. P.; Magas, V. K.; Stoecker, H.; Strottman, D. D.
2011-08-15
Substantial collective flow is observed in collisions between lead nuclei at Large Hadron Collider (LHC) as evidenced by the azimuthal correlations in the transverse momentum distributions of the produced particles. Our calculations indicate that the global v{sub 1}-flow, which at RHIC peaked at negative rapidities (named third flow component or antiflow), now at LHC is going to turn toward forward rapidities (to the same side and direction as the projectile residue). Potentially this can provide a sensitive barometer to estimate the pressure and transport properties of the quark-gluon plasma. Our calculations also take into account the initial state center-of-mass rapidity fluctuations, and demonstrate that these are crucial for v{sub 1} simulations. In order to better study the transverse momentum flow dependence we suggest a new ''symmetrized''v{sub 1}{sup S}(p{sub t}) function, and we also propose a new method to disentangle global v{sub 1} flow from the contribution generated by the random fluctuations in the initial state. This will enhance the possibilities of studying the collective Global v{sub 1} flow both at the STAR Beam Energy Scan program and at LHC.
Hypersonic combustion of hydrogen in a shock tunnel
NASA Technical Reports Server (NTRS)
Morgan, R. G.; Stalker, R. J.
1989-01-01
Results are reported on shock-tunnel experiments testing the feasibility of hypersonic combustion and thrust generation in a hydrogen scramjet model. Tests with a constant-area duct show that hypersonic combustion is possible with a central injection at static intake pressures of about 20 kPa. The results of a comparison made between model configurations with nominal combustion-chamber intake Mach numbers of 4 and 6 indicated that the hypersonic duct gives a better performance at flight enthalpies above 7 mJ/kg. It is argued that the lower temperatures associated with hypersonic flow produce more efficient combustion.
Reddy Tube Driven Table-Top Hypersonic Shock Tunnel
NASA Astrophysics Data System (ADS)
Reddy, K. P. J.; Ramesh Babu, R.; Murali, R.; Saravanan, S.
A table-top hypersonic shock tunnel driven by a manually operated shock tube of 29 mm diameter, named Reddy tube, is presented. The shock tunnel is capable of generating hypersonic flow of freestream Mach number 5.5 with a test time of 600 microseconds. Measurement of heat transfer rates for a small model is described. The measured stagnation point heat transfer rates of ~10 W/cm2 match with the theoretically estimated values within the experimental errors. The simple table-top hypersonic shock tunnel presented here is ideally suited for the class room experiments in the hypersonic aerodynamics course.
NASA Astrophysics Data System (ADS)
Shen, Chun; Qiu, Zhi; Heinz, Ulrich
2015-07-01
In ultracentral heavy-ion collisions, anisotropic hydrodynamic flow is generated by density fluctuations in the initial state rather than by geometric overlap effects. For a given centrality class, the initial fluctuation spectrum is sensitive to the method chosen for binning the events into centrality classes. We show that sorting events by total initial entropy or by total final multiplicity yields event classes with equivalent statistical fluctuation properties, in spite of viscous entropy production during the fireball evolution. With this initial entropy-based centrality definition we generate several classes of ultracentral Pb + Pb collisions at Cern Large Hadron Collider energies and evolve the events using viscous hydrodynamics with nonzero shear but vanishing bulk viscosity. Comparing the predicted anisotropic flow coefficients for charged hadrons with CMS data we find that both the Monte Carlo Glauber (MC-Glb) and Monte Carlo Kharzeev-Levin-Nardi (MC-KLN) models produce initial fluctuation spectra that are incompatible with the measured final anisotropic flow power spectrum, for any choice of the specific shear viscosity. In spite of this failure, we show that the hydrodynamic model can qualitatively explain, in terms of event-by-event fluctuations of the anisotropic flow coefficients and flow angles, the breaking of flow factorization for elliptic, triangular, and quadrangular flow measured by the CMS experiment. For elliptic flow, this factorization breaking is large in ultracentral collisions. We conclude that the bulk of the experimentally observed flow factorization breaking effects are qualitatively explained by hydrodynamic evolution of initial-state fluctuations, but that their quantitative description requires a better understanding of the initial fluctuation spectrum.
A numerical method for predicting hypersonic flowfields
NASA Technical Reports Server (NTRS)
Maccormack, Robert W.; Candler, Graham V.
1988-01-01
The flow about a body traveling at hypersonic speed is energetic enough to cause the atmospheric gases to react chemically and reach states in thermal nonequilibrium. In this paper, a new procedure based on Gauss-Seidel line relaxation is shown to solve the equations of hypersonic flow fields containing finite reaction rate chemistry and thermal nonequilibrium. The method requires a few hundred time steps and small computer times for axisymmetric flows about simple body shapes. The extension to more complex two-dimensional body geometries appears straightforward.
A numerical method for predicting hypersonic flowfields
NASA Technical Reports Server (NTRS)
Maccormack, Robert W.; Candler, Graham V.
1988-01-01
The flow about a body traveling at hypersonic speed is energetic enough to cause the atmospheric gases to react chemically and reach states in thermal nonequilibrium. In this paper, a new procedure based on Gauss-Seidel line relaxation is shown to solve the equations of hypersonic flow fields containing finite reaction rate chemistry and thermal nonequilibrium. The method requires a few hundred time steps and small computer times for axisymmetric flows about simple body shapes. The extension to more complex two-dimensional body geometries appears straightforward.
Unstart coupling mechanism analysis of multiple-modules hypersonic inlet.
