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Most spacecraft have at least one pressurized vessel on board. Because of the serious damage that might result following a high-speed on-orbit space debris particle impact, a primary design consideration is the anticipation and mitigation of that damage. Depending on pressure vessel design and impact / operating conditions, a pressure vessel impacted at hypervelocity may experience either only relatively shallow damage; a through-hole, perhaps with localized liner cracking or composite peeling; or catastrophic failure (rupture). While a puncture and the resulting leak could de-stabilize an orbiting spacecraft, an on-orbit rupture could not only lead to spacecraft loss, but for human missions, possibly loss of life. Herein we present the development of a ballistic limit equation (BLE) and a rupture limit equation (RLE) for composite overwrapped pressure vessels. Similar to a BLE that can be used to characterize whether or not a pressure vessel would be punctured by a high-speed impact, an RLE is designed to differentiate between regions of operating and impact conditions that, given a perforation, would result in either a rupture or would result only in a relatively small hole or crack. In a risk assessment that considers the various failures that might occur following debris impact, both types of equations are required. Comparisons of the RLE and the BLE developed herein with experimental results shows that both equations are able to cleanly separate the regions of rupture from non-rupture, and perforation from non-perforation. As such, the equations presented are both highly accurate in predicting the response of the COPVs and impact conditions considered.

Most spacecraft have at least one pressurized vessel on board. In addition to a hole, it is possible that a pressure vessel may experience catastrophic failure (i.e., rupture) as a result of a hypervelocity impact. If a tank rupture were to occur on-orbit following a meteoroid or orbital debris particle impact, for example, not only could it lead to loss of life, but it would also generate a tremendous amount of debris that could compromise future space assets working in similar orbits. As a result, NASA and other space faring nations have put in place spacecraft design requirements to prevent additional sizable debris from being created in the event of pressure vessel rupture or catastrophic failure. In general, these requirements state that a spacecraft's stored energy devices are to be passivated at the end of a spacecraft's mission or useful life. Programs whose spacecraft designs are not able to comply with some aspects of those requirements employ an alternative, so-called “soft passivation” or “make safe”, option. This paper provides a summary of a project performed with the intent of providing some possible guidelines and considerations that can be used by satellite programs to help satisfy passivation requirements using a “soft passivation” approach, that is, when not able to perform complete, or “hard”, passivation, with specific regard to mitigating the threat posed by on-orbit meteoroid or orbital debris particle impacts.

Most spacecraft have a pressurized tank on board, such as a liquid propellant tank or a breathable air supply tank. Because of the serious damage that might result following an on-orbit micro-meteoroid or orbital debris (MMOD) particle impact, a primary design consideration of such (often highly) pressurized tanks is the mitigation of the damage that might occur in the event of such an impact. While catastrophic failure (i.e. rupture) following a high-speed MMOD particle impact would be disastrous, a puncture and the resulting thrust caused by the expulsion of fluids or gas from the perforated tank could result in the de-stabilization of the spacecraft's orbit which could, in turn, also lead to loss of the spacecraft. As such, in the event that catastrophic failure does not occur, risk assessments must still consider whether or not a perforation of a pressurized tank might occur and, in the event of a perforation, what size hole is created by the impact. This paper presents the development of a ballistic limit equation for highly pressurized COPVs that would differentiate between combinations of impact parameters and operating conditions that would result in a hole (without rupture) in the front side of the impacted pressure vessel from those that would not. Empirical models for the hole-out area in the composite overwrap and in the interior liner material in the event of a front side perforation are also developed and discussed.

A main design consideration for spacecraft destined to operate in the meteoroid or orbital debris (MOD) environment is the prevention of possible damage that might occur following an MOD im...

In a previous study, a set of nine existing penetration models for blunt projectiles were identified and reviewed for their ability to predict the plugging-mode ballistic limit velocity of monolithic titanium alloy, aluminium alloy, and steel plates. Assessed with a database of more than 650 experimental ballistic limit measurements, often at conditions beyond that for which they had been originally derived or validated, it was found that all nine models could predict the ballistic limit to within +/- 50% for more than 50% of the database entries. In this paper, simple modifications to these nine models are proposed that introduce empirical adjustments, reformulations of the target strength dependency, or a combination of both with the pragmatic goal of identifying a model suitable for application across the wide range of monolithic metallic targets impacted by different calibre fragment-simulating projectiles across a full range of ordnance and sub-ordnance velocities in our database. The modifications were able to improve the performance of all models across this range, with the best performing models able to predict the experimental ballistic limit to within +/- 10% for 50% of the database entries and to within +/- 20% for 76% of the database entries.

All spacecraft are subject to damage from on-orbit impacts by micro-meteoroids and orbital debris particles. Typical shielding consists of a “bumper” that stands off from the main “inner wall” of the spacecraft element. Because the calculation of the size of the hole and the length of the crack that would occur in a module wall following a perforating on-orbit impact is a key ingredient in the risk assessment process of any spacecraft, it is important to correctly and accurately calculate those quantities. In two previous studies, empirical hole diameter and crack length equations were developed for the thirteen common ISS module wall configurations. In this paper we present the development of a single hole diameter equation and a single crack length equation that would unify the thirteen independent hole diameter and crack length equations developed previously. The advantages of these new equations over those developed previously are that they (1) are no longer tied to or dependent on the parameters of any single ISS module wall configuration, and (2) allow the consideration of the impact of non-spherical projectiles on multi-wall systems. We also clarify some issues with respect to the original thirteen hole diameter and crack length equations as originally presented.

