Patent Publication Number: US-11021957-B2

Title: Gas turbine engine rotor disc retention assembly

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the US National Stage of International Application No. PCT/EP2018/063206 filed 18 May 2018, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP17173117 filed 26 May 2017. All of the applications are incorporated by reference herein in their entirety. 
     FIELD OF INVENTION 
     The present invention relates to gas turbine engines, and more particularly to rotor discs of gas turbine engines. 
     BACKGROUND OF INVENTION 
     Turbine blades in various modern gas turbine engines are arranged on rotor discs. A plurality of the blades is arranged circumferentially on the rotor disc. The rotor disc has a central hole, i.e. a central bore through which a tension bolt passes when the rotor disc along with the circumferentially assembled turbine blades is positioned within the gas turbine engine. A shaft is connected to the rotor disc by generally using a Hirth joint or Hirth coupling. When the gas turbine engine is operated, in such a rotor disc with the central hole and the Hirth coupling an unsymmetrical stress distribution is produced with peak stress around the central bore of the hub at the side opposite to where the bolt load is applied. The aforementioned rotor disc and its arrangement within a gas turbine are explained hereinafter in further details with respect to  FIGS. 2 and 3 . 
       FIG. 2  schematically depicts a conventionally known rotor disc  99 , and  FIG. 3  schematically depicts the conventionally known rotor disc  99  when positioned within a gas turbine. The conventionally known rotor disc  99 , hereinafter also referred to as the rotor disc  99 , has a hub  60 , a web  70  and a blade retention arrangement  80 . The hub  60  is region or part of the rotor disc  99  that surrounds a central hole  11  or central bore  11 . The central bore  11  is arranged around a rotational axis  15  of the rotor disc  99  when the rotor disc  99  is positioned inside the gas turbine, as depicted in  FIG. 3 . From the hub  60  extends radially outwards the web  70  which is section of the rotor disc  99  that connects the hub  60  to the blade retention arrangement  80 . The blade retention arrangement  80  usually comprises slots (not shown in  FIGS. 2 and 3 ) into which roots (not shown in  FIGS. 2 and 3 ) of a plurality of turbine blades (not shown in  FIGS. 2 and 3 ) are arranged or fixed. Thus the turbine blades are circumferentially arranged on the rotor disc  99  and extend radially outwards from the rotor disc  99 , and particularly from the blade retention arrangement  80  of the rotor disc  99 . 
     As shown in  FIG. 3 , a tension bolt  4  of the gas turbine passes through the central bore  11  and is physically contacted at a first axial side  91  of the rotor disc  99 . The tension bolt  4  bears the load of the rotor disc  99  along with the turbine blades arranged on the rotor disc  99  when the rotor disc  99  along with the turbine blades are rotated while operating the gas turbine. On a second axial side  92  of the rotor disc  99 , the rotor disc  99  is contacted or coupled with a drive shaft  3  of the gas turbine via generally Hirth coupling  2 . The location of the Hirth coupling  2  is also depicted in  FIG. 2  although  FIG. 2  does not schematically depict the Hirth coupling  2  in its entirety with the drive shaft  3 . The drive shaft  3  rotationally couples the gas turbine to a downstream load for example a generator (not shown). 
     In such conventionally known rotor disc  99  having the central bore  11 , which is subject to offset loading, the rotor disc  99  is subjected to dishing, and a high stress is created in the hub  60  around the central bore  11  of the rotor disc  99 , generally with peak stress around an edge  93  of the hub  60  around the central bore  11  at the side opposite to where the bolt load is applied i.e. at the second side  92  in the examples of  FIGS. 2 and 3 .  FIG. 10  schematically depicts a stress location  65  in the hub  60  of the conventionally known rotor disc  99  when functioning within the gas turbine and connected to the drive shaft  3  and the tension bolt  4  as aforementioned with respect to  FIG. 3 . Due to the bolt load transmission, a peak in the stress occurs at the edge  93  of the hub  60 , which is not desirable due to the high stress concentration factor. The peak stress concentration at the edge  93  of the hub  60  in the conventionally known rotor disc  99  increases chances of failure of the rotor disc  99  and reduces life of the rotor disc  99 . Furthermore, dishing of the rotor disc  99  may be undesirable due to the effect it may have on the turbine blade position during turbine blade rotation. Therefore, a technique is desired to reduce concentration of the aforementioned stress at the edge  93  of the hub  60  which occurs in the conventionally known rotor disc  99 . 
