US20130028735A1 - Blade cooling and sealing system - Google Patents

Blade cooling and sealing system Download PDF

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Publication number
US20130028735A1
US20130028735A1 US13/546,678 US201213546678A US2013028735A1 US 20130028735 A1 US20130028735 A1 US 20130028735A1 US 201213546678 A US201213546678 A US 201213546678A US 2013028735 A1 US2013028735 A1 US 2013028735A1
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United States
Prior art keywords
blade
flow
cooling
vane
sealing
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.)
Abandoned
Application number
US13/546,678
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English (en)
Inventor
Alexander J. Burt
Keith C. Sadler
Kevin Gorton
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
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Rolls Royce PLC
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Filing date
Publication date
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Assigned to ROLLS-ROYCE PLC reassignment ROLLS-ROYCE PLC ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BURT, ALEXANDER JOHN, Gorton, Kevin, SADLER, KEITH CHRISTOPHER
Publication of US20130028735A1 publication Critical patent/US20130028735A1/en
Abandoned legal-status Critical Current

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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
    • 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
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
    • 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
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F05D2250/70—Shape
    • F05D2250/75—Shape given by its similarity to a letter, e.g. T-shaped
    • 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
    • F05D2260/00—Function
    • F05D2260/60—Fluid transfer

Definitions

  • the present invention concerns cooling and sealing arrangements in a gas turbine engine.
  • the present invention concerns a method and apparatus for using flow to cool components in a gas turbine engine and seal between components of a gas turbine engine.
  • the performance of a gas turbine engine may be improved by increasing the turbine gas temperature. It is therefore generally desirable to operate the turbine at the highest possible temperature. For a given gas turbine configuration (for example in terms of engine cycle compression ratio or bypass ratio), increasing the turbine entry gas temperature will produce more specific thrust (eg engine thrust per unit of air mass flow). However as turbine entry temperatures increases, the life of an uncooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
  • internal cooling passages are generally formed within the aerofoils. These internal cooling passages allow cooling air to be passed through the blades (or vanes) to remove heat through convection.
  • the cooling air is typically taken from flow through a cooler part of the engine. For example, cooling air may be bled from the compressor, prior to combustion, for example from the HP compressor. Typical cooling air temperatures (for example at maximum take off condition) are between 700 and 900 K. Gas temperatures in the turbine can be in excess of 2100 K.
  • cooling flow may enter the aerofoil via a fixture portion of the blade (at the radially inner end of the blade), pass through the blade in a substantially radial direction towards the the tip, and then exit through the tip (at the radially outer end of the blade).
  • Such cooling flow 100 is shown in the FIG. 2 arrangement, and discussed in greater detail below.
  • sealing flow is required to seal between rotor and stator rims.
  • sealing air is provided from elsewhere in the engine.
  • the sealing air may also be bled from the compressor.
  • Such sealing flow 122 , 124 is also shown in the FIG. 2 arrangement, and discussed in greater detail below.
  • a blade or vane for a gas turbine engine.
  • the blade or vane comprises a cooling fluid inlet configured to allow cooling fluid into the blade or vane.
  • the blade or vane comprises a cooling fluid outlet configured to allow cooling fluid out of the blade or vane.
  • the blade or vane comprises an aerofoil that, in use, is gas-washed by a working fluid of the gas turbine engine.
  • the blade or vane comprises a platform from which the aerofoil extends (in a direction that may be said to be radially outward).
  • the blade or vane comprises an internal cooling flow passage between the cooling fluid inlet and the cooling fluid outlet configured to channel cooling fluid through the interior of the blade or vane (for example at least through the aerofoil portion of the blade or vane).
  • the blade or vane comprises an internal bleed flow passage in fluid communication with the internal cooling flow passage at a bleeding position, and configured to bleed a portion of the cooling fluid from the internal cooling flow passage for use as a sealing flow.
  • the blade or vane comprises a sealing flow outlet through which fluid from the internal bleed flow passage exits the blade or vane, the sealing flow outlet being positioned to allow the sealing flow to be used in a seal during operation.
  • the internal cooling flow passage extends through at least a part of the aerofoil before the sealing flow is bled off at the bleeding position (for example during operation of the gas turbine engine).
  • the aerofoil may include aerodynamic surfaces that, in operation, are gas-washed by the working fluid, such as a suction surface, a pressure surface, and a platform from which the suction surface and pressure surface extend.
  • the blade or vane may comprise sealing features which may combine or interact with (for example interlock with, abut, or nearly abut) neighbouring surfaces (for example of neighbouring blades or vanes) of a gas turbine engine to form a seal when assembled.
  • the blade or vane may optionally include a shroud at its tip.
  • the shroud may include sealing features to prevent or reduce over-tip leakage flow, for example in blade embodiments.
  • the shroud may extend around the entire outer circumference of the blades.
  • the blade or vane may have a partial shroud, or winglet, at its tip (for example extending around a portion of the segment between the blades).
  • the blade or vane may have no shroud at its tip.
