C. Scharlemann

Liquid-metal-ion source development for space propulsion at ARC

The Austrian Research Centers have a long history of developing indium Liquid-Metal-Ion Source (LMIS) for space applications including spacecraft charging compensators, SIMS and propulsion. Specifically the application as a thruster requires long-term operation as well as high-current operation which is very challenging. Recently, we demonstrated the operation of a cluster of single LMIS at an average current of 100muA each for more than 4800h and developed models for tip erosion and droplet deposition suggesting that such a LMIS can operate up to 20,000h or more. In order to drastically increase the current, a porous multi-tip source that allows operation up to several mA was developed. Our paper will highlight the problem areas and challenges from our LMIS development focusing on space propulsion applications.

Development of Electric and Chemical Microthrusters

The increasing application of microsatellites (from 10 kg up to 100 kg) as well as CubeSats for a rising number of various missions demands the development of miniaturized propulsion systems. Fotec and The University of Applied Sciences at Wiener Neustadt is developing a number of micropropulsion technologies including both electric and chemical thrusters targeting high performance at small scales. Our electric propulsion developments include a series of FEEP (field emission electric propulsion) thrusters, of which the thrust ranges from μN to mN level. The thrusters are highly integrated into clusters of indium liquid-metal-ion sources that can provide ultralow thrust noise and long-term stability. We are also developing a micro PPT thruster that enables pointing capabilities for CubeSats. For chemical thrusters, we are developing novel micromonopropellant thrusters with several hundred mN as well as a 1–3 N bipropellant microrocket engine using green propellants and high specific impulse performance. This paper will give an overview of our micropropulsion developments at Fotec, highlighting performance as well as possible applications.

Indium FEEP Micropropulsion Subsystem for LISA Pathfinder

ARC-sr together with EADS Astrium and EADS Space Transportation is jointly developing a flight design for an Indium FEEP Microthruster suitable for LISA Pathfinder. The paper reviews all basic design elements and performance parameters and shows that the Indium FEEP thruster is capable of meeting all key requirements for ultra-precise attitude and orbit control.

Development of a µPPT for CubeSat Applications

An analytical model based on an oscillatory circuit equation and dynamic equation for a uniform plasma slug has been developed in order to provide a design tool which is able to predict the dependence of µPPT performance on electrode configuration and circuit paramters. The results of the model show good aggrement with experimental results and have been implemented to improve µPPT performance. Thruster performance has been experimentally characterised for a capacitance range of 2 to 31 µF and propellant surface areas of 0.25 to 1.5 cm2 for low energy operation. Results indicate that the impulse bits required for picosatellite missions are achievable with the application of miniaturised thruster heads and energy densities of approximatley 5 to 10 Jcm2.

Test of a Turbo-Pump Fed Miniature Rocket Engine

The increasing application of microsatellites (from 10 kg up to 100 kg) for a rising number of various missions requires the development of suitable propulsion systems. Microsatellites have special requirements for a propulsion system such as small mass, reduced volume, and very stringent electrical power constraints. Existing propulsion systems often can not satisfy these requirements. The present paper discusses the development and test of a bipropellant thruster complying with these requirements. The main development goal of this effort was the utilization of ethanol in combination with hydrogen peroxide (H2O2) as a non-toxic propellant combination. The Turbo-Pump Fed Miniature Rocket Engine (TPF-MRE) is a bipropellant thruster consisting of four subsystems: the propellant pumps, a decomposition chamber with a monolithic catalyst, a turbine, and the thruster itself. The turbine is driven by the decomposed hydrogen peroxide and magnetically coupled with a power generator. The power produced is then used to generate a pressure head in order to deliver the propellant into the combustion chamber. This system therefore constitutes a self-sustaining system and does not rely on the limited power supply of a micro-satellite. Previous test have shown that although the thruster can be operated with ethanol and oxygen, it was not possible to ignite the thruster when utilizing hydrogen peroxide in a 70% concentration by weight. A minor redesign of the thruster and the test facility was therefore initiated. This redesign together with the use of hydrogen peroxide in higher concentration was speculated to improve this behavior. However, even though the monolithic catalysts were able to decompose hydrogen peroxide in a concentration of 87.5 % with nearly 100 % efficiency, it was not possible to ignite or operate the thruster. Subsequently, a thorough investigation of the baseline design and operational conditions of the thruster was conduced. It was found that the failure of the thruster to ignite is due to a combination of reasons. The combustion chamber length is too short to facilitate sufficient mixing of the propellants, making an ignition impossible or very difficult at least. Additionally, the combustion chamber pressure which was chosen such that it accommodates the performance of commercially available mircopumps is considered too low. This further deteriorates the conditions for which an ignition is feasible.

