(19)
(11)EP 3 102 808 B1

(12)EUROPEAN PATENT SPECIFICATION

(45)Mention of the grant of the patent:
06.05.2020 Bulletin 2020/19

(21)Application number: 15755240.7

(22)Date of filing:  29.01.2015
(51)Int. Cl.: 
F02C 7/18  (2006.01)
F01D 25/12  (2006.01)
(86)International application number:
PCT/US2015/013472
(87)International publication number:
WO 2015/130425 (03.09.2015 Gazette  2015/35)

(54)

GAS TURBINE ENGINE WITH COOLING FLUID COMPOSITE TUBE

GASTURBINE MIT VERBUNDROHR FÜR KÜHLUNG

TUBE COMPOSITE DE FLUIDE DE REFROIDISSEMENT D'UNE TURBINE À GAZ


(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: 03.02.2014 US 201461935109 P

(43)Date of publication of application:
14.12.2016 Bulletin 2016/50

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

(72)Inventors:
  • HARRIS, Meggan
    Colchester, Connecticut 06415 (US)
  • WILLIAMS, Alexander W.
    Santa Clara, California 95054 (US)
  • SWENSON-DODGE, Sheree R.
    Lebanon, Connecticut 06249 (US)

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


(56)References cited: : 
EP-A2- 2 584 152
US-A- 4 314 794
US-A1- 2005 056 020
US-A1- 2011 070 095
US-B1- 6 200 092
US-A- 4 167 097
US-A1- 2005 056 020
US-A1- 2008 199 661
US-A1- 2013 064 681
US-B1- 7 713 029
  
      
    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 a turbine engine comprising ceramic-based composite tube for transferring a cooling fluid within a hot environment of the gas turbine engine.

    [0002] One type of gas turbine engine includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.

    [0003] In one example hot gas turbine engine environment, exit vanes are provided at the end of the compressor section before the combustor section. Turbine vanes are provided at the end of the combustor section and provide the first fixed stage of the turbine section. This area of the engine experiences some of the hottest temperatures within the engine. Cooling fluid is provided to this area to reduce component temperatures and increase durability of the engine. Cooling fluid is typically provided to various locations throughout the engine using nickel alloy tubing, such as INCONEL 718.

    [0004] A turbine engine comprising the features of the preamble of claim 1 is disclosed in US4167097.

    SUMMARY



    [0005] According to the invention, there is a turbine engine as set forth in claim 1.

    [0006] In an embodiment of the above, the inner and outer case structures provide a gas flow path. The cooling tube is in the gas flow path.

    [0007] In a further embodiment of any of the above, the cooling source is compressor bleed air.

    [0008] In a further embodiment of any of the above, the second case structure is a bearing compartment. The cooling tube is configured to provide a lubricant to the bearing compartment.

    [0009] In a further embodiment of the above, the cooling source includes the lubricant.

    [0010] In a further embodiment of any of the above, the cooling tube is a ceramic matrix composite or an organic matrix composite.

    [0011] In a further embodiment of any of the above, the cooling tube is a non-metallic structure free from insulation.

    [0012] In a further embodiment of any of the above, the cooling tube has opposing ends. A retaining feature is arranged at one of the ends.

    [0013] In a further embodiment of the above, the retaining feature is a collar that has an outer dimension that is greater than an intermediate portion of the cooling tube provided between the ends.

    [0014] In a further embodiment of the above, a retainer cooperates with the collar to secure the cooling tube to one of the first and second structures.

    [0015] In a further embodiment of any of the above, the cooling tube has a circular cross-section.

    [0016] In an alternative embodiment, the cooling tube has a non-circular cross-section.

    BRIEF DESCRIPTION OF THE DRAWINGS



    [0017] The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:

    Figure 1 is a schematic view of an example gas turbine engine including a combustor.

    Figure 2 is an enlarged schematic view of a gas turbine engine in the area of the combustor.

    Figure 3 is a schematic view of a ceramic-based composite cooling tube.

    Figure 4 is a first example cross-section of the cooling tube.

    Figure 5 is a second example cross-section of the cooling tube.

    Figure 6 is a schematic view of a hot engine environment that includes the cooling tube.

    Figure 7 is a schematic view of an example cooling tube retaining feature.



    [0018] The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.

    DETAILED DESCRIPTION



    [0019] 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 augmenter 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.

    [0020] The commercial-style gas turbine engine shown in Figure 1 is exemplary only. The disclosed cooling fluid composite tube may be used for any type of engine, including military and industrial gas turbine engines.

    [0021] 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.

    [0022] 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.

    [0023] 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.

    [0024] 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.

    [0025] 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 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), 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 1150 ft / second (350.5 meters/second).

