Spacecraft Mass Research Articles

Feasibility studies for a dust observatory between earth and the asteroid belt

Cosmic dust transports information about the composition and evolution of distant realms over space and time. Dust sources may be in our local neighbourhood, such as the surfaces of moons and small bodies, or further away, in the galactic environment. The interior of moons like Io or Enceladus are well known dust sources in our solar system. Dust grains are therefore like probes, giving us the opportunity to investigate these objects at a distance. Thus, a new field of activity has been born: Dust Astronomy. The ultimate tool for Dust Astronomy is a dust observatory which was never been flown so far. Until today, all our knowledge is based on smaller individual dust detectors flown on various interplanetary spacecraft. Recently, a new generation of dust telescopes, which can be combined to a full dust observatory, became available. The scientific goal of such an observatory is to characterise our micrometeoroid environment with high precision and sensitivity and to record an inventory of various dust populations from asteroidal, cometary and interstellar dust sources.Two competitive studies on a Dust Astronomy mission were performed as part of a master’s course, including mission analysis and spacecraft design. The goal was to demonstrate the feasibility of an observatory reaching the main asteroid belt with a minimum spacecraft mass using electric propulsion. Additional secondary payloads supporting the primary objective were selected as well. The results prove that small probes with less than 700kg mass are capable of placing an array of dust telescopes deep in the asteroid belt, leading to a huge leap in Dust Astronomy and our knowledge about the local dust environment.

Optimum design of nuclear electric propulsion spacecraft for deep space exploration

Future space exploration is now focused on new field-deep space planets. Deep space exploration and the development and utilization of resources are inestimable strategic significance for the country to seize the initiative and command heights of deep space exploration. Nuclear electric propulsion (NEP) systems convert heat from the fission reactor to electrical power and then use the electrical power to produce thrust. Compared with traditional propulsion technology, the NEP system with variable Isp can be used for several applications. The NEP spacecraft is more suitable for deep space exploration missions due to the advantages of high specific impulse, high power, and long life. This paper analyzes the relationship between the transfer travel time, specific mass, power, and payload ratio of the NEP spacecraft through simple performance models. Use the mass optimization and specific mass optimization models based on the NEP system composition and small thrust orbit theory to maximize the payload ratio for a given transfer time and technological characteristics. Finally, applied this NEP model to discuss the feasibility of Mars, Jupiter, and Saturn transfer missions and compared it with the Tianwen-1, Juno, and Cassini–Huygens spacecraft, respectively. Results show that when the NEP spacecraft specific mass reaches 4.77 kg/kW, the Earth–Mars transfer time can be changed to 331.31 days, the payload increases to 1650 kg, the transfer time for Jupiter and Saturn mission would be shorted to 661.31 days and 1131.31 days, the corresponding payload significantly increases which achieved to 1270 kg and 2981 kg. The nuclear electric propulsion spacecraft dramatically improves the detection capability of the spacecraft and provides a reference for the feasibility demonstration and subsequent design of the deep space and the extrasolar planet exploration.

New LISA dynamics feedback control scheme: Common-mode isolation of test mass control and probes of test-mass acceleration

The Drag-Free and Attitude Control System is a central element of LISA technology, ensuring the very high dynamic stability of spacecraft and test masses required in order to reach the sensitivity that gravitational wave astronomy in space requires. Applying electrostatic forces on test masses is unavoidable but should be restricted to the minimum necessary to keep the spacecraft-test masses system in place, while granting the optimal quality of test-mass free fall. To realize this, we propose a new test-mass suspension scheme that applies forces and torques only in proportion to any differential test mass motion observed, and we demonstrate that the new scheme significantly mitigates the amount of suspension forces and torques needed to control the whole system. The mathematical method involved allows us to derive a new observable measuring the differential acceleration of test masses projected on the relevant sensitive axes, which will have important consequences for LISA data calibration, processing, and analysis.

