De-icing device for aircraft turbojet engine nacelle air intake pipe

Abstract

The present invention discloses a deicing device for an air intake duct (2) of a nacelle of an aircraft turbojet engine. The jet engine nacelle extends along an axis X, and air flows in the nacelle from upstream to downstream. The air intake duct (2) comprises an inner cavity extending circumferentially around the axis X, the inner cavity comprising an inner wall (21) facing the axis X and an outer wall (22) opposite to the inner wall (21), the inner wall (21) and the outer wall (22) being connected by a leading edge. The deicing device comprises at least one ejector (3) for ejecting a hot air flow (FAC) into the inner cavity (20), the ejector (3) comprising a nozzle (30) extending along a nozzle axis (X2), and the nozzle (30) being configured to eject an asymmetric hot air flow (FAC) along the nozzle axis (X2).

Description

Technical Field

[0001] The present invention relates to the field of aircraft turbojet engines, and more particularly to a deicing device for an air inlet pipe of a nacelle of an aircraft turbojet engine. Background Art

[0002] As is known, aircraft include one or more turbojet engines allowing the aircraft to be propelled by accelerating the air flow flowing from upstream to downstream in the turbojet engines.

[0003] refer to Figure 1 , which shows a turbojet engine 100 extending along an axis X and including a fan 101 mounted in a nacelle including a casing 102 and rotating about the axis X to accelerate an airflow F flowing from upstream to downstream. Hereinafter, the terms "upstream" and "downstream" are defined relative to the flow direction of the airflow F. At its upstream end, the turbojet engine 100 includes an air intake duct 200 comprising an inner cavity 204 extending circumferentially about the axis X. The inner cavity 204 comprises an inner wall 201 facing the axis X and an outer wall 202 opposite the inner wall 201. The inner and outer walls 201, 202 are connected by a leading edge 203, also known as the "inlet duct edge." Thus, the air intake duct 200 allows the incoming airflow F to be separated into an inner airflow FINT guided by the inner wall 201 and an outer airflow FEXT guided by the outer wall 202. Hereinafter, the terms "inner" and "outer" are defined relative to the radial direction of the axis X of the turbojet engine 100.

[0004] As is known, during flight of an aircraft, due to temperature and pressure conditions, ice may accumulate on the leading edge 203 and inner wall 201 of the air intake duct 200 and form ice blocks that may be ingested by the turbojet engine 100. In order to increase the life of the turbojet engine 100 and reduce malfunctions, such ingestion should be avoided.

[0005] To eliminate ice buildup, refer to Figure 1 It is known that by injecting a circulating hot air flow FAC into the inner cavity 204, the inner wall 201 is heated by heat exchange, so that the ice cubes melt during the accumulation process, thereby avoiding ice accumulation.

[0006] The hot air FAC is injected into the inner cavity 204 through an ejector 300, which is generally a tube with a cylindrical cross section and is oriented in the same direction as the inner cavity 204. Figure 2 The hot air flow FAC circulates circumferentially in the inner cavity 204 to heat the inner wall 201 .

[0007] In actual application, since the hot air flow FAC will also heat the outer wall 202, which means heat loss, the energy efficiency of this heating method is low. Figure 3 When the hot air FAC is injected, it directly contacts the outer wall 202. When the hot air FAC is injected into the inner cavity 204, the outer wall absorbs most of the heat from the hot air FAC. Therefore, to achieve optimal heating of the inner wall 201, it is necessary to inject superheated hot air FAC into the inner cavity 204. To address this shortcoming, a straightforward solution is to change the orientation of the injector 300 within the inner cavity 204 so that it is aligned with the inner wall 201. However, this straightforward solution fails to achieve uniform circumferential heating of the outer wall 202. In the optimal case, this solution results in a shift in the contact area between the hot air FAC and the outer wall 202, thus resolving the problem.

[0008] One of the objectives of the present invention is to be able to heat the inner wall 201 of the turbojet engine nacelle air inlet duct 200 with higher energy efficiency.

