Aviation and spacecraft engineering. Рубрика в журнале - Siberian Aerospace Journal
Modeling of an optimal trajectory at a safe distance in a spacecraft cluster
Scientific article
This paper addresses the problem of modeling an optimal motion trajectory of a spacecraft within a spacecraft cluster consisting of a target and a deputy spacecraft during an undocking maneuver followed by a fly-around at a safe distance. The relative motion of the deputy spacecraft is described in a local orbital reference frame using a linearized dynamical model based on the Hill – Clohessy – Wiltshire equations. An integral performance index is introduced to characterize the tendency of motion near the boundary of the admissible relative position region. The original optimal control problem with an integral cost functional and bounded control inputs proves to be challenging for both analytical and numerical analysis, which limits the direct application of standard optimization techniques. To facilitate the study, a decomposition approach is employed, allowing the problem to be divided into two consecutive stages: transfer to the boundary of the safe zone and subsequent motion along this boundary. The Pontryagin maximum principle is applied to investigate the properties of optimal solutions and to derive analytical expressions for the adjoint variables, making it possible to analyze the structure of optimal control laws. Numerical analysis is carried out using both direct and indirect optimal control methods. A set of quality criteria is introduced to assess the obtained solutions, including the fulfillment of necessary optimality conditions, the Hamiltonian behavior, and the sensitivity of the solution to variations in problem parameters. The comparative analysis shows that direct optimization methods yield more stable and reproducible solutions, while the tangential boundary approach improves the agreement between the two stages of motion at the expense of a moderate increase in the transfer time to the boundary.
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Modeling of composite mesh adapter for Marathon satellite system spacecraft output
Scientific article
Composite shells of rotation of mesh type are often used in the production of rocket and space technology as power elements of structures for various purposes, including spacecraft hulls. High specific values of mechanical characteristics of composites allow to manufacture structures with a high degree of weight perfection. Usually composite mesh structure has the form of a cylindrical or conical shell of circu-lar cross-section and consists of a system of spiral and circular ribs running along the midpoints of the areas between the nodes of intersection of spiral ribs. The discrete structure of the mesh shell is relatively simple and can be manufactured by the method of continuous winding of composite fibers, which is very technologically advanced, wellestablished and therefore currently widespread. A characteristic example of anisogrid cylindrical and conical shells is the spacecraft adapter for GLONASS satellites orbit launching, different variants of which are still produced in the workshops of Reshetnev JSC. The shells differ in dimensional parameters (diameters and lengths) and bearing capacity, but they are structurally identical, which allows to develop a universal and maximally automated modeling and calculation procedure. This is extremely important for composite elements of rocket-space technology, which have numerous variable parameters. The optimal combination of these parameters is determined in the process of performing a complex numerical experiment. The composite mesh adapter considered in this paper, designed for the Marathon satellite orbit, differs from the previously used shell structures by the shape of the cross-section, which in the main part is a regular octagon. The previously developed algorithm for modeling an anisogrid structure with a system of spiral and annular ribs formed by unidirectional carbon fiber-reinforced plastic fibers is used. Numerical in-vestigation of stability, stiffness and stress-strain state of the structure is carried out in the environment of an integrated package of finite element programs while varying the main parameters of its mesh structure formation.
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Scientific article
One of the promising areas for improving the methods of manufacturing structural elements of rocket and space technology is the use of selective laser melting technology which represents a unique opportunity to manufacture metal products by melting powder and producing a one-piece solid phase structure. However, pores and other structural defects can appear in the formed element during laser sintering which causes a decrease in the strength characteristics of the parts produced. An important step in the additive technologies introduction is the development of methodology for the preliminary prediction of the strength characteristics of manufactured structural elements under the influence of mechanical loads with the help of mathematical modeling. The methodology for estimating the material strength reduction of a rocket-space technology element obtained using additive technologies by simulating a porous structure and calculating the characteristics of the stress-strain state is presented. The proposed mathematical model and the methodology for calculating the specimen loading on the basis of the distortion energy theory allow calculating the stress-strain state in the process of numerical simulation for different values of the pore diameter. The reduction in yield strength due to the material porosity of the part is estimated using a coefficient equal to the ratio of equivalent stresses arising when a load is applied to a specimen manufactured using traditional and additive technologies. The value of the introduced coefficient characterizes the structure of the grown product and is considered as a function of the random arrangement of pores in the specimen under study. The appearance of pores is the result of a combination of factors: the composition and dispersion of the original metal powder, feed rate, removal distance and laser power during sintering, part orientation and sintering direction, the height of the level of powder deposited on a special base before sintering, etc. The paper evaluates the reduction in strength for the working part of a series of tensile test specimens grown from metal powder of different dispersity. The non-linear nature of the dependence of the yield strength on the particle diameter of the original metal powder is established. The maximum value of the yield strength corresponds to the specimen with the minimum value of the total surface area of the pores.
