(19)
(11)EP 3 064 716 B1

(12)EUROPEAN PATENT SPECIFICATION

(45)Mention of the grant of the patent:
01.05.2019 Bulletin 2019/18

(21)Application number: 16158692.0

(22)Date of filing:  04.03.2016
(51)International Patent Classification (IPC): 
F01D 9/06(2006.01)
F01D 25/28(2006.01)

(54)

MID TURBINE FRAME WITH INTEGRATED INNER CASE HEAT SHIELD

MITTELTURBINENRAHMEN MIT INTEGRIERTEM INNENGEHÄUSEHITZESCHILD

CADRE DE TURBINE MOYENNE AVEC BOUCLIER THERMIQUE D'ÉTUI INTERNE INTÉGRÉ


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

(30)Priority: 06.03.2015 US 201514640323

(43)Date of publication of application:
07.09.2016 Bulletin 2016/36

(73)Proprietor: United Technologies Corporation
Farmington, CT 06032 (US)

(72)Inventors:
  • WINN, Joshua David
    Ellington, CT Connecticut 06029 (US)
  • CHEROLIS, Anthony P.
    Hartford, CT Connecticut 06114 (US)
  • PORTER, Steven D.
    Wethersfield, CT Connecticut 06109 (US)
  • TREAT, Christopher
    Farmington, CT 06032 (US)
  • ZACCHERA, Kevin
    Glastonbury, CT Connecticut 06033 (US)

(74)Representative: Dehns 
St. Brides House 10 Salisbury Square
London EC4Y 8JD
London EC4Y 8JD (GB)


(56)References cited: : 
WO-A1-2014/052007
US-A1- 2014 007 588
US-A1- 2011 081 237
  
      
    Note: Within nine months from the publication of the mention of the grant of the European patent, any person may give notice to the European Patent Office of opposition to the European patent granted. Notice of opposition shall be filed in a written reasoned statement. It shall not be deemed to have been filed until the opposition fee has been paid. (Art. 99(1) European Patent Convention).


    Description

    BACKGROUND



    [0001] This disclosure relates to shielding a component of a gas turbine engine, and more particularly to a shielding arrangement for a component of a mid-turbine frame.

    [0002] Gas turbine engines can include a fan for propulsion air and to cool components. The fan also delivers air into a core engine where it is compressed. The compressed air is then delivered into a combustion section, where it is mixed with fuel and ignited. The combustion gas expands downstream over and drives turbine blades.

    [0003] Some turbine engines include a mid-turbine frame configured to transfer loads between a bearing compartment and other portions of the engine static structure. The mid-turbine frame may include one or more static airfoils between inner and outer cases of the mid-turbine frame to direct the flow of products of combustion to the turbine blades. One or more service lines may extend through the airfoils.

    [0004] US 2011/081237 discloses sealing for vane segments, in which a seal housing is provided to substantially cover at least one duct wall of vane array duct of a gas turbine engine, and one example arrangement is employed in a mid-turbine frame.

    SUMMARY



    [0005] The present invention provides a mid-turbine frame for a gas turbine engine as defined in claim 1.

    [0006] In a further embodiment of any of the foregoing embodiments, the flange is at least partially received in the first frame case.

    [0007] A further embodiment of any of the foregoing embodiments, the mid-turbine frame includes a second frame case coaxial with the first frame case. The conduit extends through an airfoil between the first and second frame cases.

    [0008] In a further embodiment of any of the foregoing embodiments, the first and second frame cases bound a core flow path, and the airfoil is positioned in the core flow path.

    [0009] In a further embodiment of any of the foregoing embodiments, the conduit includes a distal portion extending through the airfoil and a proximal portion coupled to the flange. The distal portion is surrounded by an insulated sheath terminating at the proximal portion, and the heat shield extends a distance along the proximal portion.

    [0010] In a further embodiment of any of the foregoing embodiments, the conduit is configured to communicate fluid with a bearing compartment.

    [0011] A further embodiment of any of the foregoing embodiments, the mid-turbine frame includes a second frame case, and a plurality of spokes connecting the first and second frame cases. The heat shield is spaced in a circumferential direction from each of the plurality of spokes.

    [0012] A gas turbine engine according to an example of the present disclosure includes a mid-turbine frame of any of the foregoing embodiments, axially between a first turbine and a second turbine. The mid-turbine frame includes a plurality of airfoils radially between an inner frame case and an outer frame case, a plurality of flanges coupled to the inner frame case, and a plurality of conduits between the inner and outer frame cases. The plurality of conduits is coupled to the plurality of flanges. A plurality of localized heat shields each corresponds to one of the plurality of flanges.

