Abstract:
A gas turbine engine turbine blade ( 20 ) has cooling air holes ( 38 ) arranged in groups, the holes ( 38 ) in one group and which span that part of the leading edge ( 34 ) that spans the hottest part of the blade ( 20 ), are more closely spaced than the remainder of the holes ( 38 ), thereby ensuring the provision of the most cooling air, where it is most needed.

Description:
FIELD OF THE INVENTION 
   The present invention relates to turbine blades of the kind used in gas turbine engines, wherein the operating temperatures are such as to require that the turbine blades be provided with a flow of cooling air around their leading edges, in order to maintain their structural integrity. 
   BACKGROUND OF THE INVENTION 
   It is known to form a turbine blade with interior compartments, to which relatively cool air from a compressor of an associated gas turbine is fed, and to provide holes in the blade leading edge portion, which holes connect one of those compartments in cooling air flow series with the blade leading edge surface. 
   It is also known to arrange the holes described hereinbefore in one or more rows, the or each hole being lengthwise of the blade, ie substantially normal to the axis of the associated engine, when the blade is in situ therein, the holes being equally spaced. Further it is known to form the holes so that when the blade is in situ in the engine, the holes axes and engine axis define respective acute angles, such that the air flow through the holes has a directional component radially outwardly of the engine axis. 
   The known art fails to properly address the cooling needs of cooled turbine blades, having regard to the temperature gradients along their leading edges, and further as a consequence, remove more air than is strictly necessary from the engine system, thus reducing overall engine efficiency. 
   SUMMARY OF THE INVENTION 
   The present invention seeks to provide an improved air cooled turbine blade. 
   According to the present invention an air cooled gas turbine engine turbine blade is provided with an internal compartment for the receipt of cooling air, and cooling air exit holes which connect said compartment in flow series with the leading edge surface of said blade, said exit holes being arranged in one or more rows lengthwise of the blade, and those holes spanning that portion of the blade leading edge that experiences the most heat being more closely spaced than the remainder thereof. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention will now be described by way of example and with reference to the accompany drawings in which: 
       FIG. 1  is a diagrammatic view of a gas turbine engine including turbine blades in accordance with the present invention. 
       FIG. 2  is a graphic sketch of a typical temperature gradient over the leading edge of a turbine blade in situ in an operating gas turbine engine. 
       FIG. 3  is a view on line  3 — 3  of FIG.  4 . 
       FIG. 4  is a development view on line  4 — 4  of FIG.  3 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   Referring to  FIG. 1  a gas turbine engine  10  has a compressor  12 , combustion equipment  14 , a turbine section  16 , and an exhaust pipe  18 . Turbine section  16  includes a stage of turbine blades  20  mounted on a disk  22 , for rotation in known manner, on receipt thereby of a flow of hot combustion gases from the combustion equipment  14 . 
   Referring briefly to  FIG. 4  each turbine blade  20  contains a compartment  24  which in the present example includes a pair of wall structures  26  and  28 , which provide a serpentine flow path for a flow of cooling air from compressor  12 . The air enters the compartment  24  via a hole  30  in the root portion  32  of blade  20 , in known manner. 
   Referring now to  FIG. 2  the temperature gradient along the leading edge  34  of a turbine blade is generally of the form depicted by the parabolic line  36  and clearly shows that the maximum temperature is experienced at about half way along the leading edge  34 . Thereafter, the temperature reduces on both sides of the half length of the leading edge  34 , to respective intersection points A and B. The leading edge portion of the blade which should be regarded as typically blade  20  that needs most cooling air, is thus clearly defined as being between points A and B. 
   Referring to  FIG. 3  the last portion  36  of compartment  24  to receive the cooling air flow, in the present example, is connected to the gas flow duct of turbine section  16  ( FIG. 1 ) via two rows of holes  38  and  40 , the rows being positioned side by side along the leading edge  34  of the blade  20 , ie into and out of the plane of the drawing. 
   Referring to  FIG. 4  in this view in which only the centrelines of holes  38  are shown for reasons of clarity, a large proportion of holes  38  are closely spaced over that portion of blade  20  that corresponds to portion A-B in  FIG. 2 , whereas only three more widely spaced holes  38  are provided near the upper end of blade  20 , and only one hole  38  is provided in wide spaced relationship with the closely spaced holes at the lower end of blade  20 . By this means, cooling air flow holes  38  (and  40 ) in a manner which ensures that the whole length of the leading edge of blade  20  receives the quantity of cooling air appropriate to the temperature it experiences. 
   The closely spaced holes  38  are aligned with respect to the engine axis, such that their axes define a large, acute angle therewith, and their cooling air outlet ends are radially further outwardly of the engine axis than their inlet ends. Their angular attitude results in them having to pass through greater thickness of blade metal than if they were aligned with the gas flow over blade  20 . A benefit is derived from the arrangement in that the hot metal heats the air flowing through the holes  38 , and generates a convection flow, ie it speeds up the air flow. 
   The three widely spaced holes  38  also have an angular attitude with respect to the axis of engine  10 , which attitude however, is of smaller magnitude. The benefit derived is that the air flow has shorter, and therefore a quicker passage to reach the leading edge  34  and consequently is not so exposed to the convection affects of the hot metal. Therefore on reaching the leading edge  34 , the air flow is cooler and though less in quantity, is sufficient to achieve the desired cooling of the outer end portion of the leading edge  34  of blade  2 . 
   The arrangement of holes  38  in groups, some closely spaced and others more widely spaced, along the leading edge  34  of a turbine blade  20 , as described hereinbefore has been shown on a test rig to achieve a reduction of 100° C. in the maximum temperature. 
   Whilst the embodiment of the present invention described hereinbefore is the preferred embodiment, the expert in the field having read this specification will appreciate that the grouping of the cooling air holes  38  in a manner appropriate to the temperature gradient on blade  20  provides the main contribution to the improvement, some improvement over the prior art referred to in this specification can be achieved by varying the angular relationship of the holes  38  relative to the engine axis, in ways that differ from those described herein with respect to the accompanying drawings. Even to the extent of aligning the groups of holes  38  with the axis of engine  10 . Such an arrangement would reduce the difference in convective affect between the groups of holes  38  but this could be offset by the provision of more holes  38  near the end extremities of blade  20 .