Physical Investigation of the Potentially Hazardous Asteroid (144898) 2004 VD17

In this paper we present the observational campaign carried out at ESO NTT and VLT in April and May 2006 to investigate the nature and the structure of the Near Earth Object (144898) 2004 VD17. In spite of a great quantity of dynamical information, according to which it will have a close approach with the Earth in the next century, the physical properties of this asteroid are largely unknown. We performed visible and near--infrared photometry and spectroscopy, as well as polarimetric observations. Polarimetric and spectroscopic data allowed us to classify 2004 VD17 as an E-type asteroid. A good agreement was also found with the spectrum of the aubrite meteorite Mayo Belwa. On the basis of the polarimetric albedo (p_v=0.45) and of photometric data, we estimated a diameter of about 320 m and a rotational period of about 2 hours. The analysis of the results obtained by our complete survey have shown that (144898) 2004 VD17 is a peculiar NEO, since it is close to the breakup limits for fast rotator asteroids, as defined by Pravec and Harris (2000). These results suggest that a more robust structure must be expected, as a fractured monolith or a rubble pile in a "strength regime" (Holsapple 2002).Comment: 32 pages, 7 figure, paper accepted for publication in Icaru

Autonomous navigation of a spacecraft formation in the proximity of an asteroid

This paper presents a multi-sensor navigation approach to allow a formation of spacecraft to autonomously navigate in the proximity of a Near Earth Asteroid. Multiple measurements collected by on-board cameras, attitude sensors and LIDAR are used to estimate the state of each spacecraft with respect to the asteroid. Inter-spacecraft position measurements are then combined with spacecraft-to-asteroid position measurements to improve accuracy. The paper analyses the use of different filtering techniques to estimate the state of a 4-spacecraft formation with respect to the asteroid. Different combinations of measurements are constructed to evaluate the improvement in navigation performance offered by the data fusion of the measurements gathered by the four spacecraft. Moreover the robustness of the navigation system is tested against the occurrence of failures. Results show that the navigation performance is significantly improved by adding the inter-spacecraft position measurements. Finally, an asteroid orbit determination method is proposed that combines asteroid's line of sight measurements from multiple spacecraft and Sun Doppler shift sensor with spacecraft-to-ground tracking data. Different approach configurations are evaluated for a 2-spacecraft formation and it is shown that the integrated use of spacecraft-to-asteroid and ground-to-spacecraft measurements provides an effective way to improve the ephemerides of the asteroid

Autonomous navigation of spacecraft formations for asteroid exploration

This paper presents an autonomous multi-sensor navigation approach for a formation of spacecraft flying in the proximity of a near Earth asteroid.Each spacecraft embarks a different combination of high resolution cameras, attitude sensors and LIDAR to estimate the stateof each spacecraft in the formation. The work investigates the combination of measurements coming from multiple heterogeneous sensors and nonlinear sequential filtering technique to enable a formation to autonomously navigate in the proximity of asteroids. This work is divided into two parts. Firstly, each spacecraft employs an Unscented Kalman Filter to data fuse multi-sensor measurements of the relative position of the spacecraft with respect to the asteroid possibly combined with measurements of the relative position of the spacecraft within the formation, thus determining position and velocity of each member. Secondly, the combination of the autonomous orbit determination with absolute measurements is considered. Absolute measurements include range and range rate measurements from the ground station and pseudo range rate measurements from on board Sun Doppler shift sensor. The combination of the two sets of measurements and state estimations from on-board and ground provides an interesting mean to accurately determine the orbit of asteroids

