GB2413598A - Providing cooling gas to turbine blade and disc in gas turbine engine - Google Patents

Providing cooling gas to turbine blade and disc in gas turbine engine Download PDF

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Publication number
GB2413598A
GB2413598A GB0409847A GB0409847A GB2413598A GB 2413598 A GB2413598 A GB 2413598A GB 0409847 A GB0409847 A GB 0409847A GB 0409847 A GB0409847 A GB 0409847A GB 2413598 A GB2413598 A GB 2413598A
Authority
GB
United Kingdom
Prior art keywords
gas
turbine
cavity
nozzle
disc
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Withdrawn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
GB0409847A
Other versions
GB0409847D0 (en
Inventor
Mark Ashley Halliwell
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
Priority to GB0409847A priority Critical patent/GB2413598A/en
Publication of GB0409847D0 publication Critical patent/GB0409847D0/en
Publication of GB2413598A publication Critical patent/GB2413598A/en
Withdrawn legal-status Critical Current

Links

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
    • F01D5/085—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
    • 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
    • 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/126—Baffles or ribs
    • 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
    • F05D2250/30—Arrangement of components
    • F05D2250/31—Arrangement of components according to the direction of their main axis or their axis of rotation
    • F05D2250/314—Arrangement of components according to the direction of their main axis or their axis of rotation the axes being inclined in relation to each other

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

In a gas turbine engine, a turbine disc 16, has an axis of rotation, comprises a cavity 18 that is supplied with gas 13 (such as air from a compressor or the engine) having a radial velocity component relative to the axis, and provides for movement of the gas from the cavity to the interior of a turbine blade 12. The gas may be supplied to the cavity by means of a nozzle 14 comprising a first conduit part 24 parallel to the axis, and a second conduit part 26 inclined thereto. A projection of the cross-sectional area of the second conduit part may overlap with the cross-sectional area of a further conduit 36, extending through the disc 16 at least partly in a radial direction from the cavity, for conveying gas to a surface 38 receiving the turbine blade 12. To control movement of air within the cavity a member 44 may be provided extending either from the disc 16 or the nozzle 14, and seals 30, 32 may provided between the disc and the nozzle.

