Abstract:
A gas turbine engine component, typically either a turbine blade or vane or combustor, comprising a wall ( 40 ) with a first surface ( 39 ) which is adapted to be supplied with a flow of cooling air, and a second surface ( 38 ) which is adapted to be exposed to a hot gas stream ( 50 ). The wall ( 40 ) further having defined therein a plurality of passages ( 57 ), the passages ( 57 ) defined by passage walls ( 54 ), which interconnect a passage inlet ( 31 ) in said first surface ( 39 ) to a passage outlet ( 32 ) in said the second surface ( 38 ). The passages ( 57 ), cooling air and the hot gas stream ( 50 ) arranged such that in operation a flow ( 52 ) of cooling air is directed through said passages ( 57 ) to provide a flow ( 36 ) of cooling air over at least a portion of the second surface ( 39 ). The cross sectional area of each of the passages ( 57 ) progressively decreasing overall, in the direction of cooling air flow ( 52 ) through the passage ( 57 ), such that in use the flow of cooling air ( 52 ) through the passage ( 57 ) is accelerated. The passage walls ( 54 ) of the cooling passages ( 57 ) preferably diverging laterally across the wall ( 40 ) of the component whilst perpendicular to the wall ( 40 ) they converge so that overall the cross-sectional area decreases.

Description:
THE FIELD OF THE INVENTION 
     The present invention relates generally to cooling arrangements for gas turbine components and in particular to improvements to the arrangement and configuration of cooling passages which are provided within the walls of a component and are arranged to provide film cooling of the component. 
     BACKGROUND OF THE INVENTION 
     Certain components, in particular in the combustor and turbines, of a gas turbine engine are subject, in operation, to high temperature gas flows. In some cases the high temperature gas flows are at temperatures above the melting point of the component material. In order to protect the components, and in particular the surface of the components adjacent to the high temperature gas flows, from these high temperatures, various cooling arrangements are provided. 
     Generally such arrangements utilise relatively cool compressed air, which is bled from the compressor section of the gas turbine engine, to cool and protect the components subject to the high operating temperatures. 
     A well known method of cooling and protecting gas turbine components from the high temperature gas flows is film cooling in which a film of cooling air is provided along the surface of the component exposed to the high temperature gas flows. The film of cooling air is produced by conducting a flow of cooling air through a plurality of passages which perforate the wall of the component. The air exiting the passages is directed, by the passages, to flow in a boundary layer along surface of the component. This cools the wall of the component exposed to the high temperature gas flow and provides a protective film of cool air between the high temperature gas flow and the component surface. The protective film assists in keeping the high temperature gas flow away from the surface of the component wall. 
     The arrangement and configuration of the passages are carefully designed to provide, and ensure, an adequate boundary layer flow of cooling air along the surface of the component. The passages are accordingly generally angled in the flow direction of the hot gas stream so that the cooling air flows in a downstream direction over the surface of the component. 
     Ideally it is desired that the boundary layer should flow over substantially the entire surface of the component downstream of the passages. However it has been found that the cooling air leaving the passage exit generally forms a cooling stripe no wider than, or hardly wider than, the dimension of the exit of the passage. Limitations on the number, size, and spacing of the passages results in gaps in the protective cooling layer provided and/or areas of reduced protection/cooling. 
     To overcome this it has been proposed, in for example U.S. Pat. No. 3,527,543, to use divergent passages where the cross section of the passages increases towards the passage exit at the surface of the component exposed to the hot gas flow. The cooling air which flows through the passages is thereby partially spread out over a larger area of the surface. Whilst this is an improvement over a constant cross section passage it has been found that the air exiting the passage generally still does not spread out enough to provide a continuous film of cooling air between the typical spacing of the passages. 
     A further development of the diverging passages is to arrange the passages sufficiently close to each other such that the outlets of the adjacent passages, on the surface of the component exposed to the hot gas flows, intersect laterally to define a common outlet in the form of a laterally extending slot. The cooling air expands as it passes though the passages and exits from this common slot as a substantially continuous film. Such an arrangement is described more fully in U.S. Pat. No. 4,676,719 which also references other similar arrangements which are described in U.S. Pat. No. 3,515,499 and Japanese Patent Number 55-114806. 
