Pub Date : 2024-09-01DOI: 10.1016/j.jsse.2024.04.012
Matthew K. Brown, Sean Elvidge
Atmospheric drag is a major perturbation in Low Earth Orbit (LEO). The neutral density obtained from atmospheric models is a major source of uncertainty in drag calculations and therefore orbital propagation in LEO. Many atmospheric models are available, with fast empirical models most commonly used. We explore the challenges and benefits of using numerical models, specifically the Whole Atmosphere Community Climate Model with thermosphere and ionosphere extension (WACCM-X) as part of the Community Earth System Model (CESM). Numerical models provide higher resolution of thermospheric structures, along with more accurate neutral density forecasts through assimilative models such as the Advanced Ensemble electron Density Assimilation System (AENeAS). Solutions are presented to overcome the challenges of using numerical models for neutral densities.
{"title":"Using WACCM-X neutral densities for orbital propagation: Challenges and solutions","authors":"Matthew K. Brown, Sean Elvidge","doi":"10.1016/j.jsse.2024.04.012","DOIUrl":"10.1016/j.jsse.2024.04.012","url":null,"abstract":"<div><div>Atmospheric drag is a major perturbation in Low Earth Orbit (LEO). The neutral density obtained from atmospheric models is a major source of uncertainty in drag calculations and therefore orbital propagation in LEO. Many atmospheric models are available, with fast empirical models most commonly used. We explore the challenges and benefits of using numerical models, specifically the Whole Atmosphere Community Climate Model with thermosphere and ionosphere extension (WACCM-X) as part of the Community Earth System Model (CESM). Numerical models provide higher resolution of thermospheric structures, along with more accurate neutral density forecasts through assimilative models such as the Advanced Ensemble electron Density Assimilation System (AENeAS). Solutions are presented to overcome the challenges of using numerical models for neutral densities.</div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 411-416"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142571585","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-09-01DOI: 10.1016/j.jsse.2024.06.003
NASA and ESA are currently planning the Mars Sample Return campaign, comprising missions whose combined objective is to bring the first samples of Mars material back to Earth for detailed study. Until recently, the NASA-ESA plan was to return samples to Earth using three missions. The final component, the Earth Entry System (EES), will bring the Mars samples back to the Earth, where it will land following safe entry through the Earth's atmosphere. There is a concern regarding the risk of biological contamination of the Earth's biosphere from returned Mars samples if, for example, the structural integrity of the EES were compromised during its return mission due to a perforation of a critical surface resulting from a high-speed meteoroid impact. To assess the risks associated with such an event, NASA is developing equations that predict the damage that various EES elements will sustain as a result of such an impact, as well as equations that predict whether or not a particular system will sustain a critical failure following such an impact. In this paper, we review recent progress in the development of such equations for the EES forebody and the EES aftbody, the two elements of the EES that are most exposed to the meteoroid environment. Limitations of the BLEs are also discussed, which can also be used to further inform the next steps in the BLE development.
美国航天局和欧空局目前正在规划火星取样返回活动，其中包括一些飞行任务，其共同目标是将第一批火星物质样品带回地球进行详细研究。直到最近，NASA-ESA 的计划是通过三次飞行任务将样本送回地球。最后一个部分，即地球进入系统（EES），将把火星样本带回地球，在安全进入地球大气层后在地球着陆。有人担心，如果在执行返回任务期间，由于高速流星体撞击导致关键表面穿孔，EES 的结构完整性受到破坏，返回的火星样本就有可能对地球生物圈造成生物污染。为了评估与此类事件相关的风险，NASA 正在开发一些方程，用于预测各种 EES 元件在此类撞击中将遭受的破坏，以及预测特定系统在此类撞击后是否会出现临界故障的方程。在本文中，我们回顾了最近在为 EES 前体和 EES 后体开发此类方程方面取得的进展，EES 前体和 EES 后体是 EES 中暴露在流星体环境中最多的两个元件。本文还讨论了 BLE 的局限性，这些局限性也可用于进一步指导 BLE 开发的下一步工作。
{"title":"Development of ballistic limit equations in support of the Mars sample return mission","authors":"","doi":"10.1016/j.jsse.2024.06.003","DOIUrl":"10.1016/j.jsse.2024.06.003","url":null,"abstract":"<div><div><span>NASA and ESA are currently planning the Mars Sample Return campaign, comprising missions whose combined objective is to bring the first samples of Mars material back to Earth for detailed study. Until recently, the NASA-ESA plan was to return samples to Earth using three missions. The final component, the Earth Entry System (EES), will bring the Mars samples back to the Earth, where it will land following safe entry through the Earth's atmosphere. There is a concern regarding the risk of biological contamination<span> of the Earth's biosphere from returned Mars samples if, for example, the structural integrity of the EES were compromised during