Hall Thruster Direct-Drive Assessment and Demonstration

This thesis involves the theoretical and experimental study of the Hall thruster Direct Drive configuration: a innovative way to deliver power to the electric thrusters and candidate for future propulsion system of spacecraft. The direct connection between the solar array and the thruster anode, made possible by the current development of technology, allows a drastic simplification of the power processing unit of the propulsion system. This has an immediate impact on the propulsion system and its thermal control system which can be consequently lightened. Then, additional mass benefits can be exploited in other parts of the spacecraft such as the electric power system. The work is mainly divided in two parts. The first one assesses in terms of mass reduction the impact that the Direct Drive configuration entails in the spacecraft systems. Different kind of space missions with different level of Hall thruster power are considered. The second part of thesis concerns an experimental demonstration of a Direct Drive system supplying the HT-100, the low power electric thruster developed at Alta. The test required the procurement of the solar panel and the design of an electrical filter. By means of simulations with Pspice, a LC filter was developed and then arranged between the solar modules and thruster in order to dampen the current oscillations. This test successfully proved the correct ignition and operations of the thruster, representing in this way the first attempt in Europe of a Direct Drive demonstration

Propagation and reconstruction of re-entry uncertainties using continuity equation and simplicial interpolation

This work proposes a continuum-based approach for the propagation of uncertainties in the initial conditions and parameters for the analysis and prediction of spacecraft re-entries. Using the continuity equation together with the re-entry dynamics, the joint probability distribution of the uncertainties is propagated in time for specific sampled points. At each time instant, the joint probability distribution function is then reconstructed from the scattered data using a gradient-enhanced linear interpolation based on a simplicial representation of the state space. Uncertainties in the initial conditions at re-entry and in the ballistic coefficient for three representative test cases are considered: a three-state and a six-state steep Earth re-entry and a six-state unguided lifting entry at Mars. The paper shows the comparison of the proposed method with Monte Carlo based techniques in terms of quality of the obtained marginal distributions and runtime as a function of the number of samples used

Constrained optimisation of preliminary spacecraft configurations under the design-for-demise paradigm

In the past few years, the interest towards the implementation of design-for-demise measures has increased steadily. Most mid-sized satellites currently launched and already in orbit fail to comply with the casualty risk threshold of 0.0001. Therefore, satellites manufacturers and mission operators need to perform a disposal through a controlled re-entry, which has a higher cost and increased complexity. Through the design-for-demise paradigm, this additional cost and complexity can be removed as the spacecraft is directly compliant with the casualty risk regulations. However, building a spacecraft such that most of its parts will demise may lead to designs that are more vulnerable to space debris impacts, thus compromising the reliability of the mission. In fact, the requirements connected to the demisability and the survivability are in general competing. Given this competing nature, trade-off solutions can be found, which favour the implementation of design-for-demise measures while still maintaining the spacecraft resilient to space debris impacts. A multi-objective optimisation framework has been developed by the authors in previous works. The framework's objective is to find preliminary design solutions considering the competing nature of the demisability and the survivability of a spacecraft since the early stages of the mission design. In this way, a more integrated design can be achieved. The present work focuses on the improvement of the multi-objective optimisation framework by including constraints. The paper shows the application of the constrained optimisation to two relevant examples: the optimisation of a tank assembly and the optimisation of a typical satellite configuration.Comment: Pre-print submitted to the Journal of Space Safety Engineerin

Re-entry prediction and demisability analysis for the atmospheric disposal of geosynchronous satellites

The paper presents a re-entry analysis of Geosynchronous Orbit (GSO) satellites on disposal trajectories that enhance the effects of the Earth oblateness and lunisolar perturbations. These types of trajectories can lead to a natural re-entry of the spacecraft within 20 years. An analysis was performed to characterise the entry conditions for these satellites and the risk they can pose for people on the ground if disposal via re-entry is used. The paper first proposes a methodology to interface the long-term propagation used to study the evolution of the disposal trajectories and the destructive re-entry simulations used to assess the spacecraft casualty risk. This is achieved by revisiting the concept of overshoot boundary. The paper also presents the demisability and casualty risk analysis for a representative spacecraft configuration, showing that the casualty risk is greater than the 10-4 threshold and that further actions should be taken to improve the compliance of these satellites in case disposal via re-entry is used

Space system design for demise and survival

In the past two decades, the attention towards a more sustainable use of outer space has increased steadily. The major space-faring nations and international committees have proposed a series of debris mitigation measures to ensure the sustainability of the space environment. Among these mitigation measures, the de-orbiting of spacecraft at the end of their operational life is recommended in order to reduce the risk of collisions in orbit. However, re-entering spacecraft can pose a risk to people and property on the ground. A possible way to limit this risk is to use a design-for-demise philosophy, where the spacecraft is designed such that most of its components will not survive the re-entry process. However, a spacecraft designed for demise still must survive the space environment for many years. As a large number of space debris populates the space around the Earth, a spacecraft can suffer impacts from these particles, which can be extremely dangerous. This means that the spacecraft design has also to comply with the requirements arising from the survivability against debris impacts. The demisability and survivability of a spacecraft are both influenced by a set of common design drivers, such as the material of the structure, its shape, dimension, and position inside the spacecraft. It is important to consider such design choices and how they influence the mission’s survivability and demisability from the early stages of the mission design process. The thesis addresses these points with an increasingly higher level of detail by a continuous and interlinked development of a demisability and a survivability model, of two criteria to evaluate the level of demisability and survivability, and of a common framework where both models communicate and interact to find optimal solutions. First, the initial versions of the models, which is limited to simple geometrical shapes, uniform materials, and dimensions, is used to study the sensitivity of the demisability and of the survivability indices as a function of typical design-for-demise options. As new features are introduced, such as the capability of considering internal components and sub-component together with their position inside the spacecraft, as well as the type of shielding, also the analyses become more detailed. As the demisability and the survivability of a spacecraft configuration are closely linked, it is important to assess them in a concurrent fashion for which a multi-objective optimisation framework has been developed. Here the survivability and thedemisability requirements are considered simultaneously and trade-off solutions of spacecraft configurations can be obtained. The final part of the thesis presents a test case for the application of the framework, targeting one of the most interesting components from both a demisability and a survivability standpoint that are tank assemblies. Finally, a preliminary study concerning the development of a new demisability index is presented

Dataset for the 2017 Space Debris Conference article

The data-set contains all the data relative to the results of the constrained optimisation of tank assemblies for representative sun-synchronous missions. Different mission classes (in terms of mass) and mission lifetimes are taken into account. In addition three different demisability fitness functions are tested and compared.</span