Thruster Array Research Articles

Design and optimization of a high thrust density air-breathing pulsed plasma thruster array

For satellites in Very Low Earth Orbit (VLEO) or the upper atmosphere, Air-Breathing Electric Propulsion (ABEP) can be used to counteract the drag effects of the atmosphere or provide attitude control without the need for onboard propellant storage. The Dense Plasma Focus (DPF) inspired the design of an ABEP Pulsed Plasma Thruster (TAMU PPT) which takes advantage of the simplicity, scalability, and high energy density of the DPF. This work details the design and testing of a PPT for use on satellites in the upper atmosphere. Testing was conducted at a variety of pressures and pulsing energies to determine which conditions provide maximum thrust and thrust to power ratios (TPR). It was found through single pulse experiments that a pressure of 6.3 hPa (corresponding to an altitude of about 35 km) and an energy per pulse of 9 J yield a TPR of 21 mN/kW. High frequency operation at 20 Hz in similar conditions showed a TPR of up to 28 mN/kW with a corresponding thrust of 4.8 mN. High speed imaging of the plasma plume found plasma velocities in excess of 2 km/s and identified metal ejecta with velocities up to 350 m/s. The pinching phenomenon associated with a DPF was also observed and a plasma jet velocity up to 50 km/s was recorded. The TPR of the TAMU PPT demonstrated in this work is comparable to that of other ABEP thrusters, however the TAMU PPT has an advantage when comparing thrust density. The TAMU PPT has a thrust density between 4.8 and 19 mN/cm2 while other ABEP thrusters have thrust densities between 0.019 and 0.36 mN/cm2.

Active space debris removal by ion multi-beam shepherd spacecraft

Contactless transportation of a passive object by an ion beam generated by an active spacecraft's thruster is a promising way of space debris removal. The study proposes a modification of the ion beam shepherd concept using an array of impulse transfer thrusters. The multi-beam scheme provides significant increase the generated ion force, which reduces a space debris object deorbit time, thereby reducing the probability of its collision with other orbital objects. The aim of the work is to study the possibility of using several ion beams for effective contactless transportation of a space debris object. Controlling the angular motion of the object permits orientation of the object to maximize the generated ion force. The article proposes a control law for impulse transfer thrusters, which is based on the calculation of the energy of the object's unperturbed motion, and two control schemes that implement this law: (1) relay control, which implies turning the thruster on and off, and (2) ion beam axis direction control for one of the thrusters. For a space debris object of a cylindrical shape, a comparison of the time and fuel costs required for deorbiting the object using one, two, and three impulse transfer thrusters is made. It is shown that the addition of thrusters significantly reduces the descent time, but has little effect on the mass of required fuel. For the case of two engines, the best angle of ion beams axes direction and the most preferable angular motion mode are determined. Numerical simulation of space debris removal is carried out for the case when the space debris object is in the least favorable mode of angular motion and the proposed control schemes are used. The ion beam direction control scheme showed better results than the relay scheme. The use of multiple ion beams opens up new possibilities for creating new ion beam shepherd spacecraft design and developing new control methods and laws.

Electric thruster configuration design optimization for geostationary satellites with robotic manipulators

This paper proposes a novel electric thruster configuration design optimization scheme to implement efficient station-keeping operations for geostationary satellites. In this work, the electric thrusters are equipped with 4-DOF robotic manipulators to provide flexible control thrust in space, which consumes much less propellant than the conventional station-keeping process using fixed-position electric thrusters. To effectively determine the firing right ascensions and firing durations of each electric thruster during the station-keeping process, a novel station-keeping control strategy is developed in view of the low-thrust feature of electric thrusters and the flexible pointing capacity of robotic manipulators. A low-thrust station-keeping period includes one two-day orbit determination period and six two-day control periods, where the initial orbital elements are determined in the orbit determination period, and the electric thrusters fire to compensate the perturbations in the control periods. Based on the control strategy, the electric thruster configuration design optimization problem is formulated to minimize the total propellant consumption via optimizing the robotic manipulator joint swivel angles and firing durations simultaneously, subject to several practical engineering constraints including the station-keeping accuracy, angular momentum dumping, collision between electric thruster and solar array, and electric thruster plume. Finally, the proposed electric thruster configuration design optimization scheme is applied to real-world station-keeping problems. The simulation results show that the geostationary satellites can be successfully maintained in an assigned station-keeping window using the robotic manipulators based low-thrust station-keeping control strategy. Moreover, the optimized electric thruster configuration with robotic manipulators can save 31.58% propellant consumption compared with that of the conventional configuration owing to the favorable flexibility of control thrust, which demonstrates the effectiveness and practicality of this work.

