An investigation of the shielding properties of Moly Permalloy (trademark) shields designed for use with a hydrogen maser
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Wolf, S. A.↗
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Shielding Considerations for CubeSat Structures During Solar Maximum
The purpose of this lessons learned is to communicate the utility of shielding in small spacecraft planning for the support of mission assurance and reliability. Numerous SmallSats have been flying in polar low earth orbit for scientific, communications, technology demonstrations, and imaging with academic, commercial, and government interests. Shielding has been part of mission assurance and reliability from the advent of long duration spacecraft missions. The Shields-1 CubeSat has been operating in polar low earth orbit since 16 December 2018 with atomic number (Z)-grade radiation shielding and demonstrates shielding effectiveness. Shields-1 has collected a representative example of solar minimum data in 2019 with 8 Teledyne dosimeters over varying shielding effectiveness. It serves as current experimental data and has been compared with NOVICE Shielding estimates using the AP8 –AE8 trapped radiation model with the Shields-1 CAD and generic CubeSat 3 unit (U) models. Using NOVICE model radiation analysis coding, the shielding effectiveness's, based on a generic CubeSat 3U structure with 4 electronic boards, were estimated for aluminum wall thicknesses ranging from 0.204 cmto4.44 cm (0.550 g/cm2–12.0 g/cm2) thick aluminum. For modeled polar orbiting spacecraft, solar maximum total ionizing dose (TID)increases by nearly a magnitude for thin-walled aluminum 0.550g/cm2-0.686 g/cm2(0.204 cm –0.254 cm) typical CubeSat structures. The shielding effectiveness by NOVICE Sigma estimates, which is a shielding sphere approximation around a detector, showed a linear relationship with wall thickness, which increased over the wall thickness by a ratio of 1.43 determined by linear regression analysis. Using NOVICE Adjoint Monte-Carlo Modeling of solar minimum and solar maximum with the inclusion of a worst-case solar particle event over a one year mission without geomagnetic shielding, the TID for minimum and maximum conditions for a generic 3U with a wall thickness of 0.254 cm is 158 RAD and 1540 RAD, respectively. The modeled total solar maximum TID is over estimated, because at low orbital latitudes a spacecraft will have shielding from Earth's magnetic field. However, TID will still be significant at high latitudes over the poles, where a spacecraft is exposed in a solar particle event. In contrast, to a thin walled generic 3U CubeSat, Shields-1 electronics enclosure has a shielding effectiveness of 21.3 g/cm2from NOVICE Sigma modeling and is expected to show reduced total ionizing dose increases during the present active Solar Cycle 25 period. Because solar particle events during solar maximum increase TID on electronic parts with thin-walled shielding in short periods of time, it is a mission assurance and reliability consideration on the spacecraft’s mission value versus adding shielding for risk reduction of premature spacecraft or instrument payload loss. Since the volumes of many instruments and system electronics have reduced with small spacecraft, shielding material costs and weight penalties have diminished. A small spacecraft project budget and schedule may limit traditional radiation-hardened part use and radiation testing requirements, where shielding can contribute to mission assurance and reliability with reduced costs.
Shields-1↗
Shielding Considerations for CubeSat Structures During Solar Maximum
The purpose of this lessons learned paper is to communicate the utility of shielding in small spacecraft planning for the support of mission assurance and reliability. Numerous SmallSats have been flying in polar low Earth orbit for science, communications, technology demonstrations, and imaging with academic, commercial, and government interests. Shielding has been part of mission assurance and reliability from the advent of long duration spacecraft missions. The Shields-1 CubeSat has been operating in polar low Earth orbit since 16 December 2018 with atomic number (Z)-grade radiation shielding and demonstrates shielding effectiveness. Shields-1 has collected a representative example of solar minimum data in 2019 with eight Teledyne µDosimeters over varying shielding effectivenesses. It serves as current experimental data and has been compared with NOVICE Shielding estimates using the AP8 – AE8 trapped radiation model with the Shields-1 CAD and generic CubeSat three unit (3U) models. Using NOVICE model radiation analysis coding, the shielding effectivenesses, based on a generic CubeSat 3U structure with four electronic boards, were estimated for aluminum wall thicknesses ranging from 0.204 cm to 4.44 cm (0.550 g/cm 2 to 12.0 g/cm 2 ). For modeled polar orbiting spacecraft, solar maximum total ionizing dose (TID) increases by nearly a magnitude for thin-walled aluminum 0.550-g/cm 2 - 0.686-g/cm 2 (0.204-cm – 0.254-cm) typical CubeSat structures. The shielding effectiveness estimated by NOVICE Sigma, which is a shielding sphere approximation around a detector, showed a linear relationship with wall thickness, which increased over the wall thickness by a ratio of 1.43 determined by linear regression analysis. Using NOVICE Adjoint Monte-Carlo Modeling of solar minimum and solar maximum with the inclusion of a worst-case solar particle event over a one-year mission without geomagnetic shielding, the TID for minimum and maximum conditions for a generic 3U with a wall thickness of 0.254 cm is 158 RAD and 1540 RAD, respectively. The modeled total solar maximum TID is over estimated, because at low orbital