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In this paper, an integrated optimal flight plan with a guidance plan in a general transfer orbit is considered. Due to the need to simultaneously reduce energy consumption and flight time in some space missions, the spacecraft trajectory is optimized based on two criteria to minimize fuel consumption and flight time in the orbital transfer process. For this purpose, the total velocity impulse on both sides of the transition orbit and the elapsed time are defined as the criteria of the objective function which includes the spacecraft’s insertion conditions and orbital parameters related to the shape of the trajectory. The optimal transition (coasting) trajectory is then obtained using the multi-objective genetic algorithm (MOGA) search method. Subsequently, the velocity-to-be-gained guidance steering scheme is integrated to track the trajectory onboard and reach the final orbit. The numerical results of the proposed algorithm are extracted in three case studies and compared with other references that show its ability to find the optimal transfer.

Optimal control of nonlinear impulsive systems with free impulse instants and the number of impulses is investigated in this study. A scheme based on adaptive dynamic programming is developed, which leads to a feedback (approximate) solution to the defined optimal impulsive control problem. This is done by proposing a learning algorithm for tuning parameters of a function approximator, which, once tuned offline, provides feedback solution on-the-fly. The scheme is shown to handle single and multiple impulsive actuators with a small online computational burden. Afterward, the controller is applied to a challenging problem, namely, the orbital maneuver of spacecraft with the fixed final time using impulsive actuators. The objective is triggering the actuators in a fuel-optimal manner such that the spacecraft transfers to the desired orbit at a prescribed time. It was shown that the proposed scheme leads to simultaneous and feedback path planning and control for the maneuver. The potentials of the scheme are analyzed in different scenarios, including enforcing a shorter final time, selecting different initial states, and incorporating actuator faults.

This paper presents the mission design for a CubeSat-based active debris removal approach intended for transferring sizable debris objects from low-Earth orbit to a deorbit altitude of 100 km. The mission consists of a mothership spacecraft that carries and deploys several debris-removing nanosatellites, called Deorbiter CubeSats. Each Deorbiter is designed based on the utilization of an eight-unit CubeSat form factor and commercially-available components with significant flight heritage. The mothership spacecraft delivers Deorbiter CubeSats to the vicinity of a predetermined target debris, through performing a long-range rendezvous maneuver. Through a formation flying maneuver, the mothership then performs in-situ measurements of debris shape and orbital state. Upon release from the mothership, each Deorbiter CubeSat proceeds to performing a rendezvous and attachment maneuver with a debris object. Once attached to the debris, the CubeSat performs a detumbling maneuver, by which the residual angular momentum of the CubeSat-debris system is dumped using Deorbiter’s onboard reaction wheels. After stabilizing the attitude motion of the combined Deorbiter-debris system, the CubeSat proceeds to performing a deorbiting maneuver, i.e., reducing system’s altitude so much so that the bodies disintegrate and burn up due to atmospheric drag, typically at around 100 km above the Earth surface. The attitude and orbital maneuvers that are planned for the mission are described, both for the mothership and Deorbiter CubeSat. The performance of each spacecraft during their operations is investigated, using the actual performance specifications of the onboard components. The viability of the proposed debris removal approach is discussed in light of the results.

The construction of a space elevator would be an inspiring feat of planetary engineering of immense cost and risk. But would the benefit outweigh the costs and risks? What, precisely, is the purpose for building such a structure? For example, what if the space elevator could provide propellant-free (free release) orbital transfer to every planet in the solar system and beyond on a daily basis? In our view, this benefit might outweigh the costs and risks. But can a space elevator provide such a service? In this manuscript, we examine 3 tiers of space elevator launch system design and provide a detailed mathematical analysis of the orbital mechanics of spacecraft utilizing such designs. We find the limiting factor in all designs is the problem of transition to the ecliptic plane. For Tiers 1 and 2, we find that free release transfers to all the outer planets is possible, achieving velocities far beyond the ability of current Earth-based rocket technology, but with significant gaps in coverage due to planetary alignment. For Tier 3 elevators, however, we find that fast free release transfers to all planets in the solar system is possible on a daily basis. Finally, we show that Tier 2 and 3 space elevators can potentially use counterweights to perform staged slingshot maneuvers, providing a velocity multiplier which could dramatically reduce transit times to outer planets and interstellar destinations., Comment: updated for final submission to journal Acta Astronautica

This paper presents a novel design of micro solar sails for emerging lightweight chip scale spacecraft based on flexible electronic circuits. To acquire large deformation, the micro solar sails were designed to be bilayer beams that were able to be electro-thermally actuated by Joule heating. The concept design of the solar sail with high area-to-mass ratios allowed the solar sailing system, named as ChipSail, for efficient orbital transfer and attitude adjustment. The principle of solar sailing with ChipSail was illustrated, and the thickness of the two metals for the bilayer sails should be no more than 1 µm, so as to achieve the efficient solar sailing. Then, the fabrication and characterization of such bilayer microstructures for solar sails were introduced briefly. After that, the electro-thermal analysis of such solar sails deployed on the low earth orbit was carried out, and it was found that the balanced temperature of the sails under the effect of solar radiation and thermal reemission of the sails was 315.31 K, followed by electro-thermal modelling of the sails under the Joule heating. A nonlinear second order differential equation was derived, which allowed rapid prediction of the thermal distribution across the sail. Equivalence of the bilayer solar sail to a width-changing 980 µm long bilayer beam was proposed and validated by finite element analysis. Finally, the thermo-mechanical model on the bilayer sail was then established and solved numerically. Results showed that the maximum bending angle could reach to 94.05o by applying a voltage of 0.05 V across the sail. The electro-thermo-mechanical model laid a solid foundation for dynamic control of the configuration of the ChipSail for efficient orbital transfer and attitude adjustment in space.

