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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Regenerative Aerobraking

NASA's exploration goals for Mars and Beyond will require new power systems and in situ resource utilization technologies. Regenerative aerobraking may offer a revolutionary approach for in situ power generation and oxygen harvesting during these exploration missions. In theory, power and oxygen can be collected during aerobraking and stored for later use in orbit or on the planet. This technology would capture energy and oxygen from the plasma field that occurs naturally during hypersonic entry using well understood principles of magnetohydrodynamics and oxygen filtration. This innovative approach generates resources upon arrival at the operational site, and thus greatly differs from the traditional approach of taking everything you need with you from Earth. Fundamental analysis, computational fluid dynamics, and some testing of experimental hardware have established the basic feasibility of generating power during a Mars entry. Oxygen filtration at conditions consistent with spacecraft entry parameters at Mars has been studied to a lesser extent. Other uses of the MHD power are presented. This paper illustrates how some features of regenerative aerobraking may be applied to support human and robotic missions at Mars.

Moses, Robert W.↗
NASA NTRS2005

Plasma Assisted ISRU at Mars

NASA's exploration goals for Mars and Beyond will require new power systems and in situ resource utilization (ISRU) technologies. Regenerative aerobraking may offer a revolutionary approach for in situ power generation and oxygen harvesting during these exploration missions. In theory, power and oxygen can be collected during aerobraking and stored for later use in orbit or on the planet. This technology would capture energy and oxygen from the plasma field that occurs naturally during hypersonic entry using well understood principles of magnetohydrodynamics and oxygen filtration. This innovative approach generates resources upon arrival at the operational site, and thus greatly differs from the traditional approach of taking everything you need with you from Earth. Fundamental analysis, computational fluid dynamics, and some testing of experimental hardware have established the basic feasibility of generating power during a Mars entry. Oxygen filtration at conditions consistent with spacecraft entry parameters at Mars has been studied to a lesser extent. Other uses of the MHD power are presented. This paper illustrates how some features of regenerative aerobraking may be applied to support human and robotic missions at Mars.

Moses, Robert W.↗
NASA NTRS2015

Radiant Heating Tests on Meteoritic Material

A new effort geared toward modeling the physics of meteor entry and break-up is underway at NASA Ames Research Center. This is part of a broader interdisciplinary effort on providing physics-based risk assessment models for potentially hazardous objects. As part of the entry modeling task we are seeking to improve our understanding of, among other things, the ablation of meteoric material during high speed entry into earths atmosphere. Meteoroid entry differs greatly in some key respects from spacecraft entry modeling. First, the aerothermal environment at these high velocities (18 km/s) is dominated by radiation. Second, meteoroids less than, say, 50m in size, will likely lose a significant portion of their mass during the high-speed atmospheric entry process due to vaporization as well as melting and spallation. The mass of the object, in turn, directly affects the amount of energy deposited in the atmosphere, and therefore the amount of damage done. Thus it is important for us to understand and be able to model this process in greater detail in order to assess the hazard posed by these objects.In this presentation, we first give an overview of the simple ablation models that are typically used in meteor entry calculations. Additionally, we will present a new model which utilizes a similar approach to what is typically done for spacecraft TPS response modeling. This uses an equilibrium assumption near the surface to compute the ablation rate, as is done in heritage material response codes. Next we describe the radiant heating experiment which uses a high-powered laser to emulate the radiation dominated heating environment experienced by meteoroids during atmospheric entry. The facility the Laser Hardened Material Evaluation Laboratory (LHMEL) permits us to expose samples of meteoritic material to heating rates in excess of 100kW/sq.cm. An overview of the experimental set-up and test plan for the initial exploratory campaign at this facility will be given. Then we present both qualitative and quantitative results from this initial test series. Comparisons between predicted and measured ablation rates suggest that there may be significant blockage of the incident beam by the ablation plume. Furthermore, melt is shown to be a significant ablation mechanics, even at high heating rates (16 kW/sq.cm). Finally, comparisons between the phenomenology of the ablation of terrestrial rocks namely, basalt -- to that of meteorites show very different behavior. This is shown to likely be due in part to the effect of composition on the melt viscosity.

