Results for “Bumper” — Page 4

NASA NTRS1993

Physics of debris clouds from hypervelocity impacts

The protection scheme developed for long duration space platforms relies primarily upon placing thin metal plates or 'bumpers' around flight critical components. The effectiveness of this system is highly dependent upon its ability to break up and redistribute the momentum of any particle which might otherwise strike the outer surface of the spacecraft. Therefore it is of critical importance to design the bumpers such that maximum dispersion of momentum is achieved. This report is devoted to an in-depth study into the design and development of a laboratory instrument which would permit the in-situ monitoring of the momentum distribution as the impact event occurs. A series of four designs were developed, constructed and tested culminating with the working instrument which is currently in use. Each design was individually tested using the Space Environmental Effects Facility (SEEF) at the Marshall Space Flight Center in Huntsville, Alabama. Along with the development of the device, an experimental procedure was developed to assist in the investigation of various bumper materials and designs at the SEEF. Preliminary results were used to compute data which otherwise were not experimentally obtainable. These results were shown to be in relative agreement with previously obtained values derived through other methods. The results of this investigation indicated that momentum distribution could in fact be measured in-situ as the impact event occurred thus giving a more accurate determination of the effects of experimental parameters on the momentum spread. Data produced by the instrument indicated a Gaussian-type momentum distribution. A second apparatus was developed and it was placed before the shield in the line of travel utilized a plate to collect impact debris scattered backwards. This plate had a passage hole in the center to allow the particle to travel through it and impact the proposed shield material. Applying the law of conservation of angular momentum a backward momentum vector was determined from the angular velocity of the plate. The forward scattered and backward scattered momentum values were then analyzed to judge the distribution of debris. Loss of momentum was attributed to the inaccuracies of the means of measurement. Assumptions of symmetrical debris for the forward and backward scattered directions also contributed to this loss.

Zee, Ralph↗

NASA NTRS1995

Structural Damage Prediction and Analysis for Hypervelocity Impact: Properties of Largest Fragment Produced by Hypervelocity Impact of Aluminum Spheres with Thin Aluminum Sheets

Results of a series of hypervelocity impact tests are presented. In these tests, 1.275-g, 9.53-mm-diameter, 2017-T4 aluminum spheres were fired at normal incidence at eight thicknesses of 6061-T6 aluminum sheet. Bumper thickness to projectile diameter (t/D) ratio ranged from 0.026 to 0.424. Nominal impact velocity was 6.7 km/s. Results of five tests using 6.35, 9.53, and 12.70-mm-diameter aluminum spheres and other aluminum alloy bumpers are also given. A large chunky fragment of projectile was observed at the center of the debris clouds produced by the impacts. The equivalent diameter of this large fragment ranged from 5.5 mm for the lowest t/D ratio to a minimum of 0.6 mm for the case where maximum breakup of the projectile occurred (t/D approximately 0.2 to 0.3). When the t/D ratio was 0.42, numerous large flaky fragments were evenly distributed in the external bubble of bumper debris. Velocity of the large central fragments decreased continuously with increasing t/D ratio, ranging from about 99 percent to less than 80 percent of the impact velocity. The change in the velocity of small fragments spalling from the rear of the projectile was used to obtain a relationship showing a linear increase in the size of the central projectile fragment with decrease in the shock-induced stress in the projectile.

Piekutowski, Andrew J.↗

NASA NTRS2005

Uncertainty Considerations for Ballistic Limit Equations

The overall risk for any spacecraft system is typically determined using a Probabilistic Risk Assessment (PRA). A PRA determines the overall risk associated with a particular mission by factoring in all known risks to the spacecraft during its mission. The threat to mission and human life posed by the micro-meteoroid and orbital debris (MMOD) environment is one of the risks. NASA uses the BUMPER II program to provide point estimate predictions of MMOD risk for the Space Shuttle and the ISS. However, BUMPER II does not provide uncertainty bounds or confidence intervals for its predictions. In this paper, we present possible approaches through which uncertainty bounds can be developed for the various damage prediction and ballistic limit equations encoded within the Shuttle and Station versions of BUMPER II.

