Small Space Vehicles Research Articles

Analysis and improvement of characteristics of power supply systems for the small space vehicle of the AIST-2 type

Analysis and improvement of characteristics of power supply systems for the small space vehicle of the AIST-2 type

Selection of a Neural Network Architecture for Implementation of Cognitive Functions of Network Software for Control of a Group of Interacting Small Space Vehicles

The purpose of the research is to substantiate and select the architecture of a neural network for the possibility of implementing the cognitive functions of network software for controlling a grouping of interacting small spacecraft.Methods are based on the concepts of AI theory for managing the grouping of small spacecraft - the use of adaptive methods and tools that allow making decisions, similar to the mechanisms of human thinking. With regard to space communication systems with a heterogeneous structure, AI methods and technologies are aimed at the processes of predicting the state in communication channels between network nodes and automatic reconfiguration of the network of devices based on the learning processes of a neural network (NN).Results. In the learning and forecasting mode, it is necessary to use time series of parameters and coordinates of specific pairs of small spacecraft with non-zero line of sight. Especially for time series analysis, recurrent neural networks (RNN) are used, in particular, LSTM. The idea of RNN operation is to use as input data for the current forecast not only the state vectors of the SVs and their coordinates, but also the previous value of the communication quality, actual or predictive. The paper shows that the onboard computing power of a separate MSC does not allow performing forecasting and training on board. Therefore, a dedicated ground segment of forecasting and monitoring is required, which will collect a posteriori information, periodically train the cognitive model, use it to predict the quality of communication, and transmit the results to the network nodes to build data transmission routes.Conclusion. The analysis of modern solutions and the choice of neural network architecture for the implementation of the cognitive functions of the network software for controlling the grouping of interacting small spacecraft showed that the neural networks of the Transformer architecture, which are based on the mechanism of internal attention, most fully meet the requirements of the project. The Transformer architecture allows using the entirety of a priori data, has a high learning and forecasting speed.

Improving the technology for ensuring the magnetic cleanliness of small spacecraft

The object of this study is the technology of ensuring magnetic cleanliness when applied to small spacecraft (SC) with overall dimensions of ~1 m in terms of minimizing the effect of magnetic interference on the on-board magnetometer. As part of solving the general problem of increasing the reliability of spacecraft, the task of improving the technology of magnetic cleanliness is considered. It is shown that the use of the multi-dipole model for the calculation of ~50 nT magnetic interference is limited in the design of small space vehicles that have an increased installation density. The expediency of using a model of its spherical harmonics instead of a multi-dipole model to represent the magnetic field of satellite components was theoretically justified. For the practical application of the model, the budgeting of the projections of the magnetic induction of hindrance to the onboard magnetometer according to the experimentally determined coefficients of the dipole, quadrupole and octupole spherical harmonics is proposed. An algorithm for calculating the coordinates of the location and projections of the magnetic moment of the dipole source inside the satellite based on the results of measuring the coefficients of spherical harmonics of its components is proposed. The possibility of representing the coefficients of the spherical harmonics of the magnetic field of the entire satellite in the form of the sum of the listed corresponding coefficients obtained during field measurements of the components is theoretically justified. It is proposed to use the difference between the calculated and measured values of the corresponding coefficients as a criterion for the quality of the work to ensure magnetic cleanliness. If the proposed procedure is used in the space industry, the quality of ensuring the magnetic cleanliness of small space vehicles can be increased, thereby improving the reliability of their operation in orbit

Postflight Recovery of the Rotational Motion of a Small Space Vehicle from Solar Sensor Information

Postflight Recovery of the Rotational Motion of a Small Space Vehicle from Solar Sensor Information

Postflight Recovery of the Rotational Motion of a Small Space Vehicle from Solar Sensor Information

The use of the developed technology for solving inverse problems of parametric identification of the rotational motion of a spacecraft (SC) based on the accumulated sample of measurements of various compositions is expanded. On the example of processing telemetry data (measuring the direction vector to the Sun) of the BugSat-1 small spacecraft (SSC), the possibility of determining not only the characteristics of rotational motion but also its inertia coefficient is demonstrated. The rotational motion is determined based on the data obtained for five time intervals from January to May 2022, during which BugSat-1 made an uncontrolled movement around the direction vector toward the Sun. This type of motion does not allow applying traditional approaches to determine the dynamics of its rotation. In the process of data processing, a simplified model of the rotation of a SSC is used. At the same time, the initial conditions of rotational motion are estimated—the angular velocity vector and orientation angles, which made it possible to draw a conclusion about the precessional motion of a SSC. Using the analytical model of a regular precession, it is possible to estimate the inertia coefficient of a SSC. The results of flight information processing are presented and the errors of the obtained results are estimated.

