Patent Application
20070189896
This invention relates generally to gas turbine engines and more particularly, to methods and apparatus for cooling gas turbine engine rotor assemblies.
A typical gas turbine engine includes a rotor assembly having circumferentially-spaced rotor blades. Each rotor blade, sometimes referred to as a bucket, includes an airfoil that extends radially outward from a 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.
With respect to gas turbine operation, increasing inlet firing temperatures provides improved output and engine efficiencies. Increasing the inlet firing temperature results in increased gas path temperatures. Such increased gas path temperatures can result in added stress to the bucket platforms, including possibly oxidation, creep and cracking. Further, in gas turbines where closed loop cooling circuits are used in upstream airfoil components, there is no film cooling and therefore the downstream bucket platforms do not have the benefit from the film carryover from the upstream airfoils. This exacerbates the potential distress on the bucket platforms.
Some recent known turbine blade configurations do utilize film cooling for cooling the blade platform. Specifically, compressor discharge air is routed through an opening or openings in the platform, and a layer of cooling film forms on the platform to protect the platform from the high flow path temperatures. With such film cooling, however, there may only be sufficient pressure to film cool the aft section of the platform where the flow path air has been accelerated to drop the local static pressure.
In one aspect, a method for cooling a platform of a turbine blade is provided. The turbine blade has an airfoil connected to the platform and a dovetail extending from the platform. A main cooling circuit extends through the dovetail and into the airfoil. The main cooling circuit includes an exit for main cooling flow from the airfoil to exit out through the dovetail. The method includes the steps of extracting a portion of the coolant flowing through the main cooling circuit into a platform cooling circuit, and then returning the coolant from the platform cooling circuit back into the main cooling circuit to flow through the exit.
In another aspect, a turbine blade is provided. The turbine blade includes a platform, a dovetail and an airfoil having a leading edge, a trailing edge, a pressure sidewall, and a suction sidewall. The airfoil is connected to the platform. The turbine blade further includes a main cooling circuit extending through the dovetail and into the airfoil. The main cooling circuit includes an exit for main cooling flow from the airfoil to exit out through the dovetail. The turbine blade also includes a platform cooling circuit in flow communication with the main cooling circuit. The platform circuit includes an inlet for extracting a portion of coolant flowing through the main cooling circuit into the platform circuit, and an outlet through which coolant exits the platform cooling circuit.
In yet 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 and an airfoil having a leading edge, a trailing edge, a pressure sidewall, and a suction sidewall. The airfoil is connected to the platform. The turbine blade further includes a main cooling circuit extending through the dovetail and into the airfoil. The main cooling circuit includes an exit for main cooling flow from the airfoil to exit out through the dovetail. The turbine blade also includes a platform cooling circuit in flow communication with the main cooling circuit. The platform circuit includes an inlet for extracting a portion of coolant flowing through the main cooling circuit into the platform circuit, and an outlet through which coolant exits the platform cooling circuit.
Generally, and as set forth below in more detail, a rotor blade includes a main cooling circuit. The main cooling circuit extends through the dovetail and into the airfoil. Such main cooling circuit then extends from the airfoil back through the dovetail. In one embodiment, rotor blade platform cooling is provided by extracting a portion of coolant flow supplied to the airfoil from the main cooling circuit and running the coolant through a serpentine passage, or platform circuit, in the platform to convectively cool the platform. A portion of the platform serpentine cooling flow is bled off the platform circuit to feed an airfoil cooling circuit in the airfoil which cools a portion of the airfoil, and such coolant flow is then rejoined with the main airfoil cooling flow. The remainder of the platform serpentine coolant flow continues to convectively cool the bucket platform, and is then returned to the main cooling circuit and flows to an exit.
In the one embodiment, the platform serpentine cooling circuit is a cast-in feature integral with the platform. Alternatively, such circuit is partially cast with an attached cover plate to secure to the platform. To augment the heat transfer from the platform to the coolant, turbulators can be used in the circuit. Such platform cooling circuit can be used in connection with a closed loop steam cooled bucket as well as with an air-cooled bucket
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 top plate 54 which defines a radially outer boundary of an internal cooling circuit or chamber 56. Cooling circuit 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 circuit 56 includes a serpentine passage cooled with compressor bleed air or steam.
In one embodiment, rotor blade platform cooling is provided by extracting a portion of coolant flow supplied to the airfoil from main cooling circuit 62 and running the coolant through a serpentine passage, or platform circuit 64, in platform 66 to convectively cool platform 66. A portion of the platform serpentine cooling flow is bled off platform circuit 64 to feed an airfoil cooling circuit 68 in airfoil 42 which cools a portion of airfoil 42, and such coolant flow is then rejoined with the main airfoil cooling flow. The remainder of the platform serpentine coolant flow continues to convectively cool bucket platform 66, and is then returned to the main cooling circuit 66 and flows through main cooling circuit exit 70.
In the one embodiment, the platform serpentine cooling circuit is a cast-in feature integral with the platform. Specifically, the circuit can be formed using ceramic cores or using a wax in a lost wax casting process. In the lost wax casting process, a plate typically would be welded or brazed to the platform to totally enclose the circuit within the platform. To augment the heat transfer from the platform to the coolant, turbulators can be used in the circuit. Such platform cooling circuit can be used in connection with a closed loop steam cooled bucket as well as with an air-cooled bucket.
The above described platform cooling facilitates operating a gas turbine with increased inlet firing temperatures so that improved output and engine efficiencies that can be gained with such increased inlet firing temperatures without added stress to the bucket platforms. In addition, such platform cooling facilitates cooling the entire platform and not just aft sections of the platform, such as with film cooling under certain operating conditions.
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.
