GB2572793A

GB2572793A - Turbine component

Info

Publication number: GB2572793A
Application number: GB1805966.7A
Authority: GB
Inventors: Gordon Razzell Anthony; J Whittle Michael; Hiller Steven
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2018-04-11
Filing date: 2018-04-11
Publication date: 2019-10-16
Also published as: GB201805966D0

Abstract

A turbine vane 100, 200 or turbine blade comprising: a vane or blade body 101, 201,the vane or blade body being hollow at least in part and having an outer vane or blade body surface and an inner vane or blade body surface; a spar disposed at least partially within the vane or blade body, the spar comprising a spar body 102, 202 having an outer spar body surface; a flow passage 111, 211 in a gap between the outer spar body surface and the inner vane or blade body surface; and one or more flow restricting features 110, 210 disposed in the flow passage 111, 211. The vane or blade body may be made of a ceramic matrix composite (CMC). The flow restricting features may be configured such that the difference in static pressure between the inside and outside of the blade body is relatively uniform with distance from the leading edge to the trailing edge.

Description

The present disclosure concerns a turbine component such as a turbine blade or a turbine vane. The present disclosure concerns a turbine component such as a turbine blade or a turbine vane comprising a ceramic matrix composite (CMC). The present disclosure also concerns a gas turbine engine comprising one or more of the turbine components, e.g. turbine blades or turbine vanes, disclosed herein.

There is interest in using ceramic matrix composites (CMCs) in the hot sections of gas turbine engines, e.g. to produce turbine blades or turbine vanes. The use of ceramic matrix composites in the hot sections of gas turbine engines could help to improve engine performance, e.g. by reducing emissions and/or increasing cycle efficiency, relative to known engines with superalloy hot section components. The use of CMCs instead of superalloys may allow for the gas turbine engine to be operated at a higher temperature and/or may reduce the cooling requirements for components such as turbine blades and turbine vanes in the hot section of the gas turbine engine.

In a known CMC turbine vane, the vane comprises an at-least-partially hollow CMC vane body having an aerofoil shape. A web within the vane body reacts the internal pressure within the vane body relative to external pressure. The internal pressure within the vane body is typically higher than the external pressure.

According to a first aspect there is provided a turbine vane or turbine blade comprising: a vane or blade body, the vane or blade body being hollow at least in part and having an outer vane or blade body surface and an inner vane or blade body surface; a spar disposed at least partially within the vane or blade body, the spar comprising a spar body having an outer spar body surface; a flow passage in a gap between the outer spar body surface and the inner vane or blade body surface; and one or more flow restricting features disposed in the flow passage.

The vane or blade body may comprise, or consist essentially of, a ceramic matrix composite (CMC). The CMC may comprise silicon carbide. The CMC may comprise a silicon carbide fibre-reinforced silicon carbide matrix (SiC/SiC) composite. The CMC may comprise an oxide ceramic such as an N610- or N720-based ceramic. The CMC may comprise any CMC known to the person skilled in the art.

The vane or blade body may have a complex geometry to optimise engine efficiency. The vane or blade body may have a cross section that is tapered in the radial direction and/or cantered in the circumferential direction of the engine.

The vane or blade body may comprise a leading edge and a trailing edge. The vane or blade body may have an aerofoil shape. Accordingly, the vane or blade body may have two axial sides extending between the leading edge and the trailing edge, the two axial sides being connected by a curved surface in the region of the leading edge. In operation, one of the two axial sides may constitute a pressure side and the other of the two axial sides may constitute a suction side.

The trailing edge may comprise a sharp point or may be rounded at least in part. The trailing edge may have a radius of less than 5 mm, less than 3 mm, less than 2 mm, less than 1 mm or less than 0.5 mm.

A distance across the gap between the outer spar body surface and the inner vane or blade body surface may be up to 10 mm, up to 5 mm or up to 2 mm. The distance across the gap between the outer spar body surface and the inner vane or blade body surface may be at least 0.1 mm, at least 0.25 mm or at least 0.5 mm. The distance across the gap between the outer spar body surface and the inner vane or blade body surface may be substantially constant or may vary. The distance across the gap between the outer spar body surface and the inner vane or blade body surface may be selected depending upon the intended application of the turbine vane or turbine blade, in order to guide and/or control fluid flow in the flow passage as desired.

Any number of flow restricting features may be disposed in the flow passage.

Any given flow restricting feature may locally constrict the flow passage by up to or at least 10%, up to or at least 20%, up to or at least 30%, up to or at least 40%, up to or at least 50%, up to or at least 60%, up to or at least 70%, up to or at least 80%, up to or at least 90% or up to or at least 95%.

In embodiments, at least a portion of any given flow restricting feature may extend across the flow passage. For instance, the portion of the given flow restricting feature may extend from the outer spar body surface and contact the inner vane or blade body surface. Alternatively, the portion of the given flow restricting feature may extend from the inner vane or blade body surface and contact the outer spar body surface.

The flow restricting feature(s) may be provided on the outer spar body surface and/or on the inner vane or blade body surface. The flow restricting feature(s) may be integral to the outer spar body surface or the inner vane or blade body surface.