[0009] Incidentally, patent application CA2689195 provides a twisted hot gas jet. Patent application FR2813581 discloses an ejector for ejecting twisted hot gas and axial hot gas in a concentric manner to form a hot gas flow symmetrical along the axis of the ejector nozzle, thereby heating the wall of the inner cavity indiscriminately. Summary of the Invention

[0010] The present invention relates to a deicing device for an air intake duct of an aircraft turbojet engine nacelle. The aircraft turbojet engine extends along an axis X, and an airflow flows in the nacelle from upstream to downstream. The air intake duct includes an inner cavity extending circumferentially around the axis X, the inner cavity having an inner wall facing the axis X and an outer wall opposite the inner wall, the inner wall and the outer wall being connected by a leading edge. The deicing device includes at least one ejector for ejecting a hot airflow into the inner cavity, the ejector including a nozzle extending along a nozzle axis.

[0011] A significant feature of the present invention is that the nozzle is configured to eject an asymmetric stream of hot air along the nozzle axis, thereby generating turbulence near the outer wall while heating the inner wall. In other words, the rotation of the hot air stream around the nozzle axis is asymmetric. This advantageously enables differential heating of the inner and outer walls of the inner cavity, maximizing heat exchange with the inner wall while minimizing heat exchange with the outer wall. Consequently, heating of the inner wall is optimized.

[0012] Preferably, the nozzle includes at least one first channel for guiding a first elementary flow and at least one second channel for guiding a second elementary flow to form the hot air flow. These channels are overlapping, meaning they are not concentric with each other. The present invention enables the nozzle to form a first elementary flow specifically for the outer wall and a second elementary flow specifically for the inner wall, thereby achieving differentiated heating. Advantageously, in the region where the first elementary flow FE1 and the second elementary flow FE2 mix, the pressure differential deflects the hot air flow FAC through the Coanda effect.

[0013] Preferably, the first channel includes at least one airflow deflector. Preferably, the airflow deflector is configured to twist the first elementary flow. According to a preferred aspect, the airflow deflector is helical. Thus, the twisted first elementary flow can generate turbulence, thereby reducing heat exchange with the outer wall of the inner cavity and, by twisting the first elementary flow, promoting circumferential circulation within the annular inner cavity.

[0014] Preferably, the second channel is free of flow deflectors, so as to provide a second basic flow that flows substantially in the axial direction.

[0015] According to a preferred aspect, the nozzle includes a partition to separate the first channel from the second channel of the nozzle. In other words, the nozzle contains two channels. Preferably, the partition is a substantially straight wall, preferably parallel to the nozzle axis. Thus, these elementary flows can be independently directed in the nozzle, thereby achieving two different guidance modes. According to a preferred aspect, the nozzle includes only two channels.

[0016] Preferably, the first channel is tapered from upstream to downstream to accelerate the first basic flow. Preferably, the second channel is tapered from upstream to downstream to accelerate the second basic flow.

[0017] The present invention also relates to an air intake duct for an aircraft turbojet engine nacelle, the aircraft turbojet engine extending along an axis X, with air flowing from upstream to downstream in the nacelle. The air intake duct includes an inner cavity extending circumferentially around the axis X, the inner cavity including an inner wall facing the axis X and an outer wall opposite the inner wall, these walls being connected by a leading edge. The air intake duct includes a de-icing device as described above, wherein the first elementary flow is configured to be injected toward the outer wall of the inner cavity. Preferably, the second elementary flow is configured to be injected toward the inner wall of the inner cavity.

[0018] Preferably, the first elementary flow is located radially outside the second elementary flow.

[0019] The present invention also relates to a method for deicing an air intake duct of an aircraft turbojet engine nacelle using a deicing device as described above, wherein the aircraft turbojet engine extends along an axis X, an air flow flows in the nacelle from upstream to downstream, the air intake duct comprising an inner cavity extending circumferentially around the axis X, the inner cavity comprising an inner wall facing the axis X and an outer wall opposite the inner wall, the walls being connected by a leading edge, the method comprising the steps of: injecting an asymmetric hot air flow along the nozzle axis so as to heat the inner wall while generating turbulence near the outer wall. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the following briefly introduces the drawings required for use in the description of the embodiments. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0021] Figure 1 is a schematic diagram of an air intake pipe of a nacelle in the prior art;

