Turbine cooling air sealing
A turbine engine section has a stationary vane stage and a rotating blade stage. The blade stage is spaced from the vane stage to form an annular chamber therebetween. A manifold delivers a pressurized fluid to the chamber. An outer diameter (OD) sealing system restricts leakage of the pressurized fluid from the chamber. An inner diameter (ID) sealing system restricts leakage of the pressurized fluid from the chamber. A flow guide extends radially between the inner diameter sealing system and the manifold. The flow guide bisects the chamber to form a stationary chamber portion and a rotating chamber portion, the rotating chamber portion at least partially along a disk of the first blade stage.
1. A turbine engine section comprising:
a stationary vane stage;
a rotating blade stage, said blade stage spaced from said vane stage to form an annular chamber therebetween;
a manifold for delivering a pressurized fluid to the chamber;
an outer diameter sealing means for restricting leakage of the pressurized fluid from said chamber;
an inner diameter sealing means for restricting leakage of the pressurized fluid from said chamber, the inner diameter sealing means comprising an abradable seal material carried by a seal carrier; and
a flow guide of generally forwardly open, C-shaped section extending radially between said inner diameter sealing means and said manifold, said flow guide bisecting said chamber to form a stationary chamber portion and a rotating chamber portion, the rotating chamber portion at least partially along a disk of the first blade stage, said flow guide mounted to said seal carrier and generally closing off the stationary chamber portion so that the surfaces bounding the stationary chamber portion are essentially non-rotating with the stationary chamber portion bounded by the manifold forward and outboard of the stationary chamber portion, the flow guide aft of the stationary chamber portion, and the seal carrier forward and inboard of the stationary chamber portion.
2. The engine section of claim 1 wherein said outer diameter sealing means comprises a cover plate attached to said disk.
3. The engine section of claim 2 wherein:
the flow guide is non-structural and formed of sheet metal.
4. The engine section of claim 2 wherein said flow guide is affixed to said manifold with a first bolted joint and to said seal carrier with a second bolted joint.
5. The engine section of claim 4 wherein:
the flow guide is non-structural and formed of sheet metal.
6. The engine section of claim 1 wherein the manifold is affixed to the vane stage.
7. The turbine engine section of claim 1 wherein the rotating chamber portion is at least partially along a web of the disk.
8. The engine section of claim 1 wherein:
the flow guide is non-structural and formed of sheet metal.
9. A turbine engine section comprising:
a stationary vane stage;
a rotating blade stage, said blade stage spaced from said vane stage to form an annular chamber therebetween;
a manifold for delivering a pressurized fluid to the chamber;
an outer diameter sealing means for restricting leakage of the pressurized fluid from said chamber;
an inner diameter sealing means for restricting leakage of the pressurized fluid from said chamber, the inner diameter sealing means comprising an abradable seal material carried by a seal carrier; and
a flow guide of generally forwardly open, C-shaped section extending radially between said inner diameter sealing means and said manifold, said flow guide bisecting said chamber to form a stationary chamber portion and a rotating chamber portion, the rotating chamber portion at least partially along a disk of the first blade stage, said flow guide mounted to said seal carrier and generally closing off the stationary chamber portion so that the surfaces bounding the stationary chamber portion are essentially non-rotating, wherein said flow guide is affixed to said manifold with a first bolted joint and to said seal carrier with a second bolted joint.
10. The engine section of claim 9 wherein:
the flow guide is non-structural and formed of sheet metal.
11. A gas turbine engine comprising:
a turbine rotor stack assembly comprising:
a plurality of disks; and
a plurality of stages of blades, each stage carried by an associated disk of the plurality of disks;
a plurality of stator vane stages, interspersed with the blade stages; and
a cover plate mounted to an upstream side of a first disk of the plurality of disks; and
a cooling air outlet positioned to direct cooling air to a region adjacent the first disk, wherein:
the first disk has a perimeter array of slots;
the blades of a first blade stage of the plurality of stages of blades each have an inboard attachment portion received in an associated slot of the slots;
the cover plate has a first portion engaging the first disk to retain the cover plate against movement away from the first disk;
the cover plate has a second portion, outboard of the first portion and engaging the blades of the first blade stage to resist upstream movement of said blades;
the cover plate has a sealing portion for engaging a seal element inboard of a first stator vane stage of said plurality of stator vane stages; and
a flow guide extends between the cooling air outlet and an inboard seal with the rotor.
12. The engine of claim 11 wherein:
the cover plate has radial span between an inner radius and perimeter radius; and
a ratio of said radial span to a hub radius at the first blade stage is no more than 0.25.
13. The engine of claim 11 wherein:
the cover plate has radial span between an inner radius and perimeter radius; and
a ratio of said radial span to a hub radius at the first blade stage is no more than 0.20.
14. The engine of claim 11 wherein:
the cooling air outlet has a characteristic radius from a central longitudinal axis of the rotor stack assembly; and
a ratio of said characteristic radius to a hub radius at the first blade stage is no less than 0.60.
15. The engine of claim 11 wherein:
the cooling air outlet has a characteristic radius from a central longitudinal axis of the rotor stack assembly; and
a ratio of said characteristic radius to a hub radius at the first blade stage is no less than 0.70.
16. The engine of claim 11 wherein:
the engagement of the cover plate first portion to the first disk is via a locking ring.
17. The engine of claim 11 wherein:
the cooling air outlet is a tangential on-board injector.
