Orbit Selection Research Articles

Optimizing nanosatellites Earth observation missions: Orbit design for global coverage and pre-launch cloud detection dataset preparation

Optimizing Earth observation nanosatellite missions requires careful orbit selection to ensure global coverage and high image quality. The main challenge is to define an optimal orbit that maximizes image quality while achieving comprehensive Earth coverage. To address this, an orbit definition model that accounts for the J2 perturbation was developed, evaluating key parameters such as swath width, spatial resolution, time resolution, and tilt angle. The model was applied within the context of a 3U university nanosatellite equipped with the Gecko imager. The analysis identified 551 km as the most suitable altitude for achieving comprehensive coverage. At this altitude, the swath width is 88.16 km, and the distance between adjacent ground tracks is 86.33 km, satisfying the condition for global coverage with nadir pointing. The revisit time is 31 days, ensuring regular coverage of the Earth's surface. Results were validated through Systems Tool Kit simulations, showing minimal differences compared to the MATLAB model, underscoring the model's accuracy and reliability. To further optimize the mission's effectiveness and bandwidth utilization, images captured in this orbit can be processed onboard to ensure that only relevant and useful data are transmitted to the ground station. To support this, a dataset of 900 images was generated at the optimal orbit altitude, specifically tailored for a cloud detection application using the Gecko imager. The findings validate the effectiveness of the proposed orbit design model and demonstrate the feasibility of pre-launch dataset preparation, paving the way for future onboard implementations and in-orbit real-time applications.

ORBIT SELECTION OF THE SPACE INDUSTRIAL PLATFORM WITH DISTRIBUTED ELECTRICAL-POWER SYSTEM MODULES

Space industrialization is one of the prospective directions in modern aerospace science and engineering for space exploration of new resources and habitats. The key issue is to provide industrial space modules with the required amount of electricity needed. One type of power supply for such modules is the use of distributed power systems, which consist of constellations of spacecraft with contactless power transmission. Given this, the problem of rational orbit selection for their dislocation arises. Considering these problems, the methodology for orbits selection of the space industrial platform with distributed electrical-power system modules is proposed in the paper. This methodology includes orbital translation, attitude, relative dynamics estimation for each power satellite, and its corresponding orbit optimization algorithm. The orbit optimization algorithm includes statistical processing and elements of gradient and coordinate descent methods, allowing us to determine the most significant parameter influencing the duration of the contactless power transmission session. Also, quaternion mathematics is used to estimate the dynamics in the program parameters for targeting the transmitter spacecraft antenna to the receiver spacecraft rectenna. With the approaches mentioned above, the methodology proposed in this paper allows us to form the requirements for the power satellites’ attitude and orbit control system to improve the process of selecting corresponding design parameters of such systems. Thus, the usage of the proposed methodology can allow the designing of the power satellites’ attitude and orbit control system in the conceptual stages of designing.

Distant retrograde orbit baseline generation considering solar eclipse mitigation

Mitigating solar eclipses presents a formidable challenge in the design of spacecraft orbits, particularly in the case of a distant retrograde orbit (DRO), which is susceptible to significant solar eclipse threats. This paper undertakes a comprehensive analysis of the occurrence characteristics and patterns of both Moon and Earth eclipses on a DRO. The proposed approach involves a two-stage design procedure, encompassing initial orbit selection and orbit extension. The initial orbit selection method primarily focuses on mitigating Earth eclipses by employing the out-of-plane motion inherent in a DRO. The study meticulously explores the influence of three pivotal elements — amplitude, period, and initial phase — on the efficacy of Earth eclipse avoidance. Concurrently, when faced with unavoidable solar eclipses of the initial orbit, a series of tiny maneuvers are systematically applied to extend the baseline orbit. The parameter design space is discretized, and a tree search method is employed to identify viable maneuver sequences. The research successfully achieves the long-term stabilization of the baseline DRO, ensuring that solar eclipses do not exceed 2 h in 10 years. The proposed method’s robustness and applicability are substantiated through validation under the ephemeris model.

