Povinelli, L.A.
1990-01-01
An overview is given of research activity on the application of computational fluid dynamics (CDF) for hypersonic propulsion systems. After the initial consideration of the highly integrated nature of air-breathing hypersonic engines and airframe, attention is directed toward computations carried out for the components of the engine. A generic inlet configuration is considered in order to demonstrate the highly three dimensional viscous flow behavior occurring within rectangular inlets. Reacting flow computations for simple jet injection as well as for more complex combustion chambers are then discussed in order to show the capability of viscous finite rate chemical reaction computer simulations. Finally, the nozzle flow fields are demonstrated, showing the existence of complex shear layers and shock structure in the exhaust plume. The general issues associated with code validation as well as the specific issue associated with the use of CFD for design are discussed. A prognosis for the success of CFD in the design of future propulsion systems is offered.
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...
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.
Proximal bodies in hypersonic flow
NASA Astrophysics Data System (ADS)
Laurence, Stuart J.
The problem of proximal bodies in hypersonic flow is encountered in several important situations, both natural and man-made. The present work seeks to investigate one aspect of this problem by exploring the forces experienced by a secondary body when some part of it is within the shocked region created by 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 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 then one-sixth the primary diameter. The analytical results are compared with numerical simulations carried out using the AMROC software and good agreement is obtained 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, again with good agreement. In order to model this situation experimentally, a new force-measurement technique for short-duration hypersonic facilities has been developed, and results from the validation experiments are included. Finally, the analytical methodology is used to model two physical situations. First, the entry of a binary asteroid system into the Earth's atmosphere is simulated. Second, a model for a fragmenting meteoroid in a planetary atmosphere is developed, and simulations are carried out to determine whether the secondary scatter patterns in the Sikhote-Alin crater field may be attributed to aerodynamic interactions between fragments rather than to secondary fragmentation. It is found that while aerodynamic
Turbulence modeling for complex hypersonic flows
NASA Technical Reports Server (NTRS)
Huang, P. G.; Coakley, T. J.
1993-01-01
The paper presents results of calculations for a range of 2D turbulent hypersonic flows using two-equation models. The baseline models and the model corrections required for good hypersonic-flow predictions will be illustrated. Three experimental data sets were chosen for comparison. They are: (1) the hypersonic flare flows of Kussoy and Horstman, (2) a 2D hypersonic compression corner flow of Coleman and Stollery, and (3) the ogive-cylinder impinging shock-expansion flows of Kussoy and Horstman. Comparisons with the experimental data have shown that baseline models under-predict the extent of flow separation but over-predict the heat transfer rate near flow reattachment. Modifications to the models are described which remove the above-mentioned deficiencies. Although we have restricted the discussion only to the selected baseline models in this paper, the modifications proposed are universal and can in principle be transferred to any existing two-equation model formulation.
Nonequilibrium radiative hypersonic flow simulation
NASA Astrophysics Data System (ADS)
Shang, J. S.; Surzhikov, S. T.
2012-08-01
Nearly all the required scientific disciplines for computational hypersonic flow simulation have been developed on the framework of gas kinetic theory. However when high-temperature physical phenomena occur beneath the molecular and atomic scales, the knowledge of quantum physics and quantum chemical-physics becomes essential. Therefore the most challenging topics in computational simulation probably can be identified as the chemical-physical models for a high-temperature gaseous medium. The thermal radiation is also associated with quantum transitions of molecular and electronic states. The radiative energy exchange is characterized by the mechanisms of emission, absorption, and scattering. In developing a simulation capability for nonequilibrium radiation, an efficient numerical procedure is equally important both for solving the radiative transfer equation and for generating the required optical data via the ab-initio approach. In computational simulation, the initial values and boundary conditions are paramount for physical fidelity. Precise information at the material interface of ablating environment requires more than just a balance of the fluxes across the interface but must also consider the boundary deformation. The foundation of this theoretic development shall be built on the eigenvalue structure of the governing equations which can be described by Reynolds' transport theorem. Recent innovations for possible aerospace vehicle performance enhancement via an electromagnetic effect appear to be very attractive. The effectiveness of this mechanism is dependent strongly on the degree of ionization of the flow medium, the consecutive interactions of fluid dynamics and electrodynamics, as well as an externally applied magnetic field. Some verified research results in this area will be highlighted. An assessment of all these most recent advancements in nonequilibrium modeling of chemical kinetics, chemical-physics kinetics, ablation, radiative exchange
Speeding Convergence In Simulations Of Hypersonic Flow
NASA Technical Reports Server (NTRS)
Flores, J.; Cheung, S.; Cheer, A.; Hafez, M.
1991-01-01
Report describes study aimed at accelerating rates of convergence of iterative schemes for numerical integration of equations of hypersonic flow of viscous and inviscid fluids. Richardson-type overrelaxation method applied.
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.
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.
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.
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.
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.
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.
Internal hypersonic flow. [in thin shock layer
NASA Technical Reports Server (NTRS)
Lin, T. C.; Rubin, S. G.
1974-01-01
An approach for studying hypersonic internal flow with the aid of a thin-shock-layer approximation is discussed, giving attention to a comparison of thin-shock-layer results with the data obtained on the basis of the imposition theory or a finite-difference integration of the Euler equations. Relations in the case of strong interaction are considered together with questions of pressure distribution and aspects of the boundary-layer solution.
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.
On Numerical Methods For Hypersonic Turbulent Flows
NASA Astrophysics Data System (ADS)
Yee, H. C.; Sjogreen, B.; Shu, C. W.; Wang, W.; Magin, T.; Hadjadj, A.
2011-05-01
Proper control of numerical dissipation in numerical methods beyond the standard shock-capturing dissipation at discontinuities is an essential element for accurate and stable simulation of hypersonic turbulent flows, including combustion, and thermal and chemical nonequilibrium flows. Unlike rapidly developing shock interaction flows, turbulence computations involve long time integrations. Improper control of numerical dissipation from one time step to another would be compounded over time, resulting in the smearing of turbulent fluctuations to an unrecognizable form. Hypersonic turbulent flows around re- entry space vehicles involve mixed steady strong shocks and turbulence with unsteady shocklets that pose added computational challenges. Stiffness of the source terms and material mixing in combustion pose yet other types of numerical challenges. A low dissipative high order well- balanced scheme, which can preserve certain non-trivial steady solutions of the governing equations exactly, may help minimize some of these difficulties. For stiff reactions it is well known that the wrong propagation speed of discontinuities occurs due to the under-resolved numerical solutions in both space and time. Schemes to improve the wrong propagation speed of discontinuities for systems of stiff reacting flows remain a challenge for algorithm development. Some of the recent algorithm developments for direct numerical simulations (DNS) and large eddy simulations (LES) for the subject physics, including the aforementioned numerical challenges, will be discussed.
The Blunt Plate In Hypersonic Flow
NASA Technical Reports Server (NTRS)
Baradell, Donald L.; Bertram, Mitchel H.
1960-01-01
The sonic-wedge characteristics method has been used to obtain the shock shapes and surface pressure distributions on several blunt two-dimensional shapes in a hypersonic stream for several values of the ratio of specific heats. These shapes include the blunt slab at angle of attack and power profiles of the form yb = a)P, where 0 les than m less than 1, Yb and x are coordinates of the body surface, and a is a constant. These numerical results have been compared with the results of blast-wave theory, and methods of predicting the pressure distributions and shock shapes are proposed in each case. The effects of a free-stream conical-flow gradient on the pressure distribution on a blunt slab in hypersonic flow were investigated by the sonic-wedge characteristics method and were found to be sizable in many cases. Procedures which are satisfactory for reducing pressure data obtained in conical flows with small gradients are presented.
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.
Downstream Effects on Orbiter Leeside Flow Separation for Hypersonic Flows
NASA Technical Reports Server (NTRS)
Buck, Gregory M.; Pulsonetti, Maria V.; Weilmuenster, K. James
2005-01-01
Discrepancies between experiment and computation for shuttle leeside flow separation, which came to light in the Columbia accident investigation, are resolved. Tests were run in the Langley Research Center 20-Inch Hypersonic CF4 Tunnel with a baseline orbiter model and two extended trailing edge models. The extended trailing edges altered the wing leeside separation lines, moving the lines toward the fuselage, proving that wing trailing edge modeling does affect the orbiter leeside flow. Computations were then made with a wake grid. These calculations more closely matched baseline experiments. Thus, the present findings demonstrate that it is imperative to include the wake flow domain in CFD calculations in order to accurately predict leeside flow separation for hypersonic vehicles at high angles of attack.
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. PMID:25638120
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 N, flow over a hollow cylinder-flare with 30 deg 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 deg and aft-cone angle of 55 deg. Both sets of experiments involve 30 deg 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.
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.
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.
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
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 Prospects for Laminar Flow on Hypersonic Airplanes
NASA Technical Reports Server (NTRS)
Seiff, Alvin
1958-01-01
The factors which affect the extent of laminar flow on airplanes for hypersonic flight are discussed on the basis of the available data. Factors considered include flight Reynolds number, surface roughness, angle of attack, angle of leading-edge sweepback, and aerodynamic interference. Test data are presented for one complete configuration.
Hypersonic, nonequilibrium flow over a cylindrically blunted 6 deg wedge
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
1993-01-01
The numerical simulation of hypersonic flow in chemical nonequilibrium over cylindrically blunted 6 degree wedge is described. The simulation was executed on a Cray C-90 with Program LAURA 92-vl. Code setup procedures and sample results, including grid refinement studies and variations of species number are discussed. This simulation relates to a study of wing leading edge heating on transatmospheric vehicles.
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.
Unstructured Mesh Methods for the Simulation of Hypersonic Flows
NASA Technical Reports Server (NTRS)
Peraire, J.
1999-01-01
This report summarizes the research undertaken, at Aeronautics Department of the Massachusetts Institute of Technology, during the approximately five year period, February 94 - March 99. This work is part of a larger effort aimed at providing a reliable fast turn around capability for the prediction of hypersonic flows over complete vehicle configurations.
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.
Application of a parallel DSMC method to hypersonic rarefied flows
Wilmoth, R.G. )
1991-01-01
This paper describes a method for doing direct simulation Monte Carlo (DSMC) calculations using parallel processing and presents some results of applying the method to several hypersonic, rarefied flow problems. The performance and efficiency of the parallel method are discussed. The applications described are the flow in a channel and the flow about a flat plate at incidence. The results show significant advantages of parallel processing over conventional scalar processing and demonstrate the scalability of the method to large problems. 8 refs.
Flow visualization and spectroscopy in hypersonic flows: New trends
NASA Astrophysics Data System (ADS)
Trolinger, James; Eitelberg, Georg; Rapuc, Marc
1993-04-01
This paper is based upon a session of the NATO Advanced Research Workshop, New Trends in Instrumentation for Hypersonic Research held at the ONERA La Fauga Facility in France during the week of 27 Apr. 1992. The discussion includes some of the frontiers of the technology of flow visualization and spectroscopy as well as a discussion of the current development needs and trends. Included in the discussion are optical integrated measurements such as resonance absorption, schlieren, interferometry, and holographic methods. The discussion shows that while the technology is mature in a broad sense, a significant number of new development areas exist such as resonant holography and phase shifting holographic interferometry. The maturity of the technology makes it immediately applicable to many problems and the untapped potential offers considerable room for improvement of existing capability. The methods which are described can be used in harsh environments and have the potential for becoming flight test diagnostics for the measurement of temperature, density, constituency, and velocity.
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
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.
Drag Prediction and Transition in Hypersonic Flow
NASA Technical Reports Server (NTRS)
Reed, Helen L.; Kimmel, Roger; Schneider, Steven; Arnal, Daniel
1997-01-01
This paper discusses progress on issues such as instability studies, nose-bluntness and angle-of-attack effects, and leading-edge-contamination problems from theoretical, computational, and experimental points of view. Also included is a review of wind-tunnel and flight data, including high-Re flight transition data, the levels of noise in flight and in wind tunnels, and how noise levels can affect parametric trends. A review of work done on drag accounting and the role of viscous drag for hypersonic vehicles is also provided.
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.
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 remove 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 viscous flow over a blunt biconic.
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.
Progress with multigrid schemes for hypersonic flow problems
Radespiel, R.; Swanson, R.C.
1995-01-01
Several multigrid schemes are considered for the numerical computation of viscous hypersonic flows. For each scheme, the basic solution algorithm employs 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{sup 6} and Mach numbers up to 25. 32 refs., 31 figs., 1 tab.
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.
Hypersonic flows as related to the National Aerospace Plane
NASA Technical Reports Server (NTRS)
Kussoy, Marvin; Huang, George; Menter, Florian
1995-01-01
The object of Cooperative Agreement NCC2-452 was to identify, develop, and document reliable turbulence models for incorporation into CFD codes, which would then subsequently be incorporated into numerical design procedures for the NASP and any other hypersonic vehicles. In a two-pronged effort, consisting of an experimental and a theoretical approach, several key features of flows over complex vehicles were identified, and test bodies were designed which were composed of simple geometric shapes over which these flow features were measured. The experiments were conducted in the 3.5' Hypersonic Wind Tunnel at NASA Ames Research Center, at nominal Mach numbers from 7 to 8.3 and Re/m from 4.9 x 10(exp 6) to 5.8 x 10(exp 6). Boundary layers approaching the interaction region were 2.5 to 3.7 cm thick. Surface and flow field measurements were conducted, and the initial boundary conditions were experimentally documented.
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
Computational flow predictions for hypersonic drag devices
NASA Technical Reports Server (NTRS)
Tokarcik, Susan A.; Venkatapathy, Ethiraj
1993-01-01
The effectiveness of two types of hypersonic decelerators is examined: mechanically deployable flares and inflatable ballutes. Computational fluid dynamics (CFD) is used to predict the flowfield around a solid rocket motor (SRM) with a deployed decelerator. The computations are performed with an ideal gas solver using an effective specific heat ratio of 1.15. The results from the ideal gas solver are compared to computational results from a thermochemical nonequilibrium solver. The surface pressure coefficient, the drag, and the extend of the compression corner separation zone predicted by the ideal gas solver compare well with those predicted by the nonequilibrium solver. The ideal gas solver is computationally inexpensive and is shown to be well suited for preliminary design studies. The computed solutions are used to determine the size and shape of the decelerator that are required to achieve a drag coefficient of 5. Heat transfer rates to the SRM and the decelerators are predicted to estimate the amount of thermal protection required.
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.
Grid sensitivity in low Reynolds number hypersonic continuum flows
Rutledge, W.H. ); Hoffmann, K.A. . Dept. of Aerospace Engineering)
1991-01-01
A computational scheme is presented to solve the unsteady Navier-Stokes equations over a blunt body at high altitude, high Mach number atmospheric reentry flow conditions. This continuum approach is directed to low Reynolds/low density hypersonic flows by accounting for non-zero bulk viscosity effects in near frozen flow conditions. A significant difference from previous studies is the inclusion of the capability to model non-zero bulk viscosity effects. The grid definition for these low Reynolds number, viscous dominated flow fields is especially important in terms of numerical stability and accurate heat transfer solutions. 11 refs., 15 figs.
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
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.
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.
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
Modeling of associative ionization reactions in hypersonic rarefied flows
NASA Astrophysics Data System (ADS)
Boyd, Iain D.
2007-09-01
When vehicles reenter the Earth's atmosphere from space, the hypersonic conditions are sufficiently energetic to generate ionizing reactions. The production of a thin plasma layer around a hypersonic vehicle can block radio waves sent to and from the vehicle, leading to communications blackout. For Earth entry from orbit, the maximum energy involved in molecular collisions requires only associative ionization of air-species to be considered. In the present study, the modeling of such reactions is considered in detail using the direct simulation Monte Carlo (DSMC) method. For typical Earth entry conditions, with a velocity near 8km/s, it is shown that the average ionizing reaction probabilities are small. Special numerical techniques must therefore be used in the DSMC technique in order to numerically resolve these reactions. Additional simulation problems arise from the relatively small mass of the electrons in comparison to the other atoms and molecules in these flow fields. Artificially increasing the electron mass greatly increases computational efficiency, and the viability of this approach is investigated. Simulation results are presented for conditions corresponding to the RAM-C II hypersonic flight experiment that gathered measurements of electron number density. It is demonstrated that simulation results for electron number density in this energy regime are relatively insensitive to the mass of the electrons. Direct comparison of DSMC results with the RAM-C II measurements for electron number density shows excellent agreement. These satisfactory comparisons represent the first direct verification of the ability of the DSMC technique to successfully predict the weak plasma generated around a hypersonic vehicle.
Simplified Thermo-Chemical Modelling For Hypersonic Flow
NASA Astrophysics Data System (ADS)
Sancho, Jorge; Alvarez, Paula; Gonzalez, Ezequiel; Rodriguez, Manuel
2011-05-01
Hypersonic flows are connected with high temperatures, generally associated with strong shock waves that appear in such flows. At high temperatures vibrational degrees of freedom of the molecules may become excited, the molecules may dissociate into atoms, the molecules or free atoms may ionize, and molecular or ionic species, unimportant at lower temperatures, may be formed. In order to take into account these effects, a chemical model is needed, but this model should be simplified in order to be handled by a CFD code, but with a sufficient precision to take into account the physics more important. This work is related to a chemical non-equilibrium model validation, implemented into a commercial CFD code, in order to obtain the flow field around bodies in hypersonic flow. The selected non-equilibrium model is composed of seven species and six direct reactions together with their inverse. The commercial CFD code where the non- equilibrium model has been implemented is FLUENT. For the validation, the X38/Sphynx Mach 20 case is rebuilt on a reduced geometry, including the 1/3 Lref forebody. This case has been run in laminar regime, non catalytic wall and with radiative equilibrium wall temperature. The validated non-equilibrium model is applied to the EXPERT (European Experimental Re-entry Test-bed) vehicle at a specified trajectory point (Mach number 14). This case has been run also in laminar regime, non catalytic wall and with radiative equilibrium wall temperature.
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.
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
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.
Numerical simulation of laminar hypersonic flows about an ellipsoid
NASA Astrophysics Data System (ADS)
Riedelbauch, S.; Mueller, B.
The laminar hypersonic flow about a double ellipsoid, which idealizes the nose and cockpit of a spacecraft, were numerically simulated. The calculation method solves the three dimensional thin layer Navier-Stokes equations in a conservative formulation on a surface oriented calculation grid using an implicit/explicit finite difference technique. The conservative formulation allows the correct calculation of embedded compression shocks, while the head wave was treated with a shock-fitting procedure. The calculated flow fields about the ellipsoid show shock-shock and shock-boundary layer interactions in connection with separated flow. Wall flow lines and heat transfer agree qualitatively very well with film-of-oil and thermographic pictures.
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.
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.
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.
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.
Hypersonic flow over a multi-step afterbody
NASA Astrophysics Data System (ADS)
Menezes, V.; Kumar, S.; Maruta, K.; Reddy, K. P. J.; Takayama, K.
2005-12-01
Effect of a multi-step base on the total drag of a missile shaped body was studied in a shock tunnel at a hypersonic Mach number of 5.75. Total drag over the body was measured using a single component accelerometer force balance. Experimental results indicated a reduction of 8% in total drag over the body with a multi-step base in comparison with the base-line (model with a flat base) configuration.The flow fields around the above bodies were simulated using a 2-D axisymmetric Navier Stokes solver and the simulated results on total drag were compared with the measured results. The simulated flow field pictures give an insight into the involved flow physics.
Hypersonic Flow Control Using Upstream Focused Energy Deposition
NASA Technical Reports Server (NTRS)
Riggins David W.; Nelson, H. F.
1999-01-01
A numerical study of centerline and off-centerline power deposition at a point upstream of a two-dimensional blunt body at Mach 6.5 at 30 km altitude are presented. The full Navier-Stokes equations are used. Wave drag, lift, and pitching moment are presented as a function of amount of power absorbed in the flow and absorption point location. It is shown that wave drag is considerably reduced. Modifications to the pressure distribution in the flow field due to the injected energy create lift and a pitching moment when the injection is off-centerline. This flow control concept may lead to effective ways to improve the performance and to stabilize and control hypersonic vehicles.
Convergence acceleration of viscous and inviscid hypersonic flow calculations
NASA Technical Reports Server (NTRS)
Cheer, A.; Hafez, M.; Cheung, S.; Flores, J.
1989-01-01
The convergence of inviscid and viscous hypersonic flow calculations using a two-dimensional flux-splitting code is accelerated by applying a Richardson-type overrelaxation method. Successful results are presented for various cases; and a 50 percent savings in computer time is usually achieved. An analytical formula for the overrelaxation factor is derived, and the performance of this scheme is confirmed numerically. Moreover, application of this overrelaxation scheme produces a favorable preconditioning for Wynn's epsilon-algorithm. Both techniques have been extended to viscous three-dimensional flows and applied to accelerate the convergence of the compressible Navier-Stokes code. A savings of 40 percent in computer time is achieved in this case.
Vibrational Energy Transfer of Diatomic Gases in Hypersonic Expanding Flows.
NASA Astrophysics Data System (ADS)
Ruffin, Stephen Merrick
In high temperature flows related to vehicles at hypersonic speeds significant excitation of the vibrational energy modes of the gas can occur. Accurate predictions of the vibrational state of the gas and the rates of vibrational energy transfer are essential to achieve optimum engine performance, for design of heat shields, and for studies of ground based hypersonic test facilities. The Landau -Teller relaxation model is widely used because it has been shown to give accurate predictions in vibrationally heating flows such as behind forebody shocks. However, a number of experiments in nozzles have indicated that it fails to accurately predict the rate of energy transfer in expanding, or cooling, flow regions and fails to predict the distribution of energy in the vibrational quantum levels. The present study examines the range of applicability of the Landau -Teller model in expanding flows and develops techniques which provide accurate predictions in expanding flows. In the present study, detailed calculations of the vibrational relaxation process of N_2 and CO in cooling flows are conducted. A coupled set of vibrational transition rate equations and quasi one-dimensional fluid dynamic equations is solved. Rapid anharmonic Vibration-Translation transition rates and Vibration -Vibration exchange collisions are found to be responsible for vibrational relaxation acceleration in situations of high vibrational temperature and low translational temperature. The predictions of the detailed master equation solver are in excellent agreement with experimental results. The exact degree of acceleration is cataloged in this study for N_2 and is found to be a function of both the translational temperature (T) and the ratio of vibrational to translational temperatures (T_{vib}/T). Non-Boltzmann population distributions are observed for values of T _{vib}/T as low as 2.0. The local energy transfer rate is shown to be an order of magnitude or more faster than the Landau-Teller model
Numerical Solutions of Supersonic and Hypersonic Laminar Compression Corner Flows
NASA Technical Reports Server (NTRS)
Hung, C. M.; MacCormack, R. W.
1976-01-01
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 and a Reynolds number of 1.04 x 10(exp 5) with wedge angles of 15 deg, 18 deg, 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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
Viscous, radiating hypersonic flow about a blunt body
NASA Technical Reports Server (NTRS)
Passamaneck, R. S.
1974-01-01
The viscous, radiating hypersonic flow past an axisymmetric blunt body is analyzed based on the Navier-Stokes equations, plus a radiative equation of transfer derived from the Milne-Eddington differential approximation. The fluid is assumed to be a perfect gas with constant specific heats, a constant Prandtl number of order unity, a viscosity coefficient varying as a power of the temperature, and an absorption coefficient varying as the first power of the density and as a power of the temperature. The gray gas assumption is invoked, thereby making the absorption coefficient independent of the spectral frequency. Limiting forms of the solutions are studied as the freestream Mach number freestream Reynolds number and the temperature ratio across the shock wave, go to infinity, and as the Bouguer number and the density ratio across the shock wave go to zero. The method of matched asymptotic expansions is used in the analysis, and it is shown that there is a far-field precursor, composed of two regions, in which the fluid mechanics can be neglected for all practical purposes but included for completeness.
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.
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.
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.
Simulation of 3D flows past hypersonic vehicles in FlowVision software
NASA Astrophysics Data System (ADS)
Aksenov, A. A.; Zhluktov, S. V.; Savitskiy, D. V.; Bartenev, G. Y.; Pokhilko, V. I.
2015-11-01
A new implicit velocity-pressure split method is discussed in the given presentation. The method implies using conservative velocities, obtained at the given time step, for integration of the momentum equation and other convection-diffusion equations. This enables simulation of super- and hypersonic flows with account of motion of solid boundaries. Calculations of known test cases performed in the FlowVision software are demonstrated. It is shown that the method allows one to carry out calculations at high Mach numbers with integration step essentially exceeding the explicit time step.
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.
Hypersonic High-Enthalpy Flow in a Leading-Edge Separation
NASA Astrophysics Data System (ADS)
Deepak, N. R.; Gai, S. L.; O'Byrne, Sean; Moss, J. N.
Flow separation in/over a hypersonic space vehicle is an important phenomenon which occurs due to flow interaction with various geometric elements of the vehicle. This however can lead to adverse pressure gradient and localised intense heating resulting in detrimental consequences for the successful performance of the vehicle.
Thermal flux measurements in hypersonic flows: A review
NASA Astrophysics Data System (ADS)
Wendt, J. F.; Balageas, D.; Neumann, R. D.
1993-04-01
This contribution reviews the papers presented in the Session on 'Heat Flux' and 'Thermography' at a NATO Advanced Research Workshop entitled 'New Trends in Instrumentation for Hypersonic Research', 27 April-1 May, 1992, Le Fauga, France. The present status and problem areas associated with specific methods are discussed and recommendations for future research and development are presented.
Tests of Hypersonic Inlet Oscillatory Flows in a Shock Tunnel
NASA Astrophysics Data System (ADS)
Li, Zhufei; Gao, Wenzhi; Jiang, Hongliang; Yang, Jiming
For efficient operation, hypersonic air breathing engine requires the inlet to operate in a starting mode [1]. High backpressure induced by the combustion may cause the inlet to unstart in the engine actual operation [2].When unstarted, shock wave oscillations are typically observed in the inlet, a phenomenon known as buzz.
A new Lagrangian random choice method for steady two-dimensional supersonic/hypersonic flow
NASA Technical Reports Server (NTRS)
Loh, C. Y.; Hui, W. H.
1991-01-01
Glimm's (1965) random choice method has been successfully applied to compute steady two-dimensional supersonic/hypersonic flow using a new Lagrangian formulation. The method is easy to program, fast to execute, yet it is very accurate and robust. It requires no grid generation, resolves slipline and shock discontinuities crisply, can handle boundary conditions most easily, and is applicable to hypersonic as well as supersonic flow. It represents an accurate and fast alternative to the existing Eulerian methods. Many computed examples are given.
Direct numerical simulation of laminar-turbulent flow over a flat plate at hypersonic flow speeds
NASA Astrophysics Data System (ADS)
Egorov, I. V.; Novikov, A. V.
