(19)
(11)EP 3 524 807 B1

(12)EUROPEAN PATENT SPECIFICATION

(45)Mention of the grant of the patent:
04.11.2020 Bulletin 2020/45

(21)Application number: 19157332.8

(22)Date of filing:  23.12.2009
(51)International Patent Classification (IPC): 
F02K 1/36(2006.01)
B64C 21/06(2006.01)
F01D 9/06(2006.01)
F02K 3/02(2006.01)
F02K 3/077(2006.01)
F04D 29/64(2006.01)
F04D 29/68(2006.01)
F02K 3/075(2006.01)

(54)

APPARATUS COMPRISING A GAS TURBINE AND AN EJECTOR

VORRICHTUNG MIT EINEM GASTURBINENTRIEBWERK UND EINEM EJEKTOR

DISPOSITIF COMPRENANT UN MOTEUR À TURBINE À GAZ ET UN ÉJECTEUR


(84)Designated Contracting States:
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 SE SI SK SM TR

(30)Priority: 31.12.2008 US 20400408 P
23.10.2009 US 60475909

(43)Date of publication of application:
14.08.2019 Bulletin 2019/33

(62)Application number of the earlier application in accordance with Art. 76 EPC:
09015974.0 / 2204568

(73)Proprietor: Rolls-Royce Corporation
Indianapolis, IN 46241 (US)

(72)Inventor:
  • Khalid, Syed Jalaluddin
    Palm Beach Gardens, FI 33418 (US)

(74)Representative: Beyer, Andreas 
Wuesthoff & Wuesthoff Patentanwälte PartG mbB Schweigerstrasse 2
81541 München
81541 München (DE)


(56)References cited: : 
US-A- 3 317 162
US-A- 4 052 847
US-A- 5 586 431
US-A- 3 409 228
US-A- 5 435 127
  
      
    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

    Cross Reference to Related Applications



    [0001] The present application claims the benefit of U.S. Provisional Patent Application 61/204,004, filed December 31, 2008.

    FIELD OF THE INVENTION



    [0002] The present invention generally relates to gas turbine engines useful in removing a boundary layer of air, and more particularly, but not exclusively, to integrated aircraft engine with an ejector having fluid supplied by boundary layer bleed flow.

    BACKGROUND



    [0003] Challenges remain in providing aircraft propulsion systems capable of producing high thrust, emitting low jet noise, minimizing installation drag, minimizing fuel consumption, and ensuring fuel and lubrication temperatures are within specification requirements, among potential other challenges. The present invention addresses problems associated with existing propulsion systems by providing a novel and non-obvious contributions relating thereto.

    [0004] Document US 3 317 162 A discloses an aircraft wing having internal air passages for increasing lift of the aircraft. Air is directed from either a gate in the wing or by a blower located in a jet engine. For increased lift, a flap on the bottom of the wing allows out the air in the passage. The air may also be directed to a gated outlet leading to an exhaust opening.

    [0005] Document US 4 052 847 A discloses a ducted fan gas turbine engine power plant having an augmenter air intake and a flow mixer for selectively mixing the augmenter air with the remaining engine flows. The augmented air is introduced to an exhaust duct by pivoting flap members.

    [0006] Document US 3 409 228 A discloses a tertiary ejector nozzle for a reaction engine, having spaced outer and inner casings forming a cooling flow passage around the engine, wherein a converging member is connected to the inner casing to form a nozzle for a primary gas flow, and wherein an ejector flap is spaced substantially parallel with and supported on the converging member outwardly thereof. Document US 3 572 960 A discloses an axial flow compressor having a first row of fixed and a second row of rotatable cambered blades, said rows being axially spaced. The blades of the first row have, in their trailing edges, openings generally aligned with the median camber line of said blades.

    SUMMARY



    [0007] The present invention concerns an apparatus having the features of claim 1. A unique gas turbine engine ejector disclosed herein is provided for removing a boundary layer from a flow surface.

    [0008] In particular, the apparatus according to the present invention comprises a gas turbine engine having a nozzle ejector, an exhaust flow from the gas turbine engine forming the primary fluid of the nozzle ejector, a passageway operable to convey at least a portion of a boundary layer air withdrawn from a flow stream surface upstream of the nozzle ejector, and a moveable ejector selector having an open position and a closed position, the open position forming an open position secondary airflow path from an exterior location of the nozzle ejector to the primary fluid of the nozzle ejector, the closed position forming a closed position secondary airflow path from the flow stream surface, through the passageway, and to the primary fluid of the nozzle ejector.