latitudes a spacecraft will have shielding from the Earth’s magnetic field. However, TID will still be significant at high latitudes over the poles, where a spacecraft is exposed in a solar particle event. In contrast to a thin-walled generic 3U CubeSat, the Shields-1 electronics enclosure has a shielding effectiveness of 21.3 g/cm 2 from NOVICE Sigma modeling and is expected to show reduced total ionizing dose increases during the present active Solar Cycle 25 period. Because solar particle events during solar maximum increase TID on electronic parts with thin-walled shielding in short periods of time, it is a mission assurance and reliability consideration on the mission value of the spacecraft versus adding shielding for risk reduction of premature spacecraft or instrument payload loss. Since the volumes of many instruments and system electronics have reduced with small spacecraft, shielding material costs and weight penalties have diminished. A small spacecraft project budget and schedule may limit traditional radiation-hardened part use and radiation testing requirements, where shielding can contribute to mission assurance and reliability with reduced costs.
Shields-1↗
Decreasing Proton Single Event Effects in CubeSats with Shielding
Shields-1, NORAD ID 43850,has been operating in polar low earth orbit since December 2018. The shielding has resulted in a lower total ionizing dose over time than typical aluminum thin-walled CubeSat structures. The reduced ionizing dosage caused by Shields-1 increases the reliability of commercial parts and reduces internal charging. Furthermore, the Shields-1 shielding reduces the ionizing particle fluence inside the spacecraft that contributes to the ionizing dose. NOVICE Sigma shielding software, using the (Aerospace Proton) AP8 –(Aerospace Electron) AE8 solar minimum trapped belt environment for a 1-year mission, estimates a 21.3 g/cm2 aluminum effective shielding for the Shields-1 electronics enclosure. This high areal density reduces not only the total number of energetic protons, but also reduces the number of ionizing particles over all modeled energies from the estimated shielded fluence for a 1-year mission. NOVICE Adjoint CAD modeling of the Shields-1 structure, with the detector located within the electronics enclosure, estimates that the total number of particles is reduced from 2.20x 10exp9 protons/cm2 to 1.52x 10exp8 protons/cm2, which represents 6.90% of the remaining particles(figure 1). By slowing down approximations of the integral proton fluence, a minimum proton threshold is estimated at 151 MeV. In comparison, a 0.204-cm aluminum thin-walled 3-unit (U) structure, with a 0.907 g/cm2 effective shielding, has 25.0% remaining particles and a minimum proton threshold of 36.2 MeV(figure 1). Proton energies that contribute to single event effects in radiation tolerant or higher semiconductor hardness are typically 100 MeV and higher. The Shields-1 electronics enclosure is estimated to attenuate energies: 100 MeV by 76.5%, 200 MeV by 61.7%, and 500 MeV by 63.0%(figure 2), when comparing the space environment proton differential fluence with the shielded differential fluence. The aluminum thin-walled structure is estimated to attenuate proton energies: 100 MeV by 13.7%, 200 MeV by 12.4%, and 500 MeV by 12.6%(figure 2), which are lower than for the Shields-1 electronics enclosure. The significant differences in attenuation between the Shields-1 electronics enclosure and aluminum thin-walled 3U structure show the additional utility of increasing shielding effectiveness for reducing the numbers of energetic protons that contribute to single event effects.
Larry Thomsen↗
Alternative MMOD Shielding Concepts
The main types of meteoroid and orbital debris (MMOD) shields are single-layer “monolithic” shields, dual-wall “Whipple” shields, and multi-wall shields (“Stuffed Whipple” and “multi-shock” are common types). Aluminum alloys are typically used for the outer bumper layer and for the rear wall of the dual- and multi-wall shields, although carbon-composites are increasingly employed in MMOD shields given their low-mass and high-strength. Ceramic (NextelTM) and KevlarTM fabrics are commonly used for the intermediate layers of Stuffed Whipple shields. The NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) group is continuously working to improve NASA spacecraft MMOD shielding by evaluating new materials and shielding concepts by test and analysis. HVIT has performed many hypervelocity impact tests over several years to evaluate alternative MMOD shield materials and concepts. This paper will provide results of this work in the following areas: (1) Material substitutions to improve radiation and MMOD protection within Stuffed Whipple shields, (2) Metallic and ceramic foam bumper and intermediate layer materials, (3) Self-sealing materials. In the first two areas listed above, the candidate shield materials were tested under similar test conditions and with fixed shield standoff and mass. Damage to the rear wall was quantified and compared to determine the best performing shield materials. In the self-sealing material evaluations, the test objective was to gauge the ability of different materials and techniques to stop leaks in a rear wall with a 1-atmosphere (air) delta-pressure across the wall. This paper provides results of the investigations and describes forward work to continue the development of the most promising MMOD shield alternatives. KevlarTM is a trademark of DuPont de Nemours, Inc. NextelTM is a trademark of 3M Corporation Note, Trade names and trademarks are used in this report for identification only. Their usage does not constitute an official endorsement, either expressed or implied, by the National Aeronautics and Space Administration.