Description

AN ARRANGEMENT FOR PROVIDING GAS TO A TURBINE BLADE IN A GAS
TURBINE ENGINE
Embodiments of the present invention relate to an arrangement for providing gas to a turbine blade in a gas turbine engine.
Turbine blades for gas turbine engines usually require some form of cooling when the engine is operating because the operating temperature of the turbine is usually greater than the melting point temperature of the turbine blade.
Currently, gas is supplied to the interior of a turbine blade to reduce its temperature. The gas exits a turbine blade via a plurality of apertures and is then exhausted from the gas turbine engine.
Turbine blades extend from, and are held in place by turbine discs as illustrated in Fig.1. Currently, gas supplied to cool a turbine blade is supplied via at least one nozzle 114, to a cavity 18 within a turbine disc 16.
The turbine disc 16 is arranged to allow the passage of gas from the cavity 18, to the interior of a turbine blade 12.
The turbine disc 16 is usually arranged so that it comprises at least one conduit 36 that extends from the cavity 18 to a surface for receiving a turbine blade 38.
The gas is provided axially from a conduit 27 of the nozzle 114, into the cavity 18 and circulates, at least partially, around the cavity 18 of the turbine disc 16 before entering the conduit 36 of the turbine disc 16. This is known as 'windage' in the art and is illustrated generally in Fig. 1 by arrows 29.
Thermal energy is transferred from the turbine disc 16 to the gas. Consequently, due to the cooling effect of the gas on the turbine disc 16, a temperature differential is created across the turbine disc that distorts the turbine disc. The temperature differential and hence distortion is greatest when the engine is initially started or when the engine is operating at its maximum operating condition (Turbine Gas Temperature - T.G.T). This distortion may move the turbine blade 12 toward a casing 40 of the gas turbine engine and may result in contact between the turbine blade 12 and the casing 40. The turbine blade 12 may also comprise a blade shroud 42 which may also be damaged or removed due to the contact between the casing 40 and the blade shroud 42. Contact of a rotating component with the casing 40 is undesirable in an operating engine.
Consequently, a space 50 is left between the turbine blade 12 (or blade shroud 42) and the casing 40 to allow for the movement of the turbine blade 12.
When the gas turbine engine is operating under normal conditions, this space may be large and result in gas leakage around the turbine blades. This reduces the efficiency of the gas turbine engine.
Therefore, it is desirable to provide an alternative arrangement for providing gas to a turbine blade in a gas turbine engine.
According to one aspect of the present invention there is provided an arrangement for providing gas to a turbine blade in a gas turbine engine, the gas turbine engine having an axis, the arrangement comprising: a turbine disc, for rotating about the axis, defining a cavity, the turbine disc being arranged to allow the movement of gas from the cavity to an interior of a turbine blade; and a nozzle for providing gas to the cavity of the turbine disc, wherein the nozzle is arranged so that, in use, the gas provided to the cavity comprises a radial velocity component relative to the axis.
The turbine disc may comprise at least one conduit, arranged to allow the movement of gas from the cavity to an interior of a turbine blade, wherein at least one conduit of the turbine disc may be orientated having a radial component, relative to the axis. The nozzle may comprise a conduit, at least a portion of which may be orientated towards an opening of the at least one conduit of the turbine disc. The nozzle may comprise a conduit that has a portion orientated at an angle to the axis, for providing gas to the cavity.
The nozzle and at least one conduit may each have a cross sectional area and may be aligned such that at least a portion of the projection of the cross sectional area of the nozzle overlaps with the cross sectional area of one of the conduits of the turbine disc.
One benefit provided by embodiments of the present invention is that there may be a reduced transfer of thermal energy between the gas and the turbine disc. A reduced temperature differential may be created across the turbine disc. Consequently, the distortion of a turbine disc may be reduced. As a result, the spacing between a casing of the gas turbine engine and the turbine blade may be reduced.
This may reduce the leakage of air around a turbine blade when the gas turbine engine is in operation. This may improve the efficiency of a gas turbine engine.
Additionally, a further benefit is that there may be less 'windage'. One benefit provided by this is that there may be a reduced transfer of thermal energy from the turbine disc to the gas. Therefore, cooler gas is provided to the turbine blade. This may reduce the extrusion of the turbine blade due to centrifugal forces, when the gas turbine engine is in operation. Once again, this may result in the spacing between the turbine blade and a casing of the gas turbine engine to be reduced. This may reduce the leakage of gas around a turbine blade when the gas turbine engine is operating. This may improve the efficiency of the gas turbine engine and result in reduced operating costs.
Additionally, a reduced quantity of gas may be required to provide the same cooling as a prior art arrangement.
Therefore, less energy may be consumed in the transportation of the gas to the turbine blade.
The gas may be provided from a compressor of the gas turbine engine.
The arrangement may further comprise a member for controlling the flow of gas in the cavity. The member may extend from the turbine disc and protrude into the cavity.
Alternatively, the member may extend from the nozzle and protrude into the cavity. The member may be a flow dissuader.
The flow dissuader effectively reduces the volume of the cavity in which the air may move and consequently, a smaller surface of the turbine disc is presented to the air.
The flow dissuader may help to reduce the transfer of thermal energy from the turbine disc to the gas. The flow dissuader may also help to reduce 'windage'. The benefits provided in the reduction of the above are discussed in the previous paragraphs.
According to a further aspect of the present invention there is provided a method for providing gas to a turbine blade in a gas turbine engine, the gas turbine engine having a turbine disc mounted for rotation about an axis, and having a cavity that allows the movement of gas to an interior of a turbine blade; the method comprising: providing gas to the cavity with a radial velocity component relative to the axis.
For a better understanding of the present invention reference will now be made by way of example only to the accompanying drawings in which: Fig. 1 illustrates a cross sectional schematic diagram of a prior art arrangement for providing gas to a turbine blade; 20Fig.2 illustrates a cross sectional schematic diagram of one embodiment of an arrangement for providing gas to a turbine blade; Fig. 3 illustrates a cross sectional schematic diagram of another embodiment of an arrangement for providing gas to a turbine blade; Fig. 4 illustrates a cross sectional schematic diagram of a further embodiment of an arrangement for providing gas to a turbine blade.
Figures 2, 3 and 4 illustrate an arrangement 10 for providing gas to a turbine blade 12 in a gas turbine engine, the gas turbine engine having an axis 22, the arrangement 10 comprising: a turbine disc 16, for rotating about the axis 22, defining a cavity 18, the turbine disc 16 being arranged to allow the movement of gas from the cavity 18 to an interior of a turbine blade 12; and a nozzle 14 for providing gas to the cavity 18 of the turbine disc 16, wherein the nozzle 14 is arranged so that, in use, the gas provided to the cavity 18 comprises a radial velocity component relative to the axis 22.
