Abstract:
A turbine blade tip clearance control system has a rigid two part outer casing ( 42 ) which sandwiches a control ring ( 48 ) therebetween, and an air pressurised flexible inner casing ( 28 ) which carries shroud segments ( 22 ) within it. Struts ( 40 ) span the annular space between the casings ( 42, 28 ) and prevent flexing of casing ( 28 ) until blade tip clearance needs adjusting, whereupon, ring ( 48 ) is heated, along with the adjacent portion of outer casing ( 42 ) and expands, allowing casing ( 28 ) to flex outwards, thus lifting the shroud segments ( 22 ) away from the blade tips ( 24 ). 
     Closure of the tip clearance is achieved by cooling ring ( 48 ), the resulting contraction thereof, via the struts ( 40 ), flexing the inner casing ( 28 ) and shroud segments ( 22 ) inwards, against the air pressure.

Description:
FIELD OF THE INVENTION 
   The present invention relates to a structure within which a stage of turbine blades rotates, during operation of an associated gas turbine engine. 
   More specifically, the structure is of the kind which may be caused to expand and contract along lines radial to the axis of rotation of the stage of turbine blades, so as to at least reduce the magnitude of blade tip rub on structure immediately surrounding them. 
   BACKGROUND OF THE INVENTION 
   Devices are known, which are designed to expand radially about a stage of turbine blades, so as to maintain a desirable clearance therebetween. A first example is described and illustrated in published patent specification 1484936. In that example, non rotating shrouds surround a stage of turbine blades. The downstream ends of the shrouds are hooked on a first expandable ring, which is located by radial dowels. The shrouds ends are also hooked in a ring of different expansion and contraction characteristics from those of the first ring. The upstream end of each shroud has an arm fixed thereto by one end, the other end having a ball thereon, which pivots in a socket in fixed structure when the first ring expands as a result of being heated, thus enabling, the first ring to lift the shrouds away from the tips of the blades. The other ring prevents too rapid movement of the shrouds towards the tips of the blades when cooling occurs. 
   A further example is illustrated and described in published patent specification 1605403. A turbine casing surrounds a stage of turbine blades, which again, include spaced, non rotatable shrouds. A polygonal member surrounds the turbine casing, and has radially arranged bolts fixed thereto so as to project radially inwards, towards the shrouds. The bolts heads locate in the opposing ends of expandable segments which surround the shrouds, which segments in turn, are hooked via their centre portions, to the opposing ends of the respective shroud segments. When the expandable segments are heated, they expand about their centres, into arched forms, thus lifting the shroud segments away from the tips of the blades. 
   Both examples of prior art disclosed hereinbefore rely entirely on expansion, and are comprised of a multiplicity of parts, which are extremely expensive to produce, and results in complexity of assembly. In the former example, there are provided valve mechanisms which themselves must be expanded, so as to enable heat to reach the shroud moving mechanism. In the latter example, accurate movement of the blade shroud segments about the pivot point of their respective arms, raises the need for, possibly, undesirably large clearances between their downstream extremities and structure adjacent thereto, and thus would reduce turbine efficiency through gas leakage. 
   SUMMARY OF THE INVENTION 
   The present invention seeks to provide an improved gas turbine blade tip clearance control structure. 
   According to the present invention, a gas turbine engine turbine blade tip clearance control system comprises a rigid outer casing connectable to a variable temperature air supply, a flexible inner casing having an inner surface connectable to a pressurised air supply, and supporting a circumferential array of shroud segments therewithin, an equi-angular array of struts separating said casings, whereby, in operation in a gas turbine engine, said outer casing is expandable and contractable by application of hot or cold air thereto, to allow or prevent, via said struts, pressurised air acting on said inner casing inner surface, to flex said inner casing. 

   
     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 representation of a gas turbine engine incorporating blade tip clearance control structure in accordance with the present invention. 
       FIG. 2  is an enlarged, cross sectional view of the encircled portion in FIG.  1 . 
       FIG. 3  is a view on line  3 — 3  of FIG.  2 . 
   

