Orbital Maneuvering Research Articles

An intelligent hierarchical recognition method for long-term orbital maneuvering intention of non-cooperative satellites

An intelligent hierarchical recognition method for long-term orbital maneuvering intention of non-cooperative satellites

Devising a deorbitation strategy for Kazakhstani’s KazEOSat-1 spacecraft

The object of this study is the process of deorbiting the KazEOSat-1 spacecraft, which has completed its active service life in low Earth orbit. The main problem is the lack of an effective technique to deorbit KazEOSat-1, taking into account its technical characteristics, orbital parameters, and the need to minimize risks to the environment and other objects in orbit. As part of the work, a software model was built that takes into account the initial orbital parameters of the device, which are essential for planning and performing deorbiting maneuvers. The model is designed to accurately calculate the descent trajectory, taking into account the laws of celestial mechanics and the influence of atmospheric conditions. The optimal deorbiting strategy was selected based on an analysis of various methods for calculating orbital maneuvers aimed at reducing fuel consumption and minimizing environmental risks. This included a comparative analysis of existing approaches and the selection of the most suitable ones under the given mission parameters. The results of the simulation using precise modeling methods in the MATLAB software environment allowed us to determine the main deorbiting parameters, such as the altitude at which the maneuvers begin, the required velocity impulses, the total fuel consumption, and the expected time before entering the dense layers of the atmosphere. Based on the obtained data, practical recommendations were formulated for the KazEOSat-1 deorbit. The first stage, the active controlled deorbit, is carried out by operating the low-thrust engine and braking by the Earth’s atmosphere, allowing the spacecraft to descend from 758 km to 444 km in 2.5 days. The second stage, the passive uncontrolled deorbit, continues the descent to 103 km in 969 days, using only atmospheric braking. The third stage, the uncontrolled drop, begins after reaching 103 km and ends with a drop to the Earth in 834 seconds

Design And Development of a Robotic Arm for Space Debris Mitigation Using Electrostatic and Gecko-Inspired Gripping Technologies

As space debris poses a growing threat to satellite operations, there is an urgent need for advanced robotic systems capable of effective debris capture in low Earth orbit. This paper presents the development and optimization of a robotic arm equipped with an electrostatic adhesion mechanism, designed specifically for microgravity environments. The primary objective is to engineer a versatile, lightweight robotic arm that can securely capture and hold various types of debris, including non-magnetic and composite materials. Key features include electrostatic adhesive pads for adaptable grip, a telescopic extension arm to increase reach with minimal mechanical complexity, and a retractable storage profile to streamline debris retrieval and handling. Through detailed calculations, we establish the required adhesive force to counter both inertial and gravitational forces acting on debris, ensuring secure capture even during minor satellite manoeuvres. An electrostatic charging system is designed to induce sufficient adhesive force, with calculations for charge requirements and pad dimensions optimized for secure adhesion. This paper details the design, force calculations, and component selection that make the robotic arm efficient, lightweight, and adaptable, contributing to safer, more effective space debris removal.

Optimizing Consider-Covariance Models for Realistic Orbit Errors and Application to Collision Avoidance

Precise knowledge about the orbit state vectors and the associated uncertainties are crucial for reliable conjunction assessment and efficient mitigation of on-orbit collision risks through collision avoidance maneuvers by active satellites. The decision-making process heavily relies on the orbit accuracy and the uncertainty realism of the encountering objects, primarily space debris, and the maneuvering satellite itself. The process of determining an orbit creates an estimate of the orbit state and its associated state covariance matrix. This estimate serves as the initial value for predicting the orbit’s uncertainty at the time of closest approach. To determine and numerically propagate realistic orbit errors, the uncertainties of the dynamic model must be taken into account. One method that fulfils this purpose is consider covariance propagation, provided that suitable consider parameters and their variances have been established. This paper presents a novel method for optimizing the consider covariance parameters from past mission data. Real orbit data from multiple low Earth orbit satellites controlled by the DLR, German Aerospace Center’s German Space Operations Center are used to calibrate and evaluate different error models, including covariance parameters and other error terms. The results demonstrate improved uncertainty realism in orbit predictions and subsequent benefits for collision avoidance. Finally, this paper presents the required steps for integrating the new method into the existing collision avoidance system, along with initial results from conjunction assessment.

