Orbital Maneuvering Research Articles

Influence of non-equilibrium reactions on the optimization of aerothrust aeroassisted maneuver with orbital change

The spaceplane is perspective vehicle due to wide maneuverability in comparison with a space capsule. Its maneuverability is expressed by the larger flight range and also by a possibility to rotate orbital inclination in the atmosphere by the aerodynamic and thrust forces. Orbital plane atmospheric rotation maneuvers can significantly reduce fuel costs compared to rocket-dynamic non-coplanar maneuver. However, this maneuver occurs at Mach numbers about 25, and such velocities lead to non-equilibrium chemical reactions in the shock wave. Such reactions change a physicochemical air property, and it affects aerodynamic coefficients. This paper investigates the influence of non-equilibrium reactions on the aerothrust aeroassisted maneuver with orbital change. The approach is to solve an optimization problem using the differential evolution algorithm with a temperature limitation. The spaceplane aerodynamic coefficients are determined by the numerical solution of the Reynolds-averaged Navier-Stokes equations. The aerodynamic calculations are conducted for the cases of perfect and non-equilibrium gases. A comparison of optimal trajectories, control laws, and fuel costs is made between models of perfect and non-equilibrium gases. The effect of a chemically reacting gas on the finite parameters is also evaluated using control laws obtained for a perfect gas.

Automatic orbital maneuver for mega-constellations maintenance with electrodynamic tethers

This paper focuses on the feasibility of constellation maintenance for large-scale satellite constellations (i.e., mega-constellations) by the electrodynamic tethered multi-agent system with consideration of electric-magnetic-dynamic energy coupling in the orbital maneuver process. The electrodynamic tethered multi-agent system consists of electrodynamic tethers and agents (satellites, robots, or spacecrafts). The nonlinear dynamic characteristics of the electrodynamic tethered multi-agent system are investigated here, to analyze the performance of orbital maneuver. The electrodynamic tether is divided into two parts, such that, the fully insulated tether and the fully bare tether, which used to boost and deboost the electrodynamic tethered multi-agent system respectively. In the orbital boost phase, the real-time electric current generated from the segment of the fully insulated tether. Similarly, in the orbital deboost phase, the electric current generated from the fully bare tether segment. The numerical results show that the system could accelerate or decelerate the motion of the electrodynamic tethered multi-agent system by regulating the magnitude and polarity of electric current carrying tether to generate a Lorentz force, leading to orbital boost or deboost. The electrodynamic tether propulsion is useful for the constellation maintenance, such that, the Starlink, LeoSat and Project Kuiper satellite constellations.

The Geometric Imaging Model for High-Resolution Optical Remote Sensing Satellites Considering Light Aberration and Atmospheric Refraction Errors

With advances in satellite maneuvering imaging capability, stereoscopic images with large roll and pitch angles can be captured to improve the efficiency of observations. At the same time, the influences of light aberration and atmospheric refraction on image positioning accuracy will be more significant. However, these errors are not accounted for in the traditional imaging and calibration model for optical agile satellites. In this study, the formation mechanisms of the aberration and atmospheric refraction errors in optical remote sensing satellite Earth observation imaging were analyzed quantitatively, and correction models were constructed. From this, the traditional geometric imaging model was refined by introducing a correction model for aberration and atmospheric refraction errors to create a more comprehensive geometric imaging model. The feasibility of an extended rational function model, based on the constructed more comprehensive geometric imaging model, was verified quantitatively.

Dynamics of orbital boost maneuver of low Earth orbit satellites by electrodynamic tethers

This paper studied the dynamic characteristics of electrodynamic tethered satellite with consideration of coupled multiphysics field for orbital boost maneuver by a fully insulated electrodynamic tether (EDT). We proposed a simplified analytical nonlinear current–voltage circuit method based on the Parker–Murphy model to evaluate the real-time electric current-carrying tether. Furthermore, the performances of orbital boost maneuver by a fully insulated electrodynamic tether is investigated. This current–voltage model is influenced by the electron density, the intensity of Earth’s magnetic field and the space plasma temperature. The hollow plasma contactor located on the sub-satellite to attract the electrons from the ambient plasma. An on-board power supply is adopted to change the direction of electric current for boosting the system. The direction of electric current in tether is from main-satellite to sub-satellite. The efficiency of orbital boost maneuver is compared with different sizes of anodic contactors and the given voltage of power supply. The simulation results showed that electric current is easily affected by the fluctuation of electron density, the Earth’s magnetic field and plasma temperature. The electrodynamic tether system can stable boost its orbit altitude from 400 to 1200 km in four cases by a fully insulated electrodynamic tether.

