Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2001-09-26
2003-02-04
Verdier, Christopher (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C416S001000, C416S09700R
Reexamination Certificate
active
06514037
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to cooled turbine elements for gas turbine engines, and more particularly, to a method of lowering a stress in a cooled turbine element and the element made thereby.
FIG. 1
illustrates a portion of a gas turbine engine, generally designated by the reference number
10
. The gas turbine engine
10
includes cooled turbine elements such as a high pressure turbine nozzle
12
, a high pressure turbine blade (generally designated by
14
), and a first stage low pressure turbine nozzle
16
. As illustrated in
FIG. 2
, each of these cooled elements (e.g., blade
14
) includes one or more airfoils
20
, and one or more flowpath boundary members (e.g., a blade platform, generally designated by
22
). In the case of the turbine blade
14
, the element also includes a conventional dovetail
24
for connecting the blade to a turbine disk
26
(FIG.
1
), and a shank
28
extending between the dovetail and the blade platform
22
. Interior cooling passages
30
extend from openings (not shown) at the inner end of the blade dovetail
24
to cooling holes
32
in the airfoil
20
. The passages
30
convey cooling air through the blade to remove heat from the blade. The cooling air passing through the cooling holes
32
in the airfoil
20
provides a film cooling barrier around the exterior surface of the airfoil.
Each flowpath boundary member
22
has a flowpath face
34
which faces the flowpath of the engine
10
and an outside face
36
opposite the flowpath face. As will be appreciated by those skilled in the art, the flowpath face
34
of each flowpath boundary member
22
runs hotter than the outside face
36
during engine operation. This difference in temperature results in the flowpath face
34
tending to grow more as a result of thermal growth than the outside face
36
. Because the boundary member
22
is constrained by the airfoil
20
, the tendency for the flowpath face
34
to grow more than the outside face
36
produces thermal stresses in the boundary member and the airfoil. More particularly, tensile stresses are produced in a trailing edge
38
of the airfoil
20
due to the tendency for the flowpath face
34
to grow more than the outside face
36
. Experience has shown that fatigue cracks form and propagate as a result of the tensile stresses in the trailing edge
38
of the airfoil
20
, resulting in a shortened life of the blade
14
. Thus, there is a need for a method of lowering these stresses in colled turbine elements.
SUMMARY OF THE INVENTION
Briefly, apparatus of this invention is a cool turbine element for use in a flowpath of a gas turbine engine. The element comprises an airfoil having a pressure side and a suction side opposite the pressure side. The pressure side and the suction side extend axially between a leading edge and a trailing edge opposite the leading edge and radially between an inboard end and an outboard end opposite the inboard end. Further, the element comprises a flowpath boundary member extending laterally from at least one of the inboard end and the outboard end. The boundary member has a flowpath face and an outside face opposite the flowpath face. The outside face runs cooler than the flowpath face during engine operation thereby creating a tendency for the member to deflect in a direction away from the flowpath face and causing a thermally induced tensile radial stress in a region of the trailing edge of the airfoil. In addition, the element comprises an interior cooling passage extending through the airfoil from a cooling air source for transporting cooling air through the airfoil and at least one cooling hole extending from the interior cooling passage to an opening located on one of the suction side and the pressure side in an area upstream from the stressed region of the trailing edge to cool the area to a temperature below that of the trailing edge so that the airfoil thermally deflects during engine operation to a shape corresponding to that of the flowpath boundary member thereby lowering the thermally induced tensile radial stress in the airfoil at the trailing edge thereof.
In another aspect, the invention includes a method of lowering a tensile stress at a trailing edge of an airfoil of a cooled blade adjacent a platform of the blade. The method comprises the step of forming at least one cooling hole in the airfoil from an interior cooling air passage to an exterior surface of the airfoil to deliver cooling air to the exterior surface to cool an area of the exterior surface immediately adjacent the cooling hole thereby shifting tensile thermal loading from regions of the airfoil adjacent the area of the exterior surface to the cooled area.
In yet another aspect, the present invention includes a method of lowering a thermal stress at a trailing edge of an airfoil of a cooled turbine blade adjacent a platform of the blade. The method comprises the step of forming at least one cooling hole positioned upstream from the trailing edge of the airfoil and extending from an interior cooling air passage to an exterior surface of the airfoil for delivering cooling air to the exterior surface to cool the airfoil in an area of the exterior surface upstream from the trailing edge so that a thermal deflection of the airfoil more closely corresponds to a thermal deflection of the platform thereby lowering thermally induced stresses in the airfoil at the trailing edge thereof.
Other features of the present invention will be in part apparent and in part pointed out hereinafter.
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Brainch Gulcharan Singh
Danowski Michael Joseph
General Electric Company
Sonnenschein Nath & Rosenthal
Verdier Christopher
Young Rodney M.
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