18. A gas turbine engine comprising:
a turbine rotor stack comprising:
a plurality of disks; and
a plurality of stages of blades, each stage carried by an associated disk of the plurality of disks;
a plurality of stator vane stages, interspersed with the blade stages; and
a cooling air outlet positioned to direct cooling air to a region adjacent a first disk, wherein:
the first disk of the plurality of disks has a perimeter array of slots;
the blades of a first blade stage of the plurality of stages of blades each have an inboard attachment portion received in an associated slot of the slots; and
at least one of:
a first portion of the cooling air is directed radially outward and between the first blade stage and a first of said stator vane stages into a core flowpath; and
a second portion of the cooling air is directed radially inward and to an inboard seal;
the engine includes means for limiting voracity of at least one of said first portion and said second portion;
the turbine further comprises a cover plate mounted to an upstream side of the first disk;
the cover plate has a first portion engaging the first disk to retain the cover plate against movement away from the first disk;
the cover plate has a second portion, outboard of the first portion and engaging the blades of the first blade stage to resist upstream movement of said blades; and
the cover plate has a sealing portion for engaging a seal element inboard of the first stator vane stage.
19. The engine of claim 18 wherein:
the cooling air outlet is a tangential on-board injector.
20. The engine of claim 18 wherein:
the means is positioned radially outboard of a seal, the seal being radially outboard of the outlet.
21. The engine of claim 18 wherein:
the means comprises an annular flow guide.
22. The engine of claim 21 wherein:
the flow guide reduces a volume of a chamber ahead of the first disk;
the flow guide has a shape approximating a shape of an upstream face of the first disk; and
the flow guide is between inboard and outboard seals.
23. The engine of claim 21 wherein:
the flow guide is securely attached to a static structure at one of an inboard end of the flow guide and an outboard end of the flow guide; and
at the other of the inboard end of the flow guide and the outboard end of the flow guide, the flow guide is either:
securely attached to said static structure; or
slip joint coupled to said static structure.
24. The engine of claim 18 wherein:
the means comprises a non-structural annular flow guide mounted to static structure at inboard and outboard portions.
25. The engine of claim 18 wherein:
the means comprises an annular, generally C-sectioned, flow guide.
26. The engine of claim 18 wherein:
the means comprises first and second annular flow guides respectively inboard and outboard of the outlet.
27. A gas turbine engine comprising:
a turbine rotor stack comprising:
a plurality of disks; and
a plurality of stages of blades, each stage carried by an associated disk of the plurality of disks;
a plurality of stator vane stages, interspersed with the blade stages;
a cover plate mounted to an upstream side of a first disk of the plurality of disks; and
a cooling air outlet positioned to direct cooling air to a region adjacent the first disk, wherein:
the first disk has a perimeter array of slots;
the blades of a first blade stage of the plurality stages of blades each have an inboard attachment portion received in an associated slot of the slots;
the cover plate has first portion engaging the first disk to retain the cover plate against movement away from the first disk;
the cover plate has a second portion, outboard of the first portion and engaging the blades of the first blade stage to resist upstream movement of said blades;
the cover plate has a sealing portion for engaging a seal element inboard of a first stator vane stage of said plurality of stator vane stages;
a flow guide extends between the cooling air outlet and an inboard seal with the rotor;
the cover plate has radial span between an inner radius and perimeter radius; and
a first ratio of said radial span to a hub radius at the first blade stage is no more than 0.25;
the cooling air outlet has a characteristic radius from a central longitudinal axis of the rotor stack; and
a second ratio of said characteristic radius to said hub radius at the first blade stage is no less than 0.60.
28. The engine of claim 27 wherein:
the blades of the first blade stage have cooling passageways positioned to receive at least a portion of the cooling air.
29. The engine of claim 27 wherein:
the first ratio is no more than 0.20; and, the second ratio is no less than 0.70.
30. A turbine engine section comprising:
a stationary vane stage;
a rotating blade stage, said blade stage spaced from said vane stage to form an annular chamber therebetween;
a manifold for delivering a pressurized fluid to the chamber;
an outer diameter sealing means for restricting leakage of the pressurized fluid from said chamber;
an inner diameter sealing means for restricting leakage of the pressurized fluid from said chamber, the inner diameter sealing means comprising an abradable seal material carried by a seal carrier, the seal carrier extending aftward and radially inward from a forward portion comprising a flange bolted to the manifold to a rear portion carrying the inner diameter sealing means; and
a flow guide of generally forwardly open, C-shaped section extending radially between said inner diameter sealing means and said manifold, said flow guide bisecting said chamber to form a stationary chamber portion and a rotating chamber portion, the rotating chamber portion at least partially along a disk of the first blade stage, said flow guide mounted to said seal carrier.
31. A turbine engine section comprising:
a stationary vane stage;
a rotating blade stage, said blade stage spaced from said vane stage to form an annular chamber therebetween;
a manifold for delivering a pressurized fluid to the chamber;
an outer diameter sealing means for restricting leakage of the pressurized fluid from said chamber;
an inner diameter sealing means for restricting leakage of the pressurized fluid from said chamber, the inner diameter sealing means comprising an abradable seal material carried by a seal carrier; and
a flow guide of generally forwardly open, C-shaped section extending radially between said inner diameter sealing means and said manifold, said flow guide bisecting said chamber to form a stationary chamber portion and a rotating chamber portion, the stationary chamber portion not bounded by a rotor of the engine associated with the rotating blade stage, the rotating chamber portion at least partially along a disk of the first blade stage, said flow guide mounted to said seal carrier and generally closing off the stationary chamber portion so that the surfaces bounding the stationary chamber portion are essentially non-rotating with the stationary chamber portion bounded by the manifold forward and outboard of the stationary chamber portion, the flow guide aft of the stationary chamber portion, and the seal carrier forward and inboard of the stationary chamber portion.