Annotation: The article considers the present-day aspects of predicting and establishing the guaranteed service life of liquid rocket engines. The phases and sequence of works to specify the guarantee periods are presented. The physical ground for determining thermal ageing test modes and increased humidity test modes is set forth. The mathematical dependencies are given to determine the time of testing at increased temperature, increased relative humidity, equivalent article storage temperature. The generalized list and sequence of tests to establish a guarantee period are set forth. The content of works is considered to seek for and introduce new components of inter-industry application, in particular, replacing materials for rubber mixtures during production of rubber technical goods, polymer materials (fluoroplastics, polyamides), isotropic pyrographite, ozone friendly means for degreasing liquid rocket engine assemblies being in contact with propellant component – oxygen and to seek for alternative bearings.

Annotation: In the pneumohydraulic systems of liquid rocket engines and propulsion systems, electromagnetic valves that allow making the pneumohydraulic systems more simple and ensuring multiple ignition of liquid rocket engines have found wide application. The Yuzhnoye-developed electromagnetic valves are designed according to two schemes – of direct and indirect action. In the direct-action electromagnetic valves, the shutting-off device opens (closes) the throat with the force developed by electric magnet. They have gained acceptance in the pneumohydraulic systems with the working medium pressure of ~8.5 MPa, they are of simple design and have high operating speed (0.001…0.05 s). In the electromagnetic valves with amplification, the electromagnet armature is connected with control valve and the main shutting-off device moves due to the force from working medium pressure drop on it. They are used in the operating pressure range of 0.5…56 MPa, at that, the action time is 0.025…0.15 s. For the European Vega launch vehicle fourth stage main engine assembly that has pressure propellant feeding system, the electrohydraulic valve with amplification and drainage was developed. The dependence of this electrohydraulic valve high speed from the line’s output length is decreased to the maximum due to the installation of Venturi nozzle at the output connecting branch. This electrohydraulic valve is operable at the pressure below 8 MPa, the action time is 0.08…0.12 s. The present-day spacecraft gas-jet orientation and stabilization systems use as propulsion devices the electromagnetic valves with nozzles whose thrust is, as a rule, not more than 30 N and the working medium pressure is up to 24 MPa. Yuzhnoye State Design Office developed for 15B36 gas-jet system the electropneumatic valve with amplification and nozzle, which is operable at the pressure below 45 MPa, ensures the action frequency of up to 10 Hz and is capable of creating the thrust of 100 N on gaseous argon. To solve the task of decreasing the dependence of operability and high speed of electromagnetic valves with drainage and amplification on geometry of lines in which a valve is installed, the electropneumatic valve was developed that has spool elements ensuring reliable and quick action with long input lines of 0.004 m diameter. Its mass is 2…2.5 times lower than the mass of analogs. Recently, Yuzhnoye State Design Office develops the apogee RD840 LRE with 400 N thrust, for the conditions of which the direct-action electrohydraulic valve was developed and tested with the following characteristics: pressure – up to 2.15 MPa, consumed power in operation mode – less than 7.1 W, action time – not more than 0.02 s, mass – 0.19 kg. The presented electromagnetic valves by their technical and operational characteristics meet the highest world requirements and have found wide utility in liquid rocket engines and propulsion systems.

Annotation: The paper shows the peculiarities of modeling the vacuum conditions to simulate space environment when conducting various test types. Based on generalized experience, the practical criteria are recommended with consideration for accompanying physical phenomina.

Annotation: The calculation procedure is proposed, the analysis is made and the ranges of optimal probability values of no-failure operation of pneumohydraulic propellant supply system and its elements are determined based on general requirements to integrated launch vehicle.

Annotation: The paper presents the results of comparative investigation into characteristics of methane, kerosene and hydrogen in pair with oxygen. The peculiarities of each of these components are shown. The optimal parameters are considered of pneumohydraulic supply system, including the designs of tanks, pressurization system and engine supply lines cooling system.

Annotation: The paper presents a complex of analytical calculation measures that allow increasing tanks useful volume and ensure engines’ additional operational lifetime by the example of a launch vehicle, one of Yuzhnoye SDO developments. The calculated-experimental confirmation is set forth of the operability of pneumohydraulic supply system in the changed conditions that ensure considerable increase of launch vehicle power and mass characteristics.

Annotation: The main directions of upgrading the pneumohydraulic systems are considered. Some methods of increasing their operability and reliability are analyzed.