How well a spacecraft component resists the damage that might occur following a high-speed impact by a meteoroid or piece of orbital debris is one of the main design considerations for earth orbiting spacecraft. Most spacecraft have at least one pressurized tank on board, such as a liquid propellant tank. In addition to a hole, it is possible that, for certain tank designs, construction materials, operating conditions, and impact parameters, a pressurized vessel may experience catastrophic failure (i.e. rupture) as the result of such a high-speed impact. Risk assessments performed during the design phase of a spacecraft typically consider and weigh the likelihoods and consequences of various failures that might occur following a MOD particle impact, including whether or not such a tank rupture might occur. In this paper, the datasets from several test programs involving shielded and unshielded composite overwrapped pressure vessels (COPVs) are combined to yield a single, new equation that can be used for this purpose. Comparison with test data shows that this new equation is able to correctly discriminate between the COPV rupture and non-rupture data points for the full high-speed impact test database considered herein.

In keeping with the tradition of the Hypervelocity Impact Society, this paper is a written version of the keynote address I presented at the start of 2015 Hypervelocity Impact Symposium as the recipient of the Society’s Distinguished Scientist Award. It covers the highlights of my nearly 30 year career working in the area of hypervelocity impact phenomenology, mainly as it is applied to the protection of humans and spacecraft that work an operate in the micrometeoroid and orbital debris (MMOD) environment.

One of the primary design considerations of earth-orbiting spacecraft is the mitigation of the damage that might occur from an on-orbit MMOD impact. Traditional damage-resistant design consists of a ‘bumper’ that is placed a small distance away from a spacecraft component or from the wall of the element in which it is housed. The performance of such a multi-wall structural element is typically characterized by its ballistic limit equation (BLE), which defines the threshold particle size that results in a failure of the spacecraft element. BLEs are also key components of any micro-meteoroid/orbital debris (MMOD) risk assessment calculations. However, these assessments often call for BLEs to predict impact response for projectiles made of materials not used in the development of those BLEs. The question naturally arises regarding how close are the predictions of such BLEs when used in impact scenarios involving projectiles made of materials not necessarily considered in their development. In an effort to address this issue, a study was performed with the objective of assessing the validity of the NNO BLE for non-aluminum particles. Particle materials considered included steel, copper, and Al 2 O 3 (i.e. particles that are made of materials that are more dense than aluminum). Comparisons are made between actual test results involving these non-aluminum projectiles and the predictions of the NNO BLE. In nearly all cases, the NNO BLE was found not to work very well in the predicting failure / no failure response of these non-aluminum projectiles. A new NNO-type BLE is then developed that can be used to more reliably predict the response of dual-wall systems under the hypervelocity impact of such “heavier” non-aluminum projectiles.

With the improvements and advances made in the field of renewable energy, this alternative method has become competitive with more traditional sources of energy generation techniques, including hydro power or fossil fuels. This has led to construction of wind turbines in areas prone to seismic activity. To increase the power production from wind energy, wind turbines have increased in size and mass, which makes them more vulnerable to lateral loads such as seismic induced forces, wind loads, and in the case of offshore wind turbines, wave loads. For this reason, computational analysis in the field are recently focusing on considering the interaction between lateral loads to present more realistic and cost effective designs. In this paper, the nonlinear dynamic behavior of a 5-MW NREL wind turbine is evaluated considering different earthquake and wind intensities using a newly developed finite element model. The model is first calibrated and verified with simplified models using modal and static pushover analysis. Engineering demand parameters (EDP) and intensity measures (IM) are then obtained from nonlinear incremental dynamic analysis (IDA) and used to assess the probability of exceeding different damage states (DS) using fragility curves. From the findings in this research, it is shown that earthquake loads have considerable effects on the design and analysis of wind turbines. To increase the power production from wind energy, wind turbines have increased in size and mass, which makes them more vulnerable to seismic induced forces, wind loads, and in the case of offshore wind turbines, wave loads.The nonlinear dynamic behavior of a 5-MW NREL wind turbine is evaluated considering different earthquake and wind intensities using a newly developed finite element model.Engineering demand parameters (EDP) and intensity measures (IM) are then obtained from incremental dynamic analysis (IDA) and used to assess the probability of exceeding different damage states (DS).Tower design for the 5-MW turbine was found to be acceptable in the case of simultaneous earthquake and wind loads applied with different intensities.The results of fragility analysis was performed for 4 limit states which are top displacement, top rotation, material yielding, and total collapse (buckling).Fragility analysis showed that probability of failure due to top displacement and top rotation are more likely to occur than failure due to material yielding and total collapse.