Schonberg, W. P.↗

NASA NTRS2018

Extensible Column Compliance System

This research investigates a possible solution to lower the stiffness of the Extensible Columns, which provide support to the Mobile Launcher. Material_X was determined to be the material best suited to lower the stiffness while still supporting the load. The only vendor to supply sufficient data that could prove their product would work was Vendor_A. The data they provided on the Product_Y General Bumper was analyzed and a compliance system using the bumper was designed. A test plan was developed and bumpers were ordered for testing. Testing is currently in progress to determine if the Product_Y is truly a solution.

Vigneau, Alexander J.↗

NASA NTRSDate not supplied

The NASA JSC Hypervelocity Impact Technology (HVIT) Office

- HVIT’s operational objectives for NASA, commercial, and international partners are to 1) help to define MMOD risk requirements 2) provide assessments that identify risk drivers for a specific spacecraft design, and 3) work to design and optimize shielding technologies based on increasing performance and reducing mass. - The use of Bumper-code underpins all risk assessments provided by HVIT – by using a complete FEM model, Bumper can provide a wholistic view of risk for a given design and mission profile. - A full risk assessment, however, requires an iterative approach, where Bumper is first used to identify potential risk drivers. Subsequent analysis can focus on iterating the shield design in highest risk areas. And may include the refinement of BLEs through testing, design modification, or reexamination of the failure criteria. - HVIT utilizes a wide-range of expertise to provide the most accurate risk assessments possible – areas of active research are highlighted below.

hypervelocity impact↗

NASA NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

Introduction The Lockheed Martin Orion spacecraft conducted the Artemis I flight around the Moon from November 16 through December 11, 2022. After the flight, an engineer from the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group performed a Micrometeoroid and Orbital Debris MMOD analysis using the Bumper 3 risk assessment tool to predict the number of small impacts that would likely have occurred during the mission. Separately, a team from the same group inspected the Orion capsule for hypervelocity impact damage features. The results of the inspection were compared to those of the analysis to aid in improving the analysis, including the environment models. Scope of Work The spacecraft geometry model was created by Lockheed Martin during construction of the Artemis I Orion vehicle based on Computer Aided Design (CAD) models of the vehicle. New hypervelocity impact testing was performed to verify Ballistic Limit Equations (BLEs) used in the analysis to link impactor size and damage to the Thermal Protection System (TPS). The exact trajectory flown was recorded during the flight, including vehicle attitude. This data was used in conjunction with the ORDEM 3.2 and MEM 3 environment modeling tools to create models of particle flux impacting the spacecraft throughout the mission. Meteoroid shower forecast information was also included to account for additional particle flux associated with meteoroid showers. Inspection of the Orion capsule included the Backshell thermal tiles and the tape covering it, windows, fabric thermal materials, and small areas of other materials. Potential MMOD damage found was characterized using various techniques, including optical microscopy, computed tomography scanning, and X-ray spectroscopy. Findings The number of craters found in the Backshell tile, and their size distribution, matches well with the Bumper analysis prediction. Tape, window, and other material impacts recorded similarly align to Bumper analysis predictions. Conclusions and Recommendations This comparison of analysis with inspection of the hardware provides valuable insight into the MMOD environment and how accurately analysis tools assess the impact risk to spacecraft. As this was the first large, non-ablative returned surface from a lunar mission, this analysis extends the MMOD community’s insight beyond low Earth orbit into cis-lunar space.

