NASA Standard Breakup Model Research Articles

Fragmentation characterization in the circular restricted three body problem for cislunar space domain awareness

With heightened international interest in spacecraft activities in the vicinity of the Moon, cislunar space debris is likely to follow. Even one fragmentation event can have catastrophic and far-reaching consequences, which drives the need for appropriate debris characterization tools. How a single fragmentation plays out is highly dependent on any given initial condition in the near-chaotic cislunar region. This paper offers a means of structuring the cislunar region in terms of dynamical flow, which enables global characterization of fragmentation events without propagation of every possible case. This work investigates patterns in fragment behaviour as a function of energy, Δv, and orbit location, and explores emergent dynamic structures in the vicinity of the Earth-Moon L2 Lagrange point. Subsequent findings are applied to analysis of a realistic breakup event for a 500 kg satellite on an L2 Lyapunov orbit with a Jacobi constant of 3.0165, modeled using an in–house modified version of the NASA Standard Breakup Model.

Analysis of accidental spacecraft break-up events in cislunar space

Fragmentation events are the most common source of artificial debris in space. Such events can happen for multiple reasons, ranging from collisions with active or inactive objects to propellant tanks or battery explosions. They are an inherent part of human activities in space, regardless of the location. At the same time, the interest in the Moon and the space around it increases. Multiple planned missions want to leverage the peculiar dynamics of the Earth-Moon system to build an infrastructure likely to support human exploration of space for centuries to come. The present research puts a step forward in the direction of Space Debris Mitigation techniques for future cislunar missions. After being tuned with novel implementation techniques, the NASA Standard Break-up Model has been employed to simulate explosion events in multiple locations in cislunar space. Notably, the Circular Restricted Three-Body Problem model has been employed to generate orbits, corrected in a higher-fidelity model to generate initial conditions for long-term propagations. Moreover, multiple starting positions along each orbit have been crossed with several possible event dates, providing a broader understanding of the aftermath of explosion events. Three case studies have been considered: an EML2 southern Near Rectilinear Halo Orbit (NRHO), candidate for the Gateway station, a larger Halo orbit of the same family, and a Distant Retrograde Orbit (DRO). These locations have been chosen to characterize dynamically different trajectories used in current applications. Results provided include the number, locations, and characteristics of collision with the Moon and the Earth, together with considerations on the interactions with Earth’s protected orbital regions of Low Earth Orbit (LEO) and Geostationary Orbit (GEO). A novel analysis is also introduced to determine the immediate danger for the Gateway station in the case of such an event. This research lays the foundation to develop a cislunar Space Debris Mitigation (SDM) framework in the hopes that awareness is raised when there is still time for proper implementation of guidelines.

Analysis of fragments larger than 2 mm generated by a picosatellite fragmentation experiment

Satellite breakup models rely on laboratory tests and in-space collision observations; current models can match fragments distributions generated by traditional satellites but may need to be improved for small spacecraft and modern satellites employing new configurations and materials. In the last years, ground tests have been employed to assess the influence of dimensions, materials and internal configurations on fragments distributions and to define the limits of the current models. In this context, an impact test was performed at the impact facility of the University of Padova to characterize the fragmentation of a picosatellite mock-up; more than 7000 fragments were collected, classified and analysed with automatic image recognition algorithms. It was observed that the experimental characteristic length distribution is line with the prediction of the NASA SBM even for the smallest size classes, while the fragments shape distribution is strongly affected by the materials employed in the picosatellite manufacturing.The subset of the collected fragments larger than 2 mm was recently subjected to a more detailed analysis: each fragment was individually weighed, and its three main dimensions were measured. In this paper, resulting fragments distributions are compared with literature data and the NASA Standard Breakup Model; in particular, an analytic relation between fragments characteristic length and size is found. In addition, results show that characteristic length and area-to-mass distributions are affected by the target materials and are clearly influenced by the size resolution of the analysed fragments.

Characterization of the fragments generated by a Picosatellite impact experiment

Satellites breakup models are based on observations of in-space events and on impact tests performed in ground facilities. In the last years the scientific community showed a growing interest in updating such models to include the influence of modern spacecraft design and divers materials in satellites breakup. In this context, an impact test was performed at the impact facility of the University of Padova to characterize the fragmentation of a picosatellite mock-up. The kinetic energy to the involved mass ratio of the impact was of about 80 kJ/kg, well above the commonly considered catastrophic threshold of 40 kJ/kg, leading to a complete fragmentation of the impacted body. The resulting fragments were collected and characterized in terms of their size, shape and area-to-mass ratio.In this paper the impact experiment and the fragments characterization are presented; results are compared with the NASA Standard Breakup Model. It is observed that the obtained characteristic length distribution is line with the prediction of the NASA SBM even for the smallest size classes, while the area-to-mass distribution is strongly affected by the materials employed in the picosatellite manufacturing.

