Gas turbine engine compressor impeller cooling air sinks

Abstract

A gas turbine engine includes devices, systems, and methods for providing bleed air from the compressor impeller to the turbine for cooling and/or other use. The bleed air may include compressor cooling air that is routed through the diffuser and external to an outer bypass duct and/or internally to a forward wheel cavity of the turbine.

What is claimed is: 
     
         1 . A gas turbine engine comprising
 an engine core defining a rotating axis, the engine core includes a compressor having an impeller arranged to rotate about the axis to compress air with an impeller tip and a diffuser for collecting compressed air from the impeller tip, a combustor fluidly connected to receive compressed air from the diffuser for combustion, and a turbine fluidly connected to receive exhaust products from the combustor, the impeller, the diffuser, the combustor, and the turbine collectively defining a core flow path;   an outer shell disposed about the engine core to house the engine core therein; and   a bleed circuit fluidly connected between the compressor and the turbine for communicating bleed of impeller air from the impeller to the turbine, the bleed circuit including a bleed inlet arranged at the impeller tip and configured to bleed a stream of impeller air out from the core flow path and to communicate the stream to a deposit junction of the turbine.   
     
     
         2 . The gas turbine engine of  claim 1 , wherein the deposit junction includes a forward wheel cavity of the turbine and the stream of impeller air enters and purges the forward wheel cavity. 
     
     
         3 . The gas turbine engine of  claim 2 , wherein the stream of impeller air pressurizes the forward wheel cavity and leaks into the core flow path at a location between a first stage vane and a first stage blade of the turbine. 
     
     
         4 . The gas turbine engine of  claim 1 , wherein the deposit junction includes a vane cooling path of a second stage vane of the turbine and the stream of impeller air passes through the vane cooling path to cool the second stage vane. 
     
     
         5 . The gas turbine engine of  claim 4 , wherein the bleed circuit comprises at least one inlet passage defined through at least one blade of the diffuser and in communication with the bleed inlet to receive bleed of impeller air. 
     
     
         6 . The gas turbine engine of  claim 5 , wherein the bleed circuit comprises at least one transport passage in fluid communication with the at least one inlet passage, the transport passage penetrating through the outer shell and extending along the axis outside of the outer shell to a location near the turbine and into a plenum of the second stage vane of the turbine. 
     
     
         7 . The gas turbine engine of  claim 5 , wherein the stream passes through the at least one inlet passage in a direction towards a forward end of the gas turbine engine. 
     
     
         8 . The gas turbine engine of  claim 1 , wherein the bleed circuit includes a bleed inlet formed at least partially within a clearance of the impeller. 
     
     
         9 . A gas turbine engine comprising
 an engine core defining a rotating axis, the engine core including a compressor for compressing air, a combustor fluidly connected to receive compressed air from the compressor for combustion, and a turbine fluidly connected to receive exhaust products from the combustor; the compressor, the combustor, and the turbine collectively defining a core flow path; and   a bleed circuit fluidly connected between the compressor and the turbine for communicating bleed of air from an impeller of the compressor, the bleed circuit including a bleed inlet arranged at a tip of the impeller and configured to bleed a stream of air out from the core flow path and to communicate the stream to a deposit junction of the turbine.   
     
     
         10 . The gas turbine engine of  claim 9 , wherein the deposit junction includes a forward wheel cavity of the turbine and the stream of air enters the forward wheel cavity for purging the same. 
     
     
         11 . The gas turbine engine of  claim 10 , wherein the stream pressurizes the forward wheel cavity and leaks into the core flow path at a location between a first stage vane and a first stage blade of the turbine. 
     
     
         12 . The gas turbine engine of  claim 9 , wherein the deposit junction includes a vane cooling path of a second stage vane of the turbine and the stream of air passes through the vane cooling path to cool the second stage vane. 
     
     
         13 . The gas turbine engine of  claim 12 , wherein the compressor includes a diffuser having a number of blades arranged to collect compressed air from the compressor and the bleed circuit comprises at least one inlet passage defined through at least one of the number of blades of the diffuser and in communication with the bleed inlet to receive the stream of air. 
     
     
         14 . The gas turbine engine of  claim 13 , wherein the bleed circuit comprises at least one transport passage in fluid communication with the at least one inlet passage, the transport passage penetrating through an outer shell of the engine and extending along the axis outside of the outer shell to a location near the turbine and into a plenum of the second stage vane of the turbine. 
     
     
         15 . The gas turbine engine of  claim 13 , wherein the stream passes through the at least one inlet passage in a direction towards a forward end of the gas turbine engine. 
     
     
         16 . A method of operating a gas turbine engine, the method comprising
 flowing an engine core flow through each of a compressor, a combustor, and a turbine fluidly,   bleeding a stream of air in a bleed circuit from a tip of an impeller of the compressor out from the core flow path, and   depositing the stream to a deposit junction of the turbine.   
     
     
         17 . The method of  claim 16 , wherein the deposit junction includes a vane cooling path of a second stage vane of the turbine and the stream of air passes through the vane cooling path to cool the second stage vane. 
     
     
         18 . The gas turbine engine of  claim 16 , wherein bleeding includes flowing the stream of air through at least one inlet passage defined through at least one of a number of blades of a diffuser of the compressor. 
     
     
         19 . The gas turbine engine of  claim 18 , wherein bleeding includes flowing the stream of air through at least one transport passage that extends along the engine outside of an outer shell to a location near the turbine and into a plenum of the second stage vane of the turbine. 
     
     
         20 . The gas turbine engine of  claim 16 , wherein bleeding includes passing the stream through the at least one inlet passage in a direction towards a forward end of the gas turbine engine.