The flow restricting feature(s) may comprise one or more of: baffles, fins, lattices, impingement plates, ridges, dams, protrusions, pins, grooves, honeycomb material, slotted honeycomb material, foam or a mesh such as a wire mesh, deformed plates, braid seals and felted material. The flow restricting feature(s) may comprise, or consist essentially of, a ceramic, a composite ora metal.

The vane or blade body may comprise one or more vents to provide fluid communication from the flow passage to outside the vane or blade body. The vent(s) may comprise one or more openings such as apertures or slots. The vent(s) may be disposed at or in the vicinity of the or a trailing edge of the vane or blade body. The vent(s) may be disposed at or in the vicinity of the or a leading edge of the vane or blade body. The vent(s) may be disposed at one or more positions between the leading edge and the or a trailing edge of the vane or blade body, e.g. on a pressure surface and/or a suction surface.

The turbine vane or turbine blade may be configured such that the spar occupies more than 50%, more than 60%, more than 70%, more than 80%, more than 90% or more than 95% of an internal volume of the vane or blade body.

The spar body may be solid or hollow at least in part.

The spar body may have at least one internal cavity. The spar body may comprise at least one primary outlet to provide fluid communication from the internal cavity to the flow passage.

The spar body may comprise, or consist essentially of, a metal. The metal may comprise an alloy. Suitable alloys may include cobalt- or nickel-based alloys. Examples of suitable alloys may include Haynes 25, Haynes 188, MarM-002, C263,

C1023, CMSX3 or CMSX4, HASTX or other alloys compatible with CMC and/or the intended operating environment known to the person skilled in the art. The spar body may comprise, or consist essentially of, a composite material such as a CMC.

The spar body may comprise at least one secondary outlet to provide further fluid communication from the internal cavity to the flow passage. The secondary outlet(s) may be spaced apart from the primary outlet(s). The or each secondary outlet may be smaller than the primary outlet(s).

The primary outlet(s) may be provided in a forward portion of the turbine vane or turbine blade, e.g. near the or a leading edge of the vane or blade body.

The secondary outlet(s) may be provided in a rearward portion of the turbine vane or turbine blade. For instance, a secondary outlet may be provided near the or a trailing edge of the vane or blade body.

The spar body may comprise more than one internal cavity. The spar body may comprise two, three, four, five or more internal cavities.

The primary outlet(s) may provide fluid communication from a first internal cavity to the flow passage. The secondary outlet(s) may provide fluid communication from one or more further internal cavities, e.g. from a second internal cavity to the flow passage.

The first internal cavity may be disposed forward of the one or more further internal cavities.

The spar body may comprise one or more intermediate outlets to provide fluid communication from the internal cavity, or from one or more of the internal cavities, to the flow passage, e.g. at a location between the leading edge and the trailing edge of the vane or rotor body. Each of the one or more intermediate outlets may be disposed at a location between the primary outlet(s) and secondary outlet(s).

The outer spar body surface may comprise one or more channels for conveying a fluid.

The turbine vane or turbine blade may comprise one or more flow dividers disposed in the flow passage. The or each flow divider may provide a solid barrier to ensure a desired division and direction of flow of fluid within the flow passage. The or each flow divider may comprise a wall extending across the gap between the outer spar body surface and the inner vane or blade body surface. For instance, a flow divider may extend across the gap between the outer spar body surface and the inner vane or blade body surface and a primary outlet may be located either side of the flow divider. Additionally or alternatively, one or more of the flow dividers may extend around the perimeter of the spar body, thereby dividing the flow passage into one or more discrete portions in a radial direction.

The flow divider(s) may be integral to the spar body and/or the vane or blade body. The flow divider(s) may be separate components joined by any suitable means to the spar body and/or the vane or blade body. The flow divider(s) may comprise, or consist essentially of, a ceramic, a composite or a metal.

A second aspect provides a turbine comprising one or more turbine vanes and/or turbine blades of the first aspect.

A third aspect provides a gas turbine engine comprising one or more turbines of the second aspect.

The turbine(s) may be disposed in a hot section of the gas turbine engine.

A fourth aspect provides a spar for use in a turbine vane or turbine blade, e.g. a turbine vane or turbine blade according to the first aspect, comprising a spar body having an outer spar body surface, wherein one or more flow restricting features are provided on the outer spar body surface.

The spar body may be solid or hollow at least in part.

The spar body may have at least one internal cavity.

The spar body may have at least one primary outlet to provide fluid communication from the internal cavity to outside the spar body.

The flow restricting feature(s) may be integral to the outer spar body surface.

The spar body may comprise, or consist essentially of, a metal. The metal may comprise an alloy. Suitable alloys may include cobalt- or nickel-based alloys. Examples of suitable alloys may include Haynes 25, Haynes 188, MarM-002, C263, C1023, CMSX3 or CMSX4, HASTX or other alloys compatible with CMC and/or the intended operating environment known to the person skilled in the art. The spar body may comprise, or consist essentially of, a composite material such as a CMC.

The spar body may comprise at least one secondary outlet to provide further fluid communication from the internal cavity to outside the spar body. The secondary outlet(s) may be spaced apart from the primary outlet(s). The or each secondary outlet may be smaller than the primary outlet(s).

The primary outlet(s) may be provided in a forward portion of the spar.

The secondary outlet(s) may be provided in a rearward portion of the spar.

