Thruster Model Research Articles

Performance and lifetime testing of a pulsed plasma thruster for Cubesat applications

This paper describes the design and testing of the engineering model of a pulsed plasma thruster for Cubesat applications (PPTCUP-EM). It has been developed by Mars Space Ltd, Clyde Space Ltd and the University of Southampton with the main aim of increasing the in-orbit lifetime of Cubesats by providing drag compensation. Nevertheless, this thruster can be also utilized to perform formation flying, small orbit changes and Cubesat end of life deorbiting. A test campaign has been carried out to prove that the thruster and the conditioning electronics lifetimes are long enough to utilize all the propellant stored on-board. From the results of the test, the PPTCUP-EM can deliver a total impulse of 42.9±3.9 N s in about 1,125,000 shots. Moreover during the test campaign, a total of more than 1,800,000 shots have been achieved providing a safety factor of about 60% with respect to the number of required shots. The results gathered, including a preliminary characterization of the electromagnetic noise generated by the unit, are presented and show that the overall thruster performance is not influenced by the thruster aging.

Particle-in-cell simulation for different magnetic mirror effects on the plasma distribution in a cusped field thruster**Project supported by the National Natural Science Foundation of China (Grant No. 51006028) and the Foundation for Innovative Research Groups of the National Natural Science Foundation of China (Grant No. 51121004).

Magnetic mirror used as an efficient tool to confine plasma has been widely adopted in many different areas especially in recent cusped field thrusters. In order to check the influence of magnetic mirror effect on the plasma distribution in a cusped field thruster, three different radii of the discharge channel (6 mm, 4 mm, and 2 mm) in a cusped field thruster are investigated by using Particle-in-Cell Plus Monte Carlo (PIC-MCC) simulated method, under the condition of a fixed axial length of the discharge channel and the same operating parameters. It is found that magnetic cusps inside the small radius discharge channel cannot confine electrons very well. Thus, the electric field is hard to establish. With the reduction of the discharge channel’s diameter, more electrons will escape from cusps to the centerline area near the anode due to a lower magnetic mirror ratio. Meanwhile, the leak width of the cusped magnetic field will increase at the cusp. By increasing the magnetic field strength in a small radius model of a cusped field thruster, the negative effect caused by the weak magnetic mirror effect can be partially compensated. Therefore, according to engineering design, the increase of magnetic field strength can contribute to obtaining a good performance, when the radial distance between the magnets and the inner surface of the discharge channel is relatively big.

Optimization of a coaxial electron cyclotron resonance plasma thruster with an analytical model

A new cathodeless plasma thruster currently under development at Onera is presented and characterized experimentally and analytically. The coaxial thruster consists of a microwave antenna immersed in a magnetic field, which allows electron heating via cyclotron resonance. The magnetic field diverges at the thruster exit and forms a nozzle that accelerates the quasi-neutral plasma to generate a thrust. Different thruster configurations are tested, and in particular, the influence of the source diameter on the thruster performance is investigated. At microwave powers of about 30 W and a xenon flow rate of 0.1 mg/s (1 SCCM), a mass utilization of 60% and a thrust of 1 mN are estimated based on angular electrostatic probe measurements performed downstream of the thruster in the exhaust plume. Results are found to be in fair agreement with a recent analytical helicon thruster model that has been adapted for the coaxial geometry used here.

A quantitative optimisation model for a horizontal MEMS solid propellant thruster with experimental verification

This paper studies the effect of multiple geometric parameters of a micronozzle on a MEMS solid propellant thruster (MSPT) and establishes a quantitative optimisation model. Influences of the structural parameters of the micronozzle on performance are accurately calculated. The significance and contribution ratio of each parameter are evaluated by F distribution. Correspondingly, a horizontal MSPT is fabricated and tested to prove the optimisation model. In the implementation, a complete microigniter is formed by a single layer of Au, achieving an ultra-low ignition voltage of 2.1 V with an extremely low ignition resistance of 4 Ω.

Fault Reconstruction Approach for Distributed Coordinated Spacecraft Attitude Control System

This work presents a novel fault reconstruction approach for a large-scale system, that is, a distributed coordinated spacecraft attitude control system. The attitude of all the spacecrafts in this distributed system is controlled by using thrusters. All possible faults of thruster including thrust magnitude error and alignment error are investigated. As a stepping stone, the mathematical model of thruster is firstly established based on the thruster configuration. On the basis of this, a sliding mode observer is then proposed to reconstruct faults in each agent of the coordinated control system. A Lyapunov-based analysis shows that the observer asymptotically converges to the actual faults. The key feature of this fault reconstruction approach is that it can achieve a faster reconstruction of the fault in comparison with the conventional fault reconstruction schemes. It can globally reconstruct thruster faults with zero reconstruction error, and this is accomplished within finite time. The effectiveness of the proposed approach is analytically authenticated via simulation study.