To transfer a satellite or a spacecraft from a low parking orbit to another orbit requires one of the many orbital transfers. These orbital transfers need to determine some orbital elements of the initial and final orbits as perigee and apogee distances. The transfers compete to achieve the transition with minimal consumption of energy, transfer time, as well as the highest accuracy of transition. In the present research, certain mathematical procedures implementable with the help of computers will be employed to investigate the two important non-coplanar tangential transfers between circular orbits called the Hohmann and bi-elliptic transfers. Also, a comparative study between Hohmann and bi-elliptic transfers will be established. At the end of present study, we will be able to determine the lowest value of the velocity change and the best transfer between the Hohmann and bi-elliptic transfers with minimal fuel consumption.

The past few decades have seen many meaningful trials on development of solar sail spacecraft with high area-to-mass ratios and thus allowing efficient, propellant-free photon propulsion. However, the restriction to bending stiffness of thin solar films has proven to be an obstacle for building solar sails with high area-to-mass ratios. This paper presents a novel design of micro solar sails for emerging lightweight chip scale spacecraft based on flexible electronics. A representative concept design of the chip scale solar sail spacecraft with area-to-mass ratios over 100 m2/kg is proposed, which enables efficient orbital transfer and attitude adjustment. To acquire large deformation for adjustment of forces, the micro solar sails consisting of bilayer beams are actuated by Joule heating. Electro-thermal analysis on the sails after deployed into space should be the prioritized and fundamental problem for following evaluation of deflection and thus change of solar radiation pressure on the reflective sails. Solar radiation, thermal reemission, and Joule heating are incorporated into the electro-thermal model for such solar sails deployed in geospace. A nonlinear second order differential equation is derived, which is solved numerically for temperature distribution across the sail. The proposed electro-thermal model is validated by finite element analysis and lays a solid foundation for the thermo-mechanical modelling, as well as flight modelling, of the chip scale spacecraft with such micro solar sails.

Made available in DSpace on 2021-06-25T11:05:12Z (GMT). No. of bitstreams: 0 Previous issue date: 2020-11-01 Fundação de Amparo à Pesquisa do Estado de São Paulo (FAPESP) The main objective of the present paper is to find the modifications that a tethered slingshot maneuver (TSSM) can make in the orbit of a spacecraft, both in terms of energy and inclination. The TSSM is a maneuver where a tether fixed in a celestial body, like a moon or an asteroid, makes a rotation in the velocity vector of a spacecraft to modify its orbit. In particular, the present paper concentrates in showing the potential savings in fuel consumption for orbital maneuvers that use this technique in an elliptic system of primaries, which gives advantages over the circular problem. To make this study more complete, analytical approximations are derived to provide a general view of the behavior of the maneuver in terms of variations of energy and inclination as a function of different conditions for the geometry, length, and location of the tether. Among the main results, it is showed the best location to place the tether and the best moment and duration to perform the maneuver, as a function of the parameters involved, like the orbits of the primaries, the incoming velocity of the spacecraft, etc. The solutions are shown in maps giving the variations of energy and inclination for different locations of this device and assuming different incoming orbits for the spacecraft. Regions that maximize those variations are indicated. Based on those results, it is possible to find the best solution for several particular problems. The results show that this maneuver has a large potential to be explored, helping a spacecraft to make journeys to the exterior planets and out of the Solar System. Those results are arguments in favor of developing efforts to solve the technological problems involved in real applications of this technique. The main advantage of the proposed technique is the energy gain given by the maneuver, in particular when using the higher velocity of the asteroid, at the periapsis of its orbit around the Sun. The main disadvantages are the technical challenges involved in the implementation of the maneuver and the fact that to get maximum benefits, there are time restrictions to apply the maneuver, because the asteroid must be passing by its periapsis. Universidade Estadual Paulista – UNESP Instituto Nacional de Pesquisas Espaciais- INPE Universidade Estadual Paulista – UNESP FAPESP: 2016/23542-1 FAPESP: 2016/24561-0 FAPESP: 2019/15180-0

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.

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/or 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., Comment: 9 pages, 12 figures, submitted to Advances in Space Research

Low-thrust spacecraft trajectory optimization for the many-revolution orbital transfer problem is especially challenging due to the high problem dimension and perturbing accelerations that prevent ...

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.

