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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Multiphysics Thermal-Fluid Design Analysis of a Non-Nuclear Tester for Hot-Hydrogen Materials and Component Development

The objective of this effort is to perform design analyses for a non-nuclear hot-hydrogen materials tester, as a first step towards developing efficient and accurate multiphysics, thermo-fluid computational methodology to predict environments for hypothetical solid-core, nuclear thermal engine thrust chamber design and analysis. The computational methodology is based on a multidimensional, finite-volume, turbulent, chemically reacting, thermally radiating, unstructured-grid, and pressure-based formulation. The multiphysics invoked in this study include hydrogen dissociation kinetics and thermodynamics, turbulent flow, convective, and thermal radiative heat transfers. The goals of the design analyses are to maintain maximum hot-hydrogen jet impingement energy and to minimize chamber wall heating. The results of analyses on three test fixture configurations and the rationale for final selection are presented. The interrogation of physics revealed that reactions of hydrogen dissociation and recombination are highly correlated with local temperature and are necessary for accurate prediction of the hot-hydrogen jet temperature.

Wang, Ten-See↗
NASA NTRS2006

Analysis of Material Sample Heated by Impinging Hot Hydrogen Jet in a Non-Nuclear Tester

A computational conjugate heat transfer methodology was developed and anchored with data obtained from a hot-hydrogen jet heated, non-nuclear materials tester, as a first step towards developing an efficient and accurate multiphysics, thermo-fluid computational methodology to predict environments for hypothetical solid-core, nuclear thermal engine thrust chamber. The computational methodology is based on a multidimensional, finite-volume, turbulent, chemically reacting, thermally radiating, unstructured-grid, and pressure-based formulation. The multiphysics invoked in this study include hydrogen dissociation kinetics and thermodynamics, turbulent flow, convective and thermal radiative, and conjugate heat transfers. Predicted hot hydrogen jet and material surface temperatures were compared with those of measurement. Predicted solid temperatures were compared with those obtained with a standard heat transfer code. The interrogation of physics revealed that reactions of hydrogen dissociation and recombination are highly correlated with local temperature and are necessary for accurate prediction of the hot-hydrogen jet temperature.

Wang, Ten-See↗
NASA NTRS1991

Numerical investigation of an internal layer in turbulent flow over a curved hill

A numerical investigation of incompressible and compressible turbulent flows over strongly curved surfaces is presented. The turbulent flow equations are solved by a pressure based Navier-Stokes equations solver. In the method, the conservation of mass equation is replaced by a pressure correction equation applicable for both compressible and incompressible flows. The turbulence is described by a multiple time scale turbulence model supplemented with a near-wall turbulence model. The numerical results show that the internal layer is a strong turbulence field which is developed beneath the external boundary layer and is located very close to the wall. The development of the internal layer is attributed to the enormous mean flow strain rate caused by the streamline curvature. The external boundary layer flow responds rather slowly to the streamline curvature. Thus, the turbulence field of the forward corner of the curved hill is characterized by two turbulence fields interacting with each other. The turbulence intensity of the internal layer is much stronger than that of the external boundary layer, so that the development of a new boundary layer in the downstream region of the curved hill depends mostly on the internal layer. These numerical results are in good agreement with the measured data, and show that the turbulence model can resolve the turbulence field subjected to the strong streamline curvature.

Kim, S.-W.↗
NASA NTRS1960

Effect of Base Bleed and Terminal Fairings on the Performance of Exhaust-Nozzle-Afterbody Combinations at Mach Numbers of 1.93, 2.55, and 3.05

An investigation of four exhaust-nozzle-afterbody combinations has been conducted in the Langley 9- by 12-inch blowdown tunnel at Mach numbers of 1.93, 2.55, and 3.05. The models were tested on a pylon-mounted nacelle and the jet exhaust was simulated with cold air. Base bleed w a s varied from 0 to about 12 percent of the primary jet weight flow and was discharged in to the base region through either a sonic or supersonic bleed nozzle. The models were tested at zero degree angle of attack and the Reynolds number range was from 8 x 10(exp 6) to 9 x 10(exp 6) per foot. The results indicate that the base pressure and the performance of the exhaust-nozzle-afterbody combinations were little affected gy the high-velocity base bleed. The efficiency of the terminal-fairing model was only slightly less than that of the convergent-divergent nozzle-afterbody combinations; this difference indicates the loss associated with improved transonic efficiency at higher Mach numbers.

