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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Science objectives and performance of a radiometer and window design for atmospheric entry experiments

The Radiative Heating Experiment, RHE, aboard the Aeroassist Flight Experiment, AFE, (now cancelled) was to make in-situ measurements of the stagnation region shock layer radiation during an aerobraking maneuver from geosynchronous to low earth orbit. The measurements were to provide a data base to help develop and validate aerothermodynamic computational models. Although cancelled, much work was done to develop the science requirements and to successfully meet RHE technical challenges. This paper discusses the RHE scientific objectives and expected science performance of a small sapphire window for the RHE radiometers. The spectral range required was from 170 to 900 nm. The window size was based on radiometer sensitivity requirements including capability of on-orbit solar calibration.

Craig, Roger A.↗
NASA NTRS2005

A Genesis breakup and burnup analysis in off-nominal Earth return and atmospheric entry

The Genesis project conducted a detailed breakup/burnup analysis before the Earth return to determine if any spacecraft component could survive and reach the ground intact in case of an off-nominal entry. In addition, an independent JPL team was chartered with the responsibility of analyzing several definitive breakup scenarios to verify the official project analysis. This paper presents the analysis and results of this independent team.

off-nominal entry↗
NASA NTRS2007

Lightweight Thermal Protection System for Atmospheric Entry

TUFROC (Toughened Uni-piece Fibrous Reinforced Oxidation-resistant Composite) has been developed as a new thermal protection system (TPS) material for wing leading edge and nose cap applications. The composite withstands temperatures up to 1,970 K, and consists of a toughened, high-temperature surface cap and a low-thermal-conductivity base, and is applicable to both sharp and blunt leading edge vehicles. This extends the possible application of fibrous insulation to the wing leading edge and/or nose cap on a hypersonic vehicle. The lightweight system comprises a treated carbonaceous cap composed of ROCCI (Refractory Oxidation-resistant Ceramic Carbon Insulation), which provides dimensional stability to the outer mold line, while the fibrous base material provides maximum thermal insulation for the vehicle structure.

Stewart, David↗
NASA NTRS2014

Guidance and Control Architecture Design and Demonstration for Low Ballistic Coefficient Atmospheric Entry

We propose to develop a robust guidance and control system for the ADEPT (Adaptable Deployable Entry and Placement Technology) entry vehicle. A control-centric model of ADEPT will be developed to quantify the performance of candidate guidance and control architectures for both aerocapture and precision landing missions. The evaluation will be based on recent breakthroughs in constrained controllability/reachability analysis of control systems and constrained-based energy-minimum trajectory optimization for guidance development operating in complex environments.

Entry↗
NASA NTRS2014

Visualization of Flow Separation Around an Atmospheric Entry Capsule at Low-Subsonic Mach Number Using Background-Oriented Schlieren (BOS)

This paper presents the results of visualization of separated flow around a generic entry capsule that resembles the Apollo Command Module (CM) and the Orion Multi-Purpose Crew Vehicle (MPCV). The model was tested at flow speeds up to Mach 0.4 at a single angle of attack of 28 degrees. For manned spacecraft using capsule-shaped vehicles, certain flight operations such as emergency abort maneuvers soon after launch and flight just prior to parachute deployment during the final stages of entry, the command module may fly at low Mach number. Under these flow conditions, the separated flow generated from the heat-shield surface on both windward and leeward sides of the capsule dominates the wake flow downstream of the capsule. In this paper, flow visualization of the separated flow was conducted using the background-oriented schlieren (BOS) method, which has the capability of visualizing significantly separated wake flows without the particle seeding required by other techniques. Experimental results herein show that BOS has detection capability of density changes on the order of 10(sup-5).