The primary outlet(s) may provide fluid communication from a first internal cavity to outside the spar body. The secondary outlet(s) may provide fluid communication from one or more further internal cavities, e.g. from a second internal cavity to outside the spar body.

The spar body may comprise one or more intermediate outlets to provide fluid communication from the internal cavity, or from one or more of the internal cavities, to outside the spar body, e.g. at a location between the primary outlet(s) and secondary outlet(s).

One or more flow dividers may be provided on the outer spar body surface. The or each flow divider may provide a solid barrier to ensure a desired division and direction of flow of fluid around the spar body. The or each flow divider may comprise a wall protruding from the outer spar body surface. A primary outlet may be located either side of a given flow divider. One or more of the flow dividers may extend around the perimeter of the spar body.

The flow divider(s) may be integral to the spar body. The flow divider(s) may be separate components joined by any suitable means to the spar body. The flow divider(s) may comprise, or consist essentially of, a ceramic, a composite or a metal.

A fifth aspect provides a vane or blade body for use in a turbine vane or turbine blade, e.g. a turbine vane or turbine blade according to the first aspect, the vane or blade body being hollow at least in part and having an outer vane or blade body surface and inner vane or blade body surface, wherein one or more flow restricting features are provided on the inner vane or blade body surface.

The flow restricting feature(s) may be integral to the inner vane or blade body surface.

The vane or blade body may comprise one or more vents to provide fluid communication from inside to outside the vane or blade body. The vent(s) may comprise one or more openings such as apertures or slots. The vent(s) may be disposed at or in the vicinity of the or a trailing edge of the vane or blade body. The vent(s) may be disposed at or in the vicinity of the or a leading edge of the vane or blade body. The vent(s) may be disposed at one or more positions between the leading edge and the or a trailing edge of the vane or blade body, e.g. on a pressure surface and/or a suction surface.

One or more flow dividers may be provided on the inner vane or blade body surface. The or each flow divider may provide a solid barrier to ensure a desired division and direction of flow of fluid across the inner vane or blade body surface. The or each flow divider may comprise a wall protruding from the inner vane or blade body surface. One or more of the flow dividers may extend around the perimeter of the inner vane or blade body surface.

The flow divider(s) may be integral to the vane or blade body. The flow divider(s) may be separate components joined by any suitable means to the vane or blade body. The flow divider(s) may comprise, or consist essentially of, a ceramic, a composite or a metal.

The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein.

Embodiments will now be described by way of example only, with reference to the Figures, in which:

Figure 1 is a sectional side view of a gas turbine engine;

Figure 2 is a sectional view of an example of a turbine vane and a graph showing static pressure as a function of position along the chord length of the turbine vane;

Figure 3 is a sectional view of another example of a turbine vane and a graph showing static pressure as a function of position along the chord length of the turbine vane;

Figure 4 is a side view of another example of a turbine vane;

Figure 5 is a side view of another example of a turbine vane;

Figure 6 is a side view of another example of a turbine vane;

Figure 7 is a side view of another example of a turbine vane; and

Figure 8 is a side view of another example of a turbine vane.

With reference to Figure 1, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a highpressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 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 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high 17, intermediate 18 and low 19 pressure turbines drive respectively the high pressure compressor 15, intermediate pressure compressor 14 and fan 13, each by suitable interconnecting shaft.

The high 17, intermediate 18 and low 19 pressure turbines each comprise a rotor assembly. Each rotor assembly comprises a plurality of rotor blades. Each of the high 17, intermediate 18 and low 19 pressure turbines comprises a stator assembly located downstream of the rotor assembly. Each stator assembly comprises a plurality of turbine vanes arranged to straighten at least partially the flow of the hot combustion products before they reach the rotor assembly of the next turbine in the axial flow series.

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.

A hot section of a gas turbine engine may be understood to be located downstream of the combustion equipment and upstream of the exhaust nozzle. The turbine or turbines, e.g. high-, intermediate- and low-pressure turbines, are located in the hot section of the gas turbine engine.

Figure 2 includes a cross-section through an example of a turbine vane 100.

The turbine vane 100 comprises a vane body 101. The vane body 101 is made from a ceramic matrix composite (CMC). The vane body 101 may be made using any suitable manufacturing process.

The vane body 101 has an aerofoil shape. Accordingly, the vane body 101 comprises a first axial side 103 and a second axial side 104, the first and second axial sides 103, 104 extending between a leading edge 105 and a trailing edge 106. The first and second axial sides 103, 104 are connected by a curved surface in the region of the leading edge 105. The first axial side 103 of the vane body 101 constitutes a suction side of the vane 100. The second axial side 104 of the vane body 101 constitutes a pressure side of the vane 100.

The vane body 101 comprises a vent 112 in the region of the trailing edge 106. The vent 112 allows fluid to flow from the internal volume to outside the vane body 101. The vent 112 comprises a plurality of slots and/or holes.

The vane body 101 is hollow and has an internal volume.

The vane 100 includes a spar comprising a spar body 102 disposed within the vane body 101. The spar body 102 occupies most of the internal volume of the vane body 101. For instance, the spar body 102 may occupy more than 50%, more than 60%, more than 70%, more than 80%, more than 90% or more than 95% of the internal volume of the vane body 100.

