WO2010062428A2 - Integrated service tube and impingement baffle for a gas turbine engine - Google Patents

Integrated service tube and impingement baffle for a gas turbine engine Download PDF

Info

Publication number
WO2010062428A2
WO2010062428A2 PCT/US2009/055147 US2009055147W WO2010062428A2 WO 2010062428 A2 WO2010062428 A2 WO 2010062428A2 US 2009055147 W US2009055147 W US 2009055147W WO 2010062428 A2 WO2010062428 A2 WO 2010062428A2
Authority
WO
WIPO (PCT)
Prior art keywords
service tube
tube
service
turbine frame
frame assembly
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.)
Ceased
Application number
PCT/US2009/055147
Other languages
French (fr)
Other versions
WO2010062428A3 (en
Inventor
John Alan Manteiga
Robert John Parks
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Priority to DE112009003550T priority Critical patent/DE112009003550T5/en
Priority to JP2011538608A priority patent/JP5414805B2/en
Priority to CA2744226A priority patent/CA2744226C/en
Priority to GB1107527.2A priority patent/GB2477453B/en
Publication of WO2010062428A2 publication Critical patent/WO2010062428A2/en
Publication of WO2010062428A3 publication Critical patent/WO2010062428A3/en
Anticipated expiration legal-status Critical
Ceased 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
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/14—Casings modified therefor
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/18—Lubricating arrangements
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • This invention relates generally to gas turbine engine turbines and more particularly to structural members of such engines.
  • Gas turbine engines frequently include a stationary turbine frame (also referred to as an inter-turbine frame or turbine center frame) which provides a structural load pathfrom bearings which support the rotating shafts of the engine to an outer casing, which forms a backbone structure of the engine.
  • Turbine frames commonly include an annular, centrally-located hub surrounded by an annular outer ring, which are interconnected by a plurality of radially-extending struts, as well as one or more service tubes which cam- fluids to and from the hub.
  • the turbine frame crosses the combustion gas flowpath of the turbine and is thus exposed to high temperatures in operation.
  • a service tube apparatus for a gas turbine engine includes a service tube assembly including, (a) an elongated, hollow service tube: and (b) a service tube baffle surrounding the service tube which is pierced with a plurality of impingement cooling holes.
  • a turbine frame assembly for a gas turbine engine includes: (a) a turbine frame including: (i) an outer ring; (ii) a hub; and (ii) a plurality of struts extending between the hub and the outer ring; (b) at least one service tube apparatus extending between the hub and the outer ring, comprising a service tube assembly including: (i) an elongated, hollow service tube; and (ii) a service tube baffle surrounding the service tube which is pierced with a plurality of impingement cooling holes.
  • Figure 1 a schematic half-sectional view of a gas turbine engine constructed in accordance with an aspect of the present invention
  • Figures 2A and 2B are an exploded perspective view of a turbine frame assembly of the gas turbine engine of Figure 1;
  • FIGS 3 A, 3 B, and 3C are cross-sectional views of the turbine frame assembly of Figure 2;
  • Figure 4 is a perspective view of the turbine frame assembly in a partially- assembled condition
  • FIG. 5 is a perspective view of a service tube assembly constructed according to an aspect of the present invention.
  • Figure 6 is a side view of a service tube fairing.
  • Figures 1 and 2 depict a portion of a gas turbine engine 10 having, among other structures, a compressor 12, a combustor 14, and a gas generator turbine 16.
  • the engine is a t ⁇ rboshaft engine.
  • the principles described herein are equally applicable to turboprop, turbojet, and turbofan engines, as well as turbine engines used for other vehicles or in stationary applications.
  • the compressor 12 provides compressed air that passes into the combustor 14 where fuel is introduced and burned to generate hot combustion gases.
  • the combustion gases are discharged to the gas generator turbine 16 which comprises alternating rows of stationary vanes or nozzles 18 and rotating blades or buckets 20.
  • the combustion gases are expanded therein and energy is extracted to drive the compressor 12 through an outer shaft 22.
  • a work turbine 24 is disposed downstream of the gas generator turbine 16. It also comprises alternating rows of stationary vanes or nozzles 26 and rotors 28 carrying rotating blades or buckets 30. The work turbine 24 further expands the combustion gases and extracts energy to drive an external load (such as a propeller or gearbox) through an inner shaft 32.
  • an external load such as a propeller or gearbox
  • the inner and outer shafts 32 and 22 are supported for rotation in one or more bearings 34.
  • One or more turbine frames provide structural load paths from the bearings 34 to an outer casing 36, which forms a backbone structure of the engine 10.
  • a turbine frame assembly which comprises a turbine frame 38 that integrates a first stage nozzle cascade 40 of the work turbine 24, is disposed between the gas generator turbine 16 and the work turbine 24.
  • FIGs 2-4 illustrate the construction of the turbine frame assembly in more detail.
  • the turbine frame 38 comprises an annular, centrally-located hub 42 with forward and aft faces 44 and 46, surrounded by an annular outer ring 48 having forward and aft flanges 50 and 52.
  • the hub 42 and the outer ring 48 are interconnected by a plurality of radially -extending struts 54. In the illustrated example there are six equally-spaced struts 54.
  • the turbine frame 38 may be a single integral unit or it may be built up from individual components.
  • Ae struts 54 is hollow and terminates in a bleed air port 56 at its outer end, outboard of the outer ring 48.
  • a plurality of service tube assemblies 58 are mounted in the turbine frame 38, positioned between the struts 54, and extend between the outer ring 48 and the hub 42.
  • Figures 3C and 5 show the service tube assembly in more detail.
  • Each service tube assembly 58 includes a hollow service tube 60.
  • the service tube 60 has a central section 55 disposed between reduced-diameter outer and inner ends 57 and 59.
  • the inner end 59 includes a generally cylindrical male fitting 61 which forms a plug-in connection in cooperation with a female receptacle 63 of a sump 65 located within the turbine frame 38.
  • the service tube 60 may be used to transport air or oil from between the sump 65 and an external conduit (not shown) such as an oil supply- or scavenge line, or sump pressurization or vent line, which is coupled to the outer end 57.
  • the service tube 60 is surrounded by a hollow housing 71 which is an integral component that comprises a service tube baffle 62 pierced with impingement cooling holes 64, a mounting bracket 66, and a manifold 68 with an inlet tube 70.
  • the outer end 73 of the housing 71 is attached to an annular flange 75 at the outer end 57 of the service tube 60, for example by brazing or welding.
