EP3938626B1 - Secondary flow rectifier with integrated pipe - Google Patents

Secondary flow rectifier with integrated pipe Download PDF

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Publication number
EP3938626B1
EP3938626B1 EP20725890.6A EP20725890A EP3938626B1 EP 3938626 B1 EP3938626 B1 EP 3938626B1 EP 20725890 A EP20725890 A EP 20725890A EP 3938626 B1 EP3938626 B1 EP 3938626B1
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EP
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Prior art keywords
flow
vane
turbomachine
downstream
section
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EP20725890.6A
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German (de)
French (fr)
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EP3938626A2 (en
Inventor
Florent Matthieu Jacques NOBELEN
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Safran Aircraft Engines SAS
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Safran Aircraft Engines SAS
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/125—Fluid guiding means, e.g. vanes related to the tip of a stator vane

Definitions

  • the field of the invention relates to multiple-flow turbomachines, and more specifically the flow rectifiers of a separate multiple-flow turbomachine.
  • a multi-flow turbomachine as shown in figure 1 conventionally comprises a fan 1, a fan casing 2 and a casing 3 extending along a longitudinal axis X.
  • the casing 3 houses the compression, combustion and expansion elements of the turbomachine.
  • the fan casing 2 extends radially outside the fan 1 and the casing 3 so as to delimit the flow entering the fan 1.
  • the fan 1 compresses and accelerates the flow of air entering the fan casing 2, this air flow then circulating in a primary circuit 4 and a secondary circuit 5, the primary circuit 4 being located inside the casing 3 and traversing the various compression, combustion and expansion elements, the secondary circuit 5 being delimited radially internally by the casing 3 and externally by the fan casing 2.
  • the rotation of the fan 1 inducing a gyration in the flow which it accelerates it is known to arrange a flow rectifier 6 in the secondary circuit 5, the rectifier 6 comprising a plurality of vanes 7 configured to modify the direction of circulation of the flow in order to obtain an axial flow downstream of the rectifier 6.
  • the profile of the nacelles 2 is conventionally configured to form a nozzle downstream of the rectifiers and to accelerate and relax the secondary flow so as to generate the thrust, the section of the secondary circuit 5 decreasing downstream (in the case of a convergent nozzle), then can possibly increase again in the case of a convergent-divergent nozzle.
  • each flow is ejected through a nozzle.
  • the nozzle (primary and secondary) transforms potential energy into energy kinetic, that is to say, it converts the pressure of the flow into the velocity of ejection, which will generate the thrust.
  • the secondary flow nozzle surrounds and is conventionally placed upstream of the primary flow nozzle.
  • the primary flow nozzle is delimited by a cone whose tip is directed downstream and by an annular casing having a trailing edge directed downstream.
  • the cone and the casing define a circuit of convergent or convergent-divergent section according to the architectural choices made.
  • the secondary nozzle is delimited by a duct belonging to the fan casing (commonly called OFD or OFS abbreviated to “ Outer Fan Duct / Shroud ”) and to the turbomachine casing (commonly called IFD or IFS abbreviated to “ Inner Fan Duct / Shroud ”) .
  • the two casings define a convergent or convergent-divergent section according to the architecture of the rest of the engine.
  • This decrease in section is conventionally located downstream of the rectifier 6, so as to accelerate the secondary flow when it flows axially, the secondary flow then being ejected around the primary flow.
  • the air inlet must be extremely short, and the fan casing 2 must be as short as possible after the exit of the vanes 7 of the stator 6.
  • the document GB-2 546 422 discloses an assembly according to the preamble of claim 1.
  • An object of the invention is to reduce the pressure drops induced by the fan casing.
  • Another object of the invention is to accelerate the secondary flow.
  • Another object is to limit the load losses induced by the rectifier.
  • Another object of the invention is to increase the bypass ratio of the turbomachine.
  • Another object is to reduce the compression ratio of the fan.
  • the invention proposes an assembly for a turbomachine in accordance with claim 1.
  • the invention proposes a turbomachine comprising such an assembly.
  • the turbomachine extends along an axis X of the turbomachine, and the terms axial, radial and tangential refer to the axis X of the turbomachine.
  • An axial direction follows axis X of the turbomachine, a radial direction is perpendicular to axis X of the turbomachine, and a tangential direction is orthogonal to a radial direction and an axial direction.
  • the turbomachine is a dual-flow turbomachine further comprising a fan 1, housed in the fan casing 2, and rotatable about a longitudinal axis X, an inner shroud 31 configured to delimit a primary stream 4 of a primary gas flow from the turbomachine, the shroud 32 and the fan casing 2 delimiting a so-called secondary flow path for an air flow propelled by the fan 1.
  • the shroud 32 is located in the upstream extension of the casing 3 of the turbomachine.
  • the shroud 32 can be part of the casing 3, and thus form the upstream portion of the casing 3.
  • the shroud 32 and the inner shroud 31 may form a single piece and form the leading edge of the casing 3.
  • the inlet section 14a, the ejection section 14b and the outlet section 14c extending respectively from the radially inner limit to the radially outer limit of the blades 7.
  • the inlet section 14a thus corresponds to a radial section of the flow channel 13 which coincides with the leading edge 9 of the second blade 7b, and the ejection section 14b corresponds to a radial section extending downstream of the input section 14a.
  • the ejection section 14b has a surface smaller than a surface of the inlet section 14a and smaller than a surface of the outlet section 14c.
  • the flow channel 13 has a radial section 14 which is defined as a virtual plane extending from the extrados wall 10a of the first blade 7a to the intrados wall 11b of the second blade 7b while being normal to an average direction of the flow at a central streamline F and extending substantially radially relative to the longitudinal axis X.
  • central streamline is meant the streamline located equidistant from the first blade 7a and the second blade 7b.
  • the radial section 14 of the flow channel 13 has a surface which gradually decreases between the inlet section 14a and the ejection section 14b.
  • the radial section 14 has a width L defined as a distance between the upper surface 10a of the first blade 7a and the lower surface 11b of the second blade 7b for a constant distance from the axis X, and in which the width of the radial section 14 is decreasing according to the circulation of the flow in the flow channel 13 between the inlet section 14a and the ejection section 14b.
  • the extrados wall 10a of the first blade 7a and the intrados wall 11b of the second blade 7b are increasingly close to each other, for a given distance from the X axis, as the as the flow travels from upstream to downstream in the flow channel 13.
  • a radial section 14 has a shape comparable to an angular portion of a disc and has a dimension in a transverse direction and a dimension in a radial direction.
  • the radial section 14 is delimited by the first vane 7a and the second vane 7b.
  • the distance separating the first blade 7a and the second blade 7b, the width L, is a function of the distance from the axis X of the turbomachine at which the width L considered. Indeed the distance between the first blade 7a and the second blade 7b increases with the distance to the X axis.
  • the width of a radial section 14 is a function of the radius or of a distance from the axis X of the turbomachine, and increases as a function of the distance from the axis X of the turbomachine.
  • the radial section 14 is delimited radially internally by the outer shroud 32 and extends over the entire height of a blade 7.
  • the radial section 14 has a radially inner limit and a radially outer limit each substantially forming an arc of a circle.
  • the width L decreases, and optionally the dimension in the radial direction also decreases.
  • the decrease in the flow section 14 causes relaxation and therefore an acceleration of the secondary flow.
  • the lower surface 11a of the first blade 7a and the upper surface 10b of the second blade 7b are therefore configured so that the width L of a radial section 14, for a given distance from the axis X of the turbomachine, decreases as the flow moves downstream.
  • the width L of the radial section 14 will be less than the width of the inlet section 14a.
  • This width L can optionally be the length of a straight segment joining at mid-height the first blade 7b and the second blade 7a.
  • the length of the straight segment joining the first vane 7b and the second vane 7a at mid-height gradually decreases between the inlet section 14a and the ejection section 14b.
  • the ejection section 14b has the minimum area for a radial section 14.
  • the width L of a radial section 14 decreases as it moves from upstream to downstream as far as a median plane 15, the median plane 15 therefore including the ejection section 14b.
  • the median plane 15 is normal to the axis X of the turbomachine, and delimits the flow channel 13 into two parts, an upstream or intake portion 16 and a downstream or ejection portion. 17.
  • the transverse dimension of the radial section 14 is less than the transverse dimension of the inlet section 14a and greater than the transverse dimension of the ejection section 14b.
  • the flow channel 13 is convergent, the radial section 14 having a decreasing surface from upstream to downstream.
  • Inlet portion 16 of flow channel 13 is configured to do the work of changing flow direction and flow acceleration.
  • the flow channel 13 therefore has an inlet section 14a defining a plane normal (or orthogonal) to the flow direction of the flow diverted by the fan, this plane therefore not being normal to the axis X of the turbomachine, and an ejection section 14b defining a plane normal to the axis X of the turbomachine. This makes it possible to eject a flow circulating in a direction substantially parallel to the axis of the turbomachine.
  • the intake portion 16 straightens the flow while relaxing it and accelerating it until ejection at the level of the median plane 15.
  • the ejection portion 17 is configured to minimize the aerodynamic drag of the rectifier 6.
  • the angle of incidence of the profile with respect to the flow is low so as to avoid the separation of the air flow, while having the shortest possible length to minimize viscous friction.
  • Part of the ejection portion 17 is located downstream of the trailing edge 8 of the fan casing 2. Thus, this makes it possible to slow down the flow in the ejection portion 17 down to flight speed.
  • the profile of the blades 7 is configured to minimize the drag of each blade 7, the blades 7 therefore extending axially to their trailing edge 12.
  • the section of the flow channel 13 therefore increases downstream in the ejection portion 17.
  • the blades 7 have a line of camber 71 which may comprise a point of inflection, the line of camber or mean line being defined in that it extends from the leading edge 9 to the trailing edge 12 and that it is halfway between the extrados 10 and the intrados 11.
  • the line of camber 71 has an inclination with respect to the axis X of the turbomachine corresponding to the gyration of the flow at the leading edge 9, and is substantially parallel to the motor axis from the median plane 15 to the trailing edge 12.
  • the median plane 15, and therefore the ejection section 14b coincides with the trailing edge 8 of the fan casing.
  • the ejection portion 17 is therefore not streamlined.
  • the length of the fan casing 2 can be reduced to a minimum without penalizing the operation of the intake portion 16 which is streamlined by the fan casing 2, nor the operation of the ejection portion 17 whose sole role is to reduce drag.
  • a portion of the blades 7, in particular the trailing edge 12, is then located downstream of the trailing edge 8 of the fan casing 2, and is therefore not streamlined.
  • the fan casing 2 can extend axially beyond the median plane 15.
  • the trailing edge 8 of the fan casing 2 is located downstream of the median plane 15 and upstream of the trailing edges blades 12, at the level of a fairing plane 18.
  • This configuration makes it possible to form a converging then diverging profile in the streamlined part of the flow channels 13 (that is to say covered by the fan casing 2). This improves performance depending on the flight envelope.
  • each pair of adjacent vanes 7 of the straightener 6 defines a flow channel 13 configured to straighten and accelerate the flow simultaneously, the vanes of the straightener 6 thus defining a plurality of flow channels 13 distributed circumferentially.
  • the pressure losses are reduced by the reduction in the length of the fan casing 2 and the profile of the blades 7, more particularly the trailing edge 12 and the profile of the ejection portion 17 make it possible to reduce the drag and thus to limit separations and pressure drops.
  • Conventional nozzles form a converging channel which accelerates the flow without deflecting it.
  • the profile of the blades 7 ending in a trailing edge makes it possible to avoid flow separation at the outlet of the assembly.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Description

