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.

METHODS FOR COOLING THE HOT ZONE OF HEAT ENGINES AND TECHNOLOGICAL TECHNIQUES FOR IMPLEMENTING THE METHODS

2020 ◽  
pp. 10-14
Author(s):  
E. V. Panichev ◽  
V. P. Smolentsev ◽  
A. V. Shchednov
Keyword(s):  
Combustion Chamber ◽  
Protective Coatings ◽  
Rocket Engines ◽  
Internal Cooling ◽  
Fuel Component ◽  
Cooling Systems ◽  
Cooling Medium ◽  
Aerospace Technology ◽  
Liquid Rocket ◽  
Hot Zone

Various types of hot zone cooling systems for liquid rocket engines are considered. The analysis of systems for external and internal cooling of the hot zone of the combustion chamber and the jet co-PLA is performed. The area of primary use of external and internal impact of components on the combustion chamber and the jet nozzle of a temporary product with a high thermal load on the material is given. Recommendations on the choice of technological methods for processing heat-protective coatings were developed, where the advantages of combined electroabrasive treatment were revealed. In this case, the possibility of finishing the transition sections of the path from areas that are difficult to access for the tool, which have coatings with ceramic granules that have been used in the latest products of aerospace technology, is shown. The mechanism of heat transfer by the cooling medium from the areas of the greatest heating, as a rule, to the flow of the liquid fuel component of the fuel is shown. Examples of implementation of modern cooling systems on typical designs of combustion chambers of modern rocket engines are considered. It is shown that the cooling efficiency will be higher the more belts of the cooling medium curtain are located along the length of the combustion chamber. At the same time, it should be taken into account that the use of an excessively large number of belts means a significant complication of the camera design, its manufacturing technology, and an increase in the cost of the product.

Innovative methods for obtaining artificial roughness on the surfaces of heat-loaded parts of the liquid rocket engine combustion chambers

2020 ◽  
Author(s):  
A.Yu. Ryazantsev ◽  
S.S. Yukhnevich ◽  
A.A. Shirokozhukhova
Keyword(s):  
Combustion Chamber ◽  
Rocket Engine ◽  
Empirical Studies ◽  
Aerospace Industry ◽  
Rocket Engines ◽  
Artificial Roughness ◽  
Special Equipment ◽  
Liquid Rocket Engines ◽  
Engine Combustion ◽  
Liquid Rocket

The paper shows the applications of combined processing in the manufacture of parts and assembly units of liquid rocket engines in the aerospace industry. The most effective methods of obtaining artificial roughness on the surfaces of special equipment products are considered. Empirical studies of changes in the physical and mechanical properties of the material are performed using various methods of combined processing. Qualitative and quantitative relationships between the hydraulic characteristics of the rocket engine combustion chamber manufactured using the combined method, and the quality of the surface layer of the product are described and formalized. The analysis of modern processing methods is performed, and the latest methods for obtaining artificial roughness on the surfaces of rocket engine parts are presented. The relevance and need for the use of high-end technology in obtaining surface layers of products included in the structure of the combustion chamber of liquid rocket engines are proved. The results obtained allow significant expanding the technological capabilities of production, as well as appreciable improving the technical characteristics of special equipment products in the aerospace industry.

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.

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.

On Computing Losses in Blading Sections of Liquid Rocket Engine Pressurisation Stations

2020 ◽  
pp. 21-34
Author(s):  
D.A. Zhuykov ◽  
A.A. Zuev ◽  
M.I. Tolstopyatov
Keyword(s):  
Boundary Layer ◽  
Rocket Engine ◽  
Significant Proportion ◽  
Friction Loss ◽  
Liquid Rocket Engine ◽  
Rocket Engines ◽  
Initial Region ◽  
Friction Laws ◽  
Liquid Rocket Engines ◽  
Liquid Rocket