     In these prior art arrangements the passages are divergent and the cross sectional area of the passage increases towards the exit. This slows down, and diffuses, the flow of cooling air therethrough. As is taught in the prior art this slowing of the flow is important in assisting in spreading the flow of cooling air, in a boundary layer, along and over the surface of the component. Another important consideration in the design of such film cooling arrangements is to ensure that a stable boundary layer is provided over the surface of the component, and that this boundary layer remains attached to the surface of the component to thereby protect the surface from the high temperature gas stream. This boundary layer flow of cooling air is also required to withstand fluctuations and variations in the hot gas stream, that may occur during operation, to ensure that adequate cooling and protection is provided throughout the operation of the engine. In addition the flow through the passages and along the surface of the component should be as aerodynamically efficient as possible. 
     In an additional variation slots within the walls of the component can be used to direct the cooling air to the outer surface of the component. Such an arrangement is described in U.S. Pat. Nos. 2,149,510, 2,220,420 and 2,489,683. 
     Although such arrangements provide a good flow of cooling air along and over the surface of the component the structural strength of the walls of the component is reduced. This is also true, albeit to a lesser extent, with the arrangements where the passages intersect at their exits to form a common exit slot. 
     It is therefore desirable to provide an improved gas turbine engine component cooling arrangement and configuration, and in particular to provide an improved arrangement and configuration of cooling passages that address the above mentioned problems and/or offers improvements to such cooling arrangements generally. 
     SUMMARY OF THE INVENTION 
     According to the present invention there is provided a gas turbine engine component comprising a wall with a first surface which is adapted to be supplied with a flow of cooling air, and a second surface which is adapted to be exposed to a hot gas stream, the wall further has passage walls which define therein a plurality of passages, which interconnect passage inlets in said first surface of the component to passage outlets in said the second surface, the passages, passage walls defining the passages, cooling air and the hot gas stream arranged such that in operation a flow of cooling air is directed from the passage inlets to the passage outlets through said passages to provide a flow of cooling air over at least a portion of the second surface; wherein a cross sectional area of each of the passages in a direction of cooling air flow through a passage, progressively decreases overall from the passage inlets to the passage outlets such that in use the flow of cooling air from the passage inlets to the passage outlets through each passage is accelerated. 
     Preferably the passage outlet in said second surface comprises a slot defined by the passage in said second surface. The passage inlet in said first surface preferably has a different shape to the passage outlet slot. 
     The passage outlets of at least two of the plurality of passages may be combined to produce a common outlet. 
     Preferably at the passage outlet of at least two adjacent passages, at least part of the passage walls defining the adjacent passages substantially intersect the second surface of the wall exposed to the hot gas stream. 
     The cross section, substantially perpendicular to the direction of flow through the passage, of the passage inlet may be substantially circular or elliptical or rectangular 
     Preferably the passage walls, which define the passages through the walls of the component, are profiled such that in a first direction substantially perpendicular to a cooling flow direction through the passage they converge towards a centre line through the passage, and in a second direction also perpendicular to a flow direction through the passage they diverge from the centre line of the passage. Furthermore the first direction in which the passage walls diverge may be substantially parallel to the first and second surfaces of the wall of the component, and the second direction may be substantially perpendicular to the first direction and the centre line through the passage, such that from the passage inlet to the passage outlet the passage walls that define the passages are configured to diverge in the first direction laterally across the wall of the component and also simultaneously converge in the second direction. 
     The passages through the walls of the component may be angled in a flow direction of the hot gas stream that is arranged in operation to flow adjacent to the second surface of the component. 
     Preferably at the passage inlets, where the walls of the passages and the first surface of the wall of the component intersect, a rounded profile is defined between the passage walls and the first surface. Furthermore at the passage outlets, where the walls of the passages and the second surface of the wall of the component intersect, a rounded profile is defined between the passage walls and second surface. 
     A portion of the second surface of the wall exposed to hot gas stream downstream of a passage outlet may be lower than a portion of the second surface upstream of the passage outlet. 
     The passages may be curved as they pass through the wall of the component. The passage walls that define the passages may have a curved profile. 
     The component is part of a turbine section of a gas turbine engine. Furthermore the component may be a hollow turbine blade or a hollow turbine vane. 