ХОЛЛОВСКИЙ ДВИГАТЕЛЬ ST-25 ДЛЯ МАЛЫХ КОСМИЧЕСКИХ АППАРАТОВ

Results of development and experimental tests of Hall Thruster ST-25 are presented. This thruster is intended for application on small space vehicles. With the purpose of decrease of input electric power in the magnetic system of thruster a permanent magnet is applied. In the process of the thruster development his laboratory tests were conducted with the laboratory prototypes of electric power supplies. Final tests were conducted with the flying prototype of discharge power supply. Laboratory tests have confirmed the rightness of the technical decisions stopped up in the construction of Thruster. Methodology of the thruster ST-25 start at the use of preheated cathode and flying prototype of the discharge power supply were improved. Possibility of application of such type of Hall Thrusters on space vehicles with the size of on-board electric power of that is limited to the size 300 - 500 W were confirmed.

State of research and development of small spacecraft with optical communication

The current state of scientific and technical research and development of small space vehicles using optical communication systems is considered. The problems standing in the way of creating intersatellite optical communication are discussed as well as the mutual position of satellites determination and it tracking. Scientific tasks for the implementation of the development of an optimal reliable control system for satellite communication are defined.

HALL-EFFECT THRUSTER ST-25 WITH PERMANENT MAGNET

Results of development and experimental tests of Hall-effect Thruster ST-25 are presented. This thruster is intended for application on small space vehicles. With the purpose of decrease of input electric power in the magnetic system of the thruster a permanent magnet is applied. The permanent magnet is located at the base of the central magnetic core. The permanent magnet provides most of the magnetic induction in the accelerating channel of the thruster, while the side magnets provide precise magnetic field adjustment. Laboratory testing of the ST-25 thruster were realized with application of the preheated hollow cathode which is operated in auto mode at the discharge current 0.5 … 0.9 A. In the process of developing the thruster, its laboratory tests were carried out with a laboratory prototype of the Xenon system for storing and supplying a working substance. The laboratory supply system provided the supply of the working gas to the anode unit 0.65 ... 0.90 mg /s and 0.07 ... 0.10 mg / s to the hollow cathode. The characteristics and parameters of the thruster were obtained using a discharge power supply that has the properties of a voltage source and a power source. The discharge power supply is the part of the flight prototype of the power processing unit. Laboratory tests have confirmed the rightness of the technical decisions stopped up in the structure of the thruster. Methodology of the thruster ST-25 start at the use of preheated cathode and flying prototype of the discharge power supply were improved. Possibility of application of such type of Hall Thrusters on space vehicles with the size of on-board electric power of that is limited to the size 200 - 300 W were confirmed.

Boundary Layer Calculation in Gas Flows in Micronozzles

Tasks performed by small space vehicles (SSVs) require orbit and position correction of the space vehicle in outer space. Strict requirements for SSV precision position control necessitate the development of microengines equipped with micronozzles to ensure ultra-small thrust values (about 1 mN). A distinguishing feature of the gas flow in micronozzles is the decisive effect of the boundary layer. The paper introduces a method for calculating the laminar and turbulent boundary layer in micronozzles. The increased non-uniformity of the velocity in the flow cross-section and the increased pressure gradients result in the unreliable gas-dynamic calculation for micronozzles using potential-flow methods. The paper shows the practicability of applying the calculation method based on the flow separation into an isentropic core and a viscous boundary layer with the use of the boundary-layer momentum integral equation. When calculating the boundary layer, the methods of the theory of viscous fluid motion are used, and when calculating the potential flow, the methods of the potential flow theory are used. Considering that, due to the extremely small size of micronozzles, it is possible to reliably determine by experiment the integral parameters of gas and not the local ones, the comparison of the calculation results with the experimental results shows their good convergence.

Review of electric thrusters with low consumption power for corrective propulsion system of small space vehicles

This paper presents an overview of current world developments in the field of low-energy plasma propulsion systems for small micro - and nanoclass spacecraft. A promising direction of thrusters for devices of nanoclass are propulsion ion propulsion. The use of electrical energy and high-frequency electromagnetic radiation eliminates the problem of significant power losses, because at this level, the laws of optics (reflection and focusing of waves by metal surfaces and dielectrics) and electrical conductivity (energy transfer through metal conductors) work. This fact is an advantage of plasma installations over thermal ones.