The spar body 102 is hollow and has a shape that generally corresponds to that of the vane body 101.

A flow passage 111 is defined between the vane body 101 and the spar body 102.

The spar body 102 has an outer spar body surface with a plurality of dams 110 arranged thereon. Each dam 110 protrudes from the outer spar body surface. Each dam 110 is thus disposed in the flow passage 111. Each dam 110 locally constricts the width of the flow passage 111.

Only one of the dams 110 is labelled in figure 2 for the sake of clarity. 12 dams 110 are shown in figure 2. The 12 dams 110 are distributed at intervals around the perimeter of the spar body 102.

It should be appreciated that the dam 110 is merely an example of a flow restricting feature. Equally, the vane may comprise any number and arrangement of dams or other flow restricting features.

The spar body 102 has an internal cavity 107. A primary outlet 108 provides fluid communication from the internal cavity 107 to the flow passage 111 in the vicinity of the leading edge 105 of the vane body 101. The primary outlet 108 may comprise one or more apertures.

A secondary outlet 109 provides fluid communication from the internal cavity 107 to the flow passage 111 in a region proximal to the vent 112. The secondary outlet 109 may comprise one or more apertures. The secondary outlet 109 is smaller than the primary outlet 108.

The spar body 102 is made from a metal. The metal may comprise an alloy. The spar body 102 may be made by any suitable manufacturing process or combination of manufacturing processes, e.g. a subtractive manufacturing process and/or an additive manufacturing process.

When the vane 100 is installed in a stator assembly in the hot section of a gas turbine engine, the internal cavity 107 is coupled to a source of cooling fluid. The cooling fluid may comprise air. The cooling fluid may be provided at any suitable pressure, e.g. ambient pressure or at higher pressures.

During operation of the gas turbine engine, the cooling fluid, e.g. air, is supplied to the internal cavity 107. For instance, the cooling fluid may comprise relatively cool air that is fed via a system of internal passageways from a compressor upstream of the combustion equipment to the components such as turbine vane 100 operating in the hot section of the gas turbine engine.

A major portion of the cooling fluid flows out of the internal cavity 107 through the primary outlet 108, as indicated by arrow 113. The cooling fluid thus enters the flow passage 111 in the vicinity of the leading edge 105 of the vane body 101. The cooling fluid provides impingement cooling of the leading edge 105 of the vane body.

Within the flow passage 111, the cooling fluid then flows in a chordwise direction towards the trailing edge 106 of the vane body 101, as indicated by arrows 114. The dams 110 act to locally restrict flow of the cooling fluid. The cooling fluid flow may be directed radially upwards or downwards on its route to the trailing edge as required by the cooling scheme and/or pressure field.

A minor portion of the cooling fluid flows out of the secondary outlet 109, as indicated by arrow 116. The flow of cooling fluid through the secondary outlet 109 is less than through the primary outlet 108. The flow of cooling fluid through the secondary outlet 109 may be considered a bleed flow. The flow of cooling fluid through the secondary outlet enhances local cooling proximal to the trailing edge 106 of the vane body 101.

In embodiments, the spar body may not include a secondary outlet.

Cooling fluid exits the flow passage 111 through the vent 112.

Figure 2 also includes a graph. The graph shows schematically the variation in static pressure (plotted on the y-axis) with distance along the second axial side 104 (the pressure side) from the leading edge 105 to the trailing edge 106 during operation of a gas turbine engine comprising the vane 100.

A first line 117 on the graph represents the variation in external static pressure on the second axial side 104 (the pressure side). The external static pressure decreases from the leading edge to the trailing edge.

A second line 118 on the graph represents the variation in internal static pressure on the second axial side 104 (the pressure side). A pressure difference, ΔΡ, exists across the second axial side 104 of the vane body 101. As the cooling fluid flows along the flow passage 111 between the second axial side 104 of the vane body 101 and the outer spar body surface in a chordwise direction from the leading edge 105 to the trailing edge 106, the cooling fluid encounters seven dams 110. Each dam 110 locally restricts the flow of the cooling fluid resulting in a decrease in internal static pressure. The decreases in internal static pressure caused by each dam 110 can be seen in the second line 118. A consequence is that the pressure difference, ΔΡ, does not increase significantly with distance from the leading edge 105 to the trailing edge 106. The pressure difference, ΔΡ, may be relatively uniform with distance from the leading edge 105 to the trailing edge 106. Thus, the dams 110 (and/or other flow restricting features) disposed in the flow passage 111 may control the chordwise flow of cooling fluid such that the variation in internal static pressure approximates the variation in external static pressure.

The flow restricting features, e.g. dams 110, are configured such that a similar effect is achieved in respect of the pressure difference across the first axial side 103 (the suction side) of the vane body 101.

A consequence of controlling the chordwise flow of cooling fluid such that the variation in internal static pressure approximates the variation in external static pressure is that the vane need not comprise a web within the vane body. Thus, manufacture of the vane may be simpler and/or cost-effective.

Figure 3 includes a cross-section of another example of a turbine vane 200.

The turbine vane 200 comprises a vane body 201. The vane body 201 is made from a ceramic matrix composite (CMC). The vane body 201 may be made using any suitable manufacturing process.

