Speeding up the integration of Ukraine into the European railway transportation is an important task in the current development of the Ukrainian railway transport. Currently, the most effective way to travel across borders between countries with different track gauges is the use of gauge-changeable wheelsets. Continuous traffic on the Ukrainian (1520 mm gauge) and European (1435 mm gauge) railways calls not only for gauge changing facilities, but also for the compatibility of the wheel-rail contact pair on both railways: R65 rails and a cant of 1/20 in Ukraine and UIC60 rails and a cant of 1/40 in Europe. At the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine, a new wear-resistant wheel profile, ITM-73ER, was proposed. As predicted, its use in passenger cars will offer dynamic stability and a high dynamic performance throughout the range of operating speeds on the Ukrainian and European railways and acceptable indices of wheel – unworn rail interaction on both railways. In service, the shapes of the wheel and rail contact surfaces change due to wear, especially in curves. Because the Ukrainian and European railways mostly use wheel and rail profiles of their own, the use of the new wheel profile may impair the wheel–rail interaction process, enhance wheel flange wear, and shorten the wheel life. The goal of this work is to study the effect of the in-service rail head shape change in curves of the Ukrainian and European railways on the wheel–rail interaction indices of a passenger car with ITM-73ER profile wheels. The head shapes of outer rails of the Ukrainian and European railways’ circular curves were predicted for a side flange wear changing from 0 to 8 mm in 2 mm increments. The calculations were made for two circular curves of radius 300 m with UIC60 rails and a cant of 1/40 (Europe) and R65 rails and a cant of 1/20 (Ukraine). To speed up the prediction, it was assumed that the curves were traveled by four-axle fully loaded freight cars, which maximizes the rail wear. The freight car wheels were assumed to be unworn and machined to the S1002 profile (for the European railways) and to the standard profile specified by the Ukrainian State Standard GOST 10791:2016 (for the Ukrainian railways). Using the computed head shapes of R65 and UIC60 rails differing in wear degree, a study was conducted into their effect on the wheel–rail pair strain and stress field and the dynamic indices of car–track interaction for passenger cars with ITM73-ER profile wheels negotiating a circular curve of radius 300 m. It was shown that the use of the ITM-73ER wheel profile in passenger cars will offer improved indices of car–track interaction, for worn rails too, both on the Ukrainian railways and in the combined operation on the Ukrainian and European railways.
{"title":"Wheel¬–rail interaction for a passenger car with the ITM-73ER new wheel profile in curves","authors":"T. Mokrii, I. Malysheva, L. Lapina, S.S. Pasichyk","doi":"10.15407/itm2023.02.084","DOIUrl":"https://doi.org/10.15407/itm2023.02.084","url":null,"abstract":"Speeding up the integration of Ukraine into the European railway transportation is an important task in the current development of the Ukrainian railway transport. Currently, the most effective way to travel across borders between countries with different track gauges is the use of gauge-changeable wheelsets. Continuous traffic on the Ukrainian (1520 mm gauge) and European (1435 mm gauge) railways calls not only for gauge changing facilities, but also for the compatibility of the wheel-rail contact pair on both railways: R65 rails and a cant of 1/20 in Ukraine and UIC60 rails and a cant of 1/40 in Europe. At the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine, a new wear-resistant wheel profile, ITM-73ER, was proposed. As predicted, its use in passenger cars will offer dynamic stability and a high dynamic performance throughout the range of operating speeds on the Ukrainian and European railways and acceptable indices of wheel – unworn rail interaction on both railways. In service, the shapes of the wheel and rail contact surfaces change due to wear, especially in curves. Because the Ukrainian and European railways mostly use wheel and rail profiles of their own, the use of the new wheel profile may impair the wheel–rail interaction process, enhance wheel flange wear, and shorten the wheel life. The goal of this work is to study the effect of the in-service rail head shape change in curves of the Ukrainian and European railways on the wheel–rail interaction indices of a passenger car with ITM-73ER profile wheels. The head shapes of outer rails of the Ukrainian and European railways’ circular curves were predicted for a side flange wear changing from 0 to 8 mm in 2 mm increments. The calculations were made for two circular curves of radius 300 m with UIC60 rails and a cant of 1/40 (Europe) and R65 rails and a cant of 1/20 (Ukraine). To speed up the prediction, it was assumed that the curves were traveled by four-axle fully loaded freight cars, which maximizes the rail wear. The freight car wheels were assumed to be unworn and machined to the S1002 profile (for the European railways) and to the standard profile specified by the Ukrainian State Standard GOST 10791:2016 (for the Ukrainian railways). Using the computed head shapes of R65 and UIC60 rails differing in wear degree, a study was conducted into their effect on the wheel–rail pair strain and stress field and the dynamic indices of car–track interaction for passenger cars with ITM73-ER profile wheels negotiating a circular curve of radius 300 m. It was shown that the use of the ITM-73ER wheel profile in passenger cars will offer improved indices of car–track interaction, for worn rails too, both on the Ukrainian railways and in the combined operation on the Ukrainian and European railways.