     Alternatively the component is part of a combustor section of a gas turbine engine. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The present invention will now be described by way of example with reference to the following figures in which: 
     FIG. 1 shows a schematic illustration of a gas turbine engine; 
     FIG. 2 is an illustration of a turbine blade from the engine shown in FIG. 1 incorporating an embodiment of the present invention; 
     FIG. 3 is a cross sectional view of the turbine blade shown in FIG.  2  through line  3 — 3 ; 
     FIG. 4 is a more detailed view of the wall of the turbine blade of FIG. 3 showing a coolant passage therethrough; 
     FIG. 5 a  is a view on arrow A of FIG. 4; 
     FIG. 5 b  is a sectional view of the wall of the turbine blade on a plane passing through the centreline  5 A— 5 A of the passage of FIG. 4; 
     FIG. 6 is a similar view to that of FIG. 4 but of an alternative embodiment of the present invention; 
     FIG. 7 is a sectional view of the wall of the turbine blade on a plane passing though the centreline  66  of the passage of FIG. 6; 
     FIG. 8 is a similar view to that of FIG. 4 but of another alternative embodiment of the present invention; 
     FIG. 9 is a similar view to that of FIG. 4 but of a further embodiment of the present invention; 
     FIG. 10 is a similar view to that of FIG. 4 but of a yet further embodiment of the present invention; 
     FIG. 11 is a sectional view of the wall of the turbine blade on a notional surface passing through the centreline  10 — 10  of the passage of FIG.  10 . 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to FIG. 1 an example of a gas turbine engine comprises a fan  2 , intermediate pressure compressor  4 , high pressure compressor  6 , combustor  8 , high pressure turbine  9 , intermediate pressure turbine  12  and low pressure turbine  14  arranged in flow series. The fan  2  is drivingly connected to the low pressure turbine  14  via a fan shaft  3 ; the intermediate pressure compressor  4  is drivingly connected to the intermediate pressure turbine  12  via a intermediate pressure shaft  5 ; and the high pressure compressor is drivingly connected to the high pressure turbine via a high pressure shaft  7 . In operation the fan  2 , compressors  4 , 6 , turbine  9 , 12 , 14  and shafts  3 , 5 , 7  rotate about a common engine axis  1 . Air, which flows into the gas turbine engine  10  as shown by arrow B, is compressed and accelerated by the fan  2 . A first portion of the compressed air exiting the fan  2  flows into and within an annular bypass duct  16  exiting the downstream end of the gas turbine engine  10  and providing part of the forward propulsive thrust produced by the gas turbine engine  10 . A second portion of the air exiting the fan  2  flows into and through the intermediate pressure  4  and high pressure  6  compressors where it is further compressed. The compressed air flow exiting the high pressure compressor  6  then flows into the combustor  8  where it is mixed with fuel and burnt to produce a high energy and temperature gas stream  50 . This high temperature gas stream  50  then flows through the high pressure  9 , intermediate pressure  12 , and low pressure  14  turbines which extract energy from the high temperature gas stream  50 , rotating the turbines  9 , 12 , 14  and thereby providing the driving force to rotate the fan  2  and compressors  4 , 8  connected to the turbines  9 , 12 , 14 . The high temperature gas stream  50 , which still possesses a significant amount of energy and is travelling at a significant velocity, then exits the engine  10  through an exhaust nozzle  18  providing a further part of the forward propulsive thrust of the gas turbine engine  10 . As such the operation of the gas turbine engine  10  is conventional and is well known in the art. 
     It will be appreciated that in operation the combustor  8  and the turbines  9 , 12 , 14 , in particular the high pressure turbine  9 , are subjected to the high energy and temperature gas stream  50 . In order to improve the thermal efficiency of the gas turbine engine  10  it is desirable that the temperature of this stream  50  is as high as possible, and in many cases may be above the melting point of the engine  10  materials. Consequently cooling arrangements are provided for these components subjected to these high temperatures, to protect these components. 
     The turbines  9 , 12 , 14  comprise a plurality of blades mounted in an annular array from a disc structure. One of these individual turbine blades  20  from the high pressure turbine  9 , which is subject to the high energy and temperature gas stream  50  is shown, diagramatically, in FIG.  2 . The blade  20  comprises an aerofoil section  22 , a platform section  24 , and a root portion  26 . When the blade  20  is mounted within the engine  10  the aerofoil section  22  is disposed within, and exposed to, the high temperature gas stream  50 . The platform section  24  co-operates with the platform sections  24  of the other blades  20  within the array to define an annular inner ring structure which defines part of an annular turbine duct  25  through which the gas stream flows. This annular turbine duct  25  is shown by phantom lines  25 ′ in FIG.  2 . The root portion  26  attaches the turbine blade  20  to a turbine disc. 
