Satellite Breakups Research Articles

From perturbative methods to machine learning techniques in space science

AbstractPerturbation theory is a very useful tool to investigate the dynamics of models in space science. We start by presenting some results obtained implementing classical perturbation theory to investigate the motion of space debris, which are objects that populate the sky around the Earth after a satellite break-up event. When dealing with two or more break-up events, a clusterization of the fragments can be computed using machine learning techniques. We also present the celebrated KAM theory for symplectic and conformally symplectic systems. We recall several computer-assisted results in Celestial Mechanics in conservative and dissipative settings. Finally, we consider the spin-orbit problem and we show how machine learning methods can be conveniently used to classify regular and chaotic motions.

Modeling Medium-Term Debris Cloud of Satellite Breakup via Probabilistic Method

Accurately modeling the evolution of debris clouds from satellite breakups is necessary for assessing the collision risk to space assets. Generally, the evolution of debris clouds can be divided into short-, medium-, and long-term phases. Compared to the short- and long-term phases, the study on the medium-term phase was insufficient. This paper proposes a probabilistic method to model the medium-term evolution of a debris cloud. First, a satellite breakup event can be characterized by a probabilistic density distribution of the initial velocity vectors of all fragments. Then, the distribution in the initial velocity can be transformed to that in three classic orbital elements, i.e., semimajor axis, eccentricity, and inclination. Given that the driving force of medium-term evolution is the J2 perturbation and that J2 has no secular contribution to these three classic orbital elements, the distribution of classic orbital elements remains stationary during the evolution. This feature can be exploited to express the probabilistic density distribution of any given target position in terms of the density distribution of the classic orbital elements by changing variables. Finally, 16 one-variable equations with analytical first derivatives were formulated to find the proper inverse mapping from the given position to the classic orbital elements, and the Jacobian matrix of the inverse mapping was derived. Medium-term debris clouds resulting from two typical breakup velocity distributions were tested using the proposed method. Results show that the proposed method can accurately simulate the evolution of the clouds’ geometrical extent and density flow.

Uncontrolled reentries of space objects and aviation safety

There is a growing risk to aircraft in flight from collisions with debris produced by the breakup of uncontrolled rocket bodies and satellites during atmospheric reentry. On one hand, the aviation industry has grown to almost 39 million flights per year, with a 63 % increase between 2004 and 2019. On the other hand, the number of successful rocket launches more than doubled between 2015 and 2023, from 87 to 212, respectively. During those 212 launches in 2023, 128 rocket bodies were abandoned in orbit, left to reenter uncontrollably. Such uncontrolled space object reentries are hard to predict, making mitigation measures difficult. In November 2022, a reentering Long March 5B rocket caused the closure of airspace over Europe, delaying 645 flights and having a plausible economic impact of millions of Euros. There are international and domestic laws that might enable the recovery of economic losses resulting from uncontrolled reentries, but such losses should not be allowed to occur in the first place. Instead of leaving the location of a reentry to chance, controlled reentries can be achieved with existing technologies and mission designs, directing reentries away from areas of high aircraft traffic. Moving to a controlled reentry regime would create a cost to space operators – but that cost is currently being externalized to the aviation industry. Multilateral solutions to create a controlled reentry regime should be pursued, as recommended in the 2023 Montreal Recommendations on Aviation Safety and Uncontrolled Space Object Reentries, before there is a tragedy caused by an aircraft collision with debris from a rocket body or satellite.

Simulating impact-induced satellite breakups with a discrete element method

The Discrete Element Method is uniquely suited to modeling the extreme fragmentation typical of impact-induced satellite breakups. In this paper, we introduce a discrete orthotropic material model for simulating carbon fiber reinforced polymers (CFRP). We calibrate and validate the CFRP model with hypervelocity impact experiments performed at Fraunhofer EMI on thin CFRP face sheets. We demonstrate the suitability of the model for simulating spacecraft breakups by comparing simulation results to a ground-based experiment of a 15 x 15 × 15 cm CFRP micro-satellite from literature. Fragment size distributions show excellent agreement between experiment and literature. Comparisons are also made with the NASA breakup model.

