Spacecraft Control System Research Articles

Динаміка електрокерованих паливних клапанів реактивних двигунів малої тяги на «зеленому» паливі

Recently, the leading space countries have paid much attention to searching for and using new types of untoxic “green” propellants. Developments of 0.1 N to 220 N “green”-propellant thrusters have been reported, and some of them have been successfully tested in flight conditions. Liquid-propellant spacecraft thrusters are actuators of a spacecraft in-orbit motion control system. They are actuated from command signals of the spacecraft control system by applying a voltage to the electromagnet coils of the propellant valves of each of the engines that are to feed the propellant during a specified time interval to produce a required thrust impulse. A problem common to liquid-propellant thrusters irrespective of the propellant type is a thruster start-up delay with respect to the command signal and a thruster shut-down delay with respect to the electromagnet coil de-energizing. In other words, the thruster response is shifted with respect to the command signal. Because of this, the valve parameters must be chosen in such a way as to minimize the delay effect. The goal of this work is to determine design parameters of electrically controlled propellant valves that would provide an efficient in-orbit operation of “green”-propellant thrusters with a minimum of electric energy consumption. To determine these parameters, use was made of a branch standard, the procedure of work with which was based on empirical data and verified by numerical simulation. The key factor in shortening the thruster start-up delay is to reduce the seat-to-tip valve plate travel on condition that a required propellant amount is fed to the reaction chamber. However, this extends the shut-down delay, which has a harmful effect in the short-impulse operation mode. It is shown by calculation that after the electromagnetic parameters have reached their steady-state values, the voltage can be reduced by 30 to 40 per cent of its nominal value with the valve kept in an open state, after which the valve closes with a shorter delay. In addition, this way of valve energizing reduces the total electric energy consumption for valve operation. The mathematical model of valve dynamics allows one to determine the valve operation parameters in the case of short command signals probable in flight conditions. In such a case, a valve does not open to a full extent, which results in a low specific impulse. The results of this study may be used in the development of “green”-propellant thrusters.

Research on spacecraft environment control and astronaut health evaluation based on Modelica

Abstract The environmental control system of a manned spacecraft plays a pivotal role in maintaining the internal environment and temperature control, and it constitutes a fundamental component of the spacecraft. In this study, a Modelica-based model for the environmental control and health assessment of astronauts was examined with the objective of analysing the impact of the internal environment of the spacecraft on human physiological changes. The model employs multi-domain simulation techniques, integrating thermodynamic and hydrodynamic models to describe airflow and temperature changes, while incorporating human physiology models to capture the body’s response to these environmental changes. In particular, thermodynamic models are employed to calculate the transfer and distribution of heat within the cabin air, while hydrodynamic models are utilised to simulate airflow and pressure changes, which subsequently influence the breathing patterns and thermoregulatory mechanisms of the astronauts. The human physiology models include simulations of the cardiovascular, respiratory, and thermoregulatory systems. The coupling of these models enables the accurate prediction of the physiological state of astronauts under various environmental conditions. By modifying the environmental control parameters within the model, including air flow rate, temperature set point, and humidity level, the impact of disparate environmental control strategies can be evaluated. This method has the potential to be employed in the prediction of the health status of astronauts during various flight phases.

Academician V. H. Serheiev – Control System Designer, Founder of the Scientific and Design School, Honored Citizen of Kharkiv (on His 110th Birthday)

