EP2746536A1 - Rotor stage of a turbine - Google Patents
Rotor stage of a turbine Download PDFInfo
- Publication number
- EP2746536A1 EP2746536A1 EP13196186.4A EP13196186A EP2746536A1 EP 2746536 A1 EP2746536 A1 EP 2746536A1 EP 13196186 A EP13196186 A EP 13196186A EP 2746536 A1 EP2746536 A1 EP 2746536A1
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- EP
- European Patent Office
- Prior art keywords
- tip
- turbine
- rib
- cooling
- range
- 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.)
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- 238000005304 joining Methods 0.000 claims abstract description 8
- 238000001816 cooling Methods 0.000 claims description 101
- 230000001419 dependent effect Effects 0.000 claims 1
- 239000007789 gas Substances 0.000 description 30
- 239000002826 coolant Substances 0.000 description 17
- 238000000926 separation method Methods 0.000 description 6
- 241000826860 Trapezium Species 0.000 description 4
- 238000002485 combustion reaction Methods 0.000 description 4
- 230000001141 propulsive effect Effects 0.000 description 3
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- 230000007423 decrease Effects 0.000 description 2
- 238000010438 heat treatment Methods 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 238000004513 sizing Methods 0.000 description 2
- 230000002411 adverse Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000015556 catabolic process Effects 0.000 description 1
- 230000001010 compromised effect Effects 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 238000006731 degradation reaction Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000003628 erosive effect Effects 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 238000002844 melting Methods 0.000 description 1
- 230000008018 melting Effects 0.000 description 1
- 238000000034 method Methods 0.000 description 1
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- 230000002093 peripheral effect Effects 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/20—Specially-shaped blade tips to seal space between tips and stator
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/10—Two-dimensional
- F05D2250/13—Two-dimensional trapezoidal
Definitions
- the present invention relates to an aerofoil structure, typically of a turbine blade for a gas turbine engine, and in particular a structure of the tip of the aerofoil.
- a 'squealer' blade 114 has a tip plate 116 extending over the pressure and suction surface walls. Tip rib, fins or fences 118, 120 extend from the tip plate 116 and define a tip cavity 122.
- This blade 124 is known as a 'winglet' type and comprises tip wings 126 and 128 that overhang the pressure and suction surfaces 130, 132 respectively.
- the tip wings 126 and 128 define a channel 134, which can have an inlet 136 and/or an outlet 138 at the leading and trailing edges respectively.
- the function of the cavity 122 or channel 134 is to trap working gas that leaks over the peripheral wall on the pressure side of the blade.
- the trapped gas forms one or more vortices within the cavity or channel and/or over the outward surface of the tip fins 118, 120 or wings 126, 128. These vortices inhibit the over-tip leakage flow continuing to the suction side.
- These configurations serve to avoid losses in efficiency caused by working gas leakage over the turbine blade tips and also to avoid losses caused by flow disturbances set up by the leakage flow.
- a disadvantage of these configurations is that they are complex and as a result are difficult to manufacture.
- the increased surface area associated with these configurations leads to an increased heat load in the tip of the blade and it is difficult to achieve effective cooling of these blade configurations. Thus both designs are prone to degradation and short in-service life caused by high temperatures and severe thermal gradients.
- a further disadvantage of the winglet configuration is the additional weight of the blade tip; this is a particular issue considering the very high centrifugal forces present in a modern gas turbine engine.
- the squealer configuration particularly suffers from relatively high heat transfer into the blade and has limited cooling; therefore in-service life of the blade is compromised.
- FIG 4 is a sectional view A-A shown on Figure 2 .
- the over-tip leakage 144 flows over the tip from a pressure surface 146 to a suction surface 148.
- An over-tip leakage gap G is defined between the tip outer surface 150 and the casing surface 140.
- the squealer design relies on flow separation 154 at the pressure surface corner 152 of the squealer tip 120. The flow separation or vortex causes a blockage leading to a reduced overall discharge over the tip.
- the present invention seeks to address the issues of minimizing over tip leakage and cooling the tips of blades.
- a rotor stage of a turbine comprising a rotational axis, a shroud and radially inward thereof a turbine blade defined partly by a pressure side wall, a suction side wall and a tip portion, the tip portion has a pressure side tip rib and a tip cavity floor defining a tip cavity, the pressure side tip rib has a width W ps , a height H ps above the tip cavity and defines a tip gap G ps with the shroud, the pressure side tip rib has a sloping side joining the tip cavity floor and having an angle ⁇ ps to a radial line ZZ, the width W ps is in the range G ps to 5G ps , the height H ps is in the range 5G ps to 15G ps and the angle ⁇ ps is in the range 20° to 70°.
- the sloping side provides more flexibility for the positioning of cooling chambers within the blade so as to achieve improved cooling of the blade tip.
- An angle in the range of 20° to 70° is selected because this provides a balance between providing an increased space for cooling chambers within the blade tip and the need to ensure good aerodynamic performance. It has been found that with this angle range open type flow separation occurs over the rib, leading to a large vortical flow structure in the tip cavity, the above selected angle ensuring that there is a sufficiently large cavity volume to ensure such a vortical flow structure.
- Aerodynamic performance has been found to be improved with a smaller width W ps .
- a problem with squealer tips of the prior art is that a smaller width W ps results in a higher specific heat load per unit volume, and hence increased problems of internal cooling.
- the present inventors have found that providing a sloping side enables the space within which a cooling cavity can be located to be increased compared to the prior art, so that cooling of the tip can be improved whilst having an aerodynamically beneficial width W ps . In some cases it may be possible to reduce the width W ps to a width narrower than the prior art.
- the width W ps may be in the range 2G ps to 3G ps
- the height H ps may be in the range 7G ps to 8G ps
- the angle ⁇ ps may be in the range 40° to 60°.
- An angle between 40 to 60° increases the flexibility for the provision of cooling chambers, whilst maintaining a large tip cavity.
- the width W ps may be approximately 2.5G ps
- the height H ps may be approximately 7G ps
- the angle ⁇ ps may be approximately 45°.
- the sloping side may extend to a maximum radial extent of the turbine blade. That is, the pressure side tip rib may have a radially outer surface and the sloping side may extend directly between the radially outer surface and the tip cavity floor.
