Computational studies to optimize the geometry of the low-thrust rocket combustion chamber using gaseous propellants

2021 ◽  
Author(s):  
A.V. Novikov ◽  
E.A. Andreev ◽  
E.I. Bardakova
Keyword(s):  
Combustion Chamber ◽  
Thermal State ◽  
Rocket Engine ◽  
Environmental Safety ◽  
Rocket Engines ◽  
Space Technology ◽  
Low Thrust ◽  
Mathematical Experiment ◽  
Space Objects ◽  
Calculation Results

Due to the tough requirements for the environmental safety of the space objects operation, the use of methane-based fuel together with oxygen is a promising direction in developing a new generation of rocket and space technology, including low-thrust rocket engines. When developing low-thrust rocket engines running on oxygen-methane fuel, a mathematical experiment helps to identify the determining factors that affect the quality of the working process in the combustion chamber and to make a calculated optimization of the parameters for supplying fuel components to the combustion chamber. This contributes to a better understanding of the physics of the ongoing processes and leads to recommendations for the design of individual components of the combustion chamber. The numerical simulation enables us to optimize the geometry of the combustion chamber in order to obtain the maximum value of the chamber coefficient, which for an isobaric combustion chamber can be equal to the coefficient of the flow complex. This approach can significantly reduce the number of expensive bench tests. The paper introduces a physical and mathematical model of the workflow in the combustion chamber of a low-thrust rocket engine and gives a comparative analysis of the calculation results for various modifications of the original geometry of the low thrust rocket chamber. Recommendations are given for changing the initial geometry of the combustion chamber in order to increase the coefficient of the flow complex while maintaining a satisfactory thermal state of this chamber.

Numerical study of the influence of the physicochemical liquid fuel properties on the efficiency of the low-thrust rocket engine working process

2021 ◽  
Author(s):  
A.V. Novikov ◽  
E.A. Andreev ◽  
E.I. Bardakova
Keyword(s):  
Combustion Chamber ◽  
Numerical Study ◽  
Rocket Engine ◽  
Propulsion System ◽  
Rocket Engines ◽  
Space Technology ◽  
Low Thrust ◽  
Working Process ◽  
Engine Combustion ◽  
Parametric Studies

Low-thrust rocket engines are widely used in rocket and space technology for correcting the position of a spacecraft in orbit, for controlling motion along a trajectory, etc. Their number in the propulsion system can be from one to tens of units. Accordingly, the efficiency of their work significantly affects the perfection of the propulsion system as a whole. The object of the study was the low-thrust rocket engine combustion chamber operating according to the gas-liquid scheme. There were performed computational and parametric studies of various factor effects on the characteristics of the working process in the combustion chamber. The dependences of the coefficient of the consumable complex and parameters of the working process of the low-thrust rocket engine chamber on the influencing factors when using ethanol and kerosene as a fuel were calculated. A comparative analysis of the results of using these two components under similar conditions was carried out, which made it possible to reveal the influence of the physicochemical properties of the combustible component on the efficiency of the working process organization. The results obtained can be used in the design of low-thrust engines operating on the kerosene–oxygen and ethanol–oxygen propellants.

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.

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.

scholarly journals Сalculated analysis of the influence of operation and design factors on the parameters of oxygen-hydrogen low-thrust rocket engines

2019 ◽  
Vol 18 (3) ◽  
pp. 131-142
Author(s):  
V. V. Ryzhkov ◽  
I. I. Morozov
Keyword(s):  
Combustion Chamber ◽  
Gas Dynamics ◽  
Selection Criterion ◽  
Design Parameters ◽  
Rocket Engines ◽  
Low Thrust ◽  
Gaseous Oxygen ◽  
Computational Gas Dynamics ◽  
Design Factors ◽  
Selection Of