MMOD↗

NASA NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

Introduction The Lockheed Martin Orion spacecraft conducted the Artemis I flight around the Moon from November 16 through December 11, 2022. After the flight, an engineer from the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group performed a Micrometeoroid and Orbital Debris MMOD analysis using the Bumper 3 risk assessment tool to predict the number of small impacts that would likely have occurred during the mission. Separately, a team from the same group inspected the Orion capsule for hypervelocity impact damage features. The results of the inspection were compared to those of the analysis to aid in improving the analysis, including the environment models. Scope of Work The spacecraft geometry model was created by Lockheed Martin during construction of the Artemis I Orion vehicle based on Computer Aided Design (CAD) models of the vehicle. New hypervelocity impact testing was performed to verify Ballistic Limit Equations (BLEs) used in the analysis to link impactor size and damage to the Thermal Protection System (TPS). The exact trajectory flown was recorded during the flight, including vehicle attitude. This data was used in conjunction with the ORDEM 3.2 and MEM 3 environment modeling tools to create models of particle flux impacting the spacecraft throughout the mission. Meteoroid shower forecast information was also included to account for additional particle flux associated with meteoroid showers. Inspection of the Orion capsule included the Backshell thermal tiles and the tape covering it, windows, fabric thermal materials, and small areas of other materials. Potential MMOD damage found was characterized using various techniques, including optical microscopy, computed tomography scanning, and X-ray spectroscopy. Findings The number of craters found in the Backshell tile, and their size distribution, matches well with the Bumper analysis prediction. Tape, window, and other material impacts recorded similarly align to Bumper analysis predictions. Conclusions and Recommendations This comparison of analysis with inspection of the hardware provides valuable insight into the MMOD environment and how accurately analysis tools assess the impact risk to spacecraft. As this was the first large, non-ablative returned surface from a lunar mission, this analysis extends the MMOD community’s insight beyond low Earth orbit into cis-lunar space.

MMOD↗

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 NTRSDate not supplied

Micrometeoroid and Orbital Debris (MMOD) Testing, Ballistic Limit Equation 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. This paper will expand on the hypervelocity impact (HVI) testing and ballistic limit equation (BLE) definition work that was partially presented at the 2nd International Orbital De-bris (IOC-II) Conference held in Sugar Land, TX in December 2023. 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 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 co-efficients 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 space-craft 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 sur-face assessment again used the MEM-3 engineering model to define the meteoroid environ-ment 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 Mar-shall 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↗

DOE OSTI2024

Survey of tungsten gross erosion from main plasma facing components in WEST during a L-mode high fluence campaign

An initial high fluence campaign was performed in WEST, in 2023, on the newly installed actively cooled tungsten divertor composed of ITER-grade monoblocks. The campaign consisted in the repetition of a 60 s long Deuterium L-mode pulse in attached divertor conditions, cumulating over 10000s of plasma exposure. A maximum deuterium fluence of approximately 5⋅1⁢026 m−2 was reached in the outer strike point region, representative of a few high performance ITER pulses. Gross tungsten erosion inferred from visible spectroscopy shows that the most eroded plasma facing component is the inner divertor target with rates ten times larger than on the outer divertor target. The outer midplane tungsten bumpers, located a few centimeters from the plasma, show gross erosion rates two times lower than at the outer divertor. We conclude that the outer midplane bumpers have a negligible contribution to the long range tungsten migration and deposition onto the lower divertor. The cumulated gross erosion rate on the inner divertor translates in an effective gross erosion thickness of about 20μ⁢m, while it is about 2μ⁢m for the outer divertor. Strikingly, these orderings coincide with the thickness of deposits found locally on the divertor: the exposed surfaces of high field side monoblocks are covered with several tens of μ⁢m tungsten deposits, while on the lower field side, few μ⁢m thin tungsten deposits are only found on the magnetically shadowed parts of monoblocks. The strong impact of those deposits on WEST operation, namely perturbation of surface temperature measurement with infra-red thermography, and the emission of flakes causing radiative perturbation of the confined plasma, calls for anticipating similar issues in ITER. In particular, the start of research operation shall consider the definition of a divertor erosion budget in order to anticipate the formation of deleterious deposits.