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"326 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114382032","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The goal of this paper is to identify and classify the functional features of the shell of a space industrial platform. Further space exploration is limited by the difficulty of launching large-sized and massive objects into orbit. At the initial stage of the industrialization of near space, space industrial platforms can be placed therein. The configurations of existing orbital stations as a prototype of space industrial platforms are analyzed, and the ranges of the main parameters of their modules are determined. The structural layout of an industrial platform depends significantly on the technological processes implemented thereon. The configuration of a platform depends largely on a number of criteria that have an effect on its functional features. The paper identifies a number of criteria that have an effect on the functional features and configuration of a platform. They are as follows: structural modularity, the shell frame type, the shell shaping method, the sealing type, the need for a special atmosphere, the need for special process modules, the orientation and stabilization type, the power system type, the thermal control type, the need for a microclimate, and the type of preprocessing of raw materials and their components. Using these criteria, the paper proposes a classification of the functional features of an industrial platform shell. For classification, a space industrial platform is decomposed down to the level of the structural elements of its modules to be used in the development of a comprehensive mathematical model of platform operation. A set of parameters of industrial platforms is formed to ensure the technological processes implemented thereon. The set will be used in platform shell formation according to process parameters. The problem of mass optimization of a space industrial platform is formulated.
{"title":"Classification of the functional features of the shell of a space industrial plat-form","authors":"O. Palii","doi":"10.15407/itm2023.02.064","DOIUrl":"https://doi.org/10.15407/itm2023.02.064","url":null,"abstract":"The goal of this paper is to identify and classify the functional features of the shell of a space industrial platform. Further space exploration is limited by the difficulty of launching large-sized and massive objects into orbit. At the initial stage of the industrialization of near space, space industrial platforms can be placed therein. The configurations of existing orbital stations as a prototype of space industrial platforms are analyzed, and the ranges of the main parameters of their modules are determined. The structural layout of an industrial platform depends significantly on the technological processes implemented thereon. The configuration of a platform depends largely on a number of criteria that have an effect on its functional features. The paper identifies a number of criteria that have an effect on the functional features and configuration of a platform. They are as follows: structural modularity, the shell frame type, the shell shaping method, the sealing type, the need for a special atmosphere, the need for special process modules, the orientation and stabilization type, the power system type, the thermal control type, the need for a microclimate, and the type of preprocessing of raw materials and their components. Using these criteria, the paper proposes a classification of the functional features of an industrial platform shell. For classification, a space industrial platform is decomposed down to the level of the structural elements of its modules to be used in the development of a comprehensive mathematical model of platform operation. A set of parameters of industrial platforms is formed to ensure the technological processes implemented thereon. The set will be used in platform shell formation according to process parameters. The problem of mass optimization of a space industrial platform is formulated.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"27 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126481036","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper presents algorithms to calculate supersonic flow about a prospective ring wing launch vehicle by the marching method and the relaxation method. The feature of the algorithms is the introduction of two computational subregions in the ring wing zone over the rocket airframe. In the marching algorithm, the computation region is reconstructed according to the position of the marching cross-section relative to the leading and trailing edge of the ring wing. When it finds itself at the leading edge of the ring wing, the computational region is split into a lower subregion between the rocket airframe and the downstream face of the ring wing and an upper subregion between the upstream face of the ring wing and the bow shock front. When the marching cross-section finds itself at the trailing edge of the ring wing, the lower and the upper computational subregions are merged into a single computational region. Based on the marching algorithm and using the authors’ rocket flow calculation program, software is developed for a fast numerical calculation of supersonic flow about ring wing rockets. For a particular ring wing rocket configuration, the paper presents the results of comparative calculations of supersonic flow about the rocket in the form of gas-dynamic parameter isolines in the flow field and the pressure distribution over the rocket airframe and the ring wing. The results for the marching method and the relaxation method are compared. It is shown that the ring wing is responsible for an undulatory pressure distribution between the rocket airframe and the downstream face of the ring wing. The marching method simulates the flow pattern between the rocket airframe and the downstream face of the ring wing more adequately, and its computation time is two orders of magnitude shorter than that of the relaxation method. The relaxation method should be used in the case of subsonic flows between the rocket airframe and the downstream face of the ring wing. The algorithm and software developed are recommended for parametric calculations of supersonic flow about ring wing rockets.