its return mission due to a perforation of a critical surface resulting from a high-speed meteoroid impact. To assess the risks associated with such an event, NASA is developing equations that predict the damage that various EES elements will sustain as a result of such an impact, as well as equations that predict whether or not a particular system will sustain a critical failure following such an impact. In this paper, we review recent progress in the development of such equations for the EES </span></span>forebody and the EES aftbody, the two elements of the EES that are most exposed to the meteoroid environment. Limitations of the BLEs are also discussed, which can also be used to further inform the next steps in the BLE development.</div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 417-424"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141404831","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Large-scale structures have a non-negligible collision probability with micrometeoroid and orbital debris (MMOD) due to their massive size, even in geostationary orbit (GEO) with low debris flux. When MMOD impact the spacecraft surfaces at high velocity, secondary debris called ejecta are generated, and they may remain semi-permanently and accumulate because there is no atmospheric drag at high altitudes such as GEO. To evaluate the amount of ejecta generation, hypervelocity impact tests were conducted for the material for future large-scale structures or material commonly used in conventional spacecraft, such as CFRP honeycomb panels and solar cells. The effect of impact energy on ejecta generation was evaluated by changing the impact velocity and projectile density. Impact tests were also conducted on irradiated samples to investigate the effects of environmental degradation due to long-term exposure to orbit. The results showed that the amount of ejecta increased with impact energy and may have been affected by radiation-induced degradation. Next, hypervelocity impact tests were conducted to investigate the measures to reduce ejecta, and it was shown that the ejecta generation could be reduced by using low-density materials such as polyimide foam and silica aerogel.
{"title":"Estimation of ejecta generation and mitigation measures for large-scale structures on geostationary orbit","authors":"Satomi Kawamoto , Ryusuke Harada , Daisuke Joudoi , Yugo Kimoto , Taku Izumiyama , Yasuhiro Akahoshi","doi":"10.1016/j.jsse.2024.06.005","DOIUrl":"10.1016/j.jsse.2024.06.005","url":null,"abstract":"<div><div>Large-scale structures have a non-negligible collision probability<span> with micrometeoroid<span><span><span><span> and orbital debris (MMOD) due to their massive size, even in geostationary orbit (GEO) with low debris flux. When MMOD impact the spacecraft surfaces at high velocity, secondary debris called </span>ejecta<span> are generated, and they may remain semi-permanently and accumulate because there is no atmospheric drag at high altitudes such as GEO. To evaluate the amount of ejecta generation, </span></span>hypervelocity impact tests were conducted for the material for future large-scale structures or material commonly used in conventional spacecraft, such as CFRP honeycomb panels and solar cells. The effect of impact energy on ejecta generation was evaluated by changing the </span>impact velocity<span><span> and projectile density. Impact tests were also conducted on irradiated samples to investigate the effects of environmental degradation due to long-term exposure to orbit. The results showed that the amount of ejecta increased with impact energy and may have been affected by radiation-induced degradation. Next, hypervelocity impact tests were conducted to investigate the measures to reduce ejecta, and it was shown that the ejecta generation could be reduced by using low-density materials such as </span>polyimide<span> foam and silica aerogel.</span></span></span></span></div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 507-517"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141704579","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Every on-orbit collision or explosion can pose a threat, not only to the existing satellite population but also to the long-term usability of Earth orbit. This threat exists even if satellites can actively maneuver to avoid trackable debris fragments, since an estimated 96 % of potentially mission-ending (>1 cm) debris is untrackable [1]. Prevention of every on-orbit breakup may not be possible. However, armed with an understanding of the likely causes of fragmentation events, satellite developers and operators can take actions to mitigate such events in the future. Astrodynamics forensic analyses, the sleuthing techniques used to gather an event's known details and estimate its unknown parameters, can be used to develop theories about the causes of a breakup and to predict its consequences.