An IMM Algorithm for Tracking Maneuvering Space Target under Low-Rate Data

Under low data rates, the observational data is sparse, and the noise have a significant impact. Under such circumstance, filtering algorithms tend to rely more heavily on predictions. However, the absence of maneuver information usually results in significant model errors, ultimately affecting the accuracy of states estimation and potentially leading to filter divergence. To address this issue, we have constructed a model set specific to satellite maneuvers and proposed an Improved Interacting Multiple Model (IMM) algorithm based on a designed maneuver detection model and estimation model. We have developed a maneuver estimation model to estimate the maneuvers, and to perform pruning on the maneuver models before estimation. The algorithm dynamically controls the introduction of maneuver models via the detection model and estimation model. When a maneuver is detected, the algorithm introduces the maneuver model set for fusion estimation; otherwise, it degrades into a Kalman filter based on a non-maneuver model. Simulation results indicate that the proposed algorithm effectively tracks targets in low data rate scenarios while maintaining high estimation accuracy.

Absolute Intercalibration of Spaceborne Microwave Radiometers

Abstract Absolute calibration of spaceborne microwave radiometer observations consists of accurate determination of antenna cold space spillover, cross-polarization contamination, and nonlinearity coefficients of the receivers. We deem the GMI sensor to be the most accurate calibrated spaceborne microwave radiometer due to its unique calibration design features and its carefully planned orbit maneuvers. We demonstrate how to transfer the GMI calibration to other spaceborne radiometers, whose operations have sufficient time overlap with GMI. Specifically, we show results for WindSat and AMSR2. The sensor intercalibration is based on brightness temperature matchups between GMI and the other instruments over both open ocean and rainforest scenes. To assess the calibration accuracy, we compare the intercalibrated brightness temperatures with radiative transfer model calculations. In addition, we provide in situ validation results for wind speed and water vapor retrievals from the intercalibrated sensors. The intercalibration methodology allows for the creation of a multidecadal climate data record from passive microwave satellite observations. Significance Statement Creating a long-term climate data record of satellite observations of ocean winds, water vapor, and other variables requires careful and accurate calibration of the various sensors that are used. In particular, it is important to achieve the best possible consistency between the measurements from all the different instruments. This is a challenging task as the configuration and accuracy of these instruments can differ widely. The purpose of our paper is to demonstrate and validate the basic methodology for performing this intercalibration. The backbone of our method is data observed by a well-calibrated sensor that measures the passive microwave emission from Earth’s surface and atmosphere. We show how to transfer its calibration standard to other sensors.

Mega-constellation satellite maneuver forecast via network with attention mechanism

Mega-constellation satellite maneuver forecast via network with attention mechanism

Nonlinear pitch oscillation and control of a gravitational stabilized sailcraft perturbed by solar pressure torque

Nonlinear pitch oscillation and control of a gravitational stabilized sailcraft perturbed by solar pressure torque