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Numerical hydrodynamic analysis liquid dynamometer for ground run and aircraft piston engine tests
Scientific article
During designs of new internal combustion engines (ICE), modernizing existing ICEs, tuning of fuel systems and ignition systems, the use of dynamometric braking devices provides significant assistance. Such devices can also be used when individually tuned stock internal combustion engines. Many types and designs of such devices are known, including mechanical, electrical and hydrodynamic variants. If an engine torque measuring element is provided in the design of the braking devices, such braking devices are called dynamometers. Most dynamometers are significant in size and weight and are designed for testing, running in and tuning engines with a maximum power of at least 80 hp. and operating speeds of no more than 5000 per minute. Dynamometer stands are also known for high-speed sports engines that take energy from the wheel of a vehicle. The use of dynamometers with rotating elements having a large moment of inertia leads to large errors in tuning. For piston engines of UAVs with a relatively small power of 5-60 hp with operating speed range 4000-10000 per minute the most promising is the use of hydrodynamic braking devices of test and rolling benches. Compared to other devices, they are characterized by smaller dimensions, cost and a wide range of absorbed power. The work includes the design of a liquid dynamometer for testing and tuning the internal combustion engine of an unmanned aerial vehicle with a capacity of up to 100 hp. Optimal geometric parameters of hydrodynamic brake were selected. A three-dimensional solid model of a liquid dynamometer was simulated. A numerical hydrodynamic analysis was carried out with a fixed stator and a movable rotor completely filled with working fluid.
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On an alternative method for testing the dynamic strength of a small spacecraft structure
Scientific article
This article presents an analysis of the possibility of applying an alternative approach to testing the mechanical effects of the design of a small spacecraft for remote sensing of the Earth, which has an analog product that has passed a full cycle of ground experimental testing. However, despite the similar power scheme and the maximum borrowing of onboard equipment with minimal modifications, the spacecraft planned for testing has a number of significant differences. The application of the main alternative strategies in foreign and domestic practice in the ground-based experimental development of space technology is considered, their advantages and disadvantages are described. Some criteria for decision-making on the rejection of the use of traditional methods of ground-based experimental testing of space technology for mechanical effects are given. The analysis of the normative and technical documentation adopted in the domestic industry in terms of clarifying the list of development tests of spacecraft, the assumptions of applying the computational and experimental method to the development of dynamic (vibration) strength and the analysis of the design of the spacecraft planned for testing in comparison with an analog product showed that the most preferred method of testing dynamic (vibration) strength is the strategy “protocol qualifications”. In accordance with the chosen strategy, a list of tasks was defined that will clarify the nomenclature of the development tests of the research object.
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On the issue of hydrodynamic braking efficiency dur-ing high-velocity tests on a rocket-rail track
Scientific article
At present, the creation of high-velocity aircraft is a promising direction in the development of aircraft and armament both in Russia and abroad. The increase in velocity characteristics of newly developed sam-ples imposes new requirements on test bench equipment, including rocket-rail tracks. The requirements are growing both for the acceleration and braking means, which ensure the tested materiel safety. The proposed work deals with a hydrodynamic braking method used in high-velocity dynamic tests on a rocket-rail track at the Federal State Enterprise “State Governmental Scientific-Testing Area of Aircraft Systems named after L.K. Safronov". The paper gives the description of the braking devices, presents the dependencies determining the calculated values of the braking force developed by them, and describes the braking intensity control methods, which increase the efficiency and safety of braking as well as expand the permissible speed range of the hydrodynamic braking device application. The method of increasing the efficiency of the braking devices functioning by using a special form of its working part profile is presented. The corresponding examples of the braking modes are given for a comparative assessment of the braking efficiency parameters when using braking devices with special and triangular profiles. The working part profile of the hydrodynamic braking device calculated according to the proposed method provides more efficient and safe braking compared with the previously used triangular profile, by maintaining a constant stopping force in a wide velocity range.