March 5, 2024 marks the 110th birthday of Volodymyr Hryhorovych Serheiev, an outstanding scientist, chief designer of control systems for strategic combat missiles, launch vehicles, spacecraft and transport modules of orbital complexes, Academician of the National Academy of Sciences of Ukraine and an Honored Citizen of Kharkiv. Despite his significant scientific achievements in the field of rocket and space engineering, the silhouette of V. H. Serheiev is hardly reflected in the modern history of science. Based on representative sources, the article highlights the main achievements of the scientist in the field of dynamics of automatic control systems and the design of rocket and space engineering control systems. The summary of the most prominent space projects is based on the memories of witnesses and direct participants. The personal qualities of V. H. Serheiev as the head of the Special Design Office No. 692 are noted. The historical retrospective describes technical developments and inventions meant to improve the control systems of intercontinental ballistic missiles R-16, R-16У, R-36, R-36П, R-36 orbital, R-36М, R-36М UTTH, UR-100N, UR-100HU, and R-36М2. At the initiative of V. H. Serheiev, complex stands were developed, which worked non-stop in all modes of operation of the necessary equipment and solved complex scientific and technical problems of creating an on-board digital computer. V. H. Serheiev was the chief control system designer of space launch vehicles Kosmos, Kosmos-2, Cyclone-2, and Cyclone-3 and Tselina series spacecraft. The article features innovative solutions implemented under Serheiev’s leadership during the creation of the Energia launch vehicle, supply vehicles of the Almaz space rocket system and target orbital modules of the Saliut and Myr space stations, and international launch vehicles Dnipro, Rokit and Strila. The article considers the issues of V. H. Serheiev’s direct participation in the organization of the training system for scientific personnel in the relevant specialties at Ukraine’s higher education institutions, as well as the development of international cooperation. It is substantiated that the scientist created one of the leading science and construction schools for the development of control systems for rocket and space technology, which is internationally recognized. The authors emphasize that V. H. Serheiev devoted a lot of attention to people. At his initiative and with his support, a district named after the founder of aerodynamics, M. Y. Zhukovskyi, was created in Kharkiv for the employees of “Hartron” enterprise. The district was not far from the enterprise and had a well-developed infrastructure, such as kindergartens, schools, shops, a swimming pool and a complex of residential buildings.

Evolution law of flow characteristics for straight pipeline after leaking

Aiming at the leakage problem of the fluid loop system for the thermal control system of spacecraft caused by the impact of micro-meteoroid and orbital debris, we investigate the dependence of the leak rate, pressure drop and flow characteristics before and after leaking on leak position, inlet pressure, and leaking aperture, calculated in a straight pipeline with single leakage, using stationary and transient three-dimensional computational fluid dynamics simulations with commercial software (FLUENT). The working fluid adopts single-phase water without gravity. It is found that the pressure increases obviously near the leak point, which is caused by the local peak of pressure. As the length between the leak point and the inlet becomes larger, the local peak of pressure and the dimensionless leak rate are smaller. When the inlet pressure increases, the leaked mass flow rate increases; however, the ratio of the leaked flow rate to inlet flow rate decreases. Furthermore, it is observed that the dependence of dimensionless leaking rate on dimensionless leaking aperture has a scaling relation with the scaling exponent approximating to 2, which may be related to the proportion of the vortex formed by the backward flow at the leak hole, the amplitude, and the affected length along the pipe of the local peak of pressure. This study is instructive and meaningful to the leak detection and plugging of fluid loop in space.

Power engineering of the tangential supply device of the microturbine of the thermal control system of a promising spacecraft

This paper presents an overview of the current technical problem related to two-phase spacecraft thermal control systems and possible technical applications of thermal energy recovery in the organic Rankine cycle as an integral part of thermal management systems. The design solution involves the integration of a steam microturbine behind an evaporator radiator. The microturbine is a tangential supply device and a radially centripetal impeller of low speed nst40. In this area, there is no reliable data on the design and energy of both the supply device and the impeller. The energy (loss of enthalpy) of the supply device mainly determines the transport of the swirling flow to the impeller and, as a result, the circumferential operation on the turbine. A prototype of a radial microturbine has been developed and presented in order to evaluate the design of the flow part of both the supply device and the impeller. As a result of the analysis, the main determining hydrodynamic areas necessary for hydrodynamic analysis and mathematical elaboration of the flow calculation algorithm with an assessment of energy losses are identified: the flow of a swirling flow of a radial-annular slit; axial-annular slit and tangential supply device. The first two algorithms assume computational modeling, the model of energy losses in a tangential supply device is not amenable to analytical modeling because it includes a sequence (or compatibility) of flows under boundary conditions defined as “local resistances”: the sudden expansion, reversal of the flow, together with a section of radially circumferential flow, the mutual influence of these boundary conditions assumes only an expe rimental assessment of energy losses in a tangential supply device through the loss coefficient of local resistance in the range of changes in geometric and operating parameters. As a result of experimental studies, a database has been proposed on the loss coefficient of tangential microturbine supply devices in the field of the practical range of the existence of operating and design parameters.

