De-icing circuit for a turbomachine separation nozzle

Abstract

This description concerns an assembly for a turbomachine. The assembly comprises: - an annular separation nozzle, and - an annular housing fixed to the nozzle downstream of the nozzle; the housing comprising at least one de-icing air duct extending from downstream to upstream, an outlet of the air duct opening onto the nozzle, and a diffusion chamber, consisting of an annular duct positioned downstream of the air duct, the diffusion chamber comprising at least one inlet supplied with hot air from an engine and an outlet connected to the air duct, the air duct and the diffusion chamber both being formed with the inner wall of the housing. Figure for the abstract: Fig. 3

Description

Title of the invention: De-icing circuit for a turbomachine separation nozzle Technical field

[0001] This presentation concerns the field of turbomachines and more particularly concerns the defrosting of flow separation nozzles equipping turbomachines. STATE OF THE ART

[0002] In an aeronautical turbomachine of the double-body and double-flow type, a flow vein of the primary flow (or hot flow) and a flow vein of the secondary flow (or cold flow) are separated downstream of a fan by a separation nozzle.

[0003] The separation nozzle is an internal element of the turbomachine which is particularly exposed to the formation of ice in icing flight conditions. Indeed, in such conditions, humid air and rain can penetrate inside the turbomachine and deposit on the separation nozzle in the form of blocks of ice. To avoid the formation of blocks of ice on the separation nozzle, it is known to inject hot air taken from a stage of a compressor of the turbomachine into the separation nozzle. This air is typically conveyed by added conduits which pass through a casing of the turbomachine on which the separation nozzle is fixed and are fixed on the separation nozzle to discharge the taken air therein.

[0004] However, in order to optimize the mass and aerodynamic performance of the turbomachine while maintaining the same ratio between the sections of the flow veins of the primary and secondary flows, engine manufacturers seek to reduce the radii of these veins and the inter-vein distance (i.e. the thickness of the nozzle).

[0005] Document FR3095229 describes conduits which cross the primary vein from downstream to upstream to open into an opening in an internal wall of the nozzle. Such an arrangement of the channels makes it possible to reduce the radial thickness of the nozzle but requires thickening an arm positioned in the primary vein and which the conduit crosses. Thus, this arrangement induces modifications in the flow of the stream circulating in the primary vein.

[0006] Document FR3004485 describes a casing fixed downstream of the nozzle comprising an internal ferrule. A conduit is directly integrated into the internal ferrule. But this other arrangement requires the installation of an additional circumferential pipe, attached to the casing and fixed to the casing.

[0007] There is therefore a need to reduce the radial size and to reduce the number of parts in order to facilitate the assembly of the turbomachine and to increase its performances. GENERAL STATEMENT

[0008] One aim of the disclosure is to solve at least one of the problems previously mentioned.

[0009] For this purpose, according to one aspect of the present disclosure, a turbomachine assembly is proposed. The assembly comprises:

[0010] - an annular separation nozzle, configured to be positioned downstream of a turbomachine fan, in a direction of circulation of a gas flow from the fan, and for separating said flow into a primary flow and a secondary flow, the nozzle comprising an annular interior cavity, and

[0011] - an annular casing fixed to the spout downstream of the spout, the casing comprising a wall internal annular externally delimiting a primary flow vein in which the primary flow circulates.

[0012] The housing comprises:

[0013] - at least one defrost air duct extending from downstream to upstream, an outlet of the air duct opening into the interior cavity of the nozzle, and

[0014] - a diffusion chamber, consisting of an annular conduit positioned downstream of the air duct, the diffusion chamber comprising at least one inlet supplied with hot air from a motor and an outlet connected to the air duct.

[0015] Both the air duct and the diffusion chamber are formed with the inner wall of the casing.

[0016] Such an assembly makes it possible to do without the presence of an added circumferential pipe in order to reduce the size of the turbomachine at the separation nozzle. The absence of an added circumferential pipe on the casing makes it possible to significantly reduce the number of parts because it is no longer necessary to use screw and holding elements. This results in a significant saving of time when assembling the assembly, as well as a saving on the mass of the assembly.

[0017] This assembly also solves the problem of the risks of rupture of a pipeline caused by overpressure or leaks linked to a lack of sealing.