DOMAINE TECHNIQUE GÉNÉRAL ET ART ANTÉRIEURGENERAL TECHNICAL FIELD AND PRIOR ART

Le domaine de l'invention concerne les turbomachines à flux multiples, et plus précisément les redresseurs de flux d'une turbomachine à flux multiples séparés.The field of the invention relates to multiple-flow turbomachines, and more specifically the flow rectifiers of a separate multiple-flow turbomachine.

Une turbomachine à flux multiples telle qu'illustrée en figure 1 comporte classiquement une soufflante 1, un carter de soufflante 2 et un carter 3 s'étendant selon un axe longitudinal X.A multi-flow turbomachine as shown in figure 1 conventionally comprises a fan 1, a fan casing 2 and a casing 3 extending along a longitudinal axis X.

Le carter 3 loge les éléments de compression, de combustion et de détente de la turbomachine.The casing 3 houses the compression, combustion and expansion elements of the turbomachine.

Le carter de soufflante 2 s'étend radialement extérieurement à la soufflante 1 et au carter 3 de manière à délimiter le flux entrant dans la soufflante 1.The fan casing 2 extends radially outside the fan 1 and the casing 3 so as to delimit the flow entering the fan 1.

La soufflante 1 comprime et accélère le flux d'air entrant dans le carter de soufflante 2, ce flux d'air circulant ensuite dans un circuit primaire 4 et un circuit secondaire 5, le circuit primaire 4 étant situé à l'intérieur du carter 3 et parcourant les différents éléments de compression, de combustion et de détente, le circuit secondaire 5 étant délimité radialement intérieurement par le carter 3 et extérieurement par le carter de soufflante 2.The fan 1 compresses and accelerates the flow of air entering the fan casing 2, this air flow then circulating in a primary circuit 4 and a secondary circuit 5, the primary circuit 4 being located inside the casing 3 and traversing the various compression, combustion and expansion elements, the secondary circuit 5 being delimited radially internally by the casing 3 and externally by the fan casing 2.

La rotation de la soufflante 1 induisant une giration dans le flux qu'elle accélère, il est connu de disposer un redresseur 6 de flux dans le circuit secondaire 5, le redresseur 6 comportant une pluralité d'aubes 7 configurées pour modifier la direction de circulation du flux afin d'obtenir un écoulement axial en aval du redresseur 6.The rotation of the fan 1 inducing a gyration in the flow which it accelerates, it is known to arrange a flow rectifier 6 in the secondary circuit 5, the rectifier 6 comprising a plurality of vanes 7 configured to modify the direction of circulation of the flow in order to obtain an axial flow downstream of the rectifier 6.

Le profil des nacelles 2 est classiquement configuré pour former une tuyère à l'aval des redresseurs et accélérer et détendre le flux secondaire de manière à engendrer la poussée, la section du circuit secondaire 5 diminuant vers l'aval (dans le cas d'une tuyère convergente), puis peut éventuellement ré-augmenter dans le cas d'une tuyère convergente-divergente.The profile of the nacelles 2 is conventionally configured to form a nozzle downstream of the rectifiers and to accelerate and relax the secondary flow so as to generate the thrust, the section of the secondary circuit 5 decreasing downstream (in the case of a convergent nozzle), then can possibly increase again in the case of a convergent-divergent nozzle.

