Vertical aircraft turbomachine diffuser and its method of use

Abstract

An aircraft turbomachine diffuser (1) (60) extending along a longitudinal axis (Z) oriented vertically in a direction opposite to gravity and comprising an upstream flange and a downstream flange together defining a circulation channel of an airflow (F), the channel being configured to be fluidically fed by a centrifugal impeller (21) and to open into a peripheral chamber (30) surrounding a combustion chamber (31) of the aircraft turbomachine (60), the downstream flange being configured to define a bottom of the peripheral chamber (30), the diffuser (1) comprising at least one drain conduit (10) passing through the downstream flange, the channel and the upstream flange, the drain conduit (10) comprising a downstream end configured to open into the peripheral chamber (30) and an upstream end opening out of the channel, so as to drain a liquid (L) by gravity out of the peripheral chamber (30). Figure from the summary: Figure 1

Description

Title of the invention: Vertical aircraft turbomachine diffuser and its method of use technical field

[0001] The present invention relates to the field of vertical aircraft turbomachinery and aims at a diffuser of such a turbomachine.

[0002] In a known manner, an aircraft turbomachine comprises, from upstream to downstream, an air inlet, one or more compressors, a combustion chamber, and one or more turbines. The air inlet is configured to draw an airflow from outside the turbomachine. The compressor compresses the airflow and supplies the combustion chamber. The combustion chamber is also supplied with fuel by one or more injectors to ensure its chemical reaction with the airflow supplied by the compressor.

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

[0004] It is known to equip such a vertical turbomachine with a centrifugal compressor to reduce its longitudinal dimensions. The centrifugal compressor comprises an impeller and a diffuser. The impeller includes rotating, movable blades configured to receive the airflow longitudinally, accelerate it, and release it radially into the diffuser. The diffuser includes an upstream flange and a downstream flange extending radially around the longitudinal axis and together defining an airflow channel. From upstream to downstream, the channel comprises a radial portion, into which the impeller opens, a bent portion, and a longitudinal portion.

[0005] As is known, the combustion chamber is surrounded by an outer casing defining a peripheral chamber into which the longitudinal portion of the diffuser opens. The peripheral chamber is delimited upstream by the downstream flange of the diffuser and externally by the turbomachine casing. The peripheral chamber allows the airflow to circulate in contact with the wall of the combustion chamber to cool it. Part of the airflow enters the combustion chamber through inlet ports provided in the wall of the combustion chamber.

[0006] In practice, in the event of an interrupted start-up of the aircraft turbomachine, unburned fuel is likely to enter the peripheral chamber through the combustion chamber intake ports. The unburned fuel tends to accumulate in liquid form at the bottom of the peripheral chamber, specifically against the downstream flange of the diffuser, under the effect of gravity. Similarly, during turbomachine washing, the washing fluid is likely to accumulate in the peripheral chamber against the downstream flange of the diffuser under the effect of gravity.

[0007] Such an accumulation of liquid is undesirable because it can, in some cases, locally damage the downstream flange of the diffuser through corrosion or the appearance of stray flames generating a hot spot. Replacing such a part is costly and complex.

[0008] In the case of a horizontal turbomachine, the bottom of the peripheral chamber is formed by the turbomachine casing. It is known to drill openings in the turbomachine casing to drain the fluid to the outside by gravity. However, such a solution is not effective in the case of a vertical turbomachine.

[0009] The invention thus aims to protect the diffuser of a vertical aircraft turbomachine from damage caused by the accumulation of liquid in the peripheral chamber surrounding the combustion chamber. PRESENTATION OF THE INVENTION

[0010] The invention relates to an aircraft turbomachine diffuser extending along a longitudinal axis oriented upstream to downstream, the longitudinal axis being oriented vertically in a direction opposite to gravity in nominal use, the diffuser comprising an upstream flange and a downstream flange extending radially around the longitudinal axis and each comprising an opposing inner face together delimiting a circulation channel of an airflow, the channel being configured to be fluidically fed by a centrifugal impeller and to open into a peripheral chamber surrounding a combustion chamber of the aircraft turbomachine, the downstream flange comprising an outer face configured to define a bottom of the peripheral chamber.