Norton, Harry T., Jr.↗
NASA NTRSDate not supplied

Space Launch System Base Aerodynamics Post-Flight Reconstruction for Artemis I

Artemis I was the first uncrewed integrated test flight of the NASA heavy-lift, human-rated, exploration-class launch vehicle, Space Launch System (SLS), and Orion spacecraft. Artemis I successfully launched from Pad39B at NASA Kennedy Space Center on November 16th, 2022. The integrated test flight was composed of launch and ascent of SLS vehicle from lift-off to RS-25 main engine cut-off (MECO), interim cryogenic propulsion stage (ICPS) in-space flight and Orion’s trajectory around the moon and landing in the Pacific Ocean which occurred on December 11th, 2022. The SLS total thrust of 8,800,000 lbf was powered by four LOX/LH2 RS-25 engines and two 5-segment solid rocket boosters. As a result, the base flow field had highly complex phenomena and regimes. SLS base aerodynamics covers vehicle base pressure and integrated axial force during powered flight from lift-off to MECO. This work discusses the SLS base flow physics observed during Artemis I and comparisons of post-flight reconstruction with pre-flight models and Space Shuttle data. This was the first time in-depth base aerodynamics flight reconstruction has been investigated for an exploration-class launch vehicle since the Saturn V Program.

aerothermodynamics↗
NASA NTRSDate not supplied

Base Aerodynamics Post-Flight Reconstruction for Artemis I

Artemis I was the first uncrewed integrated test flight of the NASA heavy-lift, human-rated, exploration-class launch vehicle, Space Launch System (SLS), and Orion spacecraft. Artemis I successfully launched from Pad39B at NASA Kennedy Space Center on November 16th, 2022. The integrated test flight was composed of launch and ascent of SLS vehicle from lift-off to RS-25 main engine cut-off (MECO), interim cryogenic propulsion stage (ICPS) in-space flight and Orion’s trajectory around the moon and landing in the Pacific Ocean which occurred on December 11th, 2022. The SLS total thrust of 8,800,000 lbf was powered by four LOX/LH2 RS-25 engines and two 5-segment solid rocket boosters. As a result, the base flow field had highly complex phenomena and regimes. SLS base aerodynamics covers vehicle base pressure and integrated axial force during powered flight from lift-off to MECO. This work discusses the SLS base flow physics observed during Artemis I and comparisons of post-flight reconstruction with pre-flight models and Space Shuttle data. This was the first time in-depth base aerodynamics flight reconstruction has been investigated for an exploration-class launch vehicle since the Saturn V Program.

aerothermodynamics↗
NASA NTRS1999

Tail Lobe Revisited: Magnetic Field Modeling Based on Plasma Data

Plasma data from the ISEE-1 and -2 spacecraft during 1977-1980 have been used to determine the distribution of data points in the magnetotail in the range of distances -20 < XGSM < --15, i.e. which of the records that were located in the current sheet, in the tail lobe, in the magnetosheath and in the boundary layers respectively. The ISEE-1 and -2 magnetic field data for the records in the tail lobe were then used to model the tail lobe magnetic field dependence on the solar wind dynamic pressure, on the Interplanetary Magnetic Field (IMF) and on the Dst index. The tail lobe magnetic field was assumed to be dependent on the square root of the dynamic pressure based on the balance between the total magnetic pressure in the tail lobes and the dynamic pressure of the solar wind. The IMF dependent terms, added to the pressure term, were sought in many different forms while the Dst dependence of the tail lobe magnetic field was assumed to be linear. The field shows a strong dependence on the square root of the dynamic pressure and the different IMF dependent terms all constitute a significant contribution to the total field. However, the dependence on the Dst index turned out to be very weak at those down-tail distances. The results of this study are intended to be used for parameterizing future versions of the data-based models of the global magnetospheric magnetic field.