    [0009] In accordance with the present invention a boundary layer ejector may advantageously fluidically connect boundary layer bleed slots from an external surface of an aircraft/nacelle/pylon to reduce aircraft drag, reduce jet noise and decrease thrust specific fuel consumption.

    [0010] Particular embodiments of the invention are defined by the dependent claims 2 to 14.

    BRIEF DESCRIPTION OF THE DRAWINGS



    [0011] The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein:
    FIG. 1
    is a schematic cross-sectional view of a prior art gas turbine engine with a nozzle ejector;
    FIG. 2
    is a schematic cross-sectional view of a gas turbine engine with a boundary layer ejector according to an embodiment of the present application;
    FIG. 3
    is a schematic cross-sectional view of a portion of a gas turbine engine with a boundary layer ejector according a further embodiment of the present application;
    FIG. 4a
    is a perspective view of one form of aircraft using the boundary layer ejector of the present application;
    FIG. 4b
    is a perspective view of another form of aircraft using the boundary layer ejector of the present application;
    FIG. 4c
    is a perspective view of yet another form of aircraft using the boundary layer ejector of the present application; and
    FIG. 4d
    is a front view of an aircraft having one embodiment of the present application.

    DETAILED DESCRIPTION OF THE ILLUSTRATIVE EMBODIMENTS



    [0012] A propulsion nozzle ejector disclosed herein utilizes boundary layer bleed to supply airflow to an ejector passageway. The disclosed ejector is operable to increase take-off thrust, reduce in-flight thrust specific fuel consumption, reduce installation and aircraft drag, reduce in-flight jet noise through a decrease in exhaust velocity, and provide a source of relatively low temperature airflow for thermal management of aircraft systems.

    [0013] Boundary layer fluid flow is defined by the fluid in a layer adjacent to a body within which the major effects of viscosity are concentrated. In simplistic terms, viscosity can be thought of as the thickness of the low velocity fluid or the resistance of the fluid to flow along a pathway. Viscous fluids create a boundary layer adjacent a body wherein the velocity of the fluid is approximately zero at the surface of the body and increases proportionally until it reaches the bulk fluid velocity outside of the boundary layer. The boundary layer produces losses due to aerodynamic drag on the aircraft. The boundary layer can be thought of as a dead zone wherein minimal useful fluid work can be done. The present invention advantageously removes at least a portion of boundary layer fluid and converts the boundary layer fluid into a useful means of increasing system efficiency and providing a source of thermal management.

    [0014] Referring to FIG. 1, a prior art gas turbine engine 10A is illustrated. Components of the prior art engine 10A that are similar to the inventive engine 10B in FIG. 2, will have the same numerical description. The gas turbine engine 10A, as well as the inventive engine 10B, can take a variety of forms and, in application, can be used to provide power to an aircraft. The gas turbine engine 10A and/or 10B can be an adaptive cycle engine or a variable cycle engine. Alternatively, the gas turbine engine 10A and/or 10B can have any number of spools, including just one. Still further, the gas turbine engine 10A and/or 10B can be a turbojet, turboprop, turbofan, or a turboshaft engine.

    [0015] As used herein, the term "aircraft" includes, but is not limited to, helicopters, airplanes, unmanned space vehicles, fixed wing vehicles, variable wing vehicles, rotary wing vehicles, unmanned combat aerial vehicles, tailless aircraft, hover crafts, and other airborne and/or extraterrestrial (spacecraft) vehicles. Further, the present invention is contemplated for utilization in other applications that may not be coupled with an aircraft such as, for example, industrial applications, power generation, pumping sets, naval propulsion, weapon systems, security systems, perimeter defense/security systems, and the like known to one of ordinary skill in the art.

    [0016] The gas turbine engine 10A includes an inlet section 12, a compression section 14, a combustion section 16, an expansion or turbine section 18, and an exhaust section 20. In operation, air is drawn in through the inlet 12 and compressed to a high pressure relative to ambient pressure in the compression section 14. The air is mixed with fuel in the combustion section 16 wherein the fuel/air mixture burns and produces a high temperature and pressure working fluid from which the turbine section 18 extracts power. The turbine section 18 is mechanically coupled to the compression section 14 via a shaft 21. The shaft 21 rotates about a centerline axis that extends axially along the longitudinal axis of the engine 10A, such that as the turbine section 18 rotates due to the forces generated by the high pressure working fluid the compression section 14 is rotatingly driven by the turbine section 18 to produce compressed air.