A multiple-vehicle strategy for near-Earth asteroid capture

The use of asteroid resources could benefit future space missions. Instead of lifting the necessary materials off the Earth, they can be then sourced directly in space. Asteroid capture missions aim to bring asteroids closer to Earth, where they can be permanently accessed. This paper introduces a new strategy for asteroid capture missions, where two spacecraft are used for capturing near-Earth asteroids. These spacecraft act together as a ‘pitcher’ spacecraft and ‘catcher’ spacecraft, where the pitcher spacecraft hops from asteroid to asteroid and deflects them towards an orbit in the vicinity of Earth, while the catcher spacecraft is stationed at the Earth and captures the incoming asteroids. This novel two-spacecraft strategy is compared to a conventional one-spacecraft strategy using three analyses; a preliminary analysis using coplanar and circular orbits to define the problem, a statistical analysis using fictional near-Earth asteroids to obtain a large set of data, and an analysis where real near-Earth asteroids are used as mission targets. A mass model is developed to compute the retrieved asteroid mass per unit of initial wet spacecraft mass for missions aiming to retrieve multiple asteroids. Results show that the two-spacecraft strategy is capable of returning more asteroid mass and often at a shorter mission duration.

Drag-free design based on Embedded Model Control for TianQin project

With the ability to detect low-frequency gravitational waves (GWs), space-borne detectors will play an important role in exploring the universe in the future. The TianQin project proposed in China participates in this challenge by aiming to detect millihertz GWs. The TianQin GW detector consists of three spacecraft forming a triangular constellation. Each spacecraft carries a pair of test masses (TM) which, inside shielding cages, are left to free fall along the local geodesic. In this way, a pair of TMs on two separate spacecraft can be each other aligned to become the gravitational references of the GW detector, made by an inter-satellite laser interferometer. Each free-falling TM is limited by parasitic forces to be appropriately bounded. The relevant acceleration bound has been fixed to 10−15 m/s2/Hz within the TianQin measurement bandwidth (MBW) ranging from 0.1mHz to 0.1Hz. In turn, TM to cage fluctuations must be kept below 4nm/Hz to limit the stiffness coupling with the spacecraft displacement caused by non-gravitational disturbances, such as solar radiations and the thruster noise. Suppression of such disturbances calls for a challenging drag-free control technology, since the center of mass (CoM) of a single spacecraft cannot track the separated CoMs of two TMs simultaneously, and consequently tracking must be limited to three degrees of freedom (DoF): two non-orthogonal sensitive axes (one for each TM) and the perpendicular direction to their plane. Control design and simulated tests of this paper will be restricted to the drag-free control along the sensitive axes. The remaining nine DoFs of the two TMs (position and attitude) are controlled by electrostatic suspensions, not to be treated here, but accounted for in the simulated trials. The design of the two-DoF drag-free control relies on a model-based control methodology, the Embedded Model Control (EMC), capable of predicting and suppressing unknown disturbances within the required bandwidth and of decoupling the TM dynamics along the non-orthogonal sensitive axes. The paper starts with the nonlinear model of the TM to cage dynamics, followed by the relevant EMC design, restricted to the sensitive axes. Numerical simulations are employed to validate closed-loop performance and robust stability. Simulated results show that the residual TM to cage fluctuations can be kept below 3nm/Hz, which leaves a margin within the required spectral bound. The EMC methodology discussed in the paper can provide a reference for future developments and implementations.

Event-triggered adaptive terminal sliding mode tracking control for drag-free spacecraft inner-formation with full state constraints

This paper investigates the reconstruction and maintenance of the inner-formation system for an innovative drag-free spacecraft with double spherical test masses in consideration of unknown model uncertainties, external disturbances, and actuator saturation. First, spherical test mass 1 is prescribed as the ‘leader’ of the inner-formation system, which is freely floating in the nearly pure gravitational orbit without any control. The nonlinear relative orbit dynamics of the outer spacecraft and spherical test mass 2 with respect to spherical test mass 1 is deduced and the micro configuration of the inner-formation system is introduced. For the outer spacecraft, the aim is to track the spherical test mass 1 in real-time, such that both the center of masses coincided with each other. For the spherical test mass 2, the goal is tracking the reference trajectory, which is designed as a space circular fly-around spherical test mass 1 based on the Clohessy-Wiltshire equation. The state constraints caused by the limit cavity size are analyzed. Second, to achieve these goals with satisfied performance, an event-triggered adaptive terminal sliding mode tracking control strategy is developed for the inner-formation. The adaptive terminal sliding mode technique is implemented to overcome the model uncertainties and external disturbances and to accomplish the reconstruction process in a finite time. Two kinds of barrier functions are utilized to tackle the state constraints. The relative threshold event-triggered mechanism is embedded to save the computation source onboard, evidently. The simulation results demonstrate the efficacy and superiority of the proposed method.

