Patent Document

FIELD OF THE INVENTION 
       [0001]    The present invention relates to a labyrinth seal for forming a seal between a first and a second component which rotate relative to each other. 
       BACKGROUND OF THE INVENTION 
       [0002]    With reference to  FIG. 1 , a ducted fan gas turbine engine generally indicated at  10  has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake  11 , a propulsive fan  12 , an intermediate pressure compressor  13 , a high-pressure compressor  14 , combustion equipment  15 , a high-pressure turbine  16 , and intermediate pressure turbine  17 , a low-pressure turbine  18  and a core engine exhaust nozzle  19 . A nacelle  21  generally surrounds the engine  10  and defines the intake  11 , a bypass duct  22  and a bypass exhaust nozzle  23 . 
         [0003]    The gas turbine engine  10  works in a conventional manner so that air entering the intake  11  is accelerated by the fan  12  to produce two air flows: a first air flow A into the intermediate pressure compressor  14  and a second air flow B which passes through the bypass duct  22  to provide propulsive thrust. The intermediate pressure compressor  13  compresses the air flow A directed into it before delivering that air to the high pressure compressor  14  where further compression takes place. 
         [0004]    The compressed air exhausted from the high-pressure compressor  14  is directed into the combustion equipment  15  where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines  16 ,  17 ,  18  before being exhausted through the nozzle  19  to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors  14 ,  13  and the fan  12  by suitable interconnecting shafts. 
         [0005]    Labyrinth seals are used throughout a gas turbine engine, and are designed to seal two components together whilst permitting a flow of air through the sealed boundary. An example of such a seal is between a casing component of the combustion equipment  15  and a cover plate protecting components of the high pressure turbine  16 . The operating temperature of the high pressure turbine components needs to be kept at a safe level to maintain component integrity. This is achieved using a labyrinth seal to permit a purging flow of cooling air from the high pressure compressor  14  to the high pressure turbine  15  components and thereby preventing ingestion of hot working gas. 
         [0006]    Labyrinth seals have two abutting surfaces; one surface having an abradable lining and the other having a series of fins. The fins provide resistance to air flow by forcing the air to traverse around the fins along a labyrinthal path. This resistance to air flow minimises performance penalties from air leakage. 
         [0007]    During operation, thermal and mechanical movements of the gas turbine engine structure cause relative movement of the sealed components. Thus, the distance between the two abutting surfaces of the labyrinth seal changes throughout operation. This can result in periods during operation where the lining and fins are sufficiently close that the air flow through the seal is restricted to an unacceptable level. In the case where the seal has to allow a certain level of purging air flow through the seal, restriction of the flow through the seal can lead to hot gas ingestion causing damage or failure of engine components. 
         [0008]    A conventional solution to this problem is to position the lining and fins sufficiently apart so they never run close enough during operation to over-restrict the air flow through the seal. However, this results in periods of operation where the distance between the lining and fins is larger than necessary, and has the effect of reducing performance efficiency of the engine. 
       SUMMARY OF THE INVENTION 
       [0009]    Accordingly, an aim of the present invention is to provide a labyrinth seal in which air flow through the seal is better regulated. 
         [0010]    In a first aspect, the present invention provides a labyrinth seal for forming a seal between a first and a second component which rotate relative to each other, the seal having: an abradable lining mounted to the first component, and a plurality of fins projecting from the second component and arranged in abutment with the abradable lining to form a labyrinthal path for a flow of seal air through the seal; wherein the seal further has a bypass passage which extends through the abradable lining to allow a portion of the seal air to flow through the seal and bypass the labyrinthal path. 
         [0011]    Advantageously, the labyrinth seal of the present invention allows a metered flow of air independent of the relative positions of the first and second components. This means that the flow area of the bypass passage can be unaffected by the thermal and mechanical relative movements of the first and second components. Thus the bypass passage can ensure sufficient air flow through the labyrinth seal throughout operation. 
         [0012]    Furthermore, because of the flow of air bypassing the labyrinthal path, precise regulation of the amount of air flowing though the labyrinthal path can be less critical. Accordingly, the fins and abradable material of the labyrinth seal can be run in a position that provides greater engine performance efficiency. 
         [0013]    The labyrinth seal may have any one or, to the extent that they are compatible, any combination of the following optional features. 
         [0014]    Typically, one of the components is a static component. An example of this type of seal is a seal with one static component and one rotating component. 
         [0015]    Typically, the abradable lining is mounted to the static component. In this case, the plurality of fins project from the rotating component and abut the abradable lining as they rotate. 
         [0016]    Preferably, the abradable lining is a honeycomb abradable lining. Typically, the cells of the honeycomb abradable lining extend across the thickness of the lining. Suitably, the cross section of the cells may be a regular polygon, such as a hexagon. A honeycomb abradable lining is advantageous because it can be lightweight. Alternatively, however, the abradable lining can be formed of e.g. sintered metal. 