Hypervelocity↗
Regolith Derived Heat Shield for Planetary Body Entry and Descent System with In Situ Fabrication
This NIAC project investigated an innovative approach to provide heat shield protection to spacecraft after launch and prior to each EDL thus potentially realizing significant launch mass savings. Heat shields fabricated in situ can provide a thermal-protection system for spacecraft that routinely enter a planetary atmosphere. By fabricating the heat shield with space resources from materials available on moons and asteroids, it is possible to avoid launching the heat-shield mass from Earth. Regolith has extremely good insulating properties and the silicates it contains can be used in the fabrication and molding of thermal-protection materials. Such in situ developed heat shields have been suggested before by Lewis. Prior research efforts have shown that regolith properties can be compatible with very-high temperature resistance. Our project team is highly experienced in regolith processing and thermal protection systems (TPS). Routine access to space and return from any planetary surface requires dealing with heat loads experienced by the spacecraft during reentry. Our team addresses some of the key issues with the EDL of human-scale missions through a highly innovative investigation of heat shields that can be fabricated in space by using local resources on asteroids and moons. Most space missions are one-way trips, dedicated to placing an asset in space for economical or scientific gain. However, for human missions, a very-reliable heat-shield system is necessary to protect the crew from the intense heat experienced at very high entry velocities of approximately 11 km/s at approximately Mach 33 (Apollo). For a human mission to Mars, the return problem is even more difficult, with predicted velocities of up to 14 km/s, at approximately Mach 42 at the Earth-atmosphere entry. In addition to human return, it is very likely that future space-travel architecture will include returning cargo to the Earth, either for scientific purposes or for commercial reasons. Platinum, titanium, helium 3, and other metals, elements and minerals are all high-value commodities in limited supply on Earth, and it may be profitable to mine these substances throughout the Solar System and return them to Earth, if an economical method can be found. To date, several private corporations have been launched to pursue these goals. Because the heat shield is the last element to be used in an Earth-return mission, a high penalty is paid in the propellant mass required to carry the heat shield to the destination and back. If the heat shield could be manufactured in space, and then outfitted on the spacecraft prior to the reentry at Earth, then significant propellant and mass savings could be achieved during launch and space operations. Preliminary mission architecture scenarios are described, which explain the potential benefits that may be derived from using an in-situ fabricated regolith heat shield. In order to prove that this is a feasible technology concept, this project successfully fabricated heat shield materials from mineral simulant materials of lunar and Martian regolith by two methods: 1) Sintering and 2) Binding the simulant with a "room-temperature vulcanizing" (RTV) silicone formulated to withstand high temperatures. Initially a third type of fabrication was planned using the hot waste stream from regolith ISRU processes. This fabrication method was discarded since the resulting samples would be too dense and brittle for heat shields. High temperature flame tests at KSC and subsequent arc jet tests at Ames Research Center (ARC) have proved promising. These coupon tests show favorable materials properties and have the potential to be a new way of fabricating heat shields for space entry into planetary atmospheres.
