1,721,129 research outputs found
Constellations of inclined heliotropic orbits for enhanced earth coverage
No full textThis paper proposes a mission design of a constellation for improved Earth coverage using a family of 'heliotropic' orbits. The secular effects of the Earth's oblateness is exploited to maintain the spacecraft on a family of inclined orbits with a constant Sun-pointing apogee, so that enhanced coverage is provided during daylight hours for visible light imaging, or providing communications services during peak local demand. The initial orbital elements of the constellation are designed to meet increased coverage requirements. The orbit semi-major axis is selected in the heliotropic family for repeated groundtrack on specific locations on the Earth's surface. It is also shown that a small solar sail allows having these orbits at higher eccentricity, allowing the spacecraft to spend more time in the dayside of the Earth, w hen orbiting through the apogee. Copyright© (2012) by the International Astronautical Federation
Wave-like patterns in an elliptical satellite ring
Full text linkSatellite constellations are families of orbits selected to provide useful coverage patterns for telecommunications, Earth observation and navigation services. Such constellations are often assembled from families of circular orbits, which ensures a uniform spacing between satellites in each circular ring. However, there is a large class of elliptical orbits which are of practical interest including Molniya-like orbits and so-called Magic orbits [1,2]. Constellations of satellites using such elliptical orbits will then exhibit a time varying spacing between satellites as the orbital angular velocity experienced by each satellites varies around the elliptical ring. While current constellations use relatively modest numbers of satellites, future microspacecraft [3] or ‘smart dust’ type devices [4,5] may enable constellations with extremely large numbers of nodes. In this Note a continuum approach is used to model the dynamics of such constellations. A continuity equation is formed to describe the evolution of the number density of nodes as a function of both true anomaly and time. For small eccentricities, the continuity equation can be solved analytically to provide closed-form solutions which describe the evolution of the constellation for some initial distribution of nodes. The closed-form solutions can then be used to investigate pattern formation in elliptical rings. Wave-like patterns are found which circulate around the elliptical ring, with peaks in density which can in principle be used to provide enhanced coverage. A similar continuum approach with a continuity equation has been used in previous studies to develop closed-form solutions which model the time evolution of the radial distribution of constellations of microspacecraft under the action of air drag [6,7]
Orbit design for future SpaceChip swarm missions
No full textThe effect of solar radiation pressure and atmospheric drag on the orbital dynamics of satellites-on-a-chip (SpaceChips) is exploited to design long-lived orbits about the Earth. The orbit energy gain due to asymmetric solar radiation pressure, considering the Earth shadow, is used to balance the energy loss due to atmospheric drag. Future missions for a swarm of SpaceChips are proposed, where a number of small devices are released from a conventional spacecraft to perform spatially distributed measurements of the conditions in the ionosphere and exosphere. It is shown that the orbit lifetime can be extended and indeed selected through solar radiation pressure and the end-of-life re-entry of the swarm can be ensured, by exploiting atmospheric drag
Orbit evolution, maintenance and disposal of SpaceChip swarms through electro-chromic control
Full text linkThe combined effect of solar radiation pressure, Earth oblateness and atmospheric drag on the orbital dynamics of satellites-on-a-chip (SpaceChips) is investigated for future swarm mission concepts. The natural evolution of the swarm is exploited to perform spatially distributed measurements of the upper layers of the atmosphere. The energy gain from asymmetric solar radiation pressure can be used to balance the energy dissipation from atmospheric drag. An algorithm for long-term orbit control is then designed, based on changing the reflectivity coefficient of the SpaceChips. The subsequent modulation of the solar radiation pressure allows stabilisation of the swarm in the orbital element phase space. It is shown that the orbit lifetime for such devices can be extended through the interaction of solar radiation pressure and atmospheric drag and indeed selected and the end-of-life re-entry of the swarm can be ensured, by exploiting atmospheric drag
A passive satellite deorbiting strategy for MEO using solar radiation pressure and the J2 effect
No full textThe growing population of space debris poses a serious risk to the future of space flight. To effectively manage the increase of debris in orbit, end-of life disposal has become a key requirement for future missions. This poses a challenge for Medium Earth Orbit (MEO) spacecraft which require a large Δv to re-enter the atmosphere or reach the geostationary graveyard orbit. This paper further explores a passive strategy based on the joint effects of solar radiation pressure and the Earth’s oblateness acting on a high area-to-mass ratio object. The concept was previously presented as an analytical planar model. This paper uses a full 3D model to validate the analytical results numerically for equatorial circular orbits first, then investigating higher inclinations. It is shown that for higher inclinations the initial position of the Sun and right ascension of the ascending node become increasingly important. A region of very low required area-to-mass ratio is identified in the parameter space of a and inclination which occurs for altitudes below 10,000 km
Orbital dynamics of high area-to-mass ratio spacecraft under the influence of J2 and solar radiation pressure
