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In order to avoid harmful interference in the radio-frequency spectrum, geosynchronous (GEO) satellite operators coordinate their plans with one another by publicly describing their preferred orbital positions and frequency bands in an assignment process managed by the International Telecommunication Union (ITU), a specialized agency of the United Nations. But once assignments are determined, how well do satellite operators comply with them? This work describes a method for assessing GEO satellite operators' compliance with the physical component of the ITU's satellite network assignments in the geostationary belt using publicly available data. The positions of real, cataloged GEO satellites, measured in longitudinal degrees, are compared with data describing ITU satellite networks published by the Union's Space Services Department to determine whether satellites are operating near an active network assigned to one of their corresponding ITU administrations. Compliance for three satellites that perform relatively frequent longitudinal-shift maneuvers in GEO — Telesat's Nimiq 2 , Russia's Luch (Olymp) , and China's SJ-17 — is assessed over a multi-year study period and discussed. This work illustrates how historical compliance assessment can contribute to the identification and characterization of anomalous satellite behaviors, such as re-positioning the same satellite repeatedly to bring many ITU satellite networks into use, performing non-cooperative close approaches with nearby neighbors, and rendezvous operations, in the Nimiq , Luch , and SJ-17 examples, respectively. • Geosynchronous orbit is the only orbital regime with an existing slotting scheme. • Keeping to assigned orbital positions and frequencies prevents harmful interference. • Once assignments are determined, satellite operators do not always comply with them. • Publicly available data can be used to assess compliance with orbital assignments. • Patterns of violation to guidelines in GEO should inform new rules in LEO. [ABSTRACT FROM AUTHOR]

Space elevators are a promising technology for future space transportation. The primary component of a space elevator is a tether deployed from a geostationary orbit towards the Earth and into deep space. Payloads can be transported using a climber moving along the tether. However, obtaining a deeper comprehension of the precise dynamics of a space elevator is challenging. The tether could have various deformations such as elongation, bending, and torsion owing to its flexible nature and the low level of vibration damping in space. Therefore, understanding those effects on the tether is essential for the design concept of a space elevator. Reproducing the behavior of space tether systems on the ground is challenging, which is primarily due to the scale of the system and the difficulty of accurately replicating a space environment. Thus, high-fidelity numerical analysis methods are required to analyze the dynamic response of space elevators precisely. Previously proposed analysis models of space tether systems with climbers have focused on elongation, bending, and rigid body motion of the tether, and the effect of torsional deformation has been neglected. However, a tether can be twisted easily owing to its low torsional stiffness. This study established high-fidelity numerical simulation models for space elevators considering large deformations including the torsion of a tether. This study developed a flexible multibody model with a moving climber using the absolute nodal coordinate formulation with 14 degrees of freedom. Further, the gravitational perturbation caused by the non-sphericity of the Earth was considered. Using the constructed model, the motion and deformation of nonequatorial space elevators caused by the climber motion and the disturbance in the space environment were investigated. The obtained results indicate that gravitational perturbation can induce torsional deformation, particularly at higher anchor latitudes. • Nonlinear finite element method for numerical simulation model for noneuqatorial space elevators. • Elongation, bending, torsion, and rigid body rotation of a tether considered. • Flexible multibody model with moving climber that considered torsional deformation. • Gravitational perturbation caused by the non-sphericity of the Earth considered. • Model used to investigate motion and deformation of nonequatorial space elevators. [ABSTRACT FROM AUTHOR]

The Japan Aerospace Exploration Agency has developed a power processing unit for a 6 kW Hall thruster, a candidate for propulsion systems for geostationary satellites. Hall thrusters have a discharge instability of tens of kHz, which may damage thrusters, power processing units, and satellite systems. This paper proposes our system for detecting discharge current amplitudes that exceed the specified limit of the thruster. The discharge current is converted into a voltage by the current sensor and is output to a digital controller through an AD converter. The system's innovation includes a frequency prediction model and a damping correction algorithm, which allow the system to accurately detect the peak amplitude of the discharge current without testing to measure the frequency characters of the Hall thruster. The design of the amplitude detection system considers the frequencies of the ionization oscillation and the switching frequency of the power processing unit. The frequency prediction model estimates the frequency of discharge current oscillations from simple parameters: the output voltage of the discharge power supply and the thruster dimensions. The amplitude of the discharge current is damped slightly by the frequency characteristics of the system, but the algorithm corrects this. The results of the frequency prediction model agreed well with the actual measurements. The predicted value was 13.2 kHz, while the actual measurement was 8–15 kHz. The integration test showed that the amplitude detection system identified 36 peaks exceeding the specified limit of 80% peak-to-peak during a 100 ms segment, although the actual number was 34. This difference is acceptable. The proposed amplitude detection system can accurately count the peaks in the discharge current that exceed specifications. • An amplitude detection system detects discharge current oscillations that exceed the specifications for protecting satellites. • The algorithm and frequency response of the system are designed considering the frequency of ionization oscillations. • The frequency is estimated by a simple model with only the output voltage of the discharge power supply and thruster dimensions. • The system detected the amplitude of the discharge current oscillations with an error of 5.8%. [ABSTRACT FROM AUTHOR]

Attention to the method of removing large space debris objects by impacting them with an ion beam has not waned over the past 10 years. The application of this method to the space debris removal from the protected area of the Geostationary Earth Orbit (GEO) seems to be the most effective. This is due to a low energy consumption for removing an object to a safe orbit with ΔV≃11 m/s, which allows the removal of several objects by a single spacecraft. A service spacecraft can be developed for this purpose based on the TelSat geostationary communication satellite technology. At the same time, one of the problematic issues remains that is the development of an injector of a weakly divergent ion beam. The ion beam divergence during injection into the vacuum occurs under the insuperable influence of the ambipolar electric field acting in the beam volume. At the same time, the beam behavior depends on the initial divergence angle stipulated by the conditions of beam formation by the electrostatic acceleration system. This paper considers the results of calculated and experimental studies for the peculiarities of the formation of an ion beam in an ion extraction system with slit and round apertures. The dependence of the initial divergence half-angle on the beam perveance for both geometries was defined. In the operation of the injector designed for the space debris removal, it is necessary to ensure stability of the geometry of the ion-extraction system, which is subject to uneven thermal loading. The resulting thermomechanical deformations of the grids in the technology of ion thrusters are parried by giving the grids a plate shape. Therefore, flat grids made of a carbon-carbon composite material of a new type were used in experiments, which avoided noticeable deformations. The results of experimental investigation of the prototype of ion injector with two configurations of carbon grids (with round and slit perforation) and with the initial divergence half-angle of 2°-4° are given in this paper. The thrust control of an ion beam by changing of the screen grid potential under the condition of a constant beam perveance is considered. It is shown that such regulation is associated with a change in the Mach number, which affects the beam divergence. As an example, the calculation of the dependence of the coefficient k on the distance to the removing object is given, which is considered to be a spent satellite SUPERBIRD-A. • The minimum angle of the ion beam divergence of the ion injector were achieved. • A comparison of ion injector with slit-type and round-type aperture perforation. • The value of coefficient of thrust transfer to the object to be removed. [ABSTRACT FROM AUTHOR]

