Vertical aircraft turbine engine diffuser and method of using same

Abstract

The invention relates to a diffuser (1) for an aircraft turbine engine (60), said turbine engine (60) extending along a longitudinal axis (Z) oriented vertically in a direction opposite to gravity, and comprising an upstream flange and a downstream flange, which together define a flow duct for an air flow (F) wherein the flow duct is configured to be supplied with fluid by a centrifugal impeller (21), the invention relates to a diffuser (1) for an aircraft turbine engine (60), comprising a downstream flange (30) and opening into a peripheral chamber (30) surrounding a combustion chamber (31) of the aircraft turbine engine (60), where the downstream flange is configured to define a bottom of said peripheral chamber (30), and the diffuser (1) comprises at least one discharge channel (10) running through the downstream flange, the flow duct and the upstream flange, wherein the discharge channel (10) comprises a downstream end configured to open to the peripheral chamber (30), and an upstream end open to the outside of the flow conduit, thereby discharging the liquid (L) from the peripheral chamber (30) by gravity.

Description

Technical Field

[0001] This invention relates to the field of turbine engines for vertical aircraft, and to diffusers for such turbine engines. Background Technology

[0002] Knownly, an aircraft turbine engine, from upstream to downstream, comprises: an air inlet, one or more compressors, a combustion chamber, and one or more turbines. The air inlet is configured to draw in an airflow from outside the turbine engine. The compressor compresses the airflow and supplies it to the combustion chamber. The combustion chamber also supplies fuel through one or more injectors to ensure that the fuel chemically reacts with the airflow supplied by the compressor.

[0003] In the case of a vertical turbine engine, the longitudinal axis oriented from upstream to downstream of the turbine engine extends vertically upward under standard operating conditions.

[0004] It is known that such vertical turbine engines are equipped with centrifugal compressors to reduce the overall longitudinal dimension. The centrifugal compressor includes an impeller and a diffuser. The impeller includes driven rotating blades configured to receive airflow longitudinally, accelerate the airflow, and discharge the airflow radially into the diffuser. The diffuser includes an upstream flange and a downstream flange extending radially around a longitudinal axis, both of which together define a flow duct for airflow. The flow duct, from upstream to downstream, includes a radial portion through which the impeller passes, a curved portion, and a longitudinal portion.

[0005] Knownly, the combustion chamber is surrounded by an outer shell that defines a peripheral chamber to which a longitudinal portion of the diffuser opens. The upstream of the peripheral chamber is defined by a downstream flange of the diffuser, and its exterior is defined by the turbine engine casing. The peripheral chamber allows airflow to pass in contact with the walls of the combustion chamber, thereby cooling the walls. A portion of the airflow enters the combustion chamber through intake vents in the walls of the combustion chamber.

[0006] In practice, when the turbine engine of an aircraft is interrupted during startup, unburned fuel may enter the outer chamber through the air intake of the combustion chamber. Under the influence of gravity, the unburned fuel tends to accumulate in liquid form at the bottom of the outer chamber, specifically at the downstream flange of the diffuser. Similarly, during turbine engine cleaning, cleaning fluid may also accumulate in the outer chamber at the downstream flange of the diffuser under the influence of gravity.

[0007] Such liquid accumulation is undesirable because, in some cases, it can cause localized damage to the downstream flange of the diffuser through corrosion or the generation of parasitic flames and hot spots. Replacing such components is costly and complex.

[0008] In horizontal turbine engines, the bottom of the outer chamber is formed by the turbine casing. Openings are drilled in the turbine casing to allow liquid to be expelled by gravity. However, this approach is ineffective in vertical turbine engines. Horizontal diffusers are known from patent applications FR2961867A1, US2008 / 056892A1, EP2961991B1, CZ307347B6, US2012 / 167595A1, and EP2108846A2.

[0009] As is known from patent application FR3015567A1, the exhaust system in a helicopter allows for the detection of abnormal suction.

