Orbit Transfer Mission Research Articles

Adaptive tracking guidance for powered aeroassisted orbital maneuver under system uncertainties and input constraints

The powered aeroassisted orbital maneuver (AOM) is an emerging type of AOM manner, which can fulfill various complicated orbital transfer missions more flexibly and efficiently by virtue of aerodynamic forces and synergetic thrust. In this paper, an adaptive control strategy is presented for the reference trajectory tracking guidance of powered AOM under system uncertainties and input constraints. The presented controller is designed with an attack angle control law and a bank angle control law. The constrained control inputs are smoothly described through a hyperbolic tangent-type function. Then, the attack angle and bank angle control laws are separately developed by utilizing the first-order and second-order backstepping control methods. Besides, the parametric adaptive mechanism is incorporated to identify the squares of the upper bounds of the total uncertain items. Benefiting from this design, the presented controller can not only resist system uncertainties, but also accommodate input constraints. The stability evaluation shows that the radial distance and velocity tracking errors under the presented controller can eventually stabilize to the bounded minor fields around zero. Lastly, the derived results are verified through comparative studies with an existing controller and Monte Carlo simulations under different conditions.

Optimal debris removal using debris-tether-tug system considering attitude–orbit coupling effects

Towing space debris with the utilization of a tethered space tug, also referred to as debris-tether-tug (DTT) system, is considered to be one of the most promising methods to remove space debris. However, as the space tug only carries limited amounts of fuel, the number of debris that a DTT system can remove is limited. To increase the capacity of the DTT system, fuel consumption per towing cycle should be minimized. This study proposes an optimal deorbit scheme with two stages: optimal transfer and swing control. First, a dynamics model of the DTT system is formulated considering attitude–orbit coupling effects and tether flexibility. In the fuel optimal transfer stage, the dynamics model as well as two simplified models is employed to calculate the optimal transfer orbit, subject to a continuous thrust from the tug. By comparing the optimal results, it can be found that the attitude–orbit coupling effect of the DTT system is non-negligible, whereas the tether flexibility can be ignored with a large tether elasticity coefficient during the optimal orbit transfer mission. For the swing control stage at the destination, a control law is applied by regulating the velocity of tether releasement to increase the debris velocity. The parametric analysis of the control law suggests that the tether should be completely deployed before separation to achieve a larger orbit altitude increment. Appropriate parameters can be assigned to attain maximum orbit altitude increment in the swing control based on this analysis. The conclusions could serve as an effective reference for transfer orbit calculation and swing control design in future deorbit missions using the DTT system.

Two trajectory configurations for the low-thrust transfer between northern and southern halo orbits in the Earth-Moon system

This paper presents a study of the low-thrust transfer problem between L2 northern and southern halo orbits using the Pontryagin minimum principle (PMP) in the Earth-Moon circular restricted three-body problem (CRTBP). The northern and southern halo families exist in the vicinity of the Moon, and have the same orbital period distribution, which can be used for space missions such as lunar navigation constellations, lunar relay communications and lunar space stations. Traditional low-thrust transfer trajectories are multi-revolution structures, which are modeled in this problem as multiple rotations around the libration point. The characteristics of the transfer are affected by the number of revolutions. Moreover, we propose a novel approach using invariant manifolds associated with periodic orbits in the CRTBP to construct L2 north–south halo orbit transfers. This approach results in superior transfer performance, which involves shorter time of flight, lower fuel consumption and reduced engine thrust requirements. A complete set of low-thrust time-, energy- and fuel-optimal transfer algorithms using moving point strategies and continuation methods is proposed. The computational procedures and analysis presented in this paper provide a reliable reference for the design of related orbital transfer missions.