The Role of Micro Propulsion in Enabling Autonomous Operations of Nanosatellites

Micro-Electro-Mechanical Systems (MEMS)-based propulsion devices have significant potential for providing high specific power. The application of Micro Power Generation technology to space propulsion systems can enhance orbit and station-keeping capabilities, potentially at lower costs. New-generation spacecraft and satellites are becoming smaller and require micro-propulsion systems for all aspects of their operations. The development of micro-propulsion systems with associated combustors for micro-electricity generation, utilizing liquid fuel, presents major challenges in fuel injection, mixing, combustion, and chemical reactor systems. A micro-spacecraft with high-accuracy station-keeping and attitude control capabilities needs to have low mass and deliver small impulses. Each nanosatellite will weigh a maximum of 10 kg, including the propellant mass. Provisions for orbital manoeuvres, attitude control, multiple sensors, and instruments, along with full autonomy, will result in a highly capable miniaturized satellite. All onboard electronics will endure a total radiation dose of 100 k rads over a two-year mission lifetime. Nanosatellites designed for in-situ measurements will be spin-stabilized and equipped with a complement of particles and fields instruments. Nanosatellites intended for remote measurements will be three-axis stabilized and equipped with imaging and radio wave instruments. Key technologies under development include: advanced, miniaturized chemical propulsion systems; miniaturized sensors; highly integrated, compact electronics; autonomous onboard and ground operations; miniaturized onboard orbit determination methods; onboard RF communications capable of transmitting data to Earth from significant distances; lightweight and efficient solar array panels; lightweight, high-output battery cells; a miniaturized heat transport system; lightweight yet strong composite materials for nano-satellite and deployer-ship structures; and simple, reusable ground systems.

A low earth orbit satellite descent strategy

The object of the article is the method of descending a satellite from low Earth orbit (LEO) at the end of its service lifetime for its disposal. The main problem is to ensure a safe and cost-effective disposal of the satellite while minimizing risks to the environment and other objects in orbit. Based on the method, calculations were made for lowering the orbit and directing the satellite to a remote area of the Earth’s ocean. The calculations included braking pulses, fuel consumption, descent time to the burial orbit, and the time for natural descent before reaching the Earth's surface. The descent process is divided into three stages: active descent using low-thrust engines (LTE), passive descent due to natural braking in the out-of-atmosphere part, and fall through the dense atmosphere. Dependency graphs were provided to illustrate orbit altitude, satellite speed, reentry angle changes, and flight range over time. The findings revealed that combining active orbital maneuvers with passive orbital decay optimizes the deorbiting procedure by minimizing fuel requirements and operational complexity. The proposed strategy ensures the satellite-controlled re-entry and safe disposal, significantly reducing the risks of collisions and environmental impact. The results solved the problem by defining a staged satellite descent process. Duration were computed for active descent with braking impulses for uniform altitude reduction and passive descent due to atmospheric drag. The trajectory and duration of fall in atmosphere. The process ensures controlled re-entry and safe disposal. This method is suitable for satellite operators and space agencies handling LEO missions

Earth shadow time/accelerometer/GNSS-based geosynchronous transfer orbit determination method for electric propulsion satellite

Earth shadow time/accelerometer/GNSS-based geosynchronous transfer orbit determination method for electric propulsion satellite

Fuel-Efficient and Fault-Tolerant CubeSat Orbit Correction via Machine Learning-Based Adaptive Control

The increasing deployment of CubeSats in space missions necessitates the development of efficient and reliable orbital maneuvering techniques, particularly given the constraints on fuel capacity and computational resources. This paper presents a novel two-level control architecture designed to enhance the accuracy and robustness of CubeSat orbital maneuvers. The proposed method integrates a J2-optimized sequence at the high level to leverage natural perturbative effects for fuel-efficient orbit corrections, with a gated recurrent unit (GRU)-based low-level controller that dynamically adjusts the maneuver sequence in real-time to account for unmodeled dynamics and external disturbances. A Kalman filter is employed to estimate the pointing accuracy, which represents the uncertainties in the thrust direction, enabling the GRU to compensate for these uncertainties and ensure precise maneuver execution. This integrated approach significantly enhances both the positional accuracy and fuel efficiency of CubeSat maneuvers. Unlike traditional methods, which either rely on extensive pre-mission planning or computationally expensive control algorithms, our architecture efficiently balances fuel consumption with real-time adaptability, making it well-suited for the resource constraints of CubeSat platforms. The effectiveness of the proposed approach is evaluated through a series of simulations, including an orbit correction scenario and a Monte Carlo analysis. The results demonstrate that the integrated J2-GRU system significantly improves positional accuracy and reduces fuel consumption compared to traditional methods. Even under conditions of high uncertainty, the GRU-based control layer effectively compensates for errors in thrust direction, maintaining a low miss distance throughout the maneuvering period. Additionally, the GRU’s simpler architecture provides computational advantages over more complex models such as long short-term memory (LSTM) networks, making it more suitable for onboard CubeSat implementations.