{"title":"Algorithms to calculate supersonic flow about a ring wing launch vehicle","authors":"V. I. Timoshenko, V. P. Halynskyi","doi":"10.15407/itm2023.02.003","DOIUrl":"https://doi.org/10.15407/itm2023.02.003","url":null,"abstract":"This paper presents algorithms to calculate supersonic flow about a prospective ring wing launch vehicle by the marching method and the relaxation method. The feature of the algorithms is the introduction of two computational subregions in the ring wing zone over the rocket airframe. In the marching algorithm, the computation region is reconstructed according to the position of the marching cross-section relative to the leading and trailing edge of the ring wing. When it finds itself at the leading edge of the ring wing, the computational region is split into a lower subregion between the rocket airframe and the downstream face of the ring wing and an upper subregion between the upstream face of the ring wing and the bow shock front. When the marching cross-section finds itself at the trailing edge of the ring wing, the lower and the upper computational subregions are merged into a single computational region. Based on the marching algorithm and using the authors’ rocket flow calculation program, software is developed for a fast numerical calculation of supersonic flow about ring wing rockets. For a particular ring wing rocket configuration, the paper presents the results of comparative calculations of supersonic flow about the rocket in the form of gas-dynamic parameter isolines in the flow field and the pressure distribution over the rocket airframe and the ring wing. The results for the marching method and the relaxation method are compared. It is shown that the ring wing is responsible for an undulatory pressure distribution between the rocket airframe and the downstream face of the ring wing. The marching method simulates the flow pattern between the rocket airframe and the downstream face of the ring wing more adequately, and its computation time is two orders of magnitude shorter than that of the relaxation method. The relaxation method should be used in the case of subsonic flows between the rocket airframe and the downstream face of the ring wing. The algorithm and software developed are recommended for parametric calculations of supersonic flow about ring wing rockets.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"211 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"116454267","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper is concerned with the reliability and safety of launch complexes. The problems to be solved in launch complex reliability evaluation are identified: calculations of the probability of no-failure operation of passive redundancy systems with equal- and nonequal-reliability elements, reliability analysis for replacement redundancy with integer multiplicity and unloaded reserve; calculations of the probability of no-failure operation of the launch complex components in launch preparation, and calculations of the reliability indices of a component part as a whole and a comparison of the calculated reliability indices with the specification requirements. Since a launch complex consists both of renewable elements and of nonrenewable ones, the reliability indices must be calculated so that one may evaluate the reliability both of individual elements and of a system of different-type elements as a whole. These indices are characterized by the nonfailure operation time and recovery time distributions and show the probability of а serviceable state or a failure state of an element and a system. On condition that the nonfailure operation time and the recovery time can be described by the Weibull distribution, expressions are obtained for the availability factor, i.e., the probability of the launch complex being operative at an arbitrary time, except for scheduled periods during which the launch complex is not envisaged for use. Launch complex safety is evaluated by the probability of hazards, the identification of main ways to mitigate their consequences, and account for weight of the consequences of possible hazards in service. Launch complex safety indices are identified. It is shown that safety must be evaluated using indices suitable for the practical solution of problems of the justification and assurance of specified safety requirements against possible threats in the development of launch complexes. The adopted safety index is the probability that each hazard that occurs in a certain time will be eliminated. A renewal process is used to describe a random number of hazard occurrences. To determine the hazard frequency, it is recommended to use statistical data on launch complex accident rate and reliability, logical methods of event tree and fault tree analysis, accident simulation models, and expert judgments.