Capturing and deorbiting Envisat with an Airbus Spacetug. Results from the ESA e.Deorbit consolidation phase study

Airbus has been developing a mission and chaser concept for capturing and deorbiting the defunct Envisat satellite with different partners in the frame of ESA studies (e.Deorbit Phases A and B1). In the last e.Deorbit study called ‘Consolidation Phase’ which aims at implementing findings from the intermediate SRR, Airbus along with their partners CBK, MDA, SENER and GMV was pursuing the mission and chaser definition, but then based on the Airbus Spacetug vehicle for GEO servicing.The assumption for this analysis was a Spacetug vehicle adapted for the specific e.Deorbit mission rather than generic electrical Spacetug for orbit transfer or GEO servicing missions. As the same Eurostar Neo bus platform would be used in both cases, the potential for the reuse of the e.Deorbit Spacetug solution for further servicing missions would be therefore very high.Starting from mission analyses, an e.Deorbit Spacetug baseline configuration was defined, bringing various options depending on the mission scenarios. As with the Airbus Spacetug vehicle, this e.Deorbit Spacetug vehicle aims at being launched with a medium or heavy launcher. The first considered scenario is the one inherited from e.Deorbit Phase B1 study, i.e. based on motion synchronization with Envisat now preceded by a contactless detumbling manoeuvres, capture via robotic arm and fixation and stabilization using clamping mechanism, before deorbiting with controlled re-entry. The Airbus Spacetug preliminary design is based on the Eurostar Neo platform from Airbus and made of a 2.5 kW electrical power system, a chemical propulsion system with associated bi-propellant tanks, a set of sensors dedicated to AOCS/GNC, a robotic arm system, gripper and clamping mechanism for servicing and an On-Board Computer and avionics bus reused from Eurostar Neo.The preliminary assessment of the synergies between the e.Deorbit and GEO Spacetug systems already shows very good opportunities for a mutual fertilization of both systems. A great opportunity pointing out here is the high flexibility of the Spacetug platform to adapt to different mission configurations.

Extending service life of hall thrusters: recent progress and future challenges

Hall thrusters are widely applied to space propulsion missions because of characteristics such as high specific impulse, great thrust-to-power ratio, high efficiency, and simple structure. Because the scope of space missions using Hall thrusters has expanded, a higher total impulse is required to extend a north–south station-keeping mission to an orbit transfer mission, requiring Hall thrusters with improved lifetimes. Existing research shows that improving the anti-sputtering capacity of the discharge channel and significantly reducing the ion energy and flux bombardment on the wall of the discharge channel are the best measures to improve the lifetime of Hall thrusters. This study summarizes the lifetime restraining factors of Hall thrusters and analyzes the magnetic shielding technology, wall-less technology, aft-magnetic technology, and cylindrical Hall thruster technology to reduce ion energy and flux bombardment on the discharge-channel wall, adopting low-sputtering yield materials to reduce the wall erosion rate, which can improve the lifetime of Hall thrusters. This paper also identifies the limitations of each technology, such as the erosion of magnetic poles (MS) and ceramic cover plates (CHT), larger plume divergence angle (wall-less), hard-to-select aft-magnetic degree (aft-magnetic technology), low compactness, and great porosity (graphite material), providing a reference for further research.

Orbital maneuvering of electric solar wind sail based on an advanced solar wind force model

The electric solar wind sail is a propulsion system that extracts the solar wind momentum for the thrust force of a spacecraft by using an interaction between solar wind protons and the electric potential structure around charged long thin conducting tethers. The system enables a spacecraft to generate a thrust force without consuming reaction mass. This paper investigates the capability of the electric solar wind sail as a propulsion system for deep space exploration missions. The shape of the conducting tether that is determined by the equilibrium of the solar wind force and centrifugal force is numerically calculated for formulating an advanced solar wind force model. The conducting tethers deviate from the ideal sail spin plane, and the maximum value of the thrust direction varies from 13∘ to 19∘. To estimate the spacecraft thrust vector, which is calculated as the sum of solar wind force vectors exerted on each tether, best-fit polynomial equations are proposed. We performed numerical simulations for a two-dimensional orbital transfer mission to investigate the capability of the electric solar wind sail. Results of numerical simulations show that the electric solar wind sail enables spacecraft to perform Earth–Venus, Earth–Mars, and Earth–Itokawa transfer missions. Additionally, this paper performs three-dimensional simulations for an Earth–Ryugu transfer mission. The electric solar wind sail achieves a more complicated orbital transfer in a reasonable mission time.

