FIRE SUPPRESSION SYSTEM FOR A TURBOMACHINE INCLUDING MEANS FOR MAINTAINING A COOLING AIR SPEED AND CORRESPONDING TURBOMACHINE

Abstract

The invention relates to a fire suppression system (30) for a turbomachine (1), such as an aircraft turbojet or turboprop engine, the turbomachine comprising at least one turbine (3) having a turbine disk (5) and an annular cavity (16) arranged upstream of the disk (5), the disk (5) comprising an internal bore (7), and the fire suppression system (30) comprising a cooling device (14) for supplying the cavity (16) with cooling air via injection means (17). According to the invention, the fire suppression system (30) comprises means configured to, on the one hand, split the annular cavity into a first cavity and a second cavity, and on the other hand, maintain a velocity of the cooling air exiting the injection means (16) and guide the cooling air in the first cavity towards the internal bore (7) of the turbine disk (5). Figure 2

Description

Title of the invention: FIRE-FIGHTING SYSTEM FOR A TURBOMACHINE COMPRISING MEANS FOR MAINTAINING A COOLING AIR SPEED AND CORRESPONDING TURBOMACHINE Field of invention

[0001] The present invention relates to the general field of turbomachines. It relates in particular to a fire-fighting system installed in a cavity supplied with cooling air, close to a hot zone of a turbomachine, and which comprises means for maintaining a speed of the cooling air entering the cavity. The invention also relates to a turbomachine comprising such a fire-fighting system. Technical background

[0002] A turbomachine for an aircraft generally comprises, from upstream to downstream and depending on the direction of flow of the gases in the turbomachine, a compressor section, a combustion chamber, a turbine section. The compressor section comprises for example a low-pressure compressor and a high-pressure compressor and the turbine section comprises for example a high-pressure turbine and a low-pressure turbine. The turbomachine may comprise a free turbine which is driven by the gases at the outlet of the high-pressure turbine or the low-pressure turbine located upstream thereof. The rotating parts of these compressors and / or turbines such as shafts and wheels are driven and / or guided in rotation using rolling bearings which are housed in cooling and lubrication enclosures.The cooling and lubrication enclosures are arranged near so-called hot zones which are generally exposed to high temperatures because these parts are crossed by the gases from the turbomachine.

[0003] For this purpose, the turbomachines are equipped with a lubrication system for lubricating and / or cooling the rolling bearing(s) required for guiding the shafts and which are mounted in the cooling and lubrication enclosures. The turbomachines are also equipped with a cooling device for cooling the parts located in hot zones of the turbomachine and for pressurizing the enclosures. Generally, the lubrication system and the cooling device are arranged in restricted spaces favoring the proximity of the lubrication circuit of the lubrication system with rotor parts located in the hot zone of the turbomachine. In the event of a leak, the lubricant may encounter in these areas, conditions of temperature, pressure, cooling air flow speed and cooling air turbulence level which are favorable to self-ignition. This risk of self-ignition could cause damage to rotor parts, or even the rupture and / or release of parts rotating at high speed.

[0004] The invention aims to avoid the aforementioned drawbacks. Summary of the invention

[0005] The objective of the invention is to provide an optimal solution making it possible to limit, or even prevent, an ignition of a combustible fluid in a hot zone of the turbomachine while being simple and economical.

[0006] We achieve this objective in accordance with the invention by means of a fire-fighting system for a turbomachine, such as an aircraft turbojet or turboprop, the turbomachine comprising at least one turbine having a turbine disk and an annular cavity which is arranged upstream of the disk, the disk comprising an internal bore and the fire-fighting system comprising a cooling device intended to supply the cavity with cooling air via injection means, the fire-fighting system comprising means configured so as to, on the one hand, split the annular cavity into a first cavity and a second cavity and, on the other hand, maintain a speed of the cooling air at the outlet of the injection means and guide the cooling air in the first cavity towards the internal bore of the turbine disk.

[0007] Thus, this solution makes it possible to achieve the aforementioned objective. In particular, such means make it possible to prevent a possible flame following a fire starting in the cavity from being stabilized. The configuration of these means makes it possible to maintain the flow rate of the cooling air from the outlet of the injection means to the internal bore of the turbine disk, and to avoid recirculation of cooling air in the cavities, in particular the first cavity close to the turbine (free or linked) so as not to stabilize a fire there.

[0008] The fire protection system comprises one or more of the following features, taken alone or in combination:

[0009] - the speed maintaining means comprise an annular diffuser extending between an upstream edge and a downstream edge along an axis D and being provided with channels which are formed in the thickness of an annular wall of the diffuser, the channels each having an evolving profile and each opening, on the one hand in the upstream edge through a plurality of ports, and on the other hand in the downstream edge in an annular opening, the ports being arranged around the axis D and being in fluid communication with the injection means of the cooling air device.

[0010] - the diffuser has a predetermined length between the upstream edge and the edge downstream, the downstream edge being intended to be substantially flush with a free, radially internal end of a first flange mounted upstream of the turbine disc and with a predetermined clearance.

[0011] - the fire protection system is made according to a geometric law defined by the formula next: "< --2..... ^LC

[0012] with a being a predetermined angle measured between a first straight line passing through a vertex of a lumen, parallel to a respective main axis of the channels, and a second straight line tangent to a flank of a channel, L being the predetermined length of the diffuser and LC being the circumferential width of each lumen.

[0013] - the turbomachine comprises a guide bearing support for a turbine shaft which comprises a first radial flange, and a cover provided with a second radial flange intended to be fixed on the first radial flange by first fixing members, the first fixing members being distributed regularly around the longitudinal axis X and the cover comprising an annular wall intended to pass through the internal bore of the turbine disk.

[0014] - the system comprises an annular cover intended to be mounted on the cover and to cover the first fixing members, the cover comprising an annular bottom wall intended to be fixed on an annular surface of the cover.

[0015] - the annular hood comprises an annular skirt extending from the bottom wall along an axis C of the hood, the annular skirt comprising a tubular portion and a frustoconical portion connecting the tubular portion to the bottom wall along the axis C, the hood comprising a predetermined radius of curvature arranged on the one hand, between the tubular portion and the frustoconical portion and on the other hand, between the frustoconical portion and the bottom wall.