Hu, Jichao; Chang, Juntao; Wang, Lei; Cao, Shibin; Bao, Wen
2013-01-01
The combination of multiplemodules in parallel manner is an important way to achieve the much higher thrust of scramjet engine. For the multiple-modules scramjet engine, when inlet unstarted oscillatory flow appears in a single-module engine due to high backpressure, how to interact with each module by massflow spillage, and whether inlet unstart occurs in other modules are important issues. The unstarted flowfield and coupling characteristic for a three-module hypersonic inlet caused by center module II and side module III were, conducted respectively. The results indicate that the other two hypersonic inlets are forced into unstarted flow when unstarted phenomenon appears on a single-module hypersonic inlet due to high backpressure, and the reversed flow in the isolator dominates the formation, expansion, shrinkage, and disappearance of the vortexes, and thus, it is the major factor of unstart coupling of multiple-modules hypersonic inlet. The coupling effect among multiple modules makes hypersonic inlet be more likely unstarted.
Hypersonic gasdynamic laser system
Foreman, K.M.; Maciulaitis, A.
1990-05-22
This patent describes a visible, or near to mid infra-red, hypersonic gas dynamic laser system. It comprises: a hypersonic vehicle for carrying the hypersonic gas dynamic laser system, and also providing high energy ram air for thermodynamic excitation and supply of the laser gas; a laser cavity defined within the hypersonic vehicle and having a laser cavity inlet for the laser cavity formed by an opening in the hypersonic vehicle, such that ram air directed through the laser cavity opening supports gas dynamic lasing operations at wavelengths less than 10.6{mu} meters in the laser cavity; and an optical train for collecting the laser radiation from the laser cavity and directing it as a substantially collimated laser beam to an output aperture defined by an opening in the hypersonic vehicle to allow the laser beam to be directed against a target.
Aerothermodynamic shape optimization of hypersonic blunt bodies
NASA Astrophysics Data System (ADS)
Eyi, Sinan; Yumuşak, Mine
2015-07-01
The aim of this study is to develop a reliable and efficient design tool that can be used in hypersonic flows. The flow analysis is based on the axisymmetric Euler/Navier-Stokes and finite-rate chemical reaction equations. The equations are coupled simultaneously and solved implicitly using Newton's method. The Jacobian matrix is evaluated analytically. A gradient-based numerical optimization is used. The adjoint method is utilized for sensitivity calculations. The objective of the design is to generate a hypersonic blunt geometry that produces the minimum drag with low aerodynamic heating. Bezier curves are used for geometry parameterization. The performances of the design optimization method are demonstrated for different hypersonic flow conditions.
Pratt, Scott; Schlichting, Soeren; Gavin, Sean
2011-08-15
Correlations of azimuthal angles observed at the Relativistic Heavy Ion Collider have gained great attention due to the prospect of identifying fluctuations of parity-odd regions in the field sector of QCD. Whereas the observable of interest related to parity fluctuations involves subtracting opposite-sign from same-sign correlations, the STAR collaboration reported the same-sign and opposite-sign correlations separately. It is shown here how momentum conservation combined with collective elliptic flow contributes significantly to this class of correlations, although not to the difference between the opposite- and same-sign observables. The effects are modeled with a crude simulation of a pion gas. Although the simulation reproduces the scale of the correlation, the centrality dependence is found to be sufficiently different in character to suggest additional considerations beyond those present in the pion gas simulation presented here.
Chattopadhyay, S.
1994-11-01
The motivation, feasibility and potential for two unconventional collider concepts - the Gamma-Gamma Collider and the Muon Collider - are described. The importance of the development of associated technologies such as high average power, high repetition rate lasers and ultrafast phase-space techniques are outlined.
Hypersonic Aerodynamics Fellowships
1991-02-11
J.D. and Capriotti , D., "Viscous Optimized Hypersonic Waveriders," AIAA Paper No. 87-0272. Appendix B: Corpening, G. and Anderson,J., "Numerical...Hypersonic Waveriders K. G. Bowcutt, J. D. Anderson and D. Capriotti , Univ. of Maryland, College Park, MD AIAA 25th Aerospace Sciences Meeting January...OPTIMIZED HYPERSONIC WAVERIDERS by Kevin G. Rowcutt,* John D. Anderson, Jr.," and Diego Capriotti "** Department of Aerospace Engineering University of
The optimum hypersonic wind tunnel
NASA Technical Reports Server (NTRS)
Trimmer, L. L.; Cary, A., Jr.; Voisinet, R. L. P.
1986-01-01
The capabilities of existing hypersonic wind tunnels in the U.S. are assessed to form a basis for recommendations for a new, costly facility which would provide data for modeling the hypervelocity aerodynamics envisioned for the new generation of aerospace vehicles now undergoing early studies. Attention is given to the regimes, both entry and aerodynamic, which the new vehicles will encounter, and the shortcomings of data generated for the Orbiter before flight are discussed. The features of foreign-gas, impulse, aeroballistic range, arc-heated and combustion-heated facilities are examined, noting that in any hypersonic wind tunnel the flow must be preheated to prevent liquefaction upon expansion in the test channel. The limitations of the existing facilities and the identification of the regimes which must be studied lead to a description of the characteristics of an optimum hypersonic wind tunnel, including the operations and productivity, the instrumentation, the nozzle design and the flow quality. Three different design approaches are described, each costing at least $100 million to achieve workability.
The optimum hypersonic wind tunnel
NASA Technical Reports Server (NTRS)
Trimmer, L. L.; Cary, A., Jr.; Voisinet, R. L. P.
1986-01-01
The capabilities of existing hypersonic wind tunnels in the U.S. are assessed to form a basis for recommendations for a new, costly facility which would provide data for modeling the hypervelocity aerodynamics envisioned for the new generation of aerospace vehicles now undergoing early studies. Attention is given to the regimes, both entry and aerodynamic, which the new vehicles will encounter, and the shortcomings of data generated for the Orbiter before flight are discussed. The features of foreign-gas, impulse, aeroballistic range, arc-heated and combustion-heated facilities are examined, noting that in any hypersonic wind tunnel the flow must be preheated to prevent liquefaction upon expansion in the test channel. The limitations of the existing facilities and the identification of the regimes which must be studied lead to a description of the characteristics of an optimum hypersonic wind tunnel, including the operations and productivity, the instrumentation, the nozzle design and the flow quality. Three different design approaches are described, each costing at least $100 million to achieve workability.
