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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 NTRS2010

Impact of Spacecraft Shielding on Direct Ionization Soft Error Rates for Sub-130 nm Technologies

We use ray tracing software to model various levels of spacecraft shielding complexity and energy deposition pulse height analysis to study how it affects the direct ionization soft error rate of microelectronic components in space. The analysis incorporates the galactic cosmic ray background, trapped proton, and solar heavy ion environments as well as the October 1989 and July 2000 solar particle events.

Pellish, Jonathan A.↗
NASA NTRS2019

Challenges in Qualification of Thermal Protection Systems in Extreme Entry Environments

Planetary entry vehicles employ ablative TPS materials to shield the aeroshell from entry aeroheating environments. To ensure mission success, it must be demonstrated that the heat shield system, including local features such as seams, does not fail at conditions that are suitably margined beyond those expected in flight. Furthermore, its thermal response must be predictable, with acceptable fidelity, by computational tools used in heat shield design. Mission assurance is accomplished through a combination of ground testing and material response modelling. A material's robustness to failure is verified through arcjet testing while its thermal response is predicted by analytical tools that are verified against experimental data. Due to limitations in flight-like ground testing capability and lack of validated high-fidelity computational models, qualification of heat shield materials is often achieved by piecing together evidence from multiple ground tests and analytical simulations, none of which fully bound the flight conditions and vehicle configuration. Extreme heating environments (>2000 W/sq. cm heat flux and >2 atm pressure), experienced during entries at Venus, Saturn and Ice Giants, further stretch the current testing and modelling capabilities for applicable TPS materials. Fully-dense Carbon Phenolic was the material of choice for these applications; however, since heritage raw materials are no longer available, future uses of re-created Carbon Phenolic will require re-qualification. To address this sustainability challenge, NASA is developing a new dual-layer material based on 3D weaving technology called Heat shield for Extreme Entry Environments (HEEET). Regardless of TPS material, extreme environments pose additional certification challenges beyond what has been typical in recent NASA missions. Scope of this presentation: This presentation will give an overview of challenges faced in verifying TPS performance at extreme heating conditions. Examples include: (1) Bounding aeroheating parameters (heat flux, pressure, shear and enthalpy) in ground facilities. How to certify TPS if environments can't be bounded or aeroheating parameters can't be simultaneously achieved. (2) Higher uncertainties in ground test environments (facility calibration and analytical predictions) at extreme conditions. (3) Testing in flows similar to planetary atmosphere composition (H2/He for Gas and Ice Giants). (4) Test sample size limitations for qualifying seam designs. (5) Lack of computational tools capable of simulating all significant aspects of TPS performance (including initiation and propagation of failures). This presentation will provide recommendations on how the EDL community can address these challenges and mitigate some of the risks involved in flying TPS materials at extreme conditions. Examples include: (1) Dedicated activity to understanding TPS failure modes. Develop computational tools capable of modelling fluid interaction with material's thermostructural response. Validate these tools through failure testing. A better understanding of failure mechanisms may eliminate the need to fully bound all aeroheating parameters in ground testing. (2) Enhancements to current testing facilities to simulate flight-like ablation mechanism (ex. testing in Nitrogen at Ames Interaction Heating Facility to limit oxidation in favor of more sublimation). (3) Improved characterization of test conditions with new diagnostic methods and determination of environment uncertainty through rigorous statistical analysis of available data. (4) Design margin policies that are directly tied to uncertainties in ground test environments and modelling fidelity

Mahzari, Milad↗
NASA NTRS1998

Anti-Matter Propulsion for Space. (Latest citations from the Aerospace Database)

The bibliography contains citations concerning techniques for the efficient production, long-term storage and effective utilization of antimatter for space propulsion. Inertial confinement fusion (ICF), magnetic mirror fusion, and liquid-propellant thermal antimatter fusion propulsion are some of the technologies discussed. Radiation shields, cryogenic confinement of plasma and single-stage-to-orbit vehicles are also cited. (Contains 50-250 citations and includes a subject term index and title list.)

ANTIPROTONS↗
NASA NTRS1984

Free jet feasibility study of a thermal acoustic shield concept for AST/VCE application: Single stream nozzles

A technology base for the thermal acoustic shield concept as a noise suppression device for single stream exhaust nozzles was developed. Acoustic data for 314 test points for 9 scale model nozzle configurations were obtained. Five of these configurations employed an unsuppressed annular plug core jet and the remaining four nozzles employed a 32 chute suppressor core nozzle. Influence of simulated flight and selected geometric and aerodynamic flow variables on the acoustic behavior of the thermal acoustic shield was determined. Laser velocimeter and aerodynamic measurements were employed to yield valuable diagnostic information regarding the flow field characteristics of these nozzles. An existing theoretical aeroacoustic prediction method was modified to predict the acoustic characteristics of partial thermal acoustic shields.