[0016] - the hood has an annular edge arranged adjacent to a downstream end of a radially inner wall of the diffuser, the diffuser comprising a radially outer wall having a downstream edge arranged radially outside the hood and at a distance from the hood.

[0017] - the turbine disc comprises a circumferentially extending annular rim around the longitudinal axis and axially along the longitudinal axis upstream, a portion of a cooling air flow vein being delimited at least in part by a radially internal surface of the annular rim and a radially external surface of the cover.

[0018] - the ratio between the predetermined radius of curvature and the height of the vein cooling airflow is greater than 1.

[0019] - the free end of the flange is annular and radially surrounds on the outside a annular rim of the turbine disc.

[0020] - the system is configured so that the ratio of the flame speed turbulent flow rate of an air / oil mixture is less than 1.

[0021] - the tubular portion of the hood comprises the annular edge.

[0022] - the predetermined length of the diffuser is intended to be less than or equal to a distance measured between a distal end of an annular rim of the turbine disc and the injection means.

[0023] - the fixing members each comprise a head and in that the bottom wall includes housings which each open into a third hole, each housing being intended to house and contain the head of the fixing members.

[0024] - the number of fixing members is between 4 and 8.

[0025] - the diffuser is produced by an additive manufacturing process or selective fusion of powder.

[0026] - each lumen is in fluid communication with an injector of the means injection.

[0027] - the predetermined angle a being between 9° and 24°.

[0028] The invention also relates to a turbomachine, in particular of an aircraft, having a longitudinal axis X and comprising a turbine connected by a shaft providing mechanical power, a cavity arranged upstream of a first stage of the turbine, a cooling device equipped with injection means, and a fire-fighting system as mentioned above, the fire-fighting system being arranged in the cavity and the injection means being in fluid communication with the channels of the cooling system.

[0029] The turbomachine comprises one or more of the following features, taken alone or in combination:

[0030] - a first flange is arranged upstream of the turbine disc, the downstream edge of the diffuser being placed near a radially inner end of the first flange with the predetermined clearance.

[0031] - the turbine is arranged downstream of a turbine section.

[0032] - the shaft is a power shaft providing mechanical power to a shaft exit.

[0033] - the turbine is a linked turbine or a free turbine.

[0034] The invention further relates to an aircraft comprising at least one turbomachine as mentioned above. Brief description of the figures

[0035] The invention will be better understood, and other objects, details, characteristics and advantages thereof will appear more clearly upon reading the description ex detailed application which will follow, of embodiments of the invention given as purely illustrative and non-limiting examples, with reference to the appended schematic drawings in which:

[0036] [Fig-1] [Fig. 1] represents an axial and partial sectional view of an example of a turbomachine comprising at least one turbine according to the invention;

[0037] [Fig.2] [Fig.2] is an axial sectional view of a fire-fighting system in a pressurized cavity according to the invention;

[0038] [Fig.3] [Fig.3] is a perspective view of an example of a guide bearing support for a power shaft of a turbomachine according to the invention;

[0039] [Fig.4] [Fig.4] is a perspective view of an example of a cover intended to be mounted on a bearing support according to the invention;

[0040] [Fig.5] [Fig.5] is a perspective view of an example of a cooling air guide member according to the invention;

[0041] [Fig.6] [Fig.6] is an axial sectional view of a fire-fighting system arranged in a pressurized cavity of a turbomachine according to the invention;

[0042] [Fig.7] [Fig.7] represents in perspective an example of a member for adapting an air flow to another member of the turbomachine according to the invention;

[0043] [Fig.8] [Fig.8] is a view along an axial section AA of a channel of the air flow adaptation member of [Fig.7];

[0044] [Fig.9] [Fig.9] is a view along a section BB of the canal according to [Fig.8]; and

[0045] [Fig. 10] [Fig. 10] is a view along a section CC of the canal according to [Fig.8]. Detailed description of the invention

[0046] [Fig.l] partially illustrates a turbomachine 1, such as a turbojet, a turboprop or a turboshaft engine. This turbomachine is intended to be mounted in an aircraft such as an airplane or a helicopter.

[0047] In the present application, the terms "upstream", "downstream", "axial" and "axially" are defined with respect to the direction of circulation of the gases in the turbomachine and also along the longitudinal axis X (and even from left to right in [Fig.l]). The terms "radial", "radially", "internal" and "external" are also defined with respect to a radial axis Z which is perpendicular to the axis X of the turbomachine.

[0048] Generally speaking, a turbomachine, in particular an aircraft turbomachine, with a longitudinal axis X, comprises, from upstream to downstream and in the direction of flow of the gas or air flows, a compressor section, a combustion chamber, and a turbine section. These elements form a gas generator 2. The compressor section may comprise a low-pressure compressor and a high-pressure compressor. The turbine section may comprise a low-pressure turbine and a high-pressure turbine. These turbines are known as “linked turbines”. Each com pressure (low or high pressure) and each turbine (low or high pressure) comprises one or more stages respectively. Each stage comprises a wheel of moving blades which is mounted upstream or downstream of a wheel of fixed blades (or stator blades). The rotors of the low pressure compressor and the low pressure turbine are connected to each other by a low pressure shaft centered on the longitudinal axis X to form a low pressure body. Similarly, the rotors of the high pressure compressor and the high pressure turbine are connected to each other by a high pressure shaft centered on the longitudinal axis X to form a high pressure body. The turbomachine may comprise upstream of the compressor section a fan (not shown). This may also comprise downstream of the turbine section a gas exhaust nozzle.Air entering the compressor section passes through the combustion chamber which emits combustion gases to the turbine section comprising at least one linked turbine.

