Orbital Maneuvering with Spinning Electrodynamic Tethers

Study on electrodynamic tether system for space debris removal

Comparison of technologies for deorbiting spacecraft from low-earth-orbit at end of mission

The article discusses a field geophysical complex that includes the Earth's gravity and magnetic fields parameters meters — an digital zenith camera (DZC), a relative gravimeter, onboard quantum magnetometer and ground magnetovariation station. This complex allows determining such parameters of the Earth's gravity field as acceleration and anomalies of gravity, plumb line deviations, horizontal components of the acceleration of gravity and gravitational gradients, and parameters of the Earth's magnetic field — absolute and anomalous values of induction. Maps of the Earth's gravity and magnetic fields can be used in various fields, in particular, in the creation and testing of integrated autonomous navigation systems using the parameters of these fields. The paper presents examples of parameter maps created using this complex. The measurement error of acceleration and gravity anomaly is 10 µGal, plumb line deviation is ≈ 0.2″, horizontal components of gravity acceleration are ≈ 1 mGal, gravitational gradients are 10 Eötvös at a distanc

The electro-dynamic tether system, which could generate electro-dynamic force during on-orbit period, is a novel and efficient spacecraft deorbiting system. In this paper, the space environment modeling and off-orbit efficiency analysis of the bare electro-dynamic tether based deorbiting system is investigated. Firstly, the geomagnetic field model and ionospheric model are established, based on which the electric force generated by bare rope is calculated. Furthermore, the models of atmospheric resistance, earth non-spherical perturbation and other environmental perturbation are also established. Then, numerical simulation for the deorbiting process of the electro-dynamic tether system is carried out under different orbital heights, orbital inclines and lengths respectively. Finally, the effectiveness of relative factors to the system is analyzed based on the simulation results.

In the present research, optimal control problem for low thrust spacecraft orbital transfer is investigated by combination of analytical method and Artificial Bee Colony (ABC) algorithm, which is a population-based method. Spacecraft trajectory optimization is defined for orbital transfer from Low Earth Orbit (LEO) to Geosynchronous Earth Orbit (GEO). In the proposed analytical method co-state variables are simplified by employing concepts of optimal control and specific conditions which are considered for the problem, consequently, an adjoint differential equation is added to set of differential equations and these equations are solved simultaneously that result in optimum transfer trajectory. Dynamic modeling of the problem is represented based on Edelbaum equations and trajectory design is carried out considering circular orbits. Optimization problem is defined for different kinds of objective functions such as minimum time, minimum control effort, minimum time with minimum control effort as well as both minimum time and minimum control effort considering different optimal weighted coefficients. These coefficients are searched by means of artificial bee colony algorithm in weighted minimum time minimum effort scheme. Initial values of control variable and terminal time for all defined objective functions are foraged through artificial bee colony algorithm. The rate of changes of the states and control variables are obtained and depicted during orbit transfer. The obtained results demonstrate considerable convergence as well as sufficient accuracy.

Satellites that use low thrust propulsion systems for maneuvering, although efficient, can take a long time to complete significant orbit changes. Determining the non-linear optimal controls for such multirevolution maneuvers using the instantaneous orbital state dynamics can be riddled with numerical errors and are often subject to long computation times due to the large number of discretization nodes required by the optimization algorithm. An approach to optimal control of an electrodynamic tether is examined using averaged orbital state dynamics as constraints instead of instantaneous dynamic constraints. A mean square libration state is introduced in the dynamic model that captures the average of the out-of-plane libration of the tether. A sample long term optimal orbit change maneuver of a librating electrodynamic tether subject to atmospheric drag is investigated. The method of averaging is employed to transform the optimal control problem from the time domain into Fourier space where the complex problem is significantly reduced to a Zermelo type problem that is solved using a pseudospectral method. To validate the dynamic model of averaged states, the instantaneous states are propagated from the initial conditions using the resulting optimal controls.