NASA Technical Reports Server (NTRS)
Riley, Christopher J.
1992-01-01
An engineering method has been developed that couples an approximate three dimensional inviscid technique with the axisymmetric analog and a set of approximate convective heating equations. The displacement effect on the boundary layer on the outer inviscid flow is calculated and included as a boundary condition in the inviscid technique. This accounts for the viscous interaction present at lower Reynolds numbers. The method is applied to blunted axisymmetric and three dimensional elliptic cones at angle of attack for the laminar hypersonic flow of a perfect gas. The method is applied to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates, pressures, and shock shapes that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes and viscous shock-layer equations. In addition, the inclusion of viscous interaction significantly improves results obtained at lower Reynolds numbers. The new technique represents a major improvement over current engineering aerothermal methods with only a modest increase in computational effort.
NASA Astrophysics Data System (ADS)
Schweigert, Irina
2013-09-01
Recently the problem of communication blackout during reentrant flight still remains unsolved. The spacecrafts enter the upper atmospheric layers with a hypersonic speed and the shock heated air around them becomes weakly ionized. The gas ionization behind the shock front is associative in nature and occurs through chemical reactions between fragments of molecules. The formation of a plasma layer near the surfaces of spacecraft causes serious problems related to the blocking of communication channels with the Earth and other spacecrafts. A promising way of restoring the radio communications is the application of electrical and magnetic fields for controlling the plasma layer parameters. Nevertheless the flux of electrons and ions on the surface charges it that essentially decrease the effect of electro-magnetic control of local plasma density. In Ref. it is shown that there is the way to remove the surface charge using the lateral diode string structures. Based on two dimensional kinetic Particle in cell Monte Carlo collision simulations, we study the possibility of local control the plasma layer parameters near a flat surface of two different types. The gas velocity distribution is set with a model profile. We apply DC voltage up to 4 kV and magnetic field B up to 200 G.
Research and educational initiatives at the Syracuse University Center for Hypersonics
NASA Technical Reports Server (NTRS)
Spina, E.; Lagraff, J.; Davidson, B.; Bogucz, E.; Dang, T.
1995-01-01
The Department of Mechanical, Aerospace, and Manufacturing Engineering and the Northeast Parallel Architectures Center of Syracuse University have been funded by NASA to establish a program to educate young engineers in the hypersonic disciplines. This goal is being achieved through a comprehensive five-year program that includes elements of undergraduate instruction, advanced graduate coursework, undergraduate research, and leading-edge hypersonics research. The research foci of the Syracuse Center for Hypersonics are three-fold; high-temperature composite materials, measurements in turbulent hypersonic flows, and the application of high-performance computing to hypersonic fluid dynamics.
Research and educational initiatives at the Syracuse University Center for Hypersonics
NASA Technical Reports Server (NTRS)
Spina, E.; Lagraff, J.; Davidson, B.; Bogucz, E.; Dang, T.
1995-01-01
The Department of Mechanical, Aerospace, and Manufacturing Engineering and the Northeast Parallel Architectures Center of Syracuse University have been funded by NASA to establish a program to educate young engineers in the hypersonic disciplines. This goal is being achieved through a comprehensive five-year program that includes elements of undergraduate instruction, advanced graduate coursework, undergraduate research, and leading-edge hypersonics research. The research foci of the Syracuse Center for Hypersonics are three-fold; high-temperature composite materials, measurements in turbulent hypersonic flows, and the application of high-performance computing to hypersonic fluid dynamics.
Whole-field velocity measurements of flow around colliding barchan dunes
NASA Astrophysics Data System (ADS)
Bristow, N.; Blois, G.; Kim, T.; Best, J.; Christensen, K. T.
2015-12-01
Barchan dunes are crescentic bedforms located in environments with unidirectional flow and limited sediment supply, including deserts, river beds, the continental shelves and the craters of Mars. Barchans are commonly observed in fields rather than in isolation, with the evolution of, and interactions between, bedforms being highly dynamic, involving feedback mechanisms between the fluid flow, morphological change and sediment transport. A series of experiments were undertaken to discretely simulate the collision of a smaller barchan with a larger, downstream one using fixed bedform models, with each experiment representing a successive snapshot in the dune collision process. These experiments thus capture the turbulent flow over fixed-bed morphologies that correlate with rapid morphological change and high rates of sediment transport using time-resolved particle image velocimetry (PIV) in the wall-parallel plane. The use of a Refractive Index Matching (RIM) flow facility allows for the light to pass through the model, capturing areas which are otherwise obscured, such as around the horns of the dune, the sheltered region behind the crest, and areas in which the bedforms are deformed during the collision. This paper will present the results of Dynamic Mode Decomposition that has been used to identify the most dominant modes contributing to flow dynamics in each collision stage.
Hypersonic missile propulsion system
Kazmar, R.R.
1998-11-01
Pratt and Whitney is developing the technology for hypersonic components and engines. A supersonic combustion ramjet (scramjet) database was developed using hydrogen fueled propulsion systems for space access vehicles and serves as a point of departure for the current development of hydrocarbon scramjets. The Air Force Hypersonic Technology (HyTech) Program has put programs in place to develop the technologies necessary to demonstrate the operability, performance and structural durability of an expendable, liquid hydrocarbon fueled scramjet system that operates from Mach 4 to 8. This program will culminate in a flight type engine test at representative flight conditions. The hypersonic technology base that will be developed and demonstrated under HyTech will establish the foundation to enable hypersonic propulsion systems for a broad range of air vehicle applications from missiles to space access vehicles. A hypersonic missile flight demonstration is planned in the DARPA Affordable Rapid Response Missile Demonstrator (ARRMD) program in 2001.
