The U.S. Government Orbital Debris Mitigation Standard Practices (ODMSP) released in December 2019 include recommendations for disposal orbits, including options for disposal above geosynchronous orbit (GEO) and disposal between low Earth orbit (LEO) and GEO. Demonstration of compliance with these recommendations requires the ability to propagate disposal orbit parameters for 100-200 years and analyze whether the disposal orbit meets a maximum threshold of collision probability with other trackable orbital objects. To facilitate the design of missions that are ODMSP-compliant by launch service providers, a study was performed that generated a document and associated data tables containing relevant metrics needed for demonstration of ODMSP compliance for parametric orbital data sets. Results for two disposal orbit classes are presented in this paper: above-GEO and above-GPS. The study was focused on generic upper stages of typical medium and heavy class launch vehicles, and therefore the scope was limited to storage disposal orbits rather than reentry options since these upper stages generally have high reentry debris casualty area. The compliance metrics that were computed include minimum perigee altitude and maximum apogee altitude over 100 years, time spent in the +/- 300 km semi-synchronous zone over 200 years, and collision probability with other trackable orbital objects over 100 years. The resulting look-up tables and spreadsheets allow a user to input the classical orbital elements of a candidate disposal orbit and retrieve corresponding compliance metrics and total compliance status. This paper presents how the tables were generated, how they can be used, and plots of compliance metrics and flags indicating total compliance status for a variety of sample above-GEO and above-GPS disposal orbits.

Large constellations of hundreds or thousands of satellites in low Earth orbit are a recent development that is both creating new space-based services and creating challenges for space traffic management especially considering the number of new constellations and corresponding satellites proposed. The large numbers of satellites, orders of magnitude beyond historic norms and the new modes of operations present a range of technical, policy and communications challenges. These challenges are discussed along with proposed and some implemented solutions. Several recommendations are made to improve the safety of large constellations while enabling continued innovation.

Collision avoidance, the process of planning and possibly executing a manoeuvre to mitigate the risk of a collision in orbit, is becoming increasingly important as the amount of space traffic increases. This paper discusses different types of conjunction events, the technical processes involved in identifying higher risk conjunctions and possible mitigation techniques, and gaps and limitations in the processes. Possible solutions to these gaps are addressed including improved communication and coordination, more accurate and precise data, and improved education of operators. Several recommendations are made to improve the collision avoidance process and effectiveness.

Several commercial companies, as well as various nations, have proposed to deploy or are deploying many satellites in Low Earth Orbit (LEO). These large constellations will greatly increase the number of satellites operating in relatively narrow altitude regions of space. The added space traffic in these regions will create many close approaches between the members of the large constellations and other space operators. These close approach situations can necessitate maneuver(s) to avoid a potential collision. Should both satellites have maneuvering capability, the question of how the overall collision avoidance procedures should be executed is raised. Some constellations may employ automated collision avoidance systems which interact differently than conventional human-in-the-loop systems. Interactions between an automated system and another operational satellite, between two automated systems or two nonautonomous systems present new challenges for executing effective collision avoidance. Additionally, the existence of non-maneuverable satellites and space debris continues to pose additional challenges. This paper is the first of several papers that will be documenting an International Academy of Astronautics study on this topic.

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.

The U.S. Government Orbital Debris Mitigation Standard Practices (ODMSP) released in December 2019 include several disposal options that are applicable to GPS satellites. One option is to leave a satellite on a near-circular disposal orbit with limited long-term eccentricity growth. Many currently operational GPS satellites are planned to move to near-circular disposal orbits above the GPS operational constellation and below the BeiDou Satellite System (BDS). The conditions in the ODMSP specify that GPS disposal orbits avoid crossing the semi-synchronous zone (altitude range 20,182 + /-300 km, occupied by GPS) or the BDS operational altitude for 100 years. An analysis was performed to determine ways to achieve ODMSP-compliance of near-circular GPS disposal orbits and limit associated collision risk. Disposal orbit propagations over 100 years were used to determine the available initial semi-major axis (SMA) range for compliant disposal orbits. Next, the GPS constellation and satellite disposals were simulated over a 200-year time frame. The disposal orbits were propagated and the collision probability between disposed satellites was then evaluated. Results show an available initial SMA range of 363 km for an initial eccentricity of 0.0 0 03 and a range of only 3 km for an initial eccentricity of 0.001, which cover worst-case eccentricity growth scenarios. Results also show that a disposal strategy with initial eccentricity of 0.0 0 03 and initial SMA uniformly spread over 363 km has the lowest collision probability for future disposed GPS satellites.(c) 2022 International Association for the Advancement of Space Safety. Published by Elsevier Ltd. All rights reserved.

