Self-adaption compressor for improving flow of rotor blade tip and stator angular region

Abstract

The invention relates to the technical field of aircraft gas turbines and discloses a self-adaption compressor for improving flow of a rotor blade tip and a stator angular region. The self-adaption compressor comprises a stator blade, a rotor blade and a casing with a gas introducing circulation cavity. The gas introducing circulation cavity is formed inside the casing and composed of a gas introducing circulation cavity inlet, a high-pressure gas pressure-stabilizing cavity and a main gas introducing pipe. A flexible inlet side wall hinged to the inlet can regulate and control the open degree under the pressure differential effect in a self-adaption mode. The ultimate open degree position of the inlet is restrained by a hydraulically-controlled limiting supporting rod and a check block in the ultimate position of the flexible inlet side wall. The different ultimate open degrees can be set under different work states and are regulated within the open-degree range in a self-adaption mode. Through a self-circulation gas introducing structure, leakage flow of the rotor blade tip of the compressor and separation flow of the stator angular region of the compressor are improved. Through a foreign matter separating device, foreign matter entering the compressor is removed timely, and stable operation of the compressor is guaranteed. Through the high-pressure gas pressure-stabilizing cavity, needed high-pressure gas is provided for an aircraft and an engine.

Description

technical field [0001] The invention relates to the technical field of aviation gas turbines, in particular to an adaptive gas compressor for improving the flow in the rotor blade tip and stator corner area. Background technique [0002] In the flow inside the aero-engine, the flow inside the compressor is subjected to a strong reverse pressure gradient due to the small space, which makes the flow field have a complex vortex structure; the corner separation structure located at the stator end area and the rotor blade tip The leakage flow structure of the compressor is the main secondary flow structure inside the compressor, and it is the main source of flow loss and blockage inside the compressor. Cause compressor stall and surge, bring catastrophic consequences; after decades of research, many researchers have had a deep understanding of compressor stator corner flow and rotor tip flow, but due to its Due to the limitation of space and the complexity of the flow, the effec...

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Problems solved by technology

[0002] In the flow inside the aero-engine, the flow inside the compressor is subjected to a strong reverse pressure gradient due to the small space, which makes the flow field have a complex vortex structure; the corner separation structure located at the stator end area and the rotor blade tip The leakage flow structure of the compressor is the main secondary flow structure inside the compressor, and it is the main source of flow loss and blockage inside the compressor. Cause compressor stall and surge, bring catastrophic consequences; after decades of research, many researchers have had a deep understanding of compressor stator corner flow and rotor tip flow, but due to its Due to the limitation of space and the complexity of the flow, the effective control of the flow in the compressor has not been well realized according to the existing research results; Self-circulation adjustment mechanism, while using and improving the flow structure that is not conducive to performance, to achieve the purpose of improving the performance of the compressor

[0003] The flow channel inside the compressor is a constricted flow channel. The closer to the back stage, the denser the distribution of the blades, the smaller the flow space. If the foreign matter entering the engine is not removed in time, it will cause damage to the closely arranged blades, and even cause engine damage in severe cases. In addition, during the flight of the aircraft, the air conditioning and pressurization of the cockpit, the anti-icing of the wings and the engine cowling, and the cooling of the hot-end parts of the engine all need to be compressed air. The machine stage introduces high-pressure gas, and with the change of the flight state of the aircraft, the demand for bleed air and the working state of the compressor will change. Improper bleed air will interfere with the mainstream flow field in the compressor, making the working condition of the compressor changes, performance drops dramatically

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