[0018] Advantageously, but optionally, the assembly disclosed comprises at least one of the following characteristics, taken alone or in any combination: - the casing comprises an annular radial flange positioned downstream of the internal wall and radially external to the internal wall, the diffusion chamber forming a junction between the internal wall and the radial flange;

[0019] - the inner wall comprises at least one plate-shaped stiffener extending from the inner wall to the radial flange along a longitudinal plane comprising a longitudinal axis of the turbomachine;

[0020] - the stiffener is positioned on the internal wall at the level of the air duct, so as to form a wall of the air duct;

[0021] - the casing, the air duct and the diffusion chamber are produced by manufacturing additive or by casting;

[0022] - a plurality of air ducts are distributed regularly around the internal wall, the diffusion chamber comprising as many outlets as air ducts;

[0023] - at least one inlet of the diffusion chamber is directly supplied with air taken downstream in the primary flow vein;

[0024] - the diffusion chamber comprises several inlets.

[0025] According to another aspect, there is provided an aircraft turbomachine comprising: a fan and an assembly as previously described, the assembly being positioned downstream of the fan.

[0026] According to a third aspect, there is provided an aircraft comprising an airframe and a turbomachine as previously described, in which the turbomachine is fixed to the airframe. DESCRIPTION OF FIGURES

[0027] Other characteristics, aims and advantages will emerge from the following description, which is purely illustrative and non-limiting, and which must be read in conjunction with the appended drawings in which:

[0028] [Fig.l] illustrates an aircraft schematically, according to an embodiment of the present disclosure;

[0029] [Fig.2] illustrates a schematic sectional view of a propulsion unit for an aircraft, according to an embodiment of the present disclosure;

[0030] [Fig. 3] schematically illustrates an assembly according to an embodiment of the present disclosure;

[0031] [Fig.4] schematically illustrates a perspective view of a casing of the assembly, according to an embodiment of the present disclosure;

[0032] [Fig.5] schematically illustrates a sectional view of a casing of the assembly, according to an embodiment of the present disclosure.

[0033] Throughout the figures, similar elements bear identical references. DETAILED DESCRIPTION Aircraft

[0034] [Fig.l] illustrates an aircraft 100 is a device configured to rise and move in the air, and may, for example, be an airplane, civil or military, or even a helicopter. An aircraft 100 comprises an airframe which, in the case of an airplane, is composed of a fuselage, a wing structure comprising two wings, empennages, flight control surfaces and landing gear. Propulsion unit

[0035] A propulsion unit 1 as illustrated by [Fig.2] comprises an engine 2 (or turbomachine) and a nacelle 3, and has a main direction extending along a longitudinal axis XX. The propulsion unit 1 is configured to be fixed to the airframe of the aircraft 100, for example under its wings, in the case of an airplane, and this by means of a pylon (or mast). The propulsion unit 1 can also be mounted on the wing of the airplane or at the rear of its fuselage, or even be integrated into its fuselage.

[0036] The engine 2 may be a twin-spool, twin-flow, direct-drive ducted turbojet engine, as described below, but may also have a different number of spools and / or flows, and / or be another type of turbojet engine, such as a geared turbojet engine or a turboprop, with or without afterburner, ducted or unducted.

[0037] Unless otherwise specified, the terms “upstream” and “downstream” are used with reference to the overall direction of airflow through the propulsion unit 1 in operation. Similarly, an axial direction corresponds to the direction of the longitudinal axis XX and a radial direction is a direction perpendicular to the longitudinal axis XX and intersecting the longitudinal axis XX. Furthermore, an axial plane is a plane containing the longitudinal axis XX and a radial plane is a plane perpendicular to the longitudinal axis XX. A circumference is understood to be a circle belonging to a radial plane and whose center belongs to the longitudinal axis XX. A tangential or circumferential direction is a direction tangent to a circumference: it is perpendicular to the longitudinal axis XX but does not pass through the longitudinal axis XX.Finally, the adjectives "inner" (or "internal") and "outer" (or "external") are used with reference to a radial direction so that the inner part of an element is, in a radial direction, closer to the longitudinal axis XX than the outer part of the same element.

[0038] The engine 2 comprises, from upstream to downstream, a fan 20, a compressor section 22, comprising a low pressure compressor 220 and a high pressure compressor 222, a combustion chamber 24 and a turbine section 26, comprising a high pressure turbine 262 and a low pressure turbine 260.

[0039] The compressor section 22 comprises a succession of stages each comprising a wheel of moving blades (rotor) rotating in front of a wheel of fixed blades (stator). The turbine section 26 also comprises a succession of stages each comprising a wheel of fixed blades (stator) behind which a wheel of moving blades (rotor) rotates.

