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NASA NTRSDate not supplied

Analysis of Hypervelocity-Impacted Thin Films for Space Applications

Exposure to the harsh space environment and hypervelocity impacts from micrometeoroids and small orbital debris can affect space operations through long-term degradation of spacecraft materials, surfaces, or systems. Multi-layer insulation (MLI) and coated, polyimide films are spacecraft materials commonly found on the external surfaces of spacecraft for thermal control and protection. Characterizing the damage to these exposed materials with in situ and laboratory measurements can better inform spacecraft designers and operators in understanding and mitigating surface or system degradation. In this paper, we examine flown Hubble Space Telescope (HST) Equipment Bay 5 MLI and coated, polyimide ground test articles to characterize hypervelocity impact damage. Impact feature cavities are inspected to identify damage, which can validate long-term degradation models, improve thermal management systems, or improve accuracy of damage predictions. An overview of the impact characterization using optical microscopy, narrowband spectroscopy, and Scanning Election Microscopy/Energy Dispersive X- ray (SEM/EDX) of the two materials is presented. Observations of robustness in the space environment, as well as a detailed assessment of cratering and penetration statistics for the space-exposed sample, also are discussed.

Melissa A. Ward↗

NASA NTRSDate not supplied

Micrometeoroid and Orbital Debris (MMOD) Testing, Ballistic Limit Equation Definition and Risk Assessment of the Exploration Extravehicular Mobility Unit (XEMU)

Micrometeoroid↗

NASA NTRSDate not supplied

Analysis of Hypervelocity-Impacted Thin Films for Space Applications

Exposure to the harsh space environment and hypervelocity impacts from micrometeoroids and small orbital debris can affect space operations through long-term degradation of spacecraft materials, surfaces, or systems. Multi-layer insulation (MLI) and coated polyimide films are space-craft materials commonly found on the external surfaces of spacecraft for thermal control and protection. Characterizing the damage to these exposed materials with in-situ and laboratory measurements can better inform spacecraft designers and operators in understanding and mitigating surface or system degradation. In this paper, we examine flown Hubble Space Telescope (HST) Electronic Bay 5 MLI and coated polyimide ground test articles to characterize hyper-velocity impact damage. Impact feature cavities are inspected to identify damage, which can validate long-term degradation models, improve thermal management systems, or improve accuracy of damage predictions. An overview of the impact characterization using optical microscopy, narrow-band spectroscopy, and scanning election microscopy/energy dispersive x-ray (SEM/EDX) of the two materials is presented. Observations of robustness in the space environment, as well as a detailed assessment of cratering and penetration statistics for the space-exposed sample, also are discussed.

Melissa A Ward↗

NASA NTRS2005

In-Space Repair of Reinforced Carbon-Carbon (RCC) Thermal Protection System Structures

Advanced repair and refurbishment technologies are critically needed for the RCC-based thermal protection system of current space transportation system as well as for future Crew Exploration Vehicles (CEV). The damage to these components could be caused by impact during ground handling or due to falling of ice or other objects during launch. In addition, in-orbit damage includes micrometeoroid and orbital debris impact as well as different factors (weather, launch acoustics, shearing, etc.) during launch and re-entry. The GRC developed GRABER (Glenn Refractory Adhesive for Bonding and Exterior Repair) material has shown multiuse capability for repair of small cracks and damage in reinforced carbon-carbon (RCC) material. The concept consists of preparing an adhesive paste of desired ceramic with appropriate adhesives and then applying the paste to the damaged/cracked area of the RCC composites with adhesive delivery system. The adhesive paste cures at 100-120 C and transforms into a high temperature ceramic during simulated entry conditions. A number of plasma torch and ArcJet tests were carried out to evaluate the crack repair capability of GRABER materials for Reinforced Carbon-Carbon (RCC) composites. For the large area repair applications, PLASTER (Patch Laminates and Sealant Technology for Exterior Repair) based systems have been developed. In this presentation, critical in-space repair needs and technical challenges as well as various issues and complexities will be discussed along with the plasma performance and post test characterization of repaired RCC materials.