The design, integration, testing and launch of the first Finnish satellite Aalto-1 is briefly presented in this paper. Aalto-1, a three-unit CubeSat, launched into Sun-synchronous polar orbit at an altitude of approximately 500 km, is operational since June 2017. It carries three experimental payloads: Aalto Spectral Imager(AaSI), Radiation Monitor (RADMON) and Electrostatic Plasma Brake (EPB). AaSI is a hyperspectral imager in visible and near-infrared (NIR) wavelength bands, RADMON is an energetic particle detector and EPB is a de-orbiting technology demonstration payload. The platform was designed to accommodate multiple payloads while ensuring sufficient data, power, radio, mechanical and electrical interfaces. The design strategy of platform and payload subsystems consists of in-house development and commercial subsystems. The CubeSat Assembly, Integration & Test (AIT) followed Flatsat-Engineering-Qualication Model (EQM)-Flight Model (FM) model philosophy for qualification and acceptance. The paper briefly describes the design approach of platform and payload subsystems, their integration and test campaigns and spacecraft launch. The paper also describes the ground segment & services that were developed by Aalto-1 team. • The design and development of multi-payload CubeSat, Aalto 1. • The miniaturized hyperspectral imager based on piezo-actuated tunable FPI. • The de-orbiting payload consisting of coulomb drag force. • The miniaturized low power radiation monitor. • The platform design approach consisting of inhouse development and commercial subsystems. [ABSTRACT FROM AUTHOR]

In this paper, we present an algorithm developed to detect Earth-orbiting object streaks within astronomical images. The NASA Wide-field Infrared Survey Explorer (WISE) database of imagery was selected, in order to assess the viability of infrared wavelengths for the detection and analysis of satellite, rocket body and debris objects. The algorithm was designed to scan a high volume of images for the presence of streaks and then to correlate these to known Earth-orbiting objects without the use of a-priori knowledge of position or orbit parameters of the detected object. As part of the streak analysis process, the algorithm performed a photometric characterisation of the detected objects in all four infrared bands observed by the WISE satellite. The performance of the photometric analysis module was validated against known stellar magnitudes as well as the results of previous published research. • Multi-band infrared measurements analysis of satellites and space debris. • Satellite detection algorithm requiring no prior knowledge of position of orbit parameters. • Infrared colour indices of geostationary satellites are similar to those of main-belt asteroids. [ABSTRACT FROM AUTHOR]

In nominal mission scenarios, geostationary satellites perform end-of-life orbit maneuvers to reach suitable disposal orbits, where they do not interfere with operational satellites. This research investigates the long-term orbit evolution of decommissioned geostationary satellite under the assumption that the disposal maneuver does not occur and the orbit evolves with no control. The dynamical model accounts for all the relevant harmonics of the terrestrial gravity field at the typical altitude of geostationary orbits, as well as solar radiation pressure and third-body perturbations caused by the Moon and the Sun. Orbit propagations are performed using two algorithms based on different equations of motion and numerical integration methods: (i) Gauss planetary equations for modified equinoctial elements with a Runge-Kutta numerical integration scheme based on 8–7th-order Dorman and Prince formulas; (ii) Cartesian state equations of motion in an Earth-fixed frame with a Runge-Kutta Fehlberg 7/8 integration scheme. The numerical results exhibit excellent agreement over integration times of decades. Some well-known phenomena emerge, such as the longitudinal drift due to the resonance between the orbital motion and Earth's rotation, attributable to the J 22 term of the geopotential. In addition, the third-body perturbation due to Sun and Moon causes two major effects: (a) a precession of the orbital plane, and (b) complex longitudinal dynamics. This study proposes an analytical approach for the prediction of the precessional motion and show its agreement with the (more accurate) orbit evolution obtained numerically. Moreover, long-term orbit propagations show that the above mentioned complex longitudinal dynamics persists over time scales of several decades. Frequent and unpredictable migrations toward different longitude regions occur, in contrast with the known effects due only to the perturbative action of J 22. • Gravity harmonics of degree 8, third-body effects and radiation pressure included in propagations. • Orbit plane precession, due to Earth oblateness and third-body effects, is predictable analytically. • Three longitudinal dynamics modes due to degree 3 of the geopotential were found. • An additional dynamic state is enabled by third-body effects. • Irregular transitions between states exist due to lunisolar gravitational perturbations. [ABSTRACT FROM AUTHOR]

The maintenance of a catalog of space objects is integral for Space Surveillance and Tracking (SST). In order to enable safe space operations, this catalog shall be complete, accurate and traceable. One of the challenges to keep custody of active satellites comes from their unknown maneuvers plan that adds significantly to the uncertainty on their precise dynamics. This fact can imply the failure of conventional measurement-to-orbit association methods leading at the best to an inconsistent traceability in the catalog (a physical object represented by two or more catalog objects) or, at worst, to the loss of the tracking of an object. This paper presents a method to tackle the measurement-to-orbit association in the presence of unknown maneuvers. It is based on optimal control theory coupled with the admissible region. In particular, the orbit-to-orbit association problem is solved with an energy-optimal indirect approach, which is then generalized to the measurement-to-orbit association problem using the admissible regions concept. Measurements are supposed to be a collection of time-close optical measurements or tracklet that can be represented by an attributable vector. A broad bench of simulated tests is presented in the case of a geostationary (GEO) maneuvering satellite, and the method is finally applied to a real case of a longitude-shift GEO satellite using the data provided by the TAROT telescope network. • A formulation to solve the tracklet-to-orbit association problem for maneuvering space objects is proposed. • An approach combining optimal control theory and the admissible region's concept is developed. • An indirect optimal control approach is applied to the orbit-to-orbit association considering J 2 effects. • The approach is applied to the case of a geostationary satellite using real data from TAROT telescopes. [ABSTRACT FROM AUTHOR]

In recent years, On-orbit servicing (OOS) is in the ascendant, attracting attentions of a growing number of institutions and commercial space companies. The sustainable development of OOS is heavily relied on its economic benefits. However, the economic feasibility of OOS is not sufficiently analyzed in the existing studies. Based on this, a general economic value analysis framework for OOS is proposed in this paper. Firstly, the maximum price that clients would be willing to pay is obtained by real option method through comparing values of OOS by an alternative scheme. Secondly, the minimum price that the servicer could afford is predicted by Discounted Cash Flow (DCF) method, which is based on a service provider value model. Six necessary factors are considered in this model at the same time, including servicing architecture and its initial construction cost, mission related depreciation, income tax, operating expense, monetary value, as well as technical risks and the corresponding insurance premium. Then, the above two prices are compared to analyze the economic feasibility of OOS. Finally, mission cost, pricing strategies and OOS value drivers are modeled and discussed within the estimated acceptable price range. The results show that the OOS is commercially viable when the initial costs of geosynchronous communication satellites and the servicing architecture are more than $242 M and less than $140 M respectively. The applicability and usefulness of the proposed method are demonstrated by an actual business case analysis of the Mission Extension Vehicle-1 (MEV-1). • An economic value analysis framework for on-orbit servicing (OOS) is proposed. • A modified cost model for service provider is established. • It discusses acceptable service charging, value-based pricing strategy and key driving factors of OOS. • The satellite servicing for GEO satellites is a promising industry. • Appropriate value sharing between clients and service providers is essential for sustainable development of OOS. [ABSTRACT FROM AUTHOR]

To reduce the transfer time and radiation damage of all-electric geostationary satellites, this article develops a metamodel based multi-objective optimization scheme for low-thrust geostationary Earth orbit transfers. A multi-stage control strategy is employed to convert the optimal Earth orbital transfer problem into a parameterized multi-objective optimization problem with minimum total transfer time and proton displacement dose objectives. Based on the notion of multi-model fusion, the multi-fidelity system dynamics and radiation damage models are established to reduce the optimization cost. Furthermore, a novel adaptive Co-Kriging based multi-objective optimization method is proposed to effectively solve the multi-objective optimization problem with multi-fidelity models. In this approach, Co-Kriging metamodels are constructed using the multi-fidelity data to approximate the time-consuming low-thrust transfer models for optimization. And the Co-Kriging metamodels are adaptively refined by the infill samples to explore the Pareto frontier efficiently. Finally, a real-world all-electric geostationary satellite transfer example is investigated. The results show that a number of non-dominated low-thrust trajectories are successfully obtained with limited computational budget. Moreover, the total transfer time and displacement damage dose are significantly reduced after optimization, which illustrates the effectiveness of the metamodel based multi-objective low-thrust transfer optimization scheme. • The system evaluation cost is saved using the multi-fidelity transfer models. • The total transfer time and proton displacement dose are reduced via the multi-fidelity and multi-objective optimization. • The effectiveness and practicality of the proposed method are demonstrated via an engineering example. [ABSTRACT FROM AUTHOR]