The vane body 201 has an aerofoil shape. Accordingly, the vane body 201 comprises a first axial side 203 and a second axial side 204, the first and second axial sides 203, 204 extending between a leading edge 205 and a trailing edge 206. The first and second axial sides 203, 204 are connected by a curved surface in the region of the leading edge 205. The first axial side 203 of the vane body 201 constitutes a suction side of the vane 200. The second axial side 204 of the vane body 201 constitutes a pressure side of the vane 200.

The vane body 201 is hollow and has an internal volume.

The vane body 201 comprises a vent 212 in the region of the trailing edge 206. The vent 212 allows fluid to flow from the internal volume to outside the vane body 201. The vent 212 comprises a plurality of slots.

The vane 200 includes a spar comprising a spar body 202 disposed within the vane body 201. The spar body 202 occupies most of the internal volume of the vane body 201. For instance, the spar body 202 may occupy more than 50%, more than 60%, more than 70%, more than 80%, more than 90% or more than 95% of the internal volume of the vane body 200.

The spar body 202 is hollow and has a shape that generally corresponds to that of the vane body 201.

A flow passage 211 is defined between the vane body 201 and the spar body 202.

The spar body 202 has an outer spar body surface with a plurality of dams 210 arranged thereon. Each dam 210 protrudes from the outer spar body surface. Each dam 210 is thus disposed in the flow passage 211. Each dam 210 locally constricts the width of the flow passage 211.

Only one of the dams 210 is labelled in figure 2 for the sake of clarity. 12 dams 210 are shown in figure 2. The 12 dams 210 are distributed at intervals around the perimeter of the spar body 202.

It should be appreciated that the dam 210 is merely an example of a flow restricting feature. Equally, the vane may comprise any number and arrangement of dams or other flow restricting features.

The spar body 202 contains a first internal cavity 207a and a second internal cavity 207b. The first and second internal cavities 207a, 207b are separated by a dividing wall 219 extending across the spar body 202 from the first axial side 203 to the second axial side 204. A primary outlet 208 provides fluid communication from the first internal cavity 207a to the flow passage 211 in the vicinity of the leading edge 205 of the vane body 201. The primary outlet 208 may comprise one or more apertures.

Fluid communication from the second internal cavity 207b to the flow passage 211 is provided by a pair of intermediate outlets 220 and a secondary outlet 209.

The intermediate outlets 220 each provide fluid communication from the second internal cavity 207b to the flow passage 211 at a location between the leading edge 205 and the trailing edge 206. The location between the leading edge 205 and the trailing edge 206 may be at least 10% or at least 25% and/or up to 75% or up to 90% of the chordwise distance between the leading edge 205 and the trailing edge 206. The intermediate outlets 220 may be located at the same or different chordwise distances between the leading edge 205 and the trailing edge 206. The intermediate outlets 220 are provided in opposite sides of the spar body 202. Each of the intermediate outlets 220 may comprise one or more apertures. The intermediate outlets 220 may not be the same size as each other and/or may comprise a different number of apertures. The spar body 202 may comprise no intermediate outlets, one intermediate outlet, two intermediate outlets, or more than two intermediate outlets.

The secondary outlet 209 provides fluid communication from the second internal cavity 207b proximate the trailing edge 206 of the vane body 201 to the flow passage 211 in a region proximal to the vent 212. The secondary outlet 209 may comprise one or more apertures. The secondary outlet 209 may be smaller than, the same size as, or larger than, the primary outlets 208 and/or the intermediate outlets 220.

It should be appreciated that the internal cavities 207a, 207b are merely an example of an arrangement of internal cavities in the spar body 202. Equally, the spar body 202 may comprise any number of internal cavities separated by any number and arrangement of walls. The spar body 202 may also comprise any number or arrangement of outlets to provide fluid communication from the internal cavities to the flow passage 211.

The spar body 202 is made from a metal. The metal may comprise an alloy. The spar body 202 may be made by any suitable manufacturing process or combination of manufacturing processes, e.g. a subtractive manufacturing process and/or an additive manufacturing process.

When the vane 200 is installed in a stator assembly in the hot section of a gas turbine engine, the first and second internal cavities 207a, 207b are each coupled to a source of cooling fluid. The cooling fluid may comprise air. The cooling fluid may be provided at any suitable pressure, e.g. ambient pressure or at higher pressures. The first and second internal cavities 207a, 207b may be coupled to a common source of cooling fluid. The first and second internal cavities 207a, 207b may be coupled to separate sources of cooling fluid. The separate sources of cooling fluid may provide cooling fluid, e.g. air, at the same pressure or at different pressure. For example, the cooling fluid supplied to the first internal cavity 207a may have a higher pressure than the cooling fluid supplied to the second internal cavity 207b. The cooling fluid supplied to the first internal cavity 207a may constitute compressor discharge air, e.g. HP6 air. The cooling fluid supplied to the second internal cavity 207b may constitute lower pressure compressor bleed air, e.g. HP3 air.

During operation of the gas turbine engine, the cooling fluid, e.g. air, is supplied to the internal cavities 207a, 207b. For instance, the cooling fluid may comprise relatively cool air that is fed via a system of internal passageways from a compressor or compressors upstream of the combustion equipment to the components such as turbine vane 200 operating in the hot section of the gas turbine engine.