Preliminary Debris Risk Analysis for Mega-Constellation Architectures After Orbit-Raising Fragmentation Event

This paper presents a theoretical examination of the potential debris conjunction dangers faced by mega-constellations in both low Earth orbit (LEO) and medium Earth orbit (MEO). The analysis focuses on the risks posed by debris fields created by breakups occurring during an orbit-raising maneuver for vehicle replacement and/or capability reconstitution, using current telecommunications mega-constellations, such as Starlink and OneWeb, as examples. The mega-constellation designs consist of 750 and 150 satellites arranged using the Walker-Delta design for the LEO and MEO cases, respectively. The research employs physics-based orbital propagation and Monte Carlo simulations to evaluate the potential consequences of a single satellite breakup during orbit raising. The results of the simulations are used to calculate the probability of catastrophic collision and compare the debris risk between the LEO and MEO mega-constellations, with a bimodality analysis conducted for the MEO constellation. Monte Carlo analysis indicates that the LEO mega-constellation features the highest percentage of potential catastrophic collisions between debris fragments and the mega-constellation. Specifically, satellite breakup events starting within an altitude range of 100–199 km below the LEO mega-constellation, or approximately at the midpoint of a Hohmann transfer from a 300-km parking orbit, pose the greatest risk to the constellation.

Determining origins of satellite breakup events in LEO region

Determining origins of satellite breakup events in LEO region

Updates on the DebriSat hypervelocity experiment and characterization of fragments in support of environmental models

To develop, maintain, and update orbital debris environmental and break-up models, such as the NASA Orbital Debris Engineering Model (ORDEM) and the NASA Standard Satellite Breakup Model (SSBM), the NASA Orbital Debris Program Office (ODPO) relies on the most recent validated datasets from various measurement sources, both laboratory and environmental. One key project that will provide insight for break-up events using modern-day spacecraft materials and construction techniques is the DebriSat laboratory hypervelocity-impact test. Based on the mass of the target, the projectile, and the impact velocity, the expected number of fragments greater than 2 mm, generated using the NASA SSBM, was estimated to be close to 85,000. To date, the DebriSat fragment database continues to grow, with over 200,000 fragments collected that will help inform updates to the SSBM. Additionally, the growing fragment ensemble will support key parameters for the next release of NASA's environmental models, employing fragment shapes, densities, and size distributions. To further support these environmental models, specifically the size parameter used in ground-based optical measurements, optical characterization on a subset of DebriSat fragments is being conducted in ODPO's Optical Measurement Center (OMC). Broadband bidirectional reflectance distribution (BRDF) measurements will provide insight into the optical-based NASA Size Estimation Model (OSEM). The OSEM equates an object's brightness to size (e.g., diameter of a disk or sphere) given several assumed parameters, including a defined phase function, albedo, and range. To address the first defined parameter, the ODPO has been using ray-tracing software to simulate light conditions in the OMC and to generate phase functions (i.e., specular, Lambertian, and experimentally derived) of known shapes, materials, and sizes. This simulated data, in addition to the experimentally derived measurements collected in the OMC, will aid in determining whether a new phase function would be suitable for an updated OSEM. The OSEM also assumes a single-value albedo, thus pre-impact spectral measurements on a subset of DebriSat materials were acquired for baseline material characterization and to provide insight into spacecraft material taxonomies. This DebriSat spectral data, along with spectral measurements from other known spacecraft material samples, will allow for better analyses of albedo variations and the effect on size calculations of defined laboratory targets, thus further supporting OSEM updates. This paper provides an overview of DebriSat, the status of the project, updates on the parameter distributions, an overview of the NASA SSBM, and ongoing fragment characterization efforts within the OMC.