Hogue, Michael D.↗
Regolith Derived Heat Shield for Planetary Body Entry and Descent System with In Situ Fabrication
This NIAC project investigated an innovative approach to provide heat shield protection to spacecraft after launch and prior to each EDL thus potentially realizing significant launch mass savings. Heat shields fabricated in situ can provide a thermal-protection system for spacecraft that routinely enter a planetary atmosphere. By fabricating the heat shield with space resources from materials available on moons and asteroids, it is possible to avoid launching the heat-shield mass from Earth. Regolith has extremely good insulating properties and the silicates it contains can be used in the fabrication and molding of thermal-protection materials. Such in situ developed heat shields have been suggested before by Lewis. Prior research efforts have shown that regolith properties can be compatible with very-high temperature resistance. Our project team is highly experienced in regolith processing and thermal protection systems (TPS). Routine access to space and return from any planetary surface requires dealing with heat loads experienced by the spacecraft during reentry. Our team addresses some of the key issues with the EDL of human-scale missions through a highly innovative investigation of heat shields that can be fabricated in space by using local resources on asteroids and moons. Most space missions are one-way trips, dedicated to placing an asset in space for economical or scientific gain. However, for human missions, a very-reliable heat-shield system is necessary to protect the crew from the intense heat experienced at very high entry velocities of approximately 11 km/s at approximately Mach 33 (Apollo). For a human mission to Mars, the return problem is even more difficult, with predicted velocities of up to 14 km/s, at approximately Mach 42 at the Earth-atmosphere entry. In addition to human return, it is very likely that future space-travel architecture will include returning cargo to the Earth, either for scientific purposes or for commercial reasons. Platinum, titanium, helium 3, and other metals, elements and minerals are all high-value commodities in limited supply on Earth, and it may be profitable to mine these substances throughout the Solar System and return them to Earth, if an economical method can be found. To date, several private corporations have been launched to pursue these goals. Because the heat shield is the last element to be used in an Earth-return mission, a high penalty is paid in the propellant mass required to carry the heat shield to the destination and back. If the heat shield could be manufactured in space, and then outfitted on the spacecraft prior to the reentry at Earth, then significant propellant and mass savings could be achieved during launch and space operations. Preliminary mission architecture scenarios are described, which explain the potential benefits that may be derived from using an in-situ fabricated regolith heat shield. In order to prove that this is a feasible technology concept, this project successfully fabricated heat shield materials from mineral simulant materials of lunar and Martian regolith by two methods: 1) Sintering and 2) Binding the simulant with a "room-temperature vulcanizing" (RTV) silicone formulated to withstand high temperatures. Initially a third type of fabrication was planned using the hot waste stream from regolith ISRU processes. This fabrication method was discarded since the resulting samples would be too dense and brittle for heat shields. High temperature flame tests at KSC and subsequent arc jet tests at Ames Research Center (ARC) have proved promising. These coupon tests show favorable materials properties and have the potential to be a new way of fabricating heat shields for space entry into planetary atmospheres.
Hogue, Michael D.↗
Predicting the Penetration of a Shielded TPS Tile
All spacecraft are subject to the possibility of high-speed particle impacts during their mission life. In low earth orbit, those impacts could be the result of collisions with pieces of orbital debris or with meteorites. Beyond LEO, and especially beyond GEO, those impacts will likely be caused by meteorites. Such high-speed impacts on spacecraft surfaces create debris clouds that travel towards and eventually impact other downstream spacecraft components. In addition to the impulsive load that such debris clouds would impart to the spacecraft elements with which they subsequently collide, the largest fragment in these debris clouds poses a significant threat on its own to those spacecraft elements. In order to be able to assess the severity of the threat posed by such a fragment, it is important to be able to predict the extent of damage sustained by the impacted spacecraft element. In this paper, we present a new process for determining the penetration depth in a shielded TPS tile system. The requirements for the process were that it be applicable across a full spectrum of impact low velocities (i.e. from ~ 7 km/s to above 50 km/s), that it be applicable over a wide range of projectile materials densities (i.e. ranging from water to steel, for example), and that it be applicable over a wide range of trajectory obliquities (i.e. not just normal impacts). Furthermore, the process developed should be sufficiently flexible so that it can be used for shield design parameter and / or system configuration trade studies (i.e. be written in terms of as many shield and TPS tile system parameters as possible). Following the initial impact of the projectile on the outer wall of the dual-wall shield, a debris cloud (i.e. the primary debris cloud) is created that travels towards and impacts the inner wall of the shield. Within this debris cloud is a combination of solid, liquid, and vaporized material, depending on the impact velocity and the impedance mismatch of the projectile and outer wall materials. The impact of this debris cloud on the inner wall creates another debris cloud (i.e. the secondary debris cloud) that then travels towards and impacts the TPS tile. Within the particulate distribution of the material in each of these debris clouds is a so-called largest fragment. The process used to calculate the penetration depth in a shielded TPS tile system developed herein consists of the following three steps: 1. Does the initial projectile perforate the outer wall of the shield? If NO, the TPS tile remains undamaged. If YES, calculate the size and speed of the largest fragment exiting the outer wall of the shield and traveling towards the shield’s inner wall. 2. Does the largest fragment in the debris cloud exiting the outer wall of the shield perforate the inner wall of the shield? If NO, the TPS tile again remains undamaged. If YES, calculate the size and speed of the largest fragment exiting the rear of the shield inner wall. 3. Calculate the penetration depth of the largest fragment in the debris cloud exiting the inner wall into the TPS. Perforation of the inner and outer shield walls is predicted using the Cour-Palais penetration depth equation and an associated failure criterion [1]. The largest debris cloud fragment diameters and associated velocities are calculated using regression equations developed from a mix of hydrocode and experimental diameter and velocity information [2]. Finally, TPS penetration depths are calculated using an empirical penetration depth predictor equation based on high-speed impact test data and are calibrated using a limited number of hydrocode penetration depth predictions for water, nylon, aluminum, Dunite, and steel projectiles The predictions of the penetration depth calculation process are compared against the predictions of more than 60 SPHC hydrocode runs. These comparisons show that the process used to calculate TPS tile penetration depths usually yields values that are within 1 mm of the predictions of hydrocode simulations. If a maximum allowable TPS penetration depth is known, the process developed herein can now be used to develop a failure limit equation that would predict whether or not an impacting particle would penetrate deeply enough into the TPS to result in a “failed” or “not failed” end state. References [1] E.L. Christiansen, “Design and performance equations for advanced meteoroid and debris shields”, International Journal of Impact Engineering, Vol. 14, 1993, pp. 145-156. [2] W.P. Schonberg, “Predicting the size of the largest particle fragment in a debris cloud created by an orbital debris impact and its associated velocity”, 2021 Applied Space Environments Conference, NASA Jet Propulsion Laboratory, Pasadena, California, November, 2021.
