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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS2018

Hypersonic Chemically Reacting Boundary-Layer Stability using LASTRAC

Hypersonic boundary layer transition is critical to the design of all hypersonic vehicles due to its effect on the heat transfer into the vehicle surface and potential drag enhancement or reduction during reentry. Boundary layer transition and boundary layer stability analysis under hypersonic conditions has been studied for decades, yet there is ample room for improved accuracy and further investigations into the relevant phenomena. In this work, we present a recent implementation of chemical equilibrium, finite-rate chemistry, and thermochemical nonequilibrium capabilities into LASTRAC, an existing well-established boundary-layer stability analysis code. Verification against existing numerical results in the literature are presented. LASTRAC was previously able to address calorically perfect flows. By using solutions of the Parabolized Stability Equations (PSE) with chemical and thermal nonequilibrium, we are able to investigate the effects of chemical and thermal nonequilibrium on a variety of phenomena including stationary crossflow instability on a swept wing and 2nd mode instabilities over a wedge.

Kline, H. L.↗
NASA NTRS1975

The NATA code; theory and analysis. Volume 2: User's manual

The NATA code is a computer program for calculating quasi-one-dimensional gas flow in axisymmetric nozzles and rectangular channels, primarily to describe conditions in electric archeated wind tunnels. The program provides solutions based on frozen chemistry, chemical equilibrium, and nonequilibrium flow with finite reaction rates. The shear and heat flux on the nozzle wall are calculated and boundary layer displacement effects on the inviscid flow are taken into account. The program contains compiled-in thermochemical, chemical kinetic and transport cross section data for high-temperature air, CO2-N2-Ar mixtures, helium, and argon. It calculates stagnation conditions on axisymmetric or two-dimensional models and conditions on the flat surface of a blunt wedge. Included in the report are: definitions of the inputs and outputs; precoded data on gas models, reactions, thermodynamic and transport properties of species, and nozzle geometries; explanations of diagnostic outputs and code abort conditions; test problems; and a user's manual for an auxiliary program (NOZFIT) used to set up analytical curvefits to nozzle profiles.

Bade, W. L.↗
NASA NTRS1998

Real-Gas Aerothermodynamics Test Facilities

This chapter provides an overview of the current ground-based aerothermodynamic testing capabilities in Western Europe and the United States. The focus is on facilities capable of producing real-gas effects (dissociation, ionization, and thermochemical nonequilibrium) pertinent to the study of atmospheric flight in the Mach number range of 5 < M < 50. Perceived mission needs of interest to the Americans and Western Europeans are described where such real-gas flows are important. The role of Computational Fluid Dynamics (CFD) in modern ground testing is discussed, and the capabilities of selected American and European real-gas facilities are described. An update on the current instrumentation in aerothermodynamic testing is also outlined. Comments are made regarding the use of new facilities which have been brought on line during the past 3-5 years. Finally, future needs for aerothermodynamic testing, including instrumentation, are discussed and recommendations for implementation are reported.

James 0. Arnold↗
NASA NTRS1997

Investigations of Arcjet Flow Thermochemistry

An experimental and numerical investigation of the thermochemical state of arcjet flows is currently being conducted at NASA Ames Research Center. The experimental approach relies on the use of laser- and emission-spectroscopic diagnostic techniques in three regions of the flow. A fiber optic sensor is used to record spectrally resolved emission signals from the electrode package region, where the flow is most likely to be in thermochemical equilibrium. A second emission diagnostic measurement is made in the shock layer formed over a blunt-body test article placed in the stream, and a CCD camera is used to simultaneously record spectral emission from several measurement locations along the stagnation streamline. Downstream of the nozzle exit, but upstream of the test article, Laser-Induced Fluorescence (LIF) of atomic nitrogen is used to assess the nonequilibrium distribution of flow enthalpy in the free stream. Results from the measurements are compared with predictions from a two-temperature, axisymmetric flow model that solves the nozzle and shock-layer flows.

