WO2017125289A1

WO2017125289A1 - Aerofoil arrangement

Info

Publication number: WO2017125289A1
Application number: PCT/EP2017/050449
Authority: WO
Inventors: Anthony Davis
Original assignee: Siemens Aktiengesellschaft
Priority date: 2016-01-19
Filing date: 2017-01-11
Publication date: 2017-07-27

Abstract

An aerofoil (100) for a gas turbine (10) comprises: - an external surface (128) extending along a first axial direction between a leading edge (133) and a trailing edge (135) and extending along a second radial direction between a base (134) and a tip (136) of the aerofoil (100); - an internal passage (150) for channelling radially a cooling medium from the tip (136) towards the base (134), the internal passage (150) being shaped in such a way that at the base (134) the cooling medium exits the aerofoil (100) according to a flow direction (F2) having at least a component orthogonal to the radial direction.

Description

DESCRIPTION

AEROFOIL ARRANGEMENT

Field of invention

The present invention relates to guide vanes for gas turbines and, more particularly, to guide vanes for gas turbines comprising aerofoils provided with internal cooling passages.

Art Background

In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot

combustion gases. The hot gases are then channelled towards a turbine which transforms the energy from the hot gases into work for powering the compressor and other devices which converts power, for example an upstream fan in a typical aircraft turbofan engine application, or a generator in power generation application.

The turbine stages include stationary turbine nozzles having a row of vanes which channel the combustion gases into a corresponding row of rotor blades extending radially

outwardly from a supporting rotor disk. The vanes and blades may have corresponding hollow aerofoils. Aerofoils may be designed and manufactured hollow in order to save weight, to change its eigenfrequency or to include a cooling circuit therein. In the latter case, the cooling gas which circulates inside the cooling circuits is typically bleed air from the compressor discharge.

The present invention relates to a turbine hollow aerofoil which can be used in a gas turbine guide vane and which include an internal passage for cooling the aerofoil.

In the following, the terms upstream and downstream refer to the flow direction of the airflow and/or working hot gas flow through the gas turbine engine. The terms axial and radial are made with reference to a rotational axis of the gas turbine engine.

Each aerofoil includes a generally concave pressure sidewall and, an opposite, generally convex suction sidewall extending radially outwardly along a span from an aerofoil base to an aerofoil tip and axially in a chordwise direction between a leading and a trailing edge. For a gas turbine nozzle guide vane, the aerofoil extends from a root base integral with a radially inner platform of a turbine stator to a radially outer tip integral with an outer platform of the outer casing of the stator.

When it is required to cool a platform region, for example the inner platform, of a nozzle guide vane including an hollow aerofoil, it is commonly known to use a cooling passage provided in the aerofoil and oriented radially, for example from the outer tip towards the base and the inner platform. The flow of the cooling gases is required to turn through 90 degrees within a short distance, in order to provide a flow of cooling air along the platform surface. This sharp turn generates a large pressure loss, high local turbulence, and poor control of the flow, leading to

inefficient cooling.

It is therefore desirable to provide a new design for nozzle guide vanes of gas turbines and, in particular to aerofoils of nozzle guide vanes, in order to avoid the inconveniences of the prior art.

Summary of the Invention It may be an object of the present invention to provide a gas turbine guide vane achieving an efficient cooling for a platform of the guide vane. It may be a further object of the present invention to provide a gas turbine guide vane including a hollow aerofoil with a passage for cooling a platform of the vane, such passage being characterized, with respect to the prior art, by :

- smaller pressure loss,

- lower local turbulence,

- higher control of the flow.

In order to achieve the objects defined above, an aerofoil for a guide vane of a gas turbine and a guide vane for a gas turbine are provided in accordance to the independent claims. The dependent claims describe advantageous developments and modifications of the invention.

