Description of the passive magnetic attitude control system of the TNS-0 # 2 nanosatellite is presented. The parameters of the main components of the attitude control system are given and their choice is justified. Using telemetry data, the passive angular motion of TNS-0 # 2 was determined after processing measurements of the on-board sensors after launching from the ISS on August 17, 2017. The damping time of the initial angular velocity was estimated. The accuracy of magnetic stabilization after the end of transient processes was determined.

The paper considers a method of the dose exposure decrease from the charged particles of Earth’s radiation belts (ERBs) affecting a reusable orbital transfer vehicle, which is launched from a low circular orbit into geostationary orbit using a nuclear electric propulsion system. The main idea of the method consists in numerical continuation of a solution of the minimum time problem with respect to an ionizing radiation dose accumulated at the end of a transfer. To do this, equations of motion of the orbital transfer vehicle are supplemented by an additional equation for the radiation dose, and the boundary condition for the dose at the right end is introduced. In calculating the dose, the AE8/AP8 MIN, AE8/AP8 MAX, and AE9/AP9 models of fluxes of charged particles of ERBs were used. By changing the insertion trajectory shape, it became possible to lower the radiation dose by 25–38% relative to the minimum time trajectory. At the same time, the transfer time increased no more than by 7% of the minimum time of launching into geostationary orbit, and the characteristic velocity expenses increased by 320–560 m/s.

Results of spectroscopic studies of electric discharge stages in the channel of an ablative pulsed plasma thruster (APPT) are presented. The initial stage has been shown to affect further discharge development. Shock surface ionization of the propellant has been found to be the dominant mechanism for “parasitic” propellant consumption of an APPT. The results of experimental studies of the dependence of propulsion efficiency and consumption on the distance between operating surfaces of APPT propellant bars are given.

The paper considers the problem of calculating direct and low-energy, low-thrust trajectories to the libration points of the Earth–Moon system and to halo-orbits. A method for solving the problem is proposed. It consists of calculating stable manifolds of libration points or halo-orbits and calculating a low-thrust trajectory from an initial circular Earth orbit to the given point of this manifold using sub-optimal feedback control. With a fixed final mass of a spacecraft, the last stage of calculations is reduced to solving the Cauchy problem. The numerical examples are given for the calculation of direct and low-energy trajectories to libration points and to halo-orbits and the optimization of entry points to stable manifolds for low-energy trajectories.

A method of reducing the radiation power degradation of spacecraft solar array during combined geostationary orbit insertion using a booster and an electric propulsion system is considered in this paper. The essence of the method is to optimize the shape of the transfer trajectory and the perigee argument of the intermediate orbit. The maximum principle is applied to the problem of optimizing the SA electrical power at the end of 15 year spacecraft’s operational lifetime (EOL). For this, the equation for the current SA power and the constraint on this power at the EOL is added to the equations of spacecraft motion. The closed-form solution to the adjoint equation to SA power is obtained. Calculation of SA radiation degradation was carried out using the models of charged particle fluxes of Earth’s radiation belts, AE8 MAX and AP8 MAX. To increase the EOL SA power by 0.16–0.66% of the SA power at the beginning of the transfer depending on the parameters of intermediate orbits. The additional characteristic velocity increased relative to the minimum time trajectories by 13–1087 m/s depending on the parameters of intermediate orbits.

A technique of searching for optimal trajectories with gravity assisted maneuvers (GAMs) for interplanetary transfers of spacecraft (SC) with an electric propulsion system (EPS) is proposed. In this case, the indirect optimization method is used. A distinctive feature of this technique is the combination of optimality conditions at the point of GAMs within a single boundary value problem for two cases, when the height of the flyby hyperbola with the GAM is less than or equal to the maximum one. This approach makes it possible to considerably reduce the volume of necessary calculations in optimizing SC interplanetary trajectories that include GAMs. It considers end-to-end trajectory optimization with an analysis of the full set of optimality conditions at the point of the GAM. The efficiency of the proposed approach is demonstrated by the example of optimization of interplanetary trajectories from Earth to Mercury with a GAM in the vicinity of Venus and from Earth to Jupiter with a GAM near Earth.

The expediency of using the true longitude or the associated angle as an independent variable in the problems of optimizing multi-revolution, low-thrust trajectories of spacecraft is considered. The auxiliary longitude is introduced, which is used further as a new independent variable instead of time. The advantage of using the auxiliary longitude as an independent variable in the problem of optimizing the trajectories with a fixed angular range and optimum transfer time is demonstrated. The problems of optimizing the trajectories of spacecraft with an ideally controlled, limited-power engine and a limited-thrust engine with a constant exhaust velocity are considered. The possibility of improving the convergence of the methods of solving the problems of optimizing the trajectories with a limited-thrust engine is considered. This improvement is achieved by joint utilization of the smoothing approximation of a relay thrust-switching function and the regularizing transformation of an independent variable, which extends the switching function values in the vicinity of thrust switching instants. Based on the results, the method for solving the problems of optimizing multi-revolution transfers is proposed; the numerical examples of its application are presented.

