EP2746536A1 - Rotorstufe einer Turbine - Google Patents

Rotorstufe einer Turbine Download PDF

Info

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
Authority
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.)
Withdrawn
Application number
EP13196186.4A
Other languages
English (en)
French (fr)
Inventor
Li He
Qiang Zhang
Anthony Rawlinson
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Publication of EP2746536A1 publication Critical patent/EP2746536A1/de
Withdrawn legal-status Critical Current

Links

Images

Classifications

    • 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
    • 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
    • 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
    • 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.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP13196186.4A 2012-12-21 2013-12-09 Rotorstufe einer Turbine Withdrawn EP2746536A1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GBGB1223193.2A GB201223193D0 (en) 2012-12-21 2012-12-21 Turbine blade

Publications (1)

Publication Number Publication Date
EP2746536A1 true EP2746536A1 (de) 2014-06-25

Family

ID=47682419

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13196186.4A Withdrawn EP2746536A1 (de) 2012-12-21 2013-12-09 Rotorstufe einer Turbine

Country Status (3)

Country Link
US (1) US20140178207A1 (de)
EP (1) EP2746536A1 (de)
GB (1) GB201223193D0 (de)

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2014186022A1 (en) * 2013-03-13 2014-11-20 General Electric Company Turbine blade for improving dust mitigation
CN104775854A (zh) * 2015-04-23 2015-07-15 华能国际电力股份有限公司 一种具有抑制叶顶泄漏和减小叶顶温度的动叶顶部结构
CN105909315A (zh) * 2015-02-25 2016-08-31 通用电气公司 涡轮转子叶片
US10107108B2 (en) 2015-04-29 2018-10-23 General Electric Company Rotor blade having a flared tip
EP3428398A1 (de) * 2017-07-13 2019-01-16 General Electric Company Schaufelprofil und zugehöriges verfahren zur kühlung einer spitzenwand
WO2019212478A1 (en) * 2018-04-30 2019-11-07 Siemens Aktiengesellschaft Turbine blade tip with multi-outlet cooling channels
EP3575555A1 (de) * 2018-05-07 2019-12-04 Rolls-Royce Corporation Laufschaufel eines gasturbinentriebwerks und zugehöriges herstellungsverfahren
EP3623577A1 (de) * 2018-09-12 2020-03-18 United Technologies Corporation Turbinenschaufelspitzenkühlanordnung mit spülteilung
EP4056806A3 (de) * 2021-03-09 2022-11-16 Mechanical Dynamics & Analysis LLC Turbinenlaufschaufel mit versorgungsplenum für schaufelspitzenkühlöffnungen
JP7370823B2 (ja) 2018-12-03 2023-10-30 ゼネラル・エレクトリック・カンパニイ 先端レール冷却インサートを含むタービンブレード先端冷却システム