[0040] The fan 20, the rotor part of the low pressure compressor 220, and the rotor part of the low pressure turbine 260 are connected to each other by a low pressure shaft 280 extending along the longitudinal axis XX, thus forming a low pressure body. pressure 20, 220, 260, 280. The rotor part of the high pressure compressor 222 and the rotor part of the high pressure turbine 262 are connected to each other by a high pressure shaft 282 extending along the longitudinal axis XX, thus forming a high pressure body 222, 262, 282. The low pressure shaft 280 is generally housed, over a section of its length, in the high pressure shaft 282 and is coaxial with the high pressure shaft 282.

[0041] The compressor section 22, the combustion chamber 24 and the turbine section 26 are surrounded by a motor casing 23, otherwise called hubs (or "fan frame" or "fan case" according to English terminology), to which the stator parts of the low pressure compressor 220, the high pressure compressor 222, the high pressure turbine 262 and the low pressure turbine 260 are connected, while the fan 20 is surrounded by a fan casing 25. The motor casing 23 and the fan casing 25 are connected to each other by profiled arms 27 forming rectifiers (or OGV for "Outlet Guide Vanes" in English terminology) distributed circumferentially all around the longitudinal axis XX. At least some of these arms 27 can be provided structural.The longitudinal axis XX defines the axis of rotation for the fan 20, the rotor parts of the compressor section 22 and the rotor parts of the turbine section 26, in other words for the low pressure body 20, 220, 260, 280 and the high pressure body 222, 262, 282, which are each capable of being driven in rotation about the longitudinal axis XX relative to the engine casing 23 and the fan casing 25.

[0042] The nacelle 3 extends radially outside the engine 2, all around the longitudinal axis XX, so as to surround both the fan casing 25 and the engine casing 23, and to define, with a downstream portion of the engine casing 23, a downstream portion of a secondary duct B, the upstream portion of the secondary duct B being defined by the fan casing 25 and an upstream portion of the engine casing 23. The upstream portion of the nacelle 3 further defines an air inlet 29 through which the fan 20 sucks in the air flow circulating through the propulsion unit 1. The nacelle 3 is integral with the fan casing 25 and attached and fixed to the aircraft 100 by means of the mast.

[0043] The engine 2 may also comprise at least one accessory gearbox, called AGB (for “Accessory gear box” in English terminology), for example housed in a cavity provided within the nacelle 3. The accessory gearbox comprises a set of rotating elements, such as gears, making it possible to drive a plurality of shafts in rotation around their own axis, accessories being mounted on these shafts to derive useful mechanical power from their rotation. The set of gears is itself driven using a power take-off shaft (or RDS for “Radial Drive Shaft” in English terminology) connecting, possibly via a transfer case, the gearbox of accessories to at least one of the high pressure body 222, 262, 282 and the low pressure body 20, 220, 260, 280, typically by being meshed with at least one of the high pressure shaft 282 and the low pressure shaft 280. In this regard, the power take-off shaft may extend inside a longitudinal cavity provided within one of the arms 27. In this way, mechanical power is capable of being taken from at least one of the high pressure body 222, 262, 282 and the low pressure body 20, 220, 260, 280 to be delivered to the at least one of the accessories via the accessory housing.

[0044] Engine crankcase

[0045] The engine casing 23, as illustrated in [Fig. 3], is annular and comprises an inner wall 231 (or inner shell) and an outer wall 232. For simplicity, the engine casing 23 is hereinafter simply called casing 23. A primary vein A is located radially internal to the inner wall 231 of the casing 23 and a secondary vein B extends between the outer wall 232 and an inner wall of the fan casing 25. The casing 23 thus forms an inter-vein compartment.

[0046] The casing 23 further comprises a radial flange 233 positioned downstream of the inner wall 231 and radially external to the inner wall 231. In other words, the radial flange 233 extends in the casing 23 between the inner wall 231 of the casing 23 and the outer wall 232 of the casing 23. The radial flange 233 is for example configured to reinforce the rigidity of the casing 23 and also provides a fire wall function.

[0047] Furthermore, and as illustrated by Figures 4 and 5, the casing 23 comprises one or more plate-shaped stiffeners 234 extending axially outward between the inner wall 231 of the casing 23 and the flange of the casing 23. The stiffeners 234 are preferably distributed uniformly in the casing 23 all around the longitudinal axis XX of the casing 23. The stiffeners 234 are configured to ensure better rigidity of the casing 23.

