Disposal Orbit Research Articles

Reducing reentry casualty risk for spacecraft on long-term reentering eccentric inclined geosynchronous (Tundra) disposal orbits

This paper considers disposal of spacecraft and upper stages on eccentric inclined geosynchronous orbits (eccentric IGSOs), specifically the class known as Tundra orbits. 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 that could be used for Tundra orbits. A condition of this disposal option is that reentry casualty risk be limited. An analysis was performed to assess reentry casualty risk for a generic Tundra spacecraft and typical upper stages on Tundra disposal orbits. From propagation sweeps accounting for eccentricity growth of Tundra disposal orbits, several Tundra disposal orbit cases were selected for reentry analysis. A reentry risk analysis for these cases assuming reentry from near-circular orbits was performed. Results show that predicted casualty risk for the generic Tundra spacecraft and typical upper stages well exceed allowable risk limits in the ODMSP. An analysis of reentry from the selected Tundra disposal orbits accounting for the high eccentricity due to eccentricity growth just before reentry was then performed. Results show that the distribution of reentry points on the Earth can be concentrated over ocean in the southern hemisphere where there is less human population. The generic Tundra spacecraft and one of the upper stages considered were then found to be compliant with the limit of 1 in 10,000 expected casualties in the ODMSP. These results indicate promise for wider usage of the long-term re-entry option in the ODMSP.

Probabilistic assessment of disposal orbit lifetime for CubeSat rideshares on resonant decaying geosynchronous transfer orbits

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.

Altitude optimality boundary of two variants of large space debris removal to disposal orbits

The article considers the mission design problem for removal of large space debris objects from near-Earth orbits. It is assumed that a single active spacecraft-collector (SC) should de/re-orbit several objects. The upper planning level calls for a scenario that specifies the type of repeating operations of an SC. Despite the apparent diversity of these scenarios, they can be reduced to two variants of objects removal to disposal orbits (DOs). In variant I, transfers between objects are effected by an SC with detachable propulsion modules (PMs) onboard. Once the next-in-line object is captured, one module is accommodated on its surface, and, after demating, this PM ensures the relocation of the object to a DO. In variant II, the SC operates as a tug vehicle and transfers the object to DO on its own, and then returns to the next object.A reasonable choice of the removal variant depends on the SC launch mass which should be compared for the same number of de/re-orbited derelict targets. To this end, one should estimate the SC dry mass, the payload mass, and the SC propellant mass. Each of these three terms is nonlinear and depends explicitly on each other, the mass of the captured object, and on the SC maneuvers. Analytical models are proposed for estimation of the SC launch mass for each of the two removal variants. It is also shown that there exists an altitude optimality boundary (the radius of a spherical layer) above which the removal variant II is more advantageous than variant I. Such models are constructed based on several assumptions capable of breaking the iterative search of the SC launch mass.The position of the altitude optimality boundary depends on seven input parameters of the mission: the SC dry mass; the PM dry mass; the mass of the object, the required variation of the semi-major axis Δa to reach the DO, the number of objects removed by a single SC, the average ΔV for a transfer between two objects, and the effective exhaust velocity of SC main engine. In the present paper, four types of objects are considered: heavy and light LV stages in low orbits, upper stages in GEO and GNSS regions. By choosing concrete objects of study one can substantially reduce the solution space by excluding the impossible combinations of input data. By using the developed SC launch mass models, it is possible to formulate removal recommendations for concrete groups of space debris objects. These models were shown as being adequate on the examples of Elsa-M and MRV projects which are due to be implemented in the next one-two years.

