Abstract:
A turbine assembly ( 35 ) for a gas turbine engine ( 10 ) comprises a rotatable support arrangement ( 38 ) which comprises means for mounting thereon a plurality of turbine blades ( 36 ). The turbine assembly ( 35 ) defines flow path means ( 43 ) for a flow of cooling fluid therethrough. The flow path means ( 43 ) is connectable to a supply of relatively cold cooling fluid. The flow path means  43  is arranged such that the relatively cold cooling fluid is driven radially outwardly through the flow path means ( 43 ) substantially wholly by the centrifugal force generated the rotation of the turbine assembly ( 35 ) in operation. Relatively hot cooling fluid is displaced by the relatively cold cooling fluid radially inwardly through the flow path means ( 43 ).

Description:
FIELD OF THE INVENTION 
     This invention relates to turbine blade cooling systems. More particularly, but not exclusively the invention relates to turbine blade cooling systems and turbine assemblies for gas turbine engines. 
     BACKGROUND OF THE INVENTION 
     It is sometimes necessary to provide the intermediate pressure turbine of a gas turbine engine with a moderate cooling. Known techniques for cooling turbine blades in gas turbine engines use air from a pre-swirl system. However such systems for cooling are costly and inefficient and there are significant energy losses associated with such systems. 
     SUMMARY OF THE INVENTION 
     According to one aspect of this invention there is provided a turbine assembly comprising a rotatable support arrangement, a plurality of turbine blades extending radially outwardly from the support arrangement, and flow path means extending radially in each of the blades for a flow of cooling fluid therethrough, and the flow path means being connectable to a supply of relatively cold cooling fluid, wherein the flow path means is arranged such that the relatively cold cooling fluid is driven radially outwardly through the flow path means substantially wholly by the centrifugal force generated by rotation of the assembly in operation, to drive relatively hot cooling fluid radially inwardly through the flow path means. 
     Preferably, the flow path means comprises a first flow path through which said relatively cold cooling fluid can pass and a second flow path through which said relatively hot cooling fluid can pass. 
     According to another aspect of this invention there is provided a method of cooling a turbine assembly, the assembly comprising a rotatable support arrangement and a plurality of turbine blades extending radially outwardly from the support arrangement, and flow path means extending radially in each of the blades for a flow of cooling fluid therethrough, wherein the method comprises arranging the flow path means in fluid communication with a supply of relatively cold cooling fluid and rotating the support arrangement to drive the relatively cold cooling fluid radially outwardly through the flow path means substantially wholly by the centrifugal force generated by rotation of the assembly in operation, and allowing said cooling fluid to be heated in said blades, whereby relatively hot cooling fluid is displaced radially inwardly through the cooling path means by the flow of said relatively cold cooling fluid. 
     The support arrangement may define a second flow path means in fluid communication with the first mentioned flow path means. The second flow path means may comprise a feed flow path extending from an inlet to the first flow path and an exhaust flow path from the second flow path to an outlet. The inlet and outlet may be provided in substantially the same region. 
     The preferred embodiment of the turbine assembly is an intermediate pressure turbine assembly. In the preferred embodiment, fluid flowing along the feed flow path can pass into the first flow path in each blade to extract heat therefrom and thereafter can flow into the second flow path to pass into the exhaust flow path to be exhausted via the outlet. 
     Preferably, the inlet of the cooling path means is defined at a central region of the support arrangement. The outlet of the cooling path means may also be defined at the central region of the support arrangement. In one embodiment, substantially all the cooling fluid entering the first mentioned flow path means is delivered to the second flow path means. Substantially all the cooling fluid entering the feed flow path may be delivered to the first mentioned flow path means, and substantially all the cooling fluid entering the exhaust flow path may be exhausted from the outlet. 
     The support arrangement may comprise a support disc upon which said plurality of turbine blades can be mounted and said support arrangement may further include a cover member arranged over a face of the disc. The cover member may be adapted to hold the turbine blades on the disc. 
     In one embodiment, at least a part of the flow path means may extend generally radially along the support disc. A further part of the flow path means may extend generally circumferentially of the disc. In one embodiment, part of the feed flow path extends generally radially of the disc and part of the exhaust flow path extends generally radially of the disc. A further part of the feed flow path may extend generally circumferentially of the disc, and a further part of the exhaust flow path may also extend generally circumferentially of the disc. 
     The flow path means may be defined by the cover member. Preferably, the flow path means is defined between the cover member and the disc. In one embodiment, the feed and exhaust flow paths are provided generally in a plane, said plane being generally parallel to the plane of the disc. In another embodiment, the feed and exhaust flow paths are provided in a plane generally transverse to the plane of the disc. 
