EP0916808A2 - Turbine - Google Patents

Turbine Download PDF

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Publication number
EP0916808A2
EP0916808A2 EP98308747A EP98308747A EP0916808A2 EP 0916808 A2 EP0916808 A2 EP 0916808A2 EP 98308747 A EP98308747 A EP 98308747A EP 98308747 A EP98308747 A EP 98308747A EP 0916808 A2 EP0916808 A2 EP 0916808A2
Authority
EP
European Patent Office
Prior art keywords
turbine
aerofoil
disc
cooling air
blades
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.)
Granted
Application number
EP98308747A
Other languages
English (en)
French (fr)
Other versions
EP0916808B1 (de
EP0916808A3 (de
Inventor
Alec George Dodd
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 EP0916808A2 publication Critical patent/EP0916808A2/de
Publication of EP0916808A3 publication Critical patent/EP0916808A3/de
Application granted granted Critical
Publication of EP0916808B1 publication Critical patent/EP0916808B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/082—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades on the side of the rotor disc
    • 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/085—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
    • 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
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms

Definitions

  • This invention relates to a turbine and is particularly concerned with minimising the effects of cooling air leakage in a turbines which is air cooled.
  • cooling air from the compression section of the gas turbine engine flows along the radially inner regions of the engine before being deflected in radially outward directions between the disc and structure adjacent thereto. The air is then directed into cooling passages provided within turbine blades carried by one of the discs.
  • annular gas seal is positioned between the disc and the structure adjacent thereto.
  • the seal is of the labyrinth type comprising annular, axially extending parts provided on both the disc and the adjacent structure which cooperate to define a barrier in the form of a tortuous path for air attempting to flow in a radially outward direction. While such seals are partially effective in providing a barrier to air flowing in radially outward directions, there remains a certain degree of undesirable leakage of cooling air into the hot gas stream.
  • a turbine comprises at least one rotatable disc carrying an annular array of aerofoil blades, each of said blades having an aerofoil portion operationally located in an annular gas passage extending through said turbine for the flow of gas through said turbine, means being provided to direct cooling air into passages provided internally of said aerofoil blades to provide cooling thereof, said cooling air operationally flowing, at least partially, in radially outward directions over at least part of the upstream external surface of said disc prior to a part thereof being diverted to provide cooling of said aerofoil blades, means being provided radially inwardly of said aerofoil portions to direct at least some of the remaining cooling air into a region downstream of said disc in a direction having a circumferential component generally opposite to that in which said disc operationally rotates.
  • Said means to direct at least some of said remaining cooling air into said region downstream of said disc preferably comprises a plurality of passages, each interconnecting said region downstream of said disc with the region upstream of said disc.
  • Each of said blades is preferably provided with a radially inner platform to define a part of said annular gas passage, in which case one of said passages may be provided within each of said platforms, each passage being so disposed as to direct cooling air exhausted therefrom in said direction having a circumferential component.
  • a plurality of lock plates may be provided on the downstream side of said disc to provide locking of said blades on said disc, each of said lock plates having an aperture therein which is in communication with one of said passages, deflection means being provided on each of said lockplates and associated with said aperture in said lockplate to deflect cooling air from said passage associated therewith in said direction having a circumferential component.
  • Each of said deflector means may be in the form of a cowling attached to its associated lockplate.
  • Each of said blades may be provided with a shank radially inwardly of its aerofoil portion, the shanks of adjacent aerofoil blades being so configured that they cooperate to define said passages.
  • Figure 1 is a partially broken away perspective view of part of turbine in accordance with the present invention.
  • Figure 2 is a view on arrow A of Figure 1.
  • Figure 3 is a view on section line B-B of Figure 2.
  • Figure 4 is a view similar to that shown in Figure 2 of an alternative embodiment of the present invention.
  • Figure 5 is a perspective view of a portion of the embodiment shown in Figure 4.
  • a turbine 10 for a gas turbine engine (not shown) is shown in a partial, broken away view. It is of generally conventional configuration comprising an annular array of stator vanes 11 which are located upstream of an annular array of aerofoil rotor blades 12.
  • the turbine 10 is provided with several more axially alternate annular arrays of stator vanes and aerofoil blades, but these have been omitted in the interests of clarity.
  • the stator vanes 11 each comprise an aerofoil portion 13 which is situated in an annular gas passage 14 which extends through the turbine 10.
  • the radially inner and outer extents of the gas passage 14 in the region of the vane aerofoil portions 13 are respectively defined by inner and outer platforms 15 and 16 which are integral with the aerofoil portion 13.
  • the inner platforms 15 of circumferentially adjacent vanes 11 abut to define a generally continuous gas passage-defining surface as do the outer platforms 16.
  • Each stator vane 11 is respectively supported at its radially inner and outer extents by the turbine casing 17 and an inner support structure 18.
  • the aerofoil rotor blades 12 are mounted on a common disc 19 which is mounted for rotation within the turbine 10.
  • Each aerofoil rotor blade 12 comprises an aerofoil portion 20 which, like the aerofoil portions 13 of the stator vanes 11, is situated in the annular gas passage 14.
  • Radially inner and outer platforms 21 and 22 respectively on each blade 20 serve to define local portions of the gas passage 14.
  • Each aerofoil blade 12 is provided with a shank 23 radially inwardly of its inner platform 21 which interconnects the remainder of the blade 12 with a firtree root portion 24.
  • the firtree portion 24 locates in a correspondingly shaped cut-out portion 25 provided in the periphery of the disc 19, thereby providing radial constraint for the aerofoil blade 12.
  • the shanks 23 are circumferentially narrower than their associated firtree root portions 24 so that a circumferential gap 23a is defined between adjacent shanks 23.
  • each lockplate 40 is planar and locates at its radially outer extent in a radially inwardly directed groove 41 defined by its adjacent aerofoil blade 12 and at its radially inner extent in a radially outwardly directed annular groove 42 defined by the disc 19.
  • the lockplates 40 are well known as such in the construction of turbines.
  • the aerofoil blades 12 are cooled by a flow of cooling air into their interiors which is exhausted through a large number of small holes 28 in their aerofoil portions 20.
  • the cooling air is directed into the aerofoil blade 12 interiors from their radially inner extents.
  • the air flows in a radially outward direction over the upstream surface 29 of the disc 19 to enter a plurality of generally radially extending passages 30 in the disc 19 periphery.
  • One passage 30 is associated with each firtree root cut-out portion 25 so that a flow of cooling air is directed to the root portion 25 of each of the aerofoil blades 12.
  • a passage (not shown) in each root portion 25 directs cooling air into the blade 12 interior to provide convection cooling of the blade 12. It then flows through the small holes 28 to provide film cooling of the aerofoil portion. The cooling air then mixes with the gases flowing through the annular gas passage 14.
  • annular seal 31 is provided between the upstream face 29 of the disc 19 and the downstream face 32 of the fixed turbine structure 34 which supports the radially inner extents of the vanes 11.
  • the seal 31 is of the well known labyrinth type comprising a generally axially extending element 35 carried by the disc 19 and a corresponding reception element 36 carried by the fixed turbine support structure 34.
  • labyrinth seals such as that described above are not as efficient at providing a barrier to gas flow as would normally be desirable. Consequently, some cooling air inevitably leaks through the labyrinth seal 31 into the region 37 between the firtree root portions 24 and fixed turbine support structure 34. Under normal circumstances, this leaked cooling air would pass into the annular gas passage 14 and have a prejudicial effect upon the gases operationally flowing through that passage 14. However, in accordance with the present invention, the leaked cooling air is utilised in a more effective and efficient manner.
  • each of the passages 37 is circumferentially angled so that the air is exhausted from it in a generally circumferential direction. That direction is generally opposite to the operational direction of rotation 39 of the disc 19. As a consequence, the exhausted cooling air assists in the driving of the disc 19.
  • each of the lockplates which in modified form as depicted in Figures 4 and 5, is designated 40a, is provided with an aperture 43.
  • Each aperture 43 is partially enclosed by a cowling 44 which is bonded to its associated lockplate 40a and is of part-oval configuration in plan view.
  • the centre portion 45 of each cowling 44 is raised so as to define an outlet 46 adjacent one edge of its associated lockplate 40a.
  • cooling air from the region 37 flows through the gaps 23a between the blade shanks 23 as described earlier. However, that cooling air then flows through the apertures 43 in the lockplates 40a.
  • Each cowling 44 is so configured that the cooling air flow is deflected in a generally circumferential direction which is opposite to the direction of rotation 39 of the disc 19. Consequently, the deflected airflow serves the same function as the airflow exhausted from the passages 37 in improving overall turbine efficiency.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP98308747A 1997-11-05 1998-10-26 Turbine Expired - Lifetime EP0916808B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB9723268 1997-11-05
GBGB9723268.0A GB9723268D0 (en) 1997-11-05 1997-11-05 Turbine