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On the possibility of flight of a single-stage rocket to the planets of the Solar system
Scientific article
A new concept for building a small-sized rocket engine containing a special gas ionizer in the combus-tion chamber to increase its conductivity to an optimal value with a corresponding improvement in the fuel combustion process is proposed. A simplified calculation for the relative velocity of gases in a conical noz-zle is given when heating the supersonic flow of gases by plasma in a conical nozzle by means of a power-ful, electromagnetic, high-frequency field, and the influence of some technical parameters on the efficiency of the rocket's flight is also considered. A comparison of the flight altitude of a rocket with plasma heating of the gas flow in a conical nozzle with a scale model corresponding to the well-known single-stage Zenit rocket with the same weight and geometry, taking into account air resistance for the cargo version of the rocket with one small-sized engine, is made. The result is a significant reduction in fuel consumption and an increase in the maximum flight altitude by 2 times with an increase in specific impulse by 2.7 times, other things being equal. It is estab-lished that under certain conceptual parameters, it is possible to rapidly accelerate and fly a single-stage cargo rocket with a launch weight of 17.25-20.00 tons to the planets of the solar system directly from the Earth's surface using a bunch of engines of the same type. The use of many of the same type of small-sized engines allows you to abandon the Laval nozzle in favor of a simple conical nozzle, which reduces the size of the rocket as a whole. This is determined by the need to reduce the diameter of the conical nozzle in order to achieve a greater specific heating power of the plasma compared to the specific power of the fuel burned in the combustion chamber. It is also proposed a complete rejection of the steering engines, the function of which will be performed by part of the engines located closer to the diameter of the rocket. As electric generators, it is proposed to use promising prototypes of electric generators MEG-6NS, MEG-15NS and others, the company "NaukaSoft", with good weight indicators that allow in the future producing such a liquid propellant rocket engine (LPE) of small dimensions. The redistribution of part of the fuel used to produce electricity is compensated by a significant increase in the specific impulse of the LPE to increase the speed and overall efficiency of the flight with an optimal ratio of the amount of fuel to the weight of the rocket before refueling.
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Scientific article
One of the promising types of spacecrafts are large-size transformable reflectors. Such apparatuses are delivered to a target orbit folded, and then deployed to a working condition. The large aperture allows sig-nificantly expanding the capabilities of the antenna. In this case, the tasks arise of a smooth and reliable deployment, adjusting the shape of a radio-reflecting net, and adjusting the orbital position. Due to the fact that the deployment process takes a long time, accounting for disturbing influences is an important prob-lem. The presence of radiation, large temperature differences, solar wind affect the entire system and main-ly on the directional diagram. It is also necessary to smoothly deploy the structural elements, since with an increase in the diameter of the radio-reflecting surface, the moments of inertia of the antenna increase, which leads to prolonged oscillations. In this paper, the process of deployment of the reflector spokes in the presence of disturbances and measurement errors is considered. The solution to the problem is presented using the separation theorem. To estimate the parameters of the system in the presence of measurement noise, the Kalman filter is applied. Its performance is shown at various values of the noise intensity. A ran-dom process such as white noise was selected as external disturbances and measurement noises. The con-trol problem is solved using the optimal control algorithm according to the hierarchy of target criteria. The possibility of minimizing energy costs by means of interval switching on of measuring sensors is shown. The results of numerical simulation are presented.
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Optimization the position of the spacecraft instrument panel mounting points based on modal analysis
Scientific article
The paper presents optimization of the location of interface points of the spacecraft instrument panel using modal analysis, as well as a quasi-static calculation of the panel under study, confirming effective-ness of proposed changes in the panel design. The instrument panel is a three-layer honeycomb structure consisting of two aluminum plates and a honeycomb filler. Cellular panels have a number of advantages: a small weight of the structure, high rigidity, specific strength. Using finite element modeling, the range of natural frequencies and vibration patterns of the instrument panel was determined, which made it possible to determine optimal location of the panel fixing points to the spacecraft body to increase the lower limit of natural frequency range and increase its carrying capacity.
