Abstract:
Shrouds ( 40 ) surrounding a stage of turbine blades ( 44 ) are cooled by a compressor airflow which is led to the downstream end of the shrouds before contacting them. The airflow passes through apertures ( 56 ) in plates ( 50 ) then over the shrouds ( 40 ) in an upstream direction, to exit from apertures ( 62 ) in the shrouds ( 40 ) in parallel with and in the same direction as the gasflow. Airflow needed is reduced relative to prior art needs, resulting in improved engine efficiency, and ejection of the air does not disturb the gasflow.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     Not applicable. 
     The present invention relates to turbine machinery, particularly of the kind utilised in gas turbine propulsion engines. 
     More specifically, the invention relates to the improved cooling of such devices, especially in those engines used for the propulsion of aircraft, though not restrictively so. 
     BACKGROUND OF THE INVENTION 
     Field of the Invention 
     It is the common practice to provide compressor air for the purpose of cooling a multiplicity of turbine machinery parts. Thus, it is know to provide compressor air to shrouds, which in situ, surround a stage of turbine blades, and thereby form an outer wall of an associated turbine annulus in which in operation, the stage of blades rotates. 
     A first drawback to the know system is that it necessitates the provision of duel airflows, one for cooling the downstream ends of the shrouds, and another for cooling the upstream ends thereof. This results in the use of a quantity of air which consequently cannot be used for combustion, and further results in a noticeable drop in engine efficiency. 
     A second drawback in known structures is that the air used for cooling the upstream ends of the shrouds, by virtue of structure, could not be finally ejected into the gas stream in the annulus, without disturbing the flow. This added further to efficiency losses. 
     SUMMARY OF THE INVENTION 
     The present invention seeks to provide an improved mode of cooling blade shrouds in turbine machinery. 
     According to the present invention, turbo machinery for a gas turbine engine comprises a turbine blade shroud capped by a cover which is in spaced relationship therewith intermediate the shroud ends, said space being connectable via the downstream end of said cover, to a cooling airflow supply from a compressor of a said engine, and connectable via said shroud near the shroud upstream end, to the gas annulus of a said engine so as to, on driven connection to a said engine, eject said cooling airflow when effected, into the gas annulus, in parallel with, and in the direction of, the gas flow therethrough. 
    
    
     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 view of a gas turbine engine including turbine machinery of the present invention. 
     FIG. 2 is an axial cross section view through the turbine section of the engine of FIG.  1 . 
     FIG. 3 is a view in the direction of arrow  3  in FIG.  2 . 
     FIG. 4 is an enlarged pictorial part view of FIG.  2 . 
     FIG. 5 is a pictorial part view of an alternative configuration to the device of FIGS. 2 and 3. 
     FIG. 6 is a pictorial part view of a further alternative configuration to the devices of FIGS. 2,  3 ,  4  and  5 . 
     FIG. 7 is a cross sectional part view of a mode of retention of the cover of FIG.  2 . 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to FIG. 1. A gas turbine engine which includes turbine machinery in accordance with the present invention, has a compressor section  10 , a combustion and fuel entry section  12  and a turbine section  14 . The engine terminates in an exhaust nozzle  16 .  20  Referring now to FIG.  2 . The turbine section  12 , has an outer casing  18  which includes internal annular flanges  20 ,  22 . Each flange  20 ,  22  have respective annular grooves  24 ,  26 . 
     Groove  24  supports the outer rim  28  of an annular groove  30  in the upstream end of the integral shroud  32  of a non rotatable guide vane  34 . 
     By ‘upstream’ and ‘downstream’ is meant with respect to the direction of flow of gases through the engine of FIG.  1 . 
     The inner rim  36  of the guide vane shroud  32  supports the downstream end  38  of a turbine blade shroud  40  in sliding relationship. Shaped sealing strips  39 ,  41  are fitted therebetween. The shroud  40  is spaced from the tip  42  of a turbine blade  44 , and extends upstream and downstream thereof. 
     The inner rim  26  of the flange  22  supports the upstream end of the shroud  40 , which in turn, carries an annular airflow restrictor  46 , the operation of which is explained later in this specification. A shaped sealing strip  47  is fitted therebetween. 