IMPACT is a hypervelocity collision and explosion model developed by The Aerospace Corporation to predict the characteristics of debris generated by fragmentation events. Used for over 30 years, it is a semi-empirical model that has supported a wide range of orbital debris analyses including satellite risk assessment, on-orbit event analyses, flight test and mission planning, debris mitigation and spacecraft design studies, and long-term debris environment evolution and space safety analyses. The model interconnects empirical expressions with conservation laws and boundary conditions to model debris generation from individual fragmentation events. A series of efforts to study available on-orbit data and newly-generated ground-test data during the past decade has added to the understanding of on-orbit fragmentation events, particularly with regard to explosion events, sub-catastrophic breakups, smaller fragment characteristics, and more diverse modern satellite construction materials. These insights have led to several upgrades of the IMPACT collision and explosion model. The model's capabilities have been expanded to represent sub-catastrophic and asymmetric fragmentation events, more accurately model fragment material differences and their relationships to area-to-mass ratios, model smaller fragments, and relate ground-based and on-orbit data. Additional insights have been gained through the integration of individual model improvements while maintaining overall model consistency. This paper provides an overview of the current IMPACT model, discussing the design and expressions used in the current collision and explosion algorithms, changes over the past decade, and implications of insights from both on-orbit and ground test data. Modelling techniques discussed include representation of fragmenting object material differences, sub-catastrophic fragmentations, and small fragment modelling. Interactions between different model improvements are discussed, including small versus large fragment characteristics and sub-catastrophic versus catastrophic fragmentation events. Model updates are considered in the context of fragmentation event data. Future directions for development also are considered based on the insights from the aggregated data and model updates.

The feasibility of a novel lidar-based sensor for space debris detection is tested in a laboratory setting using a low-power transmitter and simplified receiver. The transmitter employs a Powell lens to spread transmitted laser light into an evenly distributed fan shape which detects particles or objects moving through the beam and measures their approximate size. Link budget calculations based on a theoretical model are validated. Additionally, several important considerations for future deployment on a satellite are identified.

The U.S. Government Orbital Debris Mitigation Standard Practices (ODMSP) released in December 2019 include a disposal option to use orbital eccentricity growth for long-term reentry within 200 years. Long-term reentry is beneficial for orbital debris mitigation because it offers a large reduction in long-term collision risk in the regions above low Earth orbit (LEO) compared to use of storage disposal orbits. As a condition for use of this option, the ODMSP specifies a 25-year limit on cumulative time a disposed spacecraft can spend in designated altitude zones that are frequently used by operational spacecraft. An analysis was performed to determine the time spent by disposed spacecraft in the geosynchronous (GEO) zone, the LEO zone, and the semi-synchronous zone for three classes of reentering disposal orbits above LEO: (1) near-circular inclined geosynchronous orbits (IGSOs), (2) eccentric IGSOs (Tundra orbits), and (3) GPS orbits. Long-term disposal orbit propagations over 200-years were performed using the Aerospace high-precision integration code TRACE. Collision probability with operational satellites in the various zones is determined using the Aerospace Debris Environment Projection Tool suite in order to check the trend with time in zone. Results of the study showed that time in zones is less than 25 years when orbital lifetime is less than 200 years, thereby clearing the way for using the new long-term re-entry option relative to this condition. The absolute numerical value of collision probability with operational satellites was found to vary substantially with disposal orbit case and zone. For most cases, results show a clear trend between time in zone and collision probability with some spread. An exception is the GEO zone, for which results show a trend but a wide spread. This may be because most GEO satellites are confined to a ring instead of the shell represented by an altitude zone.

On-orbit hypervelocity breakup models are of increasing importance as the space environment becomes more congested and issues of debris lethality and lifetime become more relevant to space operators. One such fragmentation model is The Aerospace Corporation's IMPACT, which simulates debris from explosions and hypervelocity collisions. IMPACT is a mass-based, semi-empirical model combining empirical distributions for parameters such as number, spreading velocity, fragment dimensions, and area-to-mass ratio (AMR), with physical conservation laws and boundary conditions. The original empirical relationships were based upon limited ground-test data during the early 1990s and were then updated during the late 2000′s following an extensive evaluation of over 11,000 pieces of debris from more than three-dozen historical on-orbit fragmentation events. The availability of extensive fragmentation data from the DebriSat project, including data for smaller debris fragments, now enables a fresh look at the overall model. This paper addresses observations from evaluation of specific components of the model, such as mass and density distributions, fragment shapes as a function of material, and AMR distributions, and discusses the integration of these components into a cohesive model. The combination of detailed ground test results with key parameters available from on-orbit observations provides a fuller understanding of the implications of any on-orbit fragmentation event.