  • the inner end 77 of Ae housing 71 is free to move thermally in operation, and has an opening that closely surrounds the central section 55 so as to leave a small gap for cooling air flow, as explained in more detail below.
  • the central section 55 may include an annular collar 79 about its outer periphery to define the gap in cooperation with the housing 71.
  • the service tube assemblies 58 plug into aligned openings in the outer ring 48 and the hub 42, and are secured to the outer ring 48 using bolts passing through the mounting bracket 66.
  • the nozzle cascade 40 comprises a plurality of actively-cooled airfoils. In this particular example there are 48 airfoils in total. This number may be varied to suit a particular application. Some of the airfoils, in this case 12, are axially elongated and are incorporated into fairings (see Figure 4) which protect the struts 54 and service tube assemblies 58 from hot combustion gases. Some of the fairings, in this case 6, are strut fairings 72 which are of a split configuration. The remainder of the fairings are service tube fairings 74 which are a single piece configuration. The remaining airfoils, in this case 36, are arranged into nozzle segments 76 having one or more vanes each.
  • FIG. 6 shows one of the service tube fairings 74 in more detail
  • ft includes an airfoil-shaped hollow vane 120 that is supported between an arcuate outer band 122 and an arcuate inner band 124.
  • the inner and outer bands 124 and 122 are axially elongated and shaped so that they define a portion of the flowpath through the turbine frame 38.
  • a forward hook 126 protrudes axially forward from the outer face of the outer band 122
  • an aft hook 128 protrudes axially forward from the outer face of the outer band 122.
  • the vane 120 is axially elongated and includes spaced-apart sidewalls 132 extending between a leading edge 134 and a trailing edge 136.
  • the sidewalls 132 are shaped so as to form an aerodynamic fairing for the service tube assembly 58.
  • a forward section 138 of the vane 120 is hollow and is impingement cooled, in a manner described in more detail below.
  • An aft section 140 of the vane 120 is also hollow and incorporates walls 142 that define a multiple-pass serpentine flowpath.
  • the service tube fairings 74 are cast from a metal alloy suitable for high- temperature operation, such as a cobalt- or nickel-based "superalloy", and may be cast with a specific crystal structure, such as directionalry-solidified (OS) or single-crystal (SX), in a known manner.
  • a metal alloy suitable for high- temperature operation such as a cobalt- or nickel-based "superalloy”
  • OS directionalry-solidified
  • SX single-crystal
  • An example of one suitable material is a nickel-based alloy commercially known as RENE N4.
  • the strut fairings 72, service tube fairings 74, and nozzle segments 76 are ail supported by forward and ait hangers 164 and 166 which are fastened to the forward and aft flanges SO and 52 of the turbine frame 38, respectively, for example using bolts or other suitable fasteners.
  • the forward nozzle hanger 164 is generally disk-shaped and includes an outer flange 168 and an inner flange 170, interconnected by an aft-extending arm 172 having a generally "V-shaped cross-section.
  • the inner flange 170 defines a mounting rail 174 with a slot 176 which accepts the forward hooks 126of the service tube fairings 74 and similar hooks of the strut fairings 72 and nozzle segments 76.
  • the outer flange 168 has bolt holes therein corresponding to bolt holes in the forward flange 50 of the turbine frame 38.
  • the forward nozzle hanger 164 supports the nozzle cascade 40 radially in a way that allows compliance in the axial direction.
  • the aft nozzle hanger 166 is generally disk-shaped and includes an outer flange 175 and an inner flange 177, interconnected by forward-extending arm 180 having a generally "IT-shaped cross-section.
  • the inner flange 177 defines a mounting rail 182 with a slot 184 which accepts the aft hooks 128 of the service tube fairings 74 and similar hooks of the strut fairings 72 and nozzle segments 76.
  • the outer flange 175 has bolt holes therein corresponding to bolt holes in the aft flange 52 of the turbine frame 38.
  • the aft nozzle hanger 166 supports the nozzle cascade 48 radially while providing restraint in the axial direction.
  • An annular outer balance piston (OPB) seal 188 is attached to the aft face of the hub 42, for example with bolts or other suitable fasteners.
  • the OBP seal 188 has a generally "L"-shaped cross-section with a radial arm 190 and an axial arm 192.
  • a forward sealing lip 194 bears against the hub 42.
  • an aft, radially -oufwardk -extending sealing lip 1% captures an annular, "M"-shaped seal 198 against the nozzle cascade 40.
  • a similar "NT-shaped seal 200 is captured between the forward end of the nozzle cascade 40 and another sealing lip 202 on an stationary engine structure 204.
  • the hub 42 and the OBP seal 188 define an inner manifold 206 which communicates with the interior of the hub 42.
  • the inner bands of the strut fairings 72, service tube fairings 74, and nozzle segments 76 cooperate with the hub 42 of the turbine frame 38, the OBP seal 188, and (he seals 198 and 200 to define an annular inner band cavity 208.
  • One or more cooling holes 210 pass through the radial arm 190 of the OBP seal 188. In operation, these cooling holes 210 pass cooling air from the hub 42 to an annular seal plate 212 mounted on a front face of the downstream rotor 28. The cooling air enters a hole 214 in the seal plate 212 and is then routed to the rotor 28 in a conventional fashion.
  • Theaxialarm l92 c ⁇ the OBP $eal l88 carriesanabradablemate ⁇ al216(such as a metallic hones comb) which mates with a seal tooth 218 of the seal plate 212.
  • cooling of the service tube fairings 74 is as follows. Cooling air bled from a source such as the compressor 12 (see Figure 1) is fed into the inlet tubes 70, as shown by the arrow "A".
  • Air from the outer band cavity 186 which is as combination of purge air and post-impingement flows denoted O and E in Figure 6, enters the serpentine passages in the aft sections of the vanes 120as shown at "G". It is then used therein for convective cooling in a conventional manner and subsequently exhausted through the trailing edge cooling passages.
  • the turbine frame assembly described above has multiple advantages over prior art designs.
  • the engine 10 can run hotter and longer without oil coked sump services.
  • the service tube assemblies 58 are "plug -in" components permitting inspection or cleaning without engine disassembly.
  • integration of the service tube and liner cooling improves packaging by moving the service tubes 60 away from the struts 54. There is a potential for less flowpath blockage and better engine performance than with conventional designs. Furthermore, this frees up (he struts 54 for use in providing cooling air to downstream turbine rotors or other components.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Lubrication Details And Ventilation Of Internal Combustion Engines (AREA)