{"title":"Reliability and safety of launch complexes","authors":"V. Poshyvalov, Yu.F. Daniiev","doi":"10.15407/itm2023.02.013","DOIUrl":"https://doi.org/10.15407/itm2023.02.013","url":null,"abstract":"This paper is concerned with the reliability and safety of launch complexes. The problems to be solved in launch complex reliability evaluation are identified: calculations of the probability of no-failure operation of passive redundancy systems with equal- and nonequal-reliability elements, reliability analysis for replacement redundancy with integer multiplicity and unloaded reserve; calculations of the probability of no-failure operation of the launch complex components in launch preparation, and calculations of the reliability indices of a component part as a whole and a comparison of the calculated reliability indices with the specification requirements. Since a launch complex consists both of renewable elements and of nonrenewable ones, the reliability indices must be calculated so that one may evaluate the reliability both of individual elements and of a system of different-type elements as a whole. These indices are characterized by the nonfailure operation time and recovery time distributions and show the probability of а serviceable state or a failure state of an element and a system. On condition that the nonfailure operation time and the recovery time can be described by the Weibull distribution, expressions are obtained for the availability factor, i.e., the probability of the launch complex being operative at an arbitrary time, except for scheduled periods during which the launch complex is not envisaged for use. Launch complex safety is evaluated by the probability of hazards, the identification of main ways to mitigate their consequences, and account for weight of the consequences of possible hazards in service. Launch complex safety indices are identified. It is shown that safety must be evaluated using indices suitable for the practical solution of problems of the justification and assurance of specified safety requirements against possible threats in the development of launch complexes. The adopted safety index is the probability that each hazard that occurs in a certain time will be eliminated. A renewal process is used to describe a random number of hazard occurrences. To determine the hazard frequency, it is recommended to use statistical data on launch complex accident rate and reliability, logical methods of event tree and fault tree analysis, accident simulation models, and expert judgments.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"10 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114164924","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Truncated nozzles are used for tight packing of the rocket engine. Such nozzles have a profiled tip to maximize the filling of space and reduce the overall weight. This paper is concerned with the study the effect of the tip geometry of a truncated supersonic nozzle on its characteristics. The features of the gas flow at different initial pressures and different environmental conditions in the supersonic area of a nozzle with a bell-shaped tip of different lengths are considered. The flow inside the nozzle followed by the jet outflow into the surrounding space was simulated. The flow simulation for tips at sea level showed a similar structure of the Mach number isolines, and the only difference was in the intensity of the vortex structure near the tip wall. As the pressure at the nozzle inlet increases, the length of the first “barrel” increases proportionally, and the vortex structure near the tip walls decreases. For the upper atmosphere, the flow pattern is different. The supersonic flow in the nozzle does not undergo separation, and therefore there are no vortex structures from the external environment. The flow downstream of the tip exit deflects from the axis through the angle determined by the Prandtl–Meier flow at the corner point of the tip exit, and the shape of the first “barrel” is distorted by a hanging shock. An analysis of the obtained results shows that the ambient pressure downstream the nozzle exit significantly affects the flow pattern in the nozzle. It is established that the thrust coefficient of both circuits at sea level decreases with increasing pressure at the nozzle inlet, which is explained by a decrease in the effect of the ambient pressure on the tip wall. In the upper atmosphere, the flow is adjacent to the tip wall, and the thrust coefficient for nozzles of different lengths has almost the same constant value at different inlet pressures. It is shown that a decrease in the length of the nozzle, all other geometrical dimensions of the nozzle being equal, does not significantly affect the impulse characteristics.
{"title":"Effect of the tip geometry of a truncated supersonic nozzle on its characteristics","authors":"K. Ternova","doi":"10.15407/itm2023.02.032","DOIUrl":"https://doi.org/10.15407/itm2023.02.032","url":null,"abstract":"Truncated nozzles are used for tight packing of the rocket engine. Such nozzles have a profiled tip to maximize the filling of space and reduce the overall weight. This paper is concerned with the study the effect of the tip geometry of a truncated supersonic nozzle on its characteristics. The features of the gas flow at different initial pressures and different environmental conditions in the supersonic area of a nozzle with a bell-shaped tip of different lengths are considered. The flow inside the nozzle followed by the jet outflow into the surrounding space was simulated. The flow simulation for tips at sea level showed a similar structure of the Mach number isolines, and the only difference was in the intensity of the vortex structure near the tip wall. As the pressure at the nozzle inlet increases, the length of the first “barrel” increases proportionally, and the vortex structure near the tip walls decreases. For the upper atmosphere, the flow pattern is different. The supersonic flow in the nozzle does not undergo separation, and therefore there are no vortex structures from the external environment. The flow downstream of the tip exit deflects from the axis through the angle determined by the Prandtl–Meier flow at the corner point of the tip exit, and the shape of the first “barrel” is distorted by a hanging shock. An analysis of the obtained results shows that the ambient pressure downstream the nozzle exit significantly affects the flow pattern in the nozzle. It is established that the thrust coefficient of both circuits at sea level decreases with increasing pressure at the nozzle inlet, which is explained by a decrease in the effect of the ambient pressure on the tip wall. In the upper atmosphere, the flow is adjacent to the tip wall, and the thrust coefficient for nozzles of different lengths has almost the same constant value at different inlet pressures. It is shown that a decrease in the length of the nozzle, all other geometrical dimensions of the nozzle being equal, does not significantly affect the impulse characteristics.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129950341","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
