Failure handling system for a turbomachine component cooling air circuit

Abstract

The invention relates to a failure treatment system (10) for a cooling air circuit (20) of turbomachine components, the circuit (20) being designed to communicate with a flow (V1) of the turbomachine so that pressurized cooling air is able to escape from said circuit (20) through the flow (V1). The treatment system (10) comprises a pressurized cavity (40), a fluid circulation duct (30) connected downstream to the circuit (20) and upstream to this cavity (40), a temperature-sensitive element (50) that closes the duct (30) when its temperature remains below a reference temperature (TR) and that degrades above this temperature to allow the circulation of a by-flow (FD) of gas through the duct (30) from the cavity (40) to the circuit (20), and a failure detection device (60) that emits a signal when the temperature-sensitive element (50) is degraded. Figure 3

Description

Title of the invention: Failure handling system for a cooling air circuit of turbomachine parts

[0001] An aircraft turbomachine comprises at least one compressor and one turbine coupled in rotation. Air is admitted into a duct of the gas turbine and compressed by the upstream compressor, then sent to a combustion chamber where it is mixed with fuel before being burned. The hot gases from this combustion are then expanded in a duct in a turbine located downstream of this chamber and rotate this turbine. The rotation of this turbine drives a shaft which in turn rotates the compressor.

[0002] In a turbomachine such as a twin-spool, twin-flow turbomachine extending along a longitudinal axis, the shaft driven by the turbine's rotation in turn drives not only the compressor but also the fan blades located upstream of the turbomachine. The rotation of the fan blades, along with the high-speed ejection of gases from the combustion chamber, contributes to the propulsion of the turbomachine.

[0003] In a turbomachine of the turbomotor type such as a helicopter, part of the kinetic energy of the hot gases (that which is not used to turn the compressor) turns a free turbine which in turn drives, directly or via a reducer, a receiving element (such as the main transmission box) which drives in particular a rotor shaft which carries the blades of the helicopter.

[0004] Aircraft turbomachinery includes cooling air circuits whose primary function is to cool certain turbomachine components. Typically, cooling air, also called fresh air, is drawn from a turbomachine compressor or from an external source and then injected towards the sensitive components to be cooled. In some cases, the cooling air circuit connects to a gas turbine duct, allowing pressurized cooling air to escape from the circuit through this duct. This helps to limit or prevent the intake of some of the gas flow circulating in the duct towards the cooling air circuit.

[0005] A failure of the cooling air circuit may occur. This failure may have several causes, such as a burst pipe or its blockage, or any other case of failure of the cooling air supply.

[0006] The present invention relates to a failure management system for a cooling air circuit of parts of an aircraft turbomachine, the circuit being intended to communicate with a gas turbine stream so that pressurized cooling air is able to escape from the circuit through the stream.

[0007] Indeed, a failure in the cooling air circuit leads to degraded operation of the turbomachine, with a change in the cooling air flow rate. This results in a decrease in the cooling performance of sensitive parts, as well as a risk of gas ingestion from the intake stream into the cooling circuit due to the pressure loss in the cooling air flow. This ingestion of hot gas from the intake stream, also known as gas reinjection, further impairs the cooling of sensitive parts.

[0008] Reduced cooling performance has a thermal impact on the parts being cooled, potentially shortening their lifespan. However, this type of failure in a cooling air circuit is generally undetectable during turbomachine operation. It is therefore a case of a dormant failure or degraded operation, the detection of which usually only occurs during turbomachine disassembly for maintenance.

[0009] To overcome this situation, parts of the turbomachine must be oversized to ensure they can operate without compromising flight safety throughout their entire service life. This safety principle has significant impacts on the sizing of the cooling air circuit and / or the size of the parts to be cooled. This results in suboptimal air intake and / or overall mass, negatively affecting the overall performance of the turbomachine. Description of the invention

[0010] The present invention aims to remedy these drawbacks.

[0011] The invention aims to provide a system for dealing with failures in a cooling air circuit for parts of an aircraft turbomachine that makes it possible to minimize or even eliminate the consequences of a failure in the cooling air circuit.