    [0026] One example hot engine environment is shown in Figure 2. An area of the combustor section 26 is shown in more detail in Figure 2. The combustor section 26 includes a combustor 56 having a combustor housing 60. An injector 62 is arranged at a forward end of the combustor housing 60 and is configured to provide fuel to the combustor housing 60 where it is ignited to produce hot gases that expand through the turbine section 54.

    [0027] A diffuser case 64 is secured to the combustor housing 60 and fixed relative to the engine static structure 36. Exit vanes 66 are arranged downstream from the compressor section 52 and upstream from the combustor section 26. A fluid source 68, such as bleed air from a compressor stage, provides cooling fluid F through a ceramic-based cooling tube T that extends through the exit vanes 66 to various locations within the gas turbine engine 20. Other fluids may be provided, such as diffuser air or cooled air, for example. A circumferential array of exit vanes 66 are used, which include a corresponding circumferential array of cooling tubes.

    [0028] The compressor section 52 includes a compressor rotor 70 supported for rotation relative to the engine static structure 36. The turbine section 54 includes a turbine rotor 76 arranged downstream from a tangential onboard injector (TOBI) module 78. The TOBI module 78 supports a circumferential array of vanes 82 that are arranged upstream from the turbine rotor 76. The vanes 82 provide the first fixed stator stage of the turbine section 54.

    [0029] The fluid F is distributed to various locations within the gas turbine engine 20 for a variety of uses. Because the cooling tube T is provided in a hot environment with extreme temperatures, the cooling fluid F within the cooling tube T may become heated from the surrounding environment to where the effectiveness of the cooling fluid F is significantly diminished. To this end, it is desirable to provide a ceramic-based composite cooling tube T that is light weight, rather than, for example, a heavier metallic tube with insulation. However, the ceramic-based composite may be used as a heat shield for one or more metallic tubes that are arranged internally or externally with respect to the composite. The tube may also include sealing features.

    [0030] An example cooling tube T is shown in Figures 3 and 4. The cooling tube T is a ceramic-based composite such as organic matrix composite (OMC) or ceramic matrix composite (CMC). Layers of the ceramic-based composite material can be layed up on a mandrel M to provide a desired length, shape and cross-section. The cross-section may be circular (Figure 4) or non-circular (Figure 5), for example, an elliptical shaped cooling tube T'.

    [0031] Referring to Figure 6, a cooling arrangement for a gas turbine engine schematically illustrates a cooling fluid F from a cooling source. First and second structures 108, 110, which are outer and inner case structures are spaced apart from one another. The cooling tube T is used provide a fluid connection between the first and second structures 108, 110, such that the cooling tube T transfers the cooling fluid F from the cooling source to the second structure 110.

    [0032] The inner and outer case structures may be arranged in a compressor section, such as that depicted in Figure 2. In that arrangement, the exit vanes 66 interconnect the inner and outer case structures, and the cooling tube T is disposed within the exit vane 66. The inner and outer case structures provide a gas flow path, and the cooling tube is arranged, indirectly, in the gas flow path, exposing the cooling fluid F to high heat as it travels through the exit vane 66. The ceramic-based composite material keeps the cooling fluid F at suitably low temperatures.

    [0033] Returning to Figure 6, the second case structure 110 may be a bearing compartment. The cooling source includes the lubricant, and the cooling tube T is configured to provide a lubricant to the bearing compartment.

    [0034] Referring to Figures 3 and 7, the cooling tube T has opposing ends. A retaining feature 102, 104, such as a collar, is arranged at at least one of the ends, and in the example, at both ends. The collar has an outer dimension that is greater than an intermediate portion 100 of the cooling tube T provided between the ends 102, 104. A retainer 112 cooperates with the collar to secure the cooling tube T to one of the first and second structures, for example, the first structure 108.

    [0035] It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 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 invention.

    [0036] Although the different examples have specific components shown in the illustrations, embodiments of this invention 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.

    [0037] Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.


    Claims

    1. A turbine engine (20) comprising:

    a compressor section (24);

    a combustor section (26); and

    a cooling arrangement, the cooling arrangement comprising:

    a cooling source;

    first and second structures (108, 110); and

    a cooling tube (T) fluidly providing a fluid connection between the first and second structures (108, 110), the cooling tube (T) configured to transfer a cooling fluid (F) from the cooling source to the second structure (110); wherein

    the first structure (108) is an outer case structure, and the second structure (110) is an inner case structure;

    the engine further comprises an exit vane (66) interconnecting the inner and outer case structures (108, 110) and arranged downstream from the compressor section (24) and upstream from the combustor section (26); and

    the cooling tube (T) is disposed within the exit vane (66), characterized in that:
    the cooling tube (T) is a ceramic-based composite cooling tube (T).