- the pressure side tip rib could be considered to be trapezoidal in shape with the side of the trapezium on the pressure side of the turbine blade forming a right angle with what could be considered the base of the trapezium.
- Such an arrangement of the sloping side increases the volume of the pressure side tip rib which increases the flexibility for positioning and sizing of cooling chambers within the turbine blade.
- the tip portion may have a suction side tip rib
- the suction side tip rib may have a width W ss and may define a tip gap G ss with the shroud
- the suction side tip rib may have a sloping side joining the tip cavity floor and having an angle ⁇ ss to a radial line ZZ
- the width W ss may be in the range G ss to 5G ss
- the height H ss may be in the range 5G ss to 15G ss
- the angle ⁇ ss may be in the range 20° to 70°.
- Such a construction and dimensions of the suction side tip rib has similar advantages to the pressure side tip rib, but on the suction side of the turbine blade.
- the width W ss may be in the range 2G ss to 3G ss
- the height H ss may be in the range 7G ss to 8G ss
- the angle ⁇ ss may be in the range 40° to 60°.
- the width W ss may be approximately 2.5G ss
- the height H ss may be approximately 7G ss
- the angle ⁇ ss may be approximately 45°.
- the sloping side may extend to a maximum radial extent of the turbine blade. That is, the suction side tip rib may have a radially outer surface and the sloping side may extend directly between the radially outer surface and the tip cavity floor.
- the suction side tip rib could be considered to be trapezoidal in shape with the side of the trapezium on the suction side of the turbine blade forming a right angle with what could be considered the base of the trapezium.
- Such an arrangement of the sloping side increases the volume of the suction side tip rib which increases the flexibility for positioning and sizing of cooling chambers within the turbine blade.
- the turbine blade may comprise a rib that extends around the tip of the turbine blade, wherein the rib comprises the pressure side tip rib and the suction side tip rib.
- the width of the tip rib, and/or the width W ps of the pressure side tip rib and/or the width W ss of the suction side tip rib may vary around the turbine blade. Such an arrangement permits the width of the rib(s) to be varied so that the width of the rib(s) can be reduced where the cooling requirements are less.
- the width W ss may be made thinner than the width W ps because there are less cooling requirements on the suction side. Reducing the width W ss advantageously reduces the weight of the turbine blade.
- the rib that extends around the tip of the turbine blade may have a sloping side joining the tip cavity floor.
- the angle of the sloping side of the rib, and/or the angle ⁇ ps of the pressure side tip rib and/or the angle ⁇ ss of the suction side tip rib may vary around the turbine blade.
- the angle ⁇ ps of the pressure side tip rib may be approximately 40° and the angle ⁇ ss of the pressure side tip rib may be approximately 60°.
- the angle may be selected to suit the local heat load at the respective position of the rib.
- a cooling gallery may be defined at least partly within the tip rib (e.g. the pressure side tip rib, the suction side tip rib, and/or the entire tip rib that extends around the tip of the turbine blade and comprises the pressure side tip rib and the suction side tip rib). Extending the cooling gallery into the tip rib improves cooling of the blade tip. Improved cooling can contribute to extended blade life span. Furthermore, improving the cooling of the blade tip means that it is possible to further reduce the width W ps so as to improve aerodynamic performance.
- At least 30% of the cooling gallery may be defined in the tip rib.
- at least 40% or at least 50% of the cooling gallery may be defined in the tip rib.
- the blade may comprise an internal cooling passage.
- the blade may comprise a cooling hole extending from the internal cooling passage to the cooling gallery.
- the cooling gallery may be an extension of the internal cooling passage.
- a cooling gallery may be defined wholly within the tip rib (e.g. the pressure side tip rib, the suction side tip rib, and/or the entire tip rib).
- a cooling gallery may be located at least partly above the tip cavity floor.
- the cooling gallery may extend into the tip rib in a radial direction by approximately a distance of at least one third of H ps radially outward from (or above) the floor of the tip cavity, alternatively by a distance of at least one half of H ps . The further the cooling gallery extends into the tip rib the greater the cooling of the tip rib.
- At least a major portion of the cooling gallery may be situated radially outward of the floor of the cavity.
- a cooling gallery may be located above or radially outwardly of the floor of the tip cavity, e.g. entirely above (or radially outward from) the tip cavity floor.
- At least one cooling hole may extend from the cooling gallery to an external surface of the tip rib, an inlet of the cooling hole may be located radially outwardly of the floor of the tip cavity.
- a first cooling hole may extend from the cooling gallery to the pressure side of the pressure side tip rib and a second cooling hole may extend to a position at or near a junction between an outer radial surface of the pressure side tip rib and the sloping side.
- the first and second cooling holes provide the function of film cooling, and the second cooling hole additionally contributes to controlling flow separation, which can further improve aerodynamic performance.
- the cooling gallery may extend from the pressure side to the suction side.
- the cooling gallery may be formed of an array of cooling sub-galleries.
- the tip cavity may be defined by curved surfaces forming the outer surface of a wall of generally constant thickness.
- 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 low-pressure 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.
- the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air-cooling of these airfoil components.
- the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
- Figure 6 shows an isometric view of a typical single stage cooled high-pressure turbine.
- the present tip arrangement can also be applied to other turbine or even compressor stages of a gas or steam turbine, whether for an aero, industrial or marine engine. Cooling air-flows are indicated by arrows.
- High-pressure turbine nozzle guide vanes 31 consume the greatest amount of cooling air on high temperature engines.
- High-pressure blades 32 typically use about half of the NGV coolant flow.
- the intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
- the high-pressure turbine airfoils are cooled by using high-pressure air from one of the compressors 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 cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air as effectively as possible.
- a turbine blade 32 has a longitudinally extending aerofoil portion 36 with facing suction side 37 and pressure side 38 walls.
- the aerofoil portion 36 extends across the working gas annulus, with the longitudinal direction of the aerofoil portion being arranged generally along a radial direction of the engine.
- the turbine blade 32 has a root portion 44 radially inward of the aerofoil and a tip portion 46 radially outward of the aerofoil.
- the suction side 37 and pressure side 38 walls meet at a leading edge 48 and a trailing edge 50.