Dans une turbomachine à flux séparés chaque flux est éjecté par une tuyère. La tuyère (primaire comme secondaire) transforme l'énergie potentielle en énergie cinétique, c'est-à-dire qu'elle convertit la pression du flux en vitesse d'éjection, ce qui engendrera la poussée.In a split-flow turbomachine each flow is ejected through a nozzle. The nozzle (primary and secondary) transforms potential energy into energy kinetic, that is to say, it converts the pressure of the flow into the velocity of ejection, which will generate the thrust.

La tuyère du flux secondaire entoure et est classiquement placée à l'amont de la tuyère du flux primaire. La tuyère de flux primaire est délimitée par un cône dont la pointe est dirigée vers l'aval et par un carter annulaire présentant un bord de fuite orienté vers l'aval. Le cône et le carter définissent un circuit de section convergente ou convergente-divergente selon les choix d'architecture faits.The secondary flow nozzle surrounds and is conventionally placed upstream of the primary flow nozzle. The primary flow nozzle is delimited by a cone whose tip is directed downstream and by an annular casing having a trailing edge directed downstream. The cone and the casing define a circuit of convergent or convergent-divergent section according to the architectural choices made.

La tuyère secondaire est délimitée par un conduit appartenant au carter de soufflante (couramment appelé OFD ou OFS abrégé de l'anglais « Outer Fan Duct/Shroud ») et au carter de turbomachine (couramment appelé IFD ou IFS abrégé de l'anglais « Inner Fan Duct/Shroud »). Les deux carters définissent une section convergente ou convergente-divergente selon l'architecture du reste du moteur.The secondary nozzle is delimited by a duct belonging to the fan casing (commonly called OFD or OFS abbreviated to " Outer Fan Duct / Shroud ") and to the turbomachine casing (commonly called IFD or IFS abbreviated to " Inner Fan Duct / Shroud ”) . The two casings define a convergent or convergent-divergent section according to the architecture of the rest of the engine.

Cette diminution de section est classiquement située en aval du redresseur 6, de manière à accélérer le flux secondaire lorsqu'il s'écoule axialement, le flux secondaire étant ensuite éjecté autour du flux primaire.This decrease in section is conventionally located downstream of the rectifier 6, so as to accelerate the secondary flow when it flows axially, the secondary flow then being ejected around the primary flow.

Afin de gagner en rendement propulsif, on cherche à maximiser le taux de dilution, c'est-à-dire le rapport des débits massiques du flux secondaire et du flux primaire, et donc à minimiser le rapport de compression de la soufflante 1 pour une poussée donnée.In order to gain in propulsive efficiency, it is sought to maximize the dilution rate, that is to say the ratio of the mass flow rates of the secondary flow and of the primary flow, and therefore to minimize the compression ratio of the fan 1 for a thrust given.

L'augmentation du taux de dilution augmente le diamètre de la soufflante pour une même poussée ce qui entraine l'augmentation du volume et du poids du carter de soufflante. Donc pour limiter cet inconvénient on cherche à réduire le carter de soufflante 2 à son strict minimum, afin de réduire sa masse et les pertes de charge du circuit secondaire 5, l'effet des pertes de charges sur le flux secondaire étant d'autant plus important que le débit est grand, nécessaire pour un taux de dilution important, et la pression faible, nécessaire pour un faible rapport de compression de la soufflante 1.Increasing the bypass ratio increases the diameter of the fan for the same thrust, which leads to an increase in the volume and weight of the fan casing. So to limit this drawback, it is sought to reduce the fan casing 2 to its strict minimum, in order to reduce its mass and the pressure drops of the secondary circuit 5, the effect of the pressure drops on the secondary flow being all the more important that the flow rate is large, necessary for a high bypass rate, and the pressure low, necessary for a low compression ratio of the fan 1.

Ainsi, l'entrée d'air doit être extrêmement courte, et le carter de soufflante 2 doit être la plus courte possible après la sortie des aubes 7 de redresseur 6.Thus, the air inlet must be extremely short, and the fan casing 2 must be as short as possible after the exit of the vanes 7 of the stator 6.

Le document GB-2 546 422 divulgue un ensemble conforme au préambule de la revendication 1.The document GB-2 546 422 discloses an assembly according to the preamble of claim 1.

PRÉSENTATION GÉNÉRALE DE L'INVENTIONGENERAL PRESENTATION OF THE INVENTION

Un but de l'invention est de réduire les pertes de charges induites par le carter de soufflante.An object of the invention is to reduce the pressure drops induced by the fan casing.

Un autre but de l'invention est d'accélérer le flux secondaire.Another object of the invention is to accelerate the secondary flow.

Un autre but est de limiter les pertes de charges induites par le redresseur.Another object is to limit the load losses induced by the rectifier.

Un autre but de l'invention est d'augmenter le taux de dilution de la turbomachine.Another object of the invention is to increase the bypass ratio of the turbomachine.

Un autre but est de réduire le rapport de compression de la soufflante.Another object is to reduce the compression ratio of the fan.

Afin d'y parvenir, l'invention propose un ensemble pour turbomachine conforme à la revendication 1.In order to achieve this, the invention proposes an assembly for a turbomachine in accordance with claim 1.

Cela permet de redresser et d'accélérer le flux propulsé par la soufflante et transitant dans un canal d'écoulement.This makes it possible to straighten and accelerate the flow propelled by the fan and passing through a flow channel.

Avantageusement, l'invention peut être complétée par les caractéristiques suivantes, prises seules ou en combinaison :