     
    2. The turbine engine according to claim 1, wherein the inner and outer case structures (106, 108) provide a gas flow path, the cooling tube (T) in the gas flow path.
     
    3. The turbine engine according to claim 1 or 2, wherein the cooling source is compressor bleed air.
     
    4. The turbine engine according to claim 1 or 2, wherein the second structure (110) is a bearing compartment, and the cooling tube (T) is configured to provide a lubricant to the bearing compartment.
     
    5. The turbine engine according to claim 4, wherein the cooling source includes the lubricant.
     
    6. The turbine engine according to any preceding claim, wherein the cooling tube (T) is a ceramic matrix composite or an organic matrix composite.
     
    7. The turbine engine according to claim 6, wherein the cooling tube (T) is a non-metallic structure free from insulation.
     
    8. The turbine engine according to any preceding claim, wherein the cooling tube (T) has opposing ends, and a retaining feature (102, 104) is arranged at at least one of the ends.
     
    9. The turbine engine according to claim 8, wherein the retaining feature (102, 104) is a collar that has an outer dimension that is greater than an intermediate portion (100) of the cooling tube (T) provided between the ends.
     
    10. The turbine engine according to claim 9, comprising a retainer (112) cooperating with the collar to secure the cooling tube (T) to one of the first and second structures (108, 110).
     
    11. The turbine engine according to any preceding claim, wherein the cooling tube (T) has a circular cross-section.
     
    12. The turbine engine according to any of claims 1 to 10, wherein the cooling tube (T) has a non-circular cross-section.
     
    13. The turbine engine according to any preceding claim, wherein the cooling tube (T) acts as a heatshield for one or more internal or external metallic tubes.
     


    Ansprüche

    1. Turbinentriebwerk (20), umfassend:

    einen Verdichterabschnitt (24);

    einen Brennkammerabschnitt (26); und

    eine Kühlanordnung, wobei die Kühlanordnung Folgendes umfasst:

    eine Kühlquelle;

    eine erste und eine zweite Struktur (108, 110); und

    ein Kühlrohr (T), das fluidmäßig eine Fluidverbindung zwischen der ersten und der zweiten Struktur (108, 110) bereitstellt, wobei das Kühlrohr (T) dazu konfiguriert ist, ein Kühlfluid (F) von der Kühlquelle zur zweiten Struktur (110) zu übertragen; wobei

    die erste Struktur (108) eine Außengehäusestruktur ist und die zweite Struktur (110) eine Innengehäusestruktur ist;

    das Triebwerk ferner eine Austrittsleitschaufel (66) umfasst, die die Innen- und die Außengehäusestruktur (108, 110) verbindet und stromabwärts des Verdichterabschnitts (24) und stromaufwärts des Brennkammerabschnitts (26) angeordnet ist; und

    das Kühlrohr (T) innerhalb der Austrittsleitschaufel (66) angeordnet ist, dadurch gekennzeichnet, dass:
    das Kühlrohr (T) ein Verbundkühlrohr auf Keramikbasis (T) ist.


     
    2. Turbinentriebwerk nach Anspruch 1, wobei die Innen- und die Außengehäusestruktur (106, 108) einen Gasströmungspfad bereitstellen, wobei sich das Kühlrohr (T) in dem Gasströmungspfad befindet.
     
    3. Turbinentriebwerk nach Anspruch 1 oder 2, wobei die Kühlquelle Verdichterzapfluft ist.
     
    4. Turbinentriebwerk nach Anspruch 1 oder 2, wobei die zweite Struktur (110) eine Lagerkammer ist und das Kühlrohr (T) dazu konfiguriert ist, ein Schmiermittel zu der Lagerkammer zu liefern.
     
    5. Turbinentriebwerk nach Anspruch 4, wobei die Kühlquelle das Schmiermittel beinhaltet.
     
    6. Turbinentriebwerk nach einem der vorstehenden Ansprüche, wobei das Kühlrohr (T) ein keramischer Matrixverbundstoff oder ein organischer Matrixverbundstoff ist.
     
    7. Turbinentriebwerk nach Anspruch 6, wobei das Kühlrohr (T) eine nichtmetallische Struktur ohne Isolierung ist.
     
    8. Turbinentriebwerk nach einem der vorstehenden Ansprüche, wobei das Kühlrohr (T) gegenüberliegende Enden aufweist und ein Haltemerkmal (102, 104) an mindestens einem der Enden angeordnet ist.
     