Abstract

A service tube apparatus for a gas turbine engine includes a service tube assembly having: (a) an elongated, hollow service tube, and (b) a service tube baffle surrounding the service tube which is pierced with a plurality of impingement cooling holes.

Description

INTEGRATED SERVICE TUBE AND IMPINGEMENT BAFFLE FOR A GAS
TURBINE ENGINE
BACKGROUND OF THE INVENTION
[0001] This invention relates generally to gas turbine engine turbines and more particularly to structural members of such engines.
[0002| Gas turbine engines frequently include a stationary turbine frame (also referred to as an inter-turbine frame or turbine center frame) which provides a structural load pathfrom bearings which support the rotating shafts of the engine to an outer casing, which forms a backbone structure of the engine. Turbine frames commonly include an annular, centrally-located hub surrounded by an annular outer ring, which are interconnected by a plurality of radially-extending struts, as well as one or more service tubes which cam- fluids to and from the hub. The turbine frame crosses the combustion gas flowpath of the turbine and is thus exposed to high temperatures in operation.
[0003] From a thermodynamic standpoint it is desirable to increase operating temperatures within gas turbine engines as much as possible to increase both output and efficiency. However, as engine operating temperatures are increased, increased active cooling for turbine frame, turbine nozzle, and turbine blade components becomes necessary.
10004] Conventional service tubes are mounted internal Io the struts of the frame and are inseparable from the frame. High temperature operation tends to cause undesirable oil coking within the service tubes.
BRIEF SUMMARY OF THE INVENTION
[0005] These and other shortcomings of the prior art are addressed by the present invention, which provides a service tube assembly for a gas turbine engine that incorporates active cooling.
[0006] According to one aspect, a service tube apparatus for a gas turbine engine includes a service tube assembly including, (a) an elongated, hollow service tube: and (b) a service tube baffle surrounding the service tube which is pierced with a plurality of impingement cooling holes.
[0007] According to another aspect of the invention, a turbine frame assembly for a gas turbine engine includes: (a) a turbine frame including: (i) an outer ring; (ii) a hub; and (ii) a plurality of struts extending between the hub and the outer ring; (b) at least one service tube apparatus extending between the hub and the outer ring, comprising a service tube assembly including: (i) an elongated, hollow service tube; and (ii) a service tube baffle surrounding the service tube which is pierced with a plurality of impingement cooling holes.
10008] BRIEF DESCRIPTION OF THE DRAWINGS
[0009] The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:
[0010} Figure 1 a schematic half-sectional view of a gas turbine engine constructed in accordance with an aspect of the present invention;
[0011] Figures 2A and 2B are an exploded perspective view of a turbine frame assembly of the gas turbine engine of Figure 1;
[00121 Figures 3 A, 3 B, and 3C are cross-sectional views of the turbine frame assembly of Figure 2;
[0013] Figure 4 is a perspective view of the turbine frame assembly in a partially- assembled condition;
[00141 Figure 5 is a perspective view of a service tube assembly constructed according to an aspect of the present invention; and
[ 0015] Figure 6 is a side view of a service tube fairing.
DETAILED DESCRIPTION OF THE INVENTION [0016 j Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, Figures 1 and 2 depict a portion of a gas turbine engine 10 having, among other structures, a compressor 12, a combustor 14, and a gas generator turbine 16. In the illustrated example, the engine is a tυrboshaft engine. However, the principles described herein are equally applicable to turboprop, turbojet, and turbofan engines, as well as turbine engines used for other vehicles or in stationary applications.
[0017] The compressor 12 provides compressed air that passes into the combustor 14 where fuel is introduced and burned to generate hot combustion gases. The combustion gases are discharged to the gas generator turbine 16 which comprises alternating rows of stationary vanes or nozzles 18 and rotating blades or buckets 20. The combustion gases are expanded therein and energy is extracted to drive the compressor 12 through an outer shaft 22.
10018] A work turbine 24 is disposed downstream of the gas generator turbine 16. It also comprises alternating rows of stationary vanes or nozzles 26 and rotors 28 carrying rotating blades or buckets 30. The work turbine 24 further expands the combustion gases and extracts energy to drive an external load (such as a propeller or gearbox) through an inner shaft 32.
((X)19] The inner and outer shafts 32 and 22 are supported for rotation in one or more bearings 34. One or more turbine frames provide structural load paths from the bearings 34 to an outer casing 36, which forms a backbone structure of the engine 10. In particular, a turbine frame assembly, which comprises a turbine frame 38 that integrates a first stage nozzle cascade 40 of the work turbine 24, is disposed between the gas generator turbine 16 and the work turbine 24.
[0020] Figures 2-4 illustrate the construction of the turbine frame assembly in more detail. The turbine frame 38 comprises an annular, centrally-located hub 42 with forward and aft faces 44 and 46, surrounded by an annular outer ring 48 having forward and aft flanges 50 and 52. The hub 42 and the outer ring 48 are interconnected by a plurality of radially -extending struts 54. In the illustrated example there are six equally-spaced struts 54. The turbine frame 38 may be a single integral unit or it may be built up from individual components. In the illustrated example it is cast in a single piece from a metal alloy suitable for high-temperature operation, such as a cobalt- or nickel-based "superalloy" An example of a suitable material is a nickel-based alloy commercially known as IN 718. Each of Ae struts 54 is hollow and terminates in a bleed air port 56 at its outer end, outboard of the outer ring 48.
(0021] A plurality of service tube assemblies 58 are mounted in the turbine frame 38, positioned between the struts 54, and extend between the outer ring 48 and the hub 42. In this example there are six service tube assemblies 58. Figures 3C and 5 show the service tube assembly in more detail. Each service tube assembly 58 includes a hollow service tube 60. The service tube 60 has a central section 55 disposed between reduced-diameter outer and inner ends 57 and 59. The inner end 59 includes a generally cylindrical male fitting 61 which forms a plug-in connection in cooperation with a female receptacle 63 of a sump 65 located within the turbine frame 38. The service tube 60 may be used to transport air or oil from between the sump 65 and an external conduit (not shown) such as an oil supply- or scavenge line, or sump pressurization or vent line, which is coupled to the outer end 57.
(0022] The service tube 60 is surrounded by a hollow housing 71 which is an integral component that comprises a service tube baffle 62 pierced with impingement cooling holes 64, a mounting bracket 66, and a manifold 68 with an inlet tube 70. The outer end 73 of the housing 71 is attached to an annular flange 75 at the outer end 57 of the service tube 60, for example by brazing or welding. The inner end 77 of Ae housing 71 is free to move thermally in operation, and has an opening that closely surrounds the central section 55 so as to leave a small gap for cooling air flow, as explained in more detail below. The central section 55 may include an annular collar 79 about its outer periphery to define the gap in cooperation with the housing 71.