- the root portion engages a rotor disc 52 via complimentary dovetail or in this example, fir-tree fixtures 54.
- Radially outward of the tip portion 46 is the casing shroud 35.
- the blade, disc and casing shroud form a rotor stage 56.
- a multi-pass cooling passage 38 is fed cooling air 42 by a feed passage 40 formed in the root portion 44 of the blade.
- a second cooling air feed 60 can supply additional coolant for the trailing edge 50 of the blade that can be particularly prone to thermal erosion. Cooling air leaves the multi-pass cooling passage through effusion holes 62, 64 in the aerofoil surfaces and particularly in the leading and trailing edges 48, 50 of the blade to create a film of cooling air over the aerofoil surfaces 37, 38.
- the block arrows in Figure 6 show the general direction of cooling airflow.
- the arrows in Figure 7 show in more detail the internal cooling flow directions in this multipass cooling passage system.
- a tip clearance control arrangement 66 capable of cooling or heating the shroud segment 35 to dilate or contract the shroud segments to maintain a desired position relative to the blade tip.
- cooling or heating fluid can be fed via holes 68 to impinge onto the shroud segment.
- Other mechanical tip clearance control systems whether active or scheduled, can be used and are well known in the art.
- the turbine blade 32 is defined partly by the pressure side wall 38, the suction side wall 37 and the tip portion 46.
- the tip portion 46 has a pressure side tip rib 70 and a tip cavity floor 74 partly defining a tip cavity 76.
- the pressure side tip rib 70 has a radially outer surface 71 having a width W ps , a height H ps above the tip cavity floor and defines the tip gap G ps with the shroud segment 35.
- the pressure side tip rib has a sloping side 78 joining the tip cavity floor 76 and the radially outer surface 71.
- the sloping side 78 has an angle ⁇ ps to the radial line ZZ.
- a particularly effective envelope of the pressure side tip rib dimensions is considered to be that the width W ps is in the range 2G ps to 5G ps , the height H ps is in the range 7G ps to 15G ps and the angle ⁇ ps is in the range 40° to 60°.
- One example of a tip rib has its dimensions where the width W ps is approximately 2.5G ps , the height H ps is approximately 7G ps and the angle ⁇ ps is approximately 45°.
- the gap G ps is a nominal gap size at an engine or a rotor stage design point.
- This design point is can be at a number of engine or rotor operational conditions such as at cruise or take-off. During a normal flight cycle the engine and rotor stage are usually at the cruise condition for the longest period of time and therefore the design point is chosen here to provide the greatest efficiency.
- the design point can also be at take-off where the engine is producing its greatest thrust and therefore the working gas is at its hottest and greatest flow. Thus at take-off conditions it can be advantageous to design the gap G ps to occur.
- tip clearance control systems can be used in an attempt to maintain the gap G ps to its design point.
- the suction side tip rib is also instrumental in reducing over tip leakage and heat loading in to the blade.
- the presently described blade comprises the tip portion 46 having a suction side tip rib 72.
- the suction side tip rib has a width W ss and defines a tip gap G ss with the shroud 35.
- the suction side tip rib 72 has a sloping side 80 joining the tip cavity floor 74 with its radially outer surface 73.
- the sloping side 80 has an angle ⁇ ss to a radial line ZZ.
- the width W ss is in the range G ss to 5G ss
- the height H ss is in the range 5G ss to 10G ss
- the angle ⁇ ss is in the range 20° to 70°.
- a particularly effective envelope of the suction side tip rib dimensions is considered to be that the width W ss is in the range 2G ss to 3G ss , the height H ss is in the range 7G ss to 15G ss and the angle ⁇ ss is in the range 40° to 60°.
- One example of a tip rib has its dimensions where the width W ss is approximately 2.5G ss , the height H ss is approximately 7G ss and the angle ⁇ ss is approximately 45°.
- the pressure side tip rib 70 described above can be used in conjunction with a conventional rectangular tip rib on the suction side or the above described suction side tip rib 72.
- the suction side tip rib 72 described above can also be used in conjunction with a conventional rectangular tip rib on the pressure side.
- the sloping side of the suction side rib may be angled at a different angle ⁇ ss to the angle ⁇ ps of the sloping side of the pressure side rib.
- the angle ⁇ ps may be 40° and the angle ⁇ ss may be 60°, or any other angle within the range specified, the angle being selected to suit local heat loading.
- the angle of the sloping side may continuously vary around the aerofoil to further tune the sloping sides to local heat loading.
- the angle ⁇ ps of the sloping side of the pressure side rib may be equal to the angle ⁇ ss of the sloping side of the suction side rib.
- the dimensions of the tip ribs 70, 72 may be constant along the chord of the blade from the leading to the trailing edge. Alternatively, the dimensions of the tip ribs 70, 72 may vary in any one of more of their dimensions for the best possible aerodynamic shape to minimise over tip leakage. Any variation in the dimensions can be because of the curvature of the blade between leading and trailing edges and/or the direction of the working gases and/or the mass flow of the over tip leakage flow. Further, varying the dimensions of the tip rib can reduce the weight of the blade.
- a cooling gallery 82 is defined at least partly within the tip rib 70 itself. Preferably, at least a major part of the cooling gallery 82 is situated radially outward of the floor of the cavity 76. With certain tip rib configurations, the whole of the cooling gallery 82 is positioned radially outwardly of the floor of the tip cavity.
- the cooling holes 87, 88, 89 can be arranged such that their inlets are located radially outward of the floor of the cavity 76.
- Coolant 84 is fed from one of the main multipass cooling passages 38 through a supply passage 86 and into the tip rib cooling gallery 82.
- the gallery 82 can supply any one or more cooling holes 87, 88, 89 that can be provided which in turn then supply any one or more of the external surfaces of the tip rib with coolant. Not only can the coolant in the gallery 82 remove heat by convection from the bulk mass of the blade tip rib, but also the coolant passing through the supply passage 86 can impinge on a surface of the gallery to reduce heat load.
- the cooling holes 87, 88, 89 can form a film of coolant over the surrounding surface to help prevent hot gases directly impinging on the surface and can also locally add to and cool the hot working gases which spill over the tip.