The paper presents the results of the investigation of the possibility to develop a stationary plasma thruster with increased thrust. This research was carried out by the Research Institute of Applied Mechanics and Electrodynamics, Moscow Aviation Institute (RIAME MAI) jointly with Fakel Design Bureau as an industrial partner. During research, we developed and investigated a laboratory model, prototype, and experimental model of the SPT-100VT thruster, whose main dimensions are close to those of the well-known SPT-100 serial thruster, which is produced by the Fakel Design Bureau and is successfully operating in space. As a result, it has been shown that the SPT-100VT thruster can effectively and for a long time operate with a discharge voltage of 300 V and power of up to 3 kW producing thrust more than twice that of the SPT-100 with a thrust efficiency above 60% and specific impulse over 1800 s. The increased thrust is achieved by an increase in the xenon flow rate; the efficiency of the thruster is increased due to the modernization of the thruster magnetic system and geometry of the accelerating channel.

The possibility of implementing the solar probe project by launching a research satellite into the system of heliocentric orbits with a relatively small perihelion radius and a sufficiently large inclination to the solar equator (the inclination of the last heliocentric orbit to the plane of the solar equator should be at least 30°) is analyzed. A comparative design and ballistic analysis of the possibility of using chemical and electric propulsion systems (EP) when launching a spacecraft into the considered system of heliocentric orbits is conducted. The transport systems being analyzed assume the use of the Soyuz-2-1b launch vehicle and the chemical upper stage Fregat when launching the spacecraft from Earth. The propulsion systems of the spacecraft itself are different. In one case, a chemical propulsion system is used, in the other, a solar EP based on one stationary plasma thruster of the SPT-140 type. The time of launching the spacecraft into the last heliocentric orbit of the considered orbit system is limited to 5 years from above. It was shown that using EP can significantly increase the spacecraft mass in operational orbits (from 910 to 1600 kg).

A complete cycle of studies is considered, which are related to obtaining knowledge on spectral and time-response characteristics of emission from stationary plasma thrusters (SPTs), to the development of mathematical and simulation models of such emission, and assessment of its influence on the interference immunity of space communication systems. The reliability of data transmission via Earth–spacecraft digital communication channels is assessed based on the results of simulation modeling. The bit error rate is studied as a function of signal-to-noise and signal-to-SPT-interference ratios. Power loss of the communication channel in the presence of SPT emission relative to the standard system is considered as its interference immunity indicator. Calculations are made for the power budget of deep- and near-space radio links. Power loss is assessed for the Earth–Mars radio links and within the geostationary orbit. Recommendations on reducing the SPT emission effect are presented.

Results of experimental studies of the electron concentration and near-solar plasma velocity dependences on the heliocentric distance obtained in 1970–2012 by the method of radio transillumination by signals of spacecraft for the region of 3–50 solar radii are presented. A comparative analysis of different approximations of these dependences is given and it is shown that joint analysis of the data on the plasma velocity and concentration with allowance for independence of the plasma integral flux on the heliocentric distance makes it possible to find analytical expressions corresponding well to experimental data. Analytical approximations and graphs of dependences of the velocity, concentration, acceleration, force, and kinetic energy on the distance in the regions of solar wind acceleration for low and moderate solar activity are presented.

An approach based on the so-called K-means method has been used to analyze the approximation of gravitational potential of irregularly shaped celestial bodies as the potential of three gravitating balls. Relevant distributions have been constructed for asteroids Bacchus (2063), Kleopatra (216), and Eros (433). The proposed models have been compared with models that are based on alternative approaches [1-3].

Kinetic effects of the dynamics of protons in plasmoids with a non-zero longitudinal (By) magnetic field component in a current sheet (CS) of a geomagnetic tail are considered. The results of modeling proton dynamics and a description of the mechanism of emergence of “north-south” density asymmetry are presented. The mechanism that is possibly responsible for maintaining the longitudinal magnetic field component is described. The obtained parameters are evaluated and the results are compared with observations of the Cluster mission.

An approach based on the so-called K-means method has been used to analyze the approximation of gravitational potential of irregularly shaped celestial bodies as the potential of three gravitating balls. Relevant distributions have been constructed for asteroids Bacchus (2063), Kleopatra (216), and Eros (433). The proposed models have been compared with models that are based on alternative approaches [13].

