Low-thrust Propulsion Systems Research Articles

Методика балістичного планування місій низькоорбітального сервісного обслугову-вання з малою постійною тягою рушійних систем

The current stage of space exploration is characterized by an increased interest in the development, deployment, and operation of low-orbit satellite constellations (LOSC) for Earth and near-Earth space remote sensing for military and civilian purposes and for global and regional satellite communications. Reusable space launch vehicles have significantly reduced the orbital injection cost. As a result, satellite operators are developing and deploying large-scale LOSCs of various orbital structures with a large number of spacecraft. According to current estimates, more than 70% of all the operating satellites operate in low-Earth orbits (LEOs) at altitudes between 160 km and 2,000 km. Since LEO satellites are generally much cheaper than satellites in geostationary orbits, the possibility of their on-orbit servicing (OOS) has not been the focus of research. However, the use of LEO OOS has prospects for growth. Techniques for ballistic planning of LEO OOS missions have been and are being developed. The disadvantages of approximate techniques include the use of simplified flight dynamics models. Most of the existing exact techniques are based on the use of full mathematical models of flight dynamics and the shooting method to solve the boundary value problem of an orbit transfer. Using the shooting method requires a sufficiently accurate initial guess, which is difficult to determine. To obtain a second approximation, use is mainly made of optimization methods, which do not always find a global minimum. In this regard, there is a need to develop new techniques that would be free from the above disadvantages. The goal of the article is to develop a ballistic planning technique for low-orbit servicing missions with low constant thrust propulsion systems. The technique includes the identification of LEO areas promising for OOS, a mathematical model of the dynamics of perturbed OOS orbit transfers in modified equinoctial orbital elements, and a solution algorithm for the boundary value problem of determining the control parameters of perturbed OOS low-orbit transfers. The problem is solved using methods of statistical analysis, flight dynamics, shooting, genetic optimization, and mathematical simulation. The novelty lies in the identification of LEO areas promising for OOS and the development of a mathematical model of orbit transfer dynamics in modified equinoctial orbital elements and a solution algorithm for determining the control parameters of perturbed OOS low-orbit transfers. The results of the work may be used in the justification and planning of LEO OOS missions and the formulation of requirements for LEO OSS mission propulsion systems.

Low-thrust trajectory optimization of deep space exploration based on discrete pulse method

Deep space exploration is gradually becoming the hot mission of human space activities, and this kind of mission has the characteristics of a long mission cycle, complex space environment, high fuel demand, and difficult telemetry track and command, etc. It is the future trend to carry out deep space exploration activities using low-thrust propulsion systems, and the trajectory optimization of interplanetary missions is one of the significant difficulties. Considering the characteristics and difficulties of interplanetary missions, a method of transfer orbit preliminary design in the multi-body system is proposed in this paper. The simulation results illustrate that the transfer orbit obtained by the method in this paper can provide a reliable initial guess for the optimization of high-precision orbit and decrease convergence sensitivity. The research results can provide a significant reference for the implementation of deep space exploration missions in China in the future.

Orbital Facility Location Problem for Satellite Constellation Servicing Depots

This work proposes an adaptation of the facility location problem for the optimal placement of on-orbit servicing depots for satellite constellations in high-altitude orbit. The high-altitude regime, such as medium Earth orbit, is a unique dynamic environment where low-thrust propulsion systems can provide the necessary thrust to conduct plane-change maneuvers between the various orbital planes of the constellation. As such, on-orbit servicing architectures involving servicer spacecraft that conduct roundtrips between servicing depots and the client satellites of the constellation may be conceived. To this end, a new orbital facility location problem formulation is proposed based on binary linear programming, in which the costs of operating and allocating the facility(ies) to satellites are optimized in terms of the sum of the effective mass to low Earth orbit (EMLEO). The low-thrust transfers between the facilities and the clients are computed using a parallel implementation of a Lyapunov feedback controller. The total launch cost of the depot, along with its servicers, propellant, and payload, is taken into account as the cost to establish a given depot. The proposed approach is applied to designing an on-orbit servicing depot architecture for the Galileo and the GPS constellations.