This paper describes a Markov decision process approach to a robust spacecraft mission control policy that maximizes the expected value of science reward assuming a circular orbit. The control policy that governs mission steps can be computed off-board or onboard depending upon the availability of communication bandwidth and on-board computational resources. This paper considers a sample science mission, where the spacecraft collects data from celestial objects viewable only within a certain orbit true anomaly window. Science data collection requires the spacecraft to slew its instrument(s) toward each target, and continue pointing in the direction of the target while the spacecraft traverses its orbit. Robustness and stochastic optimization of scientific reward, is achieved at the cost of computational complexity. Approximate dynamic programming (ADP) is exploited to reduce the computational time and effort to manageable levels and to treat larger problem sizes. The proposed ADP algorithm partitions the state-space based on true anomaly regions, enabling grouping of adjacent science targets. Results of a simulation case study demonstrate that our proposed ADP approach performs quite well for reasonable ranges of key problem parameters.

This paper introduces an Optimal Fuzzy Sliding Mode (OFSM) controller for the control of fuel sloshing in a spacecraft during orbital maneuvers. In order to optimize controller's gains for minimizing control inputs, a Particle Swarm Optimization (PSO) algorithm is employed. A novel gain scheduling of the OFSM controller is developed for the coupled spacecraft-fuel system. Fuzzy gains are produced on-line based-on rules that utilize dynamic system variables and the sign of saturation function of sliding surfaces. The fuel slosh is modeled as a fixed spherical mass and two spherical pendulum masses moving in a spherical tank. Computer simulations illustrate the effectiveness of the proposed controller in maneuvering spacecraft while suppressing fuel sloshing. The performance of the OFSM controller is compared to that of the optimum State-Dependent Riccati Equation (SDRE) controller.

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.

In this paper, it was studied the optimization of the cost of interplanetary missions with emphasis on reducing fuel consumption. To achieve this goal, a genetic algorithm was implemented to optimize the total impulse of orbital transfer. It was implemented a case of sending a space vehicle from Earth to a another planet using a gravity assist maneuver (swing by), in this paper it was chose sending a spacecraft from Earth to Mars with a close approach to the Venus. The method employed can be used for impulsive interplanetary missions in general, and so the solution found can become an initial solution for numerical methods of optimization of low thrust maneuvers.

This paper lies in the framework of mission scenarios, such as Active Debris Removal and On-Orbit Servicing, which require an active spacecraft (chaser) to orbit in close-proximity with respect to a space target. Specifically, these activities involve relative orbital maneuvers, such as monitoring, rendezvous and docking, in which the target-chaser distance ranges from a few tens of meters (depending on the target size) up to contact (in the case of docking). A critical challenge related to the realization of these maneuvers is the need to minimize the risk of collision, considering that the target is a non-cooperative object which may be characterized by uncontrolled rotational dynamics. This goal can be achieved by designing relative trajectories which satisfy specific constraints in terms of safety and stability, on one side, as well as by exploiting relative navigation technologies and algorithms which provide highly accurate estimates of the target-chaser relative motion parameters thus allowing to relax the control requirements. Both these aspects are addressed by this paper with focus on Geostationary Earth Orbits since they represent a particularly crowded orbital region in which the possibility to remove large debris and to extend the operative life of spacecraft, such as telecommunication ones, may have a significant scientific and economic benefit. Hence, an original method is presented to design safety ellipses for target monitoring around GEO targets, which, simultaneously, can provide optimal relative observation geometry for relative navigation (pose determination) using Electro-Optical sensors. The design approach is formulated in mean orbit parameters and it is based on a relative motion model relevant to two-satellite formations which includes the non-Keplerian perturbations due to secular Earth oblateness, as well as the possibility of considering targets moving along a small-eccentricity orbit. An example of trajectory design is shown considering a GEO target as test case. Given this trajectory, pose determination performance is also evaluated within a numerical simulation environment capable of realistically reproducing target-chaser relative dynamics, the operation of a scanning LIDAR selected on board the chaser as relative navigation sensor, and pose estimation algorithms based on the processing of 3D point clouds.

This paper considers the nonlinear interaction of the rigid body dynamics of a spacecraft with the sloshing dynamics of the liquid propellant in its partially filled spherical tank during an orbital transfer. Fuel slosh dynamics are described by a three-dimensional two-pendulum model characterizing the first two sloshing modes, and the coupled equations of motion of the vehicle and the fuel masses are derived by means of quasi-Lagrangian equations. In addition, this paper studies active nonlinear control methods to control the attitude of the spacecraft while suppressing the sloshing of fuel. Simulations are conducted to illustrate the effectiveness of these control methods.