[0011] The invention is remarkable in that the diffuser comprises at least one drainage conduit passing through the downstream flange, the vein and the upstream flange, the drainage conduit comprising a downstream end configured to open into the peripheral chamber and an upstream end opening out of the vein, so as to drain a liquid by gravity out of the peripheral chamber in nominal use.

[0012] The invention advantageously proposes to provide drainage channels through the diffuser to prevent the accumulation of liquid, particularly unburned fuel or washing fluid, in the bottom of the peripheral chamber of a vertical aircraft turbomachine. The drainage channels advantageously allow the liquid to be collected and guided out of the peripheral chamber to the outside of the turbomachine. This helps to protect the downstream flange of the diffuser from any damage, particularly from corrosion or the generation of hot spots.

[0013] According to one aspect of the invention, the drainage conduit is mounted in a sealed manner within the vein. Advantageously, the drainage conduit and the vein are not in fluidic communication.

[0014] According to one aspect of the invention, the drainage duct has an internal diameter of less than 5 mm, and preferably greater than 1 mm. Such a drainage duct advantageously has a small footprint and allows the aerodynamic performance of the diffuser's flow to be maintained. This also minimizes airflow losses through the drainage ducts. Advantageously, it is not necessary to provide a closing device or an actuable cover.

[0015] According to one aspect of the invention, said at least one drainage conduit is in the form of a plurality of drainage conduits distributed radially around the longitudinal axis. This allows the liquid to be evacuated regardless of the aircraft's attitude, in particular if it is inclined with respect to the vertical direction of gravity.

[0016] According to one aspect of the invention, the diffuser comprises a radial portion in which the centrifugal impeller is configured to discharge, the radial portion comprising a plurality of blades extending into the stream and comprising a base fixed to the upstream flange and a head fixed to the downstream flange, said drainage channel being formed in one of the plurality of blades. The drainage channels are advantageously formed by drilling through the blades along their height, which makes it possible not to affect the aerodynamic performance of the stream. Furthermore, the overall size is minimal.

[0017] According to one aspect of the invention, the diffuser comprises: • a radial portion in which the centrifugal impeller is configured to discharge, the radial portion comprising a plurality of blades extending into the stream and comprising a foot fixed on the upstream flange and a head fixed on the downstream flange, • a longitudinal section configured to open into the peripheral chamber, • an angled section connecting the radial and longitudinal sections and through which the drainage conduit extends.

[0018] Such an embodiment avoids the need to drill through the blades and reduce their mechanical strength. Such a drainage conduit, in the form of a tube, preferably has a small diameter to limit its size and minimize airflow disturbances within the channel.

[0019] According to one aspect of the invention, the diffuser comprises at least one mounting flange configured to be fixed to an aircraft turbomachine housing, the drainage duct extending through the mounting flange. This enhances the fit. The mechanics of the drainage channels, particularly when they are not formed within the blades. Furthermore, the liquid is discharged directly from the turbomachine.

[0020] The invention also relates to an assembly of a diffuser as described above and an aircraft turbomachine casing, wherein the casing and an outer face of the upstream flange together define a drainage space into which the upstream end of the drain duct opens, the casing comprising at least one drain orifice in fluidic communication with the drainage space and opening to the outside of the aircraft turbomachine. This allows the fluid to be drained from the turbomachine in a simple and practical manner.

[0021] The invention also relates to an aircraft turbomachine extending along a longitudinal axis oriented upstream to downstream, the longitudinal axis being oriented vertically in a direction opposite to gravity in nominal use, the aircraft turbomachine comprising a diffuser as described above, a centrifugal impeller, a combustion chamber and a peripheral chamber surrounding the combustion chamber and comprising a bottom defined by the outer face of the downstream flange of the diffuser, the downstream end of the drain duct opening into the peripheral chamber.

