SIMULATION OF THE CURRENT-VOLTAGE CHARACTERISTICS OF SOLAR BATTERIES WITH CONSIDERING DECREASE OF LIGHTING DUE TO THE INFLUENCE OF JETS OF THE ELECTRIC PROPULSION

2019 ◽  
pp. 56-64
Author(s):  
Ya.N. MIGUNOV ◽  
V.V. ONUFRIEV
Keyword(s):  
Solar Cells ◽  
Temperature Gradient ◽  
Electric Propulsion ◽  
Protective Coatings ◽  
Rocket Engine ◽  
Solar Array ◽  
Rocket Engines ◽  
Lighting Conditions ◽  
Main Equation ◽  
Voltage Current Characteristic

A model for calculating the voltage-current characteristic of a solar array in the presence of a temperature gradient by its photovoltaic converters and their variable illumination due to possible pollution under the action of space factors, including operation of electric rocket engines, is described. The model is based on the main equation of a solar cell. In this case both a one-dimensional and a two-dimensional temperature gradients are taken into account. The principles of constructing a model are given, and the initial data selection is made. To simulate the lighting conditions of the solar array such a concept as effective illumination is used, i.e. the density of the radiation flux which falls on photovoltaic converters passing through the protective coatings. The features of simulation of the temperature distribution in the solar array and the effective illumination of its surface in cases of parallel, serial and mixed switching of solar cells are described. The calculation procedures and examples of solar cells are given. The construction of the model in universal mathematical package Mathcad is described. Some simulation results are discussed. Key words: solar array, mathematical simulation, illumination, temperature gradient, electric rocket engine, spacecraft, Mathcad.

Investigation of the Properties of Heat-Protective Coatings to Improve the Performance of Products

2021 ◽  
Vol 1037 ◽  
pp. 516-521
Author(s):  
Vladislav Smolentsev ◽  
Nikolay Nenahov ◽  
Natalia Potashnikova
Keyword(s):  
Combustion Chamber ◽  
Protective Coatings ◽  
Rocket Engine ◽  
Multilayer Coating ◽  
Surface Preparation ◽  
Rocket Engines ◽  
Initial Surface ◽  
Technological Parameters ◽  
Materials Used ◽  
Liquid Rocket

The heat-loaded part of the combustion chamber of a liquid rocket engine are Considered. The proposed coating has several layers: an internal metal coating that contacts the part or substrate, and an external coating made of a mixture of ceramic granules and metal powder. At the same time, to obtain the initial surface for coating with the required surface layer roughness, it is proposed to use the method of sand blasting. The article analyzes possible mechanisms of material formation for "base-coating" transition zones, as well as the influence of their chemical composition on the adhesive strength of layers.. The choice of brand and combination of materials used for coating is justified. Technological modes that have been tested in production conditions when applying heat-resistant coatings to parts of modern rocket engines are proposed. The influence of technological parameters of the initial surface preparation process and the geometry of the resulting micro-relief of the substrate on the adhesion characteristics of a multilayer coating made of heat-protective materials operating in the high-temperature zone of the combustion chamber of liquid rocket engines is revealed.

COMPARATIVE ANALYSIS OF THE EFFICIENCY OF CONSTANT POWER THROTTLED ROCKET ENGINES FOR INTERORBITAL FLIGHTS TO GEOSTATIONAR

2021 ◽  
Vol 6 ◽  
pp. 66-77
Author(s):  
Igor Vasiliev ◽  
◽  
Boris Kiforenko ◽  
Yaroslav Tkachenko ◽  
◽  
...  
Keyword(s):  
Mathematical Models ◽  
Electric Propulsion ◽  
Rocket Engine ◽  
Power Level ◽  
Specific Impulse ◽  
Rocket Engines ◽  
Low Thrust ◽  
Constant Power ◽  
Control Functions ◽  
Thrust Control

