No description available.

Spacecraft that orbit in Low Earth Orbit travel through a tenuous atmosphere and hence experience aerodynamic forces that can become quite significant, specially at low altitudes. The presence of these forces can become a major design driver for missions that fly at very low altitudes. Unfortunately, spacecraft aerodynamics are not well understood. In this dissertation, a CubeSat mission is proposed which will study rarefied-gas aerodynamics, with the objective of determining the effect of surface composition, surface finishing and flow incidence angle on the drag and lift coefficients with an error of less than 5% using a novel method. The CubeSat, has been named ΔDsat, because this study, will be performed using differential measurements of drag and lift coefficients in order to eliminate any measurement bias. ΔDsat carries 4 deployable fins that can rotate independently and expose different surface types to the flow at different incident angles. In addition, in the dissertation four methods to exploit the aerodynamic forces for the missions advantage are proposed and described in detail. The first one is aerostability, which by shaping the spacecraft appropriately, the resulting aerodynamic torques stabilise the attitude spacecraft with respect to the flow. The second method uses aerodynamic drag and lift to change de inclination of a decaying spacecraft in order to maintain the Sun-synchronous aspect of an orbit whilst decaying. The required lift to drag ratio is in the order of 1.0-1.6, which is not currently achievable (it is theoretically possible), but it could be achieved if drag compensating propulsion is used (thus becoming a fuel saving strategy). The third method controls the atmospheric re¬entry interface (the location of the burn-up) by modulating the drag, hence controlling the decay profile. When applied to ΔDsat an error of less than 200 km 3cr on the re-entry location is achieved. Finally, aerostable spacecraft can be used to perform in-situ measurements of the atmospheric winds, by observing their attitude evolution. The aerostable ΔDsat CubeSat would be capable of determining the cross-track winds with an error of less 4 m/s 3cr.

The High Resolution Dynamics Limb Sounder (HIRDLS) is part of NASA's Earth Observing System (EOS). The system provides satellite instrumentation to observe the extent, cause and consequences of climate change. HIRDLS is a radiometer designed to monitor the chemistry and dynamics of the atmosphere at a high vertical and horizontal resolution from the upper troposphere to the mesosphere. Its radiometric, spectral and field of view pre-flight calibrations have taken place at Oxford in a purpose-built facility. This thesis describes the radiometric calibration. Radiance measurement traceability (the absolute radiometric uncertainty) is increasingly important for monitoring long term climate change. This area has been given particular attention. This thesis shows how this traceability has been achieved. The In Flight Calibrator (IFC) has been subjected to extensive testing and calibration which are described. Two large external blackbody targets have been used to provide calibration radiance to HIRDLS in the test facility. The uncertainties in these radiances have been assessed. The targets have facilitated the pre-flight measurement of the radiometric calibration parameters that are required before flight and allowed the in-flight calibration algorithm to be tested. The pre-flight radiometric calibration has proved invaluable to HIRDLS. If the calibration had not taken place and the instrument was assumed to be linear and "self calibrating" the atmospheric temperature retrieval accuracies would have been limited to at best 4.3 K. Using the calibration data and a simple non-linear correction this work has improved the highest accuracy to 0.8 K. This result justifies the effort put into the pre-flight radiometric calibration of HIRDLS.

Actuator failures onboard spacecraft often lead to unsatisfied control performances or even the lose of the whole mission. The aim of this thesis is to find an optimal and generic backup strategy for attitude control in the case of actuator failure. With UoSAT-12 as an example, two existing backup strategies are compared using different combinations of the remaining magnetorquers and reaction wheels. The" constant gain like" control law shows poor performances when actuator fails. A Lyapunov- based discontinuous control law is modified with the consideration on capacity of actuators and adopted for underactuated attitude control with two reaction wheels. The simulation results indicate that Lyapunov-based underactuated attitude control law can achieve better performances and is promising to deal with actuator failures. To improve the performances, two new under actuated attitude control laws are proposed based on feedback linearization technology. The two new under actuated attitude control laws can provide higher accuracy and faster attitude response. However, robustness and attitude tracking are still problems for underactuated attitude control design. Nonlinear H (X) theory is utilized in this thesis to further improve the performances of underactuated attitude control since it can deal with external disturbances and unmodeled errors. A new energy-compensation based approach, which can guarantee a solution of Hamiltonian-Jacobi-Isaacs (HJI) equation for a controllable system, is proposed. Its applications on UoSAT-12 fully active attitude control show that the control law via this approach needs smaller control inputs but against larger disturbances compared with existing approaches. It is also proven both mathematically and numerically with cascade discontinuous control law as an example in this thesis that nonlinear H (X) theory can be used to improve control accuracy against disturbances and to estimate the region of convergence. Based on that, new under actuated attitude control laws are proposed via different approaches of solving HJI inqualities. The simulations with two reaction wheels onboard UoSAT-12 indicate that nonlinear H (X) control design is a much simpler way for underactuated attitude control problem. It can not only stabilize the attitude of the satellite, but also solve the problem of attitude tracking. It also shows that the control laws based on nonlinear H (X) theory can attenuate the disturbances from both external environment and un-modelled system errors. It provides a systematic approach with well defined methods for specifying performances and capacities of onboard actuators.