Since 2010 the International Organization for Standardization (ISO) has been publishing a comprehensive set of international standards on space debris mitigation based on guidelines and best practices from the IADC and other bodies. This paper describes the structure and content of the ISO debris standards and how they are evolving to address the needs of space environment sustainability within the context of a rapidly changing space industry. Particular emphasis is placed on ISO 24113 (Space systems – Space debris mitigation requirements), which was published as a third edition during July 2019. The new requirements are compared with those in the previous edition and the rationale for the various changes is discussed. Most notably, the requirement for a spacecraft or orbital stage to exceed a specified threshold for its probability of successful disposal has been made more demanding. With the launch of numerous small spacecraft into LEO and the imminent launch of constellation systems comprising thousands of satellites, it is likely that the ISO debris standards will have to implement even stricter requirements in the future. The challenge is to do this in a way that is fair and reasonable for the entire space industry.

Tundra orbits are inclined, moderately eccentric orbits with a 24-h period. These orbits undergo large excursions in eccentricity due to luni-solar gravity perturbations. For inclinations above 50°, eccentricity can grow to a value that causes perigee to reach the Earth's atmosphere, resulting in vehicle re-entry. In the current study, a range of potential disposal options for Tundra orbits were compared. Disposal options considered include moving to a disposal orbit near the Tundra mission orbit, lowering apogee just below GEO, and moving to near-circular orbits below and above GEO. Comparison metrics include delta-V cost (a measure of maneuver effort) and long-term collision risk. For the eccentric disposal orbit options, initial inclination and right ascension of ascending node (RAAN) were parametrically varied. For the near-circular disposal orbit options, initial argument of perigee and RAAN were parametrically varied. Study results indicate that a near-Tundra disposal orbit option appears to be the favorable selection since it has much lower delta-V cost than the other options and still has the second lowest overall collision risk, well below the threshold of 0.001 that is cited in U.S. standards on debris mitigation and space safety. In terms of orbital sustainability, for all RAAN values there are inclination values that enable re-entry within 200 years.

Existing DOD and NASA satellite breakup models are based on a key laboratory test, the 1992 Satellite Orbital debris Characterization Impact Test (SOCIT), which has supported many applications and matched on-orbit events involving older satellite designs reasonably well over the years. To update and improve these models, the NASA Orbital Debris Program Office, in collaboration with the Air Force Space and Missile Systems Center, The Aerospace Corporation, and the University of Florida, conducted a hypervelocity impact test using a high-fidelity mock-up satellite, DebriSat, in controlled and instrumented laboratory conditions. DebriSat is representative of present-day LEO satellites, having been constructed with modern spacecraft materials and techniques. The DebriSat fragment ensemble provided a variety of shapes, bulk densities, and dimensions. Fragments down to 2 mm in size are being characterized by their physical and derived properties. A subset of fragments will be analyzed further in laboratory radar and optical facilities to update the existing radar-based NASA Size Estimation Model (SEM) and develop a comparable optical-based SEM. Thoroughly understanding size estimates from ground-based optical and radar sensors is one of the key parameters used in assessing the environment and the risks that debris present to operational spacecraft. The data will inform updates to the current NASA Standard Satellite Breakup Model (SSBM);, which was formulated using laboratory and ground-based measurements of on-orbit fragmentation events to describe an average breakup for spacecraft and upper stage collisions and explosions. DebriSat will extend the laboratory data ensemble. The DebriSat shape and density categories provide a baseline for non-spherical projectile hypervelocity impact testing for damage assessment. The data from these tests, simulations, and analyses will be used to update the NASA Orbital Debris Engineering Model (ORDEM) with more realistic simulations of catastrophic fragmentation events for modern satellites and to assess the risk posed by the orbital debris environment. This paper provides an overview of the project, updates on the characterization process, and the NASA analysis status.

