Vol 15, No 4 (2016)

One of the promising trends in the improvement of stand test data management and control systems and emergency protection systems involves the development of intellectual peripheral devices that provide communication with test articles. Due to the incorporation of their own processors in such devices they can solve problems of receiving data from sensors and initiating lower level control actions (independently from the master controller) on the actuators of controlled objects, which enhances the reliability and quality of control. The paper presents the latest developments directly related to the control of liquid-propellant rocket engine (LRE) parameters during stand tests, i.e. control of electro-hydraulic tilt actuators of LRE combustion chambers; control of stepping motors for the regulation of mixture ratio  and control of vernier engines; receiving, processing and recording of signals from turbine flow meters and turbopump RPM sensors.

The paper presents a calculation procedure based on the mathematical model of a pilot operated gas pressure regulator. The model takes into account the parameters of the feedback channel, the gas damper, the flow force at the main valve poppet and the dynamics of the fuel tank pressurization system of a launch vehicle. We analyzed the influence of the gas damper incorporated in the pressure relief valve on the pressure control system performance. Stability domains in the space of the regulator parameters are calculated. We assessed the control system stability margin for varying gas damper parameters. The proposed procedure of regulator design ensures the required stability margin of the gas pressure control system. The static accuracy of the system remains unchanged.

The article discusses ways of utilizing exhaust gas heat to improve the power efficiency of heat engines. The concept of an air turbine engine is chosen to make a generator drive with the capacity of 150 kW intended to produce energy for a gas-pumping unit.  In an air turbine engine atmospheric air heated by the exhaust gases of the gas turbine drive of a gas pumping unit serves as the working fluid. Three thermodynamic cycle schematics for an air turbine engine (ATE) are analyzed. The optimal design of an ATE comprising a two-stage axial compressor, a recuperative air heater (RAH), an ejector and a single-stage axial turbine is determined on the basis of thermodynamic calculations. Compressed air from the ATE compressor unit is used in the ejector as the active flow, while the combustion products of the NK-16ST engine are used as the passive flow. The recuperative air heater and ejector are located in the exhaust line of the gas pumping unit. The chosen design concept of the air turbine engine corresponds to the technical requirements of gas transmission enterprises.

The first airplane working on cryogenic fuel was developed in the USSR in early 1980s on the basis of the passenger aircraft Tu-154. It was then modified using a hydrogen engine installed instead of one of the three kerosene engines. That bypass turbo-jet engine was developed in Russia (Samara) under the guidance of General Designer academician N.D. Kuznetsov. The engine working on liquid hydrogen was designated NK-88 and the one working on liquid methane – NK-89. A unique hydrogen turbopump unit was created for the supply of cryogenic fuel. The unit can also be used for liquid methane due to the possibility of readjusting the rotor frequency. To minimize the financial expenditures it is proposed to use the turbo-pump in low-thrust rocket engines that use liquid methane as fuel.

The article describes a thermo-gas-dynamic model of small gas turbine engines. The model takes into account the influence of the engine size on the efficiency of work processes in the crucial components. Gas turbine engines are classified according to their size depending on the value of the gas generator mass flow rate corrected by the compressor exit parameters. An important feature of the working process in small gas turbine engines is that hydraulic losses in the flow section increase with the decrease in the engine’s size due to the increase of the boundary layer relative thickness. The efficiency of the compressor and turbine also decrease because of the increase in relative radial clearances. These factors are taken into account in computer modeling by making allowances for the initial values of compressor efficiency, fuel combustion efficiency, the total pressure loss coefficient and turbine efficiency. The suggested approaches were used to improve computer models of gas turbine engines. It is shown that reducing the engine size results in considerable decrease of the work process optimal parameters and specific parameters. Taking into account the influence of the engine size on the efficiency of its components widens the range of its applicability and improves the adequacy. Thus, the models provide a more adequate solution for the optimization of working process parameters and can be used for conceptual designing of small gas turbine engines.