A constraint programming framework for preliminary mission analysis: Applications for constellation-servicing active debris removal

Constraint Programming is a classical artificial intelligence paradigm characterised by its flexibility for the modelling of complex problems. In the field of space operations, this approach has been successfully used for mission planning and scheduling. This manuscript proposes a framework that leverages the strengths of Constraint Programming for the preliminary analysis of space missions, introducing some modifications to tailor it to the application at hand. Specifically, it uses constraint propagation and search techniques to thoroughly explore the configuration space of a mission in an efficient manner. Consequently, it is able to quantify the performance of precomputed mission choices with respect to the mission requirements, as well as generate new ones that optimise such performance. The proposed methodology has been particularised for two application cases involving active debris removal missions for large constellations in low Earth orbit, namely, a chaser case and a mothership case. The chaser case considers a servicing satellite that rendezvouses with the failed satellites of the constellation and directly transports them to a disposal orbit. The mothership case comprises a servicing satellite that installs deorbiting kits in each of the failed satellites, except for the one removed in the last place. This way, the servicing satellite will only transport this object, while the deorbiting kits will carry out the disposal of the rest of them. This methodology has been successfully used to evaluate a preliminary mission analysis of both application cases developed under ESA’s Sunrise project.

Evaluation of the effectiveness of the 5-year rule — impact on the orbital environment at each altitude by reducing the post-mission disposal lifetime

This paper examines the effect of shortening the post-mission disposal (PMD) orbit lifetime (remaining orbit lifetime after PMD) from 25 years to 5 or 1 year. Using NEODEEM, the debris evolutionary model developed by Kyushu University and JAXA, the change in the orbital environment at various altitude bands for long-term stability and short-term safety are discussed. It was confirmed that the PMD compliance rate is more important than shortening the orbit lifetime for improving the long-term stability of the orbital environment. In the short term, the 5- or 1-year rule reduced the collision rate and collision avoidance frequency at an altitude below 700 km, more than the 25-year rule, thus shortening the PMD orbit lifetime improves short-term safety, especially at low altitudes. The paper also evaluates the effect of large constellations (LCs) and it was shown that shortening the PMD orbit lifetime of LCs improves short-term safety, but the long-term impact is dominated by the number of LCs that remain at high operational altitudes due to PMD failure. It shows that the post-mission disposal lifetime and required compliance rate should be set according to the number and size of objects expected to be launched in the future.

Experimental study of ion beam interaction with target surface aimed at developing a contactless method for space debris removal by an ion beam

Research and development in the field of space debris (SD) removal technologies is one of the most important and urgent tasks facing all spacefaring nations. By now, a large number of different concepts of removal have already been proposed. One of them is the SD removal by ion beam generated by an ion source installed on board a service spacecraft (SSC). A detailed understanding of the process of ion beam interaction with the SD surface is required to develop an optimal control algorithm for the virtual SSC-SD mated configuration in order to improve this method of SD removal. In addition, when designing the SSC, it is necessary to take into account back flows of the material sputtered from the surface of removed SD by the ion beam, which can contaminate the SSC operating surfaces during the removal maneuver. Within the framework of the presented work, experimental studies were performed to determine the coefficient of accommodation of the ion impulse during the ion beam interaction with the surface of the target simulating the external surface of the removed object. Besides, the flow density of the sputtered target material was experimentally evaluated during its exposure to the ion beam. To do this, we used the radio-frequency ion source (injector) with a beam diameter of 160 mm, which is supposed to be used as an actuating unit on board the SSC designed for transporting the SD objects to the disposal orbit or to the orbits with limited life. When measuring the ion beam accommodation coefficient, the presence of additional force arising on the target due to the flow of sputtered material leaving its surface, and the ion-electron emission process on the target, which could distort the experimental data, were taken into account. To estimate the magnitude of the flows of sputtered target material under the influence of the ion beam, a special experimental design was developed for simulating the presence of a point source of sputtered material. The results obtained in this work will be used for modeling the impact of the ion beam on the surface of real SD objects with complex geometry using the finite element method, for refining the control algorithm of the virtual SSC-SDO mated configuration and for developing proposals on the SSC layout, which should reduce or completely eliminate the negative impact associated with the material sputtering from the SDO surface onto the SSC elements. Successful implementation of the developed method of SDO removal will contribute to improving the safety of space activities for all countries and to the development of the rocket-and-space industry in the Russian Federation, in particular.