K. Avramov, I. Biblik, I.V. Hrebennik, I. Urniaieva
A system of nonlinear partial differential equations is derived to describe the vibrations of a multi-walled nanotube. The system reduces to a nonlinear dynamic system with а large number of degrees of freedom (DOFs). To reduce its dimension, the nonlinear modal analysis method is used to give 2-DOF dynamic system, which is studied by the asymptotic multiple scale method. This gives a system of modulation equations, whose fixed points describe the free vibrations of the nanotube. The fixed points are described by nonlinear algebraic equations, whose solutions are given on a backbone curve. Use is made of the Sanders–Koiter shell model to describe the nonlinear deformation of the nanotube and Hook’s nonlocal anisotropic law to simulate its vibrations. Notice that the elastic constants of the nanotube walls differ. The nanotube model is a system of nonlinear ordinary differential equations, which is obtained by applying the weighed residuals method to the nonlinear partial equations. Three types of nonlinearities are accounted for in the nanotube model. First, the Van der Waals forces are nonlinear functions of the radial displacements. Second, the displacements of the nanotube walls are assumed to be moderate, which is described by a geometrically nonlinear model. Third, since the resultant forces are nonlinear functions of the displacements, the use of natural boundary conditions in the weighted residuals method results in additional nonlinear terms. A finite-DOF nonlinear dynamical system is derived. The free nonlinear vibrations of the nanotube are analyzed. The calculated results are shown on a backbone curve.
{"title":"Reducing the dimension of a nonlinear dynamic system to simulate a multi-walled nanotube","authors":"K. Avramov, I. Biblik, I.V. Hrebennik, I. Urniaieva","doi":"10.15407/itm2023.02.105","DOIUrl":"https://doi.org/10.15407/itm2023.02.105","url":null,"abstract":"A system of nonlinear partial differential equations is derived to describe the vibrations of a multi-walled nanotube. The system reduces to a nonlinear dynamic system with а large number of degrees of freedom (DOFs). To reduce its dimension, the nonlinear modal analysis method is used to give 2-DOF dynamic system, which is studied by the asymptotic multiple scale method. This gives a system of modulation equations, whose fixed points describe the free vibrations of the nanotube. The fixed points are described by nonlinear algebraic equations, whose solutions are given on a backbone curve. Use is made of the Sanders–Koiter shell model to describe the nonlinear deformation of the nanotube and Hook’s nonlocal anisotropic law to simulate its vibrations. Notice that the elastic constants of the nanotube walls differ. The nanotube model is a system of nonlinear ordinary differential equations, which is obtained by applying the weighed residuals method to the nonlinear partial equations. Three types of nonlinearities are accounted for in the nanotube model. First, the Van der Waals forces are nonlinear functions of the radial displacements. Second, the displacements of the nanotube walls are assumed to be moderate, which is described by a geometrically nonlinear model. Third, since the resultant forces are nonlinear functions of the displacements, the use of natural boundary conditions in the weighted residuals method results in additional nonlinear terms. A finite-DOF nonlinear dynamical system is derived. The free nonlinear vibrations of the nanotube are analyzed. The calculated results are shown on a backbone curve.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"78 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"128321213","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The ever-increasing clogging of near-Earth space by space debris objects of various sizes significantly limits the possibilities of space activities and poses a great danger to the Earth’s objects. This is especially true for low orbits with altitudes up to 2,000 km. The risk of collision of operating spacecraft with space debris threatens their functioning in near-Earth space. To control space debris, use is made of active and passive methods of space debris removal from operational orbits. At present, promising means of space debris removal are a space debris transfer to low-Earth orbits with a lifetime of less than twenty-five years, a transfer to a junk obit, and in-orbit utilization. According to the latest recommendations, space debris objects moved to low-Earth orbits should have a lifetime of less than twenty-five years. In the dense atmosphere, small space debris objects usually burn up completely, while large ones burn up only partially and may reach the Earth. Since space debris motion in the atmosphere can only be predicted with large errors, a timely and accurate prediction of the place and time of fall of large space debris objects onto the Earth is impossible. Space debris objects can remain in junk orbits for hundreds of years without interfering with space projects. This method of space debris removal reduces the risk of collision with space debris objects in the initial orbit, but increases it in the junk one. According to the concept of in-orbit utilization, space debris is considered a resource for the in-orbit industry. An active space debris removal involves high energy expenditures of service spacecraft. In this regard, the task of their estimation becomes important. The goal of this paper is a comparative assessment of the energy expenditures for moving space debris objects into utilization orbits using service spacecraft with electrojet propulsion systems. The problem is solved using methods of flight dynamics, averaging, and mathematical simulation. The novelty of the obtained results lies in the development of a ballistic scheme and a fast procedure to calculate energy expenditures for moving space debris objects to a disposal orbit using service spacecraft with constant low-thrust electrojet propulsion system. The procedure may be used in substantiating and planning space debris transfer from low-eccentricity low-Earth orbits to utilization orbits.