Karlsson, S. B. P.↗
NASA NTRS1999

Flight-Determined Subsonic Lift and Drag Characteristics of Seven Lifting-Body and Wing-Body Reentry Vehicle Configurations With Truncated Bases

This paper examines flight-measured subsonic lift and drag characteristics of seven lifting-body and wing-body reentry vehicle configurations with truncated bases. The seven vehicles are the full-scale M2-F1, M2-F2, HL-10, X-24A, X-24B, and X-15 vehicles and the Space Shuttle prototype. Lift and drag data of the various vehicles are assembled under aerodynamic performance parameters and presented in several analytical and graphical formats. These formats unify the data and allow a greater understanding than studying the vehicles individually allows. Lift-curve slope data are studied with respect to aspect ratio and related to generic wind-tunnel model data and to theory for low-aspect-ratio planforms. The proper definition of reference area was critical for understanding and comparing the lift data. The drag components studied include minimum drag coefficient, lift-related drag, maximum lift-to-drag ratio, and, where available, base pressure coefficients. The effects of fineness ratio on forebody drag were also considered. The influence of forebody drag on afterbody (base) drag at low lift is shown to be related to Hoerner's compilation for body, airfoil, nacelle, and canopy drag. These analyses are intended to provide a useful analytical framework with which to compare and evaluate new vehicle configurations of the same generic family.

Saltzman, Edwin J.↗
NASA NTRS2014

Development, Verification and Validation of Parallel, Scalable Volume of Fluid CFD Program for Propulsion Applications

There are many instances involving liquid/gas interfaces and their dynamics in the design of liquid engine powered rockets such as the Space Launch System (SLS). Some examples of these applications are: Propellant tank draining and slosh, subcritical condition injector analysis for gas generators, preburners and thrust chambers, water deluge mitigation for launch induced environments and even solid rocket motor liquid slag dynamics. Commercially available CFD programs simulating gas/liquid interfaces using the Volume of Fluid approach are currently limited in their parallel scalability. In 2010 for instance, an internal NASA/MSFC review of three commercial tools revealed that parallel scalability was seriously compromised at 8 cpus and no additional speedup was possible after 32 cpus. Other non-interface CFD applications at the time were demonstrating useful parallel scalability up to 4,096 processors or more. Based on this review, NASA/MSFC initiated an effort to implement a Volume of Fluid implementation within the unstructured mesh, pressure-based algorithm CFD program, Loci-STREAM. After verification was achieved by comparing results to the commercial CFD program CFD-Ace+, and validation by direct comparison with data, Loci-STREAM-VoF is now the production CFD tool for propellant slosh force and slosh damping rate simulations at NASA/MSFC. On these applications, good parallel scalability has been demonstrated for problems sizes of tens of millions of cells and thousands of cpu cores. Ongoing efforts are focused on the application of Loci-STREAM-VoF to predict the transient flow patterns of water on the SLS Mobile Launch Platform in order to support the phasing of water for launch environment mitigation so that vehicle determinantal effects are not realized.

West, Jeff↗
NASA NTRS1992

Comment on 'Direct observation of the intermittency of intense vorticity filaments in turbulence'

In their discussion of the behavior of vorticity and dissipation in turbulent flow, Douady et al. (1991) proposed an electrostatic analogy for the pressure based on the divergence of the incompressible Navier-Stokes equations. It is argued here that their proposed interpretation of the pressure source for incompressible flow is unnecessarily narrow and should be extended to compressible and noise-producing flow on the basis of existing theory.