Fedorczak, N.↗

NASA NTRS1972

Multimaterial lamination as a means of retarding penetration and spallation failures in plates

Experimental data are presented which show that hypervelocity impact spallation and penetration failures of a single solid aluminum plate and of a solid aluminum plate spaced a distance behind a Whipple meteor bumper may be retarded by replacing the solid aluminum plate with a laminated plate. Four sets of experiments were conducted. The first set of experiments was conducted with projectile mass and velocity held constant and with polycarbonate cylinders impacted into single plates of different construction. The second set of experiments was done with single plates of various construction and aluminum spherical projectiles of similar mass but different velocities. These two experiments showed that a laminated plate of aluminum and polycarbonate or aluminum and methyl methacrylate could prevent spallation and penetration failures with a lower areal density than either an all-aluminum laminated plate or a solid aluminum plate. The aluminum laminated plate was in turn superior to the solid aluminum plate in resisting spallation and penetration failures. In addition, through an example of 6061-T6 aluminum and methyl methacrylate, it is shown that a laminated structure ballistically superior to its parent materials may be built. The last two sets of experiments were conducted using bumper-protected main walls of solid aluminum and of laminated aluminum and polycarbonate. Again, under hypervelocity impact conditions, the laminated main walls were superior to the solid aluminum main walls in retarding spallation and penetration failures.

Dibattista, J. D.↗

NASA NTRS1982

Thermal management for large space platforms

This paper provides an evaluation of heat rejection techniques applicable to multihundred-kilowatt space platforms. A number of promising heat rejection concepts were parametrically weight-optimized over a wide range of conditions to provide a 99% reliability of achieving a 10-yr life for the multihundred-kilowatt space platform. Three panel designs were considered: (1) an advanced meteoroid-bumpered hybrid heat pipe concept, (2) a bumpered liquid concept, and (3) a space constructable heat pipe radiator. The following are some of the significant findings from the study: (1) A single subsystem approach can be used with the heat pipe system, whereas several smaller subsystems are required for the pumped fluid systems. (2) The space constructable radiator approach offers a 10% weight reduction and operational advantages over the conventionally deployed panels.

Oren, J. A.↗

NASA NTRS1984

A review of micrometeoroid flux measurements and models for low orbital altitudes of the Space Station

A review of meteoroid flux measurements and models for low orbital altitudes of the Space Station has been made in order to provide information that may be useful in design studies and laboratory hypervelocity impact tests which simulate micrometeoroids in space for design of the main wall of the Space Station. This report deals with the meteoroid flux mass model, the defocusing and shielding factors that affect the model, the probability of meteoroid penetration of the main wall of a Space Station. Whipple (1947) suggested a meteoroid bumper, a thin shield around the spacecraft at some distance from the wall, as an effective device for reducing penetration, which has been discussed in this report. The equations of the probability of meteoroid penetration, the average annual cumulative total flux, and the equations for the thickness of the main wall and the bumper are presented in this report.

Susko, M.↗

NASA NTRS1987

Coaxial tube array space transmission line characterization

The coaxial tube array tether/transmission line used to connect an SP-100 nuclear power system to the space station was characterized over the range of reactor-to-platform separation distances of 1 to 10 km. Characterization was done with respect to array performance, physical dimensions and masses. Using a fixed design procedure, a family of designs was generated for the same power level (300 kWe), power loss (1.5 percent), and meteoroid survival probability (99.5 percent over 10 yr). To differentiate between vacuum insulated and gas insulated lines, two different maximum values of the E field were considered: 20 kV/cm (appropriate to vacuum insulation) and 50 kV/cm (compressed SF6). Core conductor, tube, bumper, standoff, spacer and bumper support dimensions, and masses were also calculated. The results of the characterization show mainly how transmission line size and mass scale with reactor-to-platform separation distance.