The cooling fluid supplied to the first internal cavity 207a flows through the primary outlet 208, as indicated by arrow 213. Cooling fluid thus enters the flow passage 211 in the vicinity of the leading edge 205 of the vane body 201. The cooling fluid entering the flow passage 211 via the primary outlet 208 provides impingement cooling of the leading edge 205 of the vane body 201.

Within the flow passage 211, the cooling fluid then flows in a chordwise direction towards the trailing edge 206 of the vane body 201, as indicated by arrows 214. The dams 210 act to locally restrict flow of the cooling fluid.

A portion of the cooling fluid supplied to the second internal cavity 207b flows through the intermediate outlets 220, as indicated by arrows 221. Cooling fluid thus enters the flow passage 211 on either side of the vane body 202 at an intermediate chordwise location, i.e. a location between the leading edge 205 and the trailing edge 206.

A portion of the cooling fluid supplied to the second internal cavity 207b flows through the secondary outlet 209, as indicated by arrow 216. The intermediate outlets 220 and the secondary outlet 209 may be configured such that the flow of cooling fluid through the secondary outlet 209 is less than through the intermediate outlets 220. The flow of cooling fluid through the secondary outlet 209 may be considered a bleed flow. The flow of cooling fluid through the secondary outlet 209 enhances local cooling proximal to the trailing edge 206 of the vane body 201.

The portions of the cooling fluid supplied to the second internal cavity 207b that flow through the intermediate outlet(s) 220 and the secondary outlet 209 may be equal or different. For instance, a major portion of the cooling fluid supplied to the second internal cavity 207b may flow through the intermediate outlet(s) 220.

In embodiments, the spar body 202 may not include a secondary outlet 209.

Cooling fluid exits the flow passage 211 through the vent 212.

Figure 3 also includes a graph. The graph shows schematically the variation in static pressure (plotted on the y-axis) with distance along the second axial side 204 (the pressure side) from the leading edge 205 to the trailing edge 206 during operation of a gas turbine engine comprising the vane 200.

A first line 217 on the graph represents the variation in external static pressure on the second axial side 204 (the pressure side). The external static pressure decreases from the leading edge 205 to the trailing edge 206.

A second line 218 on the graph represents the variation in internal static pressure on the second axial side 204 (the pressure side). A pressure difference, ΔΡ, exists across the second axial side 204 of the vane body 201. As the cooling fluid flows along the flow passage 211 between the second axial side 204 of the vane body 201 and the outer spar body surface in a chordwise direction from the leading edge 205 to the trailing edge 206, the cooling fluid encounters seven dams 210. Each dam 210 locally restricts the flow of the cooling fluid resulting in a decrease in internal static pressure. The decreases in internal static pressure caused by each dam 210 can be seen in the second line 218. A consequence is that the pressure difference, ΔΡ, does not increase significantly with distance from the leading edge 205 to the trailing edge 206. The pressure difference, ΔΡ, may be relatively uniform with distance from the leading edge 205 to the trailing edge 206. Thus, the dams 210 (and/or other flow restricting features) disposed in the flow passage 211 may control the chordwise flow of cooling fluid such that the variation in internal static pressure approximates the variation in external static pressure.

In embodiments, the cooling fluid supplied to the second internal cavity 207b may be of a lower pressure than the cooling fluid supplied to the first internal cavity 207a. Thus, relatively less higher pressure cooling fluid (e.g. HP6 air) may be utilised. Such higher pressure cooling fluid may be a relatively scarce and/or precious resource compared with lower pressure cooling fluid (e.g. HP3 air). Another consequence may be that the number and arrangement of flow restricting features, e.g. dams, required to control the chordwise flow of cooling fluid such that the variation in internal static pressure approximates the variation in external static pressure may differ. For instance, fewer dams may be required than if only higher pressure cooling fluid were used.

The flow restricting features, e.g. dams 210, are configured such that a similar effect is achieved in respect of the pressure difference across the first axial side 203 (the suction side) of the vane body 201.

Figure 4 shows a side view of an example of a spar body 400.

The spar body 400 comprises a leading edge 401, a trailing edge 403, a top 404, a bottom 405, and an axial side wall 402 extending between the top 404 and the bottom 405 and also extending in a chordwise direction between the leading edge 401 and the trailing edge 403.

The spar body 400 has a honeycomb array 406 located on, and protruding from, a portion of the axial side wall 402. The honeycomb array 406 acts as a flow restricting feature. The honeycomb array 406 is located between the leading edge 401 and the trailing edge 403 of the spar body 400, and extends from the top 404 to the bottom 405 of the spar body 400. The honeycomb array 406 may extend across substantially the entire axial side wall 402, or a portion of the axial side wall 402. The height of the honeycomb array 406 may not be uniform across the portion of the axial side wall 402 on which the honeycomb array 406 is located. The honeycomb array 406 may be manufactured by any suitable manufacturing process. Examples of suitable manufacturing processes may include additive layer manufacturing, or by attachment of a foil-based honeycomb structure to the axial side wall 402. The other axial side wall (not shown) of the spar body 400 may also have flow restricting features located thereon, for example a honeycomb array 406.

Figure 5 shows a side view of another example of a spar body 500.