Horne, W. C.↗
NASA NTRS1994

Wall pressure fluctuations in the reattachment region of a supersonic free shear layer

The primary aim of this research program was to investigate the mechanisms which cause the unsteady wall-pressure fluctuations in shock wave turbulent shear layer interactions. The secondary aim was to find means to reduce the magnitude of the fluctuating pressure loads by controlling the unsteady shock motion. The particular flow under study is the unsteady shock wave interaction formed in the reattachment zone of a separated supersonic flow. Similar flows are encountered in many practical situations, and they are associated with high levels of fluctuating wall pressure. The free shear layer is formed by the flow over a backward facing step, using an existing model, with the base pressure on the step adjusted so that there is no pressure discontinuity at the lip. The shear layer therefore develops in a zero pressure gradient. The primary advantage of this flow configuration is that the reattachment process can be studied in the absence of a separation shock. The mean flow data, and some preliminary hot-wire measurements of the mass-flux fluctuations were made by Baca and Settles, Baca, Williams and Bogdonoff, who showed that the shear layer became self-similar at about 17 delta(sub 0) downstream of the lip, and that it grew at a rate typical of the observed Mach number difference (about 1/3rd the incompressible growth rate). The turbulence measurements were later extended by Hayakawa, Smits and Bogdonoff under NASA Headquarters support.

Smits, Alexander J.↗
NASA NTRS1976

Results of an investigation of jet plume effects on an 0.010-scale model (75-OTS) of the space shuttle integrated vehicle in the 9 x 7-foot leg of the NASA/Ames unitary wind tunnel (IA82B), volume 1

The base pressure environment was investigated for the first and second stage mated vehicle in a supersonic flow field from Mach 1.55 through 2.20 with simulated rocket engine exhaust plumes. The pressure environment was investigated for the orbiter at various vent port locations at these same freestream conditions. The Mach number environment around the base of the model with rocket plumes simulated was examined. Data were obtained at angles of attack from -4 deg through +4 deg at zero yaw, and at yaw angles from -4 deg through +4 deg at zero angle of attack, with rocket plume sizes varying from smaller than nominal to much greater than nominal. Failed orbiter engine data were also obtained. Elevon hinge moments and wing panel load data were obtained during all runs. Photographs of the tested configurations are shown.

Hawthorne, P. J.↗
NASA NTRS2002

Limb neurovascular control during altered otolithic input in humans

Head-down rotation (HDR), which activates the vestibulosympathetic reflex, increases leg muscle sympathetic nerve activity (MSNA) and produces calf vasoconstriction with no change in either cardiac output or arterial blood pressure. Based on animal studies, it was hypothesized that differential control of arm and leg MSNA explains why HDR does not alter arterial blood pressure. Fifteen healthy subjects were studied. Heart rate, arterial blood pressure, forearm and calf blood flow, and leg MSNA responses were measured during HDR in these subjects. Simultaneous recordings of arm and leg MSNA were obtained from five of the subjects. Forearm and calf blood flow, vascular conductances, and vascular resistances were similar before HDR, as were arm and leg MSNA. HDR elicited similar significant increases in leg (Delta 6 +/- 1 bursts min(-1); 59 +/- 16 % from baseline) and arm MSNA (Delta 5 +/- 1 bursts min(-1); 80 +/- 28 % from baseline). HDR significantly decreased calf (-19 +/- 2 %) and forearm vascular conductance (-12 +/- 2 %) and significantly increased calf (25 +/- 4 %) and forearm vascular resistance (15 +/- 2 %), with ~60 % greater vasoconstriction in the calf than in the forearm. Arterial blood pressure and heart rate were not altered by HDR. These results indicate that there is no differential control of MSNA in the arm and leg during altered feedback from the otolith organs in humans, but that greater vasoconstriction occurs in the calf than in the forearm. These findings indicate that vasodilatation occurs in other vascular bed(s) to account for the lack of increase in arterial blood pressure during HDR.