  • le canal d'écoulement comprend d'amont en aval une portion d'admission se rétrécissant d'amont vers l'aval et une portion d'éjection s'évasant d'amont vers l'aval ;
  • la première aube présente une première surface, la deuxième aube présente une deuxième surface en regard de la première surface, la première surface se rapprochant de la deuxième surface d'amont vers l'aval ;
  • le carter de soufflante s'étend autour d'un axe longitudinal et comprend une extrémité aval formant bord de fuite, et dans lequel la portion d'éjection s'étend en aval du bord de fuite du carter de soufflante ; cela permet de ralentir le flux dans la portion d'éjection jusqu'à la vitesse de vol ;
  • une ligne de cambrure de chaque aube présente un point d'inflexion ;
  • chaque aube comprend un bord d'attaque, un bord de fuite opposé au bord d'attaque, et des parois d'intrados et d'extrados reliant le bord d'attaque au bord de fuite, et le canal d'écoulement présente, d'amont en aval dans le sens d'écoulement des fluides,
  • une section d'entrée s'étendant de la première aube à la deuxième aube en étant normale à une direction moyenne de l'écoulement et tangente au bord d'attaque de l'une des aubes et présentant une première aire,
  • une section d'éjection s'étendant de la première aube à la deuxième aube en étant normale à une direction moyenne de l'écoulement et présentant une deuxième aire, et
  • une section de sortie s'étendant de la première aube à la deuxième aube en étant normale à une direction moyenne de l'écoulement et tangente au bord de fuite d'au moins l'une des aubes et présent une troisième aire, la première aire étant supérieure à la deuxième aire, la deuxième aire étant inférieure à la troisième aire ;
  • le canal d'écoulement présente une section d'entrée définissant un plan normal à la direction d'écoulement du flux détourné par la soufflante, non parallèle à l'axe de la turbomachine, et une section d'éjection définissant un plan normal à l'axe de la turbomachine ;
  • le carter de soufflante se prolonge axialement au-delà du plan médian, le bord de fuite du carter de soufflante étant situé en aval du plan médian et en amont des bords de fuite des aubes, au niveau d'un plan de carénage.
Advantageously, the invention may be supplemented by the following characteristics, taken alone or in combination:
  • the flow channel comprises, from upstream to downstream, an inlet portion narrowing from upstream to downstream and an ejection portion widening from upstream to downstream;
  • the first vane has a first surface, the second vane has a second surface opposite the first surface, the first surface approaching the second surface from upstream to downstream;
  • the fan casing extends around a longitudinal axis and comprises a downstream end forming a trailing edge, and in which the ejection portion extends downstream from the trailing edge of the fan casing; this allows the flow in the ejection portion to be slowed down to flight speed;
  • a camber line of each blade has a point of inflection;
  • each blade comprises a leading edge, a trailing edge opposite the leading edge, and lower and upper surfaces connecting the leading edge to the trailing edge, and the flow channel has, d upstream downstream in the direction of fluid flow,
  • an inlet section extending from the first blade to the second blade while being normal to an average direction of the flow and tangent to the leading edge of one of the blades and having a first area,
  • an ejector section extending from the first vane to the second vane normal to an average flow direction and having a second area, and
  • an exit section extending from the first vane to the second vane while being normal to a mean direction of the flow and tangent to the trailing edge of at least one of the vanes and having a third area, the first area being greater than the second area, the second area being less than the third area;
  • the flow channel has an inlet section defining a plane normal to the flow direction of the flow diverted by the fan, not parallel to the axis of the turbomachine, and an ejection section defining a plane normal to the shaft of the turbomachine;
  • the fan casing extends axially beyond the median plane, the trailing edge of the fan casing being located downstream of the median plane and upstream of the trailing edges of the blades, at the level of a fairing plane.

Cela permet d'accélérer le flux dans une première portion du canal d'écoulement puis de ralentir le flux dans une deuxième portion de canal d'écoulement.This accelerates the flow in a first portion of the flow channel and then slows the flow in a second portion of the flow channel.

Selon un autre aspect, l'invention propose une turbomachine comportant un tel ensemble.According to another aspect, the invention proposes a turbomachine comprising such an assembly.

PRÉSENTATION DES FIGURESPRESENTATION OF FIGURES

D'autres caractéristiques et avantages de l'invention ressortiront encore de la description qui suit, laquelle est purement illustrative et non limitative, et doit être lue en regard des figures annexées sur lesquelles :

  • La figure 1 est un schéma en vue de coupe de profil d'une turbomachine comportant une nacelle et un redresseur de flux secondaire selon l'art antérieur ;
  • La figure 2 est un schéma en vue de coupe de profil qui représente un ensemble comportant une nacelle et un redresseur de flux secondaire selon l'invention ;
  • La figure 3 est une projection sur un plan d'une coupe réalisée à rayon constant de deux aubes adjacentes d'un redresseur selon l'invention.
Other characteristics and advantages of the invention will emerge from the description which follows, which is purely illustrative and not limiting, and must be read in conjunction with the appended figures in which:
  • The figure 1 is a diagram in profile sectional view of a turbomachine comprising a nacelle and a secondary flow rectifier according to the prior art;
  • The picture 2 is a diagram in profile sectional view which represents an assembly comprising a nacelle and a secondary flow rectifier according to the invention;
  • The picture 3 is a projection on a plane of a section made at a constant radius of two adjacent vanes of a stator according to the invention.

DESCRIPTION D'UN OU PLUSIEURS MODES DE MISE EN ŒUVRE ET DE RÉALISATIONDESCRIPTION OF ONE OR MORE MODES OF IMPLEMENTATION AND REALIZATION

L'invention s'applique à une turbomachine comprenant :

  • une virole 32 configurée pour délimiter intérieurement une veine de soufflante 5 d'un flux de gaz de ladite turbomachine,
  • un carter de soufflante 2, entourant radialement la virole 32 et délimitant avec la virole 32 la veine de soufflante 5,
  • un redresseur 6 comprenant une pluralité d'aubes 7 configurées pour redresser un flux secondaire circulant dans la veine de soufflante 5, dans lequel la pluralité d'aubes 7 comprend une première aube 7a et une deuxième aube 7b adjacente à la première aube 7a délimitant entre elles un canal d'écoulement 13, la première aube 7a et la deuxième aube 7b étant configurées pour redresser et accélérer le flux circulant dans le canal d'écoulement 13.
The invention applies to a turbomachine comprising:
  • a shroud 32 configured to internally delimit a fan stream 5 of a flow of gas from said turbomachine,
  • a fan casing 2, radially surrounding the shroud 32 and delimiting with the shroud 32 the fan duct 5,
  • a straightener 6 comprising a plurality of vanes 7 configured to straighten a secondary flow circulating in the fan stream 5, in which the plurality of vanes 7 comprises a first vane 7a and a second vane 7b adjacent to the first vane 7a delimiting between they a flow channel 13, the first vane 7a and the second vane 7b being configured to straighten and accelerate the flow circulating in the flow channel 13.

Le flux circulant ainsi dans le redresseur 6 est accéléré de telle sorte qu'il n'est plus nécessaire de former une tuyère en aval du redresseur 6 entre le carter de soufflante 2 et la virole 32.The flow thus circulating in the rectifier 6 is accelerated so that it is no longer necessary to form a nozzle downstream of the rectifier 6 between the fan casing 2 and the shroud 32.

Il est donc possible de raccourcir fortement le carter de soufflante 2, et donc de réduire sa masse, ou de permettre une augmentation de son diamètre tout en conservant une masse sensiblement similaire à un carter de soufflante 2 de l'art antérieur.It is therefore possible to greatly shorten the fan casing 2, and therefore to reduce its mass, or to allow an increase in its diameter while retaining a mass substantially similar to a fan casing 2 of the prior art.

Cela permet également de réduire les pertes de charges causées par le carter de soufflante 2.This also makes it possible to reduce the pressure drops caused by the fan casing 2.

Dans tout le texte de cette demande, les notions d'amont et d'aval sont définies dans le sens de l'écoulement des gaz dans la turbomachine.Throughout the text of this application, the notions of upstream and downstream are defined in the direction of the gas flow in the turbomachine.

La turbomachine s'étend selon un axe X de turbomachine, et les termes axial, radial et tangentiel se réfèrent à l'axe X de la turbomachine. Une direction axiale suit l'axe X de la turbomachine, une direction radiale est perpendiculaire à l'axe X de la turbomachine, et une direction tangentielle est orthogonale à une direction radiale et une direction axiale.The turbomachine extends along an axis X of the turbomachine, and the terms axial, radial and tangential refer to the axis X of the turbomachine. An axial direction follows axis X of the turbomachine, a radial direction is perpendicular to axis X of the turbomachine, and a tangential direction is orthogonal to a radial direction and an axial direction.