Low-Thrust Nonlinear Orbit Control for Very Low Lunar Orbits

In the next decades, both space agencies and private competitors are targeting the lunar environment as a scientific and technological resource for future space missions. In particular, the confirmed existence of water-ice deposits in the vicinity of the poles (predominantly the south pole) makes polar or near-polar low lunar orbits attractive for the purpose of designing space missions that could search for suitable Lunar base sites. However, traveling very-low-altitude orbits is very challenging, as they are strongly perturbed by the Moon’s gravity field as well as third- and fourth-body effects due to the Earth and the Sun. Several studies demonstrate that these orbits are expected to impact the lunar surface in a few months. Therefore, the definition and implementation of an effective station-keeping strategy represents a crucial issue in order to extend satellites’ lifetime. In this paper, a feedback nonlinear control law is employed in order to perform corrective maneuvers aimed at keeping the state of the satellite within acceptable margins. The satellite is assumed to be equipped with a steerable and throttleable low-thrust propulsion system. The control law is based on the Lyapunov stability theory and does not require any reference path to track, with a considerable decrease in the computational cost. The proposed real-time control law includes control saturation, related to the maximum available thrust magnitude, and is developed employing modified equinoctial elements, in order to avoid singularities and extend its range of application. Finally, the strategy at hand is tested in the presence of all the relevant perturbations (i.e., harmonics of the selenopotential, third- and fourth-body effects) in order to show its effectiveness and efficiency.

Distributed safe trajectory optimization for large-scale spacecraft formation reconfiguration

This paper presents a distributed trajectory optimization algorithm for the reconfiguration of large-scale spacecraft formation, which uses a low thrust propulsion system to safely guide the spacecraft formation to a desired formation. Formation reconfiguration is formulated as a trajectory optimization problem with complex constraints, and the relative motion is accurately described by a nonlinear relative dynamics considering J2 perturbation. The collision avoidance constraint is convexified into a tangent plane to ensure the accuracy of the solution. The resulting nonlinear trajectory optimization problem is solved by applying the hp-adaptive pseudospectral method to convert it into a nonlinear programming problem. In order to overcome the disadvantage of huge computation of centralized algorithm considering collision avoidance, “predicted trajectory” is introduced to transmit information between spacecraft, and the parallel computing is implemented. Finally, a numerical simulation is given to verify the computational efficiency and collision avoidance effectiveness of the proposed distributed algorithm.

Earth-Venus Mission Analysis via Weak Capture and Nonlinear Orbit Control

Exploration of Venus is recently driven by the interest of the scientific community in understanding the evolution of Earth-size planets, and is leading the implementation of missions that can benefit from new design techniques and technology. In this work, we investigate the possibility to implement a microsatellite exploration mission to Venus, taking advantage of (i) weak capture, and (ii) nonlinear orbit control. This research considers the case of a microsatellite, equipped with a high-thrust and a low-thrust propulsion system, and placed in a highly elliptical Earth orbit, not specifically designed for the Earth-Venus mission of interest. In particular, to minimize the propellant mass, phase (i) of the mission was designed to inject the microsatellite into a low-energy capture around Venus, at the end of the interplanetary arc. The low-energy capture is designed in the dynamical framework of the circular restricted 3-body problem associated with the Sun-Venus system. Modeling the problem with the use of the Hamiltonian formalism, capture trajectories can be characterized based on their state while transiting in the equilibrium region about the collinear libration point L1. Low-energy capture orbits are identified that require the minimum velocity change to be established. These results are obtained using the General Mission Analysis Tool, which implements planetary ephemeris. After completing the ballistic capture, phase (ii) of the mission starts, and it is aimed at driving the microsatellite toward the operational orbit about Venus. The transfer maneuver is based on the use of low-thrust propulsion and nonlinear orbit control. Convergence toward the desired operational orbit is investigated and is proven analytically using the Lyapunov stability theory, in conjunction with the LaSalle invariance principle, under certain conditions related to the orbit perturbing accelerations and the low-thrust magnitude. The numerical results prove that the mission profile at hand, combining low-energy capture and low-thrust nonlinear orbit control, represents a viable and effective strategy for microsatellite missions to Venus.