    [0017] The compression section 14 can optionally include a low-pressure fan 23, which produces a bypass flow stream represented by arrow 26. The bypass flow stream 26 flows between an outer case 22 and an inner fairing 24 that encompasses the core 25 of the engine 10A. The core 25 produces a core exhaust flow represented by arrow 28. Mixed engine exhaust flow represented by arrows 30 is the combined mass flow of the bypass flow 26 and the core flow 28.

    [0018] A nacelle 32 encompasses the outer case 22 of the engine 10A. A variable flap 34 is positioned adjacent the aft end of the nacelle 32.Alternatively, the variable flap 34 can be located at positions other than adjacent the aft end of the nacelle 32. A nozzle 36 is positioned further aft of the variable flap 34. When the variable flap 34 is in an open position, a nozzle ejector flow 38 can be entrained through a passageway 39 formed between the variable flap 34 and the nozzle 36.Alternatively, however, the passageway 39 can be formed between the variable flap 34 and a structure or structures other than the nozzle 36. A total exhaust flow represented by arrow 40 includes the mixed engine exhaust flow 30 and the nozzle ejector flow 38.

    [0019] Referring to FIG. 2, a gas turbine engine 10B is illustrated having novel features relative to the gas turbine engine 10A. The gas turbine engine 10B includes boundary layer bleed flow represented by arrow 50. The boundary layer bleed flow 50 can be transported through bleed ports (not shown in FIG. 2 but described and shown further below) from any surface of the engine 10B or aircraft (also not shown in FIG. 2 but described and shown further below) and directed through a nacelle passageway 52. The bleed ports can take any variety of forms including slots or holes, among potential others. As a matter of convenience of description, therefore, the term "bleed port" can refer to any of these types of configurations. The surface from which the boundary layer bleed flow 50 originates from can be any surface which is exposed to a moving fluid. Such surfaces include the outside portion of a nacelle exposed to the free stream, the inside surface forming part of the passageway through which the bypass flow stream passes, among potential other surfaces.

    [0020] One or more heat exchangers 54 can be disposed within the passageway 52 to remove heat from system components (not shown). The one or more heat exchangers 54 can take a variety of forms and, in those applications in which multiple heat exchanges 54 are used, not all need be the same. The temperature of the boundary layer bleed flow 50 can be approximately the same as the ambient temperature, therefore at altitude the boundary layer bleed flow 50 can provide a large heatsink. If one or more heat exchangers are present, after the boundary layer bleed flow 50 passes the heat exchanger 54, the flow can be accelerated through a channel 51 between the nozzle 36 and the outer case 22. Whether or not heat exchangers are present, however, it will be understood that the channel 51 can be of an orientation to permit the flow to be accelerated. For example, the channel 51 can have a smaller cross sectional area relative to the passageway 52 at locations upstream of the channel 51.

    [0021] In some applications a channel member 53 can be used to selectively vary the cross sectional area of the channel 51, and in particular can be used to change the exit area of the passageway 52. The channel member 53 can be actuated at a variety of rates which can depend on properties of a total exhaust flow 56 which is the boundary layer bleed flow 50 mixed with the bypass flow 26 and the core flow 28. Removing boundary layer flow from external surfaces reduces the drag and the additional mass flow added into the total exhaust flow 56 increases the thrust of the engine 10B.