- the gallery 82 can extend part of or all the chord length of the blade between leading edge to the trailing edge.
- the gallery can extend around the leading edge and/or trailing edge so that it forms a single gallery.
- the blade can include a number of galleries 82', 82", 82'" aligned in the tip rib and supplied with different sources of coolant as can be seen in Figure 7 .
- the galleries 82', 82", 82'" can be supplied coolant from different sources or parts of the multipass passage 38 and blade coolant supplies 42, 60 via supply passages 86', 86", 86"' respectively.
- the gallery 82 is shown having a cross-sectional profile generally corresponding to the profile of the tip rib so that the walls have an approximately similar thickness; however, other cross-sectional profiles can be used such as circular or elliptical.
- Figure 10 shows an alternative blade 32 where the galleries 82 are now extensions of the main coolant passage 38 arranged within the main aerofoil of the blade.
- This embodiment has common features as described with reference to Figure 8 .
- the relevant angles of the surfaces 78 and 80 can be equated with the dashed lines 90, 92 which represent approximate centre-lines of the wall defining these surfaces.
- the same cooling holes 87, 88, 89 as shown in Figure 9 can supply coolant to the external surfaces of the tip ribs 70, 72 in similar manner.
- This alternative blade also has a tip cavity 76 defined by curved surfaces 78, 74 and 80 that can help prevent localised recirculation of hot gases that spill over the tip ribs.
- the curved surface also helps to reduce stress at otherwise sharp geometric changes between sloping surfaces 78, 80 and the tip cavity floor.
- the wall forming the tip cavity floor 74 and the surfaces 78, 80 can be formed having a generally constant thickness T.
- the aerofoil can be lighter, less complex to manufacture and can result in lower pressure loss for the coolant and supply of coolant through cooling holes 87, 88, 89.
- the cooling gallery 82 is an extension of the coolant passage 38.
- the sides of the tip are substantially straight angled sides, instead of curved sides.
- the cooling gallery 82 substantially follows the profile of the tip.
- the cooling gallery may take any suitable form.
- cooling gallery is shown.
- the cooling gallery is elliptical.
- one cooling passage 87 is shown extending to the pressure side of the turbine blade to provide the function of film cooling.
- a further passage 89 is shown. Said further passage is located at a position where the radially outer surface connects with the sloping side. The further passage 89 provides both film cooling and contributes to flow separation control.
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Abstract
Description
- The present invention relates to an aerofoil structure, typically of a turbine blade for a gas turbine engine, and in particular a structure of the tip of the aerofoil.
- For turbine rotor blades and particularly high pressure (HP) turbine blades, there is an industry wide and an ever-important object to minimise over-tip leakage of hot working gases between a tip of the blades and a casing. In general, there are three types of tip geometry configurations which attempt to minimise over tip leakage: un-shrouded, partially shrouded and fully shrouded.
- As shown in
Figure 1 , the simplest un-shroudedrotor blade tip 112 of arotor blade 110 is a flat-tip arrangement having a generally planar radially outward facing surface. However, this flat-tip design is typically associated with considerable aerodynamic and heat losses due to the over-tip leakage flow of the main working gas. To reduce over-tip leakage, other blade tip configurations have been proposed. For example, inFigure 2 a 'squealer'blade 114 has atip plate 116 extending over the pressure and suction surface walls. Tip rib, fins orfences tip plate 116 and define atip cavity 122. - A further example is shown in
Figure 3 . Thisblade 124 is known as a 'winglet' type and comprisestip wings suction surfaces tip wings channel 134, which can have aninlet 136 and/or anoutlet 138 at the leading and trailing edges respectively. - In either the squealer or winglet blade configurations, the function of the
cavity 122 orchannel 134 is to trap working gas that leaks over the peripheral wall on the pressure side of the blade. The trapped gas forms one or more vortices within the cavity or channel and/or over the outward surface of thetip fins wings - A further disadvantage of the winglet configuration is the additional weight of the blade tip; this is a particular issue considering the very high centrifugal forces present in a modern gas turbine engine.
- The squealer configuration particularly suffers from relatively high heat transfer into the blade and has limited cooling; therefore in-service life of the blade is compromised.
-
Figure 4 is a sectional view A-A shown onFigure 2 . Considering the squealer configuration in more detail and with reference toFigure 4 where there is shown asurface 140 of acasing 142; the over-tipleakage 144 flows over the tip from apressure surface 146 to asuction surface 148. An over-tip leakage gap G is defined between the tipouter surface 150 and thecasing surface 140. To reduce the over-tip leakage flow, the squealer design relies onflow separation 154 at thepressure surface corner 152 of thesquealer tip 120. The flow separation or vortex causes a blockage leading to a reduced overall discharge over the tip. This flow separation or vortex is more effective if the pressure surface squealer step is made thinner (i.e. with a smaller W). On the other hand, a thinner squealer will have a larger area-tovolume ratio and hence would be subject to higher specific heat load per unit volume. As a result, current squealer designs are mostly unable to explore the full aerodynamic potential of the squealer working mechanism due to the restrictions arisen from the enhanced heat transfer and the poor 'coolability'. - Thus the current squealer configuration leads to conflicting performance characteristics between the aerodynamics and heat transfer.
- The present invention seeks to address the issues of minimizing over tip leakage and cooling the tips of blades.
- In a first aspect of the invention there is provided a rotor stage of a turbine comprising a rotational axis, a shroud and radially inward thereof a turbine blade defined partly by a pressure side wall, a suction side wall and a tip portion, the tip portion has a pressure side tip rib and a tip cavity floor defining a tip cavity, the pressure side tip rib has a width Wps, a height Hps above the tip cavity and defines a tip gap Gps with the shroud, the pressure side tip rib has a sloping side joining the tip cavity floor and having an angle α ps to a radial line ZZ, the width Wps is in the range Gps to 5Gps, the height Hps is in the range 5Gps to 15Gps and the angle α ps is in the range 20° to 70°.
- The sloping side provides more flexibility for the positioning of cooling chambers within the blade so as to achieve improved cooling of the blade tip.