Real-Time Optimal Control for Spacecraft Orbit Transfer via Multiscale Deep Neural Networks

This study is motivated by the requirement of on-board trajectory optimization with guaranteed convergence and real-time performance for optimal spacecraft orbit transfers. To this end, a real-time optimal control approach is proposed using deep learning technologies to obtain minimum-time trajectories of solar sail spacecraft for orbit transfer missions. First, the minimum-time two-dimensional orbit transfer problem is solved by an indirect method, and the costate normalization technique is introduced to increase the probability of finding the optimal solutions. Second, by making novel use of deep learning technologies, three deep neural networks are designed and trained offline by the obtained optimal solutions to generate the guidance commands in real time during flight. Consequently, the long-standing difficulty of on-board trajectory generation is resolved. Then, an interactive network training strategy is presented to increase the success rate of finding optima by supplying good initial guesses for the indirect method. Moreover, a multiscale network cooperation strategy is designed to deal with the recognition deficiency of deep neural networks (DNNs) with small input values, which helps achieve highly precise control of terminal orbit insertion. Numerical simulations are given to substantiate the efficiency of these techniques, and illustrate the effectiveness and robustness of the proposed DNN-based trajectory control for future on-board applications.

On-line orbit planning and guidance for advanced upper stage

Purpose This paper aims to investigate the problem of on-line orbit planning and guidance for an advanced upper stage. Design/methodology/approach The double impulse optimal transfer orbit is planned by the Lambert algorithm and the improved particle swarm optimization (IPSO) method, which can reduce the total velocity increment of the transfer orbit. More specially, a simplified formula is developed to obtain the working time of the main engine for two phases of flight based on the theorem of impulse. Subsequently, the true anomalies of the start position and the end position for both two phases are planned by the Newton iterative algorithm and the Kepler equation. Finally, the first phase of flight is guided by a novel iterative guidance (NIG) law based on the true anomaly update with respect to the geometrical relationship. Also, a completely analytical powered explicit guidance (APEG) law is presented to realize orbital injection for the second phase of flight. Findings Simulations including Monte Carlo and three typical orbit transfer missions are carried out to demonstrate the efficiency of the proposed scheme. Originality/value A novel on-line orbit planning algorithm is developed based on the Lambert problem, IPSO optimization method and Newton iterative algorithm. The NIG and APEG are presented to realize the designed transfer orbit for the first and second phases of flight. Both two guidance laws achieve higher orbit injection accuracies than traditional guidance laws.

Cost analysis of solar thermal propulsion systems for microsatellite applications

In recent years, satellite design has extended towards miniaturisation to reduce associated cost with launching and conducting space missions. Small satellites provide low-cost platforms for space missions. However, this lower cost comes at the expense of the removal of key sub-systems, such as the propulsion system, due to the small available onboard volume and mass restrictions. For this reason, small, lightweight, high-performing and affordable propulsion systems are necessary. However, there is limited research available on the comparison of propulsion technologies with regards to cost. Motivated by the above challenges the objective of this paper is to provide a comparison of propulsion technologies that are compatible with small satellites with respect to cost and application. The different propulsion systems are investigated for three mission scenarios, a small on-orbit manoeuvre, a station-keeping, and a lunar orbit transfer mission. Each system is evaluated in terms of a total figure of merit which incorporates nine variables such as propellant mass, safety, and hardware price, that affect the total cost of a propulsion system. This figure of merit is used to quantitatively compare the propulsion systems to identify cost-effective solutions as a function of various mission scenarios. Solar thermal propulsion has been proposed for small satellite applications, but information regarding the concepts are not available in a single report. Therefore, a secondary objective of this paper is to provide the reader with a review of the current status of solar thermal propulsion. An important finding of this research is the classification of propulsion systems in terms of thrust, specific impulse, cost, and application.

Performance and Applicability of Orbital Transfer with Bare Electrodynamic Tether

Electrodynamic tether technology is a novel way to perform spacecraft orbital transfers. This paper focuses on a bare electrodynamic tether (BEDT) system and evaluates its orbital transfer performance and mission applicability. A nonsingular dynamic model of the BEDT orbital transfer system is developed. In particular, the ambient perturbation forces, including the atmospheric drag and Earth oblateness effect, are taken into account. Numerical simulations are then performed to investigate the orbital transfer performance under various mission parameters involving the initial orbital altitude, initial orbital inclination, tether length, tether diameter, spacecraft mass, and power voltage. An orbital transfer scheme is designed to raise and regulate the orbit. Furthermore, to verify its practicability, the mission applicability of the BEDT system is analyzed for Ariane-4’s H10 upper stage by comparison with the conventional chemical propulsion method. Finally, the conclusions and some discussions are presented to support conducting BEDT-based orbital transfer mission planning in practice.

