Earth-Mars Transfer Research Articles

Three-Dimensional Guidance Laws for Spacecraft Propelled by a SWIFT Propulsion System

This paper discusses the optimal control law, in a three-dimensional (3D) heliocentric orbit transfer, of a spacecraft whose primary propulsion system is a Solar Wind Ion Focusing Thruster (SWIFT). A SWIFT is an interesting concept of a propellantless thruster, proposed ten years ago by Gemmer and Mazzoleni, which deflects, collects, and accelerates the charged particles of solar wind to generate thrust in the interplanetary space. To this end, the SWIFT uses a large conical structure made of thin metallic wires, which is positively charged with the aid of an electron gun. In this sense, a SWIFT can be considered as a sort of evolution of the Janhunen’s E-Sail, which also uses a (nominally flat) mesh of electrically charged tethers to deflect the solar wind stream. In the recent literature, the optimal performance of a SWIFT-based vehicle has been studied by assuming a coplanar orbit transfer and a two-dimensional scenario. The mathematical model proposed in this paper extends that result by discussing the optimal guidance laws in the general context of a 3D heliocentric transfer. In this regard, a number of different forms of the spacecraft state vectors are considered. The validity of the obtained optimal control law is tested in a simplified Earth–Venus and Earth–Mars transfer by comparing the simulation results with the literature data in terms of minimum flight time.

Parallel observations process of Tianwen-1 orbit determination

Abstract To meet various requirements of the future deep space missions of China, State Key Laboratory of Astronautic Dynamics constructs a new orbit determinatin software with parallel observations process. By using 32 threads, the computational efficiency per iteration (including spacecraft’s integration) could be promoted to 10 times as that of single-threaded orbit determination. Suppose the number of observations is p p , the number of estimated parameters (including spacecraft’s state) is q q , the amount of computation of one observation is x x , the amount of computation of one Givens transformation is y y and the best number of threads is proved to be [ p ⋅ ( x + y ) / ( q ⋅ y ) ] 1 / 2 {\left[p\cdot \left(x+y)\text{/}\left(q\cdot y)]}^{1\text{/}2} for one-step threads combination. The root mean square of the postfit residuals of China’s deep space monitoring network (CDSMN) and China’s Very Long Baseline Interferometry (VLBI) network (CVN) observations in the Earth-Mars transfer phase and the Mars-orbiting phase are almost the same: about 0.3 m for Ranging, about 0.3 mm/s for Doppler, about 3 cm for VLBI delay and about 0.5 mm/s for VLBI delay rate. It is also found that all the four types of observations of CDSMN and CVN are needed in orbit determination for deep space maneuver and braking at periareion calculation. In the Mars-orbiting phase, the position accuracy after orbit determination under CDSMN-only tracking mode can reach about 1 km.

High-Order Guidance for Time-Optimal Low-Thrust Trajectories with Accuracy Control

This paper develops a high-order guidance scheme based on the expansion of the solution to a time-optimal low-thrust optimal control problem by using differential algebraic techniques. The result allows new optimal trajectories to be autonomously computed from the simple evaluation of polynomials, including updated times of flight. Furthermore, a differential algebraic technique known as automatic domain splitting is implemented to improve the guidance scheme by adaptively and automatically splitting the uncertainty domain into smaller subdomains, in which new guidance polynomials are then generated which are better equipped to handle deviations in their respective subdomains. The result is not only a more effective handling of deviations in the trajectory compared to previous algorithms which use just a single set of polynomials but one that can provide the control and time-of-flight updates to a desired accuracy across the entire domain. Combined with its high-order nature, the guidance scheme is capable of dealing with nonlinearity and large deviations with respect to the reference trajectory, thus being robust to mismodeling and unplanned events. The guidance scheme is evaluated in three scenarios: simplified rendezvous using Clohessy–Wiltshire equations to demonstrate automatic domain splitting in the guidance problem, an Earth–Mars transfer, and an asteroid landing trajectory. The results show the ability of the differential algebraic-based guidance to provide optimal control and time-of-flight updates and improvements of several orders of magnitude over previous differential algebraic-based schemes in handling large uncertainty experienced in spaceflight by leveraging automatic domain splitting.