Most spacecraft have at least one pressurized vessel on board. One of the primary design considerations for earth-orbiting spacecraft is the anticipation and mitigation of the possible damage that might occur in the event of a micrometeoroid or orbital debris (MMOD) particle impact. To prevent mission failure and possibly loss of life, protection against perforation by such high-speed impacts must be included. In addition to a hole, it is possible that, for certain pressure vessel designs, materials, impact parameters, and operating conditions, a pressure vessel may experience catastrophic failure (i.e. rupture) as a result of a hypervelocity impact. If such a tank rupture were to occur on-orbit following an MMOD impact, not only could it lead to loss of spacecraft, but quite possibly, for human missions, it could also result in loss of life. In this paper we present an update to a Rupture Limit Equation, or RLE, for composite overwrapped pressure vessels (COPVs) that was presented previously. The update consists of modified RLE parameters and coefficients that were obtained after the RLE was re-derived using new / additional data. The updated RLE functions in a manner similar to that of a ballistic limit equation, or BLE, that is, it differentiates between regions of operating and impact conditions that, given a tank wall perforation, would result in either tank rupture or only a relatively small hole or crack. This is an important consideration in the design of a COPV pressurized tank – if possible, design parameters and operating conditions should be chosen such that additional sizable debris (such as that which would be created in the event of tank rupture or catastrophic failure) is not created as a result of an on-orbit MMOD particle impact.

The fundamental components of any meteoroid/orbital debris (MOD) risk assessment calculation are environment models, damage response predictor equations, and failure criteria. In the case of a spacecraft operating in low earth orbit, the response predictor equation typically takes the form of a ballistic limit equation (BLE) that defines the threshold particle sizes that cause failure of a spacecraft wall or component. Spacecraft risk assessments often call for BLEs for spacecraft components that do not exist. In such cases, it is a common procedure to use an existing BLE after first equivalencing the actual materials and/or wall thicknesses to the materials that were used in the development of the existing BLE. The question naturally arises regarding how close are the predictions of such an ‘adapted BLE’ to the response characteristics of the actual materials/wall configurations under high speed projectile impacts. This paper presents the results of a study that compared the predictions of a commonly used BLE when adapted to the Soyuz OM wall configuration against those of a new BLE that was developed specifically for that Soyuz wall configuration. It was found that the critical projectile diameters predicted by the new Soyuz OM wall BLE can exceed those predicted by the adapted use of the existing BLE by as much as 50% of the existing BLE values. Thus, using the adapted version of the existing BLE in this particular case would contribute to a more conservative value of assessed risk. If the same trends were to hold true for other spacecraft wall configurations, then it is also possible that using existing BLEs, even after they have been adjusted for differences in materials, etc., may result in predictions of smaller critical diameters (i.e., increased assessed risk) than would using BLEs purposely developed for actual spacecraft configurations of interest.

The wide deployment of wind turbines in locations with high seismic hazard has led engineers to take into account a more comprehensive seismic design of such structures. Turbine specific guidelines usually use simplified methods and consider many assumptions to combine seismic demand with the other operational loads effecting the design of these structures. As the turbines increase in size and capacity, the interaction between seismic loads and aerodynamic loads becomes even more important. In response to the need for a computational tool that can perform coupled simulations of wind and seismic loads, a seismic module is developed for the FAST code and described in this research. This platform allows engineers working in this industry to directly consider interaction between seismic and other environmental loads for turbines. This paper details the practical application and theory of this platform and provides examples for the use of different capabilities. The platform is then used to show the suitable earthquake and operational load combination with the implicit consideration of aerodynamic damping by estimating appropriate load factors.

All earth-orbiting spacecraft are susceptible to impacts by orbital debris particles, which can occur at extremely high speeds and can damage flight- and mission-critical systems. The traditional damage mitigating shield design for this threat consists of a “bumper” that is placed at a relatively small distance away from the main “inner wall” of the spacecraft. The performance of a hypervelocity impact shield is typically characterized by its ballistic limit equation, which is typically drawn as a line of demarcation between regions of rear-wall perforation and no perforation; when graphically represented, it is often referred to as a ballistic limit curve. Once developed, these equations and curves can be used to optimize the design of spacecraft wall parameters so that the resulting shields can withstand a wide variety of high-speed impacts by orbital debris. This paper presents some comments and observations on the development of the three-part ballistic limit equation used to predict the response of dual-wall structural systems under hypervelocity projectile impact. The paper concludes with some insights into the limitations of NASA’s current MMOD risk analysis code, and offers several suggestions regarding how it could be modified so that, for example, it could be used as an integral part of a probabilistic risk assessment exercise.

Horizontal axis wind turbines can experience significant time varying aerodynamic loads that has the potential to cause adverse effects on structural, mechanical, and power production. The progress in the wind industry has caused the construction of wind farms in areas prone to high seismic activity. With the advances in computational tools, a more realistic representation of the behavior of wind turbines should be performed. One of the simulation platforms was developed using the 5 MW NREL utility scale reference turbine model. The performed simulations will be used to evaluate the effects of aerodynamic and seismic load coupling on the power generation and structural dynamics behavior of this structure. Different turbine operational scenarios such as (i) normal operational condition with no earthquake, (ii) idling condition with the presence of seismic loads, (iii) normal operational condition with earthquake, and (iv) earthquake-induced emergency shutdown will be simulated with various loading conditions to show the differences in generated power and dynamic response. The results of this paper provide formulations for calculating generated power and design deriving parameters by considering different intensity measures. Moreover, the effects of aerodynamic damping and pitch control system are presented to shows reduction in the resulting design demand loads.

