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A gas turbine engine component comprises a shroud, a U-channel, an internal cooling air passage and a U-channel cooling hole. The shroud comprises a forward face, an aft face, a first side face and a second side face. The U-channel is disposed in the aft face of the shroud. A gas path surface connec

A gas turbine engine component comprises a shroud, a U-channel, an internal cooling air passage and a U-channel cooling hole. The shroud comprises a forward face, an aft face, a first side face and a second side face. The U-channel is disposed in the aft face of the shroud. A gas path surface connects the forward face, aft face, first side face and second side face. A cooled surface connects the forward face, aft face, first side face and second side face opposite the gas path face. The internal cooling air passage extends through the shroud. The U-channel cooling hole extends into the first side face of the shroud adjacent the U-channel to intersect the internal cooling passage.

대표청구항 ▼

1. A turbine blade comprising: an airfoil;a platform surrounding a base of the airfoil;a U-channel disposed in an aft face of the platform;a root extending from the platform opposite the airfoil;an internal cooling passage extending through the turbine blade;a U-channel cooling hole extending in a d

1. A turbine blade comprising: an airfoil;a platform surrounding a base of the airfoil;a U-channel disposed in an aft face of the platform;a root extending from the platform opposite the airfoil;an internal cooling passage extending through the turbine blade;a U-channel cooling hole extending in a downstream direction from the internal cooling passage to a mate face of the platform upstream of the U-channel;a forward cooling passage extending through the turbine blade upstream from the internal cooling passage; anda first auxiliary cooling hole extending in an upstream direction from the forward cooling passage to the mate face of the platform, wherein the first auxiliary cooling hole is upstream from the U-channel cooling hole;wherein the U-channel cooling hole and the first auxiliary cooling hole are configured to impinge cooling air onto an adjacent platform face and to provide film cooling along radially inner and outer faces of the U-channel with at least a portion of the cooling air after the portion of the cooling air has impinged on the adjacent platform face. 2. The turbine blade of claim 1 wherein the airfoil comprises: a leading edge;a trailing edge;a pressure side extending between the leading edge and the trailing edge with a predominantly concave curvature;a suction side extending between the leading edge and the trailing edge with a predominantly convex curvature; anda span extending radially from an inner diameter base to an outer diameter tip;wherein the U-channel cooling hole extends into a pressure side mate face of the platform. 3. The turbine blade of claim 1 wherein the U-channel cooling hole is positioned radially inward of a trailing edge of the airfoil. 4. The turbine blade of claim 1 wherein the platform comprises: the aft face;a forward face opposite the aft face;an upper surface defining an end wall from which the airfoil extends;a lower surface opposite the upper surface and from which the root extends;a first side face; anda second side face comprising the mate face into which the U-channel cooling hole extends. 5. The turbine blade of claim 1 wherein the U-channel comprises: a first flange comprising: a first proximate end extending from the platform; anda first distal end opposite the first proximate end;a base extending radially inward from the first proximate end; anda second flange comprising: a second proximate end extending from the base; anda second distal end opposite the second proximate end. 6. The turbine blade of claim 5 wherein the second flange comprises an angel wing seal and is longer than the first flange. 7. The turbine blade of claim 5 wherein the base is arcuate. 8. The turbine blade of claim 5 wherein the U-channel cooling hole is positioned at an apex between the base, the first flange and the second flange. 9. The turbine blade of claim 1 wherein the internal cooling channel further comprises: first and second feed channels extending through the root and joining to the forward cooling passage; andthird and fourth feed channels extending through the root and joining to the internal cooling passage. 10. The turbine blade of claim 1 wherein the U-channel cooling hole extends straight between an inlet and an outlet. 11. The turbine blade of claim 1 wherein the U-channel cooling hole extends from the internal cooling passage to the mate face of the platform with a downstream vector component. 12. The turbine blade of claim 1 and further comprising: a second auxiliary cooling hole extending from the forward cooling passage to the mate face, wherein the second auxiliary cooling hole is disposed between the first auxiliary cooling hole and the U-channel cooling hole. 13. A method for cooling a U-channel in a gas turbine engine shroud, the method comprising: flowing cooling air through an internal cooling passage of the turbine engine shroud;flowing cooling air through a forward cooling passage of the turbine engine shroud;directing a first portion of the cooling air through a U-channel cooling hole extending in a downstream direction from the internal cooling passage to a mate face of the gas turbine engine shroud upstream of the U-channel so that the first portion of the cooling air impinges on an adjacent platform face;directing a second portion of the cooling air through a first auxiliary cooling hole extending in an upstream direction from the forward cooling passage to the mate face;passing the first portion of the cooling air into the U-channel to provide film cooling to the U-channel, wherein the auxiliary cooling hole is configured such that the second portion of the cooling air augments the film cooling of the first portion of the cooling air. 14. The method of claim 13 and further comprising: forming an air dam above the U-channel with the first portion of the cooling air to prevent hot combustion gas from entering the U-channel. 15. The method of claim 13 and further including: directing a third portion of the cooling air through a second auxiliary cooling hole extending from the forward cooling passage to the mate face. 16. A gas turbine engine component comprising: a shroud comprising a forward face, an aft face, a first side face and a second side face;a U-channel disposed in the aft face of the shroud;a gas path surface connecting the forward face, aft face, first side face and second side face;a cooled surface connecting the forward face, aft face, first side face and second side face opposite the gas path face;an internal cooling air passage extending through the shroud; anda U-channel cooling hole extending in a downstream direction into the first side face of the shroud adjacent the U-channel to intersect the internal cooling passage;a forward cooling passage extending through the shroud upstream from the internal cooling passage; anda first auxiliary cooling hole extending in an upstream direction from the forward cooling passage to the first side face, wherein the first auxiliary cooling hole is upstream from the U-channel cooling hole;wherein the U-channel cooling hole has an outlet positioned at an apex of the U-channel such that cooling air discharging therefrom impinges onto an adjacent platform face and flows along radially inner and outer faces of the U-channel after impinging on the adjacent platform face. 17. The gas turbine engine component of claim 16 wherein the U-channel comprises: a first flange comprising: a first proximate end extending from the aft face of the platform; anda first distal end opposite the first proximate end;a base extending radially inward from the first proximate end; anda second flange comprising: a second proximate end extending from the base; anda second distal end opposite the second proximate end. 18. The gas turbine engine component of claim 16 wherein: an airfoil extending radially outward from the gas path surface, the airfoil having a leading edge, a trailing edge, a pressure side, a suction side, an outer diameter end and an inner diameter end; anda root extending radially inward from the cooled surface. 19. The gas turbine engine of claim 16 and further comprising: a second auxiliary cooling hole extending from the forward cooling passage to the first side face, wherein the second auxiliary cooling hole is disposed between the first auxiliary cooling hole and the U-channel cooling hole.