Risk Assessment of a Large Constellation of Satellites in Low-Earth Orbit

This paper aims to assess the risk of mission termination for a large constellation of satellites in a low-Earth orbit. Many large constellations will be deployed to provide broadband network services using thousands of satellites. There is concern that such large constellations will have a serious impact on the long-term sustainability of outer space activities due to the rapid increase in population. First, therefore, the authors conducted an assessment under nominal activities (referred to as “business-as-usual”) on the basis of a prediction by ESA’s MASTER-2009 and NASA standard breakup model 2001 revision. The assessment found that nearly one catastrophic collision may happen in a large constellation, generating more than two million fragments as small as 1 mm in size. Second, the authors conducted a further assessment assuming a hypothetical collision of a satellite in a large constellation using the NASA standard breakup model and a spherical finite element model adopted in ESA’s MASTER-2009. In consequence, another catastrophic collision may happen to a large constellation, generating approximately a half-million fragments as small as 1 mm in size. Therefore, such catastrophic collisions and resulting secondary collisions should be prevented for large constellations.

Numerical simulations of hypervelocity collisions scenarios against a large satellite

This paper describes a set of numerical simulations of hypervelocity collisions involving the ESA LOFT spacecraft, with CubeSat impactors ranging from 1 U to 48 U and diverse collision scenarios. The study was done with the purpose to investigate (1) the transition between sub-catastrophic and catastrophic collision with increasing impactor's kinetic energy, and (2) the effect of impact point and encounter configuration on the collision severity.The simulations were performed with the new semi-empirical Collision Simulation Tool developed by CISAS-University of Padova (prime contractor) and etamax GmbH in the framework of an ESA contract recently completed. The CST makes possible to model a large variety of collision scenarios with spacecraft design details included, and is capable to provide statistically accurate impact fragments distributions with low computational expense. It is shown that (1) the ratio between the kinetic energy of the impactor and the mass of the target (Energy-to-Mass Ratio, EMR) is not sufficient to discriminate between sub-catastrophic and catastrophic impacts, and (2) the damage suffered by the target and the number of collision fragments are limited when impacts are not central and the ratio between impactor size and target size is small. In particular, results suggest that the NASA Standard Breakup Model overpredicts the fragments for glancing impacts and/or small impactors, showing that both the impact point and the size of the impactor at a given energy play a significant role to determine the severity of the event.

Tuning of NASA Standard Breakup Model for Fragmentation Events Modelling

The continuous growth of space debris motivates the development and the improvement of tools that support the monitoring of a more and more congested space environment. Satellite breakup models play a key role to predict and analyze orbital debris evolution, and the NASA Standard Breakup Model represents a widely used reference, with current activities relevant to its evolution and improvements especially towards fragmentation of small mass spacecraft. From an operational perspective, an important point for fragmentation modelling concerns the tuning of the breakup model to achieve consistency with orbital data of observed fragments. In this framework, this paper proposes an iterative approach to estimate the model inputs, and in particular, the parents’ masses involved in a collision event. The iterative logic exploits the knowledge of Two Line Elements (TLE) of the fragments at some time after the event to adjust the input parameters of the breakup model with the objective of obtaining the same number of real fragments within a certain tolerance. Atmospheric re-entry is accounted for. As a result, the breakup model outputs a set of fragments whose statistical distribution, in terms of number and size, is consistent with the catalogued ones. The iterative approach is demonstrated for two different scenarios (i.e., catastrophic collision and non-catastrophic collision) using numerical simulations. Then, it is also applied to a real collision event.