Dans le mode de réalisation représenté en figure 2, la turbomachine est une turbomachine à double flux comportant en outre une soufflante 1, logée dans le carter de soufflante 2, et mobile en rotation autour d'un axe longitudinal X, une virole interne 31 configurée pour délimiter une veine primaire 4 d'un flux de gaz primaire de la turbomachine, la virole 32 et le carter de soufflante 2 délimitant une veine dite secondaire d'écoulement d'un flux d'air propulsé par la soufflante 1.In the embodiment shown in figure 2 , the turbomachine is a dual-flow turbomachine further comprising a fan 1, housed in the fan casing 2, and rotatable about a longitudinal axis X, an inner shroud 31 configured to delimit a primary stream 4 of a primary gas flow from the turbomachine, the shroud 32 and the fan casing 2 delimiting a so-called secondary flow path for an air flow propelled by the fan 1.

Dans le mode de réalisation représenté, la virole 32 se situe dans le prolongement amont du carter 3 de la turbomachine.In the embodiment shown, the shroud 32 is located in the upstream extension of the casing 3 of the turbomachine.

Dans d'autres modes de réalisation, la virole 32 peut faire partie du carter 3, et ainsi former la portion amont du carter 3.In other embodiments, the shroud 32 can be part of the casing 3, and thus form the upstream portion of the casing 3.

La virole 32 et la virole interne 31 peuvent ne former qu'une pièce et former le bord d'attaque du carter 3.The shroud 32 and the inner shroud 31 may form a single piece and form the leading edge of the casing 3.

Dans le mode de réalisation représenté en figure 3, chaque aube 7 comprend un bord d'attaque 9, un bord de fuite 12 opposé au bord d'attaque 9, et des parois d'intrados 11 et d'extrados 10 reliant le bord d'attaque 9 au bord de fuite 12, et le canal d'écoulement 13 présente, d'amont en aval dans le sens d'écoulement des fluides :

  • une section d'entrée 14a s'étendant de la première aube 7a à la deuxième aube 7b en étant normale à une direction moyenne de l'écoulement et tangente au bord d'attaque 9 de l'une des aubes 7,
  • une section d'éjection 14b s'étendant de la première aube 7a à la deuxième aube 7b en étant normale à une direction moyenne de l'écoulement, et
  • une section de sortie 14c s'étendant de la première aube 7a à la deuxième aube 7b en étant normale à une direction moyenne de l'écoulement et tangente au bord de fuite 12 d'au moins l'une des aubes 7.
In the embodiment shown in picture 3 , each blade 7 comprises a leading edge 9, a trailing edge 12 opposite to the leading edge 9, and intrados 11 and extrados 10 walls connecting the leading edge 9 to the trailing edge 12, and the flow channel 13 presents, from upstream to downstream in the direction of fluid flow:
  • an inlet section 14a extending from the first blade 7a to the second blade 7b while being normal to an average direction of the flow and tangent to the leading edge 9 of one of the blades 7,
  • an ejection section 14b extending from the first vane 7a to the second vane 7b while being normal to an average flow direction, and
  • an outlet section 14c extending from the first vane 7a to the second vane 7b while being normal to an average direction of the flow and tangent to the trailing edge 12 of at least one of the vanes 7.

La section d'entrée 14a, la section d'éjection 14b et la section de sortie 14c s'étendant respectivement de la limite radialement intérieure à la limite radialement extérieure des aubes 7.The inlet section 14a, the ejection section 14b and the outlet section 14c extending respectively from the radially inner limit to the radially outer limit of the blades 7.

La section d'entrée 14a correspond ainsi à une section radiale du canal 13 d'écoulement qui coïncide avec le bord d'attaque 9 de la deuxième aube 7b, et la section d'éjection 14b correspond à une section radiale s'étendant en aval de la section d'entrée 14a.The inlet section 14a thus corresponds to a radial section of the flow channel 13 which coincides with the leading edge 9 of the second blade 7b, and the ejection section 14b corresponds to a radial section extending downstream of the input section 14a.

La section d'éjection 14b présente une surface inférieure à une surface de la section d'entrée 14a et inférieure à une surface de la section de sortie 14c.The ejection section 14b has a surface smaller than a surface of the inlet section 14a and smaller than a surface of the outlet section 14c.

Cette diminution de section du canal d'écoulement 13 permet d'accélérer le flux secondaire lorsqu'il circule dans le redresseur 6.This reduction in section of the flow channel 13 makes it possible to accelerate the secondary flow when it circulates in the rectifier 6.

Le canal d'écoulement 13 présente une section radiale 14 qui est définie comme un plan virtuel s'étendant de la paroi extrados 10a de la première aube 7a à la paroi intrados 11b de la deuxième aube 7b en étant normal à une direction moyenne de l'écoulement au niveau d'une ligne de courant centrale F et s'étendant sensiblement radialement par rapport à l'axe longitudinal X.The flow channel 13 has a radial section 14 which is defined as a virtual plane extending from the extrados wall 10a of the first blade 7a to the intrados wall 11b of the second blade 7b while being normal to an average direction of the flow at a central streamline F and extending substantially radially relative to the longitudinal axis X.

Il est entendu par ligne de courant centrale la ligne de courant située à équidistance de la première aube 7a et de la deuxième aube 7b.By central streamline is meant the streamline located equidistant from the first blade 7a and the second blade 7b.

La section radiale 14 du canal d'écoulement 13 présente une surface qui diminue progressivement entre la section d'entrée 14a et la section d'éjection 14b.The radial section 14 of the flow channel 13 has a surface which gradually decreases between the inlet section 14a and the ejection section 14b.

Plus précisément la section radiale 14 présente une largeur L définie comme une distance entre l'extrados 10a de la première aube 7a et l'intrados 11b de la deuxième aube 7b pour une distance à l'axe X constante, et dans lequel la largeur de la section radiale 14 est décroissante suivant la circulation de l'écoulement dans le canal d'écoulement 13 entre la section d'entrée 14a et la section d'éjection 14b.More specifically, the radial section 14 has a width L defined as a distance between the upper surface 10a of the first blade 7a and the lower surface 11b of the second blade 7b for a constant distance from the axis X, and in which the width of the radial section 14 is decreasing according to the circulation of the flow in the flow channel 13 between the inlet section 14a and the ejection section 14b.

Dit autrement, la paroi extrados 10a de la première aube 7a et la paroi intrados 11b de la deuxième aube 7b sont de plus en plus proches l'une de l'autre, pour une distance à l'axe X donnée, au fur et à mesure que le flux circule d'amont en aval dans le canal d'écoulement 13.In other words, the extrados wall 10a of the first blade 7a and the intrados wall 11b of the second blade 7b are increasingly close to each other, for a given distance from the X axis, as the as the flow travels from upstream to downstream in the flow channel 13.

Cela permet de faire diminuer la surface de la section radiale 14, ce qui permet d'engendrer une accélération du flux.This makes it possible to reduce the surface of the radial section 14, which makes it possible to generate an acceleration of the flow.

Cela permet notamment de faire diminuer la surface de la section radiale 14 tout en évitant des variations fortes du profil du carter de soufflante 2 et de la virole extérieure 32, de sorte que les perturbations et éventuels décollements aérodynamiques pouvant être engendrés par de telles variations sont évités.This makes it possible in particular to reduce the surface of the radial section 14 while avoiding strong variations in the profile of the fan casing 2 and of the outer shroud 32, so that the disturbances and possible aerodynamic separations which may be generated by such variations are avoided.