Three-dimensional particle-in-cell simulation of a miniature plasma source for a microwave discharge ion thruster

We have developed a three-dimensional particle model for a miniature microwave discharge ion thruster to elucidate the mechanism of ECR discharges confined in a small space. The model consists of a particle-in-cell simulation with a Monte Carlo collision algorithm (PIC-MCC) for the kinetics of charged particles, a finite-difference time-domain method for the electromagnetic fields of 4.2 GHz microwaves, and a finite element analysis for the magnetostatic fields of permanent magnets. The PIC-MCC results have shown that the electrons are well confined owing to the mirror magnetic fields and can be effectively heated in the ECR layer downstream of a ring-shaped antenna. The confinement results in the ring-shaped profiles of the plasma density along the antenna. The visual appearance of the plasma discharge of the thruster in operation was also ring-shaped. Moreover, the ions are expected to be accelerated effectively through the grid electrode without a large loss of ions toward side walls, that is, the plasma source developed here would be desirable in ion thrusters.

Dynamic Thrust Allocation of Dynamic Positioning Vessel Based on Model Predictive Control

The dynamic positioning (DP) vessel maintains its position and heading by active thrusters. For safety reasons, DP vessels are typically designed with redundancy thrusters more than needed for motion control. Optimization theories are useful in finding thrust allocation solutions that minimize fuel consumption and reduce “wear and tear” on a thruster. But several challenges exist such as uncertain thruster model, thruster dynamics characters and the individual limitations of the thrusters. In this paper, a dynamic thrust allocation scheme is presented based on model predictive control (MPC) that directly takes thrusters with dynamics characters and various constraints into account. It is shown in simulations that the MPC dynamic thrust allocation scheme performs better in comparison with an existing static allocation method. Its main advantage is the ability to handle thruster dynamics characters and various constraints.

Formation of charge-exchange ion flows near the exit from the stationary plasma thruster acceleration channel

Ion currents onto the exit plane of the acceleration channel of a stationary plasma thruster model were measured using electrostatic probes the collecting surfaces of which could be oriented either upstream or downstream with respect to the thruster plume. Using the results of measurements, the so-called “back” flows of charge-exchange ions onto the exit plane are estimated. It is shown that the back ion flows are the most intense in the close vicinity of the thruster, but do not exceed 0.6% of the total ion flow from the thruster. The formation of steady-state ion flows near the exit from the acceleration channel of a stationary plasma thruster is simulated numerically by using a three-dimensional kinetic model that describes the dynamics of ions and neutral atoms exhausting from the acceleration channel and produced in the thruster plume and takes into account resonance charge exchange of ions with neutral atoms. The distribution of the back ion current density in the exit plane is determined. The effect of the flow rate of the working gas through the cathode on the distributions of the neutral atom density and charge-exchange ion flows is demonstrated. The obtained results can be used to analyze the effect of the thruster plume on the charge state of the surfaces located in the vicinity of the thruster.

Empirical modeling of rotating thruster for underwater robotic platform

A rotating thruster model is needed for the design and control of an underwater robotic platform. Thruster rotation is used to generate an additional degree of freedom for reducing the number of thrusters. The rotation of a thruster during thrusting generates an additional reaction force and torque. A rotating thruster model is required to analyze the dynamic model of a robotic platform by considering rigid body dynamics and hydrodynamics simultaneously. We considered the empirical dynamic modeling of a rotating thruster using frequency and impulse response analysis. A rotating thruster model is derived as a transfer function of longitudinal and tangential thrust forces and rotational torque during the rotation. Simulation is performed using the proposed transfer functions, and the validity of the model is verified by comparison with simulations and the results of experiments. The proposed empirical model can be applied to control an underwater robotic platform with rotating thrusters.

Electrostatic Ion Thrusters ‐ Towards Predictive Modeling

AbstractThe development of electrostatic ion thrusters so far has mainly been based on empirical and qualitative knowhow, and on evolutionary iteration steps. This resulted in considerable effort regarding prototype design, construction and testing and therefore in significant development and qualification costs and high time demands. For future developments it is anticipated to implement simulation tools which allow for quantitative prediction of ion thruster performance, long‐term behavior and space craft interaction prior to hardware design and construction. Based on integrated numerical models combining self‐consistent kinetic plasma models with plasmawall interaction modules a new quality in the description of electrostatic thrusters can be reached. These open the perspective for predictive modeling in this field. This paper reviews the application of a set of predictive numerical modeling tools on an ion thruster model of the HEMP‐T (High Efficiency Multi‐stage Plasma Thruster) type patented by Thales Electron Devices GmbH. (© 2014 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)

