Patent Application
20050276697
This invention relates generally to gas turbine engines and more particularly, to methods and apparatus for cooling gas turbine engine rotor assemblies.
At least some known rotor assemblies include at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side, and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. Known blades are hollow such that an internal cooling cavity is defined at least partially by the airfoil, platform, shank, and dovetail.
During operation, portions of the airfoil of the blades are exposed to higher temperatures than other portions of the blades. Over time, such temperature differences and thermal strain may induce thermal stresses in the blade. Such thermal strains may induce thermal deformations to the airfoil, for example, local creep deflection, and may cause other problems such as airfoil low-cycle fatigue, which may shorten the useful life of the rotor blade.
To facilitate reducing the effects of high temperatures within at least some known rotor blades, at least some of the rotor blade airfoils include a trailing edge slot and a cut back pressure-side wall with the slot divided into evenly spaced channels which discharge a film of cooling air over the exposed back surface of the airfoil. However, because of temperature differences at different points along the trailing edge, the air from the evenly spaced slots do not cool the trailing edge enough to remove the temperature differential between different points along the trailing edge of the airfoil.
In one aspect, an airfoil for a gas turbine is provided. The airfoil includes a leading edge, a trailing edge, a tip plate, a first sidewall extending in radial span between an airfoil root and the tip plate, and a second sidewall connected to the first sidewall at the leading edge and the trailing edge to define a cooling cavity therein. The sidewall extends in radial span between the airfoil root and the tip plate. The airfoil also includes a plurality of longitudinally spaced apart trailing edge cooling slots arranged in a column extending through the first sidewall. The slots are in flow communication with the cooling cavity and arranged in a non-uniform distribution along the trailing edge so that the number of slots in at least one portion of the trailing edge is greater than a different portion of the trailing edge.
In another aspect, a turbine blade is provided. The turbine blade includes a platform, a dovetail, a shank connected to the platform and the dovetail, and an airfoil comprising a leading edge, a trailing edge, a pressure sidewall, and a suction sidewall. The airfoil is connected to the platform. The turbine blade also includes at least one cooling cavity between the pressure sidewall and the suction sidewall, and a plurality of longitudinally spaced apart trailing edge cooling slots extending along the trailing edge. The trailing edge cooling slots are in flow communication with the cooling cavity and are arranged in a non-uniform distribution along the trailing edge so that the number of trailing edge cooling slots in at least one portion of the trailing edge is greater than a different portion of the trailing edge.
In another aspect, a rotor assembly for a gas turbine is provided. The rotor assembly includes a rotor shaft and a plurality of circumferentially-spaced rotor blades coupled to the rotor shaft. Each rotor blade includes a platform, a dovetail, a shank connected to the platform and the dovetail, and an airfoil comprising a leading edge, a trailing edge, a pressure sidewall, and a suction sidewall. The airfoil is connected to the platform. The turbine blade also includes at least one cooling cavity between the pressure sidewall and the suction sidewall, and a plurality of longitudinally spaced apart trailing edge cooling slots extending along the trailing edge. The trailing edge cooling slots are in flow communication with the cooling cavity and are arranged in a non-uniform distribution along the trailing edge so that the number of trailing edge cooling slots in at least one portion of the trailing edge is greater than a different portion of the trailing edge.
In another aspect, a method of cooling a trailing edge of a rotor blade airfoil is provided. The airfoil includes a leading edge, a trailing edge, a pressure sidewall and a suction sidewall, at least one cooling cavity between the pressure sidewall and the suction sidewall, and a plurality of longitudinally spaced apart trailing edge cooling slots extending along the trailing edge. The trailing edge cooling slots are in flow communication with the cooling cavity and arranged in a non-uniform distribution along the trailing edge so that the number of trailing edge cooling slots in at least one portion of the trailing edge is greater than a different portion of the trailing edge. The method includes providing cooling air to the cooling cavity, and directing a portion of the cooling air through the plurality of cooling slots.
An airfoil for a gas turbine rotor blade that includes a plurality of longitudinally spaced apart trailing edge cooling slots arranged in a column are described in detail below. The cooling slots are arranged in a non-uniform distribution along the trailing edge so that the number of slots in at least one portion of the trailing edge is greater than a different portion of the trailing edge. The non-uniform cooling slot distribution permits the cooling air to be directed to the portions of the trailing edge that are exposed to the hottest external temperatures to improve the cooling of these areas. The improved cooling of the trailing edge alleviate possible local creep, possible oxidation and possible low-cycle fatigue of the airfoil.
Referring to the drawings,
In operation, ambient air is channeled into compressor section 22 where the ambient air is compressed to a pressure greater than the ambient air. The compressed air is then channeled into combustor section 24 where the compressed air and a fuel are combined to produce a relatively high-pressure, high-velocity gas. Turbine section 28 is configured to extract the energy from the high-pressure, high-velocity gas flowing from combustor section 24. Gas turbine system 10 is typically controlled, via various control parameters, from an automated and/or electronic control system (not shown) that is attached to gas turbine system 10.
Airfoil 42 includes a first sidewall 44 and a second sidewall 46. First sidewall 44 is convex and defines a suction side of airfoil 42, and second sidewall 46 is concave and defines a pressure side of airfoil 42. Sidewalls 44 and 46 are connected at a leading edge 48 and at an axially-spaced trailing edge 50 of airfoil 42 that is downstream from leading edge 48.
First and second sidewalls 44 and 46, respectively, extend longitudinally or radially outward to span from a blade root 52 positioned adjacent dovetail 43 to a tip plate 54 which defines a radially outer boundary of an internal cooling chamber 56. Cooling chamber 56 is defined within airfoil 42 between sidewalls 44 and 46. Internal cooling of airfoils 42 is known in the art. In the exemplary embodiment, cooling chamber 56 includes a serpentine passage 58 cooled with compressor bleed air.
Cooling cavity 56 is in flow communication with a plurality of trailing edge slots 70 which extend longitudinally (axially) along trailing edge 50. Particularly, trailing edge slots 70 extend along pressure side wall 46 to trailing edge 50. Each trailing edge slot 70 includes a recessed wall 72 separated from pressure side wall 46 by a first sidewall 74 and a second sidewall 76. A cooling cavity exit opening 78 extends from cooling cavity 56 to each trailing edge slot 70 adjacent recessed wall 72. Each recessed wall 72 extends from trailing edge 50 to cooling cavity exit opening 78. A plurality of lands 80 separate each trailing edge slot 70 from an adjacent trailing edge slot 70. Sidewalls 74 and 76 extend from lands 80.
Trailing edge slots 70 are arranged in a non-uniform distribution along trailing edge 50 so that the number of slots 70 in a first portion 82 of trailing edge 50 is greater than a second portion 84 of trailing edge 50. Particularly, a distance between trailing edge cooling slots 70 located in first portion 82 of trailing edge 50 is different than a distance between trailing edge slots 70 cooling located in second portion 84 of trailing edge 50. Specifically, the number of trailing edge cooling slots 70 per inch of first portion 82 of trailing edge 50 is greater than the number of trailing edge cooling slots 70 per inch of second portion 84 of trailing edge 50. Also, the number of trailing edge slots 70 in first portion 82 of trailing edge 50 is greater than the number of trailing edge cooling slots 70 in a third portion 86 of trailing edge 50. The exemplary embodiment of airfoil 42 shown in
The non-uniform cooling slot distribution permits the cooling air to be directed to the portions of trailing edge 50 that are exposed to the hottest external temperatures to improve the cooling of these areas. The improved cooling of trailing edge 50 alleviate possible local creep, possible oxidation and possible low-cycle fatigue of the airfoil.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.