Mizukaki, Toshiharu↗
NASA NTRS2016

Validation of High Speed Earth Atmospheric Entry Radiative Heating from 9.5 to 15.5 km/s

This paper presents an overview of the analysis and measurements of equilibrium radiation obtained in the NASA Ames Research Center's Electric Arc Shock Tube (EAST) facility as a part of recent testing aimed at reaching shock velocities up to 15.5 km/s. The goal of these experiments was to measure the level of radiation encountered during high speed Earth entry conditions, such as would be relevant for an asteroid, inter-planetary or lunar return mission. These experiments provide the first spectrally and spatially resolved data for high speed Earth entry and cover conditions ranging from 9.5 to 15.5 km/s at 13.3 and 26.6 Pa (0.1 and 0.2 Torr). The present analysis endeavors to provide a validation of shock tube radiation measurements and simulations at high speed conditions. A comprehensive comparison between the spectrally resolved absolute equilibrium radiance measured in EAST and the predictive tools, NEQAIR and HARA, is presented. In order to provide a more accurate representation of the agreement between the experimental and simulation results, the integrated value of radiance has been compared across four spectral regions (VUV, UV/Vis, Vis/NIR and IR) as a function of velocity. Results have generally shown excellent agreement between the two codes and EAST data for the Vis through IR spectral regions, however, discrepancies have been identified in the VUV and parts of the UV spectral regions. As a result of the analysis presented in this paper, an updated parametric uncertainty for high speed radiation in air has been evaluated to be [9.0%, -6.3%]. Furthermore, due to the nature of the radiating environment at these high shock speeds, initial calculations aimed at modeling phenomena that become more significant with increasing shock speed have been performed. These phenomena include analyzing the radiating species emitting ahead of the shock and the increased significance of radiative cooling mechanisms.

Equilibrium↗
NASA NTRS2019

Exploration of Atmospheric Entries at Uranus & Neptune with HEEET as Heatshield TPS

One important observation from the recent Ice Giants Study sponsored by NASA was that the predicted and margined thicknesses of HEEET (new NASA TPS technology) were much greater than could be woven with the currently established loom capabilities. Since the cost of a loom upgrade would be substantial, the present work explores the entry trajectory space to determine what combinations of entry parameters would result in HEEET thicknesses that fit within the existing loom infrastructure. Toward this end, the entry trajectory space, parameterized by ballistic coefficient and entry flight path angle, was systematically explored for 45° sphere-cone geometries of 3 different radii 0.2 m, 0.3 m, and 0.4 m which covered the range from Galileo-derived probes considered in the Ice Giants Study, and a follow-on study on the possibility of using a single probe architecture (in terms of size and mass) for various destinations, including Venus, Saturn, Uranus, and Neptune. The entry velocities, latitudes, and azimuths at Uranus and Neptune used in the present work were taken from the Ice Giants Study. For each 3D OF trajectory generated by a NASA Ames in-house code, TRAJ, the material response and thickness were computed using another NASA Ames code, FIAT, along with a margins policy proposed by the HEEET project. In the present work, ballistic coefficients ranging from 200 kg/sqm to 350 kg/sqm were considered along with entry flight path angles ranging from -16° to -36° (primarily to allow deceleration loads to vary between 50 g and 200 g).

Prabhu, Dinesh K.↗
NASA NTRS2019

Exploration of Atmospheric Entries at Uranus & Neptune with HEEET as Heatshield TPS

One important observation from the Ice Giants Study was that the predicted and margined thicknesses of HEEET were greater than could be woven with the currently established loom capabilities. Since the cost of a loom upgrade would be substantial, the present work explored the entry trajectory space to determine what combinations of entry parameters would result in HEEET thicknesses that fit within the existing loom infrastructure. Toward this end, the entry trajectory space, parameterized by ballistic coefficient and entry flight path angle, was systematically explored for 45° sphere-cone geometries of 3 different radii – 0.2 m, 0.3 m, and 0.4 m – which covered the range from Galileo-derived probes considered in the Ice Giants Study, and a follow-on study [4] on the possibility of using a single probe architecture (in terms of size and mass) for various destinations, including Venus, Saturn, Uranus, and Neptune. The entry velocities, latitudes, and azimuths at Uranus and Neptune used in the present work were taken from the Ice Giants Study [1]. For each 3DOF trajectory generated by a NASA Ames in-house code, TRAJ [5], the material response and thickness were computed using another NASA Ames code, FIAT [6], along with a margins policy proposed by the HEEET project [7]. In the present work, ballistic coefficients ranging from 200 kg/m2 to 350 kg/m2 were considered along with entry flight path angles ranging from -16° to -36° (primarily to allow deceleration loads to vary between 50 g and 200 g).