    [0013] In a further embodiment of any of the foregoing embodiments, at least one of the plurality of conduits is configured to communicate cooling airflow with a bearing compartment.

    [0014] In a further embodiment of any of the foregoing embodiments, the plurality of conduits is an oil supply line and an oil scavenge line each coupled to a bearing compartment.

    [0015] In a further embodiment of any of the foregoing embodiments, opposed walls of the inner and outer frame cases bound a core flow path, and the plurality of airfoils are located in the core flow path.

    [0016] In a further embodiment of any of the foregoing embodiments, the plurality of localized heat shields are arranged between the plurality of airfoils and surfaces of the plurality of flanges.

    [0017] In a further embodiment of any of the foregoing embodiments, each heat shield of the plurality of heat shields is spaced circumferentially about the inner frame case, and each heat shield of the plurality of heat shields defines a cavity between the inner frame case and one of the plurality of flanges.

    [0018] The present invention further provides a method of cooling a portion of a gas turbine engine as defined in claim 12.

    [0019] In a further embodiment of any of the foregoing embodiments, the heat shield is arranged between an airfoil receiving the conduit and surfaces of the flange.

    [0020] In a further embodiment of any of the foregoing embodiments, the conduit is coupled to a bearing compartment.

    [0021] Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.

    [0022] The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of an embodiment. The drawings that accompany the detailed description can be briefly described as follows.

    BRIEF DESCRIPTION OF THE DRAWINGS



    [0023] 

    Figure 1 schematically shows a gas turbine engine.

    Figure 2 schematically shows a perspective view of a mid-turbine frame.

    Figure 3 illustrates a cross-sectional view of the mid-turbine frame taken along line 3-3 of Figure 2.

    Figure 4A is a top view of selected portions of a mid-turbine frame having a localized heat shield.

    Figure 4B is a perspective view of selected portions of the mid-turbine frame of Figure 4A.

    Figure 4C is a perspective view of selected portions of the mid-turbine frame of Figure 4B with the heat shield uninstalled.


    DETAILED DESCRIPTION



    [0024] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

    [0025] The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.

    [0026] The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

    [0027] The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.

    [0028] The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

    [0029] A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition -- typically cruise at about 0.8 Mach and about 10,668 m (35,000 feet). The flight condition of 0.8 Mach and 10,668 m (35,000 feet), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. "Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)]0.5. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about 350.5 m/s (1150 ft/s).

    [0030] Figure 2 is a schematic perspective view of the mid-turbine frame 57. The mid-turbine frame 57 includes an outer frame case 62 and an inner frame case 64 which can be arranged about the central or engine axis A. The mid-turbine frame 57 is arranged axially between the low pressure turbine 46 (or first turbine) and the high pressure turbine 54 (or second turbine) shown in Figure 1. One or more spokes or tie rods 66 are distributed around a circumference of the inner frame case 64 and extend radially between an inner diameter 61 of the outer frame case 62 and an outer diameter 63 of the inner frame case 64 to provide structural support between the outer and inner frame cases 62, 64. For the purposes of this disclosure, the axial and radial directions are in relation to the central axis A unless stated otherwise.

    [0031] The inner frame case 64 supports the rotor assembly via the bearing system 38 (shown in Figure 1), and distributes the force from the inner frame case 64 to the outer frame case 62 via the spokes 66. The outer frame case 62 is configured to transfer loads from the inner frame case 64 to other portions of the engine static structure 36 (shown in Figure 1). In some examples, one or more spokes 66A define cooling passage(s) 55 (shown schematically in dashed line) to communicate cooling airflow from a coolant source 67 to the bearing compartment 38 or to a rotor assembly of the turbine section 28. Coolant sources 67 can include, but are not limited to, bleed air from an upstream stage of the compressor section 24, bypass air, or a secondary cooling system aboard the aircraft, for example.

    [0032] One or more service lines 68 are distributed around the circumference of the inner frame case 64 to provide a path between the outer and inner frame cases 62, 64. The service lines 68 can be positioned between adjacent pairs of spokes 66. The service lines 68 can include, but are not limited to, a conduit, a pressurized lubrication or oil supply line 68A, a lubrication or oil scavenge line 68B, a coolant supply line 68C coupled to the coolant source 67, and a secondary lubrication or oil drain line 68D, for example.

    [0033] The lubrication supply and scavenge lines 68A, 68B are configured to communicate lubrication between the bearing compartment 38 and a lubrication source 69, which can include a reservoir and lubrication pump, for example. In this manner, the bearing compartment 38 is lubricated and cooled during operation of the engine 20. Although six spokes 66 and four service lines 68 are shown in Figure 2, the particular number of spokes 66 and service lines 68 can be determined based on the needs of a particular situation in view of the teachings herein.