2016-06-01
A method for direct numerical simulation of a laminar-turbulent flow around bodies at hypersonic flow speeds is proposed. The simulation is performed by solving the full three-dimensional unsteady Navier-Stokes equations. The method of calculation is oriented to application of supercomputers and is based on implicit monotonic approximation schemes and a modified Newton-Raphson method for solving nonlinear difference equations. By this method, the development of three-dimensional perturbations in the boundary layer over a flat plate and in a near-wall flow in a compression corner is studied at the Mach numbers of the free-stream of M = 5.37. In addition to pulsation characteristic, distributions of the mean coefficients of the viscous flow in the transient section of the streamlined surface are obtained, which enables one to determine the beginning of the laminar-turbulent transition and estimate the characteristics of the turbulent flow in the boundary layer.
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.
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
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.
NASA Astrophysics Data System (ADS)
Bhatia, Ankush
Discontinuous Galerkin (DG) methods are high-order accurate, compact-stencil methods, proven to possess favorable properties for highly efficient parallel systems, complex geometries and unstructured meshes. Coding effort is significantly reduced for compact-stencil DG methods in comparison to main stream finite difference and finite volume methods. This work successfully introduces DG methods to thermal ablation and non-equilibrium hypersonic flows. In the state-of-the-art hypersonic flow codes, surface heating predictions are very sensitive to mesh resolution in the shock. A minor misalignment can cause major changes in the heating predictions. This is due to the lack of high-order accuracy in current streamline methods and numerical errors associated with the shock capturing approach. Shock capturing methods like slope limiter or artificial viscosity, being empirical have errors in the shock region. This work employs r-p adaptivity to accurately capture the shock with p = 0 elements (first order accuracy). Smooth flow regions are captured using p greater than 0. This method is stable. Implicit methods are developed for solution advancement with high CFL numbers. Error in the shock is reduced by redistributing the elements (outside of the shock) to within the shock (r adaptivity). Inviscid and viscous hypersonic flow problems, with same accuracy as in h-p adaptivity method, are simulated with one-third elements. This methodology requires no a priori knowledge of the shock's location, and is suitable for detached shock problems. r-p adaptivity method has allowed for successful prediction of surface heating rate for hypersonic flow over cylinder. Additionally, good comparisons are made, for non-equilibrium hypersonic flows, to the published results. This tool is also used to determine the effect of micro-second pulsed sinusoidal Dielectric Barrier Discharge (DBD) plasma actuators on the surface heating reduction for hypersonic flow over cylinder. A significant
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.
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.
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.
Hypersonic Nozzle/Afterbody Experiment: Flow Visualization and Boundary Layer Experiments
NASA Technical Reports Server (NTRS)
Keener, Earl R.; Spaid, Frank W.; Arnold, James O. (Technical Monitor)
1994-01-01
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 of the NASA Ames Research Center in a cooperative experimental program involving Ames and the 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 Aero-Space Plane (NASP) program. This paper presents experimental results consisting of oil-flow and shadowgraph flow-visualization photographs, afterbody surface-pressure distributions, boundary-layer rake measurements, and Preston-tube skin-friction measurements.
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.
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.
Transition in Hypersonic Flows Including High-temperature Gas Effects
NASA Technical Reports Server (NTRS)
Stemmer, Christian
2003-01-01
Hypersonic transition poses a special challenge for direct numerical simulations. Comparable data from Wind-tunnel tests or free-flight testing are not available or not accurate enough for comparison. The wind-tunnel testing does not allow for the exact match to the free-flight conditions at such high Mach-numbers. Flat-plate boundary-layer transition at high Mach-numbers is investigated in this work. A simulation case was chosen where chemical non-equilibrium plays an important role but ionization can be neglected. The chosen case at an altitude of H=50Km lies close to one point on the descent path of the Space Shuttle. The failure of the Space Shuttle has shown that an improved vehicle for space transportation is imperative in the close future. Transition research for an improved space-transportation vehicle is crucial in order to estimate the heat load during re-entry.
NASA Technical Reports Server (NTRS)
Kussoy, Marvin I.; Horstman, Clifford C.
1989-01-01
Experimental data for a series of two- and three-dimensional shock wave/turbulent boundary layer interaction flows at Mach 7 are presented. Test bodies, composed of simple geometric shapes, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure and heat-transfer distributions as well as limited mean-flow-field surveys in both the undisturbed and the interaction regimes. The data are presented in a convenient form for use in validating existing or future computational models of these generic hypersonic flows.
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.
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.
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.
Detailed numerical modeling of chemical and thermal nonequilibrium in hypersonic flows
Riedel, U.; Maas, U.; Warnatz, J. )
1993-03-01
Interest in hypersonic flows has created a large demand for physicochemical models for air flow computations around reentry bodies. Detailed physicochemical models for air in chemical and thermal nonequilibrium are needed for a realistic prediction of hypersonic flowfields. In this paper we develop a model, based on elementary physicochemical processes, for a detailed description of chemical nonequilibrium together with the excitation of internal DOFs. This model is implemented in a 2D Navier-Stokes code in order to show the strong influence of thermal nonequilibrium on the flowfields. The algorithm presented here is based on a fully conservative discretization of the inviscid fluxes in the conservation equations and uses the chain rule conservation law form for the viscous fluxes. The large system of ordinary differential and algebraic equations resulting from the spatial discretization is solved by a time-accurate semiimplicit extrapolation method. 34 refs.
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. PMID:20168704
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.
Hypersonic, nonequilibrium flow over the FIRE 2 forebody at 1634 sec
NASA Technical Reports Server (NTRS)
Chambers, Lin Hartung
1994-01-01
The numerical simulation of hypersonic flow in thermochemical nonequilibrium over the forebody of the FIRE 2 vehicle at 1634 sec in its trajectory is described. The simulation was executed on a Cray C90 with the program Langley Aerodynamic Upwind Relaxation Algorithm (LAURA) 4.0.2. Code setup procedures and sample results, including grid refinement studies, are discussed. This simulation relates to a study of radiative heating predictions on aerobrake type vehicles.
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
NASA Astrophysics Data System (ADS)
Widodo, Agung; Buttsworth, David
2013-04-01
Stagnation temperatures at the nozzle exit of the University of Southern Queensland hypersonic wind tunnel facility have been identified using an aspirating tube device with a 0.075 mm diameter k-type butt-welded thermocouple junction positioned at its inlet. Because of the finite thermal inertia of the thermocouple, a response time correction is introduced, and uncertainties in the response time correction are assessed and minimized by operating the aspirating device over a range of different initial temperatures. Pressure measurements within the barrel of the wind tunnel facility were used to estimate a theoretical upper bound on the flow stagnation temperature by assuming isentropic compression of the test gas. Results demonstrate that for the current operating conditions, the gas which is first delivered into the hypersonic nozzle has a stagnation temperature almost identical to the isentropic compression value of around 560 K, but a cooling effect is registered for the duration of the test flow which is about 200 ms. Thermodynamic simulations based on an unsteady energy balance model with turbulent heat transfer from the test gas within the barrel demonstrate a cooling effect of a similar magnitude to that indicated by the measured temperature variation, suggesting that strong mixing of the test gas occurs within the barrel during flow discharge through the hypersonic nozzle.
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%. PMID:24672360
Toluene-based planar laser-induced fluorescence imaging of temperature in hypersonic flows
NASA Astrophysics Data System (ADS)
Estruch-Samper, D.; Vanstone, L.; Hillier, R.; Ganapathisubramani, B.
2015-06-01
Planar laser-induced fluorescence imaging is carried out in a hypersonic gun tunnel at a freestream Mach number of 8.9 and Reynolds number of ( is the test gas). The fluorescence of toluene is correlated with the red shift of the emission spectra with increasing temperature. A two-colour approach is used where, following an excitation at 266 nm, emission spectra at two different bands are captured in separate runs using two different filters. Two different flow fields are investigated using this method: (i) hypersonic flow past a blunt nose, which is characterised by a bow shock with strong entropy effects, and (ii) an attached shock-wave/boundary-layer interaction induced by a flare located further downstream on the same blunt cylinder body. Measurements from as low as the freestream temperature of K all the way up to K are obtained. The uncertainty at the higher temperature level is approximately %, while at the low end of the temperature, an additional % uncertainty is expected. Application of the technique is further challenged at high temperatures due to the exponentially reduced fluorescence quantum yields and the occurrence of toluene pyrolysis near the stagnation region ( K). Overall, results are found to be within % agreement with the expected distributions, thus demonstrating suitability of the technique for hypersonic flow thermometry applications in low-enthalpy facilities.
Borzov, V.Yu.; Rybka, I.V.; Yur`ev, A.S.
1995-06-01
Parameters of the axisymmetric flow around bodies with different bluntness are compared in the case of constant energy supply to the main hypersonic flow. Flow structures, drag coefficients, and expenditure of energy on overcoming drag are analyzed with the effect of thermal energy on the flow taken into account for different bodies with equal volume.
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.
Leading-edge receptivity to free-stream disturbance waves for hypersonic flow over a parabola
NASA Astrophysics Data System (ADS)
Zhong, Xiaolin
2001-08-01
The receptivity of hypersonic boundary layers to free-stream disturbances, which is the process of environmental disturbances initially entering the boundary layers and generating disturbance waves, is altered considerably by the presence of bow shocks in hypersonic flow fields. This paper presents a numerical simulation study of the generation of boundary layer disturbance waves due to free-stream waves, for a two-dimensional Mach 15 viscous flow over a parabola. Both steady and unsteady flow solutions of the receptivity problem are obtained by computing the full Navier Stokes equations using a high-order-accurate shock-fitting finite difference scheme. The effects of bow-shock/free-stream-sound interactions on the receptivity process are accurately taken into account by treating the shock as a discontinuity surface, governed by the Rankine-Hugoniot relations. The results show that the disturbance waves generated and developed in the hypersonic boundary layer contain both first-, second-, and third-mode waves. A parametric study is carried out on the receptivity characteristics for different free-stream waves, frequencies, nose bluntness characterized by Strouhal numbers, Reynolds numbers, Mach numbers, and wall cooling. In this paper, the hypersonic boundary-layer receptivity is characterized by a receptivity parameter defined as the ratio of the maximum induced wave amplitude in the first-mode-dominated region to the amplitude of the free-stream forcing wave. It is found that the receptivity parameter decreases when the forcing frequency or nose bluntness increase. The results also show that the generation of boundary layer waves is mainly due to the interaction of the boundary layer with the acoustic wave field behind the bow shock, rather than interactions with the entropy and vorticity wave fields.
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.
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. PMID:19693204
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. PMID:20582058
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.
Ir Thermographic Measurements of Temperatures and Heat Fluxes in Hypersonic Plasma Flow
NASA Astrophysics Data System (ADS)
Cardone, G.; Tortora, G.; del Vecchio, A.
2005-02-01
The technological development achieved in instruments and methodology concerning both flights and ground hypersonic experiment (employed in space plane planning) goes towards an updating and a standardization of the heat flux technical measurements. In fact, the possibility to simulate high enthalpy flow relative to reentry condition by hypersonic arc-jet facility needs devoted methods to measure heat fluxes. Aim of this work is to develop an experimental numerical technique for the evaluation of heat fluxes over Thermal Protection System (TPS) by means of InfraRed (IR) thermographic temperature measurements and a new heat flux sensor (IR-HFS). We tackle the numerical validation of IR-HFS, apply the same one to the Hyflex nose cap model and compare the obtained results with others ones obtained by others methodology.
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.
Automatic liquid crystal thermography for transient heat transfer measurements in hypersonic flow
NASA Astrophysics Data System (ADS)
Babinsky, H.; Edwards, J. A.
1996-08-01
A technique has been developed to measure surface heat transfer on windtunnel models in hypersonic flow based on the colour response of encapsulated thermochromic liquid crystals. The method supplies results of a superior spatial resolution at experimental uncertainties comparable to traditional methods. The approach is different from other liquid crystal applications in several key areas. It combines the calibration of the liquid crystal coating with the actual mesurement and therefore allows for an efficient experiment. The method is automated in most steps involved. Results are shown for the flow over an axisymmetric compression corner at Mach 5 and compared with surface thermocouple measurements.
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.
An unstructured shock-fitting solver for hypersonic plasma flows in chemical non-equilibrium
NASA Astrophysics Data System (ADS)
Pepe, R.; Bonfiglioli, A.; D'Angola, A.; Colonna, G.; Paciorri, R.
2015-11-01
A CFD solver, using Residual Distribution Schemes on unstructured grids, has been extended to deal with inviscid chemical non-equilibrium flows. The conservative equations have been coupled with a kinetic model for argon plasma which includes the argon metastable state as independent species, taking into account electron-atom and atom-atom processes. Results in the case of an hypersonic flow around an infinite cylinder, obtained by using both shock-capturing and shock-fitting approaches, show higher accuracy of the shock-fitting approach.
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.
NASA Astrophysics Data System (ADS)
Gülhan, A.; Braun, S.
2011-03-01
An experimental study on the efficiency of transpiration cooling in hypersonic laminar and turbulent flow regimes is carried out in the Hypersonic Windtunnel Cologne with a focus on the aerothermal problems downstream of the cooled model part. The model is made of a material of low thermal conductivity (PEEK) with an integrated probe of a porous material. The experimental setup allows the direct comparison of the thermal behavior of transpiration cooling to a well-defined and radiatively cooled reference surface. Experiments are performed at Mach number of 6 and two different Reynolds numbers. Air, argon and helium are used as coolants at various flow rates, in order to identify the influence of coolant medium on cooling efficiency. The cooling efficiency of air and argon is comparable. Helium provides significantly higher cooling efficiency at the same blowing ratio, i.e. same coolant mass flow rate. The experimental data shows that the efficiency of the transpiration cooling in turbulent flows is much lower than in laminar flow.
Liquid crystal coatings for surface shear stress visualization in hypersonic flows
Reda, D.C.; Aeschliman, D.P.
1990-01-01
Experiments were conducted to test the surface-shear-stress visualization capabilities of shear-stress-sensitive/temperature- insensitive liquid crystal compounds in hypersonic flow. Liquid crystal coatings were applied to the surface of a conical model, which was then exposed to a high-unit-Reynolds-number (2.3 {times} 10{sup 7}/m) Mach 5 flow. The coating was illuminated by white light, and its response to the various flow situations was monitored and recorded with standard video and high-speed movie cameras. Boundary layer transition to turbulence was clearly demarcated by the technique. The dynamic location of the transition front as a function of model angle of attack (for sharp and blunt cones, with and without boundary-layer trips) was recorded, and observations were found to be consistent with established (published) trends for hypersonic flows over conical bodies. Normal shock passage over the model during tunnel shutdown was recorded (at 400 frames/second), and the liquid crystal coating was observed to respond to this event in a time interval less than or equal to the time between sequential movie frame exposures ({le} 0.0025 seconds). The liquid crystal technique has thus been demonstrated as a viable diagnostic tool for use in transient/compressible flows. 18 refs., 3 figs.
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.
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.
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.
Effect of bulk viscosity in low density, hypersonic blunt body flows
Rutledge, W.H. ); Hoffmann, K.A. )
1991-01-01
A computational fluids dynamics scheme is presented to solve the unsteady Thin-Layer Navier-Stokes (TLNS) equations over a blunt body at high altitude, high Mach number atmospheric reentry flow conditions. This continuum approach is directed to low density hypersonic flows by accounting for non-zero bulk viscosity effects in near frozen flow conditions. The TLNS equations are solved over an axisymmetric body at zero incidence relative to the free stream. The time dependent axisymmetric governing equations are transformed into a computational plane, then cast into weak conservative form and solved using a first-order fully implicit scheme in time with second-order flux vector splitting for spatial derivatives. The physical domain is defined over representative sphere and sphere/cone geometries using a body-fitted clustered algebraic grid within a fixed domain (i.e., shock capturing). At the present time, nonequilibrium thermo-chemistry effects are not modeled. Catalytic wall, ionization and radiation effects are also excluded from the current analysis. However, the significant difference from previous studies is the inclusion of the capability to model non-zero bulk viscosity effects. The importance of bulk viscosity is reviewed and blunt body flow field solutions are presented to illustrate the potential contribution of this phenomena at high altitude hypersonic conditions. The current technique is compared with experimental data and other approximate continuum solutions. A variety of test cases are also presented for a wide range of free stream Mach conditions. 18 refs., 42 figs.
General Reynolds Analogy for Blunt-Nosed Bodies in Hypersonic Flows
NASA Astrophysics Data System (ADS)
Chen, Xing-Xing; Wang, Zhi-Hui; Yu, Yong-Liang
2015-08-01
In this paper, the relation between skin friction and heat transfer along windward sides of blunt-nosed bodies in hypersonic flows is investigated. The self-similar boundary layer analysis is accepted to figure out the distribution of the ratio of skin friction to heat transfer coefficients along the wall. It is theoretically obtained that the ratio depends linearly on the local slope angle of the wall surface, and an explicit analogy expression is presented for circular cylinders, although the linear distribution is also found for other nose shapes and even in gas flows with chemical reactions. Furthermore, based on the theoretical modelling of the second order shear and heat transfer terms in Burnett equations, a modified analogy is derived in the near continuum regime by considering the rarefied gas effects. And a bridge function is also constructed to describe the nonlinear analogy in the transition flow regime. At last, the direct simulation Monte Carlo method is used to validate the theoretical results. The general analogy, beyond the classical Reynolds analogy, is applicable to both flat plates and blunt-nosed bodies, in either continuous or rarefied hypersonic flows.
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.
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.
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.
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.
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.
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.
NASA Technical Reports Server (NTRS)
Gupta, R. N.; Rodkiewicz, C. M.
1975-01-01
The numerical results are obtained for heat transfer, skin-friction, and viscous interaction induced pressure for a step-wise accelerated flat plate in hypersonic flow. In the unified approach here the results are presented for both weak and strong-interaction problems without employing any linearization scheme. With the help of the numerical method used in this work an accurate prediction of wall shear can be made for the problems with plate velocity changes of 1% or larger. The obtained results indicate that the transient contribution to the induced pressure for helium is greater than that for air.
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.
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.
Accurate Navier-Stokes results for the hypersonic flow over a spherical nosetip
Blottner, F.G.
1989-01-01
The unsteady thin-layer Navier-Stokes equations for a perfect gas are solved with a linearized block Alternating Direction Implicit finite-difference solution procedure. Solution errors due to numerical dissipation added to the governing equations are evaluated. Errors in the numerical predictions on three different grids are determined where Richardson extrapolation is used to estimate the exact solution. Accurate computational results are tabulated for the hypersonic laminar flow over a spherical body which can be used as a benchmark test case. Predictions obtained from the code are in good agreement with inviscid numerical results and experimental data. 9 refs., 11 figs., 3 tabs.
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)
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.
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.
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.
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.
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.
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.
Computer simulation of hypersonic flow over the Space Shuttle Orbiter
NASA Technical Reports Server (NTRS)
Inouye, M.
1977-01-01
Computer simulations of the flow field around the Space Shuttle Orbiter are described. Results of inviscid calculations are presented for the shock wave pattern and bottom centerline pressure distribution at 30 deg angle of attack. Results of viscous calculations are presented for wall pressure and heat transfer distributions for simple configurations representative of regions where shock wave-boundary layer interactions occur. The computer codes are verified by comparisons with wind-tunnel data and can be applied to flight conditions.
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.
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.
Kinetic simulation of rarefied and weakly ionized hypersonic flow fields
NASA Astrophysics Data System (ADS)
Farbar, Erin D.
When a vehicle enters the Earth's atmosphere at the very large velocities associated with Lunar and Mars return, a strong bow shock is formed in front of the vehicle. The shock heats the air to very high temperatures, causing collisions that are sufficiently energetic to produce ionized particles. As a result, a weakly ionized plasma is formed in the region between the bow shock and the vehicle surface. The presence of this plasma impedes the transport of radio frequency waves to the vehicle, causing the phenomenon known as "communications black out". The plasma also interacts with the neutral particles in the flow field, and contributes to the heat flux at the vehicle surface. Since it is difficult to characterize these flow fields using flight or ground based experiments, computational tools play an important role in the design of reentry vehicles. It is important to include the physical phenomena associated with the presence of the plasma in the computational analysis of the flow fields about these vehicles. Physical models for the plasma phenomena are investigated using a state of the art, Direct Simulation Monte Carlo (DSMC) code. Models for collisions between charged particles, plasma chemistry, and the self-induced electric field that currently exist in the literature are implemented. Using these baseline models, steady state flow field solutions are computed for the FIRE II reentry vehicle at two different trajectory points. The accuracy of each baseline plasma model is assessed in a systematic fashion, using one flight condition of the FIRE II vehicle as the test case. Experimental collision cross section data is implemented to model collisions of electrons with neutral particles. Theoretical and experimental reaction cross section data are implemented to model chemical reactions that involve electron impact, and an associative ionization reaction. One-dimensional Particle-In-Cell (PIC) routines are developed and coupled to the DSMC code, to assess the
Hypersonic flows generated by parabolic and paraboloidal shock waves
NASA Technical Reports Server (NTRS)
Schwartz, L. W.
1974-01-01
A computer algorithm has been developed to determine the blunt-body flowfields supporting symmetric parabolic and paraboloidal shock waves at infinite free-stream Mach number. Solutions are expressed in an analytic form as high-order power series, in the coordinate normal to the shock, whose coefficients can be determined exactly. Analytic continuation is provided by the use of Pade approximations. Test cases provide solutions of very high accuracy. In the axisymmetric case for gamma equals 715 the solution has been found far downstream, where it agrees with the modified blast-wave results. For plane flow, on the other hand, a limit line appears within the shock layer, a short distance past the sonic line, suggesting the presence of an imbedded shock. Local solutions in the downstream limit are discussed.
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.
Nonreactive viscous solver for hypersonic flows over recessed cone with injection
Tai, C.H.; Lee, Y.K.
1995-01-01
A nonreactive two-species flow in which the gas is injected from the downstream wall of the recess on a hypersonic sharp cone has been investigated by solving the steady full Navier-Stokes equations. The discretization methods have combined Roe`s scheme and multiblock grids for accurate calculations. The species equation has been included in order to simulate the flowfield by injecting different gases. The flow structure and cooling effect has been investigated at M(sub infinity) = 6.0. The flow structure of the recess has been shown by the perfect gas computational results to be significantly affected by injecting cool species gases. An optimal injection rate for the cooling effect has also been shown to exist on the downstream wall of the cone. The cooling effect of injecting helium has been shown to be better than nitrogen and air for both isothermal and adiabatic wall conditions. 18 refs.
On the Numerical Convergence to Steady State of Hypersonic Flows Over Bodies with Concavities
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
2002-01-01
Two recent numerical studies of hypersonic flows over bodies with concavities revealed problems with convergence to a steady state with an oft-used application of local-time-stepping. Both simulated flows showed a time-like, periodic shedding of vortices in a subsonic domain bounded by supersonic external flow although the simulations, using local-time-stepping, were not time accurate. Simple modifications to the numerical algorithm were implemented to enable implicit, first-order accurate in time simulations. Subsequent time-accurate simulations of the two test problems converged to a steady state. The baseline algorithm and modifications for temporal accuracy are described. The requirement for sub-iterations to achieve convergence is demonstrated. Failure to achieve convergence without time accuracy is conjectured to arise from temporal errors being continuously refocused into a subsonic domain.
PNS calculations for 3-D hypersonic corner flow with two turbulence models
NASA Technical Reports Server (NTRS)
Smith, Gregory E.; Liou, May-Fun; Benson, Thomas J.
1988-01-01
A three-dimensional parabolized Navier-Stokes code has been used as a testbed to investigate two turbulence models, the McDonald Camarata and Bushnell Beckwith model, in the hypersonic regime. The Bushnell Beckwith form factor correction to the McDonald Camarata mixing length model has been extended to three-dimensional flow by use of an inverse averaging of the resultant length scale contributions from each wall. Two-dimensional calculations are compared with experiment for Mach 18 helium flow over a 4-deg wedge. Corner flow calculations have been performed at Mach 11.8 for a Reynolds number of .67 x 10 to the 6th, based on the duct half-width, and a freestream stagnation temperature of 1750-deg Rankine.
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.
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.
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.
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.
Tivanov, G.; Rom, J.
1995-12-01
The stability characteristics of the reacting hypersonic flow of the fuel/oxidizer mixture in the stagnation region of a blunt body are studied. The conditions for oscillations of the combustion front are assumed to be determined mainly by the flow conditions at the stagnation region. The density at the stagnation region is assumed to be constant at hypersonic flow conditions. By assuming a simplified flow model, the time dependent flow equations, including the heat addition due to the chemical reactions, are reduced to a second-order nonlinear differential equation for the instantaneous temperature. The solutions are analyzed assuming a one-step chemical reaction with zero-order and first-order processes using dynamical systems methods. These methods are used to determine the stability boundaries in terms of the flow and chemical reaction parameters. It is shown that the zero-order reaction has nonperiodic solutions that may lead to explosion whereas the first-order and higher-order reactions may have periodic solutions indicating oscillations. The zero-order analysis also reaffirms the requirements for a minimum size blunt body for the establishment of a detonation (in agreement with classical detonation theory) and the first-order analysis indicates a minimum body size for establishment of oscillations. The oscillation frequencies are calculated using the small perturbation approximation for the temperature oscillations. These frequencies are compared with results from published data on spheres and hemisphere cylindrical bodies fired into hydrogen-oxygen and acetylene oxygen mixtures. Very good agreement is found between the measured and calculated results.
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.
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.
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 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.
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.
Flow fields and aerodynamic characteristics for hypersonic missiles with mid-fuselage inlets
NASA Technical Reports Server (NTRS)
Hunt, J. L.; Johnston, P. J.; Riebe, G. D.
1983-01-01
A study was made to quantify forebody flow fields and to evaluate aerodynamic performance trends on a matrix of fuselage shapes for the mid-inlet/bolt-on-engine class of hypersonic airbreathing missiles for the Navy's vertical box launcher. The study indicated that inlet mass flow and pressure recovery can be increased by cambering the nose and increasing the width of the fuselage at both Mach 4 acceleration and Mach 6 cruise conditions. Aerodynamic trim predictions show that the drag at zero lift at Mach 4 decreases while the L/D max at Mach 6 increases with the nose camber, although these tendencies reverse with increasing width of maximum fuselage cross section.
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.
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.
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 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.
Hypersonic engine component experiments in high heat flux, supersonic flow environment
NASA Technical Reports Server (NTRS)
Gladden, Herbert J.; Melis, Matthew E.
1993-01-01
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. Even though progress has been made in the computational understanding of fluid dynamics and the physics/chemistry of high speed flight, there is also a need for experimental facilities capable of providing a high heat flux environment for testing component concepts and verifying/calibrating these analyses. A hydrogen/oxygen rocket engine heat source was developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to fulfill this need. This 'Hot Gas Facility' is capable of providing heat fluxes up to 450 w/sq cm on flat surfaces and up to 5,000 w/sq cm at the leading edge stagnation point of a strut in a supersonic flow stream. Gas temperatures up to 3050 K can also be attained. Two recent experimental programs conducted in this facility are discussed. The objective of the first experiment 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. Macrophotographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight. The objective of the second experiment is to assess the capability of cooling a porous surface exposed to a high temperature, high velocity flow environment and to provide a heat transfer data base for a design procedure. Experimental results from transpiration cooled surfaces in a supersonic flow environment are presented.