ANALYSIS OF THE FEASIBILITY OF USING A DISPLACEMENT FEED SYSTEM USING AMMONIA VAPOR AS A DISPLACEMENT GAS

In recent decades, there has been a significant increase in the number of space vehicles launched annually into Earth's orbit. With the growth of the number of space vehicles, the demand for low-thrust propulsion systems, which ensure their delivery to a given orbit and the orientation and stabilization of the spacecraft in orbit, also increases. This makes it necessary to constantly improve the systems of these propulsion systems in order to ensure their best characteristics while minimizing their dimensions, weight and power consumption. This is important to ensure the optimal functioning of spacecraft and reduce the costs of launching and maintaining missions. This paper considers the possibility of using ammonia as a displacer gas in low-thrust propulsion systems. The main components that should be part of the displacement system using ammonia have been determined. A comparison of the proposed system with other types of displacement systems used in propulsion systems is given, with the determination of their advantages and disadvantages. A comparison of the displacement system using ammonia with the displacement systems using pressurization gas according to the minimum mass criterion for different mass values of the working substance and different values of its consumption to the thruster is given. Based on the results of the calculations, the ranges of applicability of the displacement system using ammonia were determined, at which this system will ensure a reduction in the mass of the propulsion system and the spacecraft as a whole. The results of this work can be used when evaluating the feasibility of using a displacement system using ammonia, as a displacement gas, in propulsion systems under development.

Low-Thrust Transfer to Quasi-Synchronous Martian Elliptic Orbit via Nonlinear Feedback Control

This study considers the problem of injecting a spacecraft into an elliptic, repeating-ground-track orbit about Mars, starting from a 4-sol highly elliptical orbit, which is a typical Martian capture orbit, entered at the end of the interplanetary transfer. The final operational orbit has apoares corresponding to the maximum (or minimum) latitude, and nine nodal periods are flown in 5 Martian nodal days. The orbit at hand is proven to guarantee coverage properties similar to the Molniya orbit about Earth; therefore, it is especially suitable for satellites that form constellations. Low-thrust nonlinear orbit control is proposed as an affordable and effective option for orbit injection, capable of attaining significant propellant reduction if compared to alternative strategies based on chemical propulsion. This work introduces a new, saturated feedback law for the low-thrust direction and magnitude that is capable of driving the spacecraft of interest toward the operational orbit. Remarkable stability properties are proven to hold using the Lyapunov stability theory. Because no reference path is to be identified a priori, this technique represents a viable autonomous guidance strategy, even in the case of temporary unavailability of the low-thrust propulsion system or in the presence of widely dispersed initial conditions and errors on estimating orbit perturbations. Monte Carlo simulations prove that the feedback guidance strategy at hand is effective and accurate for injecting a spacecraft into the desired, repeating-ground-track operational orbit without requiring any reference transfer path.

Regularized direct method for low–thrust trajectory optimization: Minimum–fuel transfer between cislunar periodic orbits

Low–thrust propulsion systems of high specific impulse are advantageous in the fuel efficiency. The fuel expenditure is further reduced together with optimization techniques though fuel–optimal low–thrust trajectories result in a discontinuous bang–bang control structure that can cause numerical difficulties. The present paper introduces regularized formulations into the direct multiple shooting method of minimizing fuel expenditure for low–thrust spacecraft trajectories. Instantaneous velocity changes approximating low–thrust maneuvers are expressed by regularized variables based on the Levi–Civita or Kustaanheimo–Stiefel transformation. The new formulation removes singularities due to null thrust impulses from derivatives of an objective function and constraints. The method automatically and robustly finds fuel–efficient bang–bang control structures for multi–revolutional transfers between periodic orbits in the Earth–Moon circular restricted three–body problem.