[0049] Referring to [Fig.l], the turbomachine comprises an additional turbine 3 called a free turbine which is mounted downstream of the gas generator 2. In particular, the free turbine is mounted downstream of the high pressure turbine or the low pressure turbine (of the turbine section). The gases from the gas generator 2 are sent to the free turbine 3 which drives a power shaft 4. In this example, the free turbine 3 is placed upstream of the exhaust nozzle or as a replacement for the latter. The power shaft provides the useful mechanical power to an output shaft for the propulsion of the aircraft. The power shaft is separate from the high pressure and / or low pressure shafts and is coaxial with them. The free turbine 3 here comprises a turbine stage with a moving wheel and a fixed wheel. The moving wheel comprises an annular disc 5 from which several moving blades 6 extend radially.The latter are distributed regularly around the disc 5 and pass through a primary vein in which a primary flow circulates. The disc 5 is driven in rotation around the longitudinal axis X by a shaft (here the power shaft) which passes through its internal bore 7 (turned towards the axis X). The fixed wheel is composed of fixed blades called distributors 8 and which are arranged upstream of the moving blades 6.

[0050] The disc 5 comprises an annular rim 9 which extends from a web 10 of the disc 5 along the longitudinal axis. The annular rim 9 also extends circumferentially around the longitudinal axis X. The annular rim 9 carries a sealing device 11 intended to cooperate with a free, radially internal end 12a of a first flange 12. The sealing device 11 is advantageously a labyrinth seal which comprises several wipers or blades extending radially and circumferentially around the longitudinal axis from the wall of the annular rim 9. These wipers are also arranged parallel along the longitudinal axis X. The first flange 12 is annular and is mounted upstream of the web 10 of the disc 5. The radially internal end 12a also carries an abradable coating arranged opposite the wipers to guarantee the sealing of the sealing device 11. The first flange 12 is also connected to a radially internal platform 13 of the upstream distributor.

[0051] With reference to Figures 1 and 2, the turbomachine 1 generally comprises a device 14 for cooling the disk 5 of the free turbine 3 (or one of the linked turbines). The cooling device 14 is supplied with air which is taken upstream of the turbomachine. Generally, the cooling air is taken from the high pressure or low pressure compressor. The cooling device 14 comprises an external pipe (not shown) which bypasses the disks of the turbine section in which cooling air circulates. The external pipe is coupled to an internal pipe 18 which is also connected to a rear bearing housing 15 which comprises a first wall 15a and a second wall 15b connected to a fixed structure of the turbomachine. The housing 15 is a part of revolution around the longitudinal axis.The cooling device 14 comprises injection means intended to project the cooling air into an annular cavity 16 arranged upstream of the disk 5 of the free turbine. The injection means are formed in the rear bearing housing 15. This annular cavity 16 is in particular located upstream of the first stage of the free turbine 3. The injection means here comprise injectors 17 which are coupled to the air duct 18 and which open into the annular cavity 16. The injectors 17 are orifices. Alternatively, the injectors are nozzles. The injectors 17 are distributed regularly around the longitudinal axis and are advantageously placed substantially opposite the disk 5 of the free turbine 3. Each injector 17 has an axis parallel to the longitudinal axis X. In the present example, there are thirty-two injectors 17.

[0052] A structural part 28 is mounted on the rear bearing housing 15. The structural part 28 is an annular part which has an axis centered on the longitudinal axis X and which has a U-shaped (pin-shaped) radial section. This makes it possible to make the connection between the housing 15 and the distributor 8. In particular, this makes it possible to make the connection between the so-called “hot” part and the so-called “cold” part of the rear bearing housing 15. This is also flexible. Due to its flexibility, the structural part 28 absorbs deformations due to thermal differences.

[0053] The turbomachine 1 comprises several bearings which make it possible to guide the rotation of the low pressure shaft, the high pressure shaft and / or the power shaft of the free turbine 3. The bearings are generally arranged in different enclosures, such as the lubrication and cooling enclosure 19, which are supplied with lubricant, here oil in the form of a mist, coming from a lubrication system. fication. The enclosure 19 is pressurized to contain the oil in a sealed manner therein and also in the form of mist. The oil mist allows for a homogeneous distribution of the oil over the components of the turbomachine to be lubricated.

[0054] The pressurization of the enclosure 19 is obtained by injecting air which is taken from the compressor(s) or the blower. The air which is taken for the pressurization of the enclosures uses the same circuit as that for cooling the disks. Preferably, the cooling air is taken upstream of the blower so that the temperature of the cooling air is not too high and so that it can effectively cool the components of the turbomachine. The annular cavity 16 is arranged radially around the enclosure 19 and is separated from it by a cover 31. The cover 31 itself is arranged radially outside a bearing support which partially delimits the enclosure 19.

[0055] With reference to [Fig. 1], the power shaft is guided in rotation by a bearing 20 located downstream of the latter and which is arranged in the enclosure 19. This bearing 20 cooperates with the downstream bearing support 21 which is generally fixed to a fixed structure of the turbomachine. The downstream bearing 20 is advantageously a rolling bearing. The latter comprises an inner ring and an outer ring between which the rolling members are arranged. The rolling members are for example rollers.

[0056] In [Fig.2] and [Fig.3], the bearing support 21 has a revolution shape extending around an axis A which is coaxial with the longitudinal axis X of the turbomachine in the installation situation. The bearing support 21 is formed of two parts hereinafter referred to as the first shell 22 and the second shell 23. The first shell 22 more precisely comprises a first annular wall 22a and a first radial fixing flange 22b which extends radially outwards from a radially external surface of the first annular wall 22a. The first shell 22 is advantageously made in one piece. The first annular wall 22a carries the external ring of the downstream bearing 20. The internal ring of the downstream bearing 20 cooperates with means mounted on the power shaft. The second shell 23 comprises a second annular wall 23a which is mounted on the first shell 22a and which extends radially outside the first annular wall 22a.The second annular wall 23a passes at least partly through the internal bore 7 of the disc 5 of the free turbine 3 along the longitudinal axis X. The second annular wall 23a supports a static part of a labyrinth seal which is located between the enclosure 19 and the annular cavity 16.

