Carsten Scharlemann

Development and Test of a Miniature Hydrogen Peroxide Monopropellant Thruster

Analysis of present and future missions concluded that a miniaturised hydrogen peroxide monopropellant rocket engine is the optimum solution for the increasing demand for small and low cost propulsion systems for small satellites. The attractiveness of monopropellant thrusters is based on its operational and structural simplicity. Additionally, the utilization of hydrogen peroxide as propellant instead of hydrazine allows the reduction of the overall costs and would qualify such a system as a green propellant propulsion system. The present paper describes the development of a monopropellant thruster utilizing hydrogen peroxide and advanced catalyst beds. The utilization of a monolithic catalyst reduces the pressure loss across the catalyst bed significantly compared to formerly used pellet or gauze catalyst. This allows the use of relative lightweight tank and significantly minimizes the total weight. For Two different catalyst materials have been developed to achieve optimized decomposition. The present paper summarizes the experimental evaluation of the catalysts. Decomposition temperatures of up to 670°C and decomposition efficiencies up to 99% have been achieved. Up to 1.2 kg of hydrogen peroxide has been decomposed by a single catalyst, corresponding to about 1.25 hrs of operation. This is estimated to correspond in vacuum condition to a total delivered total impulse of 1600 Ns. A thrust balance was designed and built. Preliminary thrust measurements under atmospheric conditions have shown that the laboratory model can generate thrust in a range of at least 50 to 550 mN.

Space micropropulsion systems for Cubesats and small satellites: From proximate targets to furthermost frontiers

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.

Assessment of Catalysts for Hydrogen-Peroxide-Based Thrusters in a Flow Reactor

Hydrogen peroxide is a candidate propellant for rocket-propulsion applications with the potential to replace highly toxic propellants currently used. Decomposition of hydrogen peroxide yields a high-temperature oxygen-steam mixture, which can be used as monopropellant or as oxidizer in a bipropellant configuration. This work examines different types of cellular ceramic-based catalysts for hydrogen-peroxide decomposition at miniature scale of nominal mass flows of 0.3  g s−1. An exhaustive investigation of different catalysts in a flow reactor configuration similar to a propulsion application is conducted. The test matrix includes honeycomb monoliths with different channel geometries, densities, lengths, different carrier materials, and wash-coating procedures, as well as different types of catalysts such as pellets and foams. Thirty nine catalyst configurations with a total of 121 catalysts have been experimentally investigated based on their transient and stationary performance at design mass-flow levels...

Hydrogen Peroxide Decomposition for Micro Propulsion: Simulation and Experimental Verification

Hydrogen peroxide is under investigation with regard to its potential to replace the presently used highly toxic oxidizers such as NTO or MON-3. Catalytically decomposed hydrogen peroxide results in a steam-oxygen mixture at elevated temperature and can be used either as a monopropellant or as an oxidizer in a bipropellant system. In order to achieve high decomposition efficiencies it is essential to understand and to be able to control the decomposition processes in detail. In particular, the choice of catalyst is one of the most essential issue in designing a propulsion system based on hydrogen peroxide. However a catalyst is defined by a multidimensional parameter matrix including the catalyst nature, diameter, length, inner and outer shape, heat capacity and conductivity of the carrier material, and, manufacturing method and many others. Reliable experimental investigation of a catalyst is a time consuming effort. To guide the experimental assessment, Fotec (formerly Austrian Institute of Technology – AIT) has developed an analytical model of the decomposition implemented into a numerical thermal model. The one dimensional decomposition model coupled to a finite element structural domain of the decomposition chamber is used to investigate the impact of the catalyst and, in addition, of the chamber structure on the decomposition behavior. Special focus is laid on the transitional behavior of hydrogen peroxide conversion to facilitate immediate start-up of the thruster system. The numerical results have been validated with experimental values. The comparison shows high accuracy of the predictions not only in the general decomposition behavior but also in intrinsic details such as the transitional behavior. Major findings of the model such as the existence of a radial temperature gradient across the catalyst have been experimentally validated. These findings point to an overestimation of experimentally determined decomposition performances, in the case of temperature measurements just downstream of the catalysts cental line. Another major finding is the identification of an mass flow overload threshold by the simulation, yielding a sudden decrease in decomposition performance after surpassing the threshold. This sudden decrease in decomposition performance has been experimentally verified.

Development and Test of an Indium FEEP Micropropulsion Subsystem 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. 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.

In-FEEP Endurance Test for LISA PF

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. 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. Two competing systems, a cesium slit emitter (Alta, Italy) and the indium needle emitter (AIT, Austria) technology have been commissioned to develop this micro-propulsion system. At this point, the cesium slit emitter was chosen by ESA as baseline. The indium needle emitter technology was chosen as back-up and its development and test are still proceeding and the obtained results are documented in the present publication.

Development of Propulsion Means for Microsatellites

The increasing application of micro-satellites (from 10 kg up to 100 kg) for a rising number of various missions demands the development of new miniaturized propulsion systems. Micro-satellites have special requirements for the propulsion system such as small mass, reduced volume, and very stringent electrical power constraints. Existing propulsion systems often can not satisfy these requirements. A number of micropropulsion systems are presently under development at the Space Propulsion & Advanced Concepts department from the Austrian Research Centers. The portfolio of the systems under development includes electrical and chemical propulsion systems. The covered thrust and specific impulse of the developed propulsion systems ranges from 1µN to 1N and 150 s to 8000 s respectively. The present paper will give an overview of the propulsion development program at ARC.