Improved Monte Carlo Tree Search-based approach to low-thrust multiple gravity-assist trajectory design

In the preliminary design of low-thrust trajectories with multiple gravity-assist phases, the selection of the celestial bodies to be considered and the sequence of optimal maneuvers have a significant impact on the overall mission performance. This aspect makes the preliminary transfer trajectory analysis a very complex task. This Short Communication presents a novel method to design heliocentric multiple gravity-assist transfer trajectories for a spacecraft with a low-thrust propulsion system, where the total velocity variation is minimized. In this context, a Bézier shape-based method is used to describe the generic arc of a propelled trajectory, while an extension of the Monte Carlo Tree Search algorithm is used for a rapid analysis of the gravity-assist maneuver sequence. In particular, the proposed approach effectively solves the problem of selecting a suitable balance parameter in the Monte Carlo Tree Search-based routines, which makes the method especially useful in a preliminary mission phase. The effectiveness of the proposed approach is shown by simulating an interplanetary transfer towards the outer regions of the Solar System.

A machine learning method for the orbit state classification of large LEO constellation satellites

Recently, large LEO constellation satellites have received much attention from the aspect of Space Domain Awareness (SDA) due to their significant number and relatively complex orbit states. Low-thrust propulsion system technologies are widely used during the lifetime of such satellites, making it difficult for the work of TLE (Two Line Element) based maneuver detection. This study proposes a machine learning method for orbit state classification based on TLE data, considering the concept of the Auto-Encoder (AE) network and Decision Tree (DT). In this manuscript, the orbit states were classified into four categories: non-maneuver, controlled orbit raising, controlled orbit lowering, and controlled orbit maintenance. First, historical TLE data of multiple satellites of the STARLINK constellation system and uncontrolled objects in a similar orbital area was analyzed using an Auto Encoder network to identify nonlinear features. Second, the extracted feature was sent to the DT network for the training of the supervised classification model, which could classify the input into four categories of orbit states. Finally, extra targets that had not been considered during the modeling process were used to test the applicability of the proposed method on satellites of the same constellation system. The results suggest that the method we presented here was applicable for orbit state classification with a macro precision of 93.8% and a macro F1-score of 93.7%, according to our numerical experiments. Once set up, the proposed method presented a feasible orbit state classifier for satellites of the same constellation system, operating at a similar altitude.

Low-Thrust Trajectory Design from Lunar Rideshare to Venus Capture

Small spacecraft are likely to be launched as part of a rideshare mission rather than on a dedicated launch. These spacecraft typically use low-thrust propulsion systems as they have lower fuel mass requirements. However, without a dedicated launch, the lower thrust can make it difficult to reach interplanetary targets. Further mission design difficulties arise as the optimizers used to search for such trajectories often require a good initial guess, which can be difficult to find. This work patches together a perturbed Sims–Flanagan transcription in three different regimes, each with a different central body, to perform a preliminary search for trajectories from a lunar rideshare to a weak capture around Venus with little input required from the mission designer. Example trajectories of such a mission are presented. The ability of this mission architecture to arrive as close to Venus’s upper atmosphere as possible is explored.

NetSat—Challenges and lessons learned of a formation of 4 nano-satellites

The objective of the NetSat mission was to develop underlying technologies for small satellite formation flying and to demonstrate it at the level of four nano-satellites in three-dimensional configurations. Thus, base technologies, such as orbit determination, attitude determination and control, inter-satellite communication, and multi-satellite operations were studied and implemented on a 3U CubeSat platform. Technological challenges concerned in particular the precision 3-axis attitude control system in combination with an electric continuous low-thrust propulsion system. Different control approaches for efficient transitions between various formation topologies were analyzed, taking the limitations of the hardware platform (propellant mass, thrust level, power budget, attitude control capabilities) as well as external orbit perturbations into account and allowing for distributed control of multi-satellite formations. In a research program of over 15 years, the essential technologies for formations of small satellites were realized subsequently and then tested in orbit on multiple missions. On this basis, NetSat was launched on 28. September 2020 on a Soyuz rocket into a circular orbit of an altitude of 575km. During the operations phase several new operations concepts developed in the project, namely the Compass IoT network spanning all ground and satellite systems with more than 50 nodes, or a flexible operations frontend with auto-operations capabilities could be tested and allowed for operating such a multi-satellite network. Additionally, the implemented inter-satellite link among the satellites could be established and tested extensively, marking the first time that four 3U CubeSats were able to successfully demonstrate inter-satellite communication in orbit.