Fig 2 illustrates an arrangement 10 for providing gas to a turbine blade 12 of a gas turbine engine. The flow of gas, which is, in this example, air, is illustrated generally by the arrows 13. The air is provided to a nozzle 14 from a compressor (not illustrated in the Figures). The air is then received by a cavity 18 of a turbine disc 16 from the nozzle 14. The turbine disc 16 is arranged to allow the movement of air from the cavity 18 to an interior of the turbine blade 12. The air is exhausted from the turbine blade 12 via a plurality of apertures 20.
The air is subsequently exhausted from the gas turbine engine (not illustrated in the figures).
In more detail, Fig 2 illustrates a nozzle 14 for receiving air from a compressor of the gas turbine engine.
The gas turbine engine has an axis 22 from which a radial direction 15, axial direction 17 and azimuthal direction 19 can be defined. The axis 22 is located off the bottom of the page, at a symmetrical centre of the turbine disc 16 in all the figs, but is illustrated to provide an indication as to the orientation of the axis 22.
The nozzle 14 is, in this embodiment, a static component and does not rotate about the axis 22 of the gas turbine engine. The nozzle 14 includes a conduit 21 which comprises a first conduit 24 and a second conduit 26. The first conduit 24 is parallel to the axis 22 of the engine and receives air from the compressor. The second conduit 26 receives air from the first conduit 24. The second conduit 26 extends continuously from the first conduit 24 and is orientated so that it has at least a radial component and an axial component. The second conduit 26 defines an opening 28 for providing air to the cavity 18 of the turbine disc 16. The nozzle 14 provides air to the cavity 18 having, at least, an axial and a radial velocity component.
The turbine disc 16, in this embodiment, rotates about the axis 22 in the azimuthal direction 19. The turbine disc 16 is arranged to define an enclosed cavity 18 around the opening 28 of the nozzle 14 and comprises an inner seal 30 and an outer seal 32. The cavity 18 of the turbine disc 16 receives air from the nozzle 14 and is enclosed in order to prevent the movement of air out of the cavity 18 other than to the interior of the turbine blade 12. The cavity 18 is defined by an interior surface 34 of the turbine disc 16, which is, in this embodiment, a curved surface that extends from the inner seal 30 to the outer seal 32.
The turbine disc 16 comprises a conduit 36 that is arranged to allow the movement of air from the cavity 18 to the interior of the turbine blade 12. The conduit 36 extends from the interior surface 34 to a surface 38 for receiving the turbine blade 12 and is orientated so that it comprises at least a radial component. The second conduit 26 is orientated, at least partially, towards the conduit 36 in at least the radial direction 15.
The second conduit 26 (and hence opening 28) and the conduit 36 each have a cross sectional area. As illustrated in Fig. 2, the projection of the cross sectional area of the second conduit 26 (represented by dotted lines 29) into the cavity 18, at least partially overlaps the cross sectional area of the conduit 36. Consequently, in this example, at least a portion of the air is directly provided to the conduit 36.
The turbine blade 12 rotates about the axis 22 in the azimuthal direction 19. Air is provided to the interior of the turbine blade 12 from the conduit 36. Thermal energy is transferred from the turbine blade 12 to the air, thereby cooling the turbine blade. The air is exhausted from the turbine blade via a plurality of apertures 20 and is subsequently exhausted from the gas turbine engine.
Due to the orientation of the second conduit 26, the nozzle 14 provides air having a radial velocity component relative to the axis 22. The orientation of the second conduit 26 and location of the nozzle 14 also provides a nozzle 14 that is located at a closer proximity to the conduit 36 than in prior art arrangements. This arrangement provides a number of benefits. One benefit is that the air is provided more directly to the second conduit 36 which results in reduced windage'. Consequently, less thermal energy is transferred from the turbine disc 16 to the air.
Therefore, a smaller temperature differential is created across the turbine disc 16 which results in less distortion of the turbine disc 16. Consequently, the movement of the turbine blade 12 towards a casing 40 is reduced. Therefore, the spacing between the blade 12 and the casing 40 can be reduced. This may result in reduced leakage of air around the turbine blade 12 and improve the efficiency of the engine.
A further benefit of the reduced windage' is that because the air provided to the turbine blade 12 may have a lower temperature, the turbine blade 12 may be cooled to a greater extent than in prior art arrangements. This may result in a reduced extrusion of the turbine blade 12 while the gas turbine engine is operating. Consequently, the space between the turbine blade 12 and the casing 40 may be reduced. This may result in reduced leakage of air around the turbine blade 12 and improve the efficiency of the engine. Alternatively, if the air supplied to the turbine blade is cooler, the quantity of air drawn from the compressor to cool the turbine blade may be reduced in order
to provide the same cooling effect as in prior art
arrangements. Consequently, less electrical energy may be required to pump air from the compressor to the nozzle 14.
This may also result in a reduced number of apertures in the turbine blade 12.
The turbine blade 12, in this embodiment, comprises a blade shroud 42. The blade shroud 42 extends from the turbine blade 12 towards the casing 40. The blade shroud 42 helps to reduce the leakage of air around the turbine blade 12 when the engine is operating. In prior art arrangements, the blade shroud 42 may be damaged or destroyed due to the distortion of the turbine disc and the subsequent movement of the turbine blade towards the casing 40. Due to the reduced distortion of the turbine disc 16 and the reduced extrusion of the turbine blade 12, the space between the blade shroud 42 and the casing 40 may be reduced. This may result in a reduced leakage of air around the turbine blade 12 and blade shroud 42 and may improve the efficiency of the gas turbine engine.
The nozzle 14 can be inserted into existing turbine
arrangements and replace a prior art nozzle. Fig 1
illustrates a prior art nozzle 114 comprising a conduit 27 that is parallel to the axis 22. The first conduit 24, illustrated in Figs 2, 3 & 4 is substantially similar to the conduit 27 of the prior art and allows the replacement of nozzle 114 with nozzle 14. Advantageously, the insertion of nozzle 14 does not require modification to the turbine arrangement.
Fig 3 illustrates an alternative embodiment of the present invention. The same reference numerals are used as those in fig 1 where they correspond to the same feature. A member 44 extends from the surface 34 of the turbine disc 16 into the cavity 18. The member 44 helps to control the movement of air in the cavity 18 so that there is reduced air flow in the lower part 48 of the cavity 18. The member 44 reduces windage' and therefore also provides the above mentioned benefits. Fig 4 illustrates that a member 46 extends from the nozzle 14 into the cavity 18 and is an alternative arrangement to the arrangement illustrated in Fig 3. The members 44 & 46 are, in this example, flow dissuaders.
Although embodiments of the present invention have been described in the preceding paragraphs with reference to various examples, it should be appreciated that modifications to the examples given can be made without departing from the scope of the invention as claimed. For example, the arrangement may comprise more than one flow dissuader or use a combination of flow dissuaders (44, 46) as illustrated in figs 3 & 4. The turbine disc 16 may comprise more than one conduit 36 for providing air to an interior of a turbine blade 12. There may be more than one nozzle 14 in the cavity 18 for providing gas for cooling a turbine blade. The turbine blade 12 may not comprise a blade shroud 42.
Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and / or shown in the drawings whether or not particular emphasis has been placed thereon. .