Satellite Thermal Management Pump Impeller Design and Optimization

This study presents a numerical approach to the design and optimization of centrifugal impellers used in the pumps of active thermal control systems of spacecraft. Although launch costs have shrunk in the last decade, the performance requirements, such as efficiency and reliability, have increased, as such systems are required to work up to 15 years, depending on the mission. To that effect, our paper deals with the first step in this pump design, namely the hydraulic optimization of the impeller. Constructively, this type of impeller allows for certain balancing systems and labyrinth seals to be applied in a more effective way, as well as allowing for additive manufacturing methods to be used—however, details regarding these aspects are beyond the scope of the current paper. By combining empirical formulas, computational fluid dynamics (CFD) analysis, and artificial neural networks (ANNs), the research focuses on achieving high efficiency and fast manufacturing. A series of geometries have been sized and validated using steady-state RANS (Reynolds Averaged Navier-Stokes) simulations, leading to the identification of the most efficient configuration. Subsequent optimization using an ANN resulted in a refined impeller design with notable improvements in hydraulic performance: a 3.55% increase in efficiency and a 7.9% increase in head. Key parameters influencing impeller performance, including blade number, incidence, and backsweep angles, are identified. This approach offers a comprehensive method to address the evolving requirements of space missions and contributes to the advancement of centrifugal pump technology in the space domain.

Attitude control of combined spacecraft with fuzzy neural network disturbance observer

A finite-time control strategy based on a fuzzy neural network disturbance observer (FNNDO) and nonsingular terminal sliding mode (NTSM) is proposed for the attitude control system of a combined spacecraft in the presence of non-cooperative targets with competitive torque output. First, a mathematical model of the combined spacecraft attitude is established with the service spacecraft as the reference. Then, the external disturbances and competitive torque output from non-cooperative targets are treated as generalized disturbances, and FNNDO is employed for tracking and compensation. Subsequently, a NTSM controller is designed to achieve finite-time attitude takeover control of non-cooperative targets. The stability of the disturbance observer and controller is proven using Lyapunov stability theory. Simulation results demonstrate that this control strategy effectively handles the competitive actions of non-cooperative targets while rapidly stabilizing the combined spacecraft’s attitude at the desired values.

Numerical investigation on cooling performance and control effect of a novel evaporation-driven cooling unit for Mars extravehicular space suit application

The special environment of Mars brings new challenges to the thermal control of spacecraft on Mars missions. The outstanding advantages of water evaporation cooling technology make it promising for use in the thermal control system of the Mars spacecraft. This paper proposes a novel concept of evaporation cooling hardware that integrates heat collection, heat transfer and heat rejection, namely evaporation-driven cooling unit. The evaporation-driven cooling unit has a simple structure, high evaporation efficiency and the potential to simplify the system topology. Furthermore, an expert hierarchical coordination control strategy applied to the Mars extravehicular space suit cooling loop with an evaporation-driven cooling unit as its heat sink is proposed. The dynamics model of this cooling loop is also constructed using the lumped parameter method. The numerical results have shown that the expert hierarchical coordination control strategy can achieve precise temperature control and reduce the water consumption of the system. The average water consumption rate during extravehicular activity is 0.6975 kg/h when the mean heat load is 572 W (including member metabolic heat and electronic waste heat). Compared to the open-loop case with the same input heat load, a typical extravehicular activity lasting 8 h will save about 11.5 g water if the expert hierarchical coordination control is applied. The proposed evaporation-driven cooling unit will provide more choices for heat rejection of extravehicular space suits. It is expected that the evaporation-driven cooling unit and expert hierarchical coordination control have potential applications in thermal control systems for other space vehicles in near-Earth orbit, on the surface of the Moon and Mars.