Fragmentation model and strewn field estimation for meteoroids entry

Everyday thousands of meteoroids enter the Earth's atmosphere. The vast majority burn up harmlessly during the descent, but the larger objects survive, occasionally experiencing intense fragmentation events, and reach the ground. These events can pose a threat for a village or a small city; therefore, models of asteroid fragmentation, along with accurate post-breakup trajectory and strewn field estimation, are needed to enable a reliable risk assessment. In this work, a methodology to describe meteoroids entry, fragmentation, descent, and strewn field is presented by means of a continuum approach. At breakup, a modified version of the NASA Standard Breakup Model is used to generate the distribution of the fragments in terms of their area-to-mass ratio and ejection velocity. This distribution, combined with the meteoroid state, is directly propagated using the continuity equation coupled with the non-linear entry dynamics. At each time step, the probability density evolution of the fragments is reconstructed using GMM interpolation. Using this information is then possible to estimate the meteoroid's ground impact probability. This approach departs from the current state-of-the-art models: it has the flexibility to include large fragmentation events while maintaining a continuum formulation for a better physical representation of the phenomenon. The methodology is also characterised by a modular structure, so that updated asteroids fragmentation models can be readily integrated into the framework, allowing a continuously improving prediction of re-entry and fragmentation events. The propagation of the fragments' density and its reconstruction, currently considering only one fragmentation point, is first compared against Monte Carlo simulations, and then against real observations. Both deceleration due to atmospheric drag and ablation due to aerothermodynamics effects have been considered.

Transformation of Satellite Breakup Distribution for Probabilistic Orbital Collision Hazard Analysis

Fragmentation clouds from explosions or collision of payloads and rocket bodies in space pose a threat to objects in Earth orbit. Most of the fragments are too small to be tracked and can only be accounted for statistically. Here, a framework for the fully statistical treatment of a fragmentation cloud, its evolution and ramifications, without the need of simplifying assumptions, is presented. The cloud is modeled as an uncertainty around a single fragment, which can be propagated using any of the existing, nondeterministic, nonlinear orbital uncertainty propagation methods. This work is focused on providing the initial distribution and the estimation of the statistical collision probability. The NASA standard breakup model is revisited to derive a probability distribution of the initial fragment cloud. Two density transformation methods are discussed to obtain the distribution in a subset of orbital elements, suitable for mid- to long-term evolution. The fragment spatial density and the impact rates on targets in any orbit are obtained. The method is applied to show the fragment cloud distribution of a payload collision in low Earth orbit (LEO). Its collision probability with a satellite in LEO and a rocket body in the geostationary transfer orbit are estimated. The result is compared against, and shows the limitations of, sampling and methods based on finite differences.

NewSpace and its implications for space debris models

Until two decades ago, the dominance of the space industry by national governments shaped the key characteristics of the spacecraft population and hence the space debris models used to anticipate future orbital populations. The rise of `NewSpace’ with the growth of private sector involvement has brought innovations, disrupting the status quo and changing to the physical characteristics and mission orbits of commercial spacecraft. Analysis of shifts in mass and launch traffic for spacecraft launched between 1980 and 2019 (source: ESA's DISCOS database) suggests significant impacts for debris modelling. Results emphasised the ongoing change towards a more commercially focused space sector. 742 spacecraft were launched in the 1980s, of which 34 were labelled as commercial. By contrast, there were 1292 commercial missions out of 2325 launched in the 2010s - an increase from 4.6% to 55.6%. These increases correlate with a rise in the number of different organisations operating spacecraft and a clear trend can be seen towards smaller, lower mass spacecraft. This is likely to alter the distributions of fragments generated in collisions compared with the distributions obtained from empirical methods, such as the NASA Standard Breakup model, derived from the fragmentation of larger spacecraft in the 600-1,000 kg range. A study of observed breakup events indicates that the NASA Standard Breakup Model over-estimates the number of large debris released during fragmentations due to collisions or explosion of satellites while also under-estimating the number of small debris. It is believed that this will have a significant impact on the outcomes of simulations of the future debris environment where these fragmentation modes are expected to dominate the generation of new debris over the propulsion based explosions which have been historically prevalent.

Longitude-dependent effects of fragmentation events in the geosynchronous orbit regime

The effects of on-orbit fragmentation events on localized debris congestion in each of the longitude slots of the geosynchronous orbit (GEO) regime are evaluated by simulating explosions and collisions of uncontrolled rocket bodies in multiple orbit configurations, including libration about one or both of the gravitational wells located at 75°E and 105°W. Fragmentation distributions are generated with the NASA Standard Breakup Model, which samples fragment area-to-mass ratio and delta-velocity as a function of effective diameter. Simulation results indicate that the long-term severity and consequence of a GEO fragmentation event is strongly dependent upon parent body longitude at the epoch of fragmentation, which can spawn bi-annual “fragment storms” in high-risk longitude slots, driven by lower-energy fragments that have been captured and have started librating around the nearby gravitational well.