In the past five years, several on-orbit collisions and explosions have occurred, involving a variety of orbiting objects with varying amounts of available observational data. Techniques and tools developed over decades at The Aerospace Corporation are used to characterize key parameters of these events, including spread velocity of the debris pieces, energy involved in the breakup events, and mass and area estimates of the individual debris fragments. These forensic capabilities are enhanced by utilizing patterns identified from different classes of historical breakups and ground-test data. This paper shows the effectiveness of this methodology when used for analysis of a variety of event types including collisions, such as the Cosmos 1408 ASAT test and SL-14 rocket body breakup, rocket body fragmentations such as the 2022 Long March 6A breakup, and satellite fragmentations such as the Resur-O1 breakup. Representative models of events are developed using the IMPACT fragmentation tool, and predictions of the lifetimes of the subtrackable orbital debris are included. Where event sources are unknown, breakup parameters and trends are used to suggest possible causes. The challenges of analyzing an orbital breakup mystery with few observational clues are also discussed.
{"title":"Forensic analysis of recent debris-generating events","authors":"D.L. Mains , G.E. Peterson , J.P. McVey , J.C. Maldonado , M.E. Sorge","doi":"10.1016/j.jsse.2024.06.006","DOIUrl":"10.1016/j.jsse.2024.06.006","url":null,"abstract":"<div><div>Every on-orbit collision or explosion can pose a threat, not only to the existing satellite population but also to the long-term usability of Earth orbit. This threat exists even if satellites can actively maneuver to avoid trackable debris fragments, since an estimated 96 % of potentially mission-ending (>1 cm) debris is untrackable [<span><span>1</span></span><span>]. Prevention of every on-orbit breakup may not be possible. However, armed with an understanding of the likely causes of fragmentation events, satellite developers and operators can take actions to mitigate such events in the future. Astrodynamics forensic analyses, the sleuthing techniques used to gather an event's known details and estimate its unknown parameters, can be used to develop theories about the causes of a breakup and to predict its consequences.</span></div><div>In the past five years, several on-orbit collisions and explosions have occurred, involving a variety of orbiting objects with varying amounts of available observational data. Techniques and tools developed over decades at The Aerospace Corporation are used to characterize key parameters of these events, including spread velocity of the debris pieces, energy involved in the breakup events, and mass and area estimates of the individual debris fragments. These forensic capabilities are enhanced by utilizing patterns identified from different classes of historical breakups and ground-test data. This paper shows the effectiveness of this methodology when used for analysis of a variety of event types including collisions, such as the Cosmos 1408 ASAT test and SL-14 rocket body breakup, rocket body fragmentations such as the 2022 Long March 6A breakup, and satellite fragmentations such as the Resur-O1 breakup. Representative models of events are developed using the IMPACT fragmentation tool, and predictions of the lifetimes of the subtrackable orbital debris are included. Where event sources are unknown, breakup parameters and trends are used to suggest possible causes. The challenges of analyzing an orbital breakup mystery with few observational clues are also discussed.</div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 388-394"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141845281","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-09-01DOI: 10.1016/j.jsse.2024.08.001
Hugh G. Lewis, Georgia Skelton
The deployment of constellations of satellites within low Earth orbit (LEO) has implications for space operations and for the broader space environment. A large active satellite population will experience high numbers of conjunctions with other resident space objects (RSOs). Even if only a small proportion are high-probability events, the substantial number of conjunctions will still lead to many potentially high-risk encounters with other RSOs and a correspondingly high burden for their operators to mitigate them via maneuvers. This burden is exacerbated if the operator adopts an approach whereby risk mitigation maneuvers are conducted at collision probability levels below the widely accepted 1E-4 (1-in-10,000). Despite these significant efforts the remaining aggregate risk may still be relatively high because of the large number of conjunctions experienced by some constellations, leading to ongoing concern over the safety of these space systems. Through an analysis of conjunction assessment data, simulations using the DAMAGE computational model, and a new mapping approach, the risks from conjunctions between large constellations and other RSOs have been investigated. The results show that some existing constellations currently face more than a 10 % annual collision probability even after accounting for their robust risk mitigation approaches, with implications for the safety and long-term sustainability of large constellations and the broader LEO environment. Overall, the work emphasizes the need for new research and guidance on this aspect of space operations.