The second node of the laser-based European Data Relay System (EDRS-C) was successfully launched on 6th August 2019. The EDRS-C satellite, developed, manufactured and tested by OHB System AG as prime contractor, has been procured in the frame of a Public Private Partnership (PPP) between the European Space Agency (ESA) and Airbus Defence and Space (ADS). The primary objective of the EDRS mission is to provide a data relay service to low-Earth-orbit satellites from geostationary orbit by means of optical and radio-frequency bands. A constellation of geostationary satellites will be linked to a network of ground stations and, using innovative laser communication technology, will receive data from Earth observation satellites from lower orbits and relay it to ground stations in Europe, in near real-time and at a rate of 1.8 Gbit/s. The geostationary EDRS-C satellite features laser communication terminal that significantly differentiates it from conventional telecom satellites. Laser Communication Terminal (LCT) was designed and manufactured by TESAT Spacecom. EDRS-C has been developed on the basis of the SmallGEO platform. In order to accommodate the LCT, adaptations to the existing platform design had been required, which implied a consolidation of the existing platform design, an extension of its competitiveness with reference to the application of optical payloads, and an important milestone in the development and industrialization of the generic SmallGEO platform product line. This paper presents the main design adaptations performed on the SmallGEO platform in order to provide the LCT with an environment that guarantees its full performances. Some selected assembly, integration and test matters specific to optical payloads, and in particular to an LCT, are summarized as well in this paper. • EDRS-C, 2nd node of the European data relay system, was launched on the 6th of August 2019. • EDRS System provides data relay service up to 1.8 Gbit/s for LEO satellites. • Major design challenge was to accommodate a laser instrument on a telecom spacecraft. • Typical verification and testing approach had to be tailored to the laser instrument. • In-orbit testing has confirmed the performances of the satellite as tested on ground. [ABSTRACT FROM AUTHOR]

The rising production rate of space debris poses an increasingly severe threat of collision to satellites in the crowded Geostationary Orbit (GEO). It also presents a unique opportunity to make use of a growing supply of in-space resources for the benefit of the satellite community. "The Recycler" is a mission proposed to source replacements for failed components in GEO satellites by extracting functioning components from non-operational spacecraft in the GEO graveyard. This paper demonstrates a method of analyzing in-space re-purposing missions such as the Recycler, using real satellite data to provide a strong platform for accurate performance estimates. An inventory of 1107 satellites in the extended GEO region is presented, and a review into past GEO satellite anomalies is conducted to show that solar arrays would be in the greatest demand for re-purposing. This inventory is used as an input to a greedy selection algorithm and trajectory simulation to show that the Recycler spacecraft could harvest components for 67 client satellites with its allotted fuel budget. This capacity directly meets the levels of customer demand estimated from the GEO satellite anomaly data, placing the Recycler as a strong contender in a future second-hand satellite-component industry. Propellant mass is found to be a greater restriction on the Recycler mission than its 15-year lifetime — a problem which could be solved by on-orbit refueling. • Satellite databases are critical when designing on-orbit servicing missions. • On-orbit re-purposing missions should focus on solar arrays in Geostationary Orbit. • Multi-target trajectory optimization is used to predict the Recycler's performance. • The Recycler can visit 67 satellites using electric propulsion with its fuel budget. • On-orbit refueling is essential for sustainable satellite-servicing missions. [ABSTRACT FROM AUTHOR]

Quantum key distribution from satellites becomes particularly valuable when it can be used on a large network and on-demand to provide a symmetric encryption key to any two nodes. A constellation model is described which enables QKD-derived encryption keys to be established between any two ground stations with low latency. This is achieved through the use of low earth orbit trusted-node QKD satellites which create a buffer of keys with the ground stations they pass over, and geostationary relay satellites to transfer secure combinations of the keys to the ground stations. Regional and global network models are considered and the use of inter-satellite QKD links for balancing keys is assessed. • A Trusted-node satellite QKD network modelling architecture is described. • Different constellation topologies for delivering encryption keys are evaluated. • Inter-satellite QKD link are considered for rebalancing of keys. [ABSTRACT FROM AUTHOR]

To meet the growing demand for all-electric propulsion satellite, a 6-kW-class Hall thruster was newly fabricated and tested in laboratory. A Hall thruster that has an annular ceramic channel in 150-mm effective diameter and a centered hollow cathode was designed to optimize thrust performance with an appropriate channel-geometry and magnetic-field-topology combination. At an input power of 6 kW, both high thrust (392 mN) and high specific impulse (1940s) was demonstrated for 300 V discharge voltage with xenon propellant. In this case a thrust efficiency of 62.7% was obtained. The thruster also worked in a low power mode at 1.8 kW, in which 92.7 mN and 1870s were available for 400 V operation. In comparison with existing thrusters, this thruster enables lower-power and higher-Isp operation for station keeping with satisfactory discharge stability, which will improve the all-electric propulsion satellite not only by reducing the propellant mass during station keeping but also by increasing the electric power available for mission payloads. Analysis in this study showed that the 6-kW operation is suitable for orbit raising maneuvers of geostationary satellites from a launch orbit to a geostationary orbit because a satellite from 3-ton to 10-ton can be transferred with two- or up to four-6kW-thrusters with a high payload ratio and a relatively short transfer time of 4- to 6-months. After transferred to a geostationary orbit, station keeping operation will be executed by the lower power mode. The Hall thruster will hence play an important role in the field of geostationary satellites. • New 6-kW Hall thruster was tested to obtain 1,900s/390 mN for GEO satellites. • Lower 1.8 kW power operation for station keeping is possible by the same thruster with good discharge stability. • Using this thruster, 3-ton to 10-ton satellite can be transferred from GTO to GEO. [ABSTRACT FROM AUTHOR]

Many defunct satellites in geosynchronous earth orbit (GEO) spin rapidly or have highly evolving spin states, with some transitioning between uniform and non-principal axis rotation (tumbling). It is hypothesized that the observed evolution of some defunct GEO satellites is caused by the Yarkovsky-O'Keefe-Radzievskii-Paddack (YORP) effect and internal energy dissipation. YORP torques are generated by the absorption, reflection, and thermal re-emission of solar radiation and are known to cause secular changes in asteroid spin rates and obliquities. The defunct GOES 8 satellite is particularly notable. This satellite's uniform spin rate rapidly decreased in 2014 and it began tumbling. In this paper, newly analyzed light curve observations of the five nearly identical defunct GOES 8-12 satellites obtained between 2014 and 2018 are presented. The observations show large diversity in evolutionary time histories, with several satellites in consistent slow tumbling, GOES 10 in fast uniform rotation, and GOES 8 transitioning between both. To better understand this diversity, YORP dynamical models are investigated. They reveal that YORP driven spin state evolution is strongly dictated by end of life appendage orientations, which differ among the five satellites. The known end of life configurations are consistent with the observed evolution of GOES 8 and GOES 10. This provides a plausible explanation for the observed spin state diversity. Implications of the observed and simulated evolution for GEO debris mitigation are discussed. • Spin state histories of the retired GOES 8-12 GEO satellites are very diverse. • Simulations show that end of life solar array angles greatly impact evolution. • The observed and simulated evolution of GOES 8 and 10 are consistent. [ABSTRACT FROM AUTHOR]