{"title":"Energy expenditures for moving space debris objects from low-Earth orbits to utilization orbits","authors":"Yu.M. Holdshtein","doi":"10.15407/itm2023.02.041","DOIUrl":"https://doi.org/10.15407/itm2023.02.041","url":null,"abstract":"The ever-increasing clogging of near-Earth space by space debris objects of various sizes significantly limits the possibilities of space activities and poses a great danger to the Earth’s objects. This is especially true for low orbits with altitudes up to 2,000 km. The risk of collision of operating spacecraft with space debris threatens their functioning in near-Earth space. To control space debris, use is made of active and passive methods of space debris removal from operational orbits. At present, promising means of space debris removal are a space debris transfer to low-Earth orbits with a lifetime of less than twenty-five years, a transfer to a junk obit, and in-orbit utilization. According to the latest recommendations, space debris objects moved to low-Earth orbits should have a lifetime of less than twenty-five years. In the dense atmosphere, small space debris objects usually burn up completely, while large ones burn up only partially and may reach the Earth. Since space debris motion in the atmosphere can only be predicted with large errors, a timely and accurate prediction of the place and time of fall of large space debris objects onto the Earth is impossible. Space debris objects can remain in junk orbits for hundreds of years without interfering with space projects. This method of space debris removal reduces the risk of collision with space debris objects in the initial orbit, but increases it in the junk one. According to the concept of in-orbit utilization, space debris is considered a resource for the in-orbit industry. An active space debris removal involves high energy expenditures of service spacecraft. In this regard, the task of their estimation becomes important. The goal of this paper is a comparative assessment of the energy expenditures for moving space debris objects into utilization orbits using service spacecraft with electrojet propulsion systems. The problem is solved using methods of flight dynamics, averaging, and mathematical simulation. The novelty of the obtained results lies in the development of a ballistic scheme and a fast procedure to calculate energy expenditures for moving space debris objects to a disposal orbit using service spacecraft with constant low-thrust electrojet propulsion system. The procedure may be used in substantiating and planning space debris transfer from low-eccentricity low-Earth orbits to utilization orbits.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"8 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114872161","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The goal of this article is to develop an effective image preprocessing algorithm and a neural network model for determining the force to be transmitted to a space debris object (SDO) for its non-contact deorbit. In the development and study of the algorithm, use was made of methods of theoretical mechanics, machine learning, computer vision, and computer simulation. The force is determined using a photo taken by an onboard camera. To increase the efficiency of the neural network, an algorithm was developed for feature recognition by the SDO edge in the photo. The algorithm, on the one hand, selects a sufficient number of features to describe the properties of the figure and, on the other hand, significantly reduces the amount of data at the neural network input. A dataset with the features and corresponding reference force values was created for model training. A neural network model was developed to determine the force to be exerted on a SDO from the SDO features. The model was tested using a set of eighteen calculated cases to determine the effectiveness, accuracy, and speed of the algorithm. The proposed algorithm was compared with two existing ones: the method of central projections onto an auxiliary plane and the multilayered neural network model that calculates the force using the SDO orientation parameters. The comparison was performed using the root mean square error, the maximum absolute error, and the maximum relative error. The test results are presented as tables and graphs. The proposed approach makes it possible to develop a system of SDO non-contact removal that does not need to determine the exact relative position and orientation with respect to the active spacecraft. Instead, the algorithm uses camera-taken photos, from which the features necessary for calculation are extracted. This makes it possible to reduce the requirements for its computing elements, to abandon sensors for determining the relative position and orientation, and to reduce the cost of the system.