    [0022] In the take-off and taxi mode, the variable flap 34 can be opened as shown in the dashed outline, which permits air to enter from the rear and go around the ejector leading edge into the exhaust stream 56, similar to that depicted in Fig. 1. The integrated static pressure around the ejector results in a thrust component while the jet noise is reduced as a result of reduced exhaust velocity. In cruise mode, the variable flap 34 can be closed as depicted by the solid outline in FIG. 2. Boundary layer bleed air 50 is then pulled from passageways connected to a surface associated with any of the nacelle, pylon, wing, or fuselage or other locations that a boundary layer can build up on. The resulting removal of the boundary layer reduces the installation/aircraft drag. The RAM effect during cruise in conjunction with ejector pumping can effectively remove the boundary layer especially since the ejector is being driven by a high velocity (normally sonic but can be supersonic) and high mass flow primary nozzle 36 which will entrain a large amount of secondary flow. Ejector pumping can be further enhanced by forming the shroud, concentric surface, or other structure that separates the primary nozzle flow from the secondary boundary layer flow as a lobed mixer configuration. Such a configuration can assist in forcing mixing between the two flows and permit shortening of the engine length. Referring to FIG. 3, an alternate engine configuration 10C is depicted. It will be appreciated, however, that the alternate engine configuration can have a number of the variations described above with respect to the gas turbine engines 10A and/or 10B. A nacelle 70 can include boundary layer bleed ports 72 fluidically connected to a member 74 or the like extending at least partially across the flow path depicted. The member 74 can take the form of a strut, an outlet guide vane, or other type of member disposed within the flow path. The member includes a passage 75 through which a boundary layer, which has been pulled through boundary layer bleed ports, 72 can pass. The passage 75 can be the hollow interior of the member 74, but can also take other forms. The passage 75 can extend across the bypass flow path 77 and the core flow path 79. In some forms, the member 74 can include one or more exit apertures 81 through which the boundary layer which has been pulled through boundary layer bleed ports 72 can pass after passing through the passage 75. The one or more exit apertures 81 can be formed in or near the trailing edge of the member 74. In some forms, a conduit 76 can be provided in a member 74 which is operable to carry relatively cool boundary layer bleed air 78 to one or more components 80 within the engine 10C requiring cooling flow. For example, the component 80 can be an electronic apparatus or a relatively hot mechanical apparatus such as a combustion or turbine component. Though the conduit 76 is depicted in a member 74 downstream of the member 74 having the exit apertures 81, it will be appreciated that the conduit 76 can be provided in the member 74 shown as having the apertures 81, but otherwise lacking the apertures 81. In other words, in lieu of or in addition to the aperture 81, the member 74 can include the conduit 76. Still further, the conduit 76 can lead to an ejector such as the nozzle ejector depicted in Fig. 2.

    [0023] In some forms, the alternate engine configuration 10C can be used to remove a boundary layer through boundary layer bleed ports 72 to increase inlet pressure recovery and reduce pressure distortion seen by the fan. The resulting higher pressure recovery increases thrust and lowers specific fuel consumption. The lower pressure distortion increases stall margin.

    [0024] In some forms, the engine 10C can include a flow path protrusion 83 that can reduce the cross sectional area of the bypass flow path 77 at that location. For example, in some forms, outlet guide vanes (OGV's) are provided to eliminate circumferential swirl and can have a flow area that increases from the leading to the trailing edge. The inner and outer walls of the OGV's can be tapered to reduce the area increase thus increasing the OGV exit velocity and ejector pumping. The flow path protrusion 83 can be used to locally lower the pressure of the fluid passing through the bypass flow path 77, so that in the case of boundary layer which has been pulled through boundary layer bleed ports 72 and is discharged through the exit aperture 81 can be entrained by the flow through the bypass flow path 77. In one form, the flow path protrusion 83 can be actuated and moved to a variety of positions to selectively lower or raise the local pressure and thus vary the ability of the flow through the bypass flow path 77 to entrain the boundary layer which has been pulled through boundary layer bleed ports 72.

    [0025] Referring now to FIG. 4a, an aircraft 90 is depicted having a fuselage 92 with wings 94 extending therefrom and gas turbine engines 96 mounted thereon. The wings 94 can include a plurality of bleed slots 98 for removing boundary layer flow from the wings 94 and delivering said flow to the ejector nozzle or other components disposed on the aircraft 90, such as through any of the various embodiments discussed herein. In addition to or alternatively, nacelle bleed slots 100 and fuselage bleed slots 102 can be utilized for removing boundary layer flow. Referring briefly to Fig. 4d, in some aircraft configurations, the Mach number in the passage located between the nacelle 146 and the fuselage 148 could increase because of the blockage created by the boundary layer thus resulting in increased installation drag. Removal of this boundary layer through one or more of the bleed slots 144 depicted in Fig. 4d can result in decreased drag. The bleed slots 144 can be located on either or both of the fuselage 148 or nacelle 146.

    [0026] Referring to FIG 4b, an alternate aircraft configuration 110 is depicted with wing 114 mounted engines 116. Similar to the aircraft of FIG. 4a, the fuselage 112, the wings 114, and the engines 116 can all include boundary layer bleed slots 118, 120 and 122, respectively.