Multiple-impulse orbital maneuver with limited observation window

This paper proposes a solution for multiple-impulse orbital maneuvers near circular orbits for special cases where orbital observations are not globally available and the spacecraft is being observed through a limited window from a ground or a space-based station. The current study is particularly useful for small private launching companies with limited access to global observations around the Earth and for orbital maneuvers around other planets for which the orbital observations are limited to the in situ equipment. An appropriate cost function is introduced for the sake of minimizing the total control/impulse effort as well as the orbital uncertainty. It is subsequently proved that for a circle-to-circle maneuver, the optimization problem is quasi-convex with respect to the design variables. For near circular trajectories the same cost function is minimized via a gradient based optimization algorithm in order to provide a sub-optimal solution that is efficient both with respect to energy effort and orbital uncertainty. As a relevant case study, a four-impulse orbital maneuver between circular orbits under Mars gravitation is simulated and analyzed to demonstrate the effectiveness of the proposed algorithm.

Parallel processing for orbital maneuver detection

With the increase in the number of objects orbiting Earth, space situational awareness (SSA) has becoming an important area of research in the space sector (Lemmens and Krag, 2014). This paper outlines approaches to increase the performance of SSA algorithms. Parallel processing is implemented in orbital maneuver detection methods with the goal of increasing the computational speed and accuracy of detection by fusing two well-known detection methods: Two-Line Elements (TLE) analysis and state propagation into a hybrid detection method. Speed is an important component when considering the implementation of real time SSA, as well as, with the increase in the number of objects orbiting Earth the throughput of maneuver detection algorithms needs to increase to stay feasible. The results show that implementing parallel processing increased the computational speed of TLE analysis by approximately eight times and of state propagation by approximately four orders of magnitude. In terms of computational speed, the novel method performed worse than TLE analysis and state propagation. However, the method reduced the number of false maneuver detections.

Comparison of BDS Station Clock Short-term Prediction Models and their Applications in Precise Orbit Determination

BDS (BeiDou Navigation Satellite System) ground tracking stations are equipped with high accuracy atomic clocks, and they are synchronized with the BDS time scale (BDT) via the Precise Orbit Determination (POD) processing. During the periods of satellite maneuver and post-maneuver, station clocks are kept fixed as known values in the POD processing. To improve the real-time POD capability, station clocks need to be predicted. In this paper, the performance of three clock prediction models is evaluated, including quadratic polynomial model (QP), periodical term model (PM), and grey model (GM). The precision of clock fitting and prediction, as well as the performance of the prediction models in POD are compared. Data of six stations are used for test, and the results show that: the mean fitting accuracy of quadratic polynomial model, periodical term model, and grey model is 0.14ns, 0.05ns, 0.27ns, respectively; the 1h and 2h prediction precision of the three models is 1.17ns, 0.88ns, 1.28ns, and 2.72ns, 2.09ns, 2.53ns, respectively. Applying the 1h and 2h predicted station clocks in the POD, the 3D orbit accuracy reaches the best using the periodical term model, while the radial accuracy of satellite orbit is rather close for the three models with the difference within 3cm.

