Abstract: A method of cooling a vane of a turbine is provided. The turbine includes an airfoil, an outer shroud disposed at an outer radial end of the airfoil and an inner shroud, the airfoil including a plurality of air channels extending along the radial direction of the turbine, the air channels comprising a first air channel and a second air channel. A cooling air is caused to flow inside the first air channel to cool the first air channel, then cool one of the outer shroud and the inner shroud. A cooling air is caused to flow inside the second air channel to cool the second air channel, then cool the other one of the outer shroud and the inner shroud.

Abstract: A method of cooling a vane of a turbine is provided. The turbine includes an airfoil, a shroud disposed at an end of the airfoil, the end being a radial end along a radial direction of the turbine, the shroud comprising a shroud main body and a shroud edge disposed on a circumference of the shroud main body to surround the shroud main body, the shroud edge comprising a shroud edge passage therein. A cooling air is caused to flow inside the shroud edge passage to cool the shroud edge, and after cooling the shroud edge, the shroud main body is cooled by using the cooling air which has flowed inside the shroud edge passage.

Abstract: An apparatus and method relating to an airfoil for a turbine engine and cooling channels within the airfoil. The airfoil includes an outer wall defining an interior bound by a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. Further, the airfoil includes ribs located within the interior and a flow enhancer extending between the ribs.

Abstract: An apparatus and method relating to an airfoil for a turbine engine and cooling channels within the airfoil. The airfoil includes an outer wall defining an interior bound by a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. Further, the airfoil includes ribs located within the interior and a flow enhancer extending between the ribs.

Abstract: A turbine nozzle for a gas turbine engine includes: (a) spaced-apart arcuate inner and outer bands; (b) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the interior of the vane defining at least a forward cavity and a mid-cavity positioned aft of the forward cavity; (c) a hollow impingement insert received inside the mid-cavity, the impingement insert having walls which are pierced with at least one impingement cooling hole; (d) a passage in the turbine vane at a radially outer end of the forward cavity adapted to be coupled to a source of cooling air; and (e) a passage in the inner band in fluid communication with a radially inner end of the forward cavity and a radially inner end of the impingement insert.

Abstract: A cooling arrangement for a gas turbine engine includes: (a) a turbine nozzle having: (i) spaced-apart arcuate inner and outer bands; and (ii) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the vane disposed in a primary flowpath of the engine; (b) a supporting structure coupled to the outer band such that an outer band cavity is defined between the outer band and the stationary structure; (c) a first conduit passing through the outer band cavity and communicating with the interior of the vane, the first conduit coupled to a first source of cooling air within the engine; and (d) a second conduit communicating with the outer band cavity, the second conduit coupled to a second source of cooling air within the engine.

Abstract: A cooling arrangement for a gas turbine engine includes: (a) a turbine nozzle having: (i) spaced-apart arcuate inner and outer bands; and (ii) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the vane disposed in a primary flowpath of the engine; (b) a supporting structure coupled to the outer band such that an outer band cavity is defined between the outer band and the stationary structure; (c) a first conduit passing through the outer band cavity and communicating with the interior of the vane, the first conduit coupled to a first source of cooling air within the engine; and (d) a second conduit communicating with the outer band cavity, the second conduit coupled to a second source of cooling air within the engine.

Abstract: A turbine nozzle for a gas turbine engine includes: (a) spaced-apart arcuate inner and outer bands; (b) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the interior of the vane defining at least a forward cavity and a mid-cavity positioned aft of the forward cavity; (c) a hollow impingement insert received inside the mid-cavity, the impingement insert having walls which are pierced with at least one impingement cooling hole; (d) a passage in the turbine vane at a radially outer end of the forward cavity adapted to be coupled to a source of cooling air; and (e) a passage in the inner band in fluid communication with a radially inner end of the forward cavity and a radially inner end of the impingement insert.

Abstract: A turbine nozzle segment includes: (a) an arcuate outer band segment; (b) a hollow, airfoil-shaped turbine vane extending radially inward from the outer band segment; (c) a manifold cover secured to the outer band such that the manifold cover and the outer band segment cooperatively define an impingement cavity; and (d) an impingement blanket disposed in the impingement cavity, the impingement blanket having at least one impingement hole formed therethrough which is arranged to direct cooling air at the outer band segment. A method is provided for impingement cooling the outer band segment.

Abstract: A turbine nozzle includes a mid vane mounted between a pair of end vanes in outer and inner bands. The mid vane includes a first pattern of film cooling holes configured to discharge more cooling air than each of the two end vanes having respective second patterns of film cooling holes.

Abstract: A turbine airfoil includes an internal cooling circuit between opposite pressure and suction sidewalls. The pressure sidewall includes a row of outlet slots for the circuit separated by converging partitions. Each slot includes an inlet, throat, and diverging outlet. Each of the partitions includes a step in a converging sidewall thereof for increasing divergence of the slots to diffuse coolant discharged therethrough.

Abstract: Critical stress in a gas turbine can be estimated using one or more readily measurable temperatures in the gas turbine. First and second critical temperatures can be estimated based on the at least one measurable temperature using heat conduction and convection equations. Subsequently, the critical stress can be estimated in real time according to a stress model prediction based on the difference between the critical temperatures, and possibly the rotational speed of the turbine, and some parameter, such as air pressure, that is indicative of air flow around the turbine component. Operation of the gas turbine can thus be controlled using the estimated critical temperatures.

Abstract: Critical stress in a gas turbine can be estimated using one or more readily measurable temperatures in the gas turbine. First and second critical temperatures can be estimated based on the at least one measurable temperature using heat conduction and convection equations. Subsequently, the critical stress can be estimated in real time according to a stress model prediction based on the difference between the critical temperatures, and possibly the rotational speed of the turbine, and some parameter, such as air pressure, that is indicative of air flow around the turbine component. Operation of the gas turbine can thus be controlled using the estimated critical temperatures.

Abstract: A turbine blade includes an airfoil having a leading edge flow chamber disposed behind a leading edge in front of a pair of side channels. The leading edge channel receives coolant from the pressure side channel in isolation from the suction side channel.

Abstract: A turbine blade includes an airfoil having a leading edge flow chamber disposed behind a leading edge in front of a pair of side channels. The leading edge channel receives coolant from the pressure side channel in isolation from the suction side channel.