Fletcher, Douglas G.↗
NASA NTRS1993

Heating analysis for a Lunar Transfer Vehicle at near-equilibrium flow conditions

A heating analysis for a 15.2 m diameter Lunar Transfer Vehicle (LTV) at 0 and 10.6 deg angle of attack for a nominal trajectory through the earth's atmosphere is described. The analysis utilizes the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) with thin-layer, Navier-Stokes, thermochemical nonequilibrium options. Radiative heating levels are calculated using the Langley Optimized RAdiative Nonequilibrium (LORAN) and the Non-EQuilibrium AIr Radiation (NEQAIR) codes. At peak heating, the shock layer is substantially in equilibrium. Comprehensive spatial and spectral grid convergence studies have been implemented to quantify grid effects on the convective and radiative heating levels. Axisymmetric tests including the coupled effects of radiative energy transfer show negligible change to the convective heating and a 20 percent reduction in the radiative heating.

Gnoffo, Peter A.↗
NASA NTRS1991

Chemical-kinetic problems of future NASA missions

Thermochemical nonequilibrium in the shock layer surrounding vehicles entering the atmospheres of earth and Mars at superescape velocities is studied, deriving reaction rate coefficients that reproduce experimental data obtained in shock tubes. Thermodynamic properties and emitted radiation intensities are obtained for shock tube flow and flow in a shock layer over a blunt body. The results indicate that the viscous layer of the ablation product over an ablating heat shield is likely to be in chemical nonequilbrium. For earth entry flight, the thickness of the nonequilbrium region is between and 2 cm at the expected peak radiation point in the aerobraking trajectory, For Martian entry flight it is between 8 and 23 cm. For the earth entry case, nonequilibrium phenomena reduce radiative heating rate, while the opposite occurs for the Martian case. The radiative heat transfer rates are significant for the Mars entry conditions at entry velocities equal to or greater than 7 km/s.

Park, Chul↗
NASA NTRS1998

Arc Jet Flow Properties Determined from Laser-Induced Fluorescence of Atomic Nitrogen

An laser-spectroscopic investigation of the thermocheMical state of arcjet flows is currently being conducted in the Aerodynamic Heating Facility (AHF) Circlet at NASA Ames Research Center. Downstream of the nozzle exit, but upstream of the test article, Laser-Induced Fluorescence (LIF) of atomic nitrogen is used to assess the nonequilibriuM distribution of flow enthalpy in the free stream. The two-photon LIF technique provides simultaneous measurements of free stream velocity, translational temperature, and nitrogen number density on the flow centerline. Along with information from facility instrumentation, these measurements allow a determination of the free stream total enthalpy, and its apportionment in to thermal, kinetic, and chemical mode contributions. Experimental results are presented and discussed for two different niti-ogen/argon test gas flow runs during which the current is varied while the pressure remains constant .

Fletcher, Douglas↗
DOE OSTI2022

High-Speed Diagnostic and Simulation Capabilities for Reacting Hypersonic Reentry Flows (LDRD Final Report)

High-enthalpy hypersonic flight represents an application space of significant concern within the current national-security landscape. The hypersonic environment is characterized by high-speed compressible fluid mechanics and complex reacting flow physics, which may present both thermal and chemical nonequilibrium effects. We report on the results of a three-year LDRD effort, funded by the Engineering Sciences Research Foundation (ESRF) investment area, which has been focused on the development and deployment of new high-speed thermochemical diagnostics capabilities for measurements in the high-enthalpy hypersonic environment posed by Sandia's free-piston shock tunnel. The project has additionally sponsored model development efforts, which have added thermal nonequilibrium modeling capabilities to Sandia codes for subsequent design of many of our shock-tunnel experiments. We have cultivated high-speed, chemically specific, laser-diagnostic approaches that are uniquely co-located with Sandia's high-enthalpy hypersonic test facilities. These tools include picosecond and nanosecond coherent anti-Stokes Raman scattering at 100-kHz rates for time-resolved thermometry, including thermal nonequilibrium conditions, and 100-kHz planar laser-induced fluorescence of nitric oxide for chemically specific imaging and velocimetry. Key results from this LDRD project have been documented in a number of journal submissions and conference proceedings, which are cited here. The body of this report is, therefore, concise and summarizes the key results of the project. The reader is directed toward these reference materials and appendices for more detailed discussions of the project results and findings.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
NASA NTRS1987

