Christie Alisa Maddock

Design of a Formation of Solar Pumped Lasers for Asteroid Deflection

This paper presents the design of a multi-spacecraft system for the deflection of asteroids. Each spacecraft is equipped with a fibre laser and a solar concentrator. The laser induces the sublimation of a portion of the surface of the asteroid, and the resultant jet of gas and debris thrusts the asteroid off its natural course. The main idea is to have a formation of spacecraft flying in the proximity of the asteroid with all the spacecraft beaming to the same location to achieve the required deflection thrust. The paper presents the design of the formation orbits and the multi-objective optimisation of the formation in order to minimise the total mass in space and maximise the deflection of the asteroid. The paper demonstrates how significant deflections can be obtained with relatively small sized, easy-to-control spacecraft.

Comparison of single and multi-spacecraft configurations for NEA deflection by solar sublimation

Since the first Near Earth Object (NEO) defence system, Project Icarus, was published in 1967, where the driving factor was a very short anticipating time (i.e. time available to act on the asteroid), the Near Earth Asteroid (NEA) hazard outlook has changed drastically. In the current state of NEO research, long-term missions are becoming more realistic and as such, the options for low thrust systems are being investigated as a viable option for deviating the asteroid path. Surface ablation approaches have been previously proposed using several techniques such as lasers and nuclear explosives. The method presented here hinges on directing solar energy using mirrors onto a small area on the surface of the asteroid. This concentrated heat then sublimates the surface matter creating narrow but expanding jets of gas and dust that produce a low continuous thrust. This low thrust would then alter the orbit of the NEA by producing a change in velocity, similar to the effect of the ‘tail’ on a comet. This paper adds a new important trade-off to the problem: a comparison between a single structure and a multi-mirror system. The systems analysed include both single and dual mirror configurations, for both a single spacecraft and multiple spacecraft in formation. The criteria include ease of launch, reliability, flexibility in achieving the mission objective, in-space mass and a basic cost analysis. The concept, and potential benefits, of formation flying have long been known. The barrier however, has been the high level of complexity involved in the control of the individual spacecraft within the formation. Advances in control algorithms and associated technologies have opened the door to using spacecraft formations for specialized missions, such as small-body missions which operate in highly perturbed environments. This paper compares the complexities of deploying and operating a large rigid structure around asteroids, with the control of a swarm of smaller structures. Configurations are presented for different NEAs, time-in-advance, and achieved deflections.

Design of optimal spacecraft‐asteroid formations through a hybrid global optimization approach

Purpose – The purpose of this paper is to present a methodology and experimental results on using global optimization algorithms to determine the optimal orbit, based on the mission requirements, for a set of spacecraft flying in formation with an asteroid.Design/methodology/approach – A behavioral‐based hybrid global optimization approach is used to first characterize the solution space and find families of orbits that are a fixed distance away from the asteroid. The same optimization approach is then used to find the set of Pareto optimal solutions that minimize both the distance from the asteroid and the variation of the Sun‐spacecraft‐asteroid angle. Two sample missions to asteroids, representing constrained single and multi‐objective problems, were selected to test the applicability of using an in‐house hybrid stochastic‐deterministic global optimization algorithm (Evolutionary Programming and Interval Computation (EPIC)) to find optimal orbits for a spacecraft flying in formation with an orbit. The ...

Ascent trajectory optimisation for a single-stage-to-orbit vehicle with hybrid propulsion

This paper addresses the design of ascent trajectories for a hybrid-engine, high performance, unmanned, single-stage-to-orbit vehicle for payload deployment into low Earth orbit. A hybrid optimisation technique that couples a population-based, stochastic algorithm with a deterministic, gradient-based technique is used to maximize the nal vehicle mass in low Earth orbit after accounting for operational constraints on the dynamic pressure, Mach number and maximum axial and normal accelerations. The control search space is first explored by the population-based algorithm, which uses a single shooting method to evaluate the performance of candidate solutions. The resultant optimal control law and corresponding trajectory are then further refined by a direct collocation method based on finite elements in time. Two distinct operational phases, one using an air-breathing propulsion mode and the second using rocket propulsion, are considered. The presence of uncertainties in the atmospheric and vehicle aerodynamic models are considered in order to quantify their effect on the performance of the vehicle. Firstly, the deterministic optimal control law is re-integrated after introducing uncertainties into the models. The proximity of the final solutions to the target states are analysed statistically. A second analysis is then performed, aimed at determining the best performance of the vehicle when these uncertainties are included directly in the optimisation. The statistical analysis of the results obtained are summarized by an expectancy curve which represents the probable vehicle performance as a function of the uncertain system parameters. This analysis can be used during the preliminary phase of design to yield valuable insights into the robustness of the performance of the vehicle to uncertainties in the specification of its parameters.