NASA Technical Reports Server (NTRS)
Chou, Lynn Chen; Mach, Kervyn D.; Deng, Zheng-Tao; Liaw, Goang-Shin
1995-01-01
A two-dimensional computer code to solve the Burnett equations has been developed which computes the flow interaction between an exhausted plume and hypersonic external flow near the afterbody of a flight vehicle. This Burnett-2D code extends the capability of Navier-Stokes solver (RPLUS2D code) to include high-order Burnett source terms and slip-wall conditions for velocity and temperature. Higher-order Burnett viscous stress and heat flux terms are discretized using central-differencing and treated as source terms. Blocking logic is adopted in order to overcome the difficulty of grid generation. The computation of exhaust plume flow field is divided into two steps. In the first step, the thruster nozzle exit conditions are computed which generates inflow conditions in the base area near the afterbody. Results demonstrated that at high altitudes, the computations of nozzle exit conditions must include the effects of base flow since significant expansion exists in the base region. In the second step, Burnett equations were solved for exhaust plume flow field near the afterbody. The free stream conditions are set at an altitude equal to 80km and the Mach number is equal to 5.0. The preliminary results show that the plume expansion, as altitude increases, will eventually cause upstream flow separation.
NASA Astrophysics Data System (ADS)
Greenshields, Christopher J.; Reese, Jason M.
2012-07-01
This paper investigates the use of Navier-Stokes-Fourier equations with non-equilibrium boundary conditions (BCs) for simulation of rarefied hypersonic flows. It revisits a largely forgotten derivation of velocity slip and temperature jump by Patterson, based on Grad's moment method. Mach 10 flow around a cylinder and Mach 12.7 flow over a flat plate are simulated using both computational fluid dynamics using the temperature jump BCs of Patterson and Smoluchowski and the direct simulation Monte-Carlo (DSMC) method. These flows exhibit such strongly non-equilibrium behaviour that, following Patterson's analysis, they are strictly beyond the range of applicability of the BCs. Nevertheless, the results using Patterson's temperature jump BC compare quite well with the DSMC and are consistently better than those using the standard Smoluchowski temperature jump BC. One explanation for this better performance is that an assumption made by Patterson, based on the flow being only slightly non-equilibrium, introduces an additional constraint to the resulting BC model in the case of highly non-equilibrium flows.
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.
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)
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.
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.
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.
Gnoffo, P.A.; Gupta, R.N.; Shinn, J.L.
1989-02-01
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)
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.
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
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.
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.
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.
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.
Modeling of electronic excitation and radiation in non-continuum hypersonic reentry flows
NASA Astrophysics Data System (ADS)
Li, Zheng; Ozawa, Takashi; Sohn, Ilyoup; Levin, Deborah A.
2011-06-01
The modeling of hypersonic radiation in non-equilibrium, non-continuum flows is considered in the framework of the direct simulation Monte Carlo (DSMC) approach. The study explores the influence of electronic states on the flow chemistry and degree of ionization as well as the assumption that the electronic states can be described by a steady state solution to a system of rate equations of excitation, de-excitation, and radiative transfer processes. The work implements selected excited levels of atomic nitrogen and oxygen and the corresponding electron impact excitation/de-excitation and ionization processes in DSMC. The simulations show that when excitation models are included, the degree of ionization in the Stardust transitional re-entry flow increases due to additional intermediate steps to ionization. The extra ionization reactions consume the electron energy to reduce the electron temperature. The DSMC predicted excited state level populations are lower than those predicted by a quasi steady state calculation, but the differences can be understood in terms of the flow distribution functions.
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.
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.
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
An implicit multigrid algorithm for computing hypersonic, chemically reacting viscous flows
Edwards, J.R.
1996-01-01
An implicit algorithm for computing viscous flows in chemical nonequilibrium is presented. Emphasis is placed on the numerical efficiency of the time integration scheme, both in terms of periteration workload and overall convergence rate. In this context, several techniques are introduced, including a stable, O(m{sup 2}) approximate factorization of the chemical source Jacobian and implementations of V-cycle and filtered multigrid acceleration methods. A five species-seventeen reaction air model is used to calculate hypersonic viscous flow over a cylinder at conditions corresponding to flight at 5 km/s, 60 km altitude and at 11.36 km/s, 76.42 km altitude. Inviscid calculations using an eleven-species reaction mechanism including ionization are presented for a case involving 11.37 km/s flow at an altitude of 84.6 km. Comparisons among various options for the implicit treatment of the chemical source terms and among different multilevel approaches for convergence acceleration are presented for all simulations.
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.
Aerodynamic and inlet flow characteristics of several hypersonic airbreathing missile concepts
NASA Technical Reports Server (NTRS)
Dillon, J. L.; Marcum, D. C., Jr.; Johnston, P. J.; Hunt, J. L.
1980-01-01
Four conceptual hypersonic missile configurations were examined experimentally and theoretically. Two of the concepts employed twin module bottom-mounted engines and two were designed for upper surface inlets or engines with the intent of reducing the vehicle observables. The tests were conducted at Mach 6 and Reynolds numbers of 6 to 7.5 x 10 to the 6th per foot. Flow field surveys in the vicinity of the engine inlet were made on all configurations and force and moment tests were conducted on three of the vehicles. Stability and control characteristics of the bottom-mounted engine configurations which incorporated slender, low wings were dominated by strong vortices that promoted severe pitchup tendencies. The shock layer and flow quality in the vicinity of the bottom-mounted engine inlets were dependent on nose shape. The spatula-like upper surface engine concept demonstrated good performance and had uniform flow entering the engine inlet, while the upper surface inlet concept with a highly swept forebody incurred large gradients due to interactions with leading edge shocks.
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.
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.
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.
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.
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.
Hoffman, J.J.; Wong, R.S.; Bussing, T.R.; Birch, S.F.
1988-03-01
Phase I results include selection of the three components of the computational of the computational algorithm (a Navier-Stokes solution algorithm, a chemistry solution algorithm, and vectorization and parallel-processing requirements for both algorithms). Development of a nonequilibrium air-chemistry reaction model is included, as well as studies of leeside models, turbulence models, and wall catalysis effects appropriate to the hypersonic flows to be considered. Mach 20 test cases were performed using the Navier-Stokes and chemistry algorithms, and a comprehensive sensitivity study was completed for the selection of an air-chemistry model. Transport property calculations are also discussed. The components of the computational algorithm developed during Phase I will be assembled during Phase II into a unified computer code capable of accurately and efficiently calculating low-density real gas flows about hypersonic vehicles.
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
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
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 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.
Swept-slot film-cooling effectiveness in hypersonic turbulent flow
NASA Technical Reports Server (NTRS)
Hefner, J. N.; Cary, A. M., Jr.
1974-01-01
Measurement results are presented for the surface equilibrium temperature downstream of swept slots, with sonic tangential air injection into a thick hypersonic turbulent boundary layer. These results are compared with unswept slot results for cooling effectiveness.
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.
A study of boundary layer transition on outgassing cones in hypersonic flow
NASA Technical Reports Server (NTRS)
Stalmach, C. J., Jr.; Bertin, J. J.; Pope, T. C.; Mccloskey, M. H.
1971-01-01
Surface heat-transfer rates and pressures were measured at hypersonic speeds on sharp cones at zero angle of attack with and without gas injection. Using the non-injection results for reference data the effects on heating and transition location of surface roughness and injectant rate, distribution and composition were determined. The transition location was sensitive to the injectant distribution. The transition Reynolds numbers were significantly greater when the injectant distribution was constant than with a variable distribution. The measured heat-transfer distribution were also strongly dependent upon the injectant distribution. Transition Reynolds number results obtained during this program with a variable injectant distribution were correlated with a limited amount of data available for a degrading model tested in a different facility. Transitional data with constant injectant distribution were correlated with earlier results. An empirical correlation of heat-transfer reduction due to gas injection in turbulent flow was developed for both distributions tested. Several effects of mass addition on heating and transition, which have been earlier reported, were observed.
Laminar heat-transfer distributions on biconics at incidence in hypersonic-hypervelocity flows
NASA Technical Reports Server (NTRS)
Miller, C. G., III; Micol, J. R.; Gnoffo, P. A.
1984-01-01
Laminar heating distributions were measured at hypersonic-hypervelocity flow conditions on a 1.9-percent-scale model of an aeroassisted vehiclee proposed for missions to a number of planets. This vehicle is a spherically blunted, 12.84/7deg biconic with the fore-cone axis bent upward 7 deg relative to the aft-cone axis to provide selftrim capability. Also tested was a straight biconic (i.e., without nose bend) with the same nose radius and half-angles as the bent-nose biconic. These measurements were made in the Langley Expansion Tube at free-stream velocities from 4.5 to 6.9 km/sec and Mach numbers from 6.0 to 9.0 with helium, nitrogen, air, and carbon dioxide test gases. The range of calculated thermochemical equilibrium normal-shock density ratios for these four test gases was 4 to 19. Angles of attack, referenced to the aft-cone, varied from 0 to 20 deg. Heating distributions predicted with a parabolized Navier-Stokes (PNS) code were compared with measurement for helium and air test gases. Measured windward and leeward heating levels were generally underpredicted by the PNS code for both test gases, and agreement was poorer on the leeward side than on the windward side.
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.
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.
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.
Computational methods for hypersonic viscous flow over finite ellipsoid-cones at incidence
NASA Technical Reports Server (NTRS)
Li, C. P.
1985-01-01
A numerical method, which is simpler than others currently in use, is proposed for determining the full viscous flow over a finite body in hypersonic stream at high altitude. It treats the shock layer surrounding the blunt foebody and the near wake behind the base simultaneously by formulating the Navier-Stokes equations in conformal and azimuthal-angle coordinates. The computational domain is confined to the body wall, outflow surface and the bow shock, which is adjusted along the coordinate normal to the wall in the course of iterations. Because of the optimal grid and a well developed alternating direction implicit factorization technique for the governing equations, reasonably accurate results can be obtained on a 30 by 36 by 6 grid with 400 time-marching iterations. Results for body shapes belonging to the ellipsoid-cone family are compared with the experimental data for the Apollo command module and the Viking aeroshell. Validation of the method based on self-consistency is also discussed.
NASA Astrophysics Data System (ADS)
Cecil, Eric
Velocity fields are measured in the shock layer and boundary layer on a plate with a cylindrical fin immersed in a hypersonic, free jet of nitrogen, using laser-induced fluorescence (LIF) of iodine. A sheet beam from a single-mode argon laser at 514 nm is used to excite hyperfine components of the P(13), R(15) and P(48), P(103) blended rotational-vibrational lines in the B-X electronic transition for iodine seeded in the flow. The Doppler broadening and shift of these lines, and the relative rotational line strengths are determined for excitation spectra recorded in a planar grid. Using this measurement technique, estimates for iodine of the mass velocity component and kinetic temperature of translation in the direction of laser propagation, rotational temperature, and relative number density are determined at each point. Sectional planes of the flow over the body are investigated at a spatial resolution on the scale of the molecular mean-free-path in the free jet near the plate leading edge. Two directions within each plane are examined, to determine the velocity vector and to investigate translational non-equilibrium. Predictions from two direct simulation Monte Carlo computations of the flow are compared with the measurements. Large values of slip velocity and temperature jump at the plate surface are observed for iodine. Measurements and DSMC predictions indicate strong translational non-equilibrium effects for the iodine in the shock wave and the thick boundary layer on the plate, and are qualitatively consistent with a bimodal velocity distribution function. As a consequence of the ratio of molecular masses, the translational non-equilibrium of iodine is much greater than for nitrogen.
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.
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.
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.
Interaction theory of hypersonic laminar near-wake flow behind an adiabatic circular cylinder
NASA Astrophysics Data System (ADS)
Hinman, W. Schuyler; Johansen, C. T.
2015-12-01
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.
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)
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.
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.
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
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
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.
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.
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.
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.
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.
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
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.
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.
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)
Johnson, C. B.; Marcum, D. C., Jr.
1974-01-01
Flow angularity and static pressure measurements have been made on the lower surface of nine forebody models that simulate the bottom forward surface of a hypersonic aircraft. Measurements were made in an area of the forebody that represents the location of an inlet of a scramjet engine. A parametric variation of the forebody surface investigated the effect of: (1) spanwise curvature; (2) longitudinal curvature; and (3) planform shape on both flow angularity and static pressure distribution. Results of each of the three parametric variations of geometry were compared to those for the same flat delta forebody. Spanwise curvature results showed that a concave shape and the flat delta had the lowest flow angularity and lowest rate of increase in flow angularity with angle of attack. Longitudinal curvature results showed a convex surface to give the better flow at the higher angles of attack. The better of the two planform shapes tested was a convex elliptical shape. Limited flow field calculations were made at angles of attack using a three dimensional, method-of-characteristics program. In general, at all angles of attack there was agreement between data and theory.
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.
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)
Laufer, Gabriel; Mckenzie, Robert L.; Fletcher, Douglas G.
1990-01-01
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 less than 2 percent, 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.
NASA Technical Reports Server (NTRS)
Laufer, Gabriel; Fletcher, Douglas G.; Mckenzie, Robert L.
1990-01-01
Laser-induced fluorescence in oxygen, in combination with Raman scattering, is shown to be an accurate means by which temperature, density, and their fluctuations due 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 uncertainty in the temperature and density measurements can be less than 2 and 3 percent, respectively, 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.
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.
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. PMID:19791885
A kinetic-theory approach for computing chemical-reaction rates in upper-atmosphere hypersonic flows
NASA Astrophysics Data System (ADS)
Gallis, Michael A.; Bond, Ryan B.; Torczynski, John R.
2009-09-01
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.
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
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
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.
A comparative study of turbulence models in predicting hypersonic inlet flows
NASA Technical Reports Server (NTRS)
Kapoor, Kamlesh
1993-01-01
A computational study has been conducted to evaluate the performance of various turbulence models. The NASA P8 inlet, which represents cruise condition of a typical hypersonic air-breathing vehicle, was selected as a test case for the study; the PARC2D code, which solves the full two dimensional Reynolds-averaged Navier-Stokes equations, was used. Results are presented for a total of six versions of zero- and two-equation turbulence models. Zero-equation models tested are the Baldwin-Lomax model, the Thomas model, and a combination of the two. Two-equation models tested are low-Reynolds number models (the Chien model and the Speziale model) and a high-Reynolds number model (the Launder and Spalding model).
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.
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 Technical Reports Server (NTRS)
Povinelli, Louis A.
1990-01-01
An overview is given of research activity on the application of computational fluid dynamics (CDF) for hypersonic propulsion systems. After the initial consideration of the highly integrated nature of air-breathing hypersonic engines and airframe, attention is directed toward computations carried out for the components of the engine. A generic inlet configuration is considered in order to demonstrate the highly three dimensional viscous flow behavior occurring within rectangular inlets. Reacting flow computations for simple jet injection as well as for more complex combustion chambers are then discussed in order to show the capability of viscous finite rate chemical reaction computer simulations. Finally, the nozzle flow fields are demonstrated, showing the existence of complex shear layers and shock structure in the exhaust plume. The general issues associated with code validation as well as the specific issue associated with the use of CFD for design are discussed. A prognosis for the success of CFD in the design of future propulsion systems is offered.
NASA Technical Reports Server (NTRS)
Povinelli, Louis A.
1991-01-01
An overview is given of research activity on the application of computational fluid dynamics (CDF) for hypersonic propulsion systems. After the initial consideration of the highly integrated nature of air-breathing hypersonic engines and airframe, attention is directed toward computations carried out for the components of the engine. A generic inlet configuration is considered in order to demonstrate the highly three dimensional viscous flow behavior occurring within rectangular inlets. Reacting flow computations for simple jet injection as well as for more complex combustion chambers are then discussed in order to show the capability of viscous finite rate chemical reaction computer simulations. Finally, the nozzle flow fields are demonstrated, showing the existence of complex shear layers and shock structure in the exhaust plume. The general issues associated with code validation as well as the specific issue associated with the use of CFD for design are discussed. A prognosis for the success of CFD in the design of future propulsion systems is offered.
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.
NASA Technical Reports Server (NTRS)
Xiques, K. E.; Rawlinson, E. G.; Stalnaker, J. F.; Spradley, L. W.
1982-01-01
A three-dimensional computational technique was used to obtain flowfield solutions to the Euler equations over selected hypersonic missile configurations. The General Interpolants Method (GIM) computer code was used with interpolation functions in an algebraic approach to generate a discrete computational grid for each configuration. The spatial marching version of the GIM code, which treats the parabolized Navier-Stokes (PNS) equations or the Euler equations with a shock capturing, 'MacCormack-like' scheme, was used to advance the solution hyperbolically over each configuration. The inviscid flowfield solutions over the two three-dimensional missile configurations, calculated using the GIM hyperbolic scheme, are presented here. The flow field over a wing/body configuration at zero degree angle of attack is presented. Flow over the fuselage of a tactical missile, termed the TAME 10, at both zero degree and 7.5 degree angles of attack is presented. In addition, an inviscid, two-dimensional analysis of an inlet configuration designed to mount on the TAME 10 is included. Contour maps of velocity and pressure are included for each configuration. Comparison of calculation and data show good agreement.
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.
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 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.
Electromagnetically Activated Hypersonic Ducts
NASA Astrophysics Data System (ADS)
MacLeod, C.
This paper explores the use of Electromagnetic Radiation as an alternative to combustion in Scramjet-like hypersonic engines. The radiation is absorbed by the flow, heating it and thereby providing an alternative to the heat derived from combustion in the Scramjet. The advantages and disadvantages of this system are explored and theoretical results are presented illustrating typical radiation pathlengths at different frequencies. Suggestions for further theoretical and practical work are also made.
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. PMID:14658473
Predictive and reinforcement learning for magneto-hydrodynamic control of hypersonic flows
NASA Astrophysics Data System (ADS)
Kulkarni, Nilesh Vijay
Increasing needs for autonomy in future aerospace systems and immense progress in computing technology have motivated the development of on-line adaptive control techniques to account for modeling errors, changes in system dynamics, and faults occurring during system operation. After extensive treatment of the inner-loop adaptive control dealing mainly with stable adaptation towards desired transient behavior, adaptive optimal control has started receiving attention in literature. Motivated by the problem of optimal control of the magneto-hydrodynamic (MHD) generator at the inlet of the scramjet engine of a hypersonic flight vehicle, this thesis treats the general problem of efficiently combining off-line and on-line optimal control methods. The predictive control approach is chosen as the off-line method for designing optimal controllers using all the existing system knowledge. This controller is then adapted on-line using policy-iteration-based Q-learning, which is a stable model-free reinforcement learning approach. The combined approach is first illustrated in the optimal control of linear systems, which helps in the analysis as well as the validation of the method. A novel neural-networks-based parametric predictive control approach is then designed for the off-line optimal control of non-linear systems. The off-line approach is illustrated by applications to aircraft and spacecraft systems. This is followed by an extensive treatment of the off-line optimal control of the MHD generator using this neuro-control approach. On-line adaptation of the controller is implemented using several novel schemes derived from the policy-iteration-based Q-learning. The implementation results demonstrate the success of these on-line algorithms for adapting towards modeling errors in the off-line design.
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.
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. PMID:17068552
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.
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.
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
PNS predictions for supersonic/hypersonic flows over finned missile configurations
NASA Technical Reports Server (NTRS)
Bhutta, Bilal A.; Lewis, Clark H.
1992-01-01
Finned missile design entails accurate and computationally fast numerical techniques for predicting viscous flows over complex lifting configurations at small to moderate angles of attack and over Mach 3 to 15; these flows are often characterized by strong embedded shocks, so that numerical algorithms are also required to capture embedded shocks. The recent real-gas Flux Vector Splitting technique is here extended to investigate the Mach 3 flow over a typical finned missile configuration with/without side fin deflections. Elliptic grid-generation techniques for Mach 15 flows are shown to be inadequate for Mach 3 flows over finned configurations and need to be modified. Fin-deflection studies indicate that even small amounts of missile fin deflection can substantially modify vehicle aerodynamics. This 3D parabolized Navier-Stokes scheme is also extended into an efficient embedded algorithm for studying small axially separated flow regions due to strong fin and control surface deflections.
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.
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.
A Study of Hypersonic Compression-Corner Flow at High Reynolds Numbers
NASA Astrophysics Data System (ADS)
Vetlutsky, V. N.; Ganimedov, V. L.
2002-07-01
To numerically solve the problem for the ramp flow at high Reynolds number, it is justified to use the classical Prandtl model and to part the flow into an inviscid region and a thin boundary layer. This model is based on a rigorous asymptotic theory and describes the flow the more accurately, the higher the Reynolds number is. Since a flow separation in a compression corner may occur, the present work deals with unsteady boundary-layer equations. The approximation error in discretization of boundary-layer equations is always smaller than the value of viscid terms since the Reynolds number can be eliminated from the equations by a suitable substitution of variables. Besides, the requirement for the sufficient number of points in the cross-flow direction can be fulfilled in the case of middle-class computers since the solution is being built in a narrow near-wall region.
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.
A hybridized method for computing high-Reynolds-number hypersonic flow about blunt bodies
NASA Technical Reports Server (NTRS)
Weilmuenster, K. J.; Hamilton, H. H., II
1979-01-01
A hybridized method for computing the flow about blunt bodies is presented. In this method the flow field is split into its viscid and inviscid parts. The forebody flow field about a parabolic body is computed. For the viscous solution, the Navier-Stokes equations are solved on orthogonal parabolic coordinates using explicit finite differencing. The inviscid flow is determined by using a Moretti type scheme in which the Euler equations are solved, using explicit finite differences, on a nonorthogonal coordinate system which uses the bow shock as an outer boundary. The two solutions are coupled along a common data line and are marched together in time until a converged solution is obtained. Computed results, when compared with experimental and analytical results, indicate the method works well over a wide range of Reynolds numbers and Mach numbers.
Experimental investigation of the magnetohydrodynamic parachute effect in a hypersonic air flow
NASA Astrophysics Data System (ADS)
Fomichev, V. P.; Yadrenkin, M. A.
2013-01-01
New data on experimental implementation of the magnetohydrodynamic (MHD) parachute configuration in an air flow with Mach number M = 6 about a flat plate are considered. It is shown that MHD interaction near a flat plate may transform an attached oblique shock wave into a normal detached one, which considerably extends the area of body-incoming flow interaction. This effect can be employed in optimizing return space vehicle deceleration conditions in the upper atmosphere.
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.
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.
1993-01-01
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, and the activity in the NASA Ames Research Center's Aerothermodynamics Branch is described. 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 flowfield simulation codes are discussed. These models were 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 is sparse and 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.
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.
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.
Simulation of Hypersonic CO2 Flows in a Range of Academic and Industrial Facilities
NASA Astrophysics Data System (ADS)
Beck, J. C.; Smith, A. J.
2009-01-01
Recent studies in high enthalpy carbon dioxide flows have provided the opportunity to validate computational tools against a wide range of data from experiments in a wide range of facilities, from a classical shock tube to the industrial facility in which the ExoMars heatshield will be qualified. Computations of both nozzle flows and stagnation tests have been performed using the thermochemical non-equilibrium Navier-Stokes code TINA, and the shock tube has been simulated using the SHOCK2 code. In general, good agreement with the experimental data has been achieved. Thus, this work provides increased confidence in the thermochemical modelling of carbon dioxide flows and the applicability of tools used in the design of entry vehicles for Mars.
Monte Carlo computation of nonequilibrium flow in a hypersonic iodine wind tunnel
Boyd, I.D.; Pham-van-Diep, G.C.; Muntz, E.P. )
1994-05-01
The nonequilibrium flow formed by the interaction of a freejet of iodine vapor impinging on a blunt body is investigated using numerical and experimental techniques. The computational approach employs the direct simulation Monte Carlo method. The experimental measurements consist of rotational temperature obtained along the flow axis and include portions of both the freejet expansion and blunt-body shock for four different stagnation conditions. Direct comparisons of the numerical results and the experimental data are quite successful at moderate temperatures. Hence, the rotational collision time of iodine is estimated in the temperature range of 100-500 K. At higher temperatures, the agreement between simulation and measurement is less satisfactory. This demonstrates the requirement for the development of a more detailed approach to simulating rotational nonequilibrium in high-temperature flows of diatomic species. 20 refs.
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.
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.
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.
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.
Three dimensional PNS solutions of hypersonic internal flows with equilibrium chemistry
NASA Technical Reports Server (NTRS)
Liou, May-Fun
1989-01-01
An implicit procedure for solving parabolized Navier-Stokes equations under the assumption of a general equation of state for a gas in chemical equilibrium is given. A general and consistent approach for the evaluation of Jacobian matrices in the implicit operator avoids the use of unnecessary auxiliary quantities and approximations, and leads to a simple expression. Applications to two- and three-dimensional flow problems show efficiency in computer time and economy in storage.
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.
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. PMID:25353565
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 Astrophysics Data System (ADS)
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 N2-N interactions. The coarse-grained model is constructed by lumping the rovibrational energy levels of the N2 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.
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.
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)
Whiting, Ellis E.
1990-01-01
Future space vehicles returning from distant missions or high earth orbits may enter the upper regions of the atmosphere and use aerodynamic drag to reduce their velocity before they skip out of the atmosphere and enter low earth orbit. The Aeroassist Flight Experiment (AFE) is designed to explore the special problems encountered in such entries. A computer code was developed to calculate the radiative transport along line-or-sight in the general 3-D flow field about an arbitrary entry vehicle, if the temperatures and species concentrations along the line-of-sight are known. The radiative heating calculation at the stagnation point of the AFE vehicle along the entry trajectory was performed, including a detailed line-by-line accounting of the radiative transport in the vacuum ultraviolet (below 200 nm) by the atomic N and O lines. A method was developed for making measurements of the haze particles in the Titan atmosphere above 200 km altitude. Several other tasks of a continuing nature, to improve the technical ability to calculate the nonequilibrium gas dynamic flow field and radiative heating of entry vehicles, were completed or advanced.
NASA Astrophysics Data System (ADS)
Li, Zheng; Sohn, Ilyoup; Levin, Deborah A.; Modest, Michael F.
2011-05-01
The current work implemented excited levels of atomic N and corresponding electron impact excitation/de-excitation and ionization processes in DSMC. Results show that when excitation models are included, the Stardust 68.9 km re-entry flow has an observable change in the ion number densities and electron temperature. Adding in the excited levels of atoms improves the degree of ionization by providing additional intermediate steps to ionization. The extra ionization reactions consume the electron energy and reduce the electron temperature. The DSMC results of number densities of excited levels are lower than the prediction of quasi steady state calculation. Comparison of radiation calculations using electronic excited populations from DSMC and QSS indicates that, at the stagnation point, there is about 20% difference of the radiative heat flux between DSMC and QSS.
