Annotation: Being a final executive element of rocket control systems, a hydraulic actuator is at the same time the main source of various non-linear dependencies in rocket dynamic design whose availability dramatically com plicates theoretical analysis of their dynamics and control systems synthesis. The required accuracy and complexity of mathematical models of hydraulic servo mechanisms are different for different design phases of guided rockets. The paper deals with the simplest models of hydraulic servo actuators intended to calculate rocket controllability and to define requirements to response and power characteristics of the actuators. To calculate the rocket stability regions and to evaluate own stability of servo actuators, a linearized mathematical model of hydraulic servo actuator is used that takes into account the most important parameters having impact on stability of the servo actuator itself and on that of the rocket: hardness of working fluid, stiffness of elastic suspension of the actuator and control element, slope of mechanical characteristic of the actuator in the area of small control signals, which, as full mathematical model analysis showed, is conditioned only by dimensions of initial axial clearances of slide’s throats. The full mathematical model constructed based on accurate calculations of the balance of fluid flow rate through the slide’s throats allows, as early as at designing phase, determining the values of most important static and dynamic characteristics of a future hydraulic actuator, selecting optimal characteristics of slides based on specified degree of stability and response of servo actuator and conducting final modeling of rocket flight on the integrated control system test benches without using real actuators and loading stands. It is correct and universal for all phases of rockets and their control systems designing and testing. Using this mathematical model, the powerful actuators of a line of intercontinental ballistic missiles with swinging reentry vehicle and the main engines actuators of Zenit launch vehicle first stage were developed. The results of their testing separately and in rockets practically fully comply with the data of theoretical calculations.

Annotation: The main parameters are understood as: carrying capacity mг, range L and fuel efficiency qт, which largely determine the competitiveness of aircraft of this type, including military transport aircraft. The reason for the creation of modifications of transport category aircraft is the requirement for a constant increase in their flight performance by increasing the carrying capacity and range. Among the main goals of implementing such decisions there is an indispensable increase in the fuel efficiency of modifications, since the cost of fuel reaches 80 % of the cost of an airplane hour during operation. There are a number of models that make it possible to assess the influence of the weight and aerodynamic parameters of the airframe of the aircraft and the fuel performance of the power plant (specific engine consumption) on the integral indicator of the fuel efficiency of the modification at cruising mode and the average hourly fuel consumption at the certification stage, when all parameters of the airframe and engine are fixed and consideration of options is not possible. A new model is proposed for the stage of designing modifications, in which deep modification changes are made in the geometry of wing and in the power plant with various variants of their correlation and coordination. The parameters of new model: specific fuel efficiency – specific route productivity, in order to form the relative carrying capacity and relative range of action for the required specific fuel efficiency. An analysis of such dependencies showed: – with an increase in relative range L , fuel costs per flight also increase; – the adequacy of changes in route performance is observed only at L < 0.5. At L > 0.5 productivity is constantly decreasing, while the specific indicator of fuel consumption per unit of work increases exponentially; – if in the analysis we take into account the specific value of transport efficiency, that is, the characteristic “load – range” ( mп.н  f L ), it becomes obvious that the most favorable (from the point of view of fuel efficiency) are relative ranges of 0,3 < L < 0,5. In this range L , not only acceptable fuel efficiency values are realized, but also the maximum value of the route performance, that is the main parameters for which modifications are developed.

Annotation: The shell structures widely used in space rocket hardware feature, along with decided advantage in the form of optimal combination of mass and strength, inhomogeneities of different nature: structural (different thicknesses, availability of reinforcements, cuts-holes et al.) and technological (presence of defects arising in manufacturing process or during storage, transportation and unforseen thermomechanical effects). The above factors are concentrators of stress and strain state and can lead to early destruction of structural elements. Their different parts are deformed according to their program and are characterized by different levels of stress and strain state. Taking into consideration plasticity and creeping of material, to determine stress and strain state, the approach is effective where the calculation is divided into phases; in each phase the parameters are entered that characterize the deformations of plasticity and creeping: additional loads in the equations of equilibrium or in boundary conditions, additional deformations or variable parameters of elasticity (elasticity modulus and Poisson ratio). Then the schemes of successive approximations are constructed: in each phase, the problem of elasticity theory is solved with entering of the above parameters. The problems of determining the lifetime of space launch vehicles and launching facilities should be noted separately, as it is connected with damages that arise at alternating-sign thermomechanical loads of high intensity. The main approach in lifetime determination is one that is based on the theory of low-cycle and high-cycle fatigue. Plasticity and creeping of material are the fundamental factors in lifetime substantiation. The article deals with various aspects of solving the problem of strength and stability of space rocket objects with consideration for the impact of plasticity and creeping deformations.