    [0027] In yet another aircraft configuration, FIG 4c illustrates a military style aircraft 130 wherein the engines 136 are embedded within a fuselage 132. Again, similar to the previous aircraft of FIGS. 4a and 4b the fuselage 132, a nacelle 141 surrounding the engine 136 and the wings 134, can all include boundary layer bleed slots 142, 140, and 138 respectively. The slots can also be located between the inlet and the fuselage (sometimes referred to as the "arm pit") where the fuselage boundary layer can thicken.

    [0028] It should be understood that the boundary layer bleed slots can be positioned anywhere on the aircraft and/or engine and are not limited by the examples disclosed in the present application.


    Claims

    1. An apparatus comprising:

    a gas turbine engine (10C) having a nozzle ejector (36), an exhaust flow (56) from the gas turbine engine (10C) forming the primary fluid of the nozzle ejector (36);

    a passageway (39) operable to convey at least a portion of a boundary layer air withdrawn from a flow stream surface upstream of the nozzle ejector (36); and

    a moveable ejector selector (34) having an open position and a closed position, the open position forming an open position secondary airflow path from an exterior location of the nozzle ejector (36) to the primary fluid of the nozzle ejector (36), the closed position forming a closed position secondary airflow path from the flow stream surface, through the passageway (39), and to the primary fluid of the nozzle ejector (36).


     
    2. The apparatus of claim 1, which further includes an aircraft (90) and wherein the flow stream surface is at least one of a nacelle (70) and a pylon.
     
    3. The apparatus of claim 1, which further includes an area control member (53) capable of varying an exit area of at least one of the open position secondary airflow path and the closed position secondary airflow path as the withdrawn boundary layer air is pumped by the exhaust flow.
     
    4. The apparatus of claim 1, which further includes a heat exchanger (54) disposed within the closed position secondary airflow path.
     
    5. The apparatus of claim 2, wherein the aircraft (90) includes a plurality of bleed slots (98, 100, 102, 118, 120, 122, 138, 142, 144) for removing boundary layer air and delivering the boundary layer air to the nozzle ejector (36).
     
    6. The apparatus of claim 1, wherein the passageway (39) includes a portion through a duct member (74) that extends across an airflow path of the gas turbine engine (10C).
     
    7. The apparatus of claim 6, wherein the gas turbine engine (10C) is a turbofan and the duct member (74) is an outlet guide vane downstream of a fan blade.
     
    8. The apparatus of claim 6, wherein the gas turbine engine (10C) is a turbofan and the duct member (74) is a strut.
     
    9. The apparatus of claim 6, wherein the flow stream surface is a nacelle (70) that encompasses the gas turbine engine (10C) and wherein the nacelle (70) includes boundary layer bleed ports (72) fluidly connected to the duct member (74) or extending at least partially across the airflow path.
     
    10. The apparatus of claim 9, wherein the duct member (74) includes one or more exit apertures (81) through which the boundary layer air which has been pulled through the boundary layer bleed ports (72) passes after passing through a passage (75) formed in the duct member (74).
     
    11. The apparatus of claim 10, wherein the passage (75) extends across a bypass flow path (77) and a core flow path (79).
     
    12. The apparatus of claim 11, wherein the gas turbine engine (10C) includes a flow path protrusion (83) that can reduce the cross sectional area of the bypass flow path (77) to locally lower the pressure of the fluid passing through the bypass flow path (77) so that in the case of boundary layer air which has been pulled through boundary layer bleed ports (72) and is discharged through the exit apertures (81) can be entrained by the flow through the bypass flow path (77).
     
    13. The apparatus of claim 6, wherein the duct member (74) includes a conduit (76) which is operable to carry relatively cool boundary layer bleed air to one or more components (80) within the gas turbine engine (10C).
     
    14. The apparatus of claim 13, wherein the one or more components (80) is an electronic apparatus.
     


    Ansprüche

    1. Vorrichtung, umfassend:

    ein Gasturbinentriebwerk (10C) mit einem Düsenejektor (36), wobei ein Abgasstrom (56) aus dem Gasturbinentriebwerk (10C) das Primärfluid des Düsenejektors (36) bildet,

    einen Durchgang (39), der dazu betriebsfähig ist, zumindest einen Teil einer Grenzschichtluft zu befördern, die aus einer Strömungsflussoberfläche stromaufwärts des Düsenejektors (36) abgezogen wurde, und

    einen bewegbaren Ejektorwähler (34) mit einer offenen Stellung und einer geschlossenen Stellung, wobei die offene Stellung einen Offenstellungs-Sekundärluftströmungspfad von einer äußeren Stelle des Düsenejektors (36) zu dem Primärfluid des Düsenejektors (36) bildet, und wobei die geschlossene Stellung einen Schließstellungs-Sekundärluftströmungspfad von der Strömungsflussoberfläche durch den Durchgang (39) und zu dem Primärfluid des Düsenejektors (36) bildet.