Thse powered aero-gravity-assist is an orbital maneuver that combines three basic components: a gravity-assist with a passage by the atmosphere of the planet during the close approach and the application of an impulse during this passage. The mathematical model used to simulate the trajectories is the Restricted Three-Body Problem including the terms coming from the aerodynamic forces. The present paper uses this type of maneuver considering that the trajectory of the spacecraft is in the ecliptic plane and the presence of the atmospheric Drag and Lift forces. The maneuver in the ecliptic plane can be done due to technologies that provides spacecraft with high values for the Lift to Drag ratio. The main advantage is that this maneuver allows the modification of the semi-major axis of the orbit of the spacecraft using the gravity of the planet and, at the same time, to change the inclination, using the high Lift that is perpendicular to the ecliptic plane. So, it is a combined maneuver that changes two important orbital parameters at the same time. The Lift is applied orthogonal to the initial orbital plane to generate an inclination change in the trajectory of the spacecraft, which is a very expensive maneuvers when made using propulsion systems. The Lift to Drag ratio used in the present paper goes up to 9.0, because there are vehicles, like waveriders, designed to have these values. When the spacecraft is passing by the periapsis of its orbit, an instantaneous impulse is applied to increase or decrease the variation of energy given by the aero-gravity-assist maneuver. The planets Venus and Mars are selected to be the bodies for the maneuver, due to their atmospheric density and strategic location in the Solar System to provide possible uses for future missions. Results coming from numerical simulations show the maximum changes in the inclination obtained by the maneuvers, as a function of the approach angle and direction of the impulse; the Lift to Drag ratio and the ballistic coefficient. In the case of Mars, inclination changes can be larger than 13°, and for Venus larger than 21°. The energy and inclination variations are shown for several selected orbits. The powered aero-gravity-assist maneuver generates inclination changes that are higher than the ones obtained from the powered maneuver and/or the aero-gravity maneuver.

As large numbers of increasingly smaller spacecraft continue to be launched, means of efficient and reliable orbital maneuvering and orbit disposal have become increasingly necessary. For spacecraft that do not contain thrusters, aerodynamic drag modulation using a retractable drag device or attitude changes presents itself as an efficient way to perform orbital maneuvers and control the re-entry location. This paper introduces an aerodynamically based re-entry guidance generation algorithm for low Earth orbit spacecraft that exhibits significant accuracy, robustness, and efficiency. The paper also presents a novel guidance tracking algorithm whereby the drag device of a spacecraft is deployed or retracted relative to a nominal deployment profile (given in the guidance) based on the difference between the actual and desired state of the spacecraft. A full state feedback linear-quadratic-regulator control scheme is utilized with the Schweighart Sedgwick equations of relative motion to drive the relative position and velocity between the spacecraft and the guidance trajectory to zero. A problem-specific Extended Kalman Filter implementation is also introduced to remove noise from the GPS-derived relative motion estimate. One thousand Monte Carlo simulations of the guidance generation algorithm with randomized initial conditions and desired re-entry locations are conducted, resulting in an average guidance error of 12.5 k m and a maximum error below 106 k m . The tracking of these aerodynamic decay guidances with the aforementioned algorithms is also simulated with drag force uncertainties up to a factor of two and navigation errors (noise and bias) comparable to that expected from a CubeSat GPS unit. Despite these simulated errors and uncertainties, this approach provides guidance tracking down to a re-entry altitude of 120 k m with a final position error under 6 k m for all cases. The algorithms detailed in this paper provide a way for any spacecraft capable of modulating its drag area to autonomously perform orbital maneuvers and execute a precise re-entry.

The paper gives examples of orbital maneuver execution of spacecraft (SC) using orbit correction propulsion system (OCPS) based on electric propulsion (EP). The high velocity of the propellant flow achieved in EP allows for orbital maneuvers with significantly lower propellant flow rate than in conventional propulsion systems. The time required to perform the orbital maneuvers using EP is connected with the available on-board power. It is typically much greater than such time needed for conventional jet propulsions. The advantages of EP are realized if mission allows a long-time operation of the propulsion system. Since the early 1970s stationary plasma thrusters (SPTs) developed on the base on the concept proposed by A.I. Morozov are used on the SC Meteor. With the help of OCPS based on SPT EOL-1, SC Meteor was installed on the conventionally synchronous orbit, which provides a fixed grid of tracks with a period of T = 102.31 min. In this case, a complete overview of all daily Earth’s surface with the bandwidth of 2900 km is obtained. In recent decades, there has been a steady trend toward miniaturization of space equipment, which requires the development of acceptable thrusters to meet new requirements. Typical total impulses of thrust required for OCPS are reduced several times. At present, a number of space constellations are being developed in Russia based on small SC with a mass from 60 to 500 kg. VNIIEM Corporation creates a space constellation IONOZOND, designed to monitor the geophysical conditions. The description of the IONOZOND constellation is given and the options for using of various OCPS, in particular, based on SPT, ion and pulsed plasma thrusters are considered. It is shown that their use in small SC can significantly increase the economic efficiency of remote sensing orbital constellations.

The primary goal of this work is to extend the work done in, [1], to provide high accuracy satellite attitude and orbit estimates needed for imaging purposes and also before execution of spacecraft orbital maneuvers for the next Egyptian scientific satellite. The problem of coarse satellite attitude and orbit estimation based on magnetometer measurements has been treated in the literature. The current research expands the field of application from coarse and slow converging estimates to accurate and fast converging attitude and orbit estimates within 0.1o, and 10 m for attitude angles and spacecraft location respectively (1-σ). The magnetometer is used for both spacecraft attitude and orbit estimation, aided with gyro to provide angular velocity me a su r eme n t s , star sensor to provide attitude quaternion, and GPS receiver to provide spacecraft location. The spacecraft under consideration is subject to solar radiation pressure forces and moments, aerodynamic forces and moments, earth’s oblateness till the fourth order (i.e. 4 J ), gravity gradient moments, and residual magnetic dipole moments. The estimation algorithm developed is powerful enough to converge quickly (actually within 10 seconds) despite very large initial estimation errors with sufficiently high accuracy estimates.

