Turbomachine including a sliding discharge valve

Abstract

This aircraft turbomachine (2) comprises at least one compressor (12, 14, 16), a combustion chamber (18), at least one turbine (20, 22), an external flow channel (32), and at least one internal flow channel (30, 36) through which flows a flow from at least one compressor (12, 14, 16) and which is provided with at least one discharge channel (38). This turbomachine (2) further comprises a movable panel (40) between an upstream position in which said movable panel (40) allows air to flow through said discharge channel (38) and a downstream position in which said movable panel (40) closes said discharge channel (38). Figure for the abbreviation: Fig 1

Description

Title of the invention: Turbomachine comprising a sliding relief valve. Technical field

[0001] The invention has as its technical domain the operability of turbomachines.

[0002] In particular, the invention relates to an aircraft turbomachine comprising a movable panel in a discharge channel of a flow channel of an internal flow. Previous techniques

[0003] A turbomachine generally comprises a primary gas flow path and a secondary gas flow path. Recently, triple-flow turbomachines have been designed; such a turbomachine is described, for example, in document WO2023 / 198962A1.

[0004] In order to increase the flow of air circulating through a compressor of the turbomachine, it is known to produce an air flow channel, called a discharge channel, which is connected downstream of this compressor and which includes a mechanism capable of closing this discharge channel and allowing air to circulate through this discharge channel.

[0005] It is thus possible to increase the flow of air circulating through this compressor by allowing the air to flow through this discharge channel in order to reduce the air pressure downstream of the compressor.

[0006] A known solution for blocking this discharge vein consists of a variable section nozzle, for example such as described in document WO2008 / 045062A1.

[0007] However, this solution requires a significant amount of space to be implemented, particularly a large space downstream of the discharge stream, which is incompatible with certain turbomachinery architectures, notably the generally frustoconical shape of a downstream section of some turbomachines. Furthermore, this solution is complex to implement. Description of the invention

[0008] The present invention therefore aims to overcome all or part of the aforementioned drawbacks, and in particular to provide a turbomachine having at least one compression stage whose air flow can be controlled according to the state of a discharge channel of the turbomachine.

[0009] The present invention relates to an aircraft turbomachine comprising at least one compressor, one combustion chamber, at least one turbine, one external flow duct, and at least one internal flow duct. in which circulates a flow from at least one compressor and which is equipped with at least one discharge channel, a connection to said internal flow channel of which is located longitudinally downstream of said compressor.

[0010] The turbomachine further comprises at least one movable panel at least between an upstream position in which said movable panel allows air to circulate through said discharge channel and a downstream position in which said movable panel closes said discharge channel.

[0011] The longitudinal direction of the turbomachine corresponds to an axis of rotation of its rotating stages. The radial and tangential directions define, with the longitudinal direction, an orthogonal frame specific to the turbomachine. The axial direction is parallel to the longitudinal direction.

[0012] The terms upstream and downstream are defined with respect to the main direction of gas flow along the longitudinal axis of the turbomachine.

[0013] Said discharge vein makes it possible to reduce the air pressure downstream of said compressor when said movable panel is in the upstream position, and therefore to increase the air flow through said compressor.

[0014] Said movable panel is a simple solution to implement and represents a compact solution, allowing its integration into many turbomachine architectures. In particular, the movement of said movable panel between the upstream and downstream positions, such that in the upstream position said movable panel allows air circulation through said discharge channel and such that in the downstream position said movable panel closes said discharge channel, disturbs the turbomachine flow less than a reverse movement between these two positions.

[0015] Said compressor may be any rotor stage which is equipped with blades and whose function is to increase the pressure of the air directed towards the combustion chamber.

[0016] In a first embodiment, the turbomachine includes a nozzle for separating the flow stream of the external flow, called secondary flow stream, and said flow stream of the internal flow, called primary flow stream, which extends through the combustion chamber.

[0017] Optionally, according to the first embodiment, the turbomachine further comprises at least one other internal flow stream in which a flow from said compressor circulates, called the third flow stream, and comprises a nozzle for separating the primary flow stream and the third flow stream, the third flow stream being located at least partly radially between the primary flow stream and the secondary flow stream.