Most spacecraft have at least one pressurized vessel on board. For robotic spacecraft, it is usually a liquid propellant tank. One of the design considerations of such spacecraft is the possible damage that might occur in the event of an on-orbit impact by a micro-meteoroid or orbital debris (MMOD) particle. While considerable effort has been expended in the study of the response of non-pressurized spacecraft components to these kinds of impacts, relatively few studies have been conducted on the pressurized elements of such spacecraft. This paper presents the results of a study performed to address the following aspects of this problem: how well current ballistic limit equations predict impact-induced perforation damage that might occur in pressurized spacecraft components such as tanks; and, the development of data-driven rupture limit equations that can be used to differentiate between impact conditions that would result in only a small hole or crack, from those that would cause catastrophic tank failure. This information would be useful to a design engineer who might be able to select tank materials to avoid catastrophic tank failure in the event of a perforating on-orbit MMOD particle impact. Tank perforation would typically result in a failed mission; and catastrophic tank failure might generate multiple pieces of new debris, increasing the risk to other spacecraft.

Most spacecraft have at least one pressurized vessel on board. For robotic spacecraft, it is usually a liquid propellant tank. In some satellite or spacecraft designs, the fuel tank or some other pressurized vessel is necessarily exposed to the hazards of space, including the micro-meteoroid and orbital debris (MMOD) environment. Because of the potential of serious mission-threatening damage that might result following an on-orbit MMOD impact, a primary design consideration of such spacecraft is the anticipation and mitigation of the possible damage that might occur in the event of such an impact. While considerable energy and effort has been expended in the study of the response of non-pressurized spacecraft components to MMOD impacts, technical and safety challenges have limited the number of tests that have been conducted on pressurized elements of such spacecraft. As a result, several investigators have taken to modelling pressurized vessels as flat plates under uniaxial or bi-axial tensile loading conditions. This paper presents the development of a data-driven equation for initially stressed flat composite material plates that differentiates between impact conditions which, given a plate perforation, would result in only a hole or crack from those which would cause catastrophic plate failure / rupture. If this equation were subsequently shown to also model the rupture/no-rupture behavior of actual COPVs, then it could also be used to appropriately tailor the design parameters and/or operating conditions of a pressurized tank. Tank designs could be adjusted so that if that tank were to be struck and perforated by the impact of an MMOD particle, then only a hole would occur and additional sizable debris would not be created as a result of that impact.

Most spacecraft have at least one pressurized vessel on board. For robotic spacecraft, it is usually a liquid propellant tank or battery. Manned spacecraft use these elements as well, and there are also pressurized living quarters and life-support systems. One of the design considerations of such spacecraft is the possible damage that might occur in the event of an on-orbit impact by a micro-meteoroid or orbital debris (MMOD) particle. While considerable energy and effort has been expended in the study of the response of non-pressurized spacecraft components to these kinds of impacts, relatively few studies have been conducted on the pressurized elements of such spacecraft. In addition, the design criteria currently used by NASA for pressurized tanks operating in the MMOD environment have not been tested or scrutinized since they were first proposed nearly 50 years ago. In this paper we review current NASA design criteria for pressurized vessels and offer suggestions for next steps in their development. We also review the development of rupture limit equations that can be used to differentiate between impact conditions that would result in only a small hole or crack in a pressurized tank, from those that would cause catastrophic tank failure. This information would be useful to a design engineer who might be able to select tank materials and tailor operating conditions to avoid catastrophic tank failure in the event of a perforating on-orbit MMOD particle impact.

The process used to calculate and reduce the consequences of meteoroid and orbital debris penetrations and their link to catastrophic failure has evolved over time. As the threat of the orbital debris population increased in the 1980s and early 1990s, NASA developed a tool to determine what percentage of space station penetrations might be survivable (referred to as the probability of no catastrophic failure, or PNCF) and how to improve that survivability percentage. The quantity PNCF is directly related to the PNP, or probability of no penetration, as calculated by Bumper , the code used by NASA to perform MOD risk assessments. Part of the process in determining PNCF involves calculating the size of the holes and cracks caused by any penetrations. A review of the original techniques used to calculate hole size and crack length as well as current equations revealed some serious concerns that needed to be addressed. As a result, a study was undertaken to develop revised models that would address these concerns. In this paper, the features of new generic hole- and crack-size prediction equations, as well as the phenomenology involved in the formation of holes and cracks in habitable space station modules are presented and discussed. By comparing the predictions of the new equations against the predictions of current hole and crack size models as well as against empirical data, we found that (1) the predictions of the new equations fit the empirical data just as well as, if not better than, the current models and that (2) the new equations displayed the appropriate phenomenological response characteristics as the diameters of impacting projectiles increased beyond the ballistic limit diameters. Based on the results, we believe that when the new hole and crack size equations are used in survivability assessments the fidelity of the PNCF calculations and predictions will also increase dramatically.