Stern, Eric↗
NASA NTRS2014

Potential for Integrating Entry Guidance into the Multi-Disciplinary Entry Vehicle Optimization Environment

The state-of-the-art in vehicle design decouples flight feasible trajectory generation from the optimization process of an entry spacecraft shape. The disadvantage to this decoupled process is seen when a particular aeroshell does not meet in-flight requirements when integrated into Guidance, Navigation, and Control simulations. It is postulated that the integration of a guidance algorithm into the design process will provide a real-time, rapid trajectory generation technique to enhance the robustness of vehicle design solutions. The potential benefit of this integration is a reduction in design cycles (possible cost savings) and increased accuracy in the aerothermal environment (possible mass savings). This work examines two aspects: 1) the performance of a reference tracking guidance algorithm for five different geometries with the same reference trajectory and 2) the potential of mass savings from improved aerothermal predictions. An Apollo Derived Guidance (ADG) algorithm is used in this study. The baseline geometry and five test case geometries were flown using the same baseline trajectory. The guided trajectory results are compared to separate trajectories determined in a vehicle optimization study conducted for NASA's Mars Entry, Descent, and Landing System Analysis. This study revealed several aspects regarding the potential gains and required developments for integrating a guidance algorithm into the vehicle optimization environment. First, the generation of flight feasible trajectories is only as good as the robustness of the guidance algorithm. The set of dispersed geometries modelled aerodynamic dispersions that ranged from +/-1% to +/-17% and a single extreme case was modelled where the aerodynamics were approximately 80% less than the baseline geometry. The ADG, as expected, was able to guide the vehicle into the aeroshell separation box at the target location for dispersions up to 17%, but failed for the 80% dispersion cases. Finally, the results revealed that including flight feasible trajectories for a set of dispersed geometries has the potential to save mass up to 430 kg.

spacecraft guidance↗
NASA NTRS2008

Ongoing Capabilities and Developments of Re-Entry Plasma Ground Tests at EADS-ASTRIUM

During re-entry, spacecrafts are subjected to extreme thermal loads. On mars, they may go through dust storms. These external heat loads are leading the design of re-entry vehicles or are affecting it for spacecraft facing solid propellant jet stream. Sizing the Thermal Protection System require a good knowledge of such solicitations and means to model and reproduce them on earth. Through its work on European projects, ASTRIUM has developed the full range of competences to deal with such issues. For instance, we have designed and tested the heat-shield of the Huygens probe which landed on Titan. In particular, our plasma generators aim to reproduce a wide variety of re-entry conditions. Heat loads are generated by the huge speed of the probes. Such conditions cannot be fully reproduced. Ground tests focus on reproducing local aerothermal loads by using slower but hotter flows. Our inductive plasma torch enables to test little samples at low TRL. Amongst the arc-jets, one was design to test architecture design of ISS crew return system and others fit more severe re-entry such as sample returns or Venus re-entry. The last developments aimed in testing samples in seeded flows. First step was to design and test the seeding device. Special diagnostics characterizing the resulting flow enabled us to fit it to the requirements.

Jullien, Pierre↗
NASA NTRS1979

Proposed methods for the reconstruction of the Pioneer/Venus probe entry trajectories

In connection with the Pioneer Venus Project, an orbiter and an array of atmospheric probes are sent to Venus in separate launches during 1978. The four probes, which are transported to the planet on an instrumented Bus are released approximately 20 days prior to encounters. There are three identical small probes and a large probe. Each is designed to make in-situ atmospheric measurements. One of the principal scientific objectives of the Pioneer Venus (P/V) multiprobe mission is the determination of the structure and composition of the Venusian atmosphere down to the planet surface. To enhance the science return it is of paramount importance to be able to accurately determine the spacecraft entry trajectories. In addition to the in-situ spacecraft measurements, earth-based Doppler data are available. The process felt to be most applicable for the P/V multiprobe entry trajectory reconstruction is considered, with emphasis placed on the highly dynamic upper entry segment of the trajectory.

Findlay, J. T.↗
NASA NTRSDate not supplied

Advances in Design Capabilities for Planetary Missions from the NASA Entry Systems Modeling and Instrumentation Portfolio

