Turbine engine with improved primary flow supply

Abstract

The invention relates to a turbine engine (1) comprising an air inlet (2), which is provided with a fan (3) rotatably mounted relative to a nacelle (9) about a main axis (AX) and through which an inlet air flow (100) is intended to pass, 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 in a splitter (23) for splitting the air inlet (2) into a primary duct (30) for circulating a primary flow (Fi) and a secondary duct (31) for circulating a secondary flow (Fn), the primary duct (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 turbine engine (1) comprises a pipe which comprises an outlet which opens into the primary duct and an inlet which opens into the secondary duct (31) in order to fluidically connect the secondary duct (31) and the primary duct (30), the pipe being provided with isolation means for selectively opening or closing the pipe. The invention also relates to a method for supplying a turbine engine (1).

Description

[0001] Description

[0002] Title of the invention: IMPROVED PRIMARY FLOW TURBOMACHINE

[0003] Technical field of the invention

[0004] 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.

[0005] 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).

[0006] Prior art

[0007] 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.

[0008] The primary airflow is compressed by stages of compressors before reaching a combustion chamber, after which it expands as it passes through turbines, before being expelled and generating thrust. The secondary airflow, on the other hand, is propelled directly by the fan to generate the main thrust.

[0009] The turbojet also includes a nacelle that supports the turbojet components and connects the turbojet to the aircraft. The nacelle includes a fan casing with an air intake attached to its upstream end.

[0010] The turbofan engine'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 streams. The leading-edge radius here refers to the measurement of the leading edge curvature of the separation nozzle. A smaller radius indicates a greater curvature, resulting in a sharper nozzle, allowing for better airflow management.The compressor case contains the compression stages for the primary airflow. The combustion case defines the combustion chamber where the 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 the blower. Finally, the exhaust case directs the exhaust gases outward, generating thrust.

[0011] 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 airflow. Cascade thrust reversers use a rearward-sliding fairing to reveal grids that redirect the secondary airflow.

[0012] 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 airflow can be redirected forward, creating an effective braking force. This technique allows for a rapid and controlled thrust reversal without adding extra mechanical components, thus reducing the weight and complexity of the overall system.

[0013] When thrust reversal is activated by changing the fan blade pitch, the turbojet's operation 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. Part of this flow must then bypass the separation nozzle located in the intermediate casing to be reintroduced into the low-pressure compressor. The low-pressure compressor's air supply then comes partly from the upstream end and partly from the secondary flow coming from the downstream end. Thrust reversal by changing the fan blade pitch presents several problems:

[0014] - Flow Interactions: Reintroducing secondary flow into the main channel can cause turbulence and inconsistencies in the airflow, which can disrupt optimal engine operation.

[0015] - Primary Flow Detachment: Disturbances in the primary flow, caused by the reversal of the secondary flow, can lead to the detachment of the latter from the internal walls of the engine, thus affecting the stability of the flow.

[0016] - Compressor surging: Pressure imbalances can lead to surging phenomena in the compressors, reducing the efficiency of the turbojet and potentially leading to serious malfunctions or even complete engine failure, thus compromising the safety of the aircraft and its occupants.

[0017] Presentation of the invention

[0018] The present invention aims to improve the operational safety of a thrust-reversing turbomachine by modifying the fan blade pitch. 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 carry an inlet airflow. The fan comprises a plurality of blades with variable pitch, and the nacelle comprises an intermediate casing defined by an inner and an outer fairing that meet at a nozzle separating the air inlet into a primary flow path and a secondary flow path. The primary flow path includes a low-pressure compressor connected to a low-pressure turbine, the fan being functionally connected to the low-pressure turbine.The turbomachine includes a conduit comprising an outlet opening into the primary flow and an inlet opening into the secondary flow to fluidly connect the secondary and primary flows. The conduit is provided with isolation means for selectively opening or closing it. According to the invention, the isolation means comprise a movable inlet gate that isolates the conduit at the inlet and is mounted to slide and / or pivot relative to the intermediate housing. The isolation means also comprise a movable outlet gate that isolates the conduit at the outlet and is mounted to slide and / or pivot relative to the intermediate housing to slide upstream of the turbomachine.