Effects of Earth’s Oblateness on Black Hole Imaging through Earth–Space and Space–Space Very Long Baseline Interferometry

Earth-based very long baseline interferometry (VLBI) has made rapid advances in imaging black holes. However, due to the limitations imposed on terrestrial VLBI by the Earth’s finite size and turbulent atmosphere, it is imperative to have a space-based component in future VLBI missions. This paper investigates the effect of the Earth’s oblateness, also known as the J 2 effect, on orbiters in Earth–space and space–space VLBI. The paper provides an extensive discussion on how the J 2 effect can directly impact orbit selection for black hole observations and how, through informed choices of orbital parameters, the effect can be used to a mission’s advantage, a fact that has not been addressed in previous space VLBI investigations. We provide a comprehensive study of how the orbital parameters of several current space VLBI proposals will vary specifically due to the J 2 effect. For black hole accretion flow targets of interest, we demonstrate how the J 2 effect leads to a modest increase in shorter-baseline coverage, filling gaps in the (u, v) plane. Subsequently, we construct a simple analytical formalism that allows isolation of the impact of the J 2 effect on the (u, v) plane without requiring computationally intensive orbit propagation simulations. By directly constructing (u, v) coverage using J 2-affected and J 2-invariant equations of motion, we obtain distinct coverage patterns for M87* and Sgr A* that show extremely dense coverage on short baselines as well as long-term orbital stability on longer baselines.

A Survey of Orbit Design and Selection Methodologies

A Survey of Orbit Design and Selection Methodologies

Energy Balance Analysis for Water Resources Satellite Operation Orbit Selection

In order to supply enough power for the satellite mission and at the same time to suppress cost increase through over-design, it is necessary to select an appropriate solar array and battery capacity. In the initial stage of satellite design, the required capacity must be analyzed to determine the solar array and battery model, which will be reflected throughout the overall satellite design. This study verifies that the CAS500 satellite platform can provide the power required for the mission in the initial stage of water resources satellite, and furthermore, it found the solar panel and battery capacity required for the water resources satellite. To this end, it was confirmed that the energy balance was satisfied by selecting the worst case one-day mission scenario of the water resources satellite under various conditions.

A reference architecture for orbiting solar reflectors to enhance terrestrial solar power plant output

Orbiting reflectors offer the possibility of illuminating large terrestrial solar power plants to enhance their output, particularly at dawn and dusk when their output is low but energy spot prices can be high. While the concept of orbiting solar reflectors has been considered in various forms in the past, there is now a timely overlap of rapidly growing global demand for clean energy services, falling launch costs through reusability and the emergence of in-orbit manufacturing technologies to enable the fabrication of large, ultra-lightweight space structures. This paper provides an end-to-end analysis of a possible minimum initial architecture to deliver such global clean energy services. The analysis will cover orbit selection, attitude control requirements, structural analysis and economical viability, followed by a discussion on regulatory issues, future improvements and further applications.

A Study of Autonomous Small Satellite Constellations for Disaster Management and Deep Space Strategy

The complex and dynamic space environment is both exciting and challenging in this NewSpace era. In particular, Low-Earth Orbits are being realigned and reinvented for various purposes using suitable technological advancement. This paper is focused on the major parameters that can be analyzed to attain orbit control and autonomy of small-satellite constellations for real-time applications. By applying industry experience to graduate research, this work addresses the related concerns in a realistic manner. Currently, global small-satellite constellation solutions are too expensive and inaccessible for many nations to help in their data reception requirements. This issue was addressed, and some of the main aspects relating to low-cost and high-benefit technical synthesis, in addition to utilization for deep-space missions, were also discussed in detail. In conclusion, this paper demonstrated a strategic approach to optimize the coverage and performance, and reduce the cost of small-satellite constellations, compared with present day constellations, allowing the data to be relayed faster and with precision. This will benefit the industry in the development of low-cost constellations, and effectively assist in disaster management and deep-space communication relays. Autonomous orbit selection and navigation can be established for better path alignment for satellites to efficiently propagate and deliver the required data.

