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This paper presents a multi-objective trajectory optimisation under uncertainty for the ascent of a two-stage, semi-reusable space launch system. Using Orbital Access’ Orbital 500- R launcher as a test case, robust multi-disciplinary design optimisation is used to analyse the trade-offs with both the vehicle and system design, and operation. An area of focus is on the predicted performance and its impact on the design and planned mission scenarios. The atmospheric model uncertainties are quantified for the atmospheric surrogate model and integrated into the optimal control solver MODHOC, extended by the addition of an unscented transformation to handle the uncertainties. A multi-objective optimisation under uncertainty is run, examining the Pareto-optimal sets for the ascent trajectory and vehicle design.

The paper will focus, on the navigation and station keeping of one or more spacecraft around the dynamic equivalent of the Lagrange points L4 and L5 of a binary asteroid system. In a recent work by one of the authors (thiry2016) it was demonstrated that this location can be conveniently used by a spacecraft equipped with a small scale laser to demonstrate laser ablation as a viable asteroid deflection technology. Furthermore, it was shown that the set of periodic solutions around L4 and L5 can be of particular interest for the exploration of a binary system and to extract and analyse surface and subsurface material. Binary asteroids are systems composed of two asteroids orbiting around their common barycentre. Some of them are part of the Near Earth Object population, like 2000 DP107, or 65803 Didymos (1996 GT),an Apollo asteroid discovered on April 11, 1996, that is the target of the DART mission. Navigating in the proximity of an asteroid is in itself a challenging task due to the complex dynamics induced by the irregular gravity field of the asteroid, the gravity of the Sun and solar radiation pressure. Even more challenging is to navigate a formation of spacecraft. Recent work by one of the authors demonstrated the possibility to autonomously navigate a formation of spacecraft in the proximity of a single asteroid with a distributed fault-tolerant autonomous system (vetrisano2016b). The complexity increases even further in the case of a binary due the interaction between the primary and the secondary. However, a combination of optical and laser ranging measurement was shown to be sufficient to maintain one or more spacecraft at L4 of the Didymos system (torre2017). Different from Torre et al. 2017 (torre2017) this paper proposes an approach using optical navigation only and exploiting the knowledge of the dynamics of the secondary asteroid around the primary. The dynamical model includes the gravity of the two asteroids, the gravity of the Sun and solar radiation pressure. Each spacecraft is equipped with only a camera, a good attitude determination and control system and an inter satellite link that allows sharing information with other spacecraft in the formation and provides a complete spacecraft-to-spacecraft position vector. Two scenarios will be considered in this study: one in which one single spacecraft uses only information from the camera on both the primary and the secondary asteroid, the other in which a small formation of spacecraft autonomously navigates with cameras and inter-satellite link only. Measurements and dynamics are processed in a form of Unscented H-infinity Filter that can effectively estimate the state of the system and allows maintaining the formation at L4.

This paper presents a preliminary analysis of the de-orbiting and re-entry dynamics of space objects with a large area to mass ratio. Two different classes of objects are considered: fragments of satellites (like solar panels and pieces of thermal blankets) and complete nano-satellites with passive de-orbiting devices like a drag sail. Different sources of uncertainty are considered including atmospheric density variability with latitude and solar cycles, aerodynamic properties of the object, light pressure, initial conditions. The coupling of the uncertainty in the aerodynamic forces and attitude motion is investigated to understand if low fidelity three degrees of freedom models can be used in place of more expensive high fidelity models to make predictions on the re-entry time. Modern uncertainty propagation and quantification tools are used to assess the effect of uncertainty on the re-entry time and study the dependency on a number of key parameters.

A deployable space web is a flexible structure that can act as a lightweight platform for construction of large structures in space. In order to save space and energy for small deployable structure ...