[0057] The bearing support 21 is fixed to the fixed structure via the rear bearing housing 15. In particular, the first radial flange 22b is intended to be fixed to the second wall 15b of the rear bearing housing 15. For this, the first radial flange 22b comprises a plurality of first holes 24 passing through the wall of the first flange radial 22b on either side along an axis parallel to the longitudinal axis X. The first holes 24 are distributed regularly around the longitudinal axis. The second wall 15b of the rear bearing housing 15 also comprises a radial portion 25 in which several blind holes 26 are formed. The blind holes 26 extend along an axis parallel to the longitudinal axis X. These are intended to cooperate with the first holes 24 of the first radial flange 22b. There are as many blind holes 26 as there are first fixing holes 24. First fixing members 27 such as screws, rods and / or bolts, etc. make it possible to fix the bearing support 21 to the housing 15. In the example shown, the first fixing members 27 comprise screws with a head 27a and a rod 27b.

[0058] The first radial flange 22b comprises a first bearing surface 29a which is defined in a radial plane perpendicular to the axis A of the bearing support. The first radial flange 22b also comprises a second bearing surface 29b which is defined in a plane perpendicular to the axis A. The first and second bearing surfaces 29a, 29b are opposite along the longitudinal axis X.

[0059] The radial portion 25 also comprises a downstream bearing surface 25a which is defined in a plane perpendicular to the longitudinal axis X. The first bearing surface 29a of the first radial wall 22b is intended to bear against the downstream bearing surface 25a. These create a support-plane connection. In the present example, the first radial flange 22b comprises an annular edge which has an annular surface 22c flush with a radially external surface 25c of the radial portion 25 so as not to disturb the flow of cooling air leaving the cooling device 14.

[0060] In [Fig. 4] the cover 31 is shown. The latter is intended to be mounted on the bearing support 21. The cover 31 is installed in the annular cavity 16 and the cooling air leaving the injectors circulates around it and towards the bore 7. The cover 31 comprises a shape of revolution around an axis B. The axis B is coaxial with the longitudinal axis X in the installation situation. The cover 31 comprises a second radial flange 33 which is intended to be fixed on the first radial flange 22b of the bearing support 21. The fixing of the first radial flange 22b on the second radial flange 33 is carried out by means of the first fixing members 27 (shown in [Fig. 5]). The cover 31 also comprises an annular wall 32 which extends along the axis B and at least partly through the internal bore of the turbine disc.The second radial flange 33 extends radially outward from a radially outer surface 34 of the annular wall 32. The annular wall 32 has an internal diameter which is greater than the external diameter of the second shell 23 (in particular the second annular wall 22a) of the bearing support 21. These first fixing members 27 are distributed regularly around the longitudinal axis X. At this . In effect, the second radial flange 33 comprises a plurality of second holes 35 which each pass through the wall of the radial flange 33 on either side along an axis parallel to the longitudinal axis X. The second holes 35 are intended to cooperate with the first holes 24 of the first radial flange 22b. In the installation situation the first and second holes 24, 35 are coaxial.

[0061] The second radial flange 33 comprises an upstream bearing surface 33a (see [Fig.2]) which is defined in a radial plane perpendicular to the axis B of the cover 31. The upstream bearing surface 33a is intended to bear against the second bearing surface 29b of the first flange 22b so as to create a support-plane connection. The second radial flange 33 also comprises an annular edge which has an annular surface 33c substantially flush with the surface 22c of the first flange.

[0062] As shown in [Fig. 4], the cover 31 comprises an annular bearing surface 36 extending radially outwardly and downstream of the second radial flange 33. In particular, the cover 31 comprises a plurality of projections 37 which extend radially outwardly from the radially external surface 34 of the annular wall 32 and which are regularly distributed around the axis of the cover 31. The projections 37 also extend from a downstream surface 33b of the second radial flange 33 along the longitudinal axis X. The downstream surface 33b is axially opposite the upstream bearing surface 33a. The annular bearing surface 36 is formed by the projections 37. The projections 37 are defined by recesses 38 into which the second holes 35 for fixing the second radial flange 33 open. In other words, the projections 37 are arranged alternately with the second holes 35 around the longitudinal axis.The recesses 38 have an orientation substantially parallel to the longitudinal axis and each have a semi-cylindrical section.

[0063] The projections 37 are circumferentially delimited by lateral panels 37a, 37b (which are also the walls of the recesses). These are also each delimited downstream by a downstream bearing surface 39. The latter forms an annular downstream bearing surface of the annular bearing surface 36. The downstream bearing surface 39 is defined in a plane which is perpendicular to the longitudinal axis X.

[0064] The turbomachine 1 also comprises a fire-fighting system 30 configured so as to annihilate a fire starting in the cavity 16 or even to eliminate any possibility of ignition in the cavity 16. Indeed, the enclosure 19 with the oil in suspension is close to the cavity 16 which receives cooling air via the cooling device 14. The oil, the cooling air and the heat prevailing in this zone could allow combustion and a fire starting generating a flame in this zone and in the cavity 16 in the event of oil leaking from the enclosure into the cavity. A fire could cause significant damage to the rotor parts, or even the release of high-energy parts.

[0065] When a flame is generated in a cavity exposed to the flow of a mixture of a first fluid (air) and a second combustible fluid (oil), it can be stabilized if the flame speed is sufficient to exceed the flow speed of the cooling air in the cavity. The flame speed depends on the temperature of the cooling air, the pressure and the oil concentration.

[0066] The maximum possible flame speed is the turbulent flame speed ST at the maximum pleating ratio. It is considered that the cooling air / oil mixture is stoichiometric to maximize the laminar flame speed, that there is no heat transfer to the cavity walls to maximize the laminar flame speed and that the pleating ratio is at saturation to maximize the flame pleating factor. The laminar flame speed is a constituent of the fuels (gasoline, kerosene, diesel, etc.) which form the combustibles.

[0067] The system 30 is configured so that the ratio of the flame speed to the flow speed of the cooling air is less than 1. In particular, the system 30 comprises maintaining means configured so as to maintain a speed of the cooling air at the outlet of the injection means and to guide the cooling air through the bore of the turbine disc of the first turbine stage. By maintaining the outlet speed of the cooling air from the injection means and by avoiding recirculations, it is possible to control and avoid stabilization of the flame generated in the cavity.