         [0017]    Conveniently, the abradable lining is stepped to further restrict the flow of air through the labyrinthal path, each fin abutting the abradable lining at a respective step. Advantageously, this allows a tighter seal to be formed, and thus limits performance losses from air leakage. 
         [0018]    Preferably, the bypass passage has a sleeve for directing air flowing through the bypass passage. The sleeve can provide a direct channel for the bypass air, from the entrance to the exit of the bypass passage. When a honeycomb abradable lining is used, this helps to avoid bypass air escaping from the passage into adjacent honeycomb cells. The sleeve also allows the internal diameter of the passage to be easily selected to provide an aerodynamically efficient length over diameter ratio. 
         [0019]    The entrance to the bypass passage can form a bell mouth. Such an entrance can improve the efficiency of air intake to the bypass passage, reducing aerodynamic losses and increasing the flow of air at the exit of the bypass passage. 
         [0020]    The bypass passage may taper along its length. The taper can be used to control the velocity or pressure of the bypass air. Thus the taper can be changed to suit the required needs of the air flow system. A taper maybe provided such that it increases the diameter of the bypass passage at the passage exit, compared to the passage entrance, which can have the effect of decreasing the velocity at the exit compared to the entrance. Conversely, a taper maybe provided to increase the diameter of the bypass passage at the passage entrance compared to the passage exit. This can have the effect of increasing the velocity and decreasing the pressure of the air at the exit compared to the entrance. 
         [0021]    Conveniently, the bypass passage can be angled relative to the axis of rotation of the first and second components to impart swirl to the air exiting the bypass passage. Advantageously, the swirl imparted to the exiting air flow can result in reduced windage losses. This in turn can lead to reduced heat pick up and increased efficiency. The bypass passage can be angled to direct the air flow to a specific location for localised cooling. 
         [0022]    Preferably, the bypass passage is formed by electromachining. Suitably, the electromachining may be electro chemical or electro discharge machining. 
         [0023]    Preferably, the labyrinth seal has a plurality of bypass passages. This allows an increased and/or distributed flow of bypass air through the seal, compared to only a single passage. For example, the labyrinth seal may have a plurality of bypass passages spaced circumferentially about the axis of rotation of the first and second components. This arrangement can help to reduce the risk of localised high temperatures. 
         [0024]    Typically, first and second components are components of a gas turbine engine. For example, the first component may be a high pressure turbine static component such as a combustor rear inner case, and the second component may be a high pressure turbine rotating component such as a disc rim cover plate. The seal between these components is important because it controls a purging air flow from the high pressure compressor to critical components of the high pressure turbine. The structure of the labyrinth seal allows cooling air to be directed to these critical components whilst maintaining a close contact, and therefore tight seal, between the fins and the abradable material. 
         [0025]    The bypass passage may extend through the abradable lining from a downstream end of the abradable lining to an upstream end of the abradable lining. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0026]    Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which: 
           [0027]      FIG. 1  shows schematically a longitudinal section through a ducted fan gas turbine engine; 
           [0028]      FIG. 2  shows schematically a longitudinal section of the region between the combustion equipment and the high pressure turbine of a gas turbine engine, a labyrinth seal being located between a combustor rear inner case and a rim cover plate; 
           [0029]      FIG. 3  shows schematically a closer view of the labyrinth seal of  FIG. 2 ; 
           [0030]      FIG. 4  shows a section of the abradable lining of the labyrinth seal of  FIGS. 2 and 3  viewed from the exit side of the seal; 
           [0031]      FIG. 5  shows the same section of abradable lining as  FIG. 4  viewed from a position radially inside the seal; and 
           [0032]      FIG. 6  shows the same section of abradable lining as  FIGS. 4 and 5  in a perspective view from the exit side of the seal. 
       
    
    
     DETAILED DESCRIPTION 
       [0033]      FIG. 2  shows schematically a longitudinal section of the region between the combustion equipment and the high pressure turbine of a gas turbine engine, a labyrinth seal  24  being located between a combustor rear inner case  34  and a rim cover plate  28 .  FIG. 3  shows schematically a closer view of the labyrinth seal  24  of  FIG. 2 . The rim cover plate  28  is positioned between the combustor rear inner casing  34  and a high pressure turbine disc  30  to protect the high pressure turbine disc  30 , to which high pressure turbine blades  32  are attached. The rim cover plate  28  rotates about the axis of the gas turbine engine. The combustor rear inner case  34  is static, and has high pressure nozzle guide vanes  31  extending therefrom. 