Hyperion: Artificial gravity reusable crewed deep space transport

We present a design to actively simulate gravity for crew members traveling to and from Mars combining tested technology with leading research. The Hyperion mission and spacecraft architecture can be the start of the next generation of spacecraft for deep-space, but to truly be the springboard for future deep space transports, it was designed to require no assistance from resupply depots at destination, prioritize human comfort for long duration transits, and have multiple configurations to adjust for the mission. The design is focused around generating artificial gravity by rotating the fixed-distance habitat and variable-distance counterweight about the propellant system in the center (X axis), while each subsystem is individually developed to support this operation. Through various configurations, the Hyperion spacecraft contains the unique ability to adjust to the mission phase for aerobraking and transport needs, making the concept particularly adept at reducing propellant needs, maintaining artificial gravity with minimal power, and keeping the center of mass in line with the thrust vector. This is accomplished using de-spun solar panels, Ion engines for spin-up and orbital maneuvers, an extendable truss for mass balance, and a phased array communication system. Hyperion was designed to meet NASA 2017–18 Revolutionary Aerospace Systems Concepts - Academic Linkage (RASC-AL) requirements of developing an Artificial Gravity Reusable Crewed Deep Space Transport and is expected to cost no more than $3.0B per year over 11 years. It is expected to be operational by 2032.

Orbital maneuver strategy design based on piecewise linear optimization for spacecraft soft landing on irregular asteroids

Recently, asteroid exploration becomes an important branch of human’s deep space activities. In this paper, a piecewise linear optimal orbital maneuver strategy is designed for a spacecraft soft landing on irregular-shaped asteroids. First, the space around an irregular asteroid is converted into several grid units, and the gravitational field of the asteroid is linearly fitted in each unit. Then, the soft-landing orbital maneuver strategy design problem is formulated as a piecewise linear optimal problem, and further transferred into a family of two-point boundary value problems, which can be solved using collocation method. Finally, a corresponding algorithm is developed to obtain the piecewise linear optimal maneuver strategy, which is proved to be able to achieve the soft-landing mission well. Simulation results show that the error of the model linearization is small enough, while the calculation efficiency is remarkably improved, and the robustness of maneuver strategy is also improved.

Attitude scheduling and verification for dynamic imaging of agile satellites

Drift angle is a major factor which limits the quality of remote sensing images. This paper presents an attitude scheduling algorithm aimed to eliminate drift angle and its verification for dynamic imaging of agile satellites. First, the attitude scheduling algorithm was derived intuitively based on two dynamic imaging criterions. Then a drift angle model based on differentiation was described for empirical verification of the proposed algorithm. For dynamic imaging mode, satellite manoeuvre velocity was taken into consideration in drift angle model. Finally, dynamic imaging simulation along a inclined target intersecting with sub-satellite track was performed based on ellipsoid model of globe. Comparison was carried out between the proposed approach and conventional compensatory approach, the results of which demonstrated the effectiveness and accuracy of the proposed approach in eliminating the drift angle.

Satellite proximate pursuit-evasion game with different thrust configurations

This paper investigates the proximity satellite pursuit-evasion game where the pursuer carries three orthogonal thrusters, each of which can exert a magnitude-limited force along one axis, while the evader has a single thruster which can produce a bounded acceleration in any direction. The pursuer or the evader tries their best to capture or escape by performing orbital maneuvers. By establishing a local moving coordinate frame on the originally revolving orbit of the evader, we reduce the dynamics of each player to the linear Clohessy-Wiltshire equations. We then transform the problem of finding the optimal controls associated with the saddle point solution of the game into the two-point boundary value problem, which is solved by combining the heuristic searching and Newton method. At last, by numerical simulations, we discuss the effectiveness of the proposed algorithm in finding the open-loop solution to the game. We show that, in contrast with pursuit-evasion games where each player has one single thruster, the problem considered in this paper may not be solved efficiently by the indirect method, since there exist some initial states of the players such that the proposed algorithm fails to solve the open-loop control.