    [0034] Each of the spokes 66 and service lines 68 extends within an airfoil 59 or static vane (shown in Figures 1 and 3) arranged in the core flow path C (shown in Figure 1). Accordingly, portions of the spokes 66 and service lines 68 are adjacent to areas of the airfoil 59 that are subject to relatively high temperatures in the core flow path C.

    [0035] Figure 3 is a highly schematic view of the mid-turbine frame 57 including a localized heat shield 60. Although the localized heat shield 60 is shown with service line 68, such as the lubrication supply or scavenge lines 68A, 68B shown in Figure 2, it should be understood that any of the spokes 66 or service lines 68 can be provided with a heat shield 60. Other portions of the engine 20 can also benefit from the teachings herein, including conduits coupled to portions of the low pressure turbine 46, the high pressure turbine 54, an intermediate turbine for a three-spool engine, or adjacent structures.

    [0036] The service line 68 extends in a radial direction from an outer frame case 62 through the airfoil 59 to an inner frame case 64. The airfoil 59 includes an airfoil section 65 extending between from an outer platform 70 to an inner platform 71. In the illustrated example, the airfoil section 65 and the outer and inner platforms 70, 71 are integrally formed, and are all positioned radially inward of the outer frame case 62 and radially outward of the inner frame case 64. In alternative examples, the airfoil section 65 and the outer and inner platforms 70, 71 include multiple components attached together. The outer and inner platforms 70, 71 bound or otherwise define a portion of the core flow path C at the mid-turbine frame 57. The airfoil section 65 extends axially from a leading edge 72 to a trailing edge 73 to direct the flow of combustion products in the core flow path C.

    [0037] The airfoil 59 defines an airfoil cavity 74 for receiving a spoke 66 or a service line 68. The airfoil cavity 74 can be configured to communicate cooling airflow from the coolant source 67 (shown in Figure 2) to select portions of the airfoil 59 to provide impingement or film cooling, for example. Radially inward portions of the airfoil cavity 74 can be bounded or otherwise defined by the outer and inner frame cases 62, 64. In alternative embodiments, separate cavities are defined between the airfoil 59 and the outer and inner frame cases 62, 64.

    [0038] An insulated sheath 75 can be arranged about portions of the service line 68 to reduce exposure of radiation or heat from the airfoil 59. In the illustrated example, the service line 68 includes a distal (or first) portion 98 extending through the airfoil 59 and a proximal (or second) portion 99 coupled to the flange 78. The distal portion 98 is surrounded by the insulated sheath 75 which extends to, or terminates at, the proximal portion 99 or at the transfer tube 80.

    [0039] The service line 68 is coupled to a tube foot or flange 78 adjacent the inner frame case 64. In one embodiment, a transfer tube 80 couples the service line 68 and the flange 78 to each other. An inner path 79 within flange 78 is configured to communicate flow or lubricant L between an inner passage 76 of the service line 68 and a bearing cavity 39 within the bearing compartment 38. The lubricant L is communicated between the flange 78 and one or more bearings 41 via the bearing cavity 39. The bearing 41 is configured to support a shaft such as the inner shaft 40 (shown in Figure 1) radially inward of the bearing 41. As shown, the flange 78 abuts against, or is coupled to, a bearing support 82 coupled to or integral with the frame case 64. The flange 78 extends radially outward through the inner frame case 64 toward the airfoil cavity 74. In alternative examples, the flange 78 is coupled or attached to one of the outer and inner frame cases 62, 64 of the mid-turbine frame 57.

    [0040] In prior arrangements, surfaces 83 of the flange 78 bound or otherwise define a portion of the airfoil cavity 74. Accordingly, the surfaces 83 may be exposed to radiation from the core flow path C via airfoil 59, thereby elevating the temperature of lubricant L communicated from the service line 68 to the bearing cavity 39 to an undesirable level.

    [0041] In the illustrated example, the heat shield 60 includes a first portion 77 configured to receive the service line 68 and a second portion 85 configured to be arranged adjacent to a perimeter 90 of the flange 78 (shown in Figure 4C). The second portion 85 of the heat shield 60 abuts against, or is coupled to, one of the outer and inner frame cases 62, 64.

    [0042] The heat shield 60 is located radially between the flange 78 and a portion of the airfoil 59. At least a portion of the heat shield 60 is positioned in a line-of-sight between surfaces 83 of the flange 78 and portions of the airfoil 59 such that the flange 78 is at least partially shielded from radiation or heat emitted from the airfoil 59.