[0012] This objective is achieved by the fact that the fault handling system of a cooling air circuit comprises a pressurized cavity, a fluid circulation duct which is connected downstream to the cooling air circuit and upstream to the pressurized cavity, a temperature-sensitive element for closing the fluid circulation duct, and a fault detection device. The fault handling system is configured such that the temperature-sensitive element closes the duct as long as its temperature remains below a reference temperature, and that it degrades when its temperature reaches the reference temperature to allow the circulation of a by-flow of gas through the duct from the pressurized cavity. towards the cooling air circuit, the fault detection device being configured to emit a fault signal when the heat-sensitive component is degraded.

[0013] Thanks to these provisions, a failure in the cooling air circuit is addressed as soon as it occurs. The consequences of this failure, namely undesirable heating of turbomachine components, are minimized or even eliminated, and the turbomachine can continue to operate without undesirable degradation of these components. Furthermore, the sizing of these components can be optimized, and maintenance operations can be anticipated. In addition, a failure in the cooling air circuit is detected as soon as it occurs and during turbomachine operation, and an operator, such as a pilot, is able to be informed of this failure in real time by the fault signal that is emitted.

[0014] For example, the heat-sensitive closing element is a hot-melt shutter comprising a material that flows through the shutter from upstream to downstream and whose melting temperature is less than or equal to the reference temperature

[0015] For example, the reference temperature is between the temperature of the cooling air circuit and the temperature of the gases in the stream at an orifice of the circuit in the stream during the operation of the turbomachine.

[0016] For example, the fault detection device is located in the conduit upstream of the thermosensitive element and is configured to emit the fault signal in the presence of the derived gas flow.

[0017] For example, the failure detection device includes a pressure sensor that is capable of detecting the pressure drop that occurs after the degradation of the thermosensitive element.

[0018] For example, the pressurized cavity is intended to be located at the level of the turbomachine compressor.

[0019] For example, the cooling air circuit is intended to cool elements of a turbine of the turbomachine whose blades are located in the first vein.

[0020] The invention also relates to an aircraft turbomachine comprising a failure handling system for a cooling air circuit of parts of the turbomachine according to the invention.

[0021] The invention also relates to a method for treating failures in an air cooling circuit for parts of an aircraft turbomachine.

[0022] According to the invention, the method comprises the following steps (a) A system for handling failures of the cooling air circuit is provided according to the invention; (b) The fault detection device is monitored.

[0023] The invention will be better understood and its advantages will become more apparent upon reading the following detailed description of embodiments shown by way of non-limiting examples. The description refers to the accompanying drawings in which:

[0024] [Fig-1] Fig. 1 is a schematic longitudinal cross-sectional view of a twin-body, twin-flow turbomachine.

[0025] [Fig.2] The [Fig.2] is a schematic view of region R of the [Fig.1] illustrating a failure handling system of a turbomachine cooling air circuit in normal operation of this circuit.

[0026] [Fig.3] The [Fig.3] is an enlarged schematic view of region R of the [Fig.1] illustrating the failure handling system of a turbomachine cooling air circuit in the event of failure of this circuit.

[0027] [Fig.4] Fig.4 is a longitudinal section view of an embodiment of the thermosensitive sealing element of the failure handling system of the turbomachine cooling air circuit.

[0028] [Fig.5] The [Fig.5] is a schematic longitudinal cross-sectional view of a turboprop-type turbomachine.

[0029] [Fig.6] The [Fig.6] is an enlarged schematic view of region S of the [Fig.5] illustrating a failure handling system of a turbomachine cooling air circuit in normal operation of this circuit. Detailed description of the invention

[0030] Consider a turbomachine 1 with longitudinal axis X, which is its axis of rotation. In the description below, the terms "internal" and "inner" refer to an element oriented towards the longitudinal axis X or located closer to this axis. The terms "external" and "outer" refer to an element oriented in the opposite direction to the longitudinal axis X or located further from this axis. The terms "upstream" and "downstream" are relative to the direction of air and gas flow during the operation of the turbomachine, i.e., from left to right in the figures. The term "radial" refers to a position or direction in a transverse plane perpendicular to the longitudinal axis X.