Discrete search-based determination of a local orbital frame in unknown environments

Discrete search-based determination of a local orbital frame in unknown environments

Orbit determination for spacecraft with continuous low-thrust using nonsingular Thrust-Fourier-Coefficients and filtering-through approach

Orbit determination for spacecraft with continuous low-thrust using nonsingular Thrust-Fourier-Coefficients and filtering-through approach

Swarm-to-swarm orbital pursuit method under delta-v maneuver for space pursuit-evasion

Swarm-to-swarm orbital pursuit method under delta-v maneuver for space pursuit-evasion

Observability-Aware Differential Dynamic Programming with Impulsive Maneuvers

In autonomous space systems, the reliability of navigation systems is essential. Observability in autonomous orbit determination techniques depends on the spacecraft’s orbital motion, making the design of autonomous navigation systems and orbital maneuvers a coupled process. This study develops a stable and efficient algorithm based on differential dynamic programming to design maneuver sequences that improve navigation performance. Our approach incorporates the Fisher information matrix into a cost function to quantify state observability and facilitates its convergence using a semi-analytic gradient and Hessian derived under impulsive maneuvers. Two numerical examples show the validity and effectiveness of our algorithm. The results indicate the stability and efficiency in determining maneuver sequences and the improvement of state estimation accuracy along an optimized trajectory. It is also applicable to other observability-aware optimal control problems because the algorithm is independent of specific systems.

The Trajectory Prediction of Spacecraft under the Influence of Gyroscopic Effect Generated during Non-Keplerian Motion

Due to perturbation forces and control forces, trajectories of spacecraft around the Earth are usually non-Keplerian orbits, which may result in a gyroscopic effect. To meet the complex demands of space operations in the future, the trajectory prediction of spacecraft under the influence of the gyroscopic effect generated during non-Keplerian motion needs to be studied in depth. The paper investigated the trajectory of spacecraft under the gyroscopic effect generated during non-Keplerian motion. Firstly, according to the similarity between the spacecraft precession motion and the gyroscopic precession, as well as the definition of the “gyroscopic effect” of high-speed rotating bodies, the “gyroscopic effect” generated during the non-Keplerian motion of spacecraft around the earth was defined. Then, taking a continuous radial thrust orbit as an example, the dynamics equations of spacecraft under the influence of gyroscopic effect were deduced. Through theoretical analysis and numerical simulation, the trajectory of spacecraft under the influence of the gyroscopic effect generated during non-Keplerian motion was investigated. Finally, the paper simulated the examples and tested the performance of the proposed method. Simulation results show that a large gyroscopic moment may be generated in some non-Keplerian motion of the spacecraft. The greater the rotational angular velocity of the orbital plane, the greater the gyroscopic moment. Due to the gyroscopic effect, there is a significant deviation in the orbit and the orbital elements compared to those without considering the gyroscopic effect, which indicates that the influence of the gyroscopic effect generated during non-Keplerian motion on the orbit of the spacecraft cannot be ignored. It can be seen from the simulation results that the gyroscopic effect has a significant influence on the trajectory of spacecraft. In some special cases, the gyroscopic effect can be utilized reasonably to save fuel and realize low-energy orbit maneuver control technology in actual space missions; but the control should be considered for the spacecraft to bring it back to the desired orbit in most cases. It is necessary to study the trajectory of spacecraft under the influence of gyroscopic effect. The method and conclusions proposed can provide a theoretical reference for spacecraft trajectory prediction and future large-scale fast orbital maneuvers to meet the needs of complex space operations.