     The shroud  40  has a number of fences  48  extending over the major portion of its length, intermediate its thickened ends, which fences are covered by a plate  50  which, with the fences  48 , forms a number of closed, elongated passages  52  lying axially of the engine. 
     The passages  52  are best seen in FIG.  3 . Only two passages  52  are shown therein, but in practice, there would be sufficient fences  48  to provide passages which would span the major portion of the width of the shroud  40  in a direction circumferentially of the turbine stage of which blade  44  forms a part. 
     The upstream ends of the fences  48  are forked, for reasons which are explained hereinafter. 
     Referring back to FIG.  2 . In operation of the gas turbine engine, air is bled from the compressor  10  (FIG. 1) and led via a circumferential row of holes  54  in the restrictor  46 , the space defined by the turbine casing  18  and plate  50 , to a further row of holes  56 , spanning the plate  50  at its downstream end. The air passes inwardly through the holes  56 , into the passages  52 , reversing its direction of flow, to flow along the passages  52 , to their upstream end portions. 
     On reaching the upstream end portions of the passages  52 , the airflow is constricted by narrowed passages defined by the forked portions of the fences  48 . As a result, the airflow is re-energised at least in some small degree, prior to reaching a cutout  58  in the end extremity of one leg  60  in each fork. The cut out  58  is more clearly, seen in FIG.  4 . The air passes through the cut outs  58  and again reverses its direction of flow, to exit from a row of holes  62  in the shroud  40 , in the same direction as the gas flow through the turbine section  14 , as is indicated by the arrow  64 . 
     It will be seen from the foregoing description how a single compressor air supply can be utilised to cool both outer and inner surfaces of a turbine blade shroud, and further, be ejected therefrom into a region of the gas flow annulus, without disturbing the gas flow itself. 
     The present invention has been described with reference to only one shroud and an associated blade. However, the turbine stage will of course comprise a ring of turbine blades  44 , and a ring of shrouds  40 . Each shroud  40  may span one or more blades  44  in known manner, and, in accordance with the example of the particular present invention, will be provided with its own plate  50 . 
     The fences  48  may be cast on the shroud  40  at the manufacturing stage thereof. However, the shroud  40  is not a rotating part, and consequently, to achieve minimum weight, appropriately shaped thin metal strips may be brazed or otherwise fixed thereon, to form the fences. 
     An alternative structure comprising a honeycomb core  66  (FIG. 5) could be substituted for the fences  48 . The walls of the core  66  would have holes  70  in them, for the passage of cooling air through the core, towards the upstream end of the shroud  40 . 
     A further alternative to fences would provide pillars  72  depicted in FIG.  6  and which would separate the shroud  40  and plate  50 . Air passing through the holes  56  would flow around the pillars  72  in a generally upstream direction, until it reached the exit holes  62  in the shroud  40 . The pillars  72  are illustrated in straight form, but they could have any form, regular or irregular. 
     It is the common practice, to seal the gap between adjacent shrouds against gas leakage by providing opposing slots in opposing shroud edges, and fitting a metal strip  74  therein, to bridge the gap. This is depicted in FIG.  7  and per se forms no part of the present invention. 
     However, FIG. 7 also depicts a mode of retaining each plate  50  on its respective shroud  40 , and consists of grooved edge portions  76  being provided along the side edges of each shroud  40 , into which the side edges of the plate  50  are sprung. 
     The grooved edge portions need not extend the full length of the respective shrouds, but preferably would extend over a substantial portion of the edges of the aforementioned intermediate part thereof. 
     Despite the serpentine path which the compressor driven air has to follow between holes  54  and  62 , a positive flow is maintained, helped by a pressure drop created at the exit ends of holes  62 , by the passage of the high velocity gases thereby, in the direction indicated by arrow  64 . 
     The cooling air is bled from one stage of compressor blades (not shown) in the associated engine, into a plenum (not shown) from where the cooling air passes to the apertures  54 , then to the apertures  56 . However, the cooling air could be piped from the plenum (not shown) by pipes (not shown) equal in number to apertures  56 , and connected thereto, one pipe to one aperture  56 . 
     Hereinbefore, the cover is represented by plate  50 . It could however, be a frusto conical member (not shown) having an axial cross-sectional shape identical with that of plate  50 , as illustrated in FIG.  2 .