Annotation: Long-term experience in development, development testing and use of generating systems of fuel tanks pressurization for rockets powered by nitrogen tetroxide and unsymmetrical dimethylhydrazine is summarized. Replacement of gas bottle pressurization systems with generating ones on such launch vehicles as 15A14, 15A15, 11K68 (8K67), 15A18M substantially simplified operation, reduced the pneumohydraulic feed system mass at least twice and its cost – by five times. Typical stages of development and introduction of the pressurization generating systems are shown: development of generators, their development testing, study of the composition and parameters of gas. The important steps were the development of methodology for pressurization system parameters calculation, which enabled achievement of the substantial improvements of their characteristics, appearance of the high-performance hightemperature (up to ~ 1000o C) unsymmetrical dimethylhydrazine tank pressurization system, study of the degree of impact of each of the pressurization system parameters on the tank pressure. Accounting of the correlation between the flow rate and the generator gas temperature improved the output performance, as well as simplified and reduced the amount of development testing of the pressurization system. Important role of the gas sprayer design in pressurization system parametric configuration is described, and the advanced versions are shown taking into account g-loads, changes in temperature, pressure and propellant level inside the tank. Significant phase in the development of the generating pressurization system was the effective use of the high-temperature pressurization of the fuel tank with submerged propulsion system. Besides for the first time the effect of mechanical temperature destratification of the propellant in the tanks was observed, which occurs during the propulsion systems shutdown. Due to this effect, the Dnepr LV payload capability enhanced. Successful engineering solutions in the design of the pressurization system were defended by ~80 copyright certificates and patents of invention, ~40 of which were successfully implemented.

Annotation: One of the main design parameters of the automatic equipment in the launch vehicle’s pneumohydraulic systems is the flow friction characteristic, which represents the proportionality factor between the automatic equipment pressure differential and velocity head. The flow friction characteristic of the completely open automatic device should have very small value with required dimensions and mass. With decrease of the pressure losses, the required upstream pressure of the propulsion system is ensured with smaller pressurization of the tanks. It results in the decrease of the required pressurization gas volume, which boosts reduction of the performance of the launch vehicle as a whole. This paper describes the method of reduction of the flow friction characteristic of the dividing valve, optimizing the geometric shape of the flow passage. The problem of minimization of the valve’s flow friction characteristic is considered with the specified mass and design dimensions restrictions. The initial design of the valve was developed, taking into account the specified requirements, literature references and parameters of the analogue units. With the goal of optimization various options of valve design were considered, different from the initial design in configuration of the inlet and discharge nozzles, notably various angle sizes, forming the stream profile, and lengths of the direct-flow sections. Four options of the valve design were calculated using numerical methods of ANSYS CFX software. Navier – Stokes equations and k-ω SST turbulence model were used. Based on the calculations results the optimal design was selected. Initial design of the valve was compared with the optimal one. The flow friction characteristic of the optimal valve design decreased by 26 % in comparison with initial design with insignificant change of mass and dimensions. The design of the developed dividing valve can be involved in the design of the new launch vehicles.

Annotation: The general information on gas pressure reducers, on their purpose in launch vehicles and spacecraft pneumohydraulic systems is set forth. The impact of different operating conditions on physical characteristics of these devices is considered. The main and auxiliary parametric characteristics of the reducer are presented and the physical process of gas pressure reduction in it is explained. The error of output pressure regulation is evaluated using full differential of function, whose arguments (input pressure, flow rate, temperature) have scatter. The reducer temperature curve is shown and the impact of structural temperature on the value of dynamic (with flow rate) and static (without flow rate) pressure in reducer output cavity is explained. The difference between the excess pressure reducer and absolute pressure reducer is shown. The brief review of the designs of liquid and bimetal thermal compensators is presented, their advantages and disadvantages are described and the experience of reducers testing with regulating springs made of elinvar is analyzed. Attention is focused on operating temperature and its impact on stability of reducer adjustment. The formulas that describe thermodynamic processes occurring in the reducer are presented. Special attention is given to the properties of regulating spring of the reducer because of change of elasticity modulus coefficient at different temperatures, the expected pressure scatter at reducer output is evaluated and the necessity of measures to reduce this error is explained. To compensate for temperature disturbance, the formula of gas pressure in closed volume of sensitive element is derived. The essence of original technique of pneumocorrection of initial pressure in sensitive element cavity that was proposed and introduced on Yuzhnoye SDO-developed reducers is set forth.