Annotation: Basic dynamic properties of the reentry part of the aircraft-type first stage were examined when turbojet engine is used in the recovery phase. Such configuration can be of interest because turbojets have considerably smaller rate of flow in comparison to rocket engines. Moreover, they are launched in the lower stratosphere or in the troposphere so that there is no need to place oxidizer supply on board. This recovery plan differs from an alternative rocket recovery system and, from our point of view, provides more efficient usage of the fuel stores because it doesn’t require the main propulsion to be started in the recovery phase. Besides the analysis of qualitative characteristics of the descend phase for this stage, the efficiency of a wing with moderate values of maximum aerodynamic characteristics and a turbojet was studied. In this case three ways for stage recovery were investigated. The first one implied unguided descend with zero angle of attack assuming that the stage is statically stable. This descend trajectory was considered as standard and was used to evaluate the efficiency of the wing and turbojet with relatively small propulsion. The second and the third design cases offered the gliding guided descend with turbojet being started only in the lower stratosphere. The last two cases used the same program for the angle of attack. The possibility to ensure permissible overload values at the critical points of the descend trajectory and acceptable values of kinematic characteristics at the earth surface tangency point are also of great interest. Thereby the program for the angle of attack was developed in a way that allowed kinematic characteristics on touchdown be as close as possible to the corresponding values, shown by civil and/or military-transport heavy aircraft. Simulation was conducted on Microsoft Visual Studio 2010.

Annotation: The main scientific and methodological propositions for designing single-stage guided missiles with main solid rocket motors that are intended for delivering payload to the given spatial point with required and specified kinematic motion parameters are defined. The aim of the article is to develop methodology for the early design phase to improve the basic characteristics of guided missiles, including formalization of complex problem to optimize design parameters, trajectory parameters and motion control programs for guided missiles capable of flying along the ballistic, aeroballistic or combined trajectories. The task is defined as a problem of the optimal control theory with limitations in form of equality, inequality and differential constraints. An approach to program forming is proposed for motion control in the form of polynomial that brings the problem of the optimal control theory to a simpler problem of nonlinear mathematical programming. When trajectory parameters were calculated the missile was regarded as material point of variable mass and the combined equations for center-of-mass motion of the guided missile with projections on axes of the terrestrial reference system were used. The structure of the mathematical model was given along with the calculation sequence of criterion functional that was used for optimization of design parameters, control programs and basic characteristics of the guided missile. The mathematical model of the guided missile provides adequate accuracy for design study to determine: overall dimensions and mass characteristics of the guided missile in general and its structural components and subsystems; power, thrust and consumption characteristics of the main engine; aerodynamic and ballistic characteristics of the guided missile. The developed methodology was tested by solving design problems. Applications of the developed program were studied to present the research results in a user-friendly form.

Annotation: At Yuzhnoye State Design Office, the Cyclone-4 launch vehicle 3rd stage engine has been developed and is under testing. For adjustment of the engine and test bench systems, in the first firing tests the radiation-cooled nozzle extension was replaced with a steel water-cooled one. It was planned to start the engine with water-cooled nozzle extension without vacuumizing and without gad dynamic pipe, which conditioned operation with flow separation at the output edge of water-cooled nozzle extension. Therefore, the calculation of flow in the nozzle with water-cooled extension, flow separation place, and thermal load on watercooled nozzle extension during operation in ground conditions is an important task. Selection of turbulent flow model has a noticeable impact on prediction of flow characteristics. The gas dynamic analysis of the nozzle with water-cooled extension showed the importance of using the turbulent flow model k-ω SST for the flows with internal separation of boundary layer and with flow separation at nozzle section. The use the flow model k-ω SST for calculation of nozzle with flow separation or with internal transitional layer allows adequately describing the flow pattern, though, as the comparison with experimental data showed, this model predicts later flow separation from the wall than that obtained in the tests. The calculation allows obtaining a temperature profile of the wall and providing the recommendations for selection of pressure measurement place in the nozzle extension for the purpose of reducing sensors indication error. With consideration for the special nature of the nozzle extension wall temperature field, the cooling mode was selected. The tests of RD861K engine nozzle with water-cooled extension allow speaking about its successful use as a required element for testing engine start and operation in ground conditions without additional test bench equipment.