Q³Sat: quantum communications uplink to a 3U CubeSat: feasibility and design (Conference Presentation)

In the absence of technically mature quantum repeaters, losses in optical fibers limit the distance for ground-bound quantum key distribution. One way to overcome these losses is via optical links to satellites, which has been demonstrated in course of the Chinese-Austrian QUESS mission. Though its findings were impressive, such a large-scale project requires massive financial and time resources. We propose a 32x10x10cm³ nanosatellite orders of magnitude cheaper which is able to perform quantum key distribution (QKD) in a trusted-node scenario, using only commercially available components. We have performed a detailed analysis of such a CubeSat mission (“Q³Sat”), finding that cost and complexity can be reduced by sending the photons from ground to satellite, i.e. using an uplink. Calculations have been done for a prepare-and-send protocol (BB84 with decoy pulses) and for a protocol exploiting quantum entanglement (E91), both using polarization as information carrier. We specified the minimum requirements for the sender stations for these two different protocols. Possible orbits have been assessed, regarding both height and ellipticity to maximize link time and minimize losses. Using long-term weather data, we developed a beam model taking into account absorption, turbulence-induced beam divergence and pointing stability of sender and receiver telescope. Using light pollution measurements from space and their spectra, we arrive at maximum expectable noise count rates. We also specify the requirements for clock stability, classical communication speed and computing requirements. Incorporating all these parameters into our model, we arrive at a link budget which we can use to calculate the expected quantum secure key rates. We have also created a preliminary design of such a 3U CubeSat including a detailed size, weight and power budget and a CAD to account for the assembly of the components. Deploying a 10 cm long mirror telescope covering the small surface of the satellite leaves enough space for a polarization analysis module and housekeeping, communication and computing electronics. Polarization analysis can be done via a polarizing beam splitter and single-photon detectors with a cross section small enough to rule out radiation damage. Pointing stability and detumbling is crucial especially for such a small satellite and can be achieved via spinning wheels, achieving a precision in the tilt and yaw axis of 40 mrad. For one such CubeSat, we estimate the quantum secure key to be acquired between two ground stations during one year to be about 13 Mbit when deploying a decoy protocol. A Bell test between ground and satellite would also be feasible. The uplink design allows to keep the more sensitive, computation-intensive and expensive devices on ground. The experiment proposed by us therefore poses a comparably low-threshold quantum space mission. For a two-year lifetime of the satellite, the price per kilobit would amount to about 20 Euro. In large constellations, Q³Sats could be used to establish a global quantum network, which would further lower the cost. Summarizing, our detailed design and feasibility study can be readily used as a template for global-scale quantum communication.

Influence of water content in an ADN based liquid monopropellant on performance characteristics

Hydrazine is advantageous for attitude control systems of satellites because it is space storable for long times and the developed thrusters are reliable for long term operations. Unfortunately hydrazine is very difficult to handle on ground due to its very high toxicity. Especially with regard to the REACH regulation of the European Community, which has passed some years ago, Hydrazine is on the candidate list of substances whose use could be limited in future. Thus strong research efforts have to be conducted today to find and to qualify alternative propellants, which have significantly simpler handling characteristics, are less toxic, environmentally benign and have similar or even better performance characteristics. One of the most promising candidates to replace hydrazine is the ADN based ionic liquid FLP-106, which is a monopropellant and has been developed by FOI. This monopropellant has a higher Isp in comparison to hydrazine and to LMP-103S which is currently used by the ECAPS company from Sweden. Also FLP-106 has a lower vapor pressure than LMP-103S. Thrusters using ionic liquid propellants are working with heated catalysts to decompose the propellant. One of the main drawbacks of this design is the lack of cold-start capability. Additionally, the high combustion temperature is a concern, possibly requiring high-temperature alloys like Iridium/Rhenium. FLP-106 consists of 64.6 % ADN (ammonium dinitramide), 23.9 % water and 11.5 % of a low volatile hydrocarbon fuel. In order to use simpler materials with lower melting points the combustion temperature of FLP-106 can be decreased by increasing the water content in the propellant. However, the increased water content will decrease the specific impulse and may also influence the ignition properties.

Micropropulsion Development at the ARC Seibersdorf Research

The increasing application of micro-satellites (from 10 kg up to 100 kg) for a rising number of various missions demands the development of new miniaturized propulsion systems. Micro-satellites have special requirements for the propulsion system such as small mass, reduced volume, and very stringent electrical power constraints. Existing propulsion systems often can not satisfy these requirements. The Space Propulsion Department of the ARC Seibersdorf research dedicated itself to the development and test of various micropropulsion systems for present and future space missions. The portfolio of the systems under development includes electrical and chemical propulsion systems. The covered thrust and specific impulse of the developed propulsion systems ranges from 1μN to 1N and 500 s to 8000 s respectively. Based on the large experience obtained over several decades in the development of Field Emission Electric Propulsion systems (FEEP), several microstructured FEEPs have been developed. The design of these systems is presented as well as preliminary test results and a summarization of the experience obtained during the process of miniaturizing such systems. The development of miniaturized chemical propulsion systems includes a bipropellant and a monopropellant thruster. The bipropellant thruster constitutes the smallest existing 1N thruster utilizing hydrogen peroxide. The thruster system consists of two micopumps for the propellant feed and a microturbine to generate the power for operating the pumps. The monopropellant thruster is a derivative of the bipropellant thruster. It offers a lower specific impulse than the bipropellant system but due to its reduced system complexity it represents also a promising candidate for several future space missions. Both systems utilize rocket grade hydrogen peroxide (green propellant), which is decomposed with the help of an advanced monolithic catalyst. The present paper discusses the design methods and the physical limitations of such chemical propulsion systems with regard to their miniaturization and summarizes their performance evaluation.© 2006 ASME