Multiwall shield designs for spacecraft have been studied extensively in the last four decades as a means of reducing the perforation threat of the near-Earth particulate environment over equivalent single-wall structures. The performance of an impact shield is typically characterized by its ballistic limit equation, which is obtained through high-speed impact tests that typically use spherical projectiles fired in light gas guns. Traditional firings employ only one projectile per launch package. However, some facilities have taken to launching multiple projectiles of different sizes in a single launch package. In light of the significant role played by a ballistic limit equation in risk assessment, a study was performed to examine whether or not this method of simultaneously launching several projectiles at a single target could give rise to some issues which may compromise the impact test data obtained in this manner. It was found that interference between debris clouds created by the impacting (simulta...

The process used to calculate and reduce the consequences of meteoroid and orbital debris (MOD) penetrations and their link to catastrophic failure has evolved over time. As the threat of the orbital debris population increased in the 1980s and early 1990s, NASA developed a tool to determine the probability of no catastrophic failure, or PNCF, for the space station and to assess possible changes in station design and/or operations to improve that survivability percentage. PNCF is directly related to the PNP, or probability of no penetration, as calculated by Bumper , the code used by NASA to perform MOD risk assessments. Part of the process in determining PNCF involves calculating the size of the holes and cracks caused by any penetrations. In this paper, the features of new generic hole- and crack-size prediction equations, as well as the phenomenology involved in the formation of holes and cracks in habitable space station modules are presented and discussed. When these new hole and crack size equations are used in survivability assessments, the fidelity of the PNCF calculations and predictions are expected to increase dramatically.

Most spacecraft have at least one pressurized vessel on board. For robotic spacecraft, it is usually a liquid propellant tank or battery. For human spacecraft, there are also pressurized living quarters and life-support systems. One of the design considerations of such spacecraft is the possible damage that might occur in the event of an on-orbit impact by a micrometeoroid or orbital debris (MMOD) particle. While considerable energy and effort has been expended in the study of the response of nonpressurized spacecraft components to these kinds of impacts, relatively few studies have been conducted on the pressurized elements of such spacecraft. In addition, the design criteria currently used by the National Aeronautics and Space Administration (NASA) for pressurized tanks operating in the MMOD environment have not been tested or scrutinized since they were first proposed nearly 45 years ago. This paper reviews current NASA design criteria for pressurized vessels and offers suggestions for next steps in their further development.

One of the primary mission risks tracked in the development of all spacecraft is that due to micro-meteoroids and orbital debris (MMOD). Both types of particles, especially those larger than 0.1 mm in diameter, contain sufficient kinetic energy due to their combined mass and velocities to cause serious damage to crew members and spacecraft. The process used to assess MMOD risk consists of three elements: environment, damage prediction, and damage tolerance. Orbital debris risk assessments for the Orion vehicle, as well as the Shuttle, Space Station and other satellites use ballistic limit equations (BLEs) that have been developed using high speed impact test data and results from numerical simulations that have used spherical projectiles. However, spheres are not expected to be a common shape for orbital debris; rather, orbital debris fragments might be better represented by other regular or irregular solids. In this paper we examine the general construction of NASA’s current orbital debris (OD) model, explore the potential variations in orbital debris mass and shape that are possible when using particle characteristic length to define particle size (instead of assuming spherical particles), and, considering specifically the Orion vehicle, perform an orbital debris risk sensitivity study taking into account variations in particle mass and shape as noted above. While the results of the work performed for this study are preliminary, they do show that continuing to use aluminum spheres in spacecraft risk assessments could result in an over -design of its MMOD protection systems. In such a case, the spacecraft could be heavier than needed, could cost more than needed, and could cost more to put into orbit than needed. The results obtained in this study also show the need to incorporate effects of mass and shape in mission risk assessment prior to first flight of any spacecraft as well as the need to continue to develop/refine BLEs so that they more accurately reflect the shape and material density variations inherent to the actual debris environment.

Spacecraft that are launched to operate in Earth orbit are susceptible to impacts by meteoroids and pieces of orbital debris (MMOD). The effect of a MMOD particle impact on a spacecraft depends on where the impact occurs, the size, composition, and speed of the impacting object, the function of the impacted system. In order to perform a risk analysis for a particular spacecraft under a specific mission profile, it is important to know whether or not the impacting particle (or its remnants) will exit the rear of an impacted spacecraft wall. A variety of different ballistic limit equations (BLEs) have been developed for many different types of structural wall configurations. BLEs can be used to optimize the design of spacecraft wall parameters so that the resulting configuration is able to withstand the anticipated variety of on-orbit high-speed impact scenarios. While the level of effort exerted in studying the response of metallic multi-wall systems to high speed particle impact is quite substantial, the extent of the effort to study composite material and composite structural systems under similar impact conditions has been much more limited. This paper presents an overview of the activities performed to assess the resiliency of composite structures and materials under high speed projectile impact. The activities reviewed will be those that have been aimed at increasing the level of protection afforded to spacecraft operating in the MMOD environment, and more specifically, on those activities performed to mitigate the mechanical and structural effects of an MMOD impact.

