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.

scholarly journals Development Trend of Korean Staged Combustion Cycle Rocket Engine

2018 ◽  
Vol 22 (3) ◽  
pp. 109-119 ◽  
Author(s):  
Chae-hyoung Kim ◽  
Yeoung Min Han ◽  
Namkyung Cho ◽  
Seung-Han Kim ◽  
Byungil Yu ◽  
...  
Keyword(s):  
Rocket Engine ◽  
Development Trend ◽  
Staged Combustion ◽  
Combustion Cycle

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).

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.

Combustion instability in a small liquid rocket motor

1999 ◽  
Vol 103 (1023) ◽  
pp. 245-252 ◽  
Author(s):  
A. Osherov ◽  
B. Natan
Keyword(s):  
Combustion Chamber ◽  
High Efficiency ◽  
Dynamic Pressure ◽  
Rocket Engine ◽  
Combustion Instability ◽  
Acoustic Resonator ◽  
Vibration Data ◽  
Engine Design ◽  
Data Processing Method ◽  
Liquid Rocket

Abstract An experimental investigation of high frequency combustion instability in a liquid rocket engine of 3kN thrust was conducted. The diagnostic method of detecting combustion instability was based on the measurement of the dynamic pressure following an artificial disturbance in the combustion chamber. Test data spectral analysis was performed by using the stochastic vibration data processing method. Although the engine has demonstrated an absence of tendency to spontaneous instability, insufficient stability of the original engine design was evident during tests with artificial triggers. A Helmholtz type resonator in form of partitioned cavities tuned to a few different, close frequencies was designed and installed in the combustion chamber wall to avoid spontaneous or triggered combustion instability. The experimental results from hot tests with artificial triggers confirmed the high efficiency of the applied acoustic resonator.

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.

scholarly journals A Study on Anti-oxidization Coating for Staged Combustion Cycle Rocket Engines

2018 ◽  
Vol 22 (5) ◽  
pp. 125-131
Author(s):  
Young-June Kim ◽  
Byong-ho Rhee ◽  
Yong-Oh Noh ◽  
Byung-Hyun Bae ◽  
Seong-Yoon Hyun ◽  
...  
Keyword(s):  
Rocket Engines ◽  
Staged Combustion ◽  
Combustion Cycle

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.

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.

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