Meteroid↗
Exploratory Environmental Tests of Several Heat Shields
Exploratory tests have been conducted with several conceptual radiative heat shields of composite construction. Measured transient temperature distributions were obtained for a graphite heat shield without insulation and with three types of insulating materials, and for a metal multipost heat shield, at surface temperatures of approximately 2,000 F and 1,450 F, respectively, by use of a radiant-heat facility. The graphite configurations suffered loss of surface material under repeated irradiation. Temperature distribution calculated for the metal heat shield by a numerical procedure was in good agreement with measured data. Environmental survival tests of the graphite heat shield without insulation, an insulated multipost heat shield, and a stainless-steel-tile heat shield were made at temperatures of 2,000 F and dynamic pressures of approximately 6,000 lb/sq ft, provided by an ethylene-heated jet operating at a Mach number of 2.0 and sea-level conditions. The graphite heat shield survived the simulated aerodynamic heating and pressure loading. A problem area exists in the design and materials for heat-resistant fasteners between the graphite shield and the base structure. The insulated multipost heat shield was found to be superior to the stainless-steel-tile heat shield in retarding heat flow. Over-lapped face-plate joints and surface smoothness of the insulated multi- post heat shield were not adversely affected by the test environment. The graphite heat shield without insulation survived tests made in the acoustic environment of a large air jet. This acoustic environment is random in frequency and has an overall noise level of 160 decibels.
Goodman, George P.↗
The Feasibility of Multipole Electrostatic Radiation Shielding
Although passive shielding appears to be the only workable solution for galactic cosmic radiation (GCR), active shielding may play an important augmenting role to control the dose from solar particle events (SPEs). It has been noted that, to meet the guidelines of NCRP Report No. 98 through the six SPEs of 1989, a crew member would need roughly double the passive shielding that is necessary to control the GCR dose . This would dramatically increase spacecraft mass, and so it has been proposed that a small but more heavily shielded storm shelter may be used to protect the crew during SPEs. Since a gradual SPE may last 5 or more days, staying in a storm shelter may be psychologically and physiologically distressing to the crew. Storm shelters do not provide shielding for the spacecraft itself against the SPE radiation, and radiation damage to critical electronics may result in loss of mission and life. Single-event effects during the radiation storm may require quick crew response to maintain the integrity of the spacecraft, and confining the crew to a storm shelter prohibits their attending to the spacecraft at the precise time when that attention is needed the most. Active shielding cannot protect against GCR because the particle energies are too high. Although lower energy particles are easier to stop in a passive shield, such shielding is more satisfactory against GCR than against SPE radiation because of the tremendous difference in their initial fluences. Even a small fraction of the SPE fluence penetrating the passive shielding may result in an unacceptably high dose. Active shielding is more effective than passive shielding against SPE radiation because it offers 100% shielding effectiveness up to the cutoff energy, and significant shielding effectiveness beyond the cutoff as well.