[0022] The invention also relates to a method of using an aircraft turbomachine as described above, in which the drain duct drains the liquid present in the peripheral chamber by gravity during nominal operation. Advantageously, such a method is implemented passively, without electronic or manual control. PRESENTATION OF THE FIGURES

[0023] The invention will be better understood upon reading the following description, given by way of example, and referring to the following figures, given by way of non-limiting examples, in which identical references are given to similar objects.

[0024] The [Fig.1] is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine according to one embodiment of the invention.

[0025] The [Fig.2] is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine diffuser according to a first embodiment of the invention.

[0026] The [Fig.3] is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine diffuser according to a second embodiment of the invention.

[0027] The [Fig.4] is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine diffuser according to a third embodiment of the invention.

[0028] It should be noted that the figures set out the invention in detail to implement the invention, said figures being of course able to serve to better define the invention where appropriate. DETAILED DESCRIPTION OF THE INVENTION

[0029] With reference to [Fig. 1], an aircraft turbomachine 60 is shown extending along a longitudinal axis Z oriented from upstream to downstream and configured for vertical mounting in an aircraft. Under standard operating conditions of the aircraft turbomachine 60, namely during nominal use, the longitudinal axis Z of the aircraft turbomachine 60 extends vertically from bottom to top in a direction opposite to gravity.

[0030] With reference to [Fig. 1], the aircraft turbomachine 60 comprises, 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 an airflow F from outside the aircraft turbomachine 60. The compressor 20 compresses the airflow F and supplies the combustion chamber 31. The combustion chamber 31 is also supplied with fuel by one or more injectors 32 to ensure its chemical reaction with the airflow F supplied by the compressor 20.

[0031] With reference to Figures 1 and 2, the combustion chamber 31 is surrounded by an outer casing 50 defining a peripheral chamber 30 into which the compressor 20 opens. The peripheral chamber 30 allows the airflow F to circulate in contact with the wall of the combustion chamber 31 to cool it. Part of the airflow F enters the combustion chamber 31 through inlet ports 33 (see [Fig. 2]) provided in the wall of the combustion chamber 31.

[0032] With further reference to Figures 1 and 2, the compressor 20 is of the centrifugal type and extends radially around the longitudinal axis Z. The centrifugal compressor 20 comprises one or more stages, each comprising an impeller 21 and a diffuser 1, which respectively form the moving and stationary parts of the compressor 20. As illustrated in [Fig. 2], the impeller 21 comprises movable vanes 22 driven in rotation about the longitudinal axis Z, which are configured to receive the airflow F longitudinally, accelerate it, and release it radially into the diffuser 1. The diffuser 1, in turn, increases the pressure of the airflow F and straightens it at the outlet of the impeller 21. In the case of a multi-stage compressor 20, the impeller 21 and the diffuser 1 of the downstream stage of the compressor 20 are subsequently considered.

[0033] With further reference to [Fig. 2], the diffuser 1 comprises an upstream flange 2 and a downstream flange 5 extending radially around the longitudinal axis Z, namely, at least along a radial component with respect to the longitudinal axis Z. The upstream flange 2 and the downstream flange 5 each comprise an inner face 3, 6 which together define an airflow channel 8, which is fluidically supplied by the impeller 21 and opens into the peripheral chamber 30. The outer face 7 of the downstream flange 5, opposite the inner face 6, defines the bottom of the peripheral chamber 30. The peripheral chamber 30 is thus delimited upstream by the downstream flange 5 of the diffuser 1 and externally by the housing 50 of the turbomachine 60. The diffuser 1 is preferably fixed to one or more housings by one or more mounting flanges. 19, extending in this example in projection from the outer face 4 of the upstream flange 2.