Majjigi, R. K.↗
NASA NTRS2018

Optimization of the Magnetic Field Topology in the Hall Effect Rocket with Magnetic Shielding

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kilowatt Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The TDU-1 thruster implements a magnetically shielded field topology and has demonstrated the elimination of the discharge channel erosion. Extensive wear testing of the TDU Hall thrusters has identified the thruster front pole covers as the next life limiting component. This effort aims to explore and investigate alternate magnetic field topologies to assess whether reductions in the front pole cover erosion can be attained while still maintaining very low erosion rates on the discharge channel walls. NASA GRC (Glenn Research Center) and JPL (Jet Propulsion Laboratory) have begun a magnetic field topology characterization and optimization study by designing four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the pole covers. Three of the four candidate magnetic field topologies have been manufactured subjected to an extensive test campaign that includes performance, plume, and stability characterization. In the Phase I test segment, the thruster's oscillation magnitude and laser-induced fluorescence (LIF) measurements of the plasma plume were performed for the three candidate topologies. In Phase I, the thruster's oscillation magnitude and LIF measurements were performed for the three candidate topologies. Phase I test results found that the B1 configuration attained lower oscillation levels than the baseline topology (B0). Additionally, laser induced fluorescence measurements along the discharge chamber centerline found that upstream retraction of the thruster’s peak magnetic field does result in an upstream shift of the acceleration zone but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment will include performing performance, stability, plume, and erosion measurements for the various candidate magnetic field topologies..

Kamhawi, Hani↗
NASA NTRS2018

Optimization of the Magnetic Field Topology in the Hall Effect Rocket with Magnetic Shielding

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The TDU-1 thruster implements a magnetically shielded field topology and has demonstrated the elimination of the discharge channel erosion. Extensive wear testing the TDU Hall thrusters has identified the thruster front pole covers as the next life limiting component. This effort aims to explore and investigate alternate magnetic field topologies to assess whether reductions in the front pole cover erosion can be attained while still maintaining very low erosion rates on the discharge channel walls. NASA GRC and JPL have begun a magnetic field topology characterization and optimization study by designing four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the pole covers. Three of the four candidate magnetic field topologies have been manufactured subjected to an extensive test campaign that includes performance, plume, and stability characterization. In Phase I of the testing campaign, the thruster's oscillation magnitude and Laser Induced fluorescence (LIF) measurements of the plasma plume were performed for the three candidate topologies. In Phase I, the thruster's oscillation magnitude and LIF measurements were performed for the three candidate topologies. Phase I test results found that the B1 configuration attained lower oscillation levels than B0. Additionally, LIF measurements along the discharge chamber centerline found that upstream retraction of the thruster's peak magnetic field does result in an upstream shift of the acceleration zone but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment will include performing performance, stability, plume, and erosion measurements for the various candidate magnetic field topologies.

Kamhawi, Hani↗
NASA NTRS2019

HERMeS Thruster Magnetic Field Topology Optimization Study: Performance, Stability, and Wear Results

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The TDU-1 thruster implements a magnetically shielded field topology and has demonstrated the elimination of the discharge channel erosion as a life limiting mechanism. Extensive wear testing the TDU Hall thrusters has identified the thruster front pole covers as the next life limiting component. This effort aims to explore and investigate alternate magnetic field topologies to assess whether reductions in the front pole cover erosion can be attained while still maintaining very low erosion rates on the discharge channel walls. Four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the front pole covers were designed. Three of the four candidate magnetic field topologies (B1, B2, and B4) have been manufactured and were subjected to an extensive test campaign that includes laser induced fluorescence (LIF), performance, stability, wear, plume, and thermal characterization. In Phase I, LIF measurements along the discharge chamber centerline found that upstream retraction of the thruster's peak magnetic field does result in an upstream shift of the acceleration zone but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment results found at a normalized thruster magnetic field setting of 1, the thruster performance was very similar for all configurations. Discharge current waveforms ripple data indicated that configurations Bo, B1, and B2 has similar oscillatory profiles with the B2 configuration transitioning to higher oscillatory mode at 400V instead of the 450V observed for configurations B0 and B1. Configuration B4 ripple data indicates that the thruster was operating in a very oscillatory mode above 350V. Inner front pole cover erosion rates for configuration B1 were approximately 65% relative to B0 and the erosion rates for configuration B2 were 40% relative to B0. Analysis of the performance, stability, and wear test results of configurations B0, B1, B2, and B4 indicates that configuration B2 may present an alternative option for B0 because it mostly maintained the same performance and stability as configuration B0 but with 40% of the inner front pole cover erosion rate.