     As shown in FIG. 3 the turbine blade  20  is hollow, with an outer wall  40  enclosing, and defining, a compartmentalised internal cavity  34 . Passages  28 , 30  within the turbine blade root  26  interconnect the internal cavity  34  with cooling air ducts (not shown) in the engine  10 . In operation pressurised cooling air, which is conventionally bled from the compressors  4 , 6  (primarily the high pressure compressor  6 ) is supplied via the engine cooling ducts and the turbine blade root passages  28 , 30  to the internal cavity  34  of the turbine blade  20 . The pressurised cooling air cools the walls  40  of the turbine blade  20  and flows through, as shown by arrows  52  and  36 , passages  57  provided within the walls  40 . This flow  36  of cooling air exiting the passages  57  flows in a boundary layer, in a downstream direction, along the surface  38  of the turbine blade  20  exposed to the high temperature gas stream  50 . The boundary layer of cooling air provides a protective film of cool air along the surface  38  of the blade  20  and provides film cooling of the blade surface  38  exposed to the high temperature gas stream  50 . 
     It will be appreciated that in a typical turbine blade there may be a number of passages  57 , generally in rows, within the entire extent of walls  40  of the blade  20  on both a suction side and pressure side of the blade  20  and at the leading and trailing edges of the blade  20 . However for the purposes of clarity and simplification only one such row of passages  57  has been shown. 
     The configuration and shape of the passages  57  is shown in more detail in FIGS. 4,  5   a , and  5   b . A plurality of discrete inlets  31  are provided in the surface of the wall  40  adjacent to cavity  34 . The inlets  31  are arranged in a row extending (spanwise) along the length of the blade  20 . The individual passages  57 , which are defined by passage walls  54 , extend through the walls  40  of the blade  20  from the inlet  31  to an outlet  32  in the surface  38  of the wall  40  exposed to the high temperature gas stream  50 . 
     A central axis  58  passes through the geometric centre of each of the passages  57 , and, as shown, the passages  57  are angled in the direction of the flow of the high temperature gas stream  50 . In operation this angling directs the flow  36  of cooling air, as it exits the passages  57 , in a downstream direction along the surface  38  of the blade  20 . The angle  0  of the central axis  58 , and so of the passages  57 , to the wall surface  39  is typically between  20  and  70  degrees. 
     The inlet  31  to the passages  57  has a substantially circular cross section in the flow  52  direction (perpendicular to the central axis  58 ). It being appreciated that due to the angle θ of the passage  57  relative to the wall surface  39 , as shown by the central axis  58 , a circular cross section inlet  31  forms an elliptical hole in the wall surface  39 , as shown in FIGS. 5 a  and  5   b.    
     The walls  54  of the passages  57  define the passages  57  as they pass through the wall  40  of the blade  20  as shown in FIGS. 4, and  5   a . As shown in FIG. 5 a , which is a view on arrow A of the surface  38  of the wall  40 , from the passage inlet  31  to the outlet  32  on the wall surface  38  the walls  54  of the individual passages  57  diverge laterally within the wall  40  in a direction generally parallel to the wall surfaces  38 , 39 . At or near the blade wall surface  38  the walls  54  of adjacent passages  57  intersect to define a common outlet slot  32  in the wall surface  38 . This outlet slot  32  is most clearly seen in FIG.  2 . In a cross sectional plane through the wall  40  from the cooling air surface  39  of the wall to the exposed surface  38  of the wall, and containing the passage central axis  58 , the walls  54  however converge on the central axis  58  from the inlet  31  to the outlet  32 , as shown in FIG.  4 . From the inlet  31  to the outlet slot  32  the walls  54  of the passages  57  therefore diverge in one direction (laterally) whilst also converging in a second substantially orthogonal direction (substantially perpendicular to the wall surfaces  38 , 39 ). 
     The cross section of the passages  57  in the flow direction  52  through the passages is generally circular at the inlet  31 . Then, as the passage  57  passes through the wall  40 , and due the profiling of the walls  54 , the cross section is smoothly developed into a generally rectangular shape, in the form of a common outlet slot  32 , at the passage outlet. It will be appreciated though that the inlet  31  cross section is not critical and the inlet  31  could be elliptical, circular, rectangular or any other shape. 
     The profiling of the passage walls  54  is such that the convergence of the walls  54  (as shown in cross sectional side view in FIG. 4) is greater than the divergence of the walls  54  (as shown in plan view in FIG. 5 a ). Therefore overall the configuration of the passages  57  converges and the cross sectional area of the passages  57  reduces, in the flow  52  direction, from the inlet  31  to the outlet  32 . 