[0034] In the example of [Fig. 2], the diffuser 1 comprises a radial portion 13 into which the impeller 21 opens, a longitudinal portion 16 opening into the peripheral chamber 30, and an angled portion 15 connecting the radial portion 13 and the longitudinal portion 16. The radial portion 13 comprises blades 14 extending in the channel 8 and fixed at their base to the upstream flange 2 and at their tip to the downstream flange 5. The blades 14 of the radial portion 13 define diverging passage sections to reduce the velocity of the airflow F and increase its pressure. The angled portion 15 allows the airflow F to be directed in the longitudinal direction. The longitudinal portion 16 preferably comprises straightening blades to reduce the rotation of the airflow F.

[0035] According to the invention and with reference to Figures 1 and 2, the diffuser 1 also includes one or more drainage conduits 10 passing through the downstream flange 5, the vein 8 and the upstream flange 2. Also according to the invention, the drainage conduit 10 includes a downstream end 11 configured to open into the peripheral chamber 30 and an upstream end 12 opening out of the vein 8, so as to drain a liquid L by gravity out of the peripheral chamber 30 in nominal use.

[0036] The liquid L may, for example, be in the form of unburned fuel from the combustion chamber 31. This can result from an interrupted start of the aircraft turbomachine 60, during which the fuel injected by the injectors 32 into the combustion chamber 31 is not completely consumed. The unburned fuel enters the peripheral chamber 30 via the inlet ports 33 and, under the effect of gravity, reaches the bottom of the peripheral chamber 30, formed by the downstream flange 5 of the diffuser 1. The liquid L may also be in the form of a washing fluid used to clean the aircraft turbomachine 60.

[0037] Thanks to the invention, in a vertical turbomachine 60, such liquids L reach the bottom of the peripheral chamber 30 and then enter the downstream end 11 of the Drainage conduits 10 operate under the influence of gravity. Also under the influence of gravity, the liquid L flows from downstream to upstream through the diffuser 1 via the drainage conduits 10 and exits at the upstream end 12 from the peripheral chamber 30 and the flow channel 8. The invention thus allows any liquid L present in the peripheral chamber 30 to be evacuated from it, taking advantage of gravity in a simple and practical manner. This prevents the accumulation of liquid L against the wall of the diffuser 1 in a vertical turbomachine. This avoids corrosion or the formation of flames that could create a hot spot at the diffuser 1. The diffuser 1 is thus protected from the risk of damage.

[0038] As illustrated in Figures 2 to 4, each drainage conduit 10 extends along at least one vertical component from downstream to upstream to allow drainage of the liquid L by gravity. In the example in [Fig. 2], the drainage conduit 10 extends along the vertical direction of gravity. In the example in [Fig. 4], the drainage conduit 10 is inclined with respect to the vertical direction of gravity but has a downstream end 11 located further downstream than the upstream end 12.

[0039] As illustrated in Figures 2 to 4, the drainage conduit 10 is mounted in a sealed manner in the vein 8. The liquid L circulating in the drainage conduit 10 is thus not in fluidic communication with the air flow F in the vein 8. This is ensured by the wall of the blades 14 in the example of Figures 2 and 3, and by the wall of the tube in the example of [Fig.4], as will be described later.

[0040] According to a first embodiment of the invention illustrated in [Fig. 2], the drainage duct 10 extends into the radial portion 13 of the diffuser 1 and is formed in a blade 14 of the radial portion 13. The drainage duct 10 thus takes the form of an opening drilled in one of the blades 14 along its height, namely from its tip to its base. Such a drainage duct 10, thanks to its internal positioning in a blade 14, thus passes through the channel 8 of the diffuser 1 without disturbing the airflow F.

[0041] With further reference to [Fig. 2], the downstream end 11 of the drainage conduit 10 is formed within the thickness of the downstream flange 5, so as to open into the peripheral chamber 30. The upstream end 12 is formed within the thickness of the upstream flange 2 and opens into a discharge space 51, delimited by the external face 4 of the upstream flange 2, by the mounting flange 19, and by the housing 50 of the turbomachine 60. The housing 50 preferably includes one or more orifices 52 providing fluidic communication between the discharge space 51 and the exterior of the turbomachine 60. The fluid L is thus drained from the peripheral chamber 30 to the discharge space 51 via the drainage conduits 10, and then to the exterior of the turbomachine 60 via the orifices 52 of the housing 50.

