Solid Propellant Microthruster Research Articles

Lab-on-PCB solid propellant microthruster with multi-mode thrust capabilities.

In the realm of nano/microsatellite clustering, the demand for microthrusters is steadily growing. Solid propellant microthrusters, recognized for their lightweight build and structural simplicity, carry significant commercial promise. However, existing solid propellant microthrusters manufactured using MEMS technology encounter notable issues such as inconsistent thrust generation positions, limited thrust profiles, and issues related to productivity, scalability, and durability. In this study, we propose a novel shared-chamber solid-propellant microthruster design that consistently produces thrust at a designated position and accommodates multiple thrust modes. The components and fabrication of this thruster were developed using lab-on-printed-circuit-board (PCB) technology and PCB surface mount technology, showcasing enhanced structural stability, scalability, and potential for mass production. Our ignition and combustion experiments confirmed the repeatability of the unit operation, a fundamental feature of this innovative microthruster. Furthermore, we successfully implemented and evaluated the power mode for increased thrust and the continuous mode for prolonged operational duration. Integrating the lab-on-PCB-based shared-chamber solid propellant microthruster with propulsion and electronic control systems holds promising potential for future satellite missions.

A review on solid propellant micro-thruster array based on MEMS technology

With the development of micro-spacecraft technology, micro-nano satellites have the advantages of small size, low power consumption, short development cycle, formation networking, etc., and can complete many complex space tasks at a lower cost. Micro-nano satellites require a micropropulsion system with the capability of performing precise total impulse and thrust to execute maneuvers, such as attitude control, orbital transfer, and gravitation compensation. In contrast to other micropropulsion systems, solid propellant microthrusters (SPM) arrays based on micro-electromechanical system (MEMS) technology possess a simple structure and quick response, which is a potential micropropulsion system. In recent years, many research groups have done a lot of research on SPM arrays. In this paper, the latest progress of SPM arrays is summarized from the aspects of structure design, propellant selection, bonding technology, ignition unit type and micro-thrust test, and some suggestions for the future development direction are given.

A STUDY ON THE RECENT ADVANCES IN THE DESIGN OF MEMS-BASED SOLID MICROTHRUSTERS

In the past few decades, space technology has moved toward using many small satellites working in a cluster instead of making individual satellites for space missions. Nano or microsatellites increase the reliability of the mission while minimizing cost, rather than having one large satellite increasing both the cost and the failure risk for any given space mission. Considering the case of a small satellite cluster, it is easy to manage mission requirements with less complexity during any hazard, compared to an individual satellite. All of this has led to the increasing application of micro and nanosatellites in space engineering. These small satellites are propelled by micropropulsion created by microthrusters. Currently, liquid, gas, and electrical thrusters are the most common. This paper mainly describes the recent advancements in the field of MEMS microthrusters. MEMS (microelectromechanical system) thrusters are used in small satellites in a variety of ways that require very little thrust. In the paper discussion, MEMS-based solid-propellant microthrusters (SPMs) are delineated. All advancements in the MEMS solid thrusters are indicated in the article, dedicated to mounting schemes, designing approaches, and various performance testing experiments. The main challenge in this technology is combustion, which takes place in an exceptionally small volume.

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.

Active Vibration Control of Flexible Satellites Using Solid Propellant Microthruster Array

AbstractThe paper studies a method which uses a solid propellant microthruster array (SPMA) as an actuator to control vibration of solar panels and the attitude of a satellite simultaneously. To so...

Performance evaluation and thermography of solid-propellant microthrusters with laser-based throttling

Among propulsion devices, conventional solid-propellant thrusters require neither a tank nor a valve and therefore have relatively high reliability and simple structures. However, the interruption and restarting of thrust production in solid-propellant thrusters is difficult because combustion is autonomously sustained after ignition. Hence, we propose a new solid-propellant microthruster that can be throttled through laser heating. The proposed thruster uses a combustion-controllable hydroxyl-terminated poly butadiene/ammonium perchlorate-based solid propellant, wherein combustion is maintained only while the burning surface is heated using a semiconductor laser. In this study, thrust was measured at various laser-head traverse velocities to enhance the performance of the prototype. Moreover, propellant holders were fabricated from polycarbonate (PC) and polymethyl methacrylate (PMMA) because the melting point of PC is higher than that of PMMA, which could increase the local pressure in the interface between the propellant holder and the solid propellant and consequently reduce ignition delay. The prototype microthruster with the PMMA propellant holder yielded a thrust of 0.06 N and Isp efficiency of 70% at a laser power density of 0.83 W/mm2 and produced stable thrust at laser-head traverse velocities ranging from 0.85 to 0.95 mm/s, whereas that with the PC propellant holder exhibited unstable thrust in all tests. Firing tests showed an ignition delay of approximately 3 s and yielded a peak in Pc after ignition. Temperature profiles measured through thermography revealed that the burning-surface temperature was dependent on the laser-head traverse velocity and propellant-holder material. For stable combustion, the burning-surface temperature would be kept constant at 1040 K, an ignition temperature was 1250 K in vacuum, and one-half the ignition delay (1.5 s) is the time required for heating the burning surface to the ignition temperature.

