GB2570652A - A cooling arrangement for a gas turbine engine aerofoil component platform - Google Patents

A cooling arrangement for a gas turbine engine aerofoil component platform Download PDF

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Publication number
GB2570652A
GB2570652A GB1801551.1A GB201801551A GB2570652A GB 2570652 A GB2570652 A GB 2570652A GB 201801551 A GB201801551 A GB 201801551A GB 2570652 A GB2570652 A GB 2570652A
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GB
United Kingdom
Prior art keywords
platform
aerofoil
cooling
surface
edge
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
GB1801551.1A
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GB201801551D0 (en
Inventor
Hunt David
Matthews Alexander
Ceremuga Tomasz
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Rolls Royce PLC
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Rolls Royce PLC
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Filing date
Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Priority to GB1801551.1A priority Critical patent/GB2570652A/en
Publication of GB201801551D0 publication Critical patent/GB201801551D0/en
Publication of GB2570652A publication Critical patent/GB2570652A/en
Pending legal-status Critical Current

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Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187Convection cooling
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02Blade-carrying members, e.g. rotors
    • F01D5/08Heating, heat-insulating or cooling means
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/30Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • F01D5/3015Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type with side plates
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/80Platforms for stationary or moving blades
    • F05D2240/81Cooled platforms
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/30Arrangement of components
    • F05D2250/31Arrangement of components according to the direction of their main axis or their axis of rotation
    • F05D2250/314Arrangement of components according to the direction of their main axis or their axis of rotation the axes being inclined in relation to each other
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/20Heat transfer, e.g. cooling
    • F05D2260/202Heat transfer, e.g. cooling by film cooling

Abstract

An aerofoil component of a gas turbine engine, wherein the inlets of the platform cooling passages are spaced closer than the outlets. The aerofoil has leading and trailing edges with pressure and suction surfaces and at least one aerofoil cooling passage for providing internal and external cooling. The platform extending laterally from a root of the aerofoil at the pressure surface suction surface includes platform cooling passages each of which extend from an inlet which is open to the aerofoil cooling passage(s) to an outlet in a surface of the platform, which may be on the leading edge of the aerofoil, upstream, in a surface of an overhang such as a seal fin. The platform may include a cooling passage which extends across a lock plate engagement. The inlets may be provided in clusters and radially distributed along an aero foil cooling passage. More outlets may be provided on the pressure surface or the aft edge than the suction surface edge. This may be used in a turbine blade or to a nozzle guide vane.

Description

(54) Title of the Invention: A cooling arrangement for a gas turbine engine aerofoil component platform Abstract Title: Cooling passages for aerofoil in gas turbine engine (57) An aerofoil component of a gas turbine engine, wherein the inlets of the platform cooling passages are spaced closer than the outlets. The aerofoil has leading and trailing edges with pressure and suction surfaces and at least one aerofoil cooling passage for providing internal and external cooling. The platform extending laterally from a root of the aerofoil at the pressure surface suction surface includes platform cooling passages each of which extend from an inlet which is open to the aerofoil cooling passage(s) to an outlet in a surface of the platform, which may be on the leading edge of the aerofoil, upstream, in a surface of an overhang such as a seal fin. The platform may include a cooling passage which extends across a lock plate engagement. The inlets may be provided in clusters and radially distributed along an aero foil cooling passage. More outlets may be provided on the pressure surface or the aft edge than the suction surface edge. This may be used in a turbine blade or to a nozzle guide vane.

At least one drawing originally filed was informal and the print reproduced here is taken from a later filed formal copy.

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A Cooling Arrangement for a Gas Turbine Engine Aerofoil Component Platform

Technical Field of Invention

The present invention relates to a cooling arrangement for an aerofoil component platform. The aerofoil component will typically be a turbine blade of a high pressure turbine section of a gas turbine engine but may be vane.

Background of Invention

With reference to Figure 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.

The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 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 lowpressure turbines 16,17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.

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.

The turbine is the hottest part of the engine and will typically have a main gas path entry temperature well in excess of the melting point of the materials used to make the gas path turbine components. The gas path turbine components, i.e. the nozzle guide vanes, blades and seal segments, may therefore be cooled internally and externally using an appropriate supply of cooling air. Using cooled components allows the turbine entry temperature to be increased which generally allows for an increase in thrust from the engine. However, the cooling air is typically provided by the high pressure compressor and represents a loss for the engine cycle as a whole. Hence, there is significant challenge to increase cooling performance whilst minimising the cooling air burden.

