1. Guidance, navigation and control system of launch vehicles, booster stages and satellites
2. Ballistic-navigational support of launch vehicles, booster stages and satellites
3. Orbital debris problem.
17
Scopus Publications
Scopus Publications
ESTIMATING THE DEGREE OF DISPOSAL OF A LAUNCH VEHICLE CASING MADE FROM POLYOLEFINS IN THE EARTH’S ATMOSPHERE Mykola Dron, Erik Lapkhanov, Aleksandr Golubek, Andrii Dreus, Olena Kositsyna, et al. Eastern European Journal of Enterprise Technologies, 2025 The object of this study is the process of disposing of the upper stage body of a launch vehicle made of polyolefins by burning in the dense layers of the Earth’s atmosphere during removal from Earth orbit. The task addressed was to determine the possibility of disposal of the upper stage bodies of launch vehicles made of polyolefins during deorbiting. The mathematical model built makes it possible to take into account the effect of ballistic and aerothermodynamic aspects at the same time. The application of this model makes it possible to estimate the degree of disposal of the upper stage bodies of launch vehicles made of polyolefins in the Earth’s atmosphere at the stage of scientific research. In turn, this makes it possible to rationally choose the design parameters of materials for launch vehicle bodies, taking into account the disposal phase in the dense layers of the atmosphere, as well as rationally select the initial parameters for deorbiting orbits. This makes it possible to maximize the level of disposal and minimize the probability of debris falling on uninhabited areas of the Earth. The results of the study showed that launch vehicle bodies made of polymer materials such as polyethylene and polypropylene could burn up in the atmospheric part of the trajectory by 90–100 %, depending on the mass-dimensional characteristics and the type of orbit. In turn, increasing the ellipticity of the orbit makes it possible to increase the steepness of the entry of the upper body of the launch vehicle into the dense layers of the atmosphere, and hence, to increase the heat flows that contribute to the combustion of the body. With this in mind, methodological recommendations have been compiled for choosing orbits of the necessary ellipticity, taking into account the place of fall of fragments of the upper bodies of carrier rockets that did not burn up in the atmosphere
CORRECTING MEASUREMENTS OF LAUNCH VEHICLE’S ANGULAR MOTION PARAMETERS OF A STRAPDOWN INERTIAL NAVIGATION SYSTEM WITH THE USE OF A CELESTIAL NAVIGATION SYSTEM O. GOLUBEK Science and Innovation, 2024 Introduction. One of the tasks of developing strapdown inertial navigation systems with microelectromechanical sensors of launch vehicles is to ensure stringent requirements for the accuracy in determining the linear and angular motion with a guarantee of the successful completion of the satellite injection mission. parameters of launch vehicle’s strapdown inertial navigation system built with the use of microelectromechanical sensors. One of the ways to compensate for this degradation is the integrated use of inertial and celestial navigation systems.Purpose. The purpose is to increase the satellite injection accuracy by launch vehicle using a strapdown inertial navigation system with microelectromechanical sensors due to a celestial navigation system with star tracker.Material and Methods. Development of Kalman fi lter providing an integrated processing of measurements of angular motion parameters of a launch vehicle by a strapdown inertial navigation system and a star tracker. Statistical modelling of launch vehicle flight under the infl uence of various stochastic disturbances. Statistical processing of modelling results. Analysis of the proposed solution eff ectiveness.Results. A workable mathematical model for correcting measurements of angular motion parameters of launch vehicle with the use of a Kalman fi lter has been developed. Its performance has been tested by the example of a launch vehicle injection into a sun synchronous orbit 700 km high with the use of a two-pulse injection scheme. It has been demonstrated that the proposed solution makes it possible to improve the accuracy of the angular orientation up to 90% and the satellite injection up to 5% in terms of altitude and orbit inclination.Conclusions. The proposed development can be used to build navigation systems for advanced launch vehicles.