Air-breathing electric propulsion: Flight envelope identification and development of control for long-term orbital stability

Air-breathing electric propulsion (ABEP) enables long duration missions at very low orbital altitudes through the use of drag compensation. A system-level spacecraft model is developed, using the interaction between thruster, intake and solar arrays, and coupled to a calculation of the drag. A quadratic solution is found for specific impulse and evaluated to identify the thruster performance required for drag-compensation at varying altitudes. An upper altitude limit around 190 km is based on a minimum thruster propellant density, resulting in required thruster performance values of Isp>3000s and T/P>8mN/kW for a realistic ABEP spacecraft. The orbit of an air-breathing spacecraft is propagated with time, which highlights the prescribed orbit eccentricity due to non-spherical gravity and therefore an increased variability in the atmospheric conditions. A thruster control law is introduced which avoids a divergent altitude behaviour by preventing thruster firings around the orbit periapsis, as well as adding robustness against atmospheric changes due to season and solar activity. Through the use of an initial frozen orbit, thruster control and an augmented T/P, a stable long-term profile is demonstrated based on the performance data of a gridded-ion thruster tested with atmospheric propellants. An initial mean semi-major axis altitude of 200 km relative to the equatorial Earth radius, a spacecraft mass of 200 kg, Isp=5455s and T/P=23mN/kW, results in an altitude range of around 10 km at altitudes of 160–183km during a period of medium to high solar activity.

Orbit and Attitude Control of a GEO Satellite in Station Keeping Flight Mode

In this study, the performance requirements influencing the orbital and attitude control system of a geostationary satellite in the station-keeping flight mode considering the coupling effect of both attitude and orbital motion is determined. Controlling and keeping the satellite in its orbital window have been done using a set of four thrusters located on one side of the satellite body, with considering the coupling effect of the attitude motion on orbital motion. The satellite’s orbital motion could be disturbed by the attitude motion in the allowable orbital window. The main factors conducting this behavior are derived utilizing the satellite attitude and orbital dynamic equations of motion. In the mathematical analysis of this study, the effects of environmental perturbations originating from the oblateness of Earth, third mass gravity like sun and moon, and solar radiation pressure on the satellite dynamic behavior are also considered. Afterward, the condition of using four installed thrusters on one side of the satellite and the reaction wheels in order to control the satellite orbital and attitude motion is investigated. To reduce the satellite attitude’s error, a proportional-derivative controller is employed to activate the reaction wheels properly. The satellite positions in north-south and east-west directions are controlled by a specific array of thrusters in order to maintain in its predefined orbital window. The required amount of velocity variations for a duration of one year via some simulation may demonstrate the effectiveness of the proposed approach in enhancing the orbital maintenance procedure of the satellite.

Electrospray Propulsion Engineering Toolkit (ESPET)

We report on the development of a software tool, the Electrospray Propulsion Engineering Toolkit (ESPET), that is currently being shared as a web application with the purpose to accelerate the development of electrospray thruster arrays for space propulsion. ESPET can be regarded as a database of microfluidic properties and electrohydrodynamic scaling models that are combined into a performance estimation tool. The multiscale model integrates experimental high-level physics characterization of microfluidic components in a full-scale electrospray propulsion (ESP) microfluidic network performance solution. ESPET takes an engineering model approach that breaks the ESP system down into multiple microfluidic components or domains that can be described by either analytical microfluidic or reduced order numerical solutions. ESPET can be divided into three parts: a central database of critical microfluidic properties, a microfluidic domain modeler, and a microfluidic network solver. Two options exist for the network solution, a detailed multi-domain solver and a QuickSolver designed for rapid design and testing of simple three-domain reservoir-feed-emitter arrays. The multi-domain network solver exploits the Hagen–Poiseuille/Ohm’s law analogy by using the publicly available SPICE (Simulation Program with Integrated Circuit Emphasis) electric circuit simulation software to solve the flow properties of the microfluidic network. Both the multi-domain and QuickSolver solutions offer Monte Carlo analysis of arrays based on user supplied tolerances on design parameters. Benchmarking demonstration examples are provided for experimental work in the literature, as well as recent experimental work conducted at Busek Co. The demonstration examples include ionic liquid propelled systems using active and passive capillary emitters, externally wetted emitter needles, and porous glass emitters, as well as a liquid metal system based on an externally wetted emitter needle.