  • the cooling flow may pass into a least a part of the inside of the aerofoil portion of the blade or vane before being bled off at the bleeding position.
  • This may mean that at least a part of the gas washed surface which may be exposed to the higher temperatures in the turbine may be cooled before the flow is bled off to be used as sealing flow.
  • the bleeding position may still be within a different part of the blade or vane (i.e. the bleeding position need not be within the aerofoil portion of the blade or vane, although in some embodiments it may be).
  • the total amount (i.e. mass flow rate) of flow (for example air) that is bled, for example from the compressor, to achieve a given level of blade cooling may thus be reduced (for example minimized), for example compared with conventional arrangements. Additionally or alternatively, the level of blade cooling may be increased for a given amount of air that is bled from the compressor.
  • the internal cooling flow passage may be a multipass cooling flow passage. Each pass may be arranged to carry cooling fluid in either a substantially radially outward direction or a substantially radially inward direction.
  • the bleeding position may be after at least one pass through the internal cooling passage from the cooling fluid inlet. For example, the bleeding position may be after two passes through the internal cooling passage, e.g. one pass in the radially outward direction and one pass in the radially inward direction.
  • Arranging the aerofoil in this manner ensures that the sealing flow that exits through the sealing flow outlet has already passed through a part of the internal cooling flow passage that extends along the chord of the aerofoil, and thus removed heat from along the chord of the aerofoil, at least, before it is used as a sealing flow. This may be advantageous in removing heat evenly from the aerofoil, for example to avoid localised hot-spots being observed during operation.
  • the internal cooling flow passage may comprise an s-shape. Such a shape may be referred to as a serpentine shape.
  • the blade or vane may comprise a fixture for attaching the blade or vane to a gas turbine engine.
  • the fixture may comprise an attachment portion, which may be a fir tree for attaching a blade to a disc, for example.
  • the fixture may comprise a shank.
  • the shank may be between the fir tree and the aerofoil.
  • the fixture may comprise the platform.
  • the aerofoil may extend from the platform. The surface of the platform from which the aerofoil extends may be exposed to the working fluid of the gas turbine engine during operation.
  • the cooling fluid inlet may be located in the fixture, for example in the shank of a rotor blade.
  • the cooling fluid inlet is radially inward of the platform. This may be convenient, for example, if the cooling flow is provided from towards the axial centreline of the engine. Additionally or alternatively, it may be convenient if the blade or vane is a rotor blade.
  • the sealing flow outlet may be located in the fixture, for example in the shank of a rotor blade.
  • the sealing flow outlet is radially inward of the platform. This may be convenient if, for example, the sealing flow is to be used in a seal towards the radially inner side of the vane or blade, for example a rim seal.
  • the bleeding position and the internal bleed flow passage may both be provided within the fixture. This may allow all of the features relating to the sealing flow to be located in close proximity, thereby allowing maximum cooling and/or a simple arrangement of features.
  • the blade or vane may have an upstream side and a downstream side defined relative to a flow direction through the gas turbine engine.
  • the upstream side may correspond to the leading edge of the blade or vane and the downstream side may correspond to the trailing edge of the blade.
  • the sealing flow outlet may be located on the downstream side of the blade or vane. Alternatively, the sealing flow outlet may be provided on the upstream side of the blade or vane.
  • the blade or vane may comprise two sealing flow outlets.
  • One of the sealing flow outlets may be located at the downstream side of the aerofoil, and the other may be located at the upstream side of the aerofoil. This may enable sealing flow to be provided to one or both of an upstream seal and a downstream seal.
  • One or both of the seals may form at least a part of a circumferential seal.
  • the blade or vane may comprise effusion cooling flow outlets configured to use cooling flow, for example from the internal cooling flow passage, to supply surface cooling to the aerofoil. This may be an efficient way to cool the blade or vane, for example by forming a cooler air film over the surface of the blade or vane to minimize its interaction with the hot working gas.
  • the cooling flow outlet itself may be an effusion cooling flow outlet.
  • rotor-stator stage for a gas turbine, wherein: the rotor stage comprises at least one blade according to any one of the preceding claims and/or the stator stage comprises at least one vane according to any one of the preceding claims; and the sealing flow exiting from the sealing flow outlet is used as a seal between the rotor stage and the stator stage.
  • the rotor stage may have blades as described herein that provide a sealing flow for a seal between the rotor stage and a neighbouring stator stage, for example the neighbouring downstream stator stage.
  • the sealing flow may thus provide a gas (for example air) seal between a rotatable rotor stage and a stationary stator stage.
  • the sealing flow may provide a rim seal.
  • a method of using a flow in a gas turbine engine to cool a blade or a vane in a stage of the gas turbine engine and to seal between a rotor stage and a stator stage of the gas turbine engine comprises an aerofoil that, in use, is gas-washed by a working fluid of the gas turbine engine.
  • the method comprises using at least a part of the flow both to cool the aerofoil and to seal between the stages.
  • Such a method may provide any one or more of the advantages described herein.
  • the method may comprise cooling the aerofoil by passing a cooling flow into the blade or vane, through an internal cooling passage inside the aerofoil, and out of the blade or vane through a cooling fluid outlet.
  • the method may comprise bleeding at least a part of the cooling flow passing through the internal cooling passage through a sealing flow outlet in the blade or vane before it reaches the cooling fluid outlet.