The article presents the results of theoretical and experimental studies of the peculiarities of the working process of low-thrust hypergolic- propellant liquid rocket engines (LTLRE) with the thrust of less than 1 N. We also discuss ways of improving their efficiency, reliability and stability of parameters. To study the stability of the hydraulic characteristics of LTLRE capillary nozzle elements we experimentally investigated the hydrodynamic characteristics of the capillaries under conditions of isothermal flow of water and heat supply and proposed a method of their calculation. Using the proposed calculation method we studied the changes of the capillary hydraulic resistance for the engine injector head and showed that the effect of thermal factors can lead to significant changes in the hydraulic characteristics of nozzle elements of engines with the thrust of less than 1 N and, consequently, to the instability of its parameters and off-design engine operation. A theoretical and simulation study was carried out to determine the region of total liquid-phase mixing depending on the capillary diameter and the thrust of the engine.  It is shown that low efficiency of liquid-phase mixing of the components is the main reason for the low efficiency of LTLRE with the thrust of less than 1 N.  We investigated experimentally the variation of energy parameters of a jet-mixing engine with the thrust of less than 1 N depending on the thrust level. We propose to use a precombustion chamber as a possible way of intensifying the intrachamber workflow. The influence of prechamber on the energy parameters of a LTLRE with the thrust of less than 1 N is analyzed.  The possibility of intensifying the intrachamber workflow with the use of a prechamber is shown. Recommendations on the choice of the geometry of the prechamber are given.

The paper discusses improving the weight and size characteristics of propulsion systems with liquid thrusters due to the replacement of conventional compressed-gas systems by hot gas systems where a gas generator operating on propellants is used for hot gas production. Possible algorithms of the operation of such feed systems are discussed. A mathematical model has been created for the purpose of computational and theoretical study of operating processes in the considered feed system. The model represents system units as lumped volumes. Differential equations of weight, internal energy, and combustion gases concentration are written for every lumped volume. The differential equation of unit structure temperature is also added. The differential equations of mass and internal energy are converted into equations of pressure and temperature. The model is expressed through mass flow ratios and heat exchanges between the system units, dependences of thermo-physical properties on the temperature, combustion gas concentrations and the ratio of fuel components in the gas generator. The temperature, pressure, concentration of combustion gases in the system units and the unit structure temperature are specified as input parameters. The simulation results showed satisfactory agreement with the experimental data, which makes it possible to use the developed model for computational and theoretical studies of the hot gas pressure feed systems.

The paper justifies the possibility of using cold-flow test data to predict the combustion efficiency for a 13.34 N thruster combustion chamber. The results of modeling working liquid (water) flowing through hydraulic passages of single injection elements and coaxial swirl injectors are presented. We performed the simulation using the Reynolds-averaged Navier-Stokes equations that describe a turbulent flow of a two-phase non-compressible liquid. We used both one-velocity and two-velocity flow models. We modeled the turbulence by the BSL isotopic turbulence model on a computational block-structure grid. The computational grid consisted mostly of HEXA elements. The orientation of the elements coincided with the flow direction in the injector hydraulic channels, which improved the flow resolution accuracy in the boundary layer near the solid wall. We compared the results of calculation with the cold-flow test data. The use of the one-velocity flow model in the design area of a single injector and the two-velocity model in the design area of the mixing element yielded the results that agreed satisfactorily with the cold-flow data.

The aim of the present work was to develop a methodology for assessing reliability on the basis of physical characteristics of an element. To this goal, we used the structure-energy power failure theory. The use of the theory makes it possible to determine the probability of failure using the information on the internal damage of the material. The gas turbine engine blades of the first, second and third turbine stages, new and reconstructed high-pressure compressor blades of stage 5 were the object of the analysis.  We used an industrial computer tomograph XTH 450 LC to carry out non-destructive testing. We carried out destructive tests on stands of the electrodynamic type. The probabilities of failure calculated by using the structure-energy failure theory correspond to the values of reliability indices obtained by the destructive experiments. We developed a computer program based on a statistical method for estimating the reliability of an aircraft engine. The program takes into account the reliability index of rotor blades calculated with the use of the structure-energy failure theory.