Kamhawi, Hani↗
NASA NTRS2009

Handbook for Designing MMOD Protection

Spacecraft are subject to micro-meteoroid and orbital debris (MMOD) impact damage which have the potential to degrade performance, shorten the mission, or result in catastrophic loss of the vehicle. Specific MMOD protection requirements are established by NASA for each spacecraft early in the program/project life, to ensure the spacecraft meets desired safety and mission success goals. Both the design and operations influences spacecraft survivability in the MMOD environment, and NASA considers both in meeting MMOD protection requirements. The purpose of this handbook is to provide spacecraft designers and operations personnel with knowledge gained by NASA in implementing effective MMOD protection for the International Space Station, Space Shuttle, and various science spacecraft. It has been drawn from a number of previous publications [10-14], as well as new work. This handbook documents design and operational methods to reduce MMOD risk. In addition, this handbook describes tools and equations needed to design proper MMOD protection. It is a living report, in that it will be updated and re-released periodically in future with additional information. Providing effective and efficient MMOD protection is essential for ensuring safe and successful operations of spacecraft and satellites. A variety of shields protect crew modules, external pressurized vessels and critical equipment from MMOD on the International Space Station (ISS). Certain Space Shuttle Orbiter vehicle systems are hardened from MMOD impact, and operational rules are established to reduce the risk from MMOD (i.e., flight attitudes are selected and late inspection of sensitive thermal protection surfaces are conducted to reduce MMOD impacts). Science spacecraft include specific provisions to meet MMOD protection requirements in their design (for example, Stardust & GLAST). Commercial satellites such as Iridium and Bigelow Aerospace Genesis spacecraft incorporate MMOD protection. The development of low-weight, effective MMOD protection has enabled these spacecraft missions to be performed successfully. This handbook describes these shielding techniques. For future exploration activities to the Moon and Mars, implementing high-performance MMOD shielding will be necessary to meet protection requirements with minimum mass penalty. A current area of technology development in MMOD shielding is the incorporation of sensors to detect and locate MMOD impact damage. Depending on the type of sensor the signals from the sensor can be processed to infer the location of the impact and the extent of damage. The objective of the sensors is to locate critical damage that would endanger the spacecraft or crew immediately or during reentry (such as an air leak from crew module or critical damage to thermal protection system of reentry vehicles). The information from the sensors can then be used with repair kits, patch kits, hatch closure or other appropriate remedial techniques to reduce MMOD risk.

Arnold, Jim↗
NASA NTRS2016

Fabrication of Regolith-Derived Radiation Shields: Preliminary Results

Unlike the Earth, Mars and asteroids do not have a magnetosphere to protect humans, mechanisms and electronics from damaging Galactic Cosmic Radiation (GCR) and solar particle events (SPE). This presents one of the highest risks to crew and onboard electronics during interplanetary journeys. The goal of this project is to evaluate the effectiveness of carbonaceous asteroids and other hydrogen-rich materials as potential radiation shielding materials, which ultimately could be tested during planned crewed missions to a captured asteroid fragment (ARM). This type of investigation represents an initial effort to develop radiation shield material compositions, production methods and technologies, and optimization methodology for manufacturing radiation shields in deep space for large exploration human missions or by emerging new industries seeking to stage their spacecraft for the exploitation of the resources of asteroids. Carbonaceous chondrites (C-type) are of particular interest as sources of compounds such as water ice and hydrogen-rich carbon molecules, which can provide sufficient low Z element density to provide radiation protection at adequate shield thicknesses.