Basic performance and future developments of BeiDou global navigation satellite system

The core performance elements of global navigation satellite system include availability, continuity, integrity and accuracy, all of which are particularly important for the developing BeiDou global navigation satellite system (BDS-3). This paper describes the basic performance of BDS-3 and suggests some methods to improve the positioning, navigation and timing (PNT) service. The precision of the BDS-3 post-processing orbit can reach centimeter level, the average satellite clock offset uncertainty of 18 medium circular orbit satellites is 1.55 ns and the average signal-in-space ranging error is approximately 0.474 m. The future possible improvements for the BeiDou navigation system are also discussed. It is suggested to increase the orbital inclination of the inclined geostationary orbit (IGSO) satellites to improve the PNT service in the Arctic region. The IGSO satellite can perform part of the geostationary orbit (GEO) satellite’s functions to solve the southern occlusion problem of the GEO satellite service in the northern hemisphere (namely the “south wall effect”). The space-borne inertial navigation system could be used to realize continuous orbit determination during satellite maneuver. In addition, high-accuracy space-borne hydrogen clock or cesium clock can be used to maintain the time system in the autonomous navigation mode, and stability of spatial datum. Furthermore, the ionospheric delay correction model of BDS-3 for all signals should be unified to avoid user confusion and improve positioning accuracy. Finally, to overcome the vulnerability of satellite navigation system, the comprehensive and resilient PNT infrastructures are proposed for the future seamless PNT services.

Observing Mercury by a quasi-propellantless mission

In this paper, the feasibility of exploiting the Sun’s gravitational force to design the final phase (capture and orbit circularisation) to Mercury with a low propellant consumption has been investigated. The initial conditions, on the eccentricity and argument of pericentre, for the circularisation phase are obtained from a prime integral of motion, which takes place when the probe moves over a polar orbit. A numerical analysis has highlighted how these initial conditions allow, under the third-body effect, an optimal reduction of eccentricity (circularisation phase) with limited variations of the other orbit elements, without requiring orbital manoeuvres. Then, by introducing a small velocity variation, the phase of ballistic capture has been joined to the circularisation phase. The obtained results can encourage the exploitation of this phenomenon in future missions for the exploration of Mercury.

Design of Satellite Maneuvers for Inertia Parameter Estimation

This paper addresses the design of satellite maneuvers for the inertia estimation. The experiment design is an important step in system identification, since the choice of the excitation signal has a great influence on the precision of the parameter estimates. In order to design an optimal maneuver, the proposed method uses a cubic B-spline representation of the trajectory. The optimized maneuver is obtained through minimization of a functional based on the Fisher information matrix. The optimization process considers the physical constraints due to the saturation of the actuators. The effectiveness of the generated trajectory is evaluated via Monte Carlo simulations using the model of a satellite-type platform. Moreover, the optimized maneuver has been also tested on a real platform in a zero-gravity experiment where, due to the limited duration of the tests, the achievement of the maximum excitation is of great importance.

Point-ahead analysis and pre-pointing link stability study of intersatellite laser communication

The static bias error angle obviously affects pre-pointing links’ stability in the presence of vibration in intersatellite laser communication. The 2nd order point-ahead angle is a source of misalignment which was ignored in most solutions, and this is the concern of our paper. In this study, we present a further analytical investigation into the point-ahead angle in complex satellite maneuvering environment. Static bias error angle induced by the 2nd order point-ahead angle has been studied under different intersatellite links. The probability density function of the pre-pointing links’ outage has been derived in the presence of pointing jitter taking consideration of the static bias angle, and the link budget has also been analyzed. Simulation model of link stability has been established to verify the numerical results by the Monte Carlo method in Matlab-Simulink environment. The results have shown that the 2nd order point-ahead angle has a significant detrimental impact on link stability in long distance links. It is a neglectable factor. This work is dedicated to intersatellite laser communication system design.

Application of differential evolution algorithm in the problem of aeroassisted orbital plane change maneuver

The paper deals with application of the differential evolution method in optimizing aeroassisted orbital plane change maneuver performed by an aerospace vehicle with a high lift-to-drag ratio. The vehicles motion relative to the Earth is described by a system of differential equations in the flight path axis system. The angle-of-attack and the air path bank angle control programs are presented in the form of Fourier series and thrust control has the form of flip-flop control. The results of solving problems without phase parameter constraints obtained by the differential evolution algorithm were compared with the results obtained using Pontryagins maximum principle. The solution of the optimization problem taking into account temperature limitations at the stagnation point was obtained on the basis of the differential evolution algorithm.

