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Design, development, and performance of an ammonia self-managed vaporization propulsion system for micro-nano satellites
Abstract
An ammonia self-managed vaporization propulsion (ASVP) system for micro-nano satellites is presented. Compared with a normal cold gas or liquefied gas propulsion system, a multiplex parallel sieve type vaporizer and related vaporization control methods are put forward to achieve self-managed vaporization of liquefied propellant. The problems of high vaporization latent heat and incomplete vaporization of liquefied ammonia are solved, so that the ASVP system takes great advantage of high theoretical specific impulse and high propellant storage density. Furthermore, the ASVP operation procedure and its physical chemistry theories and mathematical models are thoroughly analyzed. An optimal strategy of thrust control is proposed with consideration of thrust performance and energy efficiency. The ground tests indicate that the ASVP system weighs 1.8 kg (with 0.34-kg liquefied ammonia propellant) and reaches a specific impulse of more than 100 s, while the power consumption is less than 10 W. The ASVP system meets multiple requirements including high specific impulse, low power consumption, easy fabrication, and uniform adjustable thrust output, and thus is suitable for micro-nano satellites.
Relative motion control with six degrees of freedom (6-DOF) for spacecraft proximity operations in prescribed performance control (PPC) framework has become a hot issue in recent years, but actuator failure is seldom involved in controller design. In this paper, we introduce a complete thruster model to describe the actuator characteristics for thruster-only spacecraft, considering efficiency loss, thrust fluctuation and saturation. Besides, the barrier Lyapunov function (BLF) method and homeomorphic mapping method are often used in the PPC framework to constrain transformed errors. However, the two methods have singularity and infinite control effort problem once the constraints are not satisfied because of the actuator failure or other disturbance. In this paper, a novel bounded BLF (BBLF) is proposed to solve this problem. The proposed BBLF can still maintain bounded control effort and guarantee the system stability even if the transformed errors exceed the boundary. Further, the model uncertainty, actuator output uncertainty and external disturbance are summarized as lumped disturbances. A finite-time extended state observer (FTESO) is constructed to estimate the lumped disturbances. Finally, based on the estimated information from FTESO, an adaptive backstepping controller is proposed to track the desired trajectory. Numerical simulation results show the excellent dynamic response and steady-state accuracy of the proposed control strategy.
Rapid evolution of miniaturized, automatic, robotized, function-centered devices has redefined space technology, bringing closer the realization of most ambitious interplanetary missions and intense near-Earth space exploration. Small unmanned satellites and probes are now being launched in hundreds at a time, resurrecting a dream of satellite constellations, i.e., wide, all-covering networks of small satellites capable of forming universal multifunctional, intelligent platforms for global communication, navigation, ubiquitous data mining, Earth observation, and many other functions, which was once doomed by the extraordinary cost of such systems. The ingression of novel nanostructured materials provided a solid base that enabled the advancement of these affordable systems in aspects of power, instrumentation, and communication. However, absence of efficient and reliable thrust systems with the capacity to support precise maneuvering of small satellites and CubeSats over long periods of deployment remains a real stumbling block both for the deployment of large satellite systems and for further exploration of deep space using a new generation of spacecraft. The last few years have seen tremendous global efforts to develop various miniaturized space thrusters, with great success stories. Yet, there are critical challenges that still face the space technology. These have been outlined at an inaugural International Workshop on Micropropulsion and Cubesats, MPCS-2017, a joint effort between Plasma Sources and Application Centre/Space Propulsion Centre (Singapore) and the Micropropulsion and Nanotechnology Lab, the G. Washington University (USA) devoted to miniaturized space propulsion systems, and hosted by CNR-Nanotec—P.Las.M.I. lab in Bari, Italy. This focused review aims to highlight the most promising developments reported at MPCS-2017 by leading world-reputed experts in miniaturized space propulsion systems. Recent advances in several major types of small thrusters including Hall thrusters, ion engines, helicon, and vacuum arc devices are presented, and trends and perspectives are outlined.
The increasingly successful use of nanosatellites has made CubeSat an attractive form factor for a number of missions that require both an attitude control system (ACS) and primary propulsion. The CubeSat High Impulse Propulsion System (CHIPS), under development by CU Aerospace with partner VACCO Industries, is designed to enable CubeSat mission operations beyond simple orbit maintenance, including significant altitude changes, formation flying, and proximity operations such as rendezvous and docking. CHIPS integrates primary propulsion, ACS and propellant storage into a single bolt-on module that is compatible with a variety of non-toxic, self-pressurizing liquid propellants. The primary propulsion system uses micro-resistojet technology developed by CU Aerospace to superheat the selected propellant before subsequent supersonic expansion through a micro-nozzle optimized for frozen-flow efficiency. The 1U+ baseline design is capable of providing an estimated 563 N-s total impulse at 30 mN thrust using R134a propellant, giving an impulse density of 552 N-s/liter. The ACS is a cold gas, 4-thruster array to provide roll, pitch, yaw, and reverse thrust with a minimum impulse bit of 0.4 mN-s. An engineering prototype, with integrated pressure control, power control, and data logging, was extensively tested in cold and warm gas modes on the University of Illinois thrust stand under a range of conditions. This paper presents thrust and specific impulse data for the resistojet thruster using two different propellants. Resistojet data include cold and warm gas performance as a function of mass flow rate, plenum pressure, geometry, and input power.