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.
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
Molding Phonon Flow with Symmetry: Rational Design of Hypersonic Phononic Crystals
NASA Astrophysics Data System (ADS)
Koh, Cheong Yang; Thomas, Edwin L.
2009-03-01
Phononic crystals structured at appropriate length scales allow control over the flow of phonons, leading to new possibilities in applications such as heat-management, sound isolation and even energy transfer and conversion. Symmetry provides a unified framework for the interpretation 1D to 3D phononic band structures, allowing utilization of a common set of principles for designing band structures of phononic crystals as well as actual purposeful defects such as waveguide location and boundary termination in finite devices. In this work, we explore the band structure properties of phononic crystals with non-symmorphic space groups, as well as those having quasi-crystalline approximants. We demonstrate gap opening abilities from both anti-crossing and Bragg scattering, as well as unique features like ``sticking'' bands. Symmetry concepts are also powerful means to tune the density of states of the structures. Importantly, we fabricate various theoretical designs and measure their experimental dispersion diagrams for comparison with theoretical calculation. This affords an elegant approach toward a design blueprint for fabricating phononic structures for applications such as opto-acoustic coupling.
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.
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.
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}).
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}).
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.
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.
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.
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.
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.
Extension of hypersonic, high-incidence, slender-body similarity
NASA Technical Reports Server (NTRS)
Barnwell, Richard W.
1987-01-01
The Sychev (1960) analysis for inviscid hypersonic flow past slender bodies at large angle of attack is shown to be applicable to all slender-body flows whose crossflow Mach numbers are greater than sonic; it is therefore not restricted to flows with hypersonic crossflow Mach number values, as indicated elsewhere in the literature. It is also noted that the Sychev similarity applies to a number of slender-body flows with subsonic crossflow Mach numbers, including incompressible flow.
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.
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.
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.
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 Astrophysics Data System (ADS)
Kumar, C. S.; Singh, T.; Reddy, K. P. J.
2014-12-01
Heat transfer rate and pressure measurements were made upstream of surface protuberances on a flat plate and a sharp cone subjected to hypersonic flow in a conventional shock tunnel. Heat flux was measured using platinum thin-film sensors deposited on macor substrate and the pressure measurements were made using fast acting piezoelectric sensors. A distinctive hot spot with highest heat flux was obtained near the foot of the protuberance due to heavy vortex activity in the recirculating region. Schlieren flow visualization was used to capture the shock structures and the separation distance ahead of the protrusions was quantitatively measured for varying protuberance heights. A computational analysis was conducted on the flat plate model using commercial computational fluid dynamics software and the obtained trends of heat flux and pressure were compared with the experimental observation. Experiments were also conducted by physically disturbing the laminar boundary layer to check its effect on the magnitude of the hot spot heat flux. In addition to air, argon was also used as test gas so that the Reynolds number can be varied.
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 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.
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.
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.
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.
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.
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.
Alfyorov, V.I.; Yegorov, I.V.; Shcherbakov, G.I.
1995-12-31
The paper contains the results of applying a hypervelocity MHD-gas acceleration wind tunnel to investigations of flows over bodies. Consideration is given to the conditions of re producing gas dynamic and thermochemical flow parameters as applied to different types of tests: pressure and heat flux distributions, determination of shock wave positions and shapes. The measured heat fluxes towards the leading edge of swept wings are presented for sweep angles ranging from 0{degrees} to 60{degrees} at a flow velocity of U{approximately}6000 m/s. An appreciable influence of the surface nonequilibrium and catalyticity on their values is indicated. Possible investigations of flows over bodies at ultra high heat fluxes, q {approximately} 10 kW/m{sup 2} are discussed. The results of applying the facility to the verification of calculation codes and thermodynamic gas models are analyzed for flows over a hemisphere, a cone and a wedge. The calculated and measured surface pressure distributions are in good agreement for a hemisphere and satisfactory for a cone and a wedge. The shock wave positions and shapes are compared. It is shown that respective gas glow is impossible to use for this purpose.
Uncertainty Propagation in Hypersonic Vehicle Aerothermoelastic Analysis
NASA Astrophysics Data System (ADS)
Lamorte, Nicolas Etienne
Hypersonic vehicles face a challenging flight environment. The aerothermoelastic analysis of its components requires numerous simplifying approximations. Identifying and quantifying the effect of uncertainties pushes the limits of the existing deterministic models, and is pursued in this work. An uncertainty quantification framework is used to propagate the effects of identified uncertainties on the stability margins and performance of the different systems considered. First, the aeroelastic stability of a typical section representative of a control surface on a hypersonic vehicle is examined. Variability in the uncoupled natural frequencies of the system is modeled to mimic the effect of aerodynamic heating. Next, the stability of an aerodynamically heated panel representing a component of the skin of a generic hypersonic vehicle is considered. Uncertainty in the location of transition from laminar to turbulent flow and the heat flux prediction is quantified using CFD. In both cases significant reductions of the stability margins are observed. A loosely coupled airframe--integrated scramjet engine is considered next. The elongated body and cowl of the engine flow path are subject to harsh aerothermodynamic loading which causes it to deform. Uncertainty associated with deformation prediction is propagated to the engine performance analysis. The cowl deformation is the main contributor to the sensitivity of the propulsion system performance. Finally, a framework for aerothermoelastic stability boundary calculation for hypersonic vehicles using CFD is developed. The usage of CFD enables one to consider different turbulence conditions, laminar or turbulent, and different models of the air mixture, in particular real gas model which accounts for dissociation of molecules at high temperature. The system is found to be sensitive to turbulence modeling as well as the location of the transition from laminar to turbulent flow. Real gas effects play a minor role 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
Pegasus hypersonic flight research
NASA Technical Reports Server (NTRS)
Curry, Robert E.; Meyer, Robert R., Jr.; Budd, Gerald D.
1992-01-01
Hypersonic aeronautics research using the Pegasus air-launched space booster is described. Two areas are discussed in the paper: previously obtained results from Pegasus flights 1 and 2, and plans for future programs. Proposed future research includes boundary-layer transition studies on the airplane-like first stage and also use of the complete Pegasus launch system to boost a research vehicle to hypersonic speeds. Pegasus flight 1 and 2 measurements were used to evaluate the results of several analytical aerodynamic design tools applied during the development of the vehicle as well as to develop hypersonic flight-test techniques. These data indicated that the aerodynamic design approach for Pegasus was adequate and showed that acceptable margins were available. Additionally, the correlations provide insight into the capabilities of these analytical tools for more complex vehicles in which design margins may be more stringent. Near-term plans to conduct hypersonic boundary-layer transition studies are discussed. These plans involve the use of a smooth metallic glove at about the mid-span of the wing. Longer-term opportunities are proposed which identify advantages of the Pegasus launch system to boost large-scale research vehicles to the real-gas hypersonic flight regime.
Model formulation of non-equilibrium gas radiation for hypersonic flight vehicles
NASA Technical Reports Server (NTRS)
Chang, Ing
1989-01-01
Several radiation models for low density nonequilibrium hypersonic flow are studied. It is proposed that these models should be tested by the 3-D VRFL code developed at NASA/JSC. A modified and optimized radiation model may be obtained from the testing. Then, the current VRFL code could be expanded to solve hypersonic flow problems with nonequilibrium thermal radiation.
Inviscid-viscous interaction on triple-deck scales in a hypersonic flow with strong wall cooling
NASA Technical Reports Server (NTRS)
Brown, S. N.; Cheng, H. K.; Lee, C. J.
1990-01-01
The influence of wall temperatures on the flow structure in a region near a laminar separation is presently studied on the triple-deck scales, for the case of a high supersonic flow's inviscid-viscous interactions. A critical wall-temperature range is identified whose lower deck pressure-displacement relation departs from that of the classical formulation; below it, the pressure-displacement relation undergoes still greater transformations in conjunction with drastic scale-changes in the triple deck. The reduced lower-deck problem falls into supercritical, transcritical, and subcritical domains. A computational study is conducted for the compressive free-interaction solutions, and solutions are obtained for a sharp-corner ramp in the three wall-temperature ranges.
Rarefaction Effects in Hypersonic Aerodynamics
NASA Astrophysics Data System (ADS)
Riabov, Vladimir V.
2011-05-01
The Direct Simulation Monte-Carlo (DSMC) technique is used for numerical analysis of rarefied-gas hypersonic flows near a blunt plate, wedge, two side-by-side plates, disk, torus, and rotating cylinder. The role of various similarity parameters (Knudsen and Mach numbers, geometrical and temperature factors, specific heat ratios, and others) in aerodynamics of the probes is studied. Important kinetic effects that are specific for the transition flow regime have been found: non-monotonic lift and drag of plates, strong repulsive force between side-by-side plates and cylinders, dependence of drag on torus radii ratio, and the reverse Magnus effect on the lift of a rotating cylinder. The numerical results are in a good agreement with experimental data, which were obtained in a vacuum chamber at low and moderate Knudsen numbers from 0.01 to 10.
NASA Technical Reports Server (NTRS)
Palmer, Grant; Venkatapathy, Ethiraj
1993-01-01
Three solution algorithms, explicit underrelaxation, point implicit, and lower upper symmetric Gauss-Seidel (LUSGS), are used to compute nonequilibrium flow around the Apollo 4 return capsule at 62 km altitude. 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, 23, and 30, the LUSGS method produces an eight order of magnitude drop in the L2 norm of the energy residual in 1/3 to 1/2 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 23 and above. At Mach 40 the performance of the LUSGS algorithm deteriorates to the point it is out-performed by the point implicit method. The effects of the viscous terms are investigated. Grid dependency questions are explored.
Integrated numerical methods for hypersonic aircraft cooling systems analysis
NASA Technical Reports Server (NTRS)
Petley, Dennis H.; Jones, Stuart C.; Dziedzic, William M.
1992-01-01
Numerical methods have been developed for the analysis of hypersonic aircraft cooling systems. A general purpose finite difference thermal analysis code is used to determine areas which must be cooled. Complex cooling networks of series and parallel flow can be analyzed using a finite difference computer program. Both internal fluid flow and heat transfer are analyzed, because increased heat flow causes a decrease in the flow of the coolant. The steady state solution is a successive point iterative method. The transient analysis uses implicit forward-backward differencing. Several examples of the use of the program in studies of hypersonic aircraft and rockets are provided.
Holden, M.S.; Bergman, R.; Harvey, J.; Duryea, G.; Moselle, J.
1988-12-02
The first of these 2 studies examined the detailed structure of the hypersonic boundary layer over a large cone/flare configuration. Emphasis was on development and use of instrumentation with which to obtain flow-field measurements of the mean and fluctuating properties of the attached and separated shear layers. Development and use of holographic interferometry and electron-beam techniques in the high Mach number and Reynolds number environment developed in the shock tunnel are described. In the second study, detailed measurements of heat transfer, pressure, and skin friction were made on a unique 'blowing and roughness' model constructed to simulate the aerothermal phenomena associated with a rough ablating maneuverable reentry vehicle. In this study emphasis was placed on development and use of unique heat transfer and skin-friction instrumentation to obtain measurements of the combined effects of blowing and roughness and to understand how such effects influence boundary-layer separation in regions of shock wave/boundary layer interaction. Each focused around providing information with which to construct and evaluate the modeling required in time-averaged Navier-Stokes equations to predict the structure of compressible hypersonic boundary layers in regions of strong pressure gradient, shock-wave/boundary-layer interaction and flow separation over smooth, rough, and ablating surfaces.
NASA Astrophysics Data System (ADS)
Bradley, Marty Keith
A new parallel implicit adaptive mesh refinement (AMR) algorithm is developed for the prediction of unsteady behaviour of laminar flames. The scheme is applied to the solution of the system of partial-differential equations governing time-dependent, two- and three-dimensional, compressible laminar flows for reactive thermally perfect gaseous mixtures. A high-resolution finite-volume spatial discretization procedure is used to solve the conservation form of these equations on body-fitted multi-block hexahedral meshes. A local preconditioning technique is used to remove numerical stiffness and maintain solution accuracy for low-Mach-number, nearly incompressible flows. A flexible block-based octree data structure has been developed and is used to facilitate automatic solution-directed mesh adaptation according to physics-based refinement criteria. The data structure also enables an efficient and scalable parallel implementation via domain decomposition. The parallel implicit formulation makes use of a dual-time-stepping like approach with an implicit second-order backward discretization of the physical time, in which a Jacobian-free inexact Newton method with a preconditioned generalized minimal residual (GMRES) algorithm is used to solve the system of nonlinear algebraic equations arising from the temporal and spatial discretization procedures. An additive Schwarz global preconditioner is used in conjunction with block incomplete LU type local preconditioners for each sub-domain. The Schwarz preconditioning and block-based data structure readily allow efficient and scalable parallel implementations of the implicit AMR approach on distributed-memory multi-processor architectures. The scheme was applied to solutions of steady and unsteady laminar diffusion and premixed methane-air combustion and was found to accurately predict key flame characteristics. For a premixed flame under terrestrial gravity, the scheme accurately predicted the frequency of the natural
NASA Technical Reports Server (NTRS)
Creager, Marcus O.
1959-01-01
An investigation of the effects of variation of leading-edge sweep and surface inclination on the flow over blunt flat plates was conducted at Mach numbers of 4 and 5.7 at free-stream Reynolds numbers per inch of 6,600 and 20,000, respectively. Surface pressures were measured on a flat plate blunted by a semicylindrical leading edge over a range of sweep angles from 0 deg to 60 deg and a range of surface inclinations from -10 deg to +10 deg. The surface pressures were predicted within an average error of +/- 8 percent by a combination of blast-wave and boundary-layer theory extended herein to include effects of sweep and surface inclination. This combination applied equally well to similar data of other investigations. The local Reynolds number per inch was found to be lower than the free-stream Reynolds number per inch. The reduction in local Reynolds number was mitigated by increasing the sweep of the leading edge. Boundary-layer thickness and shock-wave shape were changed little by the sweep of the leading edge.
Hypersonic Materials and Structures
NASA Technical Reports Server (NTRS)
Glass, David E.
2016-01-01
Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this presentation is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components.
Hypersonic propulsion. [supersonic combustion ramjet engines
NASA Technical Reports Server (NTRS)
Beach, H. L., Jr.
1979-01-01
Research on hydrogen fueled scramjet engines for hypersonic flight is reviewed. Component developments, computational methods, and preliminary ground tests of subscale scramjet engine modules at Mach 4 and 7 are emphasized. Airframe integration, structures, and flow diagnostics are also discussed. It is shown that mixed-mode perpendicular and parallel fuel injection controls heat release over a wide Mach range and the fixed geometry inlet gives good performance over a wide range of Mach numbers.
Mcdaniel, J.C.; Fletcher, D.G.; Hartfield, R.J.; Hollo, S.D. NASA, Ames Research Center, Moffett Field, CA )
1991-12-01
A spatially-complete data set of the important primitive flow variables is presented for the complex, nonreacting, 3D unit combustor flow field employing transverse injection into a Mach 2 flow behind a rearward-facing step. A unique wind tunnel facility providing the capability for iodine seeding was built specifically for these measurements. Two optical techniques based on laser-induced-iodine fluorescence were developed and utilized for nonintrusive, in situ flow field measurements. LDA provided both mean and fluctuating velocity component measurements. A thermographic phosphor wall temperature measurement technique was developed and employed. Data from the 2D flow over a rearward-facing step and the complex 3D mixing flow with injection are reported. 25 refs.
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.
NASA Astrophysics Data System (ADS)
Andrade, R. P. G.; Noronha, J.; Denicol, Gabriel S.
2014-08-01
In this paper we investigate how the energy and momentum deposited by partonic dijets in the quark-gluon plasma (QGP) may affect the direct, elliptic, and triangular flow of low (and intermediate) pT hadrons in central Au +Au collisions at the BNL Relativistic Heavy Ion Collider. The dijets are modeled as external sources in the energy-momentum conservation equations for hydrodynamics, which are solved on an event-by-event basis within the ideal-fluid approximation. We focus our investigation at midrapidity and solve the hydrodynamic equations by imposing boost invariance. Differential anisotropic flow coefficients for pT≳1GeV are found to be significantly enhanced if the dijets deposit on average more than 12 GeV in the QGP (or more than 6 GeV per jet). Because this extra energy and momentum added to the medium perturbs the geometry-induced hydrodynamic expansion, the correlation between the v2 and v3 coefficients (for pT≳1GeV) and their corresponding initial eccentricities are considerably weakened. In addition, we argue that the extra amount of direct flow induced by dijets may be quantified by comparing the azimuthal dependence of dihadron correlations in dijet events with the corresponding quantity obtained in events without dijets. This comparison could be used to give a rough estimate of the magnitude of the effective coupling between the jets and the medium.
Experiments in hand-operated, hypersonic shock tunnel facility
NASA Astrophysics Data System (ADS)
Sudhiesh Kumar, Chintoo; Reddy, K. P. J.
2015-12-01
Experiments were conducted using the newly developed table-top, hand-operated hypersonic shock tunnel, otherwise known as the Reddy hypersonic shock tunnel. This novel instrument uses only manual force to generate the shock wave in the shock tube, and is designed to generate a freestream flow of Mach 6.5 in the test section. The flow was characterized using stagnation point pressure measurements made using fast-acting piezoelectric transducers. Schlieren visualization was also carried out to capture the bow shock in front of a hemispherical body placed in the flow. Freestream Mach numbers estimated at various points in the test section showed that for a minimum diameter of 46 mm within the test section, the value did not vary by more than 3 % along any cross-sectional plane. The results of the experiments presented here indicate that the device may be successfully employed for basic hypersonic research activities at the university level.
Hypersonic propulsion research
NASA Technical Reports Server (NTRS)
Northam, G. Burton
1990-01-01
The development of technology for the modular airframe integrated scramjet has been the focus of hypersonic propulsion research for several years. An part of this research, a variety of inlet concepts have been explored and characterized. The emphasis of the inlet program has been the development of the short (light weight), fixed geometry, side wall compression inlets that operate efficiently over a wide Mach number range. As hypersonic combustion tunnels were developed, programs to study the parameters controlling fuel mixing and combustion with single and multiple strut models were conducted using direct connect test techniques. These various tests supported the design of subscale engine test hardware that integrated inlet and combustor technology and allowed the study of the effect of heat release on thrust and combustor/inlet interaction. A number of subscale engine tests have shown predicted performance levels at Mach 4 and 7 simulated flight conditions. A few of the highlights from this research program are summarized.
Hypersonic propulsion research
NASA Technical Reports Server (NTRS)
Northam, G. Burton
1987-01-01
The NASA Langley Research Center has conducted hypersonic propulsion research since the 1960s. A variety of inlet concepts were explored and characterized. The emphasis of the inlet program was the development of the short (light weight), fixed geometry, side-wall-compression inlets that operate efficiently over a wide Mach number range. As hypersonic combustion tunnels were developed, programs to study the parameters controlling fuel mixing and combustion with single and multiple strut models were conducted using direct connect test techniques. These various tests supported the design of subscale engine test hardware that integrated inlet and combustor technology and allowed the study of the effect of heat release on thrust and combustor/inlet interaction. A number of subscale engine tests have demonstrated predicted performance levels at Mach 4 and 7 simulated flight conditions.
NASA Technical Reports Server (NTRS)
Kniskern, Marc W.
1990-01-01
The thermal effects of simulant gas injection and aerodynamic heating at the model's surface on the measurements of a non-watercooled, flow through balance were investigated. A stainless steel model of a hypersonic air breathing propulsion cruise missile concept (HAPCM-50) was used to evaluate this balance. The tests were conducted in the 20-inch Mach 6 wind tunnel at NASA-Langley. The balance thermal effects were evaluated at freestream Reynolds numbers ranging from .5 to 7 x 10(exp 6) ft and angles of attack between -3.5 to 5 deg at Mach 6. The injection gases considered included cold air, hot air, and a mixture of 50 percent Argon and 50 percent Freon-12. The stagnation temperatures of the cold air, hot air, and Ar-Fr(12) reached 111, 214, and 283 F, respectively within the balance. A bakelite sleeve was inserted into the inner tube of the balance to minimize the thermal effects of these injection gases. Throughout the tests, the normal force, side force, yaw moment, roll moment, and pitching moment balance measurements were unaffected by the balance thermal effects of the injection gases and the wind tunnel flow. However, the axial force (AF) measurement was significantly affected by balance heating. The average zero shifts in the AF measurements were 1.9, 3.8, and 5.9 percent for cold air, hot air, and Ar-Fr(12) injection, respectively. The AF measurements decreased throughout these tests which lasted from 70 to 110 seconds. During the cold air injection tests, the AF measurements were accurate up to at least ten seconds after the model was injected into the wind tunnel test section. For the hot air and Ar-Fr(12) tests, the AF measurements were accurate up to at least five seconds after model injection.
A Numerical Study of Hypersonic Forebody/Inlet Integration Problem
NASA Technical Reports Server (NTRS)
Kumar, Ajay
1991-01-01
A numerical study of hypersonic forebody/inlet integration problem is presented in the form of the view-graphs. The following topics are covered: physical/chemical modeling; solution procedure; flow conditions; mass flow rate at inlet face; heating and skin friction loads; 3-D forebogy/inlet integration model; and sensitivity studies.
Combined LAURA-UPS hypersonic solution procedure
NASA Technical Reports Server (NTRS)
Wood, William A.; Thompson, Richard A.
1993-01-01
A combined solution procedure for hypersonic flowfields around blunted slender bodies was implemented using a thin-layer Navier-Stokes code (LAURA) in the nose region and a parabolized Navier-Stokes code (UPS) on the after body region. Perfect gas, equilibrium air, and non-equilibrium air solutions to sharp cones and a sharp wedge were obtained using UPS alone as a preliminary step. Surface heating rates are presented for two slender bodies with blunted noses, having used LAURA to provide a starting solution to UPS downstream of the sonic line. These are an 8 deg sphere-cone in Mach 5, perfect gas, laminar flow at 0 and 4 deg angles of attack and the Reentry F body at Mach 20, 80,000 ft equilibrium gas conditions for 0 and 0.14 deg angles of attack. The results indicate that this procedure is a timely and accurate method for obtaining aerothermodynamic predictions on slender hypersonic vehicles.
NASA Technical Reports Server (NTRS)
Griffith, Wayland C.
1989-01-01
Possible experimental facilities appropriate to a university environment that could make meaningful contributions to the solution of problems in hypersonic aerodynamics are investigated. Needs for the National Aerospace Plane and interplanetary flights with atmospheric aerobraking are used to scope the problem. Relevant events of the past two decades in universities and at the national laboratories are examined for their implications regarding both problems and prospects. Most striking is the emergence of computational fluid dynamics, which is viewed here as an equal partner with laboratory experimentation and flight test in relating theory with reality. Also significant are major advances in instrumentation and data processing methods, especially optical techniques. The direction of the study was guided by the concept of a companion program, i.e., the university effort should complement a major area of endeavor at NASA-Langley. Through this, both faculty and student participants gain a natural and effective working relationship. Existing and proposed major hypersonic aerodynamic facilities in industry and at the national laboratories are examined by type; hypersonic wind tunnels, arc-heated tunnels, shock tubes and tunnels, and ballistic ranges. Of these, the free piston tunnel and shock tube/tunnel are most appropriate for a university.
NASA Technical Reports Server (NTRS)
Arnold, James O.; Deiwert, George S.
1997-01-01
This paper surveys the use of aerothermodynamic facilities which have been useful in the study of external flows and propulsion aspects of hypersonic, air-breathing vehicles. While the paper is not a survey of all facilities, it covers the utility of shock tunnels and conventional hypersonic blow-down facilities which have been used for hypersonic air-breather studies. The problems confronting researchers in the field of aerothermodynamics are outlined. Results from the T5 GALCIT tunnel for the shock-on lip problem are outlined. Experiments on combustors and short expansion nozzles using the semi-free jet method have been conducted in large shock tunnels. An example which employed the NASA Ames 16-Inch shock tunnel is outlined, and the philosophy of the test technique is described. Conventional blow-down hypersonic wind tunnels are quite useful in hypersonic air-breathing studies. Results from an expansion ramp experiment, simulating the nozzle on a hypersonic air-breather from the NASA Ames 3.5 Foot Hypersonic wind tunnel are summarized. Similar work on expansion nozzles conducted in the NASA Langley hypersonic wind tunnel complex is cited. Free-jet air-frame propulsion integration and configuration stability experiments conducted at Langley in the hypersonic wind tunnel complex on a small generic model are also summarized.
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. PMID:24348146
Unstart Coupling Mechanism Analysis of Multiple-Modules Hypersonic Inlet
Wang, Lei; Cao, Shibin
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. PMID:24348146
Simulation of free-space optical guiding structure based on colliding gas flows.
Kaganovich, D; Palastro, J P; Chen, Y-H; Gordon, D F; Helle, M H; Ting, A
2015-11-01
Preformed plasma channels with parabolic radial density profiles enable the extended and stable optical guiding of high-intensity laser pulses. High-voltage discharge capillaries, commonly used for channel formation, have limited guiding length and opaque walls, complicating the diagnosis of the plasma within. This paper proposes a free-space gas channel produced by the collision of several gas flows. The collision of the gas flows forms an on-axis density depression surrounded by higher density walls. By offsetting the flows, we demonstrated the creation of what we believe is a novel vortex structure that exhibits a long-lived parabolic density profile. Once ionized, the resulting plasma density profile has a near-parabolic dependence appropriate for guiding. We then performed detailed two-dimensional (2D) fluid dynamics simulations to examine the properties and stability of the guiding structure. PMID:26560601
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.
Gronberg, J
2002-10-07
A photon collider interaction region has the possibility of expanding the physics reach of a future TeV scale electron-positron collider. A survey of ongoing efforts to design the required lasers and optics to create a photon collider is presented in this paper.
NASA's hypersonic fluid and thermal physics program (Aerothermodynamics)
NASA Technical Reports Server (NTRS)
Graves, R. A.; Hunt, J. L.
1985-01-01
This survey paper gives an overview of NASA's hypersonic fluid and thermal physics program (recently renamed aerothermodynamics). The purpose is to present the elements of, example results from, and rationale and projection for this program. The program is based on improving the fundamental understanding of aerodynamic and aerothermodynamic flow phenomena over hypersonic vehicles in the continuum, transitional, and rarefied flow regimes. Vehicle design capabilities, computational fluid dynamics, computational chemistry, turbulence modeling, aerothermal loads, orbiter flight data analysis, orbiter experiments, laser photodiagnostics, and facilities are discussed.
