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Christiansen, Eric L.

Publications and source records attributed to Christiansen, Eric L..

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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 NTRS2009

Ballistic Limit Equation for Single Wall Titanium

Hypervelocity impact tests and hydrocode simulations were used to determine the ballistic limit equation (BLE) for perforation of a titanium wall, as a function of wall thickness. Two titanium alloys were considered, and separate BLEs were derived for each. Tested wall thicknesses ranged from 0.5mm to 2.0mm. The single-wall damage equation of Cour-Palais [ref. 1] was used to analyze the Ti wall's shielding effectiveness. It was concluded that the Cour-Palais single-wall equation produced a non-conservative prediction of the ballistic limit for the Ti shield. The inaccurate prediction was not a particularly surprising result; the Cour-Palais single-wall BLE contains shield material properties as parameters, but it was formulated only from tests of different aluminum alloys. Single-wall Ti shield tests were run (thicknesses of 2.0 mm, 1.5 mm, 1.0 mm, and 0.5 mm) on Ti 15-3-3-3 material custom cut from rod stock. Hypervelocity impact (HVI) tests were used to establish the failure threshold empirically, using the additional constraint that the damage scales with impact energy, as was indicated by hydrocode simulations. The criterion for shield failure was defined as no detached spall from the shield back surface during HVI. Based on the test results, which confirmed an approximately energy-dependent shield effectiveness, the Cour-Palais equation was modified.

Ratliff, J. M.↗
NASA NTRS2009

Micrometeoroid and Orbital Debris Threat Mitigation Techniques for the Space Shuttle Orbiter

An overview of significant Micrometeoroid and Orbital Debris (MMOD) impacts on the Payload Bay Door radiators, wing leading edge reinforced carbon-carbon panels and crew module windows will be presented, along with a discussion of the techniques NASA has implemented to reduce the risk from MMOD impacts. The concept of "Late Inspection" of the Nose Cap and Wing leading Edge (WLE) Reinforced Carbon Carbon (RCC) regions will be introduced. An alternative mated attitude with the International Space Station (ISS) on shuttle MMOD risk will also be presented. The significant threat mitigation effect of these two techniques will be demonstrated. The wing leading edge impact detection system, on-orbit repair techniques and disabled vehicle contingency plans will also be discussed.

Hyde, James L.↗
NASA NTRS2009

Mitigation of EMU Glove Cut Hazard by MMOD Impact Craters on Exposed ISS Handrails

Recent cut damages to crewmember extravehicular mobility unit (EMU) gloves during extravehicular activity (EVA) onboard the International Space Station (ISS) has been found to result from contact with sharp edges or pinch points rather than general wear or abrasion. One possible source of cut-hazards are protruding sharp edged crater lips from impact of micrometeoroid and orbital debris (MMOD) particles on external metallic handrails along EVA translation paths. During impact of MMOD particles at hypervelocity an evacuation flow develops behind the shock wave, resulting in the formation of crater lips that can protrude above the target surface. In this study, two methods were evaluated to limit EMU glove cut-hazards due to MMOD impact craters. In the first phase, four flexible overwrap configurations are evaluated: a felt-reusable surface insulation (FRSI), polyurethane polyether foam with beta-cloth cover, double-layer polyurethane polyether foam with beta-cloth cover, and multi-layer beta-cloth with intermediate Dacron netting spacers. These overwraps are suitable for retrofitting ground equipment that has yet to be flown, and are not intended to protect the handrail from impact of MMOD particles, rather to act as a spacer between hazardous impact profiles and crewmember gloves. At the impact conditions considered, all four overwrap configurations evaluated were effective in limiting contact between EMU gloves and impact crater profiles. The multi-layer beta-cloth configuration was the most effective in reducing the height of potentially hazardous profiles in handrail-representative targets. In the second phase of the study, four material alternatives to current aluminum and stainless steel alloys were evaluated: a metal matrix composite, carbon fiber reinforced plastic (CFRP), fiberglass, and a fiber metal laminate. Alternative material handrails are intended to prevent the formation of hazardous damage profiles during MMOD impact and are suitable for flight hardware yet to be constructed. Of the four materials evaluated, only the fiberglass formed a less hazardous damage profile than the baseline metallic target. Although the CFRP laminate did not form any noticeable crater lip, brittle protruding fibers are considered a puncture risk. In parallel with EMU glove redesign efforts, modifications to metallic ISS handrails such as those evaluated in this study provide the means to significantly reduce cut-hazards from MMOD impact craters.

