A Proton Flux Model Dedicated to Solar Arrays Degradations on Electric Orbit Raising Missions

<p>Electric Orbit Raising (EOR) for telecommunication satellites has allowed significant reduction in onboard fuel mass, at the price of extended transfer durations. These relatively long orbital transfers, which can take up to a few months, equatorially cross most of the radiation belts, resulting in significant exposure of the spacecraft to space radiations. Since there are not covered by many spacecrafts, the radiation environment of intermediate regions of the radiation belts is less known than on popular orbits such as LEO or GEO. In particular, there is a need for more specific models for the MeV energy range proton fluxes, responsible for solar arrays degradations. We present a model of proton fluxes dedicated for EOR missions that was developped as part of the ESA ARTES program. This model is able to estimate the average proton fluxes between 60 keV and 10MeV on arbitrary trajectories on the typical durations of EOR transfers. A global statistical model of the radiation belts was extracted from the Van Allen Probes (RBSP) RBSPICE data and enriched by simulation results from the Salammbô radiation belt model were used. A special care was taken to model the temporal dynamics of the proton belt, allowing to compute analytically the distribution of the average fluxes on arbitrary EOR missions.</p>

On the choice of a parking orbit for a service spacecraft

At present, the requirements for increasing spacecraft active life and operational reliability and reducing spacecraft operation costs become more and more stringent. Because of this, on-orbit servicing becomes more and more attractive. One of the most promising ways to increase the efficiency of transport operations in space is to carry out on-orbit servicing using reusable spacecraft with low-thrust solar electrojet engines. The aim of this paper is to develop a mathematical model for the choice of an optimal low near-Earth parking orbit for a reusable service spacecraft. The case of noncoplanar near-circular orbits of spacecraft and a shuttle scenario of their servicing is considered. The solution of the problem of choosing an optimal parking orbit for a reusable service spacecraft involves repeated solutions of the problem of determining the delta-velocity of the service spacecraft’s orbital transfers between its parking orbit and the orbits of the serviced spacecraft. In this connection, using the averaging method, a mathematical model is developed for the analytical determination of orbital transfer program controls and trajectories and assessing orbital transfer energy expenditures. With its use, a mathematical model is developed for the choice of a service spacecraft’s optimal parking orbit. The objective function is the total delta-velocity of the service spacecraft’s orbital transfers from its parking orbit to the orbits of the serviced spacecraft and vice versa with the inclusion of the orbital transfer frequency. The optimizable parameters are the service spacecraft parking orbit parameters. The use of the proposed models is illustrated by an example of service spacecraft parking orbit optimization. What is new is the mathematical models developed. The results obtained may be used in the preliminary planning of on-orbit servicing operations.

Reachable domain of spacecraft with a single normal impulse

Purpose The purpose of this paper is to obtain the reachable domain (RD) for spacecraft with a single normal impulse while considering both time and impulse constraints. Design/methodology/approach The problem of RD is addressed in an analytical approach by analyzing for either the initial maneuver point or the impulse magnitude being arbitrary. The trajectories are considered lying in the intersection of a plane and an ellipsoid of revolution, whose family can be determined analytically. Moreover, the impulse and time constraints are considered while formulating the problem. The upper bound of impulse magnitude, “high consumption areas” and the change of semi-major axis and eccentricity are discussed. Findings The equations of RD with a single normal impulse are analytically obtained. The equations of three scenarios are obtained. If normal impulse is too large, the RD cannot be obtained. The change of the semi-major axis and eccentricity with large normal impulse is more obvious. For long-term missions, the change of semi-major axis and eccentricity leaded by multiple normal impulses should be considered. Practical implications The RD gives the pre-defined region (all positions accessible) for a spacecraft under a given initial orbit and a normal impulse with certain magnitude. Originality/value The RD for spacecraft with normal impulse can be used for non-coplanar orbital transfers, emergency evacuation after failure of rendezvous and docking and collision avoidance.

Fast solution of minimum-time low-thrust transfer with eclipses

Optimizing low-thrust orbital transfers with eclipses by indirect methods raises several issues, namely the costate discontinuities at the eclipse entrance and exit, the initial costate guess sensitivity and the numerical accuracy required by the shooting method. The discontinuity issue is overcome by detecting the eclipse within the simulation and applying the costate jump derived analytically from the shadow constraint function. By fixing completely the targeted final position and velocity, the transversality conditions are removed and the shooting problem is recast as an unconstrained nonlinear programming problem. The numerical sensitivity issues are alleviated by using a derivative-free algorithm. The search space is reduced to four angles taking near zero values. This procedure yields a quasi-optimal solution from scratch in few minutes without requiring any specific user’s guess or tuning. The method is applicable whatever the thrust level and the eclipse configuration, as illustrated on transfers towards the geostationary orbit.