Existing active debris removal (ADR) methods that require physical contact with the target have applicability limitations depending on the maximum angular momentum that can be absorbed. Therefore, a de-tumbling phase prior to the capturing phase may be necessary. The aim of this article is to present the on-going work on the control module of a contactless de-tumbling subsystem based on eddy currents ('Eddy Brake'). This research is being carried out in the framework of the Agora mission (Active Grabbing & Orbital Removal of Ariane), which employs a robotic spacecraft concept to demonstrate technologies to autonomously de-tumble, capture and de-orbit an Ariane rocket body. The article first presents the 'Eddy Brake' method and the Guidance, Navigation and Control (GNC) architecture of a chaser spacecraft. Furthermore, the linear and rotational dynamics based on the Magnetic Tensor Theory (MTT) are explained. Then, a control strategy is presented to keep a constant relative distance between the two objects and a suitable relative pointing of the coil towards the target object. A simplified analytical solution for the control of the two objects in the 2D problem is presented and the stability of the system in the vicinity of a stable asymptotic state is analysed. Finally, two case studies are presented on the Ariane-4 H10 and Ariane-5 EPS upper stages

We present work towards the development of a new tool, FOSTRAD (Free Open Source Tool for Re-entry of Asteroids and Debris) that will use a combination of the spacecraftand object-oriented approaches and will incorporate uncertainty quantification. The current work presents progress towards incorporating uncertainty quantification by performing aeroand aerothermo-dynamic sensitivity analyses for a sphere, cube, and a cylinder with respect to the uncertain atmospheric and object parameters. The sensitivity analysis is performed using high fidelity tools such as the Direct Simulation Monte Carlo (DSMC) and the Computational Fluid Dynamics (CFD) methods. In FOSTRAD, aerodynamic computations in the continuum and free molecular regimes are performed using Modified Newtonian Theory and the free molecular analytical models of Schaaf and Chambre, respectively. For the rarefied transition regime, recently developed sigmoid bridging function is used for better tracking of the data. Analytical aerothermodynamic relations for strong shocks exist only for blunt objects that have a finite radius of curvature at the nose (like a sphere), but not for sharp-edged objects such as the cube and cylinder. We use the results from the sensitivity analyses to develop mathematical relations for the variation of aerothermodynamic properties for cube and cylinder. We also develop mathematical relations for the heat flux distribution on the surface of these objects.

This paper presents a low-cost, low mass, mission design to successfully intercept and deflect a small and faint, 4 m in diameter asteroid. Intended to be launched after 2025, the laser-ablating mission, LightTouch2 will be used to deflect the orbit of the asteroid by at least 1 m/s. This will be achieved with a total mission lifetime of less than three years. Analysis includes the initial approach of the spacecraft, the operations of the laser at an optimal spacecraft-to-asteroid distance of 50 m and the relative orbit of the spacecraft that flies in formation with the asteroid. Analysis includes line-of-sight measurements with radiometric tracking from ground station to improve the trajectory estimate and observability of the spacecraft, collision avoidance and mapping strategies. The spacecraft will also need optimal discrete control. This is achieved by impulse-bit manoeuvres used to account for the perturbations caused by the resultant thrust on the asteroid, plume impingement, laser recoil and solar radiation pressure. The spacecraft controls its trajectory within a 1 m box from the reference trajectory to enable the laser to optimally focussing the laser beam. The proposed approach uses an unscented Kalman filter to estimate the relative spacecraft-asteroid position, velocity and perturbative acceleration.

The paper presents a method for the solution of optimal control problems based on a direct finite element transcription of the differential constraints, the objective function and the algebraic constraints. The finite elements are generated on spectral basis and the proposed finite elements formulation allows for the selection of different basis for state and control. The paper briefly reviews the transcription of optimal control problems with direct final elements in time generated on spectral basis. Then, it derives the necessary conditions for optimality and an estimation of the adjoint variables. A mesh grid adaptation scheme is then proposed together with a theoretical estimation of the accuracy of the transcription scheme. Experimental results demonstrate the validity of the theoretical derivations and prove that the properties of the transcription scheme depend on the selection of the collocation points. The paper concludes with a preliminary discussion on the analogies between the direct finite elements transcription method and other methods with which it shares some commonalities.

