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1. Field of the Invention
The present invention relates generally to a turbo-machine, and more specifically to an axial flow compressor with a rotor blade having boundary layer control.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a turbo-machine, such as an axial flow compressor in a gas turbine engine, a compressor includes a row of rotor blades that compress the air or other compressible fluid. The rotor blades include a blade tip that forms a gas seal with an inner surface of a stationary shroud or casing of the turbo-machinery. A compressor blade will form a boundary layer on its surface from the compressed gas as the gas flows over the blade surface. The boundary layer is a low velocity gas on the airfoil surface that will lower the performance of the blade.
It is an object of the present invention to provide for an axial flow compressor with a rotor blade in which the boundary layer formed on the airfoil surface is significantly reduced or eliminated.
It is another object of the present invention to provide for an axial flow compressor with a rotor blade that has improved tip sealing capability.
These objectives and more are achieved in the axial flow compressor with rotor blades in which the blade tip section includes a cavity connected by an array of holes that open onto the pressure side surface of the blade to deliver gas to the cavity, and a row of blade tip holes that connect the cavity and open onto the blade tip and extend along the pressure side wall of the blade tip to discharge the air (or gas) from the cavity in a direction toward an oncoming gas flow over the blade tip. Rotation of the blade forces some of the compressed gas on the pressure side wall of the blade into the cavity and then out through the blade tip holes to reduce or eliminate the boundary layer developed around this region of the blade, and to provide for a gas flow to block oncoming compressed gases and prevent or reduce leakage across the blade tip gap.
The present invention is intended for a rotor blade in an axial flow compressor, but can also be used in a turbine rotor blade as well if the blade tip region of the blade with the cavity and the pressure side holes and blade tip holes can be included without requiring additional cooling passages for the turbine blade to provide needed cooling for the blade tip region of the turbine blade.
The array of holes 13 on the pressure sidewall in the tip region is arranged around this surface so that the compressed gas forming on this surface will flow through the holes and into the cavity 15. The size and spacing of the pressure sidewall holes 13 will depend upon the size of the blade and the composition of the compressible fluid that the blade is compressing in the turbo-machine. Also, the depth of the pressure wall side holes 13 will depend upon the diameter of each of the holes 13 and the amount of gas required to pass into the cavity 15. The blade tip holes 14 are connected to the cavity 15 and are slanted toward the pressure side wall (as opposed to the suction side wall) to block the oncoming compressed air that can pass over the blade tip and through the tip gap formed with the stationary outer shroud or blade outer air seal (BOAS). The pressurized gas discharged from the cavity through the tip holes 14 will restrict and counter the leakage of gas from the pressure side to the suction side of the blade to improve the performance of the compressor.
The cavity 15 and tip holes 14 will be charged with compressed gas from the rotation of the rotor blade 10 by allowing some of the compressed gas forming on the pressure side wall to pass through the pressure side wall holes 13 and into the cavity. The discharge pressure of tip holes 14 will be substantially lower due to acceleration of the gas flow into the clearance gap formed between the moving blade tip and stationary outer shroud. The pressure side wall holes 13 and tip holes 14 are relatively sized to maintain the pressure in cavity 15 at some desired intermediate pressure between that of the pressure side and the clearance gap. Rotation of the blade will also force the air within the cavity out through the tip holes due to high centrifugal forces developed during the blade rotation.
The blade tip with the holes 13 and 14 and cavity 15 can be used in a turbine rotor blade for the same reasons as in the compressor blade if the turbine blade does not require cooling, or if it can still be cooled in the blade tip region. Since the turbine rotor blade is typically exposed to a higher gas flow temperature than in a compressor blade, high levels of cooling might be required in the turbine blade, especially in the tip region. Later stages of turbine blade would be more acceptable for using the boundary layer control structure of the present invention because the environmental heat load is lower. First and maybe second stage turbine rotor blades of modern turbo machines are typically exposed to too high of a gas flow temperature to allow for the cavity to be filled with the hot gas flow acting on the pressure side wall surface of the rotor blade to be passed into the cavity and then through the tip holes.
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