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Orbit accuracy of the transfer orbit and the mission orbit is the basis for the orbit control of all-electric-propulsion Geostationary Orbit (GEO) satellites. Global Navigation Satellite System (GNSS) simulation data are used to analyze the main factors affecting GEO satellite orbit prediction accuracy under the no-thrust condition, and an electric propulsion calibration algorithm is designed to analyze the orbit determination and prediction accuracy under the thrust condition. The calculation results show that the orbit determination accuracy of mission orbit and transfer orbit without thrust is better than 10 m using onboard GNSS technology. The calibration accuracy of electric thrust is about 10−9 m/s2 and 10−7 m/s2 with 40 h and 16 h arc length, respectively, using the satellite self-positioning data of 100 m accuracy to calibrate the electric thrust. If satellite self-positioning data accuracy is at the 10 m level, the electric thrust calibration accuracy can be improved by about one order of magnitude, and the 14-day prediction accuracy of the transfer orbit with thrust is better than 1 km. [ABSTRACT FROM AUTHOR]

In this paper, a three-maneuver strategy is presented to study the transfer from a cislunar L 2 halo orbit to a low lunar orbit (LLO). Both the cislunar L 2 point orbits and lunar orbits are generally considered as habitats for space vehicles of various missions, so a transfer strategy between these orbits that combines flexibility with minimal fuel consumption is worth considering for future complex space applications in the Earth-Moon system. In terms of the three-maneuver transfer method, the first maneuver is carried out along the direction of the unstable invariant manifold toward the moon, and each point in the halo orbit can be a departure position of the transfer. Then, the spacecraft is allowed to drift for a period of time, but in the meantime we are constantly searching the window for the second maneuver that towards the target LLO and figuring out the transfer trajectory for optimal maneuver cost. The third maneuver is a deceleration into the target LLO. The transfer procedure is first established in the circular restricted three-body problem (CRTBP) model, with a feasibility analysis of not constraining orbital inclination of LLOs. After that, the results of the transfer to both polar and equatorial LLOs in the CRTBP model are used as initial solutions for the design procedure of transfer trajectories in the ephemeris model. Finally, optimized results of transfer window of up to 180 days with corresponding overall maneuver costs in the ephemeris model are given for a quasi-halo orbit, which demonstrate the flexibility and reliability of the three-maneuver transfer strategy. [ABSTRACT FROM AUTHOR]

This research dealt with choosing the best satellite parking orbit and then the transition of the satellite from the low Earth orbit to the geosynchronous orbit (GEO). The aim of this research is to achieve this transition with the highest possible efficiency (lowest possible energy, time, and fuel consumption with highest accuracy) in the case of two different inclination orbits. This requires choosing a suitable primary parking orbit. All of the methods discussed in previous studies are based on two orbits at the same plane, mostly applying the circular orbit as an initial orbit. This transition required the use of the advanced technique of the Hohmann transfer method for the elliptical orbits, as we did in an earlier research, namely the transition from the perigee of the initial orbit to the final orbit and then conducting the rotation of the orbit plane to match the plane for the desired final orbit. The effect of the perigee altitude of the initial orbit on the transition efficiency calculated for the values between 300 to 3000 km. It was found that increasing the altitude reduces the energy and fuel needed for transportation, but the time required for transportation increases, into account that the increased height of the initial or parking orbit also implies the requirement of higher energy to reach it. The effects of eccentricity (e) values of the initial orbit between 0.01 to 0.2 on the transition efficiency were calculated. It was found that the increase in (e) reduces the energy and fuel, but does not affect the time, required for transportation. [ABSTRACT FROM AUTHOR]

We present a study of the transfer orbits connecting landing points of irregular-shaped asteroids. The landing points do not touch the surface of the asteroids and are chosen several meters above the surface. The ant colony optimization technique is used to calculate the multiple-hopping trajectories near an arbitrary irregular asteroid. This new method has three steps which are as follows: (1) the search of the maximal clique of candidate target landing points; (2) leg optimization connecting all landing point pairs; and (3) the hopping sequence optimization. In particular this method is applied to asteroids 433 Eros and 216 Kleopatra. We impose a critical constraint on the target landing points to allow for extensive exploration of the asteroid: the relative distance between all the arrived target positions should be larger than a minimum allowed value. Ant colony optimization is applied to find the set and sequence of targets, and the differential evolution algorithm is used to solve for the hopping orbits. The minimum-velocity increment tours of hopping trajectories connecting all the landing positions are obtained by ant colony optimization. The results from different size asteroids indicate that the cost of the minimum velocity-increment tour depends on the size of the asteroids. [ABSTRACT FROM AUTHOR]

In this paper, a novel optimized algorithm of two-impulse Earth–Moon transfers in the circular restricted three-body problem (CR3BP) is proposed based on differential correction approach. In this innovative approach, the initial and final states of transfer orbit are estimated reasonably by fully considering the periapsis map (PM) condition. Then the Newton–Raphson method is utilized iteratively to deduce the differential equations of transfer orbit and further to yield the accurate initial state. Thus, the planar two-impulse Earth–Moon transfer can be accomplished. In addition, a simple design procedure is developed to solve the problem that arises from more unknown parameters in the spatial CR3BP. Therefore, the proposed method can be applied not only for the planar CR3BP but also the spatial CR3BP. Numerical results show the two-impulse Earth–Moon transfers need less energy than the Hohmann transfers and shorter flight time than the Weak Stability Boundary (WSB) transfers. Moreover, the proposed method can effectively enable a large set of two-impulse Earth–Moon transfers to be computed numerically. [ABSTRACT FROM AUTHOR]

The two-body two-impulse orbit transfer problem between two coplanar elliptic orbits using only along-track thrust is studied in this paper. Both impulse vectors are constrained to be on the along-track directions at the impulse points. First, the relationship between transfer angle and final true anomaly is derived. For a given final true anomaly, the range of the transfer angle is obtained for different cases. Then, the transfer angle is solved by the secant method in its feasible range. With the transfer angle and the transfer semilatus rectum, the velocity vectors of the transfer orbit at impulse points are obtained. Two impulse points and the corresponding impulse magnitudes for a 180 $$^{\circ }$$ transfer are also analyzed. Finally, the minimum-energy and minimum-transit-time solutions are obtained by optimizing the final true anomaly. Numerical tests demonstrate the effectiveness of the proposed method for two-impulse along-track orbit transfers. The results show that the minimum-energy transfer with two along-track impulses is close to the 'true' minimum-energy transfer with two free-direction impulses. [ABSTRACT FROM AUTHOR]

A commercial ultraviolet erasable programmable read-only memory (UVPROM) was used to demonstrate a dosimetry technique for both ground and space applications. An equivalent amount of UV to reproduce the same amount of erasure was used to calibrate dose. The new method of readout, unlike other methods, does not require the evidence of exposure to be destroyed in the course of a measurement. It requires power only during readout. Results from an experiment using this technique aboard the Microelectronics and Photonics Test Bed (MPTB) satellite are discussed. Application of Extreme Value Theory is used to analyze whether early failures in the device were statistically feasible or more likely due to large, rare energy depositions. A new dosimeter approach using a change injection method that will eliminate the need for UV as a metric is also presented as well new experiments for the devices under test aboard MPTB. [ABSTRACT FROM AUTHOR]