[0023 { The service tube assemblies 58 plug into aligned openings in the outer ring 48 and the hub 42, and are secured to the outer ring 48 using bolts passing through the mounting bracket 66. [0024 j The nozzle cascade 40 comprises a plurality of actively-cooled airfoils. In this particular example there are 48 airfoils in total. This number may be varied to suit a particular application. Some of the airfoils, in this case 12, are axially elongated and are incorporated into fairings (see Figure 4) which protect the struts 54 and service tube assemblies 58 from hot combustion gases. Some of the fairings, in this case 6, are strut fairings 72 which are of a split configuration. The remainder of the fairings are service tube fairings 74 which are a single piece configuration. The remaining airfoils, in this case 36, are arranged into nozzle segments 76 having one or more vanes each.
[0025 j For the purposes of the present invention only the service tube fairings 74 will be described in detail. The other components of the nozzle cascade 40 are described in co- pending application by J. A. Manteiga et al. entitled "Turbine Frame Assembly and Method for a Gas Turbine Engine", which is which is incorporated herein by reference.
[00261 Figure 6 shows one of the service tube fairings 74 in more detail, ft includes an airfoil-shaped hollow vane 120 that is supported between an arcuate outer band 122 and an arcuate inner band 124. The inner and outer bands 124 and 122 are axially elongated and shaped so that they define a portion of the flowpath through the turbine frame 38. A forward hook 126 protrudes axially forward from the outer face of the outer band 122, and an aft hook 128 protrudes axially forward from the outer face of the outer band 122. The vane 120 is axially elongated and includes spaced-apart sidewalls 132 extending between a leading edge 134 and a trailing edge 136. The sidewalls 132 are shaped so as to form an aerodynamic fairing for the service tube assembly 58. A forward section 138 of the vane 120 is hollow and is impingement cooled, in a manner described in more detail below. An aft section 140 of the vane 120 is also hollow and incorporates walls 142 that define a multiple-pass serpentine flowpath. A plurality of trailing edge passages 144, such as slots or holes, pass through the trailing edge 136 of each vane 120.
(0027] The service tube fairings 74 are cast from a metal alloy suitable for high- temperature operation, such as a cobalt- or nickel-based "superalloy", and may be cast with a specific crystal structure, such as directionalry-solidified (OS) or single-crystal (SX), in a known manner. An example of one suitable material is a nickel-based alloy commercially known as RENE N4. [0028| As shown in Figure 2 and 3, the strut fairings 72, service tube fairings 74, and nozzle segments 76 are ail supported by forward and ait hangers 164 and 166 which are fastened to the forward and aft flanges SO and 52 of the turbine frame 38, respectively, for example using bolts or other suitable fasteners.
[0029 J The forward nozzle hanger 164 is generally disk-shaped and includes an outer flange 168 and an inner flange 170, interconnected by an aft-extending arm 172 having a generally "V-shaped cross-section. The inner flange 170 defines a mounting rail 174 with a slot 176 which accepts the forward hooks 126of the service tube fairings 74 and similar hooks of the strut fairings 72 and nozzle segments 76. The outer flange 168 has bolt holes therein corresponding to bolt holes in the forward flange 50 of the turbine frame 38. The forward nozzle hanger 164 supports the nozzle cascade 40 radially in a way that allows compliance in the axial direction.
[0030} The aft nozzle hanger 166 is generally disk-shaped and includes an outer flange 175 and an inner flange 177, interconnected by forward-extending arm 180 having a generally "IT-shaped cross-section. The inner flange 177 defines a mounting rail 182 with a slot 184 which accepts the aft hooks 128 of the service tube fairings 74 and similar hooks of the strut fairings 72 and nozzle segments 76. The outer flange 175 has bolt holes therein corresponding to bolt holes in the aft flange 52 of the turbine frame 38. The aft nozzle hanger 166 supports the nozzle cascade 48 radially while providing restraint in the axial direction.
|00311 When assembled, the outer bands of the strut fairings 72, service tube fairings 74, and nozzle segments 76 cooperate with the outer ring 48 of the turbine frame 38 to define an annular outer band cavity 186 (see Figure 3).
[0032} An annular outer balance piston (OPB) seal 188 is attached to the aft face of the hub 42, for example with bolts or other suitable fasteners. The OBP seal 188 has a generally "L"-shaped cross-section with a radial arm 190 and an axial arm 192. A forward sealing lip 194 bears against the hub 42. and an aft, radially -oufwardk -extending sealing lip 1% captures an annular, "M"-shaped seal 198 against the nozzle cascade 40. A similar "NT-shaped seal 200 is captured between the forward end of the nozzle cascade 40 and another sealing lip 202 on an stationary engine structure 204. Collectively, the hub 42 and the OBP seal 188 define an inner manifold 206 which communicates with the interior of the hub 42. Also, the inner bands of the strut fairings 72, service tube fairings 74, and nozzle segments 76 cooperate with the hub 42 of the turbine frame 38, the OBP seal 188, and (he seals 198 and 200 to define an annular inner band cavity 208. One or more cooling holes 210 pass through the radial arm 190 of the OBP seal 188. In operation, these cooling holes 210 pass cooling air from the hub 42 to an annular seal plate 212 mounted on a front face of the downstream rotor 28. The cooling air enters a hole 214 in the seal plate 212 and is then routed to the rotor 28 in a conventional fashion.
|OO33] Theaxialarm l92 c^the OBP $eal l88carriesanabradablemateπal216(such as a metallic hones comb) which mates with a seal tooth 218 of the seal plate 212.
[0034] Referring to Figures 4 and 6 , cooling of the service tube fairings 74 is as follows. Cooling air bled from a source such as the compressor 12 (see Figure 1) is fed into the inlet tubes 70, as shown by the arrow "A".
10035] One portion of this flow exits impingement cooling holes 64 in the service tube baffles 62 and is used for impingement cooling the service tube fairings 74, as shown by arrows "C (see Figure 6). After impingement cooling, the air passes to the outer band cavity 186, as shown at "D". Another portion of air exits the service tube baffles 62 and enters the outer band cavity 186 directly, as shown by arrows "E". Finally, a third portion of the air from the service lube baffles 62 exits the between the service tube baffle 62 and the service tube 60 and purges the inner band cavity 208 (see arrow "P).
(0036] Air from the outer band cavity 186, which is as combination of purge air and post-impingement flows denoted O and E in Figure 6, enters the serpentine passages in the aft sections of the vanes 120as shown at "G". It is then used therein for convective cooling in a conventional manner and subsequently exhausted through the trailing edge cooling passages.
[0037] The turbine frame assembly described above has multiple advantages over prior art designs. The engine 10 can run hotter and longer without oil coked sump services. The service tube assemblies 58 are "plug -in" components permitting inspection or cleaning without engine disassembly. Also, integration of the service tube and liner cooling improves packaging by moving the service tubes 60 away from the struts 54. There is a potential for less flowpath blockage and better engine performance than with conventional designs. Furthermore, this frees up (he struts 54 for use in providing cooling air to downstream turbine rotors or other components.
(0038] The foregoing has described a turbine frame assembly for a gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of die preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation, the invention being defined by the claims.