Carrying out low-thrust transfers of spacecrafts in the near-earth space from intermediate elliptic to the geostationary orbit using electric rocket engines seems to be one of the most important tasks of modern cosmonautics. Electric rocket engines, whose specific impulse of the reactive jet is an order of magnitude more than in chemical RD, are preferable for interorbit flights with a maximum payload in the case when a significant increase in the duration of the maneuver is permissible. Ability to throttling the rocket engine thrust is traditionally considered as one of the ways to reduce both the engine mass and the required fuel assumptions for performing the specified maneuver. Using the concept of an ideal-rocket engine provides the upper estimates of the payload mass of interborbital flights for the given power level. Accounting for the properties of real engines leads to the need of considering the mathematical models with more strict limits on control functions. A study of the efficiency of three modes of thrust control of an electric propulsion rocket engine was carried out when performing practically interesting spacecraft flights from highly elliptical intermediate near-earth orbits to geostationary orbits. A mathematical model of constant power relay rocket engine has been built. The formulation of the variational problem of the Maer type is given about the execution of a given dynamic maneuver for the throttled and unregulated electric rocket engines of constant power. Using the Pontryagin maximum principle, an analysis of the optimal control functions was carried out, for which the final relations were written out, which allowed to write down the system of differential equations of the optimal movement of the spacecraft, equipped with relay electric rocket engine. The obtained numerical and quality results of the study of the effectiveness of various modes of thrust control of an electric propulsion engine to increase the payload of a given orbital maneuver confirmed the correctness of mathematical models of throttled and relay engines and, in general, the efficiency of using solutions of the averaged equations of optimal motion of a spacecraft for numerical solution of the corresponding boundary value problems in an exact formulation.

Graphic method for assessing the margins of resistance to acoustic vibrations in the chambers of rocket engines by combustion noise

2020 ◽  
pp. 15-21
Author(s):  
R.A. Tsarapkin ◽  
V.N. Ivanov ◽  
V.I. Biryukov
Keyword(s):  
Rocket Engine ◽  
Spectral Characteristics ◽  
Combustion Process ◽  
Combustion Noise ◽  
Rocket Engines ◽  
Acoustic Vibrations ◽  
Statistical Regression ◽  
Unknown Parameters ◽  
Damping Decrement ◽  
Resonance Frequencies

An experimental method is proposed for estimating the damping decrements of pressure fluctuations in the combustion chambers of forced rocket engines. The method is based on the statistical processing of noise pressure pulsations in the vicinity of natural resonance frequencies for normal modes of acoustic vibrations of the reaction volume and the subsequent prediction of the instability of the combustion process relative to acoustic vibrations. Based on the theory of statistical regression for multidimensional experimental data, the problem of predicting unknown parameters of sample distributions is solved by asymptotic determination of the correlation coefficient of the damping decrement of pressure vibrations through optimal linear predictors and the Kolmogorov distribution. Keywords rocket engine, combustion chamber, acoustic vibrations, combustion noise, spectral characteristics, Kolmogorov criterion, damping decrement. [email protected]

Investigation of the attenuation of the fluxes of the low-energy plasma component of electric rocket engines by the ventilation holes of the unpressurized instrument compartment of the spacecraft

2021 ◽  
Author(s):  
I.A. Maximov ◽  
A.B. Nadiradze ◽  
R.R. Rakhmatullin ◽  
V.A. Smirnov ◽  
R.E. Tikhomirov ◽  
...  
Keyword(s):  
Electric Propulsion ◽  
Plasma Source ◽  
Low Energy ◽  
Rocket Engines ◽  
Plasma Component ◽  
Plasma Flows ◽  
Energy Component ◽  
Honeycomb Filler ◽  
Semi Empirical ◽  
Made In