The present study used scanning and transmission electron microscopy, energy dispersive analysis of X-rays (EDAX) and spectroscopy (FTIR, UV-Visible and fluorescence) to examine terrestrial materials of possible astronomical significance (Oedogonium sp., Enteromopha intestinalis, Pelvetia canaliculata, Fucus vesiculosus, Bacillus cereus, Staphyllococcus aureus, poppy seed, chlorophylls 'a' and 'b', Panicum maximum, anthracite, bituminous coal, naphthalene), the Tagish lake and Carancas meteorites, a Kerala red rain sample and stratospheric air particles collected at altitudes of 38-41 km. The study was designed to determine if any of the terrestrial samples could be proposed as an effective model for the interpretation of astronomical spectroscopic observations. The study also set out to search for evidence to shed light on the origin of these meteorites, red rain and stratospheric air particles. The spectra of all the terrestrial samples (including the meteorites) exhibited absorptions in the Mid-IR region, similar to astronomical features displayed by a variety of galactic sources. Algae (Odeogonium sp.) in particular produced the largest number of absorption peaks, most of which matched those of the astronomical emission spectra of PPNe and also the UIBs. Based on these observations, algae could be defended as a biological model for the interpretation of UIBs and PPNe, and a potential candidate for interstellar material. Coal and semi anthracite, that can be regarded as steps in the degradation of biomaterial, preserve the UIB-PPNe spectral features to varying degrees. The results are consistent with the panspermia theory of Hoyle and Wickramasinghe. UV-Visible studies were also conducted on all these materials. The main absorption feature was one close to 217.5 nm (2175 A). The normalized (averaged) spectrum of the whole sequence of biological materials and their degradation products absorption feature at 217.5 nm (2175 A) further support the contention that aromatic molecules in biological materials are responsible for the interstellar absorption feature centred at 217.5 nm.

The CubeSats platform has become a common basis for the development and flight of very small, low cost spacecraft-particularly amongst University groups. The smallest CubeSats are just 1 litre in volume-comprising a 10.

The Earth is surrounded by inoperative objects generated from past and current space missions. Because of the high orbital speed, even the impact with small fragments is a hazard to operational spacecraft as it could lead to the partial or complete loss of the mission. Therefore, it is important to assess the collision risk due to space debris considering small fragments, which are usually not included in space debris modelling because their large number would make simulations extremely complex. In this work, an analytical approach is developed to describe the evolution of debris clouds created by fragmentations in Low Earth Orbit. In contrast to traditional approaches, which follow the trajectory of individual fragments, with the proposed method the cloud behaviour is studied globally, so that the presence of small fragments can be modelled. This give a deeper insight into the dynamics of debris clouds and reduces the computational effort needed to estimate the consequences of a collision. A standard breakup model is used to describe the dispersion of the fragments in terms of characteristic length, area-to mass ratio and velocity. From the velocity distribution, the fragment spatial dispersion is derived. The cloud density is expressed by a continuous function that depends on the altitude and that is set as initial condition for the orbit propagation. Based on an analytical approach proposed in the literature for interplanetary dust and spacecraft swarms, the fragment cloud evolution in time is derived through the continuity equation, which is used to describe the variation of debris density considering the effect of atmospheric drag. The approach has been extended to express the cloud density as a function of multiple orbital parameters and to model additional perturbations such as the Earth’s oblateness. The method has been validated through the comparison with the traditional numerical propagation and then applied to study many breakup scenarios. The proposed approach proves to be flexible and able to study the collision risk coming from several breakup events and to evaluate the vulnerability of different targets. It is also applied to derive an index of the environmental criticality of spacecraft.