Claims

WHAT IS CLAIMED IS:
1. A service tube apparatus for a gas turbine engine, comprising: a service tube assembly including:
(a) an elongated, hollow service tube; and
(b) a service tube baffle surrounding the service tube which is pierced with a plurality of impingement cooling holes.
2. The service tube apparatus of claim 1 further comprising a hollow housing surrounding the service tube which includes:
(a) a manifold including an inlet tube:
(b) a mounting bracket; and
(c) the service tube baffle.
3. The service tube apparatus of claim 2 wherein the housing is a single integral component.
4. The service tube apparatus of claim 2 wherein an outer end of the housing is rigidly connected (o the service tube, and an inner end of the housing is free to move in a radial direction relative to the service tube.
5. The service tube apparatus of claim 4 wherein an annular gap is defined between an inner end of the housing and the service tube.
6. The service tube apparatus of claim 1 wherein the service tube includes inner and outer ends, the inner end terminating in a generally cylindrical male fitting.
7. The sen ice tube apparatus of claim 6 wherein the service tube incorporates an enlarged-diameter central portion disposed between the inner and outer ends.
8. The service tube apparatus of claim 1 further including a service tube fairing surrounding the service tube assembly, the service tube fairing comprising:
(a) an arcuate outer band:
(b) an arcuate inner band; and
(c) an airfoil-shaped vane; wherein the vane defines a continuous fairing around the service tube assembly.
9. The servi ce tube apparatus of claim 8 wherein the vane of the service lube fairing includes walls defining a serpentine flow path therein, the serpentine flow path in fluid communication with at least one trailing edge passage disposed at a trailing edge of the vane.
10. A turbine frame assembly for a gas turbine engine, comprising:
(a) a lurbine frame including:
(i) an outer ring;
(ii) a hub; and
(ii) a plurality of struts extending between the hub and the outer ring;
(b) at least one sen ice tube apparatus extending between the hub and the outer ring, comprising a sen ice tube assembly including:
(i) an elongated, hollow service tube; and
(ii) a service tube baffle surrounding the service tube which is pierced with a plurality of impingement cooling holes.
11. The turbine frame assembly of claim 10 wherein the outer ring, the hub, and the struts are a single integral casting.
12. The turbine frame assembly of claim 10 further comprising a hollow housing surrounding the service tube which includes:
(a) a manifold including an inlet tube:
(b) a mounting bracket; and
(c) the service tube baffle.
13. The turbine frame assembly of claim 12 wherein the housing is a single integral component.
14. The turbine frame assembly of claim 12 wherein an outer end of the housing is rigidly connected to the service tube, and an inner end of the housing is free to move in a radial direction relative to the service tube.
15. The turbine frame assembly of claim 14 wherein an annular gap is defined between an inner end of the housing and the service tube.
16. The turbine frame assembly- of claim 10 wherein the service tube includes inner and outer ends, the inner end terminating in a generally cylindrical male fitting.
17. The turbine frame assembly of claim 16 wherein the service tube incorporates an enlarged-diameter central portion disposed between the inner and outer ends.
18. The turbine frame assembly of claim IO further including a service tube fairing surrounding the service tube assembly, the service tube fairing comprising:
(a) an arcuate outer band;
(b) an arcuate inner band; and
(c) an airfoil-shaped vane, wherein the vane defines a continuous fairing around the sen ice lube assemblv.
19. The turbine frame assembly of claim 18 wherein the vane of the service tube fairing includes wails defining a serpentine flow path therein, the serpentine flow path in fluid communication with at least one trailing edge passage disposed at a trailing edge of the vane.
20. The turbine frame assembly of claim 10 wherein the sen ice lube fairings are secured to the turbine frame by spaced-apan annular forward and aft nozzle hangers which engage the outer bands of the service tube fairings.
PCT/US2009/055147 2008-11-29 2009-08-27 Integrated service tube and impingement baffle for a gas turbine engine Ceased WO2010062428A2 (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
DE112009003550T DE112009003550T5 (en) 2008-11-29 2009-08-27 Integrated supply pipe and baffle for a gas turbine engine
JP2011538608A JP5414805B2 (en) 2008-11-29 2009-08-27 Integrated service tube and impingement baffle for gas turbine engines
CA2744226A CA2744226C (en) 2008-11-29 2009-08-27 Integrated service tube and impingement baffle for a gas turbine engine
GB1107527.2A GB2477453B (en) 2008-11-29 2009-08-27 Integrated service tube and impingement baffle for a gas tubine engine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US12/325,175 2008-11-29
US12/325,175 US8177488B2 (en) 2008-11-29 2008-11-29 Integrated service tube and impingement baffle for a gas turbine engine