According to a first aspect of the present invention, an aerofoil for a gas turbine comprises:

- an external surface extending along a first axial direction between a leading edge and a trailing edge and extending along a second radial direction between a first radial end and a second radial end of the aerofoil,

- an internal passage for channelling radially a cooling medium from the second radial end towards the first radial end. The internal passage is shaped in such a way that at the first radial end the cooling medium exits the aerofoil according to a flow direction having at least a component orthogonal to the radial direction. According to a second aspect of the present invention, a guide vane for a gas turbine includes:

- an aerofoil as described above,

- a platform connected to the first radial end of the

platform for receiving the cooling medium from the aerofoil according to a flow direction having at least a component orthogonal to the radial direction. The design provided by the present invention reduces the pressure loss coefficient of the internal cooling passage with respect to the prior art solutions in which the cooling medium ejected radially from the aerofoil impinges onto a plate, so causing large dynamic pressure losses.

According to possible exemplary embodiments of the present invention, the first radial end of the aerofoil is a base of the aerofoil, connected to an inner platform of the stator of the turbine. The second radial end of the aerofoil is a tip of the aerofoil, connected to an outer platform of the stator of the turbine. In such embodiments the internal passage inside the aerofoil extends from the tip to the base for cooling the inner platform of the turbine stator.

According to other exemplary embodiments of the present invention, the first radial end of the aerofoil is a tip of the aerofoil, connected to an outer platform of the stator of the turbine. The second radial end of the aerofoil is a base of the aerofoil, connected to an inner platform of the stator of the turbine. In such embodiments the internal passage inside the aerofoil extends from the base to the tip for cooling the outer platform of the turbine stator. According to exemplary embodiments of the present invention, the aerofoil further comprises an inner insert or deflector provided in the internal passage and extending from the first radial end towards the second radial end, the inner insert being connected to the internal passage for receiving the cooling medium, the inner insert being shaped in such a way that at the first radial end the cooling medium exits the aerofoil according to a flow direction having at least a component orthogonal to the radial direction. The deflector may be manufactured separately to the aerofoil and inserted to the aerofoil after the aerofoil has been manufactured, for example the deflector might not be formed during a casting process of the aerofoil. Advantageously, the use of an insert inside the aerofoil cooling internal passage permits to simplify the

manufacturing of the cooling passage, of the aerofoil and hence of the turbine guide vane itself.

According to exemplary embodiments of the present invention, the insert comprises at least a surface inclined with respect to the radial direction. Particularly, the insert may

comprise a spiral surfaces. The spiral surface may extend from to the other of the first radial end and the second radial end.

Advantageously, the insertion of a spiral surface inside the hollow aerofoil gives the cooling medium a swirling motion as it passes along the aerofoil cooling internal passage. When the cooling medium exits from the first radial end of the aerofoil the flow has a velocity largely perpendicular to the radial axis of the turbine and so parallel to the platform surface. This allows an efficient cooling of that surface by convection .

According to exemplary embodiments of the present invention, the insert comprising the inclined surface has a triangular section in a sectional plane parallel to the radial

direction. In particular, the triangular section of the insert may include a side lying along the first radial end of the aerofoil and a vertex oriented towards the second radial end of the aerofoil. In these embodiments a simple flow deflector is used at the first radial end of the aerofoil for turning the flow of the cooling medium as it exits the aerofoil .

Brief Description of the Drawings

The above mentioned attributes and other features and advantages of this invention and the manner of attaining them will become more apparent and the invention itself will be better understood by reference to the following description of embodiments of the invention taken in conjunction with the accompanying drawings, wherein:

FIG. 1 shows part of a turbine engine in a sectional view and in which the present inventive aerofoil is incorporated, FIG. 2 shows an exploded view of a guide vane aerofoil for a gas turbine according to the present invention,

FIG. 3 shows an isometric partial view of a guide vane for a gas turbine including the aerofoil of FIG. 2,

FIG. 4 shows a top view of the guide vane of Fig. 3,

FIG. 5 shows an isometric partially-sectioned view of the guide vane of Fig. 3,

FIG. 6 shows a front view of the guide vane of Fig. 3,

FIG. 7 shows an exploded view of another embodiment of a guide vane aerofoil for a gas turbine according to the present invention.