This paper describes the orbital dynamics characteristics of three strategies for performing multi -kilometer optic al interferometry on astronomical sources in earth orbit, and compares the three •strategies from several points of view, which relate usable data to system costs. The three strategies that are thought to be the most likely to succeed are: 1, the "Hanging Tether" approach in which two collectors are suspended at the ends of a gravity gradient stabilized tether with the central station located on an elevator that moves up and down the tether to keep the optical path lengths equal; 2, the "Free Flyer" approach in which the two collectors and the central station are independent satellites in stationkeeping mode; and 3, the "Spinning Tether" approach in which three collectors are located at the corners of an equilateral tnangle connected by three tethers and rotating about their common center of mass. In each case it is assumed that data is recorded as the baselines continuously traverse the U-v plane in either elliptical or circular paths. It is determined on the basis of previous work in each case that continuous thrusting which is small enough to be provided by ion thrusters is the best method of compensating for gravity gradient disturbances. The fuel requirements for these thrusters is determined for the three cases and compared, based on the assumption of uniform fringe sampling density in the U-V plane. The main conclusion is that all of the systems are feasible in synchronous orbit, but questionable in low earth orbit. Experimental data on the interactions between optical systems, ion thrusters, and tether vibrations is needed before a final choice can be made.

The problem of modeling the braking process of a nanosatellite using a current-conducting tether is considered. The task is related to the actual problem of removing nanosatellites from space that have exhausted their life without the use of jet engines. The interaction of the conducting current of the tether with the magnetic field of the Earth leads to the appearance of the ampere (Lorentz) force that provides braking of the system and a steeper trajectory of the nanosatellite entry into the dense layers of the atmosphere. Modeling and analysis of the braking system is carried out in two models: the model for a straight tether and the model with distributed parameters where the tether is represented by a set of material points. Comparison of the considered equations of motion of the system in terms of estimating the increments of the orbital parameters of the system is considered. The relative influence of aerodynamic forces and ampere forces on the braking process of the system is estimated by the method of mathematical modeling. As an example, a numerical simulation of an electrodynamic tether system motion with a bare tether in the earth's magnetic field is considered.

Mit castor satellite: Design, implementation, and testing of the communication system

Current state of work on Electric Propulsion under Russian Space Agency programs is presented here. The main part of investigation on EP takes place at Keldysh Research Center, EDB «Fakel», TsNIIMash, RIAME MAI, MAI and NIIEM. To create thrusters prototypes with high efficiency and lifetime thermal plasma, arcjet, pulsed plasma, Hall and ion thrusters are working out in the range of power from tens watt to 10 kW. In most organisations the main attention is given to Hall thrusters with closed drift of electrons: with long acceleration zone (stationary plasma thrusters - SPT) and with short acceleration zone (thrusters with anode layer - TAL). Basically, the general approach of this work is to widen developed power utilisation range 0,5 - 1,5 kW down to 100 - 500 W and up to 10 kW (in one module). Works on design of Hall thrusters with power rate up to 50 kW are also carrying out.

A new concept of ion thrusters, metal ion thrusters using magnetron electron-beam bombardment (MIT-MEB) is demonstrated. Ion thrusters provide thrusts via electrostatic fields to accelerate ions. They are widely used in spacecrafts due to the high exhaust speed. Although the thrust is very little, the final velocity of the spacecraft is much higher than using traditional rockets. For propellant, inert gases are used in conventional ion thrusters. Contrarily, a metal propellant is used in MIT-MEB. The metal propellant is solid-state, high density, easy to store, and cheap. The size of the ion thrusters can be reduced dramatically. Therefore, they can be used in a small spacecraft for both attitude and orbit control, deep space exploration, and low earth orbit. The concept of electron-beam physical vapor deposition (EB-PVD) is used to generate metal ions. A metal target is bombarded and thus heated and evaporated. An electric potential accelerates thermal-emitted electrons, which ionize the metal vapor via electron-impacting ionization. The magnetic field of a permanent magnet is used to guide the accelerated electrons towards the center of the target. This increases the efficiency of the process. Particularly but not necessarily, zinc is used for propellant due to its higher vapor pressure compared with other metals at the same temperature. This means that a lower temperature is required for zinc to be vaporized. To demonstrate our concept, an ion thruster of 10.3 ± 0.7 μ N with a power of 26.2 ± 0.7 W was constructed as a prototype. Its mass is less than 500 g, and its diameter is . The accelerating electric potential is 1 kV. Although an optimized design was not developed yet, we demonstrated the feasibility of building metal ion thrusters for the first time. It is a new design space and unexplored method of using a metal propellant.