Figure 2 shows an isometric view of a typical single stage cooled turbine in which there is a nozzle guide vane in flow series with a turbine rotor. The nozzle guide vane includes an aerofoil 31 which extends radially between inner 32 and outer 33 platforms. The turbine rotor includes a blade mounted to the peripheral edge of a rotating disc. The blade includes an aerofoil 32 which extends radially outwards from an inner platform. The radially outer end of the blade includes a shroud which sits within a seal segment 35. The seal segment is a stator component and attached to the engine casing. The arrows in Figure 2 indicate cooling flows.

The main gas path extends from an upstream direction through the nozzle guide vane which accelerates and swirls the hot gas in the direction of the turbine blade rotation. The orientation of hot gas reacts with the aerodynamic shape of the turbine blades to drive the rotor, shaft and compressor (or fan as the case may be). The vanes and blades are arranged in flow series pairs throughout the turbine section of the engine.

The high-pressure turbine components typically receive cooling air taken from the high pressure compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.

The arrangement of cooling passages within the vanes and blades deliver cooling air to the requisite locations and must be accommodated within the body of the aerofoil without compromising the resilience of the component.

The provision of internal cooling and external cooling is well known in the art and the arrangement of cooling holes and passageways has been the subject of extensive research and development for many years. Figure 3 shows a typical high pressure turbine blade 310 having internal and external air cooling features. The blade 310 includes a root portion 311 and an aerofoil portion 313 and is one of an annular array of similar blades which are circumferentially distributed around the turbine disc when assembled in a turbine module.

The root portion 311 includes a fir tree root 315 which axially engages with a corresponding slot on periphery of the disc. A shank 317 extends between the fir tree root 315 and the aerofoil portion 313. The aerofoil portion includes a platform, an aerofoil and a tip shroud.

The aerofoil portion 313 includes a leading edge 314 and a trailing edge 316 which extend radially in span between the platform 322 and tip shroud 319. Pressure and suction surfaces extend in chord between the leading and trailing edges 314, 316.

The platform 322 provides the radially inner end wall of the main gas path and forms part of a full annulus with the other blade platforms 322 mounted around the disc 54. The platforms 322 extend laterally from the root end of the aerofoil 312 thereby having portions on the pressure side, suction side, leading edge and trailing edge. The fore part of the platform terminates in one or more seal fins 348 which engage with a similar structure of the trailing edge of the upstream nozzle guide vane, the two cooperating to provide a labyrinth seal between the static and rotating components. The circumferential edges of the platform 322 may include grooves 321 for seals or dampers which may oppose or engage with corresponding features on adjacent blade platforms.

The tip shroud lies radially inwards of the seal segment and provides a rotating platform similar to the platform 322 at the root end of the blade. Thus, there is an end wall which extends laterally from the tip end of the aerofoil 312 between leading and trailing edges 314, 316 in the axial direction, and circumferential edges which lie proximal to corresponding tip shrouds of adjacent blades to provide a full annular shroud. The radially outer surface of the tip shroud may be provided with one or more air restricting features, e.g. fins, which operate aerodynamically with the seal segment to reduce over-tip leakage of the main gas path air.

The high pressure turbine blade shown in Figure 3 is an air cooled component having a plurality of internal and external cooling features. The interior of the aerofoil 312 is provided with a number of radial passages which extend from an inlet 334 in the root to an outlet 338 at the tip. Each internal aerofoil cooling passage 336 has a number of outlets 338 distributed along and cross the pressure surface, around the leading edge 314 and at or towards the trailing edge 316. Cooling holes may also be provided on the suction surface.

The internal aerofoil cooling passages 336 may be multi-pass or meandering in that they extend multiple times in span in longitudinal radially extending sections 323 joined by returns such as the ubend portions 325 shown. Alternatively or additionally, the internal aerofoil cooling passages 336 may be single pass 327 as demonstrated by two leading edge 314 passages in Figure 3.

The internal aerofoil cooling passages 336 may also include surface features which increase the surface area and/or cause turbulation in the cooling air flow thereby increasing the heat transfer from the wall of the component to the air flow. The heat transfer surface features typically include projections in the form of elongate strips or columns, so-called pedestals.

The cooling holes in the flanks and leading edge 314 of the blade are known as film cooling holes which create a thin film of cooling air which ideally forms a boundary layer which protects the surface of the component from the hot gas flow path. The exhausts in the tip shroud may provide a cooling benefit and a disruptive flow to help further reduce the over-tip leakage.

The design of the cooling geometry is typically governed by performance requirements such as temperature and pressure distributions, manufacturing capabilities and efficiency.

Air cooled turbine components are typically made from specialist alloys cast using advanced investment casting techniques prior to machining and coating as known in the art. The investment casting generally includes encapsulating a ceramic core which will define the internal cooling passages within a wax moulding having the external shape of a turbine blade. This is then invested within a ceramic shell and the wax removed. Once the metal is cast within the shell, the ceramic core is removed from within the metal blade to leave internal cooling passages in the form of the ceramic core. External features can be cast into the surface but machining is often required to provide a desired shape and features such as the film cooling holes.