- An angle in the range of 20° to 70° is selected because this provides a balance between providing an increased space for cooling chambers within the blade tip and the need to ensure good aerodynamic performance. It has been found that with this angle range open type flow separation occurs over the rib, leading to a large vortical flow structure in the tip cavity, the above selected angle ensuring that there is a sufficiently large cavity volume to ensure such a vortical flow structure.
- Aerodynamic performance has been found to be improved with a smaller width Wps. However, a problem with squealer tips of the prior art is that a smaller width Wps results in a higher specific heat load per unit volume, and hence increased problems of internal cooling. The present inventors have found that providing a sloping side enables the space within which a cooling cavity can be located to be increased compared to the prior art, so that cooling of the tip can be improved whilst having an aerodynamically beneficial width Wps. In some cases it may be possible to reduce the width Wps to a width narrower than the prior art.
- The width Wps may be in the range 2Gps to 3Gps, the height Hps may be in the range 7Gps to 8Gps and the angle α ps may be in the
range 40° to 60°. An angle between 40 to 60° increases the flexibility for the provision of cooling chambers, whilst maintaining a large tip cavity. - The width Wps may be approximately 2.5Gps, the height Hps may be approximately 7Gps and the angle α ps may be approximately 45°.
- The sloping side may extend to a maximum radial extent of the turbine blade. That is, the pressure side tip rib may have a radially outer surface and the sloping side may extend directly between the radially outer surface and the tip cavity floor. For example, the pressure side tip rib could be considered to be trapezoidal in shape with the side of the trapezium on the pressure side of the turbine blade forming a right angle with what could be considered the base of the trapezium. Such an arrangement of the sloping side increases the volume of the pressure side tip rib which increases the flexibility for positioning and sizing of cooling chambers within the turbine blade.
- The tip portion may have a suction side tip rib, the suction side tip rib may have a width Wss and may define a tip gap Gss with the shroud, the suction side tip rib may have a sloping side joining the tip cavity floor and having an angle α ss to a radial line ZZ, the width Wss may be in the range Gss to 5Gss, the height Hss may be in the range 5Gss to 15Gss and the angle α ss may be in the range 20° to 70°. Such a construction and dimensions of the suction side tip rib has similar advantages to the pressure side tip rib, but on the suction side of the turbine blade.
- The width Wss may be in the range 2Gss to 3Gss, the height Hss may be in the range 7Gss to 8Gss and the angle α ss may be in the
range 40° to 60°. - The width Wss may be approximately 2.5Gss, the height Hss may be approximately 7Gss and the angle α ss may be approximately 45°.
- The sloping side may extend to a maximum radial extent of the turbine blade. That is, the suction side tip rib may have a radially outer surface and the sloping side may extend directly between the radially outer surface and the tip cavity floor. For example, the suction side tip rib could be considered to be trapezoidal in shape with the side of the trapezium on the suction side of the turbine blade forming a right angle with what could be considered the base of the trapezium. Such an arrangement of the sloping side increases the volume of the suction side tip rib which increases the flexibility for positioning and sizing of cooling chambers within the turbine blade.
- The turbine blade may comprise a rib that extends around the tip of the turbine blade, wherein the rib comprises the pressure side tip rib and the suction side tip rib.
- The width of the tip rib, and/or the width Wps of the pressure side tip rib and/or the width Wss of the suction side tip rib may vary around the turbine blade. Such an arrangement permits the width of the rib(s) to be varied so that the width of the rib(s) can be reduced where the cooling requirements are less. For example, the width Wss may be made thinner than the width Wps because there are less cooling requirements on the suction side. Reducing the width Wss advantageously reduces the weight of the turbine blade.
- The rib that extends around the tip of the turbine blade may have a sloping side joining the tip cavity floor. The angle of the sloping side of the rib, and/or the angle α ps of the pressure side tip rib and/or the angle α ss of the suction side tip rib may vary around the turbine blade. For example, the angle α ps of the pressure side tip rib may be approximately 40° and the angle α ss of the pressure side tip rib may be approximately 60°. The angle may be selected to suit the local heat load at the respective position of the rib.
- A cooling gallery may be defined at least partly within the tip rib (e.g. the pressure side tip rib, the suction side tip rib, and/or the entire tip rib that extends around the tip of the turbine blade and comprises the pressure side tip rib and the suction side tip rib). Extending the cooling gallery into the tip rib improves cooling of the blade tip. Improved cooling can contribute to extended blade life span. Furthermore, improving the cooling of the blade tip means that it is possible to further reduce the width Wps so as to improve aerodynamic performance.
- At least 30% of the cooling gallery may be defined in the tip rib. For example, at least 40% or at least 50% of the cooling gallery may be defined in the tip rib.
- The blade may comprise an internal cooling passage. The blade may comprise a cooling hole extending from the internal cooling passage to the cooling gallery. Alternatively, the cooling gallery may be an extension of the internal cooling passage.
- A cooling gallery may be defined wholly within the tip rib (e.g. the pressure side tip rib, the suction side tip rib, and/or the entire tip rib).
- A cooling gallery may be located at least partly above the tip cavity floor. For example, the cooling gallery may extend into the tip rib in a radial direction by approximately a distance of at least one third of Hps radially outward from (or above) the floor of the tip cavity, alternatively by a distance of at least one half of Hps. The further the cooling gallery extends into the tip rib the greater the cooling of the tip rib.
- At least a major portion of the cooling gallery may be situated radially outward of the floor of the cavity.
- A cooling gallery may be located above or radially outwardly of the floor of the tip cavity, e.g. entirely above (or radially outward from) the tip cavity floor.
- At least one cooling hole may extend from the cooling gallery to an external surface of the tip rib, an inlet of the cooling hole may be located radially outwardly of the floor of the tip cavity.
- A first cooling hole may extend from the cooling gallery to the pressure side of the pressure side tip rib and a second cooling hole may extend to a position at or near a junction between an outer radial surface of the pressure side tip rib and the sloping side. In such an embodiment, the first and second cooling holes provide the function of film cooling, and the second cooling hole additionally contributes to controlling flow separation, which can further improve aerodynamic performance.
- The cooling gallery may extend from the pressure side to the suction side.
- The cooling gallery may be formed of an array of cooling sub-galleries.
- The tip cavity may be defined by curved surfaces forming the outer surface of a wall of generally constant thickness.