Mathematical modeling of the ion extraction system of a radio-frequency ion source

In order to ensure space debris removal from the near-Earth space, the world scientific community is developing various technical means potentially capable of transporting the spent spacecraft to disposal orbits. The so-called Ion Beam Shepherd is among them. According to this method, in order to transfer momentum efficiently from an ion beam generated by a radio-frequency ion source mounted on board a service spacecraft, it is necessary to ensure the generation of an ion beam with the minimum divergence angle. In order to provide generation of a highly collimated ion beam for a system of contactless removal of technogenic space debris, an ion extraction system was modeled with the aim to study the influence of its geometric parameters on the angle of ion beam divergence at the ion source exit plane. Based on the preliminary study, a model of 28 full factorial experiment was developed for two types of electrodes in an ion extraction system. The regression coefficients have been defined that allow, without performing a large array of numerical studies, to assess the influence of various geometric parameters on the ion beam divergence angle.

Contactless space debris removal from the geostationary orbit protected region

To verify the possibility of space debris removal by an ion beam, the problem of mission analysis for the motion of virtual configuration comprising a service spacecraft and space debris was solved. It is shown that using the state-of-the-art technologies it is possible to develop a service spacecraft equipped with a system for contactless removal of space debris based on a high-frequency ion source and a compensating electric propulsion system comprising a pair of stationary plasma thrusters. Such a system could provide removal of up to ten space debris from the protected area of the geostationary orbit to disposal orbits.

Promising methods for improving the radio-frequency ion thruster performance

Improving the reliability and efficiency of available electric propulsion schemes is one of the priority areas of space technology development. In particular, this should be done for ensuring the guaranteed removal of spacecraft to a disposal orbit at the end of their active life, which, in turn, will slow down the growth in the number of space debris objects in the most exploited regions of near-Earth space. The paper considers ways to improve the power and mass efficiency of one of the modern electric propulsion types - a radio-frequency ion thruster with an inductive radio-frequency discharge. Using an engineering physico-mathematical model, computational studies were performed to optimize the shape of the radio-frequency ion thruster main components, namely the discharge chamber and the grids of ion-extraction system. The thrust of the thruster was used as the optimization criterion; it was determined from the calculated ion density distributions in the radio-frequency discharge plasma. Such an approach to refine the performance improvement methods has not previously been applied to radio-frequency ion thruster. The ion density distributions, as well as the thruster integral characteristics determined from them, such as the thrust, total ion current from the thruster, propellant utilization factor, ionization cost and a number of others, were calculated for two actual standard sizes of radio-frequency ion thruster with the beam diameters of 80 mm and 160 mm, which can be used to solve the problems listed above. Besides, for the defined optimal configurations of thruster with two standard sizes the calculations were made to assess the possibility for increasing the thruster integral characteristics by applying an additional magnetostatic field in the radio-frequency discharge region. The study revealed the possibility of noticeable performance improvement compared to the basic radio-frequency ion thruster design with hemispherical discharge chamber and flat grids of the ion-extraction system.

Analysis of ETRSS-1 on-orbit performance and anomaly management

The Ethiopian remote sensing microsatellite, weighing 65 kg, was successfully launched into sun-synchronous orbit at an altitude of 628 km in 2019. The satellite has a three-year lifespan and employs a maneuver that minimizes the orbit perigee without adjusting the orbit apogee, resulting in an eccentric disposal orbit, with the perigee altitude selected to ensure re-entry into the Earth’s atmosphere within 25 years. This study presents an overview of the ETRSS-1 satellite system, including its subsystems alongside the hardware utilized during their development, as well as an analysis of its on-orbit performance. Furthermore, the spacecraft’s electro-optical multispectral camera and its ability to capture remote sensing data while adhering to appropriate operational constraints, as well as its imaging mission techniques, various types of failure modes, and anomaly detection detection techniques, will be investigated.

