Plasma Clearance Controlled Compressor

Abstract

A plasma leakage flow control system for a compressor is disclosed, comprising a circumferential row of compressor blades, an annular casing surrounding the tips of the blades, located radially apart from the tips of the blades and at least one annular plasma generator located on the annular casing. The annular plasma generator comprises an inner electrode and an outer electrode separated by a dielectric material. A gas turbine engine having a plasma leakage flow control system further comprises an engine control system which controls the operation of the annular plasma generator such that the blade tip leakage flow can be changed.

Description

BACKGROUND OF THE INVENTION[0001]This invention relates generally to gas turbine engines, and, more specifically, to the enhancement of stable flow range of compression systems therein, such as fans, boosters and compressors using plasma actuators.[0002]In a turbofan aircraft gas turbine engine, air is pressurized in a fan module, a booster module and a compression module during operation. The air passing through the fan module is mostly passed into a by-pass stream and used for generating the bulk of the thrust needed for propelling an aircraft in flight. The air channeled through the booster module and compression module is mixed with fuel in a combustor and ignited, generating hot combustion gases which flow through turbine stages that extract energy therefrom for powering the fan, booster and compressor rotors. The fan, booster and compressor modules have a series of rotor stages and stator stages. The fan and booster rotors are typically driven by a low pressure turbine and the...

AI Technical Summary

Benefits of technology

[0009]An aircraft gas turbine engine may be operated using a method for operating the plasma generator system for improving the stable flow range of the compression systems in the engine. In another aspect of the invention, an aircraft gas turbine engine may be operated using a method for reducing the tip leakage flow by reducing effective clearance between the tip of the rotating blades and a casing or shroud surrounding the blade tips. An aircraft gas turbine engine may be operated using a method for operating the plasma generator system for changing the operating efficiency of a compressor.

Problems solved by technology

The conflicting requirements of stall margin and efficiency are particularly demanding in high performance jet engines that operate under operating conditions such as severe inlet distortions and increased auxiliary power extractions, while still requiring high a level of stall margin in conjunction with high compression efficiency.

Compressor system stalls are commonly caused by flow breakdown at the tip of the compressor rotor.

These leakage flows may cause vortices to form at the tip region of the blade.

The vortices may grow in intensity and size, causing blockage and loss when the compression system is throttled and may ultimately lead to a compression system stall and reduction of efficiency.

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