Algorithms ensuring high-speed and reliable data transmission under lognormal amplitude fluctuations (described by Fraunhofer diffraction) in the spacecraft–ground tracking station line for coherent and incoherent reception of signals have been considered. An advantage of the coherent reception of millimeter-range signals with a random error-correcting code has been indicated.

An approach based on the so-called K-means method has been used to analyze the approximation of gravitational potential of irregularly shaped celestial bodies as the potential of three gravitating balls. Relevant distributions have been constructed for asteroids Bacchus (2063), Kleopatra (216), and Eros (433). The proposed models have been compared with models that are based on alternative approaches [1–3].

The paper theoretically studies the process of formation of volume charge in a spacecraft body under action of electron fluxes of Earth’s radiation belts and other cosmic particles, as well as the generation of induction currents caused by geomagnetic variations. The BLITS and BLITS-M satellites, which have a spherical configuration manufactured entirely of dielectric materials as well as the metal WESTPAC satellite, are considered. The relative simplicity of the form of a dielectric satellite allows one to obtain an analytical solution to the problem and calculate the distribution of fields and charges inside and on its surface. This solution shows that after the establishment of the stationary mode, charges accumulate mainly in a narrow layer near the satellite surface. According to the obtained estimates for low orbits, the electric field strength in this layer is below the threshold value corresponding to the breakdown of uniform dielectric in laboratory conditions. Nevertheless, one can expect the appearance of local electrical breakdowns and microdestructions in the dielectric during increasing solar activity accompanied by an increase in cosmic-ray fluxes. From this point of view, space experiments, in which dielectric destruction was observed during long-term exposure to radiation, can be interpreted as the result of accumulation of microdestructions similarly to fatigue destruction of materials under long-term load. For passive satellites similar to WESTPAC manufactured of conductive materials, the magnetic moment of induction currents in the spacecraft body is estimated. According to these estimates, the interaction of this moment with the geomagnetic field leading to the precession of the axis of the satellite’s rotation is most significant for Pc5 geomagnetic pulsations.

Climatic characteristics of the Crimean peninsula were analyzed on the basis of climatic parameters measured every day onboard NASA satellites for the period of 1983−2005. An accelerated increase in the insolation of Earth’s surface at the Kara-Dag site (as compared to that at other sites in Crimea) and an abnormal decrease in the surface temperature at the Fonar site have been revealed. The results of frequency−time wavelet analysis of local data on the insolation and temperature of the terrestrial surface enabled us to isolate periodic oscillations, the periods of which coincide with astronomical cycles, and determine an era of anomalous deviations from general trends of 1990−1995. From the analysis of corrections for vertical deformations of Earth’s surface at the Katsiveli site, a seasonal component was detected and parameters of this oscillation were calculated. Coherent variations with periods of approximately 10−12 and 60−70 years, which were found here, are inherent in the processes of different physical nature. They can be attributed to the global cycles in the Solar System, the manifestation of which in climatic and geophysical local processes is a result of the general trend towards synchronization. In this process, coherent variations of multiple or commensurate frequencies, may appear.

This study shows that oxygen atoms can be released from a crystal lattice of silicon dioxide in the lunar regolith as parts of silver hydroxide molecules. In turn, silver hydroxide can relatively easily react with hydrogen to generate water and silver. This means that the formation of water molecules involved in near-surface lunar soil is possible. The presence of water molecules in lunar soil can affect the photoelectric properties of the lunar regolith and the parameters of the dusty plasma system over the Moon.

The paper presents the results of the calculation of cosmic ionizing radiation doses for conditions that simulate ISS crew quarters outfitted with additional protection in the form of a “protective shutter” kit used in the Matryoshka-R experiment; the kit is filled with high-pressure polyethylene instead of regular means of cosmonaut personal hygiene (wet wipes and towels). The calculation was performed using the ray tracing method, which was previously verified using the experimental data on the dose loads in the space station sections. The protective shutter kit has a thickness of 10 cm and covers the exterior wall of the crew quarters; its volume is filled with high-pressure polyethylene. The calculation is performed for the locations of experimental dosimetric assemblies for typical ISS orbits in the phase of minimum and maximum solar activity. The effect from the use of the additional protection (AP) in the ISS crew quarters in terms of the equivalent dose varies between 22 and 60% depending on the orbital altitude, phase of solar activity cycle, and initial conditions of protection. The results indicate a higher efficiency of the additional protection of polyethylene in comparison with water-containing materials. This AP from cosmic radiation has the potential to be used in long-term and long-range spaceflights.