Regolith↗
NASA NTRS2006

A Practical Approach to Starting Fission Surface Power Development

The Prometheus Power and Propulsion Program has been reformulated to address NASA needs relative to lunar and Mars exploration. Emphasis has switched from the Jupiter Icy Moons Orbiter (JIMO) flight system development to more generalized technology development addressing Fission Surface Power (FSP) and Nuclear Thermal Propulsion (NTP). Current NASA budget priorities and the deferred mission need date for nuclear systems prohibit a fully funded reactor Flight Development Program. However, a modestly funded Advanced Technology Program can and should be conducted to reduce the risk and cost of future flight systems. A potential roadmap for FSP technology development leading to possible flight applications could include three elements: 1) Conceptual Design Studies, 2) Advanced Component Technology, and 3) Non-Nuclear System Testing. The Conceptual Design Studies would expand on recent NASA and DOE analyses while increasing the depth of study in areas of greatest uncertainty such as reactor integration and human-rated shielding. The Advanced Component Technology element would address the major technology risks through development and testing of reactor fuels, structural materials, primary loop components, shielding, power conversion, heat rejection, and power management and distribution (PMAD). The Non-Nuclear System Testing would provide a modular, technology testbed to investigate and resolve system integration issues.

Mason, Lee S.↗
NASA NTRS2019

Magnet Architectures and Active Radiation Shielding Study (MAARSS)

This report describes the analysis of several magnetic radiation shield architectures and the spacecraft systems associated with them, such as quench detection, thermal management, and compensation coils. This work was performed as part of the NASA Institute of Advanced Concepts (NIAC) Grant Phase II effort. Because radiation exposure during long duration space missions poses a significant risk for crewed space-flight, Advanced Magnetic Lab, Inc. (AML) and NASA are studying magnetic radiation shielding, which may generate low-mass protection. Recent technological developments in high-temperature superconducting technology suggest the possibility of such shielding. This report analyzes some of the technical difficulties involved in such a system and identifies technology areas where further investment would be warranted. Most analysis was performed on a baseline configuration (8-meter diameter coils, with 1 tesla field). However, a baseline design with an expandable coil configuration (16 meter diameter coils, 1.5 tesla field) has a significant potential to increase shielding efficiency. The magnetic forces on the magnetic shield's components are large enough to require detailed structural design. Analysis of the thermal structural response of a yttrium-barium-copper-oxide (YBCO) high-temperature, superconducting (HTS) tape in a large-scale solenoid magnet showed that structural and thermal performance improved by replacing Hastelloy® with graphene in the tape. Integrating graphene into the high-strength fiber (HSF) support structure brings even greater benefits. The energy stored in a shield system's magnetic field must be safely dissipated if a quench-type failure occurs. Analysis of quench detection by fiber-optic thermal sensing and improvement of quench characteristics using graphene showed quench is manageable. Low-temperature superconductors can quench when subjected to movement, but lab tests of coil expansion testing showed that HTS materials do not quench when subject to the movement associated with coil expansion. Compensation coils are necessary to reduce the magnetic field in the crew habitat. Analysis showed that subdividing the compensation coil into individually controlled sections allows for a robust, adjustable system. AML analyzed the forces on an approaching capsule that were a result of eddy currents in its structure induced by its motion through the magnetic field from the coils. AML also analyzed coil-to-coil forces. Thermal control to maintain the HTS material at a suitable temperature is challenging. The sunshield technology review revealed a greater possibility of thermal insulation of the coils and use of passive cooling. The availability of suitable cryocoolers was also reviewed. A parametric graph estimating mass and shielding effectiveness for strength and thickness was developed, and allows spacecraft designers to assess the tradeoffs of using a magnetic radiation shield.

Spacecraft↗
NASA NTRS1992

Mars: Past, present, and future; Proceedings of the Conference, Williamsburg, VA, July 16-19, 1991

The present conference discusses the terraforming of Mars, the history of the Viking lander efforts, prospective Mars exploration management and research initiatives in view of the Viking Missions, the rationale and constituencies for NASA's Space Exploration Initiative, Mars mission benefits, and a Congressional perspective on the Space Exporation initiative. Also discussed are a risk-based approach to robotic mission requirements, a Mars Environment Survey Mission concept, a Mars network of surface stations, Mars sample-return mission options, an integrated STS for Mars exploration, nuclear propulsion technology development by NASA/DOE, radiation shielding en route to Mars and on the surface, advanced technologies for nuclear propulsion, and 'smart' space suits for space exploration.

