Patent Publication Number: US-2022213808-A1

Title: Module of an aircraft turbine engine

Description:
TECHNICAL SCOPE OF THE INVENTION 
     The present invention relates to an aircraft turbine engine module. 
     TECHNICAL BACKGROUND 
     The prior art includes, in particular, the documents WO-A1-2014/140493, FR-A1-3 017 693, EP-A2-3 069 821, EP-A1-2 940 324, DE-A1-10 2011 108957 and EP-A1-3 444 441. 
     An aircraft turbine engine, for example of an aeroplane or helicopter, comprises an air inlet feeding a gas generator which comprises from upstream to downstream, by reference to the gas flow, at least one compressor, an annular combustion chamber, and at least one turbine. 
     A turbine of a turbine engine comprises one or more expansion stages comprising a bladed nozzle forming a stator, and a bladed wheel forming a rotor. The nozzle is attached to a casing and the wheel comprises a disc with blades on its periphery. The wheel rotates within the casing and it is known to provide a sealing ring around this wheel to limit the passage of gas between the tops of the blades and the casing and thus to ensure that as much of the combustion gas leaving the chamber as possible passes through the wheel to optimise the efficiency of the turbine engine. 
     A sealing ring typically comprises an annular body extending around an axis of revolution and comprising an outer surface and an inner surface which is coated with an annular sealing layer of abradable material on which the tops of the blades can rub in operation. 
     There are currently two sealing ring technologies. The ring according to the first technology comprises a single-piece annular body which is attached by suitable means to a casing. The ring according to the second technology comprises a sectorised annular body, the sectors of the body being attached independently of each other to the casing. 
     In both technologies, the ring is surrounded by an annular wall which includes orifices for the passage of impact cooling air on the outer surface of the body of the ring. This cooling allows for better control of the thermal behaviour of the ring during operation and thus optimises the radial clearances between the body of the ring and the tops of the blades of the wheel during operation. 
     The first technology is interesting from the point of the view of the optimisation of the mass and dimensions, while the second is interesting from a point of view of the ability to optimise the cooling and therefore the adjustment of the clearances with the tops of the blades, as well as the maintenance and easy replacement of each ring sector. 
     The present invention provides an improvement to these existing techniques. In particular, it aims to reduce the number of elements for the construction of a ring and a turbine engine module, so as to limit the number of fastening systems (screws, bolts, flanges, etc.), the risks of leakage between these elements, the mass of the turbine engine, etc. 
     SUMMARY OF THE INVENTION 
     The present invention relates to a module for an aircraft turbine engine, this module comprising: 
     at least one annular casing of an annular combustion chamber, 
     at least one sealing ring for a turbine wheel, and 
     at least one annular bearing support, 
     characterised in that this module is made in one piece. 
     The production of the module in one piece allows to simplify its design and manufacture, this production being preferably carried out by additive manufacturing. It is no longer necessary to provide systems for fixing the elements of the module, which simplifies and lightens the module. 
     The module according to the invention may comprise one or more of the following features, taken alone or in combination with each other:
         the module comprises two annular casings, respectively inner and outer, defining between them an annular recess configured to receive an annular combustion chamber,   the outer casing comprises at its upstream end an annular flange for fixing the module,   the module comprises two sealing rings, upstream and downstream respectively, each of these sealing rings comprising an annular body and an annular wall extending around the annular body and at a radial distance from this body,   the upstream ring, and in particular its annular wall, is connected to said inner annular casing,   said bearing support(s) comprises two annular shrouds, respectively inner and outer, connected together by arms extending substantially radially with respect to an axis of the module,   the outer shroud extends between the upstream and downstream rings and is connected to the downstream ring by an elastically deformable annular part,   the inner shroud extends around an inner ring gear and is connected to this inner ring gear, this ring gear comprising at least one cylindrical surface for mounting a roller bearing,   the downstream ring is connected by an outer ring gear to a junction zone between said inner and outer annular casings,   the outer ring gear comprises at its downstream end an annular flange for fixing the module.       

     The present invention also relates to a method for manufacturing a module as described above, characterised in that it is obtained by additive manufacturing. 
    