An Earth-Mars microsatellite mission leveraging low-energy capture and low-thrust propulsion

Growing interest in deep-space exploration using microsatellites is leading the development of novel trajectory design techniques compatible with the constraints of this class of spacecraft. The mission design discussed in this paper examines the possibility of injecting a microsatellite into a desired orbit around Mars taking advantage of multibody dynamics, performing Earth and Moon flybys, which concludes in a low-energy capture at Mars, followed by nonlinear control for orbit injection maneuvers. If compared to a traditional Earth-Mars transfer, the proposed solution leads to relevant propellant savings and extended launch windows, but the capture orbit established has chaotic behavior and time-varying osculating orbital parameters which may not satisfy the mission requirements. Consequently, the microsatellite is capable of transferring from the original low-energy capture orbit to the desired operative one. For this phase, the microsatellite is assumed to operate using a low-thrust propulsion system, which is ignited with the intent of driving the microsatellite toward two operational orbits: (a) a quasi-synchronous repeating 2-SOL orbit and (b) a sun-synchronous orbit with altitude of 200 km. To do this, nonlinear orbit control is employed. Convergence toward the desired operational orbits is investigated, and can be guaranteed - using the Lyapunov stability theory, in conjunction with the LaSalle invariance principle - under certain conditions related to the orbit perturbing accelerations and the low-thrust magnitude. The numerical simulations prove that the combination of multiple-flyby transfer (simulated with real ephemeris, with the use of GMAT) and low-thrust nonlinear orbit control represents a viable and effective strategy for microsatellite missions to Mars.

Multi-objective optimization of low-thrust propulsion systems for multi-target missions using ANNs

Multi-target missions are an attractive solution to visit multiple bodies, increasing the scientific return and reducing the cost, compared to multiple missions to individual targets. Examples of multi-target missions are multiple active debris removals (MADR) and multiple near-Earth asteroids rendezvous (MNR) missions. MADR missions allow for the disposal of inactive satellites, preventing the build-up of space junk, while MNR missions allow to reduce the expenses of each asteroid observation. Since those missions are long and highly demanding in terms of energy, it is paramount to select the most convenient propulsion system so that the propellant mass and the duration of the mission are minimized. To this end, this paper proposes the use of a multi-objective optimization and artificial neural networks. The methodology is assessed by optimizing trajectories for MADR and MNR sequences with off-the-shelf thrusters. Multiple Pareto-optimal solutions can be identified depending on the propulsion system characteristics, enabling mission designers to trade-off the different options quickly and reliably.

Lifetime Extension of Ultra Low-Altitude Lunar Spacecraft with Low-Thrust Propulsion System

Due to the non-spherical perturbation of the Moon, the lifetime of ultra low-altitude Lunar spacecraft may be quite short. In this paper, we analyze the lifetime of about 50 km-altitude Lunar spacecraft with different initial orbit. The lifetime in low inclination orbits is much shorter than the ones in the near polar orbits. To extend the lifetime and keep the spacecraft in an appropriate range, an orbit maintenance strategy based on low-thrust propulsion system is proposed. The influence of the orbit initial conditions (e.g., semi-major axis, inclination, right ascension of the ascending node) on lifetime extension are discussed and the effect of the low-thrust magnitude in orbit maintenance is analyzed. According to the numerical simulation results, the lifetime of about 50 km-altitude 100 kg Lunar spacecraft with 10 kg fuel and 20 mN thruster can be extended from 7.958 days to over a 109.1725 days, which demonstrates the effectiveness of the strategy. Furthermore, a global perspective for ultra low-altitude Lunar spacecraft lifetime extension problem is provided in this paper, which can be applied to Moon mission designs extensively.

