TURBOMACHINE WITH IMPROVED PRIMARY FLOW FEED

Abstract

Turbomachine (1) comprising an air inlet (2) provided with a fan (3) mounted for rotation relative to a nacelle (9) about a main axis (AX) and intended to be traversed by an inlet airflow (100), the fan (3) comprising a plurality of variable pitch blades (11), and the nacelle (9) comprising an intermediate casing (20) defined by an inner fairing (21) and an outer fairing (22) which meet at a separating nozzle (23) of the air inlet (2) into a primary flow (30) for circulation of a primary flow (FI) and a secondary flow (31) for circulation of a secondary flow (FII), the primary flow (30) comprising a low-pressure compressor (4) connected to a low-pressure turbine (8), the fan (3) being functionally connected to the low-pressure turbine (8),The turbomachine (1) comprises a conduit which includes an outlet opening into the primary flow and an inlet opening into the secondary flow (31) to fluidly connect the secondary flow (31) and the primary flow (30), the conduit being provided with isolation means allowing the conduit to be selectively opened or closed. Method of feeding a turbomachine (1). Figure for the abstract: Fig. 1,

Description

Title of the invention: IMPROVED PRIMARY FLOW TURBOMACHINE Technical field of the invention

[0001] The present invention relates to the field of turbomachinery, and more specifically to turbomachinery equipped with a thrust reversal device by adjusting the fan pitch.

[0002] The invention applies to all turbomachine designs, for example, turbojets with a fan driven directly by a low-pressure unit, indirectly by a reduction gear, single-spool, twin-spool, single-flow, and twin-flow turbojets. The turbomachines of the invention may have shrouded (with a nacelle around it) or unshrouded fans, and may be counter-rotating or not (in English, propfan, open rotor, or contra-rotating open rotor). Prior art

[0003] In a turbomachine, here a central axis AX turbojet, air is admitted, following a longitudinal direction parallel to the AX axis, into an inlet sleeve to pass through a fan comprising a series of rotating blades before splitting into a central primary flow which circulates in a so-called primary airflow circulation channel and a secondary flow surrounding the primary flow.

[0004] The primary flow is compressed by stages of compressors before reaching a combustion chamber, after which it expands as it passes through turbines, before being discharged, generating thrust. The secondary flow, on the other hand, is propelled directly by the blower to generate the main thrust.

[0005] The turbojet also includes a nacelle that supports the turbojet components and provides the connection between the turbojet and the aircraft. The nacelle includes a fan casing on the upstream end of which an air inlet is attached.

[0006] The turbojet's air intake structure incorporates several housings optimized for specific functions. The fan case surrounds the fan and guides the generated airflow. The intermediate compressor case (ICC) separates the primary airflow destined for the compressors from the secondary airflow that bypasses the engine. The intermediate case incorporates a separation nozzle located upstream of the intermediate case, which has a small leading-edge radius to minimize aerodynamic disturbances during the separation of the inlet airflow into the primary and secondary flows. The leading-edge radius here refers to the measurement of the curvature of the leading edge of the separation nozzle. A smaller radius indicates a greater curvature, resulting in a more "sharp" nozzle. allowing for better airflow management. The compressor case contains the primary airflow compression stages. The combustion case defines the combustion chamber where compressed air and fuel mix and ignite. The turbine case channels the high-energy combustion gases through the turbine, which extracts the energy needed to drive the compressors and blower. Finally, the exhaust case directs the exhaust gases outward, generating thrust.

[0007] In addition to these structural components, the turbojet engine has thrust reversers to slow the aircraft during landing. Several types of thrust reversers are commonly used. Grid-type reversers use retractable grids to redirect the airflow. Flap-type reversers use movable flaps to reverse the direction of the secondary flow. Cascade thrust reversers use a rearward-sliding fairing to reveal grids that redirect the secondary airflow.

[0008] It is also possible to reverse the thrust of the turbojet by modifying the fan blade pitch. By adjusting the angle of the fan blades, the secondary flow can be redirected forward, thus creating an effective braking force. This technique allows for a rapid and controlled reversal of thrust without adding any extra mechanical components, thereby reducing the weight and complexity of the overall system.