         [0034]    In operation, cooling combustion feed air from the high pressure compressor enters the combustion equipment of the engine at specified locations. In particular, air flow C (dashed arrowed line) from the high pressure compressor enters the combustor rear inner case  34 . This air flow C passes through the labyrinth seal  24  to regulate the temperature of the rim of the high pressure turbine disc  30  by purging the air surrounding the rim and preventing ingestion of hot working gas. 
         [0035]    The labyrinth seal  24  has an abradable honeycomb lining  38  which is attached to the combustor rear inner case  34 . The sealing surface of the abradable lining  38  is formed as a series of steps  40 . The honeycomb cells have metal foil walls and are aligned with their length direction extending across the thickness of the lining. The skilled person is familiar with the use of honeycomb abradable linings in labyrinth seal applications. 
         [0036]    Fins  46  project from the rim cover plate  28  and abut the abradable lining  38 . The arrangement of the steps  40  and the fins  46  is such that each fin  46  abuts a respective step  40  to form a labyrinthal path  48  for the flow of air between the lining  38  and the fins  46 . The labyrinthal path  48  produces resistance to the flow of air D through the seal. In operation, the abutment of the fins  46  to the steps  40  is such that the fins  46  rub into the steps  40 . The comparatively soft nature of the abradable material means that this rubbing removes material primarily from the abradable lining  38  rather than the fins, creating a tight seal without causing damage to the gas turbine components. 
         [0037]    A plurality of circumferentially spaced bypass passages  36  extend through the abradable lining  38 . The entrances  42  to the bypass passages  36  are on the combustion equipment side of the labyrinth seal  24 , and the exits  44  are on the high pressure turbine side of the labyrinth seal  24 . The passages  36  are separate from and do not interfere with the labyrinthal path  48 . The bypass passages  36  provide a route for a further, metered, independent flow of air E through the labyrinth seal  24 . The bypass passages preferably extend through the abradable lining from the entrance  42  at a downstream end  42   a  of the abradable lining to an exit  44  on an upstream end  44   a  of the abradable lining to bypass the seal fins. 
         [0038]    Advantageously, in operation the majority of the air flow through the labyrinth seal  24  can be through the bypass passages  36 . Thus the air flow E can provide most of the air necessary to regulate the temperature of the high pressure turbine disc  30 . As there is therefore a reduced requirement for the air flow D through the labyrinthal path  48 , the fins  46  and the steps  40  of the abradable lining  38  can operate in close abutment, thereby improving the efficiency of the engine by reducing air leakage through the seal and maximising feed pressure to the blade  32 . 
         [0039]    The abradable lining  38  extends circumferentially around the combustor rear inner case  34 , and  FIG. 4  shows a section of the abradable lining  38  viewed along the axial direction from the exit side of the seal  24 . The exits  44  from three of the circumferentially spaced bypass passages  36  are visible.  FIG. 5  shows the same section of the abradable lining  38  but viewed from a position radially inside the lining.  FIG. 6  shows the same section of the abradable lining  38  in a perspective view from the exit side of the seal  24 . As best shown in  FIG. 5  the bypass passages  36  are angled relative to the axis of rotation to impart swirl on the air flow E as it exits the passages  36 , the swirl being in the same direction as the direction of rotation of the high pressure turbine disc  30 . The swirl has the effect of reducing windage losses, which in turn reduces heat pickup and increases efficiency. The angling also allows the flow, where necessary, to be directed to specific regions of the high pressure turbine disc  30  or the high pressure turbine blade  32 . This can be significant if there is a risk of localised overheating. 
         [0040]    The bypass passages  36  are formed in the honeycomb lining  38  before assembly to the gas turbine engine  10 , using electro chemical or electro discharge machining. 
         [0041]    The bypass passages  36  are lined with respective sleeves  50 , although only one such sleeve is shown in  FIGS. 4 ,  5  and  6 . The sleeve  50  extends from the entrance  42  to the exit  44  of the bypass passage, and can be formed as a smooth cylindrical metal tube. The outside diameter of the tube is dimensioned to fit securely in the bypass passage  36 . The inner diameter of the tube is dimensioned to provide a length to diameter ratio which best improves aerodynamic efficiency. Advantageously, the sleeve prevents air escaping from the passage into the cells  52  of the honeycomb. 
         [0042]    The sleeve  50  of the bypass passage  36  can be inserted into the bypass passage, and then affixed using brazing or welding. If brazing is used, the sleeve can be inserted and brazed to the abradable lining  38  at the same time as the abradable lining  38  is brazed to the combustor rear inner case  34  of the gas turbine engine  10 . If welding is used, the abradable lining can first be brazed to the combustor rear inner case  34 , and then the sleeve  50  can be inserted into the bypass passage  36  and welded to the abradable lining  38 . 
         [0043]    While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.

Technology Category: f