FR3153637A1 - Vertical aircraft turbomachine diffuser and method of using same - Google Patents

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 delimiting a flow path for circulation of an air flow (F), the flow path being configured to be fluidically supplied 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 drainage duct (10) passing through the downstream flange, the flow path and the upstream flange, the drainage duct (10) comprising a downstream end configured to open into the peripheral chamber (30) and an upstream end opening out of the flow path, so as to drain a liquid (L) by gravity out of the peripheral chamber (30). Abstract figure: Figure 1

Description

Vertical aircraft turbomachine diffuser and method of using same

The present invention relates to the field of vertical aircraft turbomachines and aims at a diffuser for such a turbomachine.

As is known, 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 take an air flow from outside the turbomachine. The compressor compresses the air flow and feeds the combustion chamber. The combustion chamber is also supplied with fuel by one or more injectors to ensure its chemical reaction with the air flow provided by the compressor.

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

It is known to equip such a vertical turbomachine with a centrifugal type compressor to reduce the longitudinal size. The centrifugal compressor comprises an impeller and a diffuser. The impeller comprises rotating moving blades which are configured to receive the air flow longitudinally, accelerate it and release it radially into the diffuser. The diffuser comprises an upstream flange and a downstream flange extending radially around the longitudinal axis and together delimiting a flow path for the air flow. The path comprises, from upstream to downstream, a radial part, into which the impeller opens, a bent part and a longitudinal part.

As is known, the combustion chamber is surrounded by an outer casing defining a peripheral chamber into which the longitudinal part of the diffuser opens. The peripheral chamber is delimited upstream by the downstream flange of the diffuser and externally by the casing of the turbomachine. The peripheral chamber allows the air flow to circulate in contact with the wall of the combustion chamber to cool it. Part of the air flow enters the combustion chamber through intake orifices provided in the wall of the combustion chamber.

In practice, in the event of an interrupted start of the aircraft turbomachine, unburned fuel may enter the peripheral chamber through the combustion chamber intake ports. Unburned fuel tends to accumulate in the liquid state at the bottom of the peripheral chamber, namely against the downstream flange of the diffuser, under the effect of gravity. Similarly, during a turbomachine wash, the washing liquid may accumulate in the peripheral chamber against the downstream flange of the diffuser under the effect of gravity.

Such a build-up of liquid is undesirable because in some cases it can locally damage the downstream flange of the diffuser by corrosion or by the appearance of parasitic flames generating a hot spot. Replacing such a part is expensive and complex.

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 liquid to the outside by gravity. However, such a solution is not effective in the case of a vertical turbomachine.

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

The invention relates to an aircraft turbomachine diffuser extending along a longitudinal axis oriented from 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 inner face facing each other together delimiting a flow path for circulation of an air flow, the flow path being configured to be fluidically supplied 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.

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.

The invention advantageously proposes providing drainage ducts passing through the diffuser to prevent the accumulation of liquid, in particular unburned fuel or washing liquid, in the bottom of the peripheral chamber of a vertical aircraft turbomachine. The drainage ducts advantageously make it possible to collect and guide the liquid out of the peripheral chamber towards the outside of the turbomachine. This makes it possible to protect the downstream flange of the diffuser from any damage, in particular by corrosion or generation of hot spots.

According to one aspect of the invention, the drainage conduit is mounted in a sealed manner in the vein. Advantageously, the drainage conduit and the vein are not in fluid communication.

According to one aspect of the invention, the drainage duct comprises an internal diameter of less than 5 mm, and preferably greater than 1 mm. Such a drainage duct advantageously has a small footprint and makes it possible to maintain the aerodynamic performance of the diffuser vein. This also makes it possible to minimize air flow losses via the drainage ducts. It is advantageously not necessary to provide an actuable closing or covering member.

According to one aspect of the invention, said at least one drainage duct is in the form of a plurality of drainage ducts distributed radially around the longitudinal axis. This makes it possible to evacuate the liquid regardless of the attitude of the aircraft, in particular if it is inclined relative to the vertical direction of gravity.

According to one aspect of the invention, the diffuser comprises a radial portion into which the centrifugal impeller is configured to discharge, the radial portion comprising a plurality of blades extending into the vein and comprising a foot fixed to the upstream flange and a head fixed to the downstream flange, said drainage duct being formed in one of the plurality of blades. The drainage ducts are advantageously formed by piercing the blades along their height, which makes it possible not to affect the aerodynamic performance of the vein. The size is also minimal.