[0009] When thrust reversal is activated by changing the fan blade pitch, the operation of the turbojet engine changes radically. The secondary flow, in addition to being redirected forward to generate reverse thrust, flows up the secondary duct from the downstream end of the engine. A portion of this flow must then bypass the separation nozzle located in the intermediate casing to be reintroduced into the low-pressure compressor. The air supply to the low-pressure compressor then comes partly from the upstream end and partly from the secondary flow originating downstream.

[0010] Reversing thrust by modifying the fan blade pitch poses several problems:

[0011] -Flow Interactions: Reintroducing secondary flow into the main flow can generate turbulence and inconsistencies in the airflow, which can disrupt the optimal operation of the engine.

[0012] - Primary Flow Detachment: Disturbances in the primary flow, caused by reversing the secondary flow, can cause the latter to detach from the internal walls of the engine, thus affecting the stability of the flow.

[0013] - Compressor Pumping: Pressure imbalances can lead to pumping phenomena in the compressors, reducing the efficiency of the turbojet engine and potentially resulting in serious malfunctions or even complete engine failure, thus compromising the safety of the aircraft and its occupants. Presentation of the invention

[0014] The present invention aims to improve the operational safety of a thrust reversing turbomachine by modifying the fan blade pitch.

[0015] To this end, a turbomachine is provided comprising an air inlet equipped with a fan mounted for rotation relative to a nacelle about a main axis and intended to be traversed by an inlet airflow. The fan comprises a plurality of variable-pitch blades, and the nacelle comprises an intermediate casing defined by an inner and an outer fairing which meet at a nozzle separating the air inlet into a primary circulation channel and a secondary circulation channel. The primary channel comprises a low-pressure compressor connected to a low-pressure turbine, the fan being functionally connected to the low-pressure turbine.According to the invention, the turbomachine comprises a conduit which includes an outlet which opens into the primary flow and an inlet which opens into the secondary flow to fluidly connect the secondary flow and the primary flow, the conduit being provided with isolation means allowing the conduit to be selectively opened or closed.

[0016] According to other specific, non-exclusive and optional embodiments of the invention:

[0017] - the conduit is supported by the intermediate housing;

[0018] - the outlet opens into the primary vein between the separating nozzle and the low pressure compressor;

[0019] - the conduit has a radius of curvature of the conduit which is greater than a radius of leading edge curvature of the separating beak;

[0020] - the inlet is provided with a deflector arranged to selectively adopt a position active position in which the deflector directs part of the secondary flow towards the duct and a passive position in which the flow circulation is left free;

[0021] - the isolation means include a movable entry door that isolates the conduit at the inlet level and which is mounted to slide and / or pivot relative to the intermediate casing;

[0022] - the deflector is attached to the entrance door;

[0023] - the isolation means include a movable exit door that isolates the conduit at the outlet level and which is mounted to slide and / or pivot relative to the intermediate casing;

[0024] - the isolation means are actuated by a control ring mounted at rotation around the main axis;

[0025] The invention relates to a method for feeding a turbomachine turbine as defined above, comprising the following steps: - modify the fan blade pitch so as to switch the fan into a thrust reversal configuration; order the isolation devices in order to open the conduit.

[0026] Other features and advantages of the invention will become apparent from the following description of particular, non-limiting embodiments of the invention. Brief description of the figures

[0027] The invention will be better understood upon reading the following description, given by way of non-limiting example, and made with reference to the figures which represent:

[0028] [Fig-1] [Fig.1] is a schematic cross-sectional representation of a turbojet engine according to a first embodiment of the invention;

[0029] [Fig.2] [Fig.2] is a partial schematic cross-sectional detail representation of the turbojet engine of [Fig.1];

[0030] [Fig.3] [Fig.3] is a detailed cross-sectional representation of the turbojet engine of [Fig.1] in a first state;