The complex dynamics which describe the motion of a spacecraft far from a massive planetary body or in a highly perturbed environment close to minor celestial objects raises two fundamental but related problems. The first is represented by the difficulty to accurately predict the evolution of its orbit even over short period when its initial conditions are known with a small degree of confidence. The second is given by the need for precise real time estimation of the trajectory when the spacecraft orbits near the asteroid’s surface to avoid impacting on it. The main example of the first problem is the perturbed four body problem for the Earth-Sun-Moon system. Earth-Sun Lagrangian Point Orbits (LPOs) are often selected for astrophysics and solar terrestrial missions while low cost missions aim at exploiting the so called Weak Stability Boundaries (WSB) to move at low propellant expense within the Earth sphere of influence. As current and future missions are planned to be placed on LPOs, it is a critical aspect to clear these regions at the end of operations to avoid damages to other spacecraft. For the second problem, we have a great number of asteroids and comets orbiting the inner solar system; they represent the so-called minor celestial objects which are very interesting for science since they preserve the remnants of the early formation of the planets and could shed light on the origins of life. At the same time they are very appealing for future commercial applications for the high content of precious ore. Among these celestial objects, the family of Near Earth Objects (NEOs) follows trajectories which lie close to, and sometimes cross, the Earth’s orbit. The impact hazard with the Earth has started to become considered as serious threat. Over the last three decades a number of missions have flown to and explored asteroids and comets, relying heavily on ground support with limited autonomy. In order to perform either asteroid’s exploration or collision hazard protection, autonomous navigation is needed, also to deal with the uncertain environment. Then the manipulation of asteroids’ orbit and attitude for deflection purposes is therefore required and an interesting problem to be studied. The aim of the research presented in this dissertation is to identify and develop methodologies for uncertainty propagation for spacecraft orbit and the application to orbit determination for complex nonlinear space mechanics problems, with particular care paid to the case of close proximity operations which are required when performing missions to minor celestial objects. The results are not limited only to this kind of problem but can be applied also to different scenarios. A first set of results focuses on the prediction of the trajectory evolution under initial condition uncertainties. The accuracy of the propagation of uncertainties is intimately related to the process of trajectory estimation, which relies on the use of the covariance matrix. The covariance matrix gives an idea of the dispersion of the spacecraft in terms of position and velocity. Different techniques to propagate the covariance matrix are used to predict the evolution of the trajectory when the initial conditions are known only to a certain degree of accuracy. They are compared under a highly nonlinear scenario where a spacecraft is injected into a disposal orbit towards an impacting trajectory with the Moon from a Lagrangian Point Orbit. A second set of results focuses on the identification of the estimation techniques applied to a single spacecraft. The estimation process performs well depending on the capability to propagate the covariance matrix and to incorporate the new information. A number of filtering techniques based on the Kalman and H∞ filters, employing different methods to handle the propagation of the covariance matrix, are presented and tested in typical nonlinear environments, i.e. a WSB transferan asteroid proximity, to draw precious information on their performance. The analyses demonstrate that only a hybrid Kalman- H∞ filter can enable the spacecraft to estimate its trajectory with a good balance between accuracy and computational costs. Then this method is applied to the navigation of spacecraft formation about a NEO showing that the navigation performance is significantly improved by sharing relative information among the spacecraft and the overall system is shown to be fault-tolerant. Finally the orbit’s and attitude manipulation of a small asteroid using a laser ablation system is analysed. An on-board state estimation and control algorithm is presented that simultaneously provides an optimal proximity control and control of the rotational motion of the asteroid. During the deflection, the proximity motion of the spacecraft is coupled with the orbital and rotational motion of the asteroid. The combination of the deflection acceleration, solar radiation pressure, and gravity field and plume impingement will force the spacecraft to drift away from the asteroid. In turn, a variation of the motion of the spacecraft produces a change in the modulus and direction of the deflection action which modifies the rotational and orbital motion of the asteroid.

In the past 50 years the scientists have been developing and analysing methods and new algorithms that optimise an interplanetary trajectory according to one or more objectives. Within this field, in 1963 Lawden derived, from Pontryagin's minimum principle, the so-called `primer vector theory'. The main goal of this thesis is to develop a theoretical understanding of Lawden's theory, getting an insight into the optimality of a trajectory when mid-course corrections need to be applied. The novelty of the research is represented by a different approach to the primer vector theory, which simplifies the structure of the problem.

No description available.