    
     
       BRIEF DESCRIPTION OF THE FIGURES 
       Other features and advantages of the invention will become apparent in the course of the detailed description which follows, for the understanding of which reference is made to the appended drawings in which: 
         FIG. 1  is a schematic half-view in axial section of a part of an aircraft turbine engine, 
         FIG. 2  is a very schematic half-view in axial section of a part of an aircraft turbine engine, according to the prior art, 
         FIG. 3  is a very schematic half-view in axial section of a part of an aircraft turbine engine, according to an aspect of the invention, 
         FIG. 4  is a schematic half-view in axial section of a turbine engine module, according to an aspect of the invention, and 
         FIG. 5  is an enlarged schematic view of a detail of  FIG. 4  and shows another aspect of the invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  shows part of an aircraft turbine engine  10  such as a helicopter turbojet engine. 
     The turbine engine  10  comprises from upstream to downstream, with reference to the direction of gas flow (see arrows), an air inlet  12 , at least one compressor  14 , an annular combustion chamber  16 , and at least one turbine  18 . 
     The air entering the engine through the air inlet  12  is compressed in the compressor  14 , which is a centrifugal compressor. The compressed air exits radially outwards and feeds the combustion chamber  16  via an annular assembly forming a rectifier  20  and a diffuser  22 . 
     The combustion chamber  16  comprises two annular walls, respectively inner  16   a  and outer  16   b , which extend around each other and which are themselves arranged inside an outer casing  24  of the combustion chamber  16 . 
     This casing  24  comprises at its upstream end an annular flange  24   a  for attachment to annular flanges of the rectifier-diffuser assembly  20 - 22  as well as a casing  25  of the compressor  14  and the air inlet  12 . 
     The compressed air is mixed with fuel and burned in the combustion chamber  16 , generating combustion gases which are then injected into the turbines  18 . 
     A high-pressure turbine stage  18   a  is located just downstream of the outlet of the combustion chamber  16  and comprises a stator nozzle  28  and a rotor wheel  26 . A low-pressure turbine stage  18   b  is located downstream of the stage  18   a  and also comprises a nozzle  30  and a rotor wheel  26 . 
     A turbine nozzle comprises an annular row of fixed blades for straightening the gas stream, and a turbine wheel comprises an annular row of blades carried by a rotor disc. 
     The casing  24  further comprises at its downstream end an annular flange  24   b  for attachment to support flanges for sealing rings  36 ,  38 . 
     A casing  32  extends within the wall  16   a  and carries at its upstream end the sealing ring  36  which extends around the wheel  26  of the stage  18   a , and at its downstream end a flange  32   a  for attachment to the flange  24   b . A ring gear  34  carries the sealing ring  38  which extends around the wheel  26  of the stage  18   b . This ring gear  34  comprises a flange  34   a  for attachment to the flanges  32   a ,  24   b.    
     Each sealing ring  36 ,  38  comprises an inner cylindrical surface which is coated with an annular abradable layer configured to rub against the tops of the blades of the wheel  26  to minimise gas leakage in that zone. However, the abradable layer is optional. It may also be present on one of the rings, for example the ring  36 , and be absent from the other ring  38 . This abradable layer may have a thermal protection function. 
     The wheels  26  are connected to each other by a shaft  40  which is further connected to the impeller of the centrifugal compressor  14 . The shaft  40  is guided in rotation by rolling bearings  41  which are carried by an annular support  42  interposed between the two stages  18   a ,  18   b.    
     The bearing support  42  comprises two annular shrouds, respectively inner  42   a  and outer  42   b , connected together by an annular row of arms  44  extending substantially radially with respect to the axis A of rotation of the shaft  34 . The arms  44  are tubular and may be used for the passage of utilities  46  such as fluid lines or electrical cables. 
     The bearing support  42  is mounted inside the casing  32  and carries a bearing housing which comprises a ring gear  48  for supporting the outer bushings  41   a  of the bearings  41 . The bearings  41  are here two in number, an upstream roller bearing and a ball bearing, the inner bushings  41   b  of which are mounted directly on the shaft  40 . 
       FIG. 2  shows very schematically the current state of the art in the manufacture and assembly of several elements visible in  FIG. 1 . 
     Firstly, the sealing rings  36 ,  38  are made independently of each other and of the other surrounding parts. They are attached by flanges or hooks to casings  32 ,  34  which are themselves attached by flanges to the outer casing  24  of the chamber. The bearing support  42  is also attached by a flange  42   c  to this casing  24 . 
       FIG. 3  illustrates an aspect of the invention which consists in providing a module  50  which is single-piece, i.e. formed in one piece preferably by additive manufacturing, and including several of the aforementioned elements. 
     In the example shown, the module  50  comprises the casing  24 , the sealing rings  36 ,  38  and at least part of the bearing support  42 . 
       FIG. 4  represents a more concrete embodiment of this module  50  and  FIG. 5  is a detail view of  FIG. 4  and more specifically illustrates another aspect of the invention relating to the sealing rings  36 ,  38 . 
     Each sealing ring  36 ,  38  advantageously comprises an annular body  51  extending about the axis A and comprising an outer surface  51   a  and an inner surface  51   b  which is coated with an annular layer  53  of an abradable material. 
     The ring  36 ,  38  further comprises an annular wall  52  extending around the annular body  51  and at a radial distance from this body. This annular wall  52  comprises orifices  54  for the passage of impact cooling air over the outer surface  51   a . As can be clearly seen in  FIG. 5 , the body  51  and the wall  52  are formed in one piece and these elements and the orifices  54  are advantageously obtained by additive manufacturing. 
     The body  51  and the wall  52  define between them an annular space  56  which is closed at a downstream end and open at an upstream end. This space  56  has a radial thickness E 1  less than or equal to that E 2  of the body  51  and/or less than or equal to that E 3  of the wall  52 . 
     The following description applies more specifically to the ring  36  illustrated in  FIG. 5 . 
     The wall  52  comprises at its upstream end an annular groove open radially inwards and in which is mounted an annular sealing member  58  which is configured to cooperate with the nozzle  28  of the upstream stage  18   a  mentioned above. 
     The wall  52  extends downstream and is connected to or forms the radially inner annular casing  32  of the combustion chamber  16 . 
     The body  51  extends downstream and is connected to or forms the outer shroud  42   b  of the bearing support  42 . 
     The wall  52  comprises at least one annular row of air passage orifices  54  evenly spaced around the axis A and oriented in directions substantially radial to that axis. Each of these orifices  54  includes a constriction  54   a  at its radially inner end to accelerate the air stream through the orifice and improve the impact cooling of the body  51  of the ring  36 . 
     The rings  36 ,  38  and the assembly of the module  50  may be made of a metal alloy. The layers  53  are advantageously made of ceramic. 
     Another aspect of the invention relates to a method for manufacturing a ring  36 ,  38  as well as the module  50  by additive manufacturing. 
     In the embodiment of the module shown in  FIG. 4 , the latter comprises:
         the two inner  32  and outer  24  casings,   the two sealing rings  36 ,  38 , and   the bearing support  42 .       

