For solving non-traditional problems of rocket flight control, in particular, for the conditions of impact of a nuclear explosion, non-traditional approaches to the organization of the thrust vector control of a rocket engine are required. Various schemes of gas-dynamic thrust vector control systems that counteract impact actions on the rocket were studied. It was found that the dynamic characteristics of traditional gas-dynamic thrust vector control systems do not allow one to solve the problem of counteracting impact actions on the rocket. Appropriate dynamic characteristics can provide a perturbation of the supersonic flow by injecting into the nozzle the detonation products with the main shock wave propagating in the supersonic flow. This way to perturb the supersonic flow in a rocket engine nozzle is investigated in this paper. In order to identify the principles of producing control forces and provide a perturbation of the supersonic flow by injecting into the nozzle the detonation products with the main shock wave propagating in the supersonic flow, a computer simulation of the nozzle flow was performed. The nozzle of the 11D25 engine developed by Yuzhnoye State Design Office and used in the third stage of the Cyclone-3 launch vehicle was taken as a basis. The thrust vector control scheme relies on the use of the main fuel component detonation. The evolution of the detonation wave in the supersonic flow of the combustion chamber nozzle was simulated numerically. According to the nature of the perturbation propagation in the nozzle, the lateral force from the perturbation has an alternating character with the perturbation stabilization in sign and magnitude when approaching the critical nozzle section. The value of the relative lateral force is sufficient for counteracting large disturbing moments of short duration. Thus, the force factors that can be used to control the rocket engine thrust vector are identified. Further research should focus on finding the optimal location of the detonation product injection in order to prevent mutual compensation of force factors.

In the new conditions of application of launch vehicle boosters, space tugs, etc., modern rocket engines often do not satisfy the current stringent requirements. This calls for fundamental research into processes in rocket engines for improving their efficiency. In this regard, for the past 5 years, the Department of Thermogas Dynamics of Power Plants of the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine has conducted research on gas flow control in rocket engines to improve their efficiency and functionality. Mechanisms of flow perturbation in the nozzle of a rocket engine by liquid injection and a solid obstacle were investigated. A mathematical model of supersonic flow perturbation by local liquid injection was refined, and new solutions for increasing the energy release rate of the liquid were developed. A numerical simulation of a gas flow perturbed by a solid obstacle in the nozzle of a rocket engine made it possible to verify the known (mostly experimental) results and to reveal new perturbation features. In particular, a significant increase in the efficiency of flow perturbation by an obstacle in the transonic region was shown up, and some dependences involving the distribution of the perturbed pressure on the nozzle wall, which had been considered universal, were refined. The possibility of increasing the efficiency of use of the generator gas picked downstream of the turbine of a liquid-propellant rocket engine was investigated, and the advantages of a new scheme of gas injection into the supersonic part of the nozzle, which provides both nozzle wall cooling by the generator gas and the production of lateral control forces, were substantiated. A new concept of rocket engine thrust vector control was developed: a combination of a mechanical and a gas-dynamic system. It was shown that such a thrust vector control system allows one to increase the efficiency and reliability of the space rocket stage flight control system. A new liquid-propellant rocket engine scheme was developed to control both the thrust amount and the thrust vector direction in all planes of rocket stage flight stabilization. New approaches to the process organization in auxiliary elements of rocket engines on the basis of detonation propellant combustion were developed to increase the rocket engine performance.

The fairing serves to protect the payload against exposure to external factors during the rocket flight. It must withstand considerable force and thermal loads and safely detach and move away from the rocket. This paper deals with the process of fairing flap separation from the rocket in the Earth's dense atmosphere under conditions of considerable aerodynamic loads in the range of supersonic flight speeds. It is proposed that flap removal from the rocket structure be done using a detonation corded rocket engine, which develops a considerable thrust at a low mass and a short operation period. This significantly reduces the fairing mass. The forces arising in this process were determined by computer simulation. A technique for calculating the basic parameters of a detonation corded rocket engine for fairing flap removal is presented. The mathematical model of flap motion relative to the rocket during the process of detachment and removal consists of two parts: the calculation of the separation and acceleration of the flaps while in mechanical contact with the rocket and the calculation of the inertial motion of the flaps separated from the rocket. The computer simulation gives the projections of the aerodynamic forces and torque and the air pressure distribution for the most characteristic angles. Five protective partition shapes were simulated: conical, concave conical, spherical, concave spherical, and flat. The concave spherical shape was found to be optimal in terms of minimum energy consumption. The optimal shape, dimensions, and placement of the partition were calculated. The minimum thrust of the detonation corded engine required for flap removal from the rocket was determined, and effects that allow one to reduce this thrust were found. The calculated pressure distributions may be used in flap strength analysis.

Стаття присвячена проблемам змішування компонентів палива в камерах згорання детонаційних ракетних двигунів. Основною ідеєю, що спонукає вчених до пошуків у цьому напрямку, є вищий термодинамічний коефіцієнт корисної дії детонації в порівнянні зі звичайним горінням з дозвуковими швидкостями. Також детонаційний процес може відбуватися при відносно низьких значеннях тисків компонентів палива, що дозволяє відмовитись від важких систем живлення, а використати просту витискувальну систему подачі. Відомі експериментальні дослідження використовуються для подальших наукових пошуків шляхів вирішення проблем із сумішоутворенням.Для оцінки ефективності процесу змішування використовується комп’ютерне моделювання. Визначено масштаб турбулентності в різних схемах форсуночних головок. Проведено класифікацію схем в порядку збільшення масштабу турбулентності і, відповідно зниження ефективності двигуна. Запропоновано перехід до використання форкамер з попереднім перемішуванням компонентів палива в одному об’ємі і детонацією їх суміші в іншому.

To solve the problem of satellite control and stabilization in emergencies, it is proposed to use a detonation rocket engine, which enables active maneuvering to avoid a collision with space debris. The goal of this work is to study a new way of rocket engine thrust vector control by acting with a detonation shock wave on the gas flow in the nozzle. A detonation wave in a supersonic flow in a nozzle was numerically simulated. The simulation was conducted in a non-stationary plane formulation at different angles of inclination of the detonation gas generator that initiates a detonation shock wave to the combustion chamber axis with the use of SolidWorks application software for the 11D25 engine of the Cyclone-3 third stage. The simulation results were used to pre-optimize the location of the detonation gas generator on the nozzle wall. It was found that the effect of the detonation wave on the main gas flow in the nozzle is caused by two force factors: the first is due to the reactive force produced by the detonation product injection into the nozzle and a high-pressure zone on the wall where the detonation gas generator is mounted, and the second is due to a change in pressure distribution over the nozzle surface. In order to increase the effect of the shock wave, the detonation products must be injected parallel to the main gas flow in the nozzle or at some angle. The simulation showed the drawbacks and advantages of detonation product injection at different angles. The detonation wave effect on a supersonic nozzle flow was studied experimentally. A system was developed to record the shock detonation wave propagation using a heat meter. A special nozzle model and a gas generator were developed to initiate a detonation wave interacting with a supersonic air flow. It was found out how the detonation wave separates the main flow from the nozzle walls in the overexpanded mode. The results may be used in the space-rocket industry to provide upper stage maneuvering to avoid a collision with space debris.