Creation of a low-noise water flow control systems apparatus is the main objective of the research. The technique of carrying out the tests is based on the principles of acoustics, flow mechanics and the control theory. The main schemes of regulating the water flow rate are presented in the paper. It is shown that the scheme with the use of a flowmeter is the most promising in terms of improvement of vibronoise characteristics. A modular principle of constructing feed water rate control systems is presented. The design features of application of multistage rotary units of the apparatus meeting the hydrodynamic noise requirements for low rates without application of dissipative elements (e.g. mufflers) are formulated for the first time.  Energy comparison of various control system schemes is carried out. It is shown that further improvement of operating devices in terms of vibronoise characteristics is possible with the use of the principle of  serial-parallel water flow regulation. The results help to improve the mass-dimensional characteristics of systems and to increase their energy efficiency, to provide a low level of noise and to reduce the cost of production. The modularity of the apparatus ensures a high level of standardization and interchangeability.

Ultralow power turbine drives are used in aerospace and other transport equipment as energy for various units, mostly supplementary ones. The main components of the drives are an input device, a very low power turbine and an output device. Improving the efficiency of low-power turbine drives is a vital task. Energy indicators, such as power efficiency and specific working medium consumption of the working fluid are some of the most important performance indicators. The article presents  configurations of axial and centripetal turbine ultra-low power, describes their basic operating and geometrical parameters that influence the energy efficiency of the drive. We present mathematical models of the efficiency of the turbines of two types, obtained from the results of gas-dynamic computational experiments. The results of statistical correlation and regression analysis of computational experiments presented in the article show the adequacy of the mathematical models. We analyzed the degree of influence of the model factors and their interactions on the changes in the efficiency of using the influence coefficients and graphical analysis of the results of the optimization parameters of the turbine.

The paper presents the results that allowed obtaining the dependence of laminar flame propagation speed Sl on the equivalence ratio for a wide range of pressures and temperatures during methane combustion. A literature review was carried out to summarize the experimental data on the measurement of the Sl. The Sl was calculated using a kinetic mechanism GRI 3.0 within the required pressure and temperature range. The calculation results were generalized in the MATLAB software product to verify the Sl power dependencies on pressure and initial temperature. The results of calculation on the basis of the obtained approximating dependence were compared with the experimental data and results obtained by other authors. It was found that the exponents of power for the dependency on pressure and temperature are described not by constants or linear relations, but by second-degree equations on the fuel-air ratio. The results can be used in three-dimensional simulation of combustion processes and in calculations performed using engineering practices.

The article deals with the task of increasing the efficiency of rotary piston engines by implementing an economical diesel cycle in a module with a three-arc epitrochoid and an external gearing kinematics scheme. The necessity of the diesel cycle is caused by the need of reducing specific fuel consumption in rotary piston engines that are currently widely used in light aviation, mostly for remotely piloted aircraft. The Wankel rotary piston engine with a two-arc epitrochoid, though having a number of advantages over traditional piston engines, is inferior to them in efficiency, particularly in partial load modes. The implementation of the diesel cycle in а cycloidal piston engine with a three-arc epitrochoid makes it possible to reduce specific fuel consumption by 30% while retaining the listed positive qualities, as well as to use diesel fuel which is preferable for high-altitude aircraft due to lesser tendency to gas formation in pipelines. The engine stator inner loop is made up by three arcs of the epitrochoid that form three working chambers equally-spaced along the perimeter, which provides a decrease of the stator thermal deformation during the operation of the engine. A method is developed to determine the geometric parameters of a rotary piston engine with a three-arc epitrochoid, and design work is carried out in the КОМPAS-3D code for a 50 h. p. diesel engine to be used in light aviation and as a drive of electrical generators, pumps and compressors for ground application.