{"title":"Safety considerations for large constellations of satellites","authors":"Hugh G. Lewis, Georgia Skelton","doi":"10.1016/j.jsse.2024.08.001","DOIUrl":"10.1016/j.jsse.2024.08.001","url":null,"abstract":"<div><div>The deployment of constellations of satellites within low Earth orbit (LEO) has implications for space operations and for the broader space environment. A large active satellite population will experience high numbers of conjunctions with other resident space objects (RSOs). Even if only a small proportion are high-probability events, the substantial number of conjunctions will still lead to many potentially high-risk encounters with other RSOs and a correspondingly high burden for their operators to mitigate them via maneuvers. This burden is exacerbated if the operator adopts an approach whereby risk mitigation maneuvers are conducted at collision probability levels below the widely accepted 1E-4 (1-in-10,000). Despite these significant efforts the remaining aggregate risk may still be relatively high because of the large number of conjunctions experienced by some constellations, leading to ongoing concern over the safety of these space systems. Through an analysis of conjunction assessment data, simulations using the DAMAGE computational model, and a new mapping approach, the risks from conjunctions between large constellations and other RSOs have been investigated. The results show that some existing constellations currently face more than a 10 % annual collision probability even after accounting for their robust risk mitigation approaches, with implications for the safety and long-term sustainability of large constellations and the broader LEO environment. Overall, the work emphasizes the need for new research and guidance on this aspect of space operations.</div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 439-445"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142571559","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper proposes less fuel strategies for space debris removal. To mitigate the risk of space debris cost-efficiently, multi-rendezvous missions are under development. On the other hand, multi-rendezvous missions often require changing orbital planes of removal satellites, which requires a huge amount of ΔV. Therefore, this study focuses on exploiting the J2 perturbation force as an auxiliary force and aims to establish maneuver rules that minimize ΔV consumption while maximizing the benefit of the J2 perturbation. The J2 perturbation equation is explored analytically, which clarifies whether the change in the semi-major axis or the inclination dominates the efficiency of the exploitation. A straightforward criterion is extracted which determines the efficient maneuver based on the initial inclination of the satellite.
{"title":"Less fuel strategies for space debris removal in Low Earth Orbit","authors":"Yuki Itaya , Yasuhiro Yoshimura , Toshiya Hanada , Tadanori Fukushima","doi":"10.1016/j.jsse.2024.08.002","DOIUrl":"10.1016/j.jsse.2024.08.002","url":null,"abstract":"<div><div>This paper proposes less fuel strategies for space debris removal. To mitigate the risk of space debris cost-efficiently, multi-rendezvous missions are under development. On the other hand, multi-rendezvous missions often require changing orbital planes of removal satellites, which requires a huge amount of ΔV. Therefore, this study focuses on exploiting the J<sub>2</sub> perturbation force as an auxiliary force and aims to establish maneuver rules that minimize ΔV consumption while maximizing the benefit of the J<sub>2</sub> perturbation. The J<sub>2</sub> perturbation equation is explored analytically, which clarifies whether the change in the semi-major axis or the inclination dominates the efficiency of the exploitation. A straightforward criterion is extracted which determines the efficient maneuver based on the initial inclination of the satellite.</div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 476-480"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142571562","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-09-01DOI: 10.1016/j.jsse.2024.04.007
This study proposes a relative orbit control law for laser debris removal missions considering the uncertainties of laser ablation and atmospheric drag. A removal spacecraft irradiates laser pulses to a target debris to generate the ablation force for deorbiting. The deorbiting force lowers the target altitude, and the removal spacecraft must follow it to maintain its relative position for continuous laser irradiation. The difficulty stems from uncertainties of the magnitude of laser ablation and external disturbances such as atmospheric drag. To tackle this problem, this study derives an adaptive control method using the Gaussian process regression to cancel the uncertainties with a nonparametric regression model. Numerical simulations verify the proposed control law under the uncertainties of laser ablation and atmospheric drag. The proposed control law can contribute to the realization of a safer and more secure mission not only for laser debris removal missions, but also for other on-orbit services.
