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.

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 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).

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.

Numerical Analysis of Tip Clearance Effects in an Unshrouded Impeller for Liquid Rocket Engines

2018 ◽  
Author(s):  
Lucrezia Veggi ◽  
Mattia Reganaz ◽  
Julian D. Pauw ◽  
Oskar J. Haidn ◽  
Bernd Wagner
Keyword(s):  
Rocket Engine ◽  
New Technology ◽  
Tip Clearance ◽  
Tip Vortex ◽  
Rocket Engines ◽  
Relative Movement ◽  
Blade Passage ◽  
Design Concepts ◽  
Liquid Rocket ◽  
The Impact

For rocket engine applications an unshrouded impeller is a relatively new technology compared to the traditional shrouded impeller. An unshrouded impeller has advantages in blade optimisation, easier manufacturing, and — moreover — a potential for weight reduction of the turbopump and consequentely of the entire rocket engine. However, it also results in tip clearance during operation which has detrimental effects on the performance. Therefore, a better understanding of the impact of tip clearance on the overall flow would help to develop and adapt new design concepts. In this study a comparison between shrouded and unshrouded impellers has been numerically undertaken. The shrouded impeller L17, designed within the framework of the research project KonRAT at the Technical University of Munich, is chosen as reference geometry for this study. The unshrouded impellers are derived from the L17 design by simply removing the shroud. Three configurations of the unshrouded L17 impeller have been numerically investigated by varying the tip clearance: geometry G4 with a constant tip clearance of 4% and geometry G7 with 7% of the blade passage height at impeller outlet. The third configuration is a fictitious unshrouded impeller without any tip clearance (geometry G0). The intention of the comparison between the shrouded impeller L17 and the geometry G0 is to isolate the effects on the flow caused by the relative movement between casing and blade to allow a quantification of the tip vortex losses. The obtained performance data as well as the mechanisms of the tip vortex inside the blade passage are discussed in detail. The results show that the effects of the shear forces due to the relative movement and the effects of the tip vortex do not undergo a linear coupling and their interaction inhibits a precise quantification of the tip vortex losses. The analysis also shows that the crucial factor for the deterioration of the performance is the change of the inflow due to the influence of the tip clearance on the flow.

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.

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