Singh, Mrityunjay↗

NASA NTRS2006

In-Space Repair of Reinforced Carbon-Carbon Thermal Protection System Structures

Advanced repair and refurbishment technologies are critically needed for the thermal protection system of current space transportation system as well as for future Crew Exploration Vehicles (CEV). The damage to these components could be caused by impact during ground handling or due to falling of ice or other objects during launch. In addition, in-orbit damage includes micrometeoroid and orbital debris impact as well as different factors (weather, launch acoustics, shearing, etc.) during launch and re-entry. The GRC developed GRABER (Glenn Refractory Adhesive for Bonding and Exterior Repair) material has shown multiuse capability for repair of small cracks and damage in reinforced carbon-carbon (RCC) material. The concept consists of preparing an adhesive paste of desired ceramic with appropriate additives and then applying the paste to the damaged/cracked area of the RCC composites with adhesive delivery system. The adhesive paste cures at 100-120 C and transforms into a high temperature ceramic during simulated entry conditions. A number of plasma torch and ArcJet tests were carried out to evaluate the crack repair capability of GRABER materials for Reinforced Carbon-Carbon (RCC) composites. For the large area repair applications, integrated system for tile and leading edge repair (InSTALER) have been developed. In this presentation, critical in-space repair needs and technical challenges as well as various issues and complexities will be discussed along with the plasma performance and post test characterization of repaired RCC materials.

Singh, Mrityunjay↗

NASA NTRS2007

In-Space Repair and Refurbishment of Thermal Protection System Structures for Reusable Launch Vehicles

Advanced repair and refurbishment technologies are critically needed for the thermal protection system of current space transportation systems as well as for future launch and crew return vehicles. There is a history of damage to these systems from impact during ground handling or ice during launch. In addition, there exists the potential for in-orbit damage from micrometeoroid and orbital debris impact as well as different factors (weather, launch acoustics, shearing, etc.) during launch and re-entry. The GRC developed GRABER (Glenn Refractory Adhesive for Bonding and Exterior Repair) material has shown multiuse capability for repair of small cracks and damage in reinforced carbon-carbon (RCC) material. The concept consists of preparing an adhesive paste of desired ceramic with appropriate additives and then applying the paste to the damaged/cracked area of the RCC composites with an adhesive delivery system. The adhesive paste cures at 100-120 C and transforms into a high temperature ceramic during reentry conditions. A number of plasma torch and ArcJet tests were carried out to evaluate the crack repair capability of GRABER materials for Reinforced Carbon-Carbon (RCC) composites. For the large area repair applications, Integrated Systems for Tile and Leading Edge Repair (InSTALER) have been developed and evaluated under various ArcJet testing conditions. In this presentation, performance of the repair materials as applied to RCC is discussed. Additionally, critical in-space repair needs and technical challenges are reviewed.

Singh, M.↗

NASA NTRSDate not supplied

TPS Certification by Analysis: Model-Driven Characterization of Properties and Failure in Woven Thermal Protection Systems

Woven, ablative thermal protection system (TPS) materials provide a robust option for aggressive (re)entries and thus have been baselined for the upcoming Mars Sample Return (MSR) mission's Earth Entry System (EES). The reliability requirements for MSR-EES necessitate understanding of material property variability, which could be significant given the complex structure and anisotropic nature of properties in TPS weaves, as well as the response to potential impact with micrometeoroids or orbital debris during the EES re-entry. The TPS Certification by Analysis effort within the Entry Systems Modeling project seeks to provide computational models and analyses that support the certification against such material-based risks. For the present talk, focus will be given to the characterization of baseline woven TPS material properties and mechanical failure limits, which entails (1) use of computational techniques (e.g., machine learning) to interpret computed tomography images of the weave to generate representative structural models and (2) application of multiscale material modeling approaches to characterize thermomechanical and failure properties.

