(19)
(11)EP 3 159 482 B1

(12)EUROPEAN PATENT SPECIFICATION

(45)Mention of the grant of the patent:
30.03.2022 Bulletin 2022/13

(21)Application number: 16194110.9

(22)Date of filing:  17.10.2016
(51)International Patent Classification (IPC): 
F01D 5/18(2006.01)
(52)Cooperative Patent Classification (CPC):
F01D 5/187; F05D 2240/81

(54)

BLADE ASSEMBLY , CORRESPONDING ROTOR ASSEMBLY AND GAS TURBINE ENGINE

SCHAUFELANORDNUNG, ZUGEHÖRIGE ROTORANORDNUNG UND GASTURBINENKRAFTWERK

AGENCEMENT D'AUBE, AGENCEMENT DE ROTOR ET MOTEUR À TURBINE À GAZ ASSOCIÉS


(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: 19.10.2015 US 201514886201

(43)Date of publication of application:
26.04.2017 Bulletin 2017/17

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

(72)Inventors:
  • ENGLEHART, Joseph F.
    Gastonia, NC 82056-8581 (US)
  • PIETRASZKIEWICZ, Edward F.
    Southington, CT 06489 (US)
  • CHLUS, Wieslaw A.
    Wethersfield, CT 06109 (US)
  • KONOPKA, David M.
    Stuart, FL 34997 (US)
  • HMIEL, Luke A.
    Tequesta, FL 33469 (US)
  • BOUCHER, Kenneth
    Branford, CT 06405 (US)

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


(56)References cited: : 
EP-A1- 2 037 081
US-A1- 2006 093 484
EP-A2- 2 228 518
  
      
    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 gas turbine engines, and more particularly to thermal management of turbine components of gas turbine engines. This invention relates to a blade assembly, a rotor assembly for a gas turbine engine and a gas turbine engine.

    [0002] Gas turbines hot section components, in particular turbine vanes and blades in the turbine section of the gas turbine are configured for use within particular temperature ranges. Such components often rely on cooling airflow to maintain turbine components within this particular temperature range. For example, stationary turbine vanes often have internal passages for cooling airflow to flow through, and additionally may have openings in an outer surface of the vane for cooling airflow to exit the interior of the vane structure and form a cooling film of air over the outer surface to provide the necessary thermal conditioning. Other components of the turbine often also require such thermal conditioning to reduce thermal gradients that would otherwise be present in the structure and which are generally undesirable. Thus, ways to increase thermal conditioning capability in the turbine are desired.

    [0003] Means for providing cooling of a blade platform, in which cooling passages are provided in gussets between the platform and the blade, are disclosed in EP 2228518, US 2006/0093484 and EP 2037081.

    SUMMARY



    [0004] In one aspect of the invention, a blade assembly for a gas turbine engine is provided as claimed in claim 1.

    [0005] Optionally, a blade air passage is disposed at the blade, the gusset air passage connecting the blade air passage to the platform air passage.

    [0006] Optionally, the platform air passage is configured to convey the airflow toward the platform leading edge from a platform air passage entrance.

    [0007] Optionally, the gusset is located at about midchord of the blade.

    [0008] Optionally, the gusset is located at a pressure side of the blade.

    [0009] Optionally, the gusset air passage is formed integral with the platform air passage.

    [0010] Optionally, the gusset air passage is formed by casting.

    [0011] Optionally, a rotor assembly for a gas turbine engine includes a rotor disc and a plurality of rotor blades extending radially outwardly from the rotor disc. Each rotor blade is the blade assembly.

    [0012] Optionally, the rotor assembly is a turbine rotor assembly.

    [0013] In yet another aspect of the invention, a gas turbine engine includes a combustor and a plurality of gas turbine engine components located in fluid communication with the combustor. The gas turbine engine component includes an airfoil portion and a platform secured to the airfoil portion. The platform includes a platform air passage positioned therein. A gusset extends from the airfoil portion to the platform. The gusset includes a gusset air passage fluidly connected to the platform air passage to convey an airflow to the platform air passage. The gas turbine engine is characterized in that the platform air passage is configured to direct the airflow in an axially upstream direction toward a platform leading edge, then axially rearwardly toward a platform trailing edge, before exiting the platform air passage at a platform air passage exit, wherein the platform air passage has a spiral shape.

