Compressor endwall bleed system

Details Compressor endwall bleed system Compressor endwall bleed system

2000-03-16

2002-08-06

Kwon, John (Department: 3754)

Rotary kinetic fluid motors or pumps

Including means for handling portion separated from working...

C415S914000

Reexamination Certificate

active

06428271

ABSTRACT:

BACKGROUND OF THE INVENTION
The present invention relates generally to the field of gas turbine engines and more particularly in one embodiment the present invention defines an endwall bleed system to remove a separated boundary layer downstream of a rotating compressor tip shroud. Although the present invention was developed for a gas turbine engine, certain applications may be outside of this field.
A gas turbine engine is typical of the type of turbomachinery in which the present inventions described herein may be advantageously employed. It is well known that a gas turbine engine conventionally comprises a compressor for compressing the inlet air to an increased pressure for combustion in a combustor chamber. The mixture of fuel and the increased pressure air is burned in the combustor chamber to generate a high temperature gaseous flow stream for causing rotation of the turbine blades within a turbine. Further, the high temperature gaseous flow stream may be used directly as a thrust for providing motive power such as in a turbine jet engine.
A gas turbine engine including a shrouded compressor rotor is believed well known to gas turbine engine designers. Shrouded compressor rotors can be likened to a compressor rotor with no tip clearance. Tip clearance is generally defined as a space between the tip of the compressor blade and an opposing wall member. It has been shown that a rotor with no tip clearance has poorer performance than a rotor with a small amount of tip clearance, such as a tip clearance equal to about one percent of the compressor blade span. In many prior compressor systems utilizing a shrouded rotor, a large three dimensional boundary layer separation occurs at the tip endwall and suction surface corner. This large boundary layer separation dominates the losses and stall inception mechanism for the compressor rotor.
Heretofore, there has been a need for a method and apparatus for removing at least a portion of the separated boundary layer downstream of a rotating compressor tip shroud. The present invention satisfies this and/or other needs in a novel and unobvious way.
SUMMARY OF THE INVENTION
One form of the present invention contemplates an endwall bleed system to remove at least a portion of a separated boundary layer downstream of a rotating compression system tip shroud.
Another form of the present invention contemplates a method for bleeding off a separated boundary layer from a rotating compression system tip shroud.
Yet another form of the present invention contemplates an endwall bleed system to remove a separated boundary layer downstream of a compression system tip shroud and relieve the back pressure associated therewith.
One aspect of the present invention contemplates a compression system for a gas turbine engine. The compression system, comprising: a mechanical housing; a wheel rotatable within the housing and having a plurality of blades coupled thereto; a shroud coupled to the plurality of blades so as to separate the fluid flow within the compression system into a core stream and a bypass stream, the shroud having a fore edge and an aft edge; a static wall member coupled to the mechanical housing and having a portion aligned with the aft edge of the shroud; and at least one bleed aperture formed in the wall member adjacent the aft edge of the shroud to allow the passage of fluid into the bypass system.
One object of the present invention is to provide a unique endwall bleed system for a gas turbine engine compression system.
These and other objects will become more apparent from the following description of the preferred embodiment.


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