Енергетичні витрати на переміщення об’єктів космічного сміття з низьких навколоземних орбіт на орбіти утилізації

The ever-increasing clogging of near-Earth space by space debris objects of various sizes significantly limits the possibilities of space activities and poses a great danger to the Earth’s objects. This is especially true for low orbits with altitudes up to 2,000 km. The risk of collision of operating spacecraft with space debris threatens their functioning in near-Earth space. To control space debris, use is made of active and passive methods of space debris removal from operational orbits. At present, promising means of space debris removal are a space debris transfer to low-Earth orbits with a lifetime of less than twenty-five years, a transfer to a junk obit, and in-orbit utilization. According to the latest recommendations, space debris objects moved to low-Earth orbits should have a lifetime of less than twenty-five years. In the dense atmosphere, small space debris objects usually burn up completely, while large ones burn up only partially and may reach the Earth. Since space debris motion in the atmosphere can only be predicted with large errors, a timely and accurate prediction of the place and time of fall of large space debris objects onto the Earth is impossible. Space debris objects can remain in junk orbits for hundreds of years without interfering with space projects. This method of space debris removal reduces the risk of collision with space debris objects in the initial orbit, but increases it in the junk one. According to the concept of in-orbit utilization, space debris is considered a resource for the in-orbit industry. An active space debris removal involves high energy expenditures of service spacecraft. In this regard, the task of their estimation becomes important. The goal of this paper is a comparative assessment of the energy expenditures for moving space debris objects into utilization orbits using service spacecraft with electrojet propulsion systems. The problem is solved using methods of flight dynamics, averaging, and mathematical simulation. The novelty of the obtained results lies in the development of a ballistic scheme and a fast procedure to calculate energy expenditures for moving space debris objects to a disposal orbit using service spacecraft with constant low-thrust electrojet propulsion system. The procedure may be used in substantiating and planning space debris transfer from low-eccentricity low-Earth orbits to utilization orbits.

Application of additional inflatable aerodynamic device to ensure the required degradation of the disposal orbit of large-size space debris

The paper studies the effectiveness of joint application of active and passive space debris removal technologies in low-Earth orbits.An active spacecraft-collector (SC), equipped with several detachable thruster deorbiting kits (TDKs), ensures the capture of the rocket stage. The TDK of dry mass about 90 kg should be fixed in the stage nozzle. After undocking from the SC, the TDK executes a deorbiting velocity impulse in order to form an elliptical disposal orbit with a predicted lifetime of 25 years in compliance with ISO standard requirements. However, many experts have recently called for a significant enhancement of these requirements.One of the ways to reduce the lifetime of the resulting disposal orbit is to equip the TDK with an aerodynamic decelerator, which has a shape of a spherical balloon. After the deorbiting impulse is applied, the balloon is supercharged by the gas remaining in the pressurized-propellant feed system of the TDK. Rigidization of the balloon shell is carried out by supercharging it above the yield point. The most rational balloon diameters are chosen based on the analysis of its effect on the lifetime of the disposal orbit.If the detachable TDK is equipped with an inflatable balloon, one can consider the combined and hybrid braking schemes. According to the first scenario, the balloon of diameter 12–14 m (“Zenith-2” case) or 7–8 m (“Kosmos-3M” case) is inflated in the already formed 25-year disposal orbit reducing its lifetime to about 5 years, while the dry mass of each TDK increases by 28 kg and 10 kg, respectively. Within the second scenario, a noticeably smaller deorbiting impulse results in a hybrid disposal orbit, which exists for 25 years due to subsequent inflation of the balloon of almost the same size. An analysis shows that the combined braking scheme is more preferable: with an insignificant increase in the SC launch mass, it does not create additional risks of failure and quickly deorbits the object from the most populated altitudes.