Detailed Description

Hereinafter, above-mentioned and other features of the present invention are described in details. Various

embodiments are described with reference to the drawings, wherein the same reference numerals are used to refer to the same elements throughout. The illustrated embodiments are intended to explain, and not to limit the invention.

FIG. 1 shows an example of a gas turbine engine 10 in a sectional view. The gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20. The gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10. The shaft 22 drivingly connects the turbine section 18 to the

compressor section 14.

In operation of the gas turbine engine 10, air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16. The burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28. The combustion chambers 28 and the burners 30 are located inside the burner plenum 26. The compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.

This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for

channelling the combustion gases to the turbine 18.

The turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22. In the present example, two discs 36 each carry an annular array of turbine blades 38. However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, nozzle guiding vanes 40, which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.

The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22. The guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.

The turbine section 18 drives the compressor section 14. The compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48. The rotor blade stages 48 comprise a rotor disc supporting an annular array of blades. The compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48. The guide vane stages include an annular array of

radially extending vanes that are mounted to the casing 50.

The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.

The casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14. A radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48. The present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications. The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated. The terms forward and rearward refer to the general flow of gas through the engine. The terms axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.

Inventive embodiments of a gas turbine aerofoil arrangement 100 are shown in Figs. 2 to 7. The turbine aerofoil arrangement 100 may be used in the turbine guiding vanes 40. The aerofoil arrangement 100 may also be used in turbine blades 48 and other aerofoil

structures in a gas turbine engine such as struts. The term aerofoil arrangement is intended to mean a component having an aerofoil 101, at least one platform 124 and at least one retention feature 102 such as a (fir-tree) dovetail root 102 or pinned fixture to which the aerofoil arrangement 100 is secured within the gas turbine engine. With reference to Figs. 2 to 6 a first preferred embodiment of a gas turbine aerofoil arrangement 100 is shown, applied to a nozzle guiding vane 40. In the following description and figures the exemplary embodiments refer to a nozzle guide vane having a single and radially inner platform on to which the aerofoil extends radially outwardly therefrom. It should be appreciated that present arrangement may be also applied to an aerofoil attached to a radially outer platform. In this case the terms radially inward / outward may be

interchanged by radially outward / inward.

The guiding vane 40 (partially represented in the attached figures) comprises:

- an inner platform 124 on a radial inner side of the stator 42. The inner platform 124 extends circumferentially to abut platforms of adjacent guide vanes and axially to seal with other engine components such as casings and rotor stages. The platform 124 has an aperture 125 through which an inner insert 170 extends. At least a part of the inner insert 170, 171 is radially inward of the platform 124. A panel 126 may be attached to the radially innermost part of the inner insert 170, 171. The panel 126 and the platform 124are generally parallel to each other. The platform 124 is more distant from the rotational axis 20 than the panel 126.

Between the platform 124 and the panel 126 a cooling medium (for example compressed air from the compressor section 14) is circulated and which cools the platform 124. The inner platform 124 has a radially outer or gas-washed surface 123 and a radially inner or cooled surface 121.

The nozzle guide vane may further have an outer platform (not represented) on a radial outer side of the stator 42. The outer platform is more distant from the rotational axis 20 than the inner platform 124. The cooled surface 121 of the radially outer platform is the radially outer surface of the outer platform.

The hollow aerofoil 101 extending along the radial direction between a first radial end 134 and a second radial end 136.

The first radial end 134 of the aerofoil 101 is a base of the aerofoil 101, connected to the inner platform 124. The second radial end 136 of the aerofoil 101 is a tip of the aerofoil 101, connected to the outer platform.