霍尔推力器热模型研究

To provide energy loss loading conditions for the thermal analysis of Hall thruster, a thermal model of Hall thruster is developed. Based on the plasma theory, the mechanism of various energy losses in discharge chamber is analyzed, and related functions are established between these energy losses and operation, performance parameters and structural parameters. At last the complete thermal model of Hall thruster is obtained. The thermal model for LHT100 thruster is calculated. The following results under the steady operation are obtained: the beam energy loss is about 889 W, the energy loss at wall is about 300 W, the anode energy loss is about 44 W, the ionization energy loss is 43 W, and the radiation energy loss is 34 W, respectively. Taking these energy losses as loading conditions to perform finite element thermal analysis and using thermal balance to validate experimentally, the results indicate that the calculated values agree well with the experimental ones, the maximum error is no more than 5%.

Quadrotor-tracking controller design using adaptive dynamic feedback-linearization method

Design of an adaptive dynamic feedback-linearization control law for a quadrotor unmanned aerial vehicle under uncertain parameters is presented. Because the quadrotor carries rotational speed-varying thrusters, it has the advantage of simple mechanism compared to the pitch-varying thrusters. However, it is subjected to slow dynamics in thruster and suffers from uncertainties in efficiency due to power subsystem. Additionally, parametric uncertainties tend to exist such as thruster misalignment, mass, and inertia. The control law is targeted to tracking reference trajectories under such uncertainties. Dynamic feedback-linearization method is employed primarily to produce the small-bandwidth thruster signal. A dynamic observer is used to estimate the states of feedback-linearized system, and Lyapunov-based update laws are derived to compensate for uncertain parameters. The controller and its performance are evaluated using a nonlinear, six-degree-of-freedom dynamic model of a quadrotor unmanned aerial vehicle with a thruster model in the simulation. The results illustrate that the proposed control law enhances tracking performance even with slow and misaligned thruster.

Dynamic characteristics of a resonant gas-dynamic system for ignition of a fuel mixture

Results of an experimental study of specific features of operation of a resonant gasdynamic ignition system, as applied to initiation of liquid-propellant thrusters, are reported. Dynamic characteristics of the initiation process are determined. The behavior of the spectral characteristics of pressure oscillations in the combustor of a testbench model of a liquid-propellant thruster with ignition of the fuel mixture by the gas-dynamic system is considered.

Inductive pulsed plasma thruster model with time-evolution of energy and state properties

A model for pulsed inductive plasma acceleration is presented that consists of a set of circuit equations coupled to both a one-dimensional (1D) equation of motion and an equation governing the partitioning of energy. The latter two equations are obtained for the plasma current sheet by treating it as a single element of finite volume and integrating the governing equations over that volume. The integrated terms are replaced where necessary by physically equivalent approximations that are calculated through the solution of other parts of the governing equation set. The model improves upon previous 1D performance models by permitting the time-evolution of the temperature consistent with the time-varying energy flux into the plasma. The plasma state properties are also more realistically modelled and evolved in time, allowing for the tailoring of the model to different gases that may be chosen as propellants. Computational results for argon propellant are presented to demonstrate the efficacy of the model. The model produces a result where efficiency is maximized at a given value of the electrodynamic scaling term known as the dynamic impedance parameter. The scaling of different energy sinks as a function of the dynamic impedance parameter provides insight into the global energy partitioning in these types of accelerators. Results from the present model deviate from the previous version where temperature is selected as an input without regard for the energy that would be deposited to heat the gas to that temperature. Qualitatively and quantitatively, the model predicts specific impulse values that compare favourably with those measured for two separate inductive pulsed plasma thrusters. Efficiency is underpredicted in the regime where data are available, but the trends in the data and simulations follow similar trajectories that appear to be converging towards a predicted peak efficiency as the dynamic impedance parameter is increased.

One-dimensional model of a pulsed plasma thruster

AbstractThis paper presents a modified one dimensional model for the pulsed plasma thruster. The model takes into account the impact of thruster geometry and capacitor initial energy on plasma inductance and resistance, as well as the ablated mass per pulse. These three factors have significant impact on the performance of the thruster and have not been considered concurrently before. The model is verified by comparing the estimated specific impulse, impulse bit and thrust-to-power ratio with experimental measurements for space qualified thrusters. The maximum error is estimated to be about 8% when using the modified model, while the conventional snowplow model produces error up to about 70%. The modified model is then employed to generate design charts that would help in selecting the proper combination of thruster geometry and energy level in the preliminary design phase of a pulsed plasma thruster for different space missions.