[0031] By way of example, the invention is described in the case where the turbomachine 1 is a twin-spool, twin-flow turbomachine. Figure 1 illustrates such a turbomachine 1, in longitudinal view. This turbomachine 1 comprises a nacelle 2 with a fan 3 having blades. The turbomachine 1 has a hub 8m, one of whose rotor elements carries the ring of blades forming the propeller of the fan 3. The hub 8m consists of alternating rotor and stator elements along the longitudinal axis X. These rotor elements of the hub 8m are carried by a rotor shaft 8 which extends along the longitudinal axis X. Radially outside the hub 8m and downstream of the blower propeller 3 is an internal casing 9 which is coaxial with the hub 8m.

[0032] In normal operation of the turbomachine 1, an airflow (called secondary flow F2) circulates in an annular channel, called the second channel V2, which extends between the inner casing 9 and the nacelle 2. Another annular channel, called the first channel VI (in which an airflow Fl circulates), extends between the hub 8m and the inner casing 9. The first channel V1 includes, downstream of the fan 3, a compressor 4 and a combustion chamber 5. This compressor 4 includes, upstream, a low-pressure compressor 4a and, downstream, a high-pressure compressor 4b. The air compressed by the compressor 4 is admitted downstream into the combustion chamber 5 and mixed with fuel before being burned. The hot gases from this combustion are then expanded in a turbine in the first channel VI and rotate this turbine.This turbine comprises a high-pressure turbine 6 located downstream of the combustion chamber 5 and a low-pressure turbine 7 located downstream of the high-pressure turbine 6. The movable turbine blades of the high-pressure turbine 6 and the low-pressure turbine 7 are mounted on discs which are themselves mounted on the rotor shaft 8, and which therefore drive the rotor shaft 8 in rotation. The rotor shaft 8 drives the rotor elements of the hub 8m and the fan 3. The rotation of the fan 3, together with the high-speed ejection of gases exiting the combustion chamber 5, contributes to the propulsion of the turbomachine 1.

[0033] The high-pressure turbine 6 comprises at least one stage E and the low-pressure turbine 7 comprises a plurality of coaxial stages E, each stage E consisting of a ring of fixed turbine blades (stator) Ts, also called a distributor, followed by a ring of rotating turbine blades (rotor) TR spaced circumferentially around the longitudinal axis X, which thus drive the rotor shaft 8 in rotation. The stages E are located in the first flow VL

[0034] Fig. 2 is an enlarged schematic representation of region R of Fig. 1, which includes the high-pressure turbine 6 and the first upstream stage E of the low-pressure turbine 7, which is therefore located immediately downstream of the high-pressure turbine 6. Between these two turbines (6, 7) is a space 90, which is part, for example, of a casing called the inter-turbine casing.

[0035] The turbomachine 1 includes a cooling air circuit 20 for cooling several parts of this turbomachine, in particular the fixed turbine blade ring Ts-6 and the moving turbine blade ring TR-6 of the high-pressure turbine 6 and the fixed turbine blade rings Ts-7 and moving turbine blade rings TR-7 of the upstream first stage E of the low-pressure turbine 7. In this respect, the air circuit Cooling unit 20 comprises one or more air supply lines, schematically represented by arrow F20 in [Fig. 2]. This air supply line is, for example, connected upstream to one of the compressors (4a, 4b) so that air drawn from this compressor can serve as cooling air, also called ventilation air or fresh air, in comparison to the much higher temperature of the gas flow circulating in the primary stream VI through the turbines (6, 7). Alternatively, the air supply line is connected to other conventional fresh air sources.

[0036] This supply pipe opens directly downstream of the disc carrying the rotating turbine blade ring TR-6, so that the cooling air propagates radially outwards into a first upstream cooling air cavity 21, belonging to the cooling air circuit 20. The first upstream cooling air cavity 21 is located in the space 90 between the turbines (6, 7). This first cavity 21 is delimited upstream by the disc carrying the rotating turbine blade ring TR-6 of the high-pressure turbine 6, and downstream by (fixed) stator elements 19 of the turbomachine 1 on which the moving parts of the turbine (6, 7) bear. These stator elements 19 are located in the space 90 and are schematically represented by walls.