The present invention seeks to provide an improved cooling geometry for a platform of a turbine blade or vane.

Statements of Invention

According to an aspect there is provided an aerofoil component of a gas turbine engine, comprising an aerofoil having leading and trailing edges with pressure and suction surfaces extending therebetween and at least one aerofoil cooling passage for providing internal and external cooling for the aerofoil component; a platform extending laterally from a root of the aerofoil at the pressure surface suction surface, leading edge and trailing edge, and having fore, aft and circumferential peripheral edges; the platform including a plurality of platform cooling passages each of which extend from an inlet which is open to the at least one aerofoil cooling passages to an outlet in a surface of the platform and, wherein the inlets of the platform cooling passages are spaced closer than the outlets.

The majority rather than all of the inlets may be spaced closer than the outlets. A minority of the outlets may be spaced closer than the inlets.

At least one platform cooling passage may have a leading edge platform portion outlet on the upstream side of the leading edge of the aerofoil.

The leading edge platform portion outlet may be located in a surface of an overhang which extends axially fore of the leading edge of the aerofoil where the aerofoil joins the platform.

The leading edge platform portion outlet may be located in a radially inner surface of the overhang. The overhang may be a seal fin.

The platform may include a lockplate engagement feature on the radially inner surface of the platform.

The leading edge platform cooling passage may extend across the lockplate engagement feature.

The inlets may be provided in clusters.

The inlets within each cluster may be radially distributed along the at least one aerofoil cooling passage. Each cluster of inlets may open to different aerofoil cooling passages.

The outlets of the platform cooling passages may be provided on one or more of the circumferential, fore or aft edges of the platform.

Within the scope of this application it is expressly envisaged that the various aspects, embodiments, examples and alternatives, and in particular the individual features thereof, set out in the preceding paragraphs, in the claims and/or in the following description and drawings, may be taken independently or in any combination. For example features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.

Description of Drawings

Embodiments of the invention will now be described with the aid of the following drawings of which:

Figure 1 shows a schematic longitudinal section of a gas turbine engine.

Figure 2 shows a partial perspective view of a high pressure turbine stage which may be found in the gas turbine engine of Figure 1.

Figure 3 shows perspective views of a conventional gas turbine blade including a cut away portion in the pressure surface wall to reveal some exemplary cooling flow passages.

Figure 4 shows a radially inward facing view of a transverse section of a turbine blade and platform cooling passages.

Figure 5 is a circumferential view of a partial longitudinal section showing the leading edge platform fin seal arrangement.

Figure 6 shows a ghost view of a section of the platform 422 trailing edge 416

Detailed Description of Invention

Unless otherwise stated, geometric references to axial, radial, circumferential, fore and aft, and longitudinal will be in relation to the principal axis of the engine (XX, Figurel), with upstream and downstream, in relation to the main gas path flow direction. Chord relates to the separation between the leading edge and trailing edge of an aerofoil, and span is used in relation to the radial extent of the aerofoil. The stagger angle is the angle between the aerofoil chord line and principal axis of the engine. The axial, radial and circumferential dimensions are marked on the drawings with respective A, R and C.

Figure 4 shows a radially inward facing view of a transverse section of an aerofoil 412 component in the form of a turbine blade 410 for a gas turbine. The aerofoil 412 component, comprising: an aerofoil 412 having a leading edge 414 and a trailing edge 416 with a pressure surface 418 and a suction surface 420 extending therebetween and at least one aerofoil cooling passage 436 for providing internal and external cooling for the aerofoil 412 component. A platform 422 extends laterally from a root of the aerofoil 412 at the pressure surface 418, suction surface 420, leading edge 414 and trailing edge 416, and having fore 426, aft 428 and circumferential peripheral edges 428, 430. The platform 422 includes a plurality of platform cooling passages 432 each of which extend from an inlet 434 which is open to the at least one aerofoil cooling passage 436 to an outlet 438 in a surface of the platform 422. The inlets 434 of the platform cooling passages 432 are spaced closer than the outlets 438 to provide a starburst configuration in which the cooling passages appear to radially explode from one or more locations.

The blade is similar to the blade described in connection with Figure 3 in that there is an aerofoil 412 having leading and trailing edges 414, 416 with pressure 418 and suction surface 420 extending laterally therebetween in chord. The aerofoil 412 radially extends from a root to a tip in span. The root of the aerofoil 412 joins a platform 422 which forms the base of the aerofoil 412 and extends laterally therefrom to provide a radially inner end wall which defines the main gas path around the aerofoil 412. The blade is one of a circumferential distribution which are attached to the radial periphery of a turbine rotor disc (see Figure 2). Hence, in a working engine, adjacent platforms form a full annulus around the axis of rotation, thereby providing the radially inner boundary wall of the hot gas path annulus.