Исследование проектно-конструкторского облика декартовой двигательной установки автономного стыковочного модуля для захвата мишени

The concept of building an orbital complex for operational interception of a target is proposed, including in its composition the method of ballistic construction of operational interception, recommendations to the design appearance of the Cartesian propulsion system of the autonomous docking module, functioning at the stage of close-in guidance during tether delivery from the upper stage to the target, at the stage of capture and transfer of the target to the disposal orbit as part of the rotating tether space system. As a criterion for selecting the Cartesian propulsion system design (fuel reserves, thrust of each chamber and their number, minimum change in the coordinates of the autonomous docking module center of mass over the entire interval of Cartesian propulsion system operation), the minimum mass of the autonomous docking module is taken. For comparison, two methods of constructing the ballistic scheme of target intercept are considered: the classical method, with the orbital complex entering the target orbit with zero relative velocities, and the method based on the use of rotating tether space system, with the orbital complex entering the target intercept orbit with velocities at the rendezvous point of up to 200 m/s.

Space debris removal – Review of technologies and techniques. Flexible or virtual connection between space debris and service spacecraft

The term “space debris” refers to spacecraft not removed from orbit at the end of their service life, upper stages in the geostationary orbit region, as well as to the fragments of spacecraft and upper stages formed as a result of deliberate or accidental collision of spacecraft and upper stages with each other or with natural space debris.The problem of removing space debris from outer space is a global problem. Many countries are realizing projects on space debris cataloging, various technical means are being studied for space debris removal into graveyard orbits with parameters agreed upon by the international community. Various countries conducting space exploration have adopted special standards and guidelines for preventing the space debris formation [1,2,3,4].The report [5] emphasizes that contamination of the near-Earth space is growing steadily. The probability of spacecraft collisions with each other is significantly increasing. This can lead to the functioning spacecraft characteristics degradation and even to complete loss of performance.The article presents a brief overview of the methods and techniques that can be used to space debris removal into disposal orbits using flexible or virtual connection between space debris and service spacecraft.

Analysis of ODMSP-compliant near-circular GPS disposal orbits and resulting long-term collision risk

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 0.0003 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 0.0003 and initial SMA uniformly spread over 363 km has the lowest collision probability for future disposed GPS satellites.

Long-Term Orbit Prediction and Deorbit Disposal Investigation of MEO Navigation Satellites

With the increase in satellites in the medium Earth orbit (MEO) region, there should be a focus on orbit safety in the MEO region. A safe orbit disposal strategy is necessary to maintain the sustainability of the MEO region. This paper focuses on long-term evolution modeling, safety analysis of MEO objects, and different disposal techniques for end-of-life BDS-2 MEO satellites. On the one hand, a long-term numerical evolution model is established, and mean equinoctial elements are adopted to propagate a long-term orbit. Long-term evolution for the MEO region over 100 years is carried out, including the Galileo, BDS, GPS, and GLONASS constellations. The earliest orbit intersection time with other global navigation satellite system (GNSS) constellations is put forward. On the other hand, a dynamic model and an optimization model for disposal orbit are established, which minimize the eccentricity growth within 200 years and the fuel consumption for maneuvering to the disposal orbit. The bounds for the disposal region of BDS MEO satellites are also proposed, which consider the measurement and control error of BDS MEO satellites and the eccentricity bounds for end-of-life BDS MEO satellites. A genetic algorithm is adopted to optimize the orbital elements for end-of-life BDS MEO satellites. In addition, two disposal cases, namely, upraising and reducing the orbit, for end-of-life BDS MEO satellites are simulated. The long-term evolutions for the disposal of orbital elements within 200 years are implemented, and the fuel consumption is calculated. The results show that the current MEO region is relatively safe and that the eccentricity is the most important factor that influences the long-term evolution of safety analysis for BDS MEO disposal orbits. Upraising the orbit is safe for end-of-life BDS MEO satellites. This investigation provides the theoretical foundation for investigating the long-term evolutionary mechanisms of the MEO region and references disposal strategy analysis for decommissioned navigation satellites, and the spent upper stages for other GNSS constellations.