A case study of a potential Earth-Mars interplanetary communications relay, designed to ensure continuous communications, is detailed. The relay makes use of orbits based on artificial equilibrium points via the application of continuous low thrust, which allows a spacecraft to hover above the orbital plane of Mars and thus ensure communications when the planet is occulted with respect to the Earth. The artificial equilibria of two different low-thrust propulsion technologies are considered: solar electric propulsion, and a solar sail/solar electric propulsion hybrid. In the latter case it is shown that the combination of sail and solar electric propulsion may prove advantageous, but only under specific circumstances of the relay architecture suggested. The study takes into account factors such as the spacecraft's power requirements and communications band utilized to determine the mission and system architecture. A detailed contingency analysis is considered for recovering the relay after increasing periods of spacecraft motor failure, and combined with a consideration for how best to deploy the relay spacecraft to maximise propellant reserves and mission duration.

An autonomous control system of a rover for planetary exploration can be subdivided in two distinct parts, the rst devoted to planning and scheduling and the second to the plan execution. In this paper the concept of a hybrid intelligent system composed by three layers will be presented, focusing on the plan execution part which is embodied by a Behavioral and a Re exive layer. The plan generated by the Planning layer is composed by a sequence of logical actions that can not be directly executed by the rover hardware. The Behavioral layer translates these actions in executable actions through a Fuzzy subsystem, and prepares the input for the Re exive layer. This layer is in charge of the low level control of the rover hardware which is performed by an Arti cial Neural Network. In this paper, the whole Execution system will be presented, focusing on the path following problem and reporting the results of the rst tests executed on a real prototype.

This paper introduces the mission analysis and design of the Hevelius - Lunar Microsatellite Mission. The objective of the mission is to place at least three landers on the dark side of the Moon, to perform some scientific experiments. A microsatellite orbiter is required to support the net lander as data-relay to the Earth. Moreover, another spacecraft, a carrier, has been designed in order to bring landers to the surface of the Moon, to map the landing site and to measure the gravitational field. The Hevelius mission analysis has been driven by the need to design lowcost and low-mass space missions. Since the relay satellite must continuously see the dark side of the Moon, an operative Halo orbit around the second Lagrangian point has been designed. Three different ways have been followed to determine the optimal Halo orbit. Optimal low-cost transfers to the Halo have been designed exploiting the invariant manifolds of the Earth-Moon L1 point while a Belbruno's WSB transfer to a frozen orbit around the Moon has been chosen for the carrier. The mission analysis process has been completed with a perturbations and eclipses analysis of the final operative orbits.

In this paper trajectory optimisation for the Hevelius mission is presented. The Hevelius-Lunar Microsatellite Mission - is a multilander mission to the dark side of the Moon, supported by a relay microsatellite, orbiting on a Halo orbit around L2. Three landers, with miniaturized payloads, are transported by a carrier from a LEO to the surface of the Moon, where they perform a semi-hard landing with an airbag system. This paper will present the trajectory optimisation process, focusing, in particular, on the approach employed for Δv manoeuvre optimization. An introduction to the existing methods for trajectory optimization will be presented, subsequently it will be described how these methods have been exploited and originally combined in the Hevelius mission analysis and design.

This paper describes the mission analysis and design of the 'Hevelius - Lunar Microsatellite Mission'. The main goal of the overall mission is to place a net-lander on the far side of the Moon to perform some scientific experiments. Two different satellites have been designed to achieve this objective: a microsatellite orbiter to support the net-lander and a carrier spacecraft to transport the net-lander. An L2 Halo orbit has been selected for the orbiter in order to have a constant communication link between the landers and the Earth. The invariant manifolds of the Earth-Moon system have been used to design a low cost transfer trajectory to the L2 Halo orbit. Prior to the beginning of landing operations the carrier is parked into a frozen orbit after a WSB transfer. Finally the descent and landing phases have been designed in order to accomplish the final goals. The whole mission analysis and design process has been driven by the need for a low cost and low risk mission.

In this paper a preliminary analysis of a wide range of mission opportunities, offered by either aerogravity assist or gravity assist manoeuvres, has been carried out. After an accurate validation of aerogravity assist traditional analytical models, according to several different criteria an extensive global search for optimal trajectories has been performed for highenergy missions, resorting to gravity and aerogravity manoeuvres. To this aim, the new preliminary analysis tool PAMSIT, based on some simplified hypotheses, has been developed. This is capable of efficiently and exhaustively exploring the solutions space for this particular problem, considering the feasibility of a trajectory from both the orbital energy and phasing points of view. Then, all found solutions have been classified according to launch and arrival velocities, time of flight and planetary encounters. A comparison between the opportunities offered by gravity manoeuvres and aerogravity manoeuvres will be presented showing the advantages of the latter in all analysed cases. In particular some interesting options for missions to Jupiter and Neptune will be presented.