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Overview of unmanned aircraft cargo delivery systems
Scientific article
The article is devoted to an overview of unmanned aircraft systems (UAS according to GOST R 59517–2021, or in a number of publications – unmanned aerial vehicles or drones) capable of delivering various cargoes. The options for their application in various fields, including agriculture and forestry, fishing, wildlife protection, air quality monitoring, mining, defense and civilian use, search and rescue operations, are given and briefly analyzed. The purpose of the study was to demonstrate the identified capabilities of unmanned aircraft systems in terms of cargo delivery for various purposes, as well as the availability of infrastructures when receiving cargo for further transportation and its delivery at the place of delivery. The history of the development of unmanned vehicle technologies is given. The relevant standards of the Rus-sian Federation, classification and categorization of UAS in the Russian Federation and abroad are pre-sented. The advantages and disadvantages of UAS are described, as well as the problems of delivering mail using drones. Conclusions are drawn about the desire of developers and trends to create intelligent, fully automatic robotic aviation systems, however, it is noted that fully automatic systems do not yet exist in the world, therefore, this area of development is relevant. It is concluded that it is necessary to carry out scientific and technical developments in terms of creating automatic delivery of mail and other items, especially relevant for megacities in which the largest part of the population of developed countries is concentrated. At the same time, it is obvious that the delivery system should consist of at least three stag-es – warehouse sorting, direct flight and navigation, and cargo shipment to the appropriate post office. Each stage is an infrastructural concept with a complex system and logistics that requires artificial intelli-gence, robotic devices, and other elements and attributes of complex systems that are subject to compre-hensive study in technical, legislative, legal, and logistical terms.
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Parameters of rocket engine chambers, obtained by selective laser melting
Scientific article
When designing and testing a low-thrust rocket engine (LTRE), one of the most important tasks is to ensure the quality of materials, which, in turn, affects the reliability of the product. Currently, additive technologies for manufacturing parts from metals are actively developing. This direction is relevant for rocket and space technology products to reduce weight and increase the reliability of products. The article presents the results of studies of the chemical composition and mechanical characteristics of the material of the low-thrust rocket engine demonstrator chamber, manufactured by selective laser melting from metal powder. The properties of products made from Inconel 718 metal powder were studied. Samples were made and the chemical, mechanical and structural characteristics of the material were studied. Based on the test results, two LTRE samples were printed. LTRE chambers were tested for vibration loads, strength and tightness. Increased porosity and roughness of the test material of the engine chamber were noted. When analyzing a number of parameters of the selective laser melting technology, an experimental selection of printing parameters was carried out and the most significant factors affecting the print quality (surface roughness and porosity) were identified. Based on the results of the work carried out, four groups of controlled printing parameters were identified that affect the properties of the resulting material. The work also provides recommendations on printing modes and characteristics to obtain the highest quality parts.
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Parametric analysis of the anisogrid body of the spacecraft for cleaning the orbit of space debris
Scientific article
The article presents an approach to solving the problem of designing a spacecraft for cleaning the orbit of space debris (space garbage collector-KSM), the body of which is made in the form of a cylindrical mesh anisogrid shell. The design task is to select the optimal parameters of the anisogrid body of the KSM (the shape and cross-sectional area of the ribs, the number of annular and spiral ribs, material characteristics, etc.) that provide the necessary strength and stability of the structure with minimal weight. During the design process, a parametric analysis of the anisogrid housing of the space garbage collector was carried out. By varying the number and angle of inclination of unidirectional spiral ribs, we find the optimal design scheme that satisfies the specified safety and stability coefficients. Parametric analysis of the KSM body includes modeling of the main weight and strength parameters: determination of the stress-strain state of the structure, values of the body’s natural frequencies, determination of the bending margin from the longitudinal force, determination of the body mass. The analysis of the load-bearing capacity of the anisogrid housing of the space garbage collector was carried out by the finite element method using the MSC Nastran software package. A finite element mesh model was created from a two-node spatial finite element bundle. The disk attached to the end of the shell was modeled using a rigid finite element. The size of the final beam element for all shell models was the same and equal to 10 mm. During the parametric analysis, three variants of the mesh composite structure with a different number and angle of inclination of unidirectional spiral ribs were considered. Based on the results of parametric analysis of the spacecraft body, its geometric dimensions are determined and the mass of the spacecraft structure as a whole is minimized.
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Parametric analysis of the strength of a solid propellant rocket engine nozzle
Scientific article
The paper presents an approach to solving the problem of designing a solid propellant rocket engine (SPRE) nozzle using such a design feature as a carbon fiber insert plate. The design task is to select the optimal parameters of the plate shape and thickness, providing the required load-bearing capacity with minimal mass. During the design process, a parametric analysis of a SPRE nozzle with a carbon fiber insert plate was carried out. By varying the thickness of the plate, an optimal design scheme that corresponds to the specified safety and stability factors was selected. Parametric analysis of an insert plate made of a composite material includes modeling of its main weight and strength parameters: analysis of the stress-strain state of the structure, values of natural frequencies, determination of the buckling margin, and determination of a SPRE nozzle mass. The analysis of the load-bearing capacity of a SPRE nozzle with an insert plate made of a composite material was carried out by the finite element method using the SolidWorks Simulation software package. When conducting a parametric analysis, two variants of a SPRE nozzle with and without an insert plate were considered. According to the results of a parametric analysis of a SPRE nozzle, its geometric dimensions were determined and the structure mass was minimized.