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 performa...

In the process of on-orbit servicing, chasers achieve approach and contact with target spacecraft are via a series of orbital transfer and motion control. However, most target spacecraft are non-cooperative targets. As a matter of fact, mainly technical difficulty of non-cooperative targets still lies in unable directly obtaining their motion parameters, thereby, it is impossible to realize their autonomously rendezvous and capture. Thus, measure motion parameters of non-cooperative targets from outside is significant for research on-orbit service. In this paper, a pose measurement method coupled initial pose measurement with model-based tracking is put forward for space non-cooperative targets. Specifically, an algorithm based on SIFT matching to measure initial pose of target tracking is proposed to start subsequent target tracking. Afterwards, a target tracking algorithm is presented, which is via a priori information and the combination of edge features and point features. As well as, an M estimator is introduced to improve the accuracy and robustness of target tracking algorithm. Finally, an experimental platform is constructed to fulfill experiments to validate the effectiveness of proposed method. Also, the experimental results demonstrate that our approach has highly precision and robustness in vary illuminations. Simultaneously, our method could satisfy the real-time requirement of pose measurement for non-cooperative targets.

The usage of state transition tensors (STTs) was proved as an effective method for orbital uncertainty propagation. However, orbital maneuvers and their uncertainties are not considered in current STT-based methods. Uncertainty propagation of spacecraft trajectory with maneuvers plays an important role in spaceflight missions, e.g., the rendezvous phasing mission. Under the effects of impulsive maneuvers, the nominal trajectory of a spacecraft will be divided into several segments. If the uncertainty is piecewise propagated using the STTs one after another, large approximation errors will be introduced. To overcome this challenge, a set of modified STTs is derived, which connects the segmented trajectories together and allows for directly propagating uncertainty from the initial time to the final time. These modified STTs are then applied to analytically propagate the statistical moments of navigation and impulsive maneuver uncertainties. The probability density function is obtained by combining STTs with the Gaussian mixture model. The proposed uncertainty propagator is shown to be efficient and affords good agreement with Monte Carlo simulations. It also has no dimensionality problem for high-dimensional uncertainty propagation.

MIT Lincoln Laboratory, in collaboration with Blue Canyon Technologies LLC and Enpulsion GmbH, is developing the Agile MicroSat (AMS) as the first nanosatellite with suitable agility to enable long-duration low altitude flight and maneuvering to actively focus ground coverage in times of need. Operating at lower altitudes provides higher resolution for a given optical sensing aperture. Spacecraft agility provides the opportunity for earth remote sensing techniques observing transient, unpredictable Earth scenes such as agricultural or ecological stress, fire smoke plumes, coastal and river flooding and oil spills. AMS will use an indium-fed field effect electric propulsion thruster with thrust vector control to implement its orbital maneuvers. AMS is implementing advanced satellite control techniques including vector control momentum management during thrusting, low-drag attitude control and critical fault response in low altitude/high drag conditions. AMS will validate increasingly sophisticated and demanding guidance techniques, progressing from manual ground control to automated on-board guidance as confidence is gained through on-orbit operation. Upon reaching low altitude, AMS will maneuver into repeating satellite ground track orbits on demand to operate two optical payloads. The AMS Camera payload will collect reflected sunlight imagery of local regions of interest cued by mission operations. The AMS Beacon payload provides a near infrared laser transmitter as an artificial phase reference to demonstrate ground based adaptive optics. Mission planning is ongoing leading to anticipated launch in Q4 2021 and minimum altitude of 280 km or lower.

The use of equinoctial elements and their corresponding variational equations alleviates the problems of singularity and instability that are associated with the classical orbital elements. As such, the equinoctial elements are well suited for the long-term simulations of satellite orbit. Compared to the classical orbital elements, however, the equinoctial elements are relatively abstract as they provide little direct insight into the shape and orientation of an orbit. The motivation of this work is to derive optimum guidance laws in terms of the equinoctial elements, for the adjustment of each of the classical orbital elements. The optimum thrust laws are obtained by analyzing Gauss' variational equations through the inclusion of two steering angles in their formulations, namely azimuth (in-plane) and elevation (out-of-plane) angles. In particular, considering the classical two-body problem with the inclusion of the thrust force as the source of perturbations, the optimum thrusting strategies are derived for the modification of each orbital element. A number of case studies are presented, in which one or more elements are sought to be adjusted. Thruster characteristics and the amount of consumed propellant in each maneuver are determined using the performance parameters of a commercial-off-the-shelf, state-of-the-art low-thrust propulsion system intended for SmallSats. For the scenarios where more than one element is to be modified, the Directional Adaptive Guidance law is utilized which synthesizes multiple thrust directions into a single thrust force for inclusion in the variational equations. A detailed discussion on the tuning of the weighting factor and the adaptation law that are included in the Directional Adaptive Guidance law is presented, and some strategies for expediting convergence to the desired orbits are discussed. The results show that, for the long-term simulation of spacecraft's trajectory, the equinoctial-based, optimum guidance laws are most suitable, and they can handle the situations wherein classical orbital elements and their variational equations behave erratically due to the singularity of one or more of the elements. As illustrated in this paper, one such scenario is when the spacecraft is to be transferred to the geostationary orbit, where orbital eccentricity and inclination are zero, rendering several of the classical orbital elements undefined and Gauss variational equations unstable when the spacecraft nears the target geostationary orbit.