Mission design of DESTINY[formula omitted]: Toward active asteroid (3200) Phaethon and multiple small bodies

DESTINY+ is an upcoming JAXA Epsilon medium-class mission to fly by the Geminids meteor shower parent body (3200) Phaethon. It will be the world’s first spacecraft to escape from a near-geostationary transfer orbit into deep space using a low-thrust propulsion system. In doing so, DESTINY+ will demonstrate a number of technologies that include a highly efficient ion engine system, lightweight solar array panels, and advanced asteroid flyby observation instruments. These demonstrations will pave the way for JAXA’s envisioned low-cost, high-frequency space exploration plans. Following the Phaethon flyby observation, DESTINY+ will visit additional asteroids as its extended mission. The mission design is divided into three phases: a spiral-shaped apogee-raising phase, a multi-lunar-flyby phase to escape Earth, and an interplanetary and asteroids flyby phase. The main challenges include the optimization of the many-revolution low-thrust spiral phase under operational constraints; the design of a multi-lunar-flyby sequence in a multi-body environment; and the design of multiple asteroid flybys connected via Earth gravity assists. This paper shows a novel, practical approach to tackle these complex problems, and presents feasible solutions found within the mass budget and mission constraints. Among them, the baseline solution is shown and discussed in depth; DESTINY+ will spend two years raising its apogee with ion engines, followed by four lunar gravity assists, and a flyby of asteroids (3200) Phaethon and (155140) 2005 UD. Finally, the flight operations plan for the spiral phase and the asteroid flyby phase are presented in detail.

Artificial Neural Networks for Multiple NEA Rendezvous Missions with Continuous Thrust

The interest for near-Earth asteroids for scientific studies and, in particular, for potentially hazardous asteroids requires the space community to perform multiple-asteroid missions with close-up observations. To this end, multiple near-Earth asteroid rendezvous missions can help reduce the cost of the mission. Given the enormous number of asteroids, this work proposes a method based on artificial neural networks (ANNs) to quickly estimate the transfer time and cost between asteroids using low-thrust propulsion. The neural network output is used in a sequence-search algorithm based on a tree-search method to identify feasible sequences of asteroids to rendezvous. The rendezvous sequences are optimized by solving an optimal control problem for each leg to verify the feasibility of the transfer. The effectiveness of the presented methodology is assessed through sequences of asteroids of interest optimized using two low-thrust propulsion systems, namely solar electric propulsion and solar sailing. The results show that ANNs are able to estimate the duration and cost of low-thrust transfers with high accuracy in a modest computational time.

Trajectory options for a manned mission to Mars using high-thrust propulsion systems

An analysis has been done of performance factors (mission duration, initial mass of the interplanetary crew transfer vehicle, velocity of the re-entry into the Earth atmosphere of the descent vehicle with the crew) for a single-spacecraft manned mission to Mars using high-thrust propulsion systems. Locally optimal solutions (in terms of delta-V budgets for the transfer) were found for the Earth–Mars–Earth transfer, with varying periods of waiting in Mars orbit, minimal distance to the Sun, as well as flight paths (direct Earth–Mars–Earth transfers vs. gravity assist maneuvers at Venus during Earth–Mars or Mars–Earth transfers). The proposed classification for locally optimal solutions is applicable to both high-thrust propulsion systems and low-thrust propulsion systems. A comparison of performance factors has been done for manned Martian mission options based on liquid-propellant engines and nuclear rocket engines with a 12,5 km/s constraint on the velocity of the manned re-entry vehicle in the Earth atmosphere. Key words: Manned mission to Mars, interplanetary transfer trajectory, high-thrust, liquid-propellant rocket engine, nuclear rocket engine, mission duration, initial mass of the interplanetary vehicle, re-entry velocity of a manned vehicle for returning the crew to Earth.

Initial research on thermal decomposition of 98% concentrated hydrogen peroxide in thruster-like conditions

The current research is focused on investigating the possibility of harnessing the thermal decomposition of highly concentrated Hydrogen Peroxide 98% i.e., utilizing a decomposition process initiated and maintained solely by a temperature source for low thrust propulsion system applications. This early study investigates the influence of the initial temperature of the heater, power input condition and chamber pressure on the behaviour and stability of the prototype thruster. While it is still too early to conclude on the practicality of the research it demonstrates that a stable configuration with good transient start-ups times is indeed achievable.