Pritchard, E. B.↗
NASA NTRSDate not supplied

Prediction and Enhancement of Thermal Protection Systems from Meteoroid Damage using a Smooth Particle Hydrodynamic Code

Interplanetary spacecraft are exposed to meteoroid fluxes that range in speeds from 10 to 72 km/sec, far above the capability of today’s test facilities to provide predictions for the likelihood of spacecraft critical penetration. Of special interest are sample return missions, which (though protected by shielding) must often survive years of exposure to the meteoroid environment in order to re-enter Earth’s atmosphere with their scientific cargo. This paper describes the simulation of meteoric material damage to thermal protection systems (TPS) housed beneath protective “garage” (shielding) enclosures using the Smooth Particle Hydrodynamics Code (SPHC) operated by the Institute for Defense Analyses and Stellingwerf Consulting in support of ongoing NASA tasks. The study outlined in this paper considered the impact effect of both meteoric materials such as iron, ice, and chondrites (dunnite), and non-meteoric materials such as aluminum and nylon against both external shielding materials (single and dual aluminum bumpers) and Heat shield for Extreme Entry Environments Technology (HEEET) TPS materials, used alone and in conjunction with shielding. A general predictive damage equation to HEEET TPS is developed from these SPHC simulations for velocities up to 70 km/sec.

Hypervelocity Impact, hydrocode, meteoroid and orb↗
NASA NTRS2013

Electrodynamic Dust Shield Demonstrator

The objective of the project was to design and manufacture a device to demonstrate a new technology developed by NASA's Electrostatics and Surface Physics Laboratory. The technology itself is a system which uses magnetic principles to remove regolith dust from its surface. This project was to create an enclosure that will be used to demonstrate the effectiveness of the invention to The Office of the Chief Technologist. ONE of the most important challenges of space exploration is actually caused by something very small and seemingly insignificant. Dust in space, most notably on the moon and Mars, has caused many unforeseen issues. Dirt and dust on Earth, while a nuisance, can be easily cleaned and kept at bay. However, there is considerably less weathering and erosion in space. As a result, the microscopic particles are extremely rough and abrasive. They are also electrostatically charged, so they cling to everything they make contact with. This was first noted to be a major problem during the Apollo missions. Dust would stick to the spacesuits, and could not be wiped off as predicted. Dust was brought back into the spacecraft, and was even inhaled by astronauts. This is a major health hazard. Atmospheric storms and other events can also cause dust to coat surfaces of spacecraft. This can cause abrasive damage to the craft. The coating can also reduce the effectiveness of thermal insulation and solar panels.' A group of engineers at Kennedy Space Center's Electrostatics and Surface Physics Laboratory have developed a new technology, called the Electrodynamic Dust Shield, to help alleviate these problems. It is based off of the electric curtain concept developed at NASA in 1967. "The EDS is an active dust mitigation technology that uses traveling electric fields to transport electrostatically charged dust particles along surfaces. To generate the traveling electric fields, the EDS consists of a multilayer dielectric coating with an embedded thin electrode grid running a multiphase low frequency AC signal. Electrostatically charged particles, such as those encountered on the moon, Mars, or an asteroid, are carried along by the traveling field due to the action of Coulomb and dielectrophoretic forces."2 The technical details have been described in a separate article. This document details the design and construction process of a small demonstration unit. Once finished, this device will go to the Office of the ChiefTechnologist at NASA headquarters, where it will be used to familiarize the public with the technology. 1 NASA KSC FO Intern, Prototype Development Laboratory, Kennedy Space Center, University of Central Florida Kennedy Space

Stankie, Charles G.↗
NASA NTRS2015

Propulsion Technology Assessment: Science & Enabling Technologies to Explore the Interstellar Medium

As part of a larger effort led by the Keck Institute for Space Studies at the California Institute of Technology, the Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar system's heliopause, approximately 100 AU from the center of the sun, within 10 years and within a 2025 to 2035 launch window. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature (MaSMi) Hall thruster, a solar sail and an electric sail. Two aerial densities, and thus characteristic accelerations, 0.426 mm/s2 and 0.664 mm/s2, were analyzed for the solar sail option in order understand the impact of near and long term development of this technology. Similarly, two characteristic accelerations, 1 mm/s2 and 2 mm/s2, were also analyzed for the electric sail option in addition to tether quantities of 10 and 20, respectively, and individual tether length of 20 km. A second analysis was conducted to determine what existing solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming an earth departure characteristic energy capability provided by a Space Launch System (SLS) Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist (E-Ju), and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist (E-Ju-Su-Sa). The Oberth maneuver would need to be performed very close to the sun, wherein this study assumed a perihelion distance of approximately 11 solar radii, or 0.05 AU, away from the surface. The heat shield technology required to perform this type of ambitious maneuver was assumed to be similar to that of NASA's Solar Probe Plus mission, which is slated to launch in July 2018. With respect to a SLS Block 1B earth departure characteristic energy capability of 100 sq km/s2 for the E-Ju trajectory option, results indicated that compared to having no advanced propulsion system onboard, both the MaSMi Hall thruster and solar sail options subtract approximately 8 to 10 years from the total trip time while the electric sail outperforms all options by subtracting up to 20 years. With respect to an average kick stage velocity capability of 2.5 to 3.5 km/s at perihelion, the most sensitive segment of the E-Ju-Su-Sa trajectory option, results indicated that both the MaSMi Hall thrust and solar sail options only subtract 1 to 3 years from the total trip time whereas the electric sail again outperforms all other options by subtracting up to 5 years. In other words, if the Technology Readiness Level of an electric sail could be increased in time, this propulsion technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so even faster. Completing such an ambitious mission in that short of a timespan would be very attractive to many as it would be well within the average career span of any of those involved.

