Vol 18, No 3 (2019)

An analytical solution to the problem of determining the parameters of motion of an orbiter’s center of mass is presented on the basis on the results of measurements carried out using the optical system of a space robot located in orbit coplanar to that of the orbiter. The “space robot-orbiter” line-of-sight angle and the line-of-sight rate in the moving orbital system of coordinates of the space robot are selected as initial parameters that are measured by the space robot’s on-board optical system. Along with the known orbital parameters of the space robot they are used to solve the problem of determining the parameters of motion of the orbiter’s center of mass. When solving this task, assumptions are introduced concerning the central gravitational field of the Earth, the coplanarity of the orbits of the space robot and of the orbiter, absence of influence of the atmosphere, the moon's attraction and the pressure of the solar wind on the motion of the space robot and of the orbiter, absence of errors in the results of measurements performed by the space robot’s on-board optics. Analytical expressions are obtained to determine the unknown parameters of motion of the orbiter’s center of mass. The results presented can be used to develop methods allowing standalone determination of parameters of the orbit of unknown orbiters using on-board optics of a space robot.

This paper presents the results of theoretical and experimental research of a prototype of a propulsion system for periodical low-Earth orbit correction of research-and-educational nanosatellites. For that purpose, the prototype is to provide at least 20 m/s relative velocity for a 3U CubeSat with a mass not exceeding 4.5 kg. The personnel and environment safety were taken into account during testing and operation along with the ability to be launched as an associated payload by a “Soyuz”-series launch vehicle or from the ISS. An electro-thermal propulsion system (ResistoJet) was designed with “nonfreezing” mixture of ethanol and distilled water used as the working fluid. It is shown that a standard vehicle power system is capable of initiating one corrective thrust impulse per orbit with flight velocity change of about 0.1 m/s by introducing pulse energy storage units and allocating sufficient time for their charging in the flight profile. The propulsion system prototype was tested in atmospheric conditions. For that purpose shortened “atmospheric” nozzles were used. Testing was carried out using a zero-torque test-bench with high-speed cameras. The measured thrust value was in agreement with the calculated value of 44 mN.

The paper presents a computational method with the use of gas-dynamic functions of parameters of detonation in a one-dimensional subsonic flow of ideal gas behind the shock wave propagating in chemically active air-and-fuel mixture in a uniform-cross-section duct, where the resultant of normal pressure forces acting on its side surface is equal to zero. Stabilization of the shock wave is provided by the onset of thermal crisis with the air-and-fuel mixture combustion heat supply to the gas behind the wave. In this case the value of the specific speed of the combustion products is equal to the critical one. The solution of the total-impulse equation considering the above mentioned peculiarities of the flow in a uniform-cross-section duct establishes clear correlation of the specific speed of the stabilized shock wave to the rate of rise of the temperature of the gas behind it, which gives an opportunity to determine all detonation parameters. The shock wave can be initiated by the detonation of an explosive substance and carries a huge amount of energy. It is shown that the shock wave can be obtained only if the source of small disturbances itself moves at the supersonic speed. It is shown that total pressure behind the shock wave decreases significantly and the entropy increases due to the wave losses, whereas the static pressure increases significantly. An explanation of this effect is given. A formula for calculating the rate of gas temperature rise was derived as a function of the specific speed of the shock wave, the air-and-fuel mixture heat value and the heat availability factor that designates the dissociation of the combustion products and heat loss through the duct wall under specified initial conditions. The reliability of the method of calculating detonation was experimentally substantiated. The work is currently important for the evaluation of the detonation engine efficiency.

Despite the fact that aviation gas turbine engines (GTE) have reached a high degree of sophistication, requirements for the improvement of their efficiency are constantly increasing. Reduction of specific fuel consumption and specific weight of the engine unit makes it possible to improve aircraft performance. One of the effective means of reducing specific fuel consumption and obtaining high thermal efficiency of a gas turbine engine is the use of heat recovery, so the interest in it holds throughout the period of development of gas turbine engines. However, the use of heat recovery in aircraft gas turbine engines is faced with a contradiction: on the one hand, heat recovery allows reducing specific fuel consumption, but, on the other hand, it increases the weight of the power plant due to the presence of a heat exchanger. Moreover, with the increase in the degree of regeneration, specific fuel consumption decreases, whereas the mass of the power plant increases.To obtain the desired effect, it is necessary to optimize simultaneously both the parameters of the engine work process and the degree of regeneration of the heat exchanger according to the criteria of evaluating the engine unit in the aircraft system. For this purpose, it is necessary to have a mathematical model for estimating the weight of a highly efficient aircraft heat exchanger. The article presents a developed mathematical model for calculating the weight of a compact plate heat exchanger used to increase the efficiency of a gas turbine engine due to the heating of compressed air entering the combustion chamber by the hot gas that enters the combustion chamber from behind the turbine. We chose a rational pattern of relative motion of the working media in the heat exchanger, the optimal type of plate-type heat transfer surface in terms of minimizing the weight of the heat exchanger and the hydraulic losses in the air and gas ducts. For the selected surface type, the dependence of the specific weight of the heat exchanger on the degree of regeneration is determined for different nozzle exhaust velocities on the basis of a computational algorithm. To assess the reliability of the obtained model, comparative analysis of the effect of the degree of regeneration on the specific weight of the heat exchanger was carried out, based on the comparison of the results of calculations for the developed model with the data of other authors and with the data for the produced regenerators.

