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Cano J.L.,Elecnor Deimos | Heritier A.,Elecnor Deimos | De Zaiacomo G.,Elecnor Deimos
Proceedings of the International Astronautical Congress, IAC | Year: 2016

Returning soil samples from the Lunar South Pole region has been of great interest by scientists for decades. The samples will allow determining the age of key events and lead to a better understanding of the history of our solar system. This paper presents the mission design for the proposed Lunar Polar Sample Return, which is currently a proposed joint mission between the European Space Agency (ESA) and the Russian Federation (ROSCOSMOS) with launch foreseen in 2024. This mission is composed of a lander and an orbiter that are launched separately. The lander arrives first at the moon and extracts the samples. It remains on the lunar surface for several days before lifting off and inserting into a low lunar polar orbit. A rendezvous strategy is then initiated for the orbiter to recover the sample canister from the lander. Finally the orbiter performs a trans-Earth insertion (TEI) burn to return to Earth and release the entry capsule in Kazakhstan a few days later. An overview of the main phases of the mission is first discussed, which includes timeline and phasing strategy between the two spacecraft. One critical aspect lies in the strong requirements for the return leg to Earth, which makes this mission challenging and drives the overall design. The flight path angle at EIP is constrained by the entry corridor designed to cope with the EDL phase requirements. Moreover, landing over Kazakhstan requires that departures from the Moon are at periods of time when this is at lowest equatorial declination (thus one window of a few days each month). In addition, this needs to be performed between April and October to avoid snow cover and during the first three hours of daylight. Some minimum-energy return transfer trajectories are investigated that meet the Earth re-entry conditions and comply with the mission requirements. An optimization problem is formulated from a low lunar polar orbit to the Earth surface. The return trajectories are generated in a high fidelity model considering third body perturbations as well as Earth oblateness. The effect of the main contingencies, such as delay in the ascent of the lander or delay in the TEI implementation are also analyzed. Some strategies to deal with those contingencies are explained and return solutions are proposed to assure returns to Earth that satisfy the arrival requirements even in case of delays. Copyright © 2016 by Elecnor Deimos.


Tonetti S.,Elecnor Deimos | Cornara S.,Elecnor Deimos | Faenza M.,Nammo Raufoss AS | Verberne O.,Nammo Raufoss AS | And 2 more authors.
Proceedings of the International Astronautical Congress, IAC | Year: 2016

Hundreds of satellites populate the Earth-bounded orbits, and this number is rapidly increasing. These objects are a threat as they can collide with other satellites, in turn creating new debris objects. Several collisions have already occurred and the population of debris will keep growing if no measures are taken to mitigate the generation of space debris by implementing proper policies and mission design standards, as well as to remove space debris in the future. This paper presents the results of an ESA General Studies Programme (GSP) study, dealing with the feasibility of performing active debris removal (ADR) by using a hybrid propulsion system on the chaser spacecraft. While the study focuses mainly on the use of a hybrid rocket propulsion system on-board a chaser spacecraft that performs ADR, it also preliminarily addresses the application of this innovative propulsion technology for debris mitigation purposes. Hybrid propulsion systems seem to be a promising alternative to conventional liquid propellant in-orbit propulsion systems, in terms of complexity, cost, operational advantages, whilst also offering the use of non-toxic fuel. The study is carried out by Deimos Space, expert in mission analysis, and Nammo Raufoss, expert in hybrid propulsion. A survey of LEO in-orbit spacecraft, under current design and planned future missions, allows the identification of a set of orbits and satellites (about 130) that are interesting targets for an ADR mission. Parametric analyses, covering different target orbit altitudes and masses, are performed to compute the delta-V that a chaser spacecraft should provide to remove the identified objects. For the ADR mission analysis, three different propulsive phases are studied: 1) transfer to target orbit, considering Vega as baseline launcher and maximizing the chaser mass at target orbit; 2) approach to target, analysing far and close rendezvous; 3) de-orbit and re-entry, both controlled and un-controlled, considering different strategies and ballistic coefficients, typical thrust-to-mass ratios for hybrid propulsion system and gravity losses. The parametric delta-V collected from the propulsive phases with adequate margins is considered as input for sizing the hybrid propulsion system for selected targets with the goal of identifying the most promising debris-altitude combinations that can be removed with such kind of technology. All the main components of the propulsion system are sized and the hybrid motor performance and behaviour are assessed. A trade-off with chemical propulsion systems typically used on spacecraft is performed, identifying advantages of hybrid propulsion systems and technology gaps to cover.