Multi-Objective Evolutionary formulations for design of hybrid Earth observing constellations

With the recent advances in satellite miniaturization, communication and information technologies, and the advent of affordable small satellite launch services, there has been a paradigm shift in space exploration missions involving the transition from monolithic architectures formed by a large satellite to the concept of Distributed Spacecraft Missions (DSM), which fly multiple simpler and less costly satellites to offer increased capabilities such as better temporal, spatial, and angular sampling. Despite the potential to provide higher science return and novel data products, the orbit selection problem for Earth Observation DSMs requires a much more complex constellation design process, which involves several interrelated design variables and conflicting objectives. Because of this, prior work has focused mostly on relatively simple DSM architectures consisting of a single type of constellation, such as homogeneous Walker constellations. This paper presents a novel evolutionary formulation (i.e., chromosome and operators) in the context of Multi-Objective Evolutionary Algorithms (MOEA) that allows for the exploration of large tradespaces of non-Walker hybrid satellite constellations with diversity of orbital parameters. The idea behind this new formulation is that it allows to search through the space of different types of constellations – such as Walker formations, Sun-synchronous trains and string-of-pearls among others – and combinations thereof. This type of hybrid constellations have not been studied in detail. The methodology presented in this paper helps overcome the combinatorial explosion resulting when opening up the design space to include non-symmetrical configurations and relaxing constraints in the values of some orbital parameters (e.g. choosing a common altitude or inclination for all satellites forming the constellation). The proposed formulation is compared with a state-of-the-art evolutionary formulation using a variable-length chromosome in 5 different problems including the observation of symmetrical, asymmetrical, connected and disconnected regions of interest. Results show that the proposed formulations achieve better convergence and convergence rate than the state-of-the-art. The proposed method can reduce the effort required to design a problem formulation for each problem instance, while also reducing the risk of missing potentially good architectures due to formulations that are too restrictive and rely too much on previous experience and expertise.

Maximization of LEO Nanosatellite's Transmission Capacity to Multiple Ground Stations: Orbit Selection and Requirements on Attitude Control

This article presents the study of an optimized orbital configuration for a three-axis stabilized 6U LEO nanosatellite that accommodates a polarized calibrator designed to transmit a calibrated signal to seven ground telescopes devoted to cosmic microwave background detection with an accepted maximum absolute pointing error of 4 arcmin; visibility windows and revisit time are maximized by orbit selection and active attitude control considering no possibility of orbital mobility. The solution approach is divided into two parts: in the first part, a dedicated sensitivity study is conducted to find the optimal orbital parameters that maximize the mean contact time between the spacecraft and the list of ground telescopes; in the second part the identified promising orbital configurations are propagated to evaluate the attitude change maneuvers needed during the passage above each telescope to guarantee continuous signal transmission. The results of the sensitivity study report the maximum values for the duration and frequency of the CubeSat-to-Ground station visibility periods. The required performance in terms of pointing accuracy and repointing velocity is later used to propose an attitude control loop for precision pointing and to define a commercial off the shelf Attitude Determination and Control System architecture.

GTOC11: Methods and results from the team of Harbin Institute of Technology

This paper presents the methods and the results submitted by the team of Harbin Institute of Technology in the 11th Global Trajectory Optimization Competition (GTOC11). The mission of GTOC11 is to design the Mother Ship trajectories to execute asteroid flybys selected from over 83000 candidates, and then transfer these asteroids to 12 stations to build the “Dyson Ring” while maximizing the total performance index. The building mission is completed by five steps: “Dyson Ring” orbit selection, database creation, Mother Ship trajectory design, building station construction strategy, and asteroid transfer trajectory optimization. The database creation is mainly used to fast estimate the transfer time for the asteroid low-thrust trajectory, and it plays an essential role in the whole mission. Finally, the building of the “Dyson Ring” is realized within the given period. The final index of the submitted result is 5208.3463, which ranked seventh in GTOC11.