{"title":"Adaptive relative orbit control considering laser ablation uncertainty","authors":"","doi":"10.1016/j.jsse.2024.04.007","DOIUrl":"10.1016/j.jsse.2024.04.007","url":null,"abstract":"<div><div>This study proposes a relative orbit control law for laser debris removal missions considering the uncertainties of laser ablation<span><span><span> and atmospheric drag. A removal spacecraft irradiates laser pulses to a target debris to generate the ablation force for deorbiting. The deorbiting force lowers the target altitude, and the removal spacecraft must follow it to maintain its relative position for continuous laser irradiation. The difficulty stems from uncertainties of the magnitude of </span>laser ablation and </span>external disturbances<span><span> such as atmospheric drag. To tackle this problem, this study derives an adaptive control method using the Gaussian process regression to cancel the uncertainties with a nonparametric regression model. Numerical simulations verify the proposed control law under the uncertainties of </span>laser ablation and atmospheric drag. The proposed control law can contribute to the realization of a safer and more secure mission not only for laser debris removal missions, but also for other on-orbit services.</span></span></div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 491-497"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141057226","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-09-01DOI: 10.1016/j.jsse.2024.05.007
This paper explores the concept of space sustainability and its interconnections using systems thinking approaches. This is done by highlighting the importance of multi-disciplinary perspectives when creating policies aimed at addressing the complex challenges of sustainability for space-related activities. Causal loop diagrams are employed to highlight the presence of feedback loops and causal relationships that are typically absent in space debris models and are treated as separate systems. A systems representation of the space environment is presented along with a discussion of its role in furthering research relating to the impact of large satellite constellations on factors important for holistic sustainability. This study investigated one example feedback between the space environment and the atmosphere and found that CO2 emissions specifically emitted from launches and re-entries have no significant impact on atmospheric density below 500 km.
{"title":"A holistic systems thinking approach to space sustainability via space debris management","authors":"","doi":"10.1016/j.jsse.2024.05.007","DOIUrl":"10.1016/j.jsse.2024.05.007","url":null,"abstract":"<div><div>This paper explores the concept of space sustainability and its interconnections using systems thinking approaches. This is done by highlighting the importance of multi-disciplinary perspectives when creating policies aimed at addressing the complex challenges of sustainability for space-related activities. Causal loop diagrams are employed to highlight the presence of feedback loops and causal relationships that are typically absent in space debris models and are treated as separate systems. A systems representation of the space environment is presented along with a discussion of its role in furthering research relating to the impact of large satellite constellations on factors important for holistic sustainability. This study investigated one example feedback between the space environment and the atmosphere and found that CO<sub>2</sub> emissions specifically emitted from launches and re-entries have no significant impact on atmospheric density below 500 km.</div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 532-538"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141408580","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-09-01DOI: 10.1016/j.jsse.2024.08.005
Nicolas Billot , Stephan Hellmich , Willy Benz , Andrea Fortier , David Ehrenreich , Christopher Broeg , Alexis Heitzmann , Anja Bekkelien , Alexis Brandeker , Yann Alibert , Roi Alonso , Tamas Bárczy , David Barrado Navascues , Susana C.C. Barros , Wolfgang Baumjohann , Federico Biondi , Luca Borsato , Andrew Collier Cameron , Carlos Corral van Damme , Alexandre C.M. Correia , Thomas G. Wilson
The CHaracterising ExOPlanet Satellite (CHEOPS) is a partnership between the European Space Agency and Switzerland with important contributions by 10 additional ESA member States. It is the first S-class mission in the ESA Science Programme. CHEOPS has been flying on a Sun-synchronous low Earth orbit since December 2019, collecting millions of short-exposure images in the visible domain to study exoplanet properties.
A small yet increasing fraction of CHEOPS images show linear trails caused by resident space objects crossing the instrument field of view. CHEOPS’ orbit is indeed particularly favourable to serendipitously detect objects in its vicinity as the spacecraft rarely enters the Earth's shadow, sits at an altitude of 700 km, and observes with moderate phase angles relative to the Sun. This observing configuration is quite powerful, and it is complementary to optical observations from the ground.
To characterize the population of satellites and orbital debris observed by CHEOPS, all and every science images acquired over the past 3 years have been scanned with a Hough transform algorithm to identify the characteristic linear features that these objects cause on the images. Thousands of trails have been detected. This statistically significant sample shows interesting trends and features such as an increased occurrence rate over the past years as well as the fingerprint of the Starlink constellation. The cross-matching of individual trails with catalogued objects is underway as we aim to measure their distance at the time of observation and deduce the apparent magnitude of the detected objects.