   DESCRIPTION OF THE INVENTION 
   Referring to  FIG. 1. A  gas turbine engine  10  has a compressor  12 , a combustion section  14 , a turbine stage  16 , and an exhaust nozzle  18 , all arranged in flow series in known manner. 
   Referring now to FIG.  2 . The turbine stage  16  includes a rotary stage of turbine blades  20 , only one of which is shown. The stage of blades  20  is surrounded by a ring of shroud segments  22 , which, in, a non operative mode of engine  10 , are very closely spaced from the tips  24  of respective blades  20 . The spacing is achieved by supporting the shroud segments by cooperating hooked features  26  and  27  on their leading edges, and on the interior of a flexible casing  28  and by ‘birdmouth’ joints  30  on the interior of flexible casing  28 , cooperating with spigots  32  on the trailing edges of the shroud segments  22 . Although in this particular case a ‘birdmouth’ joint  30  is employed other fastening devices such as hooks could be employed likewise the spigots  32  could be replaced by an alternative fastening device such as a hook or lip. 
   Casing  28  is fixed in its upstream end it to further casing structure,  34 , which extends towards or over the combustion zone  14 . The downstream and of casing  28  is supported on further fixed structure  36 , via a sliding ‘bird mouth’ joint  38 , which enables some axial movement thereof, through casing  28  flexing during operation of engine  10 . Again, although a ‘bird mouth’ joint  38  is employed, other suitable joint arrangements which provide the necessary degree of sealing, may be used. 
   Casing  28  has a number of struts of substantial proportions projecting radially therefrom, in equi-angularly spaced array, the outer ends of which indirectly abut the inner surface of a rigid, low flexibility outer casing  42 , thereby supporting casing  28  against flexing under air pressure loads and mechanical generated during operation of engine  10 . 
   During at least some operating conditions of engine  10 , blades  20  will expanded radially outwards, and shroud segments  22 , must also be moved outwards, so as to eliminate or at least minimize rubbing of the blades tips  24  against them. To this end it, casing  28 , is made from a material, which is of such proportions, and is a sufficiently flexible, as to enable it to achieve the desired adequate movement. However, because struts  40  are present, that circumferential portion of rigid casing  42 , which surrounds struts  40  must also be movable. In a radially outward direction, which is explained later in this specification. The relevant portion of casing  42  is made up from two axially short casings.  44  and  46 , which are fixedly joined via flanges, which sandwich a ring  48  therebetween. Ring  48  has an inner land  50  and an outer land  52 , which overlap in their respective interfaces with the casings  44  and  46 . 
   A thin segmented ring  54  is positioned between the inner land  50  and the struts  40 , and acts as a thrust load distributor, when radial loads are experienced by struts  40  and ring  48 , as is explained hereinafter. 
   Prior to start up of engine  10 , casing  28  holds shroud segments  22  in close spaced relationship with the blades tips  24 . When engine  10  is started, and runs at idle speed, there is insufficient growth of turbine blades  20 , to require flexing of casing  28 , to cause movement of shroud segments  22  away from blades  20 . However, when an aircraft (not shown), driven by engine  10 , takes off, engine  10  is accelerated it to full thrust, at which time, its operating temperature rapidly increases, and, consequentially, so does growth of blades  20 . It then becomes necessary to flex casing  28 , to move shroud segments  22 , so as to at least reduce rubbing of blade tips  24  against them. 
   As stated hereinbefore, in order that casing  28  may flex radially outwards of the axis of engine  10 , the portion of rigid outer casing  42  which is in radial alignment with struts  40  must be caused to move in the same direction. This is achieved by heating the flanged joint and ring  48  which is sandwiched therebetween. A cowl structure  56  is provided, which surrounds the flanged joint and ring  48 , and hot air derived from an appropriate region of the compressor  12  is directed thereto via a control valve  58 , and a conduit  60 . The flanged joint and ring  48  then expand, and thus enable struts  40 , and casing  28  to follow, without losing contact therewith. 
   Flexing of casing  28  is achieved as follows. Shroud  30  segments  22 , with respective casings  28 ,  62  and  64 , form an annular space  66 , which, via a circumferential array of apertures  68 , only one of which is shown, is in permanent flow communication with a high pressure stage in the compressor  12 . As the pressure of the air delivered from compressor  12  increases during the aforementioned aircraft take off stage, it reaches a level within space  66 , at which together with thermal distortion of the casing  28  it forces casing  28  to start flexing in a radially outward v direction. Shroud segments  22  are thus lifted away from blade tips  24 . 
   When engine  10  is throttled back, as occurs when the aircraft is required to fly at cruise speeds, compressor delivery pressure will reduce, and casing  28  will begin to flex radially inwards, to the points where it attains not quite its original cold shape. This provides an appropriate spacing between shroud segments  22  and blade tips  24 . 
   In order that ring  48 , via segmented ring  54 , maintains or subsequently resumes its indirect contact with struts  40  when casing  28  flexes or has flexed radially inwards, ring  48  and associated flanges must be cooled, so as to cause them to contract at a rate which will ensure constant contact therebetween. This is achieved by directing air from the upstream, low pressure, low temperature portion of compressor  12 , via valve  58 , into cowl  56 , thus enveloping ring  48  and associated flanges therewith. 
   The appropriate actuation of valve  58 , in order to match flexing of casing  28 , and expansion of ring  48  and associated flanges, with blade tip clearance during varying engine running conditions, may be achieved in a number of ways, including developing electronic signals from any engine measurable operating parameters, such as engine revolutions, engine pressures, and engine air and/or gas pressures, and utilising those electronic signals to actuate valve  58 , so as to direct air of appropriate temperature, or pressure, to appropriate parts. 
   Casing  28  is flexed by the application of pressure to its inner surface in combination with mechanical and thermal loads, and is subjected to that pressure through all of the working regimes of engine  10 . Therefore, a counter pressure is applied to the outer surface thereof, which, combined with the inherent self supporting stiffness possessed by casing  28 , is sufficient to prevent undesirable flexing, anywhere along its length.  FIG. 3  illustrates the positional relationship between the struts  40  and the segmented load distribution ring  54 , which is seen to be split at mid point  70  between each pair of adjacent struts  40 .  FIG. 3  also depicts the angular positioning of struts  40  with respect to flexible casing  28 .