Hypersonic Wind Tunnels: Latest Citations from the Aerospace Database
NASA Technical Reports Server (NTRS)
1996-01-01
The bibliography contains citations concerning the design, construction, operation, performance, and use of hypersonic wind tunnels. References cover the design of flow nozzles, diffusers, test sections, and ejectors for tunnels driven by compressed air, high-pressure gases, or cryogenic liquids. Methods for flow calibration, boundary layer control, local and freestream turbulence reduction, and force measurement are discussed. Intrusive and non-intrusive instrumentation, sources of measurement error, and measurement corrections are also covered. The citations also include the testing of inlets, nozzles, airfoils, and other components of hypersonic aerospace vehicles. Comprehensive coverage of supersonic and blowdown wind tunnels, and force balance systems for wind tunnels are covered in separate bibliographies.
Laser-driven hypersonic air-breathing propulsion simulator
NASA Technical Reports Server (NTRS)
Joshi, Prakash B.; Lo, Edmond Y.; Pugh, Evan R.
1992-01-01
A feasibility study is presented of simulating airbreathing propulsion on small scale hypersonic models using laser energy. The laser heat addition scheme allows simultaneous inlet and exhaust flows during wind tunnel testing of models with scramjet models. The proposed propulsion simulation concept has extended the Kantrowitz (1974) idea to propulsive wind tunnel models of hypersonic aircraft. Critical issues in aeropropulsive testing of models based on a ramjet power plant are addressed which include transfer of the correct amount of energy to the flowing gas, efficient absorption of laser energy into the gas, and test performance under tunnel reservoir conditions and at reasonable Reynolds numbers.
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
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.
Analysis of cooling systems for hypersonic aircraft
NASA Technical Reports Server (NTRS)
Petley, Dennis H.; Jones, Stuart C.; Dziedzic, William M.
1991-01-01
A computer program has been written to analyze cooling systems of hypersonic aircraft. This computer program called NASP/SINDA is written into the SINDA'85 command structure and uses the SINDA'85 finite difference subroutines. Both internal fluid flow and heat transfer must be analyzed, because increased heating causes a decrease in the flow of the coolant. Also local hot spots will cause a redistribution of the coolant in the system. Both steady state and transient analyses have been performed. Details of empirical correlations are presented. Results for two cooling system applications are given.
Joint computational and experimental aerodynamics research on a hypersonic vehicle
Oberkampf, W.L.; Aeschliman, D.P.; Walker, M.M.
1992-01-01
A closely coupled computational and experimental aerodynamics research program was conducted on a hypersonic vehicle configuration at Mach 8. Aerodynamic force and moment measurements and flow visualization results were obtained in the Sandia National Laboratories hypersonic wind tunnel for laminar boundary layer conditions. Parabolized and iterative Navier-Stokes simulations were used to predict flow fields and forces and moments on the hypersonic configuration. The basic vehicle configuration is a spherically blunted 10{degrees} cone with a slice parallel with the axis of the vehicle. On the slice portion of the vehicle, a flap can be attached so that deflection angles of 10{degrees}, 20{degrees}, and 30{degrees} can be obtained. Comparisons are made between experimental and computational results to evaluate quality of each and to identify areas where improvements are needed. This extensive set of high-quality experimental force and moment measurements is recommended for use in the calibration and validation of computational aerodynamics codes. 22 refs.
The NASA Glen Research Center's Hypersonic Tunnel Facility. Chapter 16
NASA Technical Reports Server (NTRS)
Woike, Mark R.; Willis, Brian P.
2001-01-01
The NASA Glenn Research Center's Hypersonic Tunnel Facility (HTF) is a blow-down, freejet wind tunnel that provides true enthalpy flight conditions for Mach numbers of 5, 6, and 7. The Hypersonic Tunnel Facility is unique due to its large scale and use of non-vitiated (clean air) flow. A 3MW graphite core storage heater is used to heat the test medium of gaseous nitrogen to the high stagnation temperatures required to produce true enthalpy conditions. Gaseous oxygen is mixed into the heated test flow to generate the true air simulation. The freejet test section is 1.07m (42 in.) in diameter and 4.3m (14 ft) in length. The facility is well suited for the testing of large scale airbreathing propulsion systems. In this chapter, a brief history and detailed description of the facility are presented along with a discussion of the facility's application towards hypersonic airbreathing propulsion testing.
Hypersonic Shock/Boundary-Layer Interaction Database
NASA Technical Reports Server (NTRS)
Settles, G. S.; Dodson, L. J.
1991-01-01
Turbulence modeling is generally recognized as the major problem obstructing further advances in computational fluid dynamics (CFD). A closed solution of the governing Navier-Stokes equations for turbulent flows of practical consequence is still far beyond grasp. At the same time, the simplified models of turbulence which are used to achieve closure of the Navier-Stokes equations are known to be rigorously incorrect. While these models serve a definite purpose, they are inadequate for the general prediction of hypersonic viscous/inviscid interactions, mixing problems, chemical nonequilibria, and a range of other phenomena which must be predicted in order to design a hypersonic vehicle computationally. Due to the complexity of turbulence, useful new turbulence models are synthesized only when great expertise is brought to bear and considerable intellectual energy is expended. Although this process is fundamentally theoretical, crucial guidance may be gained from carefully-executed basic experiments. Following the birth of a new model, its testing and validation once again demand comparisons with data of unimpeachable quality. This report concerns these issues which arise from the experimental aspects of hypersonic modeling and represents the results of the first phase of an effort to develop compressible turbulence models.
Studies in hypersonic aeroelasticity
NASA Astrophysics Data System (ADS)
Nydick, Ira Harvey
2000-11-01
This dissertation describes the aeroelastic analysis of a generic hypersonic vehicle, focusing on two specific problems: (1) hypersonic panel flutter, and (2) aeroelastic behavior of a complete unrestrained generic hypersonic vehicle operating at very high Mach numbers. The panels are modeled as shallow shells using Marguerre nonlinear shallow shell theory for orthotropic panels and the aerodynamic loads are obtained from third order piston theory. Two models of curvature, several applied temperature distributions, and the presence of a shock are also included in the model. Results indicate that the flutter speed of the panel is significantly reduced by temperature variations comparable to the buckling temperature and by the presence of a shock. A panel with initial curvature can be more stable than the flat panel but the increase in stability depends in a complex way on the material properties of the panel and the amount of curvature. At values of dynamic pressure above critical, aperiodic motion was observed. The value of dynamic pressure for which this occurs in both heated panels and curved panels is much closer to the critical dynamic pressure than for the flat, unheated panel. A comparison of piston theory aerodynamics and Euler and Navier-Stokes aerodynamics was performed for a two dimensional panel with prescribed motion and the results indicate that while 2nd or higher order piston theory agrees very well with the Euler solution for the frequencies seen in hypersonic panel flutter, it differs substantially from the Navier-Stokes solution. The aeroelastic behavior of the complete vehicle was simulated using the unrestrained equations of motion, utilizing the method of quasi-coordinates. The unrestrained mode shapes of the vehicle were obtained from an equivalent plate analysis using an available code (ELAPS). The effects of flexible trim and rigid body degrees of freedom are carefully incorporated in the mathematical model. This model was applied to a
Experimental aerodynamics research on a hypersonic vehicle
Oberkampf, W.L.; Aeschliman, D.P.; Tate, R.E.; Henfling, J.F.
1993-04-01
Aerodynamic force and moment measurements and flow visualization results are presented for a hypersonic vehicle configuration at Mach 8. The basic vehicle configuration is a spherically blunted 10[degree] half-angle cone with a slice parallel with the axis of the vehicle. On the slice portion of the vehicle, a flap could be attached so that deflection angles of 10[degree], 20[degree] and 30[degree] could be obtained. All of the experimental results were obtained in the Sandia Mach 8 hypersonic wind tunnel for laminar boundary layer conditions. Flow visualization results include shear stress sensitive liquid crystal photographs, surface streak flow photographs (using liquid crystals), and spark schlieren photographs and video. The liquid crystals were used as an aid in verifying that a laminar boundary layer existed over the entire body. The surface flow photo-graphs show attached and separated flow on both the leeside of the vehicle and near the flap. A detailed uncertainty analysis was conducted to estimate the contributors to body force and moment measurement uncertainty. Comparisons are made with computational results to evaluate both the experimental and numerical results. This extensive set of high-quality experimental force and moment measurements is recommended for use in the calibration and validation of relevant computational aerodynamics codes.
Experimental aerodynamics research on a hypersonic vehicle
Oberkampf, W.L.; Aeschliman, D.P.; Tate, R.E.; Henfling, J.F.
1993-04-01
Aerodynamic force and moment measurements and flow visualization results are presented for a hypersonic vehicle configuration at Mach 8. The basic vehicle configuration is a spherically blunted 10{degree} half-angle cone with a slice parallel with the axis of the vehicle. On the slice portion of the vehicle, a flap could be attached so that deflection angles of 10{degree}, 20{degree} and 30{degree} could be obtained. All of the experimental results were obtained in the Sandia Mach 8 hypersonic wind tunnel for laminar boundary layer conditions. Flow visualization results include shear stress sensitive liquid crystal photographs, surface streak flow photographs (using liquid crystals), and spark schlieren photographs and video. The liquid crystals were used as an aid in verifying that a laminar boundary layer existed over the entire body. The surface flow photo-graphs show attached and separated flow on both the leeside of the vehicle and near the flap. A detailed uncertainty analysis was conducted to estimate the contributors to body force and moment measurement uncertainty. Comparisons are made with computational results to evaluate both the experimental and numerical results. This extensive set of high-quality experimental force and moment measurements is recommended for use in the calibration and validation of relevant computational aerodynamics codes.
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)
1987-01-01
A hypersonic transport aircraft design project was selected as a result of interactions with NASA Lewis Research Center personnel and fits the Presidential concept of the Orient Express. The Graduate Teaching Assistant (GTA) and an undergraduate student worked at the NASA Lewis Research Center during the 1986 summer conducting a literature survey, and relevant literature and useful software were collected. The computer software was implemented in the Computer Aided Design Laboratory of the Mechanical and Aerospace Engineering Department. In addition to the lectures by the three instructors, a series of guest lectures was conducted. The first of these lectures 'Anywhere in the World in Two Hours' was delivered by R. Luidens of NASA Lewis Center. In addition, videotaped copies of relevant seminars obtained from NASA Lewis were also featured. The first assignment was to individually research and develop the mission requirements and to discuss the findings with the class. The class in consultation with the instructors then developed a set of unified mission requirements. Then the class was divided into three design groups (1) Aerodynamics Group, (2) Propulsion Group, and (3) Structures and Thermal Analyses Group. The groups worked on their respective design areas and interacted with each other to finally come up with an integrated conceptual design. The three faculty members and the GTA acted as the resource persons for the three groups and aided in the integration of the individual group designs into the final design of a hypersonic aircraft.
NASA Technical Reports Server (NTRS)
Alkamhawi, Hani; Greiner, Tom; Fuerst, Gerry; Luich, Shawn; Stonebraker, Bob; Wray, Todd
1990-01-01
A hypersonic aircraft is designed which uses scramjets to accelerate from Mach 6 to Mach 10 and sustain that speed for two minutes. Different propulsion systems were considered and it was decided that the aircraft would use one full scale turbofan-ramjet. Two solid rocket boosters were added to save fuel and help the aircraft pass through the transonic region. After considering aerodynamics, aircraft design, stability and control, cooling systems, mission profile, and landing systems, a conventional aircraft configuration was chosen over that of a waverider. The conventional design was chosen due to its landing characteristics and the relative expense compared to the waverider. Fuel requirements and the integration of the engine systems and their inlets are also taken into consideration in the final design. A hypersonic aircraft was designed which uses scramjets to accelerate from Mach 6 to Mach 10 and sustain that speed for two minutes. Different propulsion systems were considered and a full scale turbofan-ramjet was chosen. Two solid rocket boosters were added to save fuel and help the aircraft pass through the transonic reqion. After the aerodynamics, aircraft design, stability and control, cooling systems, mission profile, landing systems, and their physical interactions were considered, a conventional aircraft configuration was chosen over that of a waverider. The conventional design was chosen due to its landing characteristics and the relative expense compared to the waverider. Fuel requirements and the integration of the engine systems and their inlets were also considered in the designing process.
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.
Hypersonic Interceptor Performance Evaluation Center aero-optics performance predictions
NASA Astrophysics Data System (ADS)
Sutton, George W.; Pond, John E.; Snow, Ronald; Hwang, Yanfang
1993-06-01
This paper describes the Hypersonic Interceptor Performance Evaluation Center's (HIPEC) aerooptics performance predictions capability. It includes code results for three dimensional shapes and comparisons to initial experiments. HIPEC consists of a collection of aerothermal, aerodynamic computational codes which are capable of covering the entire flight regime from subsonic to hypersonic flow and include chemical reactions and turbulence. Heat transfer to the various surfaces is calculated as an input to cooling and ablation processes. HIPEC also has aero-optics codes to determine the effect of the mean flowfield and turbulence on the tracking and imaging capability of on-board optical sensors. The paper concentrates on the latter aspects.
Progress in hypersonic combustion technology with computation and experiment
NASA Technical Reports Server (NTRS)
Anderson, Griffin Y.; Kumar, Ajay; Erdos, John I.
1990-01-01
Design of successful airbreathing engines for operation at near-orbital speeds presents significant challenges in all the disciplines involved, including propulsion. This paper presents a discussion of the important physics of hypersonic combustion and an assessment of the state of the art of ground simulations with pulse facilities and with computational techniques. Recent examples of experimental and computational simulations are presented and discussed. The need for continued application of these tools to establish the credibility and fidelity of engineering design methods for practical hypersonic combustors is emphasized along with the critical need for improved diagnostic methods for hypervelocity reacting flows.
Surface pressure measurements on a hypersonic vehicle
Oberkampf, W.L.; Aeschliman, D.P.; Henfling, J.F.; Larson, D.E.; Payne, J.L.
1996-02-01
Extensive surface pressure measurements were obtained on a hypersonic vehicle configuration at Mach 8 for the purpose of computational fluid dynamics code validation. Experiments were conducted in the Sandia National Laboratories hypersonic wind tunnel. All measurements were made for laminar flow conditions at a Reynolds number (based on model length) of 1.81 x 10{sup 6} and perfect gas conditions. The basic vehicle configuration is a spherically blunted, 10{degree} half- angle cone, with a slice parallel to the axis of the vehicle. To the aft portion of the slice could be attached flaps of varying angle; 10, 20, and 30{degree}. Surface pressure measurements were obtained for angles of attack from -10 to +18{degree}, for various roll angles, at 96 locations on the body surface. All three deflected flap angles produced separated flow on the sliced portion of the body in front of the flap. Because of the three-dimensional expansion over the slice, the separated flow on the slice and flap was also highly three- dimensional. The results of the present experiment provide extensive surface pressure measurements for the validation of computational fluid dynamics codes for separated flow caused by an embedded shock wave.
NASA Technical Reports Server (NTRS)
Spina, Eric F.
1995-01-01
The primary objective in the two research investigations performed under NASA Langley sponsorship (Turbulence measurements in hypersonic boundary layers using constant temperature anemometry and Reynolds stress measurements in hypersonic boundary layers) has been to increase the understanding of the physics of hypersonic turbulent boundary layers. The study began with an extension of constant-temperature thermal anemometry techniques to a Mach 11 helium flow, including careful examinations of hot-wire construction techniques, system response, and system calibration. This was followed by the application of these techniques to the exploration of a Mach 11 helium turbulent boundary layer (To approximately 290 K). The data that was acquired over the course of more than two years consists of instantaneous streamwise mass flux measurements at a frequency response of about 500 kHz. The data are of exceptional quality in both the time and frequency domain and possess a high degree of repeatability. The data analysis that has been performed to date has added significantly to the body of knowledge on hypersonic turbulence, and the data reduction is continuing. An attempt was then made to extend these thermal anemometry techniques to higher enthalpy flows, starting with a Mach 6 air flow with a stagnation temperature just above that needed to prevent liquefaction (To approximately 475 F). Conventional hot-wire anemometry proved to be inadequate for the selected high-temperature, high dynamic pressure flow, with frequent wire breakage and poor system frequency response. The use of hot-film anemometry has since been investigated for these higher-enthalpy, severe environment flows. The difficulty with using hot-film probes for dynamic (turbulence) measurements is associated with construction limitations and conduction of heat into the film substrate. Work continues under a NASA GSRP grant on the development of a hot film probe that overcomes these shortcomings for hypersonic
NASA Astrophysics Data System (ADS)
Back, B. B.; Baker, M. D.; Ballintijn, M.; Barton, D. S.; Betts, R. R.; Bickley, A. A.; Bindel, R.; Budzanowski, A.; Busza, W.; Carroll, A.; Chai, Z.; Decowski, M. P.; García, E.; Gburek, T.; George, N.; Gulbrandsen, K.; Gushue, S.; Halliwell, C.; Hamblen, J.; Hauer, M.; Heintzelman, G. A.; Henderson, C.; Hofman, D. J.; Hollis, R. S.; Hołyński, R.; Holzman, B.; Iordanova, A.; Johnson, E.; Kane, J. L.; Katzy, J.; Khan, N.; Kucewicz, W.; Kulinich, P.; Kuo, C. M.; Lin, W. T.; Manly, S.; McLeod, D.; Mignerey, A. C.; Nouicer, R.; Olszewski, A.; Pak, R.; Park, I. C.; Pernegger, H.; Reed, C.; Remsberg, L. P.; Reuter, M.; Roland, C.; Roland, G.; Rosenberg, L.; Sagerer, J.; Sarin, P.; Sawicki, P.; Seals, H.; Sedykh, I.; Skulski, W.; Smith, C. E.; Stankiewicz, M. A.; Steinberg, P.; Stephans, G. S. F.; Sukhanov, A.; Tang, J.-L.; Tonjes, M. B.; Trzupek, A.; Vale, C.; van Nieuwenhuizen, G. J.; Vaurynovich, S. S.; Verdier, R.; Veres, G. I.; Wenger, E.; Wolfs, F. L. H.; Wosiek, B.; Woźniak, K.; Wuosmaa, A. H.; Wysłouch, B.
2006-07-01
We report on measurements of directed flow as a function of pseudorapidity in Au+Au collisions at energies of sNN=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.
On Challenges for Hypersonic Turbulent Simulations
NASA Astrophysics Data System (ADS)
Yee, H. C.; Sjögreen, B.
2009-04-01
This short note discusses some of the challenges for design of suitable spatial numerical schemes for hypersonic turbulent flows, including combustion, and thermal and chemical nonequilibrium flows. Often, hypersonic turbulent flows around re-entry space vehicles and space physics involve mixed steady strong shocks and turbulence with unsteady shocklets. Material mixing in combustion poses additional computational challenges. Proper control of numerical dissipation in numerical methods beyond the standard shock-capturing dissipation at discontinuities is an essential element for accurate and stable simulations of the subject physics. On the one hand, the physics of strong steady shocks and unsteady turbulence/shocklet interactions under the nonequilibrium environment is not well understood. On the other hand, standard and newly developed high order accurate (fourth-order or higher) schemes were developed for homogeneous hyperbolic conservation laws and mixed hyperbolic and parabolic partial differential equations (PDEs) (without source terms). The majority of finite rate chemistry and thermal nonequilibrium simulations employ methods for homogeneous time-dependent PDEs with a pointwise evaluation of the source terms. The pointwise evaluation of the source term might not be the best choice for stability, accuracy and minimization of spurious numerics for the overall scheme.
On Challenges for Hypersonic Turbulent Simulations
Yee, H C; Sjogreen, B
2009-01-14
This short note discusses some of the challenges for design of suitable spatial numerical schemes for hypersonic turbulent flows, including combustion, and thermal and chemical nonequilibrium flows. Often, hypersonic turbulent flows in re-entry space vehicles and space physics involve mixed steady strong shocks and turbulence with unsteady shocklets. Material mixing in combustion poses additional computational challenges. Proper control of numerical dissipation in numerical methods beyond the standard shock-capturing dissipation at discontinuities is an essential element for accurate and stable simulations of the subject physics. On one hand, the physics of strong steady shocks and unsteady turbulence/shocklet interactions under the nonequilibrium environment is not well understood. On the other hand, standard and newly developed high order accurate (fourth-order or higher) schemes were developed for homogeneous hyperbolic conservation laws and mixed hyperbolic and parabolic partial differential equations (PDEs) (without source terms). The majority of finite rate chemistry and thermal nonequilibrium simulations employ methods for homogeneous time-dependent PDEs with a pointwise evaluation of the source terms. The pointwise evaluation of the source term might not be the best choice for stability, accuracy and minimization of spurious numerics for the overall scheme.
Advanced hypersonic aircraft design
NASA Technical Reports Server (NTRS)
Utzinger, Rob; Blank, Hans-Joachim; Cox, Craig; Harvey, Greg; Mckee, Mike; Molnar, Dave; Nagy, Greg; Petersen, Steve
1992-01-01
The objective of this design project is to develop the hypersonic reconnaissance aircraft to replace the SR-71 and to complement existing intelligence gathering devices. The initial design considerations were to create a manned vehicle which could complete its mission with at least two airborne refuelings. The aircraft must travel between Mach 4 and Mach 7 at an altitude of 80,000 feet for a maximum range of 12,000 nautical miles. The vehicle should have an air breathing propulsion system at cruise. With a crew of two, the aircraft should be able to take off and land on a 10,000 foot runway, and the yearly operational costs were not to exceed $300 million. Finally, the aircraft should exhibit stealth characteristics, including a minimized radar cross-section (RCS) and a reduced sonic boom. The technology used in this vehicle should allow for production between the years 1993 and 1995.
NASA's Hypersonic Investment Area
NASA Technical Reports Server (NTRS)
Hueter, Uwe; Hutt, John; McClinton, Charles
2002-01-01
NASA has established long term goals for access to space. The third generation launch systems are to be fully reusable and operational around 2025. The goal for third-generation launch systems represents significant reduction in cost and improved safety over the current first generation system. The Advanced Space Transportation Office (ASTP) at NASA s Marshall Space Flight Center (MSFC) has the agency lead to develop space transportation technologies. Within ASTP, under the Hypersonic Investment Area (HIA), third generation technologies are being pursued in the areas of propulsion, airframe, integrated vehicle health management (IVHM), avionics, power, operations and system analysis. These technologies are being matured through research and both ground and flight-testing. This paper provides an overview of the HIA program plans and recent accomplishments.
Hypersonic reconnaissance aircraft
NASA Technical Reports Server (NTRS)
Bulk, Tim; Chiarini, David; Hill, Kevin; Kunszt, Bob; Odgen, Chris; Truong, Bon
1992-01-01
A conceptual design of a hypersonic reconnaissance aircraft for the U.S. Navy is discussed. After eighteen weeks of work, a waverider design powered by two augmented turbofans was chosen. The aircraft was designed to be based on an aircraft carrier and to cruise 6,000 nautical miles at Mach 4;80,000 feet and above. As a result the size of the aircraft was only allowed to have a length of eighty feet, fifty-two feet in wingspan, and roughly 2,300 square feet in planform area. Since this is a mainly cruise aircraft, sixty percent of its 100,000 pound take-off weight is JP fuel. At cruise, the highest temperature that it will encounter is roughly 1,100 F, which can be handled through the use of a passive cooling system.
Homogeneous catalysts in hypersonic combustion
Harradine, D.M.; Lyman, J.L.; Oldenborg, R.C.; Pack, R.T.; Schott, G.L.
1989-01-01
Density and residence time both become unfavorably small for efficient combustion of hydrogen fuel in ramjet propulsion in air at high altitude and hypersonic speed. Raising the density and increasing the transit time of the air through the engine necessitates stronger contraction of the air flow area. This enhances the kinetic and thermodynamic tendency of H/sub 2/O to form completely, accompanied only by N/sub 2/ and any excess H/sub 2/(or O/sub 2/). The by-products to be avoided are the energetically expensive fragment species H and/or O atoms and OH radicals, and residual (2H/sub 2/ plus O/sub 2/). However, excessive area contraction raises air temperature and consequent combustion-product temperature by adiabatic compression. This counteracts and ultimately overwhelms the thermodynamic benefit by which higher density favors the triatomic product, H/sub 2/O, over its monatomic and diatomic alternatives. For static pressures in the neighborhood of 1 atm, static temperature must be kept or brought below ca. 2400 K for acceptable stability of H/sub 2/O. Another measure, whose requisite chemistry we address here, is to extract propulsive work from the combustion products early in the expansion. The objective is to lower the static temperature of the combustion stream enough for H/sub 2/O to become adequately stable before the exhaust flow is massively expanded and its composition ''frozen.'' We proceed to address this mechanism and its kinetics, and then examine prospects for enhancing its rate by homogeneous catalysts. 9 refs.
Further Investigations of Hypersonic Engine Seals
NASA Technical Reports Server (NTRS)
Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; DeMange, Jeffrey J.
2004-01-01
Durable, flexible sliding seals are required in advanced hypersonic engines to seal the perimeters of movable engine ramps for efficient, safe operation in high heat flux environments at temperatures of 2000 to 2500 F. Current seal designs do not meet the demanding requirements for future engines, so NASA's Glenn Research Center is developing advanced seals and preloading devices to overcome these shortfalls. An advanced ceramic wafer seal design and two silicon nitride compression spring designs were evaluated in a series of compression, scrub, and flow tests. Silicon nitride wafer seals survived 2000 in. (50.8 m) of scrubbing at 2000 F against a silicon carbide rub surface with no chips or signs of damage. Flow rates measured for the wafers before and after scrubbing were almost identical and were up to 32 times lower than those recorded for the best braided rope seal flow blockers. Silicon nitride compression springs showed promise conceptually as potential seal preload devices to help maintain seal resiliency.
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
The critical role of aerodynamic heating effects in the design of hypersonic vehicles
NASA Technical Reports Server (NTRS)
Wieting, Allan R.
1989-01-01
Hypersonic vehicles operate in a hostile aerothermal environment, which has a significant impact on their aerothermostructural performance. Significant coupling occurs between the aerodynamic flow field, structural heat transfer, and structural response, creating a multidisciplinary interaction. The critical role of aerodynamic heating effects in the design of hypersonic vehicles is identified with an example of high localized heating on an engine-cowl leading edge. Recent advances is integrated fluid-thermal-structural finite-element analyses are presented.