[0042] Figure 3 represents a second embodiment of the invention differing from that of Figure 2 in that the drainage conduit 10 extends into the mounting flange 19, thereby improving its mechanical strength. The drainage conduit 10 extends internally within the mounting flange 19 from a first end, fixed to the outer face 4 of the upstream flange 2, to a second end, fixed to a housing of the turbomachine 60, in particular the housing 50 in this example. The upstream end 12 of the drainage conduit 10 is formed in this example at the point where the drainage flange 19 is fixed to the housing 50 and opens to the outside of the turbomachine 60.

[0043] In the example of [Fig.3], the drainage conduit 10 is thus in the form of an opening drilled from downstream to upstream in the downstream flange 5, the blade 14, the upstream flange 2 and the fixing flange 19. The drainage conduit 10 thus puts the peripheral chamber 30 in fluidic communication directly with the outside of the turbomachine 60. The drainage conduit 10 preferably extends vertically as in this example.

[0044] According to a third embodiment illustrated in [Fig. 4], the drainage conduit 10 extends into the angled part 15 of the diffuser 1. In this example, the drainage conduit 10 is in the form of a tube extending into the channel 8 and fixed at the upstream flange 2 and the downstream flange 5. This avoids the need to drill through the blades 14 of the radial part 13. Such drainage conduits 10 can be integrated into an existing diffuser 1.

[0045] With further reference to [Fig. 4], analogously to the first embodiment, the downstream end 11 of the drainage conduit 10 is formed within the thickness of the downstream flange 5, so as to open into the peripheral chamber 30. The upstream end 12 is formed within the thickness of the upstream flange 2 and opens into a discharge space 51, delimited by the external face 4 of the upstream flange 2, by the mounting flange 19, and by the housing 50 of the turbomachine 60. The housing 50 preferably includes one or more orifices 52 providing fluidic communication between the discharge space 51 and the exterior of the turbomachine 60. The fluid L is thus drained from the peripheral chamber 30 to the discharge space 51 via the drainage conduits 10, and then to the exterior of the turbomachine 60 via the orifices. 52 of the crankcase 50.

[0046] In all the embodiments illustrated in Figures 2 to 4, the diffuser 1 preferably comprises several drainage channels 10 according to one or more of the three embodiments, which are distributed circumferentially around the longitudinal axis Z. This facilitates drainage of the peripheral chamber 30 in all aircraft attitudes, particularly when the aircraft is inclined relative to the vertical direction of gravity. Preferably, with reference to [Fig. 2], all or part of the blades 14 are provided with a drainage conduit 10. Preferably, a single blade 14 is mounted in a given blade 14 to guarantee its mechanical strength.

[0047] In all the embodiments illustrated in Figures 2 to 4, the drainage conduit 10 preferably has a diameter less than 5 mm and preferably greater than 1 mm. In the example of Figures 2 and 3, a small diameter is imposed by the geometry of the blades 14, which preferably have a thickness greater than 5 mm and preferably less than 8 mm. In the example of [Fig. 4], a small tube diameter helps to limit the disturbance of the airflow F in the channel 8.

[0048] According to a preferred embodiment, the drainage conduits 10 are open at the downstream end 11 and the upstream end 12. In other words, no closing or actuable cover element is mounted at the upstream and downstream ends 11, 12. The small diameter of the drainage conduits 10 advantageously limits fluid exchange between the peripheral chamber 30 and the exterior of the turbomachine 60 to be drained. The airflow F that could escape from the peripheral chamber 30 via the drainage conduits 10 is advantageously minimized, even in the absence of liquid L to be drained.

