Turbine coolant supply system
A gas turbine engine configured to rotate in a circumferential direction about an axis extending through a center of the gas turbine engine comprises a turbine stage. The turbine stage comprises a disk, a plurality of blades and a mini-disk. The disk comprises an outer diameter edge having slots, an inner diameter bore surrounding the axis, a forward face, and an aft face. The plurality of blades is coupled to the slots. The mini-disk is coupled to the aft face of the rotor to define a cooling plenum therebetween in order to direct cooling air to the slots. In one embodiment of the invention, the cooling plenum is connected to a radially inner compressor bleed air inlet through all rotating components so that cooling air passes against the inner diameter bore.
1. A turbine stage for a gas turbine engine configured to rotate in a circumferential direction about an axis extending through a center of the gas turbine engine, the turbine stage comprising:
a turbine disk comprising:
an outer diameter edge having slots;
an inner diameter bore surrounding the axis;
a forward face;
an aft face;
a hub extending from the inner diameter bore of the turbine disk to form an annular body; and
a plurality of holes extending through the hub;
a plurality of blades coupled to the slots;
a mini-disk comprising:
an axially extending portion disposed opposite the hub;
a radially extending portion disposed opposite the aft face of the turbine disk;
an axial retention flange disposed at a radial distal tip of the radially extending portion to engage the slots; and
a coupling disposed at an axially distal tip of the axially extending portion to engage the hub, wherein the mini-disk couples to the aft face of the turbine disk to define a cooling plenum therebetween to direct cooling air to the slots, and wherein the holes permit cooling air from within the hub to enter the cooling plenum; and
a shaft extending from the hub through the inner diameter bore coupling the turbine disk to a compressor disk, wherein the inner diameter bore and the shaft define a cooling passage fluidly coupled to the holes and the plenum.
2. The turbine stage of claim 1 and further comprising:
a cover plate coupled to the forward face of the turbine disk across the slots.
3. The turbine stage of claim 1 and further comprising:
a first stage turbine rotor coupled to the forward face of the turbine disk to define an inter-stage cavity between the first stage turbine rotor and the turbine disk; and
a first stage mini-disk coupled to a forward-facing side of the first stage turbine rotor.
4. A gas turbine engine incorporating the turbine stage of claim 3 , the gas turbine engine further comprising:
a compressor stage; and
a bleed air inlet for directing cooling air from the compressor to the cooling passage, wherein the cooling passage is radially outward of the shaft, wherein the shaft couples the compressor stage to the hub of the turbine stage.
5. The gas turbine engine of claim 4 wherein the compressor stage comprises:
a first compressor rotor having a plurality of compressor blades extending from a first rim; and
a second compressor rotor having a plurality of compressor blades extending from a second rim, the second compressor rotor coupled to the first compressor rotor;
wherein the bleed air inlet extends radially inward between the first and second rims.
6. The gas turbine engine of claim 5 and further comprising:
a compressor rotor hub connecting the second compressor rotor to the shaft; and
a tie shaft coupling the compressor rotor hub to the first stage turbine rotor.
7. A gas turbine engine comprising:
a compressor section including a bleed inlet for siphoning cooling air from the compressor section;
a turbine section comprising:
a rotor comprising:
an inner diameter bore;
an outer diameter rim;
a forward face;
an aft face;
a hub extending from the aft face; and
a first flange extending radially from the hub;
a shaft coupled to the compressor section and the turbine section, wherein the shaft extends through the inner diameter bore to join to the hub;
a plurality of blades coupled to the rotor;
a mini-disk comprising:
an axially extending portion disposed opposite the hub; and
a second flange disposed at an axially distal tip of the axially extending portion to engage the first flange, wherein the mini-disk couples to the aft face of the rotor to define a plenum; and
a cooling circuit fluidly coupling the bleed inlet of the compressor section to the plenum, the cooling circuit extending along the shaft and the aft face of the rotor, wherein a portion of the cooling circuit is defined by the inner diameter bore and the shaft.
8. The gas turbine engine of claim 7 and further comprising:
a plurality of holes in the hub to fluidly connect the cooling circuit with the plenum.
9. The gas turbine engine of claim 7 wherein:
the compressor section further comprises a rotor hub; and
the shaft comprises a tie shaft extending between the rotor hub and the turbine section.
10. The gas turbine engine of claim 7 wherein the compressor section further comprises:
a first compressor rotor having a plurality of compressor blades extending from a first rim; and
a second compressor rotor having a plurality of compressor blades extending from a second rim, the second compressor rotor coupled to the first compressor rotor;
wherein the bleed air inlet extends radially inward between the first and second rims.
11. The gas turbine engine of claim 7 wherein cooling circuit is completely defined by components configured to rotate during operation of the gas turbine engine.
12. A method of providing compressor bleed air to a turbine stage of a gas turbine engine, the method comprising:
flowing the bleed air axially along a shaft connecting a compressor stage to a turbine stage, wherein an inner diameter bore of a rotor disk and the shaft define a cavity;
passing the bleed air through the cavity;
directing the bleed air radially along an aft surface of the rotor disk; and
feeding the bleed air into a blade slot in a rim of the rotor disk.
13. The method of claim 12 and further comprising:
heating the bore of the rotor disk with the compressor bleed air to reduce a temperature gradient between the rim and the bore.
14. The method of claim 12 and further comprising:
controlling thermal growth of the rotor disk with the compressor bleed air to influence blade tip clearance.
15. The method of claim 12 and further comprising:
originating the bleed air from a rim of the compressor stage; and
routing the bleed air radially inward to the shaft.
16. The method of claim 15 wherein the bleed air is bounded from the compressor stage to the turbine stage by components of the gas turbine engine configured to rotate.
17. The method of claim 12 wherein the bleed air bypasses an inter-stage cavity defined by adjacent rotor disk in the turbine stage.