    [0014] Optionally, an air passage is disposed at the airfoil portion, the gusset air passage connecting the airfoil portion air passage to the platform air passage.

    [0015] Optionally, the platform air passage is configured to convey the airflow toward the platform leading edge from a platform air passage entrance.

    [0016] Optionally, the gusset is located at about midchord of the airfoil portion.

    [0017] Optionally, the gusset is located at a pressure side of the airfoil portion.

    [0018] Optionally, the gusset air passage is formed integral with the platform air passage.

    BRIEF DESCRIPTION OF THE DRAWINGS



    [0019] The subject matter which is regarded as the present disclosure is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:

    FIG. 1 is a schematic illustration of a gas turbine engine;

    FIG. 2 is a schematic illustration of an embodiment of a turbine rotor assembly;



    [0020] The gas turbine engine 10 further comprises a turbine section 20 for extracting energy from the combustion gases. Fuel is injected into the combustor 18 of the gas turbine engine 10 for mixing with the compressed air from the compressor 16 and ignition of the resultant mixture. The fan 12, compressor 16, combustor 18, and turbine 20 are typically all concentric about a common central longitudinal axis of the gas turbine engine 10.

    [0021] The gas turbine engine 10 may further comprise a low pressure compressor located upstream of a high pressure compressor and a high pressure turbine located upstream of a low pressure turbine. For example, the compressor 16 may be a multi-stage compressor 16 that has a low-pressure compressor and a high-pressure compressor and the turbine 20 may be a multistage turbine 20 that has a high-pressure turbine and a low-pressure turbine. In one embodiment, the low-pressure compressor is connected to the low-pressure turbine and the high pressure compressor is connected to the high-pressure turbine.

    [0022] The turbine 20 includes one or more sets, or stages, of fixed turbine vanes 22 and turbine rotors 24, each turbine rotor 24 including a plurality of turbine blades 26 (shown in FIG. 2). The turbine vanes 22 and the turbine blades 26 utilize a cooling airflow to maintain the turbine components within a desired temperature range. In some embodiments, the cooling airflow may flow internal through the turbine components to cool the components internally, while in other embodiments, the cooling airflow is utilized to form a cooling film on exterior surfaces of the components.

    [0023] FIG. 2 illustrates an example of a turbine rotor 24 structure in more detail. While the present description regards a turbine rotor 24 and turbine blades 26, it is to be appreciated that the present disclosure may be readily adapted to turbine vanes 22 and compressor 16 components. The turbine rotor 24 includes a turbine disc 28 having a disc rim 30 to which a plurality of radially-extending turbine blades 26 are mounted. Each turbine blade 26 includes an airfoil portion 32 extending from a blade platform 34. As shown in FIG. 3, a blade root 36 extends radially inboard of the blade platform 34 and is inserted into a complimentary slot 38 or other opening in the disc rim 30 to mount the turbine blade 26 to the turbine disc 28. The turbine blade 26 may be anchored in place in the turbine disc 28 by bolts, rivets, or other mechanical fastening arrangements.

    [0024] Referring now to FIG. 4, shown is a cross-sectional view of a turbine blade 26. The turbine blade 26 includes a pressure side 40 and a suction side 42, with a blade cavity 44 located between the pressure side 40 and the suction side 42 and extending along a spanwise direction 46 of the turbine blade 26 from the blade platform 34 toward a blade tip (shown in FIG. 2). A gusset 48 extends from the blade platform 34 toward the turbine blade 26, in some embodiments at the pressure side 40 of the turbine blade 26 and at a radially inboard side 50 of the blade platform 34. In some embodiments, the gusset 48 is located at about mid-chord of the turbine blade 26. The gusset 48 supports the blade platform 34 and reacts centrifugal loading on the blade platform 34, and further reduces bending stresses at the blade platform 34. It is to be appreciated that while shown at an approximately mid-chord location, the gusset 48 may be positioned at other selected locations along the turbine blade 26.

