Abstract:
A stage of guide vanes ( 20 ) are cooled by compressor air delivered via piping ( 36,38 ) and by leakage air in the space volume ( 28 ) bounded by the combustion apparatus ( 14 ) and turbine shafting. The leakage air is drawn through tubing ( 40 ) by the compressor air which is directed over the exit ends of tubing ( 40 ) to create the necessary pressure drop in the tubing ( 40 ).

Description:
FIELD OF THE INVENTION 
     The present invention relates to the cooling system of a gas turbine engine. 
     BACKGROUND OF THE INVENTION 
     Some gas turbine engines operate at temperatures which are such as to require that at least some parts of its turbine apparatus be provided with appropriate supplies of cooling air from the engine compressor. However, air taken from the compressor for turbine cooling reduces the amount available for burning in the combustion system, thus generating an engine performance penalty. That situation is further exacerbated in that the air lost to the combustion system through cooling needs, adds to air lost through unavoidable leakage thereof through seals between the static and rotating members that make up the compressor assembly, the leaked air passing into the space volume bounded by the combustion apparatus and turbine shafts. 
     SUMMARY OF THE INVENTION 
     The present invention seeks to provide a gas turbine engine with an improved cooling mode. 
     The present invention comprises a gas turbine engine including a stage of turbine guide vanes, each of which has a passage therethrough, the radially inner end of said passage, with respect to the engine axis, having a respective tubular member in nested spaced relationship therein, all said tubular members being in airflow communication with a space volume bounded by combustion apparatus and turbine shafts of said engine, and suction means connected to draw air from said space volume via said tubular members, and force said drawn air through said guide vanes. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention will now be described by way of example and with reference to the accompanying drawings in which: 
         FIG. 1  is a diagrammatic sketch of a gas turbine engine of the kind which may incorporate cooling air delivery apparatus is accordance with the present invention. 
         FIG. 2  is an enlarged part view of the turbine apparatus of  FIG. 1  including cooling air delivery apparatus in accordance with the present invention. 
         FIG. 3  is an alternative form of cooling air entry structure into the tubular members, and 
         FIG. 4  is a further alternative form of cooling entry structure into the tubular structures. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to  FIG. 1 , a gas turbine engine indicated generally by arrow  10 , has a compressor  12 , combustion apparatus  14 , a turbine section  16  and an exhaust nozzle  18 . 
     Turbine section  16  includes a stage of guide vanes  20 , immediately followed in a downstream direction by a stage of turbine blades  22 . The stage of turbine blades  22  is carried on a disk  24  in known manner. Disk  24  co-rotates with a connected shaft  26 . The combustion apparatus  14 , with shaft  26 , bound a space volume  28  that is full of air during operation of engine  10 , which air continuously leaks through seals (not shown) between the static and rotating parts (not shown) of compressor  12 . 
     Referring now to  FIG. 2 , in the present example the interior of each guide vane  20  is divided into three compartments numbered  30 ,  32  and  34  respectively. Compartment  30  is connected via piping  36  and  38 , to compressor  12  ( FIG. 1 ) for direct delivery of cooling air therein. The two opposing flows meet at the exit of pipe  36  and expand laterally around the exit end portion of a tubular member  40  into chamber  42  and into compartment  32  via a converging space  43  defined between tubular member  40  and the walls defining compartment  32 . 
     Each tubular member  40  is located in the rim  44  or an otherwise hollow annular member  46 , the radially inner portion of which is open to the space volume  28 , and thereby to air that has leaked into space volume  28  during operation of engine  10 . By this means, the compressor air flowing over the converging space  43  around the exit end of tubular members  40  creates a pressure drop within the exit ends which result in the initiation of a flow of leakage air from space volume  28 , through tubular members  40  into respective guide vanes  20 . The resulting mixture of compressor air and leakage air then flows into compartment  34 , and from there via slots  48  in the trailing edges of the guide vanes  20  into the gas annulus of turbine section  16 . 
     Referring now to  FIG. 3 , should it prove necessary to modify the relative pressures of the compressor air and leakage air in order to effect the desired flow of leakage air through tubular members  40 , a metering plate  50  may be utilised at the radially inner end  46  of annular member  44 . Metering plate  50  has a number of holes drilled in it so as to provide an appropriate flow restriction area having regard to the air flow requirements for a particular engine  10 . 
     Referring now to  FIG. 4 , this example of the present invention only differs from the example of  FIG. 2  in that the radially inner end of annular member  46  is curved towards the upstream face of the adjacent turbine disk  24 , and each wall of member  46  locates in radially spaced relationship within respective lands  54  and  56  formed on turbine disk  24 . The radial spaces are filled by annular seals  58  and  60  supported on the curved end portions of annular member  46 . An annular chamber  62  is thus formed. 
     During operation of engine  10  compressor leakage air in space volume  28  enters chamber  62  via seal  60 . However, compressor air flowing through converging space  43  sucks the air from chamber  62  and passes it through the guide vanes exactly as described with reference to FIG.  2 . 
     The present invention provides two advantages over and above prior art. One advantage which is attained by all three variants described and illustrated in this specification is that utilisation of compressor leakage air for the cooling of the stage of guide vanes  20 , enables a reduction of up to 20% of the amount of cooling air hitherto extracted directly from the compressor for that purpose. The further advantage relates only to  FIG. 4  described and illustrate herein. Leakage air is contaminated with particulate matter from the ambient atmosphere, and prior to the provision of chamber  62 , it leaked past existing seal  58  into the cooling air passages ways (not shown) in the turbine blades  22  which resulted in their blockage. The leakage air also leaked past existing seal  64  and thence through the spaced overlap  66  between the vane and blade stages, thus disturbing the gas flow. Removal of the leakage air from chamber  62  by the suction means of the present invention as described hereinbefore obviated both blockage and flow disturbance.