No full textThis paper investigates the effect of planetary oblateness and solar radiation pressure on the orbit of high area-to-mass spacecraft. A planar Hamiltonian model shows the existence of equilibrium orbits with the orbit apogee pointing towards or away from the Sun. These solutions are numerically continued to non-zero inclinations and considering the obliquity of the ecliptic plane relative to the equator. Quasi-frozen orbits are identified in eccentricity, inclination and angle between the Sun-line and the orbit perigee. The long-term evolution of these orbits is then verified through numerical integration. A set of ‘heliotropic’ orbits with apogee pointing in direction of the Sun is proposed for enhancing imaging and telecommunication on the day side of the Earth. The effects of J2 and solar radiation pressure are exploited to obtain a passive rotation of the apsides line following the Sun; moreover the effect of solar radiation pressure enables such orbits at higher eccentricities with respect to the J2 only case
Electrochromic orbit control for smart-dust devices
No full textRecent advances in MEMS (micro electromechanical systems) technology are leading to spacecraft which are the shape and size of computer chips, so-called SpaceChips, or ‘smart dust devices’. These devices can offer highly distributed sensing when used in future swarm applications. However, they currently lack a feasible strategy for active orbit control. This paper proposes an orbit control methodology for future SpaceChip devices which is based on exploiting the effects of solar radiation pressure using electrochromic coatings. The concept presented makes use of the high area-to-mass ratio of these devices, and consequently the large force exerted upon them by solar radiation pressure, to control their orbit evolution by altering their surface optical properties. The orbital evolution of Space Chips due to solar radiation pressure can be represented by a Hamiltonian system, allowing an analytic development of the control methodology. The motion in the orbital element phase space resembles that of a linear oscillator, which is used to formulate a switching control law. Additional perturbations and the effect of eclipses are accounted for by modifying the linearized equations of the secular change in orbital elements around an equilibrium point in the phase space of the problem. Finally, the effectiveness of the method is demonstrated in a test case scenario
Space debris cloud evolution in Low Earth Orbit
No full textThe Earth is surrounded by inoperative objects created by past space missions; as the orbital speed is very high, the impact with a very small fragment, down to 1 cm, can be catastrophic for operating satellites. Therefore, it is important to assess the collision risk due to space debris; this requires a reliable picture of the debris environment and a deep understanding of its evolution. In this work, an analytical approach is used to describe the evolution of a debris cloud created by a collision in Low Earth Orbit. In contrast to traditional approaches, which follow the trajectory of single fragments, here the cloud behaviour is studied globally. This reduces the computational time needed to estimate the consequence of a collision and allows simulating several what-if scenarios to understand which objects, in case of fragmentation, are more likely to pose an hazard to operational spacecraft. The NASA break-up model is used to describe fragments dispersion in terms of characteristic length, area-to-mass ratio and velocity. From the velocity distribution the fragment spatial dispersion is derived, through an estimation of the time after which the fragments create a band around the Earth. The cloud density is expressed by a distribution function that depends only on altitude and that is set as initial condition for the orbit propagation. Based on an analytical approach proposed in the literature for interplanetary dust and spacecraft swarms, the fragment cloud evolution in time is derived through the continuity equation. In this application, the continuity equation describes the variation of debris density considering Earth's gravity and atmospheric drag. The cloud evolution is compared to the numerical integration to assess the method's accuracy. The proposed approach proves to be very promising as it is able to capture the main phenomena undergoing the evolution of the semi-major axis distribution. The applicability limits are discussed and the main areas for the method improvement are identified. Copyright © 2013 by Letizia, Colombo, Lewis, Mclnnes. ©2013 by the International Astronautical Federation. All rights reserved
An optimal gains matrix for time-delay feedback control
Full text linkIn this paper we propose an optimal time-delayed feedback control (TDFC) for tracking unstable periodic orbits (UPOs). It is shown that TDFC will drive a trajectory onto a periodic orbit while minimising an integral of a cost function of the error in periodicity and the control e®ort. This optimal TDFC relies on the linearisation about the delayed trajectory not the UPO itself and therefore can be implemented without a priori knowledge of a reference orbit. This optimal TDFC is applied to the problem of tracking an unstable periodic orbit in the nonlinear equations describing the circular restricted three-body problem. The results of this investigation demonstrate that TDFC could e±ciently drive a spacecraft onto a periodic orbit in the vicinity of a (UPO) halo orbit
Solar radiation pressure-augmented deorbiting: passive end-of-life disposal from high-altitude orbits
No full textA deorbiting strategy for small satellites is proposed that exploits the effect of solar radiation pressure to increase the spacecraft orbit eccentricity so that the perigee falls below an altitude where atmospheric drag will cause the spacecraft orbit to naturally decay. This is achieved by fitting the spacecraft with an inflatable reflective balloon. Once this is fully deployed, the overall area-to-mass ratio of the spacecraft is increased; hence, solar radiation pressure and aerodynamic drag have a greatly increased effect on the spacecraft orbit. An analytical model of the orbit evolution due to solar radiation pressure and the J2 effect as a Hamiltonian system show the evolution of an initially circular orbit. The maximum reachable orbit eccentricity as a function of semimajor axis and area-to-mass ratio is found analytically for deorbiting from circular equatorial orbits of different altitudes. The analytical planar model is then adapted for sun-synchronous orbits. The model is validated numerically and verified for three test cases using a high-accuracy orbit propagator. The regions of orbits for which solar radiation pressure-augmented deorbiting is most effective are identified. Finally, different options for the design of the deorbiting device are discussed
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