МОЖЛИВОСТІ ЗАСТОСУВАННЯ СОНЯЧНОГО ВІТРИЛЬНОГО ПРИСТРОЮ ДЛЯ ВІДВЕДЕННЯ КОСМІЧНИХ ОБ’ЄКТІВ ІЗ НАПІВСИНХРОННИХ ОРБІТ

The current state of the near-Earth space environment poses a threat to spacecraft moving not only in low and geostationary but also in semi-synchronous orbits. The main methods that can be used in these orbits for space debris removal are the active method using a rocket propulsion system and the passive method using a solar sail device. The last method is of particular interest because of the low overall mass and mass characteristics of the removal system, as well as the availability of unlimited solar light radiation, which can be converted into a driving force. Solar sails are currently undergoing experimental testing in the near-Earth space, but they have not been used for medium and high orbits. Two constructive variants of solar sails construction are considered: frame and rotor. The article is devoted to the evaluation of the possibility of removing space objects from semi-synchronous orbits using a solar sail. Two variants of the removal scheme are considered: into a circular disposal orbit 200 km below the target orbit and into the dense layers of the Earth's atmosphere. A computer simulation of the removal is performed using the example of the Galileo global positioning system spacecraft with a frame-type and rotor-type solar sail. The dependence of the orbit altitude at the ending of removal on the solar sail area was obtained. The values of the solar sail area required to provide for the maneuvering and the weight of the structure were determined. It was demonstrated that the total mass of the frame-type solar sail is 9 kg (1.3% of the spacecraft mass) and that of the rotary-type solar sail is 11 kg (1.6%). As applied to disposal into dense layers of Earth's atmosphere for frame-type solar sail - 52 kg (7.4%), for rotary-type solar sail - 25.5 kg (3.6%). It follows from the obtained results that for moving to the disposal orbit 200 km below the target orbit it is more advantageous to use a frame-type solar sail, and in the dense layers of the atmosphere - the rotor type.

Design Considerations for an Earth–Mars Cycler Spacecraft Using the S1L1 Cycler

This paper discusses the design considerations for an Earth–Mars cycler spacecraft using the ballistic S1L1 cycler trajectory. The primary objective is to estimate the mass of a cycler spacecraft designed to transfer a crew of six from Earth to Mars on a nominal 154-day trajectory, with sufficient consumables and life-support capabilities to accommodate the crew on a 1032-day contingency mission, in case the transfer from the cycler to Mars is not possible. The paper develops the functional system requirements from the concept of operations and discusses the bottom-up mass and power estimates for the individual subsystems. The cycler spacecraft design includes system capabilities that are not needed for spacecraft designed for point-to-point interplanetary transfers, such as a 30-year design life and the capability to be robotically maintained, repaired, and replenished during an uncrewed Earth–Earth resupply loop. The resulting spacecraft design is similar in size, mass, and operational complexity to deep space transfer vehicles for point-to-point mission architectures. However, as a minimum number of two cyclers are needed for round-trip missions with acceptable duration, a cycler-based architecture is less suitable for initial Mars exploration than for a sustained, continuous Mars exploration campaign.

Analysis of characteristics of electric propulsion systems intended for carrying out maneuvers of maintenance of low Earth working orbit of small satellites

An analysis of the mass of the working fluid and motor operating time of electric propulsion systems applied as a part of small spacecraft to carry out maneuvers of maintenance of the low Earth working orbit is carried out. The analysis is carried out for the small spacecraft with the weight in the range from 300 to 1000 kg functioning in working orbits with the height in the range from 400 to 600 km. When carrying out the analysis the values of the specific impulse of the propulsion system in the range from 800 to 1600 sec were accepted. Procedural guidelines for assessing the value of the required characteristic speed depending on the aerodynamic drag force, as well as for assessing the value of mass of the working fluid with account for the value of the specific impulse and defining the motor operating time of the propulsion system depending on the exhaust speed of the working fluid were used. The results of calculations given in the article show that the mass of the working fluid and the motor operating time vary depending on the height of the orbit and the mass of the small spacecraft and allow making quick preliminary assessment of the main design characteristics of the electric propulsion engines used to carry out maneuvers of maintenance of the low Earth working orbit of small spacecraft with different weight dimension characteristics during the prescribed term of active existence.