    [0043] The heat shield 60 may extend a radial distance along the proximal portion 99 or the transfer tube 80 to reduce a temperature of lubricant L within an inner passage 76 of the service line 68. In some examples, the mid-turbine frame 57 includes a plurality of localized heat shields 60 (illustrated schematically in Figure 2) each corresponding to one of a plurality of flanges 78.

    [0044] The heat shield 60 can be made of various materials to shield the flange 78 from radiation including, but not limited to, a steel alloy, such as a nickel based super alloy, stainless steel, a high-temperature composite or the like. The heat shield 60 can be formed utilizing various techniques, such as stamping or casting processes. The heat shield 60 can include one or more cutouts 84 (shown in Figures 4A and 4B) to assist in formation of the heat shield 60. The cutouts 84 can be defined at 84A (Figure 4A) or 84B, for example, or another location depending on the needs of a particular situation.

    [0045] The heat shield 60 is arranged to define a localized cooling cavity 86. The localized cooling cavity 86 is provided with one or more leakage gaps such as inlet 87 and outlet 88 to communicate cooling airflow to the flange 78. In the illustrated example, the inlet 87 is defined by the inner frame case 64 and the heat shield 60, and the outlet 88 in the shape of a circumferential ring or annulus defined by the service line 68 and the heat shield 60. In another example, the inlet 87 is partially defined by the outer frame case 62.

    [0046] In the illustrated example, cooling airflow circulates or is directed through the inlet 87 where it accumulates in the localized cooling cavity 86. The cooling airflow impinges on surfaces 83 of the flange 78 and reduces a temperature of the flange 78 and the lubricant L in the inner path 79 of flange 78. Relatively warm cooling airflow is then directed from the localized cooling cavity 86 toward the outlet 88. As the cooling airflow passes through the outlet 88 in to the airfoil cavity 74, radiation or heat exposure on the flange 78 is reduced, which in turn can reduce the amount of lubricant L needed to cool the bearing compartment 38 because a temperature of the lubricant L entering bearing compartment 38 is reduced. Reductions in radiation or heat exposure to the flange 78 can reduce flange warp, the formation of oxidation deposits in the lubricant L such as "coke" or "coking", and leakage of lubricant L from the bearing compartment 38 due to thermal expansion of the flange 78.

    [0047] Figures 4A to 4C show selected portions of the mid-turbine frame 57 shown in Figure 3. As shown in Figure 4A, the localized heat shield 60 can be spaced circumferentially from an adjacent pair of spokes 66 coupled to the inner frame case 64. As shown in Figure 4C, the heat shield is removed. The flange 78 can be received in, or extend through, a cutout or recess 91 within the inner frame case 64. The flange 78 can be secured utilizing one or more fasteners 92, such as one or more bolts or mating portions to accept corresponding bolts. The recess 91 is dimensioned to facilitate removal of the flange 78 from the inner frame case 64 during maintenance operations, for example.

    [0048] The heat shield 60 can be coupled to the inner frame case 64 utilizing various techniques. In some embodiments, the heat shield 60 includes one or more tabs 93 configured to couple the heat shield 60 to the inner frame case 64. Fasteners 94 such as threaded bolts or rivets can be utilized to attach the tabs 93 to the inner frame case 64, thereby facilitating disassembly and repair. In alternative embodiments, the heat shield 60 is attached to the inner frame case 64 utilizing welding or brazing techniques.

    [0049] The heat shield 60 can be configured having various geometries to reduce the amount of radiation communicated from an adjacent airfoil 59 to the flange 78. For example, the heat shield 60 can be configured to correspond or mate to a geometry of the flange 78 and/or the recess 91. In another example, a perimeter 95 of the section portion 85 of the heat shield 60 can be configured to correspond or mate to a perimeter 96 of the recess 91. The heat shield 60 can be arranged radially outward of the recess 91. In this arrangement, the perimeter 96 or overall dimension of the recess 91 can be reduced while accommodating the flange 78 within desired manufacturing tolerances. This arrangement also allows the heat shield 60 to be relatively insensitive to manufacturing tolerances.

    [0050] A portion of the heat shield 60 can extend radially outward from the inner frame case 64 along the transfer tube 80 or the service line 68 to define the localized cooling cavity 86 (shown in Figure 3). In further examples, the heat shield 60 extends radially outward to a radially outermost portion 97 of the transfer tube 80 (shown in Figure 4C) to reduce communication of radiation from the airfoil 59 to the transfer tube 80.

    [0051] During assembly, the flange 78 is inserted through the recess 91 of the inner frame case 64 in a direction from radially outward to radially inward. The fasteners 92 are tightened or secured as desired. The heat shield 60 is positioned over the flange 78 and secured to the inner frame case 64 utilizing fasteners 94. During disassembly, fasteners 94 are removed from the heat shield 60. The heat shield 60 is removed from the inner frame case 64. The flange 78 is moved through the recess 91 in a direction from radially inward to radially outward. Portions of the mid-turbine frame 57 or bearing compartment 38 radially inward of the outer diameter 63 of the inner frame case 64 can thereafter be accessed through the recess 91.