Thermodynamics and performance of Al/CuO nanothermites with different storage time

The storage stability of energetic materials is important for their application. In our previous work, Al/CuO nanothermites was applied to a MEMS-based solid propellant micro thruster (SPM) array because of its short ignition delay and high reaction rate.In the current work, the storage stability of Al/CuO nanothermites, which were prepared by electrospray method and stored with different storage time, was systematically researched. The activation energy of Al/CuO nanothermites were measured by DSC. The ignition temperature and the curve pressure history of Al/CuO nanothermites were measured using T-jump and constant-volume pressurization tests, respectively. Further, the nanothermites were characterized by XRD, XPS, SEM and TEM.The results show that the morphology and mixture uniformity of the nanothermites after a long-time storage did not change significantly. The slow decomposition of NC and the thickening of oxide layer of Al NPs existed in the storage of Al/CuO nanothermites, analysed by XPS. The activation energy (Ea ) of Al/CuO nanothermites after storage for different time was determined using the method of Kissinger, as shown in Figure 1. Ea was decreased from 272.8 kJ/mol to 154 kJ/mol after storage for 13 months. When stored for 0, 7 and 13 months, the peak pressures of Al/CuO nanothermites were 685.8 kPa, 626.3 kPa and 625.5 kPa, respectively. In addition to the ignition temperature, it was 775 °C, 739 °C and 754 °C, respectively. This result indicated that the ignition and combustion properties of Al/CuO nanothermites are obviously reduced when stored for a long time, at room temperature.The full version of this extended abstract will appear in Defence Technology in 2020.

Control Allocation Strategy for Optimal Digital Microthruster Arrays in Orbit Control Applications

Digital microthruster arrays can be used to propel micro/nanosatellites on missions requiring high precision and low impulse. This paper presents a novel control allocation strategy for digital microthruster arrays, called segment distribution and stratification addressing (SD-SA) algorithm, which aims to find the optimal combination of microthrusters to successfully achieve a given thrusting maneuver. The SD algorithm separates a thruster array into disjoint segments, and then the SA algorithm combines these segments to provide required impulses in a control task. Compared with traditional algorithms, the SD-SA algorithm costs less time and computational storage space. It is demonstrated through simulation that the SD-SA algorithm is effective for the real-time orbit control of micro/nanosatellites.

Allocation Optimization Strategy for High-Precision Control of Picosatellites and Nanosatellites

The high-precision control of picosatellites and nanosatellites has always plagued the astronautics field. Aiming to change the status quo of the actuators not being able to meet the high-precision attitude control of picosatellites and nanosatellites, this article formulates a control allocation strategy for picosatellites and nanosatellites using the solid propellant microthruster array (SPMA). To solve the problem of the diversity and complexity of ignition combinations brought about by the high integration of the SPMA, the energy consumption factor of the optimal allocation is established, and the relationships of the array’s energy consumption factor, the control accuracy, the number, and the ignition combinations of the thruster array are deduced. The optimization objective is introduced by Sherman-Morrison formula and singular value decomposition. Thus, the energy consumption problem is transformed into an integer programming problem, acquiring the control allocation strategy and the optimal thruster energy. Simulation results show that the proposed algorithm can effectively reduce the thrust energy consumption and improve the precision control, demonstrating the feasibility and efficiency of the proposed algorithm for picosatellites and nanosatellites.

Optimal selection of appropriate fuels for a miniaturized motor and numerical simulations on its temperature distributions and thermal deformation

In this work, the energy characteristics of different propellants were thermodynamically calculated using the Minimum Gibbs free energy method. Based on such calculations and the previous experimental results using the DSC investigations on the thermal decomposition behavior, an optimized propellant formulation was selected. The numerical simulations on the temperature distributions for a prepared miniaturized motor with this formulation were conducted, and the resultant thermalstress and deformationof the thruster array were investigated. In particular, a physical micro model was established, and the effects of combustion parameters on the thermal performance were studied. The results showed thatthe highest temperature occurred at the boundary of the combustion chamber, where the thermal stress and deformation were the maximum. Furthermore, the thermal stress and deformation of silicon material was smaller than steel, and the thermal stress was the main reason for the destruction of micro-thruster stability. In addition, the maximum thermal deformation was small and had little effects on the thermalstability, whereas the thermal stress was the main reason for the poor thermal stability.