Анализ состояния разработок средств очистки орбит в околоземном космическом пространстве от объектов крупногабаритного космического мусора

The state of near-Earth space pollution is considered, including the dynamics of growth of this pollution. Presented current information on the launch of nano-satellites, which is an additional factor in the pollution of low earth orbit. On the basis of open information, the analysis of the development of Russian and foreign approaches to the most polluted orbits cleaning, including technological experiments with the use of various means of capture is carried out. For different capture devices the options of towing space debris objects to the disposal orbit are considered.

МОЖЛИВОСТІ ЗАСТОСУВАННЯ СОНЯЧНОГО ВІТРИЛЬНОГО ПРИСТРОЮ ДЛЯ ВІДВЕДЕННЯ КОСМІЧНИХ ОБ’ЄКТІВ ІЗ НАПІВСИНХРОННИХ ОРБІТ

The current state of the near-Earth space environment poses a threat to spacecraft moving not only in low and geostationary but also in semi-synchronous orbits. The main methods that can be used in these orbits for space debris removal are the active method using a rocket propulsion system and the passive method using a solar sail device. The last method is of particular interest because of the low overall mass and mass characteristics of the removal system, as well as the availability of unlimited solar light radiation, which can be converted into a driving force. Solar sails are currently undergoing experimental testing in the near-Earth space, but they have not been used for medium and high orbits. Two constructive variants of solar sails construction are considered: frame and rotor. The article is devoted to the evaluation of the possibility of removing space objects from semi-synchronous orbits using a solar sail. Two variants of the removal scheme are considered: into a circular disposal orbit 200 km below the target orbit and into the dense layers of the Earth's atmosphere. A computer simulation of the removal is performed using the example of the Galileo global positioning system spacecraft with a frame-type and rotor-type solar sail. The dependence of the orbit altitude at the ending of removal on the solar sail area was obtained. The values of the solar sail area required to provide for the maneuvering and the weight of the structure were determined. It was demonstrated that the total mass of the frame-type solar sail is 9 kg (1.3% of the spacecraft mass) and that of the rotary-type solar sail is 11 kg (1.6%). As applied to disposal into dense layers of Earth's atmosphere for frame-type solar sail - 52 kg (7.4%), for rotary-type solar sail - 25.5 kg (3.6%). It follows from the obtained results that for moving to the disposal orbit 200 km below the target orbit it is more advantageous to use a frame-type solar sail, and in the dense layers of the atmosphere - the rotor type.

Numerical simulation of high-velocity impact effect on spacecraft structure

Anthropogenic space debris pose a serious threat to operation of spacecraft (SC), including a real threat of damage and destruction of Soviet spacecraft carrying Nuclear Power Systems (NPS) onboard, which ceased their active functioning in orbit thirty years ago. Collisions of space debris with elements of SC structure may both have catastrophic consequences and cause local damage (for example, penetration of thermal control elements), resulting in the loss of spacecraft ability to function, or loss of some of their functionality. In the case of SC carrying NPS, an impact on the structure may result in debris with residual radioactivity, which calls for a continued focus on the status of this type of spacecraft in disposal orbits. In order to predict generation of space debris in near-Earth space, to evaluate SC survivability, as well as to develop engineering facilities for protection, the software package REACTOR was developed, which makes it possible to use a desktop computer to solve the problems of impact effects on complex technical objects including the function of computing the number of generated fragments (debris) and their velocity spectra. As a subject for research, a schematic SC model and a space NPS REACTOR model were constructed, calculations have been performed. As a result of calculations, qualitative and quantitative characteristics of the resulting fragments were obtained. Key words: space debris, spacecraft, collision, software package, computer, fragments, models, calculation nuclear power system.

Assessment of time spent in the LEO, GEO, and semi-synchronous zones by spacecraft on long-term reentering disposal orbits

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.