     
    2. Vorrichtung nach Anspruch 1, die ferner ein Flugzeug (90) enthält und wobei die Strömungsflussoberfläche eine solche einer Gondel (70) und/oder einer Pylone ist.
     
    3. Vorrichtung nach Anspruch 1, die ferner ein Flächensteuerbauteil (53) aufweist, welches dazu in der Lage ist, eine Austrittsfläche des Offenstellungs-Sekundärluftstrompfades und/oder des Schließstellungs-Sekundärluftstrompfades zu ändern, wenn die abgezogene Grenzschichtluft mittels des Abgasstroms gepumpt wird.
     
    4. Vorrichtung nach Anspruch 1, die ferner einen Wärmetauscher (54) aufweist, der innerhalb des Schließstellungs-Sekundärluftstrompfades angeordnet ist.
     
    5. Vorrichtung nach Anspruch 2, wobei das Flugzeug (90) mehrere Zapfschlitze (98, 100, 102, 118, 120, 122, 138, 142, 144) zum Abziehen von Grenzschichtluft und Liefern der Grenzschichtluft zu dem Düsenejektor (36) aufweist.
     
    6. Vorrichtung nach Anspruch 1, wobei der Durchgang (39) einen Abschnitt durch ein Kanalbauteil (74) aufweist, das sich quer durch einen Luftströmungspfad des Gasturbinentriebwerks (10C) erstreckt.
     
    7. Vorrichtung nach Anspruch 6, wobei das Gasturbinentriebwerk (10C) ein Mantelstromtriebwerk ist und das Kanalbauteil (74) eine Auslassleitschaufel stromabwärts eines Lüfterflügels ist.
     
    8. Vorrichtung nach Anspruch 6, wobei das Gasturbinentriebwerk (10C) ein Mantelstromtriebwerk ist und das Kanalbauteil (74) eine Strebe ist.
     
    9. Vorrichtung nach Anspruch 6, wobei die Strömungsflussoberfläche eine Gondel (70) ist, welche das Gasturbinentriebwerk (10C) umgibt, und wobei die Gondel (70) Grenzschichtzapföffnungen (72) aufweist, die fluidleitend mit dem Kanalbauteil (74) verbunden sind oder sich zumindest teilweise quer durch den Luftströmungspfad erstrecken.
     
    10. Vorrichtung nach Anspruch 9, wobei das Kanalbauteil (74) eine oder mehrere Auslassöffnungen (81) aufweist, durch welche die durch die Grenzschichtzapföffnungen (72) abgezogene Grenzschichtluft strömt, nachdem sie einen in dem Kanalbauteil (74) gebildeten Durchgang passiert hat.
     
    11. Vorrichtung nach Anspruch 10, wobei der Durchgang (75) sich über einen Bypassströmungspfad (77) und einen Kernströmungspfad (79) erstreckt.
     
    12. Vorrichtung nach Anspruch 11, wobei der Gasturbinenmotor (10C) einen Strömungspfadvorsprung (83) aufweist, der die Querschnittsfläche des Bypassströmungspfades (77) verringern kann, um den Druck des durch den Bypassströmungspfad (77) strömenden Fluids lokal abzusenken, so dass im Fall von Grenzschichtluft, die durch Grenzschichtzapföffnungen (72) abgezogen worden ist und durch die Auslassöffnungen (81) ausgestoßen worden ist, diese von der Strömung durch den Bypassströmungspfad (77) mitgerissen werden kann.
     
    13. Vorrichtung nach Anspruch 6, wobei das Kanalbauteil (74) eine Leitung (76) aufweist, welche dazu betriebsfähig ist, relativ kühle Grenzschichtzapfluft zu einem oder mehreren Bauteilen (80) innerhalb des Gasturbinentriebwerks (10C) zu tragen.
     