Designing more sophisticated contemporary liquid rocket engines requires a precise understanding of the hydrodynamics in the blading sections of the pressurisation station, which is most often a turbopump. Friction loss in blade passages and outlets forms a significant proportion of all losses. The paper shows that it is necessary to account for the initial region of hydrodynamically unbalanced flow in the boundary layer, which is most characteristic of relatively short passages in blading sections of liquid rocket engine turbopumps. We performed the analysis required to select friction drag laws for components of pressurisation station blading sections. We considered and proposed a method for numerically integrating a system of equations to determine the variation in characteristic thickness of a spatial boundary layer and friction loss, accounting for the inertial component of the flow core velocity, depending on which flow modes occur in the components of pressurisation station blading sections in a liquid rocket engine. We show that it is necessary to correctly select the friction laws and to take the initial region into account so as to precisely determine the power parameters

Conceptual Regenerative Nozzle Cooling Design for a Hydroxyl-Terminated Polybutadiene and Oxygen Hybrid Rocket Engine

2017 ◽  
Author(s):  
Luis R. Robles ◽  
Johnny Ho ◽  
Bao Nguyen ◽  
Geoffrey Wagner ◽  
Jeremy Surmi ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
High Temperatures ◽  
Cooling System ◽  
Rocket Engine ◽  
Combustion Efficiency ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Gaseous Oxygen ◽  
Rocket Nozzle ◽  
Regenerative Cycle

Regenerative rocket nozzle cooling technology is well developed for liquid fueled rocket engines, but the technology has yet to be widely applied to hybrid rockets. Liquid engines use fuel as coolant, and while the oxidizers typically used in hybrids are not as efficient at conducting heat, the increased renewability of a rocket using regenerative cycle should still make the technology attractive. Due to the high temperatures that permeate throughout a rocket nozzle, most nozzles are predisposed to ablation, supporting the need to implement a nozzle cooling system. This paper presents a proof-of-concept regenerative cooling system for a hybrid engine which uses hydroxyl-terminated polybutadiene (HTPB) as its solid fuel and gaseous oxygen (O2) as its oxidizer, whereby a portion of gaseous oxygen is injected directly into the combustion chamber and another portion is routed up through grooves on the exterior of a copper-chromium nozzle and, afterwards, injected into the combustion chamber. Using O2 as a coolant will significantly lower the temperature of the nozzle which will prevent ablation due to the high temperatures produced by the exhaust. Additional advantages are an increase in combustion efficiency due to the heated O2 being used for combustion and an increased overall efficiency from the regenerative cycle. A computational model is presented, and several experiments are performed using computational fluid dynamics (CFD).

Combustion Test of Oxidizer-Rich Single Triplex Injector Preburner for Staged Combustion Cycle Rocket Engine

2015 ◽  
Author(s):  
Su-Ji Lee ◽  
In-Sang Moon ◽  
Il-Yoon Moon ◽  
Seong-Up Ha
Keyword(s):  
Combustion Chamber ◽  
Dynamic Pressure ◽  
Rocket Engine ◽  
Rocket Engines ◽  
Fuel Injector ◽  
Staged Combustion ◽  
Combustion Pressure ◽  
The Republic ◽  
Combustion Tests ◽  
Combustion Cycle

In the Republic of Korea, research on staged-combustion cycle liquid propellant rocket engines (LPRE) is proceeding to improve efficiency of rocket engines. Recently oxidizer-rich preburner using single triplex injector is developed in relation to the main injector development and combustion tests have been performed. This preburner is designed to operate in nominal conditions with the combustion pressure of 10 MPa, OF ratio of 60. For a stable ignition, LOx is fed in two steps. Triethylaluminum-Triethylborane (TEAB) is used as hypergolic fuel for ignition, supplied through a fuel injector. Despite the small amount of fuel flow rate and high pressure condition, the combustion pressure was stably maintained around 10 MPa as designed. As a result of Fast Fourier Transform (FFT) of the combustion chamber dynamic pressure, 1L mode frequencies related to the acoustic instability and hydraulic resistance exist in the combustion chamber. But their amplitudes are less than 1% of the combustion pressure and it does not affect the combustion. Therefore combustion test is stably completed. In the near future, coupled tests with uni-element triplex injector preburner and uni-element gas/liquid injector main combustion chamber will be carried out.

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