Abstract:
An aerofoil blade or vane for a gas turbine engine comprises a body member having an inner end for mounting the blade on a shaft and an outer or tip end. A plurality of cooling passages are formed within the blade, the cooling passages comprising a plurality of inlet passages along which cooling air flows from the base towards the tip region of the blade and a plurality of return passages along which cooling air flows from the tip towards the base region of the blade. At least some of the passages are connected by a common chamber located within the tip region of the blade.

Description:
FIELD OF THE INVENTION 
   This invention relates to gas turbine aerofoil blades or vanes and is particularly concerned with the cooling of such blades or vanes. 
   BACKGROUND OF THE INVENTION 
   It is common practice to provide aerofoil blades or vanes for use in the turbines of gas turbine engines with some form of cooling in order that they are able to operate effectively in the high temperature environment of such turbines. Such cooling typically takes the form of passages within the blades or vanes which are supplied in operation with pressurised cooling air derived from the compressor of the gas turbine engine. 
   In such arrangements the cooling air is directed through passages in the blade or vane to provide convective and sometimes impingement cooling of the blade or vane&#39;s internal surfaces before being exhausted into the hot gas flogs in which the blade or vane is operationally situated. The cooling air may also be directed through small holes provided in the aerofoil surface of the blade or vane to supply a film of cooling air over the external surface of the aerofoil to provide film cooling of the aerofoil surface. 
   It is known to form such passages as one convoluted passageway which allows a length/diameter ratio to be utilised providing an acceptable degree of cooling efficiency. However, such a convoluted passageway necessarily requires bends which give rise to pressure losses without heat transfer. Also each bend requires a hole to be formed through which debris within the cooling air be exhausted. 
   SUMMARY OF THE INVENTION 
   According to the present invention there is provided an aerofoil blade or vane for a gas turbine engine comprising an elongated body member having an inner end or base by means of which the blade may be mounted on a shaft, an outer or tip end, and a plurality of cooling passages comprising a plurality of inlet passages along which cooling air flows from the base towards the tip region of the blade and a plurality of return passages along which cooling air flows from the tip towards the base region of the blade, at least some of said inlet and return passages being connected by a common chamber located within the tip region of the blade. 
   Preferably the aerofoil blade has a leading edge region and a trailing edge region wherein one of said passages is formed within the leading edge region of said blade and includes an opening at its radially inner end through which cooling fluid may be introduced into the passage. 
   Preferably at least one of said passages is in communication with the exterior of said blade to enable discharge of said cooling fluid from said blade. 
   Preferably at least one of the convex or concave walls of said blade is provided with an opening connected to the case of a cooling passage so as to provide an exhaust hole for cooling air. 
   Preferably said cooling passage is arranged to receive cooling fluid at its radially outer opening. 
   Preferably an exhaust outlet from said cooling passages is in communication with an adjacent vane or blade so as to direct cooling fluid to said adjacent blade. 
   Preferably said cooling fluid is air. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     An embodiment of the present invention will now be described by way of example only with reference to the accompanying drawings in which: 
       FIG. 1  is an illustrative view of part of a gas turbine engine; 
       FIG. 2  is a partial cross-section through a turbine blade; and 
       FIG. 3  is a cross-section on the line A—A of FIG.  2 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   With reference to  FIG. 1  a ducted fan gas turbine engine generally indicated at  10  comprises, in axial flow series, an air intake  12 , a propulsive fan  14 , an intermediate pressure compressor  16 , a high pressure compressor  18 , combustion equipment  20 , a high pressure turbine  22 , an intermediate pressure turbine  24 , a low pressure turbine  26  and an exhaust nozzle  28 . 
   The gas turbine engine  10  works in the conventional manner so that air entering the intake  12  is accelerated by the fan  14  to produce two air flows, a first air flow into the intermediate pressure compressor  16  and a second by-pass airflow which provides propulsive thrust. The intermediate pressure compressor  16  compresses the air flow directed into it before delivering the air to the high pressure compressor  18  where further compression takes place. 
   The compressed air exhausted from the high pressure compressor  18  is directed into the combustion equipment  20  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  22 ,  24  and  26  before being exhausted through the nozzle  28  to provide additional propulsive thrust. The high, intermediate and low pressure turbines  22 ,  24  and  26  respectively, drive the high and intermediate pressure compressors  16  and  18  and the fan  14  by suitable interconnecting shafts. 