NASA Astrophysics Data System (ADS)
Back, B. B.; Baker, M. D.; Ballintijn, M.; Barton, D. S.; Betts, R. R.; Bickley, A. A.; Bindel, R.; Budzanowski, A.; Busza, W.; Carroll, A.; Chai, Z.; Decowski, M. P.; García, E.; Gburek, T.; George, N.; Gulbrandsen, K.; Gushue, S.; Halliwell, C.; Hamblen, J.; Hauer, M.; Heintzelman, G. A.; Henderson, C.; Hofman, D. J.; Hollis, R. S.; Hołyński, R.; Holzman, B.; Iordanova, A.; Johnson, E.; Kane, J. L.; Katzy, J.; Khan, N.; Kucewicz, W.; Kulinich, P.; Kuo, C. M.; Lin, W. T.; Manly, S.; McLeod, D.; Mignerey, A. C.; Nouicer, R.; Olszewski, A.; Pak, R.; Park, I. C.; Pernegger, H.; Reed, C.; Remsberg, L. P.; Reuter, M.; Roland, C.; Roland, G.; Rosenberg, L.; Sagerer, J.; Sarin, P.; Sawicki, P.; Seals, H.; Sedykh, I.; Skulski, W.; Smith, C. E.; Stankiewicz, M. A.; Steinberg, P.; Stephans, G. S.; Sukhanov, A.; Tang, J.-L.; Tonjes, M. B.; Trzupek, A.; Vale, C.; van Nieuwenhuizen, G. J.; Vaurynovich, S. S.; Verdier, R.; Veres, G. I.; Wenger, E.; Wolfs, F. L.; Wosiek, B.; Woźniak, K.; Wuosmaa, A. H.; Wysłouch, B.
2005-04-01
This Letter describes the measurement of the energy dependence of elliptic flow for charged particles in Au+Au collisions using the PHOBOS detector at the Relativistic Heavy Ion Collider. Data taken at collision energies of √(sNN)=19.6, 62.4, 130, and 200 GeV are shown over a wide range in pseudorapidity. These results, when plotted as a function of η'=|η|-ybeam, scale with approximate linearity throughout η', implying no sharp changes in the dynamics of particle production as a function of pseudorapidity or increasing beam energy.
Vorticity interaction effects on blunt bodies. [hypersonic viscous shock layers
NASA Technical Reports Server (NTRS)
Anderson, E. C.; Wilcox, D. C.
1977-01-01
Numerical solutions of the viscous shock layer equations governing laminar and turbulent flows of a perfect gas and radiating and nonradiating mixtures of perfect gases in chemical equilibrium are presented for hypersonic flow over spherically blunted cones and hyperboloids. Turbulent properties are described in terms of the classical mixing length. Results are compared with boundary layer and inviscid flowfield solutions; agreement with inviscid flowfield data is satisfactory. Agreement with boundary layer solutions is good except in regions of strong vorticity interaction; in these flow regions, the viscous shock layer solutions appear to be more satisfactory than the boundary layer solutions. Boundary conditions suitable for hypersonic viscous shock layers are devised for an advanced turbulence theory.
TBCC Discipline Overview. Hypersonics Project
NASA Technical Reports Server (NTRS)
Thomas, Scott R.
2011-01-01
The "National Aeronautics Research and Development Policy" document, issued by the National Science and Technology Council in December 2006, stated that one (among several) of the guiding objectives of the federal aeronautics research and development endeavors shall be stable and long-term foundational research efforts. Nearly concurrently, the National Academies issued a more technically focused aeronautics blueprint, entitled: the "Decadal Survey of Civil Aeronautics - Foundations for the Future." Taken together these documents outline the principles of an aeronautics maturation plan. Thus, in response to these overarching inputs (and others), the National Aeronautics and Space Administration (NASA) organized the Fundamental Aeronautics Program (FAP), a program within the NASA Aeronautics Research Mission Directorate (ARMD). The FAP initiated foundational research and technology development tasks to enable the capability of future vehicles that operate across a broad range of Mach numbers, inclusive of the subsonic, supersonic, and hypersonic flight regimes. The FAP Hypersonics Project concentrates on two hypersonic missions: (1) Air-breathing Access to Space (AAS) and (2) the (Planetary Atmospheric) Entry, Decent, and Landing (EDL). The AAS mission focuses on Two-Stage-To-Orbit (TSTO) systems using air-breathing combined-cycle-engine propulsion; whereas, the EDL mission focuses on the challenges associated with delivering large payloads to (and from) Mars. So, the FAP Hypersonic Project investments are aligned to achieve mastery and intellectual stewardship of the core competencies in the hypersonic-flight regime, which ultimately will be required for practical systems with highly integrated aerodynamic/vehicle and propulsion/engine technologies. Within the FAP Hypersonics, the technology management is further divided into disciplines including one targeting Turbine-Based Combine-Cycle (TBCC) propulsion. Additionally, to obtain expertise and support from outside
Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations
NASA Technical Reports Server (NTRS)
Cockrell, Charles E., Jr.; Huebner, Lawrence D.; Finley, Dennis B.
1996-01-01
An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, including canopies, engine components, and control surfaces, with two pure waverider shapes, both conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) solutions were obtained over a Mach number range of 1.6 to 4.63. The experimental data show the component build-up effects and the aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations is not comparable to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configurations exhibit good lateral-directional stability characteristics.
Palmer, R.
2009-10-19
Parameters are given of muon colliders with center of mass energies of 1.5 and 3 TeV. Pion production is from protons on a mercury target. Capture, decay, and phase rotation yields bunch trains of both muon signs. Six dimensional cooling reduces the emittances until the trains are merged into single bunches, one of each sign. Further cooling in 6 dimensions is then applied, followed by final transverse cooling in 50 T solenoids. After acceleration the muons enter the collider ring. Ongoing R&D is discussed.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will design, build, and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604BOO02G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate the aerodynamic flight database for the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. Al these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database i n the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime, The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
Hypersonic Wake Diagnostics Using Laser Induced Fluorescence Techniques
NASA Technical Reports Server (NTRS)
Mills, Jack L.; Sukenik, Charles I.; Balla, Robert J.
2011-01-01
A review of recent research performed in iodine that involves a two photon absorption of light at 193 nm will be discussed, and it's potential application to velocimetry measurements in a hypersonic flow field will be described. An alternative seed atom, Krypton, will be presented as a good candidate for performing nonintrusive hypersonic flow diagnostics. Krypton has a metastable state with a lifetime of approximately 43 s which would prove useful for time of flight measurement (TOF) and a sensitivity to collisions that can be utilized for density measurements. Calculations using modest laser energies and experimental values show an efficiency of excited state production to be on the order of 10(exp -6) for a two photon absorption at 193 nm.
Application of CFD to a generic hypersonic flight research study
NASA Technical Reports Server (NTRS)
Green, Michael J.; Lawrence, Scott L.; Dilley, Arthur D.; Hawkins, Richard W.; Walker, Mary M.; Oberkampf, William L.
1993-01-01
Computational analyses have been performed for the initial assessment of flight research vehicle concepts that satisfy requirements for potential hypersonic experiments. Results were obtained from independent analyses at NASA Ames, NASA Langley, and Sandia National Labs, using sophisticated time-dependent Navier-Stokes and parabolized Navier-Stokes methods. Careful study of a common problem consisting of hypersonic flow past a slightly blunted conical forebody was undertaken to estimate the level of uncertainty in the computed results, and to assess the capabilities of current computational methods for predicting boundary-layer transition onset. Results of this study in terms of surface pressure and heat transfer comparisons, as well as comparisons of boundary-layer edge quantities and flow-field profiles are presented here. Sensitivities to grid and gas model are discussed. Finally, representative results are presented relating to the use of Computational Fluid Dynamics in the vehicle design and the integration/support of potential experiments.
A study of hypersonic small-disturbance theory
NASA Technical Reports Server (NTRS)
Van Dyke, Milton D
1954-01-01
A systematic study is made of the approximate inviscid theory of thin bodies moving at such high supersonic speeds that nonlinearity is an essential feature of the equations of flow. The first-order small-disturbance equations are derived for three-dimensional motions involving shock waves, and estimates are obtained for the order of error involved in the approximation. The hypersonic similarity rule of Tsien and Hayes, and Hayes' unsteady analogy appear in the course of the development. It is shown that the hypersonic theory can be interpreted so that it applies also in the range of linearized supersonic flow theory. Several examples are solved according to the small-disturbance theory, and compared with the full solutions when available.
Airbreathing Hypersonic Systems Focus at NASA Langley Research Center
NASA Technical Reports Server (NTRS)
Hunt, James L.; Rausch, Vincent L.
1998-01-01
This paper presents the status of the airbreathing hypersonic airplane and space-access vehicle design matrix, reflects on the synergies and issues, and indicates the thrust of the effort to resolve the design matrix and to focus/advance systems technology maturation. Priority is given to the design of the vision operational vehicles followed by flow-down requirements to flight demonstrator vehicles and their design for eventual consideration in the Future-X Program.
Aerothermal/FEM Analysis of Hypersonic Sharp Leading Edges
NASA Technical Reports Server (NTRS)
Kolodziej, Paul; Bull, Jeffrey D.; Kowalski, Thomas R.; Rasky, Daniel J. (Technical Monitor)
1995-01-01
Advanced hypersonic vehicles, like wave riders, will have sharp leading edges to minimize drag. These designs require accurate finite element modeling (FEM) of the thermal-structural behavior of a diboride ceramic matrix composite sharp leading edge. By coupling the FEM solver to an engineering model of the aerothermodynamic heating environment the impact of non catalytic surfaces, rarefied flow effects, and multidimensional conduction on the performance envelopes of sharp leading edges can be examined.
Boundary Layer Transition Experiments in Support of the Hypersonics Program
NASA Technical Reports Server (NTRS)
Berry, Scott A.; Chen, Fang-Jenq; Wilder, Michael C.; Reda, Daniel C.
2007-01-01
Two experimental boundary layer transition studies in support of fundamental hypersonics research are reviewed. The two studies are the HyBoLT flight experiment and a new ballistic range effort. Details are provided of the objectives and approach associated with each experimental program. The establishment of experimental databases from ground and flight are to provide better understanding of high-speed flows and data to validate and guide the development of simulation tools.
Hypersonic technology-approach to an expanded program
NASA Technical Reports Server (NTRS)
Hearth, D. P.; Preyss, A. E.
1976-01-01
An overview of research, testing, and technology in the hypersonic range. Military and civilian hypersonic flight systems envisaged, ground testing facilities under development, methods for cooling the heated airframe, and use of hydrogen as fuel and coolant are discussed extensively. Air-breathing hypersonic cruise systems are emphasized, the airframe-integrated scramjet configuration is discussed and illustrated, materials proposed for hypersonic vehicles are reviewed, and test results on hypersonic flight (X-15 research aircraft) are indicated. Major advances and major problems in hypersonic flight and hypersonic technology are outlined, and the need for a hypersonic flying-laboratory research craft is stressed.
NASA Technical Reports Server (NTRS)
Gunness, R. C., Jr.; Knight, C. J.; Dsylva, E.
1972-01-01
The unified small disturbance equations are numerically solved using the well-known Lax-Wendroff finite difference technique. The method allows complete determination of the inviscid flow field and surface properties as long as the flow remains supersonic. Shock waves and other discontinuities are accounted for implicity in the numerical method. This technique was programed for general application to the three-dimensional case. The validity of the method is demonstrated by calculations on cones, axisymmetric bodies, lifting bodies, delta wings, and a conical wing/body combination. Part 1 contains the discussion of problem development and results of the study. Part 2 contains flow charts, subroutine descriptions, and a listing of the computer program.
NASA Technical Reports Server (NTRS)
Knight, Doyle D.; Becht, Robert J.
1995-01-01
The objective of the current research is the development of an improved k-epsilon two-equation compressible turbulence model for turbulent boundary layer flows experiencing strong viscous-inviscid interactions. The development of an improved model is important in the design of hypersonic vehicles such as the National Aerospace Plane (NASP) and the High Speed Civil Transport (HSCT). Improvements have been made to the low Reynolds number functions in the eddy viscosity and dissipation of solenoidal dissipation of the k-epsilon turbulence mode. These corrections offer easily applicable modifications that may be utilized for more complex geometries. The low Reynolds number corrections are functions of the turbulent Reynolds number and are therefore independent of the coordinate system. The proposed model offers advantages over some current models which are based upon the physical distance from the wall, that modify the constants of the standard model, or that make more corrections than are necessary to the governing equations. The code has been developed to solve the Favre averaged, boundary layer equations for mass, momentum, energy, turbulence kinetic energy, and dissipation of solenoidal dissipation using Keller's box scheme and the Newton spatial marching method. The code has been validated by removing the turbulent terms and comparing the solution with the Blasius solution, and by comparing the turbulent solution with an existing k-epsilon model code using wall function boundary conditions. Excellent agreement is seen between the computed solution and the Blasius solution, and between the two codes. The model has been tested for both subsonic and supersonic flat-plate turbulent boundary layer flow by comparing the computed skin friction with the Van Driest II theory and the experimental data of Weighardt; by comparing the transformed velocity profile with the data of Weighardt, and the Law of the Wall and the Law of the Wake; and by comparing the computed results
Generic hypersonic vehicle performance model
NASA Technical Reports Server (NTRS)
Chavez, Frank R.; Schmidt, David K.
1993-01-01
An integrated computational model of a generic hypersonic vehicle was developed for the purpose of determining the vehicle's performance characteristics, which include the lift, drag, thrust, and moment acting on the vehicle at specified altitude, flight condition, and vehicular configuration. The lift, drag, thrust, and moment are developed for the body fixed coordinate system. These forces and moments arise from both aerodynamic and propulsive sources. SCRAMjet engine performance characteristics, such as fuel flow rate, can also be determined. The vehicle is assumed to be a lifting body with a single aerodynamic control surface. The body shape and control surface location are arbitrary and must be defined. The aerodynamics are calculated using either 2-dimensional Newtonian or modified Newtonian theory and approximate high-Mach-number Prandtl-Meyer expansion theory. Skin-friction drag was also accounted for. The skin-friction drag coefficient is a function of the freestream Mach number. The data for the skin-friction drag coefficient values were taken from NASA Technical Memorandum 102610. The modeling of the vehicle's SCRAMjet engine is based on quasi 1-dimensional gas dynamics for the engine diffuser, nozzle, and the combustor with heat addition. The engine has three variable inputs for control: the engine inlet diffuser area ratio, the total temperature rise through the combustor due to combustion of the fuel, and the engine internal expansion nozzle area ratio. The pressure distribution over the vehicle's lower aft body surface, which acts as an external nozzle, is calculated using a combination of quasi 1-dimensional gas dynamic theory and Newtonian or modified Newtonian theory. The exhaust plume shape is determined by matching the pressure inside the plume, calculated from the gas dynamic equations, with the freestream pressure, calculated from Newtonian or Modified Newtonian theory. In this manner, the pressure distribution along the vehicle after body
HIAD-2 (Hypersonic Inflatable Aerodynamic Decelerator)
The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) project is a disruptive technology that will accommodate the atmospheric entry of heavy payloads to planetary bodies such as Mars. HIAD over...
NASA Technical Reports Server (NTRS)
Eggers, A J , Jr; Savin, Raymond C
1955-01-01
A procedure for calculating three-dimensional steady and nonsteady supersonic flows with the method of characteristics is developed and discussed. An approximate method is deduced from the characteristics method and shown to be of practical value at high supersonic speeds.
Hypersonic Wind Tunnel Test of a Flare-type Membrane Aeroshell for Atmospheric Entry Capsules
NASA Astrophysics Data System (ADS)
Yamada, Kazuhiko; Koyama, Masashi; Kimura, Yusuke; Suzuki, Kojiro; Abe, Takashi; Koichi Hayashi, A.
A flexible aeroshell for atmospheric entry vehicles has attracted attention as an innovative space transportation system. In this study, hypersonic wind tunnel tests were carried out to investigate the behavior, aerodynamic characteristics and aerodynamic heating environment in hypersonic flow for a previously developed capsule-type vehicle with a flare-type membrane aeroshell made of ZYLON textile sustained by a rigid torus frame. Two different models with different flare angles (45º and 60º) were tested to experimentally clarify the effect of flare angle. Results indicate that flare angle of aeroshell has significant and complicate effect on flow field and aerodynamic heating in hypersonic flow at Mach 9.45 and the flare angle is very important parameter for vehicle design with the flare-type membrane aeroshell.
Hypersonic modes in nanophononic semiconductors.
Hepplestone, S P; Srivastava, G P
2008-09-01
Frequency gaps and negative group velocities of hypersonic phonon modes in periodically arranged composite semiconductors are presented. Trends and criteria for phononic gaps are discussed using a variety of atomic-level theoretical approaches. From our calculations, the possibility of achieving semiconductor-based one-dimensional phononic structures is established. We present results of the location and size of gaps, as well as negative group velocities of phonon modes in such structures. In addition to reproducing the results of recent measurements of the locations of the band gaps in the nanosized Si/Si{0.4}Ge{0.6} superlattice, we show that such a system is a true one-dimensional hypersonic phononic crystal. PMID:18851224
Systems Challenges for Hypersonic Vehicles
NASA Technical Reports Server (NTRS)
Hunt, James L.; Laruelle, Gerard; Wagner, Alain
1997-01-01
This paper examines the system challenges posed by fully reusable hypersonic cruise airplanes and access to space vehicles. Hydrocarbon and hydrogen fueled airplanes are considered with cruise speeds of Mach 5 and 10, respectively. The access to space matrix is examined. Airbreathing and rocket powered, single- and two-stage vehicles are considered. Reference vehicle architectures are presented. Major systems/subsystems challenges are described. Advanced, enhancing systems concepts as well as common system technologies are discussed.
Calculation of hypersonic shock structure using flux-split algorithms
NASA Technical Reports Server (NTRS)
Eppard, W. M.; Grossman, B.
1991-01-01
There exists an altitude regime in the atmosphere that is within the continuum domain, but wherein the conventional Navier-Stokes equations cease to be accurate. The altitude limits for this so called continuum transition regime depend on vehicle size and speed. Within this regime the thickness of the bow shock wave is no longer negligible when compared to the shock stand-off distance and the peak radiation intensity occurs within the shock wave structure itself. For this reason it is no longer valid to treat the shock wave as a discontinuous jump and it becomes necessary to compute through the shock wave itself. To accurately calculate hypersonic flowfields, the governing equations must be capable of yielding realistic profiles of flow variables throughout the structure of a hypersonic shock wave. The conventional form of the Navier-Stokes equations is restricted to flows with only small departures from translational equilibrium; it is for this reason they do not provide the capability to accurately predict hypersonic shock structure. Calculations in the continuum transition regime, therefore, require the use of governing equations other than Navier-Stokes. Several alternatives to Navier-Stokes are discussed; first for the case of a monatomic gas and then for the case of a diatomic gas where rotational energy must be included. Results are presented for normal shock calculations with argon and nitrogen.
Observation and tuning of hypersonic bandgaps in colloidal crystals.
Cheng, Wei; Wang, Jianjun; Jonas, Ulrich; Fytas, George; Stefanou, Nikolaos
2006-10-01
Composite materials with periodic variations of density and/or sound velocities, so-called phononic crystals, can exhibit bandgaps where propagation of acoustic waves is forbidden. Phononic crystals are the elastic analogue of the well-established photonic crystals and show potential for manipulating the flow of elastic energy. So far, the experimental realization of phononic crystals has been restricted to macroscopic systems with sonic or ultrasonic bandgaps in the sub-MHz frequency range. In this work, using high-resolution Brillouin spectroscopy we report the first observation of a hypersonic bandgap in face-centred-cubic colloidal crystals formed by self-assembly of polystyrene nanoparticles with subsequent fluid infiltration. Depending on the particle size and the sound velocity in the infiltrated fluid, the frequency and the width of the gap can be tuned. Promising technological applications of hypersonic crystals, ranging from tunable filters and heat management to acousto-optical devices, are anticipated. PMID:16951677
Hypersonic trans-Pacific flight
NASA Technical Reports Server (NTRS)
1987-01-01
The Advanced Aeronautics Design Program at The Ohio State University was to design a vehicle for hypersonic passenger flight across the Pacific Ocean. The specifications were as follows: (1) hypersonic flight; (2) range of 8000 nm; (3) passenger seating greater than 250; (4) operation from 15000 ft runways Mach number and altitude of operation were at the discretion of the design teams as were the propulsion system and type of fuel. The advanced aeronautics design sequence established specifically for this program consisted of a three quarter sequence as follows: Fall: ME 694 Senior Design Seminar - one quarter hour. Designers and specialists met one hour each week for ten weeks on relevant flight vehicle design topics. Winter: ME 515H Flight Vehicle Design - four quarter hours. Three design teams of six students each performed preliminary design studies of hypersonic configurations and potential propulsion systems. Each team's results were summarized in a final presentation to NASA Lewis Research Center personnel. The presentations resulted in the selection of the most promising design for additional development. Spring: AAE 516H Advanced Flight Vehicle Design - four quarter hrs. The class was reorganized to focus upon the specific design selected from the Winter configuration studies. Detailed analyses of thermal protection systems, costs, mission refinements, etc., completed the design task and final presentations were made to NASA Lewis Research Center staff.
Mace, R.E. . Bureau of Economic Geology)
1993-02-01
Numerical models are useful tools for developing an understanding of ground-water flow in sparsely characterized low-permeability aquifers. Finite-difference, cross-sectional models of Cretaceous chalk and marl formations near the Superconducting Super Collider (SSC) were constructed using MODFLOW to evaluate ground-water circulation paths and travel times. Weathered and fractured zones with enhanced permeability were included to assess the effect these features had on flow paths and times. Pump tests, slug tests, packer tests, core tests, and estimates were used to define hydraulic properties for model input. The model was calibrated with water-level data from monitor wells and from wire-line piezometers near a test shaft excavated by the SSC project. A ratio of vertical-to-horizontal permeability of 0.0085 was estimated through model calibration. A chalk-to-marl permeability ratio of 18 was needed to reproduce artesian head in a well completed in chalk beneath marl. Hydraulic head distributions and ground-water flow paths reflected local, intermediate, and regional flow systems with recharge beneath upland surface-water divides and discharge in valleys. Most of the flow (99%) occurred in the weathered zone, with average residence times of 5 to 10 years. Residence time in unweathered chalk bedrock was substantially longer, at an average of 1.7 Ma. As expected, the model demonstrated that deep and rapid ground-water circulation might occur in fracture zones. Particle paths calculated using MODPATH showed that ground-water travel times from recharge areas to the SSC subsurface facilities might be 20 to 60 years where flow is through fracture zones.
Trinks, O; Beck, W H
1998-10-20
With a first application of semiconductor lasers to absorption measurements of seeded atomic Rb in high-enthalpy flow fields, a diagnostic technique for time-resolved determination of flow velocity and gas temperature with a line-shape analysis was developed. In our measurements a GaAlAs diode laser was used to scan repetitively at 15 kHz over 1.3 cm(-1) across the D(2) resonance transition (5S(1/2) ? 5P(3/2), 780.2 nm) of seeded atomic Rb to obtain multiple absorption line shapes. The time-dependent signal contains highly resolved spectral line-shape information, which we interpret by fitting the spectrally resolved line shapes to Voigt profiles. Kinetic temperatures in the range 900-1400 K and gas velocities in the range 3900-6200 ms(-1) were obtained from the Doppler-broadened component of the line shape and from the Doppler shift, respectively, of the absorption frequency. PMID:18301526
Non-Equilibrium Effects on Hypersonic Turbulent Boundary Layers
NASA Astrophysics Data System (ADS)
Kim, Pilbum
Understanding non-equilibrium effects of hypersonic turbulent boundary layers is essential in order to build cost efficient and reliable hypersonic vehicles. It is well known that non-equilibrium effects on the boundary layers are notable, but our understanding of the effects are limited. The overall goal of this study is to improve the understanding of non-equilibrium effects on hypersonic turbulent boundary layers. A new code has been developed for direct numerical simulations of spatially developing hypersonic turbulent boundary layers over a flat plate with finite-rate reactions. A fifth-order hybrid weighted essentially non-oscillatory scheme with a low dissipation finite-difference scheme is utilized in order to capture stiff gradients while resolving small motions in turbulent boundary layers. The code has been validated by qualitative and quantitative comparisons of two different simulations of a non-equilibrium flow and a spatially developing turbulent boundary layer. With the validated code, direct numerical simulations of four different hypersonic turbulent boundary layers, perfect gas and non-equilibrium flows of pure oxygen and nitrogen, have been performed. In order to rule out uncertainties in comparisons, the same inlet conditions are imposed for each species, and then mean and turbulence statistics as well as near-wall turbulence structures are compared at a downstream location. Based on those comparisons, it is shown that there is no direct energy exchanges between internal and turbulent kinetic energies due to thermal and chemical non-equilibrium processes in the flow field. Instead, these non-equilibria affect turbulent boundary layers by changing the temperature without changing the main characteristics of near-wall turbulence structures. This change in the temperature induces the changes in the density and viscosity and the mean flow fields are then adjusted to satisfy the conservation laws. The perturbation fields are modified according to
Experimental research of the aerodynamics of nozzles and plumes at hypersonic speeds
NASA Technical Reports Server (NTRS)
Keener, Earl R.
1992-01-01
The purpose was 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 of the NASA Ames Research Center. The model design and test planning were performed in close cooperation with members of the National Aero-Space Plane (NASP) computational fluid dynamics (SFD) team, so that the measurements could be used in CFD code validation studies. Presented here is a description of the experiment, the extent of the measurements obtained, and the experimental results.
Analysis of hypersonic nozzles including vibrational nonequilibrium and intermolecular force effects
NASA Technical Reports Server (NTRS)
Canupp, Patrick W.; Candler, Graham V.; Perkins, John N.; Erickson, Wayne D.
1992-01-01
A computational fluid dynamics algorithm is developed for the study of high-pressure axisymmetric hypersonic nozzle flows. The effects of intermolecular forces and vibrational nonequilibrium are included in the analysis. The numerical simulation of gases with an arbitrary equation of state is discussed. Simulations for a high pressure nozzle (p(0) = 138 MPa) demonstrate that both intermolecular forces and vibrational nonequilibrium have a significant affect on the flow. These nonideal effects tend to increase the Mach number at the nozzle exit plane. Thus, they must be included in the design and analysis of high pressure hypersonic nozzles.
Smart structures applications for hypersonic vehicles
NASA Astrophysics Data System (ADS)
August, James A.; Joshi, Shiv P.
1996-05-01
A survey of current literature was performed and vehicle designers from the aerospace industry were polled to examine how state of the art smart structural concepts could improve the design of hypersonic vehicles. Several types of hypersonic vehicles; including winged single stage to orbit, sub-orbital cruise aircraft, and supersonic/hypersonic missiles have demanding airframe and systems requirements which may not be sufficiently met with traditional structural designs. The use of smart structures is examined to improve vehicle performance in areas such as active vibration control, noise reduction, vehicle attitude control, structural cooling, and engine performance. The operating environment of hypersonic vehicles are examined and the capabilities of currently used structural materials and actuators are compared with those of smart materials and structures. Possible smart structures applications are presented as modifications to existing designs as well as new structural concepts. Conclusions are made on the suitability of various smart structures concepts for current and future hypersonic applications.
Conceptual Design and Numerical Simulations of Hypersonic Waverider Vehicle
NASA Astrophysics Data System (ADS)
Cao, D. Y.; Zhang, J. B.; Lee, C. H.
