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• Are osynchronous and low Mars orbit satellites evaluated on their EOL opportunities. • Transferring them to graveyard orbits is the only realistic end-of-life measure. • Graveyard orbits prevent interactions with original orbits for at least 200 years. • Derelict spacecraft in areosynchronous orbits must be lowered in altitude by 400 km. • Graveyard orbits for low Mars orbit satellites are 100 km below the original orbits. • For graveyard orbit objects, free tumbling and uncertainties in state are assessed. • Graveyard orbit behaviour depends ongeometry characteristics of spacecraft-Mars-Sun. Mars is expected to become a focal point of exploration (human and robotic) in the coming century, with a very likely need for a robust space infrastructure. Be it communication and navigation satellite constellations or scientific missions in low Mars orbits (LMO) and Areosynchronous orbits (ASO), every individual satellite will have a definitive period of operation after which it becomes derelict. At the end-of-life (EOL) the satellite shall be proactively dealt with in a sustainable manner to protect our access to the space environment of Mars and opportunities to use this. Clearly, impacting Mars or escaping Mars' gravity are no viable options. This paper aims at identifying graveyard orbit solutions in circummartian space for future Mars space debris. Orbital stability for a period of 200 years is studied for Martian orbits using the symplectic integration technique. Extensive validations are performed and propagation and integration settings are tuned to suit a variety of configurations. A plethora of candidate graveyard orbit solutions are found and presented for orbits in the ASO and LMO regimes. For example, it is found that transferring an ASO satellite to 400 km below the nominal orbit altitude would ensure a stability margin of ± 25 km for at least 200 years. Multiple orbital geometry characteristics (combinations of semi-major axis, inclination, right ascension of ascending node), satellite geometries (various values of area-to-mass ratio) and uncertainties are studied to produce a comprehensive analysis of long-term stability of potential graveyard orbits around Mars, making them attractive for such purposes. The protected zones are found to be safe from debris even for an uncertainty in initial eccentricity of 0.01 and variations in cross-sectional area due to uncontrolled tumbling. The overall objective of this paper is to make designers of future missions to Mars aware of the EOL aspects and include this in their mission design proposals at an early stage already. [ABSTRACT FROM AUTHOR]

Satellite formation plays an important role in space missions such as on-orbit services and distributed space facilities. In these missions, hovering in a bounded range is one of the core tasks, but hovering technologies for the chief satellite in an elliptical orbit have been rarely investigated. Given the boundary constraints of relative motion between the chief and deputy satellites, this paper proposes two models for the design and control of hovering formation near an elliptical reference orbit through impulse: (1) the revisiting hovering formation model, which forces the deputy satellite to revisit the same relative position and maintains the formation via impulses; (2) the sequential-impulse-control hovering formation model, which uses a sequence of impulses to ensure the deputy satellite flies within the given boundary. Unlike the revisiting hovering formation model, this model does not require a fixed revisiting point, so the satellite's trajectory is more flexible. The geometrical properties, formation maintenance strategy, and impulse-control solution for both models are investigated in detail. Finally, numerical simulations are conducted to verify the advantages of the proposed methods over other conventional methods, suggesting their wide applications in practical space missions. • Two impulsive-control models are proposed for bounded hovering formation in elliptical orbit. • The proposed methods support in-plane, out-of-plane, and three-dimension hovering formations. • The in-plane and out-of-plane controls are decoupled to decrease control complexity and difficulty. • Different factors on the fuel cost are analyzed to provide a reference for practical engineering. [ABSTRACT FROM AUTHOR]

In terms of cataloged debris produced, the anti-satellite test carried out by Russia, in November 2021, at an altitude of about 480 km, leading to the destruction of the old satellite Cosmos 1408, was the second worst to date and represented the third worst fragmentation in orbit. It generated more than 1/4 of the cataloged debris produced over 55 years by all such tests and almost twice as many as were produced by all previous Soviet tests. After placing this event in its historical context, this paper analyzes in detail how the evolution of the Cosmos 1408 debris cloud affected the environment below 600 km in the first seven months, focusing on the two operational space stations and the Starlink large constellation of satellites. During the first six months following the test, the Cosmos 1408 cloud of fragments nearly doubled the average flux of cataloged objects on the International Space Station and increased by about 3/4 that on China's Tiangong. In the same period, the Starlink large constellation saw an average increase in the flux of cataloged objects of about 20%. Some orbital planes, the "counter-rotating" ones with respect to the Cosmos 1408 debris cloud, were more affected than others, and the affected planes gradually changed over time, due to the differential precession of cloud and constellation nodes. However, being the Starlink constellation 70 km higher up, the flux of Cosmos 1408 cataloged debris steadily decreased over the period analyzed, due to the cloud orbital decay, reducing to just over a quarter of its extrapolated initial value after seven months. • A review of ASAT tests as a source of orbital debris was presented. • The impact of the Cosmos 1408 destruction was put into context. • The cataloged debris flux increase on space stations was computed and discussed. • The debris flux increase and evolution on Starlink orbit planes was analyzed. • A moratorium on this type of ASAT tests was recommended, including low LEO. [ABSTRACT FROM AUTHOR]