Optimal trajectory design for interception and deflection of Near Earth Objects

Many asteroids and comets orbit the inner solar system; among them Near Earth Objects (NEOs) are those celestial bodies for which the orbit lies close, and sometimes crosses, the Earth's orbit. Over the last decades the impact hazard they pose to the Earth has generated heated discussions on the required measures to react to such a scenario. The aim of the research presented in this dissertation is to develop methodologies for the trajectory design of interception and deflection missions to Near Earth Objects. The displacement, following a deflection manoeuvre, of the asteroid at the minimum orbit intersection distance with the Earth is expressed by means of a simple and general formulation, which exploits the relative motion equations and Gauss' equations. The variation of the orbital elements achieved by any impulsive or low-thrust action on the threatening body is derived through a semi-analytical approach, whose accuracy is extensively shown. This formulation allows the analysis of the optimal direction of the deflection manoeuvre to maximise the achievable deviation. The search for optimal opportunities for mitigation missions is done through a global optimisation approach. The transfer trajectory, modelled through preliminary design techniques, is integrated with the deflection model. In this way, the mission planning can be performed by optimising different contrasting criteria, such as the mass at launch, the warning time, and the total deflection. A set of Pareto fronts is computed for different deflection strategies and considering various asteroid mitigation scenarios. Each Pareto set represents a number of mission opportunities, over a wide domain of launch windows and design parameters. A first set of results focuses on impulsive deflection missions, to a selected group of potentially hazardous asteroids; the analysis shows that the ideal optimal direction of the deflection manoeuvre cannot always be achieved when the transfer trajectory is integrated with the deflection phase. A second set of results includes solutions for the deviation of some selected NEOs by means of a solar collector strategy. The semi-analytical formulation derived allows the reduction of the computational time, hence the generation of a large number of solutions. Moreover, sets of Pareto fronts for asteroid mitigation are computed through the more feasible deflection schemes proposed in literature: kinetic impactor, nuclear interceptor, mass driver device, low-thrust attached propulsion, solar collector, and gravity tug. A dominance criterion is used to perform a comparative assessment of these mitigation strategies, while also considering the required technological development through a technology readiness factor. The global search of solutions through a multi-criteria optimisation approach represents the first stage of the mission planning, in which preliminary design techniques are used for the trajectory model. At a second stage, a selected number of trajectories can be optimised, using a refined model of the dynamics. For this purpose, the use of Differential Dynamic Programming (DDP) is investigated for the solution of the optimal control problem associated to the design of low-thrust trajectories. The stage-wise approach of DDP is exploited to integrate an adaptive step discretisation scheme within the optimisation process. The discretisation mesh is adjusted at each iteration, to assure high accuracy of the solution trajectory and hence fully exploit the dynamics of the problem within the optimisation process. The feedback nature of the control law is preserved, through a particular interpolation technique that improves the robustness against some approximation errors. The modified DDP-method is presented and applied to the design of transfer trajectories to the fly-by or rendezvous of NEOs, including the escape phase at the Earth. The DDP approach allows the optimisation of the trajectory as a whole, without recurring to the patched conic approach. The results show how the proposed method is capable of fully exploiting the multi-body dynamics of the problem; in fact, in one of the study cases, a fly-by of the Earth is scheduled, which was not included in the first guess solution

PI -- Terminal Planetary Defense

We present a practical and effective method of planetary defense that allows for extremely short mitigation time scales. The method uses an array of small hypervelocity kinetic penetrators that pulverize and disassemble an asteroid or small comet. This mitigates the threat using the Earth's atmosphere to dissipate the energy in the fragment cloud. The system allows a planetary defense solution using existing technologies. This approach will work in extended time scale modes where there is a large warning time, as well as in short interdiction time scenarios with intercepts of minutes to days before impact. In longer time intercept scenarios, the disassembled asteroid fragments largely miss the Earth. In short intercept scenarios, the asteroid fragments of maximum $\sim$10-meter diameter allow the Earth's atmosphere to act as a "beam dump" where the fragments burn up and/or air burst, with the primary channel of energy going into spatially and temporally de-correlated shock waves. It is the de-correlated blast waves that are the key to why PI works so well. The effectiveness of the approach depends on the intercept time and size of the asteroid, but allows for effective defense against asteroids in the 20-1000m diameter class and could virtually eliminate the threat of mass destruction posed by these threats with very short warning times, though longer warning is always preferred. A 20m diameter asteroid ($\sim$0.5Mt, similar to Chelyabinsk) can be mitigated with a 100s prior to impact intercept with a 10m/s disruption. With ~1m/s internal disruption, a 5 hours prior to impact intercept of a 50m diameter asteroid ($\sim$10Mt yield, similar to Tunguska), a 1 day prior to impact intercept of 100m diameter asteroid ($\sim$100Mt yield), or a 10-20 day prior to impact intercept of Apophis ($\sim$370m diameter, $\sim$4Gt yield) would mitigate these threats.Comment: 174 pages, 130 figures. Published in Advances in Space Research (ASR) 10-22; https://www.sciencedirect.com/science/article/pii/S027311772200939

Conceptual Design of a Flight Validation Mission for a Hypervelocity Asteroid Intercept Vehicle

Near-Earth Objects (NEOs) are asteroids and comets whose orbits approach or cross Earth s orbit. NEOs have collided with our planet in the past, sometimes to devastating effect, and continue to do so today. Collisions with NEOs large enough to do significant damage to the ground are fortunately infrequent, but such events can occur at any time and we therefore need to develop and validate the techniques and technologies necessary to prevent the Earth impact of an incoming NEO. In this paper we provide background on the hazard posed to Earth by NEOs and present the results of a recent study performed by the NASA/Goddard Space Flight Center s Mission Design Lab (MDL) in collaboration with Iowa State University s Asteroid Deflection Research Center (ADRC) to design a flight validation mission for a Hypervelocity Asteroid Intercept Vehicle (HAIV) as part of a Phase 2 NASA Innovative Advanced Concepts (NIAC) research project. The HAIV is a two-body vehicle consisting of a leading kinetic impactor and trailing follower carrying a Nuclear Explosive Device (NED) payload. The HAIV detonates the NED inside the crater in the NEO s surface created by the lead kinetic impactor portion of the vehicle, effecting a powerful subsurface detonation to disrupt the NEO. For the flight validation mission, only a simple mass proxy for the NED is carried in the HAIV. Ongoing and future research topics are discussed following the presentation of the detailed flight validation mission design results produced in the MDL