Superior performance of a MEMS-based solid propellant microthruster (SPM) array with nanothermites

We propose top ignition MEMS-based solid propellant microthruster (SPM) array to satisfy the precise propulsion requirements of small spacecraft, especially micro/nanosatellites. A 10 × 10 SPM array prototype was fabricated and assembled by standard microfabrication processes. Nano-Al/CuO was chosen as the solid propellant because of its high combustion heat and safety. To improve the propulsion performance of the microthruster, nanothermites mesoparticles were assembled by electrospray mixing (ES) process. The Chemical Equilibrium with Application calculation result indicates that the addition of nitrocellulose (NC) acting as binder into the nanothermites would enhance the thrust performance. The test results verified this. An impulse testing stand was employed to measure the impulse generated by the assembled SPM array, and a high-speed camera was used to record the combustion process. The specific impulse (I sp) and total impulse (I) of the binder-free nano-Al/CuO were 10.2 s and 155.9 μN s, respectively. The propulsion values of the ES nanothermites increased monotonically with increasing NC content, and the max I sp and I were 27.2 s and 346.9 μN·s, respectively. A 3D X-ray microscope was used to analyze the fired SPM array. The results show that the discontinuous bonding layer led to undesired ignition of the neighboring units.

Design and optimization of micro-semiconductor bridge used for solid propellant microthrusters array

Igniter is the basic component of MEMS-based solid propellant microthrusters (SPM) array, which should response fast with low dissipated energy. To satisfy the requirements, micro-semiconductor bridge (MSCB) with dual V-angles contributing to the reduction of dissipated energy is introduced into the array. The electrical explosion characteristics of MSCB investigated under constant voltage, in the considerations of the limit of power supply in the micro/nanosatellite, are similar to those under capacitive discharge. The bridge was completely vaporized and burst into the hot plasma under high-level voltage. While under low-level voltage, the bridge was partly vaporized without detected plasma. The 3#SCB (90° V-angle) with the smallest power density resulting in the lowest applied voltage (12.0 V) and the 2#SCB with the smallest narrow width ( W min ) leading to the lowest average power (11.3 W) were chosen to the further optimization. The ignition tests were implemented to examine the ignition reliability of the contractible igniters, which were optimized to satisfy the voltage limitation with sufficient ignition capacity. The results show that the effective SCB with lowest dissipated power is the 3#-3SCB, which function within 100 μ under 7.0 V, and the average power below 5.0 W.

A MEMS-based solid propellant microthruster array for space and military applications

Since combustion is an easy way to achieve large quantities of energy from a small volume, we developed a MEMS based solid propellant microthruster array for small spacecraft and micro-air-vehicle applications. A thruster is composed of a fuel chamber layer, a top-side igniter with a micromachined nozzle in the same silicon layer. Layers are assembled by adhesive bonding to give final MEMS array. The thrust force is generated by the combustion of propellant stored in a few millimeter cube chamber. The micro-igniter is a polysilicon resistor deposited on a low stress SiO2/SiNx thin membrane to ensure a good heat transfer to the propellant and thus a low electric power consumption. A large range of thrust force is obtained simply by varying chamber and nozzle geometry parameters in one step of Deep Reactive Ion Etching (DRIE). Experimental tests of ignition and combustion employing home made (DB+x%BP) propellant composed of a DoubleBase and Black-Powder. A temperature of 250 °C, enough to propellant initiation, is reached for 40 mW of electric power. A combustion rate of about 3.4 mm/s is measured for DB+20%BP propellant and thrust ranges between 0.1 and 3,5 mN are obtained for BP ratio between 10% and 30% using a microthruster of 100 μm of throat wide.

Design, fabrication and test of a solid propellant microthruster array by conventional precision machining

Prototypes of 10×10 and 100×100 scale solid propellant microthruster(SPM) array were fabricated with conventional precision machining, due to restriction on the manufacture of large-scale microthruster array with present MEMS manufacturing technique. The microthruster consists of ignition circuit layer, ignition powder layer, propellant layer, cavity layer and nozzle layer. Row–column driving control plan was applied in ignition circuit to realize simultaneous ignition and reduce the number of wires. The solid propellant was AP-HTPB pellet propellant with a diameter between 100μm and 200μm. A test to choose proper ignition powder was performed among black powder, modified back powder and self-made ignition powder, which demonstrated the reliability of self-made ignition powder and realize 100% ignition success rate. In the mean time, a measurement of ignition resistor’s surface temperature was taken to determine ignition delay, together with numerical simulation and ignition resistor fusing test. Thrust was also measured for the 10×10 scale microthruster array, and the results show that the specific impulse is much lower than calculated result, so a more comprehensive numerical model is in need to depict the working process of a single SPM. The test also demonstrates the feasibility of large-scale solid propellant microthruster array's fabrication by conventional fabrication methods.