     Each turbine blade may have a securing portion to secure the blade to the disc, and an opening may be defined in the securing portion through which cooling fluid can enter the first flow path in the blade. Each blade may further include a shank and an aerofoil section, the shank extending between the securing portion and the aerofoil section. A shroud member may be provided between the shank and the aerofoil section, whereby, when assembled, the shroud members of adjacent turbine blades engage each other to define a space between the shroud and the disc. In one embodiment, an opening for the second flow path in the blade may be defined in the shank, whereby cooling fluid in the second flow path in each blade can be passed from the blade into the space. 
     The exhaust path in the support arrangement may be in fluid communication with the space, whereby cooling fluid may flow from said second path means in the blade to the exhaust path means via said space. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     An embodiment of the invention will now be described by way of example only with reference to the accompanying drawings, in which: 
     FIG. 1 is a sectional side view of the upper half of a gas turbine engine; 
     FIG. 2 is a sectional side view of part of a high pressure turbine incorporated in the engine shown in FIG. 
     FIG. 3 is a schematic cross-sectional side view of part of one embodiment of the turbine assembly shown in FIG. 2; 
     FIG. 4 is a schematic rear view of another embodiment of a turbine assembly; 
     FIG. 5 is a close up sectional view of the turbine assembly shown in FIG. 4; and 
     FIG. 6 is a view along the lines VI—VI in FIG.  5 . 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to FIG. 1, a gas turbine engine is generally indicated at  10  and comprises, in axial flow series, an air intake  11 , a propulsive fan  12 , an intermediate pressure compressor  13 , a high pressure compressor  14 , a combustor  15 , a turbine arrangement comprising a high pressure turbine  16 , an intermediate pressure turbine  17  and a low pressure turbine  18 , and an exhaust nozzle  19 . 
     The gas turbine engine  10  operates in a conventional manner so that air entering the intake  11  is accelerated by the fan  12  which produce two air flows: a first air flow into the intermediate pressure compressor  13  and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor  14  where further compression takes place. 
     The compressed air exhausted from the high pressure compressor  14  is directed into the combustor  15  where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines  16 ,  17  and  18  before being exhausted through the nozzle  19  to provide additional propulsive thrust. The high, intermediate and low pressure turbines  16 ,  17  and  18  respectively drive the high and intermediate pressure compressors  14  and  13  and the fan  12  by suitable interconnecting shafts. 
     Referring to FIG. 2, there is shown a section through part of the intermediate pressure turbine  17  which is a single stage turbine and is connected to, and drives, the intermediate pressure compressor  13  via a shaft  28 . A casing  24  extends around the intermediate pressure turbine  17  and also extends around the high and low pressure turbines  16  and  18 . 
     The intermediate pressure turbine  17  comprises a stator assembly  31  comprising an annular array of fixed guide vanes  32  arranged upstream of a rotary assembly  35 . The guide vanes  32  are supported between an outer support structure  34  which extends circumferentially around the outer ends of the array of guide vanes  32  and an inner support structure  134  located radially inwardly of the guide vanes  32 . The rotary assembly comprises an annular array of turbine blades  36  mounted on a rotatable support arrangement  38  which in turn is mounted on the shaft  28 . The rotatable support arrangement  38  comprises a turbine disc  40  and a cover plate  42  mounted over the dished rear face  44  of the disc  40  to define cooling flow path means  43  (as will be explained below). The blades  36  each comprise an aerofoil section  46 , a shroud member  48  provided at the radially inner end of each aerofoil section  46 , a shank  50  extending radially inwardly of the shroud member and a securing portion  52  in the form of a fir tree root provided at the radially inner end of the shank  50 . 
     When all of the blades  36  have been assembled around the disc  40 , the shroud members  48  of adjacent blades  36  engage each other to define spaces  54  between the shroud members  48 , the disc  40  and between the shanks  50  of adjacent blades  36 . A plurality of such spaces  54  are provided, extending in an annular manner around the disc  40 . 
     The high and low pressure turbines  16  and  18  also comprise arrangements of guide vanes and rotor blades. The high pressure turbine  16  receives combustion products from the combustor  15  and is connected to and drives the high pressure compressor  14  via a shaft  26  (see FIG.  1 ). Similarly, the low pressure turbine  18  receives combustion products from the intermediate pressure turbine  17  and is connected to, and drives, the fan  12  via a shaft  30  (see FIG.  1 ). 