-
-
Figure 1 is a view of a tip portion of a known turbine blade having a flat tip surface, -
Figure 2 is a view of a tip portion of a known turbine blade having a squealer tip configuration, -
Figure 3 is a view of a tip portion of a known turbine blade having a winglet tip configuration, -
Figure 4 is a sectional view A-A shown onFigure 2 and showing a known tip fin. -
Figure 5 is a schematic longitudinal cross-section through a ducted fan gas turbine engine in which the present invention is incorporated, -
Figure 6 is an isometric view of a typical single stage cooled turbine of the gas turbine described with reference toFigure 4 , -
Figure 7 shows a chord-wise cross-section through a high-pressure turbine blade incorporating the present invention, -
Figure 8 is a schematic section of a tip portion and shroud of a rotor stage of a turbine; the tip portion includes an arrangement of tip ribs and is in accordance with the present invention, -
Figure 9 is a schematic section of a tip portion of a blade and shroud of a rotor stage of a turbine; the tip portion includes a cooling scheme in accordance with the present invention, -
Figure 10 is a schematic section of an alternative tip portion of a blade and shroud of a rotor stage of a turbine; the tip portion includes a cooling scheme in accordance with the present invention; -
Figure 11 is a schematic section of an alternative pressure side tip portion of a blade; -
Figure 12 is a schematic section of a further alternative pressure side tip portion of a blade; and -
Figure 13 is a schematic section of a yet further alternative pressure side tip portion of a blade. - With reference to
Figure 5 , 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, anair intake 11, apropulsive fan 12, anintermediate pressure compressor 13, a high-pressure compressor 14,combustion equipment 15, a high-pressure turbine 16, andintermediate pressure turbine 17, a low-pressure turbine 18 and a coreengine exhaust nozzle 19. Anacelle 21 generally surrounds theengine 10 and defines theintake 11, abypass duct 22 and abypass exhaust nozzle 23. - The
gas turbine engine 10 works in a conventional manner so that air entering theintake 11 is accelerated by thefan 12 to produce two air flows: a first air flow A into theintermediate pressure compressor 13 and a second air flow B which passes through thebypass duct 22 to provide propulsive thrust. Theintermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to thehigh pressure compressor 14 where further compression takes place. - The compressed air exhausted from the high-
pressure compressor 14 is directed into thecombustion 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 low-pressure turbines nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high andintermediate pressure compressors fan 12 by suitable interconnecting shafts. - The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
- In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air-cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
-
Figure 6 shows an isometric view of a typical single stage cooled high-pressure turbine. The present tip arrangement can also be applied to other turbine or even compressor stages of a gas or steam turbine, whether for an aero, industrial or marine engine. Cooling air-flows are indicated by arrows. - Internal convection and external coolant films are the prime methods of cooling the gas path components such as
aerofoils 36,platforms 34, shrouds 33 andcasing shroud segments 35 etc. High-pressure turbine nozzle guide vanes 31 (NGV) consume the greatest amount of cooling air on high temperature engines. High-pressure blades 32 typically use about half of the NGV coolant flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air. - The high-pressure turbine airfoils are cooled by using high-pressure air from one of the compressors 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 cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air as effectively as possible.
- Ever increasing gas temperature levels combined with a drive towards flatter combustion radial temperature profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the extremities of the blades and vanes, and the working gas annulus end-walls.
- Referring to
Figures 6 and7 , aturbine blade 32 has a longitudinally extendingaerofoil portion 36 with facingsuction side 37 andpressure side 38 walls. Theaerofoil portion 36 extends across the working gas annulus, with the longitudinal direction of the aerofoil portion being arranged generally along a radial direction of the engine. Theturbine blade 32 has aroot portion 44 radially inward of the aerofoil and atip portion 46 radially outward of the aerofoil. Thesuction side 37 andpressure side 38 walls meet at aleading edge 48 and a trailingedge 50. The root portion engages arotor disc 52 via complimentary dovetail or in this example, fir-tree fixtures 54. Radially outward of thetip portion 46 is thecasing shroud 35. The blade, disc and casing shroud form arotor stage 56. - In this exemplary embodiment, a
multi-pass cooling passage 38 is fed coolingair 42 by afeed passage 40 formed in theroot portion 44 of the blade. A secondcooling air feed 60 can supply additional coolant for the trailingedge 50 of the blade that can be particularly prone to thermal erosion. Cooling air leaves the multi-pass cooling passage through effusion holes 62, 64 in the aerofoil surfaces and particularly in the leading and trailingedges Figure 6 show the general direction of cooling airflow. The arrows inFigure 7 show in more detail the internal cooling flow directions in this multipass cooling passage system. - Radially outwardly of the turbine blade is the
casing shroud segment 35 and partially shown a tipclearance control arrangement 66 capable of cooling or heating theshroud segment 35 to dilate or contract the shroud segments to maintain a desired position relative to the blade tip. As is well known in the tip clearance control field cooling or heating fluid can be fed viaholes 68 to impinge onto the shroud segment. Other mechanical tip clearance control systems, whether active or scheduled, can be used and are well known in the art. - Referring now to
Figures 8 and 9 , theturbine blade 32 is defined partly by thepressure side wall 38, thesuction side wall 37 and thetip portion 46. Thetip portion 46 has a pressureside tip rib 70 and atip cavity floor 74 partly defining atip cavity 76. The pressureside tip rib 70 has a radially outer surface 71 having a width Wps, a height Hps above the tip cavity floor and defines the tip gap Gps with theshroud segment 35. The pressure side tip rib has asloping side 78 joining thetip cavity floor 76 and the radially outer surface 71. The slopingside 78 has an angle α ps to the radial line ZZ. - The problems of known blade tip configurations as described in the background section, are greatly resolved by the present tip configuration by a tip rib having the width Wps in the range Gps to 5Gps, the height Hps is in the range 5Gps to 10Gps and the angle α ps is in the range 20° to 60°.