Investigation of the technology of mutual navigation and orientation of small space vehicles flying in formation

The paper presents the problem of ensuring support of the flight of a group of small spacecraft (microsatellites) taking into account the small mutual distances between them. The purpose of using the orbital constellation specified is to create a radio communication system to control remote objects like unmanned aerial vehicles and ground robots located in hard-to-reach areas of the Earth from the Central ground station. To reduce the cost of microsatellite design, it was decided to rigidly fix the receiving and transmitting antennas on their housings and use the spatial orientation of the entire apparatus for antenna guidance. This seriously complicated the tasks of navigation and orientation of microsatellites in a formation and required the development of a new method for determining the orientation of a single microsatellite. The essence of the method is to process the image obtained by means of a video camera mounted on a nearby microsatellite. We used methods of computer vision. The results of mathematical modeling simulation, as well as the results of full-scale bench experiment confirming the efficiency of the proposed method are presented.

Hardware and GNC solutions for controlled spacecraft re-entry using aerodynamic drag

Traditionally, controlled spacecraft re-entries have been conducted using propulsive de-orbit burns which are risky, expensive, and may not be possible for all vehicles. Recently, the miniaturization of technology has ushered in a new class of small satellites (such as CubeSats) that are too small to host thrusters but may require a controlled de-orbit if they contain materials capable of surviving re-entry. For all space vehicles requiring a controlled re-entry, the ability to harness the naturally occurring aerodynamic drag force for orbit control provides a cheaper and more reliable alternative to chemical propulsion.This paper discusses a comprehensive method for drag-controlled re-entry that is applicable to any vehicle capable of modulating its ballistic coefficient. First, a novel guidance generation algorithm efficient enough to run onboard a CubeSat outputs a desired ballistic coefficient profile and corresponding numerically propagated trajectory that if followed, will lead the spacecraft to a desired de-orbit location. This guidance generation algorithm is based on an analytical solution that provides convergence guarantees, ensures rapid performance, and facilitates a controllability analysis. Next, the guidance tracking algorithm utilizes an extended Kalman filter and GPS measurements to estimate the position and velocity of the satellite relative to the guidance. A full state feedback linear-quadratic-regulator (LQR) control strategy is then used to drive the relative position and velocity to zero using solely aerodynamic drag. This paper also discusses a novel retractable drag de-orbit device (D3) that can be attached to existing CubeSat structures and can easily be scaled up for larger satellites. The D3 provides passive three-axis attitude stabilization using aerodynamic and gravity gradient forces and can be repeatedly modulated to perform aerodynamically-based orbital maneuvering and controlled re-entry. The design of the planned 2U CubeSat to test the D3 and control algorithms in flight is also discussed.The re-entry point targeting algorithms were validated through extensive Monte Carlo simulations which included realistic GPS measurement errors and drag force uncertainties. The algorithms were able to guide the satellite to a desired de-orbit location with an average error below 25 km and in all cases, the targeting error was low enough for debris mitigation purposes. The accuracy and reliability of these algorithms coupled with the D3 device that has successfully undergone thermal vacuum, vibration, and fatigue testing provide a cheap, reliable, and comprehensive attitude, orbit, and de-orbit control solution that can be used on large and small space vehicles, possibly replacing conventional propulsion and attitude control systems and making space more accessible to everyone.

Accuracy of autonomous navigation by the mutual method in the case of a group flight of space vehicles

The article presents the results of an investigation of the accuracy of estimating errors in the determination of the motion parameters of the center of mass of a small spacecraft performing a group flight together with other small spacecraft, among which there is a small spacecraft that acts as the leader of the group. The angles between the directions to the small spacecraft-leader, and navigational stars, one of which is in the plane of the orbit of the small spacecraft, and the direction to the second one coincides with the binormal of this plane are chosen as the primary navigation parameters measured by the onboard facilities of the small spacecraft. When estimating the errors in determining the parameters of the motion of the center of mass of a small spacecraft, assumptions are made about the central gravitational field of the Earth, the normal law of error distribution of on-board navigation measurements with known constant variances. The research was carried out on the basis of the method of analytical estimation of the accuracy of autonomous navigation of space vehicles. As a result of the studies, analytical expressions of covariance matrices are obtained that allow one to estimate the maximum achievable accuracy of the solution of the problem posed, depending on the altitude of the orbit of a group of small space vehicles, the displacement of the slave small spacecraft relative to the small spacecraft-leader by the latitude argument, the variances of measurement errors and the number of measurements during a navigation mode. The presented results can find application when it is necessary to substantiate ways of improving the accuracy of autonomous navigation of small spacecraft performing group flights.