[0049] With reference to Figures 2 to 4, the invention also relates to a method of using the aircraft turbomachine 60 consisting of draining a liquid L present in the peripheral chamber 30 via the drainage channels 10 under the effect of gravity. Such a method is advantageously passive, that is to say, it does not require any electronic or manual control. The drainage is advantageously implemented in a simple and practical manner when the turbomachine 60 is extended vertically during nominal operation.

Claims

Demands

1. Diffuser (1) of an aircraft turbomachine (60) extending along a longitudinal axis (Z) oriented from upstream to downstream, the longitudinal axis (Z) being oriented vertically in a direction opposite to gravity in nominal use, the diffuser (1) comprising an upstream flange (2) and a downstream flange (5) extending radially about the longitudinal axis (Z) and each comprising an opposing inner face (3, 6) together defining a flow path (8) for an airflow (F), the flow path (8) being configured to be fluidically fed by a centrifugal impeller (21) and to open into a peripheral chamber (30) surrounding a combustion chamber (31) of the aircraft turbomachine (60), the downstream flange (5) comprising an outer face (7) configured to define a bottom of the peripheral chamber (30), the diffuser (1) being characterized in that it comprises at least one drainage conduit (10) passing through the downstream flange (5),the vein (8) and the upstream flange (2), the drainage conduit (10) comprising a downstream end (11) configured to open into the peripheral chamber (30) and an upstream end (12) opening out of the vein (8), so as to drain a liquid (L) by gravity out of the peripheral chamber (30) in nominal use.

2. Diffuser (1) according to claim 1, in which the drainage conduit (10) is mounted in a sealed manner in the vein (8).

3. Diffuser (1) according to any one of claims 1 and 2, wherein the drainage conduit (10) comprises an internal diameter less than 5mm, and preferably greater than 1mm.

4. Diffuser (1) according to any one of claims 1 to 3, wherein said at least one drainage conduit (10) is in the form of a plurality of drainage conduits (10) distributed circumferentially around the longitudinal axis (Z).

5. Diffuser (1) according to any one of claims 1 to 4, comprising a radial portion (13) in which the centrifugal impeller (21) is configured to discharge, the radial portion (13) comprising a plurality of blades (14) extending in the channel (8) and comprising a foot (15) fixed on the upstream flange (2) and a head (16) fixed on the downstream flange (5), said drainage channel (10) being formed in one of the plurality of blades (14).

6. Diffuser (1) according to any one of claims 1 to 4, comprising: • a radial portion (13) in which the centrifugal impeller (21) is configured to open, the radial portion (13) comprising a plurality of blades (14) extending into the channel (8) and comprising a foot fixed on the upstream flange (2) and a head fixed on the downstream flange (5), • a longitudinal portion (16) configured to open into the peripheral chamber (30), • an angled portion (15) connecting the radial portion (13) and the longitudinal portion (16) and in which the drainage conduit (10) extends.

7. Diffuser (1) according to any one of claims 1 to 6, comprising at least one mounting flange (19) configured to be fixed to an aircraft turbomachine (60) casing (50), the drain duct (10) extending through the mounting flange (19).

8. Assembly of a diffuser (1) according to any one of claims 1 to 6 and of an aircraft turbomachine housing (50) (60), in which the housing (50) and an outer face (4) of the upstream flange (2) together delimit a discharge space (51) into which the upstream end (12) of the drain duct (10) opens, the housing (50) comprising at least one drain orifice (52) in fluidic communication with the discharge space (51) and opening to the outside of the aircraft turbomachine (60).

9. Aircraft turbomachine (60) extending along a longitudinal axis (Z) oriented upstream to downstream, the longitudinal axis (Z) being oriented vertically in a direction opposite to gravity in nominal use, the aircraft turbomachine (60) comprising a diffuser (1) according to any one of claims 1 to 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 face (7) of the downstream flange (5) of the diffuser (1), the downstream end (11) of the drain duct (10) opening into the peripheral chamber (30).

10. Method of using an aircraft turbomachine (60) according to claim 9, wherein the drain conduit (10) drains the liquid (L) present in the peripheral chamber (30) by gravity in nominal use.

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