Existing DOD and NASA satellite breakup models are based on a key laboratory test, the 1992 Satellite Orbital debris Characterization Impact Test, which has supported many applications and matched on-orbit events involving older satellite designs reasonably well over the years. In 2014, the NASA Orbital Debris Program Office, in collaboration with the Air Force Space and Missile Systems Center, The Aerospace Corporation, and the University of Florida, conducted a hypervelocity impact test using a high-fidelity, mock-up satellite, DebriSat, in controlled and instrumented laboratory conditions to update and improve these models. DebriSat is representative of present-day, low Earth orbit satellites, having been constructed with modern spacecraft materials and techniques. The DebriSat fragment ensemble provided a variety of shapes, bulk densities, and dimensions. Fragments down to 2 mm in size are being characterized by their physical and derived properties. The data will inform updates to the current NASA Standard Satellite Breakup Model (SSBM), which was formulated using laboratory and ground-based measurements of on-orbit fragmentation events to describe an average breakup for spacecraft and upper-stage collisions and explosions. Although individual fragment collection and characterization is still ongoing, this paper will present the status of the DebriSat fragment data analysis including cumulative characteristic length and cumulative mass distributions, area-to-mass distributions, and characteristic length versus mass distributions. Additionally, comparisons to the NASA SSBM will be presented.

Major changes are expected in the space environment due to the imminent launch of NewSpace (sometimes Space 2.0) systems, particularly large low Earth orbit (LEO) constellations that will change the space operations environment. This new activity could overwhelm current space flight safety processes. However, there are many encouraging signs that government, industry, and the space community are acting to address these issues although additional steps remain. These efforts are described, and some additional steps and issues that remain to be explored are discussed.

A growing number of satellites are operating on inclined geosynchronous orbits (IGSOs) with inclination typically much higher than that of traditional geosynchronous orbits (GEOs). Several recent studies have considered the long-term evolution of IGSOs. Unlike traditional GEO disposal orbits, IGSO disposal orbits can undergo large excursions in eccentricity due to the effect of luni-solar gravity perturbations. For specific ranges of initial orbital elements, perigee can reach the Earth's atmosphere, resulting in vehicle reentry. A previously published study by the authors on Tundra orbits, a specific class of IGSO with critical inclination and moderate eccentricity, demonstrated that orbital lifetime can be reduced below 200 years (in some cases below 25 years) and that the corresponding collision probability with background objects can be significantly reduced below that for traditional GEO disposal orbits. This paper presents a study of a broad range of IGSO disposal orbits. Three cases of IGSO disposal orbit were considered: (1) near circular orbits (motivated by the BeiDou and IRNSS constellations); (2) orbits with intermediate eccentricity (motivated by the QZS constellation); and (3) orbits with larger eccentricity (motivated by the Sirius Tundra constellation). For each case, a large number of long-term propagations using the high-precision code TRACE were performed. Disposal orbit initial inclination and right ascension of ascending node (RAAN) were parametrically varied. The Aerospace Debris Environment Projection Tool (ADEPT) suite was used to determine collision probability with inactive background objects and with operational satellites in GEO, medium Earth orbit (MEO), and low Earth orbit (LEO). Study results show that orbital lifetime of IGSOs can be reduced to less than 200 years for all three IGSO cases considered if initial inclination is high enough. The minimum required inclination depends on initial RAAN. An orbital lifetime less than 200 years offers effective reduction of probability of collision with inactive background objects (by 0.7–1.8 orders of magnitude) and can be achieved in a wider disposal orbit design space than would be needed to reduce orbital lifetime to less than 25 years. For near circular IGSOs, collision probability with operational satellites will be higher at high inclinations than at low inclinations, but there are options for mitigation. When inclination is not high enough to achieve reentry, a high eccentricity storage disposal orbit may be a consideration if collision probability is lower than for other options.

On December 22, 2018, the Orbcomm FM 16 satellite experienced a debris generating event. Numerous pieces of debris were cataloged, and this paper documents a forensic analysis of the observed debris. The Aerospace Corporation has developed various techniques to evaluate debris-generating events and determine various characteristics of those events such as spread velocity of the debris pieces, energy of the event, and area/mass estimates of the individual objects. These techniques were applied to the Orbcomm FM 16 event. Analysis indicates the event occurred at ~05:16:55 UT on December 22, 2018, when the satellite was over the Pacific Ocean headed southwards. It was found that the average velocity imparted to the debris was ~91.2 m/sec. Nearly all of the resulting orbits had higher apogees than the main satellite, although a few pieces of debris experienced only a small orbit change in the main orbit when compared to all of the others. What made this event highly unusual is the strong linear correlation in the radial/along-track plane of the delta-V distribution. Coupled with the dominance of the normal component in a single direction (except for one extraneous object), the distribution indicates that the debris pieces were given off in a distinct fan-shape. This fan-shape of debris does not fall into the normal behavior observed in previous explosions or collisions. The source of the event is currently unknown, with explosions typically showing more spherical distributions of debris while collisions are typically conical in distribution. Estimates of the collision energy and areato-mass ratio of the individual pieces are made in an attempt to identify possible causes.