[0068] The fire-fighting system 30 comprises a member for adapting the speed of the cooling air leaving the injection means of the cooling device 14. In particular, the adaptation member is configured so as to maintain the speed at the outlet of the injectors 17. In the present example, this member is a diffuser 52 which also makes it possible to guide the cooling air towards the bore 7 of the disk 5 of the free turbine. The adaptation member is part of the means for maintaining the speed.

[0069] In [Fig.7], the diffuser 52 is annular and comprises a shape of revolution around an axis D. The axis D is coaxial with the axis of the turbomachine 1 in the installation situation. The diffuser 52 comprises an annular wall 53 which extends along the axis D. The annular wall 53 is delimited upstream by an upstream edge 54 and downstream by a downstream edge 55. The upstream edge 54 and the downstream edge 55 are opposite along the axis D. The upstream edge 54 has a surface 54a which is defined in a plane which is perpendicular to the axis D. The surface 54a is intended to bear against a bearing surface 90 (see [Fig. 6]) of the rear bearing housing 15. In particular, the annular bearing surface 90 is defined in a plane which is perpendicular to the longitudinal axis X. The bearing surface 90 is formed by a shoulder which also forms the radial portion 25. Furthermore, the bearing surface 90 is located upstream of the bearing surface downstream 25a. The injectors 17 of the cooling device 14 open into the bearing surface 90.

[0070] The annular wall 53 of the diffuser 52 has an internal diameter which is greater than the external diameter of the annular edge (delimited by the annular surface 22c) of the first radial flange 22b. Similarly, the diameter of the annular wall 53 is greater than the diameter of the radial portion 25 (delimited by the surface 25c).

[0071] The diffuser 52 installed in the annular cavity splits it into a first cavity 16a and a second cavity 16b. The second cavity 16b is located radially outside the diffuser 52 and is considered a dead cavity since no cooling air is intended to circulate therein. The first cavity 16a is located radially inside the diffuser 52. The cooling air is mainly guided towards the first cavity 16a which forms a cooling air stream V.

[0072] Indeed, the diffuser 52 extends, along the longitudinal axis, between the injection means (injectors 17) and the radially internal end 12a of the first flange 12 (and / or a distal end of the annular rim 9 of the disc). The annular wall 53 has a length L1 which is greater than the length of the radial portion 25 (and in particular of the surface 25c of the radial portion 25). In the installation situation, the upstream edge 54 bears against the bearing surface 90 and at least a portion of the radially internal surface of the annular wall 53 is in contact with the radially external surface 25c of the housing 15.

[0073] As also illustrated in Figures 2 and 6, the downstream edge 55 of the diffuser is located close to the radially internal end 12a of the first flange 12. The length L1 is substantially equal to the distance between the outlet of the injection means (injectors 17) and the radially internal end 12a of the first flange 12.

[0074] A flow of cooling air circulates in the sealing device 11 and a flow of cooling air circulates in the flow stream V. A clearance J remains between the downstream edge 55 and this radially internal end 12a. However, the cooling air does not circulate through this clearance J. The clearance J makes it possible to take into account the thermal expansions of the parts and in particular of the diffuser and the flange. Advantageously, the clearance is less than or equal to 1 mm.

[0075] With reference to Figures 8 to 10, the diffuser 52 comprises a plurality of channels 56 which are formed in the thickness of the annular wall 53. The channels 56 are intended to be in fluid communication with the injection means (in particular the injectors 17) of the housing 15. There are as many channels as there are injectors 17. Here, there are therefore thirty-two channels 56 which are distributed regularly around the axis D. Each injector 17 opens into a channel 56. The channels 56 form a radially inner wall 53a and a radially outer wall 53b. The radially inner wall 53b has a length L2 less than the length of the radially outer wall 53b. The length L2 is measured between a first upstream end and a first downstream end. The length of the radially outer wall 53b corresponds to the length L1 of the annular wall 53. The radially outer wall comprises a second downstream end downstream of the first downstream end. The second downstream end forms the downstream edge 55 of the diffuser 52. Each channel 56 is delimited by the radially inner wall 53b and the radially outer wall 53a which are connected by a first flank 57a and a second flank 57b. The first and second flanks 57a, 57b are opposite in the circumferential direction (around the axis D).

[0076] The channels 56 have an evolving profile so as to be able to guide the flow and control the flow rate of the cooling air through the diffuser 52 and into the cavity. By “evolving profile” we mean the fact of having a variation in shape, thickness and / or dimensions according to one or more sections. The channels 56 each open into the upstream edge 54 respectively through a plurality of slots 58. In particular, each slot 58 opens into the surface 54a. The slots 58 are arranged around the axis of the diffuser 52. Advantageously, but not limitingly, the slots 58 are spaced and distributed regularly around the axis of the diffuser. In the present example, each slot 58 has an elongated (or oblong) shape in the circumferential direction (around the axis D). The channels 56 also open into an annular opening 59 at the downstream edge 55. The annular opening 59 is a single opening.

[0077] The evolving profile of the channels 56 is obtained by varying the radial height and the circumferential width thereof. More precisely, each channel 56 has a radial height H which varies between the upstream edge 54 and the downstream edge 55. Even more precisely, the height H varies by decreasing from the upstream edge 54 towards the downstream edge 55. In other words, the height at the level of the light 58 is greater than the height at the level of the downstream edge 55. Each channel 56 also has a circumferential width LC, measured between the first and second flanks 57a, 57b, which varies between the upstream edge 54 and the downstream edge 55. The circumferential width LC increases from the upstream edge 54 towards the downstream edge 55.

[0078] The section of each channel is substantially constant (within + / - 20%). In particular, the variation in height H and circumferential width LC is substantially constant. In this way, the section of the channels 56 changes little and the flow speed also changes little, thus remaining almost constant between the inlet and the outlet of the diffuser 52 while directing the cooling air correctly to cool the disk 5 of the free turbine.