     As shown in FIG. 5 b  and  5   a  inside the wall  40  adjacent passages  57  are separated by roughly triangular pedestals  55 , defined in part by the passage walls  54 . These pedestals  55  tie the walls together and maintain the strength of the wall  40 . This provides mechanical strength superior to a simple slot arrangement. 
     Preferably the basic shape of each of the passages  57  is generated by a family of straight lines passing through the wall  40  in a similar way to the central axis  58 . As such the passages can be manufactured by linear drilling, for example by using a laser. Other conventional methods could however be used to manufacture the passages. For example they could also be produced by electrode discharge machining or water jet drilling. Alternatively the walls  40  and cooling passages  57  could be manufactured by precision casting. 
     In operation cooling air within the cavity  34  flows into the passage inlet  31  and through the passages  57  defined by the passage walls  54 , as shown by arrow  52  in FIG.  4 . As the cooling air flows through the passages  57 , defined by the laterally diverging walls  54 , it spreads out laterally. At the outlet  32  the cooling air is combined, within the common outlet slot  32 , with cooling air flow  36  from adjacent passages  57  such that the cooling air flow  36  exits the outlet slot  32  as a film of cooling air extending along the length L of the slot  32 . Due to the shallow angle θ of the passages  57 , relative to the wall surface  38 , and the flow of the high temperature gas stream  50  along the surface of the wall  38 , the film of cooling air flow  36  exiting the outlet slot  32  flows downstream along the surface  38  in a boundary layer. This boundary layer along the surface  38  provides the required film cooling of the surface  38  and protection of the surface  38  from the high temperature gas stream  50 . As such the flow  52 , 36  through and out of the passages  57  is similar to other prior art arrangements in which cooling air flows through a slot outlet to provide a boundary layer film. 
     However according to the invention, due to the combined overall convergence and reduction in overall cross sectional area of the passages  57 , between the inlet  31  and outlet  32 , the cooling air flow  52 , 36  is accelerated as it flows through the passages  57 . The minimum throat area of the passages  57  and hence the maximum flow velocity is preferably arranged at or just before the passage outlet  32 . This acceleration of the cooling air flow through the passages  57  due to the reduction in overall cross section is an important aspect of the invention. Such an arrangement being completely against the teaching of conventional cooling passage designs which are arranged to decelerate the flow through passages which only have overall divergent and increasing cross sectional area passages. 
     It has been found that accelerating the cooling air flow  52 , 36  as it flows through the passages  57  has a number of advantages. Firstly it minimises inlet flow separations that can occur with prior art designs where the flow is decelerated. It also minimises the aerodynamic losses associated with flow  52 , 36  through the passages  57  and/or allows higher cooling air flows  52 , 36  without additional aerodynamic performance penalties, as compared to the prior art arrangements that decelerate the cooling air flow  52 , 36 . Additionally by accelerating the flow  52 , 36  of the cooling air through the passages  57  an improved, near laminar and relatively thin boundary layer film flow  36  of cooling air is provided along the surface  38  of the blade  20 . This boundary layer, produced by this arrangement, is more stable, and the cooling air flow  36  at the outlet  32  is less turbulent than that produced in the prior art methods. This inhibits mixing of the cooling air flow  36  along the surface  38  with the high temperature gas stream  50  which improves film cooling and provides an improved protective barrier over the surface  38  of the blade  20 . The overall convergence and reduction in cross section of the passages  57  also improves the lateral distribution and spreading out of the cooling air flow  52 , 36  within the passages  57  to produce a near uniform, or more uniform, cooling film across the length L of the outlet slot  32 . The arrangement according to the invention also combines these benefits with those of a slot type outlet, and/or passage, in which the cooling air flow is spread out over the surface  38  of the blade  20 . 
     In this arrangement the outlet flow  36  from the passage outlet slot  32  is also kept on the surface  38  of the wall by the Coanda Effect which is also improved by accelerating the cooling air flow  36 . This reduces the tendency of the outlet flow  36  to lift off from the surface  38  of the blade  20 , which can occur with other arrangements. Such lift off of the flow over the surface  38  of the blade  20  adversely affects the film cooling of, and protection provided to, the blade wall  40 . Consequently this arrangement can be used with higher flow rates of cooling air which provide improved film cooling. Such higher cooling air flow rates are difficult to provide with prior art arrangements due to the tendency of the flow produced along the walls to lift off. 