• une partie radiale dans laquelle le rouet centrifuge est configuré pour déboucher, la partie radiale comprenant une pluralité de pales s’étendant dans la veine et comprenant un pied fixé sur le flasque amont et une tête fixée sur le flasque aval,
• une partie longitudinale configurée pour déboucher dans la chambre périphérique,
• une partie coudée reliant la partie radiale et la partie longitudinale et dans laquelle s’étend le conduit de drainage.

According to one aspect of the invention, the diffuser comprises:

• a radial portion into which the centrifugal impeller is configured to discharge, the radial portion comprising a plurality of blades extending into the vein and comprising a foot fixed to the upstream flange and a head fixed to the downstream flange,
• a longitudinal part configured to open into the peripheral chamber,
• a bent part connecting the radial part and the longitudinal part and in which the drainage duct extends.

Such an embodiment makes it possible to avoid drilling the blades and reduce their mechanical strength. Such a drainage duct in the form of a tube preferably has a small diameter to limit the size and minimize the disturbances to the air flow caused in the vein.

According to one aspect of the invention, the diffuser comprises at least one fixing flange configured to be fixed to an aircraft turbomachine casing, the drainage duct extending through the fixing flange. This makes it possible to reinforce the mechanical strength of the drainage ducts, in particular when they are not formed in the blades. The liquid is furthermore discharged directly from the turbomachine.

The invention also relates to an assembly of a diffuser as described above and an aircraft turbomachine casing, in which the casing and an outer face of the upstream flange together delimit a discharge space into which the upstream end of the drainage duct opens, the casing comprising at least one drainage orifice in fluid communication with the discharge space and opening outside the aircraft turbomachine. This makes it possible to drain the liquid simply and practically from the turbomachine.

The invention also relates to an aircraft turbomachine extending along a longitudinal axis oriented from 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 previously, 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 drainage duct opening into the peripheral chamber.

The invention also relates to a method of using an aircraft turbomachine as described above, in which the drainage duct drains the liquid present in the peripheral chamber by gravity in nominal use. Such a method is advantageously implemented passively, without electronic or manual control.

PRESENTATION OF FIGURES

The invention will be better understood upon reading the description which follows, 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.

There FIG. 1 is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine according to one embodiment of the invention.

There FIG. 2 is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine diffuser according to a first embodiment of the invention.

There FIG. 3 is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine diffuser according to a second embodiment of the invention.

There FIG. 4 is a schematic representation in longitudinal half-section of a vertical aircraft turbomachine diffuser according to a third embodiment of the invention.

It should be noted that the figures set out the invention in detail to implement the invention, said figures can of course be used to better define the invention where appropriate.

DETAILED DESCRIPTION OF THE INVENTION

In reference to the FIG. 1 , an aircraft turbomachine 60 is presented extending along a longitudinal axis Z oriented from upstream to downstream and configured to be mounted vertically 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.

In reference to the 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 take an air flow F from outside the aircraft turbomachine 60. The compressor 20 makes it possible to compress the air flow F and to supply 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 air flow F supplied by the compressor 20.

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 air flow F to circulate in contact with the wall of the combustion chamber 31 to cool it. A portion of the air flow F enters the combustion chamber 31 through intake orifices 33 (see FIG. 2 ) provided in the wall of the combustion chamber 31.

Still with 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 respectively forming the mobile member and the fixed member of the compressor 20. As illustrated in the FIG. 2 , the impeller 21 comprises moving blades 22 driven in rotation around the longitudinal axis Z, which are configured to receive the air flow F longitudinally, accelerate it and release it radially into the diffuser 1. The diffuser 1 makes it possible to increase the pressure of the air flow F and to straighten 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.

Still referring to the 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 relative to the longitudinal axis Z. The upstream flange 2 and the downstream flange 5 each comprise an inner face 3, 6 together delimiting a vein 8 for circulation of the air flow F, 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 casing 50 of the turbomachine 60. The diffuser 1 is preferably fixed to one or more casings by one or more fixing flanges 19, extending in this example in projection from the outer face 4 of the upstream flange 2.