    9. Turbinentriebwerk nach Anspruch 8, wobei das Haltemerkmal (102, 104) ein Bund ist, der einen Außendurchmesser aufweist, der größer als ein Zwischenabschnitt (100) des Kühlrohrs (T), der zwischen den Enden bereitgestellt ist, ist.
     
    10. Turbinentriebwerk nach Anspruch 9, umfassend eine Halterung (112), die mit dem Bund zusammenwirkt, um das Kühlrohr (T) an einem aus der ersten und der zweiten Struktur (108, 110) zu fixieren.
     
    11. Turbinentriebwerk nach einem der vorstehenden Ansprüche, wobei das Kühlrohr (T) einen kreisförmigen Querschnitt aufweist.
     
    12. Turbinentriebwerk nach einem der Ansprüche 1 bis 10, wobei das Kühlrohr (T) einen nicht kreisförmigen Querschnitt aufweist.
     
    13. Turbinentriebwerk nach einem der vorstehenden Ansprüche, wobei das Kühlrohr (T) als ein Hitzeschild für eines oder mehrere innere oder äußere Metallrohre wirkt.
     


    Revendications

    1. Turbine à gaz (20) comprenant :

    une section de compresseur (24) ;

    une section de chambre de combustion (26) ; et

    un agencement de refroidissement, l'agencement de refroidissement comprenant :

    une source de refroidissement ;

    des première et seconde structures (108, 110) ; et

    un tube de refroidissement (T) fournissant fluidiquement une liaison de fluide entre les première et seconde structures (108, 110), le tube de refroidissement (T) étant configuré pour transférer un fluide de refroidissement (F) de la source de refroidissement à la seconde structure (110) ; dans laquelle

    la première structure (108) est une structure de boîtier externe et la seconde structure (110) est une structure de boîtier interne ;

    le moteur comprend en outre une aube de sortie (66) reliant les structures de boîtier interne et externe (108, 110) et disposée en aval de la section de compresseur (24) et en amont de la section de chambre de combustion (26) ; et

    le tube de refroidissement (T) est disposé à l'intérieur de l'aube de sortie (66), caractérisée en ce que :
    le tube de refroidissement (T) est un tube de refroidissement composite à base de céramique (T).


     
    2. Turbine à gaz selon la revendication 1, dans laquelle les structures de boîtier interne et externe (106, 108) fournissent un chemin d'écoulement de gaz, le tube de refroidissement (T) dans le chemin d'écoulement de gaz.
     
    3. Turbine à gaz selon la revendication 1 ou 2, dans laquelle la source de refroidissement est de l'air de prélèvement du compresseur.
     
    4. Turbine à gaz selon la revendication 1 ou 2, dans laquelle la seconde structure (110) est un compartiment de palier, et le tube de refroidissement (T) est configuré pour fournir un lubrifiant au compartiment de palier.
     
    5. Turbine à gaz selon la revendication 4, dans laquelle la source de refroidissement comporte le lubrifiant.
     
    6. Turbine à gaz selon une quelconque revendication précédente, dans laquelle le tube de refroidissement (T) est un composite à matrice céramique ou un composite à matrice organique.
     
    7. Turbine à gaz selon la revendication 6, dans lequel le tube de refroidissement (T) est une structure non métallique exempte d'isolation.
     
    8. Turbine à gaz selon une quelconque revendication précédente, dans laquelle le tube de refroidissement (T) a des extrémités opposées, et un élément de retenue (102, 104) est disposé au niveau d'au moins une des extrémités.
     
    9. Turbine à gaz selon la revendication 8, dans laquelle l'élément de retenue (102, 104) est un collier qui a une dimension externe qui est supérieure à une partie intermédiaire (100) du tube de refroidissement (T) fournie entre les extrémités.
     
    10. Turbine à gaz selon la revendication 9, comprenant un dispositif de retenue (112) coopérant avec le collier pour fixer le tube de refroidissement (T) à l'une des première et seconde structures (108, 110).
     
    11. Turbine à gaz selon une quelconque revendication précédente, dans laquelle le tube de refroidissement (T) a une section transversale circulaire.
     
    12. Turbine à gaz selon l'une quelconque des revendications 1 à 10, dans laquelle le tube de refroidissement (T) a une section transversale non circulaire.
     
    13. Turbine à gaz selon une quelconque revendication précédente, dans laquelle le tube de refroidissement (T) agit comme un écran thermique pour un ou plusieurs tubes métalliques internes ou externes.
     




    Drawing












    REFERENCES CITED IN THE DESCRIPTION



    This list of references cited by the applicant is for the reader's convenience only. It does not form part of the European patent document. Even though great care has been taken in compiling the references, errors or omissions cannot be excluded and the EPO disclaims all liability in this regard.

    Patent documents cited in the description