Publisher Summary Orbital maneuvers transfer spacecraft from one orbit to another. Orbital changes can be dramatic as the transfer from a low-earth parking orbit to an interplanetary trajectory can also be quite small, as in the final stages of the rendezvous of one spacecraft with another. Changing orbits requires the firing of onboard rocket engines. This chapter focuses on impulsive maneuvers in which the rockets fire in relatively short bursts to produce the required velocity change. Impulsive maneuvers are those in which brief firings of on-board rocket motors change the magnitude and direction of the velocity vector instantaneously. During an impulsive maneuver, the position of the spacecraft is considered to be fixed and only the velocity changes. The chapter presents the classical, energy-efficient Hohmann transfer maneuver, and generalizes it to the bi-elliptic Hohmann transfer to see if even more efficiency can be obtained. The Hohmann transfer is the most energy efficient two-impulse maneuver for transferring between two coplanar circular orbits sharing a common focus. The Hohmann transfer is an elliptical orbit tangent to both circles on its apse line and helpful in sorting out orbit transfer strategies to use the fact that the energy of an orbit depends only on its semimajor axis.

Green propellants are currently considered as enabling technology that is revolutionizing the development of high-performance space propulsion, especially for small-sized spacecraft. Modern space missions, either in LEO or interplanetary, require relatively high-thrust and impulsive capabilities to provide better control on the spacecraft, and to overcome the growing challenges, particularly related to overcrowded LEOs, and to modern space application orbital maneuver requirements. Green monopropellants are gaining momentum in the design and development of small and modular liquid propulsion systems, especially for CubeSats, due to their favorable thermophysical properties and relatively high performance when compared to gaseous propellants, and perhaps simpler management when compared to bipropellants. Accordingly, a novel high-thrust modular impulsive green monopropellant propulsion system with a micro electric pump feed cycle is proposed. MIMPS-G500mN is designed to be capable of delivering 0.5 N thrust and offers theoretical total impulse Itot from 850 to 1350 N s per 1U and &gt, 3000 N s per 2U depending on the burnt monopropellant, which makes it a candidate for various LEO satellites as well as future Moon missions. Green monopropellant ASCENT (formerly AF-M315E), as well as HAN and ADN-based alternatives (i.e., HNP225 and LMP-103S) were proposed in the preliminary design and system analysis. The article will present state-of-the-art green monopropellants in the (EIL) Energetic Ionic Liquid class and a trade-off study for proposed propellants. System analysis and design of MIMPS-G500mN will be discussed in detail, and the article will conclude with a market survey on small satellites green monopropellant propulsion systems and commercial off-the-shelf thrusters.

Application of high-power laser propulsion to explore a near-Earth space is discussed for a long time, starting from the pioneer investigations by Arthur Kantrowitz (1972) and Prof. A.M. Prokhorov (1976). A number of theoretical and experimental works on the laser propulsion were carried out in Russia, the USA, Germany, Japan, China, Brazil, and Australia since this time. Various models of spacecrafts with a laser propulsion were proposed with all this going on. One of these vehicles is the Lightcraft Technology Demonstrator developed by Prof. Leik Myrabo to be applied to launching of satellites into near-Earth space. In the chapter, we present a space mini-vehicle with laser propulsion produced by the aerospace laser-propulsion engine. The mini-vehicle optical system is designed to satisfy the principal conditions defining the vehicle application in a space. Particularly, one of these conditions is an independence of the vehicle orbital maneuver on a mutual orientation of the vehicle and laser power beam. This is achieved by using auxiliary onboard optical units such as receiver telescope, optical turret and hinges.

This paper compares possibility for the spacecraft equipped with liquid-fuelled rocket engine to change orbital plane for purposes of space debris collection. The simple method of orbital plane rotation angle calculation is described. Also recommendations for fuel components selection made using possible orbital plane rotation angle as the main factor.

This paper presents a new approach for autonomous reconfiguration of distributed space systems, which ensures safe guidance of spacecraft formations towards the desired patterns while optimizing the total propellant consumption. The orbital transfer is reduced to the form of a convex optimization problem to guarantee rapid computation of control laws. Hence, tasks are iteratively assigned to the component platforms to detect the best reconfiguration strategy. The path-planning is entrusted to a reference satellite of the cluster, that coordinates the remaining ones by means of a procedure based on genetic algorithms. Two methods are proposed, depending on the organizational architecture of the spacecraft formation. In the first one, the maneuver is completely planned by the reference satellite, that determines final tasks and control actions for the whole cluster. As an alternative to such a fully-centralized approach, a distributed version of the algorithm is proposed: tasks are sorted by the reference satellite and transfer orbits are computed by exploiting the computational resources of the whole cluster. Whatever the considered framework, both the planners ensure a safe transition of the formation towards the target geometry. Simulation results show that, when relative distances are of the order of hundreds of meters, a mean delta-v per satellite of the order of 0.1 m/s is required to reconfigure LEO clusters within one orbital period. (C) 2020 COSPAR. Published by Elsevier Ltd. All rights reserved.