In the example of the FIG. 2 , the diffuser 1 comprises a radial part 13 into which the impeller 21 opens, a longitudinal part 16 opening into the peripheral chamber 30 and a bent part 15 connecting the radial part 13 and the longitudinal part 16. The radial part 13 comprises blades 14 extending in the vein 8 and fixed at their root on the upstream flange 2 and at their head on the downstream flange 5. The blades 14 of the radial part 13 define divergent passage sections to reduce the speed of the air flow F and increase its pressure. The bent part 15 makes it possible to orient the air flow F in the longitudinal direction. The longitudinal part 16 preferably comprises straightening blades to reduce the gyration of the air flow F.

According to the invention and with reference to figures 1 and 2, the diffuser 1 also comprises one or more drainage conduits 10 passing through the downstream flange 5, the vein 8 and the upstream flange 2. Still according to the invention, the drainage conduit 10 comprises 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.

The liquid L is for example in the form of unburned fuel coming from the combustion chamber 31. This may 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 entirely consumed. The unburned fuel enters the peripheral chamber 30 via the intake orifices 33 and reaches the bottom of the peripheral chamber 30, formed by the downstream flange 5 of the diffuser 1, under the effect of gravity. The liquid L may also be in the form of a washing liquid used to clean the aircraft turbomachine 60.

Thanks to the invention, in a vertical turbomachine 60, such liquids L reach the bottom of the peripheral chamber 30 then enter the downstream end 11 of the drainage conduits 10 under the effect of gravity. Still under the effect of gravity, the liquid L circulates from downstream to upstream through the diffuser 1 via the drainage conduits 10 and emerges via the upstream end 12 outside the peripheral chamber 30 and the vein 8. The invention thus makes it possible to evacuate any liquid L present in the peripheral chamber 30 from the latter, taking advantage of gravity in a simple and practical manner. This makes it possible to avoid the accumulation of liquid L against the wall of the diffuser 1 in a vertical turbomachine. This prevents corrosion or the appearance of flames generating a hot spot at the diffuser 1. The diffuser 1 is thus protected from a risk of damage.

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 of the FIG. 2 , the drainage conduit 10 extends in the vertical direction of gravity. In the example of the FIG. 4 , the drainage conduit 10 is inclined relative to the vertical direction of gravity but has a downstream end 11 located further downstream than the upstream end 12.

As illustrated in Figures 2 to 4, the drainage duct 10 is mounted in a sealed manner in the vein 8. The liquid L circulating in the drainage duct 10 is thus not in fluid 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 the FIG. 4 , as will be described later.

According to a first embodiment of the invention illustrated in the FIG. 2 , the drainage duct 10 extends in the radial part 13 of the diffuser 1 and is formed in a blade 14 of the radial part 13. The drainage duct 10 is thus in the form of an opening pierced in one of the blades 14 according to its height, namely from its head to its foot. Such a drainage duct 10, thanks to its internal positioning in a blade 14, thus crosses the vein 8 of the diffuser 1 without disturbing the air flow F.

Still referring to the FIG. 2 , the downstream end 11 of the drainage duct 10 is formed in the thickness of the downstream flange 5, so as to open into the peripheral chamber 30. The upstream end 12 is formed in the thickness of the upstream flange 2 and opens into an evacuation space 51, delimited by the external face 4 of the upstream flange 2, by the fixing flange 19 and by the casing 50 of the turbomachine 60. The casing 50 preferably comprises one or more orifices 52 putting the evacuation space 51 into fluid communication with the exterior of the turbomachine 60. The liquid L is thus drained from the peripheral chamber 30 to the evacuation space 51 via the drainage ducts 10, then to the exterior of the turbomachine 60 via the orifices 52 of the casing 50.

There FIG. 3 represents a second embodiment of the invention differing from that of the FIG. 2 in that the drainage duct 10 extends into the fixing flange 19, which improves its mechanical strength. The drainage duct 10 extends internally in the fixing flange 19 from a first end, fixed at the external face 4 of the upstream flange 2, towards a second end, fixed to a casing of the turbomachine 60, in particular the casing 50 in this example. The upstream end 12 of the drainage duct 10 is in this example formed at the level of the fixing of the drainage flange 19 on the casing 50 and opens outside the turbomachine 60.

In the example of the FIG. 3 , the drainage duct 10 thus takes the form of an opening pierced from downstream to upstream in the downstream flange 5, the blade 14, the upstream flange 2 and the fixing flange 19. The drainage duct 10 thus places the peripheral chamber 30 in fluid communication directly with the exterior of the turbomachine 60. The drainage duct 10 preferably extends vertically as in this example.