[0018] In a second alternative embodiment, the turbomachine comprises at least said internal flow channel, referred to as the third flow channel, and another internal flow stream in which flows a stream from said compressor, and which extends through the combustion chamber, called primary flow stream, and comprising a spout for separating the primary flow stream and the third flow stream, the third flow stream being located at least partly radially between the primary flow stream and the external flow stream, called secondary flow stream.

[0019] According to the first or second embodiment, the third flow vein remains at a distance from the combustion chamber.

[0020] Advantageously, said internal flow vein is radially delimited by an internal wall and by an external wall, said discharge vein extending radially outwards from an opening formed in the external wall of said internal flow vein.

[0021] Advantageously, said movable panel in the downstream position partly forms the external wall of said internal flow vein.

[0022] Preferably, the turbomachine comprises an upstream fairing and a downstream fairing which together define an outlet of said discharge stream, said movable panel in the downstream position axially extending the upstream fairing towards the downstream fairing.

[0023] Optionally, said movable panel in the downstream position is flush with the downstream fairing.

[0024] Optionally, said movable panel in the downstream position is flush with the upstream fairing.

[0025] Preferably, said movable panel extends circumferentially along a limited angular sector.

[0026] Preferably, the turbomachine comprises a plurality of movable panels, each extending circumferentially along a limited angular sector.

[0027] The turbomachine can comprise between two and twelve movable panels, in particular ten movable panels.

[0028] The movable panels can be located at the same longitudinal position of the turbomachine.

[0029] The movable panels can be spaced from each other in the circumferential direction, in particular equidistributed.

[0030] Said discharge channel may include a plurality of air outlets, each air outlet extending circumferentially along a limited angular sector, each movable panel being able, in upstream position, to allow air to flow through the air outlets of said discharge channel and being able, in downstream position, to close one of the air outlets of said discharge channel.

[0031] Optionally, the turbomachine includes a pressurized box which is located partly radially outside said discharge channel and which partially surrounds said discharge channel, said movable panel in the upstream position being arranged at least partly in the pressurized box.

[0032] The pressurized chamber thus constitutes a pressure barrier between the air located radially inside the pressurized chamber and the air located radially outside the pressurized chamber. Therefore, the pressurized chamber, partially surrounding said discharge channel, reduces the surface area of ​​said movable panel on which the pressure difference between the air located radially inside said movable panel and the air located radially outside said movable panel is applied.

[0033] The pressurized chamber may include an opening through which said movable panel slides.

[0034] Said movable panel may include a seal which is in contact with a wall of the pressurized chamber when said movable panel is in the downstream position so as to prevent the circulation of fluid between the pressurized chamber and said vein through the opening of the pressurized chamber.

[0035] Said movable panel may include a seal which is in contact with the downstream fairing when said movable panel is in the downstream position so as to improve the sealing of said discharge channel.

[0036] Advantageously, said movable panel comprises a perforated inner skin and a honeycomb core located radially outside the perforated inner skin, the turbomachine comprising an acoustic resistive skin located radially outside said movable panel so as to form an acoustic stack consisting of the perforated inner skin, the honeycomb core and the acoustic resistive skin.

[0037] Thus, acoustic resonance phenomena in said discharge vein are reduced, or even prevented.

[0038] Optionally, the turbomachine includes a control device for said movable panel equipped with rails and actuators.

[0039] Optionally, said movable panel can be piloted stationary in any position between the upstream position and the downstream position.

[0040] Optionally, the turbomachine includes a pressure sensor configured to measure a downstream pressure of said compressor, said movable panel being piloted in upstream and downstream positions according to a pressure measurement of the pressure sensor.

[0041] Said movable panel can be steered in the upstream position so that the air pressure measured by the pressure sensor is below a predetermined lower pressure threshold. Said movable panel can be steered in the downstream position so that that the air pressure measured by the pressure sensor is above a predetermined upper pressure threshold.

[0042] According to a particular embodiment, said internal flow channel is provided with at least first and second discharge channels whose respective connections to said internal flow channel are located longitudinally downstream of said compressor, the second discharge channel being located downstream of the first discharge channel, the turbomachine further comprising at least first and second movable panels between upstream and downstream positions, the first and second movable panels in the upstream position respectively allowing air circulation through the first and second discharge channels, the first and second movable panels in the downstream position respectively closing the first and second discharge channels.

[0043] Thus, the operability of the turbomachine is further improved.