The interior of the blade 410 includes a plurality of internal aerofoil cooling passages 436 which extend radially along the aerofoil 412 from the root to the tip in a similar manner to Figure 3. The aerofoil cooling passages 436 provide a supply of cooling air for internal and external cooling of the aerofoil 412. It will be appreciated that the number and direction of the internal aerofoil cooling passages 436 may be provided as required for a particular cooling performance. For example, the internal aerofoil cooling passages 436 may be meandering or single pass and may traverse the blade in chord rather than in span, or both.

The span of the internal aerofoil cooling passages 436 may go beyond the platform 422 and into the shank or root of the blade. The aerofoil cooling passages 436 may be defined by internal walls or webs which span between the internal surfaces of the pressure and suction walls, and by portions of the pressure and suction surface walls themselves. The aerofoil cooling passages 436 may extend fully across the width blade between the pressure and suction surfaces 418, 420, or there may be separate passages for each of the pressure and suction surface walls individually. There may also be internal aerofoil cooling passages which are defined in full or partially only by internal walls which do not directly form part of the pressure and suction surface 420 walls. The aerofoil cooling passages 436 may have a suction side and a pressure side relative to the camber line of the aerofoil 412, the camber line being an imaginary line which passages mid-width between the leading and trailing edges 414, 416 of the aerofoil 412 as is well understood in the art.

In the example shown in Figure 4, the internal aerofoil cooling passages 436 provide a four pass meandering flow path in which air is received at an infeed towards the leading edge 414 before 424 spanning the aerofoil 412 between root and tip four times whilst meandering downstream. Thus, a cooling passage may be a plurality of radially extending passages which are connected in series by returns in the form of u-bends as shown in Figure 3.

The relative size of the aerofoil cooling passages 436, infeeds and exhausts, internal cooling features and cooling hole outlets 438 may be determined as required for a given cooling objective. It will be appreciated that the invention may be applicable to any suitable internal aerofoil cooling passages 436.

The platform 422 extends laterally from each part of the root end of the aerofoil to aft 428, fore 426 and circumferential edges 428, 430. More specifically there is a platform leading edge 426, a platform trailing edge 428, a platform pressure surface edge 428, and suction surface edge 430, with the four edges defining the peripheral edges to the platform 422. The platform 422 is polygonal in planform and may have four sides. The platform 422 may be a parallelogram as shown in Figure 4.

The aerofoil 412 is located towards a central portion of the platform 422 so as to be fully surrounded thereby. The aerofoil 412 is placed at a stagger angle on the platform 422 to provide the necessary angle of incidence with the upstream air exiting the nozzle guide vane as well known in the art.

In addition to the peripheral edges of the platforms 422, the radially outward facing surface of the platform 422 includes a pressure surface platform portion 440, a suction surface platform portion 442, a leading edge platform portion 444 and a trailing edge platform portion 446. The leading edge platform portion 444 may be taken to be any surface axially forward of the fore 426 most portion of the aerofoil 412 where it joins the platform 422. The trailing edge platform portion 446 may be taken to be any portion aft 428 of the trailing edge 416 where the aerofoil 412 joins the platform 422. In the described embodiment, part of the trailing edge platform portion 446 is aft of the suction surface 420 due to the stagger of the aerofoil 412. The suction side portion is the area between the a leading edge platform portion 444 and trailing edge platform portion 446 on the suction side of the aerofoil 412, and the pressure side platform portion 440 is the area between the leading and trailing edge platform portions 444, 446 on the pressure side of the aerofoil 412. The platform 422 extends axially between platform 422 leading and trailing edges 414, 416, and circumferentially between circumferential edges.

The circumferential edges 428, 430 of the platform 422 are provided by a radially orientated face which circumferentially opposes a corresponding face of an adjacent blade. The edges may be axially aligned to the principal axis of the engine to provide a rectangular platform 422 when viewed radially inwards in plan, or may be at an angle to so as to be offset from the main gas path flow, thereby providing a parallelogram shape.

The opposing circumferential edges of adjacent blade platforms 422 are typically separated by a small gap to allow for relative in-use movements between the platforms 422. The radial gap that exists between the platforms 422 may include a damper or seal as known in the art. The seal may be a strip seal housed within a groove similar to that shown in Figure 3. The leading and trailing platform 422 edges may lie in a plane which is normal to the principal axis so as to have a uniform axial position across the circumferential length. However, curved leading and trailing edges 414,416 are known. Curved circumferential edges are also known.

