ASSEMBLY FOR A DOUBLE-FLOW AIRCRAFT TURBOMACHINE, THE ASSEMBLY BEING EQUIPPED WITH INCIDENCE-DRIVEN AIR DISCHARGE VINES

Abstract

The invention relates to an assembly (32) for a dual-flow aircraft turbomachine, comprising: - an inner shroud (42) configured to externally delimit a primary gas flow path (16) of the turbomachine; - an outer shroud (40) configured to internally delimit a secondary gas flow path (18) of the turbomachine; - at least one air discharge duct (46) extending between the inner shroud (42) and the outer shroud (40), the air discharge duct (46) opening into the secondary flow path (18) through an outlet orifice (52) equipped with discharge fins (54). According to the invention, at least some of the discharge fins (54) are mounted to be movable so as to be able to be controlled in incidence between a propulsion position and a thrust reversal position. Figure for the abstract: Figure 1.

Description

Title of the invention: ASSEMBLY FOR A DOUBLE-FLOW AIRCRAFT TURBOMACHINE, THE ASSEMBLY BEING EQUIPPED WITH PILOTED AIR DISCHARGE VINES IMPACT Technical field

[0001] The present invention relates to the field of dual-flow aircraft turbomachines. It relates more particularly to compressor air discharge systems, comprising discharge vanes intended to direct the air taken from the secondary stream of the turbomachine. Such discharge vanes are for example known from document WO2016 / 156739. STATE OF PRIOR ART

[0002] On conventional twin-flow aircraft turbomachines, such as twin-flow and twin-spool turbojets, it is known to implement low-pressure compressor air discharge systems. These systems, which implement compressor air discharge valves, also called VBV valves (from the English "Variable Bleed Valve"), allow the diversion of a portion of the primary flow towards the secondary flow, at the outlet of the low-pressure compressor. The function of these discharge systems lies in adapting the air flow at the outlet of the low-pressure compressor, to ensure the operation of the high and low-pressure compressors in their respective domains, and thus avoid harmful phenomena such as surge. Another function consists of evacuating hailstones by centrifugation at the outlet of the low-pressure compressor.

[0003] Discharge systems usually comprise several air discharge ducts for introducing the air taken from the low-pressure compressor outlet into the secondary stream. At an outlet orifice of each of these ducts, discharge fins are provided having a pitch angle adapted to direct the flow of air taken with an axial component downstream, when it is introduced into the secondary stream.

[0004] Thanks to this specific orientation of the flow of air taken, it mixes easily with the secondary flow, flowing in the secondary vein from upstream to downstream in normal propulsion configuration.

[0005] Furthermore, certain aircraft turbomachine architectures provide for using the secondary vein as a flow channel for counter-thrust gases. These are, for example, turbomachines with a ducted fan, and whose fan blades rotary engines are variable-pitch. Such turbomachines, with a very high bypass ratio, are for example known as UHBR turbomachines (from the English "Ultra High By-pass Ratio"), and their bypass ratio can be greater than 15. This concept, commonly called VPF fan (from the English "Variable Pitch Fan"), is intended to further reduce the compression ratio of the fans of conventional UHBR turbomachines, thanks to an increase in the bypass ratio. This specificity provides a very significant gain in fuel consumption, and provides an extended operability range between the take-off and cruise phases.

[0006] Nevertheless, this type of high bypass ratio design results in a particularly large outside diameter for the turbomachine, which requires significant efforts in the dimensioning of the nacelle, in order to reduce its mass and drag. A consequence of the reduction in size of the nacelle may be the impossibility of installing a conventional thrust reversal system there. In such a case, the variable-pitch fan is also used to reverse the direction of flow of the secondary flow in the secondary vein, and thus create the counter-thrust in the reverse direction, necessary for the deceleration of the aircraft during landing.

[0007] In the thrust reversal configuration, the reverse flow, mainly located in a head region of the fan, bypasses the lips of the nacelle (corresponding to the air inlet in the conventional propulsion configuration), then flows over the external surface of the nacelle before being reintroduced into the turbomachine, via the secondary nozzle.