The spar body 500 comprises a leading edge 501, a trailing edge 503, a top 504, a bottom 505, and an axial side wall 502 extending between the top 504 and the bottom 505 and also extending in a chordwise direction between the leading edge 501 and the trailing edge 503.

The spar body 500 has a slotted honeycomb array 506 located on, and protruding from, a portion of the axial side wall 502. The slotted honeycomb array 506 may act as a flow restricting feature. The slotted honeycomb array 506 is located between the leading edge 501 and the trailing edge 503 of the spar body 500, and extends from the top 504 to the bottom 505 of the spar body 500. The slotted honeycomb array 506 may extend across substantially the entire axial side wall 502, or a portion of the axial side wall 502. The height of the slotted honeycomb array 506 may not be uniform across the portion of the axial side wall 502 on which the slotted honeycomb array 506 is located. The slotted honeycomb array 506 may be manufactured by any suitable process. Examples of suitable manufacturing processes may include additive layer manufacturing, or by attachment of a foil-based slotted honeycomb structure to the axial side wall 502. The other axial side wall (not shown) of the spar body 500 may also have flow restricting features located thereon, for example a slotted honeycomb array 506.

Figure 6 shows a side view of a spar body 600.

The spar body 600 comprises a leading edge 601, a trailing edge 603, a top 604, a bottom 605, and an axial side wall 602 extending between the top 604 and the bottom 605 and also extending in a chordwise direction between the leading edge 601 and the trailing edge 603.

The spar body 600 has a plurality of dams 606 located on, and protruding from, a portion of the axial side wall 602. The plurality of dams 606 acts as flow restricting features. The plurality of dams 606 is located between the leading edge 601 and the trailing edge 603 of the spar body 600, and extends from the top 604 to the bottom 605 of the spar body 600. The plurality of dams 606 may extend across substantially the entire axial side wall 602, or a portion of the axial side wall 602. The height of the plurality of dams 606 may not be uniform across the portion of the axial side wall 602 on which the plurality of dams 606 is located. The plurality of dams 606 may be manufactured by any suitable manufacturing process. Examples of suitable manufacturing processes may include additive layer manufacturing, or casting or machining into the surface of the axial side wall 602 of the spar body 600. Each of the plurality of dams 606 is generally arcuate in form with the peak of the arc disposed towards the leading edge 601. The other axial side wall of the spar body 600 may also have flow restricting features located thereon, for example a plurality of dams.

Figure 7 shows a side view of another example of a spar body 700.

The spar body 700 comprises a leading edge 701, a trailing edge 703, a top 704, a bottom 705, and an axial side wall 702 extending between the top 704 and the bottom 705 and also extending in a chordwise direction between the leading edge 701 and the trailing edge 703.

The spar body 700 has a plurality of pins 706 located on, and protruding from, a portion of the axial side wall 702. The plurality of pins 706 act as flow restricting features. The plurality of pins 706 is located between the leading edge 701 and the trailing edge 703 of the spar body 700, and extends from the top 704 to the bottom 705 of the spar body 700. The plurality of pins 706 may extend across substantially the entire axial side wall 702, or a portion of the axial side wall 702. The height of the plurality of pins 706 may not be uniform across the portion of the axial side wall 702 on which the plurality of pins

706 is located. The plurality of pins 706 may be manufactured by any suitable process. Examples of suitable manufacturing processes may include casting, or additive layer manufacturing. The plurality of pins 706 may be formed by placing inserts into holes drilled into the axial side wall 702 of the spar body 700. The plurality of pins 706 are arranged in a series of chevrons pointing towards the leading edge 701. The other axial side wall of the spar body 700 may also have flow restricting features located thereon, for example a plurality of pins.

Figure 8 shows a side view of another example of a spar body 800.

The spar body 800 comprises a leading edge 801, a trailing edge 803, a top 804, a bottom 805, and an axial side wall 802 extending between the top 804 and the bottom 805 and also extending in a chordwise direction between the leading edge 801 and the trailing edge 803.

The spar body 800 has a plurality of dams 806 located on, and protruding from, a portion of the axial side wall 802. The plurality of dams 806 acts as flow restricting features. The plurality of dams 806 is located between the leading edge 801 and the trailing edge 803 of the spar body 800, and extends from the top 804 to the bottom 805 of the spar body 800. The plurality of dams 806 may extend across substantially the entire axial side wall 802, or a portion of the axial side wall 802. The height of the plurality of dams 806 may not be uniform across the portion of the axial side wall 802 on which the plurality of dams 806 is located. The plurality of dams 806 may be manufactured by any suitable process. Examples of suitable manufacturing processes may include additive layer manufacturing, or casting or machining grooves into the surface of the axial side wall 802 of the spar body 800. The plurality of dams 806 zigzag as they extend from the top 804 to the bottom 805 of the spar body 800.

In embodiments, the plurality of dams may be straight from the top to the bottom of the spar body. The plurality of dams may be aligned parallel to the leading edge and/or the trailing edge of the spar body.

The example spar bodies 400, 500, 600, 700 may be receivable at least partially within a vane or rotor body.

In a vane or rotor comprising a vane or rotor body with a spar body (e.g. one of example spar bodies 400, 500, 600, 700) received at least partially therein, one or more additional or alternative flow restricting features may be disposed on a portion of an inner surface of the vane or rotor body.