Low Temperature Eutectic Alloy Bonding Used in MEMS-Based Solid Propellant Microthrusters Array

Considering the safety of the charge filled in the chambers and the operational reliability, a low temperature eutectic alloy bonding was preferred for the final assembling process of MEMS-based solid propellant microthrusters (SPM) array with top-side igniters. The optimum conditions of the alloy coating are the pH value of the solution is 0.5, the base metal layer is composed of Ti (60nm)/Cu (500nm) deposited by magnetron sputtering. The bonding process was conducted in an oven with air and the bonding temperature is 70°C. To predict the performance of bonding layer, the temperature distributions of eutectic alloy layer were simulated by ANSYS software when the chamber is full of high temperature reactants. The simulation results demonstrated the bonding strength wouldn’t cut down since the heat transfer induced by the combustion of charge. The assembled MEMS-SPM array was tested under constant voltage, the results indicated that the eutectic solder bonding procedure meet the requirements of the MEMS-SPM assembling.

Design and Fabrication of MEMS-Based Solid Propellant Microthrusters Array

A solid propellant microthruster array consists of three parts (five layers) was designed to form propulsion system of nanosatellite. The layers of the array were fabricated by different MEMS technique respectively. Then the array was assembled by adhesive bonding (H70E thermal epoxy). An initial test was operated with the ignition test stand composed of high speed video system, electronic power source, and oscilloscope. A constant voltage was applied to the igniters through two probes contact with the pad of ignition circuitry. The test indicates that the microthruster array can be ignited under 15V.

Modeling the Burning Rate of Heterogeneous Propellants over a Wide Range of Pressures

The variation of steady burning rate for heterogeneous propellants over a wide range of pressure is studied by experimental methods and theoretical analysis. Two kinds of heterogeneous propellants, GATo-1 CMDB propellant and 86 series composite propellant, are selected. An explanation for two inflection points observed in burning rate–pressure curve is presented. It is concluded that condensed-phase reaction dominates the burning process in the low-pressure region, which presents a low-pressure exponent, whereas heat feedback from the gas phase predominates in the high-pressure region with a high-pressure exponent. In the medium-pressure region, these two factors influence the burning process cooperatively and pressure exponent reduces first and then rises. Results calculated from the model show good agreement with experimental data. Thermal parameters analysis indicate burning rate is more sensitive to thermal conductivity of condensed phase, activation energy of condensed phase, specific heat capacity, and delay time of heat release of the condensed-phase reaction in high pressures. This model is of significance in burning rate characteristics in high pressures and interior ballistic properties in solid-propellant impulsive microthruster.

Effect of Unsteadiness and Nozzle Asymmetry on Thrust of a Microthruster

Due to the small size and short propulsion pulses of solid propellant microthrusters, unsteady characteristics and/or fabrication errors may cause appreciable effects on the total thrust. The present work simulates a 2D planar and conical nozzle of 200 μm throat diameter using the traditional momentum equations and turbulence closure model. Unsteady characteristics were found to be insensitive to the inlet mass flux but sensitive to the chamber size. The characteristic time of transient momentum variations in the nozzle was found to be on the order of 0.1 ms, and that for decay was smaller than that for development by about 25%. This work confirmed that the magnitude of thrust was not sensitive to distortions in the nozzle geometry. A mismatch of the nozzle throat changed the position of sonic plane, but its effect on thrust was almost negligible.

A novel micro-fabricated thruster for drug release in remote controlled capsule

A novel solid propellant microthruster has been developed and used as the actuation assembly of a patented remote controlled capsule (RCC). It produces sufficient gas pressure to empty the drug reservoir within the RCC. A MEMS-based microigniter works as the critical component of the miniaturized thruster that houses the detonating agent – diazodintrophenol (DDNP) and the propellant – black powder. Ignition and combustion tests were performed to justify its feasibility and reliability. Results demonstrated that 166 mW power consumption led to one successful combustion and that complete and rapid drug release was achieved when the propellant ranged from 16 mg to 20 mg. The RCC integrated with the microthruster could provide a promising alternative for site-specific drug delivery in human GI tract.