[0010] Therefore, the present invention aims to protect the diffuser of a vertical aircraft turbine engine from damage caused by the accumulation of liquid in the outer chamber surrounding the combustion chamber. Summary of the Invention

[0011] This invention relates to a diffuser for an aircraft turbine engine, the turbine engine extending along a longitudinal axis oriented from upstream to downstream, the longitudinal axis being perpendicular to the direction opposite to gravity under normal operating conditions. The diffuser includes an upstream flange and a downstream flange extending radially around the longitudinal axis, each flange having opposing inner surfaces that together define a flow channel for allowing airflow. The flow channel is configured to be fluidly supplied by a centrifugal impeller and leads to a peripheral chamber surrounding the aircraft turbine engine's combustion chamber. The downstream flange includes an outer surface configured to define the bottom of the peripheral chamber.

[0012] A significant feature of this invention is that the diffuser includes at least one discharge channel that extends through a downstream flange, a flow pipe, and an upstream flange. The discharge channel includes a downstream end configured to open into a peripheral chamber and an upstream end opening into the outside of the flow pipe, thereby allowing liquid to be discharged from the peripheral chamber by gravity under normal operating conditions.

[0013] Advantageously, the invention proposes providing a discharge channel traveling through the diffuser to prevent liquids, especially unburned fuel or cleaning fluids, from accumulating at the bottom of the peripheral chamber of the vertical aircraft turbine engine. Advantageously, the discharge channel draws liquid from the peripheral chamber and directs it towards the outside of the turbine engine. This protects the downstream flange of the diffuser from damage, particularly from corrosion or hotspot formation.

[0014] According to one aspect of the invention, the discharge channel is disposed in a sealed manner within the flow pipe. Advantageously, the discharge channel is not in fluid communication with the flow pipe.

[0015] According to one aspect of the invention, the inner diameter of the discharge channel is less than 5 mm, and preferably greater than 1 mm. Advantageously, the overall size of such a discharge channel is very small, and it maintains the aerodynamic performance of the diffuser's flow path. This also minimizes airflow losses through the discharge channel. Furthermore, there is no need for actuable closures or covers.

[0016] According to one aspect of the invention, the at least one discharge channel is a plurality of discharge channels circumferentially distributed around a longitudinal axis. This allows for the evacuation of liquid regardless of the aircraft's attitude, especially when tilted vertically relative to gravity.

[0017] According to one aspect of the invention, the diffuser includes a radial portion to which a centrifugal impeller is configured; the radial portion has a plurality of blades extending into a flow channel, and the blades include a blade root attached to an upstream flange and a blade head attached to a downstream flange, the discharge channel being formed within one of the plurality of blades. Advantageously, the discharge channel can be formed by drilling holes in the blades according to the blade height, thereby not affecting the aerodynamic performance of the flow channel. Furthermore, the overall size is extremely small.

[0018] According to one aspect of the invention, the diffuser comprises: In the radial portion, a centrifugal impeller is configured to extend into the radial portion, which includes a plurality of blades extending into the flow channel, the blades including a root attached to an upstream flange and a head attached to a downstream flange; The longitudinal portion is configured to lead to the peripheral chamber; the curved portion connects the radial portion and the longitudinal portion, and the discharge passage extends through the curved portion.

[0019] This implementation eliminates the need for drilling holes in the blades, thus preserving their mechanical strength. The tubular discharge channel preferably has a small diameter to limit the overall size and minimize interference with the airflow generated within the flow channel.

[0020] In one aspect of the invention, the diffuser includes at least one attachment bracket configured to attach to the housing of an aircraft turbine engine, through which an exhaust passage extends. This increases the mechanical strength of the exhaust passage, especially in cases where no exhaust passage is formed in the blades. Liquid can also be vented directly from the turbine engine.

[0021] The present invention also relates to an assembly consisting of the aforementioned diffuser and an aircraft turbine engine housing, wherein the outer surface of the housing and the upstream flange together define a venting space, and the upstream end of the discharge passage opens into the venting space; the housing includes at least one drain hole in fluid communication with the venting space and opening to the outside of the aircraft turbine engine. This allows liquid to be easily and conveniently discharged from the turbine engine.

[0022] The present invention also relates to an aircraft turbine engine extending along a longitudinal axis oriented from upstream to downstream; in normal operating conditions, the longitudinal axis is oriented perpendicularly in the direction opposite to gravity, the aircraft turbine engine including a diffuser, a centrifugal impeller, a combustion chamber, and a peripheral chamber surrounding the combustion chamber as described above, and including a bottom defined by the outer surface of the downstream flange of the diffuser, the downstream end of the exhaust passage leading to the peripheral chamber.