This paper details the development of a planning algorithm for multiple co-located geostationary satellites to perform station keeping and momentum unloading maneuvers concurrently. The objective is to minimize the overall fuel consumption while guaranteeing a safe separation distance between the satellites within a specific geostationary slot, as well as managing their stored angular momentum to maintain their nadir pointing orientation. The algorithm adopts the leader-follower architecture to define relative orbital elements of the satellites equipped with four gimbaled on-off electric thrusters, and solves a convex optimization problem with inequality constraints, including momentum unloading requirements, to determine the optimal maneuvers. The proposed algorithm is verified, in terms of fuel consumption, constraints enforcement and satellites performance, using numerical simulations that take into account dominant perturbations in the geostationary environment. • Concurrent station keeping and momentum dumping for collocated GEO satellites. • Satellites are equipped with 4 gimbaled on-off electric thrusters. • Convex optimization model for maneuver planning with minimal fuel consumption. • Optimization in receding horizon control form to enhance maneuvering autonomy. • More satellites can be managed in a given GEO slot with minimal fuel consumption. [ABSTRACT FROM AUTHOR]

This paper focuses on geostationary satellites with an electric propulsion system and develops two efficient station-keeping strategies in full and failure modes. High-precision mean orbital elements are calculated by an estimator and used as the inputs of two control strategies. In the full mode, an optimization-based maneuver planner, which requires the computations of future orbits, is proposed to achieve long-term minimum-fuel station-keeping. In the failure mode, a fault detection algorithm is designed to diagnose the thruster failure and compute the orbital deviation caused by faulted thruster. Subsequently, a real-time control law as well as a threshold function is formulated to produce the on-off commands of electric thrusters, which correct the orbital deviation. Finally, numerical simulations are made to validate the accuracy and efficiency of two strategies. • High-precision mean orbital element estimation method is stated. • The fuel optimal station-keeping strategy in full mode is studied. • The fault detection algorithm to diagnose the thruster failure is designed. • The real-time station-keeping strategy in failure mode is developed. [ABSTRACT FROM AUTHOR]

This paper addresses the scheduling problem arising from refueling multiple geosynchronous (GEO) satellites with multiple servicing vehicles [many-to-many (M2M) refueling]. The problem is defined by a set of potential fuel tanker locations, a homogeneous fleet of servicing vehicles with limited capacities, and a set of fuel-deficient GEO satellites with known fuel demands. The objective is to open a subset of fuel tankers, to assign GEO satellites to these tankers and to design vehicle servicing sequences, in order to minimize the total mission costs. To achieve such an economical refueling strategy, the fuel tanker location and routing decision are required to be determined simultaneously. It is shown that this problem can be formulated as a location-routing problem (LRP) which is obviously NP-hard. In order to solve it, an ant colony optimization (ACO) metaheuristic that features a special solution representation scheme is proposed. The proposed algorithm is then tested on two types of instances depending on whether the cost for opening a fuel tanker is considered or not. Numerical experiments show that both the consideration of the fuel tanker opening costs and their locations do affect the refueling schedule. Furthermore, the proposed algorithm outperforms the existing approaches in the M2M refueling mission design, and it is able to detect the optimal refueling strategy on all instances containing 85 GEO targets with 15 potential fuel tankers while previous works stopped at 15 satellites and 2 fuel tankers. • Scheduling of the multiple satellites refueling mission is presented. • Servicing vehicle routing and the fuel tanker location are considered simultaneously. • An ant colony optimization based solving methodology is proposed. • A special solution presentation scheme is applied to improve the computational burden. [ABSTRACT FROM AUTHOR]

The use of Electric Propulsion (EP) on satellites for commercial, defense, and space science missions has been increasing in recent decades, from the first successful operation in 1964 aboard the Zond-2 spacecraft to the present day. This paper provides an overview of the technological and commercial development of EP systems that have been deployed. A review of the early years of EP application ends in 1980, when the first geostationary commercial satellite using EP, Intelsat-V, was launched. Beyond 1980, all EP-based spacecraft deployment data through 2018 are presented, divided by spacecraft type: GEO-synchronous satellite, LEO satellites, deep-space missions and small satellites. To date, a total of 587 spacecraft have been launched with some variant of electric propulsion. During the 1960s and 1970s, all 48 spacecraft using EP were government missions, with the US and USSR leading in the development, production, and flight of these systems. These first platforms included a variety of pulsed plasma thrusters, resistojets, arcjets, ion thrusters and Hall thrusters. The number of GEO satellites with electric propulsion systems has increased significantly since 1981, from an average of less than 5 satellites per year during the 1980s to over 15 in recent years. The corresponding annual fraction of EP based GEO satellite launches, compared to all GEO satellite launches, has increased from 20% during the 1980s to over 40% in recent years. For LEO applications, a gradual increase in the utilization of EP has been realized. Of the 167 EP-based LEO platforms deployed, resistojets were the most prolific legacy thruster type (124 S/C) with Hall thrusters gaining traction in recent years (25 S/C), appearing on 19 of 45 satellite missions in the past decade. Of all EP-based LEO missions, approximately half served as testbeds for new technologies. Through 2018, eight deep space spacecraft with EP have been launched, with the US, Japan, and the European Union leading these efforts. Small satellites are also benefiting from this technology, with 24 EP-based small satellites launched to date. Nearly half of these were launched between 2016 and 2018, demonstrating accelerated growth and a large potential for the future of this spacecraft class. • Overview of all electric propulsion missions, divided into mission classes. • For each mission type statistics are presented on the different thruster subclasses. • Arguments are given for the historical trends of the use of electric propulsion subclasses. • Current trends, for the growth/use of classes of electric propulsion, are analyzed. [ABSTRACT FROM AUTHOR]

The goal of this paper is to investigate the on-orbit servicing (OOS) of geosynchronous satellites based on low-thrust transfers while considering the different demands of the satellites and the effects of perturbations. One or more servicing spacecraft start from a parking orbit and service the satellites in sequence via low-thrust orbit transfers, and return to the parking orbit in the end. The different demands lead to different service requirements of time and mass, while the perturbations affect the control solution of servicing spacecraft during each transfer, and induce orbit drifts of uncontrolled target satellites during the mission. An analytical method that can efficiently calculate the fuel consumption of low-thrust transfer is proposed, and methods for evaluating the orbit drifts of uncontrolled satellites are presented. The impact of perturbations on OOS mission is investigated in detail, showing that non-negligible factors include lunisolar attraction and tessral harmonic. Based on the dynamic models, a string-based simulated annealing algorithm is adopted to schedule the OOS mission. The efficiency of the proposed dynamic models and optimization method are verified through ample simulations. Results show that it is much more efficient and reasonable to take orbit drift calculations as corrections on the scheduling result assuming no orbit drifts. The proposed models and methods make it possible to schedule the OOS mission with low-thrust transfers, while achieving high accuracy under the effects of J2 perturbation, lunisolar attracion, tesseral harmonic and solar radiation pressure. • On-orbit servicing (OOS) of GEO satellites via low-thrust transfers is studied. • Analytical transfer solution is developed to calculate the fuel consumption. • J2, lunisolar attraction and tesseral harmonic affect fuel consumption of OOS. • Various servicing demands lead to different requirements of time and cargo mass. • A string-based simulated annealing algorithm is designed to schedule the OOS mission. [ABSTRACT FROM AUTHOR]

