Vol 61, No 5 (2023)

The transfer scheme that provides a parabolic entry of the spacecraft into the Earth’s atmosphere has been optimized. Such a maneuver can be of interest in experimental testing of the spacecraft reentry from the Moon or after interplanetary missions. It is assumed that the spacecraft is inserted into a low Earth orbit and is equipped with a chemical propulsion system and a limited-thrust engine, which should provide a maneuver to bring the spacecraft into the Earth’s atmosphere. The optimization criterion takes into account the characteristic velocity of the maneuver. The developed method of optimizing the transfer scheme and the spacecraft trajectory itself is based on the maximum principle. Single-revolution and multi-revolution transfer trajectories are analyzed. It is shown that for single-revolution trajectories there is an optimal time and an optimal angular distance of flight. Their values and the minimum characteristic velocity of the maneuver are evaluated. Unlike single-revolution trajectories, the characteristic velocity for multi-revolution trajectories monotonically decreases with increasing transfer duration. The dependence of the characteristic velocity on the transfer duration for single-, two-, three- and four-revolution trajectories is given. The transfer duration ranges in which it is advisable to use each type of trajectory are analyzed.

The study considers the problem of calculating the low-energy trajectories of a low-thrust spacecraft to the Moon during the ballistic capture. The transfer is carried out using a transit trajectory in the vicinity of one of the collinear libration points L1 or L2 of the Earth-Moon system. Using a transit trajectory allows us to reduce fuel consumptions for the transfer by applying spacecraft dynamic in the Earth-Moon system. After exit from the orbit of ballistic capture, depending on the goal of mission the required lunar orbit can be formed, or the maneuver can be completed for inserting into the required interplanetary trajectory. A method for solving the problem is proposed, which consists in selecting the suitable transit trajectory to ensure sufficiently long duration of staying a spacecraft in the sphere of influence of the Moon, and in calculating the optimal low-thrust trajectories from initial lunar orbit to the transit trajectory to the Moon. To solve the problem of optimal control and calculate the optimal exit point to the transit trajectory, the Pontryagin’s maximum principle is used in combination with the continuation method by parameter. Numerical examples are given for calculating low-energy trajectories to the Moon during the ballistic capture with the optimization of exit point to the transit trajectory.

Possible aspects of violation of the functional safety of spacecraft in terms of electromagnetic compatibility with electric rocket thrusters in their work on alternative working substances are considered. The procedure of experimental determination of spectral–time characteristics of own electromagnetic radiation of laboratory model of stationary plasma thruster SPT-70 developed by the Research Institute of Applied Mechanics and Electrodynamics of the Moscow Aviation Institute is described. Measurements of noise emissions were carried out on a vacuum installation with a “radiotransparent” compartment and a shielded echo-free camera in the frequency range of 1–12 GHz for typical discharge capacities (600, 800, and 1000 W), vertical and horizontal polarization, and various working substances used (krypton and xenon). The conducted studies have allowed obtaining new comparative results of the assessment of spectral characteristics of SPT-70 radiation for standard modes and prospective working bodies within the orthogonal polarization bases. The new results should include information about the radiation characteristics of SPT-70 in the time area. It is shown that the transition from xenon to krypton retains the pulsed nature of the radiation of a stationary plasma thruster, leading not only to an increase in the amplitude of pulses, but also to an increase in the frequency of repetition of “bursts” and an increase in their duration, which requires additional measures to ensure electromagnetic compatibility in order to preserve the functional safety of the spacecraft.

The aim of this study is to optimize a low-thrust interplanetary trajectory using collinear libration points L1 and L2 as the junction points of the geocentric or planetocentric segments of the trajectory with the heliocentric segment. The problem of optimizing the heliocentric segment of perturbed low-thrust interplanetary transfer is considered in the four-body ephemeris model, which includes the Sun, Earth, target planet, and spacecraft. To optimize the trajectories, an indirect approach is used based on Pontryagin’s maximum principles and the continuation method. The possibility of reducing the characteristic velocity in comparison with the estimates obtained through the method of zero sphere of influence is shown.

This article presents the results of 1000-h tests of a radiofrequency ion thruster (RFIT) with electrodes of an ion-extraction system made of carbon–carbon composite material based on the non-woven carbon frame. The quality of the surface of the thruster IES accelerating electrode being the key element of the RFIT from the lifetime point of view was assessed by visual examination and scanning electron microscopy. The maximum depth of erosion cavity on the accelerating electrode surface was determined. Electrode-surface elemental analysis was performed by the method of electron-probe microanalysis.