Space-Based Passive Orbital Maneuver Detection Algorithm for High-Altitude Situational Awareness

Orbital maneuver detection for non-cooperative targets in space is a key task in space situational awareness. This study develops a passive maneuver detection algorithm using line-of-sight angles measured by a space-based optical sensor, especially for targets in high-altitude orbit. Emphasis is placed on constructing a new characterization for maneuvers as well as the corresponding detection method. First, the concept of relative angular momentum is introduced to characterize the orbital maneuver of the target quantitatively, and the sensitivity of the proposed characterization is analyzed mathematically. Second, a maneuver detection algorithm based on the new characterization is designed in which sliding windows and correlations are utilized to determine the mutation of the maneuver characterization. Subsequently, a numerical simulation system composed of error models, reference missions and trajectories, and computation models for estimating errors is established. Then, the proposed algorithm is verified through numerical simulations for both long-range and close-range targets. The results indicate that the proposed algorithm is effective. Additionally, the sensitivity of the proposed algorithm to the width of the sliding window, accuracy of the optical sensor, magnitude and number of maneuvers, and different relative orbit types is analyzed, and the sensitivity of the new characterization is verified using simulations.

Navigation performance analysis of Earth–Moon spacecraft using GNSS, INS, and star tracker

Global Navigation Satellite System (GNSS) can provide an approach for spacecraft autonomous navigation in earth–moon space to make up for the insufficiency of earth-based tracking, telemetry, and control systems. However, its weak power and poor observation geometry near the moon causes new problems. After the GNSS signal characteristics and satellite visibility were evaluated in Phasing Orbit and Lunar Transfer Orbit, we proposed an adaptive Kalman filter based on the Carrier-to-Noise ratio (C/N0) and innovation vector to weaken the influence of GNSS accuracy attenuation as much as possible. The experimental results show that the spacecraft position and velocity accuracy are better than 10 m and 0.1 m/s near the Earth, and better than 50 m and approximately 0.2 m/s near the moon use GNSS with the proposed adaptive algorithms. Additionally, because of the deterioration of navigation performance based on the orbit filter during orbital maneuvering, we used accelerometer data to compensate for the dynamic model to maintain navigation performance. The results of the experiment provide a reference for subsequent studies.

Precise Orbit Determination for Maneuvering HY2D Using Onboard GNSS Data

The Haiyang-2D (HY2D) satellite is the fourth ocean dynamics environment monitoring satellite launched by China. The satellite operates on a re-entry frozen orbit, which necessitates orbital maneuvers to maintain its designated path once the satellite’s sub-satellite point deviates beyond a certain threshold. However, the execution of orbit maneuvers presents a significant challenge to the field of Precise Orbit Determination (POD). The thesis selects the on-board GPS data of HY2D satellite in December 2023 and five maneuvering days of that year. Employing a multifaceted approach that includes the assessment of observational data quality, orbit overlap, external orbit validation, and SLR (Satellite Laser Ranging) verification, the research delves into precise orbit determination during both maneuver and non-maneuver periods. The results indicate that: (1) The average number of satellites tracked by the receiver is 6.4; (2) During the non-maneuver periods, the average RMS (Root Mean Square) value of the radial difference in the 6-h overlapping arc segment is 0.66 cm, and the three-dimensional position difference is about 1.16 cm; (3) When compared with the precision science orbits (PSO) provided by CNES (Centre National d’Études Spatiales), the average values of the RMS values of the differences in the radial (R), transverse (T), and normal (N) directions during the non-maneuver and maneuver periods are respectively 1.32 cm, 2.31 cm, 1.92 cm and 3.04 cm, 8.78 cm, 2.72 cm. (4) The SLR verification of the orbit revealed a residual RMS of 2.24 cm. This suggests that by incorporating the modeling of maneuver forces during the maneuver periods, the impact of orbital maneuvers on orbit determination can be mitigated.