Numerical study and optimization of battery thermal management systems (BTMS) Based on Fin-Phase change material (PCM) in variable gravity environments

Phase change materials (PCMs) are widely employed in electronic thermal control systems for spacecraft because of their substantial energy storage competencies. Hence, the impact of gravitational acceleration (GA) on PCM thermal performance holds particular significance. Moreover, given considerations for spacecraft center of gravity, attitude control, and structural design, choosing an optimal battery placement orientation becomes pivotal for maximizing available space utilization. This study employs the enthalpy-porosity method to simulate and assess the heat transfer efficiency of PCM and the thermal behavior of cooling battery thermal management system (BTMS) under different gravity environments (0.05 g, 0.1 g, 1 g, 10 g, and 20 g) and different inclination angles (θ = 90°, 45°, 0°). The research findings indicate that when θ = 90°, as gravity increases from 0.05g to 20g, the complete melting time of PCM at GA = 20g is shortened by 95.05 s (15.00 %) compared to GA = 0.05g. Particularly, when GA = 10 g, increasing θ from 45° to 90° reduces PCM melting time by 59.05 s (9.66 % reduction) compared to θ = 45°, while at θ = 0°, it decreases by 65.75 s (10.75 % reduction). Notably, at θ = 45°, some unmelted solid PCM accumulates in the lower right corner of the enclosure hindering the BTMS heat transfer. Based on PCM melting characteristics, three optimization structures for different θ are proposed. At t = 240 s, PCM melting significantly improves with fins. Comparing Convective Heat Transfer Coefficients (CHTC) curves before and after optimization reveals reduced accumulated effects. However, more fins aren't always better due to BTMS weight considerations. Investigating enhanced CHTC, fin mass, and volume ratio(γ), the optimized model impacts BTMS weight differently for θ values. θ = 45°, a combination of annular and straight fins, is determined optimal using entropy-weighted TOPSIS. This paper innovatively studies PCM melting in BTMS under varied gravitational environments and θ, proposing an enhanced heat transfer scheme for aerospace BTMS design.

Design of anti-unwinding attitude coupling controller for flexible spacecraft using positive position feedback control

Flexible vibration has a great impact on flexible spacecraft attitude control, which will reduce the pointing accuracy of the spacecraft platform and bring severe challenges to the design of a high-precision attitude control system for flexible spacecraft. Aiming at the problem of attitude and vibration coupling control of flexible spacecraft, an anti-unwinding attitude coupling controller using the positive position feedback (PPF) control is proposed in this paper. The model compensation control term is constructed based on the prior information to improve the tracking performance. The proposed attitude coupling control method can effectively suppress the flexible vibration while achieving more accurate attitude control. The obtained closed-loop system under the proposed control scheme is asymptotically stable, which is proved by the Lyapunov stability theory. The numerical simulation illustrates that the proposed method can effectively improve the attitude control accuracy of the flexible spacecraft and is capable of large-angle attitude maneuver control because of its prominent performance in attitude tracking and its unwinding-free performance.

On-policy and off-policy Q-learning strategies for spacecraft systems: An approach for time-varying discrete-time without controllability assumption of augmented system

This article investigates two On-policy and Off-policy Q-learning algorithms for time-varying linear discrete-time systems (DTSs) in the presence of complete dynamic uncertainties. To handle the challenge of time-varying description, the lifting method is employed to transform the original time-varying linear DTS into time-invariant linear DTS in the absence of the conventional controllability condition, which affects to the convergence of traditional Q-learning algorithms. Based on theoretical analysis of the structure in the obtained time-invariant linear DTS, On-policy and Off-policy algorithms are proposed to guarantee the convergence of Q-learning algorithms. Both On-policy and Off-policy Q-learning algorithms guarantee model-free consideration under the data collection. Especially, the Off-policy technique is able to develop the algorithm with high data efficiency because the collected data can be utilized again after each iteration. Finally, the simulation results of two-dimensional systems and spacecraft control systems are presented to validate the effectiveness of the two proposed control schemes.