Effect of Earth-Moon’s gravity on TianQin’s range acceleration noise. II. Impact of orbit selection

The paper is a sequel to our previous work (Zhang et al. Phys. Rev. D 103, 062001 (2021)). For proposed geocentric space-based gravitational wave detectors such as TianQin, gLISA, and GADFLI, the gravity-field disturbances, i.e., the so called ``orbital noise'', from the Earth-Moon system on the sensitive intersatellite laser interferometric measurements should be carefully evaluated and taken into account in the concept studies. Based on TianQin, we investigate how the effect, in terms of frequency spectra, varies with different choices of orbital orientations and radii through single-variable studies, and present the corresponding roll-off frequencies that may set the lower bounds of the targeted detection bands. The results, including the special cases of geostationary orbits (gLISA/GADFLI) and repeat orbits, can provide a useful input to orbit and constellation design for future geocentric missions.

CASPA-ADM: a mission concept for observing thermospheric mass density

Cold Atom technology has undergone rapid development in recent years and has been demonstrated in space in the form of cold atom scientific experiments and technology demonstrators, but has so far not been used as the fundamental sensor technology in a science mission. The European Space Agency therefore funded a 7-month project to define the CASPA-ADM mission concept, which serves to demonstrate cold-atom interferometer (CAI) accelerometer technology in space. To make the mission concept useful beyond the technology demonstration, it aims at providing observations of thermosphere mass density in the altitude region of 300–400 km, which is presently not well covered with observations by other missions. The goal for the accuracy of the thermosphere density observations is 1% of the signal, which will enable the study of gas–surface interactions as well as the observation of atmospheric waves. To reach this accuracy, the CAI accelerometer is complemented with a neutral mass spectrometer, ram wind sensor, and a star sensor. The neutral mass spectrometer data is considered valuable on its own since the last measurements of atmospheric composition and temperature in the targeted altitude range date back to 1980s. A multi-frequency GNSS receiver provides not only precise positions, but also thermosphere density observations with a lower resolution along the orbit, which can be used to validate the CAI accelerometer measurements. In this paper, we provide an overview of the mission concept and its objectives, the orbit selection, and derive first requirements for the scientific payload.

Next Generation Gravity Mission Elements of the Mass Change and Geoscience International Constellation: From Orbit Selection to Instrument and Mission Design

ESA’s Next Generation Gravity Mission (NGGM) is a candidate Mission of Opportunity for ESA–NASA cooperation in the frame of the Mass Change and Geosciences International Constellation (MAGIC). The mission aims at enabling long-term monitoring of the temporal variations of Earth’s gravity field at relatively high temporal (down to 3 days) and increased spatial resolutions (up to 100 km) at longer time intervals. This implies also that time series of GRACE and GRACE-FO can be extended towards a climate series. Such variations carry information about mass change induced by the water cycle and the related mass exchange among atmosphere, oceans, cryosphere, land and solid Earth and will complete our picture of global and climate change. The main observable is the variation of the distance between two satellites measured by a ranging instrument. This is complemented by accelerometers that measure the nongravitational accelerations, which need to be reduced from ranging measurements to obtain the gravity signal. The preferred satellite constellation comprises one satellite pair in a near-polar and another in an inclined circular orbit. The paper focuses on the orbit selection methods for optimizing the spatial sampling for multiple temporal resolutions and then on the methodology for deriving the engineering requirements for the space segment, together with a discussion on the main mission parameters.