    [0025] Referring now to FIG. 5, a cross-sectional view of the turbine blade 26 through the gusset 48 is illustrated. The blade platform 34 includes a platform air passage 52, which is connected to the blade cavity 44 via a gusset air passage 54 extending through the gusset 48. The gusset air passage 54 allows for diversion of a portion of blade cooling airflow 56 from the blade cavity 44 to the platform air passage 52 to cool the blade platform 34 via, in some embodiments, a plurality of platform openings (not shown) in the blade platform forming a cooling film on the platform 34. In the embodiment of FIG. 5, the platform air passage 52 and the gusset air passage 54 may be formed concurrently with the manufacture of the turbine blade 26 by, for example, a casting process. In an alternative embodiment, illustrated in FIG. 6, the gusset air passage 54 is formed in a secondary process after formation of the turbine blade 26. For example, the gusset air passage 54 may be formed by a drilling operation, after which an entry opening 56 at the turbine blade 26 is closed via, for example, welding. It is to be appreciated that while the gusset air passage 54 is described herein as being located at the turbine blade 26, turbine vanes 22 may utilize gusset air passages 54 to cool platforms of turbine vanes 22. Further, while a single gusset 48 and gusset air passage 54 are shown, embodiments of turbine blades 26 or turbine vanes 22 may include two or more gussets 48 and/or two or more gusset air passages 54. In some embodiments, the gusset air passage 54 is circular in cross-section, while in other embodiments, other cross-sectional shapes such as elliptical or oval, may be utilized.

    [0026] Referring now to FIG. 7, the platform air passage 52 may take one of a variety of shapes. The platform air passage 52 is configured to direct the cooling flow 56 into a platform air passage entrance 58 and then in an axially upstream direction 60, relative to a general airflow direction through the turbine section 20 toward a platform leading edge 62. The cooling airflow 56 then flows axially rearwardly toward a platform trailing edge 64 before exiting the platform air passage 52 at a platform air passage exit 66. The platform air passage 52 has a spiral shape. Directing the cooling airflow 56 forward, then rearwardly, directs the highest pressure cooling airflow 56 at the platform leading edge 62, prior to pressure losses degrading the cooling effectiveness of the airflow as it flows rearwardly. In alternative embodiments, the gusset 48 and gusset air passage 54 may be located at or near a platform leading edge 62, with the cooling airflow 56 directed rearwardly along the platform air passage 52.

    [0027] While the present disclosure has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the present disclosure is not limited to such disclosed embodiments. Accordingly, the present invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.


    Claims

    1. A blade assembly for a gas turbine engine, comprising:

    a blade (26);

    a blade platform (34) secured to the blade (26), the blade (26) extending radially outwardly from the blade platform (34), the blade platform (34) including a platform air passage (52) disposed therein; and

    a gusset (48) extending from the blade (26) to the blade platform (34), the gusset (48) including a gusset air passage (54) fluidly connected to the platform air passage (52) to convey an airflow (56) to the platform air passage (52), wherein the platform air passage (52) is configured to direct the airflow (56) in an axially upstream direction toward a platform leading edge (62), then axially rearwardly toward a platform trailing edge (64), before exiting the platform air passage (52) at a platform air passage exit (66),

    characterized in that:
    the platform air passage (52) has a spiral shape.
     
    2. The blade assembly of claim 1, further comprising a blade air passage (44) disposed at the blade (26), the gusset air passage (54) connecting the blade air passage (44) to the platform air passage (52).
     
    3. The blade assembly of claim 1 or claim 2, wherein the platform air passage (52) is configured to convey the airflow (56) toward the platform leading edge (62) from a platform air passage entrance (58).
     
    4. The blade assembly of any preceding claim, wherein the gusset (48) is disposed at about midchord of the blade (26).
     
    5. The blade assembly of any preceding claim, wherein the gusset (48) is disposed at a pressure side of the blade (26).
     
    6. The blade assembly of any preceding claim, wherein the gusset air passage (54) is formed integral with the platform air passage (52).
     
    7. The blade assembly of claim 6, wherein the gusset air passage (54) is formed by casting.
     
    8. A rotor assembly (24) for a gas turbine engine, comprising:

    a rotor disc (28); and

    a plurality of rotor blades (26) extending radially outwardly from the rotor disc (26), each rotor blade being a blade assembly (26) as claimed in any preceding claim.


