Patent Application
20060127212
The invention relates generally to gas turbine engines and, more particularly, to airfoil platform impingement cooling.
Gas turbine engine airfoils, such as high pressure turbine vanes, are typically cooled by compressor bleed air. Conventional turbine vanes, such as the one shown at 9 in
One disadvantage of the above vane cooling scheme is that it requires additional cooling air to purge the turbine cavity between the adjacent rows of vanes and turbine blades. Furthermore, the film cooling holes must be sufficiently long to allow the cooling air to flow from the plenum to the gas path side of the platform, which results in greater turbine vane manufacturing costs.
It is therefore an object of this invention to provide a new airfoil platform cooling system that addresses the above problems.
In one aspect, the present invention provides an airfoil for a gas turbine engine, the airfoil comprising at least a platform having a gas path side and a back side, an airfoil portion extending from the gas path side of the platform, and a plenum located on a side of the platform opposite said airfoil portion, the plenum communicating with a source of coolant, the plenum having an outlet hole extending through a wall thereof, the outlet hole having an exit facing the back side of the platform and oriented for directing the coolant thereagainst.
In another aspect, the present invention provides a turbine vane for a gas turbine engine, comprising: a platform having a gas path side, a back side opposite said gas path side, and an overhanging portion; an airfoil portion extending from said gas path side of said platform; a plenum located on the back side of the platform; and at least one impingement hole extending through a wall of the plenum and having an axis intersecting the overhanging portion of the platform for directing coolant from the plenum onto the back side of the overhanging portion.
In another aspect, the present invention provides a turbine section for a gas turbine engine, comprising a turbine nozzle adapted to direct a stream of hot combustion gases to a turbine rotor, the turbine rotor having a plurality of circumferentially distributed blades projecting radially outwardly from a rotor disk, the rotor disk having a front rotor disk cavity, the turbine nozzle comprising a plurality of vanes extending radially between inner and outer bands forming radially inner and outer boundaries for the stream of hot combustion gases, each of a plurality of said vanes having a plenum located radially inwardly of said inner band, and at least one impingement hole oriented to cause coolant in the plenum to impinge onto a radially inwardly facing surface of the inner band and then flow into the front rotor disk cavity intermediate the turbine nozzle and the turbine rotor to at least partly purge the cavity from the hot combustion gases.
In a still further general aspect, the present invention provides a method of cooling an overhanging portion of a platform of a turbine vane, comprising the steps of: a) feeding cooling air into a plenum located underneath the platform and b) causing at least part of the cooling air in the plenum to impinge onto an undersurface of the overhanging portion of the platform.
Further details of these and other aspects of the present invention will be apparent from the detailed description and figures included below.
Reference is now made to the accompanying figures depicting aspects of the present invention, in which:
The turbine section 18 typically comprises a high pressure turbine 18a and a low pressure turbine 18b downstream of the high pressure turbine 18a. As shown in
The turbine rotor 22 includes a plurality of circumferentially spaced-apart blades 24 (only one shown in
The turbine nozzle 20 includes a plurality of circumferentially spaced vanes 32 (only one shown in
The exemplary high pressure turbine vane 32 shown in
As shown in
In operation, cooling discharge air from the compressor flows into the through a cooling air circuit to plenum 46. The cooling air, as represented by arrow 59, then flow through the cooling hole 54 and impinges onto the back side 55 of the rear overhang 50. After cooling the platform overhang back side 55, the cooling air discharged from the impingement hole 54 flows into the front rotor disk cavity 52 to purge this space in order to limit ingestion of hot gases and, thus, prevent overheating of the rotor disk 26.
It can be readily appreciated that the above described cooling scheme advantageously provides for the efficient use of cooling air by allowing the same cooling air to be used for: 1) impingement cooling on the back side of the rear overhang 50 of the inner high pressure vane inner band, and 2) purging of the high pressure turbine front cavity 52 to minimizing cooling air consumption and avoid hot gas ingestion. This dual use of the cooling air provides a benefit to the overall engine aerodynamic efficiency by reducing the amount of cooling air required to cool the engine 10.
Furthermore, impingement holes 54 are shorter in length than conventional film cooling holes (0.15 inch to 0.25 inch as compared to 0.750 inch), which contributes to lower the vane manufacturing costs.
The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without department from the scope of the invention disclosed. For example, it is understood that the impingement holes could be otherwise positioned and oriented to cool other portions of the inner vane platform. Also, while the invention as been described in the context of a high pressure turbine vane inner platform, it is understood that the same principles could be applied to other gas turbine engine airfoil structures, such as turbine blades. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.