The results of an experimental study of the attenuation of the fluxes of the low-energy component of the plasma formed during the operation of electric propulsion engines (ERE), ventilation holes (VH) of the non-sealed equipment compartment (NSEC) of the spacecraft (SC) are presented. Authors studied the attenuation of plasma fluxes by standard VHs made in honeycomb panels that form the NSEC. A Hall-effect engine of the SPT-70 type was used as a plasma source. The experiment consisted of measuring the plasma concentration at the inlet and outlet of the VH. The concentration at the inlet was measured with a flat Langmuir probe, and at the outlet with a Faraday probe, which allows collecting all ions passing through the VH. The aim of the work was to study the weakening of the fluxes of the lowenergy component of the EJE plasma when passing through the VH in the honeycomb-nels that form the NSEC. Based on the experimental data, a semi-empirical model was constructed that describes the dependence of the attenuation coefficient of plasma flows on the geometric parameters of the vent-holes. It has been established that a vent-holes of this design attenuates the plasma flows by 102 ... 104 times. The largest contribution to the weakening of plasma fluxes is made by the honeycomb filler, which is due to the recombination of ions during their collision with the channel walls. Taking into account the attenuation of the fluxes of the low-energy component of the plasma of electric rocket engines by ventilation holes is a key stage in assessing the effect of plasma on the power on-board equipment of spacecraft and should be used by spacecraft developers when analyzing the resistance to this factor.

scholarly journals Determination of receiver consumption characteristics using computer simulation

2020 ◽  
Vol 8 (2) ◽  
pp. 10-14
Author(s):  
S.S. Vasyliv ◽  
◽  
V.S. Zhdanov ◽  
M.V. Yevseyenko ◽  
◽  
...  
Keyword(s):  
Computer Simulation ◽  
Flow Rate ◽  
Rocket Engine ◽  
Flow Characteristics ◽  
Maximum Flow ◽  
Complex Configuration ◽  
Rocket Engines ◽  
Gas Leakage ◽  
Control Forces ◽  
Firing Tests

The problem of implementing the detonation mode of fuel combustion in thermal propulsion systems has been widely studied last decade. There are many works on fundamental and applied research on pulsating detonation. Solid propellant detonation engines can develop significant forces for a short time at low structural masses, and therefore they are ideal for auxiliary systems for the removal of separated rocket parts. In addition, detonation processes can be used to create control forces for correcting the trajectory of aircraft. All these facts determine the relevance of the area of work. For studying detonation installations, it is necessary to create test stands, but the design of test installations is an urgent and complex optimization problem. It is advisable to solve this problem with the help of computer simulation. In the existing experimental methods, for designing, it is necessary to determine in advance the geometric parameters of receivers and pipelines that provide the necessary gas consumption for firing tests of detonation rocket engines. The work is devoted to the development of a method for determining the flow characteristics of a receiver with a pipeline of complex configuration based on the constructed model of the stand. Based on the initial data, a computer simulation of the air leakage process from the receiver was carried out, for which the Solid Works software package was used. The places of pressure drop, maximum flow rate, and air mass flow are determined. The low value of the flow rate factor is due to the complex configuration of the pipeline with numerous bends and two bellows. Comparison of calculation results with experimental data was held. The difference between the experimental and calculated values does not exceed 3.6%. The obtained information is used to select the required value of the oxidizer excess coefficient during firing tests of detonation rocket engine models. Keywords: flow rate, gas leakage, receiver, model.

scholarly journals Liquid rocket engine performance assessment in the context of small launcher optimisation

2019 ◽  
Vol 11 (3) ◽  
pp. 135-145 ◽  
Author(s):  
Alexandru-Iulian ONEL ◽  
Oana-Iuliana POPESCU ◽  
Ana-Maria NECULAESCU ◽  
Tudorel-Petronel AFILIPOAE ◽  
Teodor-Viorel CHELARU
Keyword(s):  
Mathematical Model ◽  
Performance Assessment ◽  
Rocket Engine ◽  
Engine Performance ◽  
Data Fitting ◽  
Fuel Combustion ◽  
Liquid Rocket Engine ◽  
Rocket Engines ◽  
Optimisation Algorithm ◽  
Liquid Rocket

The paper presents a fast mathematical model that can be used to quickly asses the propulsive characteristics of liquid propelled rocket engines. The main propulsive parameters are computed using combustion surfaces obtained after a nonlinear data fitting analysis. This approach is much more time efficient than using standard codes which rely on frequent calls of the Fuel Combustion Charts and interpolating their data. The tool developed based on the proposed mathematical model can be used separately or it can be integrated in a multidisciplinary optimisation algorithm for a preliminary microlauncher design.

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