Fluid reaction surfaces (i.e. – impellers) – Rotor having flow confining or deflecting web – shroud or... – Axially extending shroud ring or casing
Reexamination Certificate
2002-07-31
2004-11-02
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Rotor having flow confining or deflecting web, shroud or...
Axially extending shroud ring or casing
C416S09700R
Reexamination Certificate
active
06811378
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates to a blade for a gas turbine, and more specifically, to the cooling of a gas turbine blade shroud.
A gas turbine is typically comprised of a compressor section, a combustor section and a turbine section. The compressor section produces compressed air. Then fuel is mixed with some of the compressed air and burned in the combustor section. The compressed, high temperature gas produced in the combustor section is then expanded through rows of stationary vanes and rotating blades in the turbine section to produce power in the form of a rotating shaft.
Each of the rotating blades has an airfoil portion and a root portion that connects it to a rotor. Since the blades are exposed to the compressed, hot gas discharging from the combustor section, the turbine blades must be cooled to prevent failure. Usually this cooling is done by taking a portion of the compressed air produced by the compressor and using it as cooling air in the turbine section to cool turbine blades. The cooling air enters each cooled turbine blade through its root, and flows through radial passageways in the airfoil portion of the blades. While in many cooled turbine blades, the radial passageways discharge the cooling air radially outward at the blade tip, some turbine blades incorporate shrouds that project outwardly from the airfoil at the blade tip. These shrouds prevent hot gas leakage past the blade tips, and may also be used to dampen blade vibration that tends to occur during normal operation of gas turbine engines. Unfortunately, excessive creep and creep failures can occur in blade shrouds due to the high operating temperatures.
While the known methods of cooling turbine blades are generally successful at cooling the airfoil portions of turbine blades, designs for cooling shrouds have produced mixed results. In some designs, cooling air discharged from the radial passages at the blade tip flows over the radially outward facing surface of the shroud. Although this provides some cooling, it is often insufficient to adequately cool the shroud due to heating of the cooling air in the airfoil passageways.
Another design includes incorporating cooling passages into each shroud, with the cooling passages extending approximately parallel to the radially inward facing surface of the shroud. These passages, which connect to one or more of the radial passageways, divert cooling air from the airfoil passageways so that it flows through the cooling passages in the shroud, thereby lowering the operating temperature of the shroud. While this method of internally cooling the shroud is generally more effective than flowing cooling air over the radially outward facing surface of the shroud, the heat transfer rate from the shroud to the cooling air in the passages may be insufficient to prevent excessive creep at certain operating conditions.
What is needed is a turbine blade having a shroud that is sufficiently cooled to prevent excessive creep at all engine operating conditions.
SUMMARY AND OBJECTS OF THE INVENTION
It is therefore an object of the present invention to provide a turbine blade having a shroud that is sufficiently cooled at all engine operating conditions to prevent the excessive creep that can occur in turbine shrouds when turbine blades are exposed to high stress and very high operating temperatures.
According to the preferred embodiment of the present invention, a turbine blade is disclosed having a root portion with a cooling fluid cavity therein, a platform connected to the root portion, an airfoil portion extending from the platform, the airfoil portion includes at least one cooling passageway extending substantially radially through the airfoil, and at least one cooling hole extending substantially radially through the airfoil, with the one cooling passageway and the cooling hole each defined by an inner wall having an inlet for receiving a flow of cooling fluid from the cavity. The turbine blade further includes a shroud projecting outwardly from the airfoil and has a radially inward facing surface, a radially outward facing surface, and a shroud edge extending therebetween, at least one cooling fluid outlet adjacent the edge, and at least one cooling passage between the radially inward facing surface and the radially outward facing surface. The cooling passage is approximately parallel to the radially inward facing surface, and a tube is located within the cooling hole. The tube has an outer wall, a first end adjacent the inlet and a second end radially outward therefrom. The cooling passage communicates with the inlet through the tube, and standoff means between the inner wall of the cooling passageway and the outer wall of the tube maintain the inner wall of said cooling passageway in spaced relation to said outer wall of the tube to minimize heat transfer between the airfoil and the tube.
The above, and other objects, features and advantages of the present invention will become apparent from the following description read in conjunction with the accompanying drawings.
REFERENCES:
patent: 3623825 (1971-11-01), Schneider
patent: 3747182 (1973-07-01), Hayes
patent: 3846041 (1974-11-01), Albani
patent: 5259730 (1993-11-01), Damlis et al.
patent: 5941687 (1999-08-01), Tubbs
patent: 6099253 (2000-08-01), Fukue et al.
patent: 6152695 (2000-11-01), Fukue et al.
patent: 6193465 (2001-02-01), Liotta et al.
Look Edward K.
Mack Brian R.
McAleenan James M.
Power Systems Mfg, LLC
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