[0008] In this thrust reversal configuration, the quality of the flow in the opposite direction in the secondary vein is essential to ensure a maximum level of counter-thrust. However, in the secondary vein, this flow of counter-thrust gas is disturbed by the counter-current introduction of the discharge air flows taken from the outlet of the low-pressure compressor. Statement of the invention

[0009] To address the drawback mentioned above, relating to the embodiments of the prior art, the invention firstly relates to an assembly for a dual-flow aircraft turbomachine, comprising: - an internal shell configured to externally delimit a primary gas flow vein of the turbomachine; - an external shell configured to internally delimit a secondary gas flow vein of the turbomachine; - at least one air discharge duct extending between the inner shell and the outer shell, the air discharge duct opening into the secondary vein through an outlet orifice equipped with discharge fins.

[0010] According to the invention, at least some of said discharge fins are mounted to be mobile so as to be able to be controlled in incidence between a propulsion position and a thrust reversal position.

[0011] Thus, thanks to the incidence control of the discharge vanes, the assembly according to the invention is able to reject the flow coming from the discharge duct in the same axial direction as the counter-thrust gas flow which flows in the secondary vein. In other words, the invention makes it possible to adapt the orientation of the discharge vanes depending on whether the turbomachine is in normal propulsion or counter-thrust configuration, so that the discharge air flow can always be introduced in axial co-current into the secondary vein.

[0012] The invention preferably provides at least any one of the following optional features, taken alone or in combination.

[0013] Preferably, in the propulsion position, the discharge fins have first acute pitch angles, oriented in a first direction, and in the thrust reversal position, the discharge fins have second acute pitch angles, oriented in a second direction opposite to the first.

[0014] Preferably, in the thrust reversal position, the movable discharge fins of the same discharge duct have second acute pitch angles of the same value, or substantially of the same value.

[0015] Alternatively, in the thrust reversal position, the movable discharge fins of the same discharge duct have second acute pitch angles with at least two different values, the two end fins located respectively closest to the two opposite axial ends of the outlet orifice of the discharge duct, preferably having second acute pitch angles of values ​​greater than that of at least one other movable discharge fin located axially between the two end fins. This differentiation in the value of the second acute pitch angles makes it possible to locally manage the orientation of the flow taken, at the level of the critical zones located at the upstream and downstream ends of the outlet orifice of the discharge duct.

[0016] Preferably, the assembly comprises means for controlling the incidence of the movable discharge fins, the control means being at least partly arranged between the inner and outer ferrules of the assembly.

[0017] Preferably, each air discharge conduit opens into the primary vein through an inlet orifice equipped with a compressor air discharge valve.

[0018] The invention also relates to a dual-flow aircraft turbomachine, comprising an assembly as described above, the latter being able for example to be an intermediate casing hub of the turbomachine.

[0019] Preferably, the turbomachine comprises a fan, with blades of variable-pitch rotary blowers.

[0020] Preferably, the turbomachine comprises a fan drive reducer.

[0021] The invention finally relates to a method for controlling such an aircraft turbomachine, comprising a step of switching from a propulsion configuration to a thrust reversal configuration, this step comprising an operation of modifying the incidence of the movable discharge vanes, consisting of moving these vanes from their propulsion position to their thrust reversal position.

[0022] Other advantages and characteristics of the invention will appear in the detailed non-limiting description below. Brief description of the drawings

[0023] This description will be made with regard to the attached drawings among which;

[0024] [Fig-1] represents a schematic side view of a turbojet according to the invention;

[0025] [Fig.2] represents an enlarged, more detailed view of an assembly forming part of integral part of the turbomachine shown in the preceding figure, the assembly being presented according to a first preferred embodiment of the invention, with its discharge fins in the propulsion position;

[0026] [Fig.3] is a view similar to that of [Fig.2], with discharge fins shown in the reverse thrust position;

[0027] [Fig.4] is a view similar to that of [Fig.2], with the assembly being presented according to another preferred embodiment of the invention; and

[0028] [Fig.5] is a view similar to that of [Fig.2], with the assembly being presented according to yet another preferred embodiment of the invention. DETAILED DISCLOSURE OF PREFERRED EMBODIMENTS

[0029] With reference to [Fig. 1], a twin-spool turbojet engine 1 is shown, having a very high bypass ratio, for example greater than 15. The turbojet engine 1 conventionally comprises a gas generator 2 on either side of which are arranged a low-pressure compressor 4 and a low-pressure turbine 12, this gas generator 2 comprising a high-pressure compressor 6, a combustion chamber 8 and a high-pressure turbine 10. Subsequently, the terms “upstream” and “downstream” are considered according to a main direction 14 of gas flow within the turbojet engine, when the latter is in normal propulsion configuration.