{"title":"Determination of the force exerted by an ion beam on a space debris object from the edges of its images using deep learning","authors":"M. Redka","doi":"10.15407/itm2023.02.051","DOIUrl":"https://doi.org/10.15407/itm2023.02.051","url":null,"abstract":"The goal of this article is to develop an effective image preprocessing algorithm and a neural network model for determining the force to be transmitted to a space debris object (SDO) for its non-contact deorbit. In the development and study of the algorithm, use was made of methods of theoretical mechanics, machine learning, computer vision, and computer simulation. The force is determined using a photo taken by an onboard camera. To increase the efficiency of the neural network, an algorithm was developed for feature recognition by the SDO edge in the photo. The algorithm, on the one hand, selects a sufficient number of features to describe the properties of the figure and, on the other hand, significantly reduces the amount of data at the neural network input. A dataset with the features and corresponding reference force values was created for model training. A neural network model was developed to determine the force to be exerted on a SDO from the SDO features. The model was tested using a set of eighteen calculated cases to determine the effectiveness, accuracy, and speed of the algorithm. The proposed algorithm was compared with two existing ones: the method of central projections onto an auxiliary plane and the multilayered neural network model that calculates the force using the SDO orientation parameters. The comparison was performed using the root mean square error, the maximum absolute error, and the maximum relative error. The test results are presented as tables and graphs. The proposed approach makes it possible to develop a system of SDO non-contact removal that does not need to determine the exact relative position and orientation with respect to the active spacecraft. Instead, the algorithm uses camera-taken photos, from which the features necessary for calculation are extracted. This makes it possible to reduce the requirements for its computing elements, to abandon sensors for determining the relative position and orientation, and to reduce the cost of the system.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"1231 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123029715","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The goal of this paper is to develop elements of a simulation algorithm for determining the controlled dynamic parameters of the sustainer stages of launch vehicles (LVs) equipped with an active control system (ACS). In this study, methods of system analysis and computational rocket dynamics were used. The paper proposes a system approach to the organization of LV ACS information support with account for specified limiting values of the controlled dynamic parameters: the pitch rate, the velocity pressure, and the angle of attack. In flight, the LV ACS uses information on these parameters to suppress bending deformations of the LV structure and form a trajectory close to the energy-optimal one. The controlled dynamic parameters were brought to a simplified form, thus making it possible to take the data needed for their calculation from the inertial sensors of the LV control system. Simulation algorithm elements were developed to determine the dynamic parameters from the actual values of the center of mass motion parameters in the launch coordinate system, which can be obtained from their calculated values and the corresponding isochronous variations of their apparent values in the inertial coordinate system. The elements of the simulation algorithm for the determination of the LV sustainer stage dynamic parameters may be used in the development of ACS methodological support. The main advantage of the proposed system approach with account for specified limiting values of the controlled dynamic parameters is that it does not require any detailed simulation of dynamic loads on the LV sustainer stages and uses nothing but information on the dynamic parameters that characterize LV trajectory motion conditions.
{"title":"System approach to the organization of information support for a launch vehi-cle’s active control system driven by the current operating conditions","authors":"V. Gorbuntsov, O. Zavoloka","doi":"10.15407/itm2023.02.021","DOIUrl":"https://doi.org/10.15407/itm2023.02.021","url":null,"abstract":"The goal of this paper is to develop elements of a simulation algorithm for determining the controlled dynamic parameters of the sustainer stages of launch vehicles (LVs) equipped with an active control system (ACS). In this study, methods of system analysis and computational rocket dynamics were used. The paper proposes a system approach to the organization of LV ACS information support with account for specified limiting values of the controlled dynamic parameters: the pitch rate, the velocity pressure, and the angle of attack. In flight, the LV ACS uses information on these parameters to suppress bending deformations of the LV structure and form a trajectory close to the energy-optimal one. The controlled dynamic parameters were brought to a simplified form, thus making it possible to take the data needed for their calculation from the inertial sensors of the LV control system. Simulation algorithm elements were developed to determine the dynamic parameters from the actual values of the center of mass motion parameters in the launch coordinate system, which can be obtained from their calculated values and the corresponding isochronous variations of their apparent values in the inertial coordinate system. The elements of the simulation algorithm for the determination of the LV sustainer stage dynamic parameters may be used in the development of ACS methodological support. The main advantage of the proposed system approach with account for specified limiting values of the controlled dynamic parameters is that it does not require any detailed simulation of dynamic loads on the LV sustainer stages and uses nothing but information on the dynamic parameters that characterize LV trajectory motion conditions.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"2 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129574383","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