Abstract In the light curve of the geostationary earth orbit satellite (GEOS), there are 3 abrupt changes (AC) that are commonly to be seen. One of the AC that appears a peak near zero phase angle has been proved the specular reflection of solar panels leading to the phenomenon. However, the reasons of another two AC produced in the light curve are still unknown. In this paper, we proposed an advanced light curve model of GEOS and theoretically explained these unknown abrupt changes. Through calculating the satellite TDRS-8's light curve and comparing its actual observation, the proposed model is verified and the causes of the unknown abrupt changes are illustrated. The importance and purpose of this discovery are also emphasized at the end of the paper. Highlights • An advanced light curve model is proposed and the simulation light curve is more in line with the actual measurement. • The proposed model can determine visible facets of a space object of any shapes. • The proposed model take the effects of perturbation forces and solar panel rotation into consideration. • The reasons for the abrupt change in the light curve of satellite TDRS-8 are discovered. [ABSTRACT FROM AUTHOR]

Abstract On-Orbit Servicing (OOS) is becoming more extensively discussed in research and commercial developments. The key point of servicing relies on the relation between two systems, the Servicer executing the tasks and the Client being serviced. With the increase of interest in OOS applications, the exploration of the Client-Servicer relationship in different scenarios becomes a necessary step in servicing implementation. Exploring such a relationship involves translating the information available about OOS into useful inputs for spacecraft design, a challenging task due to the variety of applications and possible interactions between both sides. Then, different modelling and simulation techniques can be used to model and to explore the interaction of such complex systems for different scenarios of servicing. This paper presents an agent-based framework to explore the relations of Servicer and Client for applications of On-Orbit Servicing. To do so, the case for a hypothetical satellite operator is explored, under a specific set of requirements and conditions for different servicing solutions. First, the Agent Based Modelling and Simulation (ABMS) is established based on a specific set of rules, definition metrics and operation characteristics relating Servicer and Client. Then, the basic assumptions of Client, Servicer and servicing operations are discussed and the initial simulation parameters are defined. The ABMS is used to simulate OOS for a fleet of geostationary communication satellites for the cases of Lifetime Extension , Refuel and Rescue and Recover. In the end, the ABMS outputs are used to explore the design aspects of the Servicer and Client satellites at systems level. The results highlight the advantages of having this type of framework for early assessment of OOS under different types of context for both Servicer and Client. The paper concludes with directions of how the framework can be used to explore more complex and realistic scenarios of OOS, and assess their potential benefits. Highlights • Framework to explore the Servicer-Client relationship in On-Orbit Servicing. • Concurrent approach of Client and Servicer points of view is necessary. • Systems' technical and financial metrics build a more solid analysis of servicing. • Consideration of different servicing applications and capabilities. • Parameters to assist during the decision making process towards servicing. [ABSTRACT FROM AUTHOR]

Abstract The possibility to reconfigure satellite formations with utilization of the drag and the solar radiation pressure has been established with previous works. Passive formation reconfiguration maneuvers may be achieved by controlling the attitude of the satellites and varying the area exposed to the perturbation forces. This paper deals with the development of a simple but effective model to understand the possibilities offered by this kind of maneuvers. Limits and advantages of the presented maneuvers are examined considering different reconfigurations maneuvers among close relative trajectories in Low Earth Orbits, Medium Earth Orbits and Geostationary Orbits. The theoretical investigations are confirmed by numerical simulations where the Inverse Dynamics Particle Swarm Optimization is employed. The relative dynamics is simulated with a high-fidelity orbital simulator considering all the perturbations affecting the spacecrafts orbits and the gravity gradient torque influencing the attitude dynamics. Several test cases are reported to underline all the possibilities offered by the proposed maneuvering strategy. [ABSTRACT FROM AUTHOR]

Abstract This paper presents a set of deployment and formation acquisition procedures for spacecraft swarms in perturbed near-circular orbits subject to operational constraints. Specifically, two open-loop command sequences are developed that allow a mothership to deploy a large number of deputies into passively safe formations. These formations guarantee a user-specified minimum separation between all spacecraft in either the orbit plane or the plane perpendicular to the flight direction. Following the open-loop deployment sequence, all deputies are required to passively drift for a specified time in order to accommodate commissioning operations such as de-spin and sensor and actuator calibration. At the end of the commissioning phase, a nonlinear low-thrust control law is engaged to acquire the desired formation and counteract any errors introduced in the deployment sequence. Additionally, analytical lower bounds on the duration of the passively safe relative motion after deployment are derived as functions of deployment uncertainty parameters. These bounds can be used to define requirements for the deployment system and specify the duration of the commissioning phase to ensure passive collision avoidance until actuation capabilities are established. The proposed deployment and formation acquisition procedures are validated through simulations using a high-fidelity numerical orbit propagator. The results of these simulations demonstrate that the proposed procedures provide a means of safely initializing swarm formations at low delta-v cost in low earth orbit. Highlights • Open-loop deployment procedures for spacecraft swarms are proposed. • These procedures ensure passively safe relative motion accounting for uncertainty. • Analytical lower bounds on duration of passively safe relative motion are derived. • Deployment procedures are validated using a high-fidelity orbit propagator. [ABSTRACT FROM AUTHOR]

Abstract The problem of finding the optimal thrust profile of a launcher upper stage is analyzed. The engine is non-re-ignitable and it is continuously thrusting, following either a linear or a bilevel parametric profile, until reaching the targeted coplanar orbit. This problem differs from the classical rocket problem where the thrust level is a time-dependent function varying freely between prescribed bounds. Applying the maximum principle yields an analytical closed-loop solution for the thrust direction. Furthermore the final point is found to be necessarily at an apsis, reached from above in the case of a perigee injection. The optimal control problem reduces to a nonlinear problem with only the thrust profile parameters as unknowns. This formulation eases preliminary design studies aiming at defining the optimum upper stage thrust profile. An application case targeting a geostationary transfer orbit illustrates the solution method. Highlights • The classical rocket problem is revisited considering a parametric optimized thrust profile. • The thrust profile is either linear or bilevel. • Under these thrust profile assumptions, the paper demonstrates that the optimal injection point is the targeted orbit perigee. • The paper demonstrates also that the optimal trajectory finishes with a downwards leg. • These results explain why the injection occurs from above on many launch trajectories. [ABSTRACT FROM AUTHOR]

Abstract An approximate solution to the angles-only initial relative orbit determination problem is developed and evaluated in the context of the two-body problem with J 2 perturbations. The algorithm is non-iterative and requires the singular value decomposition of a 6 × 6 matrix and the solution of a fifteenth-order polynomial. The performance of the algorithm is investigated for non-circular low-Earth orbits and near-circular geostationary orbits with and without measurement error. Highlights • An approximate solution to the angles-only relative orbit determination problem is developed. • The algorithm is developed in the context of the two-body problem with J2 perturbations. • The algorithm is non-iterative and can be used to initialize a Kalman filter. • The performance of the algorithm is evaluated for non-circular LEO and GEO orbits. [ABSTRACT FROM AUTHOR]

Abstract This paper presents the optimization method and results of National University of Defense Technology for the problem A of the 9 t h China trajectory optimization competition (CTOC9). An effective optimization approach using controlled regressive orbits are developed to design the three-satellite formation-flying trajectory for monitoring the beams of 18 GEO satellites. Firstly, a leader-follower formation based on the energy matching condition under J 2 perturbation is used to keep the formation configuration stable for long-term flight mission. Secondly, the optimal formation-flying trajectory is obtained to minimize the mission time by solving the established optimization model for maximizing the minimum monitoring profit of all GEO satellites. Finally, the controlled regressive orbits with multiple flight phases are designed to meet the fuel constraints according to the orbital drift law under J 2 perturbation, and a multi-impulse, perturbed orbital targeting technique is employed to connect the neighboring flight phases. The improved results show that it takes only 20.568 days to complete the mission, indicating that the proposed method performs particularly well on this problem. Moreover, the proposed method can be potentially used in the practical engineering due to its simplicity in optimization calculation and orbital maneuvering strategy. Highlights • A stable configuration is used in the design of formation-flying trajectory. • A trajectory optimization model is established to minimize the mission time. • Controlled regressive orbits are developed for GEO satellite beams monitoring. [ABSTRACT FROM AUTHOR]