    [0052] During operation of the gas turbine engine 20, the airfoils 59 are subjected to extreme temperatures in the core flow path C. The airfoils 59 are cooled by cooling airflow that is communicated to the airfoil cavity 74. The cooling airflow may be provided through a passage 55 (shown in Figure 2) within one or more of the spokes 66, which accumulates within the bearing cavity 39, for example. A portion of the cooling airflow is communicated or directed from passages of the spokes 66, for example, to the inlet 87 of the heat shield 60 to cool the flange 78. The cooling airflow may be pressurized or communicated between the inlet 87 and outlet 88 of the localized cooling cavity 86 through convection. The relatively warm airflow is communicated or directed from the outlet 88 to the airfoil cavity 74. A portion of the cooling airflow exits at the airfoil 59 into the core flow path C through one or more film cooling holes (not shown), for example, or at another portion of mid-turbine frame 57.

    [0053] Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.

    [0054] It should be understood that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.

    [0055] The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.


    Claims

    1. A mid-turbine frame (57) for a gas turbine engine (20), comprising:

    a first frame case (64); and

    a flange (78) coupled to the first frame case (64),

    a heat shield (60) adjacent to the flange (78) and between adjacent spokes (66), a cooling cavity (86) defined between the heat shield (60) and the flange (78), the cooling cavity (86) having an inlet (87) bounded by the first frame case (64) and an outlet (88) bounded by a conduit (68) coupled to the flange (78),

    wherein the heat shield (60) includes a first portion (77) and a second portion (85), the first portion (77) extends in a circumferential direction, the second portion (85) extends in a radial direction and abuts the first frame case (64),

    wherein a perimeter (90) of the heat shield (60) mates with a perimeter of the flange (78).


     
    2. The mid-turbine frame as recited in any preceding claim, wherein the flange (78) is at least partially received in the first frame case (64).
     
    3. The mid-turbine frame as recited in any preceding claim, comprising a second frame case (62) coaxial with the first frame case (64), and wherein the conduit (68) extends through an airfoil (59) between the first and second frame cases (62, 64), wherein the first and second frame cases (62, 64) optionally bound a core flow path, and the airfoil (59) is positioned in the core flow path.
     
    4. The mid-turbine frame as recited in claim 3, wherein the conduit (68) includes a distal portion (98) extending through the airfoil (59) and a proximal portion (99) coupled to the flange (78), the distal portion (98) surrounded by an insulated sheath (75) terminating at the proximal portion (99), and the heat shield (60) extends a distance along the proximal portion (99).
     
    5. The mid-turbine frame as recited in claim 3 or 4, wherein the conduit (68) is configured to communicate fluid with a bearing compartment (38).
     
    6. The mid-turbine frame as recited in any preceding claim, comprising:

    a second frame case (62);

    a plurality of spokes (66) connecting the first and second frame cases (62, 64); and

    wherein the heat shield (60) is spaced in a circumferential direction from each of the plurality of spokes (66).


     
    7. A gas turbine engine (20), comprising:
    a mid-turbine frame (57) as recited in any preceding claim axially between a first turbine (46) and a second turbine (54), the mid-turbine frame (57) comprising:

    a plurality of airfoils (59) radially between an inner frame case (64) and an outer frame case (62);

    a plurality of flanges (78) coupled to the inner frame case (64);

    a plurality of conduits (68) between the inner and outer frame cases (62, 64), the plurality of conduits (68) coupled to the plurality of flanges (78); and

    a plurality of localized heat shields (60) each corresponding to one of the plurality of flanges (78).


     
    8. The gas turbine engine as recited in claim 7, wherein:
    at least one of the plurality of conduits (68) is configured to communicate cooling airflow with a bearing compartment (38).
     
    9. The gas turbine engine as recited in claim 7 or 8, wherein the plurality of conduits (68) is an oil supply line and an oil scavenge line each coupled to a bearing compartment (38).
     
    10. The gas turbine engine as recited in any of claims 7 to 9, wherein opposed walls of the inner and outer frame cases (62, 64) bound a core flow path, and the plurality of airfoils (59) are located in the core flow path and wherein the plurality of localized heat shields (60) are optionally arranged between the plurality of airfoils (59) and surfaces of the plurality of flanges (78).
     
