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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 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

Studies of Cosmic Ray Modulation and Energetic Particle Propagation in Time-Dependent 3-Dimensional Heliospheric Magnetic Fields

The primary goal of this project was to perform theoretical calculations of propagation of cosmic rays and energetic particles in 3-dimensional heliospheric magnetic fields. We used Markov stochastic process simulation to achieve to this goal. We developed computation software that can be used to study particle propagation in, as two examples of heliospheric magnetic fields that have to be treated in 3 dimensions, a heliospheric magnetic field suggested by Fisk (1996) and a global heliosphere including the region beyond the termination shock. The results from our model calculations were compared with particle measurements from Ulysses, Earth-based spacecraft such as IMP-8, WIND and ACE, Voyagers and Pioneers in outer heliosphere for tests of the magnetic field models. We particularly looked for features of particle variations that can allow us to significantly distinguish the Fisk magnetic field from the conventional Parker spiral field. The computer code will eventually lead to a new generation of integrated software for solving complicated problems of particle acceleration, propagation and modulation in realistic 3-dimensional heliosphere of realistic magnetic fields and the solar wind with a single computation approach.

Zhang, Ming↗
DOE OSTI2024

Explosively driven Richtmyer–Meshkov instability jet suppression and enhancement via coupling machine learning and additive manufacturing

The ability to control the behavior of fluid instabilities at material interfaces, such as the shock-driven Richtmyer–Meshkov instability, is a grand technological challenge with a broad number of applications ranging from inertial confinement fusion experiments to explosively driven shaped charges. In this work, we use a linear-geometry shaped charge as a means of studying methods for controlling material jetting that results from the Richtmyer–Meshkov instability. A shaped charge produces a high-velocity jet by focusing the energy from the detonation of high explosives. The interaction of the resulting detonation wave with a hollowed cavity lined with a thin metal layer produces the unstable jetting effect. By modifying the characteristics of the detonation wave prior to striking the lined cavity, the kinetic energy of the jet can be enhanced or reduced. Modifying the geometry of the liner material can also be used to alter jetting properties. We apply optimization methods to investigate several design parameterizations for both enhancing or suppressing the shaped-charge jet. This is accomplished using 2D and 3D hydrodynamic simulations to investigate the design space that we consider. We also apply new additive manufacturing methods for producing the shaped-charge assemblies, which allow for the experimental testing of complicated design geometries obtained through computational optimization. We present a direct comparison of our optimized designs with experimental results carried out at the High Explosives Application Facility at Lawrence Livermore National Laboratory.

36 MATERIALS SCIENCE↗
NASA NTRS1993

Signal processing of jet noise from flyover test data

Narrow-band spectra characterizing jet noise are constructed from flyover acoustic measurements. Radar and c-band tracking systems provided the aircraft position histories which enabled directivity and smear angles from the aircraft to each microphone to be computed. These angles are based on source emission time and thus give some idea about the directivity of the radiated sound field due to jet noise. Simulated spectra are included in the paper to demonstrate spectral broadening due to smear angle. The acoustic data described in the study has application to community noise analysis, noise source characterization and validation of prediction models. Both broadband-shock noise and turbulent mixing noise are observed in the spectra. A detailed description of the signal processing procedures is provided.

Kelly, Jeffrey J.↗
NASA NTRS1993

Simulation of the prompt energization and transport of radiation belt particles during the March 24, 1991 SSC

We model the rapid (about 1 min) formation of a new electron radiation belt at L about or = 2.5 that resulted from the Storm Sudden Commencement (SSC) of March 24, 1991 as observed by the Combined Release and Radiation Effects Satellite (CRRES) satellite. Guided by the observed electric and magnetic fields, we represent the time-dependent magnetospheric electric field during the SSC by an asymmetric bipolar pulse that is associated with the compression and relaxation of the Earth's magnetic field. We follow the electrons using a relativistic guiding center code. The test-particle simulations show that electrons with energies of a few MeV at L greater than 6 were energized up to 40 MeV and transported to L about or = 2.5 during a fraction of their drift period. The energization process conserves the first adiabatic invariant and is enhanced due to resonance of the electron drift motion with the time-varying electric field. Our simulation results, with an initial W(exp -8) energy flux spectra, reproduce the observed electron drift echoes and show that the interplanetary shock impacted the magnetosphere between 1500 and 1800 MLT.