Christiansen, Eric L.↗
NASA NTRS2009

Mitigation of EMU Cut Glove Hazard from Micrometeoroid and Orbital Debris Impacts on ISS Handrails

Recent cut damages sustained on crewmember gloves during extravehicular activity (ISS) onboard the International Space Station (ISS) have been caused by contact with sharp edges or a pinch point according to analysis of the damages. One potential source are protruding sharp edged crater lips from micrometeoroid and orbital debris (MMOD) impacts on metallic handrails along EVA translation paths. A number of hypervelocity impact tests were performed on ISS handrails, and found that mm-sized projectiles were capable of inducing crater lip heights two orders of magnitude above the minimum value for glove abrasion concerns. Two techniques were evaluated for mitigating the cut glove hazard of MMOD impacts on ISS handrails: flexible overwraps which act to limit contact between crewmember gloves and impact sites, and; alternate materials which form less hazardous impact crater profiles. In parallel with redesign efforts to increase the cut resilience of EMU gloves, the modifications to ISS handrails evaluated in this study provide the means to significantly reduce cut glove risk from MMOD impact craters

Ryan, Shannon↗
NASA NTRS2009

Honeycomb vs. Foam: Evaluating Potential Upgrades to ISS Module Shielding

The presence of honeycomb cells in a dual-wall structure is advantageous for mechanical performance and low weight in spacecraft primary structures but detrimental for shielding against impact of micrometeoroid and orbital debris particles (MMOD). The presence of honeycomb cell walls acts to restrict the expansion of projectile and bumper fragments, resulting in the impact of a more concentrated (and thus lethal) fragment cloud upon the shield rear wall. The Multipurpose Laboratory Module (MLM) is a Russian research module scheduled for launch and ISS assembly in 2011 (currently under review). Baseline shielding of the MLM is expected to be predominantly similar to that of the existing Functional Energy Block (FGB), utilizing a baseline triple wall configuration with honeycomb sandwich panels for the dual bumpers and a thick monolithic aluminum pressure wall. The MLM module is to be docked to the nadir port of the Zvezda service module and, as such, is subject to higher debris flux than the FGB module (which is aligned along the ISS flight vector). Without upgrades to inherited shielding, the MLM penetration risk is expected to be significantly higher than that of the FGB module. Open-cell foam represents a promising alternative to honeycomb as a sandwich panel core material in spacecraft primary structures as it provides comparable mechanical performance with a minimal increase in weight while avoiding structural features (i.e. channeling cells) detrimental to MMOD shielding performance. In this study, the effect of replacing honeycomb sandwich panel structures with metallic open-cell foam structures on MMOD shielding performance is assessed for an MLM-representative configuration. A number of hypervelocity impact tests have been performed on both the baseline honeycomb configuration and upgraded foam configuration, and differences in target damage, failure limits, and derived ballistic limit equations are discussed.

Ryan, Shannon J.↗
NASA NTRS2009

Characteristics of Whipple Shield Performance in the Shatter Regime

Between the onset of projectile fragmentation and the assumption of rear wall failure due to an impulsive load, multi-wall ballistic limit equations are linearly interpolated to provide reasonable yet conservative predictions of perforation thresholds with conveniently simple mathematics. Although low velocity and hypervelocity regime predictions are based on analytical expressions, there is no such scientific foundation for predictions in the intermediate (or shatter) regime. As the debris flux in low earth orbit (LEO) becomes increasingly dominated by manmade pollution, the profile of micrometeoroid and orbital debris (MMOD) risk shifts continually towards lower velocities. For the International Space Station (ISS), encounter velocities below 7 km/s now constitute approximately 50% of the penetration risk. Considering that the transition velocity from shatter to hypervelocity impact regimes described by common ballistic limit equations (e.g. new non-optimum Whipple shield equation [1]) occurs at 7 km/s, 50% of station risk is now calculated based on failure limit equations with little analytical foundation. To investigate projectile and shield behavior for impact conditions leading to projectile fragmentation and melt, a series of hypervelocity impact tests have been performed on aluminum Whipple shields. In the experiments projectile diameter, bumper thickness, and shield spacing were kept constant, while rear wall thickness was adjusted to determine spallation and perforation limits at various impact velocities and angles. The results, shown in Figure 1 for normal and 45 impacts, demonstrated behavior that was not sufficiently described by the simplified linear interpolation of the NNO equation (also shown in Figure 1). Hopkins et al. [2] investigated the performance of a nominally-identical aluminum Whipple shield, identifying the effects of phase change in the shatter regime. The results (conceptually represented in Figure 2) were found to agree well with those obtained in this study at normal incidence, suggesting that shielding performance in the shatter regime could be well described by considering more complex phase conditions than currently implemented in most BLEs. Furthermore, evidence of these phase effects were found in the oblique test results, providing the basis for an empirical description of these effects that can be applied in MMOD risk assessment software. In this paper, results of the impact experiments are presented, and characteristics of target damage are evaluated. A comparison of intermediate velocity impact failure mechanisms in current BLEs are discussed and compared to the findings of the experimental study. Risk assessment calculations have been made on a simplified structure using currently implemented penetration equations and predicted limits from the experimental program, and the variation in perceived mission risk is discussed. It was found that ballistic limit curves that explicitly incorporated phase change effects within the intermediate regime lead to a decrease in predicted MMOD risk for ISS-representative orbits. When considered for all Whipple-based shielding configurations onboard the ISS, intermediate phase change effects could lead to significant variations in predicted mission risk.