[0019] According to other specific, non-exclusive and optional embodiments of the invention:

[0020] - the conduit is supported by the intermediate casing;

[0021] - the outlet opens into the primary vein between the separation nozzle and the low pressure compressor;

[0022] - the duct has a duct radius of curvature which is greater than a leading edge radius of curvature of the separating nozzle;

[0023] - the inlet is equipped with a deflector arranged to selectively adopt an 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;

[0024] - the deflector is attached to the front door;

[0025] /

[0026] - the isolation means are operated by a control ring mounted to rotate around the main axis;

[0027] The invention also relates to a method of supplying a turbomachine turbine as defined above, comprising the following steps: modifying a fan blade pitch so as to put the fan into a thrust reversal configuration; controlling the isolation means so as to open the duct.

[0028] Other features and advantages of the invention will become apparent from the following description of particular, non-limiting embodiments of the invention.

[0029] Brief description of the figures

[0030] 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: [Fig. 1] Figure 1 is a schematic cross-sectional representation of a turbojet according to a first embodiment of the invention; [Fig. 2] Figure 2 is a partial schematic cross-sectional detail representation of the turbojet of Figure 1;

[0031] [Fig. 3] Figure 3 is a detailed cross-sectional representation of the turbojet engine in Figure 1 in a first state;

[0032] [Fig. 4] Figure 4 is a detailed cross-sectional representation of the turbojet engine in Figure 1 in a second state;

[0033] [Fig. 5] Figure 5 is a detailed cross-sectional representation of a turbojet engine according to a second embodiment of the invention in a first state;

[0034] [Fig. 6] Figure 6 is a detailed cross-sectional representation of the turbojet engine in Figure 4 in a second state;

[0035] [Figure 7] Figure 7 is a schematic partial perspective representation of the turbojet engine in Figure 6.

[0036] Description of the implementation methods

[0037] Referring to Figures 1 and 2, a propulsion system 1000 comprises a turbomachine, here a turbojet engine generally designated 1, into which an airflow 100 is admitted, following a longitudinal direction parallel to the axis AX of rotation of the turbomachine 1, into an inlet 2 to pass through a fan 3 having a series of rotating blades. 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 fixed in rotation to the high-pressure compressor 5, before being discharged by passing through a low-pressure turbine 8 and generating thrust. The remainder of the airflow 100 is propelled directly by the fan 3, which is functionally connected to the low-pressure compressor 4 via an axial shaft 4.3, to generate the main thrust.In this text, the terms "internal" and "external" are used in reference to the position or orientation relative to the axis of rotation of 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 a nominal operation of the turbojet 1. 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.

[0039] 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.

[0040] 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 in itself, for modifying the angle of the blades 11 relative to the airflow 100.

[0041] 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 for the air inlet 2. The separating nozzle 23 divides the air inlet 2 into a primary flow 30 for primary flow Fl and a secondary flow 31 for secondary flow FIL. The primary flow 30 thus supplies the low-pressure compressor 4. As shown in Figure 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.

[0042] The conduit 40 is a cylindrical conduit with a radius of curvature R40 that is greater than the leading edge radius R23 of the separating nozzle 23. As shown in Figures 3 and 4, the conduit 40 is provided with isolation means 50 that allow the conduit 40 to be selectively opened or closed, thus establishing or separating the primary vein 30 and the secondary vein 31 in fluidic communication. More specifically, the isolation means 50 include 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 inlet 42.An inlet actuator 52, here in the form of a pneumatic cylinder, controls the sliding of the inlet gate 51 between a closed position (Figure 3), in which it closes the inlet 42 to shut off the duct 40, and a clearing position (Figure 4), in which it opens the inlet 42 to open the duct 40. The transition from the closed position to the clearing position is achieved by sliding the gate 51 upstream. When the turbojet engine 1 is in nominal propulsion operation, the primary flow Fi and the secondary flow Fn circulate from upstream to downstream (Figure 3). The gate 51 is in its closed position (Figure 3), and no flow circulates in the duct 40.