Publications (2)

Publication Number Publication Date
WO2010062428A2 true WO2010062428A2 (en) 2010-06-03
WO2010062428A3 WO2010062428A3 (en) 2011-03-24

Family

ID=42222957

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2009/055147 Ceased WO2010062428A2 (en) 2008-11-29 2009-08-27 Integrated service tube and impingement baffle for a gas turbine engine

Country Status (6)

Country Link
US (1) US8177488B2 (en)
JP (1) JP5414805B2 (en)
CA (1) CA2744226C (en)
DE (1) DE112009003550T5 (en)
GB (1) GB2477453B (en)
WO (1) WO2010062428A2 (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2206882A3 (en) * 2008-12-23 2015-12-02 General Electric Company Centrifugal Compressor Forward Thrust and Turbine Cooling Apparatus
EP2206902A3 (en) * 2008-12-23 2015-12-02 General Electric Company Turbine cooling air from a centrifugal compressor

Families Citing this family (69)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9410608B2 (en) 2011-06-08 2016-08-09 United Technologies Corporation Flexible support structure for a geared architecture gas turbine engine
US9239012B2 (en) 2011-06-08 2016-01-19 United Technologies Corporation Flexible support structure for a geared architecture gas turbine engine
US9631558B2 (en) * 2012-01-03 2017-04-25 United Technologies Corporation Geared architecture for high speed and small volume fan drive turbine
US8894365B2 (en) 2011-06-29 2014-11-25 United Technologies Corporation Flowpath insert and assembly
CH705514A1 (en) 2011-09-05 2013-03-15 Alstom Technology Ltd Gas channel for gas turbine, has supports, outer housing and inner housing that are equipped with refractory linings that are fastened to support structure, such that stress-free thermal expansion of linings is ensured
US10287914B2 (en) 2012-01-31 2019-05-14 United Technologies Corporation Gas turbine engine with high speed low pressure turbine section and bearing support features
WO2013128683A1 (en) * 2012-02-27 2013-09-06 三菱重工業株式会社 Gas turbine
US10125693B2 (en) 2012-04-02 2018-11-13 United Technologies Corporation Geared turbofan engine with power density range
US20150308351A1 (en) 2012-05-31 2015-10-29 United Technologies Corporation Fundamental gear system architecture
US8572943B1 (en) 2012-05-31 2013-11-05 United Technologies Corporation Fundamental gear system architecture
US9212566B2 (en) 2012-06-15 2015-12-15 General Electric Company Tripod buckle for split fairing of a gas turbine engine
US10087847B2 (en) * 2012-09-26 2018-10-02 United Technologies Corporation Seal assembly for a static structure of a gas turbine engine
WO2014105657A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Mount with deflectable tabs
WO2014105735A1 (en) * 2012-12-29 2014-07-03 United Technologies Corporation Cast steel frame for gas turbine engine
WO2014105604A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Angled cut to direct radiative heat load
US10240532B2 (en) 2012-12-29 2019-03-26 United Technologies Corporation Frame junction cooling holes
EP2938836B1 (en) 2012-12-29 2020-02-05 United Technologies Corporation Seal support disk and assembly
WO2014137444A2 (en) 2012-12-29 2014-09-12 United Technologies Corporation Multi-ply finger seal
EP2938868B1 (en) 2012-12-29 2019-08-07 United Technologies Corporation Flow diverter assembly
US9903224B2 (en) 2012-12-29 2018-02-27 United Technologies Corporation Scupper channelling in gas turbine modules
WO2014105100A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Bumper for seals in a turbine exhaust case
WO2014105800A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Gas turbine seal assembly and seal support
EP2938863B1 (en) 2012-12-29 2019-09-25 United Technologies Corporation Mechanical linkage for segmented heat shield
DE112013006258B4 (en) 2012-12-29 2026-01-15 RTX Corporation (n.d.Ges.d. Staates Delaware) Turbine frame arrangement and method for providing a turbine frame arrangement
WO2014105616A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Turbine exhaust case architecture
WO2014105599A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Heat shield for cooling a strut
US10053998B2 (en) 2012-12-29 2018-08-21 United Technologies Corporation Multi-purpose gas turbine seal support and assembly
WO2014105619A1 (en) * 2012-12-29 2014-07-03 United Technologies Corporation Multi-function boss for a turbine exhaust case
WO2014105602A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Heat shield for a casing
WO2014105603A1 (en) 2012-12-29 2014-07-03 United Technologies Corporation Multi-piece heat shield