Li, Xinlin↗
NASA NTRSDate not supplied

Dynamic Analysis of Reynolds Number Effects on Trailing Edge Transonic Vortex Shedding and Its Impact on Turbine Blade Aerodynamic Performance

Time-resolved, high-speed self-aligned focusing schlieren images were acquired in the NASA Glenn Research Center Transonic Turbine Blade Cascade facility to help understand the aerodynamic behavior of high-pressure, thick trailing edge turbine blades. The trailing edge thickness of 9% of axial chord tested represents simulated ceramic matrix composite fabrication constraints, which was verified previously to possess a high-loss flow regime at high inlet turbulence conditions over a narrow range of Reynolds numbers and at a fixed design exit Mach number of 0.74. Our high-speed images, which were acquired at 10 distinct Reynolds numbers, show a significant increase in energy from flow oscillations due to transonic vortex shedding at Reynolds numbers corresponding to the high loss conditions. For those conditions, strong acoustic waves turn into shock waves. Spectral Proper Orthogonal Decomposition of the high-speed images shows acoustic waves from trailing edge vortex shedding at all conditions, with increased spectral energy at the high-loss conditions and slightly increasing frequency (about 6%) as a function of Reynolds number. Analysis of potential feedback timing is performed using velocity fields from a previous LES simulation, considering different feedback mechanisms. Most noteworthy is the acoustic/shock-boundary layer interaction mechanism on the suction surface at the blade geometric throat, which likely plays an important role in realistic curved blade passages.

Trailing Edge↗
DOE OSTI2025

Pressure-temperature equation of state of Al 2 ⁢O 3 up to 14 Mbar and 40 kK

Sapphire (Al 2 ⁢O 3 ), known for its remarkable incompressibility at ambient conditions, plays a pivotal role in both static and dynamic compression research. Accurately characterizing its equation of state (EoS) is essential for these applications. Here, we present a complete Hugoniot of Al 2 ⁢ O 3 as locus of experimentally assessed, high-precision, pressure, density and temperature states up to 14 Mbar and 43 kK. The Hugoniot is established with single shock experiments using magnetically launched hyper velocity flyers on the Z Accelerator at Sandia National Laboratories. We explore principal Hugoniot states at very high shock 𝑇 and 𝑝 in the solid phase, tracking the solid-liquid boundary and culminating at 2.4-fold compression, where data provides a direct constraint on the liquid phase. Corresponding shock release data probe thermodynamic states complementary to the Hugoniot and place additional constraints on tabular EoS models. Our findings indicate a significant deviation from existing tabular EoS models for Al 2 ⁢ O 3 dictating a comprehensive overhaul. We develop two advanced EoSs for Al 2 ⁢ O 3 the SESAME 97412 model, featuring an extensive phase diagram that includes three solid phases and the liquid phase, and the updated LEOS 2200m2 model. EoS development is assisted with Quantum Molecular Dynamics simulations. Our experimental data allows for stringent testing of our EoSs. Both models accurately capture the Hugoniot of Al 2 ⁢O 3 up to the highest pressures and temperatures. Rigorous experimental determination of extreme pressures and temperatures, paired with sophisticated models, advances the frontier of EoS development beyond 1 terapascal.

Kalita, Patricia [Sandia National Laboratories (SN↗
NASA NTRS1992

Test description and preliminary pitot-pressure surveys for Langley Test Technique Demonstrator at Mach 6

A propulsion/airframe integration experiment conducted in the NASA Langley 20-Inch Mach 6 Tunnel using a 16.8-in.-long version of the Langley Test Technique Demonstrator configuration with simulated scramjet propulsion is described. Schlieren and vapor screen visualization of the nozzle flow field is presented and correlated with pitot-pressure flow-field surveys. The data were obtained at nominal free-stream conditions of Re = 2.8 x 10 exp 6 and a nominal engine total pressure of 100 psia. It is concluded that pitot-pressure surveys coupled to schlieren and vapor-screen photographs, and oil flows have revealed flow features including vortices, free shear layers, and shock waves occurring in the model flow field.