[0023] The invention also relates to a method for using the aircraft turbine engine described above, wherein, during normal use, the exhaust channel expels liquid present in the peripheral chamber by gravity. This method can advantageously be implemented passively, without the need for electronic or manual control. Attached Figure Description

[0024] The invention will be better understood by reading the following description, given by way of example with reference to the accompanying drawings, in which like reference numerals denote similar objects.

[0025] Figure 1 is a longitudinal half-sectional schematic diagram of a vertical aircraft turbine engine according to an embodiment of the present invention.

[0026] Figure 2 is a longitudinal half-sectional schematic diagram of the diffuser of a vertical aircraft turbine engine according to the first embodiment of the present invention.

[0027] Figure 3 is a longitudinal half-sectional schematic diagram of the diffuser of a vertical aircraft turbine engine according to a second embodiment of the present invention.

[0028] Figure 4 is a longitudinal half-sectional schematic diagram of the diffuser of a vertical aircraft turbine engine according to a third embodiment of the present invention.

[0029] It should be noted that the accompanying drawings illustrate the invention in detail for the purpose of implementing the invention; of course, these drawings can also be used to better define the invention if necessary. Detailed Implementation

[0030] Referring to Figure 1, an aircraft turbine engine 60 is shown, which extends along a longitudinal axis Z oriented from upstream to downstream and is configured to be vertically mounted in the aircraft. Under standard operating conditions of the aircraft turbine engine 60, i.e., during normal use, the longitudinal axis Z of the aircraft turbine engine 60 extends vertically from bottom to top in a direction opposite to gravity.

[0031] refer to Figure 1 The aircraft turbine engine 60 includes, from upstream to downstream, an air inlet 70, one or more compressors 20, a combustion chamber 31, and one or more turbines 40. The air inlet 70 is configured to draw in an airflow F from outside the aircraft turbine engine 60. The compressor 20 compresses the airflow F and supplies it to the combustion chamber 31. The combustion chamber 31 is also supplied with fuel by one or more fuel injectors 32 to ensure that it chemically reacts with the airflow F supplied by the compressor 20.

[0032] refer to Figure 1 and Figure 2 The combustion chamber 31 is surrounded by a housing 50, which defines an outer chamber 30 through which the compressor 20 passes. The outer chamber 30 allows an airflow F to circulate in contact with the walls of the combustion chamber 31 to cool the combustion chamber. A portion of the airflow F enters the combustion chamber 31 through an air intake 33 (see Figure 2) provided on the walls of the combustion chamber 31.

[0033] Still referencing Figure 1 and Figure 2 The compressor 20 is centrifugal and extends radially along its longitudinal axis Z. The centrifugal compressor 20 includes one or more stages, each stage including an impeller 21 and a diffuser 1, which respectively constitute the moving and stationary parts of the compressor 20. Figure 2 As shown, impeller 21 includes movable blades 22 rotatably driven about a longitudinal axis Z, the movable blades being configured to receive airflow F longitudinally, accelerate it, and release it radially into diffuser 1. Diffuser 1 is used to increase the pressure of airflow F and collimate it at the outlet of impeller 21. In the case of multi-stage compressor 20, the impeller 21 and diffuser 1 of the downstream stage of compressor 20 will be considered below.

[0034] Still referencing Figure 2The diffuser 1 includes an upstream flange 2 and a downstream flange 5, both extending radially along a longitudinal axis Z, i.e., at least along a radial component relative to the longitudinal axis Z. The upstream flange 2 and the downstream flange 5 each have inner surfaces 3 and 6, which together define a flow channel 8 through which airflow F is supplied fluid by the impeller 21 and leads to a peripheral chamber 30. The outer surface 7 of the downstream flange 5, opposite the inner surface 6, defines the bottom of the peripheral chamber 30. Thus, the peripheral chamber 30 is defined upstream by the downstream flange 5 of the diffuser 1 and externally by the housing 50 of the turbine engine 60. The diffuser 1 is preferably attached to one or more housings by one or more attachment brackets 19, which in this example extend from the outer surface 4 of the upstream flange 2.