First Breakthrough for Future Air-Breathing Magneto-Plasma Propulsion Systems

A new breakthrough in jet propulsion technology since the invention of the jet engine is achieved. The first critical tests for future air-breathing magneto-plasma propulsion systems have been successfully completed. In this regard, it is also the first time that a pinching dense plasma focus discharge could be ignited at one atmosphere and driven in pulse mode using very fast, nanosecond electrostatic excitations to induce self-organized plasma channels for ignition of the propulsive main discharge. Depending on the capacitor voltage (200-600 V) the energy input at one atmosphere varies from 52-320 J/pulse corresponding to impulse bits from 1.2-8.0 mNs. Such a new pulsed plasma propulsion system driven with one thousand pulses per second would already have thrust-to-area ratios (50-150 kN/m²) of modern jet engines. An array of thrusters could enable future aircrafts and airships to start from ground and reach altitudes up to 50km and beyond. The needed high power could be provided by future compact plasma fusion reactors already in development by aerospace companies. The magneto-plasma compressor itself was originally developed by Russian scientists as plasma fusion device and was later miniaturized for supersonic flow control applications. So the first breakthrough is based on a spin-off plasma fusion technology.

Microfabricated Electrospray Thruster Array with High Hydraulic Resistance Channels

Electrospray thrusters can deliver the efficient primary propulsion and attitude control actuation needed by small satellites. The microfabricated multiemitter array is an ideal implementation of this technology because it addresses the thrust, volume, mass, and scalability requirements across the size range of small satellites. This paper describes the microfabrication and testing of an internally fed 64-emitter array. The main goal is to decouple the design and fabrication of the emitters from the need to provide a flowpath with sufficient hydraulic resistance for each emitter. The latter has traditionally been achieved by reducing the diameter of the emitter’s inner channel or by filling it with a bed of beads; whereas in this paper, individual flow-restrictive channels are etched on the back of the emitter wafer, connecting each emitter to a common distribution manifold. The source operates steadily in vacuum. Beam interception by the extractor is negligible in a wide range of flow rates and emitter potentials. The thrust, mass flow rate, specific impulse, and propulsive efficiency are calculated using the time-of-flight technique. These parameters range between 8.1 and , and , 96 and 236 s, and 61 and 75% using the ionic liquid 1-ethyl-3-methylimidazolium bis(trifluoromethylsulfonyl)imide as propellant.

Thermo-mechanical Design for On-orbit Verification of MEMS based Solid Propellant Thruster Array through STEP Cube Lab Mission

A MEMS solid propellant thruster array shall be operated within an allowable range of operating temperatures to avoid ignition failure by incomplete combustion due to a time delay in ignition. The structural safety of the MEMS thruster array under severe on-orbit thermal conditions can also be guaranteed by a suitable thermal control. In this study, we propose a thermal control strategy to perform on-orbit verification of a MEMS thruster module, which is expected to be the primary payload of the STEP Cube Lab mission. The strategy involves, the use of micro-igniters as heaters and temperature sensors for active thermal control because an additional heater cannot be implemented in the current design. In addition, we made efforts to reduce the launch loads transmitted to the MEMS thruster module at the system level structural design. The effectiveness of the proposed thermo-mechanical design strategy has been demonstrated by numerical analysis.

Verification of MEMS fabrication process for the application of MEMS solid propellant thruster arrays in space through launch and on-orbit environment tests

One of the most significant barriers encountered to the space application of MEMS technology is its lack of reliability and flight heritage in space environments. In this study a MEMS solid propellant thruster array was selected for the verification test of MEMS technology in space. The function and performance of MEMS solid thruster have been previously verified by laboratory-level research in universities. To ensure the successful operation of the MEMS thruster module before flight demonstration on-orbit, launch and on-orbit environment tests were performed at the qualification level. In the launch test, sine burst, and random vibration loads were applied to the MEMS thruster module. The thermal vacuum tests were carried out for the on-orbit environment test. As a result of the launch vibration test and on-orbit environment test, the variations of the characteristics were less than 0.7%, and all the functional requirements were successfully verified after the vibration tests. The tests successfully verified the manufacturing process because the thruster module showed stable normal function before the ignition. The test result outputs will be helpful in establishing MEMS fabrication guidelines for space applications.