JP6271582B2 (en) 2012-12-29 2018-01-31 ユナイテッド テクノロジーズ コーポレイションＵｎｉｔｅｄ Ｔｅｃｈｎｏｌｏｇｉｅｓ Ｃｏｒｐｏｒａｔｉｏｎ Gas turbine seal assembly and seal support
WO2014105716A1 (en) 2012-12-31 2014-07-03 United Technologies Corporation Turbine exhaust case multi-piece frame
JP6249499B2 (en) * 2012-12-31 2017-12-20 ユナイテッド テクノロジーズ コーポレイションＵｎｉｔｅｄ Ｔｅｃｈｎｏｌｏｇｉｅｓ Ｃｏｒｐｏｒａｔｉｏｎ Multi-piece frame for turbine exhaust case
EP2938860B1 (en) 2012-12-31 2018-08-29 United Technologies Corporation Turbine exhaust case multi-piece frame
US10221707B2 (en) 2013-03-07 2019-03-05 Pratt & Whitney Canada Corp. Integrated strut-vane
US10330011B2 (en) * 2013-03-11 2019-06-25 United Technologies Corporation Bench aft sub-assembly for turbine exhaust case fairing
WO2015017040A2 (en) 2013-07-30 2015-02-05 United Technologies Corporation Gas turbine engine vane ring arrangement
US9835038B2 (en) 2013-08-07 2017-12-05 Pratt & Whitney Canada Corp. Integrated strut and vane arrangements
US9644495B2 (en) * 2013-08-20 2017-05-09 Honeywell International Inc. Thermal isolating service tubes and assemblies thereof for gas turbine engines
US9556746B2 (en) * 2013-10-08 2017-01-31 Pratt & Whitney Canada Corp. Integrated strut and turbine vane nozzle arrangement
US9598981B2 (en) * 2013-11-22 2017-03-21 Siemens Energy, Inc. Industrial gas turbine exhaust system diffuser inlet lip
US9587519B2 (en) * 2013-11-22 2017-03-07 Siemens Energy, Inc. Modular industrial gas turbine exhaust system
EP3140515B1 (en) 2014-05-08 2019-04-03 Siemens Energy, Inc. Airfoil cooling with internal cavity displacement features
US9920869B2 (en) * 2014-05-22 2018-03-20 United Technologies Corporation Cooling systems for gas turbine engine components
US9803502B2 (en) * 2015-02-09 2017-10-31 United Technologies Corporation Cooling passages for a mid-turbine frame
US9885254B2 (en) * 2015-04-24 2018-02-06 United Technologies Corporation Mid turbine frame including a sealed torque box
US9982600B2 (en) * 2015-05-19 2018-05-29 United Technologies Corporation Pre-skewed capture plate
FR3036442B1 (en) * 2015-05-21 2021-07-16 Snecma TURBOMACHINE WITH A VENTILATION SYSTEM
US9683458B2 (en) 2015-08-20 2017-06-20 Pratt & Whitney Canada Corp. Oil scupper system for bearing housing of gas turbine engine
JP6546481B2 (en) * 2015-08-31 2019-07-17 川崎重工業株式会社 Exhaust diffuser
US10577973B2 (en) * 2016-02-18 2020-03-03 General Electric Company Service tube for a turbine engine
DE102016217320A1 (en) * 2016-09-12 2018-03-15 Siemens Aktiengesellschaft Gas turbine with separate cooling for turbine and exhaust housing
US10422246B2 (en) 2016-10-24 2019-09-24 General Electric Company Bearing housing with impingement cooling system
US10914187B2 (en) * 2017-09-11 2021-02-09 Raytheon Technologies Corporation Active clearance control system and manifold for gas turbine engine
US10605119B2 (en) * 2017-09-25 2020-03-31 United Technologies Corporation Turbine frame assembly for gas turbine engines
US20190112946A1 (en) * 2017-10-13 2019-04-18 United Technologies Corporation Double wall service tube with annular airflow passage
FR3076860B1 (en) * 2018-01-12 2020-11-13 Safran Aircraft Engines TURBOMACHINE SERVITUDES PASSAGE ARM
US10746049B2 (en) * 2018-03-30 2020-08-18 United Technologies Corporation Gas turbine engine case including bearing compartment
FR3086339B1 (en) * 2018-09-20 2021-01-15 Safran Aircraft Engines TURBOREACTOR INTER-TURBINE CASING EQUIPPED WITH A COOLING AND VENTILATION CIRCUIT
JP7419002B2 (en) * 2019-09-12 2024-01-22 三菱重工業株式会社 Strut cover, exhaust casing and gas turbine
US11085313B2 (en) * 2019-11-22 2021-08-10 Raytheon Technologies Corporation System and method for transporting lubricant through a vane
US11261757B2 (en) * 2019-12-05 2022-03-01 Pratt & Whitney Canada Corp. Boss for gas turbine engine
US11970279B2 (en) 2020-02-21 2024-04-30 General Electric Company Control system and methods of controlling an engine-mounting link system
US11939070B2 (en) 2020-02-21 2024-03-26 General Electric Company Engine-mounting links that have an adjustable inclination angle
US11473439B1 (en) * 2021-09-23 2022-10-18 General Electric Company Gas turbine engine with hollow rotor in fluid communication with a balance piston cavity
CN116927953A (en) 2022-03-31 2023-10-24 通用电气公司 Surface with shape memory alloy particles
CN117365748A (en) 2022-07-01 2024-01-09 通用电气公司 Self-cleaning tubing for hydrocarbon fluids
CN117365747A (en) 2022-07-01 2024-01-09 通用电气公司 Self-cleaning tubing for hydrocarbon fluids
US12467366B2 (en) * 2023-12-01 2025-11-11 General Electric Company Turbine engine with a nozzle having cooling features