The future of deep space communications encompasses a challenging situation where the current facilities used to communicate with different spacecraft may become saturated as a result of an increasing number of missions and their complexity. From this forecast, the present study intends to provide a solution to saturation problems through strategically-located upgradable relays for Earth–Mars communications. The foremost goal of this paper is to quantitatively uncover the potential enhancements coming from relay placement in strategic orbits between Earth and Mars. Herein, two relay configurations – a.k.a. network topologies – are analyzed: the Lagrange-relays network topology and a circular, homogeneously-distributed satellite constellation, acknowledged here as pearl constellation. The first uses the Earth–Sun system Lagrange points L 3 , L 4 and L 5 as potential locations for the relays, whilst the second defines an optimized orbit between Earth and Mars with 3 or 4 relay satellites. To aid in the analysis, the authors developed an open-sourced piece of software that obtains the link availability as well as the data rate at which two nodes may communicate, taking as a reference the Deep Space Network (DSN) for Earth, and the Mars Reconnaissance Orbiter for Mars. For complex topologies with more than two communicating nodes, the software outputs the end-to-end bit rate and optimal communication route at each time step. Moreover, this product is extensible to analyze and optimize any network topology and could be adapted to be used for contact management and mission planning in the future. The results show that the network-topology proposals are an advantageous option to significantly increase the link availability of Earth–Mars communications. Nevertheless, the Direct-To-Earth (DTE) link always outperforms the multi-hop path due to the limited telecommunication system's capabilities of both the spacecraft and the relays. As a result of this, the study includes an analysis on the requirements of the relay's design in order to make the constellation a beneficial and comparable alternative to the DTE link. This way, the proposed network topologies become a suitable option whom to share with the DSN communications workload, providing enhanced bit rates and data volumes as well as higher availability of the communication. [Display omitted] • Deep space relays in strategic orbits help to improve Earth–Mars communications. • Solar conjunctions may be successfully avoided through relays. • Communication availability is significantly improved; up to continuous visibility. • Reasonable gain enhancements could lead to bit rates comparable to the current ones. • Total downlink data volume could be tripled. [ABSTRACT FROM AUTHOR]

Coverage analysis is an essential process when designing a satellite constellation, and the variations of global revisit characteristics are usually elusive in discontinuous coverage problems. In dealing with this, a general method for analyzing the global revisit characteristics of constellations with repeating ground track orbits is proposed in this paper. Firstly, a two-dimensional map based analytical model for the point coverage of constellations is introduced. Then, the variations of revisit characteristics along both the parallel and meridian are analyzed based on the presented analytical model. Finally, numerical examples are carried out. On the one hand, the proposed analytical model for point coverage is demonstrated to be of high precision. On the other hand, all possible positions where the discontinuous changes of revisit characteristics occur can be analytically obtained, which can reveal the discontinuous variations of revisit characteristics along both the parallel and meridian, and can also ensure that all possible global or regional revisit characteristics are counted for a given constellation. The proposed method can contribute to constellation design and coverage performance analysis. • An analytical model for the coverage calculation of satellite constellations is developed. • Conditions for the appearance of the discontinuous changes of global revisit characteristics are deduced analytically. • Variations of the revisit characteristics of satellite constellations along both the parallel and meridian are revealed. [ABSTRACT FROM AUTHOR]