Claims (13)

  1. Claims 1. An arrangement for providing gas to a turbine blade in a gas
    turbine engine, the gas turbine engine having an axis, the arrangement comprising: a turbine disc, for rotating about the axis, defining a cavity, the turbine disc being arranged to allow the movement of gas from the cavity to an interior of a turbine blade; and a nozzle for providing gas to the cavity of the turbine disc, wherein the nozzle is arranged so that, in use, the gas provided to the cavity comprises a radial velocity component relative to the axis.
  2. 2. An arrangement as claimed in claim 1, wherein the turbine disc comprises at least one conduit, arranged to allow the movement of gas from the cavity to an interior of a turbine blade, and at least one conduit of the turbine disc is orientated having a radial component, relative to the axis.
  3. 3. An arrangement as claimed in claim 2, wherein the nozzle comprises a conduit, at least a portion of which is orientated towards an opening of the at least one conduit of the turbine disc.
  4. 4. An arrangement as claimed in any one of the preceding claims, wherein the nozzle comprises a conduit that has a portion orientated at an angle to the axis, for providing gas to the cavity.
  5. 5. An arrangement as claimed in any one of claims 2 to 4, wherein the nozzle and at least one conduit of the turbine disc each have a cross sectional area and are aligned such that at least a portion of the projection of the cross sectional area of the nozzle overlaps with the cross sectional area of one of the conduits of the turbine disc.
  6. 6. An arrangement as claimed in any one of the preceding claims, wherein the gas is provided from a compressor of the gas turbine engine.
  7. 7. An arrangement as claimed in any one of the preceding claims, wherein the arrangement further comprises a member for controlling the flow of gas in the cavity.
  8. 8. An arrangement as claimed in claim 7, wherein the member extends from the turbine disc and protrudes into the cavity.
  9. 9. An arrangement as claimed in claim 7, wherein the member extends from the nozzle and protrudes into the cavity.
  10. 10. A nozzle for providing gas to a turbine blade in a gas turbine engine as claimed in any one of the preceding claims.
  11. 11. A method for providing gas to a turbine blade in a gas turbine engine, the gas turbine engine having a turbine disc mounted for rotation about an axis, and having a cavity that allows the movement of gas to an interior of a turbine blade; the method comprising: providing gas to the cavity with a radial velocity component relative to the axis.
  12. 12. An arrangement / nozzle / method substantially as hereinbefore described with reference to and / or as shown in the accompanying drawings.
  13. 13. Any novel subject matter or combination including novel subject matter disclosed, whether or not within the scope of or relating to the same invention as the preceding claims.
GB0409847A 2004-05-01 2004-05-01 Providing cooling gas to turbine blade and disc in gas turbine engine Withdrawn GB2413598A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
GB0409847A GB2413598A (en) 2004-05-01 2004-05-01 Providing cooling gas to turbine blade and disc in gas turbine engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB0409847A GB2413598A (en) 2004-05-01 2004-05-01 Providing cooling gas to turbine blade and disc in gas turbine engine