[0030] The low-pressure compressor 4 and the low-pressure turbine 12 form a low-pressure body, and are connected to each other by a low-pressure shaft 11 centered on a longitudinal central axis 3 of the turbojet. Similarly, the high-pressure compressor 6 and the high-pressure turbine 10 form a high-pressure body, and are connected to each other by a high-pressure shaft 13 also centered on the axis 3, and arranged around the low pressure shaft 11.

[0031] The turbojet 1 also comprises, upstream of the gas generator 2 and the low-pressure compressor 4, a single fan 15 which is here arranged directly at the rear of an air inlet cone of the engine. The fan 15 comprises a ring of fan blades 17 rotating around the axis 3, this ring being surrounded by a fan casing 9. In this preferred embodiment of the invention, the fan blades 17 are variable pitch, that is to say that their incidence can be controlled by a control mechanism 20 arranged at least partly in the inlet cone, and designed to pivot these blades 17 around their respective longitudinal axes 22. This control mechanism 20, of known design of the mechanical, electrical, hydraulic, and / or pneumatic type, is itself controlled by an electronic control unit (not shown), which makes it possible to order the value of the pitch angles of the blades 17 according to the needs encountered.

[0032] The blower 15, of the VPF type, is not driven directly by the low pressure shaft 11, but only driven indirectly by this shaft, via a drive reducer 24, which allows it to rotate at a slower speed. Nevertheless, a solution with direct drive of the blower 15, by the low pressure shaft 11, falls within the scope of the invention.

[0033] Furthermore, the turbojet 1 defines a primary vein 16 intended to be crossed by a primary flow 16a, as well as a secondary vein 18 intended to be crossed by a secondary flow 18a located radially outwards relative to the primary flow, the flow of the fan therefore being divided at the level of a flow separation nozzle 26.

[0034] As is known to those skilled in the art, the secondary vein 18 is delimited radially outwards in part by an external shroud 23, preferably metallic, extending the fan casing 9 towards the rear. In addition, downstream of the fan 15, in the secondary vein 18, a ring of guide vanes is provided which are here outlet guide vanes 30 (or OGV, from the English “Outlet Guide Vane”). These stator vanes 30 connect the outer shell 23 to an intermediate casing hub 32 surrounding the low-pressure compressor 4. These stator vanes 30 are circumferentially spaced from each other, and allow the secondary flow to be straightened after it has passed through the fan 15. In addition, these vanes 30 can also fulfill a structural function, by ensuring the transfer of forces coming from the reducer 24 and the rolling bearings of the motor shafts and the fan hub, to the outer shell 23.Then, these forces can pass through an engine attachment (not shown) fixed on the shroud 23, and connecting the turbojet to a mounting pylon of the aircraft. In other words, the assembly 32 forms the hub of an intermediate casing, the latter being completed by radial arms formed by the stator blades 30, and also completed by the ex- shroud. inner 23 on which the heads of these blades 30 are fixed.

[0035] The assembly 32 also fulfills the function of an inter-vein compartment, being produced at the feet of the stator blades 30, but can also extend downstream as shown diagrammatically in [Fig.l]. This assembly comprises an external shroud 40 configured to internally delimit a portion of the secondary gas flow vein 18, as well as an internal shroud 42 configured to externally delimit the primary gas flow vein 16. The two shrouds 40, 42 extend downstream from the separation nozzle 26, which connects them. Between these two shrouds 40, 42, an inter-vein compartment 44 is effectively delimited, in which various components and equipment of the turbojet can be arranged.

[0036] Downstream of the stator blades 30, the assembly 32 is equipped with a plurality of air discharge ducts 46, distributed around the axis 3. Each discharge duct 46 extends generally radially, possibly with an axial component going downstream, going from the inner shell 42 to the outer shell 40, so as to be able to communicate the primary vein 16 with the secondary vein 18. More precisely, each air discharge duct 46 opens into the primary vein 16 through an inlet orifice 48 equipped with a discharge valve VBV 50, the inlet orifice 48 being arranged axially between the low-pressure compressor 4 and the high-pressure compressor 6. Similarly, each air discharge duct 46 opens into the secondary vein 18, through an outlet orifice 52 equipped with discharge fins 54.