The paper analyses the use of a Nuclear Electric Propulsion systems for missions to the outer part of the solar system and points out the possibility to deploy a probe while performing a gravity assist manoeuvre in the vicinity of Jupiter or to insert the spacecraft in a highly elliptical orbit about Pluto. The design of the trajectories has been performed with a direct transcription method by finite elements in time. As the problem presents quite a number of possible solutions dependent on launch window, transfer time and the use of gravity assist manoeuvres, a global optimization strategy has been used to procure sets of promising initial guesses. These initial guesses have been subsequently optimized using direct transcription and NLP. INTRODUCTION Currently most of our knowledge about Pluto and its moon Charon comes from indirect clues. No spacecraft has ever visited either of them, and from the Earth (or its proximity) its angular size is close about resolving limit of the most capable ground and space-based observatories. However, there is unanimity on the scientific interest of the Pluto-Charon system. Therefore a mission to the Pluto-Charon system and eventually to a Kuiper Belt Object (KBO) will significantly increase our knowledge of the formation and evolution of the Solar System as well as the origin of volatiles and organic molecules that enabled the appearance of life on our own planet. The mission therefore has a strong exobiological interest, which could be increased exponentially by adding new elements like a Europa or Titan microprobe deployment on route to the final destination, taking advantage of the opportunities provided by the gravity assist at the giant planets. In this paper some possible mission scenarii for a mission to Pluto and the Kuiper belt are proposed: these include the utilization of advanced propulsion systems (nuclear electric propulsion) and power technologies and the possibility to deploy a probe while a gravity assist maneuver in the vicinity of Jupiter is performed. If the Jupiter option is selected the possibility of a swing-by of one of the moons is investigated. In particular a swing-by of Ganymede can be performed to brake the probe while the main spacecraft continues on its way to Pluto. In addition the possibility of advanced missions using chemical propulsion ad gravity or aero-gravity assist maneuvers have been studied. Analyses available in the literature propose to carry out a very quick flyby of Pluto and Charon with a large relative velocity, thus enabling a limited science return. Another option is therefore to study the possibility to alter the mission analysis concept in such a way that the flyby velocity can be reduced. The design of the NEP trajectories has been performed with a direct transcription method by finite elements in time. However the problem presents quite e number of possible solutions dependent on launch window, transfer time and combination of planetary encounters, therefore in order to find favourable launch windows and the optimal sequence of swing-bys a global optimization strategy has been used to procure sets of promising initial guesses. Then, these initial guesses have been optimized using direct transcription and NLP. TECHNOLOGICAL AND SCIENTIFIC OBJECTIVES A mission to the Pluto-Charon system would significantly increase our knowledge of the formation and evolution of the Solar System. Of particular interest in the Pluto-Charon system is the atmospheric transfer of methane between these bodies and their compositional difference. 54th International Astronautical Congress of the International Astronautical Federation, the International Academy of Astronautics, and the International Institute of Space Law 29 September 3 October 2003, Bremen, Germany IAC-03-A.P.14 Copyright © 2003 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Released to IAF/IAA/AIAA to publish in all forms. 2 Some of the driving scientific objectives of such a mission would include: • Surface chemical composition • Surface morphology • Atmospheric chemistry • Gravimetry A strawman payload to achieve the scientific objectives within the allocated mass limits include: • Imaging X-ray Spectrometer • Wide / Narrow Field Imager • IR-Spectrometer • Radio Science Experiment The available payload mass obviously depends on the mission scenario. However it is reasonable to assume that even a limited payload mass value (e.g. 20 kg), would be sufficient to meet a significant part of the scientific goals. This should be achievable even with current technology and considering the heritage of other planetary exploration missions like SMART-1. The scientific return from a Pluto mission would increase tremendously if either the fly-by would occur at a small relative velocity or if the spacecraft could go into orbit around the PlutoCharon system. This would not only increase the coverage but also the accuracy of the scientific investigations because of the low signal to noise ratio for certain instruments partly due to the large distance from the sun. From a technological point of view, inserting a probe in orbit around such a distant planet at a reasonable propellant expenditure, poses considerable issues both in terms of propulsion and power. Furthermore the long travel results in a long waiting time before any result can be obtained. Since access to distant targets as Pluto can be effectively achieved through a swing-by of Jupiter, the delivery of a probe in the jovian system will significantly improve the scientific return of the mission offering some intermediate results while waiting for the analysis of the Pluto-Charon system. MISSION ANALYSIS Two missions have been studied: a double probe to Jupiter and Pluto, a small probe to Europa. In both cases the spacecrafts are equipped with nuclear electric propulsion (NEP) engines and will perform a number of gravity assist manoeuvres to reach the final destination. Initial Guess Generation The aim is to find an optimal sequence of transfers from the Earth to Pluto or from Earth to Jupiter passing by a predefined number of intermediate stops (actually swingbys). Even though the propulsion system is electric and not chemical, the trajectory, which minimizes the overall cost in terms of ∆v, is regarded as optimal for both means of propulsion since a further optimization with a better model of electric propulsion will be performed using DITAN (a software tool for the design of gravity assist lowthrust trajectories, developed by Politecnico di Milano under ESA contract). Each arc connecting two subsequent bodies has a deep space ∆v manoeuvre at an unknown point in time and space, each swingby is modelled collapsing the sphere of influence in one single point in space with radius rp linking the transfer arcs ri before and ro after the swing by, thus the following relation must hold