Preliminary debris risk assessment for mega-constellations in low and medium earth orbit due to satellite breakup

This paper presents a theoretical analysis of the potential risk posed by artificial debris clouds in low Earth orbit (LEO) from mega-constellations, modeled after current communication constellations such as Starlink and OneWeb with 750 satellites each. The analysis examines three different constellation designs: a low-altitude LEO, a high-altitude LEO, and a medium Earth orbit (MEO) constellation, which will be positioned using the Walker-Delta design. The study is based on physics-based digital mission engineering and a Monte Carlo simulation framework. The simulation involves debris generated from a single breakup of one randomly selected satellite per run, but does not consider cascading debris events. This debris cloud is propagated for 1 week and how it interacts with the mega-constellation is recorded. The results show an average of 705.65 potential conjunctions within the LEO constellation, with 14.40% of those being considered catastrophic, and an average of 165.5 conjunctions in the MEO constellation, with 0.72% considered catastrophic.

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.

Simulating hypervelocity impact with a discrete element approach

We present new developments to a discrete element model designed to simulate hypervelocity impacts between on-orbit satellites and space debris. The advantage of the discrete element-based simulation method is its ability to accurately model fragmentation which always accompanies a hypervelocity impact event. As previous work showed the model's ability to accurately simulate hypervelocity impact fragmentation at very high impact velocities, our focus in this paper is predominantly on improving the model's behavior at lower velocities. These low-velocity secondary-impacts also play an important role in determining the final distribution of fragments following a hypervelocity impact event. We achieve this by enhancing our discrete element model's bonding structure to more accurately simulate a continuum material. We evaluate the enhanced model's performance with a variety of simulations ranging from tension tests to hypervelocity impacts. Parameters for the model are calibrated for the whole range of impact velocities using well know ballistic limit equations. Three large impact simulations are presented which demonstrate the model's capabilities in capturing the full range of fragment sizes. The satellite breakup simulation results are compared to the NASA breakup model and show very good agreement.

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.

Numerical investigation on the standard catastrophic breakup criteria

Hypervelocity collisions are predicted to be the dominant space debris source in the mid-term future, when a critical spatial density of satellites is reached. The NASA Standard Satellite Breakup Model (SSBM) has been adopted by major space agencies for characterizing hypervelocity spacecraft breakups for debris environment modeling. The SSBM is an empirical model based on data from ground tests and observations of on-orbit events. We propose to enhance this database by numerical simulations including a wide range of collision conditions and complex spacecraft models. We established the software tool PHILOS-SOPHIA for systematically studying the effects of on-orbit hypervelocity collisions. A particular focus was laid on the breakup criteria of the SSBM, which defines an energy-to-mass-ratio of 40 J/g being the collision condition for catastrophic fragmentations. We simulated six different scenarios of a complex spacecraft colliding with a small satellite. In the detailed fragmentation analysis, we find both good agreements and clear deviations between the hydrocode results and the SSBM predictions. Particularly, the collision geometry strongly influences the fragmentation damage and the area-to-mass distributions. Depending on the collision vector, impacts on the outer parts of a spacecraft may result in both higher and lower fragmentation in comparison with impacts on the center of mass. The simple breakup criteria does not reflect this complexity and we recommend performing more research. Numerical simulations, thoroughly backed by advanced experiments, can make a significant contribution to improve the accuracy of breakup models.

Spacecraft Electrical Passivation: From Study to Reality

Due to the increasing risk posed on satellites from space debris, requirements are now being implemented into satellites to be launched in the future to prevent the release of additional debris. One of the key requirements is the electrical passivation of satellites at end of mission (EoM) in order to prevent the break-up of satellites due to the stored energy in the power subsystem. In order to comply with these new specifications and find suitable solutions for the upcoming missions, a set of workshops and R&D activities have been initiated and conducted in the past few years. In particular it has provided a better understanding of the environmental constraints and the battery safety after EoM, together with the identification of the most appropriate implementations for electrical passivation. The latest available results of these activities and the ongoing work are presented within this paper. In addition, new spacecraft under development are currently implementing electrical passivation solutions taking into account the conclusions of these studies.