[0037] The discharge fins 54 here have the specific feature of being controlled in incidence in the outlet orifice 52, so as to adapt to the configuration of the turbojet engine. These fins 54 are thus capable of being moved from a propulsion position shown in full gear in [Fig.l], to a thrust reversal position shown in dotted lines in this same figure, and vice versa. The objective thus lies in adapting the pitch angle of these fins 54, so that the discharge air flow coming from the low pressure compressor 4 can be introduced axially cocurrently into the secondary vein 18, whether the turbojet engine is in normal propulsion configuration, or in thrust reversal configuration.

[0038] The incidence control of the discharge fins 54 is carried out using control means 58, of known design of the mechanical, electrical, hydraulic, and / or pneumatic type. These control means 58 are preferably arranged, at least in part, in the inter-vein space 44, and they are themselves controlled by an electronic control unit (not shown), which makes it possible to order the required fin position according to the configuration of the turbojet.

[0039] [Fig. 2] shows the discharge fins 54 in their propulsion position, allowing the introduction of the discharge air flow 60 in axial co-current with the flow of secondary air 18a circulating in the secondary vein 18 of the turbojet in normal propulsion configuration. To do this, the fins 54 have first acute pitch angles Al, oriented in a first direction, corresponding to the counterclockwise direction in [Fig.2]. The pitch angle of each fin 54 is defined, in a conventional manner, between a discharge plane P tangent or substantially tangent to the external shell 40 at the outlet orifice 52, and a direction of the chord 62 of a profile of this fin 54.

[0040] Here, the first acute pitch angles A1 are for all the fins 54 of identical or substantially identical value, and for example between 20 and 70°. The inclination chosen is therefore such that the discharge air flow 60, coming from the primary flow 16a, is introduced into the secondary vein 18 from the orifice 52 in an introduction direction comprising an axial component going from upstream to downstream, in relation to the main flow direction 14. This direction of introduction of the discharge air flow 60 corresponds here to the chord direction 62 of the discharge fins 54, due to their preferentially symmetrical shape.

[0041] The change of position of the discharge fins 54 is carried out via the dedicated control means 58. In this regard, it is noted that these means 58 can individually control the incidence of each of the fins, or control them in groups, or even control them all simultaneously by means of connecting mechanisms connecting these different fins of the same conduit 46. Moreover, it is noted that in the outlet orifice 52 of each discharge conduit 46, the fins 54 can be arranged within one or more fin grids (not shown) partially closing this orifice 52. The change of position consists of pivoting the fins 54 according to their incidence setting axes 64, which are preferably axes oriented in a tangential direction in relation to the axis 3, that is to say axes parallel to the discharge plane P, or inscribed in this last.

[0042] [Fig. 3] shows the discharge fins 54 in their thrust reversal position, after their pivoting caused by the control means 58 around the incidence setting axes 64. In this thrust reversal position of the discharge fins 54, these make it possible to introduce the discharge air flow 60 in axial co-current with a counter-thrust gas flow 18a' circulating in the secondary vein 18 of the turbojet in thrust reversal configuration. To do this, the fins 54 have second acute setting angles A2, oriented in a second direction opposite to the first, and corresponding to the clockwise direction in [Fig. 3]. This setting angle of each fin 54 is also defined, in a conventional manner, between the discharge plane P tangent or substantially tangent to the external shell 40 at the level of the outlet orifice 52, and the direction of the chord 62 of a profile of this fin 54.

[0043] The second acute pitch angles A2 are here, for all the fins 54, of identical or substantially identical value, and for example between 20 and 70°. The inclination chosen is therefore such that the discharge air flow 60, coming from the primary flow 16a, is introduced into the secondary vein 18 from the orifice 52 in an introduction direction comprising an axial component going from downstream to upstream, in relation to the main flow direction 14. This direction of introduction of the discharge air flow 60 corresponds here to the chord direction 62 of the discharge fins 54, due to their preferentially symmetrical shape.

[0044] Other preferred embodiments of the invention are conceivable, such as providing discharge fins 54 of non-symmetrical shape, for example adopting a blade profile as shown in [Fig.4].

[0045] Another mode is shown in [Fig. 5], in which in the thrust reversal position, the fins 54 of the same discharge duct 46 no longer have second acute pitch angles A2 of the same value or substantially the same value, but on the contrary have at least two different values. In this mode, the two fins 54 which have the largest value for the second acute pitch angle A2 are the two end fins, located respectively at the two opposite axial ends of the outlet orifice 52. The value of this angle A2 is therefore lower for the other fins 54 located more in the center, with possibly a progressive reduction in this angle value going towards the center of the outlet orifice 52.