Tether attachment point stabilization of noncooperative debris captured by a tethered space system

Before an active debris removal process is carried out, the attitude stabilization of the debris needs to be executed to ensure security. Fortunately, for the tethered system, a full despun is not always necessary; a stabilization of the motion of the tether attachment point is usually sufficient. This paper discusses the way to stabilize the motion of the tether attachment point of a captured noncooperative space debris without any auxiliary mechanisms. The debris possesses unknown inertia information and states after capture, while the tether tension is the only input to manipulate its attitude. The direction and value of the tension is regulated based on the continuous dissipation of the rotational energy of the debris. Moreover, the range of the tether direction is restricted to a small conical region, in order to reduce the tether libration motion and the orbit maneuver of the main satellite. The motion of the tether attachment point is desired to be observed instead of measuring the attitude motion of the debris in the feedback loop, which might be easier to attain in practical applications. Numerical simulations are carried out to validate the effectiveness of the control method and to check the performance of the controller with different parameters.]

Controlled Spacecraft Re-Entry of a Drag De-Orbit Device (D3)

For spacecraft containing components that survive re-entry, it is important to de-orbit the satellite over a non-populated area. Because most CubeSats (cube satellites that conform to the CubeSat form factor) do not have their own propulsion systems and cannot perform a de-orbit burn, aerodynamic drag modulation presents an attractive solution to re-enter the satellite at the desired location. The University of Florida Advanced Autonomous Multiple Spacecraft (ADAMUS) lab has developed a drag de-orbit device (D3) for CubeSats, which are affordable systems for demonstrating attitude and orbit control. The device consists of four retractable tape-spring booms that are designed and manufactured to validate the targeted re-entry of a CubeSat in Low-Earth Orbit (LEO). By modulating the D3 drag area, orbital maneuvering and controlled re-entry can be performed. This paper outlines the functional and vibration testing procedures and results for the completed device. The complete drag device was taken to NASA Ames Research Center and subjected to random vibrations at 9.6 times the force of gravity. The device survived vibration testing with no obvious damage. After vibration testing, it was determined that no major mechanical design changes to the D3 will need to be made. Future work for this project includes assembling a final, flight-ready drag device that will be attached to a CubeSat for launch.

Sun-Synchronous Orbit Transfer Scheme Based on Solar Sail

Aiming at the sun-synchronous orbit transfer task of the micro-nano satellite solar sail, the method of combining the weighted control law and the solar sail optimal direction is used for maneuvering the orbit, and a segmentation control strategy is proposed to control the local time of the descending node of target orbit. The theoretical and simulation results show that the sun-synchronous orbit transfer task can be completed by using the micro-nano satellite solar sail, and the local time of the descending node is consistent with the initial value. In addition, the local time of the descending node has a certain impact on the orbit maneuvering efficiency of the solar sail, and the task can be completed by using the segmentation control strategy on orbits with different local time of the descending node.

Are mercury emissions from satellite electric propulsion an environmental concern?**The data that support the findings of this study are available from the corresponding author upon reasonable request.

A new generation of satellites for Earth observation and telecommunications are being designed and built with off the shelf components. This is driving down costs and permitting the launch of large satellite swarms with unprecedented spatial and temporal coverage. On-orbit maneuvers are commonly performed using ion thrusters. Mercury is one of the cheapest and easiest to store propellants for electric propulsion. While some mercury released in Low Earth Orbit may escape Earth’s gravitational field, mercury emissions originating from many common orbital maneuvers will return to Earth. The environmental and human health implications of such releases have not been evaluated. Using an atmospheric chemical transport model, we simulate global deposition of mercury released from satellite propulsion systems. We estimate that 75% of the mercury falling back to Earth will be deposited in the world’s oceans, with potentially negative implications for commercial fish and other marine life. Understanding the scale of this novel mercury source in a post-Minamata Convention world is necessary to limit ecosystem exposure to mercury contamination.