A data-driven configurability evaluation method for spacecraft control systems

<p style='text-indent:20px;'>To release severe constraints of spacecraft on orbit resources (in-cluding computing resources, hardware resources and energy resources), the data-driven reconfigurability evaluation of spacecraft attitude control system is studied in this paper. Firstly, the reconfiguration capability is quantitatively described based on the data forms of stable kernel representation and stable image representation of spacecraft control system subject to noise. Then, based on the obtained reconfigurability index, the maximum boundary of system re-configurability is given, which provides theoretical guidance for the design of autonomous reconfiguration strategy. Finally, an example is employed to verify the effectiveness and correctness of the proposed method.</p>

Расчетно-экспериментальное исследование закрученного кольцевого потока

In the course of work on the energy perfection of thermal control systems for spacecraft with a twophase circuit, the issue of partial regeneration of thermal energy into electrical energy in a low-speed turbogenerator is considered, part of the design work requires computational modeling during the transport of swirling flows in the axial direction from the external input to the input plane into the impeller, which determines the need for theoretical and experimental elaboration of the problem. The paper considers transformations of equations for changing the amount of fluid motion in boundary conditions of an axial annular channel with fixed cylindrical surfaces. Assuming the symmetry axis of the flows using the integral form of writing the continuity equation, the relations are obtained in the form of two differential equations with expressed derivatives along the channel axis for the total pressure p* and the circumferential velocity constant Cu = UR (const — at the integration step). The equation forms the basis of the algorithm of integration in finite differences supplemented by a system of service equations describing the friction stress, thermodynamic parameters, etc. Test calculations are carried out using real parameters, the results are analyzed.

Intelligent fault diagnosis method of spacecraft control system based on sequence data-image mapping

<p style='text-indent:20px;'>Satellite networking, as the future development direction of aero-space, requires high-precision autonomous fault diagnosis capability for a single satellite. In this paper, aiming at the characteristics of closed-loop fault propagation and high data dimensionality of spacecraft control system, neural network algorithms are conducted to study the fault diagnosis of spacecraft high-dimensional coupled data. Based on the ground test data of a certain spacecraft, this paper converts the high-dimensional sequence data into grayscale images, and then uses Convolutional Neural Network (CNN) and Long Short-Term Memory (LSTM) to diagnose them respectively. The effectiveness of the methods in this paper is illustrated by comparing and validating them with three non-image-based machine learning algorithms, namely, K-NearestNeighbor, Bayesian classifier, and K-NearestNeighbor based on Principal Component Analysis.</p>

Simplified Maneuvering Strategies for Rendezvous in Near-Circular Earth Orbits

The development of autonomous guidance control and navigation systems for spacecraft would greatly benefit applications such as debris removals or on-orbit servicing, where human intervention is not practical. Within this context, inspired by Autonomous Vision Approach Navigation and Target Identification (AVANTI) demonstration, this work presents new guidance algorithms for rendezvous and proximity operations missions. Analytical laws are adopted and preferred over numerical methods, and mean relative orbital elements are chosen as state variables. Application times, magnitudes and directions of impulsive controls are sought to minimize propellant consumption for the planar reconfiguration of the relative motion between a passive target spacecraft and an active chaser one. In addition, simple and effective algorithms to evaluate the benefit of combining in-plane and out-of-plane maneuvers are introduced to deal with 3D problems. The proposed new strategies focus on maneuvers with a dominant change in the relative mean longitude (rarely addressed in the literature), but they can also deal with transfers where other relative orbital elements exhibit the most significant variations. A comprehensive parametric analysis compares the proposed new strategies with those employed in AVANTI and with the global optimum, numerically found for each test case. Results are similar to the AVANTI solutions when variations of the relative eccentricity vector dominate. Instead, in scenarios requiring predominant changes in the relative mean longitude, the required ΔV exhibits a 49.88% reduction (on average) when compared to the original methods. In all the test cases, the proposed solutions are within 3.5% of the global optimum in terms of ΔV. The practical accuracy of the presented guidance algorithms is also tested with numerical integration of equations of motion with J2 perturbation.