Aeroassisted-vehicle design studies for a manned Mars mission

An aerobrake design that has matured over several years of development accounting for all of the important flow phenomenology which are characteristic of aerobraking vehicles is proposed as the mission baseline. Flight regimes and aerothermal environments for both Mars and Earth entry are calculated using advanced methods to account for real-gas, thermochemical, relaxation effects. The results are correlated with thermal-protection and structural requirements and mission performance capability. The importance of nonequilibrium radiative heating for Earth aerocapture is demonstrated. It is suggested that two aerobrakes of different sizes will produce optimal performance for the three phases of the mission (i.e., one aerobrake for Mars aerocapture and descent of the surface lander and another for Earth return).

Menees, Gene P.↗
NASA NTRS1987

Aeroassisted-vehicle design studies for a manned Mars mission

An aerobrake design accounting for all of the important flow phenomenology which are characteristic of aerobraking vehicles is proposed as the mission baseline. Flight regimes and aerothermal environments for both Mars and earth entry are calculated using advanced methods to account for real-gas, thermochemical, relaxation effects. The results are correlated with thermal-protection and structural requirements and mission performance capability. The importance of nonequilibrium radiative heating for earth aerocapture is demonstrated. It is suggested that two aerobrakes of different sizes will produce optimal performance for the three phases of the mission (i.e., one aerobrake for Mars aerocapture and descent of the surface lander and another for earth return).

Menees, Gene P.↗
NASA NTRS1993

Computational flow predictions for hypersonic drag devices

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.

Tokarcik, Susan A.↗
NASA NTRS1995

Vibrational and Electronic Energy Transfer and Dissociation of Diatomic Molecules by Electron Collisions

At high altitudes and velocities equal to or greater than the geosynchronous return velocity (10 kilometers per second), the shock layer of a hypersonic flight will be in thermochemical nonequilibrium and partially ionized. The amount of ionization is determined by the velocity. For a trans atmospheric flight of 10 kilometers per second and at an altitude of 80 kilometers, a maximum of 1% ionization is expected. At a velocity of 12 - 17 kilometer per second, such as a Mars return mission, up to 30% of the atoms and molecules in the flow field will be ionized. Under those circumstances, electrons play an important role in determining the internal states of atoms and molecules in the flow field and hence the amount of radiative heat load and the distance it takes for the flow field to re-establish equilibrium. Electron collisions provide an effective means of transferring energy even when the electron number density is as low as 1%. Because the mass of an electron is 12,760 times smaller than the reduced mass of N2, its average speed, and hence its average collision frequency, is more than 100 times larger. Even in the slightly ionized regime with only 1% electrons, the frequency of electron-molecule collisions is equal to or larger than that of molecule-molecule collisions, an important consideration in the low density part of the atmosphere. Three electron-molecule collision processes relevant to hypersonic flows will be considered: (1) vibrational excitation/de-excitation of a diatomic molecule by electron impact, (2) electronic excitation/de-excitation, and (3) dissociative recombination in electron-diatomic ion collisions. A review of available data, both theory and experiment, will be given. Particular attention will be paid to tailoring the molecular physics to the condition of hypersonic flows. For example, the high rotational temperatures in a hypersonic flow field means that most experimental data carried out under room temperatures are not applicable. Also, the average electron temperature is expected to be between 10,000 and 20,000 K. Thus only data for low energy electrons are relevant to the model.

Huo, Winifred M.↗
NASA NTRS2005

On Nonequilibrium Radiation in Hydrogen Shock Layers

The influence of thermochemical nonequilibrium in the shock layer over a vehicle entering the atmosphere of an outer planet is examined qualitatively. The state of understanding of the heating environment for the Galileo Probe vehicle is first reviewed. Next, the possible reasons for the high recession in the frustum region and the low recession in the stagnation region are examined. The state of understanding of the nonequilibrium in the hydrogen flow is then examined. For the entry flight in Neptune, the possible influence of nonequilibrium is predicted.