[0035] exist Figure 2 In the example shown, diffuser 1 includes a radial portion 13, a longitudinal portion 16, and a curved portion 15. Impeller 21 leads to the radial portion 13, the longitudinal portion 16 leads to the peripheral chamber 30, and the curved portion 15 connects the radial portion 13 and the longitudinal portion 16. The radial portion 13 includes blades 14 that extend into the flow channel 8, with their roots attached to the upstream flange 2 and their heads attached to the downstream flange 5. The blades 14 of the radial portion 13 define a branching channel cross-section to reduce the velocity of the airflow F and increase its pressure. The curved portion 15 allows the airflow F to be guided longitudinally. The longitudinal portion 16 preferably includes collimating blades to reduce swirl in the airflow F.

[0036] According to the present invention and reference Figure 1 and Figure 2 The diffuser 1 also includes one or more discharge channels 10 that pass through the downstream flange 5, the flow pipe 8, and the upstream flange 2. Also according to the invention, the discharge channel 10 includes a downstream end 11 and an upstream end 12, the downstream end 11 being configured to open into the peripheral chamber 30, and the upstream end 12 being configured to open into the outside of the flow pipe 8, so that liquid L is discharged from the peripheral chamber 30 by gravity during normal use.

[0037] Liquid L may be, for example, unburned fuel from combustion chamber 31. This may be due to an interruption in the start-up of the aircraft turbine engine 60, at which point the fuel injected by injector 32 into combustion chamber 31 is not fully consumed. The unburned fuel enters the peripheral chamber 30 through intake port 33 and reaches the bottom of the peripheral chamber 30 formed by downstream flange 5 of diffuser 1 under gravity. Liquid L may also be in the form of cleaning fluid used to clean the aircraft turbine engine 60.

[0038] With the aid of this invention, in the vertical turbine engine 60, such liquid L reaches the bottom of the peripheral chamber 30 under gravity and then enters the downstream end 11 of the discharge channel 10. Also under gravity, liquid L flows from downstream to upstream through the diffuser 1 via the discharge channel 10 and is discharged from the upstream end 12 outside the peripheral chamber 30 and the flow pipe 8. Therefore, this invention allows for the simple and practical use of gravity to drain any liquid L present in the peripheral chamber 30. This prevents liquid L from accumulating on the walls of the diffuser 1 of the vertical turbine engine. This prevents corrosion or the generation of hot spots on the diffuser 1 surface. This protects the diffuser 1 from damage.

[0039] like Figures 2 to 4 As shown, each discharge channel 10 extends along at least one vertical component from downstream to upstream to allow liquid L to be discharged under gravity. In the example shown in Figure 2, the discharge channel 10 extends in the direction perpendicular to gravity. In the example shown in Figure 4, the discharge channel 10 is inclined relative to the direction perpendicular to gravity, but its downstream end 11 is located further downstream than its upstream end 12.

[0040] As shown in Figures 2 to 4, the discharge channel 10 is installed in a sealed manner within the flow pipe 8. Therefore, the liquid L flowing in the discharge channel 10 does not have fluid communication with the airflow F in the flow pipe 8. In the examples of Figures 2 and 3, this is ensured by the wall of the blade 14; in the example of Figure 4, this is ensured by the wall of the pipe, as will be described below.

[0041] According to the first embodiment of the invention shown in FIG2, the discharge channel 10 extends within the radial portion 13 of the diffuser 1 and is formed in the blades 14 of the radial portion 13. Therefore, the discharge channel 10 is formed by drilling a hole along its height (i.e., from the blade tip to the blade root) in one of the blades 14. Due to its internal position within the blade 14, this discharge channel 10 can pass through the flow channel 8 of the diffuser 1 without interfering with the airflow F.

[0042] Still referencing Figure 2The downstream end 11 of the discharge passage 10 is formed within the thickness of the downstream flange 5, thus leading to the peripheral chamber 30. The upstream end 12 is formed within the thickness of the upstream flange 2 and leads to the venting space 51 defined by the outer surface 4 of the upstream flange 2, the attachment bracket 19, and the housing 50 of the turbine engine 60. The housing 50 preferably includes one or more orifices 52 for providing fluid communication between the venting space 51 and the outside of the turbine engine 60. Thus, liquid L is discharged from the peripheral chamber 30 through the discharge passage 10 toward the venting space 51, and then through the orifices 52 in the housing 50 toward the outside of the turbine engine 60.