HEEET↗
NASA NTRS2018

Radiative Heating of Large Meteoroids During Atmospheric Entry

A high-fidelity approach for simulating the aerothermodynamic environments of meteor entries was developed, which allows the commonly assumed heat transfer coefficient of 0.1 to be assessed. This model uses chemically reacting computational fluid dynamics (CFD), coupled with radiation transport and surface ablation. Coupled radiation accounts for the impact of radiation on the flowfield energy equations, while coupled ablation explicitly models the injection of ablation products within the flowfield and radiation simulations. For a meteoroid with a velocity of 20 km/s, coupled radiation is shown to reduce the stagnation point radiative heating by over 60%. The impact of coupled ablation (with coupled radiation) is shown to provide at least a 70% reduction in the radiative heating relative to cases with only coupled radiation. This large reduction is partially the result of the low ionization energies of meteoric ablation products relative to air species. The low ionization energies of ablation products, such as Mg and Ca, provide strong photoionization and atomic line absorption in regions of the spectrum that air species do not. MgO and CaO are also shown to provide significant absorption. Turbulence is shown to impact the distribution of ablation products through the shock-layer, which results in up to a 100% increase in the radiative heating downstream of the stagnation point. To create a database of heat transfer coefficients, the developed model was applied to a range of cases. This database considered velocities ranging from 14 to 20 km/s, altitudes ranging from 20 to 50 km, and nose radii ranging from 1 to 100 m. The heat transfer coefficients from these simulations are below 0.045 for the range of cases, for both laminar and turbulent, which is significantly lower than the canonical value of 0:1. When the new heat transfer model is applied to a Tunguska-like 15 Mt entry, the effect of the new model is to lower the height of burst by up to 2 km, depending on assumed entry angle. This, in turn, results in a significantly larger ground damage footprint than when the canonical heating assumption is used.

Christopher O Johnston↗
NASA NTRSDate not supplied

ELECTRICAL SHORTING OF THERMOCOUPLES IN ATMOSPHERIC ENTRY APPLICATIONS

The Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) will test Hyper-sonic Inflatable Aerodynamic Decelerator (HIAD) technology. Temperature measurements are taken across the vehicle to address a Level 1 mission requirement regarding performance of the Flexible Thermal Protection System (F-TPS). Temperature measurements are made using dozens of thermocouples (TCs) embedded within the F-TPS at different depths and various locations shown in Figure 1.

Cole D Kazemba↗
NASA NTRS2020

NASA Ames Thermophysics Ground Test Facilities Supporting Future Planetary Atmospheric Entry

A review of the current facility capabilities for testing Thermal Protection Systems and quantifying their entry environments at NASA Ames Research Center is presented based on the expected targets of interest to the Planetary Science and Astrobiology communities. While the operational capabilities of these facilities are generally considered sufficient for supporting future missions to targets of interest, expanded ground test capabilities such as larger sample sizes, flight-relevant gas mixtures, dusty environments, and flight-relevant shear/pressure combinations would reduce future entry vehicle design uncertainties and applied margins. These reduced uncertainties may translate into reduced entry vehicle masses, increased robustness, and decreased operational risks during entry phases for science missions. Expanded ground test capabilities would also offer the ability to study material failure modes in environments even more representative of flight than are currently achievable. A list of desired future test capabilities is presented along with suggestions of possible methods of achieving each. The main recommendation of this paper is the undertaking of a detailed study of the benefits and associated costs of each of these expanded capabilities to determine the best future path.

arc jet↗
NASA NTRS1998

Mars Pathfinder Atmospheric Entry Reconstruction

The primary objective of the Mars Pathfinder mission was to demonstrate an innovative, low-cost and reliable method for placing a science payload on the surface of Mars. This paper describes the results of an effort to access the spacecraft performance during Entry, Descent and Landing.

Mars↗