In this paper an algorithm for orbit determination in deep space using only measurements of attitude and of line of sight of known celestial bodies is presented. Measurements are then processed using two different filters reconstructing part or the entire state vector: a batch least squares filter and a combination of batch and Kalman filter. The strategy and the effectiveness of the two filters are compared on two sample cases of interplanetary trajectory toward Mercury (inner part of the solar system) and toward Jupiter (outer part of the solar system). Introduction One of the driving issues for near future missions is autonomy: in fact autonomy reduces the need for support from ground, so reducing ground station development and maintenance, cost, man effort, reaction time to unexpected situations, and, avoiding dangerous delays in control loops, it is an essential concept when dealing with cheap and multiple satellites and very long missions, where long time is required for transfer. Historically navigation of deep-space satellites has involved the use of radio data types for determining trajectory of the spacecraft and then predicting its course in the future. These radio data types include Doppler data to determine the component of the line of sight velocity and ranging to determine the line of ✫ Research Fellow. ✞ Full Professor Graduated Student Copyright 2002 by the authors. Published by the American Institute of Aeronautics and Astronautics,Inc., with permission. Released to AAS/AIAA to publish in all forms. sight position. Doppler measurements are acquired on Earth in S and X band to determine the velocity of the spacecraft while position can be computed measuring the transit time of a radio signal from the ground station to the spacecraft. Therefore traditionally navigation is completely demanded to operations on Earth, with a consequent cost in terms of human resources and to the fundamental issues related to the delay in receiving and sending signal toward and from the spacecraft or related to possible occultation or failure that may jeopardise the entire mission. For some missions, during approach to target bodies, optical data taken with an onboard camera were used to measure the target-relative position of the spacecraft, providing additional information. This methodology of combining both radio and optical data has worked very well in the past (Voyager and Galileo missions). This can be regarded as an early attempt of onboard autonomy especially for Galileo. The same concept for image processing, experimented on Galileo, was implemented onboard the first mission designed to test autonomous navigation systems: Deep Space 1. In fact. for purposes of developing autonomously navigating spacecraft, the software needed to process radio navigation data is very complex. In addition a radio system onboard cannot be considered really autonomous because an uplink from the ground is needed. By its nature, optical data are not as precise as radio systems, but have the distinct advantage of being self-contained onboard the spacecraft. DS-1 system was enriched with a completely autonomous centerfinding procedure of the objects used to determine its position and, as a consequence, the position of the spacecraft using an ephemeris database. In this paper a strategy for orbit determination based on measurements of LOS (Line Of Sight) of known celestial bodies is presented. Two algorithms