Numerical simulation of SMART-1 Hall-thruster plasma interactions

SMART-1 has been the first European mission using a Hall thruster to reach the moon. An onboard plasma diagnostic package allowed a detailed characterization of the thruster exhaust plasma and its interactions with the spacecraft. Analysis of in-flight data revealed, amongst others, an unpredicted large cyclic variation of the spacecraft floating potential and a mysterious asymmetry in the plasma surrounding the spacecraft. To investigate the details of the anomalies, we developed the numerical software tool SmartPIC to characterize and predict spacecraft-plasma interactions. Technical details, such as solar arrays and onboard diagnostic devices, have been modeled with high accuracy. All basic plasma parameters, the spacecraft floating potential, backflow distributions, and ion impact energies are calculated by the code and are available in high spatial resolution throughout the computational domain containing the entire satellite. It was possible to clearly identify the rotating solar cells arrays as the source of the cyclic variation of the spacecraft floating potential. Furthermore, the asymmetry of the plasma formation around the spacecraft is linked to the location of the neutralizer causing a region of increased charge-exchange collisions particle generation. Both of these results have significant impact on the implementation of electric propulsion systems on satellites.

Analytic model for the Assessment of the Electrode Configuration of a µPPT

The analytic model presented is aimed at understanding the dependency of Micro Pulsed Plasma Thruster performance on electrode geometry parameters. An advanced, one dimensional electromechanical model, based on accurate inductance calculation and a detailed description of the inhomogeneous magnetic field distribution accelerating the plasma, will be introduced. Including effects of inhomogeneity in an analytical model of a µPPT not only enables an accurate description of the plasma acceleration process for configurations with large propellant height compared to width, but can also be a valuable design tool regarding thruster geometry and its impact on performance. The effect of electrode separation and thickness on plasma acceleration and performance will be discussed, as well as the influence of changing electrical parameters. The impact of flared electrodes is investigated and found as a possible way to increase energy transfer into the plasma and therefore thruster performance. The model is validated by comparison to data from experimental investigation on a miniaturized Pulsed Plasma Thruster.

Influence of the Solar Array s on the Floating Potential of SMART -1: Numerical Simulations

research is inves tigating this anomaly with the numerical code SmartPic. The most recent results of this investigation are presented in the present paper . They indeed show that the solar array rotation is responsible for the observed anomaly. The relative position of the s olar array, or to be more precise, of the interconnectors on the array, to the thruster plasma governs the electron current collected by the spacecraft. The variation of the available interconnector surface to collect the electron current then determines t he floating potential. SmartPic successfully predict s very similar variations of the floating potential as the measurements on -board of SMART -1. The SPEDE probes were successfully implemented and simulated. The results of the se simulation show that the difference in collected current of the two probes is due to their position on two opposite sides of the spacecraft. The one probe closer to the neutralizer is in a region of higher charge exchange number densities therefore increasing the current flux on the probe.

In-FEEP Qualification Test Program for LISA Pathfinder

The Laser Interferometer Space Antenna project (LISA) is a co-operative program between ESA and NASA to detect gravitational waves by measuring distortions in the spacetime fabric. For this task, three satellites will fly in a triangular formation. LISA Pathfinder is the precursor mission to LISA designed to validate the core technologies intended for LISA. One of the enabling technologies is the micro-propulsion system necessary to achieve the uniquely stringent propulsion requirements.

Development of Miniaturized Green Propellant Based Mono- and Bipropellant Thrusters

,The increasing application of microsatellites (from 10 kg up to 100 kg) for a rising number of various missions requires the development of suitable propulsion systems. Microsatellites have special requirements for a propulsion system such as small mass, reduced volume, and very stringent electrical power constraints. Existing propulsion systems often can not satisfy these requirements. The present paper discusses the successful development and test of a two thruster systems operating with green propellants: a monopropellant and a bipropellant thruster. The main development goal of this effort was the utilization of green propellants. Both thruster types use hydrogen peroxide in a concentration of 87.5 %. Ongoing investigation will identify the ideal fuel for the bipropellant thruster. The two main candidates are kerosene and ethanol. Previous reported difficulties to generate autoignition in the bipropellant system have been solved. Initial test have shown that the thruster reliably auto-ignites without using an additional ignition system. Based on the test results, the performance of the bipropellant thruster was calculated. With the present mass flow rates and fuel ratios, the thruster generates a thrust between 1.7 and 1.9 N with a specific impulse between 290 and 330 s. The monopropellant system developed by ARC was further improved and tested. The catalytic decomposition with its advanced monolithic catalyst has obtained efficiencies up to 99% resulting in decomposition temperatures of 650°C. Performance measurements on a thrust balance has shown that the thruster can generate thrust between 150 and 1700 mN with a specific impulse of 153 s. Additional improvements presently implemented have the potential to increase the specific impulse to 164 s.