Toward an Accurate Physics-Based UAV Thruster Model

Small unmanned aerial vehicles (UAVs) come in many types, the most common being fixed-wing and rotorcraft. Most of these are powered by brushless dc motors driving fixed-pitch propellers. Since the thrusters are typically quite powerful, relative to the weight of the aircraft, the motion of these UAVs is usually dominated by the thruster dynamics. It therefore becomes particularly important to have a good model of the thruster, which can be assembled based on simple measurements of the system properties, rather than from exhaustive testing. This paper presents such a model. The governing equations are assembled by considering, in succession, the motor electrodynamics and the propeller aerodynamics. The results of the model are compared to experimental test results for a particular thruster assembly. Agreement between the two is excellent-with an error of 4.7% in thrust and 7.6% in torque under static conditions-thereby demonstrating the validity of the proposed approach.

Analysis of underwater thruster model with ambient flow velocity using CFD

Thrust force is a very important factor for underwater vehicles. The thrust force that is determined by the pressure gradient between a propeller and a thruster can be represented by the ambient flow velocity introduced as the control volume and the axial flow velocity of a propeller. Because a change in ambient flow velocity triggers a change in the pressure gradient between a propeller and a thruster, a model taking account of the ambient flow velocity is required for an unmanned underwater vehicle (UUV) system. However, the axial flow velocity introduced into a propeller is very difficult to measure without accurate test devices. Therefore, in this study, the axial flow velocity is calculated with the computational fluid dynamics (CFD) method to use it as a basis for estimating the approximate value of the thrust force. As a result, a relatively accurate analysis of the effect of the ambient flow velocity on the thrust force can be obtained with considerable time and cost effectiveness as compared to the existing experimental methods. To evaluate the validity of the data from the CFD analysis results depending on the change in ambient flow velocity and the pressure gradient of a thruster, the resulting CFD values were compared with the thrust forces obtained in the previously performed thrust force experiment of a thruster depending on the ambient flow velocity in a circulating water channel.

Krypton charge exchange cross sections for Hall effect thruster models

Following discharge from a Hall effect thruster, charge exchange occurs between ions and un-ionized propellant atoms. The low-energy cations produced can disturb operation of onboard instrumentation or the thruster itself. Charge-exchange cross sections for both singly and doubly charged propellant atoms are required to model these interactions. While xenon is the most common propellant currently used in Hall effect thrusters, other propellants are being considered, in particular, krypton. We present here guided-ion beam measurements and comparisons to semiclassical calculations for Kr+ + Kr and Kr2+ + Kr cross sections. The measurements of symmetric Kr+ + Kr charge exchange are in good agreement with both the calculations including spin-orbit effects and previous measurements. For the symmetric Kr2+ + Kr reaction, we present cross section measurements for center-of-mass energies between 1 eV and 300 eV, which spans energies not previously examined experimentally. These cross section measurements compare well with a simple one-electron transfer model. Finally, cross sections for the asymmetric Kr2+ + Kr → Kr+ + Kr+ reaction show an onset near 12 eV, reaching cross sections near constant value of 1.6 Å2 with an exception near 70–80 eV.

Thruster Modelling for Underwater Vehicle Using System Identification Method

Abstract This paper describes a study of thruster modelling for a remotely operated underwater vehicle (ROV) by system identification using Microbox 2000/2000C. Microbox 2000/2000C is an XPC target machine device to interface between an ROV thruster with the MATLAB 2009 software. In this project, a model of the thruster will be developed first so that the system identification toolbox in MATLAB can be used. This project also presents a comparison of mathematical and empirical modelling. The experiments were carried out by using a mini compressor as a dummy depth pressure applied to a pressure sensor. The thruster model will thrust and submerge until it reaches a set point and maintain the set point depth. The depth was based on pressure sensor measurement. A conventional proportional controller was used in this project and the results gathered justified its selection.

Nonlinear Modeling and Identification of Underwater Thruster

Complexity in modeling an underwater thruster analytically, leave system identification approached a good option on developing accurate dynamic model for underwater thruster. Nonlinearity of the system due to fluid solid interaction such as, hydrodynamic force hitting the propeller blade make the system identification became a trivial task. This paper presents nonlinear modeling and identification of the underwater thruster from input–output measurement. The system was fully submerged, and current was varied as the input and the output thrust was measured using load cell. Nonlinear Hammerstein method is chosen for identification of the system. The results are numerically and graphically presented. The nonlinear system identification with second order linear dynamics gave the best result, where is the model, can fit up to 82 % of the real response of the thrust. The finding of the model can be utilized in the future to improve underwater vehicle performance by developing optimum control algorithm.