Thermoacoustic Stability Boundaries in the Model Combustion Chamber with a Gas-Centered Swirl Coaxial Injector

Yong Ho Min,Youngbin Yoon,Chae Hoon Sohn

doi:10.1155/2019/9232739

Abstract

The stability of a combustion chamber with a gas-centered swirl coaxial injector is investigated over a wide range of operating conditions in the aspect of thermoacoustic instabilities. First, flame shapes induced by the injector are analyzed for various recess lengths from experimental results. The spreading angle and flame size are reduced as the recess length increases. And, as a stability criterion, the damping factor is introduced. The operating conditions are divided into 7 sets with 9 tests, and fuel mass flow is fixed in each set. Stability boundaries are identified in terms of damping factor for the 63 tests by experiments without external perturbations. The stability map for the self-excited instability is obtained and it reveals two unstable regions. One is located at low dynamic head ratio and the other one is at high dynamic head ratio. A stable region is found at moderate dynamic head ratios. Relative stability is evaluated by various flow parameters over a wide range of operating conditions in the combustion chamber with a gas-centered swirl coaxial injector.

Highlights

The combustion chambers for rocket engines and aeroengines are designed with severe criteria
From the previous work [20], it has been found that the spray pattern varies depending on the recess length in numerical simulations
As the recess length increases, the flame is stabilized and spreading angle of the jets decreases for the same condition

Summary

Introduction

The combustion chambers for rocket engines and aeroengines are designed with severe criteria. The chamber should bear high mass flow, higher temperature, and higher combustion efficiency as well as more flexibility regarding operating conditions These constraints have to be taken into account with the potential risk of thermoacoustic instabilities [1]. A gas-centered swirl coaxial (GCSC) injector has been adopted widely in high-performance rocket engines, especially, with a staged combustion cycle In this cycle, oxidizer-rich burning takes place in a preburner or a gas generator, and as a result, a small amount of product gas is produced. The GCSC injector is an important device for highperformance propulsion, there is little information on its stability characteristics In this regard, combustion stability is investigated over a wide range of fuel and oxidizer mass flows in a chamber with the GCSC injector in this study. Stability boundaries or unstable regions for self-excited instability are found in terms of the parameters of damping factor and stability maps are demonstrated

Experimental Methods

Results and Discussion

Design point