thermal protection systems↗

NASA NTRS1994

Environmental Durability Issues for Solar Power Systems in Low Earth Orbit

Space solar power systems for use in the low Earth orbit (LEO) environment experience a variety of harsh environmental conditions. Materials used for solar power generation in LEO need to be durable to environmental threats such as atomic oxygen, ultraviolet (UV) radiation, thermal cycling, and micrometeoroid and debris impact. Another threat to LEO solar power performance is due to contamination from other spacecraft components. This paper gives an overview of these LEO environmental issues as they relate to space solar power system materials. Issues addressed include atomic oxygen erosion of organic materials, atomic oxygen undercutting of protective coatings, UV darkening of ceramics, UV embrittlement of Teflon, effects of thermal cycling on organic composites, and contamination due to silicone and organic materials. Specific examples of samples from the Long Duration Exposure Facility (LDEF) and materials returned from the first servicing mission of the Hubble Space Telescope (HST) are presented. Issues concerning ground laboratory facilities which simulate the LEO environment are discussed along with ground-to-space correlation issues.

Degroh, Kim K.↗

NASA NTRS2013

Determine ISS Soyuz Orbital Module Ballistic Limits for Steel Projectiles Hypervelocity Impact Testing

A new orbital debris environment model (ORDEM 3.0) defines the density distribution of the debris environment in terms of the fraction of debris that are low-density (plastic), medium-density (aluminum) or high-density (steel) particles. This hypervelocity impact (HVI) program focused on assessing ballistic limits (BLs) for steel projectiles impacting the enhanced Soyuz Orbital Module (OM) micrometeoroid and orbital debris (MMOD) shield configuration. The ballistic limit was defined as the projectile size on the threshold of failure of the OM pressure shell as a function of impact speeds and angle. The enhanced OM shield configuration was first introduced with Soyuz 30S (launched in May 2012) to improve the MMOD protection of Soyuz vehicles docked to the International Space Station (ISS). This test program provides HVI data on U.S. materials similar in composition and density to the Russian materials for the enhanced Soyuz OM shield configuration of the vehicle. Data from this test program was used to update ballistic limit equations used in Soyuz OM penetration risk assessments. The objective of this hypervelocity impact test program was to determine the ballistic limit particle size for 440C stainless steel spherical projectiles on the Soyuz OM shielding at several impact conditions (velocity and angle combinations). This test report was prepared by NASA-JSC/ HVIT, upon completion of tests.

Lyons, Frankel↗

NASA NTRS2015

Joint Polar Satellite System (JPSS) Micrometeoroid and Orbital Debris (MMOD) Assessment

The Joint Polar Satellite System (JPSS) Project requested the NASA Engineering and Safety Center (NESC) conduct an independent evaluation of the Micrometeoroid and Orbital Debris (MMOD) models used in the latest JPSS MMOD risk assessment. The principal focus of the assessment was to compare Orbital Debris Engineering Model version 3 (ORDEM 3.0) with the Meteoroid and Space Debris Terrestrial Environment Reference version 2009 (MASTER-2009) and Aerospace Debris Environment Projection Tool (ADEPT) and provide recommendations to the JPSS Project regarding MMOD protection. The outcome of the NESC assessment is contained in this report.