At present, satellite systems, each comprising hundreds of satellites, are, and are to be, deployed in low orbits. In addition, existing satellite systems are replenished. There has appeared a trend towards the development of modular satellites, which will lead to the development of easy-to-maintain spacecraft consisting of many small structural modules with standardized interface mechanisms. To extend the life of all these systems and reduce their maintenance cost, it is advisable to develop a system for their maintenance. Despite the relatively large number of works on the rendezvous problem, this problem is considered in a somewhat simplified formulation, which is not sufficient for spacecraft servicing in low orbits. As a rule, the consideration is limited to coplanar rendezvous problems in an impulse formulation. In real conditions, rendezvous maneuvers in low orbits are nontrivial. As is known, the orbital parameters of low-orbit spacecraft may differ significantly: the difference in the longitude of ascending nodes (LAN) may reach tens and even hundreds of degrees. Because of this, the energy consumption for an orbital plane change becomes unacceptably high for modern service spacecraft. This energy consumption can be reduced by using the precession of the line of nodes due to the non-centrality of the Earth's gravitational field. A waiting maneuver of a service spacecraft in a well-chosen orbit makes it possible to eliminate the mismatch between the LANs of the service spacecraft’s parking and destination orbits, thus significantly reducing the orbital transfer energy consumption. However, the long wait time of the service spacecraft in its parking orbit significantly increases the total orbital transfer time. The aim of this article is to develop a mathematical model of bicriteria optimization of a transfer of a service spacecraft with a low constant thrust engine between low near-circular orbits with significantly different LANs. This problem is solved by averaging the service spacecraft’s dynamics equations over a fast parameter and using a genetic algorithm of global Pareto optimization. The novelty of the results obtained lies in a formulation of a bicriteria optimization problem and the development of a mathematical model for choosing an optimal service spacecraft parking orbit. The mathematical model developed may be used in planning service spacecraft transfers between low near-circular orbits with significantly different LANs.

The performance capabilities of an experimental Hall thruster obtained experimentally for both variable and constant thrust modes at low power levels enable to implement orbit transfer under the influence of air drag with small satellites. For this purpose, the measured thruster performance was employed in calculations of transfer trajectories. As a result, by applying optimal thrust acceleration control the required propellant mass for a given LEO transfer was reduced by 15-17% as compared to that required for spacecraft operation with a constant thrust to power ratio. For this purpose, a Hall thruster with a movable anode is needed. In addition, the operation of an array of thrusters enables to increase a factor of variations of the thrust acceleration during the flight time.

The present paper describes an advanced diagnostics for erosion detection in both Hall thrusters and Ion engines. This diagnostics has been developed by Alta in the framework of a Technology Research Programme funded by ESA, with the aim of further investigating the erosion phenomenon in such thrusters. The system is installed inside the vacuum facility hosting the test, so allowing for successive measurements of the erosion during the thruster operation. By using two different optical methods, according to the type of target engine, it has been possible to reconstruct the shape of the channel profile in a Hall thruster (Alta’s HT-100) and to acquire high-res images of the grid holes of an Ion engine. An extensive test campaign has been carried out on the HT-100 thruster, measuring the erosion on channels made of three different ceramic materials.

The current status of measuring radiation using JAXA satellites is reviewed. Starting with Engineering Test Satellite‐V (ETS‐V; KIKU‐5 in Japanese) in 1987, efforts to conduct radiation measurements in space have continued using almost all Japan Aerospace Exploration Agency (JAXA formerly NASDA) satellites (ETS‐VI, ADEOS, ADEOS‐II, MDS‐1, DRTS (ongoing), and ALOS (ongoing)), in geostationary orbit (GEO), geostationary ‐transfer orbit (GTO), and low‐Earth orbit (LEO). Electrons, protons, alpha particles, and heavy ions have been the main objects of study. Future plans for radiation monitoring in JAXA, including GOSAT, Jason‐2 (in ollaboration with CNES), SmartSat (in collaboration with NICT), and ISS/JEM/Exposure Facility/SEDA‐AP, are presented.