Park, Chul↗
DOE OSTI2023

First principles simulation of reacting hypersonic flow over a blunt wedge

This article presents molecular-level analysis of a reactive, near-continuum, Mach 21 nitrogen flow over a blunt wedge using the direct molecular simulation (DMS) method. The flow conditions lead to internal energy excitation and dissociation in the flow field, resulting in thermal and chemical nonequilibrium in the flow. Thermal nonequilibrium in the vibrational mode is observed to extend to the molecular level, where the vibrational energy distributions at various points in the flow field are observed to be non-Boltzmann. Furthermore, this is the first reactive DMS calculation where the wall is assumed to be isothermal and full momentum accommodation of the particles is enforced, hence incorporating viscous wall effects. Since the DMS method uses a quantum mechanically generated interaction potential as its only modeling input, all thermochemical and transport properties of the flow field can directly be attributed to the ab initio potential energy surface. Using the DMS solution as a benchmark, this article assesses the performance of Navier–Stokes computational fluid dynamics solutions using lower fidelity two-temperature models. Two models are chosen as points of comparison: the well-known Park two-temperature model and the recently developed modified Marrone and Treanor model.

Mechanics↗
NASA NTRS2018

Overview of Recent EAST Testing, Modeling & Analysis

Shock Layer Radiation at NASA Ames - Background: Complex aerothermal and thermochemical phenomena of planetary entry define convective and radiative heating. A spacecraft's TPS (Thermal Protective System) mitigates heat transfer to substructure. Successful TPS design relies on verifiable characterization of these phenomena in the anticipated flight environment. - Approach: EAST (Electric Arc Shock Tube) [facility at Ames] simulates high-enthalpy, real-gas phenomena encountered by hypersonic vehicles entering planetary atmospheres by spectrally imaging the flow behind a moving shock wave. - Goal: Validate aerothermal models (DPLR (Data Parallel Line Relaxation Code) and NEQAIR (Nonequilibrium Radiative Transport and Spectra Program)), inform model improvements, reduce uncertainty and quantify design uncertainties. - Recent Relevant Projects: MSL (Mars Science Laboratory) & Mars 2020, InSight, OSIRIS-REx (Origins Spectral Interpretation Resource Identification Security Regolith Explorer), Orion EFT-1 (Exploration Flight-1) and EM-1 (Exploration Mission-1) and New Frontiers.

Brandis, Aaron↗
NASA NTRS1998

Computational Analysis of Shock Layer Emission Measurements in an Arc-Jet Facility

This paper reports computational analysis of radiation emission experiments in a high enthalpy arc-jet wind tunnel at NASA Ames Research Center. Recently, as part of ongoing arc-jet characterization work, spectroscopic radiation emission experiments have been conducted at the 20 MW NASA Ames arc-jet facility. The emission measurements were obtained from the arc-jet freestream and from a shock layer formed in front of flatfaced models. Analysis of these data is expected to provide valuable information about the thermodynamic state of the gas in the arc-jet freestream and in the shock layer as well as thermochemical equilibration processes behind the shock in arc-jet flows. Knowledge of the thermodynamic state of the gas in arc-jet test flows and especially within the shock layer is essential to interpret the heat transfer measurements such as in surface catalysis experiments. The present work is a continuation of previous work and focuses on analysis of the emission data obtained at relatively low-pressure conditions for which the arc-jet shock layer is expected to be in thermal and chemical nonequilibrium. Building blocks of the present computational analysis are: (1) simulation of nonequilibrium expanding flow in the converging-diverging conical nozzle and supersonic jet; (2) simulation of nonequilibrium shock layer formed in front of the flat-faced cylinder model; and (3) prediction of line-of-sight radiation from the computed flowfield. For computations of the nonequilibrium flow in the conical nozzle and shock layer, multi-temperature nonequilibrium codes with the axisymmetric formulation are used. For computations of line-of-sight radiation. a nonequilibrium radiation code (NEQAIR) is used to predict emission spectra from the computed flowfield. Computed line-of-sight averaged flow properties such as vibrational and rotational temperatures, species number densities within the shock layer will be compared with those deduced from the experimental spectra. Detailed comparisons of computational and experimental spectra will also be presented.