Air-breathing electric propulsion: Flight envelope identification and development of control for long-term orbital stability

Air-breathing electric propulsion (ABEP) enables long duration missions at very low orbital altitudes through the use of drag compensation. A system-level spacecraft model is developed, using the interaction between thruster, intake and solar arrays, and coupled to a calculation of the drag. A quadratic solution is found for specific impulse and evaluated to identify the thruster performance required for drag-compensation at varying altitudes. An upper altitude limit around 190 km is based on a minimum thruster propellant density, resulting in required thruster performance values of Isp>3000s and T/P>8mN/kW for a realistic ABEP spacecraft. The orbit of an air-breathing spacecraft is propagated with time, which highlights the prescribed orbit eccentricity due to non-spherical gravity and therefore an increased variability in the atmospheric conditions. A thruster control law is introduced which avoids a divergent altitude behaviour by preventing thruster firings around the orbit periapsis, as well as adding robustness against atmospheric changes due to season and solar activity. Through the use of an initial frozen orbit, thruster control and an augmented T/P, a stable long-term profile is demonstrated based on the performance data of a gridded-ion thruster tested with atmospheric propellants. An initial mean semi-major axis altitude of 200 km relative to the equatorial Earth radius, a spacecraft mass of 200 kg, Isp=5455s and T/P=23mN/kW, results in an altitude range of around 10 km at altitudes of 160–183km during a period of medium to high solar activity.

Combined System–Trajectory Design for Geostationary Orbit Platforms on Hybrid Transfer

A novel methodology for a combined systems–trajectory optimization for a geostationary equatorial orbit (GEO) platform is proposed to obtain comprehensive design solutions. A combined chemical–electric propulsion system is used to execute hybrid high-thrust/low-thrust trajectory transfer to GEO, thereby balancing the overall system mass and transfer time. A systematic and payload-centric mission design provides a new set of design options to deliver tailored solutions to customized payloads. The hybrid trajectory characterization and spacecraft systems design find the required platform launch mass to deliver a GEO platform with a defined final mass and operational power. Elements of the system design are combined with those of multispiral low-thrust trajectory optimization as well as radiation absorption and solar array degradation to provide a comprehensive design solution. The result is a wide set of solutions to reach GEO, where fully chemical and fully electric transfers represent the boundaries of the hybrid transfer trade space. A payload throughput power of 20 kW entails a spacecraft mass in GEO between 4000 and 4550 kg, an initial thrust-to-mass ratio range of , and a cover-glass thickness between 4 and 24 mils to guarantee a minimum end-of-life/beginning-of-life power ratio of 85%. In addition, all-electric solutions from different injection orbits yield transfers to GEO with a time of flight of 60–150 days and an initial mass for the platform of 4400–5500 kg.

Управляемый тензор инерции космического аппарата трансформируемой конструкции

A simplified model of a transformable spacecraft is considered, including a rod-type transformation mechanism with movable weights. The mechanism can be used to adapt the dynamic properties of the spacecraft to the environment or the operating conditions of on-board systems, for example, to counter the moments of external disturbances during attitude control and angular stabilization. By changing the position of the transformation mechanism, the spacecraft inertia tensor can be put in diagonal form, which makes it possible to exclude the force interconnections between the channels and to eliminate the constant component of the gravitational moment. For a simplified model of the transformation mechanism, we establish the analytical dependence of the components of the inertia tensor on the parameters determining the position of the transformation mechanism. It is shown that by adjusting the moving mass, which is 0.5% of the entire spacecraft mass, we obtain the spacecraft configuration that ensures the diagonality of the inertia tensor.

Concurrent Learning Adaptive Finite-Time Control for Spacecraft With Inertia Parameter Identification Under External Disturbance

This article investigates the postcapture control problem for the combined spacecraft formed after a service spacecraft captures a noncooperative target by manipulators. Two concurrent learning adaptive finite-time controllers are developed to achieve trajectory tracking and inertia parameter identification simultaneously. As a very common uncertainty affecting the spacecraft motion, the external disturbance should be considered, but it is ignored in most studies identifying inertia parameters. In this article, the presented controllers are designed according to the specific knowledge of the external disturbance: one controller is designed based on the disturbance bound that is assumed to be known; the other identifies the disturbance parameters together with the inertia parameters, presupposing that the nonlinearities involved in the external disturbance are known. Both of the two controllers require no a priori knowledge of the mass and inertia matrix of the combined spacecraft and take the input saturation into account. Sufficient conditions to identify inertia parameters are also given with no need of persistent excitation. The convergence of the closed-loop system is proved within the Lyapunov framework. Numerical results show that the proposed controllers can track the desired position and attitude trajectories in finite time, and, at the same time, the inertia parameters converge to their real values.