Nanoengineered thrusters for the next giant leap in space exploration

Abstract

자세제어 추력기 배기가스에 의한 태양전지판의 동적 영향 분석

본 연구의 목적은 인공위성의 자세제어 추력기 배기가스가 태양전지판에 충돌하여 생기는 동적 영향인 교란력과 교란토크를 분석하는 것이다. 두 개의 매개변수인 태양전지판과 추력기 사이의 거리와 추력기의 경사각을 사용하는 Sweep Analysis를 실시하여 교란력과 교란토크 원(原)데이터베이스를 만들었다. 이 데이터베이스를 기반으로 이 두 개의 매개변수에 의해 달라지는 교란력과 교란토크의 특성을 기술하는 3차 다항 근사식을 만들고, 각 태양전지판의 각(角)위치에 대해 근사식의 계수들을 최종 결과로서 얻는다. 이 결과들은 자세제어계의 입력 자료로서 사용되어 추력기 배치 최적화에 사용된다. 정지궤도복합위성의 후보 태양전지판 두 종류에 이 분석을 적용하여, 교란력과 교란토크 결과를 비교 분석하였다. The purpose of this study is to analyse the dynamic disturbances(disturbed forces and disturbed torques) due to attitude control thruster's plume impingement on the solar arrays. To produce database of the dynamic disturbances a sweep analysis was done, in which the two parameters are used; the distance between the thruster and solar arrays and the thruster tilt angle. Based on the database, a third order polynomial approximation is computed to represent the characteristics of the disturbed forces and torques. The final results are the coefficients of the approximation for each solar array angle position. These results as input data are used to optimize the configuration of the attitude control thrusters. This analysis is appled to the two candidate solar arrays for Geo-Kompsat-2 satellite and the results of the disturbed forces and disturbed torques are compared and analysed.

Silicon‐on‐insulator‐based microsolid propellant thruster array

A silicon-on-insulator (SOI)-based microelectromechanical system solid propellant thruster array is presented. The micronozzle layer and microigniter layer of the array are formed on both surfaces of the SOI wafer, respectively. The layout design rules can properly distribute the heat produced by the microigniter below the solid propellant inside the chamber. Test results show that the highest temperature of the microigniter is located under the propellant. When one unit in the array is ignited by a DC voltage of 7.9 V, other units around it are not ignited. Moreover, the ignition resistors can work again, which means that the array can be used repeatedly when the solid propellant is refilled.

Experimental Investigation of a Direct-Drive Hall Thruster and Solar Array System

Studies of high-power solar electric propulsion systems (i.e., tens to hundreds of kilowatts) suggest that significant mass savings may be realized by implementing direct-drive power systems. The National Direct-Drive Testbed was established to address issues associated with implementation of direct drive, and experimental results at power levels up to 10 kW are reported here. Hall thruster operation and control were shown to be simple and no different from that for operation on conventional power supplies. Thruster oscillations were the same as those for conventional power supplies, did not adversely affect solar array operation, and were independent of filter capacitance from 8 to . Solar array current and voltage oscillations were very small compared to their mean values and showed a modest dependence on capacitor size. Significantly, no instabilities or anomalous behavior were observed in the thruster or power system at any operating condition investigated, including near and at the array peak power point. Thruster startup using the anode propellant flow as the power “switch” was shown to be simple and reliable, with system transients mitigated by the proper selection of filter capacitance size. A simple electrical circuit model was developed and is shown to have good agreement with the experimental data.

Monolithic 2 DOF fully compliant space pointing mechanism

Abstract. This paper describes the conception, modeling, and development of a fully compliant two-degree-of-freedom pointing mechanism for application in spacecraft thruster, antenna, or solar array systems. The design objectives and the advantages of a compliant solution are briefly discussed. Detailed design decisions to meet project objectives are described. Analytical and numerical models are developed and subsequently verified by prototype testing and measurements in several iterations. A final design of the 3-D printed titanium monolithic pointing mechanism is described in detail and its performance is measured.

MEMS solid propellant thruster array with micro membrane igniter

A microigniter with improved membrane for MEMS solid propellant thruster array is proposed. Although existing igniters using dielectric SiNx membranes consume low power, they cannot withstand shocks during the propellant charging process. A glass wafer is selected as the material for the igniter. A glass membrane is fabricated by the anisotropic wet etching of photosensitive glass. Platinum is used for the ignition coil. A 30–40-μm thick glass membrane exhibited sufficient strength to withstand accidental impacts during the propellant charging process. Further, the microigniter using the glass membrane had an appropriate power consumption to ignite the solid propellant. The thermal, electrical, and mechanical characteristics of the fabricated microigniter were measured. The proposed microigniter provided the sufficient heat for the propellant to be ignited with the given electric power. The fracture pressure of the glass membrane was thrice higher than the conventional microigniter. A solid propellant lead styphnate was filled into the microigniter without any additional processes. Ignition tests with fully assembly MEMS thrusters were performed successfully. Thus, it is demonstrated that a simpler, robust, and low-cost fabrication process of an igniter that performs well when assembled into a micro solid propellant thruster (MSPT) array is possible.