With the continuous development of space rendezvous technology, more and more attention has been paid to the study of spacecraft orbital pursuit-evasion differential game. Therefore, we propose a pursuit-evasion game algorithm based on branching improved Deep Q Networks to obtain a space rendezvous strategy with non-cooperative target. Firstly, we transform the optimal control of space rendezvous between spacecraft and non-cooperative target into a survivable differential game problem. Next, in order to solve this game problem, we construct Nash equilibrium strategy and test its existence and uniqueness. Then, in order to avoid the dimensional disaster of Deep Q Networks in the continuous behavior space, we construct a TSK fuzzy inference model to represent the continuous space. Finally, in order to solve the complex and timeconsuming self-learning problem of discrete action sets, we improve Deep Q Networks algorithm, and propose a branching architecture with multiple groups of parallel neural Networks and shared decision modules. The simulation results show that the algorithm achieves the combination of optimal control and game theory, and further improves the learning ability of discrete behaviors. The algorithm has the comparative advantage of continuous space behavior decision, can effectively deal with the continuous space chase game problem, and provides a new idea for the solution of spacecraft orbit pursuit-evasion strategy.

This investigation deals with the problem of spacecraft relative motion control, which is typically associated with the spacecraft rendezvous and proximity maneuvers. Relative position and linear velocity are considered. A distinguishing attribute of the presented approach is consideration of definitely larger relative distance between the satellites than it is commonly addressed in the literature. The presented control method is applicable in the case where the chief satellite moves in a known, highly elliptical orbit. A quasi-optimal control is found by a model predictive control algorithm, where the nonlinear optimization problem is reduced to quadratic optimization by preliminary estimation of the future control trajectory. Significance of the method has been verified using a computer simulation.

Using thrusters for either orbital maneuvers or attitude control change the current spacecraft mass properties and results in an associated reaction force and torque. To perform orbital and attitude control using thrusters, or to obtain optimal trajectories, the impact of mass variation and depletion of the spacecraft must be thoroughly understood. Some earlier works make rocket-body specific assumptions such as axial symmetric bodies or certain tank geometries hat limit the applicability of the models. Other earlier works require further derivation to implement the provided equations of motion in simulation software. This paper develops the fully coupled translational and rotational equations of motion of a spacecraft that is ejecting mass through the use of thrusters and can be readily implemented in flight dynamics software. The derivation begins considering the entire closed system: the spacecraft and the ejected fuel. Then the exhausted fuel motion in free space is expressed using the thruster nozzle properties and the familiar thrust vector to avoid tracking the expelled fuel in the simulation. Additionally, the present formulation considers a general multi-tank and multi-thruster approach to account for both the depleting fuel mass in the tanks and the mass exiting the thruster nozzles. General spacecraft configurations are possible where thrusters can pull from a single tank or multiple tanks, and the tank being drawn from can be switched via a valve. Numerical simulations are presented to perform validation of the model developed and to show the impact of assumptions that are made for mass depletion in prior developed models.

The present work formulates an orbital transfer for an Earth-to-Earth mission between non coplanar orbits with different altitudes with a special feature: the occurrence of a lunar flyby during the transfer orbit. This lunar flyby is intended to help change the plane of motion of the spacecraft without fuel consumption. Only two-impulsive trajectories are considered with the velocity increments applied at the initial and final orbits. In order to solve this problem, a 3D patched-conic approximation associated with a two-point boundary value problem is proposed. The same transfer problem is formulated considering the spatial circular restricted three-body problem (SCR3BP). The results of the patched-conic approximation is compared with the results of the SCR3BP showing a good agreement between the models. This work also determines several trajectories in order to perform a study of the fuel consumption considering several inclinations and altitudes of both initial and final orbits around the Earth. The longitude of the ascending node of the initial orbit, and, the altitude of close approach with the Moon during the flyby are also analyzed. According to the total velocity increment analysis, the changing plane assisted by a lunar flyby can be very favorable. Despite the increase of the time of flight, the saving of fuel is considerable. Indeed, the total velocity increment of this kind of maneuver is in some cases better than the velocity increment provided by the bi-parabolic transfer.

During the spacecraft (SC) mission, some of orbital elements change during the flight, the need to study the spacecraft maneuvers become paramount importance need. A full mathematical model for the Spacecraft lifetime maneuvers is deduced, implemented in a simulation scenario with a GUI, and then desired simulation is verified during the different operations' phases such as orbit transfer, propagation, maintenance and deorbiting. In the current article, an interpretation of the carried-out model will be introduced to match a real case of mission analysis as possible for newly designed satellite. Two types of orbits (inclined and Sun synchronous Orbit SSO) are investigated for the SC mission analysis over the whole lifetime phases. The investigation is carried out using two types of power systems (electrical and chemical) using the same propellant tank model with capacity 60 kg. A numerical comparison was concluded for both studied power systems.

