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A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section co

A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section. The second flow of compressed air bypasses the combustor and arrives at the turbine section to form a cooling flow. A first portion of the cooling flow is routed to a front face of the rotor disc and a second portion of the cooling flow is routed to a rear face of the rotor disc.

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1. A gas turbine engine comprising: a compressor section, a combustor, and a turbine section in flow series;a drive arm connecting the turbine section with the compressor section; anda turbine section cooling system including: a first compressed air bleed arrangement which bleeds a first flow of com

1. A gas turbine engine comprising: a compressor section, a combustor, and a turbine section in flow series;a drive arm connecting the turbine section with the compressor section; anda turbine section cooling system including: a first compressed air bleed arrangement which bleeds a first flow of compressed air from a high pressure stage of the compressor section, the first flow of compressed air bypassing the combustor and arriving at the turbine section to form a sealing flow between a row of stator vanes and an adjacent downstream rotor disc and/or a first cooling flow for the stator vanes and/or the rotor disc; anda second compressed air bleed arrangement which bleeds a second flow of compressed air from at least one lower pressure stage of the compressor section relative to the high pressure stage, the second flow of compressed air bypassing the combustor and arriving at the turbine section to form a second cooling flow, a first portion of the second cooling flow being routed to a front face of the rotor disc radially outwardly of the drive arm and a second portion of the second cooling flow being routed to a rear face of the rotor disc. 2. The gas turbine engine according to claim 1, wherein the turbine section cooling system further includes a re-pressurizing pump which re-pressurizes at least one of the first portion and the second portion of the second cooling flow after the second flow of compressed air is bled from the at least one lower pressure stage of the compressor section. 3. The gas turbine engine according to claim 2, wherein a row of turbine blades is mounted to a rim of the rotor disc, at least a part of the re-pressurized first portion and/or second portion of the second cooling flow being subsequently routed into the turbine blades. 4. The gas turbine engine according to claim 2, wherein the re-pressurizing pump comprises at least one of a first centrifugal pump adjacent the front face of the rotor disc for pressurizing the first portion of the second cooling flow and a second centrifugal pump adjacent the rear face of the rotor disc for pressurizing the second portion of the second cooling flow. 5. The gas turbine engine according to claim 4, wherein the drive arm connects the rotor disc to the high pressure stage of the compressor section, the first centrifugal pump penetrating the drive arm to extend over the front face of the rotor disc both outboard and inboard of the drive arm. 6. The gas turbine engine according to claim 4, wherein the at least one of the first centrifugal pump and the second centrifugal pump comprises: a cover plate axially-spaced from a respective face of the rotor disc such that a channel producing radially outward delivery of a respective portion of the second cooling flow is formed between the cover plate and the respective face, anda row of circumferentially spaced vanes positioned in the channel and co-rotational with the rotor disc to pressurize the respective portion of the second cooling flow delivered through the channel. 7. The gas turbine engine according to claim 5, wherein the at least one of the first centrifugal pump and the second centrifugal pump comprises: a cover plate axially-spaced from a respective face of the rotor disc such that a channel producing radially outward delivery of a respective portion of the second cooling flow is formed between the cover plate and the respective face, anda row of circumferentially spaced vanes positioned in the channel and co-rotational with the rotor disc to pressurize the respective portion of the second cooling flow delivered through the channel. 8. The gas turbine engine according to claim 1, wherein the second portion of the second cooling flow is routed via a central bore of the rotor disc. 9. The gas turbine engine according to claim 1, wherein the second compressed air bleed arrangement comprises passages for the second flow of compressed air extending radially inwards from the at least one lower pressure stage of the compressor section, the passages co-rotating with the at least one lower pressure stage to reduce swirl pressure losses in the second flow of compressed air. 10. The gas turbine engine according to claim 1, wherein the second compressed air bleed arrangement comprises a vortex reducer to reduce a tangential velocity of the second flow of compressed air as the second flow of compressed air is bled from the at least one lower pressure stage of the compressor section and thereby reduce swirl pressure losses in the second flow of compressed air. 11. A method of cooling a turbine section of a gas turbine engine which comprises in flow series a compressor section, a combustor, and the turbine section, and a drive arm connecting the turbine section with the compressor section, the method comprising: bleeding a first flow of compressed air from a high pressure stage of the compressor section, the first flow of compressed air bypassing the combustor and arriving at the turbine section to form a sealing flow between a row of stator vanes and an adjacent downstream rotor disc and/or a first cooling flow for the stator vanes and/or the rotor disc;bleeding a second flow of compressed air from at least one lower pressure stage of the compressor section relative to the high pressure stage, the second flow of compressed air bypassing the combustor and arriving at the turbine section to form a second cooling flow;routing a first portion of the second cooling flow to a front face of the rotor disc radially outwardly of the drive arm; androuting a second portion of the second cooling flow to a rear face of the rotor disc. 12. The method according to claim 11, further comprising: re-pressurizing at least one of the first portion and the second portion of the second cooling flow after the second flow of compressed air is bled from the at least one lower pressure stage of the compressor section. 13. The method according to claim 12, wherein a row of turbine blades is mounted to a rim of the rotor disc, and the method further comprises: subsequently routing at least a part of the re-pressurized first portion and/or second portion of the second cooling flow into the turbine blades.