[0044] Advantageously, the turbomachine includes a propeller equipped with blades that are rotating about a longitudinal axis of the turbomachine.

[0045] Optionally, the propeller is unfaired.

[0046] When the propeller is unshod, the turbomachine lacks a nacelle. In this case, the secondary flow is not radially delimited externally.

[0047] Optionally, the propeller can be shrouded. In this case, the turbomachine can include a nacelle radially delimiting the secondary flow.

[0048] Advantageously, the turbomachine includes a non-rotating stage equipped with a plurality of blades that are non-rotating about the longitudinal axis of the turbomachine, the non-rotating stage being arranged downstream of the propeller.

[0049] Optionally, the plurality of blades of the non-rotating stage has variable pitch or pitch, each blade of the non-rotating stage including in particular a mechanism for adjusting its orientation around a radial axis of the turbomachine.

[0050] Advantageously, the turbomachine includes a shaft, in particular a low-pressure shaft, which is integral with the propeller, said compressor and said turbine.

[0051] In a particular embodiment, the turbomachine further comprises an intermediate fan, in particular shrouded, which constitutes said compressor, the turbomachine further comprising a low pressure compressor, a high pressure compressor, a high pressure turbine, and a low pressure turbine constituting said turbine.

[0052] Advantageously, the low pressure compressor is integral with the low pressure shaft.

[0053] According to a particular design, the intermediate blower is connected to the propeller so that the intermediate blower and the propeller have the same rotational speed.

[0054] According to another particular design, the intermediate blower is connected to the propeller via a reducer so that the intermediate blower and the propeller have a proportional rotational speed.

[0055] It is particularly advantageous to control the compression ratio of the intermediate blower according to this particular design because the range of rotational speeds is very small here and the range of flow rates between idle and full throttle is high.

[0056] Advantageously, the turbomachine includes a shaft, in particular a high-pressure shaft, which is integral with the high-pressure compressor and the high-pressure turbine.

[0057] The present invention also relates to an aircraft comprising at least one turbomachine as defined above. Brief description of the drawings

[0058] Other objects, features and advantages of the invention will become apparent from the following description, given solely by way of non-limiting example and made with reference to the accompanying drawings in which:

[0059] [Fig. 1] illustrates a turbomachine according to a first example of embodiment of the invention when a movable panel is in upstream position;

[0060] [Fig.2] is a detail view of [Fig.1];

[0061] [Fig.3] is a view analogous to that of [Fig.2] when the movable panel is in downstream position;

[0062] [Fig.4] is a detailed view of [Fig.3];

[0063] [Fig.5] is a perspective view of a movable panel according to a second example of the realization of the invention; and

[0064] [Fig. 6] is a perspective view of air outlets from a discharge duct of the turbomachine according to the second embodiment of the invention. Detailed description

[0065] Fig. 1 represents an aircraft turbomachine 2 comprising an unfaired propeller 4 which is provided with blades 6 rotating about a longitudinal axis of the turbomachine 2, a non-rotating stage 8 which is provided with a plurality of variable-pitch blades 10 which are non-rotating about the longitudinal axis, a faired intermediate fan 12, a low-pressure compressor 14, a high-pressure compressor 16, a combustion chamber 18, a high-pressure turbine 20, and a low-pressure turbine 22. The intermediate fan 12 has, in particular, the function of increasing the pressure of the air directed towards the combustion chamber 18.

[0066] The turbomachine 2 also includes a low-pressure shaft 24 integral with the propeller 4, the intermediate blower 12, the low-pressure compressor 14 and the low-pressure turbine 22, and a high-pressure shaft 26 integral with the high-pressure compressor 16 and the high-pressure turbine 20.

[0067] The turbomachine 2 includes a first separation nozzle 28 separating a first flow path of the internal airflow 30, called the primary flow path, from a flow path of the external airflow 32, called the secondary flow path. The first separation nozzle 28 is located longitudinally between the propeller 4 and the non-rotating stage 8.

[0068] The turbomachine 2 also includes a second separation nozzle 34 for the primary flow stream 30 and a second internal flow stream 36, referred to as the third flow stream. The second separation nozzle 34 is located longitudinally downstream of the non-rotating stage 8. The third flow stream 36 is located radially between the portion of the primary flow stream 30 extending downstream from the second separation nozzle 34 and the secondary flow stream 32.