In recent years, more and more attention has been paid to nozzles with an unconventional profile, which differs from that of the classical streamline-profiled Laval nozzle. In such nozzles, the flow fields typically include interacting supersonic and subsonic flows, often with recirculation regions and a complex wave structure of the flow. This work is concerned with a numerical study of the wave structure of the gas flow in a truncated supersonic nozzle with an elliptical bell-shaped tip whose length is long in comparison with the conical section upstream of the tip. The gas flow inside the nozzle and in the surrounding space was simulated using the ANSYS software package. The calculations were carried out in a non-stationary axisymmetric formulation based on the Reynolds-averaged Navier–Stokes equations closed with the use of the SST turbulence model with near-wall functions and a compressibility correction. In the calculations, the nozzle inlet pressure and the ambient pressure were varied. The correctness of the methodological approaches to the solution of the problem was confirmed in the authors’ previous works. The study showed the following. At low values of the nozzle inlet pressure (P0 < 50 bar) and an ambient pressure of 1 bar, the tip wall exhibits a developed separation zone with a large-scale vortex and a small-scale one (near the tip exit). The first "barrel" of the outflowing gas shows a "saddle" low-intensity compression wave structure. In the case of a separated flow, the tip wall pressure in the separation zone is about 15% less than the ambient pressure. At P0 > 100 bar, the tip wall pressure is nearly proportional to the nozzle inlet pressure. In the upper atmosphere, when going in a radial direction from the nozzle axis at the tip exit cross-section, the static pressure monotonically decreases, reaches a minimum, and then increases linearly to the its maximum value on the tip wall. In the case of a separated flow in the tip at a sea-level ambient pressure, the static pressure at the tip exit cross-section behaves in the same manner for inlet pressures P0 > 50 bar. At P0 = 50 bar, there exist two extrema: the pressure first deceases to its minimum value, then increases to its maximum value, and then decreases slightly to its value on the tip wall.
{"title":"Wave structure of the gas flow in a truncated nozzle with a long bell-shaped tip","authors":"G. Strelnikov, N. Pryadko, K. Ternova","doi":"10.15407/itm2023.01.040","DOIUrl":"https://doi.org/10.15407/itm2023.01.040","url":null,"abstract":"In recent years, more and more attention has been paid to nozzles with an unconventional profile, which differs from that of the classical streamline-profiled Laval nozzle. In such nozzles, the flow fields typically include interacting supersonic and subsonic flows, often with recirculation regions and a complex wave structure of the flow. This work is concerned with a numerical study of the wave structure of the gas flow in a truncated supersonic nozzle with an elliptical bell-shaped tip whose length is long in comparison with the conical section upstream of the tip. The gas flow inside the nozzle and in the surrounding space was simulated using the ANSYS software package. The calculations were carried out in a non-stationary axisymmetric formulation based on the Reynolds-averaged Navier–Stokes equations closed with the use of the SST turbulence model with near-wall functions and a compressibility correction. In the calculations, the nozzle inlet pressure and the ambient pressure were varied. The correctness of the methodological approaches to the solution of the problem was confirmed in the authors’ previous works. The study showed the following. At low values of the nozzle inlet pressure (P0 < 50 bar) and an ambient pressure of 1 bar, the tip wall exhibits a developed separation zone with a large-scale vortex and a small-scale one (near the tip exit). The first \"barrel\" of the outflowing gas shows a \"saddle\" low-intensity compression wave structure. In the case of a separated flow, the tip wall pressure in the separation zone is about 15% less than the ambient pressure. At P0 > 100 bar, the tip wall pressure is nearly proportional to the nozzle inlet pressure. In the upper atmosphere, when going in a radial direction from the nozzle axis at the tip exit cross-section, the static pressure monotonically decreases, reaches a minimum, and then increases linearly to the its maximum value on the tip wall. In the case of a separated flow in the tip at a sea-level ambient pressure, the static pressure at the tip exit cross-section behaves in the same manner for inlet pressures P0 > 50 bar. At P0 = 50 bar, there exist two extrema: the pressure first deceases to its minimum value, then increases to its maximum value, and then decreases slightly to its value on the tip wall.","PeriodicalId":287730,"journal":{"name":"Technical mechanics","volume":"36 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-04-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"124721145","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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