The present disclosure may be applied to rotors as well as to vanes.

The present disclosure may be applied to any turbine, e.g. aircraft gas turbine engines or ground-based power-generating turbines.

In a turbine vane or blade comprising a composite such as a CMC, the or each CMC surface may be left in an as-formed state or may be prepared or finished in any suitable way, e.g. to provide a smoother surface. A smoother surface may provide more consistent sealing and/or better fluid flow properties.

The flow restricting feature(s) may each comprise one or more slots or holes in a portion thereof. This may allow less-restricted fluid flow in some areas.

It will be understood that the invention is not limited to the embodiments abovedescribed and various modifications and improvements can be made without departing from the concepts herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A turbine vane (100; 200) or turbine blade comprising: a vane or blade body (101; 201), the vane or blade body (101; 201) being hollow at least in part and having an outer vane or blade body surface and an inner vane or blade body surface; a spar disposed at least partially within the vane or blade body, the spar comprising a spar body (102; 202) having an outer spar body surface; a flow passage (111; 211) in a gap between the outer spar body surface and the inner vane or blade body surface; and one or more flow restricting features (110; 210) disposed in the flow passage (111; 211).

2. A turbine vane or turbine blade according to claim 1, wherein the vane or blade body (101; 201) comprises, or consists essentially of, a ceramic matrix composite (CMC).

3. A turbine vane or turbine blade according to claim 1 or claim 2, wherein the gap between the outer spar body surface and the inner vane or blade body surface is up to 5 mm or up to 2 mm across.

4. A turbine vane or turbine blade according to claim 1, claim 2 or claim 3, wherein the flow restricting feature(s) (110; 210) is/are provided on the outer spar body surface and/or on the inner vane or blade body surface.

5. A turbine vane or turbine blade according to any one of the preceding claims, wherein the flow restricting feature(s) (110; 210) comprise one or more of: baffles, fins, lattices, impingement plates, ridges, dams, protrusions, pins, grooves, honeycomb material, slotted honeycomb material, foam or a mesh such as a wire mesh, deformed plates, braid seals and felted material.

6. A turbine vane or turbine blade according to any one of the preceding claims, wherein the vane or blade body (101; 201) comprises one or more vents (112; 212) to provide fluid communication from the flow passage (111; 211) to outside the vane or blade body (101; 201).

7. A turbine vane or turbine blade according to any one of the preceding claims, wherein the turbine vane or turbine blade is configured such that the spar occupies more than 50%, more than 60%, more than 70%, more than 80%, more than 90% or more than 95% of an internal volume of the vane or blade body (101; 201).

8. A turbine vane or turbine blade according to any one of the preceding claims, wherein the spar body (102; 202) has at least one internal cavity (107) and comprises at least one primary outlet (108) to provide fluid communication from the internal cavity (107) to the flow passage (111; 211).

9. A turbine vane or turbine blade according to any one of the preceding claims, wherein the spar body (102; 202) comprises, or consists essentially of, a metal or a composite material.

10. A turbine vane or turbine blade according to claim 8 or claim 9 when dependent on claim 8 comprising at least one secondary outlet (109; 209) to provide further fluid communication from the internal cavity (107) to the flow passage (111; 211).

11. A turbine vane or turbine blade according to claim 10, wherein the secondary outlet(s) (109; 209) is/are spaced apart from the primary outlet(s) (108; 208) and/or the or each secondary outlet (109; 209) is smaller than the primary outlet(s) (108; 208).

12. A turbine vane or turbine blade according to claim 8, or any one of claims 9 to 11 when dependent on claim 8, wherein the primary outlet(s) (108; 208) is/are provided in a forward portion of the turbine vane or turbine blade.

13. A turbine vane or turbine blade according to claim 10, or claim 11 or claim 12 when dependent on claim 10, wherein the secondary outlet(s) (109; 209) is/are provided in a rearward portion of the turbine vane or turbine blade.

14. A turbine vane or turbine blade according to claim 8, or any one of claims 9 to 13 when dependent on claim 8, wherein the spar body (202) comprises more than one internal cavity (207a, 207b).

15. A turbine vane or blade according to any one of the preceding claims, wherein the outer spar body surface comprises one or more channels for conveying a fluid.

16. A turbine vane or blade according to any one of the preceding claims comprising one or more flow dividers disposed in the flow passage (111; 211).

17. A turbine (17; 18; 19) comprising one or more turbine vanes and/or turbine blades according to any one of the preceding claims.

18. A gas turbine engine (10) comprising one or more turbines (17; 18; 19) according to claim 17.

19. A spar for use in a turbine vane or turbine blade, e.g. a turbine vane or turbine blade according to any one of claims 1 to 16, comprising a spar body (102; 202) having an outer spar body surface, wherein one or more flow restricting features (110; 210) are provided on the outer spar body surface.

20. A vane or blade body (101; 201) for use in a turbine vane or turbine blade, e.g. a turbine vane or turbine blade according any one of claims 1 to 16, the vane or blade body (101; 201) being hollow at least in part and having an outer vane or blade body surface and inner vane or blade body surface, wherein one or more flow restricting features (110; 210) are provided on the inner vane or blade body surface.