Families Citing this family (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6462332B2 (ja) * 2014-11-20 2019-01-30 三菱重工業株式会社 タービン動翼及びガスタービン
JP6565105B2 (ja) * 2015-09-29 2019-08-28 三菱日立パワーシステムズ株式会社 動翼及びこれを備えるガスタービン
CN105179022B (zh) * 2015-09-30 2017-06-27 北京大学 一种采用叶顶肋翼结构的涡轮叶片
US10006294B2 (en) * 2015-10-19 2018-06-26 General Electric Company Article and method of cooling an article
US10260355B2 (en) 2016-03-07 2019-04-16 Honeywell International Inc. Diverging-converging cooling passage for a turbine blade
US10227878B2 (en) * 2016-03-10 2019-03-12 General Electric Company Article and method of forming an article
US10519861B2 (en) * 2016-11-04 2019-12-31 General Electric Company Transition manifolds for cooling channel connections in cooled structures
US11434770B2 (en) 2017-03-28 2022-09-06 Raytheon Technologies Corporation Tip cooling design
US11021967B2 (en) * 2017-04-03 2021-06-01 General Electric Company Turbine engine component with a core tie hole
US10545057B2 (en) 2017-06-22 2020-01-28 Unison Industries, Llc Air temperature sensor and method of reducing error
US10585007B2 (en) 2017-06-22 2020-03-10 Unison Industries, Llc Air temperature sensor
US10578498B2 (en) 2017-06-22 2020-03-03 Unison Industries, Llc Air temperature sensor
USD844466S1 (en) 2017-06-22 2019-04-02 Unison Industries, Llc Temperature sensor
US10605675B2 (en) 2017-06-22 2020-03-31 Unison Industries, Llc Air temperature sensor
US10605098B2 (en) * 2017-07-13 2020-03-31 General Electric Company Blade with tip rail cooling
KR101937588B1 (ko) * 2017-09-13 2019-01-10 두산중공업 주식회사 터빈의 냉각 블레이드 및 이를 포함하는 터빈 및 가스터빈
US10570750B2 (en) 2017-12-06 2020-02-25 General Electric Company Turbine component with tip rail cooling passage
US10408065B2 (en) 2017-12-06 2019-09-10 General Electric Company Turbine component with rail coolant directing chamber
EP3765714B1 (de) 2018-03-14 2024-01-03 General Electric Technology GmbH Kühlanordnung für eine turbinenanordnung
CN108506047A (zh) * 2018-03-15 2018-09-07 哈尔滨工业大学 具有三维曲面叶顶的轴流式涡轮叶栅
US11352889B2 (en) * 2018-12-18 2022-06-07 General Electric Company Airfoil tip rail and method of cooling
KR102162969B1 (ko) * 2019-02-26 2020-10-07 두산중공업 주식회사 터빈 블레이드 및 이를 포함하는 가스 터빈
US11299991B2 (en) * 2020-04-16 2022-04-12 General Electric Company Tip squealer configurations
EP4234885A3 (de) * 2021-02-04 2023-09-06 Doosan Enerbility Co., Ltd. Schaufelblatt mit einem system zur kühlung verjüngter schaufelspitzen für eine turbinenlaufschaufel, turbinenlaufschaufel, turbinenlaufschaufelanordnung, gasturbine und verfahren zur herstellung eines schaufelblatts

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5660523A (en) * 1992-02-03 1997-08-26 General Electric Company Turbine blade squealer tip peripheral end wall with cooling passage arrangement
US5738491A (en) * 1997-01-03 1998-04-14 General Electric Company Conduction blade tip
US20040179940A1 (en) * 2003-03-12 2004-09-16 Florida Turbine Technologies, Inc. Multi-metered film cooled blade tip
US20050111979A1 (en) * 2003-11-26 2005-05-26 Siemens Westinghouse Power Corporation Cooling system for a tip of a turbine blade
US20090092500A1 (en) * 2005-06-24 2009-04-09 Snecma Hollow turbomachine blade
US20100047057A1 (en) * 2008-06-30 2010-02-25 Rolls-Royce Plc Aerofoil
KR20120048439A (ko) * 2010-11-05 2012-05-15 한국항공대학교산학협력단 스퀄러팁을 갖는 가스터빈 블레이드

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19944923B4 (de) * 1999-09-20 2007-07-19 Alstom Turbinenschaufel für den Rotor einer Gasturbine
US6602052B2 (en) * 2001-06-20 2003-08-05 Alstom (Switzerland) Ltd Airfoil tip squealer cooling construction
US7997865B1 (en) * 2008-09-18 2011-08-16 Florida Turbine Technologies, Inc. Turbine blade with tip rail cooling and sealing
US8366394B1 (en) * 2010-10-21 2013-02-05 Florida Turbine Technologies, Inc. Turbine blade with tip rail cooling channel

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5660523A (en) * 1992-02-03 1997-08-26 General Electric Company Turbine blade squealer tip peripheral end wall with cooling passage arrangement
US5738491A (en) * 1997-01-03 1998-04-14 General Electric Company Conduction blade tip
US20040179940A1 (en) * 2003-03-12 2004-09-16 Florida Turbine Technologies, Inc. Multi-metered film cooled blade tip
US20050111979A1 (en) * 2003-11-26 2005-05-26 Siemens Westinghouse Power Corporation Cooling system for a tip of a turbine blade
US20090092500A1 (en) * 2005-06-24 2009-04-09 Snecma Hollow turbomachine blade
US20100047057A1 (en) * 2008-06-30 2010-02-25 Rolls-Royce Plc Aerofoil
KR20120048439A (ko) * 2010-11-05 2012-05-15 한국항공대학교산학협력단 스퀄러팁을 갖는 가스터빈 블레이드

