Tandem rotor blades with cooling features

Abstract

A gas turbine engine has a compressor section with a downstream most blade stage and a downstream vane row positioned downstream of the downstream most blade stage. The downstream most blade stage includes a plurality of blade pairs with the majority of blade pairs being circumferentially spaced apart from others of the blade pairs, each the blade pair being operatively connected to a rotor disk disposed radially inward from the blade pairs with each blade pair including a forward blade and an aft blade. The aft blades are configured to further condition airflow with respect to the forward blade. A tangential onboard injector (TOBI) is positioned radially inwardly of the downstream vane row.

1 . A gas turbine engine comprising:
 a compressor section having a downstream most blade stage and a downstream vane row positioned downstream of said downstream most blade stage, said downstream most blade stage including a plurality of blade pairs with the majority of blade pairs being circumferentially spaced apart from others of said blade pairs, each said blade pair being operatively connected to a rotor disk disposed radially inward from said blade pairs with each said blade pair including a forward blade and an aft blade, said aft blades being configured to further condition airflow with respect to said forward blade; and   a tangential onboard injector (TOBI) being positioned radially inwardly of said downstream vane row.   
     
     
         2 . The gas turbine engine as set forth in  claim 1 , wherein there is no intervening stator vane stage between said forward blades and said aft blades. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein at least 10% of the cooling air passing through said TOBI is compressor midstream air. 
     
     
         4 . The gas turbine engine as set forth in  claim 3 , wherein said compressor midstream air is cooled in a heat exchanger before passing into said TOBI. 
     
     
         5 . The gas turbine engine as set forth in  claim 2 , wherein air passing through said TOBI is air having been compressed by downstream most blade stage of said compressor. 
     
     
         6 . The gas turbine engine as set forth in  claim 5 , wherein said air passes through a heat exchanger before passing through said TOBI. 
     
     
         7 . The gas turbine engine as set forth in  claim 2 , wherein a seal is provided between a downstream end of said downstream most blade stage and an upstream end of said downstream vane row, and said seal defining a cavity extending radially inwardly along a hub of said disk. 
     
     
         8 . The gas turbine engine as set forth in  claim 7 , wherein said seal is provided on both said downstream end of said downstream most blade stage, and said upstream end of said downstream vane row. 
     
     
         9 . The gas turbine engine as set forth in  claim 7 , wherein a static structure radially inward of said downstream vane row includes a plate extending at an angle which is non-perpendicular relative to a center axis of said engine and having a component extending radially inwardly and said disk hub also having a portion extending in a direction which is non-perpendicular to said central axis and having a component which is radially inwardly, said plate causing the air having passed through said TOBI to be confined adjacent to said disk hub portion. 
     
     
         10 . The gas turbine engine as set forth in  claim 9 , wherein said plate and said disk hub portion each having a radially inner portion which extends closer to being parallel to said central axis than does a radially outer portion. 
     
     
         11 . The gas turbine engine as set forth in  claim 10 , wherein a mini disk is positioned radially inwardly of said radially inner portion of said hub. 
     
     
         12 . The gas turbine engine as set forth in  claim 11 , wherein said downstream most blade stage has air passages between said blades and said rotor disk such that air from said cavity can pass upstream toward a vane row upstream of said downstream most blade stage. 
     
     
         13 . The gas turbine engine as set forth in  claim 1 , wherein cooling air passing through said TOBI is compressor midstream air. 
     
     
         14 . The gas turbine engine as set forth in  claim 1 , wherein air passing through said TOBI is air having been compressed by said downstream most blade stage of said compressor. 
     
     
         15 . The gas turbine engine as set forth in  claim 1 , wherein a seal is provided between a downstream end of said downstream most blade stage and an upstream end of said downstream vane row, and said seal defining a cavity extending radially inwardly along a hub of said disk. 
     
     
         16 . The gas turbine engine as set forth in  claim 15 , wherein said seal is provided on both said downstream end of said downstream most blade stage, and said upstream end of said downstream vane row. 
     
     
         17 . The gas turbine engine as set forth in  claim 1 , wherein a static structure radially inward of said downstream vane row includes a plate extending at an angle which is non-perpendicular relative to a center axis of said engine and having a component extending radially inwardly and said disk hub also having a portion extending in a direction which is non-perpendicular to said central axis and having a component which is radially inwardly, said plate causing the air having passed through said TOBI to be confined adjacent to said disk hub portion. 
     
     
         18 . The gas turbine engine as set forth in  claim 1 , wherein a mini disk is positioned radially inwardly of said radially inner portion of said hub. 
     
     
         19 . The gas turbine engine as set forth in  claim 1 , wherein said downstream most blade stage has air passages between said blades and said rotor disk such that cavity air from said cavity can pass upstream toward a vane row upstream of said downstream most blade stage. 
     
     
         20 . A gas turbine engine, comprising:
 a compressor section having a rotor disk defined between a compressor case and a centerline axis; and   
       a plurality of stages defined radially inward relative to the compressor case, wherein the plurality of stages includes at least one tandem blade stage, wherein the at least one tandem blade stage includes:
 a plurality of blade pairs, each blade pair being circumferentially spaced apart from the other blade pairs, each blade pair being operatively connected to the rotor disk, wherein each blade pair includes a forward blade and an aft blade, wherein the aft blade is configured to further condition air flow with respect to the forward blade without an intervening stator vane stage shrouded cavity therebetween.