A two-stage pre-film layered partly premixed high-temperature rise combustor structure

Abstract

The invention discloses a two-stage prefilming delamination part premixing high-temperature-rise combustion chamber structure. The flame of a main combustion region in a combustion chamber is divided into a precombustion stage and a main combustion stage which respectively adopt diffusion combustion and fuel oil direct spray combustion, and the diffusion combustion and the fuel oil direct spray combustion are coupled for realizing the delamination part premixing combustion. Both main combustion stage fuel oil and precombustion stage fuel oil adopts an in-passage prefilming air atomization technology, the precombustion stage utilizes a rotational flow and flange stabilizing device for enlarging the stable work range of an engine in the small-thrust state and providing a stable ignition source for the two-stage prefilming delamination part premixing high-temperature-rise combustion chamber under the great-thrust state of the engine, and the main combustion flame utilizes rotational flow and precombustion stage flame for providing high-temperature fuel gas stable flame. A precombustion stage spray nozzle improves the atomization quality through two-stage rotational flow prefilming air atomization, and a main combustion stage spray nozzle realizes fast uniform blending mixing of oil and gas in a way of separating the fuel oil atomization process from the blending and mixing process.

Description

technical field [0001] The invention relates to the field of aero-engines, in particular to a layered partly premixed high-temperature rise combustor structure utilizing atomization and step-by-step pre-film air atomization and rapid oil-gas mixing technology to realize uniform combustion in the main combustion zone. Background technique [0002] High temperature rise combustors are required for military aero engines. In order to meet the aerodynamic and thermal properties of the next-generation military aeroengine with a thrust-to-weight ratio of 12-15, the turbine inlet temperature of the military aeroengine should be significantly higher than that of the military aeroengine currently in service. Therefore, the design of combustion chamber components will develop in the direction of high temperature and high heat capacity combustion. The temperature rise of the combustion chamber of an active military aero-engine is about 900°C, and the next-generation military aviation t...

AI Technical Summary

Problems solved by technology

However, there are combustion instability problems such as thermoacoustic oscillation in lean premixed combustion. With the increase of the inlet temperature of the combustion chamber, the combustion instability is strengthened, and it is not suitable for high-temperature rise combustion chambers with higher pressure ratios.

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