     
    9. The rotor assembly of claim 8, wherein rotor assembly (24) is a turbine rotor assembly (24).
     
    10. A gas turbine engine (10), comprising:

    a combustor (18); and

    a plurality of gas turbine engine components disposed in fluid communication with the combustor (18), including:

    an airfoil portion (26);

    a platform (34) secured to the airfoil portion (26), the platform (34) including a platform air passage (52) disposed therein; and

    a gusset (48) extending from the airfoil portion (26) to the platform (34), the gusset (48) including a gusset air passage (54) fluidly connected to the platform air passage (52) to convey an airflow (56) to the platform air passage (52), wherein the platform air passage (52) is configured to direct the airflow (56) in an axially upstream direction toward a platform leading edge (62), then axially rearwardly toward a platform trailing edge (64), before exiting the platform air passage (52) at a platform air passage exit (66),

    characterized in that:
    the platform air passage (52) has a spiral shape.
     
    11. The gas turbine engine of claim 10, further comprising an air passage (44) disposed at the airfoil portion (26), the gusset air passage (54) connecting the airfoil portion air passage (44) to the platform air passage (52).
     
    12. The gas turbine engine of claim 10 or claim 11, wherein the platform air passage (52) is configured to convey the airflow toward the platform leading edge (62) from a platform air passage entrance (58).
     
    13. The gas turbine engine of any of claims 10 to 12, wherein the gusset (48) is disposed at about midchord of the airfoil portion (26).
     
    14. The gas turbine engine of any of claims 10 to 13, wherein the gusset (48) is disposed at a pressure side of the airfoil portion (26).
     
    15. The gas turbine engine of any of claims 10 to 14, wherein the gusset air passage (54) is formed integral with the platform air passage (52).
     


    Ansprüche

    1. Schaufelanordnung für ein Gasturbinenkraftwerk, umfassend:

    eine Schaufel (26);

    eine Schaufelplattform (34), die an der Schaufel (26) gesichert ist, wobei sich die Schaufel (26) radial auswärts von der Schaufelplattform (34) erstreckt, wobei die Schaufelplattform (34) einen Plattform-Luftdurchgang (52) beinhaltet, der darin angeordnet ist; und

    ein Winkelstück (48), das sich von der Schaufel (26) zu der Schaufelplattform (34) erstreckt, wobei das Winkelstück (48) einen Winkelstück-Luftdurchgang (54) beinhaltet, der mit dem Plattform-Luftdurchgang (52) fluidverbunden ist, um einen Luftstrom (56) zu dem Plattform-Luftdurchgang (52) zu transportieren, wobei der Plattform-Luftdurchgang (52) dazu konfiguriert ist, den Luftstrom (56) in einer axial stromaufwärtigen Richtung zu einer Plattformanströmkante (62) zu leiten, dann axial nach hinten zu einer Plattformabströmkante (64), bevor er an einem Plattform-Luftdurchgangsausgang (66) aus dem Plattform-Luftdurchgang (52) austritt,

    dadurch gekennzeichnet, dass:
    der Plattform-Luftdurchgang (52) eine Spiralform aufweist.


     
    2. Schaufelanordnung nach Anspruch 1, ferner umfassend einen Schaufel-Luftdurchgang (44), der an der Schaufel (26) angeordnet ist, wobei der Winkelstück-Luftdurchgang (54) den Schaufel-Luftdurchgang (44) mit dem Plattform-Luftdurchgang (52) verbindet.
     
    3. Schaufelanordnung nach Anspruch 1 oder Anspruch 2, wobei der Plattform-Luftdurchgang (52) dazu konfiguriert ist, den Luftstrom (56) von einem Plattform-Luftdurchgangeingang (58) zu der Plattformanströmkante (62) zu transportieren.
     
    4. Schaufelanordnung nach einem der vorstehenden Ansprüche, wobei das Winkelstück (48) an ungefähr einer Mittelsehne der Schaufel (26) angeordnet ist.
     
    5. Schaufelanordnung nach einem der vorstehenden Ansprüche, wobei das Winkelstück (48) an einer Druckseite der Schaufel (26) angeordnet ist.
     
    6. Schaufelanordnung nach einem der vorstehenden Ansprüche, wobei der Winkelstück-Luftdurchgang (54) einstückig mit dem Plattform-Luftdurchgang (52) ausgebildet ist.
     