Determination of the Heat-Controlled Accumulator Volume for the Two-Phase Thermal Control Systems of Spacecraft

<div>For spacecraft with high power consumption, it is reasonable to build the thermal control system based on a two-phase mechanically pumped loop. The heat-controlled accumulator is a key element of the two-phase mechanically pumped loop, which allows for the control of pressure in the loop and maintains the required level of coolant boiling temperature or cavitation margin at the pump inlet. There can be two critical modes of loop operation where the ability to control pressure will be lost. The first critical mode occurs when the accumulator fills with liquid at high heat loads. The second critical mode occurs when the accumulator is at low heat loads and partial loss of coolant, for example, due to the leak caused by micrometeorite breakdown. Both modes are caused by insufficient accumulator volume or working fluid charge. To analyze the loop characteristics in critical modes, experiments were conducted on a test bench with ammonia coolant, and a mathematical simulation of a two-phase mechanically pumped loop was performed. The results show that the loop can operate in critical modes in a certain range of heat loads. The conducted studies allow for the design of a heat-controlled accumulator with the minimum required volume, expand the performance range of a two-phase mechanically pumped loop, and increase the reliability of its operation in orbit during long-term missions.</div>

Nonzero-Sum Pursuit-Evasion Game Control for Spacecraft Systems: A Q-Learning Method

In this paper, the nonzero-sum pursuit-evasion (PE) game control problem is studied for a class of linear spacecraft control systems subject to the complete information case and incomplete information case. The incomplete information includes the cost functions and the control inputs. In practical confrontation situations, due to the incomplete information constraints, it is impossible for the pursuer and the evader to build up the exact opposite cost function. Hence, a nonzero-sum game framework is utilized to describe the PE game problem of the double-spacecraft system. First of all, under the complete information case, the nonzero-sum PE game control strategy is designed by solving the coupled Riccati recursions. Then, aiming at the incomplete information case, a control gain estimator is established, which lays the foundation for the control strategy design of the pursuit spacecraft. On the basis of the estimated control gain, the pursuit control strategy is solved by using the standard discrete-time Riccati recursion. In order to further get rid of the system information, a Q-learning-based control gain is designed for the pursuit spacecraft. Finally, a numerical example on the PE spacecraft system is provided to verify the effectiveness of the proposed PE game control strategies.

SPECIFICATION FORMALIZATION OF STATE CHARTS FOR COMPLEX SYSTEM MANAGEMENT

This article presents a formalization approach for the requirements of object-oriented programs with state machines, using a spacecraft control system as a case study. It proposes a state pattern implementation, where each state is represented as a class with clearly defined responsibilities, and the transitions between states are controlled by the state objects themselves. Additionally, the application of model checking, theorem proving, and code generation techniques are discussed. The effectiveness of the proposed approach in ensuring compliance with the specified requirements is demonstrated, while also identifying potential drawbacks and limitations of the approach. The implementation is validated using a range of formal verification techniques, including model checking and theorem proving. The article also discusses how the approach can be extended and applied to other complex systems. Overall, the valuable insights into the formalization of requirements for object-oriented programs with state machines are provided, offering a practical and effective approach for verifying the correctness and completeness of such implementations. The results of this work have important implications for the development of safety-critical systems and can potentially improve the quality and reliability of software systems in various domains. By using mathematical models and rigorous formal methods, it is possible to detect and eliminate errors early in the development process, leading to higher confidence in the correctness of the final product. Future research in this area could explore the use of more advanced techniques, such as model-driven development and automatic code synthesis, to further streamline the software development process. Additionally, the development of more efficient and user-friendly tools could make these techniques more accessible to a wider range of developers and organizations. Altogether, the combination of formal methods and software engineering has the potential to revolutionize the way software systems are designed, developed, and verified, leading to safer and more reliable software for critical applications.