As shown in Figure 5, the leading edge 414 of the platform 422 extends axially upstream of the leading edge 414 of the blade. The platform 422 may terminate in one or more extensions which overhang 448 a shank or root of the blade and the corresponding portions of the disc. The overhang 448 may include one or more cantilevered flanges.

The overhang 448 may be located downstream of a trailing edge 450 of the upstream nozzle guide vane platform 452 so as to provide a flow restriction therebetween. In the described embodiment, the overhang 448 includes a seal fin which projects upstream into stator-rotor gap as best seen in Figure 5. The seal fin is located radially inwards of the vane platform 452 and axially overlaps with it to provide a tortuous path which restricts the flow through the gap from the inboard side into the main gas path. In use, air is provided radially inwards of the seal fin at a higher pressure to provide a positive pressure. Thus, a leading edge 414 of the platform 422 may aerodynamically engage with a trailing edge 416 of an upstream nozzle guide vane platform 452 to provide a labyrinth path. The labyrinth path provides a restriction to and consequential reduction in the cooling flow required to prevent overheating and ingestion of the hot gas from the main flow path radially inwards of the platforms 422.

Although portions of the seal fin, and thus overhang 448, may not directly from part of the main gas path wall, they are still considered to be part of the platform 422 for the present disclosure.

Figure 5 shows the root portion of the aerofoil component 410 including part of the shank which extends between the root (not shown) and platform 422, the shank being distinguishable by not directly engaging with the disc slot. A portion of the aerofoil 412 is shown extending radially outwards from the platform 422 at the leading edge 414. A radial thickness of the platform 422 either side of the aerofoil 412 is indicated by the dotted line.

A first and a second cooling passage are shown within the aerofoil 412 section. The aerofoil cooling passages 436 may extend radially away from the platform 422 along the span of the aerofoil 412 and be one of a plurality of such passages.

The platform 422 overhang 448 is tapered from the full thickness of the platform 422 to a thinner cantilevered seal fin. The extent of the taper will be dictated by the radial and axial location of the overhang 448 tip and provides a radially inward slope away from the nominal height of the interaerofoil 412 platform 422. The slope may begin downstream, at, or upstream of the leading edge 414 of the blade which will be determined by other considerations.

The platform 422 is provided on the blade root which engages with a corresponding rotor slot as described in connection to Figures 2 and 3. The blade root may be any suitable engagement feature but will typically be a so-called fir tree root or a dove tail root.

The axial end of the blade root is capped with a plate, a lockplate 454, which may aid sealing of the turbine blade root and provide axial retention. There may be a rear lockplate and a front lockplate 454. Lockplates are well known in the art {ACH - REF}. An example of a known rear lockplate 54 can be seen in Figure 2.

The lockplate 454 is axially retained by a retention feature which may be attached to but is typically formed within the underside of the platform 422. The retention feature may be in the form of a lockplate groove 456. The lockplate groove 456 may be provided by a slot in the radial inner (underside) surface of the platform 422. The lockplate groove 456 extends circumferentially across the underside of the platform 422 and may extend fully between the circumferential edges of the blade platform 422. The lockplate groove 456 may be provided between the axially inner surface of a circumferential flange which extends radially inwards from the underside of the platform 422 and a portion of the blade shank. Alternatively, the groove may be provided by a slot which is cast or machined into the underside of the platform 422.

The lockplate 454 is located between the fore end of the blade shank and the tip 426 of the overhang 448. The lockplate 454 may be located below the platform 422 or overhang 448. In the described embodiment, the lockplate 454 is located axially fore 424 of the leading edge 414 of the aerofoil 412 where it joins the platform 422 and radially inwards of the overhang 448.

The radially inner edge of the lockplate 454 (not shown) may be retained by a suitable engagement on the disc or an associated component. The engagement may be another groove or a clamping arrangement. There is typically one lockplate 454 per blade. The lockplates are proximal to or abut adjacent corresponding lockplates to provide an annular disc around an axial face of the blade roots.

In use, the lockplate 454 is rotated with the turbine rotor and experiences a centrifugal force which loads the underside of the platform 422. This creates operational load and associated stress on the platform 422 which is generally undesirable but the deleterious effects can be reduced by placing additional material around the area to strengthen it, or by increasing the cooling of the area.

As best shown in Figure 4, the platform 422 includes a plurality of platform cooling passages 432 which extend laterally through the platform 422 from a central portion to an extremity thereof. The cooling passages are partially, possibly fully, contained within the radial thickness of the platform 422 and extend from an inlet 434 to an outlet 438.