2018 IEEE Aerospace Conference (March 3-10, 2018. Yellowstone Conference Center), Montana, USA, 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1810037000

The main idea of this paper is to propose the construction of an “Escape Portal” to send a spacecraft to the exterior planets, or even to make it escape from the Solar System, using a Tethered Sling Shot Maneuver (TSSM) with an asteroid. The construction of this portal allows an unlimited number of maneuvers with the same tether, which is very interesting when considering a possible use for small satellites. This structure would be formed by a tether that remains fixed in an asteroid. At the other end of the tether, a large net is fixed, such that the only action required from the spacecraft to make the TSSM is to hit the net. The net can have a mechanism to open a passage to release the spacecraft when the desired rotation is obtained. This technique would avoid some of the problems that appear when assuming that the spacecraft needs to carry a tether on board that would be released to hit the asteroid just before the maneuver.

This paper studies the proximate satellite interception guidance strategies where both the interceptor and target can perform orbital maneuvers with magnitude limited thrusts. This problem is regarded as a pursuit-evasion game since satellites in both sides will try their best to capture or escape. In this game, the distance of these two players is small enough so that the highly nonlinear earth-centered gravitational dynamics can be reduced to the linear Clohessy-Wiltshire (CW) equations. The system is then simplified by introducing the zero effort miss variables. Saddle solution is formulated for the pursuit-evasion game and time-to-go is estimated similarly as that for the exo-atmospheric interception. Then a vector guidance is derived to ensure that the interception can be achieved in the optimal time. The proposed guidance law is validated by numerical simulations. Keywords: Differential games, Saddle solution, Satellite interception, Time-to-go estimation, Zero effort miss trajectory

The orbital dynamics basis for orbital plane change maneuver using chemical propulsion is the impulsive orbital change theory in practical engineering. The orbital plane change theory using continuous thrust suitable for electric propulsion is studied in this paper. A set of nonsingular orbital variational equations using quaternion is used to investigate the orbital motion of spacecraft under constant normal acceleration departing from a Keplerian circular orbit firstly. Results show that the orbit of spacecraft under continuous constant normal acceleration is a circular orbit over the gravitational center on the same spherical surface as the initial orbit, and it is tangent to the initial circular orbit at the initial position. Continuous and discontinuous maneuver strategies are then designed using the previous theory, and they are validated by numerical simulations. Results indicate that the strategy can work effectively.

The thrust vector control (TVC) scheme is a powerful method in spacecraft attitude control. Since the control of a small spacecraft is being studied here, a solid rocket motor (SRM) should be used instead of a liquid propellant motor. Among the TVC methods, gimbaled-TVC as an efficient method is employed in this paper. The spacecraft structure is composed of a body and a gimbaled-SRM where common attitude control systems such as reaction control system (RCS) and spin-stabilization are not presented. A nonlinear two-body model is considered for the characterization of the gimbaled-thruster spacecraft where, the only control input is provided by a gimbal actuator. The attitude of the spacecraft is affected by a large exogenous disturbance torque which is generated by a thrust vector misalignment from the center of mass (C.M). A linear control law is designed to stabilize the spacecraft attitude while rejecting the mentioned disturbance torque. A semi-analytical formulation of the region of attraction (RoA) is developed to ensure the local stability and fast convergence of the nonlinear closed-loop system. Simulation results of the 3D maneuvers are included to show the applicability of this method for use in a small spacecraft.

A numerical optimization study of minimum-fuel low-Earth-orbit aeroglide and aerothrust aeroassisted orbital transfer of a small spacecraft subject to constraints on heating rate and heating load w...

This paper introduces a mission concept for active removal of orbital debris based on the utilization of the CubeSat form factor. The CubeSat is deployed from a carrier spacecraft, known as a mothership, and is equipped with orbital and attitude control actuators to attach to the target debris, stabilize its attitude, and subsequently move the debris to a lower orbit where atmospheric drag is high enough for the bodies to burn up. The mass and orbit altitude of debris objects that are within the realms of the CubeSat’s propulsion capabilities are identified. The attitude control schemes for the detumbling and deorbiting phases of the mission are specified. The objective of the deorbiting maneuver is to decrease the semi-major axis of the debris orbit, at the fastest rate, from its initial value to a final value of about 6471 km (i.e., 100 km above Earth considering a circular orbit) via a continuous low-thrust orbital transfer. Two case studies are investigated to verify the performance of the deorbiter CubeSat during the detumbling and deorbiting phases of the mission. The baseline target debris used in the study are the decommissioned KOMPSAT-1 satellite and the Pegasus rocket body. The results show that the deorbiting times for the target debris are reduced significantly, from several decades to one or two years.

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.

An image-based servo controller for the guidance of a spacecraft during non-cooperative rendezvous is presented in this paper. The controller directly utilizes the visual features from image frames of a target spacecraft for computing both attitude and orbital maneuvers concurrently. The utilization of adaptive optics, such as zooming cameras, is also addressed through developing an invariant-image servo controller. The controller allows for performing rendezvous maneuvers independently from the adjustments of the camera focal length, improving the performance and versatility of maneuvers. The stability of the proposed control scheme is proven analytically in the invariant space, and its viability is explored through numerical simulations.