Squire, Michael D.↗

NASA NTRS2010

Hypervelocity Impact Performance of Open Cell Foam Core Sandwich Panel Structures

Open cell metallic foam core sandwich panel structures are of interest for application in spacecraft micrometeoroid and orbital debris shields due to their novel form and advantageous structural and thermal performance. Repeated shocking as a result of secondary impacts upon individual foam ligaments during the penetration process acts to raise the thermal state of impacting projectiles ; resulting in fragmentation, melting, and vaporization at lower velocities than with traditional shielding configurations (e.g. Whipple shield). In order to characterize the protective capability of these structures, an extensive experimental campaign was performed by the Johnson Space Center Hypervelocity Impact Technology Facility, the results of which are reported in this paper. Although not capable of competing against the protection levels achievable with leading heavy shields in use on modern high-risk vehicles (i.e. International Space Station modules), metallic foam core sandwich panels are shown to provide a substantial improvement over comparable structural panels and traditional low weight shielding alternatives such as honeycomb sandwich panels and metallic Whipple shields. A ballistic limit equation, generalized in terms of panel geometry, is derived and presented in a form suitable for application in risk assessment codes.

Ryan, S.↗

NASA NTRS1993

Contamination measurements on experiment M0003

The contamination monitors on the M0003 experiment consisted of passive sample collectors to measure molecular and particulate deposition. The collectors were placed in the leading and trailing edge trays of M0003. The objective was to quantify the contaminants and determine how contamination affected the other materials in the trays. Duplicate collectors were placed inside the vacuum canisters that provided protection from the launch, ascent, deployment, and recovery environments. Analyses of the samples show small quantities of molecular deposition on both the leading and trailing edge trays. Silicone deposits are negligible. The aluminized fused silica, second surface mirrors showed small changes in solar absorptances. This is consistent with the small quantities of molecular deposits deduced from the FTIR measurements. The results of particle deposition, light scatter, and a micrometeoroid (or debris) impact are presented.

Borson, Eugene N.↗

NASA NTRSDate not supplied

TPS Certification by Analysis: Model-Driven Characterization of Properties and Failure in Woven Thermal Protection Systems

Woven, ablative thermal protection system (TPS) materials provide a robust option for aggressive (re)entries and thus have been baselined for the upcoming Mars Sample Return (MSR) mission’s Earth Entry System (EES). The reliability requirements for MSR-EES necessitate understanding of material property variability, which could be significant given the complex structure and anisotropic nature of properties in TPS weaves, as well as the response to potential impact with micrometeoroids or orbital debris during the EES re-entry. The TPS Certification by Analysis effort within the Entry Systems Modeling project seeks to provide computational models and analyses that support the certification against such material-based risks. For the present talk, focus will be given to the characterization of baseline woven TPS material properties and mechanical failure limits, which entails (1) use of computational techniques (e.g., machine learning) to interpret computed tomography images of the weave to generate representative structural models and (2) application of multiscale material modeling approaches to characterize thermomechanical and failure properties.

Justin B. Haskins↗

NASA NTRSDate not supplied

Micrometeoroid and Orbital Debris (MMOD) Testing, Ballistic Limit Definition and Risk Assessment of the Exploration Extravehicular Mobility Unit (xEMU)