Finite element reconstruction approach for on-orbit spacecraft breakup dynamics simulation and fragment analysis

Breakup model is the key area of space debris environment modeling. NASA standard breakup model is currently the most widely used for general-purpose. It is a statistical model found based on space surveillance data and a few ground-based test data. NASA model takes the mass, impact velocity magnitude for input and provides the fragment size, area-to-mass ratio, velocity magnitude distributions for output. A more precise approach for spacecraft disintegration fragment analysis is presented in this paper. This approach is based on hypervelocity impact dynamics and takes the shape, material, internal structure and impact location etc. of spacecraft and impactor, which might greatly affect the fragment distribution, into consideration. The approach is a combination of finite element and particle methods, entitled finite element reconstruction (FER). By reconstructing elements from the particle debris cloud, reliable individual fragments are identified. Fragment distribution is generated with undirected graph conversion and connected component analysis. Ground-based test from literature is introduced for verification. In the simulation satellite targets and impactors are modeled in detail including the shape, material, internal structure and so on. FER output includes the total number of fragments and the mass, size and velocity vector of each fragment. The reported fragment distribution of FER shows good agreement with the test, and has good accuracy for small fragments.

Theoretical and empirical analysis of the average cross-sectional areas of breakup fragments

This paper compares two different approaches to calculate the average cross-sectional area of breakup fragments. The first one is described in the NASA standard breakup model 1998 revision. This approach visually classifies fragments into several shapes, and then applies formulae developed for each shape to calculate the average cross-sectional area. The second approach was developed jointly by the Kyushu University and the NASA Orbital Debris Program Office. This new approach automatically classifies fragments into plate- or irregular-shapes based on their aspect ratio and thickness, and then applies formulae developed for each shape to calculate the average cross-sectional area. The comparison between the two approaches is demonstrated in the area-to-mass ratio (A/m) distribution of fragments from two microsatellite impact experiments completed in early 2008. A major difference between the two approaches comes from the calculation of the average cross-sectional area of plates. In order to determine which of the two approaches provides a better description of the actual A/m distribution of breakup fragments, a theoretical analysis in the calculation of the average cross-sectional area of an ideal plate is conducted. This paper also investigates the average cross-sectional area of multi-layer insulation fragments. The average cross-sectional area of 214 multi-layer insulation fragments was measured by a planimeter, and then the data were used to benchmark the average cross-sectional areas estimated by the two approaches. The uncertainty in the calculation of the average cross-sectional area with the two approaches is also discussed in terms of size and thickness.

Microsatellite Impact Tests to Investigate the Outcome of Satellite Fragmentations

To predict the future orbital environments, it is necessary to know outcome of the satellite fragmentation. The NASA standard breakup model is designed to describe the outcome of typical satellite fragmentation. This model is an empirical model and themajor data sources are the 1980s on-orbit satellite breakup events and the ground-based Satellite Orbit Debris Characterization Impact Test series conducted in early 1990s. The target cubic satellites ranged from 15 to 20 cm in size and about 1000 g in mass. Results from all seven impact tests carried out in 2008 are shown in this paper and compared with the NASA standard breakupmodel to demonstrate potential improvements of the model in the future.

Faster algorithm of debris cloud orbital character from spacecraft collision breakup

Space debris is polluting the space environment. Collision fragment is its important source. NASA standard breakup model, including size distributions, area-to-mass distributions, and delta velocity distributions, is a statistic experimental model used widely. The general algorithm based on the model is introduced. But this algorithm is difficult when debris quantity is more than hundreds or thousands. So a new faster algorithm for calculating debris cloud orbital lifetime and character from spacecraft collision breakup is presented first. For validating the faster algorithm, USA 193 satellite breakup event is simulated and compared with general algorithm. Contrast result indicates that calculation speed and efficiency of faster algorithm is very good. When debris size is in 0.01–0.05 m, the faster algorithm is almost a hundred times faster than general algorithm. And at the same time, its calculation precision is held well. The difference between corresponding orbital debris ratios from two algorithms is less than 1% generally.