The inlets 434 to the platform cooling passages 432 may be in fluid communication with the internal aerofoil cooling passages 436 of the blade. The outlet 438 of each platform 422 cooling passage is located in an external surface of the platform 422. The surface may be an edge. The edge may be a fore 426 edge, an aft 428 edge or one of the circumferential edges. The outlets 438 may be distributed along the edges of the platforms 422. Alternatively or additionally, the cooling passage outlets 438 may be provided in one or more radially outer or radially inner platform 422 surfaces. The radially outer and inner platform 422 surfaces may be the radially outer main gas path defining surface or a radially inwards facing surface which defines a cavity with the blade shank and/or the disc posts which define the root receiving disc slots. The platform 422 surface which includes a platform 422 cooling outlet 438 may be on an inter-blade platform 422 surface or on the leading 448 or trailing edge 446 platform portions.

In the example of Figure 4, the platform cooling passages 432 extend in flow series from an internal aerofoil cooling passage 436 to outlets 438 distributed about the periphery of platform 422. In particular, there are outlets 438 on the platform 422 trailing edge 416, the pressure surface 418 platform 422 edge and the suction surface 420 platform 422 edge. It will be appreciated that the distribution of the platform cooling passages 432may include a different distribution to the one shown and any or all of the platform 422 surfaces and edges may include one or more, or no, cooling passage outlets 438.

The leading edge platform portion 444 may include an overhang 448 as described above. A portion of the cooling passages may extend upstream of the blade towards and into the leading edge platform portion 444. The leading edge platform portion 444 cooling passages may terminate on a radially outer or radially inner surface of the overhang 448. Hence, one or more platform cooling passages 432may include an outlet 438 in the underside of the overhang 448. An outlet 438 positioned here may introduce a flow of cooling air into the seal gap 458, or the air chamber which is located radially inwards of the seal gap 458 and provides air thereto. In some instances, the outlet 438 may be placed in the overhang 448 tip end face. However, this will be dependent at least on the thickness of the overhang 448 which will need to accommodate the cooling passage and sufficient material on either side to provide the mechanical requirements and manufacturing tolerances.

In the example of Figure 5, the leading edge 414 platform 422 cooling passage extends from an internal aerofoil cooling passage 436 to an outlet 438 on the underside of the overhang 448. The passage extends over the radially outer edge of the lockplate groove 456 and lockplate 454. More specifically, the outlet 438 is located on the underside of the overhang 448 between the overhang 448 tip and lockplate 454 retention feature. The cooling passage may optionally open out into the lockplate groove 456 or the groove flange.

The angle of inclination of the leading edge 414 platform 422 cooling passage taken relative to the rotational axis of the engine may be horizontal or radially outwards away from the inlet 434, or radially inwards away from the inlet 434. Altering the angle of inclination may allow the relative position of the aerofoil cooling passage 436, lockplate groove 456, desired outlet 438 position and mass which is to be cooled to be accommodated.

The platform cooling passages 432are provided in a radial arrangement relative to the longitudinal axis or radial axis of the aerofoil 412. The distribution when viewed in plan could be described as a starburst or radially exploding arrangement.

The distribution of the cooling passages may be specified to avoid positioning the inlets in areas of high stress and to ensure adequate 'run through' distance inside the aerofoil cooling passage to avoid 'back wall impingement' defects during manufacture. In some embodiments, the distribution may be selected to manipulate the internal air flow structure for example to prevent flow recirculation, or to target a particular pressure ratio across the hole in order to, for example, prevent ingestion.

The cooling passages are in the form of elongate conduits which may be straight and may have a round transverse cross-section. The transverse cross-section may be circular or oval for example.

The platform cooling passages 432may be different sizes. There may have two diameters. The first diameter may be in the range 0.3mm to 0.7mm. The second diameter may be in the range 0.3mm to 0.7mm The platform cooling passages 432may have one or more bifurcations and thus one or more inlets 434 or outlets 438. The platform cooling passages 432 may be manufactured by boring the platform 422, using for example, electro-discharge machining.

The platform cooling passages 432 shown in Figure 4 extend from the aerofoil cooling passages 436 which are located within the root portion of the aerofoil 412 and platform 422. However, the platform cooling passages 432 may be provided elsewhere if there is a suitable or more convenient source of cooling air in the vicinity of the turbine rotor. Such a source of cooling air may be provided in the shank, root or underside of the platform 422 for example.

The platform cooling passages 432 may be non-parallel when viewed in plan from a radially inward facing direction. The platform cooling passages 432 may be provided in a fanned distribution in which the inlets 434 are spaced closer together than the outlets 438. When straight, this provides at least some of the platform cooling passages 432 with an axial and circumferential component rather than being exclusively circumferential or axial.