Dans le mode de réalisation représenté, une section radiale 14 présente une forme comparable à une portion angulaire d'un disque et présente une dimension dans une direction transversale et une dimension dans une direction radiale.In the embodiment shown, a radial section 14 has a shape comparable to an angular portion of a disc and has a dimension in a transverse direction and a dimension in a radial direction.

Dans la direction transversale, la section radiale 14 est délimitée par la première aube 7a et la deuxième aube 7b.In the transverse direction, the radial section 14 is delimited by the first vane 7a and the second vane 7b.

La distance séparant la première aube 7a et la deuxième aube 7b, la largeur L, est fonction de la distance à l'axe X de la turbomachine à laquelle la largeur L considérée. En effet la distance entre la première aube 7a et la deuxième aube 7b est croissante avec la distance à l'axe X.The distance separating the first blade 7a and the second blade 7b, the width L, is a function of the distance from the axis X of the turbomachine at which the width L considered. Indeed the distance between the first blade 7a and the second blade 7b increases with the distance to the X axis.

Il en résulte que la largeur d'une section radiale 14 est fonction du rayon ou d'une distance à l'axe X de la turbomachine, et augmente en fonction de la distance à l'axe X de la turbomachine.As a result, the width of a radial section 14 is a function of the radius or of a distance from the axis X of the turbomachine, and increases as a function of the distance from the axis X of the turbomachine.

Dans une direction radiale, la section radiale 14 est délimitée radialement intérieurement par la virole extérieure 32 et s'étend sur toute la hauteur d'une aube 7.In a radial direction, the radial section 14 is delimited radially internally by the outer shroud 32 and extends over the entire height of a blade 7.

La section radiale 14 présente une limite radialement intérieure et une limite radialement extérieure formant chacune sensiblement un arc de cercle.The radial section 14 has a radially inner limit and a radially outer limit each substantially forming an arc of a circle.

En déplaçant la section radiale 14 d'amont vers l'aval, la largeur L diminue, et optionnellement la dimension dans la direction radiale diminue également.By moving the radial section 14 from upstream to downstream, the width L decreases, and optionally the dimension in the radial direction also decreases.

Ainsi, la diminution de la section d'écoulement 14 entraîne une détente et donc une accélération du flux secondaire.Thus, the decrease in the flow section 14 causes relaxation and therefore an acceleration of the secondary flow.

Plus précisément, l'intrados 11a de la première aube 7a et l'extrados 10b de la deuxième aube 7b sont donc configurés pour que la largeur L d'une section radiale 14, pour une distance à l'axe X de la turbomachine donnée, décroisse au fur et à mesure du déplacement du flux vers l'aval.More specifically, the lower surface 11a of the first blade 7a and the upper surface 10b of the second blade 7b are therefore configured so that the width L of a radial section 14, for a given distance from the axis X of the turbomachine, decreases as the flow moves downstream.

Si on considère une section radiale 14 située en aval de la section d'entrée 14a, la largeur L de la section radiale 14 sera inférieure à la largeur de la section d'entrée 14a.If we consider a radial section 14 located downstream of the inlet section 14a, the width L of the radial section 14 will be less than the width of the inlet section 14a.

Il est évident que pour comparer la largeur de la section d'entrée 14a et la largeur L de la section radiale 14, il faut que ces deux valeurs soient exprimées pour un même rayon.It is obvious that in order to compare the width of the inlet section 14a and the width L of the radial section 14, these two values must be expressed for the same radius.

Cette largeur L peut optionnellement être la longueur d'un segment de droite joignant à mi-hauteur la première aube 7b et la deuxième aube 7a.This width L can optionally be the length of a straight segment joining at mid-height the first blade 7b and the second blade 7a.

Avantageusement, pour chaque section radiale 14 entre la section d'entrée 14a et la section d'éjection 14b, la longueur du segment de droite joignant à mi-hauteur la première aube 7b et la deuxième aube 7a diminue progressivement entre la section d'entrée 14a et la section d'éjection 14b.Advantageously, for each radial section 14 between the inlet section 14a and the ejection section 14b, the length of the straight segment joining the first vane 7b and the second vane 7a at mid-height gradually decreases between the inlet section 14a and the ejection section 14b.

La section d'éjection 14b présente la surface minimale pour une section radiale 14.The ejection section 14b has the minimum area for a radial section 14.

Dans le mode de réalisation représenté, la largeur L d'une section radiale 14 diminue en se déplaçant d'amont vers l'aval jusqu'à un plan médian 15, le plan médian 15 comportant donc la section d'éjection 14b.In the embodiment shown, the width L of a radial section 14 decreases as it moves from upstream to downstream as far as a median plane 15, the median plane 15 therefore including the ejection section 14b.

Dans le mode de réalisation représenté, le plan médian 15 est normal à l'axe X de la turbomachine, et délimite le canal d'écoulement 13 en deux parties, une portion amont ou d'admission 16 et une portion aval ou d'éjection 17.In the embodiment shown, the median plane 15 is normal to the axis X of the turbomachine, and delimits the flow channel 13 into two parts, an upstream or intake portion 16 and a downstream or ejection portion. 17.

Si la section radiale 14 est située dans la portion d'admission 16, la dimension transversale de la section radiale 14 est inférieure à la dimension transversale de la section d'entrée 14a et supérieure à la dimension transversale de la section d'éjection 14b.If the radial section 14 is located in the intake portion 16, the transverse dimension of the radial section 14 is less than the transverse dimension of the inlet section 14a and greater than the transverse dimension of the ejection section 14b.

Dit autrement, dans la portion d'admission 16, le canal d'écoulement 13 est convergent, la section radiale 14 présentant une surface décroissante d'amont vers l'aval.In other words, in the intake portion 16, the flow channel 13 is convergent, the radial section 14 having a decreasing surface from upstream to downstream.

Cela provoque une détente du flux traversant le canal d'écoulement 13, et incidemment une accélération du flux.This causes a relaxation of the flow passing through the flow channel 13, and incidentally an acceleration of the flow.

La portion d'admission 16 du canal d'écoulement 13 est configurée pour réaliser le travail de modification de la direction d'écoulement et de l'accélération du flux.Inlet portion 16 of flow channel 13 is configured to do the work of changing flow direction and flow acceleration.

Le canal d'écoulement 13 présente donc une section d'entrée 14a définissant un plan normal (ou orthogonal) à la direction d'écoulement du flux détourné par la soufflante, ce plan n'étant donc pas normal à l'axe X de la turbomachine, et une section d'éjection 14b définissant un plan normal à l'axe X de la turbomachine. Cela permet d'éjecter un flux circulant dans une direction sensiblement parallèle à l'axe de la turbomachine.The flow channel 13 therefore has an inlet section 14a defining a plane normal (or orthogonal) to the flow direction of the flow diverted by the fan, this plane therefore not being normal to the axis X of the turbomachine, and an ejection section 14b defining a plane normal to the axis X of the turbomachine. This makes it possible to eject a flow circulating in a direction substantially parallel to the axis of the turbomachine.

Dit autrement, la portion d'admission 16 redresse le flux tout en le détendant et en l'accélérant jusqu'à l'éjection au niveau du plan médian 15.In other words, the intake portion 16 straightens the flow while relaxing it and accelerating it until ejection at the level of the median plane 15.