Small spacecraft-based space system for real-time Earth surface monitoring

The paper presents basic results of exploratory design studies into a space system for real-time monitoring of the Earth surface based on small spacecraft* for Earth remote sensing, which provides acquisition in real time of highly detailed Earth surface images with resolution of 0.7–1.0 m, a swath of 20 km and 15m in-plane georeferencing accuracy from a ~300 km orbit with a mass of spacecraft ~200 kgf repeating with high frequency (at least every 1.0–1.5h). The said parameters are in line with the current state of the art and are quite feasible for our country’s industry. _________________________________________ * — for the purposes of this paper ‘small spacecraft’ is a 150-220 kg spacecraft. Key words: small spacecraft, space system, Earth remote sensing, satellite bus, payload module, electro-optical equipment, multipurpose ground facility, combined ground station, services of space monitoring of the Earth surface.

SMALL SPACECRAFT-BASED SPACE SYSTEM FOR REAL-TIME EARTH SURFACE MONITORING

The paper presents basic results of exploratory design studies into a space system for real-time monitoring of the Earth surface based on small spacecraft* for Earth remote sensing, which provides acquisition in real time of highly detailed Earth surface images with resolution of 0.7–1.0 m, a swath of 20 km and 15m in-plane georeferencing accuracy from a ~300 km orbit with a mass of spacecraft ~200 kgf repeating with high frequency (at least every 1.0–1.5h). The said parameters are in line with the current state of the art and are quite feasible for our country’s industry. _________________________________________ * — for the purposes of this paper ‘small spacecraft’ is a 150-220 kg spacecraft. Key words: small spacecraft, space system, Earth remote sensing, satellite bus, payload module, electro-optical equipment, multipurpose ground facility, combined ground station, services of space monitoring of the Earth surface.

Large-scale flexible spacecraft mass properties determination using torque-generating control

It is necessary to determine the mass properties of an on-orbit large-scale flexible spacecraft in order to achieve high-precision attitude control. The mass properties of the spacecraft include the inertia, the center of mass, and the mass. Research in this area had been mostly focused on rigid spacecrafts. The coupling between the rigid body of the spacecraft and the large flexible appendages on it, however, could have significant effect on the mass properties determination. This article proposes a momentum conservation–based determination method that takes into consideration the rigid–flexible coupling factors of large-scale flexible spacecrafts. First, a control method is designed to ensure that the spacecraft stays motionless after the motivation incurred for mass properties determination. Second, an inertia matrix determination method is developed using a Kalman filter, in which the coupling factors are added in the amendment procedure. Third, the Kalman filter with input is applied in the determination of the center of mass: on the one hand, this method can use one the accelerometer placed at the flexible appendages, if the deformation can be measured; on the other side, the method can use three accelerometers placed at three orthogonal points of the rigid part of the spacecraft, when the deformation cannot be measured. Finally, the mass can be gained by estimating the center of mass twice. Simulations were carried out on a large-scale rigid–flexible coupling spacecraft. The results demonstrated that the determination errors in all cases are less than 10%, which meet the engineering requirements.

Parametric Analysis of High-Delta-V CubeSat Missions with a Miniature Ion Thruster

The increasing capabilities of the CubeSat platform have led to growing interest in performing more complex commercial and science missions with these miniature spacecraft. In particular, electric propulsion provides unprecedented mission and enables ambitious but low-cost Earth missions as well as lunar, asteroid, and interplanetary exploration. As a case study, a 6 U (where U represents a “unit”) CubeSat using the Miniature Xenon Ion Thruster was designed for a notional mission with a 2 kg, 1.6 U payload, resulting in a spacecraft wet mass of 11.9 kg and a burn duration of 15 months. This spacecraft is capable of up to for a 0.5 U payload, as well as for a 2 U payload. Parametric analyses with generalized electric thruster properties show that mission performance is sensitive to thruster and total efficiency , with a 10% increase in efficiency resulting in a 16% decrease in burn time. This decrease in burn time is also possible by decreasing , but it incurs mass and payload volume penalties. Parametric studies of neutralizer cathode properties show that neutralizer cathode selection is critical, and that such cathodes should be designed to require low power, low to zero flow rate, and long life for high- missions.