Applying Lambert problem to association of radar-measured orbit tracks of space objects

Thousands of orbit tracks of space objects are collected by a radar each day, and many may be from uncatalogued objects. As such, it is an urgent demand to catalogue the uncatalogued objects, which requires to determine whether two or more un-correlated tracks (UCTs) are from the same object. This paper proposes to apply the Lambert problem to associate two radar-measured orbit tracks of LEO and HEO objects. A novel method of position correction is proposed to correct the secular and short periodic effects caused by the J 2 perturbation, making the Lambert problem applicable to perturbed orbit tracks. After that, an orbit selection method based on the characteristics of residuals solves the multiple-revolution Lambert problem. Extensive experiments with simulated radar measurements of LEO and HEO objects are carried out to assess the performance of the proposed method. It is shown that the semi-major axis can be determined with an error less than 200 m from two tracks separated by 4 days. The true positive (TP) rates for associating two LEO tracks apart by less than 6 days are 94.2%. The TP rate is still at 73.1% even for two tracks apart by 8–9 days. The results demonstrate the strong applicability of the proposed method to associate radar measurements of uncatalogued objects.

Conceptual design and assessment of a mars orbital logistics node for sustainable human exploration

Logistics supply chains with nodes at critical locations such as the Earth’s surface, Earth’s orbit, cis-lunar space, Mars orbit and the surface of Mars are crucial to enable sustainable human exploration of Mars. The concept of a Mars orbital logistics node is envisaged to have at least aggregation, refueling, resupply and refurbishing capabilities is defined in this paper. Its elements are analogous to platforms like the International Space Station and Lunar Gateway. The stationing orbit of the logistics node is one of the primary design drivers in determining the associated propellant requirement. We evaluate ΔV requirements across four phases of the mission—interplanetary transfer, capture and departure, deorbit and entry, and launch and orbit transfer. A comprehensive data base of interplanetary trajectories is analyzed. Perturbations due to Mars’ gravity are considered and multiple options for maneuver are considered herein. Three mission scenarios – long stay crewed, short stay crewed and cargo transfer – are evaluated along with three deorbit concepts. A broad range of Low Mars Orbits (LMO) with periods between 0.4 and 0.6 sols and Highly Elliptical Mars Orbits (HEMO) with periods between 2 and 4 sols is found to fall on the trade-off curve of total ΔV and orbital periods across seven synodic periods. Final orbit selection can be made using this data and a desired criteria and mission timeline. For example, HEMO minimizes propellant to access the stationing orbit from interplanetary trajectory whereas LMO minimizes propellant requirement to access the surface from the stationing orbit. HMO is also favorable for short stay and cargo missions.

Eclipse avoidance in TianQin orbit selection

In future geocentric space-based gravitational-wave observatory missions, eclipses due to passing through the Moon's and Earth's shadows can negatively impact the sciencecraft's thermal stability and steady power supply. The occurrence should be reduced as much as possible in orbit design. In regard to TianQin's circular high orbits, we tackle the combined challenges of avoiding eclipses and stabilizing the nearly equilateral-triangle constellation. Two strategies are proposed, including initial phase selection and orbit resizing to 1:8 synodic resonance with the Moon, where the latter involves slightly raising TianQin's preliminary orbital radius of $1\times 10^5$ km to $\sim 100900$ km. As the result, we have identified pure-gravity target orbits with a permitted initial phase range of $\sim 15^\circ$, which can maintain eclipse-free during the 3+3 month observation windows throughout a 5-year mission started in 2034, and meanwhile fulfil the constellation stability requirements. Thereby the eclipse issue for TianQin can be largely resolved.

Impact of propulsion system characteristics on the potential for cost reduction of earth observation missions at very low altitudes