Proliferated low Earth orbit constellations are becoming more common both in reality and in planning. The resilience of these constellations to the radiation effects of adverse solar weather is of interest given that some of them pass over polar regions. This paper stochastically simulates solar weather events over the course of solar cycle 25 in order to quantify radiation effects on a model of Planet Lab's Flock constellation under various design and programmatic decisions. The Flock constellation is a highly proliferated, commercial, and polar low Earth orbit earth observation architecture. Lifecycle and radiation failures are simulated along with a reconstitution system in order to quantify architectural performance as a function of constellation population. This time-based measure of performance is then used to quantify the constellation resilience and allow for programmatic and design variables to be altered in order to observe the effects on the resilience of proliferated constellations against adverse solar weather radiation effects. A sensitivity analysis is performed in order to assess which decisions most improve or degrade resilience. Broadly, the results show that commercial off the shelf components and programmatic decisions such as launch and manufacturing schedules may improve constellation resilience against adverse solar weather effects. More specifically, the results also show that commercial off the shelf components with even basic shielding levels provide high resilience to cumulative total ionizing dose effects and single event upset radiation effects. However, any further reductions to shielding would have significant impacts on single event upset rates whereas total ionizing dose would remain negligible. Furthermore, manufacturing time for Flock satellites is sufficient to maintain a resilient reconstitution strategy. Finally, results show that in order to improve the resilience afforded by a reconstitution strategy, improvements to manufacturing and launch times should be made in parallel, rather than one or the other in isolation. • COTS and programmatics may improve LEO constellation resilience to solar weather. • LEO constellations are highly resilient to even extreme solar weather radiation. • Shielding and launch times are significant for resilience to LEO solar radiation. • Manufacture and launch times should improve parallelly for constellation resilience. • Cumulative radiation in LEO is negligible, primary concern is single event effects. [ABSTRACT FROM AUTHOR]

The emergence of inter-satellite link (ISL) technology, which realizes the high-precision measurement of the inter-satellite distances and forms a spatial geometric constraint within the constellation, enables the high-precision autonomous orbit determination of the GNSS satellites. Firstly, according to the fact that ISL adjustment is singular to the satellite orbit state vector, a distributed autonomous orbit determination algorithm employing the long-term forecast ephemeris to constrain the constellation orientation is designed in this paper. Then, the autonomous orbit determination using 60-day onboard ISL observations of the whole constellation is carried out. Then, the resulting orbits accuracy is evaluated by comparing them with a precise orbit determination using a global reference network with an accuracy of 0.10 m. The results show that the mean root mean square (RMS) of the user range error (URE) over a 60-day period of the 24 BDS-3 satellites in the constellation is 3.22 m. The characteristics of orbit errors for individual satellite are very similar, which is mainly caused by the overall orientation errors of the constellation. After eliminating these orientation errors, the corresponding mean RMS of URE is reduced to 0.68 m. The prediction errors of the Earth rotation parameters and the determination errors of the parameters of the state vectors, which compose the orientation errors, are further investigated. The Earth rotation parameter prediction errors mainly cause errors in the estimates of the longitude of ascending node, of which the 60-day error can reach 2.5e-5°. The determination errors of the parameters of the state vectors -- the inclination and the right ascension of ascending node, can reach 2.4e-5° and 5.1e-5° rad, respectively. [ABSTRACT FROM AUTHOR]

This paper explores a space mission concept in the geostationary orbit (GEO) region to build a catalogue of space objects down to 10 cm and provide satellite inspection services by using a mother satellite (EaglEye) and two smallsats (Harriers). The mothercraft carries the daughter smallsats, is launched in geostationary transfer orbit (GTO), and reaches the operational sub-GEO circular equatorial orbit at 3000 km from the GEO ring. The Harriers are deployed on the same operational orbit and phased using their propulsion system to be positioned at +/- 800 km from the mothercraft. The RSO measurement phase can then start with the objective to build a complete catalogue of all RSOs in less than one year, down to 10 cm size and up to 15 deg inclination. For that purpose, each spacecraft carries the same wide-field imager, providing a field of view (FOV) of about 700 square degrees with a pixel angular resolution of about 13 arcsec. The three satellites continuously and simultaneously observe a common FOV sliding on the GEO ring. We show that the RSO position and orbit are derived from the multispacecraft measurements, and the catalogue construction can be completed in about 7.5 months. The Harriers are designed to operate in either inspection or RSO detection mode. To limit the carried mass and save delta V, each Harrier is sent back to the mothercraft for refuelling as needed. The mission profile is flexible and can alternate inspections and RSO catalogue updates or perform both in parallel. Also, the use of both Harriers for inspections allows 3D stereo imaging. The mission lifetime is eight years with chemical propulsion and exceeds 15 years with electric propulsion while complying with a medium-size launcher. [ABSTRACT FROM AUTHOR]