  • the method may comprise using the flow passing out of the blade through the sealing flow outlet in a seal between the stages.
  • the flow that is used for cooling and sealing may be bled from a compressor of the gas turbine engine before entry into a combustor.
  • using at least a portion of the flow for both cooling and sealing means that less air needs to be bled from the compressor and/or that more heat can be removed from a given blade or vane for a given amount of air bled from the compressor.
  • FIG. 1 is a sectional side view of a gas turbine engine
  • FIG. 2 is a schematic cross-section through a turbine rotor blade and sealing features, showing internal cooling flow passages through the blade;
  • FIG. 4 is a schematic cross-section through a part of a turbine blade and fixture according to an embodiment, showing an internal bleed flow passage
  • FIG. 5 is schematic cross-section through a part of a turbine blade and fixture according to an embodiment in a different plane to that of FIG. 4 , also showing an internal bleed flow passage;
  • 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 16 , 17 , 18 respectively drive the high and intermediate pressure compressors 14 , 13 and the fan 12 by suitable interconnecting shafts.
  • Gas temperatures in the turbine can be in excess of 2100 K. This may be higher than the melting point of the materials from which the turbine components are manufactured. Furthermore, as mentioned herein, it is desirable to operate the turbine at the highest possible temperature because generally, for a given gas turbine configuration, increasing the turbine entry gas temperature will produce more specific thrust.
  • the cooling air 112 , 114 , 116 passes through the passages 32 , 34 , 36 in the aerofoil portion 44 of the blade 40 , and then out 118 through the radially outer tip 46 of the blade.
  • a turbine blade 60 according to an embodiment of the invention is shown in FIG. 3 .
  • the turbine blade 60 may be any type of turbine blade.
  • the turbine blade 60 may be part of a high pressure turbine 16 , an intermediate pressure turbine 17 , or a low pressure turbine 18 .
  • the turbine blade 60 may be part of any type of gas turbine engine, for example a ducted fan gas turbine (turbofan) engine 10 such as that shown in FIG. 1 , a turbojet, a turboprop, a turboshaft, an open rotor engine, or any other gas turbine engine, for example axial flow or radial flow.
  • a ducted fan gas turbine (turbofan) engine 10 such as that shown in FIG. 1
  • a turbojet such as that shown in FIG. 1
  • turboprop such as that shown in FIG. 1
  • turboshaft such as that shown in FIG. 1
  • open rotor engine such as that shown in FIG. 1
  • any other gas turbine engine for example axial flow or radial flow.
  • the turbine blade 60 has a fixture 62 , a platform 64 , an aerofoil (or aerofoil portion) 66 , and a tip 68 .
  • An internal cooling flow passage 70 passes through the inside (interior) of the blade 60 .
  • Cooling fluid (for example cooling air) 200 passes through the internal cooling flow passage 70 , thereby cooling the turbine blade 60 .
  • the fixture 62 may allow the blade 60 to be attached to a corresponding component of a gas turbine engine, for example to a turbine disc (not shown).
  • the term fixture 62 may be used to refer to parts of the blade 60 that are radially inward of the platform 64 .
  • the fixture 62 may not be gas washed by the working fluid passing through the turbine 16 , 17 , 18 .
  • the fixture 62 (part or all of which may be referred to as a root) has an attachment portion (which may be a fir tree) 61 (partially shown) and a shank 63 .
  • the attachment portion 61 may comprise the part of the fixture 62 that allows the blade 60 to be attached to the corresponding component, such as a turbine disc.
  • the internal cooling flow passage 70 has a cooling fluid inlet 71 to the blade 60 .
  • the inlet 71 is provided in the shank 63 of the attachment portion 62 . This allows the cooling air to enter at a radially inner (or inboard) side (or portion) of the turbine blade 60 .
  • the inlet 71 to the internal cooling flow passage 70 may be in a different position in the blade 60 .
  • the inlet 71 may be in a radially outer (or tip) portion of the vane.
  • the cooling fluid 201 entering the cooling fluid inlet 71 may be provided from any suitable source of relatively cool fluid, such as from the compressor 13 , 14 prior to combustion.
  • any suitable source of relatively cool fluid such as from the compressor 13 , 14 prior to combustion.
  • the cooling fluid inlet(s) 71 may be provided from any suitable source of relatively cool fluid, such as from the compressor 13 , 14 prior to combustion.
  • the cooling fluid inlet(s) 71 may be provided from any suitable source of relatively cool fluid, such as from the compressor 13 , 14 prior to combustion.
  • the cooling fluid inlet(s) 71 may be provided from any suitable source of relatively cool fluid, such as from the compressor 13 , 14 prior to combustion.
  • 0.2% to 10% for example 0.5% to 5%, for example 0.75% to 2%
  • 1% to 1.5% of flow through the compressor may be bled off to enter the cooling fluid inlet(s) 71 for each stage that is cooled.
  • the amount of compressor air that is bled may be dependent on the stage that is cooled,
  • an intermediate pressure turbine stage may require in the range of from 0.2% to 2% of compressor air to be bled to it, and a high pressure turbine stage may require in the range of from 1% to 10% of compressor air to be bled from it.
  • other amounts of compressor air i.e. outside of these ranges may be required for a given stage.
  • the cooling flow 201 passes radially outwardly 202 through a radial part of the internal cooling flow passage 70 in the shank 63 .
  • the cooling passage 70 passes through the inside of the blade 60 , inside and through the platform 64 , and into the aerofoil 66 . This allows the surfaces of the aerofoil 66 (for example the pressure surface and the suction surface) that are exposed, in use, to the hot working gas, to be cooled by the cooling air.
  • the internal cooling flow passage 70 forms a serpentine, or ‘s’ shape 76 at least partially within the aerofoil portion 66 in the FIG. 3 embodiment.
  • the serpentine internal passage 76 comprises a first radial passage 73 through which the cooling air flows in a radially outward direction.
