scholarly journals Driving and damping mechanisms for transverse combustion instabilities in liquid rocket engines

2017 ◽  
Vol 820 ◽  
Author(s):  
A. Urbano ◽  
L. Selle
Keyword(s):  
Rocket Engine ◽  
Three Dimensional ◽  
Rocket Engines ◽  
Combustion Instabilities ◽  
Time Resolved ◽  
Physical Mechanisms ◽  
Eddy Simulation ◽  
Liquid Rocket Engines ◽  
Large Eddy ◽  
Liquid Rocket

This work presents the analysis of a transverse combustion instability in a reduced-scale rocket engine. The study is conducted on a time-resolved database of three-dimensional fields obtained via large-eddy simulation. The physical mechanisms involved in the response of the coaxial hydrogen/oxygen flames are discussed through the analysis of the Rayleigh term in the disturbance-energy equation. The interaction between acoustics and vorticity, also explicit in the disturbance-energy balance, is shown to be the main damping mechanism for this instability. The relative contributions of Rayleigh and damping terms, depending on the position of the flame with respect to the acoustic field, are discussed. The results give new insight into the phenomenology of transverse combustion instabilities. Finally, the applicability of spectral analysis on the nonlinear Rayleigh and dissipation terms is discussed.

scholarly journals Numerical Investigation of Injection, Mixing and Combustion in Rocket Engines Under High-Pressure Conditions

2020 ◽  
pp. 209-221
Author(s):  
Christoph Traxinger ◽  
Julian Zips ◽  
Christian Stemmer ◽  
Michael Pfitzner
Keyword(s):  
High Pressure ◽  
Large Eddy Simulations ◽  
Experimental Investigations ◽  
High Fidelity ◽  
Rocket Engines ◽  
Eddy Simulation ◽  
Liquid Rocket Engines ◽  
Large Eddy ◽  
Numerical Tools ◽  
Liquid Rocket

Abstract The design and development of future rocket engines severely relies on accurate, efficient and robust numerical tools. Large-Eddy Simulation in combination with high-fidelity thermodynamics and combustion models is a promising candidate for the accurate prediction of the flow field and the investigation and understanding of the on-going processes during mixing and combustion. In the present work, a numerical framework is presented capable of predicting real-gas behavior and nonadiabatic combustion under conditions typically encountered in liquid rocket engines. Results of Large-Eddy Simulations are compared to experimental investigations. Overall, a good agreement is found making the introduced numerical tool suitable for the high-fidelity investigation of high-pressure mixing and combustion.

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

scholarly journals Numerical Analysis of Self-Excited Combustion Instabilities in a Small MMH/NTO Liquid Rocket Engine

2020 ◽  
Vol 2020 ◽  
pp. 1-17
Author(s):  
Jianxiu Qin ◽  
Huiqiang Zhang
Keyword(s):  
Droplet Size ◽  
Mass Fraction ◽  
Rocket Engine ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Liquid Rocket Engine ◽  
Combustion Instabilities ◽  
Pressure Oscillations ◽  
Two Phase ◽  
Liquid Rocket

Combustion instabilities in a small MMH/NTO liquid rocket engine used for satellite attitude and course control are numerically investigated. A three-dimensional Navier-Stokes code is developed to simulate two-phase spray combustion for cases with five different droplet Sauter Mean Diameters. As the droplet size increases from 30 microns to 80 microns, pressure oscillations are stronger with larger amplitudes. But an increase of the droplet size in the range of 80 microns to 140 microns indicates a reduction in the amplitudes of pressure oscillations. This trend is the same as the Hewitt criterion. The first tangential (1T) mode and the first longitudinal (1L) mode self-excited combustion instabilities are captured in the 60-micron and 80-micron cases. Abrupt spikes occur in the mass fraction of MMH and coincide with abrupt spikes in the mass fraction of NTO at the downstream regions just adjacent to the impinging points. Thus, local combustible high-dense mixtures are formed, which result in quasiconstant volume combustion and abrupt pressure spikes. The propagation and reflection of pressure waves in the chamber stimulate the combustion instability. When the droplet size is too small or too large, it is difficult to form local high-dense premixtures and combustion is stable in the chamber.

scholarly journals Effects of LOX Particle Diameter on Combustion Characteristics of a Gas-Liquid Pintle Rocket Engine

2020 ◽  
Vol 2020 ◽  
pp. 1-16
Author(s):  
Xuan Jin ◽  
Chibing Shen ◽  
Rui Zhou ◽  
Xinxin Fang
Keyword(s):  
Rocket Engine ◽  
Oscillation Amplitude ◽  
Three Dimensional ◽  
Particle Diameter ◽  
Low Frequency ◽  
Pressure Oscillation ◽  
Combustion Efficiency ◽  
Combustion Characteristics ◽  
Rocket Engines ◽  
Liquid Rocket

LOX/GCH4 pintle injector is suitable for variable-thrust liquid rocket engines. In order to provide a reference for the later design and experiments, three-dimensional numerical simulations with the Euler-Lagrange method were performed to study the effect of the initial particle diameter on the combustion characteristics of a LOX/GCH4 pintle rocket engine. Numerical results show that, as the momentum ratio between the radial LOX jet and the axial gas jet is 0.033, the angle between the LOX particle trace and the combustor axial is very small. Due to the large recirculation zones, premixed combustion mainly occurs in the injector wake region. As the initial LOX particle diameter increases, the LOX evaporation rate and the combustion efficiency decrease until the combustion terminates with the initial LOX particle diameter greater than 110 μm. The oscillation amplitude of the combustor pressure increases significantly along with the increase of the initial LOX particle diameter, and the low-frequency unstable combustion occurs when the initial LOX particle diameter exceeds 60 μm. The combustor pressure oscillation at about 40 Hz couples with the swinging process of spray and flame, while the unsteady LOX evaporation amplifies the combustor pressure oscillations at 80 Hz and its harmonic frequency.

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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