Mirror formation control in the vicinity of an asteroid

Two strategies are presented for the positioning and control of a spacecraft formation designed to focus sunlight onto a point on the surface of asteroid, thereby sublimating the material and ejecting debris creating thrust. In the flrst approach, the formation is located at artiflcial equilibrium points around the asteroid and controlled using the force from the solar radiation pressure. The second approach determines the optimal periodic formation orbits, subject to the gravitational perturbations from the asteroid, the solar radiation pressure and the control acceleration derived from a control law.

Multi-objective optimal control of ascent trajectories for launch vehicles

This paper presents a novel approach to the solution of multi-objective optimal control problems. The proposed solution strategy is based on the integration of the Direct Finite Elements Transcription method, to transcribe dynamics and objectives, with a memetic strategy called Multi Agent Collaborative Search (MACS). The original multi-objective optimal control problem is reformulated as a bi-level nonlinear programming problem. In the outer level, handled by MACS, trial control vectors are generated and passed to the inner level, which enforces the solution feasibility. Solutions are then returned to the outer level to evaluate the feasibility of the corresponding objective functions, adding a penalty value in the case of infeasibility. An optional single level refinement is added to improve the ability of the scheme to converge to the Pareto front. The capabilities of the proposed approach will be demonstrated on the multi-objective optimisation of ascent trajectories of launch vehicles.

Vehicle and mission design of a future small payload launcher

This paper presents the conceptual design and performance analysis of a partially reusable space launch vehicle for small payloads. The system uses a multi-stage vehicle with rocket engines, with a reusable first stage capable of glided or powered flight, and expendable upper stage(s) to inject a 500 kg payload in different low Earth orbits. The space access vehicle is designed to be air-launched from a modified aircraft carrier. The aim of the system design is to develop a commercially viable launch system for near-term operation, thus emphasis is placed on the efficient use of high TRL technologies. The vehicle design are analysed using a multi-disciplinary design optimisation approach to evaluate the performance, operational capabilities and design trade-offs.

Return and abort trajectory optimisation for reusable launch vehicles

Among the future space access vehicles, the lifting body spaceplane is the most promising approach to prevent damage to both the launcher and the payload in case of loss of thrust. The glide performances of the vehicle allow the recovery in both nominal and abort cases. The approach presented is used in the investigation of the unpowered descent paths of a sample vehicle through trajectory optimisation. The vehicles downrange and crossrange limits are obtained for aborts in multiple flight conditions.

Global solution of multi-objective optimal control problems with multi agent collaborative search and direct finite elements transcription

This paper addresses the solution of optimal control problems with multiple and possibly conflicting objective functions. The solution strategy is based on the integration of Direct Finite Elements in Time (DFET) transcription into the Multi Agent Collaborative Search (MACS) framework. Multi Agent Collaborative Search is a memetic algorithm in which a population of agents performs a set of individual and social actions looking for the Pareto front. Direct Finite Elements in Time transcribe an optimal control problem into a constrained Non-linear Programming Problem (NLP) by collocating states and controls on spectral bases. MACS operates directly on the NLP problem and generates nearly-feasible trial solutions which are then submitted to a NLP solver. If the NLP solver converges to a feasible solution, an updated solution for the control parameters is returned to MACS, along with the corresponding value of the objective functions. Both the updated guess and the objective function values will be used by MACS to generate new trial solutions and converge, as uniformly as possible, to the Pareto front. To demonstrate the applicability of this strategy, the paper presents the solution of the multi-objective extensions of two well-known space related optimal control problems: the Goddard Rocket problem, and the maximum energy orbit rise problem.

Integrated 915 MHz dual-patch circularly polarized antenna for Suaineadh space web experiment

The University of Glasgow, UK, and KTH (Royal Institute of Technology), Sweden, have collaborated on the Suaineadh (Suaineadh is a Scots Gaelic word meaning “twisting”) experiment [1] to be onboard a REXUS [2] rocket. The experiment objectives are to deploy a space-web [3] [4] using centrifugal forces and to stabilize the web once full deployment has been achieved.