Analysis and Calculation on Collision Breakup Characteristics of Orbital Spacecraft

The collision breakup problem of orbital spacecraft is studied in this paper.The collision breakup algorithm of orbital spacecraft is presented based on NASA standard breakup model. Constraint conditions required for the breakup calculation are analyzed emphatically.Finally,the general utility simulation calculation program is implemented,and is verified to be correct by comparing with measured data of P-78 satellite breakup event.The research can be an important reference to predict and analyze damage characteristics of orbital collision events.

A comparison of NASA, DoD, and hydrocode ballistic limit predictions for spherical and non-spherical shapes versus dual- and single-wall targets, and their effects on orbital debris penetration risk

Abstract 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.

Investigation and comparison between new satellite impact test results and NASA standard breakup model

This paper summarizes two new satellite impact tests conducted in order to investigate on the outcome of low- and hyper-velocity impacts on two identical target satellites. The first experiment was performed at a low velocity of 1.5 km/s using a 40-gram aluminum alloy sphere, whereas the second experiment was performed at a hyper-velocity of 4.4 km/s using a 4-gram aluminum alloy sphere by two-stage light gas gun in Kyushu Institute of Technology. To date, approximately 1,500 fragments from each impact test have been collected for detailed analysis. Each piece was analyzed based on the method used in the NASA Standard Breakup Model 2000 revision. The detailed analysis will conclude: 1) the similarity in mass distribution of fragments between low and hyper-velocity impacts encourages the development of a general-purpose distribution model applicable for a wide impact velocity range, and 2) the difference in area-to-mass ratio distribution between the impact experiments and the NASA standard breakup model suggests to describe the area-to-mass ratio by a bi-normal distribution.

SOCIT4 collisional-breakup test data analysis: With shape and materials characterization

In this paper we revisit the 1995 Kaman database of the SOCIT4 fragment characteristics with added analysis of a subset of the cataloged fragments from the test. This database was compiled from the last of a series of four hypervelocity impact tests conducted under a U.S. Department of Defense program in 1991–1992. This test targeted a flight-ready, U.S. Transit navigation satellite, yielding collision fragments in the size regime of sub-millimeter through tens of centimeters. Results in this database were used in the 1998 NASA Standard Breakup Model to represent characteristic length (size) and area-to-mass distributions of fragments smaller than 10 cm. In this analysis we explore, in detail, the tabulated fragment material and shape. What emerges is a clear distinction in fragment area-to-mass ratio between primarily metal and primarily non-metal fragments. Metal fragments, which are dominated by aluminum, follow the characteristic curve of increasing area-to-mass ratio with decreasing characteristic length: objects move from the character of large irregular bent shards to that of small solid spheroids. Non-metal fragments, dominated by phenolic/plastic, do not appear to move towards solid spheroids as easily as metals. Also unlike the metals, their area-to-mass ratio curve plateaus in the midsize region (smaller than 1 cm), coinciding with a peak in plate-like, non-metal fragments. The internal structure of the Transit payload, with its phenolic surface skin and packed arrays of plastic circuit boards, certainly governs this plateau behavior. In the small fragment regime (smaller than 4 mm) phenolic/plastic ellipsoidal ‘Nuggets’ dominate the population. They outnumber aluminum ‘Nuggets’ by over four to one. Through this study we gain a more detailed understanding of the collisional-breakup process of this particular payload, but also have begun to determine how these data may apply to other breakups. Our long-term goal is to apply this new understanding to future upgrades of the NASA Standard Breakup Model.

Comparison of fragments created by low- and hyper-velocity impacts

This paper summarizes two new satellite impact experiments. The objective of the experiments was to investigate the outcome of low- and hyper-velocity impacts on two identical target satellites. The first experiment was performed at a low-velocity of 1.5 km/s using a 40-g aluminum alloy sphere. The second experiment was performed at a hyper-velocity of 4.4 km/s using a 4-g aluminum alloy sphere. The target satellites were 15 cm × 15 cm × 15 cm in size and 800 g in mass. The ratios of impact energy to target mass for the two experiments were approximately the same. The target satellites were completely fragmented in both experiments, although there were some differences in the characteristics of the fragments. The projectile of the low-velocity impact experiment was partially fragmented while the projectile of the hyper-velocity impact experiment was completely fragmented beyond recognition. To date, approximately 1500 fragments from each impact experiment have been collected for detailed analysis. Each piece has been weighed, measured, and analyzed based on the analytic method used in the NASA Standard Breakup Model (2000 revision). These fragments account for about 95% of the target mass for both impact experiments. Preliminary analysis results will be presented in this paper.