The platform cooling passages 432 may extend across multiple portions of the platform 422. Thus, as can be seen in Figure 4, the cooling passages have an inlet 434 on the suction side of the blade and an outlet 438 on the platform 422 trailing edge 416. Also shown is an example where the cooling passages extend from any of the pressure surface 418, leading edge 414 or suction surface 420 of the aerofoil 412 before 424 extending axially forwards into the a leading edge platform portion 444 . In some instances, the platform cooling passages 432extend forwards into the overhang 448.

As described below, the platform cooling passages 432may also have a radial component. Each of the circumferential, radial and axial components of the platform cooling passages 432 may be different from an adjacent cooling passage and may be different to any of the other platform cooling passages 432 within the platform 422.

The platform cooling passages 432 may be distributed along the pressure surface 418 platform 422 portions and suction surface 420 platform 422 portions. The pressure surface 418 platform 422 portion may include more cooling passages than the suction surface 420 platform 422 portion. The inlets 434 may be distributed evenly or unevenly alone the pressure and suction surface 418, 420 sides. The inlets 434 may be in fluid communication with the aerofoil cooling passages 436. The inlets 434 may be provided on an inner surface of the aerofoil cooling passages 436 along the suction side or pressure side of the aerofoil 412. By suction side and pressure side, it will be appreciated that the cooling passages are located on the suction side or pressure side of the aerofoil 412 relative to the camber line. Thus, an inlet 434 may be provided on a central aerofoil 412 passageway which is separated from a pressure or suction surface 420 wall by an internal wall such may be provided in a dual wall pressure or suction surface 420 aerofoil 412 arrangement.

The inlets 434 may be provided in clusters. Each cluster may be in fluid communication with an aerofoil cooling passage 436 such that a first cluster of platform cooling passages 432 are fed from a first aerofoil cooling passage 436, whereas a second cluster are fed from a second aerofoil cooling passage 436. Thus, the inlets 434 may be grouped together such that each of the internal aerofoil cooling passages 436 feed a plurality of the platform 422 cooling passages. The inlets 434 may be located towards or at the leading edge 414 region of the aerofoil 412, or at a mid-chord portion thereof. The aerofoil cooling passages 436 may be different sections of a single cooling passage. The sections may be in flow series. The sections may be axially separated. The sections may be separated along the flow series by at least one return in the flow path. The inlet 434 clusters may be at significantly different pressures and temperatures.

The a leading edge platform portion 444 may include a plurality of cooling passages which extend from the aerofoil 412 towards the leading edge 414. As can be seen in Figure 4, the leading edge 414 portion cooling passages may terminate downstream of the overhang 448 tip, or downstream the overhang 448 altogether. The leading edge 414 portion cooling passages may terminate at common axial location. Hence, the outlets 438 may all be provided in a circumferentially extending line.

One aim of the platform cooling passages 432 is to provide an even thermal distribution and reduce any problematic thermal gradients throughout the platform 422. The thermal gradients may be determined by a local thickness, cooling air or the main gas path temperature. The thermal loading or thermal gradient may be more problematic in some areas rather than others. For example, the centrifugal loading imparted to the platform 422 on or local to the overhang 448 can cause increased levels of stress and potential creep over the service life of the part.

The platform cooling passages 432along the trailing edge 416 of the platform 422 may have inlets 434 distributed exclusively on the suction side of the aerofoil 412. The suction side may also have inlets 434 for cooling passages for the suction side of the platform 422 and/or the leading edge 414 portion. The number of passages on the suction side may be less than the number on the pressure surface 418. The suction side outlets 438 may be distributed predominantly towards the leading and trailing edges 414, 416. The pressure surface 418 outlets 438 may be distributed along the full axial length of the platform 422.

Figure 6 shows a ghost view of a section of the platform 422 trailing edge 416. Thus, there is shown the platform 422 trailing edge 416 face, a platform 422 suction surface 420 edge face, two aerofoil cooling passages 436 and a plurality of platform cooling passages 432 extending therebetween.

The platform trailing edge face surface may be a radial face which faces axially rearwards towards the downstream nozzle guide vane of the following turbine stage. The platform suction surface 420 may be a radial face on the circumferential periphery of the platform 422 which faces the circumferential edge of the adjacent platform 422 in use.

The aerofoil cooling passages 436 may be in axial flow series and may be connected via a return as shown in Figure 6. The inlets 434 may be on the radial portion of the aerofoil cooling passage 436 or the return.

The inlets 434 may be radially and laterally distributed across the aerofoil cooling passages 436.

The outlets 438 may be placed at a common radially height or altered so as to be at different radial positions in the platform thickness at the edge end. The outlets 438 may be at mid-span of the platform edge surface thickness or be located towards the radial inner or outer surface of the platform 422. The outlets 438 may be located within a seal groove.

The platform cooling passages 432 may be inclined relative to one another to account for the radial thickness of the platform 422 and the relative positions of the inlet 434 and outlet 438.