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Scientific article
A mathematical model of the aircraft avionics thermal state describing the heat exchange of the onboard equipment housing with a honeycomb structure made of a carbon fiber composite, the process of heat transfer of the onboard equipment elements and the air is developed. The considered heat transfer process in a heterogeneous medium is described by the boundary value problem for the heat equation with boundary conditions of the third kind. To solve the direct problem of the onboard equipment housing with a honeycomb structure thermal state, the Monte- Carlo method on the basis of the probabilistic representation of the solution in the form of an expectation of the functional of the diffusion process is used. The inverse problem of the honeycomb structure heat exchange is solved by minimizing the function of the squared residuals weighted sum using an iterative stochastic quasigradient algorithm. The developed mathematical model of the onboard equipment in the unpressurized compartment thermal state is used for optimizing the temperature and airflow of the thermal control system of the blown onboard equipment in the unpressurized compartment of the aircraft.
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Article
During the period of active lifetime in different types of orbits, spacecraft (SC) are exposed to various factors of outer space. The main factor influencing the performance of radio-electronic equipment is the ionizing radiation of outer space. The main effect associated with the impact of ionizing radiation of outer space on the SC (dose effects), in its physics, is determined by the radiation dose absorbed in the components of the electronic component base (ECB) and construction materials during the entire lifetime. This effect explains the failures in the operation of radio-electronic equipment due to the degradation of the parameters of the used ECB products and materials. Calculation of levels of exposure to absorbed doses depending on various parameters of the orbit is a necessary and important task for ensuring the SC functioning during a given lifetime, since calculations of radiation resistance are based on the levels of exposure in orbit. Carrying out calculations taking into account the SC and onboard equipment (OBE) design features is the key problem from the point of view of minimizing mass protection and the scope of tests of critical ECB products. In addition to taking into account the design of the SC and the OBE, an important aspect when calculating the radiation resistance is taking into account the relative position of the OBE in the SC. The paper considers the possibility of unifying the requirements for radiation resistance for SC with dif-ferent operating orbits and lifetime, as well as the possibility of carrying out a unified calculation without taking into account the design features of the SC, OBE and relative position within the SC.
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Scientific article
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 nst<40. 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.
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Pressure measurement inside non-sealed equipment bay of the geostationary spacecraft
Scientific article
The equipment of the non-sealed spacecraft is functioning during the operation phase in the envi-ronment that includes the equipment bay inherent atmosphere. JSC “ISS” together with MAI have developed and implemented in software a mathematical model for the pressure dynamics estimation of the equipment bay inherent atmosphere and for the design parameters selection and construction of the ventilation openings of the non-hermetic equipment bay. The pressure drop dynamics estimation of the equipment bay inherent atmosphere was also carried out according to the developed model. A block of the pressure sensors (joint development of ISS and Novosibirsk State University) was in-tegrated into the geostationary spacecraft for the in-situ pressure measurement inside the non-hermetic equipment bay during the operation. This block consists of two sensors: a semiconductor sensor based on MEMS- technology (micro-electromechanical system) and an inverse magnetron sen-sor with the cold cathode. The pressure sensor unit provides the pressure measurement from 790 up to 1∙108 mm Hg. The authors present the results of pressure measurements inside the non-hermetic equipment bay during the first six months of the spacecraft operation on the geostationary orbit. This article also compares the in-situ pressure measurement results and the calculated pressure drop obtained using the mathematical model for the non-hermetic equipment bay spacecraft.