According to a possible embodiment (not represented) of the aerofoil arrangement 101, the guiding vanes 40 are arranged in a guiding vane sectors or an annular array of vanes, each comprising an inner platform, an outer platform and a

plurality of hollow aerofoil 100 extending along the radial direction between the inner platform and the outer platform. Each guiding vane sector is structurally and functionally equivalent to a plurality of the guiding vane 40 above described . The aerofoil 101 has an external surface 128. The external surface 128 axially extends between a leading edge 133 and a trailing edge 135 and radially extends between a base 134 and a tip 136. The aerofoil 101 is hollow and further comprises an internal cooling passage 150 for channelling the cooling medium from the tip 136 to the base 134 and for finally delivering the cooling medium to the inner platform 124. The cooling medium enters the cooling internal passage 150 at the tip 136 along an input radial flow direction Fl . Other aerofoils may have different cooling passages, such as serpentine passages, and which can be fed via an aperture positioned radially inward of the aerofoil. The cooling passage and therefore the cooling medium can make a number of radially inward and outward passes before being fed to the insert 170, 171 or aperture 125.

In general, according to the present invention, the internal passage 150 is shaped in such a way that at the first radial end 134 the cooling medium exits the aerofoil 101 according to an output flow direction F2 having at least a component orthogonal to the radial direction. Preferably, at the first radial end 134 the cooling medium exits the aerofoil 101 according to a flow direction F2 largely perpendicular to the radial axis and so parallel to and over the cooled surface

121 of the inner platform 124. Remotely from the exit of the internal passage 150 of the aerofoil 101, the cooling medium flows according to a final flow direction F3 parallel to the cooled surface 121of the inner platform 124. The smooth transition from the input radial flow direction Fl to the final flow direction F3, through the output flow direction F2, permits to achieve the advantages of the present

invention .

As should be understood the cooling medium is directed by the insert 170, 171 to flow over the cooled surface 121 to remove heat in the platform 124. The term ^,over' is intended to mean that the cooling medium is in contact with or flows across the cooled surface 121. Essentially, the cooling fluid or medium is intended to immediately adjacent the cooled surface 121.

With reference to the first embodiment of Figs. 2 to 6, this is obtained by providing an inner insert 170 having a spiral surface 162 in the internal passage 150. The spiral surface 162 is arranged around a stem 175 and is inclined with respect to the radial direction of the gas turbine engine 10. The insert 170 may have a single spiral surface or wall 162 or as shown two spiral walls 162 and 162A and which form two spiral channels 166 and 166A having two entries 163, 163A and two exits 164, 164A. Shown the two spiral walls 162, 162A are equally spaced from one another and approximately equal amounts of coolant pass along the two channels 166, 166A. However, it is possible for one channel to have a greater cross-sectional than the other such that preferential amounts of coolant is channelled to one side of the platform than the other e.g. a greater amount of coolant is channelled to the suction side than the pressure side or vice versa. Such preferential cooling will depend on the particular thermal characteristics of each application of the present

arrangement .

Where the deflector 170 has a single spiral wall 162 with one entry 163 and one exit 164, the exit 164 is arranged to direct the cooling medium to a preferred location, for example a hot spot, on the platform to preferentially cool that location. This is beneficial because cooling a hot spot reduces the temperature gradient across the platform and help prevent or minimise thermal gradients which can be

detrimental to the life of the component. Similarly, where there are two spiral walls 162, 162A their exits can be directed to two hot spots, one on the pressure side 124P and the other on the suction side 124S of the platform.

The insert 170 is fixed to the inner platform 124 and extends from one to the other of the base 134 and the tip 136. The insert 170 may be fixed to other parts of the aerofoil arrangement. In other embodiments the insert 170 only extends part of the span of the aerofoil from the inner platform 124 towards the tip of the aerofoil 136 or other platform. In his example, the inner insert 170 receives the cooling medium from the tip 136 and guides it along the spiral surface 162, 162A (as schematically represented by the arrows F4 of Figs. 4 and 5) to the base 134 and to the platform 124, where the cooling medium exits the aerofoil 101 according to the output flow direction F2.