Dangers of Space Debris

S pace D ebris : the H ubris of H umanity Botao Peng there. Back in 2007, China conducted an anti-satellite weapon test that destroyed a decommissioned weather satellite, smashing the object into 150,000 pieces each larger than 1 cm . Not only does space junk pose a threat to space travel, it also increases the chance of self multiplication. Space debris was not always a problem. In the early years of the space age, scientists would leave abandoned rocket stages in orbit after use. This was not such a bad idea, considering how empty space was at the time. However, over time, the number of space missions increased from simply sending a weather satellite into space to landing a man on the moon. With each mission, more waste was left behind, increasing the amount of debris orbiting the earth. The used rocket stages that once were few and unnoticeable now cloud our view of the heavens. So much in fact, that, 47.7% of debris orbiting Earth is from break-up debris [4] and used propulsion engines cause 45.4% of all satellite break-ups [4] . Such a recent collision occurred in 2005. A 31-year-old U.S. rocket body hit “A single puncture could effectively destroy all electronics , release all the oxygen , impair the escape vessels , and even kill all members aboard .” B erkeley S cientific J ournal • W aste • S pring 2015 • V olume 19 • I ssue 2 • 1 B S J centimeters is not big. 10 centimeters is the average diameter of a bagel. 10 centimeters is even shorter than the length of an average smartphone. So why is 10 centimeters a big deal? Imagine a piece of metal 10 centimeters in diameter orbiting around the Earth in space Yet this piece of metal has enough momentum and force to puncture the International Space Station. A single puncture could effectively destroy all electronics, release all the oxygen, impair the escape vessels, and even kill all members aboard. This is the danger of space debris. Space debris is junk that orbits the earth. This ranges from abandoned satellites to broken segments from rockets and missiles. It also includes defragmented rocket stages as well as broken equipment from collisions with other debris. The first real serious satellite fragmentation occurred in June of 1961 when two satellites collided with one another [4] . The collision resulted in a break-up of the satellites, sending pieces of both into orbit. This only increases the chances of another collision of occurring. Yet the fragmentation does not stop

Satellite Breakup Parameter Determination

A satellite breakup can cause collision risk to other resident satellites. If the time, location, and energy content of the breakup can be determined, The Aerospace Corporation has models that can quickly estimate the risk to other resident satellites from the debris. We have developed a method that allows determination of all required breakup parameters within about 12 h after the breakup event. If a large number of radar tracks are available, this timeline could be even shorter. The method then facilitates rapid determination of debris collision risk to resident satellites, providing satellite operators with timely space situational awareness.

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.

The Jovian Rings

AbstractA comparison of the Jovian and Saturnian rings is made by reviewing the recent advances in planetary spacecraft exploration and theoretical study. Two main issues are addressed, namely, the different structures of these two planetary ring systems and the water ice composition of the Saturnian rings. It is suggested that answers might be found by invoking tidal capture of Trans-Neptunian Objects with highly differentiated structures even though catastrophic breakup of pre-existing satellites in the ring regions remains a real possibility. Erosion mechanisms such as meteoroid impact, photo-sputtering, orbital instability of charged dust particles and thermal evaporation acting at different time scales could lead to the preservation of the Saturnian ring system but not the Jovian ring system of large mass originally.

Fragmentation Analysis of Cosmos 2421 Satellite

2008年上半年俄罗斯的Cosmos 2421卫星接连发生了三次不同程度的解体, 并产生了数百颗碎片. 本文利用美国公布的Cosmos 2421卫星及其碎片的两行根数系列, 对三次解体事件发生的时间, 解体碎片的轨道分布、速度分布和面积质量比以及寿命等进行了分析, 并对解体碎片对航天器的影响进行了评估. 分析结果表明, Cosmos2421卫星的三次解体事件分别发生在2008年3月14日、4月28日、6月9日; 解体碎片分布在200~1400 km 的高度范围内; Cosmos 2421卫星解体导致的碎片在空间三个方向上的速度增量均值分别为-8.4m/s, 8.6m/s, -8.3m/s; 67\%的解体碎片的寿命都在1年以内, 解体事件造成500 km以下的空间碎片空间密度增加, 对载人航天器产生了影响. 通过对Cosmos 2421卫星解体事件的分析可以看出, 利用解体碎片的轨道信息可以反演解体事件特性, 根据解体碎片的寿命和空间密度分布的计算结果可以评估解体事件对未来发射活动和在轨卫星的影响.

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.