La portion d'éjection 17 est configurée pour minimiser la traînée aérodynamique du redresseur 6.The ejection portion 17 is configured to minimize the aerodynamic drag of the rectifier 6.

L'angle d'incidence du profil par rapport au flux est faible de manière à éviter le décollement du flux d'air, tout en ayant une longueur la plus courte possible pour minimiser les frottements visqueux.The angle of incidence of the profile with respect to the flow is low so as to avoid the separation of the air flow, while having the shortest possible length to minimize viscous friction.

Une partie de la portion d'éjection 17 est située en aval du bord de fuite 8 du carter de soufflante 2. Ainsi, cela permet de ralentir le flux dans la portion d'éjection 17 jusqu'à la vitesse de vol.Part of the ejection portion 17 is located downstream of the trailing edge 8 of the fan casing 2. Thus, this makes it possible to slow down the flow in the ejection portion 17 down to flight speed.

Plus spécifiquement, en aval du plan médian 15 le profil des aubes 7 est configuré pour minimiser la traînée de chaque aube 7, les aubes 7 s'étendant donc axialement jusqu'à leur bord de fuite 12.More specifically, downstream of the median plane 15 the profile of the blades 7 is configured to minimize the drag of each blade 7, the blades 7 therefore extending axially to their trailing edge 12.

La section d'une aube 7 en aval du plan médian 15, plus particulièrement sa dimension dans la direction tangentielle, diminue vers l'aval jusqu'à son bord de fuite 12, la diminution de la dimension tangentielle de l'aube 7 étant configurée pour limiter les décollements aérodynamiques.The section of a blade 7 downstream of the median plane 15, more particularly its dimension in the tangential direction, decreases downstream as far as its trailing edge 12, the reduction in the tangential dimension of the blade 7 being configured to limit aerodynamic separations.

Ainsi, les flux transitant dans des canaux d'écoulement 13 situés côte à côte se rejoignent sans décollement aérodynamique.Thus, the flows transiting in the flow channels 13 located side by side meet without aerodynamic separation.

La section du canal d'écoulement 13 augmente donc vers l'aval dans la portion d'éjection 17.The section of the flow channel 13 therefore increases downstream in the ejection portion 17.

Optionnellement, les aubes 7 présentent une ligne de cambrure 71 qui peut comporter un point d'inflexion, la ligne de cambrure ou ligne moyenne étant définie en ce qu'elle s'étend du bord d'attaque 9 au bord de fuite 12 et qu'elle est à midistance de l'extrados 10 et de l'intrados 11.Optionally, the blades 7 have a line of camber 71 which may comprise a point of inflection, the line of camber or mean line being defined in that it extends from the leading edge 9 to the trailing edge 12 and that it is halfway between the extrados 10 and the intrados 11.

La ligne de cambrure 71 présente une inclinaison par rapport à l'axe X de la turbomachine correspondant à la giration du flux au bord d'attaque 9, et est sensiblement parallèle à l'axe moteur du plan médian 15 au bord de fuite 12.The line of camber 71 has an inclination with respect to the axis X of the turbomachine corresponding to the gyration of the flow at the leading edge 9, and is substantially parallel to the motor axis from the median plane 15 to the trailing edge 12.

Avantageusement, le plan médian 15, et donc la section d'éjection 14b, coïncide avec le bord de fuite 8 du carter de soufflante. La portion d'éjection 17 n'est donc pas carénée. Ainsi, la longueur du carter de soufflante 2 peut être réduite au minimum sans pénaliser le fonctionnement de la portion d'admission 16 qui est carénée par le carter de soufflante 2, ni le fonctionnement de la portion d'éjection 17 dont le seul rôle est de réduire la traînée.Advantageously, the median plane 15, and therefore the ejection section 14b, coincides with the trailing edge 8 of the fan casing. The ejection portion 17 is therefore not streamlined. Thus, the length of the fan casing 2 can be reduced to a minimum without penalizing the operation of the intake portion 16 which is streamlined by the fan casing 2, nor the operation of the ejection portion 17 whose sole role is to reduce drag.

Une portion des aubes 7, notamment le bord de fuite 12, se situe alors en aval du bord de fuite 8 du carter de soufflante 2, et n'est donc pas carénée.A portion of the blades 7, in particular the trailing edge 12, is then located downstream of the trailing edge 8 of the fan casing 2, and is therefore not streamlined.

Cela permet de minimiser la longueur du carter de soufflante 2, et de ce fait de minimiser les pertes de charge induites par le carter de soufflante 2.This makes it possible to minimize the length of the fan casing 2, and thereby to minimize the pressure drops induced by the fan casing 2.

Dans une variante, le carter de soufflante 2 peut se prolonger axialement au-delà du plan médian 15. Dans cette configuration, le bord de fuite 8 du carter de soufflante 2 est situé en aval du plan médian 15 et en amont des bords de fuite des aubes 12, au niveau d'un plan de carénage 18. Cette configuration permet de former un profil convergent puis divergent dans la partie carénée des canaux d'écoulement 13 (c'est à dire couverte par le carter de soufflante 2). Cela permet d'améliorer la performance en fonction du domaine de vol.In a variant, the fan casing 2 can extend axially beyond the median plane 15. In this configuration, the trailing edge 8 of the fan casing 2 is located downstream of the median plane 15 and upstream of the trailing edges blades 12, at the level of a fairing plane 18. This configuration makes it possible to form a converging then diverging profile in the streamlined part of the flow channels 13 (that is to say covered by the fan casing 2). This improves performance depending on the flight envelope.

Avantageusement, chaque paire d'aubes 7 adjacentes du redresseur 6 définit un canal d'écoulement 13 configuré pour redresser et accélérer le flux simultanément, les aubes du redresseur 6 définissant ainsi une pluralité de canaux d'écoulement 13 répartis circonférentiellement.Advantageously, each pair of adjacent vanes 7 of the straightener 6 defines a flow channel 13 configured to straighten and accelerate the flow simultaneously, the vanes of the straightener 6 thus defining a plurality of flow channels 13 distributed circumferentially.

Cela permet d'accélérer le flux de manière homogène sur toute la circonférence du redresseur 6.This allows the flow to be accelerated evenly around the entire circumference of the straightener 6.

Dans un tel ensemble, l'absence de tuyère formée par le carter de soufflante 2 et la virole 32 est compensée par l'effet de détente du redresseur 6, plus particulièrement par le travail de détente réalisé par la portion d'admission 16 des canaux d'écoulement 13.In such an assembly, the absence of a nozzle formed by the fan casing 2 and the ferrule 32 is compensated by the expansion effect of the rectifier 6, more particularly by the expansion work carried out by the inlet portion 16 of the channels flow 13.

Les pertes de charges sont réduites par la diminution de la longueur de le carter de soufflante 2 et le profil des aubes 7, plus particulièrement le bord de fuite 12 et le profil de la portion d'éjection 17 permettent de réduire la traînée et ainsi de limiter les décollements et les pertes de charges.The pressure losses are reduced by the reduction in the length of the fan casing 2 and the profile of the blades 7, more particularly the trailing edge 12 and the profile of the ejection portion 17 make it possible to reduce the drag and thus to limit separations and pressure drops.