[0079] In the exemplary embodiment, each lumen 58 comprises a height of between 3 and 4 mm. This comprises a width LC of between 7 and 8.5 mm. This also comprises a lumen height corresponding to the height of the channel at light level (channel input).

[0080] In [Fig. 10], each lumen 58 also comprises a connecting radius RI of the order of 1.7 mm between the first edge 58a and the second edge 58b of the lumen between which the height of the lumen is measured. This connecting radius is substantially identical between the radially inner wall and the radially outer wall at the flanks.

[0081] Each section S of a channel 56 of the diffuser 52 (in a plane perpendicular to the axis D of the diffuser) is less than the ratio D / pST. In this way, the shape of the channel makes it possible to maintain the flow rate from the inlet (light) to the outlet (opening) of each channel of the diffuser 52. S is the section of each channel 56, p is the density of the cooling air at the outlet of the injection means, ST is the flame speed and D is the mass flow rate of the cooling air. The section of each channel can be between 26 and 31 mm2.

[0082] [Fig. 9] represents a section along the section BB of [Fig. 8] (in the circumferential direction). In this section we see that a channel 56 has a section that is substantially frustoconical in shape and in particular straight. The flanks 57a, 57b of each channel 56 are inclined at a predetermined angle a (alpha). The predetermined angle a is measured between a first straight line DI passing through a vertex of a slot (corresponding to a channel respectively), parallel to the respective main axis E of the channel 56, and a second straight line D2 tangent to the flank of the channel. Advantageously, but not limitatively, the predetermined angle a is between 9° and 24°.

[0083] The shape of the diffuser 56 is obtained by applying the following formula:

[0084] a<-^

[0085] With a being the predetermined angle, L being a predetermined length of the diffuser 52 between the upstream edge 54 and the downstream edge 55 (i.e. the length L1), and “a” being the circumferential width LC (or the diameter of each light). This formula is an empirical formula which aims to identify the angle from which a separation takes place. Numerous experimental tests have been carried out on the diffuser in order to characterize the different parameters. The formula is for example described in the publication “Memento des perte de charges”, LE. Idel'cik, Ed. Eyrolles, Paris, 1986.

[0086] Advantageously, but not limitatively, the diffuser 52 comprises sealing means arranged between the channels 56 in the circumferential direction. The sealing means comprise the separation walls between the channels.

[0087] Advantageously, the diffuser 52 is produced by an additive manufacturing process or selective powder fusion. In this way, it is produced from a single piece (made from one material). Additive manufacturing makes it possible to produce complex geometries. and parts made from one piece. Preferably, but not limited to, additive manufacturing is a laser powder bed fusion process known by the acronym SLM for "Selective Laser Melting". The process is carried out using a facility in which several layers of materials, particularly in powder form, are superimposed on a manufacturing support. The powder layers from a feed tank are transferred to the manufacturing support and are then melted one after the other by means of a laser beam moving over the surface of each layer. With the additive manufacturing process, channels with such an evolving profile that is both convergent (radially) and divergent (in the azimuthal direction), as well as such small dimensions, can be produced in one piece.Likewise, since the diffuser 52 is produced by additive manufacturing, the inter-channel walls forming the sealing means are continuous and without steps, which guarantees very good sealing.

[0088] The diffuser 52 is made of a metallic material. An example of a metallic material is a nickel-based alloy such as Inconel 718. Such a material has high thermomechanical resistance.

[0089] With reference to Figures 2, 5 and 5, the fire-fighting system 30 further comprises an annular cover 40 intended to be mounted on the cover 31. The cover 40 comprises a shape of revolution around an axis C. The axis C is coaxial with the longitudinal axis X of the turbomachine in the installation situation. The cover 40 is intended to cover the first fixing members 27 in order to prevent the recirculation of air at the level of these. The recirculations create turbulence which by nature promotes the attachment of flames. The cover 40 makes it possible to avoid turbulence and to reduce the passage section under the disk and therefore to accelerate the speed of the cooling air. For this purpose, the first radial flange 22b of the bearing support 21 is mounted on the housing 15 which comprises a portion of the cooling air circuit.The cooling air leaving the injectors 17 is injected in a direction parallel to the longitudinal axis and radially outside the cover 40. The cooling air circulates in the first cavity 16a (flow air vein V).

[0090] The cover 40 and the diffuser 52 form the means for maintaining the speed and guiding the cooling air in the cavity (in particular the first cavity 16a) and in the bore of the disc.

[0091] With reference to [Fig. 6], the fact of attaching and fixing the cover 40 to the lid 31 makes it possible to reconfigure a portion of a flow vein V of cooling air between the disc 5 and the cover 40. In this way, the cooling air no longer circulates around the screws (fixing members 27) which are covered by the cover. In particular, the portion of flow vein V of cooling air which is reconfigured is delimited at least in part by a radially internal surface 9a of the annular rim 9 and a radially external surface 40a of the cover. The internal diameter delimited by the radially internal surface 9a is greater than the external diameter of the annular skirt of the cover (described below) and defined by a radially external surface. The cooling air flow path V has a height Dh which is substantially constant between the annular rim and the internal bore 7 of the disc 5.

[0092] In [Fig. 5], the cover 40 comprises a bottom wall 41 from which an annular skirt 42 rises. The bottom wall 41 is crossed by a central bore 43 along the axis C of the cover. The central bore 43 defines an internal annular surface 43a. The annular skirt 42 extends along the axis C between the bottom wall 41 and an upstream annular edge 44. The annular wall 32 of the cover 31 is intended to pass through the central bore 43 of the bottom wall 41 of the cover 40. Advantageously, the internal diameter of the annular skirt 42 of the cover 40 is greater than the external diameter of the annular wall of the cover 31 so as to cover the first fixing members 27.

[0093] According to the example shown, the diameter of the annular skirt 42 varies between the bottom wall 41 and the annular edge 44. In particular, the annular skirt 42 has a decreasing diameter from the annular edge 44 to the bottom wall 4L. In [Fig. 2], the annular edge is defined in a radial plane PI which is upstream of a radial plane P2 in which the bearing surface of a screw head 27a of the first fixing members 27 is defined. In other words, the annular edge 44 at least partially overlaps the projections 37 of the cover. This makes it possible to cover the screws well. The rod 27b connected to the screw head 27a extends through the holes 24, 35 of the flanges 22b, 33.