    7. Schaufelanordnung nach Anspruch 6, wobei der Winkelstück-Luftdurchgang (54) durch Gießen ausgebildet ist.
     
    8. Rotoranordnung (24) für ein Gasturbinenkraftwerk, umfassend:

    eine Rotorscheibe (28); und

    eine Vielzahl von Rotorschaufeln (26), die sich radial auswärts von der Rotorscheibe (26) erstreckt, wobei jede Rotorschaufel eine Schaufelanordnung (26) nach einem der vorstehenden Ansprüche ist.


     
    9. Rotoranordnung nach Anspruch 8, wobei die Rotoranordnung (24) eine Turbinenrotoranordnung (24) ist.
     
    10. Gasturbinenkraftwerk (10), umfassend:

    eine Brennkammer (18); und

    eine Vielzahl von Gasturbinenkraftwerkskomponenten, die in Fluidverbindung mit der Brennkammer (18) angeordnet ist, beinhaltend:

    einen Schaufelblattabschnitt (26);

    eine Plattform (34), die an dem Schaufelblattabschnitt (26) gesichert ist, wobei die Plattform (34) einen Plattform-Luftdurchgang (52) beinhaltet, der darin angeordnet ist; und

    ein Winkelstück (48), das sich von dem Schaufelblattabschnitt (26) zu der Plattform (34) erstreckt, wobei das Winkelstück (48) einen Winkelstück-Luftdurchgang (54) beinhaltet, der mit dem Plattform-Luftdurchgang (52) fluidverbunden ist, um einen Luftstrom (56) zu dem Plattform-Luftdurchgang (52) zu transportieren, wobei der Plattform-Luftdurchgang (52) dazu konfiguriert ist, den Luftstrom (56) in einer axial stromaufwärtigen Richtung zu einer Plattformanströmkante (62) zu leiten, dann axial nach hinten zu einer Plattformabströmkante (64), bevor er an einem Plattform-Luftdurchgangsausgang (66) aus dem Plattform-Luftdurchgang (52) austritt,

    dadurch gekennzeichnet, dass:
    der Plattform-Luftdurchgang (52) eine Spiralform aufweist.


     
    11. Gasturbinenkraftwerk nach Anspruch 10, ferner umfassend einen Luftdurchgang (44), der an dem Schaufelblattabschnitt (26) angeordnet ist, wobei der Winkelstück-Luftdurchgang (54) den Schaufelblattabschnitt-Luftdurchgang (44) mit dem Plattform-Luftdurchgang (52) verbindet.
     
    12. Gasturbinenkraftwerk nach Anspruch 10 oder Anspruch 11, wobei der Plattform-Luftdurchgang (52) dazu konfiguriert ist, den Luftstrom von einem Plattform-Luftdurchgangeingang (58) zu der Plattformanströmkante (62) zu transportieren.
     
    13. Gasturbinenkraftwerk nach einem der Ansprüche 10 bis 12, wobei das Winkelstück (48) an ungefähr einer Mittelsehne des Schaufelblattabschnitts (26) angeordnet ist.
     
    14. Gasturbinenkraftwerk nach einem der Ansprüche 10 bis 13, wobei das Winkelstück (48) an einer Druckseite des Schaufelblattabschnitts (26) angeordnet ist.
     
    15. Gasturbinenkraftwerk nach einem der Ansprüche 10 bis 14, wobei der Winkelstück-Luftdurchgang (54) einstückig mit dem Plattform-Luftdurchgang (52) ausgebildet ist.
     


    Revendications

    1. Ensemble pale pour un moteur à turbine à gaz, comprenant :

    une pale (26) ;

    une plate-forme de pale (34) fixée à la pale (26), la pale (26) s'étendant radialement vers l'extérieur depuis la plate-forme de pale (34), la plate-forme de pale (34) comportant un passage d'air de plate-forme (52) disposé dans celle-ci ; et

    un gousset (48) s'étendant depuis la pale (26) vers la plate-forme de pale (34), le gousset (48) comportant un passage d'air de gousset (54) relié fluidiquement au passage d'air de plate-forme (52) pour transporter un flux d'air (56) vers le passage d'air de plate-forme (52), dans lequel le passage d'air de plate-forme (52) est conçu pour orienter le flux d'air (56) dans une direction axialement amont vers un bord d'attaque de plate-forme (62), puis axialement vers l'arrière vers un bord de fuite de plate-forme (64), avant de sortir du passage d'air de plate-forme (52) au niveau d'une sortie de passage d'air de plate-forme (66),

    caractérisé en ce que :
    le passage d'air de plate-forme (52) a une forme en spirale.