Ferri A.,Thales Alenia | Fenoglio F.,Thales Alenia | Perino M.A.,Thales Alenia | Pelle S.,Thales Alenia | And 3 more authors.
Proceedings of the International Astronautical Congress, IAC | Year: 2016

In the last years, the interest in the Lunar South Polar region has grown significantly within the international exploration and science community, fueled by the abundance of new data acquired by the fleet of orbiter missions which have been sent to the Moon in the past decade. The next step in exploration of this important region requires direct investigation of the surface and the material there, both through in-situ measurements and via a sample return. Thales Alenia Space is leading an ESA study to design a Lunar Polar Sample Return (LPSR) mission as a joint ESA-ROSCOMOS exploration mission. The mission objective is the retrieval of water ice samples from the lunar South Pole and return them back to Earth still in a solid state. The principal objective of this activity is the definition of a feasible mission scenario and assess the European system elements of the Lunar Polar Sample Return (LPSR) mission architecture with particular regard to the sampling return phases, including: • Sample Handling preservation from the lunar surface to retrieval on Earth • Lunar Ascent Vehicle • In-orbit Rendezvous and Capture • Orbiter Module • Earth Return Vehicle • Earth Return Capsule This paper will present the pre-Phase A results of the study. Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved.


Llado N.,Elecnor Deimos | Ren Y.,York University | Masdemont J.J.,Institute Destudis Espacials Of Catalonia Ieec | Masdemont J.J.,Polytechnic University of Catalonia | And 2 more authors.
Acta Astronautica | Year: 2014

In this paper we address the feasibility of capturing small Near-Earth Asteroids (NEAs) into the vicinity of the Sun-Earth L2 libration point using a continuous-thrust propulsion system assumed to be attached to the asteroid. The vicinity of this libration point is a gateway to the Earth-Moon neighborhood and using it for capture, or for transit, small NEAs could be interesting for mining or science purposes. Due to limited maneuver capabilities and security concerns, only NEAs with very small mass, and not representing a potential hazard, are analyzed. First, the NEAs are pruned from JPL NEAs (Jet Propulsion Laboratory, 2012) [1] database and their diameter and mass are estimated using two different methods based on physical properties. Then, fuel-optimal continuous-thrust transfer orbits from the original positions of the NEAs to the Sun-Earth L2 libration point are computed. For this trajectory optimization, the initial seeds are generated by means of a global optimization procedure based on a differential evolution algorithm. Next, these initial seeds are refined with a fourth order Runge-Kutta shooting method, and finally we list the candidate NEAs to be captured using a continuous-thrust propulsion system including the key parameters defining their transfer trajectory. © 2013 IAA.


Vasconcelos J.F.,Elecnor Deimos | Rosa P.,Elecnor Deimos | Kerr M.,Elecnor Deimos | Sierra A.L.,Elecnor Deimos | And 2 more authors.
European Space Agency, (Special Publication) ESA SP | Year: 2015

This paper describes the development of a fault detection system for a model scale autonomous aircraft. The considered fault scenario is defined by malfunctions in the elevator, namely bias and stuck-in-place of the surface. The H∞ design methodology is adopted, with an LFT description of the aircraft longitudinal dynamics, that allows for fault detection explicitly synthesized for a wide range of operating airspeeds. The obtained filter is validated in two stages: in a Functional Engineering Simulator (FES), providing preliminary results of the filter performance; and with experimental data, collected in field tests with actual injection of faults in the elevator surface.


Cacciatore F.,Elecnor Deimos | Cichocki F.,Elecnor Deimos
Proceedings of the International Astronautical Congress, IAC | Year: 2013

In the frame of the studies performed for the asteroid sample retrieval mission MarcoPolo-R, special relevance is being given to the analysis and design of the mission phase in which the S/C will fly in close formation with the target body. The motion around an asteroid shows distinct peculiarities due to the dynamical environment encountered. The low mass of the asteroid implies that other perturbations may be large with respect to the body central gravity; the irregular shape of the asteroid, which in some cases can be very far from being a sphere, causes the classical spherical harmonics modelling to lose accuracy; finally, many asteroids, as the old MarcoPolo-R baseline target 1996 FG3, are binary systems, composed by two bodies of comparable size, generating complex gravitational forces on the S/C. The analysis and design of the close formation flying with an asteroid is complicated further by the limited knowledge available of the physical characteristics of the asteroid environment (mass, size, shape, rotation of the primary, secondary mass size and orbit...). The main tasks carried out by Deimos for the proximity phase analysis were aimed at assessing the delta-V cost and feasibility (stability, safety, and operations) of different types of orbit about the asteroid to execute asteroid global characterisation, radio science experiment and local characterisation prior to descent and landing. The proximity flight conditions considered were controlled polar orbits, terminator plane photo-stable orbits, inertial hovering and body fixed hovering. Two guidance modes were defined for the body fixed hovering, based on the use of full asteroid-relative position observables or only on altitude and relative surface velocity data. This latter mode was further optimised to reduce the resulting longitude drift over the hovering duration. In order to take into account both the effect of different orbit reference conditions (for instance, reference asteroid distance or orbital radii) and of the variability of the environment parameters, two main sets of simulations were carried out. In the first set in which nominal environment conditions were used and the orbit reference parameters were varied, while in the second set extensive Monte Carlo runs were used to simulate nominal orbit scenarios in different stochastic realisations of the environment, enforcing physical coherency of each shot.