Everhart, Joel L.↗
NASA NTRS1984

SSME thrust chamber simulation using Navier-Stokes equations

The capability of the PHOENICS fluid dynamics code in predicting two-dimensional, compressible, and reacting flow in the combustion chamber and nozzle of the space shuttle main engine (SSME) was evaluated. A non-orthogonal body fitted coordinate system was used to represent the nozzle geometry. The Navier-Stokes equations were solved for the entire nozzle with a turbulence model. The wall boundary conditions were calculated based on the wall functions which account for pressure gradients. Results of the demonstration test case reveal all expected features of the transonic nozzle flows. Of particular interest are the locations of normal and barrel shocks, and regions of highest temperature gradients. Calculated performance (global) parameters such as thrust chamber flow rate, thrust, and specific impulse are also in good agreement with available data.

Przekwas, A. J.↗
NASA NTRS1989

Computation of electron density in high altitude re-entry flows

The direct simulation Monte Carlo (DSMC) method has been modified to deal with the very small degree of ionization that occurs in the early stages of the 'communications blackout' phase of re-entry. The simplified model of the process assumes that the ionization is entirely a consequence of the reaction N + O yields NO(+) + e-. The measurements from the RAM C-II test vehicle at an altitude of 81 km have been chosen as the test case. It was found that the form of the electron distribution along the shock layer was in agreement with the measurements. That normal to the surface was also in agreement with the measurements if the surface is regarded as fully catalytic. However, quantitative agreement for the fully catalytic case requires a reaction rate about an order of magnitude greater than those generally associated with this reaction.

Bird, G. A.↗
NASA NTRS2000

X-38 Experimental Aerothermodynamics

The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic environment is being defined through a synergistic combination of ground based testing and computational fluid dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp 6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were made and generally compared to within 50 deg F of flight prediction.

Horvath, Thomas J.↗
NASA NTRS2019

Aerodynamic and Aerothermal Simulations of Mars Concept Vehicles Using Overset DPLR

NASA's Descent System Studies (DSS) Program is studying various concept vehicles to enable landing of heavy payloads on the surface of Mars. While it is desirable to run high-fidelity CFD simulations to accurately assess the aerodynamic and aerothermal effects of various design changes during EDL, it is usually difficult to quickly generate high-quality grids suitable for such analyses. One approach to address this bottleneck in mesh generation is through the use oversetting grids. Although the overset approach is efficient and powerful in solving partial differential equations on complex geometries, new users often find it challenging to apply overset concepts for their simulations. For example, generating hyperbolic grids with sufficient overlap; priority in hole-cutting on multiple overlapping grids; and fixes to assemble overlapping viscous grids at the body surface. The objective of this presentation is to introduce a simple process that combines the advantages of near-body, point-matched, structured grids with oversetting background grids suitable for grid alignment. This approach allows for grids that can be sequenced, reclustering of mesh spacing at the wall, and grid alignment with the bow shock. The current methodology is tested on a Mid-L/D configuration using the overset DPLR code.

Overset Grids↗
NASA NTRS1973

Heating and flow-field studies on a straight-wing hypersonic reentry vehicle at angles of attack from 20 to 80 deg with simulation of real-gas trends

Data are presented from a series of phase-change heat transfer and flow visualization tests at Mach 7.4, 8, and 10.3 in air, Mach 19.5 in nitrogen, Mach 20.3 in helium, and Mach 6 in tetrafluoromethane (CF4) on the windward surface of a straight wing hypersonic reentry configuration for angles of attack from 20 deg to 80 deg. The results indicate that: (1) for hypersonic stream Mach numbers, the flow field over the straight-wing configuration is essentially independent of Mach number, (2) transition Reynolds number decreases with increasing angle of attack, (3) at some critical angle of attack, the wing-shock standoff distance is greatly increased and the stagnation line moves downstream from the wing leading edge, (4) value of the critical angle of attack is very sensitive to the flow shock density ratio or effective gamma, and (5) at angles of attack above the critical value for all gases, the nondimensional level of heat transfer to the wing is higher for the higher shock density ratio flows.

Hunt, J. L.↗
NASA NTRS1988

Numerical simulation of hypersonic flow over an aeroassist flight experiment vehicle

Solutions of wind-tunnel and entry-flight flow around the vehicle are obtained from the Navier-Stokes equations coupled with the chemical species continuity equations if needed. The time-iterative method employs several techniques: shock fitting, chemistry-split ADI and an algebraic grid in conformal spherical-polar space. Sensitivities of the results to numerical parameters and to frozen, equilibrium and finite rate reactions are investigated in the forebody computation. Quantitative results are obtained for the shock layer and the near wake for the entire vehicle corresponding to both ground test and flight conditions. Complex flow characteristics are analyzed on the basis of the complete flowfield over the aerobrake and simplified afterbodies. The method is stable and cost effective, and has yielded shock locations and wall pressure distributions which are in good agreement with wind-tunnel data.