Gokcen, Tahir↗
NASA NTRSDate not supplied

Shock-Layer Radiation Insights Available Through Flowfield Property Binning

This paper introduces a method for assessing similarity between available shock-layer radiation measurements and the radiative heating to a vehicle of interest. This method will be referred to as the flowfield property binning (FPB) approach. The challenge in developing this approach is that radiometer measurements from a surface normal line of sight, observed radiation measurements of the entire shock-layer from a distance, or measured shock-tube nonequilibrium profiles are notably different geometrically than the radiative flux to a vehicle surface. Therefore, the relative contributions of various regions of the shock-layer to the measurement, and how these relative contributions compare with the radiative heating to a vehicle of interest, are unclear. By tracking the radiation contribution from each flowfield location, and then grouping this contribution based on two local flowfield properties, these geometric differences are removed. This provide a two dimensional group, or bin, distribution that represents the radiation environment in the same format for various types of radiation simulations, such as observed radiation or radiative flux to a vehicle surface. In the limits of optically-thin radiation transport and thermochemical equilibrium flow, the choice of pressure and temperature as the binning flowfield properties is shown to provide a bin distribution that uniquely defines the radiation simulation. This concept is found to be sufficient for establishing similarity between two cases even when the optically-thin radiation transport and thermochemical equilibrium assumptions are not satisfied. A simple metric of comparison between bin distributions from different cases is provided by the histogram intersection algorithm. This algorithm results in a histogram intersection factor (H) between 0 and 1, which represents the fraction of the radiation that is similar between the two cases. Therefore, if H evaluated between bin distributions for two completely different radiation problems is near 1, then they represent essentially the same radiation problem. This means that they depend on the same weighting of various radiation modeling parameters evaluated at the same pressure and temperature. If one of these cases agrees well with a radiation measurement, then the matching bin distributions provide justification for applying this experimental validation to the other case. Even for H values as low as 0.2, the measurement case could be considered relevant if there are no other useful measurements. This approach allows observed radiation measurements to directly inform radiative heating simulations, which allows NASA’s past and future investments in observed radiation measurements to directly inform the radiative heating uncertainty or margins for future vehicle design. This FPB approach is also used to determine shock-tube conditions for informing backshell radiative heating simulations, where the normal shock velocity and Lagrangian time from the shock is used for the binning variables instead of pressure and temperature.

Shock Layer Radiation↗
NASA NTRSDate not supplied

Boundary Layer Stability and Transition in a Chemically Reacting Martian Atmosphere using LASTRAC

The Thermal Protection System (TPS) is a significant portion of the mass of reentry vehicles, planetary probes, and Martian entry vehicles. Reducing this mass has benefits in terms of decreased fuel requirements and increased payload; however, due to the high risk and un-certainty, the TPS or heat shields are designed conservatively by assuming fully turbulent flow. Laminar flow results in reduced heat flux, and improved transition prediction has the potential to reduce TPS mass and uncertainty in aerothermodynamic predictions. Limited previous research exists examining the problem of transition prediction for Martian atmospheric entry, with studies available on transition on the Mars Science Laboratory TPS. Although transition was demonstrated in wind-tunnel tests, uncertainty in the transition location resulted in a TPS designed for fully turbulent flow, and therefore greater mass than required for a partially-laminar condition. In this work, we extend the boundary layer stability code LASTRAC, recently modified to include chemical and thermal nonequilibrium capabilities, to a model of the Martian atmosphere. LASTRAC provides Parabolized Stability Equations (PSE) as well as Linear Stability Theory (LST) to predict the stability of a boundary layer and transition with semi-empirical eN methods. Results included in this work compare disturbance growth characteristics between air and Martian atmosphere at similar nondimensional freestream conditions on a simple flat plate geometry. Both chemical nonequilibrium and thermochemical nonequilibrium, as well as both PSE and LST, are used.

Heather L Kline↗