Experimental studies of using the digital control system in power conversion equipment of high-voltage power supply systems in spacecrafts with a hydrogen energy accumulator

Solar-hydrogen energy for spacecraft, space interplanetary stations, space bases on the planets of the solar system must reliably provide energy: spacecraft during peak periods (repairs at a space facility, energy-consuming experiments, cleaning of facilities, regeneration of waste to extract oxygen and hydrogen, etc. .d.) and bases on planets, especially where the night lasts for a long time.The onboard hydrogen module of the solar-hydrogen plant includes: a container with water; distiller; deionizer; water lines with adjustable shut-off valves; electrolyzer; cryogenic side tanks with hydrogen and oxygen; fuel cell, hydrogen safety sensor systems based on leakage and concentration sensors; level gauges; cryogenic and gas pipelines of hydrogen and oxygen; onboard cryo-refrigerator; automatic control system for the hydrogen module. The article presents the results of experimental studies of static, dynamic and mass-dimensional characteristics of power conversion equipment (hereinafter - PCE) in the power supply system (hereinafter - PSS) of the spacecraft (hereinafter - spacecraft) with a digital control system. The purpose of experimental research is to verify the theoretical results obtained earlier in the design of digital PCE control systems of high-voltage PSS spacecraft, including the study of static and dynamic characteristics. We developed and manufactured an experimental PCE sample with an output power of 2500W in this research. It includes: four voltage stabilization modules and the 100V load supply output bus filter. Each voltage stabilization module includes a battery energy conversion channel, a solar battery energy conversion channel and a digital control system unit. The power conversion channel of the accumulator battery is designed and made on the basis of a bridge volt-booster inverter-transformer circuit. Voltage control in energy converter of accumulator battery is based on phase shift in control pulses of transistors in adjustable inverter rack related to unregulated one. The power conversion channel of the solar battery is designed and made on the basis of a two-phase circuit of a direct voltage converter of an increasing type with switching throttles. The digital control system of the voltage stabilization module is designed on the basis of analogues in the modern electronic component base in the Space class and consists of four identical channels. Each channel of the control system controls its own voltage stabilization module. We obtained the following results during our experimental studies: some pulsations of the output voltage of the energy conversion equipment during operation on an active load, transient processes of the output voltage of the energy conversion equipment during pulsations of the load current without changing the mode of operation, transient processes of the output voltage of the energy conversion equipment when changing the modes of its operation and weight and size characteristics for different values of the output power. The obtained results of experimental studies confirm the previously obtained theoretical results, and also showed that the use of a distributed digital control system will increase the specific energy characteristics of the newly developed PSS spacecraft while ensuring the given static and dynamic characteristics of output voltage quality. The results of the work are designed to create a new on-board power conversion equipment for power supply systems of promising spacecraft. The area of application of the project results is space instrumentation.

Failure Identification Method and LMI-Based Control System for Small Spacecraft

This study aimed to design a linear matrix inequality (LMI)-based robust control system and a failure identification method for small spacecraft. The key feature of the proposed method is its capability to withstand failure in the actuation system by means of the observer and controller gain definition. An effective approach for withstanding failure is required due to the small capabilities of the momentum exchange device (MED) and the system’s sensitivity to external perturbations. Passive fault-tolerant approaches can be included in the design process of the control algorithm. In particular, the main objective of this study was the design of an H∞ controller that started from an LMI formulation and considered the parametric uncertainties and matched failure of the actuation system. In addition, a fault detection method based on sliding mode observers was proposed to include an active disturbance correction algorithm and improve the system robustness and performance during spacecraft stabilization. The closed-loop system was evaluated for different initial conditions, including attitude positions that are far from the desired condition. The effectiveness of the proposed approach was demonstrated using extensive simulations that considered a pyramidal actuation configuration and hardware constraints.