    11. The gas turbine engine as recited in any of claims 7 to 9, wherein each heat shield (60) of the plurality of heat shields is spaced circumferentially about the inner frame case (64), and each heat shield (60) of the plurality of heat shields defines a cavity (86) between the inner frame case (64) and one of the plurality of flanges (78).
     
    12. A method of cooling a portion of a gas turbine engine (20) comprising a mid turbine frame according to any of the preceding claims, comprising:

    directing cooling airflow through the inlet (87) between the turbine first frame case (64) and the heat shield (60);

    directing cooling airflow from the inlet (87) to the cooling cavity (86), wherein the cooling cavity (86) is defined between the heat shield (60) and the flange (78), and the cavity (86) is defined between the heat shield (60) and the conduit (68) coupled to the flange (78); and

    directing cooling airflow from the cavity (86) to the outlet (88) between the conduit (68) and the heat shield (60).


     
    13. The method as recited in claim 12, wherein the heat shield (60) is arranged between an airfoil (59) receiving the conduit (68) and surfaces of the flange (78) and/or wherein the conduit (68) is coupled to a bearing compartment (38).
     


    Ansprüche

    1. Mittelturbinenrahmen (57) für ein Gasturbinentriebwerk (20), Folgendes umfassend:

    ein erstes Rahmengehäuse (64); und

    einen Flansch (78), der mit dem ersten Rahmengehäuse (64) gekoppelt ist,

    einen Hitzeschild (60), der an den Flansch (78) angrenzt und sich zwischen angrenzenden Speichen (66) befindet,

    einen Kühlungshohlraum (86), der zwischen dem Hitzeschild (60) und dem Flansch (78) definiert ist, wobei der Kühlungshohlraum (86) einen Einlass (87), der durch das erste Rahmengehäuse (64) begrenzt ist, und einen Auslass (88) aufweist, der durch eine Leitung (68) begrenzt ist, die mit dem Flansch (78) gekoppelt ist,

    wobei der Hitzeschild (60) einen ersten Abschnitt (77) und einen zweiten Abschnitt (85) beinhaltet, wobei der erste Abschnitt (77) sich in einer umlaufenden Richtung erstreckt und der zweite Abschnitt (85) sich in einer radialen Richtung erstreckt und an dem ersten Rahmengehäuse (64) anliegt,

    wobei ein Umfang (90) des Hitzeschilds (60) mit einem Umfang des Flanschs (78) zusammenpasst.


     
    2. Mittelturbinenrahmen nach einem der vorstehenden Ansprüche, wobei der Flansch (78) mindestens teilweise in dem ersten Rahmengehäuse (64) aufgenommen ist.
     
    3. Mittelturbinenrahmen nach einem der vorstehenden Ansprüche, ein zweites Rahmengehäuse (62) umfassend, das koaxial mit dem ersten Rahmengehäuse (64) ist, und wobei die Leitung (68) sich durch eine Schaufel (59) zwischen dem ersten und dem zweiten Rahmengehäuse (62, 64) erstreckt, wobei das erste und das zweite Rahmengehäuse (62, 64) gegebenenfalls einen Kernströmungspfad begrenzen und die Schaufel (59) in dem Kernströmungspfad positioniert ist.
     
    4. Mittelturbinenrahmen nach Anspruch 3, wobei die Leitung (68) einen distalen Abschnitt (98), der sich durch die Schaufel (59) erstreckt, und einen proximalen Abschnitt (99) beinhaltet, der mit dem Flansch (78) gekoppelt ist, wobei der distale Abschnitt (98) von einem isolierten Mantel (75) umgeben ist, der an dem proximalen Abschnitt (99) endet, und der Hitzeschild (60) sich um einen Abstand entlang des proximalen Abschnitts (99) erstreckt.
     
    5. Mittelturbinenrahmen nach Anspruch 3 oder 4, wobei die Leitung (68) dazu konfiguriert ist, Fluid mit einer Lagerkammer (38) zu verbinden.
     
    6. Mittelturbinenrahmen nach einem der vorstehenden Ansprüche, Folgendes umfassend:

    ein zweites Rahmengehäuse (62);

    eine Vielzahl von Speichen (66), die das erste und das zweite Rahmengehäuse (62, 64) verbindet; und

    wobei der Hitzeschild (60) in einer umlaufenden Richtung von jeder der Vielzahl von Speichen (66) beabstandet ist.