As space agencies and private companies are developing new space-based surveillance and tracking activities to catalogue and characterize the distribution of small debris, the CHEOPS experience is timely and relevant. With the first CHEOPS mission extension currently running until the end of 2026, and a possible second extension until the end of 2029, the longer time coverage will make our dataset even more valuable to the community, especially for characterizing objects with recurrent crossings.
{"title":"In-situ observations of resident space objects with the CHEOPS space telescope","authors":"Nicolas Billot , Stephan Hellmich , Willy Benz , Andrea Fortier , David Ehrenreich , Christopher Broeg , Alexis Heitzmann , Anja Bekkelien , Alexis Brandeker , Yann Alibert , Roi Alonso , Tamas Bárczy , David Barrado Navascues , Susana C.C. Barros , Wolfgang Baumjohann , Federico Biondi , Luca Borsato , Andrew Collier Cameron , Carlos Corral van Damme , Alexandre C.M. Correia , Thomas G. Wilson","doi":"10.1016/j.jsse.2024.08.005","DOIUrl":"10.1016/j.jsse.2024.08.005","url":null,"abstract":"<div><div>The CHaracterising ExOPlanet Satellite (CHEOPS) is a partnership between the European Space Agency and Switzerland with important contributions by 10 additional ESA member States. It is the first S-class mission in the ESA Science Programme. CHEOPS has been flying on a Sun-synchronous low Earth orbit since December 2019, collecting millions of short-exposure images in the visible domain to study exoplanet properties.</div><div>A small yet increasing fraction of CHEOPS images show linear trails caused by resident space objects crossing the instrument field of view. CHEOPS’ orbit is indeed particularly favourable to serendipitously detect objects in its vicinity as the spacecraft rarely enters the Earth's shadow, sits at an altitude of 700 km, and observes with moderate phase angles relative to the Sun. This observing configuration is quite powerful, and it is complementary to optical observations from the ground.</div><div>To characterize the population of satellites and orbital debris observed by CHEOPS, all and every science images acquired over the past 3 years have been scanned with a Hough transform algorithm to identify the characteristic linear features that these objects cause on the images. Thousands of trails have been detected. This statistically significant sample shows interesting trends and features such as an increased occurrence rate over the past years as well as the fingerprint of the Starlink constellation. The cross-matching of individual trails with catalogued objects is underway as we aim to measure their distance at the time of observation and deduce the apparent magnitude of the detected objects.</div><div>As space agencies and private companies are developing new space-based surveillance and tracking activities to catalogue and characterize the distribution of small debris, the CHEOPS experience is timely and relevant. With the first CHEOPS mission extension currently running until the end of 2026, and a possible second extension until the end of 2029, the longer time coverage will make our dataset even more valuable to the community, especially for characterizing objects with recurrent crossings.</div></div>","PeriodicalId":37283,"journal":{"name":"Journal of Space Safety Engineering","volume":"11 3","pages":"Pages 498-506"},"PeriodicalIF":1.0,"publicationDate":"2024-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142571564","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-09-01DOI: 10.1016/j.jsse.2024.07.009
Juan C. Maldonado, Alan B. Jenkin, John P. McVey
Recent years have seen an increase in CubeSat missions on rideshares to geosynchronous orbit. The typical practice for these missions is to deploy the CubeSat on a geosynchronous transfer orbit (GTO) which results in a perigee altitude that is low enough that atmospheric drag will cause the apogee altitude to decay and eventual reentry. However, for GTOs, demonstrating compliance with limits on orbital lifetime in orbital debris mitigation guidelines is not straightforward due to a solar resonance phenomenon that exists which can cause high variability of the orbital lifetime of the satellite. This paper presents a procedure for determining the likelihood that orbital lifetime of a rideshare CubeSat on a resonant GTO will have an orbital lifetime below a 25-year and 5-year limit. The procedure uses a Monte Carlo analysis in which uncertain parameters are randomly varied, including the launch time/initial RAAN and the drag coefficient. The Aerospace precision propagation tool TRACE is used with a high-fidelity force model to enable precision integration through the atmosphere at perigee. It is shown in the study that there is a systematic variation in likelihood of staying below a 25- and 5-year orbital lifetime limit and that drag enhancement devices may be needed to meet the 5-year limit for CubeSat rideshares. The study also presents findings on a solar radiation pressure induced resonance that was observed for high area-to-mass ratios which suggests that there can be a diminishing return when increasing the area of a drag enhancement device to quicken deorbit.
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