[0031] [Fig.4] [Fig.4] is a detailed cross-sectional representation of the turbojet engine of [Fig.1] in a second state;

[0032] [Fig.5] [Fig.5] is a detailed cross-sectional representation of a turbojet engine according to a second embodiment of the invention in a first state;

[0033] [Fig.6] [Fig.6] is a detailed cross-sectional representation of the turbojet engine of [Fig.4] in a second state;

[0034] [Fig.7] [Fig.7] is a schematic partial perspective representation of the turbojet engine of [Fig.6]. Description of the implementation methods

[0035] With reference to Figures 1 and 2, a propulsion assembly 1000 comprises a turbomachine, here a turbojet generally designated 1, in which an airflow 100 is admitted, following a longitudinal direction parallel to the axis AX of rotation of the turbomachine 1, into an inlet sleeve 2 to pass through a fan 3 comprising a series of rotating blades.

[0036] Part of the airflow 100 is compressed by a first low-pressure compressor 4 and then a second high-pressure compressor 5 before reaching a combustion chamber 6, after which it expands by passing through a high-pressure turbine 7, which is rotating with the high-pressure compressor 5, before being discharged through a low-pressure turbine 8 and generating thrust. The remainder of the airflow 100 is propelled directly by the blower 3, which is functionally connected to the low-pressure compressor 4 via an axial shaft 4.3, to generate the main thrust.

[0037] In this text, the terms "internal" and "external" are used with reference to the position or orientation relative to the axis of rotation of the turbines 7 and 8.

[0038] In this text, the terms "upstream" and "downstream" are used with reference to the position or orientation of an element with respect to the direction of flow of the airflow 100 in the turbojet 1 in nominal operation of the turbojet 1.

[0039] As a preliminary matter, an axial direction is defined, a radial direction which is orthogonal to the axial direction and a circumferential / tangential direction which is orthogonal to the axial and radial directions.

[0040] The turbojet 1 also includes a nacelle 9 which supports the elements of the turbojet 1 and provides its connection to an aircraft not shown.

[0041] The blower 3 is here a variable pitch blade blower. The blades 11 are pivotally mounted on a hub 10 which contains the mechanism, known per se, for modifying the angle of the blades 11 relative to the airflow 100.

[0042] The nacelle 9 comprises an intermediate housing 20 defined by an inner fairing 21 and an outer fairing 22 which meet at a separating nozzle 23 of the air inlet 2. The separating nozzle 23 separates the air inlet 2 into a primary flow 30 for primary flow FI and a secondary flow 31 for secondary flow FIL. The primary flow 30 thus supplies the low-pressure compressor 4. As shown in [Fig. 2], the housing 20 is traversed by a conduit 40 which includes an outlet 41 that opens into the primary flow 30 and an inlet 42 that opens into the secondary flow 31 to fluidly connect the secondary flow 31 and the primary flow 30. The outlet 41 opens into the primary flow 30 between the separating nozzle 23 and the low-pressure compressor 4.

[0043] The conduit 40 is here a cylindrical conduit which has a conduit radius of curvature R40 which is greater than the leading edge radius of curvature R23 of the separating nozzle 23.

[0044] As can be seen in Figures 3 and 4, the conduit 40 is provided with isolation means 50 allowing the conduit 40 to be selectively opened or closed, thus establishing or separating the primary vein 30 and the secondary vein 31 from the fluidic flow. More specifically, the isolation means 50 comprise a movable inlet gate 51 that isolates the conduit 40 at the inlet 42 and is slidably mounted relative to the intermediate housing 20 in a direction substantially tangent to the fairing 22 at the level of the inlet 42. An inlet actuator 52, here in the form of a pneumatic cylinder, controls a sliding of the inlet door 51 between a closing position ([Fig.3]) in which it closes the inlet 42 to close the conduit 40, and a release position ([Fig.4]) in which it releases the inlet 42 to open the conduit 40.

[0045] When the turbojet 1 is in nominal propulsion operation, the primary flow F^t and the secondary flow Fu circulate from upstream to downstream ([Fig.3]). The gate 51 is in its closed position ([Fig.3]) and no flow circulates in the duct 40.

