scholarly journals Increase in Performance of Laser Ignition Micro Solid Rocket by Control of Combustion Chamber Pressure

2013 ◽  
Vol 61 (1) ◽  
pp. 9-15 ◽  
Author(s):  
Yusuke MASUDA ◽  
Hiroyuki KOIZUMI ◽  
Tomoyuki HAYASHI ◽  
Masakatsu NAKANO ◽  
Kimiya KOMURASAKI ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
Chamber Pressure ◽  
Laser Ignition ◽  
Solid Rocket

scholarly journals Numerical Study on Response Characteristics of Solid Rocket Pintle Motor

2019 ◽  
Author(s):  
Jeevan Sapkota ◽  
Yi Hua Xu ◽  
Hai Jun Sun
Keyword(s):  
Flow Field ◽  
Combustion Chamber ◽  
Total Pressure ◽  
Numerical Study ◽  
Rocket Motor ◽  
Chamber Pressure ◽  
Head Model ◽  
Solid Rocket Motor ◽  
Rapid Changes ◽  
Solid Rocket

Pintle technology is currently a versatile technology used in a solid rocket motor (SRM) to control the desired thrust by changing the nozzle throat area, while effectively controlling the chamber pressure at the same time. The sudden movement of the pintle can induce rapid changes in the flow field and the occurrence of pressure oscillations inside the combustion chamber. The analysis of such rapid changes is essential to design an efficient controllable pintle rocket motor for a better thrust regulation. Two-dimensional axisymmetric models with mesh generation and required boundary condition were designed to analyze the effects of three different pintle head shape models in SRM thrust regulation effect. Dynamic mesh method was used with specific velocity for moving plug/pintle in the numerical analysis of SRM thrust regulation. The effects of different pintle head models on the flow field, combustion chamber pressure, mass-flow rate, thrust and Mach number were investigated. According to the analysis of total pressure response time, the simulation data revealed that circular pintle head model responded faster among three different models. According to the thrust effect, parabolic pintle has the maximum value of thrust and the greatest total pressure recovery coefficient among all pintle head models.

Evaluation of Combustion Chamber Pressure of 1N-Class Green Monopropellant Thruster with Discharge Plasma System

2017 ◽  
Vol 65 (3) ◽  
pp. 117-122
Author(s):  
Asato WADA ◽  
Hiroshi MAEDA ◽  
Takahiro SHINDO ◽  
Hiroki WATANABE ◽  
Haruki TAKEGAHARA
Keyword(s):  
Combustion Chamber ◽  
Discharge Plasma ◽  
Chamber Pressure ◽  
Plasma System

Preliminary Studies on Non-Reactive Flow Vortex Cooling

2019 ◽  
Vol 12 (3) ◽  
pp. 262-271
Author(s):  
T.N. Rajesh ◽  
T.J.S. Jothi ◽  
T. Jayachandran
Keyword(s):  
Combustion Chamber ◽  
Inner Core ◽  
Rocket Engine ◽  
Injection Pressure ◽  
Chamber Pressure ◽  
Reactive Flow ◽  
Stoichiometric Ratio ◽  
Gaseous Oxygen ◽  
Mixture Ratio ◽  
Vortex Combustion