- A particularly effective envelope of the pressure side tip rib dimensions is considered to be that the width Wps is in the range 2Gps to 5Gps, the height Hps is in the range 7Gps to 15Gps and the angle α ps is in the
range 40° to 60°. One example of a tip rib has its dimensions where the width Wps is approximately 2.5Gps, the height Hps is approximately 7Gps and the angle α ps is approximately 45°. - The gap Gps is a nominal gap size at an engine or a rotor stage design point. This design point is can be at a number of engine or rotor operational conditions such as at cruise or take-off. During a normal flight cycle the engine and rotor stage are usually at the cruise condition for the longest period of time and therefore the design point is chosen here to provide the greatest efficiency. The design point can also be at take-off where the engine is producing its greatest thrust and therefore the working gas is at its hottest and greatest flow. Thus at take-off conditions it can be advantageous to design the gap Gps to occur. As is well known, tip clearance control systems can be used in an attempt to maintain the gap Gps to its design point.
- Of primary importance is the above defined pressure side tip rib; however, the suction side tip rib is also instrumental in reducing over tip leakage and heat loading in to the blade. Thus the presently described blade comprises the
tip portion 46 having a suctionside tip rib 72. The suction side tip rib has a width Wss and defines a tip gap Gss with theshroud 35. The suctionside tip rib 72 has asloping side 80 joining thetip cavity floor 74 with its radiallyouter surface 73. The slopingside 80 has an angle α ss to a radial line ZZ. In general, the width Wss is in the range Gss to 5Gss, the height Hss is in the range 5Gss to 10Gss and the angle α ss is in the range 20° to 70°. - A particularly effective envelope of the suction side tip rib dimensions is considered to be that the width Wss is in the range 2Gss to 3Gss, the height Hss is in the range 7Gss to 15Gss and the angle αss is in the
range 40° to 60°. One example of a tip rib has its dimensions where the width Wss is approximately 2.5Gss, the height Hss is approximately 7Gss and the angle α ss is approximately 45°. - The pressure
side tip rib 70 described above can be used in conjunction with a conventional rectangular tip rib on the suction side or the above described suctionside tip rib 72. The suctionside tip rib 72 described above can also be used in conjunction with a conventional rectangular tip rib on the pressure side. - The sloping side of the suction side rib may be angled at a different angle α ss to the angle α ps of the sloping side of the pressure side rib. For example, the angle α ps may be 40° and the angle α ss may be 60°, or any other angle within the range specified, the angle being selected to suit local heat loading. The angle of the sloping side may continuously vary around the aerofoil to further tune the sloping sides to local heat loading. In alternative embodiments the angle α ps of the sloping side of the pressure side rib may be equal to the angle α ss of the sloping side of the suction side rib.
- The dimensions of the
tip ribs tip ribs - A further and important aspect of the above described dimensions of the tip ribs is that their cross-sectional profile lends itself to improved cooling or coolability over conventional tip ribs. As can be seen in
Figures 7 and9 , acooling gallery 82 is defined at least partly within thetip rib 70 itself. Preferably, at least a major part of thecooling gallery 82 is situated radially outward of the floor of thecavity 76. With certain tip rib configurations, the whole of thecooling gallery 82 is positioned radially outwardly of the floor of the tip cavity. The cooling holes 87, 88, 89 can be arranged such that their inlets are located radially outward of the floor of thecavity 76. -
Coolant 84 is fed from one of the mainmultipass cooling passages 38 through asupply passage 86 and into the tiprib cooling gallery 82. Thegallery 82 can supply any one or more cooling holes 87, 88, 89 that can be provided which in turn then supply any one or more of the external surfaces of the tip rib with coolant. Not only can the coolant in thegallery 82 remove heat by convection from the bulk mass of the blade tip rib, but also the coolant passing through thesupply passage 86 can impinge on a surface of the gallery to reduce heat load. The cooling holes 87, 88, 89 can form a film of coolant over the surrounding surface to help prevent hot gases directly impinging on the surface and can also locally add to and cool the hot working gases which spill over the tip. - The
gallery 82 can extend part of or all the chord length of the blade between leading edge to the trailing edge. The gallery can extend around the leading edge and/or trailing edge so that it forms a single gallery. Alternatively, the blade can include a number ofgalleries 82', 82", 82'" aligned in the tip rib and supplied with different sources of coolant as can be seen inFigure 7 . Thegalleries 82', 82", 82'" can be supplied coolant from different sources or parts of themultipass passage 38 and blade coolant supplies 42, 60 viasupply passages 86', 86", 86"' respectively. - The
gallery 82 is shown having a cross-sectional profile generally corresponding to the profile of the tip rib so that the walls have an approximately similar thickness; however, other cross-sectional profiles can be used such as circular or elliptical. -
Figure 10 shows analternative blade 32 where thegalleries 82 are now extensions of themain coolant passage 38 arranged within the main aerofoil of the blade. This embodiment has common features as described with reference toFigure 8 . The relevant angles of thesurfaces lines Figure 9 , can supply coolant to the external surfaces of thetip ribs - This alternative blade also has a
tip cavity 76 defined bycurved surfaces surfaces - The wall forming the
tip cavity floor 74 and thesurfaces - Referring to
Figure 11 a portion of a further alternative blade is shown. Similarly to the blade shown inFigure 10 , the coolinggallery 82 is an extension of thecoolant passage 38. However, in the embodiment shown inFigure 11 the sides of the tip are substantially straight angled sides, instead of curved sides. In the present embodiment, the coolinggallery 82 substantially follows the profile of the tip. However, in alternative embodiments the cooling gallery may take any suitable form. - Referring to
Figures 12 and 13 a further alternative configuration of cooling gallery is shown. In the examples shown inFigures 12 and 13 the cooling gallery is elliptical. In the embodiment shown inFigure 12 , onecooling passage 87 is shown extending to the pressure side of the turbine blade to provide the function of film cooling. InFigure 13 afurther passage 89 is shown. Said further passage is located at a position where the radially outer surface connects with the sloping side. Thefurther passage 89 provides both film cooling and contributes to flow separation control.