The Entry Systems Modeling project (ESM) is supported by both the NASA Space Technology and the Science Mission Directorates and focuses on developing simulation tools and validated models for characterizing the performance of entry systems tailored to planetary destinations across the Solar System. ESM is organized into six technical capability areas that together address all relevant factors related to spacecraft entry, as well as some aspects of descent: Thermal Protection System (TPS) Materials; Aerothermodynamics; Entry & Descent Vehicle Dynamics; Guidance, Navigation, and Control; Vehicle Systems Analysis; and Advanced Tools and Numerical Methods. Development within the capability areas is undertaken explicitly with a focus on transition and infusion to science missions, human exploration missions, and commercial space activities. The present talk details developments that specifically impact science missions, including simulation tool capabilities that aid in mission design and model development to understand entry system performance at a given destination. Examples of the successful infusion and transition of such project outcomes to science missions also are provided. Several simulation tool development efforts within ESM have resulted in new design capabilities for missions. One such outcome is improved toolsets for mission trajectory and concept of operations design. Specifically, an initiative to couple a leading tool for entry, ascent/descent, and orbital trajectory optimization (Program to Optimize Simulated Trajectories II or POST2) to those used within the Agency for interplanetary trajectory optimization (Copernicus and Monte) has made substantial progress, with the outcomes to date promising to allow efficient trajectory optimization across mission phases. Additionally, toolchains for the evaluation of vehicle performance during entry and descent have been developed that allow assessment of multi-dimensional aeroheating on detailed vehicle geometries, characterization of deployment and inflation of parachutes, and assessment of vehicle dynamic stability during descent. These capabilities are achieved by coupling diverse sets of physics together – material response, computational fluid dynamics, radiation, and vehicle dynamics – to suitably describe complex entry and descent phenomena. Several model development and validation efforts for specific destinations and entry regimes also are underway within the ESM project. For instance, new experimental capabilities to validate radiation models at low densities/high altitudes recently have been established with project support, specifically the Low-Density Shock Tube (LDST) at the NASA Ames Research Center Electric Arc Shock Tube (EAST) facility. The LDST is being leveraged to develop improved models of shock layer kinetics and radiation in Titan atmospheres, while future studies will be conducted in the LDST and the existing high velocity shock tube to provide validation data for radiation models of Venus, Ice Giants, and Mars atmospheres. Models describing the aerothermal and thermo-structural performance of Thermal Protection System (TPS) materials has been another focus, with multiscale modeling activities on-going for the two leading TPS materials applicable to a range of entry conditions and science missions: the Phenolic-Impregnated Carbon Ablator (PICA) and woven materials like 3D Mid-Density Carbon Phenolic (3MDCP). A continual effort is made to infuse and transition outcomes from ESM simulation tool and model development activities into relevant science missions. Significant progress has been made on this front, with missions such as Dragonfly, DAVINCI, and Mars Missions benefitting from project outcomes. The groundwork also is being laid to provide insights into forward looking missions to Gas/Ice Giants as well as for potential sample returns.

Justin Haskins↗
NASA NTRS2017

Autogenic-Feedback Training Exercise (AFTE) Mitigates the Effects of Spatial Disorientation to Simulated Orion Spacecraft Re-Entry: Individual Differences

NASA has identified a potential risk of spatial disorientation to future astronauts during re-entry of the proposed Orion spacecraft. The purpose of this study was to determine if a 6-hour physiological training procedure, Autogenic-Feedback Training Exercise (AFTE), can mitigate these effects. Twenty subjects were assigned to two groups (AFTE and Control) matched for motion sickness susceptibility and gender. All subjects received a standard rotating chair test to determine motion sickness susceptibility; three training sessions on a manual performance task; and four exposures to a simulated Orion re-entry test in the rotating chair. Treatment subjects were given two hours of AFTE training before each Orion test. A diagnostic scale was used to evaluate motion sickness symptom severity. Results showed that 2 hours of AFTE significantly reduced motion sickness symptoms during the second Orion test. AFTE subjects were able to maintain lower heart rates and skin conductance levels and other responses than the control group subjects during subsequent tests. Trends show that performance was less degraded for AFTE subjects. The results of this study indicate that astronauts could benefit from receiving at least 2 hours of preflight AFTE. In addition, flight crews could benefit further by practicing physiologic self-regulation using mobile devices.

human performance↗
NASA NTRS1973

Pioneer spacecraft for atmospheric entry missions to the outer planets.

The composition of the atmospheres of the Jovian planets is probably representative of the composition of the original matter from which the solar system was formed. An entry probe would be required for obtaining quantitative data about the atmosphere as a function of altitude. A mission involving the detachment of an atmospheric entry probe from a Pioneer spacecraft is being considered. Instruments in the probe are to measure pressure, temperature, and composition of the atmosphere during descent. The Pioneer spacecraft is to serve as a communications relay for the data transmitted from the probe.