[0043] When thrust reversal braking is desired, the fan blade pitch 11 of fan 3 is adjusted to switch fan 3 into a thrust reversal configuration. The actuator 52 is then moved upstream to its open position (Figure 4), thus opening the duct 40. The secondary flow Fn then circulates in the channel 31 from downstream to upstream, and a portion of this secondary flow Fn is diverted through the channel 40 into the channel 30. The large radius of curvature R40 of the channel 40 reduces the risk of primary flow separation Fi in the channel 30 and the potential for compressor 4 and 5 to over-slosh, thereby improving the safety of operation of the turbojet 1 in thrust reversal mode.

[0044] 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.

[0045] 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 that isolates the conduit 40 at the outlet 41. The gate 53 is slidably mounted relative to the intermediate housing 20 in a direction substantially tangent to the fairing 21 at the outlet 41. The outlet gate 53 can selectively adopt a closed position (Figure 5) in which it closes the outlet 41 to shut off the conduit 40 on the side of the primary stream 30, and a clearing position (Figure 6) in which it releases the outlet 41 to open the conduit 40 on the side of the primary stream 30. The transition from the closed position to the clearing position is achieved by sliding the gate 51 upstream.

[0046] 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 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 articulated, into which a downstream end of a first connecting rod 57 is articulated. The internal surface 51.2 of the door 51 is arranged to adopt a radius of curvature R51 substantially equal to the radius of curvature R40, or at least greater than the radius of curvature R23.

[0047] 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.

[0048] The exit gate 53 includes a second sleeve 58 in which is articulated a downstream end of a second connecting rod 59.

[0049] The upstream ends of 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-engine type lever 63, the second end of which 64 is mounted for rotation relative to the housing 20 around a second shaft 65. The median point 66 of force application of the lever 63 is mounted for rotation on a control ring 70, which is itself rotatable about the axis AX. 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.

[0050] When the turbojet engine 1 is in nominal propulsion operation, the primary flow Fi and the secondary flow Fn circulate from upstream to downstream (Figure 5). The gates 51 and 53 are in their closed positions (Figure 5), and no flow circulates in the duct 40. 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 ring 70 is rotated to move the gates 51 and 53 to their open positions (Figure 5), thus opening the duct 40.

[0051] The gate 51 rotates around the axis A51 by 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 Fn, 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.

[0052] Part of the secondary flow Fn 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 Fi 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.

[0053] By actuating the crown 70 in the opposite direction, gates 51 and 53 are moved from their open position to their closed position to close the conduit 40. The internal surface 51.2 thus moves to a passive deflection position in which the secondary flow FH is allowed to circulate freely. Gate 51 slides upstream to move into the open configuration.

[0054] 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.

[0055] Especially,

[0056] - 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;

[0057] - 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;

[0058] - although here the invention has been described in application to a propulsion assembly comprising a twin-spool turbojet engine, the invention also applies to propulsion assemblies comprising other types of turbomachinery such as for example a propulsion assembly comprising a single-spool turbojet engine;

[0059] - although here the deflector is attached to the entrance door, the invention also applies to a deflector separate from the entrance door such as for example a deflector mounted to rotate on the intermediate casing.

Claims

Demands Claim 1. 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 (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 for selectively opening or closing the conduit, wherein: the isolation means (50) comprise a movable inlet gate (51) which isolates the conduit (40) at the inlet (42) and which is pivotally mounted relative to the intermediate casing (20), and the isolation means (50) comprise a movable outlet gate (53) which isolates the conduit (40) at the outlet (41) and which is slidably mounted relative to the intermediate casing (20) to slide upstream of the turbomachine (1). Claim 2. Turbomachine according to claim 1, in which the conduit (40) is carried by the intermediate casing (20). Claim 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). Claim 4. Turbomachine according to any one of the preceding claims, wherein the duct has a duct radius of curvature which is greater than a leading edge radius of curvature (R23) of the separation nozzle (23). Claim 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 towards the conduit (40) and a passive position in which the flow circulation is left free. Claim 6. Turbomachine (1) according to claims 5 and 6, wherein the deflector (51.2) is integral with the inlet door (51). Claim 7. 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). Claim 8. A method for supplying a turbomachine turbine (1) according to any one of claims 1 to 7, comprising the following steps: modifying a blade pitch (11) of the blower (3) so as to switch the blower (3) into a thrust reversal configuration; controlling the isolation means (50) so as to open the duct (40)

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