  • the first radial passage 73 is joined, via a bend or corner in the internal passage 70 at or towards the tip (or radially outer portion) 68 , to a second radial passage 74 through which the cooling air flows in a radially inward direction.
  • the second radial passage 74 is joined, via a bend or corner in the internal passage 70 at or towards the fixture (or radially inner portion) 62 , to a third radial passage 75 through which the cooling air flows in a radially outward direction.
  • the three radial internal passages 73 , 74 , 75 together with the bends that join them together, form the serpentine internal flow passage 76 .
  • the cooling air 200 flows through the internal cooling flow passage 70 (which includes the serpentine passage 76 ) in the direction of the arrows through the passage 70 shown in FIG. 3 .
  • the internal flow passage may take a different shape. Any suitable shape could be used.
  • the internal cooling flow passage could have 1, 2, 3 (as in the case of the FIG. 3 embodiment), 4, 5, 6, 7, 8, 9, 10 or more than 10 radial flow passages, each of which may be joined together by a bend to form a continuous passage 70 .
  • the internal flow passage 70 of some embodiments may comprise features other than radial flow passages. Some embodiments may have more than one internal flow passage, any number of which may have an associated internal bleed flow passage. Some embodiments may, for example, have more than one serpentine passage. Each serpentine passage may have its own inlet(s) and outlet(s).
  • the serpentine passage may have an axial component (in relation to the engine axis X-X) that is configured to direct flow in a generally upstream (as in the example shown in FIG. 3 ), or generally downstream, direction.
  • the direction of the axial component of the flow may be different for different passages, i.e. one passage may result in an upstream axial flow component, and another passage may result in a downstream axial flow component.
  • the or each serpentine passage could flow either forwards or rearwards within the blade.
  • the bleeding position 81 is after two passes (which may be referred to as radial passes) through the blade 60 .
  • the bleeding position 81 may be at a different position along the internal cooling flow passage 70 .
  • the internal cooling flow passage 70 may take a substantially different shape to that shown in FIG. 3 . It is desirable for the internal flow passage 70 to have passed into (and optionally back out of) the aerofoil portion 66 of the blade 60 before the bleeding position 81 . This may allow heat to be removed particularly effectively from the gas washed surfaces of the blade 60 for example from the aerofoil portion 66 .
  • the proportion of the flow 210 that enters the internal bleed flow passage 80 may be determined by, for example, the relative size (for example cross-sectional area, or effective flow area) of the internal bleed flow passage 80 and the cooling flow passage 75 downstream of the bleeding position 81 . Additionally or alternatively, the proportion of the flow 210 that enters the internal bleed flow passage 80 may be determined by the connection geometry between the internal cooling flow passage 70 and the internal bleed flow passage 80 at the bleeding position 81 . For example, the angle between the internal cooling flow passage 70 and the internal bleed flow passage 80 may have an impact, as may the internal wall shape.
  • the proportion 210 of the flow 200 in the internal cooling flowing flow passage 74 upstream of the bleeding position 81 that enters the bleed flow passage 80 may be in the range of from 5% to 75%, for example on the order of 30%.
  • a blade 60 that is in the range of from 2% to 25%, for example 5% to 15%, for example 8% to 12%, for example around 10% lighter than a blade of equivalent external geometry (for example having the same external aerofoil geometry) with the separate cooling and sealing flows such as that shown in FIG. 2 .
  • This may result in further advantages, such as weight reductions in other components, such as the turbine rotor disc.
  • the angle between the axial direction and the bled sealing air 210 may be in the range of from 0 degrees to 45 degrees, for example 1 degree to 35 degrees, for example 2 degrees to 25 degrees, for example 3 degrees to 15 degrees, for example 4 degrees to 10 degrees, for example around 5 degrees (positive angles being defined as being towards the radially outer direction).
  • any suitable exit angle for the bled sealing air 210 may be chosen depending on, for example, the geometry of the seal.
  • the direction of the bled sealing air 210 that exits through the flow outlet 82 may additionally or alternatively have a circumferential component.
  • the internal cooling flow passage 70 and the internal bleed flow passage 80 may be arranged so as not to interact with each other, for example at any position other then the bleed flow position.
  • the dashed line along the internal bleed flow passage 80 represents the internal bleed flow passage 80 passing to the side of the internal cooling flow passage 70 so as to avoid the two passages interacting.
  • the internal bleed flow passage 80 is shown as passing to the side of the internal cooling flow passage 70 within the shank 63 . This may be convenient because the shank may have a sufficient thickness for the two passages to pass side-by-side, for example in the local circumferential plane (that is, a plane perpendicular to the local radial direction).
  • the two passages 70 , 80 may pass each other side-by-side in other places depending on the desired arrangement, for example of the bleed position 81 and/or the sealing flow outlet 82 .
  • the dogleg 85 has a component in the local circumferential direction to allow the internal bleed flow passage 80 to pass to the side of the internal cooling flow passage 70 when viewed along a radial direction (which direction may be referred to as a longitudinal axis of the blade 60 itself).
  • the dogleg 85 is arranged such that the internal bleed flow passage is taken towards the pressure surface of the blade 60 .
  • Other embodiments may have alternative arrangements which allow the two internal flow passages 70 , 80 to pass each other, which may or may not comprise dogleg structures. Some embodiments may not comprise a dogleg structure.