Orbital transfer strategy using only-accelerating maneuvers

Two types of new orbital transfer strategies that use only-accelerating maneuvers are proposed for a simple spacecraft with only one engine. Based on the requirement of only-accelerating maneuvers, the constraints on the orbital parameters in the entire transfer process are derived from Gauss variational equations. The explanation of these constraints from the geometric viewpoint makes it easy to determine an initial maneuver sequence without time-consuming computation. Only-accelerating maneuvers for an orbital transfer mission can also be implemented by two different approaches: impulsive maneuvers and finite-thrust propulsive maneuvers. The algorithm to determine both maneuvers are summarized in three steps. Impulsive maneuvers can accomplish a transfer mission in an orbital period, whereas finite-thrust propulsive maneuvers require several orbital periods, but a smaller thrust. Finally, numerical simulations are conducted for their application to specific transfer missions.

Neighboring optimal control for open-time multiburn orbital transfers

While a parametric approach has been proposed recently to establish the neighboring optimal control (NOC) for multiburn orbital transfer problems with fixed final time, it is not applicable once the final time is free. Indeed, it is quite common for orbital transfer missions to leave their final time free. In this paper, we continue developing the parametric approach and extend it to establishing the NOC for such problems with free final time. The key idea is to construct a new parameterized family of extremals such that every neighboring extremal around a nominal one is totally determined by its final time and a vector. Then, the NOC for free-time scenarios is readily established through deriving the Taylor expansion of the parameterized neighboring extremals; as a by-product, some prerequisite conditions, fundamental for establishing the NOC, are formulated. Moreover, a scheme for performing the NOC is proposed accordingly. Finally, to illustrate the improvement of this paper, a free-time multiburn orbital transfer problem is computed and some Monte Carlo campaigns are tested in the presence of various perturbations, showing that the final condition errors are significantly reduced by employing the NOC developed in the paper.

Mathematical modeling of spacecraft guidance and control system in 3D space orbit transfer mission

Spacecraft performance in an orbital maneuver relies on guidance and control systems which manage the thrust direction within orbit transfer. In this article, the guidance and control approach for spacecraft having a 3D orbit transfer mission is proposed. To derive the optimal variation of steering angles with initial and terminal constraints on the space orbits, a mathematics polynomial function of the guidance command with unknown coefficients is introduced, one of which is determined to achieve the transfer accuracy requirement between space orbits. Genetic Algorithm is employed in finding optimal variation of guidance command and the optimal initial states within the transfer. The attitude control system is also modeled to evaluate the spacecraft response with respect to generated commands by the guidance system. Gas thrusters are considered as attitude actuators for space mission and linear controller with pulse-width pulse-frequency modulator and unconstrained control allocation is employed for controlling steering angles. Results indicate that the presented approach for guidance and control system fairly satisfies the mission requirement.

Non-linear solution to the maximum height orbit transfer guidance problem

An optimal control solution to the highly non-linear problem of orbit transfer mission is achieved by using a newly proposed analytical perturbation technique. The problem is classified as a two-point boundary value problem in order to optimize a performance measure in a given time. Assuming a constant thrust operating in a given length of time, it is sought to find the thrust direction history of a transfer from a given initial orbit to the largest possible orbit. The system dynamical model is stated by regarding a variable mass spacecraft moving in the variable gravitational field of the Earth, based on the two-body problem. To assess the perturbation solution fidelity, a numerical solution based on the Gauss pseudospectral method has been employed. The main novelty of this work is in applying a new analytical solution strategy that is a combination of perturbation technique and backward integration to a highly non-linear problem in the calculus of variations approach.