     
    7. Gasturbinentriebwerk (20), Folgendes umfassend:
    einen Mittelturbinenrahmen (57) nach einem der vorstehenden Ansprüche, der sich axial zwischen einer ersten Turbine (46) und einer zweiten Turbine (54) befindet, wobei der Mittelturbinenrahmen (57) Folgendes umfasst:

    eine Vielzahl von Schaufeln (59), die sich radial zwischen einem inneren Rahmengehäuse (64) und einem äußeren Rahmengehäuse (62) befindet;

    eine Vielzahl von Flanschen (78), die mit dem inneren Rahmengehäuse (64) gekoppelt ist;

    eine Vielzahl von Leitungen (68) zwischen dem inneren und dem äußeren Rahmengehäuse (62, 64), wobei die Vielzahl von Leitungen (68) mit der Vielzahl von Flanschen (78) gekoppelt ist; und

    eine Vielzahl von lokalisierten Hitzeschilden (60), wobei jeder einem der Vielzahl von Flanschen (78) entspricht.


     
    8. Gasturbinentriebwerk nach Anspruch 7, wobei:
    mindestens eine der Vielzahl von Leitungen (68) dazu konfiguriert ist, Kühlungsluftströmung mit einer Lagerkammer (38) zu verbinden.
     
    9. Gasturbinentriebwerk nach Anspruch 7 oder 8, wobei die Vielzahl von Leitungen (68) eine Ölzuführleitung und eine Ölrückführleitung ist, die jeweils mit einer Lagerkammer (38) gekoppelt sind.
     
    10. Gasturbinentriebwerk nach einem der Ansprüche 7 bis 9, wobei gegenüberliegende Wände des inneren und des äußeren Rahmengehäuses (62, 64) einen Kernströmungspfad begrenzen, die Vielzahl von Schaufeln (59) sich in dem Kernströmungspfad befindet und wobei die Vielzahl von lokalisierten Hitzeschilden (60) gegebenenfalls zwischen der Vielzahl von Schaufeln (59) und Flächen der Vielzahl von Flanschen (78) angeordnet ist.
     
    11. Gasturbinentriebwerk nach einem der Ansprüche 7 bis 9, wobei jeder Hitzeschild (60) der Vielzahl von Hitzeschilden umlaufend um das innere Rahmengehäuse (64) beabstandet ist und jeder Hitzeschild (60) der Vielzahl von Hitzeschilden einen Hohlraum (86) zwischen dem inneren Rahmengehäuse (64) und einem der Vielzahl von Flanschen (78) definiert.
     
    12. Verfahren zum Kühlen eines Abschnitts eines Gasturbinentriebwerks (20), das einen Mittelturbinenrahmen nach einem der vorstehenden Ansprüche umfasst, Folgendes umfassend:

    Leiten von Kühlungsluftströmung durch den Einlass (87) zwischen dem ersten Rahmengehäuse (64) der Turbine und dem Hitzeschild (60);

    Leiten von Kühlungsluftströmung von dem Einlass (87) zu dem Kühlungshohlraum (86), wobei der Kühlungshohlraum (86) zwischen dem Hitzeschild (60) und dem Flansch (78) definiert ist und wobei der Hohlraum (86) zwischen dem Hitzeschild (60) und der Leitung (68) definiert ist, die mit dem Flansch (78) gekoppelt ist; und

    Leiten von Kühlungsluftströmung von dem Hohlraum (86) zu dem Auslass (88) zwischen der Leitung (68) und dem Hitzeschild (60).


     
    13. Verfahren nach Anspruch 12, wobei der Hitzeschild (60) zwischen einer Schaufel (59), die die Leitung (68) aufnimmt, und Flächen des Flanschs (78) angeordnet ist und/oder wobei die Leitung (68) mit einer Lagerkammer (38) gekoppelt ist.
     


    Revendications

    1. Cadre de turbine moyenne (57) pour un moteur à turbine à gaz (20), comprenant :

    un premier étui de cadre (64) ; et

    une bride (78) couplée au premier étui de cadre (64),

    un bouclier thermique (60) adjacent à la bride (78) et entre des rayons adjacents (66),

    une cavité de refroidissement (86) définie entre le bouclier thermique (60) et la bride (78), la cavité de refroidissement (86) ayant une entrée (87) délimitée par le premier étui de cadre (64) et une sortie (88) délimitée par un conduit (68) couplé à la bride (78),

    dans lequel le bouclier thermique (60) comprend une première portion (77) et une deuxième portion (85), la première portion (77) s'étend dans une direction circonférentielle, la deuxième portion (85) s'étend dans une direction radiale et vient en butée contre le premier étui de cadre (64),

    dans lequel un périmètre (90) du bouclier thermique (60) s'accouple avec un périmètre de la bride (78).


     
    2. Cadre de turbine moyenne selon une quelconque revendication précédente, dans lequel la bride (78) est reçue au moins partiellement dans le premier étui de cadre (64).
     