Modelization of Low-Cost Maneuvers for an Areostationary Preliminary Mission Design

The aim of this paper is to analyze the determination of interplanetary trajectories from Earth to Mars to evaluate the cost of the required impulse magnitudes for an areostationary orbiter mission design. Such analysis is first conducted by solving the Lambert orbital boundary value problem and studying the launch and arrival conditions for various date combinations. Then, genetic algorithms are applied to investigate the minimum-energy transfer orbit. Afterwards, an iterative procedure is used to determine the heliocentric elliptic transfer orbit that matches at the entry point of Mars’s sphere of influence with an areocentric hyperbolic orbit imposing specific conditions on inclination and periapsis radius. Finally, the maneuvers needed to obtain an areostationary orbit are numerically computed for different objective condition values at the Mars entry point to evaluate an areostationary preliminary mission cost for further study and characterization. Results show that, for the dates of the minimum-energy Earth–Mars transfer trajectory, a low value for the maneuvers to achieve an areostationary orbit is obtained for an arrival hyperbola with the minimum possible inclination and a capture into an elliptical trajectory with a low periapsis radius and an apoapsis at the stationary orbit. For a 2026 mission with a TOF of 304 for the minimum-energy Earth–Mars transfer trajectory, for a capture with a periapsis of 300 km above the Mars surface the value achieved will be 2.083 km/s.

Design and Realization Separable Observation Sensors of Tianwen-1 Mars Probe

Tianwen 1 Mars exploration mission is a high-risk mission with long cycle, high communication delay, high product reliability and harsh deep space environment. Two separate observation sensors are installed on Tianwen 1 probe. During the Earth Mars transfer phase of Tianwen 1, the first separable observation sensor was separated from Tianwen 1, which monitored the state of Tianwen 1 and its surrounding space environment. On the Mars orbit, the second separable observation sensor was separated from the Tianwen 1 probe and took a color group photograph of the Tianwen 1 probe and Mars. This paper first introduces the task requirements of separable observation sensors and the design constraints of deep space environment on products. Secondly, the composition of the separable observation sensor is introduced. Thirdly, the relevant indicators of the separable observation sensor are introduced. Finally, the ground test and in orbit test of the separable observation sensor are introduced.

Enabling Interplanetary Exploration for CubeSats with a Fully Chemical Propulsion System

Interplanetary CubeSat missions are currently becoming more popular, with a significant number of recently planned missions. The context of this paper is a Mars mission, starting from a parking orbit around Earth: the adoption of a chemical propulsion system for the Earth-Mars transfer phase is investigated, considering the recent technological developments for CubeSats. A trade-off of propulsion system type and propellant results in the choice of a mono-propellant system adopting the HAN-based propellant AF-M315E (ASCENT). The main challenge for the propulsion system is to fit inside a CubeSat standardised volume, which can range up to 24 U, for which the implementation of a suitable COTS micro-pump is considered. Finally, the complete architecture and design of the propulsion system is presented. This work demonstrates the feasibility of adopting full chemical propulsion for an interplanetary CubeSat mission, with consequent advantages in terms of transfer time and required power, at the cost of relatively small mass and volume left for the other subsystems. Even better results can be expected for interplanetary missions requiring slightly lower ΔV budgets, such as Near Earth Objects exploration or asteroid fly-by missions. Keywords: CubeSat, Chemical Propulsion, Green Mono-propellant, Interplanetary Mission, Mars Exploration

Pseudo-Theory-Based Optimization Design Method for Lunar Gravity-Assisted Human Earth-Mars Transfer Trajectory

The lunar gravity assist (LGA) is recognized as one of the potential ways to reduce the energy demand of human Mars exploration. Nevertheless, it indeed increases the difficulty of designing the exploration trajectory due to its inevitable considerations of a more complex gravitational environment and additional flying stages. Focusing on this issue, an efficient pre-design method is proposed on the basis of the pseudo-state theory. By embedding the pseudo-sweep back step into the dynamical propagation, the accuracy of two-body trajectories can be recovered to a high level, which brings significant convenience of initially designing high-precision trajectory, rapidly searching the departure window, and fast evaluating the LGA effect. A large amount of simulation results show that the application of the pseudo theory could enhance at least 88% trajectory accuracy, increase the 60% success rate, and decrease 5% the necessary cost when optimizing the Earth-Mars transfer trajectory, with comparisons to the two-body trajectory.