Nowadays there is international consensus that space activities must be managed to minimize debris generation and risk. The paper presents a method for the end-of-life (EoL) disposal of spacecraft in Medium Earth Orbit (MEO). The problem is formulated as a multiobjective optimisation one, which is solved with an evolutionary algorithm. An impulsive manoeuvre is optimised to reenter the spacecraft in Earth’s atmosphere within 100 years. Pareto optimal solutions are obtained using the manoeuvre Δ v and the time-to-reentry as objective functions to be minimised. To explore at the best the search space a semi-analytical orbit propagator, which can propagate an orbit for 100 years in few seconds, is adopted. An in-depth analysis of the results is carried out to understand the conditions leading to a fast reentry with minimum propellant. For this aim a new way of representing the disposal solutions is introduced. With a single 2D plot we are able to fully describe the time evolution of all the relevant orbital parameters as well as identify the conditions that enables the eccentricity build-up. The EoL disposal of the Galileo constellation is used as test case. [ABSTRACT FROM AUTHOR]

Navigation with Indian Constellation (NAVIC) also known as Indian Regional Navigation Satellite System (IRNSS) is India’s regional navigation system designed to provide position accuracy better than 20 m over India and the region extending to 1500 km around India. The reduced dynamic precise orbit estimation is utilized to determine the orbit broadcast parameters for IRNSS constellation. The estimation is mainly affected by the parameterization of dynamic models especially Solar Radiation Pressure (SRP) model which is a non-gravitational force depending on shape and attitude dynamics of the spacecraft. An empirical nine parameter solar radiation pressure model is developed for IRNSS constellation, using two-way range measurements from IRNSS C-band ranging system. The paper addresses the development of modified SRP empirical model for IRNSS (IRNSS SRP Empirical Model, ISEM). The performance of the ISEM was assessed based on overlap consistency, long term prediction, Satellite Laser Ranging (SLR) residuals and compared with ECOM9, ECOM5 and new-ECOM9 models developed by Center for Orbit Determination in Europe (CODE). For IRNSS Geostationary Earth Orbit (GEO) and Inclined Geosynchronous Orbit (IGSO) satellites, ISEM has shown promising results with overlap RMS error better than 5.3 m and 3.5 m respectively. Long term orbit prediction using numerical integration has improved with error better than 80%, 26% and 7.8% in comparison to ECOM9, ECOM5 and new-ECOM9 respectively. Further, SLR based orbit determination with ISEM shows 70%, 47% and 39% improvement over 10 days orbit prediction in comparison to ECOM9, ECOM5 and new-ECOM9 respectively and also highlights the importance of wide baseline tracking network. [ABSTRACT FROM AUTHOR]

Basing on the theory results considered in the previous papers of the series for traditional one-tiered constellation formed on the orbits with the same values of altitudes and inclinations for all the satellites of the constellation, the method for constellation design using compound satellite structures on orbits with different altitudes and inclinations and synchronized nodal regression is developed. Compound, multi-tiered, satellite structures (constellations) are based on orbits with different values of altitude and inclination providing nodal regression synchronization. It is shown that using compound satellite constellations for Earth periodic coverage makes it possible to sufficiently improve the Earth coverage, as compared to the traditional constellations based on the orbits with common altitude and inclination for all the satellites of the constellation, and, as a consequence, to get new opportunities for the satellite constellation design for different types of prospective space systems regarding increasing the quality of observations or minimization of the number of the satellites required. [ABSTRACT FROM AUTHOR]

The present paper is concerned with the search for orbits that have potential to require low fuel consumption for station-keeping maneuvers for constellations of satellites. The method used to study this problem is based on the integral over the time of the undesired perturbing forces. This integral measures the change of velocity caused by the perturbation forces acting on the satellite, so mapping orbits that are less perturbed, which generates good candidates for orbits that requires low fuel consumption for station-keeping maneuvers. The integral over the time depends only on the orbit of the spacecraft and the dynamical system considered. The type of engine and the control technique applied to the spacecraft are not considered to search for those orbits. It can be a good strategy to be applied for a first mapping of orbits. For this search, it is analyzed the integral of orbits with different values of the Keplerian elements in order to find the best ones with respect to this criterion. The perturbations considered are the ones caused by the third body, which includes the Sun and the Moon, and the J 2 term of the geopotential. The results presented here show numerical simulations to obtain the integral of those perturbing forces for different orbits. The GPS and the Molniya constellations are used as examples for those calculations. [ABSTRACT FROM AUTHOR]