With reference to the second alternative embodiment of Fig. 7, the advantages of the present arrangement are obtained also by providing an insert or deflector 171 having a

triangular prism shape near the end of the internal passage 150. The deflector 171 has two surfaces 173, 174 both

inclined with respect to the radial direction of the gas turbine engine 10. The two surfaces 173, 174 are inclined towards opposite sides of the aerofoil 101, e.g. towards the pressure side 137 and the suction side 138 of the aerofoil.

As a result, the deflector 171 has a triangular section in a sectional plane parallel to the radial direction, i.e. a sectional plane orthogonal to rotational axis 20. The triangular section of the deflector 171 includes a base side 177 parallel to the base 134 of the aerofoil 101 and a vertex 178 oriented towards the tip 136. In the second embodiment, the deflector 171 constitutes a flow deflector, used at the first radial end 134 of the aerofoil 101 for turning the flow of the cooling medium as it exits the aerofoil, from the radial direction Fl to the output direction F2. At least a part of the triangular prism is located below or radially inwardly of the platform 124 such that the deflected cooling medium is directed over the cooled or radially inner surface 121 of the platform 124. He deflector may be trapezoidal prism (i.e. a truncated

triangular prism) having a relatively small flat surface instead of an apex. Otherwise the trapezoidal prism operates and can have the same function as the triangular prism.

As can be seen in Fig.7 the triangular prism 171 is curved in the plane of the platform 124 and generally conforms to the curvature of the aerofoil 101. The triangular prism 171 is generally positioned central to the aperture 125 formed in the platform 124 such that an approximately equal amount of cooling medium is directed towards the pressure and suction sides of the platform and over the cooled surface 121 on each side of the platform. However, in certain circumstances the heat incurred by the platform can be greater on one side of the platform than the other and therefore the triangular prism 171 may be offset to direct more cooling medium to the hotter side of the platform than the other side of the platform. For example, if the suction side 124S of the platform is hotter than the pressure side 124P, the

triangular prism 171 is located nearer the pressure side so that more cooling medium is directed over the cooled surface of the platform's suction side 124S. This is by virtue of a greater area of the opening 125 being on the suction side of the apex 178 of the triangular prism 171. The base side 177 of the triangular prism 171 is attached to or integral to a lower panel or plate 126. The lower panel 126 helps to direct the flow F2 of cooling medium over the cooled surface 121 of the platform 124. The lower plate 126 extends part of the circumferential and/or axial extent of the platform, but may extend to overlap all the platform.

According to another embodiment of the present aerofoil arrangement, the first radial end of the aerofoil is a tip of the aerofoil, connected to an outer platform of the stator of the turbine. The second radial end of the aerofoil is a base of the aerofoil, connected to an inner platform of the stator of the turbine. In such embodiment, the internal passage inside the aerofoil extends from the base to the tip for cooling the outer platform of the turbine stator. In such embodiment, the outer platform has a similar structure of the inner platform of the embodiments in Figs. 2 to 7, i.e.

including a lower panel and an upper panel, parallel to each other, for allowing the circulation of the cooling medium in the outer platform.

Cooling of the cooled surface 121 may be further enhanced by the addition of well known surface features 141 (see dashed lines in Fig.3) such as ribs, fins, pins, pedestals or turbulators. These cooling features may be located on part or all of the cooled surface 121 both to reduce overall temperature and/or the temperature gradient of the platform. The cooling features 14 may be located only on the pressure side or the suction side of the platform. In particular, there surface features may be placed where there is a hot- spot on the platform to enhance cooling in the region of the hot-spot. The cooling features can function in two ways: firstly, to increase surface area exposed to the cooling medium and therefore increase heat transfer to the cooling medium; and secondly, to create turbulence in the cooling medium to maximise convection in the bulk cooling medium. In all the embodiments of the present aerofoil arrangement, it should be noted that the cooling medium within the aerofoil generally passes in a radial direction and that the deflector 170, 171 directs the radial flow to a plane generally perpendicular to the radial direction. This generally perpendicular plane is also defined by the coled surface 121 of the platform. The deflector 170, 171 can be arranged to direct cooling medium in the circumferential direction or the axial direction or in a combination of circumferential and axial directions as the particular application requires.