[0037] Due to the high pressure of the cooling airflow circulating in the first cavity 21, this air escapes from the circuit 20 through a first orifice (annular gap) 91 defined axially between a radially internal platform of the rotating turbine blade ring TR-6 and a radially internal stator boundary wall of the first flow VI. Thus, the pressurized cooling air enters the first flow VI through the first annular orifice 91, establishing fluid communication between the cooling air circuit 20 and the first flow VI. This airflow is indicated by arrow F21. The airflow F21 prevents, or at least limits, the ingestion of a portion of the flow Fl into the cooling air circuit 20, and allows the performance of this circuit 20 to be maintained.

[0038] Advantageously, part of the cooling air located in the first cavity 21 flows downstream to one or more other cavities of the cooling air circuit 20, and in particular to a second cavity 22 which is delimited upstream by stator elements 19 and downstream by the disc carrying the ring of movable turbine blades TR-7 of the low pressure turbine 7. This airflow is indicated by the arrow F23.

[0039] Also due to the high pressure of the cooling airflow circulating in the second cavity 22, this air escapes from the cooling air circuit 20 through a second orifice (annular gap) 92 defined axially between a platform radially internal to the rotating turbine blade ring TR-7, and a radially internal stator wall delimiting the first flow VI. Thus, pressurized cooling air enters the first flow V1 through the second annular orifice 92, establishing fluidic communication between the cooling air circuit 20 and the first flow VI. This airflow is indicated by arrow F22. The airflow F22 also prevents, or at least limits, the ingestion of a portion of the flow Fl into the cooling air circuit 20.

[0040] The cooling air circuit 20 communicates with the first flow V1 via the annular ports (91, 92). The cooling air flows F21, F22 and F23, which are part of the cooling air circuit 20, provide cooling for the disc carrying the rotating turbine blade ring TR-6 of the high-pressure turbine 6 and for the disc carrying the rotating turbine blade ring TR-7 of the upstream first stage E of the low-pressure turbine 7. The cooling air flows F21, F22 and F23 also provide cooling for other parts such as the platforms of these rotating turbine blade rings (TR-6, TR-7).

[0041] The failure treatment system 10 for the cooling air circuit 20 includes a fluid circulation duct 30 which is connected downstream to this cooling air circuit 20 and upstream to a pressurized cavity 40 containing gas or air, as illustrated in [Fig. 2]. This pressurized cavity 40 is part of the treatment system 10. This pressurized cavity 40 is located, for example, at the compressor 4 of the turbomachine 1, which allows the direct use of air compressed by the turbomachine 1.

[0042] The failure handling system 10 also includes a temperature-sensitive element 50 for closing this fluid circulation conduit 30. The temperature-sensitive element 50 closes the conduit 30 as long as its temperature remains below a reference temperature TR. This situation corresponds to normal operation of the turbomachine 1 and is illustrated in [Fig.2].

[0043] In the event of a failure of the cooling air circuit 20, the pressure in the cavities (21, 22) of this cooling air circuit 20 decreases, such that hot gases pass from the first vein V1 into these cavities through the first orifice 91 and possibly the second orifice 92, which prevents the cooling of the discs carrying the rotating turbine blade rings (TR-6, TR-7). The flow of these hot gases is represented by the white arrows FC in [Fig. 3], which illustrates the case of failure of the cooling air circuit 20. The temperature-sensitive element 50 degrades when its temperature reaches the reference temperature TR to allow the circulation of a by-flow FD of gas through the conduit 30 from the cavity Pressurized air 40 flows to the cooling air circuit 20 under the pressure present in this pressurized cavity 40. For example, duct 30 opens into the first cavity 21. This diverted flow FD of fresh gas, at a lower temperature than the hot gases from the first stream VI, opposes the hot gas flow FC and prevents or reduces the entry of these hot gases. The diverted flow FD of fresh gas thus cools the cavities (21, 22), the discs carrying the rotating turbine blade rings (TR-6, TR-7), and other adjacent components.