The aerofoil 412 will typically be integrally formed with the platform 422 and other parts of the blade to provide a monolith. The blade may be formed by as a single crystal cast component as commonly known in the art.

Although the above disclosure is predominantly concerned with a turbine blade, it will be appreciated that the invention may be applicable to a nozzle guide vane.

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

Claims (13)

Claims:
1. An aerofoil (412) component of a gas turbine engine, comprising:
an aerofoil having leading and trailing edges (414, 416) with pressure and suction surfaces (418, 420) extending therebetween and at least one aerofoil cooling passage (436) for providing internal and external cooling for the aerofoil component;
a platform (422) extending laterally from a root of the aerofoil at the pressure surface suction surface, leading edge and trailing edge, and having fore (424), aft (426) and circumferential peripheral edges (428, 430); the platform including a plurality of platform cooling passages (432) each of which extend from an inlet (434) which is open to the at least one aerofoil cooling passages (436) to an outlet (438) in a surface of the platform and, wherein the inlets of the platform cooling passages are spaced closer than the outlets.
2. An aerofoil component as claimed in claim 1, wherein at least one platform cooling passage has a leading edge platform portion outlet on the upstream side of the leading edge of the aerofoil.
3. An aerofoil component as claimed in claim 2, wherein the a leading edge platform portion outlet is located in a surface of an overhang (448) which extends axially fore of the leading edge of the aerofoil where the aerofoil joins the platform.
4. An aerofoil component as claimed in claim 3, wherein the leading edge platform portion outlet is located in a radially inner surface of the overhang.
5. An aerofoil component as claimed in claim 4, wherein the overhang is a seal fin.
6. An aerofoil component as claimed in any of claims 2 to 5, wherein the platform includes a lockplate engagement feature on the radially inner surface of the platform.
7. An aerofoil component as claimed in claim 6, wherein the leading edge platform cooling passage extends across the lockplate engagement feature.
8. An aerofoil component as claimed in any preceding claim, wherein the inlets are provided in clusters.
9. An aerofoil component as claimed in claim 8, wherein the inlets within each cluster are radially distributed along the at least one aerofoil cooling passage.
10. An aerofoil component as claimed in claim 9, having a plurality of aerofoil cooling passages 436 wherein each cluster of inlets is opened to different aerofoil cooling passages.
11. An aerofoil component as claimed in any preceding claim, wherein the outlets of the platform cooling passages are provided on one or more of the circumferential, fore or aft edges of the platform.
12. An aerofoil component as claimed in claim 11, wherein the circumferential edges include a pressure surface platform edge and a suction surface platform edge, wherein the pressure surface platform edge includes a greater number of outlets than the suction surface edge.
13. An aerofoil component as claimed in claim 12, wherein the aft edge includes more platform 5 cooling passage outlets than the suction surface edge.
Intellectual Property Office
GB1801551.1A 2018-01-31 2018-01-31 A cooling arrangement for a gas turbine engine aerofoil component platform Pending GB2570652A (en)

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Application Number Priority Date Filing Date Title
GB1801551.1A GB2570652A (en) 2018-01-31 2018-01-31 A cooling arrangement for a gas turbine engine aerofoil component platform

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GB2570652A true GB2570652A (en) 2019-08-07

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Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1528224A2 (en) * 2003-10-31 2005-05-04 General Electric Company Method and apparatus for cooling gas turbine engine rotor blade
US20110223005A1 (en) * 2010-03-15 2011-09-15 Ching-Pang Lee Airfoil Having Built-Up Surface with Embedded Cooling Passage
EP2589749A2 (en) * 2011-11-04 2013-05-08 General Electric Company Bucket assembly for turbine system
WO2014186005A2 (en) * 2013-02-15 2014-11-20 United Technologies Corporation Gas turbine engine component with combined mate face and platform cooling
EP2990608A1 (en) * 2014-06-27 2016-03-02 Mitsubishi Hitachi Power Systems, Ltd. Rotor blade and gas turbine equipped with same

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1528224A2 (en) * 2003-10-31 2005-05-04 General Electric Company Method and apparatus for cooling gas turbine engine rotor blade
US20110223005A1 (en) * 2010-03-15 2011-09-15 Ching-Pang Lee Airfoil Having Built-Up Surface with Embedded Cooling Passage
EP2589749A2 (en) * 2011-11-04 2013-05-08 General Electric Company Bucket assembly for turbine system
WO2014186005A2 (en) * 2013-02-15 2014-11-20 United Technologies Corporation Gas turbine engine component with combined mate face and platform cooling
EP2990608A1 (en) * 2014-06-27 2016-03-02 Mitsubishi Hitachi Power Systems, Ltd. Rotor blade and gas turbine equipped with same

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