Non-NASA Center↗
NASA NTRS1971

Space Shuttle: Static pressure distribution on Chrysler Corporation Space Division SERV booster configuration

A dual purpose test was conducted in the propulsion wind tunnel (PWT) to evaluate the performance of an aerospike engine, in the presence of a booster, and obtain forebody and base pressure distributions on the booster in which it is installed. The test item was a 2.5 percent scaled replica of the SERV booster employing a 5 percent spike length aerospike engine installed in the base region of the model. Cold flow air was used to simulate engine jet operation. Two booster configurations were investigated, one on which reentry aerospike engine thermal protection doors were installed, and another where the doors were removed. The data presented are representative of the latter configuration for a Mach number range of 0 to 1.25 at angles of attack of 0 and 8 degrees and 0 degrees angle of sideslip.

Price, E. A.↗
NASA NTRS1993

Navier-Stokes flow field analysis of compressible flow in a high pressure safety relief valve

The objective of this study is to investigate the complex three-dimensional flowfield of an oxygen safety pressure relieve valve during an incident, with a computational fluid dynamic (CFD) analysis. Specifically, the analysis will provide a flow pattern that would lead to the expansion of the eventual erosion pattern of the hardware, so as to combine it with other findings to piece together a most likely scenario for the investigation. The CFD model is a pressure based solver. An adaptive upwind difference scheme is employed for the spatial discretization, and a predictor, multiple corrector method is used for the velocity-pressure coupling. The computational result indicated vortices formation near the opening of the valve which matched the erosion pattern of the damaged hardware.

Vu, Bruce↗
NASA NTRS2003

Asymmetric Base-Bleed Effect on Aerospike Plume-Induced Base-Heating Environment During Power-Pack Out

A computational heat transfer design methodology was developed to study tbe dual-engine linear aerospike plume-induced base-heating environment during one power-pack out, in ascent flight. It includes a three-dimensional, finite volume, viscous, chemically reacting, and pressure-based computational fluid dynamics formulation, a special base-bleed boundary condition, and a three-dimensional, finite volume, and spectral-line-based weighted-sum-of-gray-gases absorption computational radiation heat transfer formulation. A separate radiation model was used for diagnostic purposes. The computational methodology was systematically benchmarked. in this study, near-base radiative heat fluxes were computed, and they compared well with those measured during static linear aerospike engine tests. The base-heating environment of 18 trajectory points secected from three power-pack out scenarios was computed. The computed asymmetric base-heating physics were analyzed. The power-pack out condition has the most impact on convective base heating when it happens early in flight. The soume of its impact comes from the asymmetric and reduced base bleed.

Wang, Ten-See↗
NASA NTRS2004

Asymmetric Base-Bleed Effect on Aerospike Plume-Induced Base-Heating Environment

A computational heat transfer design methodology was developed to study the dual-engine linear aerospike plume-induced base-heating environment during one power-pack out, in ascent flight. It includes a three-dimensional, finite volume, viscous, chemically reacting, and pressure-based computational fluid dynamics formulation, a special base-bleed boundary condition, and a three-dimensional, finite volume, and spectral-line-based weighted-sum-of-gray-gases absorption computational radiation heat transfer formulation. A separate radiation model was used for diagnostic purposes. The computational methodology was systematically benchmarked. In this study, near-base radiative heat fluxes were computed, and they compared well with those measured during static linear aerospike engine tests. The base-heating environment of 18 trajectory points selected from three power-pack out scenarios was computed. The computed asymmetric base-heating physics were analyzed. The power-pack out condition has the most impact on convective base heating when it happens early in flight. The source of its impact comes from the asymmetric and reduced base bleed.

Wang, Ten-See↗