NASA Technical Reports Server (NTRS)
Midden, Raymond E.; Miller, Charles G., III
1985-01-01
The Langley Hypersonic CF4 Tunnel is a Mach 6 facility which simulates an important aspect of dissociative real-gas phenomena associated with the reentry of blunt vehicles, i.e., the decrease in the ratio of specific heats (gamma) that occurs within the shock layer of the vehicle. A general description of this facility is presented along with a discussion of the basic components, instrumentation, and operating procedure. Pitot-pressure surveys were made at the nozzle exit and downstream of the exit for reservoir temperatures from 1020 to 1495 R and reservoir pressures from 1000 to 2550 psia. A uniform test core having a diameter of circa 11 in. (0.55 times the nozzle-exit diameter) exists at the maximum value of reservoir pressure and temperature. The corresponding free-stream Mach number is 5.9, the unit Reynolds number is 4 x 10 to the 5th power per foot, the ratio of specific heats immediately behind a normal shock is 1.10, and the normal-shock density ratio is 12.6. When the facility is operated at reservoir temperatures below 1440 R, irregularities occur in the pitot-pressure profile within a small region about the nozzle centerline. These variations in pitot pressure indicate the existence of flow distrubances originating in the upstream region of the nozzle. This necessitates testing models off centerline in the uniform flow between the centerline region and either the nozzle boundary layer or the lip shock originating at the nozzle exit. Samples of data obtained in this facility with various models are presented to illustrate the effect of gamma on flow conditions about the model and the importance of knowing the magnitude of this effect.
Hypersonic turbulent wall boundary layer computations
NASA Astrophysics Data System (ADS)
Kim, S. C.; Harloff, G. J.
1988-05-01
The Baldwin-Lomax algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer layer eddy viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith model for a flat plate at hypersonic speed, the new values of the coefficient were obtained. The results show that the values of C sub cp and C sub kleb are functions of both Mach number and wall temperature ratio. The C sub cp and C sub kleb variations with Mach number and wall temperature were used for the calculations of both a 4 deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.
Hypersonic turbulent wall boundary layer computations
NASA Technical Reports Server (NTRS)
Kim, S. C.; Harloff, G. J.
1988-01-01
The Baldwin-Lomax (1978) algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer-layer eddy-viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith (1974) model for a flat plate at hypersonic speed, the new values of the coefficients were obtained. The results show that the values of C(cp) and C(kleb) are functions of both Mach number and wall temperature ratio. The C(cp) and C(kleb) variations with Mach number and wall temperature were used for the calculations of both a 4-deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin-layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.
Hypersonic turbulent wall boundary layer computations
NASA Technical Reports Server (NTRS)
Kim, S. C.; Harloff, G. J.
1988-01-01
The Baldwin-Lomax algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer layer eddy viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith model for a flat plate at hypersonic speed, the new values of the coefficient were obtained. The results show that the values of C sub cp and C sub kleb are functions of both Mach number and wall temperature ratio. The C sub cp and C sub kleb variations with Mach number and wall temperature were used for the calculations of both a 4 deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.
Palmer, R.B. |; Sessler, A.; Skrinsky, A.
1996-01-01
Muon Colliders have unique technical and physics advantages and disadvantages when compared with both hadron and electron machines. They should thus be regarded as complementary. Parameters are given of 4 TeV and 0.5 TeV high luminosity {micro}{sup +}{micro}{sup {minus}}colliders, and of a 0.5 TeV lower luminosity demonstration machine. We discuss the various systems in such muon colliders, starting from the proton accelerator needed to generate the muons and proceeding through muon cooling, acceleration and storage in a collider ring. Problems of detector background are also discussed.
NASA Astrophysics Data System (ADS)
Palmer, R. B.; Sessler, A.; Skrinsky, A.; Tollestrup, A.; Baltz, A. J.; Chen, P.; Cheng, W.-H.; Cho, Y.; Courant, E.; Fernow, R. C.; Gallardo, J. C.; Garren, A.; Green, M.; Kahn, S.; Kirk, H.; Lee, Y. Y.; Mills, F.; Mokhov, N.; Morgan, G.; Neuffer, D.; Noble, R.; Norem, J.; Popovic, M.; Schachinger, L.; Silvestrov, G.; Summers, D.; Stumer, I.; Syphers, M.; Torun, Y.; Trbojevic, D.; Turner, W.; Van Ginneken, A.; Vsevolozhskaya, T.; Weggel, R.; Willen, E.; Winn, D.; Wurtele, J.
1996-05-01
Muon Colliders have unique technical and physics advantages and disadvantages when compared with both hadron and electron machines. They should thus be regarded as complementary. Parameters are given of 4 TeV and 0.5 TeV high luminosity μ+μ- colliders, and of a 0.5 TeV lower luminosity demonstration machine. We discuss the various systems in such muon colliders, starting from the proton accelerator needed to generate the muons and proceeding through muon cooling, acceleration and storage in a collider ring. Problems of detector background are also discussed.
NASA Technical Reports Server (NTRS)
Kirk, Benjamin S.; Bova, Stephen W.; Bond, Ryan B.
2011-01-01
Presentation topics include background and motivation; physical modeling including governing equations and thermochemistry; finite element formulation; results of inviscid thermal nonequilibrium chemically reacting flow and viscous thermal equilibrium chemical reacting flow; and near-term effort.