Electric propulsion (EP) technologies offer performance advantages compared to chemical systems. The implementation of electric thrusters in small spacecraft has been prohibited by the high power demands of conventional thrusters and the complexity of the required power conditioning units and propellant supply system. Advances in micro electronics have made the use of EP on small satellites feasible. Research and development on 3 technologies: 1. Helicon Double Layer Thruster (HDLT), 2. Hollow Cathode Thruster (HCT) and 3. Pulsed Plasma Thruster (PPT) has focussed on characterisation and design optimisation for small, micro and nanosatellite missions respectively. All thrusters operate in differing regimes and display varying degrees of development. Initial characterisation of the HDLT operating with argon propellant shows a maximum measured thrust of 1.7 mN for an operating power of 500 W. The current investigation has demonstrated a thrust of 1.7 mN and specific impulse of 70s at 50W for a HCT operating with krypton propellant. The development of PPTs for the STRaND-1 nanosatellite mission demonstrates the option of utilising EP to extend the lifetime of CubeSat missions. Index Terms—Electric Propulsion 1. INTRODUCTION The space sector has been reluctant to widely implement Electric Propulsion on board commercial satellites. This is due to the relative complexity of the technologies and the incumbent power consumption trading off against the high efficiency of the thrusters. Proliferation of research and development of satellite technologies from government organisations to academia and private enterprise has resulted in a focus on low cost technologies with a rapid design and development phase. While advances in microelectronics have allowed the mass, volume and cost of power supply units for EP systems to be reduced [1]. There now exists a wide variety of EP options with varying levels of complexity and technology readiness levels (TRL) which can be exploited to provide options for station keeping, orbit maintenance and orbit raising.
Micro spacecraft which have gained huge popularity in the last decade are termed as “CubeSat”, a well-known class of small satellites. The fast growing functionality and popularity of CubeSat has helped researchers to push the technology demonstration towards efficient performance and reliability needed for commercial and governmental applications. Keeping space debris and use of fossil fuels as propellant in mind, an attempt has been made to develop a green propellant to overcome the issues of space debris, toxicity of propellant and affordable propellant. All liquid propellant based propulsion requires liquid feed system and for that, uses pressurized gas (basically helium) to assist liquid to expel out of the tank. But once the liquid propellant is over, the pressurized gas is unutilized and system runs out of fuel. So the current work, apart from being used as normal cold gas propulsion, it also offers the advantage of using feed system unutilized gases in case of liquid fuel runs out. The present work focuses on cold gas micro thruster development, and its experimental testing in space and sea level condition at various pressure conditions. A computational fluid dynamics model for the same has been developed to validate its experimental results. The thrust value recorded between micro to milli Newton ranges to fulfil the requirements of attitude and station keeping for the CubeSat of 1 to 50 kg dry mass range. Thrust values of 0.8 mN at 1 bar pressure difference to 2.24 mN at 4 bar pressure difference were reported for vacuum environment testing. It’s almost just twice the value of thrust achieved in case of sea level for the corresponding pressure differences. Furthermore, the parameters like pressure, temperature, Mach number, velocity vector, specific impulse, nozzle efficiency etc. are studied and reported for 8 different cases of pressure difference ranging from 1 bar pressure difference to 4 bar pressure difference respectively in atmospheric and vacuum environment.
At present, very few CubeSats have flown in space featuring propulsion systems. Of those that have, the literature is scattered, published in a variety of formats (conference proceedings, contractor websites, technical notes, and journal articles), and often not available for public release. This paper seeks to collect the relevant publically releasable information in one location. To date, only two missions have featured propulsion systems as part of the technology demonstration. The IMPACT mission from the Aerospace Corporation launched several electrospray thrusters from Massachusetts Institute of Technology, and BricSAT-P from the United States Naval Academy had four micro-Cathode Arc Thrusters from George Washington University. Other than these two missions, propulsion on CubeSats has been used only for attitude control and reaction wheel desaturation via cold gas propulsion systems. As the desired capability of CubeSats increases, and more complex missions are planned, propulsion is required to accomplish the science and engineering objectives. This survey includes propulsion systems that have been designed specifically for the CubeSat platform and systems that fit within CubeSat constraints but were developed for other platforms. Throughout the survey, discussion of flight heritage and results of the mission are included where publicly released information and data have been made available. Major categories of propulsion systems that are in this survey are solar sails, cold gas propulsion, electric propulsion, and chemical propulsion systems. Only systems that have been tested in a laboratory or with some flight history are included.