A modularized airframe/propulsion integrated model is established by oblique shock wave theory, engineering method and method of characteristics(MOC). Based on this method, a new design methodology for hypersonic waverider vehicle which integrated scramjets with waverider airframe derived from cone-wedge flow field is presented. Integrated aero-propulsion performance of the waverider vehicle under on-design and off-design conditions is predicted using Euler equations discretized by Harten-Yee non-MUSCL TVD scheme and the combustor flow field is approximated by a quasi-ID cycle analysis, skin friction of vehicle is calculated by reference temperature method.
Status of turbulence modeling for hypersonic propulsion flowpaths
NASA Astrophysics Data System (ADS)
Georgiadis, Nicholas J.; Yoder, Dennis A.; Vyas, Manan A.; Engblom, William A.
2014-06-01
This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer methods such as large eddy simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath, including laminar-to-turbulent boundary layer transition, shock wave/turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers), and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.
Hypersonic Navier Stokes Comparisons to Orbiter Flight Data
NASA Technical Reports Server (NTRS)
Campbell, Charles H.; Nompelis, Ioannis; Candler, Graham; Barnhart, Michael; Yoon, Seokkwan
2009-01-01
Hypersonic chemical nonequilibrium simulations of low earth orbit entry flow fields are becoming increasingly commonplace as software and computational capabilities become more capable. However, development of robust and accurate software to model these environments will always encounter a significant barrier in developing a suite of high quality calibration cases. The US3D hypersonic nonequilibrium Navier Stokes analysis capability has been favorably compared to a number of wind tunnel test cases. Extension of the calibration basis for this software to Orbiter flight conditions will provide an incremental increase in confidence. As part of the Orbiter Boundary Layer Transition Flight Experiment and the Hypersonic Thermodynamic Infrared Measurements project, NASA is performing entry flight testing on the Orbiter to provide valuable aerothermodynamic heating data. An increase in interest related to orbiter entry environments is resulting from this activity. With the advent of this new data, comparisons of the US3D software to the new flight testing data is warranted. This paper will provide information regarding the framework of analyses that will be applied with the US3D analysis tool. In addition, comparisons will be made to entry flight testing data provided by the Orbiter BLT Flight Experiment and HYTHIRM projects. If data from digital scans of the Orbiter windward surface become available, simulations will also be performed to characterize the difference in surface heating between the CAD reference OML and the digitized surface provided by the surface scans.
Status of Turbulence Modeling for Hypersonic Propulsion Flowpaths
NASA Technical Reports Server (NTRS)
Georgiadis, Nicholas J.; Yoder, Dennis A.; Vyas, Manan A.; Engblom, William A.
2012-01-01
This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer meth- ods such as Large Eddy Simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath including laminar-to-turbulent boundary layer transition, shock wave / turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers) and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.
Airbreathing Hypersonic Technology Vision Vehicles and Development Dreams
NASA Technical Reports Server (NTRS)
McClinton, C. R.; Hunt, J. L.; Ricketts, R. H.; Reukauf, P.; Peddie, C. L.
1999-01-01
Significant advancements in hypersonic airbreathing vehicle technology have been made in the country's research centers and industry over the past 40 years. Some of that technology is being validated with the X-43 flight tests. This paper presents an overview of hypersonic airbreathing technology status within the US, and a hypersonic technology development plan. This plan builds on the nation's large investment in hypersonics. This affordable, incremental plan focuses technology development on hypersonic systems, which could be operating by the 2020's.
Hypersonic Airbreathing Vehicles/Technologies
NASA Technical Reports Server (NTRS)
Hunt, James L.
1996-01-01
Hypersonic airbreathing horizontal takeoff and landing (HTOL) vehicles are highly integrated systems involving many advanced technologies. The design environment is variable rich, intricately networked, and sensitivity intensive; as such, it represents a tremendous challenge. Creating a viable design requires addressing three main elements: (1) an understanding of the 'figures of merit' and their relationship, (2) the development of sophisticated configuration discipline prediction methods and a synthesis procedure, and (3) the synergistic integration of advanced technologies across the discipline spectrum. This paper will focus on the vision for hypersonic airbreathing vehicles and the advanced technologies that forge the designs. Airbreathing hypersonics encompass endoatmospheric (airplanes...missiles are a part of the matrix but will not be included in this paper since they are an air force focus) and space access vehicles with speed from Mach 4 up to Mach 25 (orbital). These vehicles can be divided into two classes...cruisers and accelerators. The cruiser designs reflect high lift-to-drag whereas the accelerators reflect low drag per unit inlet capture; thus, the cross section of the accelerator attributes a much larger percentage to propulsion. One of the more design influencing items is fuel. The hydrogen fueled vehicles must be very volumetric efficient to contain the low density fuel and thus tend to be a bit bulgy (more conducive to lifting bodies or wing bodies) whereas with hydrocarbon fueled vehicles, the concern is loading because of the high density fuel; thus, they may tend to be more towards waveriders which are not usually very volumetric efficient. Hydrocarbon fuels (endothermic) are limited in their engine cooling capacity to below Mach 8.
Transpiration cooling in hypersonic flight
NASA Technical Reports Server (NTRS)
Tavella, Domingo; Roberts, Leonard
1989-01-01
A preliminary numerical study of transpiration cooling applied to a hypersonic configuration is presented. Air transpiration is applied to the NASA all-body configuration flying at an altitude of 30500 m with a Mach number of 10.3. It was found that the amount of heat disposal by convection is determined primarily by the local geometry of the aircraft for moderate rates of transpiration. This property implies that different areas of the aircraft where transpiration occurs interact weakly with each other. A methodology for quick assessments of the transpiration requirements for a given flight configuration is presented.
Novel inlet-airframe integration methodology for hypersonic waverider vehicles
NASA Astrophysics Data System (ADS)
Ding, Feng; Liu, Jun; Shen, Chi-bing; Huang, Wei
2015-06-01
With the aim of integrating a ramjet or scramjet with an airframe, a novel inlet-airframe integration methodology for the hypersonic waverider vehicle is proposed. For this newly proposed design concept and for the specified flight conditions, not only the forebody of the vehicle but also its engine cowl and wings can ride on the bow shock wave, causing the bow shock wave to remain attached to the leading edge for the entire length of the vehicle. Thus, this integrated vehicle can take full advantage of the waverider's high lift-to-drag ratio characteristics and the ideal pre-compression surface for the engine. In this work, a novel inlet-airframe integrated axisymmetric basic flow model that accounts for both external and internal flows is first designed using the method of characteristics and the streamline tracing technique. Subsequently, the design of the inlet-airframe integrated waverider vehicle is generated from the integrated axisymmetric basic flow model using the streamline tracing technique. Then, the design methodologies of both the integrated axisymmetric basic flow model and the integrated waverider vehicle are verified by a computational numerical method. Finally, the viscous effects and performance of both the integrated axisymmetric basic flow model and the integrated waverider vehicle are analysed under the design condition using the numerical computation. The obtained results show that the proposed approach is effective in designing the integrated hypersonic waverider vehicle.
X-43 Hypersonic Vehicle Technology Development
NASA Technical Reports Server (NTRS)
Voland, Randall T.; Huebner, Lawrence D.; McClinton, Charles R.
2005-01-01
NASA recently completed two major programs in Hypersonics: Hyper-X, with the record-breaking flights of the X-43A, and the Next Generation Launch Technology (NGLT) Program. The X-43A flights, the culmination of the Hyper-X Program, were the first-ever examples of a scramjet engine propelling a hypersonic vehicle and provided unique, convincing, detailed flight data required to validate the design tools needed for design and development of future operational hypersonic airbreathing vehicles. Concurrent with Hyper-X, NASA's NGLT Program focused on technologies needed for future revolutionary launch vehicles. The NGLT was "competed" by NASA in response to the President s redirection of the agency to space exploration, after making significant progress towards maturing technologies required to enable airbreathing hypersonic launch vehicles. NGLT quantified the benefits, identified technology needs, developed airframe and propulsion technology, chartered a broad University base, and developed detailed plans to mature and validate hypersonic airbreathing technology for space access. NASA is currently in the process of defining plans for a new Hypersonic Technology Program. Details of that plan are not currently available. This paper highlights results from the successful Mach 7 and 10 flights of the X-43A, and the current state of hypersonic technology.
Multiphysics Simulation of Active Hypersonic Lip Cooling
NASA Technical Reports Server (NTRS)
Melis, Matthew E.; Wang, Wen-Ping
1999-01-01
This article describes the application of the Multidisciplinary Analysis (MDA) solver, Spectrum, in analyzing a hydrogen-cooled hypersonic cowl leading-edge structure. Spectrum, a multiphysics simulation code based on the finite element method, addresses compressible and incompressible fluid flow, structural, and thermal modeling, as well as the interactions between these disciplines. Fluid-solid-thermal interactions in a hydrogen impingement-cooled leading edge are predicted using Spectrum. Two- and semi-three-dimensional models are considered for a leading edge impingement coolant, concept under either specified external heat flux or aerothermodynamic heating from a Mach 5 external flow interaction. The solution accuracy is demonstrated from mesh refinement analysis. With active cooling, the leading edge surface temperature is drastically reduced from 1807 K of the adiabatic condition to 418 K. The internal coolant temperature profile exhibits a sharp gradient near channel/solid interface. Results from two different cooling channel configurations are also presented to illustrate the different behavior of alternative active cooling schemes.
A methodology for hypersonic transport technology planning
NASA Technical Reports Server (NTRS)
Repic, E. M.; Olson, G. A.; Milliken, R. J.
1973-01-01
A systematic procedure by which the relative economic value of technology factors affecting design, configuration, and operation of a hypersonic cruise transport can be evaluated is discussed. Use of the methodology results in identification of first-order economic gains potentially achievable by projected advances in each of the definable, hypersonic technologies. Starting with a baseline vehicle, the formulas, procedures and forms which are integral parts of this methodology are developed. A demonstration of the methodology is presented for one specific hypersonic vehicle system.
Hypersonic transports - Economics and environmental effects.
NASA Technical Reports Server (NTRS)
Petersen, R. H.; Waters, M. H.
1972-01-01
An economic analysis of hypersonic transports is presented to show projected operating costs (direct and indirect) and return on investment. Important assumptions are varied to determine the probable range of values for operating costs and return on investment. The environmental effects of hypersonic transports are discussed and compared to current supersonic transports. Estimates of sideline and flyover noise are made for a typical hypersonic transport, and the sonic boom problem is analyzed and discussed. Since the exhaust products from liquid hydrogen-fueled engines differ from those of kerosene-fueled aircraft, a qualitative assessment of air pollution effects is made.
Hypersonic transports: Economics and environmental effects
NASA Technical Reports Server (NTRS)
Petersen, R. H.; Waters, M. H.
1972-01-01
An economic analysis of hypersonic transports is presented to show projected operating costs (direct and indirect) and return on investment. Important assumptions are varied to determine the probable range of values for operating costs and return on investment. The environmental effects of hypersonic transports are discussed and compared to current supersonic transports. Estimates of sideline and fly-over noise are made for a typical hypersonic transport, and the sonic boom problem is analyzed and discussed. Since the exhaust products from liquid hydrogen-fueled engines differ from those of kerosene-fueled aircraft, a qualitative assessment of air pollution effects is made.
Hypersonic transports - Economics and environmental effects.
NASA Technical Reports Server (NTRS)
Petersen, R. H.; Waters, M. H.
1973-01-01
An economic analysis of hypersonic transports is presented to show projected operating costs (direct and indirect) and return on investment. Important assumptions are varied to determine the probable range of values for operating costs and return on investment. The environmental effects of hypersonic transports are discussed and compared to current supersonic transports. Estimates of sideline and flyover noise are made for a typical hypersonic transport, and the sonic boom problem is analyzed and discussed. Since the exhaust products from liquid hydrogen-fueled engines differ from those of kerosene-fueled aircraft, a qualitative assessment of air pollution effects is made.
NASA Astrophysics Data System (ADS)
Estorf, M.; Wolf, T.; Radespiel, R.
2005-02-01
In summer 2003 the Hypersonic Ludwieg Tube Braunschweig (HLB) has been commissioned. It has been designed for Machnumber M = 6. The operative range of the unit Reynoldsnumber is between 3 · 106 1/m and 20 · 106 1/m. The test section has 500 mm diameter and the run time with near steady flow conditions is 80 ms. First measurements of the pitot pressure within the test-section have shown good transverse uniformity, but measurements of the total temperature within the storage tube revealed a strong stratification accompanied by convective flow within the heated section of the tube. In this paper we present recent measures to attenuate the stratification. Measurements of the temperature distribution within the testsection have been performed. Further we compare first results of invisvid numerical simulations of the unsteady onset of flow in the HLB to measured data. Key words: hypersonic flow, Ludwieg tube, flow measurement techniques, numerical flow simulation.
NASA Technical Reports Server (NTRS)
Kaul, U. K.
1988-01-01
Computations of the hypersonic flow around sharp cones were carried out using the PNS code with attention given to the heat transfer predictions around the transition region. Results of calculations performed over 5, 8, and 10 deg half-angle sharp cones in the Mach number range of 7 to 10 are presented. It is noted that calculations of this type have become an integral part of the general design procedure for hypersonic vehicles such as the National Aerospace Plane and the Space Shuttle.
Key technique study and application of infrared thermography in hypersonic wind tunnel
NASA Astrophysics Data System (ADS)
LI, Ming; Yang, Yan-guang; Li, Zhi-hui; Zhu, Zhi-wei; Zhou, Jia-sui
2014-11-01
The solutions to some key techniques using infrared thermographic technique in hypersonic wind tunnel, such as temperature measurement under great measurement angle, the corresponding relation between model spatial coordinates and the ones in infrared map, the measurement uncertainty analysis of the test data etc., are studied. The typical results in the hypersonic wind tunnel test are presented, including the comparison of the transfer rates on a thin skin flat plate model with a wedge measured with infrared thermography and thermocouple, the experimental study heating effect on the flat plate model impinged by plume flow and the aerodynamic heating on the lift model.
NASA Technical Reports Server (NTRS)
Scott, Carl D.
1989-01-01
An account is given of the function of physical aspects of a gas on the characteristics of the flow and of the heating associated with hypersonic flight. At the high temperatures encountered, the thermal and chemical characteristics of the air in a hypersonic vehicle's shock layer are altered in ways which depend on the atomic and molecular structure of N and O and their ions; similar effects exist in scramjet propulsion systems. These properties in turn influence the character of shock waves and expansions, and hence the pressure, temperature, and velocity distributions. Transport properties affecting the boundary-layer structure will also affect heat flux and shear stress.
Design of a convective cooling system for a Mach 6 hypersonic transport airframe
NASA Technical Reports Server (NTRS)
Helenbrook, R. G.; Anthony, F. M.
1971-01-01
Results of analytical and design studies are presented for a water-glycol convective cooling system for the airframe structure of a hypersonic transport. System configurations and weights are compared. The influences of system pressure drop and flow control schedules on system weight are defined.
Overview of X-38 Hypersonic Aerothermodynamic Wind Tunnel Data and Comparison with Numerical Results
NASA Technical Reports Server (NTRS)
Campbell, C.; Caram, J.; Berry, S.; Horvath, T.; Merski, N.; Loomis, M.; Venkatapathy, E.
2004-01-01
A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic aerothermodynamic characteristics of the X-23/X-24A derived X-38 crew return vehicle are being evaluated in various wind tunnels in support of this effort. Aerothermodynamic data from two NASA hypersonic tunnels at Mach 6 and Mach 10 has been obtained with cast ceramic models and a thermographic phosphorus digital imaging system. General windward surface heating features are described based on experimental surface heating images and surface oil flow patterns for the nominal hypersonic aerodynamic orientation. Body flap reattachment heating levels are examined. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with this data.
Volume interchange factors for hypersonic vehicle wake radiation
NASA Technical Reports Server (NTRS)
Edwards, D. K.; Babikian, D. S.
1987-01-01
Volume interchange factors are shown to be convenient in modeling the radiative processes in the wake of a hypersonic vehicle. Use of the factors facilitates calculating not just the radiative heating rates on afterbody surfaces but also the radiative de-excitation rates from stimulated emission and re-excitation rates from absorption in rarefied nonequilibrium flows. Sample calculations of volume interchange factors are presented for volume configurations modeling wake elements, and the numerical results are compared to limiting approximations to clarify the operation of the emission, transmission, and absorption processes.
Studies of engine-airframe integrated hypersonic vehicles
NASA Technical Reports Server (NTRS)
Saland, H.; Fox, H.; Hoydysh, W.
1972-01-01
A parametric study of an integrated airframe and engine is presented for a hypersonic transport at an altitude of 70,000 feet and a free stream Mach number of 6. The engine considered is a subsonic combustion ramjet using conventional hydrocarbon fuels. The lift-to-drag ratio of the aircraft for two configurations, one with full capture and accelerated flight and the other allowing spillage of the leading shock and in unaccelerated flight, is studied. The parameters varied are the engine efficiencies, the angle of attack, the combustion rates, as well as the captured mass flow. Lift-to-drag ratios on the order of 6.5 are obtained.
Hypersonic Interplanetary Flight: Aero Gravity Assist
NASA Technical Reports Server (NTRS)
Bowers, Al; Banks, Dan; Randolph, Jim
2006-01-01
The use of aero-gravity assist during hypersonic interplanetary flights is highlighted. Specifically, the use of large versus small planet for gravity asssist maneuvers, aero-gravity assist trajectories, launch opportunities and planetary waverider performance are addressed.
Cooling/fuel system for hypersonic flight
Lander, H.R.; Schnurstein, R.E.
1993-08-17
A method is described of simultaneously providing a heat sink and reactive fuel factions production from hydrocarbons having an average molecular weight of between 100 and 1,000 in hypersonic propulsion applications comprising: (i) in a hypersonic vehicle having high heat flux structural regions, causing a hydrocarbon exposure to a high heat flux structural region and imparting a temperature increase to the hydrocarbon; (ii) reducing temperature gradients of the high heat flux structural region by heat transfer from the high flux structural region to the hydrocarbon such that a portion of the hydrocarbon pyrolyzes into olefinic fractions; and (iii) utilizing the olefinic fractions as a fuel in hypersonic propulsion in a hypersonic vehicle.
NASA Astrophysics Data System (ADS)
Palmer, R. B.; Gallardo, J. C.
INTRODUCTION PHYSICS CONSIDERATIONS GENERAL REQUIRED LUMINOSITY FOR LEPTON COLLIDERS THE EFFECTIVE PHYSICS ENERGIES OF HADRON COLLIDERS HADRON-HADRON MACHINES LUMINOSITY SIZE AND COST CIRCULAR e^{+}e^- MACHINES LUMINOSITY SIZE AND COST e^{+}e^- LINEAR COLLIDERS LUMINOSITY CONVENTIONAL RF SUPERCONDUCTING RF AT HIGHER ENERGIES γ - γ COLLIDERS μ ^{+} μ^- COLLIDERS ADVANTAGES AND DISADVANTAGES DESIGN STUDIES STATUS AND REQUIRED R AND D COMPARISION OF MACHINES CONCLUSIONS DISCUSSION
Research in robust control for hypersonic aircraft
NASA Technical Reports Server (NTRS)
Calise, A. J.
1993-01-01
The research during the second reporting period has focused on robust control design for hypersonic vehicles. An already existing design for the Hypersonic Winged-Cone Configuration has been enhanced. Uncertainty models for the effects of propulsion system perturbations due to angle of attack variations, structural vibrations, and uncertainty in control effectiveness were developed. Using H(sub infinity) and mu-synthesis techniques, various control designs were performed in order to investigate the impact of these effects on achievable robust performance.
Nonparallel instability of supersonic and hypersonic boundary layers
NASA Technical Reports Server (NTRS)
El-Hady, Nabil M.
1991-01-01
Multiple scaling technique is used to examine the nonparallel instability of supersonic and hypersonic boundary-layer flows to three dimensional (first mode) and two dimensional (second mode) disturbances. The method is applied to the flat plate boundary layer for a range of Mach numbers from 0 to 10. Growth rates of disturbances are calculated based on three different criteria: following the maximum of the mass-flow disturbance, using an integral of the disturbance kinetic energy, and using the integral of the square of the mass-flow amplitude. By following the maximum of the mass-flow disturbance, the calculated nonparallel growth rates are in good quantitative agreement with the experimental results at Mach number 4.5.
Nonparallel instability of supersonic and hypersonic boundary layers
NASA Technical Reports Server (NTRS)
El-Hady, Nabil M.
1991-01-01
Multiple scaling technique is used to examine the nonparallel instability of supersonic and hypersonic boundary-layer flows to three-dimensional (first mode) and two-dimensional (second mode) disturbances. The method is applied to the flat plate boundary layer for a range of Mach numbers from 0 to 10. Growth rates of disturbances are calculated based on three different criteria: following the maximum of the mass-flow disturbance, using an integral of the disturbance kinetic energy, and using an integral of the square of the mass-flow amplitude. By following the maximum of the mass-flow dusturbance, the calculated nonparallel growth rates are in good quantitative agreement with the experimental results of Kendall (1967) at Mach number 4.5.
Nonparallel instability of supersonic and hypersonic boundary layers
NASA Technical Reports Server (NTRS)
El-Hady, Nabil M.
1991-01-01
Multiple scaling technique is used to examine the nonparallel instability of supersonic and hypersonic boundary-layer flows to three-dimensional (first mode) and two-dimensional (second mode) disturbances. The method is applied to the flat plate boundary layer for a range of Mach numbers from 0 to 10. Growth rates of disturbances are calculated based on three different criteria: following the maximum of the mass-flow disturbance, using an integral of the disturbance kinetic energy, and using the integral of the square of the mass-flow amplitude. By following the maximum of the mass-flow disturbance, the calculated nonparallel growth rates are in good quantitative agreement with the experimental results at Mach number 4.5.
Unsteady Aerodynamic Interaction between Two Bodies at Hypersonic Speed
NASA Astrophysics Data System (ADS)
Ozawa, Hiroshi; Kitamura, Keiichi; Hanai, Katsuhisa; Mori, Koichi; Nakamura, Yoshiaki
This paper presents experimental results of unsteady aerodynamic interactions including Shock/Shock Interaction (SSI) and Shock/Boundary Layer Interaction (SBLI) between two bodies at hypersonic speed. These interactions can be seen in space vehicles consisting of multi-bodies, such as a TSTO, or at a scramjet engine inlet. The present study considers the effect of a flat plate below the SSI where a boundary-layer is developed on the plate surface. More specifically, the interacted flow for a combination of a flat plate (FP) and a hemi-circular cylinder (HCC) is examined at a hypersonic speed (M∞=8.1) the distributions of surface pressure and heat transfer rate are measured. To obtain various SSI patterns, the clearance between two bodies (FP and HCC) is changed. Results show that unsteadiness at the SSI point causes a feedback loop between the two bodies; a jet flow impinges on the FP, the effect of which propagates upstream where the jet impinges on the FP, and the aerodynamic and aerothermodynamic loads reach their maxima. Finally, we found that the feedback loop can be destroyed by installing a fence on the FP to reduce unsteadiness of flow field.
Detailed investigation of flowfields within large scale hypersonic inlet models
NASA Technical Reports Server (NTRS)
Seebaugh, W. R.; Doran, R. W.; Decarlo, J. P.
1971-01-01
Analytical and experimental investigations were conducted to determine the characteristics of the internal flows in model passages representative of hypersonic inlets and also sufficiently large for meaningful data to be obtained. Three large-scale inlet models, each having a different compression ratio, were designed to provide high performance and approximately uniform static-pressure distributions at the throat stations. A wedge forebody was used to simulate the flowfield conditions at the entrance of the internal passages, thus removing the actual vehicle forebody from consideration in the design of the wind-tunnel models. Tests were conducted in a 3.5 foot hypersonic wind tunnel at a nominal test Mach number of 7.4 and freestream unit Reynolds number of 2,700,000 per foot. From flowfield survey data the inlet entrance, the entering inviscid and viscous flow conditions were determined prior to the analysis of the data obtained in the internal passages. Detailed flowfield survey data were obtained near the centerlines of the internal passages to define the boundary-layer development on the internal surfaces and the internal shock-wave configuration. Finally, flowfield data were measured across the throats of the inlet models to evaluate the internal performance of the internal passages. These data and additional results from surface instrumentation and flow visualization studies were utilized to determine the internal flowfield patterns and the inlet performance.
Recombination Catalysts for Hypersonic Fuels
NASA Technical Reports Server (NTRS)
Chinitz, W.
1998-01-01
The goal of commercially-viable access to space will require technologies that reduce propulsion system weight and complexity, while extracting maximum energy from the products of combustion. This work is directed toward developing effective nozzle recombination catalysts for the supersonic and hypersonic aeropropulsion engines used to provide such access to space. Effective nozzle recombination will significantly reduce rk=le length (hence, propulsion system weight) and reduce fuel requirements, further decreasing the vehicle's gross lift-off weight. Two such catalysts have been identified in this work, barium and antimony compounds, by developing chemical kinetic reaction mechanisms for these materials and determining the engine performance enhancement for a typical flight trajectory. Significant performance improvements are indicated, using only 2% (mole or mass) of these compounds in the combustor product gas.
Rekindled vision of hypersonic travel
NASA Technical Reports Server (NTRS)
Colladay, Raymond S.
1987-01-01
NASA has joined with the DOD to conduct the National Aerospace Plane (NASP) program, whose experimental test vehicle will be designated the X-30. NASP will study the X-30's takeoff from a runway under its own power, acceleration to high Mach number on the basis of airbreathing propulsion, emergence into LEO, reentry into the earth atmosphere, and descent to a powered horizontal landing. NASP will thereby generate technology base data for three distinct types of aircraft: upper-atmosphere hypersonic-cruise aircraft, LEO space transports, and military transatmospheric vehicles. The current concept-validation phase of NASP focuses on airbreathing propulsion, lightweight/high-strength heat-resistant materials, and computational fluid dynamics.
Hypersonic scramjet engine fuel injector
Lee, C.P.; Venkataramani, K.S.; Lahti, D.J.; Lee, V.H.
1990-02-27
This patent describes a hypersonic scramjet engine fuel injector. It comprises: a housing having a generally horizontal top wall, an inclined bottom wall, and a generally vertical end wall attached together to define in cross-section a generally right triangle, the housing also having two generally vertical side walls having a the-generally-right-triangle shape. The side walls attached to the top, bottom, and end walls to define a fuel-tight, generally right-triangular wedge. The top wall having a fuel inlet orifice. The end wall having at least one convergent-divergent fuel outlet nozzle, and at least one wall of the bottom and side walls having a plurality of spaced-apart fuel-exit holes.
NASA Technical Reports Server (NTRS)
Rausch, J. R.