2013-07-11
for nitrogen using molecular dynamics simulation”, 28th International Symposium on Rarefied Gas Dynamics, AIP Conf. Proc. 1501, 519-526 (2012); doi...accurate for highly nonequilibrium flows relevant to rarefied flows and sharp flow features with small length scales. Currently, both CFD and DSMC use...While such simulations are not expected to overlap with the 3D near-continuum flows in the near future, they certainly overlap with rarefied DSMC
Discrete Particle Simulation Techniques for the Analysis of Colliding and Flowing Particulate Media
NASA Astrophysics Data System (ADS)
Mukherjee, Debanjan
Flowing particulate media are ubiquitous in a wide spectrum of applications that include transport systems, fluidized beds, manufacturing and materials processing technologies, energy conversion and propulsion technologies, sprays, jets, slurry flows, and biological flows. The discrete nature of the media, along with their underlying coupled multi-physical interactions can lead to a variety of interesting phenomena, many of which are unique to such media - for example, turbulent diffusion and preferential concentration in particle laden flows, and soliton like excitation patterns in a vibrated pile of granular material. This dissertation explores the utility of numerical simulations based on the discrete element method and collision driven particle dynamics methods for analyzing flowing particulate media. Such methods are well-suited to handle phenomena involving particulate, granular, and discontinuous materials, and often provide abilities to tackle complicated physical phenomena, for which pursuing continuum based approaches might be difficult or sometimes insufficient. A detailed discussion on hierarchically representing coupled, multi-physical phenomena through simple models for underlying physical interactions is presented. Appropriate physical models for mechanical contact, conductive and convective heat exchange, fluid-particle interactions, adhesive and near-field effects, and interaction with applied electromagnetic fields are presented. Algorithmic details on assembling the interaction models into a large-scale simulation framework have been elaborated with illustrations. The assembled frameworks were used to develop a computer simulation library (named `Software Library for Discrete Element Simulations' (SLIDES) for the sake of reference and continued future development efforts) and aspects of the architecture and development of this library have also been addressed. This is an object-oriented discrete particle simulation library developed in Fortran
Advanced optical diagnostics in hypersonic research
NASA Astrophysics Data System (ADS)
Cattolica, Robert J.
1988-10-01
The renewed emphasis on hypersonic research has stimulated a resurgence of interest in experimental methods for the study of high-speed flows. Improvement in the physical and chemical models used in computational fluid dynamic simulation of hypersonic flows requires a modern experimental data base. Optical diagnostics provide the capability to make nonintrusive measurements of density, temperature, velocity, and species concentration in hypersonic flows. The short test time available in hypersonic wind tunnels or flight experiments necessitates spectroscopic methods capable of producing high signal levels. Fluorescence methods based on laser or electron-beam excitation satisfy this requirement. For flight experiments, electron-beam excitation offers a number of advantages over laser excitation that include small device size, high electrical efficiency, and multiple-state and species-selective excitation. Disadvantages of the electron beam fluorescence (EBF) technique included a complex excitation mechanism and some limitations in high-density applications. Laser fluorescence methods (LIF) have been developed extensively in recent years for combustion research, but need further advances in miniaturization of lasers for application to in-flight hypersonic combustion and aerodynamic experiments. Both techniques require a fundamental understanding of the complications introduced by physical effects such as energy transfer and quenching of the fluorescence signal. With modern electro-optic instrumentation it is now possible to examine in detail the influence of these phenomena on EBF and LIF fluorescence spectra in the laboratory and to extend these measurement techniques for use in flight research. To illustrate some of the research required to develop these methods to address issues relevent to hypersonic flight, examples of experiments on the use of EBF and LIF spectroscopy for the measurement of nitric oxide concentration are presented.
NASA Astrophysics Data System (ADS)
Graham, Michael
2015-11-01
Blood is a suspension of objects of various shapes, sizes and mechanical properties, whose distribution during flow is important in many contexts. Red blood cells tend to migrate toward the center of a blood vessel, leaving a cell-free layer at the vessel wall, while white blood cells and platelets are preferentially found near the walls, a phenomenon called margination that is critical for the physiological responses of inflammation and hemostasis. Additionally, drug delivery particles in the bloodstream will also undergo segregation - the influence of these phenomena on the efficacy of such particles is unknown. This talk describes efforts to gain a systematic understanding of flow-induced segregation phenomena in blood and other complex mixtures, using a combination of theory and direct simulations. Contrasts in size, deformability and shape can all lead to segregation. A kinetic theory model based on pair collisions and wall-induced hydrodynamic migration can capture the key effects observed in direct simulations, including a ``drainage transition'' in which one component is completely depleted from the bulk of the flow. Experiments performed in the laboratory of Wilbur Lam indicate the physiological and clinical importance of these observations. This talk is based upon work supported by the National Science Foundation under Grants No. CBET- 1132579 and No. CBET-1436082.
Preliminary aerothermodynamic design method for hypersonic vehicles
NASA Technical Reports Server (NTRS)
Harloff, G. J.; Petrie, S. L.
1987-01-01
Preliminary design methods are presented for vehicle aerothermodynamics. Predictions are made for Shuttle orbiter, a Mach 6 transport vehicle and a high-speed missile configuration. Rapid and accurate methods are discussed for obtaining aerodynamic coefficients and heat transfer rates for laminar and turbulent flows for vehicles at high angles of attack and hypersonic Mach numbers.
2013-07-09
rarefied gas flow simulations by the DSMC method,” Phys. Fluids...Symposium on Rarefied Gas Dynamics, AIP Conf. Proc. 1501, 519-526 (2012); doi: 10.1063/1.4769583 Valentini, P., Zhang, C., and Schwartzentruber, T.E...method [1], which simulates the Boltzmann equation [2] and is therefore accurate for highly nonequilibrium flows relevant to rarefied flows and
NASA Technical Reports Server (NTRS)
Hefner, J. N.
1972-01-01
The lee-surface flow phenomena on a delta-wing orbiter and a straight-wing orbiter have been investigated at angles of attack between 0 deg and 50 deg at a Mach number of 6. Limited studies of the delta-wing orbiter were conducted at a Mach number of 19. Heat-transfer data, pressure distributions, and oil-flow studies were employed to experimentally examine the nature of the surface flow and the severity of the lee-surface heating. The effects of Reynolds number on the flow field and heating were investigated. Problem areas are defined and areas for further study are recommended.