    3. Cadre de turbine moyenne selon une quelconque revendication précédente, comprenant un deuxième étui de cadre (62) coaxial au premier étui de cadre (64), et dans lequel le conduit (68) s'étend à travers un profil aérodynamique (59) entre les premier et deuxième étuis de cadre (62, 64), dans lequel les premier et deuxième étuis de cadre (62, 64) délimitent en option un trajet d'écoulement central, et le profil aérodynamique (59) est positionné dans le trajet d'écoulement central.
     
    4. Cadre de turbine moyenne selon la revendication 3, dans lequel le conduit (68) comprend une portion distale (98) s'étendant à travers le profil aérodynamique (59) et une portion proximale (99) couplée à la bride (78), la portion distale (98) étant entourée par une gaine isolée (75) se terminant au niveau de la portion proximale (99), et le bouclier thermique (60) s'étend sur une distance le long de la portion proximale (99).
     
    5. Cadre de turbine moyenne selon la revendication 3 ou 4, dans lequel le conduit (68) est configuré pour communiquer un fluide avec un compartiment de support (38).
     
    6. Cadre de turbine moyenne selon une quelconque revendication précédente, comprenant :

    un deuxième étui de cadre (62) ;

    une pluralité de rayons (66) reliant les premier et deuxième étuis de cadre (62, 64) ; et

    dans lequel le bouclier thermique (60) est espacé dans une direction circonférentielle de chacun de la pluralité de rayons (66).


     
    7. Moteur à turbine à gaz (20), comprenant :
    un cadre de turbine moyenne (57) selon une quelconque revendication précédente axialement entre une première turbine (46) et une deuxième turbine (54), le cadre de turbine moyenne (57) comprenant :

    une pluralité de profils aérodynamiques (59) radialement entre un étui de cadre intérieur (64) et un étui de cadre extérieur (62) ;

    une pluralité de brides (78) couplées à l'étui de cadre intérieur (64) ;

    une pluralité de conduits (68) entre les étuis de cadre intérieur et extérieur (62, 64), la pluralité de conduits (68) étant couplée à la pluralité de brides (78) ; et

    une pluralité de boucliers thermiques localisés (60) correspondant chacun à une de la pluralité de brides (78).


     
    8. Moteur à turbine à gaz selon la revendication 7, dans lequel :
    au moins un de la pluralité de conduits (68) est configuré pour communiquer le flux d'air de refroidissement avec un compartiment de support (38).
     
    9. Moteur à turbine à gaz selon la revendication 7 ou 8, dans lequel la pluralité de conduits (68) est une ligne d'alimentation en huile et une ligne de récupération d'huile chacune couplée à un compartiment de support (38).
     
    10. Moteur à turbine à gaz selon l'une quelconque des revendications 7 à 9, dans lequel des parois opposées des étuis de cadre intérieur et extérieur (62, 64) délimitent un trajet d'écoulement central, et la pluralité de profils aérodynamiques (59) sont localisés dans le trajet d'écoulement central et dans lequel la pluralité de boucliers thermiques localisés (60) sont agencés en option entre la pluralité de profils aérodynamiques (59) et des surfaces de la pluralité de brides (78).
     
    11. Moteur à turbine à gaz selon l'une quelconque des revendications 7 à 9, dans lequel chaque bouclier thermique (60) de la pluralité de boucliers thermiques est espacé circonférentiellement autour de l'étui de cadre intérieur (64), et chaque bouclier thermique (60) de la pluralité de boucliers thermiques définit une cavité (86) entre l'étui de cadre intérieur (64) et une de la pluralité de brides (78).
     
    12. Procédé de refroidissement d'une portion d'un moteur à turbine à gaz (20) comprenant un cadre de turbine moyenne selon l'une quelconque des revendications précédentes, comprenant :

    la direction d'un flux d'air de refroidissement à travers l'entrée (87) entre le premier étui de cadre de turbine (64) et le bouclier thermique (60) ;

    la direction d'un flux d'air de refroidissement de l'entrée (87) à la cavité de refroidissement (86), dans lequel la cavité de refroidissement (86) est définie entre le bouclier thermique (60) et la bride (78), et la cavité (86) est définie entre le bouclier thermique (60) et le conduit (68) couplé à la bride (78) ; et

    la direction d'un flux d'air de refroidissement de la cavité (86) à la sortie (88) entre le conduit (68) et le bouclier thermique (60).


     
    13. Procédé selon la revendication 12, dans lequel le bouclier thermique (60) est agencé entre un profil aérodynamique (59) recevant le conduit (68) et des surfaces de la bride (78) et/ou dans lequel le conduit (68) est couplé à un compartiment de support (38).
     




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    Cited references

    REFERENCES CITED IN THE DESCRIPTION



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    Patent documents cited in the description