Optimal trajectory design for interception and deflection of Near Earth Objects

Many asteroids and comets orbit the inner solar system; among them Near Earth Objects (NEOs) are those celestial bodies for which the orbit lies close, and sometimes crosses, the Earth’s orbit. Over the last decades the impact hazard they pose to the Earth has generated heated discussions on the required measures to react to such a scenario. The aim of the research presented in this dissertation is to develop methodologies for the trajectory design of interception and deflection missions to Near Earth Objects. The displacement, following a deflection manoeuvre, of the asteroid at the minimum orbit intersection distance with the Earth is expressed by means of a simple and general formulation, which exploits the relative motion equations and Gauss’ equations. The variation of the orbital elements achieved by any impulsive or low-thrust action on the threatening body is derived through a semi-analytical approach, whose accuracy is extensively shown. This formulation allows the analysis of the optimal direction of the deflection manoeuvre to maximise the achievable deviation. The search for optimal opportunities for mitigation missions is done through a global optimisation approach. The transfer trajectory, modelled through preliminary design techniques, is integrated with the deflection model. In this way, the mission planning can be performed by optimising different contrasting criteria, such as the mass at launch, the warning time, and the total deflection. A set of Pareto fronts is computed for different deflection strategies and considering various asteroid mitigation scenarios. Each Pareto set represents a number of mission opportunities, over a wide domain of launch windows and design parameters. A first set of results focuses on impulsive deflection missions, to a selected group of potentially hazardous asteroids; the analysis shows that the ideal optimal direction of the deflection manoeuvre cannot always be achieved when the transfer trajectory is integrated with the deflection phase. A second set of results includes solutions for the deviation of some selected NEOs by means of a solar collector strategy. The semi-analytical formulation derived allows the reduction of the computational time, hence the generation of a large number of solutions. Moreover, sets of Pareto fronts for asteroid mitigation are computed through the more feasible deflection schemes proposed in literature: kinetic impactor, nuclear interceptor, mass driver device, low-thrust attached propulsion, solar collector, and gravity tug. A dominance criterion is used to perform a comparative assessment of these mitigation strategies, while also considering the required technological development through a technology readiness factor. The global search of solutions through a multi-criteria optimisation approach represents the first stage of the mission planning, in which preliminary design techniques are used for the trajectory model. At a second stage, a selected number of trajectories can be optimised, using a refined model of the dynamics. For this purpose, the use of Differential Dynamic Programming (DDP) is investigated for the solution of the optimal control problem associated to the design of low-thrust trajectories. The stage-wise approach of DDP is exploited to integrate an adaptive step discretisation scheme within the optimisation process. The discretisation mesh is adjusted at each iteration, to assure high accuracy of the solution trajectory and hence fully exploit the dynamics of the problem within the optimisation process. The feedback nature of the control law is preserved, through a particular interpolation technique that improves the robustness against some approximation errors. The modified DDP-method is presented and applied to the design of transfer trajectories to the fly-by or rendezvous of NEOs, including the escape phase at the Earth. The DDP approach allows the optimisation of the trajectory as a whole, without recurring to the patched conic approach. The results show how the proposed method is capable of fully exploiting the multi-body dynamics of the problem; in fact, in one of the study cases, a fly-by of the Earth is scheduled, which was not included in the first guess solution

Asteroid hazard mitigation: deflection models and mission analysis

Small celestial bodies such as Near Earth Objects (NEOs) have become a common subject of study because of their importance in uncovering the mysteries of the composition, formation and evolution of the solar system. Among all asteroids, NEOs have stepped into prominence because of two important aspects: they are among the easiest celestial bodies to reach from Earth, in some cases with less demanding trajectories than a simple Earth-Moon trajectory and, even more meaningful, they may pose a threat to our planet. The purpose of this thesis is to provide a comprehensive insight into the asteroid hazard problem and particularly to its mitigation. Six different concepts are fully described; specifically models for nuclear interceptor, kinetic impactor, low-thrust propulsion, mass driver, solar collector and gravity tug are developed and their efficiency is assessed for a complete set of different types of hazardous celestial objects. A multi-criteria optimization is then used to construct a set of Pareto-optimal asteroid deflection missions. The Pareto-optimality is here achieved not only by maximizing the deflection of the threatening object, but also by minimizing the total mass of the deflection mission at launch and the warning time required to deflect the asteroid. A dominance criterion is also defined and used to compare all the Pareto sets for all the various mitigation strategies. The Technology Readiness Level for each strategy is also accounted for in the comparison. Finally, this thesis will also show that impulsive deflection methods may easily catastrophically disrupt an asteroid if the required energy for a deflection reaches a certain limit threshold. A statistical model is presented to approximate both the number and size of the fragments and their initial dispersion of velocity and then used to assess the potential risk to Earth posed by the fragmentation of an asteroid as a possible outcome of a hazard mitigation mission