Switzer, Colleen A.↗

NASA NTRS1987

Coaxial tube array space transmission line characterization

The coaxial tube array tether/transmission line used to connect an SP-100 nuclear power system to the space station was characterized over the range of reactor-to-platform separation distances of 1 to 10 km. Characterization was done with respect to array performance, physical dimensions and masses. Using a fixed design procedure, a family of designs was generated for the same power level (300 kWe), power loss (1.5 percent), and meteoroid survival probability (99.5 percent over 10 yr). To differentiate between vacuum insulated and gas insulated lines, two different maximum values of the E field were considered: 20 kV/cm (appropriate to vacuum insulation) and 50 kV/cm (compressed SF6). Core conductor, tube, bumper, standoff, spacer and bumper support dimensions, and masses were also calculated. The results of the characterization show mainly how transmission line size and mass scale with reactor-to-platform separation distance.

Switzer, Colleen A.↗

NASA NTRS1990

Advanced meteoroid and debris shielding concepts

The results of the shielding screening and optimization work conducted at the NASA Johnson Space Center Hypervelocity Impact Research Laboratory (HIRL) are described. The impact tests to investigate and optimize certain aspects of the aluminum mesh double-bumper concept, which was considered the best performer in the screening test, are analyzed. Preliminary results indicate that the aluminum mesh double-bumper concept shows a 30 to 40 percent weight reduction compared to an aluminum Whipple shield in the velocity range study, besides such advantages as an increased resistance to penetration and the fact that the aluminum mesh does not produce damaging secondary ejecta particles. It is suggested that the ballistic protection can be improved even more if an intermediate layer of a high-strength cloth is attached near the rear wall.

Christiansen, Eric L.↗

NASA NTRS1990

Theoretical hypervelocity ballistic limit for single or double plates using nonlinear modal analysis

The original research on the use on nonlinear vibration technique to solve for the hypervelocity ballistic limit for double plates is examined. Such structure is commonly found in typical Space Station design where the incoming space or man-made debris would be fragmented upon hitting the outer plate (shield) and the subsequent impact on the main wall would result in a much reduced damge of the space station or spacecraft. The existing few theoretical impact equations do not agree well with each other. The existing computer code bumper used at NASA-Johnson Space Center appears to predict an unconservative ballistic limit when compared with experimental data where the velocity ranges from 3 km/s to 8 km/s. Such unconservative prediction is unacceptable from a practical safe design point of view. The bumper code is based on Wilkinson's (1968) paper and his equations have not been improved nor modified even though they are viewed with suspicion due to lack of agreement with experiments. To make matters worse, there is not another theory which is better than Wilkinson's equation and the designers are forced to use purely empirical Nysmith (1969) or semiempirical equations developed by Cour-Palais in 1969. The Cour-Palais equations were later modified empirically in 1989. The purpose of the present investigation is to examine the many assumptions of the Wilkinson equation. An attempt is made to present design charts based on the modified-Wilkinson equation so that the designer can get a feel of the ranges of the parameters which are of interest and discard a huge range of parameters, thus, significantly reducing the number of test shots required. The analysis is based on a solution of the governing nonlinear differential equations for a plate, assuming axisymmetric behavior using polar coordinates.

Hui, David↗

NASA NTRS1991

MLIBlast: A program to empirically predict hypervelocity impact damage to the Space Station

MLIBlast is described, which consists of a number of DOC PC based MIcrosoft BASIC program modules written to provide spacecraft designers with empirical predictions of space debris damage to orbiting spacecraft. The Spacecraft wall configuration is assumed to consist of multilayer insulation (MLI) placed between a Whipple style bumper and a pressure wall. Predictions are based on data sets of experimental results obtained from simulating debris impact on spacecraft. One module of MLIBlast facilitates creation of the data base of experimental results that is used by the damage prediction modules of the code. The user has a choice of three different prediction modules to predict damage to the bumper, the MLI, and the pressure wall.

Rule, William K.↗