A wireless addressing interface circuitry for microthruster array applications

PurposeThe paper aims to develop a wireless addressing interface circuitry for solid propellant microthruster array applications.Design/methodology/approachThe solid propellant microthruster is a relatively new class of micropropulsion system for microspacecraft. To produce a controlled vectored thrust, a microthruster array is needed. Realization of the addressing ability and wireless communication is the key to the development of the microthruster array. Therefore, a prototype wireless addressing circuitry was developed to realize the addressing of the microthruster array by a multiplexing system. The addressing circuitry also enables measurement of the igniter temperature variation with time by measuring the igniter resistance change and automatic control by RS232 and RF wireless communications. Operating principles, design, fabrication, and testing of the circuitry are addressed.FindingsA prototype integrated wireless addressing circuitry was designed and fabricated to realize the addressing of individual microthrusters in the microthruster array, to measure the igniter temperature variation with time, and to achieve automatic control using RS232 and RF wireless communications. Using the programmable voltage source in the circuitry, the igniter temperature could be accurately controlled in 256 steps using an 8 bit word. The 10 bit analog‐to‐digital converter feedback loop circuitry enabled real‐time monitoring of each igniter in the microthruster array and allowing each igniter to be functionally controlled.Originality/valueIn this paper, a wireless addressing interface circuitry is developed for the first time for solid propellant microthruster array applications.

Investigation on the ignition of a MEMS solid propellant microthruster before propellant combustion

A MEMS-based solid propellant microthruster with thin film Au/Ti igniter has been fabricated for micropropulsion applications in microspacecraft. Since the propellant ignition process is crucial to the development of solid propellant microthrusters, finite-element based electro-thermal modeling is performed to predict and optimize the transient propellant ignition process. The model is implemented to determine the evolution of the propellant temperature with time, space and voltage and the ignition power, ignition delay and ignition energy. The difference in property between thin film Au/Ti and bulk Au/Ti and heat transfer in the micro-scale thruster are highlighted. A custom-built interface circuitry is employed to measure the igniter temperature experimentally. The electro-thermal model can be employed to find the optimum ignition subsystem for the solid propellant microthruster.

Solid Propellant Microthrusters on Silicon: Design, Modeling, Fabrication, and Testing

The design, modeling, fabrication, and characterization of solid propellant microthrusters for space application is presented. The operational concept of solid propellant thruster is simply based on the combustion of a solid energetic material stored in a micromachined chamber. Each thruster contains four main parts (nozzle, heater, chamber, seal). Thrusters presented in this paper have a chamber area of 2.25 mm2. Throat diameters (160, 250, and 500 mum) have been calculated to generate thrusts in the range from 0.3 to 30 mN that meets station keeping requirements for micro satellites. Fabrication processes for each part of the thruster are presented as well as the assembling procedure and the electronic module implementation. Characterizations give ignition power between 80 and 150 mW depending on the energetic material contained in the igniter's cavity. The ignition success is of 100% when using a Zirconium Potassium Perchlorate propellant. Thrust balance is presented as well as the thrust measurements that validate our model

Reliability of freestanding polysilicon microheaters to be used as igniters in solid propellant microthrusters

This paper presents the design, fabrication and characterisation of surface micromachined polysilicon microheaters to be used as microigniters for micropropulsion applications. The microigniters are heated up by Joule effect and the thermal losses through the substrate are minimised by suspending the microheaters above the substrate. The developed process was compatible with the integration of the nozzle part of the microthruster. The electrical, thermal and mechanical characteristics of the microheaters were studied with the aim of evaluating their reliability. Temperatures up to 470°C could be reached with an electrical power of 45mW/beam. The current–voltage relation followed a linear characteristic at low power; at high bias voltages, a drift of the electrical resistance was measured after a few I–V cycles at power higher than 40mW/beam. The elastic and plastic deformation threshold of the microheaters in operation and their maximum deflection before rupture were measured. The microheaters could dissipate relatively high constant powers for a few minutes to hours. The fabricated microheaters are promising candidates for the ignition of solid propellant microthrusters.

Performance Prediction of a Novel Solid-Propellant Microthruster

A novel microelectromechanical-systems-based solid-propellant microthruster is developed for stationkeeping, attitude control, drag compensation, and orbit adjust of microspacecraft. Key to the development of the solid-propellant microthruster is the generation of extremely accurate thrust levels and impulse bits. Modeling and simulation can predict the propellant combustion and gas expansion processes inside the microthruster and then derive the microthruster performance. A computational-fluid-dynamics-based model is proposed in this paper to optimize the design by computing the subsonic/supersonic micronozzle flow and calculating the microthruster performance. Wall heat loss and boundary-layer effects, which are especially important for microscale thrusters, are highlighted in the model. Different solid propellants, microthruster geometries, and operating conditions are also evaluated. The computational-fluid-dynamics modeling results are compared with the one-dimensional thermodynamic modeling results and experimental testing data.