[0048] In operation, the fan 20, illustrated by [Fig. 2], draws in an air flow, a portion of which, circulating within the primary vein A passing through the engine casing 23 from one side to the other, is successively compressed within the compressor section 22, ignited within the combustion chamber 24 by combustion of fuel, and expanded within the turbine section 26 before being ejected from the engine 2. Another portion of the air flow circulates within the secondary vein B which takes an elongated annular shape surrounding the engine casing 23, the air drawn in by the fan 20 being straightened by the straighteners 27 and then ejected from the propulsion unit 1. This thrust can, for example, be used for the benefit of the aircraft 100 on which the propulsion unit 1 is attached and fixed. In this way, the propulsion unit 1 generates thrust.

[0049] Annular separation nozzle

[0050] The portion of air flow passing through the primary vein A, called the primary flow, and the portion of air flow passing through the secondary vein B, called the secondary flow, are separated from each other upstream of the casing 23 by an annular separation nozzle 21. The casing 23 and the nozzle 21 form an assembly E, as illustrated in [Fig.3]. The nozzle 21 is positioned downstream of the fan 20 and is fixed upstream of the casing 23. The nozzle 21 forms the upstream separation point between the primary flow path A and the secondary flow path B. The nozzle 21 comprises an inner wall 211, an outer wall 212 and an axial wall 213 separating the nozzle 21 from the casing 23. The inner wall 211 of the nozzle 21 and the outer wall 212 of the nozzle 21 are respectively in the extension of the inner wall 231 of the casing 23 and the outer wall 232 of the casing 23. The inner wall 211 of the nozzle 21 and the outer wall 212 of the nozzle 21 are joined to each other upstream of the nozzle 21.The axial wall 213, or downstream wall, is positioned downstream of the spout 21 and is in contact with the upstream part of the casing 23.

[0051] The separating spout 21 has an annular internal cavity 214. The internal cavity 214 is radially between the internal wall 211 of the spout 21 and the external wall 212 of the spout 21 and axially between the upstream junction of the internal wall 211 of the spout 21 and the downstream wall 213 of the spout 21. Preferably, the cavity 214 communicates with the casing 23 through the downstream wall 214 of the spout 21 and by means of orifices 215 provided for this purpose in the downstream wall 214 of the spout 21. The downstream wall 213 of the spout 21 comprises at least one orifice 215 and preferably several orifices 215 distributed uniformly all around the longitudinal axis XX of the spout 21.

[0052] Defrosting device

[0053] The casing 23 further comprises a defrosting device, as illustrated in FIGS. 3 to 5. The defrosting device is configured to convey hot air into the interior cavity of the spout 21 in order to heat it. Indeed, the spout 21 is subjected to low temperatures which can prove problematic due to the creation of pieces of ice induced by the icing of the spout 21.

[0054] The defrosting device is configured to take hot air from the engine 2, downstream in the primary flow stream A. The hot air is for example taken from a compressor of the engine 2 such as the high pressure compressor 222 and communicates it to the internal cavity 213 of the nozzle 21 via the orifices 215 provided in the downstream wall 214 of the nozzle 21.

[0055] The defrosting device preferably comprises a diffusion chamber 235 and one or more air ducts 236. The defrosting device 234 is formed in the casing 23 and forms a single piece with the casing 23. In other words, the casing 23, the diffusion chamber 235 and the air ducts 236 are manufactured from a single piece during a single manufacturing process such as for example by casting using cores in order to provide the ducts or by fa additive building. Advantageously, the diffusion chamber 235 and the air ducts 236 are formed in a wall of the casing 23 such as, for example, in the internal wall 231 of the casing 23.

[0056] The air duct 236 is a duct extending longitudinally between the nozzle 21 and the diffusion chamber 235 positioned downstream of the air duct 236. The air duct 236 fluidly connects the interior cavity 213 of the nozzle 21 to the diffusion chamber 235. The air duct 236 comprises an inlet 2361 and an outlet 2362, as illustrated in [Fig. 5]. The inlet 2361 is connected to the diffusion chamber 235 and the outlet 2362 is connected to an orifice 215 of the downstream wall 214 of the spout 21. The defrosting device 234 comprises at least one air duct 236 and preferably a plurality of air ducts 236. The air ducts 236 are advantageously positioned uniformly and at regular intervals in the casing 23 around the longitudinal axis XX. Each air duct 236 communicates upstream with an orifice 215 of the downstream wall 214 of the spout 21 and communicates downstream with the diffusion chamber 235.