1977-01-01
The effect of interaction between the reaction control system (RCS) jets and the flow over the space shuttle orbiter in the atmosphere was investigated in the NASA Langley 31-inch continuous flow hypersonic tunnel at a nominal Mach number of 10.3 and in the AEDC continuous flow hypersonic tunnel B at a nominal Mach number of 6, using 0.01 and .0125 scale force models with aft RCS nozzles mounted both on the model and on the sting of the force model balance. The data show that RCS nozzle exit momentum ratio is the primary correlating parameter for effects where the plume impinges on an adjacent surface and mass flow ratio is the parameter when the plume interaction is primarily with the external stream. An analytic model of aft mounted RCS units was developed in which the total reaction control moments are the sum of thrust, impingement, interaction, and cross-coupling terms.
A quiet tunnel investigation of hypersonic boundary-layer stability over a cooled, flared cone
NASA Technical Reports Server (NTRS)
Blanchard, Alan E.; Selby, Gregory V.; Wilkinson, Stephen P.
1996-01-01
A flared-cone model under adiabatic and cooled-wall conditions was placed in a calibrated, low-disturbance Mach 6 flow and the stability of the boundary layer was investigated using a prototype constant-voltage anemometer. The results were compared with linear-stability theory predictions and good agreement was found in the prediction of second-mode frequencies and growth. In addition, the same 'N = 10' criterion used to predict boundary-layer transition in subsonic, transonic, and supersonic flows under low freestream noise conditions was found to be applicable for the hypersonic flow regime as well. Under cooled-wall conditions, a unique set of spectral data was acquired that documents the linear, nonlinear, and breakdown regions associated with the transition of hypersonic flow under low-noise conditions.
Hypersonic Post-Shock Cavity Ring-Down Spectroscopy
NASA Astrophysics Data System (ADS)
Suas-David, Nicolas; Kassi, Samir; Benidar, Abdessamad; Georges, Robert
2015-06-01
A highly sensitive experimental set-up (αmin = 10-10 cm-1) has been developed to produce high-temperature infrared spectra of methane in the Tetradecad polyad region (1.67 μm) using cw-CRDS. A continuous flow of methane admixed to argon is initially heated at 1000 - 1500 K and then accelerated to hypersonic speeds in a vacuum chamber before being abruptly stopped by the impact on a planar screen set perpendicular to the flow axis, forming a stationary shock wave detached from the screen (bow shock). The CRD optical beam probes the very hot subsonic zone behind the shock where the gas temperature is close to the stagnation one. Computational Fluid Dynamics calculations have been performed to characterize the post-shock structure of the flow. Spectra reveal a series of new hot bands of fundamental interest for the modeling of highly excited levels of methane.
Astrophysical Jets as Hypersonic Buckshot: Laboratory Experiments and Simulations
NASA Astrophysics Data System (ADS)
Frank, A.; Ciardi, A.; Yirak, K.; Lebedev, S.
2009-08-01
Herbig-Haro (HH) jets are commonly thought of as homogeneous beams of plasma traveling at hypersonic velocities. Structure within jet beams is often attributed to periodic or ``pulsed'' variations of conditions at the jet source. In this contribution we offer an alternative to ``pulsed'' models of protostellar jets. Using direct numerical simulations and laboratory experiments we explore the possibility that jets are chains of sub-radial clumps propagating through a moving inter-clump medium. Our simulations explore an idealization of this scenario by injecting small (r < r_{jet}), dense (rho > rho_{jet}) spheres embedded in an otherwise smooth inter-clump jet flow. The spheres are initialized with velocities differing from the jet velocity by ˜ 15%. We find the consequences of shifting from homogeneous to heterogeneous flows are significant as clumps interact with each other and with the inter-clump medium in a variety of ways. We also present new experiments that, for the first time, directly address issues of magnetized astrophysical jets. Our experiments explore the propagation and stability of super-magnetosonic, radiatively cooled, and magnetically dominated bubbles with internal, narrow jets. The results are scalable to astrophysical environments via the similarity of dimensionless numbers controlling the dynamics in both settings. These experiments show the jets are subject to kink mode instabilities which quickly fragment the jet into narrow chains of hypersonic knots, providing support for the ``clumpy jet'' paradigm.
Wind-Tunnel Balance Characterization for Hypersonic Research Applications
NASA Technical Reports Server (NTRS)
Lynn, Keith C.; Commo, Sean A.; Parker, Peter A.
2012-01-01
Wind-tunnel research was recently conducted at the NASA Langley Research Center s 31-Inch Mach 10 Hypersonic Facility in support of the Mars Science Laboratory s aerodynamic program. Researchers were interested in understanding the interaction between the freestream flow and the reaction control system onboard the entry vehicle. A five-component balance, designed for hypersonic testing with pressurized flow-through capability, was used. In addition to the aerodynamic forces, the balance was exposed to both thermal gradients and varying internal cavity pressures. Historically, the effect of these environmental conditions on the response of the balance have not been fully characterized due to the limitations in the calibration facilities. Through statistical design of experiments, thermal and pressure effects were strategically and efficiently integrated into the calibration of the balance. As a result of this new approach, researchers were able to use the balance continuously throughout the wide range of temperatures and pressures and obtain real-time results. Although this work focused on a specific application, the methodology shown can be applied more generally to any force measurement system calibration.
2-D Air-Breathing Lightcraft Engine Experiments in Hypersonic Conditions
NASA Astrophysics Data System (ADS)
Salvador, Israel I.; Myrabo, Leik N.; Minucci, Marco A. S.; de Oliveira, Antonio C.; Toro, Paulo G. P.; Chanes, José B.; Rego, Israel S.
2011-11-01
Experiments were performed with a 2-D, repetitively-pulsed (RP) laser Lightcraft model in hypersonic flow conditions. The main objective was the feasibility analysis for impulse generation with repetitively-pulsed air-breathing laser Lightcraft engines at hypersonic speeds. The future application of interest for this basic research endeavor is the laser launch of pico-, nano-, and micro-satellites (i.e., 0.1-100 kg payloads) into Low-Earth-Orbit, at low-cost and on-demand. The laser propulsion experiments employed a Hypersonic Shock Tunnel integrated with twin gigawatt pulsed Lumonics 620-TEA CO2 lasers (˜ 1 μs pulses), to produce the required test conditions. This hypersonic campaign was carried out at nominal Mach numbers ranging from 6 to 10. Time-dependent surface pressure distributions were recorded together with Schlieren movies of the flow field structure resulting from laser energy deposition. Results indicated laser-induced pressure increases of 0.7-0.9 bar with laser pulse energies of ˜ 170 J, on off-shroud induced breakdown condition, and Mach number of 7.
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.
2009-01-01
The quality of simulated hypersonic stagnation region heating on tetrahedral meshes is investigated by using a three-dimensional, upwind reconstruction algorithm for the inviscid flux vector. Two test problems are investigated: hypersonic flow over a three-dimensional cylinder with special attention to the uniformity of the solution in the spanwise direction and hypersonic flow over a three-dimensional sphere. The tetrahedral cells used in the simulation are derived from a structured grid where cell faces are bisected across the diagonal resulting in a consistent pattern of diagonals running in a biased direction across the otherwise symmetric domain. This grid is known to accentuate problems in both shock capturing and stagnation region heating encountered with conventional, quasi-one-dimensional inviscid flux reconstruction algorithms. Therefore the test problem provides a sensitive test for algorithmic effects on heating. This investigation is believed to be unique in its focus on three-dimensional, rotated upwind schemes for the simulation of hypersonic heating on tetrahedral grids. This study attempts to fill the void left by the inability of conventional (quasi-one-dimensional) approaches to accurately simulate heating in a tetrahedral grid system. Results show significant improvement in spanwise uniformity of heating with some penalty of ringing at the captured shock. Issues with accuracy near the peak shear location are identified and require further study.
Entropy considerations applied to shock unsteadiness in hypersonic inlets
NASA Astrophysics Data System (ADS)
Bussey, Gillian Mary Harding
The stability of curved or rectangular shocks in hypersonic inlets in response to flow perturbations can be determined analytically from the principle of minimum entropy. Unsteady shock wave motion can have a significant effect on the flow in a hypersonic inlet or combustor. According to the principle of minimum entropy, a stable thermodynamic state is one with the lowest entropy gain. A model based on piston theory and its limits has been developed for applying the principle of minimum entropy to quasi-steady flow. Relations are derived for analyzing the time-averaged entropy gain flux across a shock for quasi-steady perturbations in atmospheric conditions and angle as a perturbation in entropy gain flux from the steady state. Initial results from sweeping a wedge at Mach 10 through several degrees in AEDC's Tunnel 9 indicates the bow shock becomes unsteady near the predicted normal Mach number. Several curved shocks of varying curvature are compared to a straight shock with the same mean normal Mach number, pressure ratio, or temperature ratio. The present work provides analysis and guidelines for designing an inlet robust to off- design flight or perturbations in flow conditions an inlet is likely to face. It also suggests that inlets with curved shocks are less robust to off-design flight than those with straight shocks such as rectangular inlets. Relations for evaluating entropy perturbations for highly unsteady flow across a shock and limits on their use were also developed. The normal Mach number at which a shock could be stable to high frequency upstream perturbations increases as the speed of the shock motion increases and slightly decreases as the perturbation size increases. The present work advances the principle of minimum entropy theory by providing additional validity for using the theory for time-varying flows and applying it to shocks, specifically those in inlets. While this analytic tool is applied in the present work for evaluating the stability
Space Shuttle and Hypersonic Entry
NASA Technical Reports Server (NTRS)
Campbell, Charles H.; Gerstenmaier, William H.
2014-01-01
Fifty years of human spaceflight have been characterized by the aerospace operations of the Soyuz, of the Space Shuttle and, more recently, of the Shenzhou. The lessons learned of this past half decade are important and very significant. Particularly interesting is the scenario that is downstream from the retiring of the Space Shuttle. A number of initiatives are, in fact, emerging from in the aftermath of the decision to terminate the Shuttle program. What is more and more evident is that a new era is approaching: the era of the commercial usage and of the commercial exploitation of space. It is probably fair to say, that this is the likely one of the new frontiers of expansion of the world economy. To make a comparison, in the last 30 years our economies have been characterized by the digital technologies, with examples ranging from computers, to cellular phones, to the satellites themselves. Similarly, the next 30 years are likely to be characterized by an exponential increase of usage of extra atmospheric resources, as a result of more economic and efficient way to access space, with aerospace transportation becoming accessible to commercial investments. We are witnessing the first steps of the transportation of future generation that will drastically decrease travel time on our Planet, and significantly enlarge travel envelope including at least the low Earth orbits. The Steve Jobs or the Bill Gates of the past few decades are being replaced by the aggressive and enthusiastic energy of new entrepreneurs. It is also interesting to note that we are now focusing on the aerospace band, that lies on top of the aeronautical shell, and below the low Earth orbits. It would be a mistake to consider this as a known envelope based on the evidences of the flights of Soyuz, Shuttle and Shenzhou. Actually, our comprehension of the possible hypersonic flight regimes is bounded within really limited envelopes. The achievement of a full understanding of the hypersonic flight
Hypersonic Propulsion at Pratt and Whitney: Overview
NASA Technical Reports Server (NTRS)
Kazmar, Richard R.
2002-01-01
Pratt & Whitney (P&W) is developing the technology for hypersonic components and engines. A supersonic combustion ramjet (scramjet) database was developed during the National Aero Space Plane (NASP) program using hydrogen fueled propulsion systems for space access vehicles and serves as a point of departure for the current emphasis on hydrocarbon scramjets. The Air Force Hypersonic Technology (HyTech) Office has put programs in place to develop the technologies necessary to demonstrate the operability, performance and structural durability of a liquid hydrocarbon fueled scramjet system that operates from Mach 4 to 8. Fuel-cooled superalloys and lightweight structures are being developed to improve thermal protection and durability and to reduce propulsion system weight. The application of scramjet engine technology as part of combined cycle propulsion systems is also being pursued under NASA and U.S. Air Force sponsorship. The combination of scramjet power and solid rocket booster acceleration is applicable to hypersonic cruise missiles. Scramjets that use gas turbines for low speed acceleration and scramjets using rocket power for low speed acceleration are being studied for application to reusable launch systems and hypersonic cruise vehicles. P&W's recent activities and future plans for hypersonic propulsion will be described.
NASA Technical Reports Server (NTRS)
Hawthorne, P. J.
1975-01-01
Data are documented which were obtained during wind tunnel tests. The test was conducted beginning 4 March and ending 6 March 1974 for a total of 24 occupancy hours. all test runs were conducted at a Mach number of 10.3 and at Reynolds numbers of 0.65, 1.0 and 1.33 million per foot. Only the complete 140A/B was tested with various elevon, speedbrake, and bodyflap settings at angles of attack from 12 to 37 degrees at 0 and -5 degrees of beta, and from 0 to -9 degrees of beta at 20 and 30 degrees angle of attack. The purpose was to obtain hypersonic longitudinal and lateral-directional stability and control characteristics of the updated space shuttle vehicle configuration.
Weakly Ionized Plasmas in Hypersonics: Fundamental Kinetics and Flight Applications
Macheret, Sergey
2005-05-16
The paper reviews some of the recent studies of applications of weakly ionized plasmas to supersonic/hypersonic flight. Plasmas can be used simply as means of delivering energy (heating) to the flow, and also for electromagnetic flow control and magnetohydrodynamic (MHD) power generation. Plasma and MHD control can be especially effective in transient off-design flight regimes. In cold air flow, nonequilibrium plasmas must be created, and the ionization power budget determines design, performance envelope, and the very practicality of plasma/MHD devices. The minimum power budget is provided by electron beams and repetitive high-voltage nanosecond pulses, and the paper describes theoretical and computational modeling of plasmas created by the beams and repetitive pulses. The models include coupled equations for non-local and unsteady electron energy distribution function (modeled in forward-back approximation), plasma kinetics, and electric field. Recent experimental studies at Princeton University have successfully demonstrated stable diffuse plasmas sustained by repetitive nanosecond pulses in supersonic air flow, and for the first time have demonstrated the existence of MHD effects in such plasmas. Cold-air hypersonic MHD devices are shown to permit optimization of scramjet inlets at Mach numbers higher than the design value, while operating in self-powered regime. Plasma energy addition upstream of the inlet throat can increase the thrust by capturing more air (Virtual Cowl), or it can reduce the flow Mach number and thus eliminate the need for an isolator duct. In the latter two cases, the power that needs to be supplied to the plasma would be generated by an MHD generator downstream of the combustor, thus forming the 'reverse energy bypass' scheme. MHD power generation on board reentry vehicles is also discussed.
Weakly Ionized Plasmas in Hypersonics: Fundamental Kinetics and Flight Applications
NASA Astrophysics Data System (ADS)
Macheret, Sergey
2005-05-01
The paper reviews some of the recent studies of applications of weakly ionized plasmas to supersonic/hypersonic flight. Plasmas can be used simply as means of delivering energy (heating) to the flow, and also for electromagnetic flow control and magnetohydrodynamic (MHD) power generation. Plasma and MHD control can be especially effective in transient off-design flight regimes. In cold air flow, nonequilibrium plasmas must be created, and the ionization power budget determines design, performance envelope, and the very practicality of plasma/MHD devices. The minimum power budget is provided by electron beams and repetitive high-voltage nanosecond pulses, and the paper describes theoretical and computational modeling of plasmas created by the beams and repetitive pulses. The models include coupled equations for non-local and unsteady electron energy distribution function (modeled in forward-back approximation), plasma kinetics, and electric field. Recent experimental studies at Princeton University have successfully demonstrated stable diffuse plasmas sustained by repetitive nanosecond pulses in supersonic air flow, and for the first time have demonstrated the existence of MHD effects in such plasmas. Cold-air hypersonic MHD devices are shown to permit optimization of scramjet inlets at Mach numbers higher than the design value, while operating in self-powered regime. Plasma energy addition upstream of the inlet throat can increase the thrust by capturing more air (Virtual Cowl), or it can reduce the flow Mach number and thus eliminate the need for an isolator duct. In the latter two cases, the power that needs to be supplied to the plasma would be generated by an MHD generator downstream of the combustor, thus forming the "reverse energy bypass" scheme. MHD power generation on board reentry vehicles is also discussed.
CARS Temperature Measurements in a Hypersonic Propulsion Test Facility
NASA Technical Reports Server (NTRS)
Jarrett, Olin, Jr.; Smith, M. W.; Antcliff, R. R.; Northam, G. Burt; Cutler, A. D.; Capriotti, D. P.; Taylor, D. J.
1990-01-01
Nonintrusive diagnostic measurements were performed in the supersonic reacting flow of the Hypersonic Propulsion Test Cell 2 at NASA-Langley. A Coherent Anti-stokes Raman Spectroscopy (CARS) system was assembled specifically for the test cell environment. System design considerations were: (1) test cell noise and vibration; (2) contamination from flow field or atmospheric borne dust; (3) unwanted laser or electrically induced combustion (inside or outside the duct); (4) efficient signal collection; (5) signal splitting to span the wide dynamic range present throughout the flow field; (6) movement of the sampling volume in the flow; and (7) modification of the scramjet model duct to permit optical access to the reacting flow with the CARS system. The flow in the duct was a nominal Mach 2 flow with static pressure near one atmosphere. A single perpendicular injector introduced hydrogen into the flow behind a rearward facing step. CARS data was obtained in three planes downstream of the injection region. At least 20 CARS data points were collected at each of the regularly spaced sampling locations in each data plane. Contour plots of scramjet combustor static temperature in a reacting flow region are presented.
Generic Hypersonic Inlet Module Analysis
NASA Technical Reports Server (NTRS)
Cockrell, Chares E., Jr.; Huebner, Lawrence D.
2004-01-01
A computational study associated with an internal inlet drag analysis was performed for a generic hypersonic inlet module. The purpose of this study was to determine the feasibility of computing the internal drag force for a generic scramjet engine module using computational methods. The computational study consisted of obtaining two-dimensional (2D) and three-dimensional (3D) computational fluid dynamics (CFD) solutions using the Euler and parabolized Navier-Stokes (PNS) equations. The solution accuracy was assessed by comparisons with experimental pitot pressure data. The CFD analysis indicates that the 3D PNS solutions show the best agreement with experimental pitot pressure data. The internal inlet drag analysis consisted of obtaining drag force predictions based on experimental data and 3D CFD solutions. A comparative assessment of each of the drag prediction methods is made and the sensitivity of CFD drag values to computational procedures is documented. The analysis indicates that the CFD drag predictions are highly sensitive to the computational procedure used.
NASA's hypersonic propulsion program: History and direction
NASA Technical Reports Server (NTRS)
Wander, Steve
1992-01-01
Research into hypersonic propulsion; i.e., supersonic combustion, was seriously initiated at the Langley Research Center in the 1960's with the Hypersonic Research Engine (HRE) project. This project was designed to demonstrate supersonic combustion within the context of an engine module consisting of an inlet, combustor, and nozzle. In addition, the HRE utilized both subsonic and supersonic combustion (dual-mode) to demonstrate smooth operation over a Mach 4 to 7 speed range. The propulsion program thus concentrated on fundamental supersonic combustion studies and free jet propulsion tests for the three dimensional fixed geometry engine design to demonstrate inlet and combustor integration and installed performance potential. The developmental history of the program is presented. Additionally, the HRE program's effect on the current state of hypersonic propulsion is discussed.
Discrete Roughness Transition for Hypersonic Flight Vehicles
NASA Technical Reports Server (NTRS)
Berry, Scott A.; Horvath, Thomas J.
2007-01-01
The importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed. The paper is organized by hypersonic vehicle applications characterized in a general sense by the boundary layer: slender with hypersonic conditions at the edge of the boundary layer, moderately blunt with supersonic, and blunt with subsonic. This paper is intended to be a review of recent discrete roughness transition work completed at NASA Langley Research Center in support of agency flight test programs. First, a review is provided of discrete roughness wind tunnel data and the resulting correlations that were developed. Then, results obtained from flight vehicles, in particular the recently flown Hyper-X and Shuttle missions, are discussed and compared to the ground-based correlations.
Issues Associated with a Hypersonic Maglev Sled
NASA Technical Reports Server (NTRS)
Haney, Joseph W.; Lenzo, J.
1996-01-01
Magnetic levitation has been explored for application from motors to transportation. All of these applications have been at velocities where the physics of the air or operating fluids are fairly well known. Application of Maglev to hypersonic velocities (Mach greater than 5) presents many opportunities, but also issues that require understanding and resolution. Use of Maglev to upgrade the High Speed Test Track at Holloman Air Force Base in Alamogordo New Mexico is an actual hypersonic application that provides the opportunity to improve test capabilities. However, there are several design issues that require investigation. This paper presents an overview of the application of Maglev to the test track and the issues associated with developing a hypersonic Maglev sled. The focus of this paper is to address the issues with the Maglev sled design, rather than the issues with the development of superconducting magnets of the sled system.
Photon collider Higgs factories
NASA Astrophysics Data System (ADS)
Telnov, V. I.
2014-09-01
The discovery of the Higgs boson (and still nothing else) have triggered appearance of many proposals of Higgs factories for precision measurement of the Higgs properties. Among them there are several projects of photon colliders (PC) without e+e- in addition to PLC based on e+e- linear colliders ILC and CLIC. In this paper, following a brief discussion of Higgs factories physics program I give an overview of photon colliders based on linear colliders ILC and CLIC, and of the recently proposed photon-collider Higgs factories with no e+e- collision option based on recirculation linacs in ring tunnels.
Boundary Layer Control for Hypersonic Airbreathing Vehicles
NASA Technical Reports Server (NTRS)
Berry, Scott A.; Nowak, Robert J.; Horvath, Thomas J.
2004-01-01
Active and passive methods for tripping hypersonic boundary layers have been examined in NASA Langley Research Center wind tunnels using a Hyper-X model. This investigation assessed several concepts for forcing transition, including passive discrete roughness elements and active mass addition (or blowing), in the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air Tunnels. Heat transfer distributions obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the Hyper-X nominal Mach 7 flight test-point of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For passive roughness, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The passive roughness study resulted in a swept ramp configuration, scaled to be roughly 0.6 of the calculated boundary layer thickness, being selected for the Mach 7 flight vehicle. For the active blowing study, the manifold pressure was systematically varied (while monitoring the mass flow) for each configuration to determine the jet penetration height, with schlieren, and transition movement, with the phosphor system, for comparison to the passive results. All the blowing concepts tested, which included various rows of sonic orifices (holes), two- and three-dimensional slots, and random porosity, provided transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model surface static pressure, which is adequate to ensure sonic jets. The present results indicate that the jet penetration height for blowing was roughly half the height required with passive roughness elements for an equivalent amount of transition movement.
Hypersonic Flight Vehicle X-43C
NASA Technical Reports Server (NTRS)
2002-01-01
An artist's rendering of air-breathing, hypersonic X-43C, part of NASA's Hyper-X series of flight demonstrator. Now in development, the X-43C is expected to accelerate to a maximum potential speed of about 5,000 mph, and could undergo flight testing as early as the year 2008. Revolutionizing the way we gain access to space is NASA's primary goal for the Hypersonic Investment Area, managed for NASA by the Advanced Space Transportation Program at the Marshall Space Flight Center in Huntsville, Alabama. The Hypersonic Investment area, which includes leading-edge partners in industry and academia, will support future generation reusable vehicles and improved access to space. These technology demonstrators, intended for flight testing by decade's end, are expected to yield a new generation of vehicles that routinely fly about 100,000 feet above Earth's surface and reach sustained speeds in excess of March 5 (3,750 mph), the point at which 'supersonic' flight becomes 'hypersonic' flight. The flight demonstrators, the Hyper-X series, will be powered by air-breathing rocket or turbine-based engines, and ram/scramjets. Air-breathing engines, known as combined-cycle systems, achieve their efficiency gains over rocket systems by getting their oxygen for combustion from the atmosphere, as opposed to a rocket that must carry its oxygen. Once a hypersonic vehicle has accelerated to more than twice the speed of sound, the turbine or rockets are turned off, and the engine relies solely on oxygen in the atmosphere to burn fuel. When the vehicle has accelerated to more than 10 to 15 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's series of hypersonic flight demonstrators include three air-breathing vehicles: the X-43A, X-43B and X-43C.
Hypersonic Flight Vehicle X-43B
NASA Technical Reports Server (NTRS)
2002-01-01
An artist's rendering of the air-breathing, hypersonic X-43B, the third and largest of NASA's Hyper-X series flight demonstrators, which could fly later this decade. Revolutionizing the way we gain access to space is NASA's primary goal for the Hypersonic Investment Area, managed for NASA by the Advanced Space Transportation Program at the Marshall Space Flight Center in Huntsville, Alabama. The Hypersonic Investment area, which includes leading-edge partners in industry and academia, will support future generation reusable vehicles and improved access to space. These technology demonstrators, intended for flight testing by decade's end, are expected to yield a new generation of vehicles that routinely fly about 100,000 feet above Earth's surface and reach sustained speeds in excess of Mach 5 (3,750 mph), the point at which 'supersonic' flight becomes 'hypersonic' flight. The flight demonstrators, the Hyper-X series, will be powered by air-breathing rocket or turbine-based engines, and ram/scramjets. Air-breathing engines, known as combined-cycle systems, achieve their efficiency gains over rocket systems by getting their oxygen for combustion from the atmosphere, as opposed to a rocket that must carry its oxygen. Once a hypersonic vehicle has accelerated to more than twice the speed of sound, the turbine or rockets are turned off, and the engine relies solely on oxygen in the atmosphere to burn fuel. When the vehicle has accelerated to more than 10 to 15 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's series of hypersonic flight demonstrators includes three air-breathing vehicles: the X-43A, X-43B and X-43C.
Optimal trajectories for hypersonic launch vehicles
NASA Technical Reports Server (NTRS)
Ardema, Mark D.; Bowles, Jeffrey V.; Whittaker, Thomas
1994-01-01
In this paper, we derive a near-optimal guidance law for the ascent trajectory from earth surface to earth orbit of a hypersonic, dual-mode propulsion, lifting vehicle. Of interest are both the optical flight path and the optimal operation of the propulsion system. The guidance law is developed from the energy-state approximation of the equations of motion. Because liquid hydrogen fueled hypersonic aircraft are volume sensitive, as well as weight sensitive, the cost functional is a weighted sum of fuel mass and volume; the weighting factor is chosen to minimize gross take-off weight for a given payload mass and volume in orbit.
Flight testing of airbreathing hypersonic vehicles
NASA Technical Reports Server (NTRS)
Hicks, John W.
1993-01-01
Using the scramjet engine as the prime example of a hypersonic airbreathing concept, this paper reviews the history of and addresses the need for hypersonic flight tests. It also describes how such tests can contribute to the development of airbreathing technology. Aspects of captive-carry and free-flight concepts are compared. An incremental flight envelope expansion technique for manned flight vehicles is also described. Such critical issues as required instrumentation technology and proper scaling of experimental devices are addressed. Lastly, examples of international flight test approaches, existing programs, or concepts currently under study, development, or both, are given.