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  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/546,678 2011-07-27 2012-07-11 Blade cooling and sealing system Abandoned US20130028735A1 (en)

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GB201112880A GB201112880D0 (en) 2011-07-27 2011-07-27 Blade cooling and sealing system

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WO2016135779A1 (fr) * 2015-02-26 2016-09-01 株式会社 東芝 Pale de rotor de turbine et turbine
US9638045B2 (en) 2014-05-28 2017-05-02 General Electric Company Cooling structure for stationary blade
US9771816B2 (en) 2014-05-07 2017-09-26 General Electric Company Blade cooling circuit feed duct, exhaust duct, and related cooling structure
US9822653B2 (en) 2015-07-16 2017-11-21 General Electric Company Cooling structure for stationary blade
US9909436B2 (en) 2015-07-16 2018-03-06 General Electric Company Cooling structure for stationary blade
US20190153885A1 (en) * 2014-11-12 2019-05-23 United Technologies Corporation Platforms with leading edge features
CN113623072A (zh) * 2021-08-23 2021-11-09 中国科学院工程热物理研究所 一种用于高压比轴流压气机的后面级盘缘冷却结构

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US9638045B2 (en) 2014-05-28 2017-05-02 General Electric Company Cooling structure for stationary blade
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JPWO2016135779A1 (ja) * 2015-02-26 2017-10-05 株式会社東芝 タービン動翼及びタービン
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