     The outer casing  24  comprises at its upstream end the annular flange  24   a  for fixing the module, for example to the aforementioned flanges of the casing  24  and of the diffuser-rectifier assembly  20 - 22  of  FIG. 2 . 
     As mentioned above, the upstream ring  36 , and in particular its annular wall  52 , is connected to the inner casing  32 . 
     The outer shroud  42   b  of the bearing support  42  extends between the rings  36 ,  38  and is connected to the downstream ring  38  by an elastically deformable annular part  60 . This part  60  is relatively flexible and is capable of elastic deformation in the axial and/or radial direction to allow for differential thermal expansions during operation in particular. This part  60 , also referred to as a pin, can be used to support the outer shroud  42   b  which is then not supported by the arms but by this flexible part. The inner shroud  42   a  can be supported in the same way by means of another flexible part. 
     The inner shroud  42   a  of the bearing support  42  extends around the inner ring gear  48  and is connected to this inner ring gear which comprises cylindrical surfaces  48   d  for mounting the outer bushings  41   a  of the bearings  41 . 
     The downstream ring  38  is connected by an outer ring gear  34  to a junction zone between the casings  24 ,  32 . This ring gear  34  comprises at its downstream end the annular flange  34   b  for fixing the module  50 . 
     The single-piece construction of each ring  36 ,  38  allows to simplify its design and manufacture and to integrate all the functions of a ring of the previous technique, including those of retention of the blades in the event of breakage, of aerothermal function, etc. 
     The ring is cooled by the impact of the air passing through the orifices  54  of the wall  52  during operation. The shape of these orifices  54  and the distance between them and the body  51  (radial thickness E 1 ) are determined to optimise the cooling of the ring and therefore the performances. 
     The single-piece module  50  allows a significant reduction in mass (in the order of 25 to 30% in the example shown) compared to the prior art. 
     The additive manufacturing allows these manufacturing and optimisation objectives to be achieved.