[0069] The intermediate blower 12 is located longitudinally between the first separation nozzle 28 and the second separation nozzle 34.

[0070] The primary flow channel 30 delimits a primary airflow which flows from the first separation nozzle 28 through the intermediate blower 12, the low pressure compressor 14, the high pressure compressor 16, the combustion chamber 18, the high pressure turbine 20 and the low pressure turbine 22. The primary flow channel 30 is radially delimited by an inner wall 30a and by an outer wall 30b.

[0071] The secondary flow vein 32 delimits a secondary air flow which flows from the first separation nozzle 28 radially outside the first separation nozzle 28. The secondary flow vein 32 is not delimited radially outside here.

[0072] The third flow vein 36 delimits a third airflow which flows from the second separation nozzle 34 radially outside the second separation nozzle 34. The third flow vein 36 is delimited radially by an internal wall 36a and by an external wall 36b.

[0073] The turbomachine 2 is therefore a three-flow aircraft engine architecture with three paths for the air that passes through the propeller 4.

[0074] The turbomachine 2 also includes a pressure sensor 37 suitable for measuring a downstream pressure of the intermediate blower 12, and here arranged in the primary flow channel 30 longitudinally between the intermediate blower 12 and the second separation nozzle 34.

[0075] As illustrated more visibly in Figures 2 and 3, a flow vein of an internal flow of the turbomachine 2, and here the third flow vein 36, is provided with a discharge vein 38 which is connected to the flow vein of the internal flow downstream of the intermediate blower 12.

[0076] The discharge vein 38 is here connected on one side to the third flow vein 36, and opens on the other side onto a radially external surface of the turbomachine 2, the discharge vein 38 extending radially outwards from an opening formed in the external wall 36b of the third flow vein 36 remaining radially outside the first flow vein 30, in particular remaining radially between the primary flow vein 30 and the secondary flow vein 32 and opening into the secondary flow vein 32.

[0077] The turbomachine 2 further comprises a movable panel 40 between an upstream position in which the movable panel 40 allows air to circulate through the discharge channel 38 and a downstream position in which the movable panel 40 closes the discharge channel 38, the upstream position of the movable panel 40 being represented in [Fig.2] and the downstream position of the movable panel 40 in [Fig.3].

[0078] The movable panel 40 in the downstream position partly forms the outer wall 36b of the third flow vein 36.

[0079] The turbomachine 2 comprises an upstream fairing 42 forming the outer surface of the turbomachine 2 upstream of the discharge channel 38 and a downstream fairing 44 forming the outer surface of the turbomachine 2 downstream of the discharge channel 38. The upstream fairings 42 and downstream fairings 44 together define an outlet of the discharge channel 38. The movable panel 40 in the downstream position extends axially the upstream fairing 42 towards the downstream fairing 44.

[0080] The turbomachine 2 further comprises a pressurized box 46 which is located partly radially outside the discharge channel 38, which partially surrounds the discharge channel 38, and which is provided with an opening for the movable panel 40. The movable panel 40 in the upstream position is arranged partly in the pressurized box 46. In particular, the movable panel 40 moves through the opening of the pressurized box 46 when it passes from the upstream position to the downstream position and vice versa.

[0081] The movable panel 40 includes a seal which forms a seal by contact with a wall of the pressurized chamber 46 when the movable panel 40 is in the downstream position.

[0082] With reference to figures 4 and 5, the movable panel 40 comprises a perforated inner skin 48 and a honeycomb core 50 located radially outside the perforated inner skin 48.

[0083] The turbomachine 2 includes an acoustic resistive skin 52 located radially outside the movable panel 40, in particular radially between the movable panel 40 and the upstream fairing 42.

[0084] The perforated inner skin 48, the alveolar core 50 and the acoustic resistive skin 52 together form an acoustic stack.

[0085] The movable panel 40 includes a seal 54 which is in contact with the downstream fairing 44 when the movable panel 40 is in the downstream position so as to improve the sealing of the blocked discharge channel 38.

[0086] The turbomachine 2 further includes a control device 56 for the movable panel 40. The control device 56 includes rails 56a and actuators 56b, and is capable of positioning the movable panel 40 in the upstream or downstream position according to the air pressure measured by the pressure sensor 37.