Annotation: During experimental investigation of the dynamic characteristics of a pneumatic test bench for testing liquid rocket engine high-flowrate automatic units, the effect was detected of 20-35% sound speed increase in the gas flow moving along the channel with corrugated wall (metal hose) which is a part of test bench drain system. The article presents the results of experiments and the task of theoretical justification of the effect is solved. It is indicated that its causes may be two mutually complementary factors – a decrease of gas compressibility at eddy motion and oscillations of metal hose wall. The physical model is considered that describes variation of gas elasticity and density in the conditions of high flow vorticity. It is supposed that in the near-wall layer of the channel, toroidal vortexes (vortex rings) are formed, which move into turbulent core of the flow where their size decreases and the velocity of rotation around the ring axis of torus increases. The spiral shape of the corrugation ensures also axial rotation, which increases vortexes stability. The intensive rotation around the ring axis creates considerable centrifugal forces; as a result, the dependence of pressure on gas density and the sound speed increase. The mathematical model has been developed that describes coupled longitudinal-lateral oscillations of gas and channel’s corrugated shell. It is indicated that in the investigated system, two mutually influencing wave types are present – longitudinal, which mainly transfer gas pressure pulses along the channel and lateral ones, which transfer the shell radial deformation pulses. As a result of modeling, it has been ascertained that because of the lateral oscillations of the wall, the propagation rate of gas pressure longitudinal waves (having the same wave length as in the experiments at test bench) turns out to be higher than adiabatic sound speed.

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 Ukrainian companies Yuzhnoye SDO and SE PA YMZ supply VG143 main engine assembly for Vega LV AVUM upper stage, which is a one-chamber LRE of 250 kg thrust with five ignitions in flight. By the present, 11 successful launches of Vega LV have been made. In the process of flight operation, there were no critical comments on engines operation. This LRE has a combination of attractive characteristics, such as high specific pulse, low mass, multiple ignitions in flight, high reliability confirmed by good results of flight test of the prototype engines. The reserve of this engine from the viewpoint of further modernization is far from being exhausted. Enhancing the capabilities of payload injection by launch vehicles into various orbits of artificial Earth satellites is the main task for the developers of ILV as a whole and for the developers of separate assemblies and systems, such as LRE being part of ILV. With consideration for the experience of prototype engines testing, we should note the following ways of main engine assembly modernization: – increasing the specific pulse due to the increase of nozzle expansion ratio; – decreasing the volume of internal manifolds and mass of chamber; – increasing the operation time; – increasing the ignitions number; – increasing the duration of pauses between ignitions and orbital functioning time. Increasing the thrust and specific pulse of Vega LV VG143 main engine assembly and AVUM stage takes place due to the use of pneumopump propellant feeding system instead of standard pressure feeding. Besides, the information is presented on RD859, RD864, RD866 and RD869 prototype engines, the data on their basic characteristics, testing and operation. The below information is of interest to LRE and LV developers.

Annotation: One of the important directions in the development of Yuzhnoye SDO liquid rocket engines is creation of the engines for launch vehicle upper stages, boosters, space tugs and takeoff-and-landing vehicles. The article presents an overview of Yuzhnoye SDO – developed liquid rocket engines, their basic characteristics, distinctive features and the current status of development and operation. The article presents the information on the following engines: RD858 and RD859 operating on storable propellants, for lunar takeoff-and-landing Block E module; RD809M and RD809K operating on liquid oxygen + kerosene created on the basis of RD8 serial control engine of Zenit launch vehicle second stage. In this case, RD809M engine is RD8 version with tight integration and RD809K engine is its one-chamber version; RD805 engine operating on liquid oxygen + kerosene created on the basis of combustion chamber of RD8 serial control engine of Zenit launch vehicle second stage.; RD835 engine operating on liquid oxygen + kerosene created for the second stages of launch vehicles of Mayak type; the engines and propulsion systems operating on storable propellants, such as RD861K (main engine of Cyclone-4 third stage and Cyclone-4M launch vehicle second stage), DU802 (liquid propulsion system of Krechet autonomous space tug of conversion Dnepr launch vehicle), RD840 (apogee liquid rocket engine of liquid propulsion system of geostationary communication satellite bus), VG143 (main engine assembly of the fourth stage of European Vega launch vehicle), RD864 and RD869 (main engines of Dnepr launch vehicle upper stages). The information presented in the article is of interest to liquid rocket engines and launch vehicles developers.