A well-known hazard associated with exposure to the space environment is the risk of failure due to an impact from a micrometeoroid and orbital debris (MMOD) particle. As NASA prepares to return astronauts to the moon with the Artemis program, the next generation of spacesuit is in development to support future extravehicular activities (EVAs.) An MMOD impact to the spacesuit is of great concern as a large leak could prevent an astronaut from safely reaching an airlock in time resulting in a loss of life. The exploration extravehicular mobility unit (xEMU) must meet MMOD requirements for multiple environments including those in low earth orbit (LEO) as well as the meteoroid and secondary lunar regolith ejecta environments found on the lunar surface. The subject of this paper is an internal xEMU configuration design developed by NASA Johnson Space Center (JSC) personnel. The xEMU shares similarities with the legacy Extravehicular Mobility Unit (EMU) spacesuit that is currently used for ISS EVAs, however differences in the layup (e.g., materials, thicknesses, and layers) of the fabric environmental protection garment (EPG), portable life support system (xPLSS) and helmet required an extensive test program to determine ballistic performance. Over 100 hypervelocity impact (HVI) tests were performed by the NASA/JSC HVIT and White Sands Test Facility (WSTF) teams on the xEMU EPG, xPLSS and helmet to generate ballistic limit equations (BLEs) for MMOD impacts. Additionally, over 50 low speed tests (< 1km/s) were performed by the NASA/JSC HVIT and Southwest Research Institute (SwRI) teams on the xEMU EPG, xPLSS and helmet to generate BLEs for lunar ejecta impacts. Post testing, ballistic limit equations (BLEs) used to define the performance of the various regions on the xEMU spacesuit were developed from a generic set of BLEs. The HVI and low speed testing was performed to establish a physical basis for the equations with the coefficients and exponents of the generic BLEs adjusted to fit the test data. The xEMU BLEs were added to the NASA/JSC software application used for spacecraft MMOD risk assessments (BUMPER-3). A finite element model (FEM) of the xEMU spacesuit, which defines the size and shape of the spacesuit as well as the locations of the various shielding configurations, was created based on a solid model provided by the xEMU program office. Using the FEM file and added xEMU BLEs, BUMPER-3 assessments of the xEMU spacesuit for probability of no penetration (PNP) were performed. For the LEO assessment of a typical ISS EVA, the orbital debris and meteoroids environments were defined using the latest engineering models, ORDEM 3.2 and MEM-3 respectively. The lunar surface assessment again used the MEM-3 engineering model to define the meteoroid environment along with the current released lunar surface ejecta model, NASA SP-8013 (developed during the Apollo Program). The Space Team in the Natural Environments Branch at Marshall Space Flight Center (MSFC) will soon release the new Lunar Meteoroid Ejecta Engineering Model (LMEEM), at which time the xEMU lunar surface EVA will be reassessed. Assessment of the MMOD risk for an 8-hour, 2-person EVA in both LEO and on the lunar surface showed that the xEMU spacesuit meets the program technical requirement of 1 in 2500 failure odds. Similar to the legacy EMU spacesuit, the majority of the MMOD risk (96% of the LEO EVA risk and 99% of the lunar surface EVA risk) is concentrated in regions of xEMU that are comprised primarily of softgoods (arms, legs, and gloves) rather than the hardgoods (xPLSS, hard upper torso and helmet).

Micrometeoroid↗

NASA NTRS2014

Physics-based Entry, Descent and Landing Risk Model

A physics-based risk model was developed to assess the risk associated with thermal protection system failures during the entry, descent and landing phase of a manned spacecraft mission. In the model, entry trajectories were computed using a three-degree-of-freedom trajectory tool, the aerothermodynamic heating environment was computed using an engineering-level computational tool and the thermal response of the TPS material was modeled using a one-dimensional thermal response tool. The model was capable of modeling the effect of micrometeoroid and orbital debris impact damage on the TPS thermal response. A Monte Carlo analysis was used to determine the effects of uncertainties in the vehicle state at Entry Interface, aerothermodynamic heating and material properties on the performance of the TPS design. The failure criterion was set as a temperature limit at the bondline between the TPS and the underlying structure. Both direct computation and response surface approaches were used to compute the risk. The model was applied to a generic manned space capsule design. The effect of material property uncertainty and MMOD damage on risk of failure were analyzed. A comparison of the direct computation and response surface approach was undertaken.