Background: The impulse for the propulsion of a rocket engine is obtained from the combustion of propellant mixture inside the combustion chamber and as the plume exhausts through a convergent- divergent nozzle. At stoichiometric ratio, the temperature inside the combustion chamber can be as high as 3500K. Thus, effective cooling of the thrust chamber becomes an essential criterion while designing a rocket engine. Objective: A new cooling method of thrust chambers was introduced by Chiaverni, which is termed as Vortex Combustion Cold-Wall Chamber (VCCW). The patent works on cyclone separators and confined vortex flow mechanism for providing high propellant mixing with improved degree of turbulence inside the combustion chamber, providing the required notion for studies on VCCW. The flow inside a VCCW has a complex structure characterised by axial pressure losses, swirl velocities, centrifugal force, flow reversal and strong turbulence. In order to study the flow phenomenon, both the experimental and numerical investigations are carried out. Methods: In this study, non-reactive flow analysis was conducted with real propellants like gaseous oxygen and hydrogen. The test was conducted to analyse the influence of mixture ratio and injection pressure of the propellants on the chamber pressure in a vortex combustion chamber. A vortex combustor was designed in which the oxidiser injected tangentially at the aft end near the nozzle spiraled up to the top plate and formed an inner core inside the chamber. The fuel was injected radially from injectors provided near the top plate and the propellants were mixed in the inner core. This resulted in enhanced mixing and increased residence time for the fuel. More information on the flow behaviour has been obtained by numerical analysis in Fluent. The test also investigated the sensitivity of the tangential injection pressure on the chamber pressure development. Results: All the test cases showed an increase in chamber pressure with the mixture ratio and injection pressure of the propellants. The maximum chamber pressure was found to be 3.8 bar at PC1 and 2.7 bar at PC2 when oxidiser to fuel ratio was 6.87. There was a reduction in chamber pressure of 1.1 bar and 0.7 bar at PC1 and PC2, respectively, in both the cases when hydrogen was injected. A small variation in the pressure of the propellant injected tangentially made a pronounced effect on the chamber pressure and hence vortex combustion chamber was found to be very sensitive to the tangential injection pressure. Conclusion: VCCW mechanism has been to be found to be very effective for keeping the chamber surface within the permissible limit and also reducing the payload of the space vehicle.

Solid Rocket Motor Combustion Chamber Aeroacoustics

AIAA Journal ◽  
1978 ◽  
Vol 16 (11) ◽  
pp. 1123-1124
Author(s):  
Warren C. Strahle ◽  
John C. Handley
Keyword(s):  
Combustion Chamber ◽  
Rocket Motor ◽  
Solid Rocket Motor ◽  
Solid Rocket

Development Results of Laser Ignition System for Solid Rocket Motor

2021 ◽  
Vol 19 (5) ◽  
pp. 807-811
Author(s):  
Keisuke MINAMI ◽  
Yoshiki MATSUURA ◽  
Koki KITAGAWA ◽  
Satoshi ARAKAWA ◽  
Naoki MORISHITA ◽  
...  
Keyword(s):  
Rocket Motor ◽  
Laser Ignition ◽  
Solid Rocket Motor ◽  
Ignition System ◽  
Solid Rocket

Solid Rocket Motor Internal Ballistics Simulation Considering Complex 3D Propellant Grain Geometries

2018 ◽  
pp. 146-169
Author(s):  
Guilherme Lourenço Mejia
Keyword(s):  
Solid Propellant ◽  
Structural Integrity ◽  
Combustion Process ◽  
Burning Surface ◽  
Chamber Pressure ◽  
Computational Tool ◽  
Solid Rocket Motor ◽  
Rocket Motors ◽  
Internal Ballistics ◽  
Solid Rocket

Solid rocket motors (SRM) are extensively employed in satellite launchers, missiles and gas generators. Design considers propulsive parameters with dimensional, manufacture, thermal and structural constraints. Solid propellant geometry and computation of its burning rate are essential for the calculation of pressure and thrust vs time curves. The propellant grain geometry changes during SRM burning are also important for structural integrity and analysis. A computational tool for tracking the propagation of tridimensional interfaces and shapes is then necessary. In this sense, the objective of this work is to present the developed computational tool (named RSIM) to simulate the burning surface regression during the combustion process of a solid propellant. The SRM internal ballistics simulation is based on 3D propagation, using the level set method approach. Geometrical and thermodynamic data are used as input for the computation, while simulation results of geometry and chamber pressure versus time are presented in test cases.

Laser ignition of an experimental combustion chamber with a multi-injector configuration at low pressure conditions

2017 ◽  
Vol 9 (3) ◽  
pp. 299-311 ◽  
Author(s):  
Michael Börner ◽  
Chiara Manfletti ◽  
Gerhard Kroupa ◽  
Michael Oschwald
Keyword(s):  
Combustion Chamber ◽  
Low Pressure ◽  
Laser Ignition ◽  
Experimental Combustion

Performance Improvement of a Laser-Ignition Micro Solid Rocket by Controlling the Combustion Wave Front

2012 ◽  
Author(s):  
Hiroyuki Koizumi ◽  
Yusuke Masuda ◽  
Tomoyuki Hayashi ◽  
Kimya Komurasaki ◽  
Yoshihiro Arakawa ◽  
...  
Keyword(s):  
Wave Front ◽  
Performance Improvement ◽  
Combustion Wave ◽  
Laser Ignition ◽  
Solid Rocket
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