Traditionally, the space industry produced large and sophisticated spacecraft handcrafted by large teams of engineers and budgets within the reach of only a few large government-backed institutions. However, over the last decade, the space industry experienced an increased interest towards smaller missions and recent advances in commercial-off-the-shelf (COTS) technology miniaturization spurred the development of small spacecraft missions based on the CubeSat standard. CubeSats were initially envisioned primarily as educational tools or low cost technology demonstration platforms that could be developed and launched within one or two years. Recently, however, more advanced CubeSat missions have been developed and proposed, indicating that CubeSats clearly started to transition from being solely educational and technology demonstration platforms to offer opportunities for low-cost real science missions with potential high value in terms of science return and commercial revenue. Despite the significant progress made in CubeSat research and development over the last decade, some fundamental questions still habitually arise about the CubeSat capabilities, limitations, and ultimately about their scientific and commercial value. The main objective of this review is to evaluate the state of the art CubeSat capabilities with a special focus on advanced scientific missions and a goal of assessing the potential of CubeSat platforms as capable spacecraft. A total of over 1200 launched and proposed missions have been analyzed from various sources including peer-reviewed journal publications, conference proceedings, mission webpages as well as other publicly available satellite databases and about 130 relatively high performance missions were downselected and categorized into six groups based on the primary mission objectives including “Earth Science and Spaceborne Applications”, “Deep Space Exploration”, “Heliophysics: Space Weather”, “Astrophysics”, “Spaceborne In Situ Laboratory”, and “Technology Demonstration” for in-detail analysis. Additionally, the evolution of CubeSat enabling technologies are surveyed for evaluating the current technology state of the art as well as identifying potential areas that will benefit the most from further technology developments for enabling high performance science missions based on CubeSat platforms.
The operating principle and process of a new type of microthrust measurement stand were presented. The measuring range of the thrust stand is 0-200 mN with ±1% measuring precision. Cold N2 and N2O thruster was employed to ground experiment in vacuum to simulate the working environment. The operating parameters which include thrust, mass flow rate of propellant, pressure and vacuity were tested. Test data was also compared with theoretical prediction and results show that the thrust stand is steady and reliable. The thrust stand can measure mN grade thrust accurately.
The international radio amateur organization AMSAT will launch their P3D satellite weighing 400 kg into a highly elliptical 16 h orbit in April 1996 with the second prototype launch of the ARIANE V rocket. On this satellite, for the first time, the orbit control will be performed by an arcjet system. This system, dubbed ATOS for arcjet thruster on OSCAR satellite, will be a 700 W ammonia arcjet system. The paper describes the project background and the different system components required for the arcjet operation and their development status as of May 1993.
Microsatellites have been suggested as a means of enhancing a variety of proposed space missions, ranging from low-Earth-orbit to solar-system exploration. With improvements in propulsion technology geared toward microsatellites, the ultimate delta-V (deltaV) capabilities of some microsatellite systems are now in the range of several km/s, opening the doors to a variety of high deltaV, fast response scenarios. This paper provides a brief overview of propulsion technologies currently available for microsatellites, and an evaluation of each technology for potential use in a demanding mission. The sample mission is that of a microsatellite inspector which, starting in a 200 km parking orbit, must be diverted to rendezvous with another satellite in orbit at a different altitude and inclination. It is found that existing bipropellant microrocket designs provide a high thrust value, combined with a 300 s specific impulse, allowing for response times of only a few hours for such an inspector mission with deltaV requirements over 1 km/s. Miniaturized electrostatic thrusters provide the largest ultimate deltaV capability, approaching 10 km/s, but with a very low thrust level and therefore a response time capability of several months. Newly developed micro-solar thermal systems fill in the middle ground for these two options, providing the moderate thrust levels and specific impulse values necessary for a response time on the order of one day and a deltaV of several km/s.
Gas Dynamics of Solid Propellant Rocket Motor
- X S Wu
- J Chen
- D Wang
Propulsion subsystem technology for YH-1 Mars probe
- S Q Guo
- H Hou
- Y Zhang
Technology of ammonia flashing jet propulsion in BX-1 satellite
- Q Wei
- Y C Li
Full elastic microthrust measurement equipment
- H B Tang
- C Liu
- M Xiang
- HB Tang