Fiber-reinforced polymer composites offer inherent advantages over traditional metallic materials in a number of different ways; however, these materials are also highly susceptible to impact damage. In this paper, we explore the response of FRP (fiber reinforced polymer) composites under impact conditions that could result in their rupture or catastrophic failure. The work performed was aimed at developing a general, data-driven equation for initially-stressed, flat, composite plates that would differentiate between impact conditions that would result in only a hole or crack and those which would cause catastrophic plate failure or rupture. If this equation were to be subsequently shown to also model the rupture/non-rupture behavior of, for example, composite overwrapped pressure vessels, then it could also be used to appropriately tailor the design parameters and/or operating conditions of such pressurized tanks.

The establishment of human habitats on the Moon and on Mars will require protecting them from the hazards of near-Earth and interplanetary space. In addition to solar radiation, another hazard to be faced by these habitats is the damage that can result from the high speed impact of a meteoroid on a critical structural component. Therefore, lunar habitats and their accompanying support facilities need to be designed with adequate levels of protection that will allow them to also withstand the damage that can result from a meteoroid impact. In this paper we discuss some approaches to shielding for lunar habitats, focusing on shielding that is intended primarily to provide protection against meteoroid impacts and on shielding approaches that use resources mined or extracted from the Moon. The Moon’s mineralogy is discussed and suggestions are presented for materials and material combinations that can be used to develop shielding for lunar habitats and which are comprised primarily or entirely of lunar materi...

dc = critical projectile diameter for failure of rear wall, cm dp = projectile diameter, cm KMLI = adjustment factor for multilayer insulation KS2 = adjustment factor for scaling standoff distance S2 in the hypervelocity regime Ktw = adjustment factor for equipment cover plate thickness tw in the hypervelocity regime K3D, K3S = ESA triple-wall ballistic limit equation fit factors for the hyperballistic and ballistic velocity regimes, respectively S1, S2 = standoff between first and second, and second and third bumper, where third bumper backwall, cm tb = thickness of the inner/second bumper (the rear face sheet in the case of a honeycomb sandwich panel), cm tob = thickness of the outer bumper (the front face sheet in the case of a honeycomb sandwich panel), cm tw = thickness of third bumper/back wall/equipment cover plate, cm V, Vn = impact velocity and its normal component (V cos ), respectively, km=s Vt1 = transition velocity between ballistic and shatter velocity regimes, km=s Vt2 = transition velocity between shatter and hypervelocity regimes, km=s Vt1;n, Vt2;n = normal component of Vt1 cos and Vt2 cos , respectively, km=s , , , , " = fit parameters for the Schafer–Ryan–Lambert ballistic limit equation = impact angle (0 deg is a perpendicular impact on the target surface), deg ob, p = volumetric density of the outer bumper and the projectile, respectively, g=cm y;ksi = yield stress of the equipment cover plate, ksi

All Earth-orbiting spacecraft are susceptible to damage that can be caused by high-speed impacts with pieces of man-made debris or naturally-occurring meteoroids, and spacecraft at locations other than near Earth are subject to similar naturally-occurring hazards. Traditional protective shield design consists of a “bumper” that is placed at a relatively small distance away from the main “inner wall” of the spacecraft component, the performance of which is typically characterized by its ballistic limit equation (BLE). This paper addresses the question of how well the NASA/JSC dual-wall BLE performs when it is used to predict inner wall response in applications other than those used for its development. The major conclusions reached as a result of the analyses performed are that (1) to be truly conservative the critical projectile diameter value as calculated by the NASA/JSC dual-wall BLE needs to be multiplied by 0.75 to accommodate results from other test databases, (2) the NASA/JSC dual-wall BLE is not as conservative for impact obliquities exceeding 60° as it is for obliquities of 45° or less, and (3) the NASA/JSC dual-wall BLE is not as conservative for impact tests with MLI between the bumper and inner wall as it is for tests without the MLI.

All earth-orbiting spacecraft are susceptible to impacts by these particles, which can occur at extremely high speeds and can damage flight- and mission-critical systems. The traditional damage mitigating shield design for this threat consists of a “bumper” that is placed several cm away from the main “inner wall” of the spacecraft. Typical orbital debris risk analyses that include ballistic limit equations (BLEs) and curves (BLCs) assume that orbital debris particles are spherical in shape. However, spheres are not a common shape for orbital debris; rather, debris fragments might be better represented by other regular or irregular solids. This paper presents the results of a study comparing BLCs developed by NASA and the DoD for velocities up to 4 km/s considering spheres, cubes, and a “flake” shape that was proposed within NASA's Standard Breakup Model to represent orbital debris. It also compares performance of these shapes using hydrocodes at higher velocities (7–12 km/s), and generates a combined BLC for these shapes for the entire orbital debris velocity regime. In addition to shape, a multi-view method is used to examine the effects of a variety of cube and flake impact orientations on BLC, as well as a “characteristic length” parameter developed by NASA to compare the particle shapes on the basis of their radar cross section. The developed non-spherical BLCs are then evaluated for overall penetration risk considering the orbital debris environment. Their predictions of risk are compared to that predicted using sphere-based BLCs. This methodology is then extended to a single-wall shield design for velocities up to 15 km/sec, and the results of DoD predictions for a sphere and cube are compared with NASA predictions for a sphere having the same characteristic length. The results indicate that we may be over-predicting orbital debris risk for dual-wall shields by a factor of two—and for single walls by a factor of four—by limiting our analyses to spheres instead of using more representative debris shapes, such as cubes and flakes, and its characteristic length as the primary particle parameter.