[0087] The embodiment illustrated in Figures 5 and 6 differs from the embodiment illustrated in Figures 1 to 4 in that the discharge stream 38 comprises a plurality of air outlets 58 spaced from one another in the circumferential direction of the turbomachine 2. The air outlets 58 are located at the same longitudinal position and each extends circumferentially over a limited angular sector. The turbomachine 2 comprises, for example, between two and twelve air outlets 58, and here comprises ten air outlets 58.

[0088] The turbomachine 2 also comprises a plurality of movable panels 40 spaced apart along the circumferential direction of the turbomachine 2. The movable panels 40 are located at the same longitudinal position and each extends circumferentially over a limited angular sector. The turbomachine 2 comprises, for example, between two and twelve movable panels 40, and here comprises ten movable panels 40.

[0089] The movable panels 40 are arranged opposite the air outlets 58 of the discharge channel 38. Each movable panel 40 is movable between an upstream position in which the movable panel 40 in question allows air to circulate through the air outlet 58 concerned of the discharge channel 38 and a downstream position in which the movable panel 40 in question closes the air outlet 58 concerned of the discharge channel 38.

Claims

Demands

1. Turbomachine (2) for aircraft comprising at least one compressor (12, 14, 16), a combustion chamber (18), at least one turbine (20, 22), an external flow channel (32), and at least one internal flow channel (30, 36) through which flows a flow from at least one compressor (12, 14, 16) and which is provided with at least one discharge channel (38) having a connection to said internal flow channel (30, 36) located longitudinally downstream of said compressor (12, 14, 16), characterized in that it further comprises at least one movable panel (40) between an upstream position in which said movable panel (40) permits air to flow through said discharge channel (38) and a downstream position in which said movable panel (40) closes said discharge channel (38).

2. Turbomachine (2) according to claim 1, comprising a separation nozzle (28) of the external flow stream (32), referred to as secondary flow stream (32), and of said internal flow stream (30), referred to as primary flow stream (30), which extends through the combustion chamber (18).

3. Turbomachine (2) according to claim 1, comprising at least said internal flow channel (36), referred to as third flow channel (36), and another internal flow channel in which flows a flow from said compressor (12, 14, 16), and which extends through the combustion chamber (18), referred to as primary flow channel (30), and comprising a separating nozzle (34) of the primary flow channel (30) and the third flow channel (36), the third flow channel (36) being located at least partly radially between the primary flow channel (30) and the external flow channel (32), referred to as secondary flow channel (32).

4. Turbomachine (2) according to any one of claims 1 to 3, wherein said internal flow channel (30, 36) is radially delimited by an internal wall (30a, 36a) and by an external wall (30b, 36b), said discharge channel (38) extending radially outwards from an opening formed in the external wall (30b, 36b) of said internal flow channel (30, 36).

5. Turbomachine (2) according to claim 4, wherein said movable panel (40) in the downstream position forms part of the outer wall (30b, 36b) of said internal flow channel (30, 36).

6. Turbomachine (2) according to any one of claims 1 to 5, comprising an upstream fairing (42) and a downstream fairing (44) which together delimit an outlet of said discharge channel (38), said movable panel (40) in the downstream position axially extending the downstream fairing (44) in the direction of the downstream fairing (42).

7. Turbomachine (2) according to any one of claims 1 to 6, wherein said movable panel (40) extends circumferentially along a limited angular sector.

8. Turbomachine (2) according to any one of claims 1 to 7, comprising a pressurized box (46) which is located partly radially outside said discharge channel (38) and which partly surrounds said discharge channel (38), said movable panel (40) in the upstream position being arranged at least partly in the pressurized box (46).

9. Turbomachine (2) according to any one of claims 1 to 8, wherein said movable panel (40) comprises a perforated inner skin (48) and a honeycomb core (50) located radially outside the perforated inner skin (48), the turbomachine (2) comprising an acoustic resistive skin (52) located radially outside said movable panel (40) so as to form an acoustic stack consisting of the perforated inner skin (48), the honeycomb core (50) and the acoustic resistive skin (52).

10. Turbomachine (2) according to any one of claims 1 to 9, comprising a control device (56) for said movable panel (40) provided with rails (56a) and actuators (56b).

Images