[0046] Alternatively, the largest value for the second acute setting angle A2 could be applied only to the upstream end fin 54, to ensure bypassing of the critical zone constituted by the upstream end portion of the spout-shaped outlet orifice 52. Indeed, in this critical zone shown in [Fig. 5], too great an axial drawdown of the discharge air flow 60 could prove inappropriate for bypassing the spout, hence the advantage of locally raising the value of the angle A2 for the discharge fin located directly axially opposite this spout.

[0047] Of course, various modifications may be made by those skilled in the art to the invention which has just been described, solely by way of non-limiting examples and the scope of which is defined by the appended claims. In particular, the various preferred embodiments described above can be combined with each other.

Claims

Claims

1. Assembly (32) for a dual-flow aircraft turbomachine, comprising: - an internal shroud (42) configured to externally delimit a primary gas flow duct (16) of the turbomachine; - an external shroud (40) configured to internally delimit a secondary gas flow duct (18) of the turbomachine; - at least one air discharge duct (46) extending between the inner shell (42) and the outer shell (40), the air discharge duct (46) opening into the secondary stream (18) through an outlet orifice (52) equipped with discharge fins (54), characterized in that at least some of said discharge fins (54) are mounted to be movable so as to be able to be controlled in incidence between a propulsion position and a thrust reversal position, in that the assembly is designed to bring the discharge fins (54) into the propulsion position, so that the discharge air flow (60),from the primary flow (16a), is introduced into the secondary flow path (18) from the outlet orifice (52) in an introduction direction comprising an axial component going from upstream to downstream, in axial co-current with the secondary air flow (18a) circulating in the secondary flow path (18) of the turbomachine in normal propulsion configuration, and in that the assembly is also designed to bring the discharge fins (54) into the thrust reversal position, so that the discharge air flow (60), from the primary flow (16a), is introduced into the secondary flow path (18) from the outlet orifice (52) in an introduction direction comprising an axial component going from downstream to upstream, in axial co-current with a counter-thrust gas flow (18a') circulating in the secondary flow path (18) of the turbomachine in thrust reversal configuration.,

2. Assembly according to claim 1, characterized in that in the propulsion position, the discharge fins (54) have first acute pitch angles (A1), oriented in a first direction, and in that in the thrust reversal position, the discharge fins (54) have second acute pitch angles (A2), oriented in a second direction opposite to the first.

3. Assembly according to claim 2, characterized in that in the position thrust reversal, the movable discharge fins (54) of the same discharge duct (46) have second acute pitch angles (A2) of the same value, or substantially of the same value.

4. Assembly according to claim 2, characterized in that in the thrust reversal position, the movable discharge fins (54) of the same discharge duct (46) have second acute pitch angles (A2) with at least two different values, the two end fins (54) located respectively closest to the two opposite axial ends of the outlet orifice (52) of the discharge duct (46), preferably having second acute pitch angles (A2) of values ​​greater than that of at least one other movable discharge fin (54) located axially between the two end fins.

5. Assembly according to any one of the preceding claims, characterized in that it comprises means (58) for controlling the incidence of the movable discharge fins (54), the control means (58) being at least partly arranged between the internal and external ferrules (42, 40) of the assembly.

6. A dual-flow aircraft turbomachine (1), comprising an assembly (32) according to any one of the preceding claims.

7. Turbomachine according to claim 6, characterized in that it comprises a fan (15), with variable-pitch rotary fan blades (17).

8. Turbomachine according to claim 6 or 7, characterized in that it comprises a drive reducer (24) for the fan (15).

9. Method for controlling an aircraft turbomachine (1) according to any one of claims 6 to 8, characterized in that it comprises a step of switching from a propulsion configuration to a thrust reversal configuration, this step comprising an operation of modifying the incidence of the movable discharge vanes (54), consisting of moving these vanes from their propulsion position to their thrust reversal position, so that the discharge air flow (60), coming from the primary flow (16a), is introduced into the secondary vein (18) from the outlet orifice (52) in an introduction direction comprising an axial component going from downstream to upstream, in axial co-current with a counter-thrust gas flow (18a') circulating in the secondary vein (18) of the turbomachine in thrust reversal configuration.

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