Publications (3)

Publication Number Publication Date
EP0916808A2 true EP0916808A2 (de) 1999-05-19
EP0916808A3 EP0916808A3 (de) 2000-01-12
EP0916808B1 EP0916808B1 (de) 2003-03-12

Family

ID=10821552

Family Applications (1)

Application Number Title Priority Date Filing Date
EP98308747A Expired - Lifetime EP0916808B1 (de) 1997-11-05 1998-10-26 Turbine

Country Status (3)

Country Link
EP (1) EP0916808B1 (de)
DE (1) DE69812044T2 (de)
GB (1) GB9723268D0 (de)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1703081A1 (de) * 2005-02-23 2006-09-20 Rolls-Royce Plc Seitenplatte
EP1703082A1 (de) * 2005-02-23 2006-09-20 Rolls-Royce Plc Seitenplatte
US7465149B2 (en) 2006-03-14 2008-12-16 Rolls-Royce Plc Turbine engine cooling

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
NL88170C (de) * 1952-10-31 1900-01-01
US3043561A (en) * 1958-12-29 1962-07-10 Gen Electric Turbine rotor ventilation system
US3501249A (en) * 1968-06-24 1970-03-17 Westinghouse Electric Corp Side plates for turbine blades
US4453888A (en) * 1981-04-01 1984-06-12 United Technologies Corporation Nozzle for a coolable rotor blade
EP0626036B1 (de) * 1992-02-10 1996-10-09 United Technologies Corporation Ejektor für kühlfluid

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1703081A1 (de) * 2005-02-23 2006-09-20 Rolls-Royce Plc Seitenplatte
EP1703082A1 (de) * 2005-02-23 2006-09-20 Rolls-Royce Plc Seitenplatte
US7465149B2 (en) 2006-03-14 2008-12-16 Rolls-Royce Plc Turbine engine cooling

Also Published As

Publication number Publication date
DE69812044D1 (de) 2003-04-17
DE69812044T2 (de) 2003-08-21
EP0916808B1 (de) 2003-03-12
EP0916808A3 (de) 2000-01-12
GB9723268D0 (en) 1998-01-07

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