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Prospects for the development of charger-discharge devices of spacecraft power supply systems
Scientific article
Power supply system (PSS) is designed to ensure uninterrupted autonomous power supply of on-board equipment in all modes and at all stages during the active life of spacecraft. Lots of PSS makes up a significant proportion of the mass spacecraft and improvement of specific energy characteristics of PSS leads to a synergistic effect when the mass and energy consumption resources available for the payload are simultaneously increased, which increases efficiency of spacecraft generally. The article considers the evolution of structural and circuit solutions for PSS spacecraft, its energyconverting equipment and the effect of these changes on energy-mass characteristics of PSS. It is shown that a significant effect on energy and mass characteristics of PSS provide structural and circuit design solutions for charging and discharging devices of energy-converting equipment and the choice of voltage value of recharging batteries (RB). The development of the element base, the creation of programmable digital devices capable of functioning under the influence of space factors and the emergence of new circuit design and management solutions for pulse converters that have occurred in the last decade opens up new opportunities for improvement of the PSS of spacecraft. In the article as a chagrining and discharging device of PSS a pulse voltage converter (PVC) with a new modulation strategy is reviewed, with the ability to reverse the flow of energy and the ability to work in a step-up mode with high efficiency. Its application as a single charger-discharge device (CDD) allows for a significant improvement in performance of CDD and PSS in general, such as efficiency, energy mass, reliability and a number of others. Ability of PVC to reverse the flow of energy and the possibility of working in a step-up mode opens up the possibility to abandon the use of RB with a voltage lower than the voltage at the main output of PSS and switch to using RB with an average discharge voltage close to the voltage at the main output of PSS. Such a structural and circuit design solution of CDD and RB will allow to increase efficiency of CDD up to 99 % and additionally improve energy and mass characteristics of PSS.
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Scientific article
Changing the low-voltage level of the output load power bus (27–28 V) in the power supply system (PSS) of the spacecraft (SC) to a high-voltage (100 V) allowed us to significantly reduce the SC mass in connection with the reduction in the mass of cables and energy converting equipment (ECE). However, a number of problems have arisen related to the difficulty of matching the increased voltage levels of energy sources and loads, taking into account the necessary level of reliability of the PSS. Therefore, the issues of choosing the PSS structure and methods for developing ECE are relevant and priority task facing their developers. To date, in the field of development and creation of high-voltage high-power PSS of SC, a promising direction is their design based on integrated ECE modules, in particular, on the basis of modules of charge-discharge regulators (CDR) of accumulator batteries (AB). In the article, a calculation and comparative analysis of the SC PSS structures with the connection of the CDR module to the solar battery (SB) bus and with the connection of the CDR module to the output load power bus is performed. In the course of analysis of the results obtained, it was found that both options for the PSS implementation can be optimal depending on the given curve of the SC load and the requirements for the PSS for specific energy, weightdimension and other characteristics. The final choice of the SC PSS structure should be made subject to the specific power of the ECE and the subsequent calculation of the weight-dimension characteristics of the alternative PSS. Simulation of two options for the implementation of the AB CDR module was carried out: a push-pull converter with one inductor and a Weinberg converter with a magnetically coupled inductor and an additional power diode. It is established that both investigated options can be used in the development and creation of the CDR module of the highvoltage PSS of spacecraft. However, the design of CDR module based on the Weinberg converter can significantly reduce the values of the used inductors and output capacitors subject to the required levels of output voltage ripple.
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Prospects of application of the combined storage space application
Scientific article
Combined energy storage systems are widely used as part of renewable sources in the manufacturing industry, transportation infrastructure, space engineering, and other industries. This trend is due to their higher reliability and efficiency than that of power supplies made of the same type of cells, which can be observed, in particular, in devices for space applications, where power cells are subject to higher requirements for the range of operating temperatures, as well as for input and output characteristics. The article describes the development of the structure of a portable combined energy storage device based on a block of supercapacitors and batteries with a charge and discharge control system, with a scalable (based on the components used) spectrum of input characteristics and a wide range of operating temperatures. The authors presented a mathematical model of a combined energy storage device developed in the Simulink environment, which makes it possible to assess the performance capabilities of the proposed structure by analyzing the different modes of operation of the circuit. Such a device can be used in conditions of extremely low charging currents. For example, if the solar panels are shaded or their spatial orientation is sub-optimal, high-capacity lithium-ion batteries cannot be charged correctly. Also, the advantages of combined electricity storage structures include their operability over a wide temperature range thanks to the ability of supercapacitors to retain their charge even at low temperatures. The article also shows the printed circuit assembly in the form of a 3D model obtained by designing the device circuit in Altium Designer 17 CAD; the results of research and performance testing of a physically implemented combined energy storage device are shown, which confirm its performance characteristics on the example of one of the component modules of the prototype satellite platform CubeSat; the article also provides recommendations for the possible application of such devices and highlights prospects for the application of combined energy storage devices in actuating elements of large flexible spacecraft.
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