Abstract From September 1st 2017 to October 15th 2017, the 9th China Trajectory Optimization Competition (CTOC) was jointly held by the Chinese Society of Theoretical and Applied Mechanics, Beijing Institute of Technology, Beijing Institute of Electronic System Engineering, and Tsinghua Space Center. The competition contains two problems, where problem A focuses on the cluster flight orbit design and optimization for GEO satellite beams monitoring, and problem B focuses on the design and deployment of a satellite navigation regional augmentation system. In this short communication, a brief history of China trajectory optimization competition is reviewed firstly. Then, description of problem A is presented. Finally, the competition results and brief discussion are given. Highlights • History of China trajectory optimization competition (CTOC) is presented. • Detailed description of problem A of 9th CTOC is presented. • The competition results and brief discussion are given. [ABSTRACT FROM AUTHOR]

Abstract This paper presents the crucial method for obtaining results of Tsinghua University in the 9th China Trajectory Optimization Competition (CTOC-9). To monitor the usage of the important frequency resources of GEO satellites, CTOC-9 focuses on the orbit design and trajectory optimization of satellite cluster flight for geostationary orbit (GEO) satellite beams monitoring missions. The three cluster satellite orbits and complex constraints cause much difficulties in the design and optimization. The search for the globally optimal mission trajectory with multi-impulse orbit maneuver is performed in two steps. The first step involves a roughly global search in the search space that has been bounded based on the fuel constraints, the monitoring conditions and the formation flying model of three monitoring cluster satellites. To bound the global search space, several legs are patched together to represent the whole trajectory of monitoring mission, and the impulses occurring in the patch points are estimated by the Hohmann transfer. And the particle swarm optimization (PSO) algorithm is modified and then employed to the multi-leg optimization. The second step involves the local optimization of the multi-impulse patched-conic legs which stand out from the roughly global search. The main task of the local optimization is to solve a series of J 2 perturbed Lambert problems by the homotopic approach. Based on the impulsive results, some specific subtasks are selected and re-optimized to reduce the number of subtasks. Of the many trajectories found, the optimal trajectory submitted for the competition is reported. A post-competition result is reported, which is with the primary objective of 19.61 days, prior to the winner of 20.43 days. Highlights • Formation flying configuration is designed to monitor 18 GEO satellite beams. • Multiple monitoring legs and nominal trajectories are used for global search. • Performance index is designed for the global search with the PSO algorithm. • The impulsive results are locally optimized with the homotopic approach. • Specific subtasks are re-optimized to reduce the monitoring times. [ABSTRACT FROM AUTHOR]

Abstract The problem A of the 9 t h China Trajectory Optimization Competition (CTOC-9) requires monitoring 18 geostationary orbit (GEO) satellite beams using a satellite cluster which has three members. This paper summarizes the solution strategies and gives the best results found at Research Center of Flight Dynamics and Control of NUDT. The visibility conditions of the satellite beams are analyzed firstly. A circular configuration is selected for the satellite cluster due to its uniform monitoring performance, and several repeat-groundtrack orbits are chosen as the cluster's working orbits for their revisit feature. Since the satellite beams points to quite different directions in the space, a combination of two or three repeat-groundtrack orbits is adopted to reduce the mission duration. The optimal combination is found out by using exhaustive method. When the satellites transfer between different working orbits, the Particle Swarm Optimization (PSO) and Sequence Quadratic Programming (SQP) algorithms are used to minimize the fuel consumption. An orbit propagation algorithm that can quickly predict the beam visibility to the cluster is introduced to improve the computational efficiency. Two results were found according to two different orbit combinations. Specifically, when the combination of two repeat-groundtrack orbits is adopted, it requires 24.931 days and 244.531 kg fuel to complete the mission. When the combination of three repeat-groundtrack orbits is adopted, it requires 21.408 days and 450.640 kg fuel. The latter is the best result found by our team for the competition. Highlights • Multiple repeat-groundtrack orbits are used to monitor 18 GEO satellite beams. • Prediction method of beam visibility is used to improve computational efficiency. • Fuel consumption of orbit maneuver is minimized by a hybrid optimization method. [ABSTRACT FROM AUTHOR]

Repeater systems in a geostationary orbit utilizing free-space optical-communication offer great potential to backup, process and archive large amounts of data collected or generated at remote locations. In contrast to existing or upcoming global satellite communication systems, such optical GEO relays are able to provide a huge return-channel data throughput with channel rates in the gigabit-per-second range. One of the most critical aspects of such data uplinks are atmospheric disturbances above the optical ground terminals used to connect to the space segment. In this study, we analyse the design drivers of optical ground stations for land-based applications. In particular, the effects of atmospheric attenuation and atmospheric turbulence are investigated. Moreover, we present implementation ideas of the necessary ground infrastructure and exemplify our results in a case study on the applicability of free-space optical satellite communication to the radio astronomy community. Our survey underpins pre-existing ventures to foster optical relay services like the Space-Data-Highway operating via the European Data Relay System. With well-designed, self-sufficient and small-sized ground terminals new user groups could be attracted, by offering alternatives to the emerging LEO mega-constellations and GEO-satellite communication systems, which operate at low return channel data rates across-the-board. [ABSTRACT FROM AUTHOR]

The use of concurrent engineering in the design of space missions allows to take into account in an interrelated methodology the high level of coupling and iteration of mission subsystems in the preliminary conceptual phase. This work presents the result of applying concurrent engineering in a short time lapse to design the main elements of the preliminary design for a lunar exploration mission, developed within ESA Academy Concurrent Engineering Challenge 2017 . During this program, students of the Master in Space Systems at Technical University of Madrid designed a low cost satellite to find water on the Moon south pole as prospect of a future human lunar base. The resulting mission, The Moon Explorer And Observer of Water/Ice (MEOW) compromises a 262 kg spacecraft to be launched into a Geostationary Transfer Orbit as a secondary payload in the 2023/2025 time frame. A three months Weak Stability Boundary transfer via the Sun-Earth L1 Lagrange point allows for a high launch timeframe flexibility. The different aspects of the mission (orbit analysis, spacecraft design and payload) and possibilities of concurrent engineering are described. [ABSTRACT FROM AUTHOR]

Considering the effects of the gravity gradient, the solar radiation pressure (SRP) and the thermal radiation, the coupling dynamical model is established for a super large flexible beam structure, which moves in the orbital plane around the earth. The effects of environmental disturbances on the orbital motion, the attitude motion and the structural vibration are analyzed simultaneously. Influences of the initial attitude angle, the structure size and other system parameters have been investigated for the coupling dynamics. The results show that the envelop shapes of the vibration responses are dependent on the attitude motion, and the perturbation of orbital motion increases with the growth of the initial attitude angle. When the structural fundamental frequency is close to the attitude frequency with the increase of the structure size, the average orbit altitude would decrease slightly. The main frequency component of the attitude motion decreases and the frequency component related to the structural vibration encounters and grows up. The resonance then leads to system instability. Moreover, the magnitude of the attitude motion can also be influenced by the slightly intense structural vibration, which might be excited by the momentum exchange devices of the attitude control system. In addition, in the presence of the SRP, the orbital radius is perturbed obviously and the eccentricity of orbital motion is generally changed. The effects of the thermal radiation and the gravity gradient are in the same order of magnitude on the structural vibration in the low orbit while the former one turns to be dominant in geostationary orbit. [ABSTRACT FROM AUTHOR]