Анализ ключевых технологий создания многоспутниковых орбитальных группировок малых космических аппаратов

One of the key areas of modern world cosmonautics is the development of cluster space systems for various purposes, consisting of a large number of functioning spacecraft. This became possible due to a decrease in the mass of spacecraft due to the creation and use of new materials, the development of electronics and microelectromechanical systems, the use of the group launch method, the development of multi-agent technologies and inter-satellite communication systems. There are projects of systems consisting of a large number of space objects, such as OneWeb, Planet, Starlink, Satellogic, etc. The main classes of devices used to create such multi-satellite systems are small satellites, including the number of micro (up to 100 kg) and nano (up to 10-15 kg) classes, which have significant advantages over heavy spacecraft, especially in terms of the timing and cost of their creation. The deployment of multi-satellite constellations, including hundreds and thousands of satellites, requires fundamentally new approaches to the creation of spacecraft and the system as a whole at all stages of the life cycle. The article discusses the key technologies used to create multisatellite orbital constellations based on small satellites at different stages of the life cycle - from the early stages of design to the stage of operation and disposal (information from orbit). The experience of a joint project of Samara University and the Progress Rocket and Space Center on the creation of a constellation of small spacecraft of the AIST series is presented.

Rigid-flexible coupling effect on attitude dynamics of electric solar wind sail

This paper investigates the modelling of rigid-flexible coupling effect on the attitude dynamics and spin control of an electric solar wind sail (E-sail) by developing a rigid-flexible coupling dynamic model. The model considers the attitude dynamics of the central spacecraft, the elastic deformation of the tethers and the rigid-flexible coupling between the spacecraft and the tether. The attitude and translation dynamics of the central spacecraft is described by the natural coordinate formulation, while the tether deformation is described by the high-fidelity nodal position finite element method. The latter enables a natural coupling between the motion of the flexible tethers and the rigid-body dynamics of the central spacecraft at the anchor points where the tethers connected to the spacecraft by Lagrange multipliers. Based on the model, the influence of the rigid-flexible coupling, E-sail orientation and geometrical configuration on the dynamic characteristics of the E-sail is investigated by a parametric analysis. It is found that the deformation motion of flexible tethers will cause the offset of centres of mass and thrust of E-sail, which generates disturbance torques on the central spacecraft. Through the nonlinear rigid-flexible coupling, the disturbance causes the tension fluctuations and the undesired fluctuations of the E-sail's attitude and spin rate. The parametric analysis indicates that the E-sail is more stable if the spin plane passes the centre of mass of the central spacecraft. Finally, the controllability of E-sail spin rate is investigated by applying simple feedback torque controls at the central spacecraft or at the central spacecraft and the remote units simultaneously. The analysis demonstrates the spin rate cannot be controlled by the central spacecraft along due to the rigid-flexible coupling and must be controlled at the remote units with finite control input.

Method of providing energy resources to a circumlunar satellite for passing longduration shadow zones

Researches of the possibilities of creating a lunar navigation and communication system have revealed features of the sun's illumination of near-lunar orbits: the annual passage of orbits through a long cycle of about 55 days, containing shadow sections of the orbit from the Moon and a pause of up to 185 days with possible overlays of penumbra and shadow sections of the orbit from the Earth. At the same time, the penumbra and shadow zones from the Earth near the Moon have significant differences in size and can cover the entire orbital grouping of the lunar navigation and communication system. As a result, the problem arises of calculating the energy resources of a circumlunar satellite for passing long-duration shadow zones and maintaining the spacecraft's thermal regime for the period of passing the shadow section of the orbit from the Earth. The energy balance of the spacecraft is modeled for various modes of its operation during the passage of long-duration shadow zones: regular, duty, storage. It is proposed to use blinds to regulate the heat radiation from the spacecraft radiators when passing the shadow sections of the orbit in the considered modes. To reduce the cost of spacecraft mass it is recommended to switch the spacecraft to storage mode, which reduces power consumption in comparison with standby mode. The efficiency of the proposed scheme for passing the shadow zones of the lunar navigation and communication system spacecraft was evaluated for three spacecraft sizes according to the criterion – the minimum amount of mass spent on increasing the energy intensity of the battery and installing blinds.