According to a third embodiment illustrated in the FIG. 4 , the drainage duct 10 extends into the bent part 15 of the diffuser 1. The drainage duct 10 is in this example in the form of a tube extending into the vein 8 and fixed at the level of the upstream flange 2 and the downstream flange 5. This makes it possible not to have to pierce the blades 14 of the radial part 13. Such drainage ducts 10 can be integrated into an existing diffuser 1.

Still referring to the FIG. 4 , similarly to the first embodiment, the downstream end 11 of the drainage duct 10 is formed in the thickness of the downstream flange 5, so as to open into the peripheral chamber 30. The upstream end 12 is formed in the thickness of the upstream flange 2 and opens into an evacuation space 51, delimited by the external face 4 of the upstream flange 2, by the fixing flange 19 and by the casing 50 of the turbomachine 60. The casing 50 preferably comprises one or more orifices 52 placing the evacuation space 51 in fluid communication with the exterior of the turbomachine 60. The liquid L is thus drained from the peripheral chamber 30 to the evacuation space 51 via the drainage ducts 10, then to the exterior of the turbomachine 60 via the orifices 52 of the casing 50.

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

Still in all the embodiments illustrated in Figures 2 to 4, the drainage conduit 10 preferably has a diameter less than 5mm and preferably greater than 1mm. 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 5mm and preferably less than 8mm. In the example of the FIG. 4 , a small tube diameter makes it possible to limit the disturbance of the air flow F in vein 8.

According to a preferred aspect, the drainage ducts 10 are open at the downstream end 11 and the upstream end 12. In other words, no actuable closing or covering element is mounted at the upstream and downstream ends 11, 12. The small diameter of the drainage ducts 10 advantageously makes it possible to limit the fluid exchanges between the peripheral chamber 30 and the exterior of the turbomachine 60 to be drained. The air flow F likely to escape from the peripheral chamber 30 via the drainage ducts 10 is advantageously minimized, even in the absence of liquid L to be drained.

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 conduits 10 under the effect of gravity. Such a method is advantageously passive, that is to say that it does not require any electronic or manual control. The drainage is advantageously implemented in a simple and practical manner when the turbomachine 60 extends vertically in nominal use.

Claims (10)

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 around the longitudinal axis (Z) and each comprising an inner face (3, 6) facing each other together delimiting a flow path (8) for circulating an air flow (F), the flow path (8) being configured to be fluidically supplied 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 duct (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. Diffuser (1) according to claim 1, in which the drainage conduit (10) is mounted in a sealed manner in the vein (8). Diffuser (1) according to one of claims 1 and 2, in which the drainage conduit (10) comprises an internal diameter less than 5 mm, and preferably greater than 1 mm. Diffuser (1) according to one of claims 1 to 3, wherein said at least one drainage duct (10) is in the form of a plurality of drainage ducts (10) distributed radially around the longitudinal axis (Z). Diffuser (1) according to one of claims 1 to 4, comprising a radial part (13) in which the centrifugal impeller (21) is configured to open, the radial part (13) comprising a plurality of blades (14) extending in the vein (8) and comprising a foot (15) fixed on the upstream flange (2) and a head (16) fixed on the downstream flange (5), said drainage conduit (10) being formed in one of the plurality of blades (14). Diffuser (1) according to one of claims 1 to 4, comprising:

• a radial part (13) into which the centrifugal impeller (21) is configured to discharge, the radial part (13) comprising a plurality of blades (14) extending in the vein (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.

Diffuser (1) according to one of claims 1 to 6, comprising at least one fixing flange (19) configured to be fixed to a casing (50) of an aircraft turbomachine (60), the drainage duct (10) extending through the fixing flange (19). Assembly of a diffuser (1) according to one of claims 1 to 6 and of a casing (50) of an aircraft turbomachine (60), in which the casing (50) and an outer face (4) of the upstream flange (2) together delimit an evacuation space (51) into which the upstream end (12) of the drainage duct (10) opens, the casing (50) comprising at least one drainage orifice (52) in fluid communication with the evacuation space (51) and opening outside the aircraft turbomachine (60). 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 aircraft turbomachine (60) comprising a diffuser (1) according to 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 drainage duct (10) opening into the peripheral chamber (30). Method of using an aircraft turbomachine (60) according to claim 9, in which the drainage conduit (10) drains the liquid (L) present in the peripheral chamber (30) by gravity in nominal use.