Optimal Interplanetary Transfer of Solar Wind Ion Focusing Thruster-Based Spacecraft

The Solar Wind Ion Focusing Thruster (SWIFT) is a highly-innovative propellantless propulsion concept, recently proposed by Gemmer and Mazzoleni. In its nominal configuration, a SWIFT consists of a conically-shaped mesh of positively-charged conducting tethers, with its vertex linked to the spacecraft and its axis oriented towards the Sun. The SWIFT collects and filters the solar wind plasma and suitably directs the positive ions, which are then accelerated by an ion thruster. Such a device is theoretically able to generate a deep-space propulsive acceleration that comes, in part, from the solar wind dynamic pressure impinging on the conical grid and, in part, from the positive ion beam. In particular, the orientation of the ion beam may be chosen in such a way as to set the resultant propulsive acceleration and steer the spacecraft. The aim of this paper is to analyze the performance of a SWIFT-propelled spacecraft in an orbit-to-orbit two-dimensional interplanetary transfer. To that end, some mission scenarios are studied, in an optimal framework, by minimizing the total flight time necessary for the spacecraft to complete the transfer as a function of the propulsion system performance parameters. Numerical simulations are used to compare the optimal flight times calculated in simplified Earth–Venus and Earth–Mars transfers with those obtained by considering other propellantless propulsion systems.

Refined MagSail thrust model for preliminary mission design and trajectory optimization

A magnetic sail is an advanced propellantless propulsion system concept for deep space missions, proposed by Dana Andrews and Robert Zubrin about 30 years ago. The thrust is generated by the electromagnetic interaction between the solar wind charged particles and an artificial magnetic field obtained through a large current-carrying coil made of superconducting material. This paper discusses an up-to-date single-loop magnetic sail thrust model based on the numerical and experimental results obtained in the last decade. In particular, a particle model is initially used to evaluate the performance of a reference configuration in terms of thrust components in a body reference frame. Then, some scaling relationships are used to estimate the magnetic sail performance as a function of the design parameters. The proposed thrust model is used to analyze the minimum-time transfer problem between two circular and coplanar heliocentric orbits. Minimum flight times necessary to complete ephemeris-free Earth-Venus and Earth-Mars transfers are calculated as a function of the magnetic sail characteristic size in a parametric way.

A Stochastic Modeling Approach for Interplanetary Supply Chain Planning

The unique and complex challenges of sustainable interplanetary travel necessitate a novel space logistics paradigm, which is hereafter referred to as the interplanetary supply chain. To lay the groundwork for this anticipated paradigm shift, this work aims to introduce principles and methodologies of supply chain (SC) planning to the field of space logistics. To this end, a stochastic multistage mixed integer linear programming model is proposed that explores the vision of a permanent human presence on Mars from a logistics standpoint. The model optimizes material flows, propellant production, and in-space infrastructure allocations to identify the optimal SC network design. Notably, demand uncertainty at the Mars base is considered, which makes it the first model to investigate the effects of stochasticity in space logistics planning. Several case studies are investigated, which incorporate data on Earth–Mars transfer windows and flight trajectories obtained through the National Aeronautics and Space Administration’s trajectory browser. The results reveal numerous interdependencies and, thus, highlight the need for holistic SC modeling approaches that go beyond the isolated mission planning of the extant literature. For instance, the findings show frequent interactions among spacecraft as they regularly engage in division of labor regarding cargo and propellant transportation. Moreover, complex propellant supply networks that involve strategic in-space infrastructure allocation and propellant production on celestial bodies are established. Overall, the developed model provides valuable insights into the planning of a sustainable interplanetary travel, offers a new SC perspective, and lays the foundation for future research in this area.

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.