descent and landing↗

NASA NTRS2019

Micro-meteoroid and orbital debris radar from Goldstone Radar observations

Micro-meteoroid and orbital debris (MMOD) refers to millions of micrometeoroids and orbital debris orbiting the Earth, which can come from remaining parts of spacecraft from previous missions. These debris objects keep colliding and creating new pieces of debris, with very different sizes ranging from millimeters to meters. About 99.3% of the debris objects have a size < 1 cm. These debris objects, traveling at speeds of 10 km/s, can cause significant damage to spacecraft, satellites, and astronauts. Despite the historic use of shielding to protect vital spacecraft components, MMOD still present a potential hazard for future missions and operations involving humans in space. Goldstone’s orbital debris radar (ODR) enhances the NASA orbital debris model with radar observation data, providing vital information on orbital debris detections: size, Doppler, range, and orbit inclination angle. These observations currently provide the only access to information about this size class (2 mm to 10 mm) of orbital debris. Goldstone’s ODR experiments operate as a bistatic radar, with overlapping beams providing a wide distribution of ranges. From 2004 to early 2018, Goldstone’s ODR has been collecting orbital debris data from altitudes covering 280 km to 3000 km. Since the decommissioning of the receive antenna in early 2018, Goldstone’s range coverage is reduced to a 300 km window within a 600 km to 1000 km potential range coverage, depending on the pointing geometry and antenna station used. We are investigating alternative transmit receive configurations at Goldstone to recover our previous wide range coverage.

Rodriguez-Alvarez, Nereida↗

NASA NTRS2009

Hypervelocity Impact Performance of Open Cell Foam Core Sandwich Panel Structures

Metallic foams are a relatively new class of materials with low density and novel physical, mechanical, thermal, electrical and acoustic properties. Although incompletely characterized, they offer comparable mechanical performance to traditional spacecraft structural materials (i.e. honeycomb sandwich panels) without detrimental through-thickness channeling cells. There are two competing types of metallic foams: open cell and closed cell. Open cell foams are considered the more promising technology due to their lower weight and higher degree of homogeneity. Leading micrometeoroid and orbital debris shields (MMOD) incorporate thin plates separated by a void space (i.e. Whipple shield). Inclusion of intermediate fabric layers, or multiple bumper plates have led to significant performance enhancements, yet these shields require additional non-ballistic mass for installation (fasteners, supports, etc.) that can consume up to 35% of the total shield weight [1]. Structural panels, such as open cell foam core sandwich panels, that are also capable of providing sufficient MMOD protection, represent a significant potential for increased efficiency in hypervelocity impact shielding from a systems perspective through a reduction in required non-ballistic mass. In this paper, the results of an extensive impact test program on aluminum foam core sandwich panels are reported. The effect of pore density, and core thickness on shielding performance have been evaluated over impact velocities ranging from 2.2 - 9.3 km/s at various angles. A number of additional tests on alternate sandwich panel configurations of comparable-weight have also been performed, including aluminum honeycomb sandwich panels (see Figure 1), Nomex honeycomb core sandwich panels, and 3D aluminum honeycomb sandwich panels. A total of 70 hypervelocity impact tests are reported, from which an empirical ballistic limit equation (BLE) has been derived. The BLE is in the standard form suitable for implementation in risk analysis software, and includes the effect of panel thickness, core density, and facesheet material properties. A comparison between the shielding performance of foam core sandwich panel structures and common MMOD shielding configurations is made for both conservative (additional 35% non-ballistic mass) and optimistic (additional mass equal to 30% of bumper mass) considerations. Suggestions to improve the shielding performance of foam core sandwich panels are made, including the use of outer mesh layers, intermediate fabric/composite layers, and varying pore density.

Ryan, Shannon↗

NASA NTRS1975

Studies of solar flares and impact craters in partially protected crystals

Solar-flare track gradients and microcrater densities were determined in two lunar rocks, and it was found that approximately 900 craters greater than or equal to 0.1 microns in diameter and between 10 and 15 times 10 to the minus sixth craters greater than or equal to 500 microns central pit diameter are produced per sq cm per yr per 2-pi steradians on the moon. There is no evidence for a change in the flux of particles producing submicron craters in the last 1,000,000 years. No anisotropy was found between lunar north and the plane of the ecliptic in the directional flux distribution of micrometeoroids producing submicron craters. In contrast to data reported by other investigators, the ratio of solar-flare track densities at 100-micron depth to the abundance of 0.1-micron craters/sq cm is determined to be approximately one for all samples studied.

Morrison, D. A.↗