[0046] When it is desired to brake the aircraft by thrust reversal, the pitch of the fan blades 11 of the fan 3 is modified to switch the fan 3 to a thrust reversal configuration. The actuator 52 is controlled to move the inlet door 51 to its clearance position ([Fig. 4]) and thus open the duct 40. The secondary flow Fu then circulates in the channel 31 from downstream to upstream, and a portion of the secondary flow Fu is drawn off and routed through the duct 40 into the channel 30. The large radius of curvature R40 of the duct 40 reduces the risk of separation of the primary flow F! in the channel 30 and the potential pumping phenomena of the compressors 4 and 5, thus improving the safety of operation of the turbojet 1 in thrust reversal mode.

[0047] Elements identical or analogous to those previously described shall bear a numerical reference identical to that in the following description of a second embodiment of the invention.

[0048] According to a second embodiment of the invention shown in Figures 5 and 6, the isolation means 50 also include a movable outlet gate 53 which isolates the conduit 40 at the outlet 4L. The gate 53 is mounted to slide relative to the intermediate housing 20 in a direction substantially tangent to the fairing 21 at the outlet 4L. The outlet gate 53 can selectively adopt a closing position ([Fig.5]) in which it closes the outlet 41 to close the conduit 40 on the side of the primary vein 30, and a release position ([Fig.6]) in which it releases the outlet 41 to open the conduit 40 on the side of the primary vein 30.

[0049] According to this second embodiment, the door 51 is mounted for rotation about an axis A51 extending in a substantially tangential direction. The door 51 has, in this instance, a substantially L-shaped axial cross-section comprising a wing 54 extending in a direction substantially tangent to the fairing 22 and a web 55 extending in a substantially radial direction. The door 51 includes an axially projecting external surface 51.1 from which a first sleeve 56 is inserted, into which is articulated a downstream end of a first connecting rod 57. The internal surface 51.2 of the gate 51 is arranged to adopt a radius of curvature R51 substantially equal to the radius of curvature R40, at least greater than the radius of curvature R23.

[0050] The external surface 51.1 is aerodynamically arranged to follow the lines of the fairing 22 when the door 51 is in its closed position.

[0051] The exit gate 53 includes a second sleeve 58 in which is articulated a downstream end of a second connecting rod 59.

[0052] The upstream ends of the connecting rods 57 and 59 are articulated on a first substantially radial shaft 60. The shaft 60 is connected at its radially external end 61 to a first end 62 of an inter-motor type lever 63, the second end 64 of which is rotatably mounted relative to the housing 20 around a second shaft 65. The median force application point 66 of the lever 63 is ball-mounted on a control ring 70 rotatably mounted about the axis AX and which is known per se. Thus, a movement of the lever 63 causes the gates 51 and 53 to move from their closed state to their open state and vice versa.

[0053] When the turbojet 1 is in nominal propulsion operation, the primary flow F^t and the secondary flow Fu circulate from upstream to downstream ([Fig. 5]). The gates 51 and 53 are in their closed positions ([Fig. 5]) and no flow circulates in the duct 40.

[0054] When it is desired to brake the aircraft by thrust reversal, the pitch of the fan blades 11 of the fan 3 is modified so as to switch the fan 3 into a thrust reversal configuration. The ring 70 is rotated so as to move the doors 51 and 53 into their clearance positions ([Fig. 5]) and thus open the duct 40.

[0055] The gate 51 rotates about the axis A51 through an angle of approximately forty-five degrees, causing the core 54 of the gate 51 to extend into the secondary channel and adopt an active position for deflecting the secondary flow Fu, which then flows in the channel 31 from downstream to upstream. The internal surface 51.2 thus acts as a flow deflector integrated into the gate 51.

[0056] Part of the secondary flow Fu is thus taken and conveyed by the conduit 40 into the vein 30. The large radius of curvature R51 of the internal surface 51.2 reduces the risks of separation of the primary flow F! in the vein 30 and the possible pumping phenomena of the compressors 4 and 5, thus improving the safety of use of the turbojet 1 in thrust reversal mode.