Optimum design of nuclear electric propulsion spacecraft for deep space exploration

Future space exploration is now focused on new field-deep space planets. Deep space exploration and the development and utilization of resources are inestimable strategic significance for the country to seize the initiative and command heights of deep space exploration. Nuclear electric propulsion (NEP) systems convert heat from the fission reactor to electrical power and then use the electrical power to produce thrust. Compared with traditional propulsion technology, the NEP system with variable Isp can be used for several applications. The NEP spacecraft is more suitable for deep space exploration missions due to the advantages of high specific impulse, high power, and long life. This paper analyzes the relationship between the transfer travel time, specific mass, power, and payload ratio of the NEP spacecraft through simple performance models. Use the mass optimization and specific mass optimization models based on the NEP system composition and small thrust orbit theory to maximize the payload ratio for a given transfer time and technological characteristics. Finally, applied this NEP model to discuss the feasibility of Mars, Jupiter, and Saturn transfer missions and compared it with the Tianwen-1, Juno, and Cassini–Huygens spacecraft, respectively. Results show that when the NEP spacecraft specific mass reaches 4.77 kg/kW, the Earth–Mars transfer time can be changed to 331.31 days, the payload increases to 1650 kg, the transfer time for Jupiter and Saturn mission would be shorted to 661.31 days and 1131.31 days, the corresponding payload significantly increases which achieved to 1270 kg and 2981 kg. The nuclear electric propulsion spacecraft dramatically improves the detection capability of the spacecraft and provides a reference for the feasibility demonstration and subsequent design of the deep space and the extrasolar planet exploration.

Onboard Orbit Determination for Deep-Space CubeSats

In this work, an orbit determination algorithm suitable for CubeSats onboard implementation is developed, which simulates optical autonomous navigation accomplished by a stand-alone platform. An extended Kalman filter featuring line-of-sight acquisitions of planets is selected as the state estimator, and its performances are tested on a Raspberry Pi, whose characteristics are comparable to a miniaturized onboard computer. An improvement of the solution accuracy is performed by correcting the planetary light-time and aberration effects as well as by exploiting the optimal beacons selection strategy to acquire the external observations. Moreover, the numerical precision of the estimator is improved through the implementation of factorization techniques and nondimensionalization strategies. The results are presented for a sample Earth–Mars transfer, where the time slot for the navigation campaign involves 2 h every 10 days. At final time, the probe position and velocity are estimated with a accuracy of 360 km and 0.04 m/s, respectively.

Multi-Objective Low-Thrust Trajectory Optimization with Robustness to Missed Thrust Events

This paper introduces a new technique for directly controlling the missed thrust recovery margin (MTRM) of a low-thrust spacecraft trajectory. MTRM is defined here as the longest amount of time a spacecraft may coast away from a nominal trajectory while still being able to reach a terminal manifold once thruster operations are resumed. The “virtual swarm” optimization technique developed here simultaneously optimizes the nominal spacecraft trajectory along with many recovery trajectories. The objective can be to maximize the MTRM of the nominal trajectory at its weakest point or to constrain the worst-case MTRM to be at or above a desired level while optimizing a different value (e.g., mass, time of flight). The technique is demonstrated for a direct Earth–Mars transfer and for a gravity assist trajectory to the asteroid Psyche. Further, a method for finding the Pareto front of MTRM, arrival mass, and arrival date is presented to address the related multi-objective optimization problem.

Использование регулярных переменных в задаче оптимизации траектории КА с малой тягой

The regularization of spacecraft motion equations by the Kustaanheimo-Stiefel transformation for coordinates and Sundman’s transformation for time in the case of interplanetary low-thrust optimal transfer is considered. From Pontryagin’s maximum principle, the thrust vector optimal control is derived under the limited power condition. The Earth-Mars transfer problem is solved in the regular variables. The comparison of calculated trajectories with the ones obtained by the parameter continuation method is performed, and the stability properties of the two-point boundary value problem in the Cartesian and regular variables are studied.

Optimization of Earth-Mars transfer trajectories with launch constraints

Optimization of Earth-Mars transfer trajectories with launch constraints

Set-oriented design of interplanetary low-thrust trajectories using Earth Gravity Assist

A method by which to incorporate the Electric Delta-V Earth Gravity Assist (EDVEGA) scheme into a patched three-body approach to design an interplanetary transfer trajectory is presented in this paper. The EDVEGA scheme is a promising technique to reduce fuel consumption using Earth gravity assist and electric propulsion. In a patched three-body approach, the dimension of the problem that connects some trajectories is reduced by the use of special attainable sets. However, because of the singularity associated with Earth's center, it is difficult to connect an EDVEGA trajectory with a trajectory designed in the three-body problem. Therefore, this paper proposes a design method for an Earth–Mars transfer trajectory that combines the EDVEGA scheme with a patched three-body approach using Levi-Civita regularization and a special contour plot. Through comparison with conventional invariant manifold techniques, we demonstrate that fuel consumption is reduced by the proposed method.