Hopkins, Randall C.↗
NASA NTRS2016

Propulsion Technology Assessment: Science and Enabling Technologies to Explore the Interstellar Medium

Led by the Keck Institute for Space Studies at the California Institute of Technology, the Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar system's heliopause, approximately 100 AU from the center of the sun, within 10 years and within a 2025-2035 launch window. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature (MaSMi) Hall thruster, a solar sail and an electric sail. Two aerial densities, and thus characteristic accelerations, 0.426 mm/sq s and 0.664 mm/sq s were analyzed for the solar sail option in order understand the impact of near and long term development of this technology. Similarly, two characteristic accelerations, 1 mm/s2 and 2 mm/sq s, were also analyzed for the electric sail option in addition to tether quantities of 10 and 20, respectively, and individual tether length of 20 km. A second analysis was conducted to determine what existing solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming an earth departure characteristic energy capability provided by a Space Launch System (SLS) Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist (E-Ju), and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist (E-Ju-Su-Sa). The Oberth maneuver would need to be performed very close to the sun, wherein this study assumed a perihelion distance of approximately 11 solar radii, or 0.05 AU, away from the surface. The heat shield technology required to perform this type of ambitious maneuver was assumed to be similar to that of NASA's Solar Probe Plus mission, which is slated to launch in July 2018. With respect to a SLS Block 1B earth departure characteristic energy capability of 100 km2/sq s for the E-Ju trajectory option, results indicated that compared to having no advanced propulsion system onboard, both the MaSMi Hall thruster and solar sail options subtract approximately 8 to 10 years from the total trip time while the electric sail outperforms all options by subtracting up to 20 years. With respect to an average kick stage velocity capability of 2.5 to 3.5 km/s at perihelion, the most sensitive segment of the E-Ju-Su-Sa trajectory option, results indicated that both the MaSMi Hall thrust and solar sail options only subtract 1 to 3 years from the total trip time whereas the electric sail again outperforms all other options by subtracting up to 5 years. In other words, if the Technology Readiness Level of an electric sail could be increased in time, this propulsion technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so even faster. Completing such an ambitious mission in that short of a timespan would be very attractive to many as it would be well within the average career span of any of those involved.

Hopkins, Randall C.↗
NASA NTRS1980

Performance of entry heat shields on Pioneer Venus probes

The Pioneer Venus probes approached Venus with high relative velocity. As they entered the atmosphere, they were rapidly decelerated by aerodynamic drag, and a great deal of heat was generated. To protect the probe structure and the scientific instruments, a carbon phenolic heat shield was placed on the front of the probes. Because the design of heat shields for planetary entry is a developing technology, thermocouples were placed in the heat shields so that actual and predicted heat shield performance could be compared. The function of the heat shield is discussed, the probe environments during entry into the Venusian atmosphere are described, and some results from the heat shield experiment are presented. It was found that for the most part, the heat shields performed better than expected.

Pitts, W. C.↗
NASA NTRS2010

Dust Removal Technology Demonstration for a Lunar Habitat

We have developed an Electrodynamic Dust Shield (EDS), an active dust mitigation technology with applications to solar panels, thermal radiators, optical systems, visors, seals and connectors. This active technology is capable of removing dust and granular material with diameters as large as several hundred microns. In this paper, we report on the development of three types of EDS systems for NASA's Habitat Demonstration Unit (HDU). A transparent EDS 20 cm in diameter with indium tin oxide electrodes on a 0.1 mm-thick polyethylene terephtalate (PET) film was constructed for viewport dust protection. Two opaque EDS systems with copper electrodes on 0.1 mm-thick Kapton were also built to demonstrate dust removal on the doors of the HDU. A lotus coating that minimizes dust adhesion was added to one of the last two EDS systems to demonstrate the effectiveness of the combined systems.

Calle, C. I.↗