We present and discuss a solution to the growing demand for satellite telecommunication coverage in the high-latitude geographical regions (beyond 55°N), where the signal from geostationary satellites is limited or unavailable. We focus on the dynamical issues associated to the design, the coverage, the maintenance and the disposal of a set of orbits selected for the purpose. Specifically, we identify a group of highly inclined, moderately eccentric geosynchronous orbits derived from the Tundra orbit (geosynchronous, eccentric and critically inclined). Continuous coverage can be guaranteed by a constellation of three satellites in equally spaced planes and suitably phased. By means of a high-precision model of the terrestrial gravity field and the relevant environmental perturbations, we study the evolution of these orbits. The effects of the different perturbations on the ground track (which is more important for coverage than the orbital elements themselves) are isolated and analyzed. The physical model and the numerical setup are optimized with respect to computing time and accuracy. We show that, in order to maintain the ground track unchanged, the key parameters are the orbital period and the argument of perigee. Furthermore, corrections to the right ascension of the ascending node are needed in order to preserve the relative orientation of the orbital planes. A station-keeping strategy that minimizes propellant consumption is then devised, and comparisons are made between the cost of a solution based on impulsive maneuvers and one with continuous thrust. Finally, the issue of end-of-life disposal is discussed.
This paper reports and discusses the design and ground tests of a CubeSat payload which allows to measure, in-situ and in real time, the degradation of a polymer of electronic interest due to atomic oxygen etching in LEO. It provides real-time information on how the degradation occurs, eliminating the need to work with samples recovered once the mission has finished. The polymer, TIPS-Pentacene, is deposited on the surface of a microelectromechanical (MEMS) cantilever, which works as a resonator embedded in a Pulsed Digital Oscillator circuit. The mass losses in the polymer due to atomic oxygen corrosion produce variations in the resonant frequency of the MEMS, which is continuously sensed by the circuit and transmitted to the ground. This way, polymer mass losses around 10-12 kg can be detected during the mission. The payload is a part of the 3Cat-1 mission, a nano-satellite aimed at carrying out several scientific experiments.
Given a limited warning time, an asteroid impact mitigation campaign would hinge on uncertainty-based information consisting of remote observational data of the identified Earth-threatening object, general knowledge of near-Earth asteroids (NEAs), and engineering judgment Due to these ambiguities, the campaign credibility could be profoundly compromised. It is therefore imperative to comprehensively evaluate the inherent uncertainty in deflection and plan the campaign accordingly to ensure successful mitigation. This research demonstrates dual-deflection mitigation campaigns consisting of primary (instantaneous/quasi-instantaneous) and secondary (slow-push) deflection missions, where both deflection efficiency and campaign credibility are taken into account The results of the dual-deflection campaign analysis show that there are trade-offs between the competing aspects: the launch cost, mission duration, deflection distance, and the confidence in successful deflection. The design approach is found to be useful for multi-deflection campaign planning, allowing us to select the best possible combination of missions from a catalogue of campaign options, without compromising the campaign credibility.
In this paper we address the feasibility of capturing small Near-Earth Asteroids (NEAs) into the vicinity of the Sun-Earth L-2 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)  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 L-2 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. (C) 2013 IAA. Published by Elsevier Ltd. All rights reserved.
One of the biggest challenges of the exploration of the Moon is the survival of the crew and the lunar assets during the lunar night. The environmental conditions on the lunar surface and its cycle, with long periods of darkness, make any long mission in need of specific amounts of heat and electricity to be successful. We have analyzed two different systems to produce heat and electricity on the Moon's surface. The first system consists of Thermal Wadis, sources of thermal power that can be used to supply heat to protect the exploration systems from the extreme cold during periods of darkness. Previous results showed that Wadis can supply enough heat to keep lunar devices such as rovers above their minimum operating temperature (approximately 243 K). The second system studied here is the Thermal Energy Storage (TES), which is able to run a heat engine during the lunar night to produce electricity. When the Sun is shining on the Moon's surface, the system can run the engine directly using the solar power and simultaneously heat a thermal mass. This thermal mass is used as a high temperature source to run the heat engine during the night. We present analytical and numerical calculations for the determination of an appropriate thermal mass for the TES system.
A theoretical analysis considering the capabilities of nano electrokinetic thrusters for space propulsion is presented. The work describes an electro-hydro-dynamic model of the electrokinetic flow in nano-channels and represents the first attempt to exploit the advantages of the electrokinetic effect as the basis for a new class of nano-scale thrusters suitable for space propulsion. Among such advantages are their small volume, fundamental simplicity, overall low mass, and actuation efficiency. Their electrokinetic efficiency is affected by the slip length, surface charge, pH and molarity. These design variables are analyzed and optimized for the highest electrokinetic performance inside nano-channels. The optimization is done for power consumption, thrust and specific impulse resulting in high theoretical efficiency ~99% with corresponding high thrust-to-power ratios. Performance curves are obtained for the electrokinetic design variables showing that high molarity electrolytes lead to high thrust and specific impulse values, whereas low molarities provide highest thrust-to-power ratios and efficiencies. A theoretically designed 100 nm wide by 1 µm long emitter optimized using the ideal performance charts developed would deliver thrusts from 5 to 43 µN, specific impulse from 60 to 210 s, and would have power consumption between 1–15 mW. It should be noted that although this is a detail analytical analysis no prototypes exist and any future experimental work will face challenges that could affect the final performance. By designing an array composed of thousands of these single electrokinetic emitters, it would result in a flexible and scalable propulsion system capable of providing a wide range of thrust control for different mission scenarios and maintaining very high efficiencies and thrust-to-power ratio by varying the number of emitters in use at any one time.
The analysis of optical navigation in an Earth–Moon libration point orbit is examined. Missions to libration points have been winning momentum during the last decades. Its unique characteristics make it suitable for a number of operational and scientific goals. Literature aimed to study dynamics, guidance and control of unstable orbits around collinear libration points is vast. In particular, several papers deal with the optimisation of the Δv budget associated to the station-keeping of these orbits. One of the results obtained in literature establishes the critical character of the Moon–Earth system in this aspect. The reason for this behaviour is twofold: high Δv cost and short optimal manoeuvre spacing. Optical autonomous navigation can address the issue of allowing a more flexible manoeuvre design. This technology has been selected to overcome similar difficulties in other critical scenarios. This paper analyses in detail this solution. A whole GNC system is defined to meet the requirements imposed by the unstable dynamic environment. Finally, a real simulation of a spacecraft following a halo orbit of the L2 Moon–Earth system is carried out to assess the actual capabilities of the optical navigation in this scenario.
The purpose of this paper is to obtain a third-order expression, for the in-plane and out-of-plane amplitudes, of the solutions of the elliptic Hill–Clohessy–Wiltshire non-linear equations. The resulting third-order solution is explicit in terms of true anomaly. The coefficients of the expansions are given as functions of the eccentricity e of the orbit of the leader (i.e., are valid for all values of e). For e=0 we recover the solution given by Richardson and Mitchell for the circular case; for e≠0 the linear terms of the solution recover the solution found by Lawden for the linearised elliptic HCW equations, also known as the Tschauner–Hempel equations. In the last part of the paper we explain how a formal series solution of the elliptic HCW non-linear equations (in powers of the two amplitudes and the eccentricity) can be obtained, using the Lindstedt–Poincaré procedure.