Overview of China’s 2020 Mars mission design and navigation

Scheduled for an Earth-to-Mars launch opportunity in 2020, the China’s Mars probe will arrive on Mars in 2021 with the primary objective of injecting an orbiter and placing a lander and a rover on the surface of the Red Planet. For China’s 2020 Mars exploration mission to achieve success, many key technologies must be realized. In this paper, China’s 2020 Mars mission and the spacecraft architecture are first introduced. Then, the preliminary launch opportunity, Earth–Mars transfer, Mars capture, and mission orbits are described. Finally, the main navigation schemes are summarized.

Mission architecture for Mars exploration based on small satellites and planetary drones

Purpose The purpose of this paper is to deal with the study of an innovative unmanned mission to Mars, which is aimed at acquiring a great amount of detailed data related to both Mars’ atmosphere and surface. Design/methodology/approach The Mars surface exploration is conceived by means of a fleet of drones flying among a set of reference points (acting also as entry capsules and charging stations) on the surface. The three key enabling technologies of the proposed mission are the use of small satellites (used in constellation with a minimum of three), the use of electric propulsion systems for the interplanetary transfer (to reduce the propellant mass fraction) and lightweight, efficient, drones designed to operate in the harsh Mars environment and with its tiny atmosphere. Findings The low-thrust Earth-Mars transfer is designed by means of an optimization approach resulting in a duration of slightly more than 27 months with a propellant amount of about 125 kg, which is compatible with the choice of considering a 500 kg-class spacecraft. Four candidate drone configurations have been selected as the result of a sensitivity analysis. Flight endurance, weight and drone size have been considered as the driving design parameters for the selection of the final configuration, which is characterized by six rotors, a total mass of about 6.5 kg and a flight endurance of 28 minutes. In the mission scenario proposed, the drone is assumed to be delivered on the Mars surface by means of a passive entry capsule, which acts also as a docking station and charging base. Such a capsule has been sized both in terms of mass (68 kg) and power (80 W), showing to be compatible with 500 kg-class spacecraft. Research limitations/implications As a general conclusion, the study shows the mission concept feasibility. Practical implications The concept would return incomparable scientific data and can be also be potentially implemented with a relatively low budget exploiting of the shelf components to the larger extent, small identical spacecraft buses and modular low-cost drones. Originality/value The innovative mission architecture proposed in this study aims at providing a complete coverage of the surface and lowest atmospheric layers. The main innovation factor of the proposed mission consists in the adoption of small multi-copter UAVs, also called “drones,” as remote-sensing platforms.

A comparative study between control strategies for a solar sailcraft in an Earth–Mars transfer

The goal of this work was a comparative study of solar sail trajectory optimization using different control strategies. Solar sailcraft is propulsion system with great interest in space engineering, since it uses solar radiation to propulsion. So there is no need for propellant to be used, thus it can remains active throughout the entire transfer maneuver. This type of propulsion system opens the possibility to reduce the cost of exploration missions in the solar system. In its simplest configuration, a Flat Solar Sail (FSS) consists of a large and thin structure generally composed by a film fixed to flexible rods. The performance of these vehicles depends largely on the sails attitude relative to the Sun. Using a FSS as propulsion, an Earth–Mars transfer optimization problem was tackled by the algorithms GEOreal1 and GEOreal2 (Generalized Extremal Optimization with real codification). Those algorithms are Evolutionary Algorithms (AE) based on the theory of Self-Organized Criticality. They were used to optimize the FSS attitude angle so it could reach Mars orbit in minimum time. It was considered that the FSS could perform up to ten attitude maneuvers during orbital transfer. Moreover, the time between maneuvers can be different. So, the algorithms had to optimize an objective function with 20 design variables. The results obtained in this work were compared with previously results that considered constant values of time between maneuvers.