Experimental studies of correction propulsion system elements for small space vehicles manufactured due to additive method

Producing ammonia correction propulsion system (CPS) elements for maneuvering satellite platforms (MSP) of small space vehicles (SSV) is a relevant problem. The investigation is devoted to the solution of the named problem with the use of direct metal laser sintering (DMLS) method. The research objective is to confirm the feasibility of manufacturing ETMT and CPS evaporator with autonomous heating elements (AHE) by DMLS method, based on the prototypes experimental testing. During the research the following tasks were solved: creating 3D models for ETMT and double-threaded evaporator and producing experimental prototypes by DMSL method. 3D models of ETMT and evaporator casings were developed following the prototypes produced by the conventional methods of turning and milling. 3D models of ETMT and evaporator casings represent complex integral parts with multiple passages for working medium flow. Experimental studies of ETMT and the evaporator were performed with nitrogen as a working medium. ETMT and evaporator temperature characteristics were determined during the experiments. The investigation was made of ETMT with nominal thrust of 30 mN and power consumption in the range of 5-60 W with and without heat insulation. AHE with embedded thermocouples, having the diameter of 6 mm and power consumption of 60 W, was used. AHE temperature was limited by 973 K. A double-threaded evaporator was investigated for power consumption of 5-30 W, the evaporator casing temperature limited by 393 K. The maximal increase in the gas temperature equaled no more than 8.6 % at the nozzle exit in the power consumption range of 10-60 W for ETMT with heat insulation. At ETMT power consumption of 5–50 W, the build-up time for ETMT was 400–600 s. While at power consumption of 50–60 W, it was 200–400 s. At power consumption of 10–30 W, the evaporator casing temperature reached 393 K in 100–340 s, AHE temperature being 400–460 K and the gas temperature at the evaporator throttle exit being no more than 290 K. At power consumption of 5 W, the maximum evaporator casing temperature of 375 K was reached in 1200 s, AHE temperature being 370 K and the gas temperature at the evaporator throttle exit being no more than 302 K.

МАТЕМАТИЧЕСКОЕ МОДЕЛИРОВАНИЕ ЭЛЕКТРОНАГРЕВНОЙ ДВИГАТЕЛЬНОЙ УСТАНОВКИ В ФОРМИРОВАНИИ МИКРОСПУТНИКОВОЙ ГРУППИРОВКИ

The subject of the study in the article is the mathematical model of the propulsion system. It is built on the basis of the formed transfer functions of the elements. The goal is to rationalize the process of heating the onboard propulsion system on the basis of mathematical model under given constraints. Tasks: formalization of processes in an electric heating engine with a working body ammonia; formalization of the model of the onboard propulsion system; formation of the structural scheme; consideration of physical processes occurring in the nodes of the propulsion system; description of gas and hydraulic processes; the description of thermodynamic and electrokinetic processes; The construction of a mathematical model based on transfer functions. The methods used are: models of transfer functions of a tank, a filter, a steam generator, a receiver and a jet, an engine with their ranges of work. The following results are obtained. A block diagram of the onboard propulsion system was added, supplemented with a control unit and a power supply system. A formalized mathematical model of an onboard propulsion system with working body ammonia is created. From it formed a model consisting of the key elements that make up the onboard propulsion system, which is used when rationalizing the heating of the working fluid. The scientific novelty of the results is as follows. The mathematical model of the electro-heating propulsion system onboard small space vehicles has been further developed through its application to calculate the traction characteristics of the dispenser, which makes it possible to use an ammonia electric heating rocket engine in the formation of a constellation of satellites. The limitations of the operating parameters of the model are introduced. It was proposed to conduct the further workability of the model in Matlab Simulink. Thus, a rational value of the current and voltage parameters will be obtained, at which the time of the system's output to the operating mode will be minimal, and the thrust is maximum for the given operating temperature and pressure ranges

Approach to a simplified numerical optimization of low-power Hall thrusters

Simple, efficient, multi-parametrical design-oriented approach to optimize key parameters of miniaturized, low-power Hall-type thrusters potentially suitable for application at Cubesats and other compact spacecraft is proposed, designed, applied to a typical thruster design, and tested experimentally. Based on the detailed calculations of magnetic field, temperature distribution, electron and ion energies and other vital parameters in the detailed design with specific materials, the key power fluxes and critical temperature points in the thruster could be determined. The obtained results were then compared with the available test data, and the model was verified. The paper thus proposes an architecture of the modular CAD-based approach for the fast assessment of the Hall thruster performance at the design and test stages, and could be useful to design and optimize low-thrust devices vitally required for propelling small satellites and space vehicles of various types, including Cubesats, micro-satellites and interplanetary space probes.