Earth observation is one of the most important satellites’ applications. Past earth observation systems have used traditional space technology to achieve the best possible performance, but have been very expensive. Recently, thanks to advancements in technology and modern microelectronics, small satellites have become more and more useful at much lower costs, even if with reduced performance. The resolution of the optical payload improves as the altitude is reduced. Space system mass is proportional to the cube of the linear dimensions. This means that by flying at lower altitudes, satellites can reduce their payload size and therefore the entire mass of the satellite, thus reducing the cost of the system dramatically. However, almost all the earth observation missions fly at the minimum altitude that provides a sufficient orbital life. The addition of a propulsion system capable of providing drag compensation for the entire satellite operative life provides the possibility to fly at very low earth orbit. In this way, the same performance can be obtained with a smaller and cheaper system. To obtain the same coverage more units are needed to replace a larger unit at higher altitude. In this paper it is confirmed that future smallsat observation systems, operating at a lower altitude than traditional systems, have the potential for comparable or better performance, much lower overall mission cost (by a significant factor), lower risk (both implementation and operations), shorter schedules, lower up-front development cost, more sustainable business model, to be more flexible and resilient, more responsive to both new technologies and changing needs, and to mitigate the problem of orbital debris. This paper focus in particular on the effect of the propulsion system parameters (performance and costs) on the cost model as a function of the altitude. It is demonstrated that new affordable chemical propulsion systems provide already significant benefits with limited constraints, allowing a useful reduction of altitude and, consequently, costs. Electric propulsion systems have the potential to allow even lower altitudes or longer lifetimes; however, they have a stronger impact on the satellite design related to their power consumption, generally requiring deployable solar panels, which can limit the flexibility in the orbit selection or the added weight and cost of batteries. The development of electric thrusters that have good performance and limited impact on the satellite architecture (particularly at small scales) is fundamental to exploit their potential for reduced mission costs through very low altitude flight.

Innovative Solar Sail Earth-Trailing Trajectories Enabling Sustainable Heliophysics Missions

The presented trajectory design study demonstrates that near-future solar sails can reach and maintain innovative science orbits about non-traditional Earth-trailing equilibrium points. The proposed configurations can enable long duration stereoscopic solar imaging while potentially alleviating telemetry requirements constraining current solar sail studies. Initial guesses for controlled periodic orbits about equilibrium points that trail the Sun-Earth line by 1.5∘-15∘ are constructed from a linearized solar sail circular restricted three body problem equations of motion. Orbits are then converged in the full nonlinear model using a high order direct collocation method. Transfers to Earth-trailing orbits are then constructed under the assumption that the solar sail is a secondary payload on a larger spacecraft en-route to the Sun-Earth L1 Lagrange point. Naturally occurring gravitational manifolds flowing towards the Lagrange point, as well as navigation data from previous Lagrange point missions, are used to generate a set of baseline trajectories for the primary spacecraft. End-to-end trajectory design and optimization is then carried out to construct time optimal solar sail transfers to each target science orbit from initial conditions constrained to lie on the carrier spacecraft’s Lagrange point trajectory. Additionally, trade studies are performed to assess how sail performance and Earth-trailing orbit selection affect transfer time of flight and instrument pointing accuracy. Results indicate that, based on current technological developments, near future solar sails are capable of station keeping about periodic Earth-trailing orbits while maintaining near-constant solar observation. Moreover, time of flights on the order of 1 year are observed for most combinations of sail performance and feasible initial conditions.

MEO SAR: System Concepts and Analysis

Existing microwave remote sensing instruments used for Earth observation face a clear tradeoff between spatial resolution and revisit times at global scales. The typical imaging capabilities of current systems range from daily observations at kilometer-scale resolutions provided by scatterometers to meter-scale resolutions at lower temporal rates (more than ten days) typical of synthetic aperture radars (SARs). A natural way to fill the gap between these two extremes is to use medium-Earth-orbit SAR (MEO-SAR) systems. MEO satellites are deployed at altitudes above the region of low Earth orbits (LEOs), ending at around 2000 km and below the geosynchronous orbits (GEOs) near 35 786 km. MEO SAR shows a clear potential to provide advantages in terms of spatial coverage, downlink visibility, and global temporal revisit times, e.g., providing moderate resolution images (some tens of meters) at daily rates. This article discusses the design tradeoffs of MEO SAR, including sensitivity and orbit selection. The use of these higher orbits opens the door to global coverage in one- to two-day revisit or continental/oceanic coverage with multidaily observations, making MEO SAR very attractive for future scientific missions with specific interferometric and polarimetric capabilities.