Ryan, Shannon↗
NASA NTRS2009

Performance of Whipple Shields at Impact Velocities above 9 km/s

Whipple shields were first proposed as a means of protecting spacecraft from the impact of micrometeoroids in 1947 [1] and are currently in use as micrometeoroid and orbital debris shields on modern spacecraft. In the intervening years, the function of the thin bumper used to shatter or melt threatening particles has been augmented and enhanced by the use of various types and configurations of intermediate layers of various materials. All shield designs serve to minimize the threat of a spall failure or perforation of the main wall of the spacecraft as a result of the impact of the fragments. With increasing use of Whipple shields, various ballistic limit equations (BLEs) for guiding the design and estimating the performance of shield systems have been developed. Perhaps the best known and most used are the "new" modified Cour-Palais (Christiansen) equations [2]. These equations address the three phases of impact: (1) ballistic (<3 km/s), where the projectile is moving too slowly to fragment and essentially penetrates as an intact projectile; (2) shatter (3 to 7 km/s), where the projectile fragments at impact and forms an expanding cloud of debris fragments; and (3) melt/vaporization (>7 km/s), where the projectile melts or vaporizes at impact. The performance of Whipple shields and the adequacy of the BLEs have been examined for the first two phases using the results of impact tests obtained from two-stage, light-gas gun test firings. Shield performance and the adequacy of the BLEs has not been evaluated in the melt/vaporization phase until now because of the limitations of launchers used to accelerate projectiles with controlled properties to velocities above 7.5 km/s. A three-stage, light-gas gun, developed at the University of Dayton Research Institute (UDRI) [3], is capable of launching small, aluminum spheres to velocities above 9 km/s. This launcher was used to evaluate the ballistic performance of two Whipple shield systems, various thermal protection system materials, and other spacecraft-related materials to the impact of 1.6-mm- to 2.6-mm-diameter, 2017-T4 aluminum spheres at impact velocities ranging from 8.91 km/s to 9.28 km/s. Test results, details of the shield systems, and nominal ballistic limits for the two Whipple shields are shown in Figures 1 and 2.

Christiansen, Eric L.↗
NASA NTRS2009

Ballistic Performance of Porous Ceramic Thermal Protection Systems at 9 km/s

Porous-ceramic, thermal-protection-systems are used heavily in current reentry vehicles like the Orbiter, and they are currently being proposed for the next generation of manned spacecraft, Orion. These materials insulate the structural components and sensitive electronic components of a spacecraft against the intense thermal environments of atmospheric reentry. Furthermore, these materials are also highly exposed to space environmental hazards like meteoroid and orbital debris impacts. This paper discusses recent impact testing up to 9 km/s on ceramic tiles similar to those used on the Orbiter. These tiles have a porous-batting of nominally 8 lb/cubic ft alumina-fiber-enhanced-thermal-barrier (AETB8) insulating material coated with a damage-resistant, toughened-unipiece-fibrous-insulation (TUFI) layer.

Miller, Joshua E.↗
NASA NTRS2009

Projectile Density Effects on Shield Performance

In the past, the orbital debris environment was modeled as consisting entirely of aluminum particles. As a consequence, most of the impact test database on spacecraft micro-meteoroid and orbital debris (MMOD) shields, and the resulting ballistic limit equations used to predict shielding performance, has been based on using aluminum projectiles. Recently, data has been collected from returned spacecraft materials and other sources that indicate higher and lower density components of orbital debris also exist. New orbital debris environment models such as ORDEM2008 provide predictions of the fraction of orbital debris in various density bins (high = 7.9 g/cu cm, medium = 2.8 g/cu cm, and low = 0.9-1.1 g/cu cm). This paper describes impact tests to assess the effects of projectile density on the performance capabilities of typical MMOD shields. Updates to shield ballistic limit equations are provided based on results of tests and analysis.