ASSESSING THE POSSIBILITY OF USING A VARIABLE-LENGTH LAUNCH VEHICLE WITH A POLYMER BODY FOR ORBITING PAYLOAD Aleksandr Golubek, Serhii Aleksieienko, Mykola Dron, Andrii Dreus Eastern European Journal of Enterprise Technologies, 2024 The object of this study was the motion of an ultralight class variable-length launch vehicle made of a polymer body along the active phase of the trajectory. The work considers the solution to the problem of designing low-cost means of delivery to orbit, namely to the assessment of the possibility of removing the payload by a carrier rocket with a polymer body of variable length beyond the dense atmosphere of the Earth. For this purpose, ballistic projection of the trajectory of the launch vehicle was carried out taking into account overloading; its aerodynamic characteristics and peculiarities of aerothermodynamic processes occurring during flight in the atmospheric phase of the trajectory were determined. The closeness (up to 10 %) of the obtained results with known experimental data is shown. The influence of the aerodynamic force on the parameters of the launch vehicle motion was studied. A flight simulation was conducted, the results of which showed the fundamental possibility of launching a CubeSat 24U class payload using a launch vehicle with a polymer body of variable length to a suborbital trajectory with an altitude of about 300 km. At the same time, the effective longitudinal overload on the body of the launch vehicle does not exceed 4 units, and the temperature on the surface of the body does not exceed 300 K. A feature of the research is the use of a multidisciplinary approach, which implies taking into account the interrelationship of aerodynamic, thermodynamic, and ballistic processes. The established motion parameters, aerodynamic characteristics, and the surface heating temperature of the launch vehicle body are key values for further research on the design and analysis of a launch vehicle with a polymer body of variable length. These data could be used to calculate the mechanical and thermal loads acting on the structure of the launch vehicle during flight
ESTIMATION OF THE POSSIBILITY OF USING ELECTRIC PROPULSION SYSTEMS FOR LARGE-SIZED ORBITAL DEBRIS POST-MISSION DISPOSAL A. V. GOLUBEK, , M. M. DRON’, O. M. PETRENKO, , , and Space Science and Technology, 2023 The steady increase in the amount of large-sized orbital debris represents a substantial threat to satellite missions. Currently, many methods of cleaning near-Earth space with the use of various means based on various physical principles are considered. Out of them all, the active method using a rocket propulsion system is the most commonly implemented. Considering the high specific impulse, small size, and mass of electric propulsion systems, they are a particularly attractive choice as means of post-mission disposal. Despite their advantages, such systems have certain peculiarities that need to be considered in the process of designing and implementing modern post-mission disposal means. These peculiarities include the maximum time of a single firing of the electric propulsion system, the maximum time of the battery charging, and the time of operation of the control system. The purpose of this work is the determination the capabilities of the modern Hall thrusters ST-25 and ST-40 developed by Space Electric Thruster Systems in solving the problem of post-mission disposal of large-sized orbital debris from low-Earth orbits taking into account the limitations on the power supply system. To achieve this goal, methods of analysis, synthesis, comparison, and computer simulation were used. In the course of the carried-out research, the following problems were solved. A scheme for post-mission disposal of large-sized orbital debris from low-Earth orbit was developed with consideration of the use of an electric propulsion system. The dependence was determined of the minimum delta-v increment required for post-mission disposal of an object within 25 years on the initial altitude of the orbit and the ballistic coefficient of the orbital debris. The upper boundary of the combinations of masses of orbital debris, the altitude of the initial orbit, and the ballistic coefficient were determined, for which post-mission disposal from near-Earth orbits is possible with the use of electric propulsion systems. The obtained results can be used in solving problems of the development of modern means of active post-mission disposal of orbital debris with the use of Hall thrusters developed by Space Electric Thruster Systems
Possibilities for Expanding the Application Areas of Suborbital Launch Vehicles Proceedings of the International Astronautical Congress Iac, 2023
DETERMINATION OF DESIGN PARAMETERS OF THE SYSTEM TO DE-ORBITING OF THE UPPER STAGE OF ZENIT-2 LAUNCH VEHICLE FROM NEAR-EARTH ORBITS Proceedings of the International Astronautical Congress Iac, 2023