     Further embodiments of the invention are shown in FIGS. 6 to  11 . These embodiments are generally similar to the embodiment described in detail above. Consequently only the differences between these embodiments and the above arrangement will be described, and like reference numerals have been used for like features. Furthermore although the additional individual features of the successive embodiments have been combined in FIGS. 6 to  11  it is contemplated that they can be used separately or in different combinations in other further embodiments. 
     In a second embodiment of the invention as shown in FIGS. 6 and 7 the inlet  31   a  to the passages  57   a  has a rounded profile. This further minimises inlet flow separations and further improves the aerodynamic efficiency of this arrangement. 
     As shown in the embodiment illustrated in FIG. 8 the outlet slot  32   b  can also be faired or rounded into the surface of the wall  38 . This reduces any exit separations of the cooling air flow  36 . Furthermore such rounding of the outlet slot  32   b  improves the Coanda effect associated with the outlet  32   b  which further reduces any tendency of the outlet flow  36  to lift off from the surface  38 . 
     In the embodiment shown in FIG. 9 the surface  38 ″ of the wall exposed to the high temperature gas stream  50  downstream of the outlet slot  32   c  is lower than the surface  38  upstream of the outlet slot  32   c . The extended position of the upstream surface  38  being shown by phantom line  38 ′. The distance d between the downstream surface  38 ″ and the position of extended surface  38 ′ is preferably equal to the displacement thickness which would accommodate the cooling flow  36  without disturbing the main flow  50 , ignoring mixing, caused by the flow  36  of cooling air flow from the outlet  32   d . By this arrangement the high temperature gas stream  50  is less disturbed by the flow  36  of cooling air from the outlet  32   d  and along the surface  38 ″ of the wall  40  while maintaining the high cooling effectiveness of the cooling near to the wall  40 . This arrangement is particularly advantageous if the high temperature gas stream  50  is flowing over the surface  38  at a high Mach number, and hence velocities, where the arrangement reduces loss inducing shock waves which may be generated by the flow  36  of cooling air from the outlet  32   c.    
     In the embodiment shown in FIG. 10 and 11 the passages  57   d  still have a laterally divergent profile in one direction (FIG.  11 ), and a convergent profile in another direction (FIG.  10 ), with the overall cross section converging and reducing towards the passage outlet  32   d  such that the cooling flow is accelerated through the passage  57   d . However the walls  54   d , and profiling of the passages  57   d  through the wall  40  are curved rather than straight sided as in the previous embodiments. The passage  57   d  is also curved as it passes through the wall  40  as shown by the curved, notional, central axis  58  of the passage  57   d . This curved profiling improves the flow  52  of cooling air through the passages  57   d . Furthermore by curving the passages  57   d , as shown by the notional central axis  58 , the angle θ of the passage outlet  32   d  relative to the wall surfaces  38  can be reduced as compared to the case with straight walled passages  57 . This improves the flow  36  of cooling air film along the downstream wall surface  38 ″ and further reduces any tendency of the film to lift off the surface  38 ″. In this embodiment the basic shape of the passages  57   d  is no longer generated by a family of straight lines, as is generally the case in the previous embodiments, and the passages  57   d  and walls  40  are typically manufactured by precision casting to achieve the curved profile. It being appreciated that other conventional methods of producing the passages are generally not applicable to producing such curved passages  57   d.    
     Although not shown it will also be appreciated that the cross section and height h of the outlet slot  32   d  can be varied along its length L, and in particular across each passage L 1  in order to improve the lateral distribution of the cooling flow  36  over the surface  38 ″. 
     The invention has been described with reference to cooling turbine blades  20 . It will be appreciated though that the invention can also be applied to, and used on, the nozzle guide vanes of a turbine to provide improved cooling to the surfaces and walls of the vanes similarly exposed to the high temperature gas stream  50 . Such nozzle guide vanes having a similar aerofoil and platform sections and also generally being hollow with an internal cavity defined by vane walls. Cooling air being supplied to the internal cavity of the vanes and passing through cooling passages within the vane walls thereby providing cooling and protection of the vanes. 
     It will further be appreciated and contemplated by those skilled in the art that the cooling passage arrangement and configuration could also equally well be applied to other components which are required to be film cooled. For example the walls of the combustor are conventionally provided with film cooling and the invention can be advantageously applied to providing film cooling of such combustor walls.