Gorishnyy, T; Ullal, C K; Maldovan, M; Fytas, G; Thomas, E L
2005-03-25
In this Letter we propose the use of hypersonic phononic crystals to control the emission and propagation of high frequency phonons. We report the fabrication of high quality, single crystalline hypersonic crystals using interference lithography and show that direct measurement of their phononic band structure is possible with Brillouin light scattering. Numerical calculations are employed to explain the nature of the observed propagation modes. This work lays the foundation for experimental studies of hypersonic crystals and, more generally, phonon-dependent processes in nanostructures.
1955-01-01
8217rinRE-DifMENSONAL HtYPERtSONIC 15.W indicated-flow-separation oin the leewardl side of (lie body for excellent agreemelnt in tlie plano of symmlletry...REIMARKS b~ound~ary layers may, inl like imanner, prove useful il- pie - A mnethod of characteristics employing p)ressure and-flow deigdrednesoa
Experimental investigation of hypersonic aerodynamics
NASA Technical Reports Server (NTRS)
Heinemann, K.; Intrieri, Peter F.
1987-01-01
An extensive series of ballistic range tests are currently being conducted at the Ames Research Center. These tests are intended to investigate the hypersonic aerodynamic characteristics of two basic configurations, which are: the blunt-cone Galileo probe which is scheduled to be launched in late 1989 and will enter the atmosphere of Jupiter in 1994, and a generic slender cone configuration to provide experimental aerodynamic data including good flow-field definition which computational aerodynamicists could use to validate their computer codes. Some of the results obtained thus far are presented and work for the near future is discussed.
Galactic scale gas flows in colliding galaxies: 3-dimensional, N-body/hydrodynamics experiments
NASA Technical Reports Server (NTRS)
Lamb, Susan A.; Gerber, Richard A.; Balsara, Dinshaw S.
1994-01-01
We present some results from three dimensional computer simulations of collisions between models of equal mass galaxies, one of which is a rotating, disk galaxy containing both gas and stars and the other is an elliptical containing stars only. We use fully self consistent models in which the halo mass is 2.5 times that of the disk. In the experiments we have varied the impact parameter between zero (head on) and 0.9R (where R is the radius of the disk), for impacts perpendicular to the disk plane. The calculations were performed on a Cray 2 computer using a combined N-body/smooth particle hydrodynamics (SPH) program. The results show the development of complicated flows and shock structures in the direction perpendicular to the plane of the disk and the propagation outwards of a density wave in both the stars and the gas. The collisional nature of the gas results in a sharper ring than obtained for the star particles, and the development of high volume densities and shocks.
NASA Technical Reports Server (NTRS)
Hunt, J. L.
1973-01-01
Data are presented from a series of phase-change heat transfer and flow visualization tests at Mach 7.4, 8, and 10.3 in air, Mach 19.5 in nitrogen, Mach 20.3 in helium, and Mach 6 in tetrafluoromethane (CF4) on the windward surface of a straight wing hypersonic reentry configuration for angles of attack from 20 deg to 80 deg. The results indicate that: (1) for hypersonic stream Mach numbers, the flow field over the straight-wing configuration is essentially independent of Mach number, (2) transition Reynolds number decreases with increasing angle of attack, (3) at some critical angle of attack, the wing-shock standoff distance is greatly increased and the stagnation line moves downstream from the wing leading edge, (4) value of the critical angle of attack is very sensitive to the flow shock density ratio or effective gamma, and (5) at angles of attack above the critical value for all gases, the nondimensional level of heat transfer to the wing is higher for the higher shock density ratio flows.
NASA Technical Reports Server (NTRS)
Glass, Christopher E.
2000-01-01
An uncoupled Computational Fluid Dynamics-Direct Simulation Monte Carlo (CFD-DSMC) technique is developed and applied to provide solutions for continuum jets interacting with rarefied external flows. The technique is based on a correlation of the appropriate Bird breakdown parameter for a transitional-rarefied condition that defines a surface within which the continuum solution is unaffected by the external flow-jet interaction. The method is applied to two problems to assess and demonstrate its validity; one of a jet interaction in the transitional-rarefied flow regime and the other in the moderately rarefied regime. Results show that the appropriate Bird breakdown surface for uncoupling the continuum and non-continuum solutions is a function of a non-dimensional parameter relating the momentum flux and collisionality between the two interacting flows. The correlation is exploited for the simulation of a jet interaction modeled for an experimental condition in the transitional-rarefied flow regime and the validity of the correlation is demonstrated. The uncoupled technique is also applied to an aerobraking flight condition for the Mars Global Surveyor spacecraft with attitude control system jet interaction. Aerodynamic yawing moment coefficients for cases without and with jet interaction at various angles-of-attack were predicted, and results from the present method compare well with values published previously. The flow field and surface properties are analyzed in some detail to describe the mechanism by which the jet interaction affects the aerodynamics.
Impact of Ion Acoustic Wave Instabilities in the Flow Field of a Hypersonic Vehicle on EM Signals
NASA Astrophysics Data System (ADS)
Mudaliar, Saba; Sotnikov, Vladimir
2016-10-01
Flow associated with a high speed air vehicle (HSAV) can get partially ionized. In the absence of external magnetic field the flow field turbulence is due to ion acoustic wave (IAW) instabilities. Our interest is in studying the impact of this turbulence on the radiation characteristics of EM signals from the HSAV. We decompose the radiated signal into coherent and diffuse parts. We find that the coherent part has the same spectrum as that of the source signal, but it is distorted because of dispersive coherent attenuation. The diffuse part is expressed as a convolution (in wavenumber and frequency) of the source signal with the spectrum of electron density fluctuations. This is a constrained convolution in the sense that the spectrum has to satisfy the IAW dispersion relation. A quantity that characterizes the flow is the mean free path (MFP). When the MFP is large compared to the thickness of the flow the coherent part is significant. If the MFP is larger than the thickness of the flow the diffuse part is the dominant part of the received signal. In the special case when the source signal frequency is close the electron plasma frequency, there can exist in the flow region Langmuir modes in addition to the EM modes. The radiation characteristics of EM source signals from the HSAV in this case are quite different.