Family Cites Families (31)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2961150A (en) 1958-12-30 1960-11-22 Gen Electric Frame structure for turbo-machine
FR1325291A (en) 1962-05-28 1963-04-26 Rolls Royce Bearing support structure for turbojet
DE2712856A1 (en) * 1977-03-21 1978-09-28 Cary Dipl Ing Brandt OIL BURNER ARRANGEMENT FOR HEATING SYSTEMS
US4369016A (en) 1979-12-21 1983-01-18 United Technologies Corporation Turbine intermediate case
US5020318A (en) 1987-11-05 1991-06-04 General Electric Company Aircraft engine frame construction
US5080555A (en) 1990-11-16 1992-01-14 General Motors Corporation Turbine support for gas turbine engine
JPH0792002B2 (en) 1991-12-26 1995-10-09 ゼネラル・エレクトリック・カンパニイ Damper assembly for gas turbine engine strut
GB2267736B (en) 1992-06-09 1995-08-09 Gen Electric Segmented turbine flowpath assembly
US5272869A (en) 1992-12-10 1993-12-28 General Electric Company Turbine frame
US5292227A (en) 1992-12-10 1994-03-08 General Electric Company Turbine frame
US5284011A (en) 1992-12-14 1994-02-08 General Electric Company Damped turbine engine frame
US5483792A (en) 1993-05-05 1996-01-16 General Electric Company Turbine frame stiffening rails
US5438756A (en) 1993-12-17 1995-08-08 General Electric Company Method for assembling a turbine frame assembly
US5851105A (en) 1995-06-28 1998-12-22 General Electric Company Tapered strut frame
US5634767A (en) 1996-03-29 1997-06-03 General Electric Company Turbine frame having spindle mounted liner
US6102577A (en) 1998-10-13 2000-08-15 Pratt & Whitney Canada Corp. Isolated oil feed
US6267553B1 (en) * 1999-06-01 2001-07-31 Joseph C. Burge Gas turbine compressor spool with structural and thermal upgrades
US6358001B1 (en) 2000-04-29 2002-03-19 General Electric Company Turbine frame assembly
US6439841B1 (en) 2000-04-29 2002-08-27 General Electric Company Turbine frame assembly
US6447248B1 (en) 2000-10-20 2002-09-10 General Electric Company Bearing support fuse
US6612807B2 (en) 2001-11-15 2003-09-02 General Electric Company Frame hub heating system
US6708482B2 (en) 2001-11-29 2004-03-23 General Electric Company Aircraft engine with inter-turbine engine frame
US6672833B2 (en) 2001-12-18 2004-01-06 General Electric Company Gas turbine engine frame flowpath liner support
US6796765B2 (en) 2001-12-27 2004-09-28 General Electric Company Methods and apparatus for assembling gas turbine engine struts
US6935837B2 (en) 2003-02-27 2005-08-30 General Electric Company Methods and apparatus for assembling gas turbine engines
US6860716B2 (en) 2003-05-29 2005-03-01 General Electric Company Turbomachine frame structure
US6983608B2 (en) 2003-12-22 2006-01-10 General Electric Company Methods and apparatus for assembling gas turbine engines
US7353647B2 (en) 2004-05-13 2008-04-08 General Electric Company Methods and apparatus for assembling gas turbine engines
US7007488B2 (en) * 2004-07-06 2006-03-07 General Electric Company Modulated flow turbine nozzle
FR2898641B1 (en) * 2006-03-17 2008-05-02 Snecma Sa CARTERING IN A TURBOJET ENGINE
FR2914017B1 (en) * 2007-03-20 2011-07-08 Snecma SEALING DEVICE FOR A COOLING CIRCUIT, INTER-TURBINE HOUSING BEING EQUIPPED AND TURBOREACTOR COMPRISING THE SAME

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2206882A3 (en) * 2008-12-23 2015-12-02 General Electric Company Centrifugal Compressor Forward Thrust and Turbine Cooling Apparatus
EP2206902A3 (en) * 2008-12-23 2015-12-02 General Electric Company Turbine cooling air from a centrifugal compressor

Also Published As

Publication number Publication date
GB201107527D0 (en) 2011-06-22
WO2010062428A3 (en) 2011-03-24
JP2012510588A (en) 2012-05-10
US20100135786A1 (en) 2010-06-03
CA2744226A1 (en) 2010-06-03
US8177488B2 (en) 2012-05-15
JP5414805B2 (en) 2014-02-12
DE112009003550T5 (en) 2012-08-23
GB2477453B (en) 2014-04-02
GB2477453A (en) 2011-08-03
CA2744226C (en) 2017-02-28

Similar Documents

Publication Publication Date Title
CA2744226C (en) Integrated service tube and impingement baffle for a gas turbine engine
CA2680634C (en) Turbine frame assembly and method for a gas turbine engine
CA2680640C (en) Split fairing for a gas turbine engine
US10822973B2 (en) Shroud for a gas turbine engine
CA2647121C (en) Multi-source gas turbine cooling
US10605168B2 (en) Interdigitated turbine engine air bearing cooling structure and method of thermal management
US10718265B2 (en) Interdigitated turbine engine air bearing and method of operation
US9845691B2 (en) Turbine nozzle outer band and airfoil cooling apparatus
US10876407B2 (en) Thermal structure for outer diameter mounted turbine blades
EP2105580B1 (en) Hybrid impingement cooled turbine nozzle
EP1384858A2 (en) Internal cooling of the case of a low pressure turbine
US10787931B2 (en) Method and structure of interdigitated turbine engine thermal management
EP2551458A2 (en) Blade Cooling and Sealing System
US20180142564A1 (en) Combined turbine nozzle and shroud deflection limiter
EP2096265A2 (en) Turbine nozzle with integral impingement blanket
EP3232009B1 (en) Blade outer air seal having retention snap ring
US11021961B2 (en) Rotor assembly thermal attenuation structure and system
EP3425164B1 (en) Tandem rotor disk apparatus and corresponding gas turbine engine

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 09791984

Country of ref document: EP

Kind code of ref document: A2

ENP Entry into the national phase

Ref document number: 1107527

Country of ref document: GB

Kind code of ref document: A

Free format text: PCT FILING DATE = 20090827

WWE Wipo information: entry into national phase

Ref document number: 1107527.2

Country of ref document: GB

WWE Wipo information: entry into national phase

Ref document number: 2744226

Country of ref document: CA

WWE Wipo information: entry into national phase

Ref document number: 2011538608

Country of ref document: JP

122 Ep: pct application non-entry in european phase

Ref document number: 09791984

Country of ref document: EP

Kind code of ref document: A2