[0043] Figure 3 illustrates a second embodiment of the present invention, which is consistent with... Figure 2 The difference in the illustrated embodiment is that the exhaust passage 10 extends through the attachment bracket 19, thereby increasing its mechanical strength. The exhaust passage 10 extends within the attachment bracket 19 from a first end attached to the outer surface 4 of the upstream flange 2 toward a second end attached to the housing of the turbine engine 60 (specifically, housing 50 in this example). In this example, the upstream end 12 of the exhaust passage 10 is formed at the attachment height between the attachment bracket 19 and the housing 50 and opens to the outside of the turbine engine 60.

[0044] In the example shown in Figure 3, the discharge passage 10 is an opening formed by drilling through the downstream flange 5, blade 14, upstream flange 2, and attachment bracket 19 sequentially from downstream to upstream. Thus, the discharge passage 10 provides direct fluid communication between the peripheral chamber 30 and the outside of the turbine engine 60. As shown in this example, the discharge passage 10 preferably extends vertically.

[0045] according to Figure 4 In the third embodiment shown, the discharge channel 10 extends within the curved portion 15 of the diffuser 1. In this example, the discharge channel 10 is tubular, extending within the flow conduit 8 and attached at the height of the upstream flange 2 and the downstream flange 5. This eliminates the need for drilling holes in the blades 14 of the radial portion 13. Such a discharge channel 10 can be integrated into an existing diffuser 1.

[0046] Referring again to Figure 4, similar to the first embodiment, the downstream end 11 of the discharge channel 10 is formed within the wall thickness of the downstream flange 5 to lead to the peripheral chamber 30. The upstream end 12 is formed within the wall thickness of the upstream flange 2 and leads to the venting space 51 defined by the outer surface 4 of the upstream flange 2, the attachment bracket 19, and the housing 50 of the turbine engine 60. The housing 50 preferably includes one or more orifices 52 to allow fluid communication between the venting space 51 and the outside of the turbine engine 60. Thus, liquid L is discharged from the peripheral chamber 30 toward the venting space 51 via the discharge channel 10, and then toward the outside of the turbine engine 60 via the orifices 52 on the housing 50.

[0047] exist Figures 2 to 4 In all the embodiments shown, and in one or more of the three embodiments described above, the diffuser 1 preferably includes a plurality of discharge channels 10 circumferentially distributed around the longitudinal axis Z. This is advantageous for discharging from the peripheral chamber 30 in all attitudes of the aircraft, especially when the aircraft is tilted relative to the vertical direction of gravity. Preferably, referring to FIG2, all or part of the blades 14 are provided with discharge channels 10. Preferably, to ensure the mechanical strength of the blades 14, one blade 14 is installed within a given blade 14.

[0048] Similarly, in all embodiments shown in Figures 2 to 4, the inner diameter of the discharge channel 10 is preferably less than 5 mm and preferably greater than 1 mm. In the examples shown in Figures 2 and 3, the small diameter is determined by the geometry of the blade 14, the thickness of which is preferably greater than 5 mm and preferably less than 8 mm. In the example shown in Figure 4, the small diameter limits interference with the airflow F within the flow channel 8.

[0049] According to a preferred aspect, the discharge passage 10 is open at the height of the downstream end 11 and the upstream end 12. In other words, no operable closure or covering is provided at the upstream and downstream ends 11, 12. The small diameter of the discharge passage 10 advantageously restricts fluid exchange between the peripheral chamber 30 and the outside of the turbine engine 60 to be discharged. Even in the absence of liquid L to be discharged, the airflow F that might escape from the peripheral chamber 30 via the discharge passage 10 is advantageously minimized.

[0050] Referring to Figures 2 to 4, the present invention also relates to a method of using an aircraft turbine engine 60, the method comprising: discharging liquid L present in a peripheral chamber 30 via a discharge channel 10 under the influence of gravity. Advantageously, the method is passive, i.e., requiring no electronic or manual control. When the turbine engine 60 is vertically arranged in normal operating conditions, the discharge process can be implemented in a simple and practical manner.