   The high pressure turbine  22  includes an annular array of cooled aerofoil blades, one of which  30  can be seen in FIG.  2 . The aerofoil portion  32  of the blade  30  includes a learning edge region  34  and a trailing edge region  36  and is of generally hollow form provided with a series of internal bridging members  38 ,  40 ,  42 ,  44 ,  46  and  48  which extend from the concave suction side  50  to the convex pressure side  52  of the aerofoil. A blade platform  53  extends outwardly from the aerofoil portion  32  of the blade  30 . 
   The bridging member  38  in the leading edge region of the blade  30  extends substantially the full radial length of the blade  30  but does not reach the tip portion  54  of the blade. The radial length of the blade  30  is that length which extends radially outwardly from the root portion to the tip portion of the blade  30  when arranged as one of any array of blades positioned circumferentially around the appropriate gas turbine engine shaft. Thus a gap is formed between the end  56  of the bridging member  38  and the tip  54  of the blade. 
   Similarly a gap is formed in the tip portion  54  of the blade as the bridging members  40 ,  42 ,  44  and  46  extend a shorter radial length than bridging member  38 . 
   A hole  66  is provided in the tip  54  of the blade  30  and provides an exit for dust particles and debris which may be carried by the cooling air as it passes through the blade  30 . 
   The bridging members divide the hollow interior of the blade  30  into a plurality of passages or channels  68 ,  70 ,  72 ,  76 ,  77 ,  78  and  84  through which cooling air may flow. 
   The bridging members  40  and  42  are formed as a pair extending radially outwardly from a shank portion  58 . Similarly the bridging members  44  and  46  also extend from a shank portion  60  located at the base  62  of the blade  30 . The bridging member  48  adjacent the trailing edge  36  of the blade  30  also extends radially outwardly from a shank portion  64 . 
   Outlet apertures  74  and  75  are formed at the radially inner ends of the passages  72  and  77  to allow cooling air to be exhausted to the mainstream airflow. 
   In operation, the interior of the blade  30  is supplied with a flow of cooling air derived from the gas turbine engine compressor. This cooling air is directed into the channels  68 ,  70 ,  76  and  78 . The direction of the cooling air flow through the blade  30  is shown by arrows C. The cooling air entering channel  68  may be partly exhausted through apertures in the aerofoil wall to form a cooling film on the exterior of the aerofoil. The remainder of the air flows radially outwardly over the tip  56  of bridging member  38  and combines with flow directed into channel  70  to provide impingement cooling of the underside of the blade tip  54 . The cooling air is then directed radially inwardly into the passage  72  located between the bridging members  40  and  44  and is discharged through outlet aperture  74  into a zone beneath the blade platform  53 . 
   Similarly cooling air directed into the channels  70 ,  76  and  78  provides impingement cooling of the undersurface of the tip portion  54  and is subsequently directed radially inwardly into channels  72  and  77  and exhausted between shanks under the blade platforms  53  via exhaust outlets  74  and  75 . The cooling air from channel  78  reaches the passage  84  through holes  80  and  82  located in the radially outer portion of the bridging member  48 . This provides cooling of the trailing edge portion of the blade which requires greater cooling than the remainder of the blade. 
   The air entering the region between the shanks is exhausted into the passage  84  through an aperture  90 , cooling the rear of the aerofoil and the platforms  53 . Air from passage  84  is exhausted through the aerofoil wall to provide film cooling. The holes  80  and  82  limit the temperature at the tip of this passage. 
   The passageways and chambers formed by the bridging members allow cooling air to flow through the internal region the blade  30  and provide impingement cooling of the underside of the blade tip  54 . 
   Advantageously, the region  86  of the hollow interior of the blade defines a chamber into which cooling air from the channels  68 ,  70 ,  76  and  78  is directed. This provides cooling of the blade tip  54  by impingement cooling of its inner surface. As the bridging members  40 ,  42 ,  44  arid  46  are foreshortened to define the chamber  86  there is a saving in weight compared with convoluted converted passage arrangements and the disadvantages associated with the bends in convoluted passage arrangements are avoided. Pressure losses are minimised due to the lack of bends and thus the pressure of the cooling air remains relatively high compared to prior art systems which utilise convoluted passageways. 
   Various modifications may be made without departing from the invention. Thus, for example, the cooling air could be used to provide film cooling through film cooling holes located across the external blade surface if required. 
   It is also envisaged that the return channels  72 ,  77  and  84  may be connected to an adjacent vane or blade so as to exhaust cooling air into the adjacent vane or blade. 
   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 nor particular emphasis has been placed thereon.