[0044] The heat-sensitive sealing element 50 is, for example, a hot-melt seal 51 comprising a material that flows through the seal 51 from upstream to downstream and whose melting temperature is less than or equal to the reference temperature TR. For example, the hot-melt seal 51 comprises a threaded ring 511 screwed onto a threaded portion of the circulation conduit 30, and includes a core 512 made of this material and surrounded by this ring 511. This embodiment is illustrated in [Fig. 4]. For example, the threaded ring 511 is steel or Inconel, and the core material 512 is a bismuth-based non-ferrous alloy.

[0045] The reference temperature TR is between the temperature of the cooling air circuit 20 and the temperature of the gases in the first flow VI at an orifice (91, 92) of the circuit 20 in the flow VI during the operation of the turbomachine. For example, the temperature of the cooling air circuit 20 is approximately 230°C and the temperature of the gases in the first flow VI at an orifice (91, 92) is approximately 900°C, and the reference temperature TR is equal to 380°C.

[0046] The failure handling system 10 also includes a failure detection device 60 configured to emit a failure signal when the temperature-sensitive element 50 is degraded. As illustrated in Figures 2 and 3, the failure detection device 60 is located in the fluid circulation duct 30 upstream of the temperature-sensitive element 50 and is configured to emit the failure signal in the presence of the by-flow FD of gas, i.e., when the gas is flowing in the duct 30. For example, the failure detection device 60 includes a pressure sensor 61 capable of detecting the pressure drop that occurs in the fluid circulation duct 30 after the temperature-sensitive element 50 is degraded. The failure detection device 60 then sends a failure signal to the avionics, for example via cable (not shown), so as to inform an operator such as the pilot.

[0047] Alternatively, the fault detection device 60 directly detects the degraded state of the thermosensitive element 50.

[0048] The invention also relates to a turbomachine of the turboshaft type, such as a helicopter. Such a turboshaft engine 1 is illustrated in [Fig. 5]. The turboshaft engine 1 is oriented along a longitudinal axis X and comprises a compressor 4 and an upstream turbine 6 coupled in rotation, a combustion chamber 5, and a downstream turbine 7. Air is admitted into a duct of the turboshaft engine and compressed by the compressor 4, then sent into the combustion chamber 5 where it is mixed with fuel before being burned. The hot gases from this combustion are then expanded in a first duct V1 in an upstream turbine 6 located downstream of this chamber 5, and rotate this upstream turbine 6. The rotation of this upstream turbine drives a shaft 8, which in turn rotates the compressor. The hot gases from the combustion are also expanded in a downstream turbine 7 located on the first duct VI downstream of the upstream turbine 6.The downstream turbine 7 in turn drives, directly or via a reduction gear, a secondary rotor shaft 8s which carries the blades of the helicopter 1.

[0049] Fig. 6 is an enlarged schematic representation of region S of Fig. 5, which includes the upstream turbine 6 and the downstream turbine 7 which is located immediately downstream of the downstream turbine 6.

[0050] Between these two turbines (6, 7) is a space 90. The turboshaft engine 1 includes a cooling air circuit 20 intended to cool the upstream turbine blade ring 6 and the downstream turbine blade ring 7. The cooling air circuit 20 includes one or more air supply ducts, schematically represented by the arrow F20, which supply fresh air.

[0051] This supply pipe opens into a first upstream cooling air cavity 21, belonging to the cooling air circuit 20. The first upstream cooling air cavity 21 is located in the space 90 between the turbines (6, 7). This first cavity 21 is delimited upstream by the upstream turbine 6, and downstream by (fixed) stator elements 19 of the turboshaft engine 1, which are located in the space 90 and are schematically represented by walls. The air circulating in the first cavity 21 escapes from the circuit 20 through a first orifice 91 in the radially internal wall of the first VL duct. This airflow is indicated by arrow F21.