Metzger, Philip T.↗
X-Ray Computed Tomography Inspection of the Stardust Heat Shield
The "Stardust" heat shield, composed of a PICA (Phenolic Impregnated Carbon Ablator) Thermal Protection System (TPS), bonded to a composite aeroshell, contains important features which chronicle its time in space as well as re-entry. To guide the further study of the Stardust heat shield, NASA reviewed a number of techniques for inspection of the article. The goals of the inspection were: 1) to establish the material characteristics of the shield and shield components, 2) record the dimensions of shield components and assembly as compared with the pre-flight condition, 3) provide flight infonnation for validation and verification of the FIAT ablation code and PICA material property model and 4) through the evaluation of the shield material provide input to future missions which employ similar materials. Industrial X-Ray Computed Tomography (CT) is a 3D inspection technology which can provide infonnation on material integrity, material properties (density) and dimensional measurements of the heat shield components. Computed tomographic volumetric inspections can generate a dimensionally correct, quantitatively accurate volume of the shield assembly. Because of the capabilities offered by X-ray CT, NASA chose to use this method to evaluate the Stardust heat shield. Personnel at NASA Johnson Space Center (JSC) and Lawrence Livermore National Labs (LLNL) recently performed a full scan of the Stardust heat shield using a newly installed X-ray CT system at JSC. This paper briefly discusses the technology used and then presents the following results: 1. CT scans derived dimensions and their comparisons with as-built dimensions anchored with data obtained from samples cut from the heat shield; 2. Measured density variation, char layer thickness, recession and bond line (the adhesive layer between the PICA and the aeroshell) integrity; 3. FIAT predicted recession, density and char layer profiles as well as bondline temperatures Finally suggestions are made as to future uses of this technology as a tool for non-destructively inspecting and verifying both pre and post flight heat shields.
McNamara, Karen M.↗
Technique for Configuring an Actively Cooled Thermal Shield in a Flight System
Broad area cooling shields are a mass-efficient alternative to conductively cooled thermal radiation shielding. The shield would actively intercept a large portion of incident thermal radiation and transport the heat away using cryogenic helium gas. The design concept consists of a conductive and conformable surface that maximizes heat transfer and formability. Broad Area Cooled (BAC) shields could potentially provide considerable mass savings for spaceflight applications by eliminating the need for a rigid thermal radiation shield for cryogen tanks. The BAC consists of a network of capillary tubes that are thermally connected to a conductive shield material. Chilled helium gas is circulated through the network and transports unwanted heat away from the cryogen tanks. The cryogenic helium gas is pumped and chilled simultaneously using a specialized pulse-tube cryocooler, which further improves the mass efficiency of the system. By reducing the thermal environment temperature from 300 to 100 K, the radiative heat load on a cryogen tank could be reduced by an order of magnitude. For a cryogenic liquid propellant scenario of oxygen and hydrogen, the boiloff of hydrogen would be significantly reduced and completely eliminated for oxygen. A major challenge in implementing this technology on large tanks is that the BAC system must be easily scalable from lab demonstrations to full-scale missions. Also, the BAC shield must be conformable to complex shapes like spheres without losing the ability to maintain constant temperature throughout. The initial design maximizes thermal conductivity between the capillary tube and the conductive radiation shielding by using thin, corrugated aluminum foil with the tube running transverse to the folds. This configuration has the added benefit of enabling the foil to stretch and contract longitudinally. This allows the BAC to conform to the complex curvature of a cryogen tank, which is key to its success. To demonstrate a BAC shield system with minimal impact to current cryogen tank designs, the shielding must be applied after the final assembly of the tank and supporting structure. One method is to pre-fabricate the shield in long strips. A spool of corrugated aluminum foil with a thermally sunk aluminum capillary running through the center could then be simply wound around the cryogen tanks and encapsulated within the multi-layer insulation (MLI) blanket. Then, on orbit, the BAC would intercept thermal radiation coming in through the MLI and transport it away from the cryogen tanks. An optimization of the design could be done to take into account mass savings from thinner MLI blankets, eliminating solid thermal shields, and ultimately, a reduction in the required cryogen tank size.