[0022] Figure 2 is a schematic cross-sectional view of the theoretical flow of hot air through an intake pipe in the prior art;

[0023] Figure 3 is a schematic cross-sectional view of the actual flow of hot air through the intake pipe in the prior art;

[0024] Figure 4 is a schematic diagram of an air intake pipe of a nacelle of the present invention;

[0025] Figure 5 1 is a schematic diagram of the flow of hot air through an intake pipe according to an embodiment of the present invention;

[0026] Figure 6 is a schematic diagram of a nozzle of an ejector of a deicing device of the present invention;

[0027] Figure 7 is a cross-sectional view of a nozzle of an ejector of a deicing device of the present invention;

[0028] Figure 8 is a perspective view of a partition wall and an airflow deflector according to the present invention; and

[0029] Figure 9 It is a side view of the partition wall and air flow deflector of the present invention.

[0030] It should be understood that the detailed description of the drawings is used to implement the present invention, and if necessary, the drawings can be used to better define the present invention. DETAILED DESCRIPTION

[0031] refer to Figure 4, which shows a turbojet engine 1 extending along an axis X and including a fan 10 mounted in a nacelle having a casing 12. The fan 10 rotates about the axis X to accelerate an airflow F flowing from upstream to downstream. Hereinafter, the terms "upstream" and "downstream" are defined relative to the flow direction of the airflow F. The turbojet engine 1 includes an air intake duct 2 at its upstream end. The air intake duct comprises an inner cavity 20 extending circumferentially about the axis X. The inner cavity comprises an inner wall 21 facing the axis X and an outer wall 22 opposite the inner wall 21. The inner wall 21 and the outer wall 2 are connected by a leading edge 23, also known as the "inlet duct edge." Thus, the air intake duct 2 allows the incoming airflow F to be separated into an inner airflow FINT guided by the inner wall 21 and an outer airflow FEXT guided by the outer wall 22. Hereinafter, the terms "inner" and "outer" are defined relative to the radial direction of the axis X of the turbojet engine 1.

[0032] Turbine jet engine 1 includes a de-icing device for eliminating ice accumulation. As is known per se, this de-icing device comprises an ejector 3 for injecting a hot air flow FAC into inner cavity 20. The flow of this hot air FAC heats inner wall 21 by heat exchange, causing ice to melt as it accumulates, thereby preventing ice accumulation.

[0033] refer to Figure 5 The ejector 3 includes a nozzle 30 extending along a nozzle axis X2 to inject a hot gas flow circumferentially into the inner cavity 20 of the intake duct 2. Preferably, the nozzle axis X2 is tangential to the axis X of the turbojet engine. In the present invention, the nozzle 30 is configured to inject an asymmetric hot gas flow FAC along the nozzle axis X2. In other words, the hot gas flow FAC is asymmetrical in rotation about the nozzle axis X2.

[0034] In this embodiment, reference Figure 6 and 7 The nozzle 30 includes a first channel 31 for guiding a first elementary flow FE1 and a second channel 32 for guiding a second elementary flow FE2, thereby forming a thermal air flow FAC. In other words, the thermal air flow FAC comprises the first elementary flow FE1 and the second elementary flow FE2. Preferably, the nozzle 30 includes only two channels, namely the first channel 31 and the second channel 32, to achieve optimal de-icing, but it goes without saying that the number of channels may be greater. The first elementary flow FE1 and the second elementary flow FE2 are different. In this embodiment, the first elementary flow FE1 is twisted about the nozzle axis X2, while the second elementary flow FE2 flows parallel to the nozzle axis X2, thereby forming an asymmetric overall flow along the nozzle axis X2.

[0035] In this embodiment, along the nozzle axis X2, the cross-sectional area of ​​the first channel 31 is greater than the cross-sectional area of ​​the second channel 32. As will be explained later, this makes it possible to twist the first elementary flow FE1 in an optimal manner.

[0036] The first channel 31 is tapered from upstream to downstream to accelerate the first elementary flow FE1. In other words, the cross-sectional area of ​​the first channel 31 narrows from upstream to downstream, and the cross-sectional area at the inlet is larger than the cross-sectional area at the outlet.