[0052] Advantageously, part of the cooling air located in the first cavity 21 flows downstream to one or more other cavities of the cooling air circuit 20, and in particular to a second cavity 22 which is delimited upstream by stator elements 19 and downstream by the downstream turbine 7. The air flowing in the second cavity 22 escapes from the circuit 20 through a second orifice 92 in the radially internal wall of the first vein VL. This airflow is indicated by the arrow F22.

[0053] The failure handling system 10 of the cooling air circuit 20 comprises a fluid circulation duct 30 which is connected downstream to this cooling air circuit 20 and upstream to a pressurized cavity 40, as illustrated in [Fig. 6] during normal operation. The failure handling system 10 also comprises a temperature-sensitive element 50 for closing this fluid circulation duct 30, which degrades above a reference temperature to allow air to flow, and a failure detection device 60 configured to emit a failure signal when the temperature-sensitive element 50 degrades. In Figures 5 and 6, the elements bearing the same reference numerals as in Figures 1 to 3 have an identical or similar function.

[0054] The invention is described above in the case of a cooling air circuit 20 which is intended in particular to cool elements of a turbine (6, 7) whose blades are located in the first VL channel. More generally, the invention applies to the cooling of any part of a turbomachine 1.

[0055] The invention also relates to a method for handling failures in a cooling air circuit 20 of parts of an aircraft turbomachine. This method uses a failure handling system 10 for the cooling air circuit 20, as described above (step (a)). In a step (b), an operator monitors the failure detection device 60 and the sending of a failure signal by this failure detection device 60.

Claims

Demands

1. A failure treatment system (10) for a cooling air circuit (20) of parts of an aircraft turbomachine, said circuit (20) being intended to communicate with a channel (VI) of said turbomachine so that pressurized cooling air is able to escape from said circuit (20) through said channel (VI), the treatment system (10) being characterized in that it comprises a pressurized cavity (40), a fluid circulation duct (30) which is connected downstream of said cooling air circuit (20) and upstream of said pressurized cavity (40), a temperature-sensitive element (50) for closing said fluid circulation duct (30) and a failure detection device (60), said treatment system (10) being configured such that said temperature-sensitive element (50) closes said duct (30) as long as its temperature remains below a reference temperature (TR),and that it degrades when its temperature reaches said reference temperature (TR) to allow the circulation of a derived flow (FD) of gas through said conduit (30) from said pressurized cavity (40) to said cooling air circuit (20), said fault detection device (60) being configured to emit a fault signal when said temperature-sensitive component (50) is degraded.

2. Failure treatment system (10) according to claim 1 wherein said thermosensitive sealing element (50) is a hot-melt shutter (51) comprising a material which passes through said shutter (51) from upstream to downstream and whose melting temperature is less than or equal to said reference temperature (TR).

3. Failure treatment system (10) according to claim 1 or 2 said reference temperature (TR) is between the temperature of said cooling air circuit (20) and the temperature of the gases in said vein (VI) at an orifice of said circuit (20) in said vein (VI) during the operation of said turbomachine.

4. Failure handling system (10) according to any one of claims 1 to 3 wherein said failure detection device (60) is located in said conduit (30) upstream of said heat-sensitive member (50) and is configured to emit said failure signal in the presence of said derived flow (DF) of gas.

5. Failure handling system (10) according to claim 4 wherein said failure detection device (60) comprises a pressure sensor (61) which is capable of detecting the pressure drop which occurs after the degradation of said thermosensitive element (50).

6. Failure treatment system (10) according to any one of claims 1 to 5 wherein said pressurized cavity (40) is intended to be located at the compressor of said turbomachine.

7. Failure treatment system (10) according to any one of claims 1 to 6 wherein said cooling air circuit (20) is intended to cool elements of a turbine (6, 7) of said turbomachine whose blades are located in said first vein (VI).

8. Aircraft turbomachine comprising a failure treatment system (10) for a cooling air circuit (20) of parts of said turbomachine according to any one of the preceding claims.

9. A method for handling failures of a cooling air circuit (20) of parts of an aircraft turbomachine characterized in that it comprises the following steps: (a) A failure handling system (10) is provided for said cooling air circuit (20) according to any one of claims 1 to 7; (b) Said failure detection device (60) is monitored.

Images