The paper presents the results of calculating thermodynamic and thermophysical properties of the combustion products of gaseous oxygen-hydrogen fuel according to the ideal LRE model taking into account the phase state of the components, as well as the parameters of a low-thrust engine according to the model of computational gas dynamics to ensure the selection of operation and design factors that define the design of a thruster for advanced aerospace objects. It is shown that ideal models can be used for the selection of some parameters, such as: the excess oxidant ratio, the pressure in the combustion chamber, the geometric degree of area expansion ratio. High-level computational gas dynamics models need to be used for the selection of some of the parameters of the engine to be designed, such as: design parameters of the propellant injection pattern, reduced length of the combustion chamber and some others. Air specific impulse was used as the selection criterion. The obtained calculation data allow one to choose the main parameters of the engine being designed with account for real processes in the combustion chamber and the nozzle of the engine.

scholarly journals Computer-aided design of low-thust rocket engines using the domain-specific knowledge database and CAE / CAD systems

2020 ◽  
Vol 18 (4) ◽  
pp. 106-116
Author(s):  
V. V. Ryzhkov ◽  
I. I. Morozov ◽  
E. A. Lapshin
Keyword(s):  
Knowledge Base ◽  
Computer Aided Design ◽  
Rocket Engine ◽  
Rocket Engines ◽  
Low Thrust ◽  
Production Environment ◽  
Gaseous Oxygen ◽  
Engine Model ◽  
Computer Aided ◽  
Aided Design

The paper presents approaches to computer-aided design of low-thrust thrust rocket engines using an extensive knowledge base that allows making basic technical decisions that determine the conceptual design of the engine, based on the developed algorithm of this process. The procedure of creating an electronic 3D-model of a low-thrust rocket engine fueled by gaseous oxygen-hydrogen in the environment of the graphical complex UNIGRAPHICS is described. 3D electronic models of the main elements of a rocket engine with a thrust of P = 25 N were obtained, with subsequent virtual assembly of all components, including the components comprised in the knowledge base, providing the development, among other things, of design documentation, creation of a production environment based on an electronic engine model, preparation for the product manufacturing and the manufacturing proper.

scholarly journals On the question of solid-propellant rocket engine design preventing unstable operation in the combustion chamber

2017 ◽  
pp. 63-72
Author(s):  
G. A. Glebov ◽  
S. A. Vysotskaya
Keyword(s):  
Combustion Chamber ◽  
Solid Propellant ◽  
Rocket Engine ◽  
Rocket Engines ◽  
Pressure Oscillations ◽  
Engine Design ◽  
Solid Propellant Rocket ◽  
Solid Propellant Rocket Engine ◽  
Flow Duct ◽  
Propellant Rocket

The paper presents results of a numerical investigation concerning the effect that the flow duct shape and combustion rate equation have on the gas dynamic vortex flow pattern and self-excited pressure oscillations in the combustion chamber of a solid-propellant rocket engine. We provide guidelines on upgrading solid-propellant rocket engines in order to decrease the magnitude of pressure pulses in the case of pulsating combustion.

scholarly journals Factors Affecting Characteristic Length of the Combustion Chamber of Liquid Propellant Rocket Engines

2019 ◽  
Vol 38 (3) ◽  
pp. 729-744
Author(s):  
Tajwali Khan ◽  
Ihtzaz Qamar
Keyword(s):  
Combustion Chamber ◽  
Characteristic Length ◽  
Rocket Engine ◽  
Droplet Diameter ◽  
Computer Code ◽  
Rocket Engines ◽  
Fuel Droplet ◽  
Liquid Propellant ◽  
Proposed Model ◽  
Propellant Rocket

Optimum characteristic length of the combustion chamber of liquid rocket engine is very important to get higher energy from the liquid propellants. Characteristic length is defined by the time required for complete burning of fuel. Combustion reactions are very fast and combustion is evaporation dependent. This paper proposes fuel droplet evaporation model for liquid propellant rocket engine and discusses the factors which can affect the required size of characteristic length of the combustion chamber based on proposed model. The analysis is performed for low temperature combustion chamber. A computer code based on proposed model is generated, which solve analytical equations to calculate combustion chamber characteristic length under various input conditions. The analysis shows that characteristic length is affected by combustion chamber temperature, pressure, fuel droplet diameter, chamber diameter, mass flow rate of propellants and relative velocity of the droplet in the combustion chamber.

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