Barkfknecht, Peter↗
Polyvinyl Alcohol and Medium Density Polyethylene Composite Materials for Neutron Radiation Shielding
Manned spacecraft venturing beyond the Earth’s Van Allen Belts will be subjected to high levels of radiation from both energetic particles emanating from the sun and from galactic atomic events, such as supernovae. The metal shell of a spacecraft offers some protection from gamma radiation and high energy particles originating from both the sun and cosmic sources. The interaction of these high energy particles and radiation with the spacecraft hull create secondary radiation and neutrons, additional hazards for astronauts and electronic systems. Shielding of neutrons is the most challenging of the high energy radiation, since neutrons have zero charge and are least affected by electrical fields, magnetic fields, and atomic forces. Previous studies showed that polyethylene is an effective neutron shielding material, due to its high hydrogen content. High energy neutrons are slowed to thermal neutrons through multiple collisions with the hydrogen atoms, increasing the effectiveness of neutron capture by boron-10 atoms. This study discusses the fabrication and testing of polyvinyl alcohol (PVOH) samples as a neutron shielding material, with comparison to medium density polyethylene (MDPE) composite samples to be used for neutron shielding for spacecraft. PVOH, an aqueous polymer, was chosen for hydrogen content and for enabling the use of organic reinforcement such as cellulose for secondary structural composites. OLTARIS (On-Line Tool for the Assessment of Radiation in Space) was used for radiation shielding simulation and compared with experimental data. Modeling indicates that the most effective shielding from radiation was in an aluminum layer backed by a polymer layer, either MDPE or PVOH. Cross-linked PVOH was also included in the OLTARIS modeling. All 3 polymer cases showed shielding improvement over aluminum alone, with similar dose equivalent reduction for galactic cosmic radiation(GCR)and for radiation from a solar particle event (SPE). Fourier Transform Infrared (FTIR) Spectroscopy was used to confirm the cross-linking in the PVOH sample and to determine the uniformity through the thickness. PVOH, cross-linked PVOH, and MDPE samples were tested for neutron shielding effectiveness using a 1 curie Americium-Beryllium neutron source. Test results indicated the most shielding in the cross-linked PVOH, followed by the pristine PVOH, and MDPE. Neutron testing suggests that polyvinyl alcohol is a good alternative to polyethylene for neutron shielding. The neutron shielding materials, PVOH and cross-linked PVOH developed under this research will enable the safe operation of spacecraft beyond the Earth’s protective Van Allen Radiation Belts. It will be crucial for NASA’s manned missions to the Moon or Mars to protect the health of astronauts and equipment from the harmful effects of excessive radiation.
neutron, shielding, radiation, polyvinyl alcohol, ↗
Development of Additive Manufacturing Technologies for 3D Printing of Spacecraft Heat Shields
Introduction: Ablative heat shields are an enabling technology for entry into planetary atmospheres. From the PICA heatshields used for several Mars rovers to the carbon phenolic material used for Galileo’s Jupiter entry probe, the heat shield manages the heat load transferred to the payload, protecting the sensitive scientific instruments carried on entry probes. The Additive Manufacturing of Thermal Protection Systems (AMTPS) project, an Early Career Initiative (ECI) funded by NASA’s Space Technology Mission Directorate and led by NASA Johnson Space Center, seeks to develop materials and processes for 3D printing ablative heat shields for spacecraft. Current methods for producing ablative heat shields are extremely labor intensive and re-quire extensive hands-on processes and quality control characterization. Additive manufacturing (AM) offers the possibility of reduced production times, improved reliability, and enhanced performance via graded compositions. Costs will also be reduced by reducing the time and labor required for heat shield production. Direct integration of the heat shield onto the structure during processing simplifies integration and reduces risk. Material Development: A critical challenge for the project is development of a material system that can (1) be printed in a near-net shape process and (2) perform well as an ablator. Achieving printability requires the material to flow under applied pressure, but maintain its shape once extruded from the printer nozzle. Ablative performance is measured by a multitude of markers, including char yield, char strength, thermal conductivity, and recession rate. Furthermore, there are several mechanical and thermal property considerations for vehicle integration including coefficient of thermal expansion (CTE) and residual stress. AM technology will be leveraged to grade the material formulation and properties through the thickness of the heat shield, an architecture not possible with current manufacturing processes. To this end, “robust” material formulations have been pre-pared with higher density for use on the surface where most ablation will occur. “Insulative” material formulations, with lower density and lower thermal conductivity, are prepared for use in the depth of the heat shield. This graded architecture will re-duce the overall mass of the heat shield and reduce costs and/or increase scientific payload capacities. To achieve a material system with the required properties, multiple resins have been investigated in collaboration with NASA Ames Research Center. To tune printability and performance, resin additives were studied to improve flexibility of the cured material while maintaining acceptable ablative performance. Material coupons were printed and studied via a suite of mechanical and thermal characterization methods. Arc jet testing was conducted at NASA Ames Research Center to evaluate ablative performance and thermal protection under conditions expected in atmospheric entry. Manufacturing Scale-Up: A partnership with Oak Ridge National Laboratory (ORNL) aims to enable full-scale fabrication of a 3D printed heat shield. Leveraging expertise in manufacturing and 3D printing at ORNL, a mid-scale manufacturing demonstration unit will be built and tested, using a dual-layer ablative system printed directly onto the titanium structure. Work on robotic system integration is ongoing and efforts to scale up material mixing with a material compound will ensure accurate and homogenous composition. Flight Test: A hypersonic sub-orbital flight test will provide a rigorous test of material performance ranging from ablation, thermal management, and mechanical integrity. Design of the capsule has taken place in collaboration with the University of Kentucky. Data collected from the flight will inform future design efforts in material formulation, printing methodology, and heat shield-capsule integration.