[0094] More precisely still and with reference to figures 5 and 6, the annular skirt 42 comprises a tubular bearing surface 45 which comprises the annular edge 44. The annular skirt 42 also comprises a substantially frustoconical portion 46 which connects the tubular portion to the bottom wall 4L.

[0095] Advantageously, the cover 40 comprises a first predetermined radius of curvature “rl” which is arranged between the tubular portion 45 and the frustoconical portion 46 and a second predetermined radius of curvature “r2” which is arranged between the frustoconical portion 46 and the bottom wall 4L. In the example shown, the first and second radii of curvature are identical.

[0096] The ratio between the predetermined radius of curvature (rl, r2) and the height Dh of the cooling air flow duct V is greater than 1. Such a ratio makes it possible to maintain the speed and avoid recirculation of the cooling air leaving the injection means towards the internal bore of the disc 5. An example of a radius of curvature is between 2 and 3 mm.

[0097] According to an exemplary embodiment illustrated in [Fig.5], the substantially frustoconical portion comprises a first frustoconical portion 46a which is connected to the wall of bottom 41 and a second frustoconical portion 46b which is arranged between the tubular portion 45 and the first frustoconical portion 46a. Alternatively, the annular skirt 42 comprises a single frustoconical portion whose diameter variation is constant. The radii of curvature rl, r2 are arranged at the same locations.

[0098] The bottom wall 41 is intended to be fixed on the annular bearing surface 36 of the cover 31 by second fixing members 48 illustrated in FIGS. 5 and 6. The second fixing members 48 may be screws, rods and / or bolts, etc. The bottom wall 41 comprises a plurality of third holes 49 which pass through the wall on either side along an axis parallel to the axis C. These third holes 49 are intended to cooperate with fourth holes 50 which are arranged in the annular bearing surface 36 of the cover 31. The bottom wall 41 comprises an upstream surface 41a (see [Fig. 6]) which is defined in a plane perpendicular to the longitudinal axis X. The upstream surface 41a bears against the annular bearing surface 39 of the cover. The fourth holes 50 pass through the wall of the projections 37 on either side along the axis B as illustrated in Figures 4, 5 and 6.

[0099] With reference to [Fig. 5], the bottom wall 41 comprises housings 63 which each open into a third hole 49. Each housing 63 also opens onto a downstream surface 41b of the bottom wall 4L. The downstream surface 41b is opposite the upstream surface 41a along the axis of the cover 40 (longitudinal axis in the installation situation). In the present example, the housing 63 is a recess made in the thickness of the bottom wall 4L. The housings 63 each also open onto the internal annular surface 43a of the central bore of the cover 40. Alternatively, the housings are made by counterbores. Here, the second fixing members are advantageously screws each having a head 48a and a shank 48b. Each head 48a is housed and contained in a housing 63. The heads do not protrude from the housing 63 (in particular when the fixing members 48 are mounted in the cover 40 and the lid 31).In other words, the height of the housing along the longitudinal axis is equal to or greater than the height of a head 48a. This prevents recirculation and obstructs the flow vein V. The number of fixing members is between 4 and 8.

[0100] Advantageously, but not limitatively, certain projections 37 comprise the fourth holes 50. In this example, first projections 370 without holes are arranged alternately with second projections 371 provided with holes in the circumferential direction. In other words, every other projection comprises a fourth hole 50.

[0101] Advantageously, the internal annular surface 43a of the central bore 43 of the cover 40 is intended to be in contact (or in support) by complementarity of shape with an annular bearing surface 31a (cf. [Fig.4]) of the cover 31. The surface of annular bearing surface 31a is defined by a shoulder 51 (visible in [Fig.6]) arranged between the annular bearing surface 36 (projections 37) of the cover 31 and the annular wall 32.

[0102] In Figures 2 and 5, the annular edge 44 of the cover 40 is located adjacent to a first downstream end of the radially inner wall 53a of the diffuser 52. A tiny clearance of a few millimeters (of the order of 5 mm) is provided between the downstream end of the radially inner wall 53a of the diffuser 52 and the annular edge 44 of the cover. The diffuser extends at least partly radially outside a portion of the rear bearing housing, the bearing support, the cover and also the cover. Likewise, the downstream edge 55 of the diffuser 52 (and carried by the radially external wall 53b) of the diffuser 52 extends radially outside the cover 31 and downstream of the annular edge 44. The downstream edge 55 of the diffuser 52 is defined in a plane P which is located downstream of the plane PL. The downstream edge 55 extends radially outside the cover 40 and at a distance from the cover 40.The downstream edge 55 (of a wall portion which extends the radially external wall 53b) covers a portion of the cover 40 so that the cooling air circulates around the cover 40. In this way, the cooling air leaving the diffuser 52 is guided at least in part by a radially external surface of the cover. According to an advantageous, but not limiting, characteristic, the radially internal wall 53a has a radially internal surface 53al which has surface continuity with the radially external surface of the cover 40.

[0103] The fire-fighting system has been described in relation to an annular cavity provided upstream of a free turbine disc but may be arranged in a cavity upstream of a disc of a linked turbine in which a flow of cooling air circulates.

[0104] We will now describe the method for mounting the fire-fighting system 30 in the turbomachine. The method comprises a step of providing the sealing system comprising the diffuser 52 and the cover 40. Beforehand, the method comprises a step of positioning and fixing the bearing support 21 in the turbomachine. The method then comprises a step of positioning the cover 31 around the bearing support 21. The cover 31 is positioned so that the first and second holes 24, 35 are opposite each other. In the present example, the sections of the first and second holes are circular. The first fixing members 27 are then mounted to fix the radial flange 22b of the bearing support 21 and the radial flange 33 of the cover 31.

[0105] The method comprises a step of fixing the rear bearing housing 15 on the bearing support 21. Advantageously, the cover 31, the bearing support 21 and the housing 15 are fixed with the same fixing members 27.