Li, C. P.↗
DOE OSTI2025

Euler equations and the Sod shock tube problem

The Euler equations are a subset of the magnetohydrodynamic (MHD) equations in the infinitely collisional, unmagnetized limit. MHD modeling is central to many areas of plasma physics, ranging from low-temperature glow discharges to inertial confinement fusion. An important aspect of the Euler equations is their ability to describe states with discontinuities, such as shock waves. A standard benchmark test for numerical implementation of the Euler equations is the Sod shock tube. In this test, the system is initialized at rest with a pressure and density discontinuity, which results in a shock wave traveling into the low-pressure region and a rarefaction wave traveling into the high-pressure region. Starting with the presentation of the Euler equations, a numerical algorithm is presented here to solve these equations in one dimension. This is followed by an overview of the Sod shock tube problem that includes the precise initial setup and the analytic solution. Finally, the analytic solution is compared with results from numerical simulations.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1985

Description and calibration of the Langley Hypersonic CF4 tunnel: A facility for simulating low gamma flow as occurs for a real gas

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.

Midden, Raymond E.↗
NASA NTRS1999

Transonic Flutter Suppression Control Law Design, Analysis and Wind-Tunnel Results

The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using classical, and minimax techniques are described. A unified general formulation and solution for the minimax approach, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.

Mukhopadhyay, Vivek↗
NASA NTRS2008

Transient Three-Dimensional Startup Side Load Analysis of a Regeneratively Cooled Nozzle

The objective of this effort is to develop a computational methodology to capture the startup side load physics and to anchor the computed aerodynamic side loads with the available data from a regeneratively cooled, high-aspect-ratio nozzle, hot-fired at sea level. The computational methodology is based on an unstructured-grid, pressure-based, reacting flow computational fluid dynamics and heat transfer formulation, a transient 5 s inlet history based on an engine system simulation, and a wall temperature distribution to reflect the effect of regenerative cooling. To understand the effect of regenerative wall cooling, two transient computations were performed using the boundary conditions of adiabatic and cooled walls, respectively. The results show that three types of shock evolution are responsible for side loads: generation of combustion wave; transitions among free-shock separation, restricted-shock separation, and simultaneous free-shock and restricted shock separations; along with the pulsation of shocks across the lip, although the combustion wave is commonly eliminated with the sparklers during actual test. The test measured two side load events: a secondary and lower side load, followed by a primary and peak side load. Results from both wall boundary conditions captured the free-shock separation to restricted-shock separation transition with computed side loads matching the measured secondary side load. For the primary side load, the cooled wall transient produced restricted-shock pulsation across the nozzle lip with peak side load matching that of the test, while the adiabatic wall transient captured shock transitions and free-shock pulsation across the lip with computed peak side load 50% lower than that of the measurement. The computed dominant pulsation frequency of the cooled wall nozzle agrees with that of a separate test, while that of the adiabatic wall nozzle is more than 50% lower than that of the measurement. The computed teepee-like formation and the tangential motion of the shocks during lip pulsation also qualitatively agree with those of test observations. Moreover, a third transient computation was performed with a proportionately shortened 1 s sequence, and lower side loads were obtained with the higher ramp rate.

Wang, Ten-See↗
NASA NTRSDate not supplied

Simulation of Dust-Laden Flow in the DLR L2K Facility

An integrated CFD-particle trajectory code called US3D-DUST is used to simulate the dusty flow environments performed in the German Aerospace Center (DLR) L2K arcjet facility. The particle trajectory code uses a Lagrangian approach, where the particle trajectories through the shock layer can be computed using a set of coupled ordinary differential equations. The L2K experiments measured mass loss to flat-face shaped samples of the Norcoat-Liège thermal protection system material. The US3D-DUST simulations were able to reproduce particle velocity and stagnation point heating rate data. When applied to the L2K dusty flow conditions, the US3D-DUST predictions of mass loss due to particle impacts and surface recession rates over-predicted the experimental values. It is recommended to perform additional dusty flow tests in the L2Kto provide and improve the surface damage model for Norcoat-Liège.

Thermal protection systems↗