Electrothermal propulsion can be seen as an intermediate concept between electrical and chemical propulsion. Propellant heating typically happens by means of a resistance (resistojet) or an electrical discharge (arcjet). For the extremely miniaturized applications that will be discussed in this chapter, resistojets are by far the most commonly used form of electrothermal thrusters.

Small satellites with increased capabilities in terms of power and propulsion are being demanded for future missions. This paper addresses an alternative bi-modal solution which consists of a solar thermal propulsion system coupled with a micro-Organic Rankine Cycle system, to co-generate thrust and electrical power. Current literature on bi-modal systems is limited to static power conversion systems such as thermionic conversion processes. Therefore, this paper expands the research of bi-modal systems to dynamic power conversion systems and latent heat storage systems. The paper documents the design process, key design parameters, and feasibility of this system for a Geostationary Transfer Orbit to Lunar Orbit insertion mission. The results of a single-objective optimisation show the system is most suitable on-board small satellites with a gross mass above 300 kg. The propellant accounts for 50% of the total system mass. The final design uses Silicon as the latent heat energy storage system due to its high specific energy of more than 250 Wh/kg. Additionally, the enthalpy method is used to describe the dynamic behaviour of the phase change material and results show the insulation thermal conductivity has the largest effect, up to 17%, on the receiver's maximum achievable steady-state temperature.

Small satellites are receiving increased recognition in the space domain due to their reduced associated launch costs and shorter lead time when compared to larger satellites. However, this advantage is often at the expense of mission capabilities, such as available electrical power and propulsion. A possible solution is to shift from the conventional solar photovoltaic and battery configuration to a micro-Organic Rankine Cycle (ORC) and thermal energy storage system that uses the waste energy from a solar thermal propulsion system. However, limited literature is available on micro-ORC systems, which are capable of producing a few hundred Watts of electrical power. This paper describes the proposed system layout and model of the integrated micro-ORC system, for various working fluids such as Toluene, Hexamethyldisiloxane (MM), and Octamethylcyclotetrasiloxane (D4). Toluene has been identified as a promising working fluid candidate resulting in a power generation system volume fraction of 18% for a 215 kg Low Earth Orbit satellite. The micro-ORC system is capable of producing 200 W of electrical power. The design provides high specific energies of at least 500 Wh/kg but, has a low shared specific power of 10 W/kg. A preliminary design of the micro-turbine provides a conservative total-to-static efficiency of 57%.

Small satellites with increased capabilities in terms of power and propulsion are being demanded for future missions. This paper proposes a possible solution which is the design of a novel integrated solar thermal system that co-generates propulsion and power on-board mini satellites. The system consists of a solar thermal propulsion system (STP) coupled with a micro-Organic Rankine Cycle (ORC) system to harness the waste heat from the STP receiver to provide electrical power and mitigate the need for solar panels. STP provides an alternative to conventional propulsion systems for missions requiring velocity changes of between 800 m/s and 2500 m/s. Additional advantages include higher specific impulses than chemical propulsion systems, throttability, re-start capabilities, and faster transfer times than electrical propulsion systems. The faster transfer times are especially useful for missions that travel across high radiation regions such as the Van Allen Belt. This unique configuration shares resources such as the concentrator and receiver to potentially extend the power and propulsion capabilities while adhering to the strict mass and volume constraints of small satellites. However, there is currently no literature available on the design process of the proposed bi-modal system. This paper therefore presents an integrated solar thermal design strategy for a Geostationary Transfer Orbit to Lunar orbit insertion mission. The design methodology is described in detail to assist with future evaluations of integrated solar thermal systems for other applications and missions. The system is designed to provide a velocity increment of 1.6 km/s. Five mini-satellite sizes were investigated with a gross wet mass of 100 kg, 200 kg, 300 kg, 400 kg, and 500 kg respectively. Each satellite requires to produce an electrical power of 1 W/kg. The STP system uses water as the propellant due to its safety and performance attributes. Toluene has been selected as the working fluid for the ORC due to its high thermal efficiency. By incorporating the use of a high-temperature receiver, propellant temperatures around 2500 K can be achieved that can produce high specific impulse values of more than 300 s. The design has been optimized for various design parameters, such as propellant temperature, nozzle area ratio, burn time, concentrator design, and ORC cycle pressures. The optimization provides an initial framework in the selection of an optimal integrated solar thermal design for the proposed Lunar mission. An analysis of variance has also been conducted to identify which system parameters, such as optical efficiency and turbine efficiency, have the most influential effect on the system. The heaviest components of the system are the propellant (40 to 50%), concentrator (8%), and insulation (8%) with respect to the gross mass of the satellite.