[0057] Advantageously, each air duct 236 is positioned in the casing 23 at the level of a stiffener 234.

[0058] The diffusion chamber 235 is an annular duct extending into the engine casing 23 and all around the primary stream. The diffusion chamber 235 comprises one or more inlets 2351 fluidly connected to a duct 237 for supplying hot air to the engine 2. The diffusion chamber 235 comprises as many outlets 2352 as the number of air ducts 236, each outlet 2352 of the diffusion chamber 235 being fluidly connected to an air duct 236. The diffusion chamber 235 thus makes it possible to distribute the hot air taken from the engine 2 throughout the interior cavity 213 of the nozzle 21 via the plurality of air ducts 236.

[0059] Advantageously, the diffusion chamber 235 is positioned in the casing 23 at the junction of the internal wall 231 of the casing 23 and the radial flange 233 of the casing 23.

[0060] Advantages

[0061] The formation of the air ducts 236 and the diffusion chamber 235 directly in the internal wall 231 of the casing 23 makes it possible to reduce the radial thickness of the casing 23 and thus to limit the radial dimension of the inter-vein compartment. This reduction in thickness makes it possible to increase the performance of the engine 2.

[0062] The formation of the air ducts 236 and the diffusion chamber 235 as previously described also makes it possible to limit the number of parts of the engine 2 and therefore to facilitate its assembly. It also makes it possible to avoid any risk of leakage in the casing 23 (or "burst duct" according to English terminology) linked to a rupture of an added pipe or to an assembly defect.

[0063] The de-icing device as described makes it possible not to modify the structure of the engine 2 and the shape of the primary flow vein A and the flow vein secondary B.

[0064] The positioning of the air ducts 236 relative to the stiffeners 234 makes it possible to facilitate and simplify the manufacture of the casing 23.

Claims

Claims

1. Assembly (E) for a turbomachine comprising: - an annular separation nozzle (21), configured to be positioned downstream of a turbomachine fan (20), in a direction of circulation of a gas flow from the fan, and to separate said flow into a primary flow and a secondary flow, the nozzle (21) comprising an annular interior cavity (213), - an annular casing (23) fixed to the nozzle (21) downstream of the nozzle (21), the casing (23) comprising an annular internal wall (231) externally delimiting a primary flow vein (A) in which the primary flow circulates, the casing (23) comprising: - at least one defrosting air duct (236) extending from downstream to upstream, an outlet of the air duct (236) opening into the interior cavity (213) of the nozzle (21), - a diffusion chamber (235), constituted by an annular duct positioned downstream of the air duct (236),the diffusion chamber (235) comprising at least one inlet (2351) intended to be supplied with hot air from a turbomachine (2) and an outlet (2352) connected to the air duct (236); the air duct (236) and the diffusion chamber (235) both being formed with the internal wall (231) of the casing (23).,

2. Assembly (E) according to claim 1, in which the casing (23) comprises an annular radial flange (233) positioned downstream of the internal wall (231) and radially external to the internal wall (231), the diffusion chamber (235) forming a junction between the internal wall (231) and the radial flange (233).

3. Assembly (E) according to claim 2, in which the internal wall (231) comprises at least one stiffener (234) in the form of a plate extending from the internal wall (231) towards the radial flange (233) along a longitudinal plane comprising a longitudinal axis (XX) of the turbomachine.

4. An assembly (E) according to claim 3, wherein the stiffener (234) is positioned on the inner wall (231) at the air duct (236), so as to form a wall of the air duct (236).

5. Assembly (E) according to any one of claims 1 to 4, in which the casing (23), the air duct (236) and the diffusion chamber (235) are produced by additive manufacturing or by casting.

6. Assembly (E) according to any one of claims 1 to 5,

7.

8.

9.

10. comprising a plurality of air ducts (236) distributed regularly around the internal wall (231), the diffusion chamber (235) comprising as many outlets (2352) as there are air ducts (236). Assembly (E) according to any one of claims 1 to 6, in which the at least one inlet (2351) of the diffusion chamber (235) is directly supplied with air taken downstream in the primary flow stream (A). Assembly (E) according to any one of claims 1 to 7, in which the diffusion chamber (235) comprises several inlets (2351). Aircraft turbomachine (1) (100) comprising: a fan (20) and an assembly (E) according to any one of claims 1 to 8, the assembly (E) being positioned downstream of the fan (20). An aircraft (100) comprising an airframe and a turbomachine (1) according to claim 9, wherein the turbomachine (1) is attached to the airframe.

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