Solid Thruster Research Articles

Design of ignition control system for failed spacecraft array solid thruster

The life of micro-nano spacecraft is generally short. In order to solve the problem of space debris endangering space safety after its failure and make it deorbit quickly, this paper has carried out the hardware system design and software system design of the ignition control system of the failed spacecraft attitude adjustment solid thruster. In order to facilitate the realization of high integration, the hardware system adopts the array arrangement circuit and selects the MOSFET as the driving switch. The software system receives the ignition command from the spacecraft platform and enables the row and column control terminals of the ignition circuit to achieve addressing ignition. The results show that the control module can correctly receive the instructions from the spacecraft platform and make the correct ignition action, successfully complete the attitude adjustment task, and improve the reliability of the deorbiting of the failed spacecraft.

Numerical analysis of separation dynamics for space debris removal payload based on solid thruster

Aiming at the current mainstream flying anchor payload for space debris removal, a solid thruster was proposed as the separation driving force, and a complete dynamic model of the spacecraft platform and payload were established. The effects of the gap between the directional button and the directional tube, thrust eccentricity and the initial attitude of the platform on the target accuracy under the flying anchor payload were studied. The analysis results showed that the initial attitude of the platform was the main factor affecting the target accuracy of the flying anchor payload. The control gap between 0.1mm ~ 0.25mm can effectively reduce the displacement of payload separation and the influence of disturbance on the platform.

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.

Piece-wise affine MPC-based attitude control for a CubeSat during orbital manoeuvres

The orbital manoeuvre of a CubeSat is an essential ability in the process of constellation deployment and other missions. Therefore, an attitude control method based on piece-wise affine (PWA) MPC was proposed for CubeSats during orbital manoeuvres. Then, a solid thruster with high thrust that was previously applied only to large and traditional satellites was innovatively applied to CubeSats for orbital manoeuvres. However, the solid thruster with high thrust has a large eccentric torque from the thrust deviation, so it will cause rapid attitude rolling and result in orbital manoeuvre mission failure. Hence, PWA-MPC strategy with the disturbance observer was designed, which can solve the eccentric torque problem and achieve high accuracy attitude control by a hybrid actuator composed of momentum wheels and a cold gas thruster. In addition, during the orbital manoeuvre process, propellant consumption results in a continuous change in the moment of inertia of a CubeSat. In order to offset this change, the guide rail of the cold gas thruster was innovatively designed to move the cold gas thruster in the working process of the solid thruster. Finally, real-time hardware-in-loop simulation results show that the PWA-MPC strategy has higher attitude stability than those of other MPC strategies in the orbital manoeuvre process of a CubeSat and deserves to be applied.

Integral design method for simple and small Mars lander system using membrane aeroshell

To execute Mars surface exploration missions, spacecraft need to overcome the difficulties of the Mars entry, descent, and landing (EDL) sequences. Previous landing missions overcame these challenges with complicated systems that could only be executed by organizations with mature technology and abundant financial resources. In this paper, we propose a novel integral design methodology for a small, simple Mars lander that is achievable even by organizations with limited technology and resources such as universities or emerging countries. We aim to design a lander (including its interplanetary cruise stage) whose size and mass are under 1 m3 and 150 kg, respectively. We adopted only two components for Mars EDL process: a “membrane aeroshell” for the Mars atmospheric entry and descent sequence and one additional mechanism for the landing sequence. The landing mechanism was selected from the following three candidates: (1) solid thrusters, (2) aluminum foam, and (3) a vented airbag. We present a reasonable design process, visualize dependencies among parameters, summarize sizing methods for each component, and propose the way to integrate these components into one system. To demonstrate the effectiveness, we applied this methodology to the actual Mars EDL mission led by the National Institute of Information and Communications Technology (NICT) and the University of Tokyo. As a result, an 80 kg class Mars lander with a 1.75 m radius membrane aeroshell and a vented airbag was designed, and the maximum landing shock that the lander will receive was 115 G.

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.

Effect of wall heat transfer characteristic on the micro solid thruster based on the AP/HTPB aerospace propellant

The combustion processes of Ammonium perchlorate/Hydroxyl-terminated polybutadiene (AP/HTPB) aerospace propellant are analytically studied to determine the effect of wall materials and thickness on the heat transfer characteristics within a solid micro thruster. In this study, the flow and combustion characteristic of the propellant are described in terms of a combustion rate equation associated with a chemical reaction kinetics model. The profiles describing changes in the combustion surface by the user-defined function (UDF) are used to obtain the distribution of temperature and content of propellant using a numerical method. The combustion time and thrust are simulated to validate the results by comparing them with the results from the literature. Moreover, we demonstrate that a smaller thermal conductivity and a larger heat capacity of the wall material can lead to improvement in the thruster performance. Finally, analyzing the significant effects of wall thickness the heat transfer characteristics within the thruster allowed us to determine an optimal wall thickness value.

큐브위성 탑재를 위한 MEMS 고체 추력기의 구조설계 및 검증

MEMS solid thruster module is composed of solid thruster and its control board. It was developed for the purpose of an academic research. Therefore, thermo-mechanical design and verification for space usage were not considered in the design phase. To mount it on a cube satellite without any design modification, technical efforts at the system level structure design is required. In this study, we proposed a structural design concept to mount the MEMS thruster module by using brackets for guaranteeing structure safety under launch loads and easier mating and de-mating of MEMS thruster module during test phase. The effectiveness of the design has been verified through structural analysis and vibration test. In addition, electrical connection method using spring pins between MEMS thruster and control board is effective for guaranteeing the structural safety under launch vibration loads.

Design, fabrication and thrust measurement of a micro liquid monopropellant thruster

A liquid monopropellant MEMS thruster was designed, fabricated and tested. For application on a nanosatellite for orbit control, a liquid propellant MEMS thruster delivers better performance than a solid thruster. Two issues must be addressed for a liquid monopropellant MEMS thruster: high energy content of the monopropellant to overcome the excessive heat loss associated with the small scale of the thruster, and repeatability of generated thrust force. The present study proposed blending 90 wt% hydrogen peroxide with ethanol at an oxidizer to fuel ratio of 30 to enhance the energy content of the propellant. The thruster structure was constructed using glass layers that were individually patterned by wet etching. The decomposition catalyst was separately prepared by wet impregnation of the active material, Pt, on the gamma alumina pellets and inserted into the thrust chamber before the UV bonding process of the glass layers. The firing test of the assembled MEMS thruster was conducted and thrust was measured both with ethanol blended hydrogen peroxide and pure hydrogen peroxide as a reference monopropellant. The measured thrusts were approximately 30 mN for both 1.7 ml min−1 flow rate of blended hydrogen peroxide and 2.0 ml min−1 flow rate of pure hydrogen peroxide. The measured thrust for 1.7 ml min−1 pure hydrogen peroxide was approximately 24 mN. The measured thrust was 40% less than the design thrust for both monopropellants. The uncertainty of the thrust was less with blended monopropellant than with pure hydrogen peroxide.

An experimental investigation of superheated subliming solid thruster performance

An experimental investigation of superheated subliming solid thruster performance