Claims

1. A diffuser (1) for an aircraft turbine engine (60) extending along a longitudinal axis (Z) oriented from upstream to downstream, the longitudinal axis (Z) being perpendicular to the direction opposite to gravity in normal operating conditions, the diffuser (1) comprising an upstream flange (2) and a downstream flange (5) extending radially around the longitudinal axis (Z), each of the flanges (2, 5) having opposing inner surfaces (3, 6) that collectively define a flow duct (8) for allowing airflow (F) to pass through, the flow duct (8) being configured to be fluidly supplied by a centrifugal impeller (21) and leading to a peripheral chamber (30) surrounding a combustion chamber (31) of the aircraft turbine engine (60), the downstream flange (5) comprising an outer surface (7) configured to define the bottom of the peripheral chamber (30), characterized in that, The diffuser (1) includes at least one discharge passage (10) that travels through the downstream flange (5), the flow pipe (8), and the upstream flange (2). The discharge passage (10) includes a downstream end (11) configured to lead to the peripheral chamber (30) and an upstream end (12) leading to the outside of the flow pipe (8), thereby discharging liquid (L) from the peripheral chamber (30) by gravity under normal operating conditions.

2. The diffuser (1) as claimed in claim 1, wherein the discharge passage (10) is disposed in the flow pipe (8) in a sealed manner.

3. The diffuser (1) as claimed in claims 1 and 2, wherein the inner diameter of the discharge channel (10) is less than 5 mm, and preferably greater than 1 mm.

4. The diffuser (1) according to any one of claims 1-3, wherein the at least one discharge channel (10) is in the form of a plurality of discharge channels (10) circumferentially distributed around the longitudinal axis (Z).

5. The diffuser (1) of any one of claims 1-4, comprising a radial portion (13), wherein the centrifugal impeller (21) is configured to extend toward the radial portion (13), the radial portion (13) comprising a plurality of blades (14) extending into the flow conduit (8), and the blades (14) comprising a root (15) attached to the upstream flange (2) and a head (16) attached to the downstream flange (5), wherein the discharge passage (10) is formed within one of the plurality of blades (14).

6. The diffuser (1) as claimed in any one of claims 1-4, comprising: The radial portion (13) is configured to lead to the centrifugal impeller (21), the radial portion (13) including a plurality of blades (14) extending into the flow conduit (8), the blades (14) including a root attached to the upstream flange (2) and a head attached to the downstream flange (5); The longitudinal portion (16) is configured to open into the peripheral chamber (30); A curved portion (15) connects the radial portion (13) to the longitudinal portion (16), and the discharge passage (10) extends through the curved portion (15).

7. The diffuser (1) of any one of claims 1-6, comprising at least one attachment bracket (19) configured to attach to the housing (50) of an aircraft turbine engine (60), the exhaust passage (10) extending through the attachment bracket (19).

8. An assembly consisting of a diffuser (1) as claimed in any one of claims 1-6 and a housing (50) of an aircraft turbine engine (60), wherein the housing (50) and the outer surface (4) of the upstream flange (2) define a venting space (51), the upstream end (12) of the venting passage (10) leading to the venting space (51), the housing (50) including at least one drain hole (52) in fluid communication with the venting space (51) and leading to the outside of the aircraft turbine engine (60).

9. An aircraft turbine engine (60) extending along a longitudinal axis (Z) oriented from upstream to downstream, wherein, in normal operating conditions, the longitudinal axis (Z) is oriented perpendicularly to the direction opposite to gravity, the aircraft turbine engine (60) comprising a diffuser (1) as claimed in any one of claims 1-7, a centrifugal impeller (21), a combustion chamber (31) and a peripheral chamber (30) surrounding the combustion chamber (31), and comprising a bottom defined by the outer surface of the downstream flange (5) of the diffuser (1), the downstream end (11) of the exhaust passage (10) opening to the peripheral chamber (30).

10. A method of using an aircraft turbine engine (60) as claimed in claim 9, wherein, during normal use, the discharge passage (10) discharges liquid (L) present in the peripheral chamber (30) by gravity.

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