The sequential use of an electric propulsion system is considered in combination with a high-thrust propulsion system for application to the propellant-optimal Geostationary Orbit insertion problem, whilst considering both temporal and radiation flux constraints. Such usage is found to offer a combined propellant mass saving when compared with an equivalent high-thrust only transfer. This propellant mass saving is seen to increase as the allowable transfer duration is increased, and as the thrust from the low-thrust system is increased, assuming constant specific impulse. It was found that the required plane change maneuver is most propellant-efficiently performed by the high-thrust system. The propellant optimal trajectory incurs a significantly increased electron flux when compared to an equivalent high-thrust only transfer. However, the electron flux can be reduced to a similar order of magnitude by increasing the high-thrust propellant consumption, whilst still delivering an improved mass fraction. [ABSTRACT FROM AUTHOR]

We present and discuss a solution to the growing demand for satellite telecommunication coverage in the high-latitude geographical regions (beyond 5 5 ° N), where the signal from geostationary satellites is limited or unavailable. We focus on the dynamical issues associated to the design, the coverage, the maintenance and the disposal of a set of orbits selected for the purpose. Specifically, we identify a group of highly inclined, moderately eccentric geosynchronous orbits derived from the Tundra orbit (geosynchronous, eccentric and critically inclined). Continuous coverage can be guaranteed by a constellation of three satellites in equally spaced planes and suitably phased. By means of a high-precision model of the terrestrial gravity field and the relevant environmental perturbations, we study the evolution of these orbits. The effects of the different perturbations on the ground track (which is more important for coverage than the orbital elements themselves) are isolated and analyzed. The physical model and the numerical setup are optimized with respect to computing time and accuracy. We show that, in order to maintain the ground track unchanged, the key parameters are the orbital period and the argument of perigee. Furthermore, corrections to the right ascension of the ascending node are needed in order to preserve the relative orientation of the orbital planes. A station-keeping strategy that minimizes propellant consumption is then devised, and comparisons are made between the cost of a solution based on impulsive maneuvers and one with continuous thrust. Finally, the issue of end-of-life disposal is discussed. [ABSTRACT FROM AUTHOR]

Space debris occupies a valuable orbital resource and is an inevitable and urgent problem, especially for large space debris because of its high risk and the possible crippling effects of a collision. Space debris has attracted much attention in recent years. A tethered system used in an active debris removal scenario is a promising method to de-orbit large debris in a safe manner. In a tethered system, the flexibility of the tether used in debris removal can possibly induce tangling, which is dangerous and should be avoided. In particular, attachment point bias due to capture error can significantly affect the motion of debris relative to the tether and increase the tangling risk. Hence, in this paper, the effect of attachment point bias on the tethered system is studied based on a dynamic model established based on a Newtonian approach. Next, a safety metric of avoiding a tangle when a tether is tensioned with attachment point bias is designed to analyse the tangling risk of the tethered system. Finally, several numerical cases are established and simulated to validate the effects of attachment point bias on a space tethered system. [ABSTRACT FROM AUTHOR]

The Inter-Agency Space Debris Coordination Committee (IADC) Space Debris Mitigation Guidelines, issued in 2002 and revised in 2007, address the post mission disposal of objects in orbit. After their mission, objects crossing the Low Earth Orbit (LEO) should have a remaining lifetime in orbit not exceeding 25 years. Objects near the Geostationary Orbit (GEO) region should be placed in an orbit that remains outside of the GEO protected region. In this paper, the impact of satellites and rocket bodies performing End-of-Life (EOL) orbital manoeuvres on the collision risk in the LEO and GEO protected regions is investigated. The cases of full or partial compliance with the IADC post mission disposal guideline are studied. ESA's Meteoroid and Space Debris Terrestrial Environment Reference (MASTER) model is used to compare the space debris flux rate of the object during the remaining lifetime estimated for the pre-EOL-manoeuvre and for the post-EOL-manoeuvre orbit. The study shows that, on average, the probability of collision can be significantly decreased by performing an EOL-manoeuver. [ABSTRACT FROM AUTHOR]

A compensation method is used in Chinese Fengyun-4 satellite to counteracting the line-of-sight influence by attitude motion during imaging. The method is acted on-board by adding the compensation amount to the instrument scanning control circuit. The mathematics simulation and the three-axis air-bearing test results show that the method works effectively. [ABSTRACT FROM AUTHOR]

State-of-the-art all-electric geostationary earth orbit (GEO) satellites use electric thrusters to execute all propulsive duties, which significantly differ from the traditional all-chemical ones in orbit-raising, station-keeping, radiation damage protection, and power budget, etc. Design optimization task of an all-electric GEO satellite is therefore a complex multidisciplinary design optimization (MDO) problem involving unique design considerations. However, solving the all-electric GEO satellite MDO problem faces big challenges in disciplinary modeling techniques and efficient optimization strategy. To address these challenges, we presents a surrogate assisted MDO framework consisting of several modules, i.e., MDO problem definition, multidisciplinary modeling, multidisciplinary analysis (MDA), and surrogate assisted optimizer. Based on the proposed framework, the all-electric GEO satellite MDO problem is formulated to minimize the total mass of the satellite system under a number of practical constraints. Then considerable efforts are spent on multidisciplinary modeling involving geosynchronous transfer, GEO station-keeping, power, thermal control, attitude control, and structure disciplines. Since orbit dynamics models and finite element structural model are computationally expensive, an adaptive response surface surrogate based optimizer is incorporated in the proposed framework to solve the satellite MDO problem with moderate computational cost, where a response surface surrogate is gradually refined to represent the computationally expensive MDA process. After optimization, the total mass of the studied GEO satellite is decreased by 185.3 kg (i.e., 7.3% of the total mass). Finally, the optimal design is further discussed to demonstrate the effectiveness of our proposed framework to cope with the all-electric GEO satellite system design optimization problems. This proposed surrogate assisted MDO framework can also provide valuable references for other all-electric spacecraft system design. [ABSTRACT FROM AUTHOR]

Areostationary satellites are considered a high interest group of satellites to satisfy the telecommunications needs of the foreseen missions to Mars. An areostationary satellite, in an areoequatorial circular orbit with a period of 1 Martian sidereal day, would orbit Mars remaining at a fixed location over the Martian surface, analogous to a geostationary satellite around the Earth. This work addresses an analysis of the perturbed orbital motion of an areostationary satellite as well as a preliminary analysis of the aerostationary orbit estimation accuracy based on Earth tracking observations. First, the models for the perturbations due to the Mars gravitational field, the gravitational attraction of the Sun and the Martian moons, Phobos and Deimos, and solar radiation pressure are described. Then, the observability from Earth including possible occultations by Mars of an areostationary satellite in a perturbed areosynchronous motion is analyzed. The results show that continuous Earth-based tracking is achievable using observations from the three NASA Deep Space Network Complexes in Madrid, Goldstone and Canberra in an occultation-free scenario. Finally, an analysis of the orbit determination accuracy is addressed considering several scenarios including discontinuous tracking schedules for different epochs and different areoestationary satellites. Simulations also allow to quantify the aerostationary orbit estimation accuracy for various tracking series durations and observed orbit arc-lengths. [ABSTRACT FROM AUTHOR]