A novel algorithm to optimize complicated low‐thrust trajectory

PurposeThis paper aims to investigate, a new optimization algorithm for complex orbit transfer missions with low‐thrust propulsion system, which minimizes the drawbacks of traditional optimization methods, such as bad convergence, difficulty of initial guesses and local optimality.Design/methodology/approachFirst, the trajectory optimization problem comes down to a nonlinear constraint parameter optimization by using the concept of traditional hybrid method. Then, one utilizes genetic algorithm (GA) to solve this parameter optimization problem after treating the constraints with the simulated annealing (SA) and random penalty function. Finally, one makes use of localized optimization to improve the precision of the final solutions.FindingsThis algorithm not only keeps the advantages of traditional hybrid method such as high precision and smooth solutions, but also inherits the merits of GA which could avoid initial guess work and obtain a globally optimal solution.Research limitations/implicationsFurther, research is required to reduce the computational complexity when the transfer trajectory is very complex and/or has many adjustable variables.Practical implicationsBy using this method, the globally optimal solutions of some complex missions, which could not be obtained by traditional method, could be found.Originality/valueThis method combines the GA with traditional hybrid method, and utilizes SA and random penalty functions to treat with constraints, and then gives out a super convergence way to find the globally optimal low‐thrust transfer orbit.

エレクトロダイナミックテザーとモーメンタムテザーを使った混成軌道間輸送システムに関する研究

An orbit transfer system that transports a payload from LEO to GEO using spinning-momentum-tether is introduced in this paper. This system is analyzed in consideration of the change of the angular velocity before and after the catch or the release of the payload for the various parameters of the tether system. The time and the propellant mass required for reboosting the stage are also calculated in detail. As the new thruster for reboosting, the electrodynamic tether (ED tether) system is considered in this study and this system is compared with the conventional ion-thruster system. In addition, the total system mass is optimized for the mission interval that the next transportation of a payload is possible. The results also show that ED tether system can reduce the total system mass compared with ion-thruster system for the same orbit transfer mission.

Performance Optimization of a Multi-Stage Spinning-Balance Tether Orbit Transfer System

An orbit transfer system that transports a payload from LEO to GEO using multi-stage spinning tethers is studied in this paper. We propose a tether configuration called “Balance Tether.” The variations in the total mass and mission duration of the system are analyzed for different numbers of stages and different lengths of tethers. In addition, the possibility of debris impact damaging each of several tether system configurations is also analyzed. As a result, we obtain the guidelines for a system design that minimizes total system mass while also taking into account the system lifetime as limited by debris impact. In addition, it is shown that the Balance Tether can reduce the total system mass compared to a single-tether system. The results also show the effectiveness of the spinning-tether system as compared to a conventional chemical propulsion system for the same orbit transfer mission.

イオン推進機の軌道間輸送機への適用検討

It is expected that an orbit transfer vehicle using an ion thruster (IOTV) can obtain a large payload ratio, because the ion thruster has a high specific impulse (several thousands seconds). However, orbit transfer missions by means of the IOTV require large trip times, therefore it is necessary for the ion thruster to have a sufficient endurance performance. In this study, the trajectory of the solar powered IOTV was analyzed in consideration of the influence of earth's eclipse, then the payload capability of the IOTV and the steady/endurance performance of the ion thruster were evaluated. For results, it was found that the IOTV could achieve approximately double the payload mass of the OTV using a chemical rocket, and a required durability to the ion thruster was comparable to one that has used for ETS-VI.

Optimal Multi-Criteria Aeroassisted Orbital Transfer Trajectories

Single and multiple atmospheric pass aeroassisted orbital transfers for three representative AOTV configurations on a geostationary to space station orbit transfer mission are investigated using multi-criteria trajectory optimization. The fuel optimal results indicate that independent of the vehicle type, the maximum heat flux is reduced by 20% and 35%, respectively, using two and three atmospheric passes in comparison to a single-pass. The multi-criteria results show that the ability to trade-off fuel consumption for a lower peak heating rate improves for all transfer modes (single- and multiple-pass) with increasing lift-to-drag ratio.

Electric propulsion: An evolutionary technology

The NASA Lewis Research Center (LeRC) conducts and directs an electric propulsion research and technology program aimed at providing high-performance electric propulsion system options for a broad range of near and far-term missions. This evolutionary program emphasizes the development of propulsion systems for three classes of missions: (1) near term auxiliary propulsion applications such as North-South Stationkeeping for next generation communications satellites and orbit maintainence for orbiting platforms such as Space Station Freedom; (2) advanced solar electric propulsion and SP-100-class nuclear electric propulsion for Earth-space orbit transfer and robotic planetary missions; and (3) very high power systems to support major space missions including the Space Exploration Initiative. To cover widely disparate mission requirements, the LeRC program includes research on electrothermal, electrostatic, and electromagnetic systems. This paper provides an overview of the LeRC program with a focus on recent progress.