Claims (15)
- A rotor stage of a turbine comprising
a rotational axis (XX),
a shroud (35) and radially inward thereof
a turbine blade (32) defined partly by a pressure side wall (38), a suction side wall (37) and a tip portion (46),
the tip portion (46) has a pressure side tip rib (70) and a tip cavity floor (74) defining a tip cavity (76),
the pressure side tip rib has a width Wps, a height Hps above the tip cavity and defines a tip gap Gps with the shroud,
the pressure side tip rib has a sloping side (78) joining the tip cavity floor and having an angle α ps to a radial line ZZ,
the width Wps is in the range Gps to 5Gps,
the height Hps is in the range 5Gps to 15Gps and
the angle α ps is in the range 20° to 70°. - A rotor stage of a turbine as claimed in claim 1, wherein
the width Wps is in the range 2Gps to 3Gps,
the height Hps is in the range 7Gps to 8Gps and
the angle α ps is in the range 40° to 60°. - A rotor stage of a turbine as claimed in claim 1 or 2, wherein the pressure side tip rib has a radially outer surface and the sloping side directly extends between the radially outer surface and the tip cavity floor.
- A rotor stage of a turbine as claimed in any previous claim, wherein the tip portion (46) has a suction side tip rib (72),
the suction side tip rib has a width Wss and defines a tip gap Gss with the shroud,
the suction side tip rib has a sloping side (80) joining the tip cavity floor and having an angle α ss to a radial line ZZ,
the width Wss is in the range Gss to 5Gss,
the height Hss is in the range 5Gss to 15Gss and
the angle α ss is in the range 20° to 70°. - A rotor stage of a turbine as claimed in claim 4, wherein
the width Wss is in the range 2Gss to 3Gss,
the height Hss is in the range 7Gss to 8Gss and
the angle α ss is in the range 40° to 60°. - A rotor stage of a turbine as claimed in any one of the previous claims, comprising a tip rib, wherein the tip rib comprises the pressure side tip rib, and wherein a cooling gallery (82) is defined at least partly within a tip rib.
- A rotor stage of a turbine as claimed in claim 6, wherein a major portion of the cooling gallery is situated radially outward of the tip cavity floor (74).
- A rotor stage of a turbine as claimed in claim 6 or 7, wherein the blade comprises an internal cooling passage (38, 60) and the cooling gallery (82) is an extension of the internal cooling passage.
- A rotor stage of a turbine as claimed in claim 8, wherein the tip cavity is defined by curved surfaces forming the outer surface of a wall of generally constant thickness (T).
- A rotor stage of a turbine as claimed in claim 6 or 7, wherein a cooling gallery (82) is defined wholly within the tip rib.
- A rotor stage of a turbine as claimed in any one of claims 6 to 10, wherein a cooling gallery (82) is located above the tip cavity floor.
- A rotor stage of a turbine as claimed in any one of claims 6 to 11, wherein at least one cooling hole (87, 88, 89) extends from the cooling gallery (82) to an external surface of the pressure side tip rib, an inlet of the cooling hole is located radially outwardly of the floor of the tip cavity.
- A rotor stage of a turbine as claimed in any one of claims 6 to 12 as dependent on claim 4 or 5, wherein the tip rib comprises the suction side rib and the cooling gallery is defined at least partly within the suction side tip rib.
- A rotor stage of a turbine as claimed in claim 13, wherein the cooling gallery (82) extends from the pressure side to the suction side.
- A rotor stage of a turbine as claimed in any one of claims 6 to 14, wherein the cooling gallery (82) is formed of an array of cooling sub-galleries.
Applications Claiming Priority (1)
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GBGB1223193.2A GB201223193D0 (en) | 2012-12-21 | 2012-12-21 | Turbine blade |
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EP2746536A1 true EP2746536A1 (en) | 2014-06-25 |
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Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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EP13196186.4A Withdrawn EP2746536A1 (en) | 2012-12-21 | 2013-12-09 | Rotor stage of a turbine |
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US (1) | US20140178207A1 (en) |
EP (1) | EP2746536A1 (en) |
GB (1) | GB201223193D0 (en) |
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WO2014186022A1 (en) * | 2013-03-13 | 2014-11-20 | General Electric Company | Turbine blade for improving dust mitigation |
US9850762B2 (en) | 2013-03-13 | 2017-12-26 | General Electric Company | Dust mitigation for turbine blade tip turns |
CN105909315A (en) * | 2015-02-25 | 2016-08-31 | 通用电气公司 | Turbine rotor blade |
EP3061914A1 (en) * | 2015-02-25 | 2016-08-31 | General Electric Company | Turbine rotor blade and corresponding gas turbine engine |
CN104775854A (en) * | 2015-04-23 | 2015-07-15 | 华能国际电力股份有限公司 | Movable vane top structure capable of restraining tip leakage and reducing tip temperature |
US10107108B2 (en) | 2015-04-29 | 2018-10-23 | General Electric Company | Rotor blade having a flared tip |
EP3428398A1 (en) * | 2017-07-13 | 2019-01-16 | General Electric Company | Airfoil and corresponding method of cooling a tip rail |
WO2019212478A1 (en) * | 2018-04-30 | 2019-11-07 | Siemens Aktiengesellschaft | Turbine blade tip with multi-outlet cooling channels |
EP3575555A1 (en) * | 2018-05-07 | 2019-12-04 | Rolls-Royce Corporation | Gas turbine engine blade and corresponding method of forming |
US11655717B2 (en) | 2018-05-07 | 2023-05-23 | Rolls-Royce Corporation | Turbine blade squealer tip including internal squealer tip cooling channel |
EP3623577A1 (en) * | 2018-09-12 | 2020-03-18 | United Technologies Corporation | Gas turbine engine airfoil tip cooling arrangement with purge partition |
US10801334B2 (en) | 2018-09-12 | 2020-10-13 | Raytheon Technologies Corporation | Cooling arrangement with purge partition |
JP7370823B2 (en) | 2018-12-03 | 2023-10-30 | ゼネラル・エレクトリック・カンパニイ | Turbine blade tip cooling system including tip rail cooling insert |
EP4056806A3 (en) * | 2021-03-09 | 2022-11-16 | Mechanical Dynamics & Analysis LLC | Turbine blade with tip cooling hole supply plenum |
US11840940B2 (en) | 2021-03-09 | 2023-12-12 | Mechanical Dynamics And Analysis Llc | Turbine blade tip cooling hole supply plenum |
Also Published As
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US20140178207A1 (en) | 2014-06-26 |
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