Christiansen, Eric L.↗
NASA NTRS2009

Recent Shuttle Post Flight MMOD Inspection Highlights

Post flight inspections on the Space Shuttle Atlantis conducted after the STS-11.5 mission revealed a 0.11 inch (2.8 mm) hole in the outer face sheet of the starboard payload bay door radiator panel #4. The payload bay door radiators in this region are 0.5 inch (12.7 mm) thick aluminum honeycomb with 0.011 in (0.279 mm) thick aluminum face sheets topped with 0.005 in (0.127 mm) silver-Teflon tape. Inner face sheet damage included a 0.267 in (6.78 mm) long through crack with measureable deformation in the area of 0.2 in (5.1 mm). There was also a 0.031 in (0.787 nun) diameter hole in the rear face sheet. A large approximately l in (25 mm) diameter region of honeycomb was also destroyed. Since the radiators are located on the inside of the shuttle payload bay doors which are closed during ascent and reentry, the damage could only have occurred during the on-orbit portion of the mission. During the August 2007 STS-118 mission to the International Space Station, a micro-meteoroid or orbital debris (MMOD) particle impacted and completely penetrated one of shuttle Endeavour's radiator panels and the underlying thermal control system (TCS) blanket, leaving deposits on (but no damage to) the payload bay door. While it is not unusual for shuttle orbiters to be impacted by small MMOD particles, the damage from this impact is larger than any previously seen on the shuttle radiator panels. One of the largest impacts ever observed on a crew module window occurred during the November 2008 STS-126 mission to the International Space Station. Damage to the window was documented by the crew on orbit. Post flight inspection revealed a 0.4 in (10.8 mm) crater in the window pane, with a depth of 0.03 in (0.76 mm). The window pane was replaced due to the damage caused by this impact. Analysis performed on residue contained in dental mold impressions taken of the site indicated that a meteoroid particle produced this large damage site. The post flight inspection after the subsequent mission, STS-119 in March of 2009, produced a large MMOD impact feature in a wing leading edge reinforced carbon-carbon panel. The crater measured 0.18 in (4.5 nun) in diameter and was nearly 0.037 in (0.93 nun) deep. The thickness of the silicon carbide coating that protects the carbon substrate is nominally 0.02 in (0.5 nun) to 0.04 in (1 mm), making this a significant impact into the RCC. The damage occurred on the upper surface of the panel, which experiences lower heat loads on re-entry. This poster will document the data collected from the impact sites and will include results of the Scanning Electron Microscope/Energy Dispersive X-ray (SEM/EDX) analysis. Evidence will be presented that suggests a source of the impacts.

Hyed, James L.↗
NASA NTRS2008

Impact-Locator Sensor Panels

Electronic sensor systems for detecting and locating impacts of rapidly moving particles on spacecraft have been invented. Systems of this type could also be useful on Earth in settings in which the occurrence of impacts and/or the locations of impacts are not immediately obvious and there are requirements to detect and quickly locate impacts to prevent or minimize damage.

Christiansen, Eric L.↗
NASA NTRS2007

Extravehicular Mobility Unit Penetration Probability from Micrometeoroids and Orbital Debris: Revised Analytical Model and Potential Space Suit Improvements

The NASA Extravehicular Mobility Unit (EMU) micrometeoroid and orbital debris protection ability has recently been assessed against an updated, higher threat space environment model. The new environment was analyzed in conjunction with a revised EMU solid model using a NASA computer code. Results showed that the EMU exceeds the required mathematical Probability of having No Penetrations (PNP) of any suit pressure bladder over the remaining life of the program (2,700 projected hours of 2 person spacewalks). The success probability was calculated to be 0.94, versus a requirement of >0.91, for the current spacesuit s outer protective garment. In parallel to the probability assessment, potential improvements to the current spacesuit s outer protective garment were built and impact tested. A NASA light gas gun was used to launch projectiles at test items, at speeds of approximately 7 km per second. Test results showed that substantial garment improvements could be made, with mild material enhancements and moderate assembly development. The spacesuit s PNP would improve marginally with the tested enhancements, if they were available for immediate incorporation. This paper discusses the results of the model assessment process and test program. These findings add confidence to the continued use of the existing NASA EMU during International Space Station (ISS) assembly and Shuttle Operations. They provide a viable avenue for improved hypervelocity impact protection for the EMU, or for future space suits.