     
    2. Ensemble pale selon la revendication 1, comprenant en outre un passage d'air de pale (44) disposé au niveau de la pale (26), le passage d'air de gousset (54) reliant le passage d'air de pale (44) au passage d'air de plate-forme (52).
     
    3. Ensemble pale selon la revendication 1 ou la revendication 2, dans lequel le passage d'air de plate-forme (52) est conçu pour transporter le flux d'air (56) vers le bord d'attaque de plate-forme (62) depuis une entrée de passage d'air de plate-forme (58) .
     
    4. Ensemble pale selon une quelconque revendication précédente, dans lequel le gousset (48) est disposé environ à mi-corde de la pale (26).
     
    5. Ensemble pale selon une quelconque revendication précédente, dans lequel le gousset (48) est disposé au niveau d'un intrados de la pale (26).
     
    6. Ensemble pale selon une quelconque revendication précédente, dans lequel le passage d'air de gousset (54) est formé d'un seul tenant avec le passage d'air de plate-forme (52).
     
    7. Ensemble pale selon la revendication 6, dans lequel le passage d'air de gousset (54) est formé par coulage.
     
    8. Ensemble rotor (24) pour un moteur à turbine à gaz, comprenant :

    un disque rotor (28) ; et

    une pluralité de pales de rotor (26) s'étendant radialement vers l'extérieur depuis le disque rotor (26), chaque pale de rotor étant un ensemble pale (26) selon une quelconque revendication précédente.


     
    9. Ensemble rotor selon la revendication 8, dans lequel l'ensemble rotor (24) est un ensemble rotor de turbine (24).
     
    10. Moteur à turbine à gaz (10), comprenant :

    une chambre de combustion (18) ; et

    une pluralité de composants de moteur à turbine à gaz disposés en communication fluidique avec la chambre de combustion (18), comportant :

    une partie de profil aérodynamique (26) ;

    une plate-forme (34) fixée à la partie de profil aérodynamique (26), la plate-forme (34) comportant un passage d'air de plate-forme (52) disposé dans celle-ci ; et

    un gousset (48) s'étendant depuis la partie de profil aérodynamique (26) vers la plate-forme (34), le gousset (48) comportant un passage d'air de gousset (54) relié fluidiquement au passage d'air de plate-forme (52) pour transporter un flux d'air (56) vers le passage d'air de plate-forme (52), dans lequel le passage d'air de plate-forme (52) est conçu pour orienter le flux d'air (56) dans une direction axialement amont vers un bord d'attaque de plate-forme (62), puis axialement vers l'arrière vers un bord de fuite de plate-forme (64), avant de sortir du passage d'air de plate-forme (52) au niveau d'une sortie de passage d'air de plate-forme (66),

    caractérisé en ce que :
    le passage d'air de plate-forme (52) a une forme en spirale.


     
    11. Moteur à turbine à gaz selon la revendication 10, comprenant en outre un passage d'air (44) disposé au niveau de la partie de profil aérodynamique (26), le passage d'air de gousset (54) reliant le passage d'air de partie de profil aérodynamique (44) au passage d'air de plate-forme (52).
     
    12. Moteur à turbine à gaz selon la revendication 10 ou la revendication 11, dans lequel le passage d'air de plate-forme (52) est conçu pour transporter le flux d'air vers le bord d'attaque de plate-forme (62) depuis une entrée de passage d'air de plate-forme (58).
     
    13. Moteur à turbine à gaz selon l'une quelconque des revendications 10 à 12, dans lequel le gousset (48) est disposé environ à mi-corde de la partie de profil aérodynamique (26).
     
    14. Moteur à turbine à gaz selon l'une quelconque des revendications 10 à 13, dans lequel le gousset (48) est disposé au niveau d'un intrados de la partie de profil aérodynamique (26) .
     
    15. Moteur à turbine à gaz selon l'une quelconque des revendications 10 à 14, dans lequel le passage d'air de gousset (54) est formé d'un seul tenant avec le passage d'air de plate-forme (52) .
     




    Drawing

















    Cited references

    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