     FIG. 3 shows a schematic part sectional side view of the intermediate pressure turbine  17 ; the same features as in FIG. 2 have been given the same reference numerals. The cooling flow path means  43  is defined in the rotatable support arrangement  38 , and comprises a feed channel  58  defined between the cover plate  42  and the disc  40 , and an exhaust channel  60  defined within the cover plate  42 . 
     The feed channel  58  extends radially outwardly of the support arrangement  38  to the blade  36 . A first channel  62  is defined inside the blade  36  which is in fluid communication with the feed channel  58 . A second channel  64  extends from, and is in fluid communication with the first channel  62 . The second channel  64  is also defined inside the blade  36  and is in fluid communication with the exhaust channel  60 . As can be seen from FIG. 3, a flow of cooling fluid, as indicated by the arrows A passes along the feed channel  58  to the first channel  62  and thereafter to the exhaust channel  60  via the second channel  64 . As the cooling fluid flows in the direction indicated by the arrows A, heat is extracted from the disc  40  and from the blades  36 . As shown, substantially all the air entering the first channel  62 , the second channel  64  and the exhaust channel  60  is exhausted therefrom. A small amount of air may be bled off from the first or second channel  62 ,  64  if desired. 
     During the operation of the intermediate pressure turbine  17 , the blades  36  are heated, which in turn heats the air in the first and second channels  62 ,  64  thereby causing the air to expand. The air in the channels  62 ,  64  is displaced by incoming cooler air of higher density driven along the feed channel  58  by centrifugal force created by the rotation of the intermediate pressure turbine  17 . The hot air in the channels  62 ,  64  displaced along the exhaust channel  60 . 
     As a result, a continuous cycle of cooling air is established through the channels  58 ,  62 ,  64 ,  60  to effect cooling of the blade  36 . 
     A pressure difference is established across the first and second channels  62 ,  64  which drives the air through the channels. Since the pressures at the channels  62 ,  64  are greater than the pressure at the inlet of the feed channel  58  and at the exhaust channel  60 , the exhaust channel  60  can exhaust to a region of the same pressure as the inlet for the feed channel  58 . 
     A further embodiment is shown in FIGS. 4,  5  and  6  in which the feed and exhaust channels are arranged such that they extend generally parallel to the rear face  44  of the disc  40 , and are generally in the same plane. In FIGS. 4,  5  and  6  in which no more than two of the blades are shown for clarity, the feed channels are designated  158 A and  158 B, and the exhaust channels are designated  160 A,  160 B. Each feed channel comprises a radial part  158 A, and a circumferentially extending part  158 B. The air flows radially outwardly along the channel  158 A, into the channel  158 B and thereafter through a plurality of openings  170  each of which communicates with the first channel in the associated blade  36 . On return from each blade  36 , the hot air passes from the second channel  64  therein into the spaces  54  between the shanks  50  of the blades  36  and into the exhaust channel  160 B and thereafter into one of the radially extending channels  160 A. As can be seen from FIG. 6 the channels  158 A,  158 B,  160 A,  160 B are defined between a cover plate  172  for the disc  40 , and the disc  40  itself, by appropriate shaped formations  174  extending from the cover plate  172 , the formations  174  being adapted to engage the blade  36  or the disc  40 . 
     It is desirable to ensure that the cooling air flows inwardly through the feed channels  58 ,  158  and outwardly via the exhaust channels  60 ,  160 , rather than in the opposite direction. To effect this, the feed channels  60 ,  160  are provided with biassing means to direct the flow of cooling air in the desired direction. An example of such a biassing means is to angle the inlet slots or to make the cooling inlet slightly narrower than the exhaust. 
     There is thus described, a system for cooling the disc  40  of a turbine assembly, and also for cooling the blades  36  mounted on the disc  40 , which relies on a thermosiphon effect to drive the cooling air through the cooling passages. Advantages of the above described embodiments are that the air passing out of the second channels  62  in the blades  36  is used to provide annular sealing, which means that no additional air is required for cooling. Similarly, since the air is driven by a thermosiphon effect created by the rotation of the turbine blades, there is no net pumping power required. An additional advantage is that the flow of air tends to increase as the temperature of the blades increases which means that there is a degree of self modulation. 
     Various modifications can be made without departing from the scope of the invention. For example, the channels could be arranged in a different configuration to that shown in FIGS. 3 and 4. 
     The preferred embodiment of the invention has the advantage that air used for cooling is destined for annulus sealing. As a consequence, no additional cooling air is required. A further advantage of the preferred embodiment is that cooling air flow increases with blade temperature which allows a degree of self-modulation of the cooling. In addition, no net work is done in the preferred embodiment so that no net pumping power is required, and the air can be returned to its supply pressure, if desired. 
     Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.