Recently, in order to develop high-strength lightweight core materials of sandwich structures for multi-functional applications, a large number of truss structures have been created, including pyramidal and tetrahedral truss cores. In this paper, a new truss structure is developed to be used as core material in sandwich panels. The X-shaped truss core consists of discrete hourglass-shaped unit cells formed by the groove-to-groove connection of two flat X-shaped truss elements made by metal plate cutting. In order to determine optimal geometrical parameters of a unit cell of X-shaped truss core, in this work it is proposed to plot the diagrams of relative density versus the angle of the rods for the required values of equivalent critical compression and lateral shear stresses, and for the required values of equivalent compression and shear stiffness of the unit cell of X-shaped truss core. The results show that with the same mechanical characteristics, the relative density of the optimal X-shaped truss core is less than the relative density of optimal pyramidal and tetrahedral truss cores.

The article presents a solution of the problem of substantiating the architecture of an interactive information model for the launch preparation of an integrated launch vehicle (ILV). The relevance of the research is due to the complexity and high dynamics of the ILV launch preparation process, as well as the latest achievements in the field of information modeling and decision support theory. The aim of the research is to raise the situational awareness of the decision maker in managing the process of preparing ILV for launch. The process of formation of situational awareness as an integral part of the decision-making process is analyzed. On the basis of this analysis a model of an indicator of the effectiveness of situational awareness formation is proposed for the first time. A model for evaluating the effectiveness of the process under study is developed on the basis of the mathematical apparatus of the theory of continuous-time Markov chains. The results of numerical experiments confirming the adequacy of the model and making it possible to state the quantitative requirements for the process of formation of situational awareness, ensuring its maximum efficiency, are presented. The tasks of the information system aimed at satisfying the requirements are stated, and the architecture of the information model of the ILV launch preparation process is developed. The research results can be used to justify the quantitative requirements for the developed information models of the launch situation at the stage of development of technical specifications.

The failure of heavily loaded rotating parts of aviation gas turbine engines may bring about dangerous consequences. The life of such parts is limited with the use of computational and experimental methods. Computational life prediction methods that are used without carrying out life-cycle tests of engine parts or assemblies should be substantiated experimentally. The best option for verifying the computational methods is to use the results of cyclic tests of model disks. These tests make it possible to reproduce loading conditions and surface conditions that correspond to those of real disks, and the data on the load history and material properties make it possible to simulate stress-strain behavior of disks under test conditions by calculation. This paper shows the process of planning such tests. It is assumed that the tests will be carried out in two stages - before and after the initiation of a low-cycle fatigue crack. A number of criteria are formulated that the geometry of model disks and their loading conditions are to satisfy. Based on these criteria, model disks were designed and the conditions for their testing were selected.

Determination of creep strain arising due to stresses acting up to the moment of time t is considered. The phenomenon of constant-stress creep is described. A method is proposed to determine the parameters of the Arutyunyan creep kernel selected to describe the behavior of a material using two creep curves of a material with viscoelastic properties and nonlinear dependence of creep strain on the stress. In addition, the constant in the expression describing nonlinear dependence of creep strain on the stress is defined. The AMg6M alloy, widely used in the design of aerospace products, was chosen as the material to be analyzed. The tests were carried out on samples 3 mm thick at strains of 65 MPa and 156.9 MPa. According to the results of testing samples of materials on the test bench of Samara University creep curves were obtained. By determining the parameters of the approximation of the Arutyunyan kernel and the parameter included in the expression of nonlinear dependence of creep strain on the stress, it is possible to determine the value of the creep strain of the material for arbitrary values of stress and time. Comparison of the experimental and calculated creep curves for the AMg6M alloy confirms the validity of determination of the rheological characteristics of the tested material.