In recent years, satellite design has extended towards miniaturisation to reduce associated cost with launching and conducting space missions. Small satellites provide low-cost platforms for space missions. However, this lower cost comes at the expense of the removal of key sub-systems, such as the propulsion system, due to the small available onboard volume and mass restrictions. For this reason, small, lightweight, high-performing and affordable propulsion systems are necessary. However, there is limited research available on the comparison of propulsion technologies with regards to cost. Motivated by the above challenges the objective of this paper is to provide a comparison of propulsion technologies that are compatible with small satellites with respect to cost and application. The different propulsion systems are investigated for three mission scenarios, a small on-orbit manoeuvre, a station-keeping, and a lunar orbit transfer mission. Each system is evaluated in terms of a total figure of merit which incorporates nine variables such as propellant mass, safety, and hardware price, that affect the total cost of a propulsion system. This figure of merit is used to quantitatively compare the propulsion systems to identify cost-effective solutions as a function of various mission scenarios. Solar thermal propulsion has been proposed for small satellite applications, but information regarding the concepts are not available in a single report. Therefore, a secondary objective of this paper is to provide the reader with a review of the current status of solar thermal propulsion. An important finding of this research is the classification of propulsion systems in terms of thrust, specific impulse, cost, and application.

The recent developments in space exploration have reinstated the Moon as a primary target for near future space missions. The principal reasons include the Moon being the closest test bed and analogue for planetary space missions and the prospect of scientific lunar bases and orbital stations within the next decade. Previous space missions have vastly improved our understanding on hazards of human spaceflights but not fully regarding the threats affecting a prospective lunar base or orbital station. The micrometeorite hazard has been partially addressed as an issue which can potentially impact both astronauts' health and safety as well as create issues for lunar bases and orbital stations, such as degradation or permanent damage of equipment and facilities. The current understanding is based partly on dust and micrometeoroid flux measurements and impact flash observations. However, observations with improved spatial and temporal resolution are imperative for advancing existing hazard models. In this paper, a mission concept of a constellation of nanosatellites is proposed that can both observe larger parts of cis-lunar and trans-lunar space while providing higher temporal resolution. Nanosatellite missions are a cost-effective solution providing data for significant improvement of our current understanding of lunar micrometeoroid flux models, and thus directly the scale of hazards caused by micrometeoroid impacts to future lunar missions. Additionally, such a distributed constellation mission will offer countless opportunities for academia, students and young scientists worldwide. The mission concept (Moon Compact Satellite for Hazard Assessment - MOOCHA) is a result of the Nordic-European Astrobiology Campus Summer School 2018 themed “Microsatellites in Planetary and Atmospheric Research” and was further developed during the 2019 follow-up summer school “Design of Small Satellite Missions for Planetary Studies”, both taking place in Tartu, Estonia and co-organized by the Stockholm University Astrobiology Centre, the University of Tartu, the European Astrobiology Campus and the Nordic Network of Astrobiology and supported by European Union's European Regional Development Fund and Estonia.

The concept of the composite monocoque chassis has been implemented in many vehicle designs; however, there is limited open literature defining the process of simulating a composite monocoque chassis. The purpose of this research is to develop a composite monocoque chassis by analysing its structural integrity through an iterative finite element analysis process with the intention of developing a lightweight solar-powered vehicle. Factors that influence this methodology include; the definition of the vehicle loading conditions, failure criteria, and important design parameters, chief among which is the torsional stiffness. The primary design criterion considered is the torsional stiffness which is determined from the application requirements and data available in the literature. The design methodology then follows an iterative process where various geometry and lay-up changes are considered. Under the same loading conditions, with the aim of increasing the torsional stiffness to achieve the required parameter. The ultimate strength of the material was also considered throughout the simulation process however, in most cases, the model failed to meet the torsional stiffness parameter before the material failure or delamination. Secondly, an analysis of the mounting points was conducted to ensure that the chassis is able to withstand the concentrated loads at the suspension mounts. This analysis is concerned with the principal stresses which gives insight into the most suitable orientation of the lay-up. The methodology presented in this paper stands to be supportive in designing a fully composite monocoque chassis for lightweight race vehicle applications.