Ballistic limit curves (BLCs) for dual- or multi-wall spacecraft wall systems impacted by spherical projectiles exist for a variety of impact conditions. However, meteoroids and orbital debris particles can take any shape, and non-spherical projectiles can be more damaging than equal mass spherical projectiles under the same impact conditions. In a recent study, a series of BLCs for a typical dual-wall configuration impacted by a variety of non-spherical projectiles were developed and presented in terms of equivalent spherical projectile diameter as a function of impact velocity. In order to be more consistent with debris environment predictions and to be useful for spacecraft design, such BLCs need to be drawn in units that are consistent with environment flux models used in that design process. In this paper, we have recast the equivalent spherical projectile diameter BLCs and now present them using radar cross-section (RCS) diameter as the characteristic length parameter. The recast BLCs are seen to be more tightly grouped and are not as spread out as they were when plotted using equivalent spherical projectile diameter. Thus, the effect of projectile shape on penetrating ability is shown to be somewhat reduced when considering characteristic length as the size parameter as opposed to an equivalent mass diameter.

Many military and civilian aircraft are expected to operate beyond their designed life expectancy. As a result, the development of reliable and cost-effective repair techniques for deteriorating components in aging aircraft is of great interest. The application of bonded composite material patches to restore cracked and corroded metallic airframes has been shown to be an effective repair technique for increasing the durability and damage tolerance of the repaired structure. Composite Repair of Aircraft Structures (CRAS) is a computer program that is currently being used to design bonded patches for damaged metallic structures. The results are presented for a study whose objective was to extend the applicability of the CRAS program to include bonded patch design for aircraft structures that have sustained adjacent cracking and corrosion damage. Patch designs were developed for three different cracking and corrosion configurations. Composite patches were manufactured according to the designs developed and applied to damaged aluminum panels using a hot-bonding device. Patched and unpatched panels were loaded to failure using a material testing system (MTS) 880 machine. In all cases the failure loads of the patched panels exceeded those of corresponding unpatched panels. These results validated the patch design procedure developed herein for the configurations considered.

Hypervelocity impact testing of elevated and room temperature bumpers was performed to establish whether or not bumper hole diameter is temperature dependent. The hole diameters in the heated bumper tests were found to besignificantly larger than those of the room temperature tests. The heated and room temperature bumper data were compared to the Housen-Schmidt, Maiden, and Nysmith bumper hole diameter models to determine how well the models predicted the data. The Housen-Schmidt model correlated well with the room temperature bumper data, but did not correlate well with the heated data sets. The Maiden and Nysmith models did not appear to correlate with any of the data sets. Empirical models were formulated to predict hole sizes from 2 to 7 km/s for heated bumpers at the various temperatures considered. Good correlation was found between the models developed and the test data. The test results obtained and the models developed suggest that bumper hole diameter is indeed temperature dependent.

Ballistic limit curves for dual-wall spacecraft systems are typically developed using ground-based hypervelocity impact tests with spherical projectiles. However, most orbital debris particles are not spherical and can take any shape. Non-spherical projectiles can also be more damaging than equal mass spherical projectiles. Although some recent efforts have examined the effects of non-spherical projectile impacts, ballistic limit curves for non-spherical projectiles are still not available. This paper presents the results of a study in which numerical simulation methods were used to develop ballistic limit curves for a variety of non-spherical projectile shapes impacting a typical dual-wall spacecraft configuration. Ballistic limit curves for non-spherical projectiles are presented and compared against that for spherical projectiles impacting the same wall system. It is found that long cylinder and long cone projectiles have a higher penetrating capability than all other projectile shapes considered, including spherical projectiles. The implication is that if a spacecraft wall system is designed using spherical projectile-based ballistic limit curves, the design will likely be non-conservative.

The problem of pollution within Earth’s orbital environment has gained considerable recognition over the past decades. Determining adequate passive protection schemes is an unending process that attempts to meet different objectives for widely varying types of missions. Significant amounts of resources have been expended toward development of numerical and analytical models that model the response of a variety of target systems under high-speed orbital debris impacts. The objective of the study whose results are presented herein was to improve upon an existing oblique hypervelocity impact model that characterizes the various secondary debris clouds created in such an impact. This was accomplished by reducing the model’s dependence on empirical user-defined parameters and by correcting an error in one of its equations. Predictions of the improved model are compared with numerical simulations generated during previous impact studies under comparable conditions. It is found that the improved model does a reasonable job of predicting the characteristics of the secondary debris clouds created in an oblique hypervelocity impact.