A numerical method for predicting hypersonic flowfields
NASA Technical Reports Server (NTRS)
Maccormack, Robert W.; Candler, Graham V.
1989-01-01
The flow about a body traveling at hypersonic speed is energetic enough to cause the atmospheric gases to chemically react and reach states in thermal nonequilibrium. The prediction of hypersonic flowfields requires a numerical method capable of solving the conservation equations of fluid flow, the chemical rate equations for specie formation and dissociation, and the transfer of energy relations between translational and vibrational temperature states. Because the number of equations to be solved is large, the numerical method should also be as efficient as possible. The proposed paper presents a fully implicit method that fully couples the solution of the fluid flow equations with the gas physics and chemistry relations. The method flux splits the inviscid flow terms, central differences of the viscous terms, preserves element conservation in the strong chemistry source terms, and solves the resulting block matrix equation by Gauss Seidel line relaxation.
A numerical method for predicting hypersonic flowfields
NASA Technical Reports Server (NTRS)
Maccormack, Robert W.; Candler, Graham V.
1989-01-01
The flow about a body traveling at hypersonic speed is energetic enough to cause the atmospheric gases to chemically react and reach states in thermal nonequilibrium. The prediction of hypersonic flowfields requires a numerical method capable of solving the conservation equations of fluid flow, the chemical rate equations for specie formation and dissociation, and the transfer of energy relations between translational and vibrational temperature states. Because the number of equations to be solved is large, the numerical method should also be as efficient as possible. The proposed paper presents a fully implicit method that fully couples the solution of the fluid flow equations with the gas physics and chemistry relations. The method flux splits the inviscid flow terms, central differences of the viscous terms, preserves element conservation in the strong chemistry source terms, and solves the resulting block matrix equation by Gauss Seidel line relaxation.
Fischer, W.
2011-12-01
Ion colliders are research tools for high-energy nuclear physics, and are used to test the theory of Quantum Chromo Dynamics (QCD). The collisions of fully stripped high-energy ions create matter of a temperature and density that existed only microseconds after the Big Bang. Ion colliders can reach higher densities and temperatures than fixed target experiments although at a much lower luminosity. The first ion collider was the CERN Intersecting Storage Ring (ISR), which collided light ions [77Asb1, 81Bou1]. The BNL Relativistic Heavy Ion Collider (RHIC) is in operation since 2000 and has collided a number of species at numerous energies. The CERN Large Hadron Collider (LHC) started the heavy ion program in 2010. Table 1 shows all previous and the currently planned running modes for ISR, RHIC, and LHC. All three machines also collide protons, which are spin-polarized in RHIC. Ion colliders differ from proton or antiproton colliders in a number of ways: the preparation of the ions in the source and the pre-injector chain is limited by other effects than for protons; frequent changes in the collision energy and particle species, including asymmetric species, are typical; and the interaction of ions with each other and accelerator components is different from protons, which has implications for collision products, collimation, the beam dump, and intercepting instrumentation devices such a profile monitors. In the preparation for the collider use the charge state Z of the ions is successively increased to minimize the effects of space charge, intrabeam scattering (IBS), charge change effects (electron capture and stripping), and ion-impact desorption after beam loss. Low charge states reduce space charge, intrabeam scattering, and electron capture effects. High charge states reduce electron stripping, and make bending and acceleration more effective. Electron stripping at higher energies is generally more efficient. Table 2 shows the charge states and energies in the
NASA Technical Reports Server (NTRS)
Drozda, Tomasz G.; Quinlan, Jesse R.; Pisciuneri, Patrick H.; Yilmaz, S. Levent
2012-01-01
Significant progress has been made in the development of subgrid scale (SGS) closures based on a filtered density function (FDF) for large eddy simulations (LES) of turbulent reacting flows. The FDF is the counterpart of the probability density function (PDF) method, which has proven effective in Reynolds averaged simulations (RAS). However, while systematic progress is being made advancing the FDF models for relatively simple flows and lab-scale flames, the application of these methods in complex geometries and high speed, wall-bounded flows with shocks remains a challenge. The key difficulties are the significant computational cost associated with solving the FDF transport equation and numerically stiff finite rate chemistry. For LES/FDF methods to make a more significant impact in practical applications a pragmatic approach must be taken that significantly reduces the computational cost while maintaining high modeling fidelity. An example of one such ongoing effort is at the NASA Langley Research Center, where the first generation FDF models, namely the scalar filtered mass density function (SFMDF) are being implemented into VULCAN, a production-quality RAS and LES solver widely used for design of high speed propulsion flowpaths. This effort leverages internal and external collaborations to reduce the overall computational cost of high fidelity simulations in VULCAN by: implementing high order methods that allow reduction in the total number of computational cells without loss in accuracy; implementing first generation of high fidelity scalar PDF/FDF models applicable to high-speed compressible flows; coupling RAS/PDF and LES/FDF into a hybrid framework to efficiently and accurately model the effects of combustion in the vicinity of the walls; developing efficient Lagrangian particle tracking algorithms to support robust solutions of the FDF equations for high speed flows; and utilizing finite rate chemistry parametrization, such as flamelet models, to reduce
Center of Excellence for Hypersonics Research
2012-01-25
computational aerodynamics . These numerical flux functions are very effective at providing stable, robust simulations of steady-state hypersonic flows...novel second-order modified Crank-Nicholson approach was developed and tested. For certain problems (e.g. jets in supersonic crossflow), the method...re-entry vehicles, and to fuel-air mixing represented by a jet or jets in a supersonic cross-flow. This approach was also used to perform
Stability of hypersonic compression cones
NASA Astrophysics Data System (ADS)
Reed, Helen; Kuehl, Joseph; Perez, Eduardo; Kocian, Travis; Oliviero, Nicholas
2012-11-01
Our activities focus on the identification and understanding of the second-mode instability for representative configurations in hypersonic flight. These include the Langley 93-10 flared cone and the Purdue compression cone, both at 0 degrees angle of attack at Mach 6. Through application of nonlinear parabolized stability equations (NPSE) and linear parabolized stability equations (PSE) to both geometries, it is concluded that mean-flow distortion tends to amplify frequencies less than the peak frequency and stabilize those greater by modifying the boundary-layer thickness. As initial disturbance amplitude is increased and/or a broad spectrum disturbance is introduced, direct numerical simulations (DNS) or NPSE appear to be the proper choices to model the evolution, and relative evolution, because these computational tools include these nonlinear effects (mean-flow distortion). Support from AFOSR/NASA National Center for Hypersonic Research in Laminar-Turbulent Transition through Grant FA9550-09-1-0341 is gratefully acknowledged. The authors also thank Pointwise, AeroSoft, and Texas Advanced Computing Center (TACC).