[0057] By actuating the crown 70 in the opposite direction, the doors 51 and 53 are moved from their open position to their closed position in order to close the conduit 40. The internal surface 51.2 therefore passes into a passive deflection position in which the circulation of the secondary flow Fu is left free.

[0058] Of course, the invention is not limited to the embodiments described but encompasses any variant falling within the scope of the invention as defined by the claims.

[0059] In particular,

[0060] - although here a single conduit 40 has been described, the invention also applies, and preferably, to a turbomachine comprising a plurality of conduits such as for example between two and sixteen;

[0061] - although here the entrance door is mounted to pivot or slide relative to the casing, the invention also applies to an entrance door mounted to slide and pivot relative to the casing;

[0062] - although here the invention has been described in application to a propulsion system including a twin-body turbofan engine, the invention also applies to propulsion assemblies comprising other types of turbomachinery such as for example a propulsion assembly comprising a single-body twin-body turbofan engine;

[0063] - although here the deflector is integral with the entrance door, the invention applies also to a deflector separate from the entrance door such as for example a deflector mounted on rotation on the intermediate casing.

Claims

Demands

1. Turbomachine (1) comprising an air inlet (2) provided with a fan (3) mounted for rotation relative to a nacelle (9) about a principal axis (AX) and intended to be traversed by an inlet airflow (100), the fan (3) comprising a plurality of variable-pitch blades (11), and the nacelle (9) comprising an intermediate casing (20) defined by an inner fairing (21) and an outer fairing (22) which meet at a separating nozzle (23) of the air inlet (2) into a primary flow (30) for circulation of a primary flow (FO) and a secondary flow (31) for circulation of a secondary flow (Fu), the primary flow (30) comprising a low-pressure compressor (4) connected to a low-pressure turbine (8), the fan (3) being functionally connected to the low-pressure turbine (8),characterized in that the turbomachine (1) comprises a conduit (40) which includes an outlet (41) opening into the primary flow and an inlet (42) opening into the secondary flow (31) to fluidly connect the secondary flow (31) and the primary flow (30), the conduit (40) being provided with isolation means allowing the conduit to be selectively opened or closed.

2. Turbomachine according to claim 1, wherein the conduit (40) is carried by the intermediate casing (20).

3. Turbomachine according to claim 1 or 2, wherein the outlet (41) opens into the primary channel (30) between the separation nozzle (23) and the low pressure compressor (4).

4. Turbomachine according to any one of the preceding claims, wherein the duct has a duct radius of curvature that is greater than a leading edge radius of curvature (R23) of the separation nozzle (23).

5. Turbomachine (1) according to any one of the preceding claims, wherein the inlet (42) is provided with a deflector (51.2) arranged to selectively adopt an active position in which the deflector (51.2) directs a portion of the secondary flow to the conduit (40) and a passive position in which the flow circulation is left free.

6. Turbomachine (1) according to any one of the preceding claims, wherein the isolation means (50) include

7.

8.

9.

10. a movable entry door (51) which isolates the conduit (40) at the level of the entry (42) and which is mounted to slide and / or pivot relative to the intermediate casing (20). Turbomachine (1) according to claims 5 and 6, wherein the deflector (51.2) is integral with the inlet door (51). Turbomachine (1) according to any one of the preceding claims, wherein the isolation means (50) comprise a movable outlet door (53) which isolates the conduit (40) at the outlet (41) and which is mounted to slide and / or pivot relative to the intermediate casing (20). Turbomachine (1) according to any one of the preceding claims, wherein the isolation means (50) are actuated by a control ring (70) mounted to rotate about the main axis (AX). Method for feeding a turbomachine turbine (1) according to any one of claims 1 to 9, comprising the following steps: - modify the blade pitch (11) of the fan (3) so as to put the fan (3) into a thrust reversal configuration; - order the isolation means (50) so as to open the conduit (40).

Images