Chase, Thomas D.↗
NASA NTRS2007

Hypervelocity Impact Evaluation of Metal Foam Core Sandwich Structures

A series of hypervelocity impact (HVI) tests were conducted by the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology Facility (HITF) [1], building 267 (Houston, Texas) between January 2003 and December 2005 to test the HVI performance of metal foams, as compared to the metal honeycomb panels currently in service. The HITF testing was conducted at the NASA JSC White Sands Testing Facility (WSTF) at Las Cruces, New Mexico. Eric L. Christiansen, Ph.D., and NASA Lead for Micro-Meteoroid Orbital Debris (MMOD) Protection requested these hypervelocity impact tests as part of shielding research conducted for the JSC Center Director Discretionary Fund (CDDF) project. The structure tested is a metal foam sandwich structure; a metal foam core between two metal facesheets. Aluminum and Titanium metals were tested for foam sandwich and honeycomb sandwich structures. Aluminum honeycomb core material is currently used in Orbiter Vehicle (OV) radiator panels and in other places in space structures. It has many desirable characteristics and performs well by many measures, especially when normalized by density. Aluminum honeycomb does not perform well in Hypervelocity Impact (HVI) Testing. This is a concern, as honeycomb panels are often exposed to space environments, and take on the role of Micrometeoroid / Orbital Debris (MMOD) shielding. Therefore, information on possible replacement core materials which perform adequately in all necessary functions of the material would be useful. In this report, HVI data is gathered for these two core materials in certain configurations and compared to gain understanding of the metal foam HVI performance.

Yasensky, John↗
NASA NTRS2007

Code Assesses Risks Posed by Meteoroids and Orbital Debris

BUMPER II version 1.92e is a computer code for assessing the risk of damage from impacts of micrometeoroids and orbital debris on the International Space Station (ISS), including those parts of the ISS covered by shielding that affords partial protection against such impacts. (Other versions of BUMPER II have been written for other spacecraft.) Bumper II quantifies the probability of penetration of shielding and the damage to spacecraft equipment as functions of the size, shape, and orientation of the spacecraft; the parameters of its orbit; failure criteria that quantify impact damage at the threshold of failure for each spacecraft surface; and the impact-damage resistance of each spacecraft surface as defined by "ballistic limit equations" that return the size of a failure-causing particle as a function of target parameters (including materials, configurations, thicknesses, and gap distances) and impact conditions (impact velocity and the density and shape of the impactor). BUMPER II version 1.92e contains several dozen ballistic limit equations that are based on results from thousands of hypervelocity impact tests conducted by NASA on ISS shielding and other hardware, and on results from numerical simulations of impacts.

Christiansen, Eric L.↗
NASA NTRS2007

Space Shuttle MMOD Threat Mitigation Techniques

Prior to each shuttle mission, threat assessments are performed to determine the risk of critical penetration, payload bay door radiator tube leak and crew module window replacement from Micrometeoroid and Orbital Debris (MMOD). Mission parameters, such as vehicle attitude, exposure time and altitude are used as inputs for the analysis. Ballistic limit equations, based on hypervelocity impact testing of shuttle materials are used to estimate the critical particle diameters of the outer surfaces of the vehicle. The assessments are performed using the BUMPER computer code at the NASA/JSC Hypervelocity Impact Technology Facility (HITF). The most critical involves the calculation of Loss of Crew and Vehicle (LOCV) risk. In recent years, NASA has implemented several techniques to reduce the risk to the Shuttle from MMOD impacts. This paper will describe on-orbit inspection of the reinforced carbon-carbon (RCC) regions and the methods used discern hypervelocity impact damage. Impact damage contingency plans and on-orbit repair techniques will also be discussed. The wing leading edge impact detection system (WLEIDS) and it's role in the reduction of on-orbit risk reduction will be presented. Finally, an analysis of the effectivity of alternative shuttle flight attitudes on MMOD risk will be demonstrated.

Hyde, Justin L.↗
NASA NTRS2006

Metal Foam Shields

This paper compares the ballistic performance of metallic foam sandwich structures with honeycomb structures. Honeycomb sandwich structures, consisting of metallic or composite facesheets and honeycomb cores, are often used in spacecraft construction due to their light-weight and structural stiffness. Honeycomb panels, however, are considered rather poor candidates for protection from micrometeoroid orbital debris (MMOD) particles because the honeycomb channels the debris cloud from MMOD impacts on outer facesheet causing a concentrated load on the second facesheet. Sandwich structures with light-weight, open-cell metallic cores and metal or composite facesheets provide improved MMOD protection because channeling does not occur and because the core is more effective at disrupting hypervelocity impacts then honeycomb. This paper describes hypervelocity impact tests on metallic foam sandwich structures (aluminum and titanium) with metallic facesheets, compare them to equivalent mass and thickness honeycomb panels, based on the results of hypervelocity impact tests.

Christiansen, Eric L.↗