[0106] The method comprises a step of placing the diffuser 52 in the cavity 16. During this step the diffuser 52 is mounted on the rear bearing housing 15. Advantageously, the diffuser 52 is fixed beforehand on the housing 15 before the mounting the housing 15 in the cavity 16. The fixing comprises a weld or a bolted connection.

[0107] The housing 15 is also mounted with the structural part 28.

[0108] The method also comprises a step of placing the cover 40 on the lid 31. The cover 40 is installed so that the bottom wall 41 is in contact with the annular bearing surface 36 of the lid 31. The fourth holes 50 of the cover 40 are also positioned so as to be opposite the holes 49 of the lid 31.

[0109] The method further comprises a step of fixing the cover 40 on the lid 31. During this step, the fixing members 48 such as screws are engaged in the holes of the lid 31 and the cover 40.

[0110] The cooling air leaving the injectors 17 enters each channel 56 and is then guided towards the flow vein V of cooling air and maintaining its speed from the outlet of the injectors 17. The flow speed of the cooling air is greater than any flame speed generated in the cavity by a possible fire.

Claims

Claims

1. Fire-fighting system (30) for a turbomachine (1), such as an aircraft turbojet or turboprop, the turbomachine comprising at least one turbine (3) having a turbine disk (5), an annular cavity (16) arranged upstream of the disk (5) which comprises an internal bore (7) and a cooling device (14) intended to supply the cavity (16) with cooling air via injection means (17), the fire-fighting system (30) comprising means configured so as to, on the one hand, split the annular cavity (16) into a first cavity (16a) and a second cavity (16b) and, on the other hand, maintain a speed of the cooling air at the outlet of the injection means (17) and guide the cooling air in the first cavity (16a) towards the internal bore (7) of the turbine disk (5),characterized in that the means comprise a member for adapting the speed of the cooling air leaving the injection means (17) and an annular cover (40) intended to be mounted on a cover (31) of the turbomachine and to cover first fixing members (27) so that the cooling air leaving the adaptation member (52) is guided at least in part by a radially external surface (40a) of the cover (40).,

2. Fire-fighting system (30) according to claim 1, characterized in that the adaptation member is an annular diffuser (52) extending between an upstream edge (54) and a downstream edge (55) along an axis D and being provided with channels (56) which are formed in the thickness of an annular wall (53) of the diffuser (52), the channels each having an evolving profile and each opening, on the one hand in the upstream edge (54) through a plurality of ports (58), and on the other hand in the downstream edge (55) in an annular opening (59), the ports (58) being arranged around the axis D and being in fluid communication with the injection means (17) of the cooling air device (14).

3. Fire-fighting system (30) according to claim 2, characterized in that the diffuser (52) has a predetermined length (L1) between the upstream edge (54) and the downstream edge (55), the downstream edge (55) being intended to be substantially flush with a radially internal end (12) of a flange (12) mounted upstream of a web (10) of the turbine disk (5) and with a predetermined clearance (J).

4. Fire-fighting system (30) according to one of claims 2 to 3, characterized in that it is produced according to a geometric law defined by the following formula: a < ..... with a being a predetermined angle measured between a first straight line (Dl) passing through a vertex of a light (58), parallel to a respective main axis (E) of the channels (56), and a second straight line (D2) tangent to a flank (57a, 57b) of a channel (56), L being the predetermined length (Ll) of the diffuser (52) and LC being the circumferential width of each light (58).

5. Fire-fighting system (30) according to the preceding claim, characterized in that the cover (40) comprises an annular bottom wall (41) intended to be fixed on an annular bearing surface (36) of the cover (31).

6. Fire-fighting system (30) according to the preceding claim, characterized in that the annular cover (40) comprises an annular skirt (42) extending from the bottom wall (41) along an axis C of the cover (40), the annular skirt (42) comprising a tubular portion (45) and a frustoconical portion (46) connecting the tubular portion to the bottom wall (41) along the axis C, the cover (40) comprising a predetermined radius of curvature (rl, r2) arranged on the one hand, between the tubular portion (45) and the frustoconical portion (46) and on the other hand, between the frustoconical portion (46) and the bottom wall (41).

7. Fire-fighting system (30) according to one of claims 5 and 6, characterized in that the cover (40) has an annular edge (44) arranged adjacent to a downstream end of a radially inner wall (53a) of the diffuser (52), the diffuser (52) comprising a radially outer wall (53b) having a downstream edge (55) which extends radially outside the cover (40) and at a distance from the cover (40).

8. Fire-fighting system (30) according to any one of claims 5 to 7, characterized in that the turbine disc (5) comprises an annular rim (9) extending circumferentially around the longitudinal axis and axially along the longitudinal axis upstream, and in that the fire-fighting system (30) comprises a portion of the cooling air flow vein (V) which is delimited at least in part by a radially internal surface of the annular rim (9) and the radially external surface (40a) of the cowl (40).

9. Fire-fighting system (30) according to claim 6 and 8, characterized in that the ratio between the predetermined radius of curvature (rl, r2) and the height (Dh) of the flow vein (V) of cooling air is greater than 1.

10. Turbomachine (1) in particular of an aircraft, having a longitudinal axis X and comprising a turbine (3) connected by a shaft providing mechanical power, the annular cavity (16) arranged upstream of a first stage of the turbine (3), a cooling device (14) equipped with injection means (17), and a fire-fighting system (30) according to any one of the preceding claims, the fire-fighting system (30) being arranged in the annular cavity (16) and the injection means being in fluid communication with the channels (56) of the cooling system (14).

11. Turbomachine (1) according to the preceding claim, characterized in that it comprises a bearing support (21) for guiding a shaft of the turbine (3) which comprises a first radial flange (22b), and a cover (31) provided with a second radial flange (33) intended to be fixed on the first radial flange (22b) by first fixing members (27), the first fixing members being distributed regularly around the longitudinal axis X and the cover (31) comprising an annular wall (32) intended to pass through the internal bore (7) of the turbine disk.

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