Patent Publication Number: US-2022213800-A1

Title: Sealing ring for a wheel of a turbomachine turbine

Description:
TECHNICAL SCOPE OF THE INVENTION 
     The present invention relates to a sealing ring for a wheel of a turbomachine turbine. 
     TECHNICAL BACKGROUND 
     The technical background comprises, in particular, the documents GB 1 484 288 A and GB 2 125 111 A. 
     An aircraft turbomachine, 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 turbomachine turbine comprises one or more expansion stages comprising a bladed turbine stator forming a stator, and a bladed wheel forming a rotor. The turbine stator is fastened 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 turbomachine. 
     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 first technology comprises a one-piece annular body which is fastened by suitable means to a casing. The second technology ring comprises a sectorised annular body, the sectors of the body being fastened independently of each other to the casing. 
     In both technologies, the ring is surrounded by an annular wall which comprises openings for the passage of air for impact cooling 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 view of the mass and space optimisation, while the second is interesting from the 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 turbomachine 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 turbomachine, etc. 
     SUMMARY OF THE INVENTION 
     The present invention relates to a sealing ring for a wheel of an aircraft turbomachine turbine, this ring comprising an annular body extending around an axis of revolution and comprising an outer surface and an inner surface which is coated with an annular layer of an abradable material, the ring further comprising an annular wall extending around the annular body and at a radial distance from this body, this annular wall comprising openings for the passage of air for impact cooling on said outer surface, characterized in that the body and the wall are integrally formed. 
     The production of the body and the wall of the ring 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 fastening the body to the wall, which simplifies and lightens the ring. 
     The ring according to the invention may comprise one or more of the following features, taken alone or in combination with each other:
         the body and the wall define between them an annular space which is closed at a downstream end and open at an upstream end by reference to the flow of gases through the ring in operation.   said space has a radial thickness less than or equal to that of the body and/or less than or equal to that of the wall.   said wall comprises at its upstream end an annular groove open radially inwards and in which an annular sealing member is mounted.   said wall extends downstream to form a radially inner annular casing of a combustion chamber.   said body extends downstream to form a radially outer annular shroud of an annular bearing support.   said wall comprises at least one annular row of air passage openings evenly spaced around said axis and oriented in substantially radial directions with respect to this axis.   each of the air passage openings comprises a constriction at its radially inner end.   the ring is made of a metal alloy and coated at least in part with a ceramic thermal barrier; the thermal barrier can be constituted by the aforementioned abradable layer; this abradable layer then has a dual function of abradability (improving the rotor/stator behaviour during contacts) and of thermal barrier (controlling the temperature of the ring against the hot gases of the vein).       

     The present invention also relates to a method of manufacturing a ring as described above, characterised in that the body, the wall and the openings of this wall are obtained by additive manufacturing. 
    
    
     
       BRIEF DESCRIPTION OF THE FIGURES 
       Further features and advantages of the invention will become apparent from the following detailed description, for the understanding of which reference is made to the annexed drawings in which: 
         FIG. 1  is a schematic half-view in axial section of a portion of an aircraft turbomachine, 
         FIG. 2  is a very schematic half-view in axial section of a portion of an aircraft turbomachine, according to the prior art, 
         FIG. 3  is a very schematic half-view in axial section of a portion of an aircraft turbomachine, according to an aspect of the invention, 
         FIG. 4  is a schematic axial sectional half-view of a turbomachine 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 a part of an aircraft turbomachine  10  such as a helicopter turbojet engine. 
     The turbomachine  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 fastening 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 turbine stator  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 turbine stator  30  and a rotor wheel  26 . 
     A turbine stator comprises an annular row of fixed straightening blades of 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 fastening to sealing ring support flanges  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 fastening 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 fastening to the flanges  32   a ,  24   b.    
     Each sealing ring  36 ,  38  comprises an inner cylindrical surface which is coated with an abradable annular layer configured to wear by friction with the tops of the blades of the wheel  26  to minimise the gas leakage in that area as much as possible. This abradable layer advantageously has a thermal barrier function. Alternatively, the rings could not comprise such a layer or only one of the rings  36 ,  38  (e.g. the ring  36 ) could comprise such a layer. 
     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 easements  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 rings  41   a  of the bearings  41 . The bearings  41  are here two in number, an upstream roller bearing and a ball bearing, the inner rings  41   b  of which are mounted directly on the shaft  34 . 
       FIG. 2  shows very schematically the current state of the art in manufacturing and assembling of several elements visible in  FIG. 1 . 
     Firstly, the sealing rings  36 ,  38  are made independently of each other and of the other surrounding pieces. They are fastened by flanges or hooks to the casings  32 ,  34  which are themselves fastened by flanges to the outer casing  24  of the chamber. The bearing support  42  is also fastened 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 one-piece, i.e. integrally formed, 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 a portion 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 around 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 and radially spaced from the annular body  51 . This annular wall  52  comprises openings  54  for the passage of air for impact cooling on the outer surface  51   a . As can be clearly seen in  FIG. 5 , the body  51  and the wall  52  are integrally formed and these elements and the openings  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 of E 2  of the body  51  and/or less than or equal to that of 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 which is open radially inwards and in which an annular sealing member  58  for the upstream stage  18   a  referred to above is mounted. 
     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 openings  54  evenly spaced around the axis A and oriented in directions substantially radial with respect to that axis. Each of these openings  54  comprises a constriction  54   a  at its radially inner end to accelerate the stream of air flowing through the opening and improve the cooling by impact 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 fastening 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 portion  60 . This portion  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 portion  60 , also known 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 portion. The inner shroud  42   a  can be supported in the same way by means of another flexible portion. 
     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 rings  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 fastening the module  50 . 
     The one-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, aerothermal function, etc. 
     The ring is cooled by the impact of the air flowing through the openings  54  of the wall  52  during operation. The shape of these openings  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 performance. 
     The one-piece module  50  can significantly reduce its mass (in the order of 25 to 30% in the example shown) compared to the previous technique. 
     Additive manufacturing allows these manufacturing and optimisation objectives to be achieved.