This paper aims to quantify the responsiveness of two on-orbit servicing infrastructures providing services to multiple serviceable platforms in coplanar medium Earth orbit (MEO) and geostationary orbit (GEO). The platforms to be serviced are assumed to be made of elementary units (EUs). EUs are small standardized structural units capable of aggregating with each other and gathering the key functions of a typical satellite within the size of a 6U cubesat. The first servicing infrastructure considered in this paper, called “ Without Depot” (WoD), includes a launch vehicle and a robotic servicer. The second servicing infrastructure, called “ With Depot” (WD) includes a launch vehicle, a robotic servicer and an orbital depot of EUs. The responsiveness of these infrastructures is quantified using a queueing theory-based stochastic simulation on Simulink taking into account the launch and platform random failures. The metrics used to quantify the responsiveness are the service completion rate and the average waiting time before an EU is replaced over a 10-year period of operation. With respect to those metrics, it is demonstrated that WD is more responsive than WoD. However, WD is also shown to be likely to cost more than WoD. Finally, the sensitivity of the responsiveness of WD to the capacity of the orbital depot is analyzed to find the minimum storage capacity for which WD is the most responsive. This last result could be of interest for space designers dealing with the sizing of on-orbit servicing infrastructures. We believe that the concept introduced in this paper will be a critical milestone in the design of a responsive integrated space infrastructure dedicated to the development and prosperity of a new MEO/GEO economy. [ABSTRACT FROM AUTHOR]

An analysis is performed on the space debris flux determination of geosynchronous (GEO) orbit based on earth-fixed coordinate system. As part of the examination, flux contribution from different components of space debris is presented here in order to clarify the linkage between orbital characteristic and resulting flux distribution. The analysis of GEO flux determination based on earth-fixed coordinate system reveals several new insights: 1) Near the GEO equatorial plane, average flux mainly centralizes in two azimuth ranges: the higher priority one, local azimuth close to±90°, mainly contributed from objects with relatively larger inclination; the lower priority one, local azimuth close to 0°, mainly contributed from objects with relatively smaller inclination. 2) Impact flux against geostationary satellites is not unified by longitude. Longitude-dependent flux determination based on earth-fixed frame reveals that impact flux against geostationary satellite near the geopotential wells (105°W and 75°E longitude zones) is about two to three times larger than average. 3) The comparison of flux determination against certain targets based on earth-fixed coordinate system and inertially-fixed coordinate system shows significant difference. [ABSTRACT FROM AUTHOR]

Dynamic analysis and trajectory tracking of a Tethered Space Robot (TSR) is investigated in this paper. A hybrid controller is used to perform the control task. It consists of two components, the first one deals with librational motion of the tether, while the second one takes care of the manipulator motion. A Nonlinear Model Predictive Control (NMPC) approach is used to control the tether libration; for this purpose, the libration is described by a single degree of freedom and the tether length rate is employed as the input to suppress the librational motion. A modified Computed Torque Method (CTM) is used to control the manipulator motion. The dynamic interaction between the manipulator motion and the librational motion is considered both in the system dynamics and control of the system. Using numerical simulations, performance of the proposed control system is evaluated for end-effector positioning as well as for trajectory tracking for two cases: a Low Earth Orbit (LEO) and the Geostationary Earth Orbit (GEO). [ABSTRACT FROM AUTHOR]

A stable, ultra long-duration high-altitude balloon (HAB) platform which can maintain stationary position would represent a new paradigm for telecommunications and high-altitude observation and transmission services, with greatly reduced cost and complexity compared to existing technologies including satellites, telecom towers, and unmanned aerial vehicles (UAVs). This contribution proposes a lightweight superpressure balloon platform for deployment to an altitude of 25 km. Electrohydrodynamic (EHD) thrusters are presented to maintain position by overcoming stratospheric winds. Critical to maintaining position is a continual supply of electrical power to operate the on-board propulsion system. One viable solution is to deliver power wirelessly to a high-altitude craft from a ground-based transmitter. Microwave energy, not heavily attenuated by the atmosphere, can be provided remotely from a ground-based generator (magnetron, klystron, etc.) and steered electrically with an antenna array (phased array) at a designated frequency (such as 2.45 or 5.8 GHz). A rectifying antenna (“rectenna”) on the bottom of the balloon converts waves into direct current for on-board use. Preliminary mission architecture, energy requirements, and safety concerns for a proposed system are presented along with recommended future work. [ABSTRACT FROM AUTHOR]

The dynamics and lifetime reduction of geostationary transfer orbits (GTOs) are of great importance to space debris mitigation. The orbital dynamics, subjected to a complex interplay of multiple perturbations, are complicated and sensitive to the initial conditions and model parameters. In this paper, a simple but effective non-singular orbital dynamics model in terms of Milankovitch elements is derived. The orbital dynamics, which include the Earth oblateness, luni-solar perturbations, and atmospheric drag, are averaged over the orbital motion of the GTO object, or, as needed, also over the orbital motions of the Moon and Sun, to eliminate the short-period terms. After the averaging process, the effect of the atmospheric drag assumes a simple analytical form. The averaged orbital model is verified through a numerical simulation compared with commercial orbit propagators. GTO lifetime reduction by using the luni-solar perturbations is studied. It is shown that the long-period luni-solar perturbation is induced by the precession of the GTO orbital plane and apsidal line, whereas the short-period perturbation is induced by the periodic luni-solar orbital motions. The long- and short-period perturbations are isolated and studied separately, and their global distribution with respect to the orbital geometry is given. The desired initial orbital geometry with a short orbital lifetime is found and verified by a numerical simulation. [ABSTRACT FROM AUTHOR]

Satellite operations are becoming an increasingly private industry, requiring increased profitability. Efficient and safe operation of satellites in orbit will ensure longer lasting and more profitable satellite services. This paper focuses on the use of a multi-objective evolutionary algorithm to schedule the maneuvers of a hypothetical satellite operating at geosynchronous altitude, by seeking to minimize the propellant consumed through the execution of stationkeeping maneuvers and the time the satellite is displaced from its desired orbital plane. Minimization of the time out of place increases the operational availability and minimizing the propellant usage which allows the spacecraft to operate longer. North–South stationkeeping was studied in this paper, through the use of a set of orbit inclination change maneuvers each year. Two cases for the maximum number of maneuvers to be executed were considered, with four and five maneuvers per year. The results delivered by the algorithm provide maneuver schedules which require 40–100 m/s of total Δv for two years of operation, with the satellite maintaining the satellite's orbital plane to within 0.1° between 84% and 96% of the two years being modeled. [ABSTRACT FROM AUTHOR]

The robust optimal sun-pointing control strategy for a large geostationary solar power satellite (SPS) is addressed in this paper. The SPS is considered as a huge rigid body, and the sun-pointing dynamics are firstly proposed in the state space representation. The perturbation effects caused by gravity gradient, solar radiation pressure and microwave reaction are investigated. To perform sun-pointing maneuvers, a periodically time-varying robust optimal LQR controller is designed to assess the pointing accuracy and the control inputs. It should be noted that, to reduce the pointing errors, the disturbance rejection technique is combined into the proposed LQR controller. A recursive algorithm is then proposed to solve the optimal LQR control gain. Simulation results are finally provided to illustrate the performance of the proposed closed-loop system. [ABSTRACT FROM AUTHOR]

The objective of this paper is to find the minimum-fuel station change for geostationary satellites with low-thrust while considering significant perturbation forces for geostationary Earth orbit (GEO). The effect of Earth's triaxiality, lunisolar perturbations, and solar radiation pressure on the terminal conditions of a long duration GEO transfer is derived and used for establishing the station change model with consideration of significant perturbation forces. A method is presented for analytically evaluating the effect of Earth's triaxiality on the semimajor axis and longitude during a station change. The minimum-fuel problem is solved by the indirect optimization method. The easier and related minimum-energy problem is first addressed and then the energy-to-fuel homotopy is employed to finally obtain the solution of the minimum-fuel problem. Several effective techniques are employed in solving the two-point boundary-value problem with a shooting method to overcome the problem of the small convergence radius and the sensitivity of the initial costate variables. These methods include normalization of the initial costate vector, computation of the analytic Jacobians matrix, and switching detection. The simulation results show that the solution of the minimum-fuel station change with low-thrust considering significant perturbation forces can be obtained by applying these preceding techniques. [ABSTRACT FROM AUTHOR]