Dixon, W. J.↗
NASA NTRS2021

Enabling Mars Radio Occultation by Smallsats

We introduce a Mars Radio Occultation (RO) mission concept leveraging on small satellite (smallsat) technologies that will enable measurements that address both Mars science and exploration priorities. The RO technique measures the Doppler shift of radio signals of a spacecraft occulting behind a planet’s limb that contains information about the planet’s atmospheric density, temperature, and pressure. We design a smallsat constellation and simulate “crosslink” RO observations between the smallsats to determine the expected accuracy and spatiotemporal coverage of Mars RO atmospheric profiles. Such measurements are key to spacecraft Entry, Descent, and Landing (EDL) and Ascent from the Surface (AST) of Mars. Today, there are limited number of Mars atmospheric profiles with high vertical resolution, especially measurements near the surface that are difficult to obtain from satellite-based passive instruments due to atmospheric absorption. We find that crosslink ROs between a constellation of six smallsats provide global and diurnal cycle coverage with dozens of occultations per day, providing temperature information from near-surface up to ~45 km with altitude-dependent accuracy that ranges between <0.5 K at lower altitude to <5 K in the middle-to-upper atmosphere. Ionospheric ROs reveal an expected electron density of ~5 – 10% at the peak ionospheric height between 100 and 140 km. We conclude that observations obtained from a smallsat constellation can augment existing observing platforms by reducing observational gaps and yield high resolution measurements required for safe spacecraft operations during EDL and AST.

Martinez, German↗
NASA NTRS1972

Thermal design of composite materials high temperature attachments

The thermal aspects of using filamentary composite materials as primary airframe structures on advanced atmospheric entry spacecraft such as the space shuttle vehicle were investigated to identify and evaluate potential design approaches for maintaining composite structures within allowable temperature limits at thermal protection system (TPS) attachments and/or penetrations. The investigation included: (1) definition of thermophysical data for composite material structures; (2) parametric characterization and identification of the influence of the aerodynamic heating and attachment design parameters on composite material temperatures; (3) conceptual design, evaluation, and detailed thermal analyses of temperature limiting design concepts; and (4) the development of experimental data for assessment of the thermal design methodologies and data used for evaluation of the temperature-limiting design concepts. Temperature suppression attachment concepts were examined for relative merit. The simple isolator was identified as the most weight-effective concept and was selected for detail design, thermal analysis, and testing. Tests were performed on TPS standoff attachments to boron/aluminum, boron/polyimide and graphite/epoxy composite structures.

Source record↗
NASA NTRS1972

Space vehicle gyroscope sensor applications

Considerations which form the basis for the specification, design and evaluation of gyroscopes for spaceborne sensor applications are presented. The applications are distinguished by basic vehicle category: launch vehicles, spacecraft, entry vehicles and sounding rockets. Specifically excluded from discussion are gyroscope effector applications. Exotic or unconventional gyroscopes for which operational experience is nonexistent are mentioned only briefly to alert the reader of future trends. General requirements for testing and evaluation are discussed.

Source record↗
NASA NTRS1992

Fast aurora zone analysis

The Flight Dynamics Facility (FDF) of the Flight Dynamics Division (FDD), of the Goddard Space Flight Center provides acquisition data to tracking stations and orbit and attitude services to scientists and mission support personnel. The following paper explains how a method was determined that found spacecraft entry and exit times of the aurora zone.

Booker, Mattie↗
NASA NTRS1993

Computational reacting gas dynamics

In the study of high speed flows at high altitudes, such as that encountered by re-entry spacecrafts, the interaction of chemical reactions and other non-equilibrium processes in the flow field with the gas dynamics is crucial. Generally speaking, problems of this level of complexity must resort to numerical methods for solutions, using sophisticated computational fluid dynamics (CFD) codes. The difficulties introduced by reacting gas dynamics can be classified into three distinct headings: (1) the usually inadequate knowledge of the reaction rate coefficients in the non-equilibrium reaction system; (2) the vastly larger number of unknowns involved in the computation and the expected stiffness of the equations; and (3) the interpretation of the detailed reacting CFD numerical results. The research performed accepts the premise that reacting flows of practical interest in the future will in general be too complex or 'untractable' for traditional analytical developments. The power of modern computers must be exploited. However, instead of focusing solely on the construction of numerical solutions of full-model equations, attention is also directed to the 'derivation' of the simplified model from the given full-model. In other words, the present research aims to utilize computations to do tasks which have traditionally been done by skilled theoreticians: to reduce an originally complex full-model system into an approximate but otherwise equivalent simplified model system. The tacit assumption is that once the appropriate simplified model is derived, the interpretation of the detailed numerical reacting CFD numerical results will become much easier. The approach of the research is called computational singular perturbation (CSP).

Lam, S. H.↗
NASA NTRS1998

Thermal Protection of Future Transatmospheric Vehicles

The field of hypersonic vehicle thermal protection is described from its beginning to the significant advancements that have been made in developing materials and systems used for thermally protecting vehicles which enter transit atmospheres. These include launch vehicles, entry and hypersonic vehicles, and planetary entry spacecraft, which as a group can be called transatmospheric vehicles which travel through the atmospheres of Earth and other planets at hypervelocities, with Mach numbers ranging from 5 to 50.

Rasky, Daniel J.↗