Back, B.B.; Wuosmaa, A.H.; Baker, M.D.; Barton, D.S.; Carroll, A.; Chai, Z.; Gushue, S.; Hauer, M.; Heintzelman, G.A.; Holzman, B.; Pak, R.; Remsberg, L.P.; Seals, H.; Sedykh, I.; Stankiewicz, M.A.; Steinberg, P.; Sukhanov, A.; Ballintijn, M.; Busza, W.; Decowski, M.P.
2006-07-07
We report on measurements of directed flow as a function of pseudorapidity in Au+Au collisions at energies of {radical}(s{sub NN})=19.6, 62.4, 130 and 200 GeV as measured by the PHOBOS detector at the BNL Relativistic Heavy Ion Collider. These results are particularly valuable because of the extensive, continuous pseudorapidity coverage of the PHOBOS detector. There is no significant indication of structure near midrapidity and the data surprisingly exhibit extended longitudinal scaling similar to that seen for elliptic flow and charged particle pseudorapidity density.
NASA Technical Reports Server (NTRS)
Cheng, H. K.; Wong, Eric Y.; Dogra, V. K.
1991-01-01
Grad's thirteen-moment equations are applied to the flow behind a bow shock under the formalism of a thin shock layer. Comparison of this version of the theory with Direct Simulation Monte Carlo calculations of flows about a flat plate at finite attack angle has lent support to the approach as a useful extension of the continuum model for studying translational nonequilibrium in the shock layer. This paper reassesses the physical basis and limitations of the development with additional calculations and comparisons. The streamline correlation principle, which allows transformation of the 13-moment based system to one based on the Navier-Stokes equations, is extended to a three-dimensional formulation. The development yields a strip theory for planar lifting surfaces at finite incidences. Examples reveal that the lift-to-drag ratio is little influenced by planform geometry and varies with altitudes according to a 'bridging function' determined by correlated two-dimensional calculations.
NASA Technical Reports Server (NTRS)
Cheng, H. K.; Wong, Eric Y.; Dogra, V. K.
1991-01-01
Grad's thirteen-moment equations are applied to the flow behind a bow shock under the formalism of a thin shock layer. Comparison of this version of the theory with Direct Simulation Monte Carlo calculations of flows about a flat plate at finite attack angle has lent support to the approach as a useful extension of the continuum model for studying translational nonequilibrium in the shock layer. This paper reassesses the physical basis and limitations of the development with additional calculations and comparisons. The streamline correlation principle, which allows transformation of the 13-moment based system to one based on the Navier-Stokes equations, is extended to a three-dimensional formulation. The development yields a strip theory for planar lifting surfaces at finite incidences. Examples reveal that the lift-to-drag ratio is little influenced by planform geometry and varies with altitudes according to a 'bridging function' determined by correlated two-dimensional calculations.
2010-03-31
provides a low-strain-rate flameholding region (Johnson 2005, Bergthorson eta!. 2007). Figure 2 shows a composite (spliced) schlieren image from a pair...by varying the mass injected through the perforated ramp. Figure 3 is a composite schlieren image for a lower stream injection of UR = 45 m/s and...exit/exhaust and inlet pressures, respectively, and p u2/2 is the dynamic pressure. 3 Fig. 2. Composite Schlieren image of the expansion-ramp flow
Flux-split algorithms for the multi-dimensional Euler equations with real gases. [in hypersonic flow
NASA Technical Reports Server (NTRS)
Grossman, B.; Walters, R. W.
1989-01-01
Upwind algorithms are developed for the numerical solution of the multidimensional Euler equations for real gases. Flux-splitting methods are derived which account for a general equation of state. Approximations to the state equation based on physical arguments result in simplified algorithms which may be implemented into existing perfect-gas codes. Applications of the method to several high-Mach-number high-temperature flows are presented for two and three space dimensions.
Theoretical calculation of mid-infrared spectra from hypersonic non-ablative sphere
NASA Astrophysics Data System (ADS)
Wu, Jie; Yu, Xilong; Zhu, Xijuan; Ma, Jing; Mao, Hongxia
2016-10-01
Hypersonic body moving in the atmosphere will suffer high temperature reacting flows which will emit complex radiation. Theoretical calculation was taken in this paper for a hypersonic non-ablative sphere. Hypersonic flow around the sphere was simulated using 9 species chemical kinetic and two temperature thermal non-equilibrium model. Based on this simulated flow field, the LOS method is used to solve radiative transfer and line-by-line model is used to calculate the spectrum from molecular and atoms in mid-infrared. The spectra from different components have been analyzed one by one. The calculation founds out that atom N and O diatomic molecule NO and bremsstrahlung will be important radiation source in this pure air hypersonic flow field. The radiation from hypersonic flow field has been analyzed in both high pressure environment and low pressure environment.