additive manufacturing↗
Alternative MMOD Shielding Concepts
A core objective of the National Aeronautics and Space Administration (NASA) Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) group is to improve spacecraft meteoroid and orbital debris (MMOD) shielding by evaluating new materials and shielding concepts by test and analysis. This paper describes hypervelocity impact tests that were performed over several years to evaluate alternative MMOD shielding concepts. Specific objectives of the work were to incorporate improved MMOD and radiation protection within Stuffed Whipple shields, and to evaluate various self-sealing MMOD shielding concepts. The tests were performed under similar impact conditions and with fixed shield standoff and comparable overall shield mass. Damage to the rear wall was quantified and compared to determine the best performing shield concepts. Generally, the best shielding for both MMOD and radiation protection included polymer materials rich in hydrogen introduced into the intermediate and final layers of the shielding. In the self-sealing material evaluations, the test objective was to gauge the ability of different materials and techniques to stop leaks in a rear wall with a delta-pressure across the wall typical of spacecraft crew cabins (i.e., up to 1 atmosphere pressure). Shielding containing ionomer films and elastomer materials satisfied this objective for pressure shell holes up to 6mm diameter. This paper provides results of the investigations.
Orbital Debris↗
Shielding against galactic cosmic rays
Ions of galactic origin are modified but not attenuated by the presence of shielding materials. Indeed, the number of particles and the absorbed energy behind most shield materials increases as a function of shield thickness. The modification of the galactic cosmic ray composition upon interaction with shielding is the only effective means of providing astronaut protection. This modification is intimately conntected with the shield transport porperties and is a strong function of shield composition. The systematic behavior of the shield properites in terms of microscopic energy absorption events will be discussed. The shield effectiveness is examined with respect to convectional protection practice and in terms of a biological endpoint: the efficiency for reduction of the probability of transformation of shielded C3H1OT1/2 mouse cells. The relative advantage of developing new shielding technologies is discussed in terms of a shield performance as related to biological effect and the resulting uncertainty in estimating astronaut risk.
Schimmerling, W.↗
Analytic Shielding Optimization to Reduce Crew Exposure to Ionizing Radiation Inside Space Vehicles
A sustainable lunar architecture provides capabilities for leveraging out-of-service components for alternate uses. Discarded architecture elements may be used to provide ionizing radiation shielding to the crew habitat in case of a Solar Particle Event. The specific location relative to the vehicle where the additional shielding mass is placed, as corroborated with particularities of the vehicle design, has a large influence on protection gain. This effect is caused by the exponential- like decrease of radiation exposure with shielding mass thickness, which in turn determines that the most benefit from a given amount of shielding mass is obtained by placing it so that it preferentially augments protection in under-shielded areas of the vehicle exposed to the radiation environment. A novel analytic technique to derive an optimal shielding configuration was developed by Lockheed Martin during Design Analysis Cycle 3 (DAC-3) of the Orion Crew Exploration Vehicle (CEV). [1] Based on a detailed Computer Aided Design (CAD) model of the vehicle including a specific crew positioning scenario, a set of under-shielded vehicle regions can be identified as candidates for placement of additional shielding. Analytic tools are available to allow capturing an idealized supplemental shielding distribution in the CAD environment, which in turn is used as a reference for deriving a realistic shielding configuration from available vehicle components. While the analysis referenced in this communication applies particularly to the Orion vehicle, the general method can be applied to a large range of space exploration vehicles, including but not limited to lunar and Mars architecture components. In addition, the method can be immediately applied for optimization of radiation shielding provided to sensitive electronic components.
Gaza, Razvan↗
A Launch Requirements Trade Study for Active Space Radiation Shielding for Long Duration Human Missions
A trade study for an active shielding concept based on magnetic fields in a solenoid configuration versus mass based shielding was developed. Monte Carlo simulations were used to estimate the radiation exposure for two values of the magnetic field strength and the mass of the magnetic shield configuration. For each field strength, results were reported for the magnetic region shielding (end caps ignored) and total region shielding (end caps included but no magnetic field protection) configurations. A value of 15 cSv was chosen to be the maximum exposure for an astronaut. The radiation dose estimate over the total shield region configuration cannot be used at this time without a better understanding of the material and mass present in the end cap regions through a detailed vehicle design. The magnetic shield region configuration, assuming the end cap regions contribute zero exposure, can be launched on a single Space Launch System rocket and up to a two year mission can be supported. The magnetic shield region configuration results in two versus nine launches for a comparable mass based shielding configuration. The active shielding approach is clearly more mass efficient because of the reduced number of launches than the mass based shielding for long duration missions.
Singleterry, Robert C., Jr.↗