Cano J.L.,Elecnor Deimos | Bellei G.,Elecnor Deimos | Martin J.,Elecnor Deimos
Proceedings of the International Astronautical Congress, IAC | Year: 2013

Protecting Earth from the threat implied by the Near Earth Objects (NEO) is gaining momentum in recent years. In the last decade a number of mitigation methods have been pushed forward as a possible remedy to that threat, including nuclear blasts, kinetic impactor, gravity tractors and others. Tools are required to evaluate the NEO deflection performances of each of the different methods, coupled with the orbital mechanics associated to the need to transfer to the target orbit and maybe rendezvous with it. The present suite of tools do provide an integral answer to the need of determining if an asteroid is to collide with Earth (NIRAT tool), compute the required object deflection (NEODET tool) and assess the design features of the possible mitigation space missions (RIMISET tool). The tools are presented, their design analyzed as well as the methods and architecture implemented. Results are provided for two asteroids 2011 AG5 (using the orbit determination solution where this asteroid still was a risk object) and 2007 VK184 and the obtained data discussed in comparison to other results. Copyright© (2013) by the International Astronautical Federation.


Bombardelli C.,Technical University of Madrid | Amato D.,Technical University of Madrid | Cano J.L.,Elecnor Deimos
Acta Astronautica | Year: 2016

Based on a hypothetical asteroid impact scenario proposed during the 2015 IAA Planetary Defense Conference (PDC), we study the deflection of fictitious asteroid 2015 PDC starting from ephemeris data provided by the conference organizers. A realistic mission scenario is investigated that makes use of an ion beam shepherd spacecraft as a primary deflection technique. The article deals with the design of a low-thrust rendezvous trajectory to the asteroid, the estimation of the propagated covariance ellipsoid and the outcome of an ion beam slow-push deflection starting from three worst case scenarios (impacts in New Delhi, Dhaka and Tehran). Displacing the impact point towards an extremely low-populated, easy-to-evacuate region, as opposed to full deflection, is found to be a more effective mitigation approach. Mission design, technical and political aspects are discussed. © 2015 IAA.


De Filippis L.,Elecnor Deimos | Kerr M.,Elecnor Deimos | Haya R.,Elecnor Deimos
Astrophysics and Space Science Proceedings | Year: 2016

In this paper a new trajectory generator for the Terminal Area Energy Management phase of a Reusable Launch Vehicle is presented. During this phase the vehicle has to glide at low Mach to reach the point close to the runway where automatic approach and landing starts. The algorithm presented here is based on the concept of Energy Corridor management and it is composed of two main elements: a trajectory propagator and a ground track generator. Imposing a dynamic pressure profile as function of the altitude, a heading path is selected in order to steer the vehicle toward the runway, putting to zero cross and down track errors. © Springer International Publishing Switzerland 2016.


Becedas J.,Elecnor Deimos
Proceedings - 2014 8th International Conference on Innovative Mobile and Internet Services in Ubiquitous Computing, IMIS 2014 | Year: 2014

Earth Observation (EO) is considered a key element in the European Research Roadmap and an opportunity market for the next years. However, this field presents some critical challenges to cover the current demand of services: i) there is massive and large-sized data from Earth Observation recordings; ii) On demand storage, processing and distribution of geoinformation generated with the recorded data are required. Conventional infrastructures have the risks of over/under size the infrastructure when big data is used, they are not flexible to cover sudden changes in the demand of services and the access to the information presents large latencies. These aspects limit the use of EO technology for real time use. The use of cloud computing technology can overcome the previously defined limitations. The GEO-Cloud experiment emerged to find viable solutions to provide highly demanding EO services by using future internet technologies. It is a close to reality experiment, part of the FP7 Fed4FIRE project. GEO-Cloud consists of the design, implementation and testing in cloud a complete EO system, from the acquisition of geo-data with a constellation of satellites to its on demand distribution to end users with remote access. This paper presents the GEO-Cloud experiment design, architecture and foreseen research activity. © 2014 IEEE.

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