Un tel ensemble permet donc de redresser et accélérer le flux transitant dans les canaux d'écoulement 13, à la différence d'éléments de déviation de flux classiques.Such an assembly therefore makes it possible to straighten and accelerate the flow passing through the flow channels 13, unlike conventional flow deflection elements.

Les redresseurs classiques redressent le flux et le ralentissent.Conventional straighteners straighten the flow and slow it down.

Les distributeurs classiques accélèrent le flux tout en le déviant, c'est-à-dire que le flux arrive dans le distributeur avec une direction d'écoulement sensiblement parallèle à l'axe X de la turbomachine et sort du distributeur avec une direction d'écoulement inclinée par rapport à l'axe de la turbomachine.Conventional distributors accelerate the flow while deviating it, that is to say the flow arrives in the distributor with a direction of flow substantially parallel to the axis X of the turbomachine and leaves the distributor with a direction of flow inclined with respect to the axis of the turbomachine.

Les tuyères classiques forment un canal convergent qui accélère le flux sans le dévier.Conventional nozzles form a converging channel which accelerates the flow without deflecting it.

En outre, le profil des aubes 7 se terminant par un bord de fuite permet d'éviter le décollement de flux à la sortie de l'ensemble.In addition, the profile of the blades 7 ending in a trailing edge makes it possible to avoid flow separation at the outlet of the assembly.

Claims (9)

  1. Assembly for a turbomachine extending along an axis (X) and comprising :
    - a shroud (32) configured to delimit a fan flow path (5) of a gas flow of said turbomachine,
    - a fan casing (2), radially surrounding the shroud (32) and delimiting with the shroud (32) the fan flow path (5), and
    - outlet guide vanes comprising a plurality of vanes (7) configured to straighten a secondary flow circulating in the fan flow path (5), wherein the plurality of vanes (7) comprises a first vane (7a) and a second vane (7b) adjacent to the first vane (7a) delimiting between them a flow channel (13) configured to straighten the flow,
    characterized in that
    the flow channel (13) is convergent and configured to accelerate the flow by means of an inlet section (14a) lying in a plane that is not perpendicular to the axis of the turbomachine and an outlet section lying in a plane (14b) that is perpendicular to the axis (X) of the turbomachine, the first vane (7a) and the second vane (7b) each having a downstream, unshrouded part forming a trailing edge.
  2. The turbomachine assembly of claim 1, wherein the flow channel (13) comprises from upstream to downstream an inlet portion (16) narrowing from upstream to downstream and an ejection portion (17) flaring from upstream to downstream.
  3. The turbomachinery assembly of any of claims 1 or 2, wherein the first vane (7a) has a first surface, the second vane (7b) has a second surface facing the first surface, the first surface approaching the second surface from upstream to downstream.
  4. The turbomachine assembly of claim 1-3, wherein the fan case (2) extends about a longitudinal axis (X) and includes a trailing edge downstream end (8), and wherein the ejection portion (17) extends downstream of the trailing edge (8) of the fan casing (2).
  5. The turbomachinery assembly of any of claims 1 to 4, wherein a camber line of each vane (7) has an inflection point.
  6. An assembly according to any one of claims 1 to 5, wherein each vane (7) comprises a leading edge (9), a trailing edge (12) opposite the leading edge (9), and pressure (11) and suction (10) walls connecting the leading edge (9) to the trailing edge (12), and the flow channel (13) has, from upstream to downstream in the direction of fluid flow,
    - an inlet section (14a) extending from the first vane (7a) to the second vane (7b), being normal to a mean direction of flow and tangent to the leading edge (9) of one of the vanes (7) and having a first area ,
    - an ejection section (14b) extending from the first vane (7a) to the second vane (7b), being normal to a mean direction of flow and having a second area, and
    - an exit section (14c) extending from the first vane (7a) to the second vane (7b), being normal to a mean direction of flow and tangent to the trailing edge (12) of at least one of the vanes (7) and having a third area,
    and wherein the first area is greater than the second area, the second area being less than the third area.
  7. An assembly according to any of claims 1 to 6, wherein the flow channel (13) has an inlet section (14a) defining a plane normal to the flow direction of the flow diverted by the fan, not parallel to the axis (X) of the turbomachine, and an ejection section (14b) defining a plane normal to the axis (X) of the turbomachine.
  8. Assembly according to one of claims 1 to 7, in which the fan case (2) extends axially beyond the median plane (15), the trailing edge (8) of the fan casing (2) being located downstream of the median plane (15) and upstream of the trailing edges (12) of the vanes, at the level of a fairing plane (18).
  9. A turbomachine comprising a turbomachine assembly according to any of claims 1 to 8.
EP20725890.6A 2019-03-15 2020-03-12 Secondary flow rectifier with integrated pipe Active EP3938626B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1902662A FR3093756B1 (en) 2019-03-15 2019-03-15 secondary flow rectifier has integrated nozzle
PCT/FR2020/050524 WO2020188197A2 (en) 2019-03-15 2020-03-12 Secondary flow rectifier with integrated pipe

Publications (2)

Publication Number Publication Date
EP3938626A2 EP3938626A2 (en) 2022-01-19
EP3938626B1 true EP3938626B1 (en) 2022-11-16

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EP20725890.6A Active EP3938626B1 (en) 2019-03-15 2020-03-12 Secondary flow rectifier with integrated pipe

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US (1) US11434773B2 (en)
EP (1) EP3938626B1 (en)
CN (1) CN114286886B (en)
CA (1) CA3130189A1 (en)
FR (1) FR3093756B1 (en)
WO (1) WO2020188197A2 (en)

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US7730715B2 (en) * 2006-05-15 2010-06-08 United Technologies Corporation Fan frame
US8016561B2 (en) * 2006-07-11 2011-09-13 General Electric Company Gas turbine engine fan assembly and method for assembling to same
US20080159856A1 (en) * 2006-12-29 2008-07-03 Thomas Ory Moniz Guide vane and method of fabricating the same
US9957918B2 (en) * 2007-08-28 2018-05-01 United Technologies Corporation Gas turbine engine front architecture
FR2961565B1 (en) * 2010-06-18 2012-09-07 Snecma AERODYNAMIC COUPLING BETWEEN TWO ANNULAR ROWS OF AUBES FIXED IN A TURBOMACHINE
FR3032480B1 (en) * 2015-02-09 2018-07-27 Safran Aircraft Engines AIR RECOVERY ASSEMBLY WITH IMPROVED AERODYNAMIC PERFORMANCE
FR3032495B1 (en) * 2015-02-09 2017-01-13 Snecma RECOVERY ASSEMBLY WITH OPTIMIZED AERODYNAMIC PERFORMANCE
FR3046811B1 (en) * 2016-01-15 2018-02-16 Snecma DAUGHTER OUTPUT DIRECTOR FOR AIRCRAFT TURBOMACHINE, HAVING AN IMPROVED LUBRICANT COOLING FUNCTION
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Also Published As

Publication number Publication date
FR3093756A1 (en) 2020-09-18
WO2020188197A3 (en) 2020-11-26
WO2020188197A2 (en) 2020-09-24
CN114286886A (en) 2022-04-05
US20220186624A1 (en) 2022-06-16
FR3093756B1 (en) 2021-02-19
EP3938626A2 (en) 2022-01-19
CA3130189A1 (en) 2020-09-24
US11434773B2 (en) 2022-09-06
CN114286886B (en) 2024-05-10

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