Publications (2)

Publication Number Publication Date
GB0409847D0 GB0409847D0 (en) 2004-06-09
GB2413598A true GB2413598A (en) 2005-11-02

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2011968A2 (en) * 2007-07-02 2009-01-07 United Technologies Corporation Angled on-board injector
US11506072B2 (en) * 2020-03-03 2022-11-22 Itp Next Generation Turbines S.L. Blade assembly for gas turbine engine

Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB988541A (en) * 1962-03-06 1965-04-07 Ruston & Hornsby Ltd Gas turbine rotor cooling
US3575528A (en) * 1968-10-28 1971-04-20 Gen Motors Corp Turbine rotor cooling
US3883263A (en) * 1972-12-21 1975-05-13 Ausburg Nuremberg Aktiengesell Device for cooling rotor blades with solid profile of motor vehicle gas turbines
GB1472570A (en) * 1973-08-02 1977-05-04 Gen Electric Gas turbines
GB2054046A (en) * 1979-07-12 1981-02-11 Rolls Royce Cooling turbine rotors
US4882902A (en) * 1986-04-30 1989-11-28 General Electric Company Turbine cooling air transferring apparatus
GB2225063A (en) * 1988-10-21 1990-05-23 Mtu Muenchen Gmbh Turbine cooling arrangement
US5575616A (en) * 1994-10-11 1996-11-19 General Electric Company Turbine cooling flow modulation apparatus
EP0921272A2 (en) * 1997-12-03 1999-06-09 Rolls-Royce Plc Turbine rotor disc assembly
EP1260673A2 (en) * 2001-05-21 2002-11-27 General Electric Company Turbine cooling circuit

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB988541A (en) * 1962-03-06 1965-04-07 Ruston & Hornsby Ltd Gas turbine rotor cooling
US3575528A (en) * 1968-10-28 1971-04-20 Gen Motors Corp Turbine rotor cooling
US3883263A (en) * 1972-12-21 1975-05-13 Ausburg Nuremberg Aktiengesell Device for cooling rotor blades with solid profile of motor vehicle gas turbines
GB1472570A (en) * 1973-08-02 1977-05-04 Gen Electric Gas turbines
GB2054046A (en) * 1979-07-12 1981-02-11 Rolls Royce Cooling turbine rotors
US4882902A (en) * 1986-04-30 1989-11-28 General Electric Company Turbine cooling air transferring apparatus
GB2225063A (en) * 1988-10-21 1990-05-23 Mtu Muenchen Gmbh Turbine cooling arrangement
US5575616A (en) * 1994-10-11 1996-11-19 General Electric Company Turbine cooling flow modulation apparatus
EP0921272A2 (en) * 1997-12-03 1999-06-09 Rolls-Royce Plc Turbine rotor disc assembly
EP1260673A2 (en) * 2001-05-21 2002-11-27 General Electric Company Turbine cooling circuit

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2011968A2 (en) * 2007-07-02 2009-01-07 United Technologies Corporation Angled on-board injector
EP2011968A3 (en) * 2007-07-02 2012-08-29 United Technologies Corporation Angled on-board injector
US11506072B2 (en) * 2020-03-03 2022-11-22 Itp Next Generation Turbines S.L. Blade assembly for gas turbine engine

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