[0037] According to one aspect of the present invention, the second channel 32 is tapered from upstream to downstream to accelerate the second elementary flow FE2. In other words, the cross-sectional area of ​​the second channel 32 narrows from upstream to downstream, with the cross-sectional area at the inlet being larger than the cross-sectional area at the outlet.

[0038] The narrowing ratios of the first channel 31 and the second channel 32 , ie, the ratios of their inlet cross-sectional areas to their outlet cross-sectional areas, are equal or different.

[0039] like Figure 6 and Figure 7 As shown, the nozzle 30 includes a partition 4 and an airflow deflector 5. Preferably, the partition 4 is a substantially straight partition wall so as to divide the nozzle 30 into a first channel 31 and a second channel 32. Of course, the partition 4 can be of different shapes, specifically concave or convex.

[0040] Preferably, the partition 4 extends parallel to the nozzle axis X2 to achieve differentiated guidance between the first channel 31 and the second channel 32. In this embodiment, the first channel 31 includes a single airflow deflector 5. However, it goes without saying that it may include multiple airflow deflectors connected in series or in parallel. Preferably, the length of the partition 4 matches the length of the airflow deflector 5.

[0041] In this embodiment, the air flow deflector 5 is configured to twist the first elementary flow FE1. Figure 8 and 9 As shown, the airflow deflector 5 is spirally shaped, creating three airflow twists. Preferably, the number of twists ranges from two to five to optimally twist the first elementary flow FE1. The first channel 31 preferably has a cross-section that tapers toward its outlet to twist and accelerate the first elementary flow FE1, thereby increasing turbulence within the inner cavity 20 and preventing direct impact with the outer wall 22. Furthermore, because the first elementary flow FE1 is injected in a twisted manner, the airflow deflector 5 naturally promotes circumferential flow in the intake duct 2. Unlike the prior art, which uses only the outer wall 22 of the inner cavity 20 for circumferential guidance, resulting in heat loss, the present invention achieves circumferential guidance through twisting of the hot air flow FAC, thereby reducing heat loss. The second channel 32 lacks an airflow deflector, allowing the first and second elementary flows FE1 and FE2 to be injected at different angles (axial / spiral trajectories) and velocities.

[0042] The first elementary flow FE1 and the second elementary flow FE2 are injected in an overlapping manner. In the mixing area of ​​the first elementary flow FE1 and the second elementary flow FE2, the pressure difference deflects the hot air flow FAC through the Coanda effect. The positions, angles, flow ratios, and relative velocities of the first elementary flow FE1 and the second elementary flow FE2 are determined to achieve optimal deflection, i.e., deflection that conforms to the annular shape of the inner cavity 20. To control the deflection, the ejector 3 preferably includes a static or dynamic device for adjusting heat loss.

[0043] refer to Figure 5 , the first elementary flow FE1 is configured to be ejected toward one side of the outer wall 22 of the inner cavity 20, and the second elementary flow FE2 is configured to be ejected toward one side of the inner wall 21 of the inner cavity 20. In other words, the first channel 31 is located radially outside the second channel 32. As will be explained in an exemplary embodiment of the present invention, this deicing device makes it possible to optimize the heating of the inner wall 21 of the inner cavity 20 while reducing heat losses.

[0044] Alternatively, the second channel 32 is located radially outside and the first channel 31 is located radially inside, depending on the supply pressures of the first channel 31 and the second channel 32 .

[0045] The method includes the steps of injecting an asymmetric overall hot air flow FAC into the inner cavity 20 along the nozzle axis X2 via the injector 3. In this embodiment, a first radially outer channel 31 injects a first elementary flow FE1. This first elementary flow FE1 is twisted, thereby twisting the entire hot air flow FAC and reducing contact with the outer wall 22. Furthermore, the first elementary flow FE1 generates turbulence near the outer wall 22, thereby reducing the circulation velocity near the outer wall 22 and thus controlling the amount of heat captured. In other words, heat exchange between the outer wall 22 and the first elementary flow FE1 of the hot air flow FAC is significantly reduced.