Claims

1. An aerofoil arrangement (100) for a gas turbine (10) comprising an aerofoil (101) and a platform (124), the aerofoil (101) attached to and extending from a first surface

(123) of the platform (124),

the aerofoil (101) comprising:

an external surface (128) extending between a leading edge (133) and a trailing edge (135) and extending in a radial direction between a first radial end (134) and a second radial end (136) of the aerofoil (101) and

an internal passage (150) for channelling a cooling medium towards the first radial end (134), the platform (126) extending in a plane perpendicular to the radial direction and comprising

a second cooled surface (121), and

defining an aperture (125), the aperture (125) is aligned with the internal passage (150) to allow the passage of cooling medium (Fl),

the aerofoil arrangement (100) further comprising a deflector (170, 171) located at least partly on the second cooled surface side of the platform (124), the deflector (170, 171) arranged to deflect the cooling medium to flow (F2) over the second cooled surface (121) of the platform

(124) to cool the platform in use.

2. The aerofoil arrangement (100) as claimed in claim 1 wherein the deflected cooling medium (F2) has a flow

direction (F2) having at least a component orthogonal to the radial direction and preferably mostly in the orthogonal direction to the radial or plane of the cooled surface (121) .

3. The aerofoil arrangement (100) of any one of claims 1-2, wherein the deflector (170, 171) comprises at least a surface (162, 162A, 173, 174) inclined with respect to the radial direction .

4. The aerofoil arrangement (100) of any one of claims 1-3, wherein the deflector (170) comprises at least one or

preferably two spiral surfaces (162, 162A) .

5. The aerofoil arrangement (100) of any one of claims 1-4, wherein the deflector (170) also extends between one to the other of the first radial end (134) and the second radial end (136) .

6. The aerofoil arrangement (100) of any one of claims 1-5 wherein the deflector (170) comprises two spiral walls (162, 162A) which form two spiral channels (166, 166A) having two entries (163, 163A) and two exits (164, 164A) .

7. The aerofoil arrangement (100) of claim 6 wherein the two spiral walls (162, 162A) are equally spaced from one another and approximately equal amounts of cooling medium pass along the two channels (166, 166A) .

8. The aerofoil arrangement (100) of claim 6 wherein the two spiral walls (162, 162A) are not equally spaced such that one channel has a greater cross-sectional area than the other and a greater amount of cooling medium passes along one channel (166, 166A) than the other so that a greater amount of coolant is directed to one of the suction side (124S) or pressure side (124P) of the platform (124) .

9. The aerofoil arrangement (100) of any one of claims 1-3, wherein the deflector (171) has a triangular section or trapezoidal section in a sectional plane parallel to the radial direction.

10. The aerofoil arrangement (100) of claim 9, wherein the triangular section of the deflector (171) in the sectional plane parallel to the radial direction includes a side along the first radial end (134) of the aerofoil (100) and a vertex oriented towards the second radial end (136) of the aerofoil (100) .

11. An aerofoil arrangement (100) as claimed in any one of claims 9-10, wherein the triangular or trapezoidal prism (171) is curved in the plane of the platform and preferably curved along at least a part of a mean camber line of the aerofoil (101) .

12. An aerofoil arrangement (100) as claimed in any one of claims 9-10, wherein the triangular or trapezoidal prism (171) is off-set from a centre-line of the aperture (125) and/or a mean camber line of the aerofoil (101) so that a greater amount of cooling medium is directed to one of the suction side (124S) or pressure side (124P) of the platform (124) .

13. An aerofoil arrangement (100) as claimed in any one of claims 9-12, wherein the triangular or trapezoidal prism (171) has a base side (177) which is attached to or integral with a lower panel (126), the lower plate (126) extends at least a part of the circumferential and/or axial extent of the platform (124) .

14. An aerofoil arrangement (100) as claimed in any one of claims 1-12, wherein at least a part of the cooled surface (121) comprises surface features (141) to enhance cooling.