    14. Vorrichtung nach Anspruch 13, wobei das eine oder die mehreren Bauteile (80) ein elektronischer Apparat ist.
     


    Revendications

    1. Appareil comprenant:

    un moteur à turbine à gaz (10C) comprenant un injecteur de buse (36), un flux d'échappement (56) provenant du moteur à turbine à gaz (10C) formant le fluide primaire de l'injecteur de buse (36) ;

    un passage (39) servant à transporter au moins une partie de l'air de couche limite retiré de la surface d'écoulement en amont de l'injecteur de buse (36) ; et

    un sélecteur (34) d'éjecteur mobile ayant une position ouverture et une position fermée, la position ouverte formant un trajet d'écoulement d'air secondaire en position ouverture à partir d'un emplacement extérieur de l'éjecteur de buse (36) en direction du fluide primaire de l'éjecteur de buse (36), la position fermée formant un trajet d'écoulement d'air secondaire en position fermée à partir de la surface d'écoulement, au travers du passage (39), et en direction du fluide primaire de l'éjecteur de buse (36).


     
    2. Appareil selon la revendication 1, qui comprend en outre un aéronef (90) et dans la surface d'écoulement est au moins une nacelle (70) et/ou mât d'accrochage.
     
    3. Appareil selon la revendication 1, qui comprend en outre un élément de commande de zone (53) apte à modifier une zone de sortie du trajet d'écoulement d'air seconde en position ouverture et du trajet d'écoulement secondaire en position fermée alors que l'air de la couche limite retiré est pompé par le flux d'échappement.
     
    4. Appareil selon la revendication 1, qui comprend en outre un échangeur de chaleur (54) disposé à l'intérieur du trajet d'écoulement secondaire en position fermée.
     
    5. Appareil selon la revendication 2, dans lequel l'aéronef (90) comprend une pluralité de fentes de prélèvement (98, 100 ; 102, 118, 120, 122, 138, 142, 144) pour retirer l'air de la couche limite et pour alimenter e air de couche limite l'éjecteur de buse (36).
     
    6. Appareil selon la revendication 1, dans lequel le passage (39) comprend une partie traversant un élément de conduite (74) qui passe dans tout le trajet de flux d'air du moteur à turbine à gaz (10C).
     
    7. Appareil selon la revendication 6, dans lequel le moteur à turbine à gaz (10C) est un turboréacteur et l'élément de conduite (74) est un aubage directeur de sortie en aval d'une aube mobile de soufflante.
     
    8. Appareil selon la revendication 6, dans lequel le moteur à turbine à gaz (10C) est un turboréacteur et l'élément de conduite (74) est un mât.
     
    9. Appareil selon la revendication 6, dans lequel la surface d'écoulement est une nacelle (70) qui entoure le moteur à turbine à gaz (10C) et dans lequel la nacelle (70) comprend des trous de prélèvement de couche limite (72) en communication fluidique avec l'élément de conduite (74) ou s'étendant au moins en partie dans tout le trajet de flux d'air.
     
    10. Appareil selon la revendication 9, dans lequel l'élément de conduite (74) comprend une ou plusieurs ouvertures de sortie (81) que traverse l'air de couche limite qui a été aspiré par les trous de prélèvement de couche limite (72) après avoir traversé un passage (75) formé dans l'élément de conduite (74).
     
    11. Appareil selon la revendication 10, dans lequel le passage (75) s'étend dans tout un trajet d'écoulement de dérivation (77) et dans un trajet d'écoulement central (79).
     
    12. Appareil selon la revendication 11, dans lequel le moteur à turbine à gaz (10C) comprend une saillie de trajet d'écoulement (83) qui peut réduire la zone traversable du trajet d'écoulement de dérivation (77) pour abaisser localement la pression du fluide traversant le trajet d'écoulement de dérivation (77) de sorte que, dans le cas où l'air de couche limite qui aurait été aspiré par les trous de prélèvement de couche limite (72) et aurait été évacué par les ouvertures de sortie (81), puisse être entraîné par l'écoulement dans le trajet d'écoulement de dérivation (77).
     
    13. Appareil selon la revendication 6, dans lequel l'élément de conduite (74) comprend une conduite (76) qui peut servir à porter l'air prélevé de couche limite relativement tiède dans au moins un composant (80) à l'intérieur du moteur à turbine à gaz (10C).
     
    14. Appareil selon la revendication 13, dans lequel ledit au moins un composant (80) est un appareil électronique.
     




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    Cited references

    REFERENCES CITED IN THE DESCRIPTION



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    Patent documents cited in the description