[0046] Conversely, the radially inner second channel 32 ejects an axial second elementary flow FE2. This allows a large amount of heat to be transferred to the inner wall 21, heating it and keeping it dry. Advantageously, the first elementary flow FE1 enables the second elementary flow FE2 to flow circumferentially without removing heat from it, which is advantageous. The asymmetry of the hot air flow FAC advantageously enables differential heating of the inner wall 21 and the outer wall 22. By improving the nozzle 30 of the ejector 3, the present invention improves the de-icing effect of the intake duct 2 in a practical manner. The efficiency achieved in practice exceeds 75%, which is higher than the efficiency of commercially available de-icing devices, which does not exceed 70%.

[0047] After injection, the hot air flow FAC has flowed through the inner chamber 20 for a distance approximately 15 times the total diameter of the nozzle 30 of the ejector 3. The entire air volume in the inner chamber 20 is displaced, heating the intake duct rim by convection. Advantageously, due to the initial asymmetry, the temperature of the air volume is uniform across the transverse direction, thus achieving optimal heat exchange.

Claims

1. A deicing device for an air intake duct (2) of a nacelle of an aircraft turbojet engine (1), the aircraft turbojet engine extending along an axis (X), an air flow (F) flowing in the nacelle from upstream to downstream, the air intake duct (2) comprising an inner cavity (20) extending circumferentially around the axis (X), the inner cavity comprising an inner wall (21) facing the axis (X) and an outer wall (22) opposite the inner wall (21), the inner wall (21) and the outer wall (22) being connected by a leading edge (23), the deicing device comprising at least one ejector (3) for injecting a hot air flow (FAC) into the inner cavity (20), the ejector (3) comprising a nozzle (30) extending along a nozzle axis (X2), characterized in that The nozzle (30) is configured to eject an asymmetric hot air flow (FAC) along a nozzle axis (X2) to generate turbulence near the outer wall (22) while heating the inner wall (21); The nozzle (30) comprises at least one first channel (31) configured to guide a first elementary flow (FE1) and at least one second channel (32) configured to guide a second elementary flow (FE2) to form a hot air flow (FAC); the first elementary flow (FE1) and the second elementary flow (FE2) are ejected in an overlapping manner; The nozzle (30) includes a partition (4) to divide the nozzle (30) into the first channel (31) and the second channel (32); The first channel (31) comprises at least one air flow deflector (5) configured to distort the first elementary flow (FE1).

2. The deicing device according to claim 1, wherein: The airflow deflecting member (5) is spiral-shaped.

3. The deicing device according to claim 1, wherein: The first channel (31) tapers from upstream to downstream.

4. An air intake duct (2) of a nacelle of an aircraft turbojet engine (1), the aircraft turbojet engine extending along an axis (X), an air flow (F) flowing in the nacelle from upstream to downstream, the air intake duct (2) comprising an inner cavity (20) extending circumferentially around the axis (X), the inner cavity comprising an inner wall (21) facing the axis (X) and an outer wall (22) opposite the inner wall (21), the inner wall (21) and the outer wall (22) being connected by a leading edge (23), characterized in that The air intake duct (2) comprises the de-icing device according to any one of claims 1 to 3, wherein the first elementary flow (FE1) is configured to be injected toward the outer wall (22) of the inner cavity (20).

5. The air intake pipe (2) according to claim 4, characterized in that The second elementary flow (FE2) is configured to be injected toward the inner wall (21) of the inner cavity (20).

6. A method for deicing an air intake duct (2) of a nacelle of an aircraft turbojet engine (1) using a deicing device according to any one of claims 1 to 3, wherein the aircraft turbojet engine extends along an axis (X), an air flow (F) flows from upstream to downstream in the nacelle, the air intake duct (2) comprising an inner cavity (20) extending circumferentially around the axis (X), the inner cavity comprising an inner wall (21) facing the axis X and an outer wall (22) opposite the inner wall (21), the inner wall (21) and the outer wall (22) being connected by a leading edge (23), characterized in that The method comprises the steps of injecting an asymmetric hot air flow (FAC) along a nozzle axis (X2) to generate turbulence near the outer wall (22) while heating the inner wall (21).

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