Results for “Bumper” — Page 4

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↗

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.↗

NASA NTRS1991

Empirical predictions of hypervelocity impact damage to the space station

A family of user-friendly, DOS PC based, Microsoft BASIC programs written to provide spacecraft designers with empirical predictions of space debris damage to orbiting spacecraft is described. The spacecraft wall configuration is assumed to consist of multilayer insulation (MLI) placed between a Whipple style bumper and the pressure wall. Predictions are based on data sets of experimental results obtained from simulating debris impacts on spacecraft using light gas guns on Earth. A module of the program facilitates the creation of the data base of experimental results that are used by the damage prediction modules of the code. The user has the choice of three different prediction modules to predict damage to the bumper, the MLI, and the pressure wall. One prediction module is based on fitting low order polynomials through subsets of the experimental data. Another prediction module fits functions based on nondimensional parameters through the data. The last prediction technique is a unique approach that is based on weighting the experimental data according to the distance from the design point.

Rule, W. K.↗

NASA NTRS1990

Augmentation of orbital debris shielding for Space Station Freedom

Augmentation concepts for the Space Station Freedom (SSF) are described in detail and advantages and disadvantages of various proposals are evaluated. It is pointed out that early augmentation of SSF debris shielding, which would take place 6-9 years after the First Element Launch, could provide the greatest advantages to the program in terms of weight, cost, and safety benefits. Potential augmentation concepts include an attached, rigid aluminum second bumper; the multi-shock shield concept; a mesh double bumper concept; a debris shield augmentation element; a multi-shock airbag protection concept for 2-10 cm debris; and a debris sweeper. Requirements for and operation and deployment of these concepts are described in detail and an example of the augmentation effect on cumulative SSF probability of no-failure from debris and meteoroid impacts is given. An estimate of predicted encounter rates between debris and a 100 m radius sphere around SSF is presented.

Christiansen, Eric L.↗

NASA NTRS1991

Shield sizing and response equations

A consolidated list is presented of meteoroid debris shield equations which have been given in the referenced memorandums. In some cases, equations have been updated; thus, this memorandum supersedes reference 1. The equations are presented in two parts: (1) shield sizing equations which are used to produce preliminary estimates of shielding weights; and (2) response equations to describe the impact conditions (projectile size as a function of velocity, density, and impact angle) causing failure of a given shield that are to be used for probability analyses (such as in the modified BUMPER program). Specific equations are given that are applicable for the following types of shields: aluminum Whipple shields; Nextel multishock (MS) shields; and mesh double bumper (MDB) shields. These equations will be updated in the future as warranted by the results of additional HVI tests, analyses, and shield modeling.

Christiansen, Eric L.↗

NASA NTRS1992

Ballistic limit regression analysis for Space Station Freedom meteoroid and space debris protection system

Relationships defining the ballistic limit of Space Station Freedom's (SSF) dual wall protection systems have been determined. These functions were regressed from empirical data found in Marshall Space Flight Center's (MSFC) Hypervelocity Impact Testing Summary (HITS) for the velocity range between three and seven kilometers per second. A stepwise linear least squares regression was used to determine the coefficients of several expressions that define a ballistic limit surface. Using statistical significance indicators and graphical comparisons to other limit curves, a final set of expressions is recommended for potential use in Probability of No Critical Flaw (PNCF) calculations for Space Station. The three equations listed below represent the mean curves for normal, 45 degree, and 65 degree obliquity ballistic limits, respectively, for a dual wall protection system consisting of a thin 6061-T6 aluminum bumper spaced 4.0 inches from a .125 inches thick 2219-T87 rear wall with multiple layer thermal insulation installed between the two walls. Normal obliquity is d(sub c) = 1.0514 v(exp 0.2983 t(sub 1)(exp 0.5228). Forty-five degree obliquity is d(sub c) = 0.8591 v(exp 0.0428) t(sub 1)(exp 0.2063). Sixty-five degree obliquity is d(sub c) = 0.2824 v(exp 0.1986) t(sub 1)(exp -0.3874). Plots of these curves are provided. A sensitivity study on the effects of using these new equations in the probability of no critical flaw analysis indicated a negligible increase in the performance of the dual wall protection system for SSF over the current baseline. The magnitude of the increase was 0.17 percent over 25 years on the MB-7 configuration run with the Bumper II program code.

Jolly, William H.↗

NASA NTRS1992

Properties of largest fragment produced by hypervelocity impact of aluminum spheres with thin aluminum sheets

Results are presented from hypervelocity impact tests in which 1.275 g spheres of 2017-T4 Al alloy were fired at normal incidence at eight thicknesses of 6061-T6 Al alloy sheets, with impact velocity of about 6.7 km/sec; additional data are presented for smaller and larger spheres than these, in the cases of other Al alloy impact bumpers. A large fragment of the projectile is observable at the center of the debris clouds generated upon impact. The velocity of these large fragments decreased continuously with increasing bumper thickness/projectile diameter ratio, from 99 percent to less than 80 percent of impact velocity; there is a linear increase in the size of the central projectile fragment with decreasing shock-induced stress in the projectile.

Piekutowski, Andrew J.↗

NASA NTRS1992

Momentum distribution in debris cloud during hypervelocity impact

The long term operation of the Space Station Freedom requires a scheme to protect it from high velocity impacts by both man-made particles and micrometeor fragments. One such scheme is the use of metal plates to serve as shields against such orbital debris. These 'bumper' plates, as they are referred to, serve to break up any incident particle and redistribute its momentum over a larger area. It is therefore necessary to determine the momentum distribution within the debris cloud produced by such collisions in order to evaluate a materials effectiveness at accomplishing this task. This paper details the design and development of an innovative device which has made this possible. Momentum profiles were obtained for a series of test conditions. Total momentum values in the debris cloud were then calculated from these profiles. These results indicated that a momentum amplification exists with a multiplication factor of between 2 and 3. Thus the role of the bumper to serve as a means for momentum redistribution and not reduction was verified.

Lemaster, P.↗