A simulation of the thermal environment of a plastic body of a new type of launch vehicle at the atmospheric phase of the trajectory Andrii Dreus, Vitaly Yemets, Mykola Dron, Mykhailo Yemets, Aleksandr Golubek Aircraft Engineering and Aerospace Technology, 2022 Purpose Leading developers and providers in the modern space launch market note a splash in the development of ultralight launch vehicle (LV), driven by the growing demand for small satellites for large constellations in low Earth orbits. One of the promising ways to solve the problem of the quick launch of such satellites is to use a new type of ultralight launch vehicle with a plastic body. The project of such a launch vehicle was proposed by Oles Honchar Dnipro National University (Ukraine). Along with that, there is a need for appropriate research studies on the thermal resistance of the plastic shell, as the physical, mechanical and thermophysical characteristics of polymers significantly differ from traditional aerospace materials. The purpose of this study is to validate the design and ballistic parameters of such a launch vehicle in terms of providing an acceptable thermal environment at the atmospheric phase of the trajectory. Design/methodology/approach The workability of a new type of propulsion system is being investigated experimentally in bench conditions. To study the process of aerodynamic heating of a plastic shell, numerical modeling based on the integration of the flight dynamics and heat transfer equations is used. Findings Brief information about the design of a new type of ultra-light autophage launch vehicle with a plastic body is presented. A mathematical model for the movement of the launch vehicle at the atmospheric phase of the trajectory, and for the heating of the polyethylene body of the launch vehicle, taking into account the dynamic change in the atmospheric parameters is proposed. The influence of the motion trajectory on the thermal environment of the rocket body is investigated, rational motion trajectories and corresponding permissible g-loads are determined. Originality/value The fundamental possibility of using plastic (polyethylene) as a structural material and fuel for bodies of a new type of ultralight launch vehicles has been substantiated. It is shown that to ensure acceptable thermal conditions of a plastic body, it is necessary to use thermal insulation. It is proposed to use a polymeric Teflon coating as such thermal insulation. The results are important for the development of technologies for launching small satellites into orbit, as the use of plastic as the main structural material of the rocket body will significantly reduce the launch cost.
DETERMINING THE DEGREE OF EFFECT OF HEAT FLOWS ON THE DEFORMATION OF THE SHELL OF A SPACE INFLATABLE PLATFORM WITH A PAYLOAD Erik Lapkhanov, Oleksandr Palii, Aleksandr Golubek Eastern European Journal of Enterprise Technologies, 2022 This paper reports a study into the influence exerted by the thermal flows of space environment on the deformation of the shell of a space inflatable platform with a payload. The mathematical model of the effect of temperature fluctuations on the mass-inertial characteristics of the space inflatable platform of an ellipsoidal shape has been improved. The following assumptions were introduced to the model. The temperature distribution on the illuminated part and the unlit part of the shell is uniform. The gradient of the temperature difference between the illuminated and unlit parts is the same for all points of the shell. To determine deformations, a moment-free theory was used. The model of the space inflatable platform is a «rubber bullet» that works only for stretching and compression. All deformations are elastic. The advantages and limitations of the use of the developed mathematical model have been determined. Computer simulation of the orbital motion of a space inflatable platform with a payload in a sun-synchronous orbit was carried out. The material of the platform shell is Kapton. Estimates of temperature fluctuations in the illuminated and unlit part of the shell and the temperature of the gas inside it were obtained. The dependence of elastic deformations on temperature was determined, taking into account the Young’s modulus of the material. The influence of changes in gas pressure on the movement of payload attachment points and the change in the inertia tensor have been determined. The obtained results showed that the inertia tensor varies within the order of 10–5 kgm2. The maximum deviation of the fastening points of the payload from the initial position on the illuminated part of the shell was about 10–6 m. Considering the stability of the structure to the effects of heat flows of the space environment, the possibility of using space inflatable platforms as a means for separating a grouping of satellites has been shown
Thermal Optimization of Trajectories of Space Debris Removal into the Earth's Atmosphere Proceedings of the International Astronautical Congress Iac, 2022
DETERMINATION OF DESIGN PARAMETERS OF THE SYSTEM FOR COMBINED DEORBITING OF THE UPPER STAGES OF CYCLONE-3 LAUNCH VEHICLE FROM NEAR-EARTH ORBITS Proceedings of the International Astronautical Congress Iac, 2022
Investigation of aerodynamic heating of space debris object at reentry to earth atmosphere Proceedings of the International Astronautical Congress Iac, 2018