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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 (zh) * 2015-02-25 2016-08-31 通用电气公司 涡轮转子叶片
EP3061914A1 (de) * 2015-02-25 2016-08-31 General Electric Company Turbinenlaufschaufel und zugehöriges gasturbinenkraftwerk
CN104775854A (zh) * 2015-04-23 2015-07-15 华能国际电力股份有限公司 一种具有抑制叶顶泄漏和减小叶顶温度的动叶顶部结构
US10107108B2 (en) 2015-04-29 2018-10-23 General Electric Company Rotor blade having a flared tip
EP3428398A1 (de) * 2017-07-13 2019-01-16 General Electric Company Schaufelprofil und zugehöriges verfahren zur kühlung einer spitzenwand
WO2019212478A1 (en) * 2018-04-30 2019-11-07 Siemens Aktiengesellschaft Turbine blade tip with multi-outlet cooling channels
EP3575555A1 (de) * 2018-05-07 2019-12-04 Rolls-Royce Corporation Laufschaufel eines gasturbinentriebwerks und zugehöriges herstellungsverfahren
US11655717B2 (en) 2018-05-07 2023-05-23 Rolls-Royce Corporation Turbine blade squealer tip including internal squealer tip cooling channel
EP3623577A1 (de) * 2018-09-12 2020-03-18 United Technologies Corporation Turbinenschaufelspitzenkühlanordnung mit spülteilung
US10801334B2 (en) 2018-09-12 2020-10-13 Raytheon Technologies Corporation Cooling arrangement with purge partition
JP7370823B2 (ja) 2018-12-03 2023-10-30 ゼネラル・エレクトリック・カンパニイ 先端レール冷却インサートを含むタービンブレード先端冷却システム
EP4056806A3 (de) * 2021-03-09 2022-11-16 Mechanical Dynamics & Analysis LLC Turbinenlaufschaufel mit versorgungsplenum für schaufelspitzenkühlöffnungen
US11840940B2 (en) 2021-03-09 2023-12-12 Mechanical Dynamics And Analysis Llc Turbine blade tip cooling hole supply plenum

Also Published As

Publication number Publication date
GB201223193D0 (en) 2013-02-06
US20140178207A1 (en) 2014-06-26

Similar Documents

Publication Publication Date Title
EP2746536A1 (de) Rotorstufe einer Turbine
EP2725195B1 (de) Turbinenlaufschaufel und zugehörige Rotorstufe
EP2489838B1 (de) Gekühlte Komponente für eine Turbine eines Gasturbinenmotors
US8840371B2 (en) Methods and systems for use in regulating a temperature of components
EP2825748B1 (de) Kühlkanal für ein gasturbinentriebwerk und gasturbinentriebwerk
US10247012B2 (en) Aerofoil blade or vane
US9382811B2 (en) Aerofoil cooling arrangement
US7850428B2 (en) Aerofoils
EP2235328B1 (de) Schaufelkühlung
US9683449B2 (en) Stator vane row
EP1424467A2 (de) Turbinenschaufelreihe bestehend aus Schaufeln mit langer und kurzer Sehnenlänge
EP2568119B1 (de) Schaufel einer Gasturbine mit verbesserter Austrittskanten-Kühlanordnung
US8297925B2 (en) Aerofoil configuration
US9512782B2 (en) Gas turbine engine end-wall component
US20130028735A1 (en) Blade cooling and sealing system
US8419366B2 (en) Blade
US9376918B2 (en) Aerofoil cooling arrangement
EP2604795B1 (de) Laufschaufel oder -leitschaufel

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 20131209

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

AX Request for extension of the european patent

Extension state: BA ME

R17P Request for examination filed (corrected)

Effective date: 20141219

RBV Designated contracting states (corrected)

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: ROLLS-ROYCE PLC

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: EXAMINATION IS IN PROGRESS

17Q First examination report despatched

Effective date: 20191004

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: ROLLS-ROYCE PLC

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION IS DEEMED TO BE WITHDRAWN

18D Application deemed to be withdrawn

Effective date: 20200215