Optimal trajectory design for interception and deflection of Near Earth Objects

Many asteroids and comets orbit the inner solar system; among them Near Earth Objects (NEOs) are those celestial bodies for which the orbit lies close, and sometimes crosses, the Earth’s orbit. Over the last decades the impact hazard they pose to the Earth has generated heated discussions on the required measures to react to such a scenario. The aim of the research presented in this dissertation is to develop methodologies for the trajectory design of interception and deflection missions to Near Earth Objects. The displacement, following a deflection manoeuvre, of the asteroid at the minimum orbit intersection distance with the Earth is expressed by means of a simple and general formulation, which exploits the relative motion equations and Gauss’ equations. The variation of the orbital elements achieved by any impulsive or low-thrust action on the threatening body is derived through a semi-analytical approach, whose accuracy is extensively shown. This formulation allows the analysis of the optimal direction of the deflection manoeuvre to maximise the achievable deviation. The search for optimal opportunities for mitigation missions is done through a global optimisation approach. The transfer trajectory, modelled through preliminary design techniques, is integrated with the deflection model. In this way, the mission planning can be performed by optimising different contrasting criteria, such as the mass at launch, the warning time, and the total deflection. A set of Pareto fronts is computed for different deflection strategies and considering various asteroid mitigation scenarios. Each Pareto set represents a number of mission opportunities, over a wide domain of launch windows and design parameters. A first set of results focuses on impulsive deflection missions, to a selected group of potentially hazardous asteroids; the analysis shows that the ideal optimal direction of the deflection manoeuvre cannot always be achieved when the transfer trajectory is integrated with the deflection phase. A second set of results includes solutions for the deviation of some selected NEOs by means of a solar collector strategy. The semi-analytical formulation derived allows the reduction of the computational time, hence the generation of a large number of solutions. Moreover, sets of Pareto fronts for asteroid mitigation are computed through the more feasible deflection schemes proposed in literature: kinetic impactor, nuclear interceptor, mass driver device, low-thrust attached propulsion, solar collector, and gravity tug. A dominance criterion is used to perform a comparative assessment of these mitigation strategies, while also considering the required technological development through a technology readiness factor. The global search of solutions through a multi-criteria optimisation approach represents the first stage of the mission planning, in which preliminary design techniques are used for the trajectory model. At a second stage, a selected number of trajectories can be optimised, using a refined model of the dynamics. For this purpose, the use of Differential Dynamic Programming (DDP) is investigated for the solution of the optimal control problem associated to the design of low-thrust trajectories. The stage-wise approach of DDP is exploited to integrate an adaptive step discretisation scheme within the optimisation process. The discretisation mesh is adjusted at each iteration, to assure high accuracy of the solution trajectory and hence fully exploit the dynamics of the problem within the optimisation process. The feedback nature of the control law is preserved, through a particular interpolation technique that improves the robustness against some approximation errors. The modified DDP-method is presented and applied to the design of transfer trajectories to the fly-by or rendezvous of NEOs, including the escape phase at the Earth. The DDP approach allows the optimisation of the trajectory as a whole, without recurring to the patched conic approach. The results show how the proposed method is capable of fully exploiting the multi-body dynamics of the problem; in fact, in one of the study cases, a fly-by of the Earth is scheduled, which was not included in the first guess solution.EThOS - Electronic Theses Online ServiceGBUnited Kingdo

Space Science

The all-encompassing term Space Science was coined to describe all of the various fields of research in science: Physics and astronomy, aerospace engineering and spacecraft technologies, advanced computing and radio communication systems, that are concerned with the study of the Universe, and generally means either excluding the Earth or outside of the Earth's atmosphere. This special volume on Space Science was built throughout a scientifically rigorous selection process of each contributed chapter. Its structure drives the reader into a fascinating journey starting from the surface of our planet to reach a boundary where something lurks at the edge of the observable, light-emitting Universe, presenting four Sections running over a timely review on space exploration and the role being played by newcomer nations, an overview on Earth's early evolution during its long ancient ice age, a reanalysis of some aspects of satellites and planetary dynamics, to end up with intriguing discussions on recent advances in physics of cosmic microwave background radiation and cosmology