Peculiarities of using additional low-power energy sources in the power supply system of small spacecraft

We propose a method of selecting the orbit for spin-stabilized spacecraft in order to obtain maximum electric energy on board due to the spacecraft motion in the earth magnetic field. The article describes the peculiarities of using additional low-power energy sources for power supply systems of small space vehicles. A method is proposed for taking into account the complex dependence of the position of the resultant magnetic induction vector on the spatial attitude of the spacecraft for converting the energy of the earth magnetic field into electric power.

IMPLEMENTATION OF A SMALL SPACE VEHICLE DEVELOPMENT PROJECT AS AN ILLUSTRATION OF THE INNOVATIVE UNIVERSITY EVOLUTION

Innovative development of the national economy, that has been declared in the Conception of long-term social and economic development of the Russian Federation, has become especially important in the modern situation characterized by geopolitical challenges and threats. Russia is looking for new possibilities to strengthen its competitive position in the global economy. One of the most promising innovative development areas is space science and technologies. Space technologies of manned flights developed in Russia have always been exceptionally competitive. However, in the last decade the onset of new aerospace industries in such countries as China, India or the USA has jeopardized the Russian Federation position of the world leader in this field. In this context such new trends in space technologies as the design, manufacture and launching of small-size space vehicles may be interesting both in terms of advanced engineering and technological solutions and in terms of commercial application of innovative low-cost space products. There are already several very positive practical examples of this engineering field development that involves the adoption of the so-called three-spiral theory describing the interaction between the state, science and business. One of such examples is the Southwest State University that has accumulated an extensive experience of long-time cooperation with Roscosmos (Russian Space Agency) core engineering businesses in the area of small space vehicle design and manufacture. This is a unique experience that has no parallels in Russia or abroad. In terms of economics such integration presents a good object for the investigation of system interaction between national authorities, businesses, educational and scientific institutions. It also makes a sound basis for the successful dissemination of the know-how.

Анализ нормируемых параметров электромагнитной совмести-мости блоков бортовой аппаратуры малых космических аппаратов

Анализ нормируемых параметров электромагнитной совмести-мости блоков бортовой аппаратуры малых космических аппаратов

Exploring the Possibilities for Improving the Performance of the Adapters used for Launching Multiple Small Space Vehicles on a Single Launch Vehicle

Background/Objectives: The research is aimed at determining the ways to improve the effectiveness of reconfigurable and extendable multipurpose adapters for multiple launch of small space vehicles by a launch vehicle. Methods: The structural design method is selected for studying design and constructive shape of the SSV adapters. The structural approach was used to define the principle for optimization of adapters of various constructive schemes, excluding modifications of launch vehicle used for multiple launches and enabling to extend and reduce adapter configuration with respect to a universal basic structure used for all tasks for SSV launch to be solved. Findings: To obtain quasi-optimal design and engineering parameters for the entire target set of tasks solved by the adapter, based on the structural approach we have considered design schemes of multi-purpose adapters for multiple and concurrent launching of SSVs of rod, shell, beam and platform types with side and end attachment configuration, the schemes with shock absorbers mounted in the places of adapters and MCAs installation, and scheme configuration combining adapter + small upper stage. The base and extension structures have been determined to form multi-purpose adapter with flexible and extensible design. As a result of the proposed technical solutions, weight of the adapter is reduced by 30%, configuration density is increased by 30-40 %, the number of the launched SSVs is increased, the fairing length and weight is reduced by 50%, root-mean-square amplitudes of accidental vibration decreased by 2.5-3 times, level of shock loads during SSV separation decreases by 6-10 dB, provision is made for shockless separation of SSV required lateral alignment of LV, cost reduction for ground and flight tests. Application/Improvements: Further improvement of the multi-purpose adapters is connected with the co-design of “Adapter + SSV” system, taking into account the mutual influence of design and engineering parameters of the SSV, adapter and LV used to launch them.