Microdebris particles generated by hypervelocity impacts between complex projectiles and ballistic missile targets can affect sensitive components of such aerospace systems. Current models assume an isotropic dispersion of this microdebris about the postimpact debris cloud center of mass. Data collected from recent tests indicate that a nonisotropic microdebris dispersion model is needed. The results are presented of a study aimed at improving the capabilities of an existing model that characterizes the microdebris environment created in a high-speed impact. Following a review of test data collection techniques, a nonisotropic microdebris dispersion model is developed as a series of modifications to the existing microdebris size distribution model. Relevant test data are then used to evaluate the proposed nonisotropic model. It is found that the modified model is more correct in accounting for the microdebris fragment population created in a high-speed impact than the current isotropic dispersion model.

The static and dynamic indentation of structural elements such as beams and plates continue to be intriguing problems, especially for scenarios where large area contacts are expected to occur. Standard methods of indentation analyses use a beam theory solution to obtain an overall load–displacement relationship and then a Hertzian contact solution to calculate local stresses under the indenter. However, these techniques are only applicable in a fairly limited class of problems: the stress distribution in the contact region will differ significantly from a Hertzian one when the contact length exceeds the thickness of the beam. The indentation models developed herein are improvements over existing GLOBAL/LOCAL models for static and dynamic indentation of cantilever beams. Maximum contact stresses, beam displacements, and contact force time histories are obtained and compared with the predictions of current static and dynamic indentation models. The validity of the solutions presented herein is further assessed by comparing the results obtained to the predictions of modified beam theory solutions.

The orientation of a single-crystal material is known to affect the strength and life of structural component parts. Results are presented of an investigation of the effects of secondary axis orientation angles on the failure of the first-stage of the space shuttle main engine alternate turbopump development of the high-pressure fuel turbopump. First, the correlation of different failure models with low-cycle fatigue data for nickel-base single-crystal test specimens was analyzed. Then the models with the highest correlation coefficients were used to study the actual single-crystal blade structure. Based on the results obtained, a new failure model was proposed. A detailed finite element model for the first-stage blade was used to calculate the stresses and strains at all blade nodes for different material orientations. Results of the analysis showed that the critical value of the failure model could vary by up to a factor of 3 by changing the primary and secondary material orientations. A comparison between analytical results and engine test results showed good correlation and also demonstrated the dependence of cracking location on crystal orientation.

All spacecraft in low earth orbit are subject to high-speed impacts by meteoroids and orbital debris particles. These impacts can damage flight-critical systems, which can in turn lead to catastrophic failure of the spacecraft. In addition to threatening the operation of the spacecraft itself, on-orbit impacts also generate a significant amount of damaging ricochet ejecta particles. These high-speed particles can destroy critical external spacecraft subsystems, which also poses a threat to the spacecraft and its inhabitants. Since the majority of on-orbit debris impacts are expected to occur at oblique angles, the characterization of ricochet debris created in an orbital debris particle impact is an issue that must be addressed. This paper presents the results of a study performed to develop an empirical model that characterizes the secondary ejecta created by a high speed impact on a typical aerospace structural surface. Specifically, the model predicts the spread and trajectory of ricochet debris particles created in a hypervelocity impact as well as the size an velocity of the most damage particle in the ricochet debris cloud. Results obtained using the model are compared with experimental results and predictions obtained in a previous study.

All earth-orbiting spacecrafts are susceptible to impacts by meteoroids and pieces of orbital debris. These impacts occur at extremely high speeds and can damage flight-critical systems, which can in turn lead to catastrophic spacecraft failure. With the advent of many new high-strength composite materials and their proliferation in aircraft applications, it is necessary to evaluate their potential for use in long-duration space and aerospace structural systems. This paper presents the results of several experimental investigations in which several different composite materials were tested for their ability to prevent the perforation of the inner walls in multi-wall systems subjected to hypervelocity projectile impact. Damage to multi-wall systems that employ composite materials is compared to damage in traditional all-aluminum multi-wall structures of similar geometry and weight under similar impact conditions. The advantages and disadvantages of using composite materials in perforation-resistant multi-wall structures are presented and discussed. The analyses performed show that the effectiveness of multi-wall systems with composite materials varies significantly depending upon the location within the multi-wall system of the composite materials. Based on the results obtained, it appears that composite materials are most effective when used as the middle wall of a triple-wall system.

Standard methods of beam indentation analysis use a beam theory solution to obtain the load-displacement relationship and a Hertz solution to calculate local stresses. However, when the contact length exceeds the thickness of the beam point contact can no longer be assumed and Hertzian relations are no longer valid. This paper presents an improved superposition solution technique that uses a true elasticity solution to obtain the load-displacement relationship in non-Hertzian indentation problems.

Summary During their missions in space, spacecraft are subjected to high velocity impacts by orbital debris particles. Such impacts are expected to occur at non-normal incidence angles and can cause severe damage to the spacecraft as well as to its external flight-critical components. In order to ensure crew safety as well as the proper function of internal and external spacecraft systems, the characteristics of debris clouds generated by such impacts must be known. In this paper, a first-principles-based analytical model is developed for the characterization of the penetration and ricochet debris clouds created by the hypervelocity impact of a spherical projectile on a thin aluminum plate. This model employs normal and oblique shock wave theory to characterize the penetration and ricochet processes. The predictions of the model are compared against numerical and experimental results.