     U.S. Pat. No. 4,844,694 discloses a fastening spindle and a method of attaching the rotor elements together utilizing the spindle. The system permits the visual inspection of the rotor assembly and to determine if it is properly tightened without the need for any additional post assembly inspection. The system and method is used for fastening a plurality of rotor elements together. 
     SUMMARY OF INVENTION 
     Thus the object of the present invention is to provide a technique for reducing stress concentration in a gas turbine rotor disc. It is desirable that the present technique provides reduction in stress concentration at the edge, opposite to the side of the rotor disc where the tension bolt load is applied, of the hub of the rotor disc. 
     The above objects are achieved by a gas turbine engine rotor disc of the present technique, a rotor disc assembly of the present technique and a gas turbine engine of the present technique. Advantageous embodiments of the present technique are provided in dependent claims. Features of the independent claims can be combined with features of dependent claims, and features of dependent claims can be combined together. 
     In the present technique a gas turbine engine rotor disc for a gas turbine engine is presented. The rotor disc includes a hub, a web, a blade retention arrangement, a rotational axis, a first axial side and a second axial side. The hub includes a central bore around the rotational axis. The web is integrally formed with the hub. The web extends radially outwards from the hub to the blade retention arrangement. The blade retention arrangement has a centre of mass. A radial plane passes through the centre of mass. The radial plane is perpendicular to the rotational axis. The first axial side is adapted for engaging a tension bolt of the gas turbine engine. The radial plane intersects the hub defining a first axial side portion and a second axial side portion. The first axial side portion is towards the first axial side and the second axial side portion is towards the second axial side. The second axial side portion has an axial extent which is between 10% and 30% greater than an axial extent of the first axial side portion. 
     The aforementioned design of the rotor disc, i.e. wherein the second axial side portion is axially longer than the first axial side portion by 10% to 30%, optimizes the stress profile within the hub and thereby reduces stress concentration at the edge of the hub. The added material, due to greater axial length of the second side of the hub, in the region of the high edge stress, offsets the peak stress and reduces the dishing. Thus, the aforementioned rotor disc experiences reduction in dishing of the rotor disc. The rotor disc of the present technique is particularly beneficial for use in turbine designs with thin discs that are prone to dishing, and that have a centre bolt or tension bolt design that can cause dishing of the end disc, that is the disc that is directly physically contacted with the centre bolt or the tension bolt, due to the staggered load transmission of the bolt-load. 
     In an embodiment of the gas turbine rotor disc, the second axial side portion has the axial extent which is between 20% and 25% greater than the axial extent of the first axial side portion. 
     In an embodiment of the gas turbine engine rotor disc, to determine the axial extents for the gas turbine rotor disc, measurements of the first axial extent and the second axial extent are limited to a region of the hub that has geometric similarity at the first axial side and the second axial side. In another embodiment of the gas turbine engine rotor disc, the region of the hub is free from an integrally formed connection projecting out from the hub and contacting one or more components of the gas turbine engine. In another embodiment of the gas turbine engine rotor disc, measurement of the first axial extent and the second axial extent are defined at an axial surface of the hub. The aforementioned embodiments provide simple ways of fixing or deciding the first and the second axial extents. 
     In another embodiment of the gas turbine engine rotor disc, the hub at the first axial side includes a chamfered recess adapted for engaging the tension bolt of the gas turbine engine. This provides a simple construct for positioning and integrating the rotor disc of the present technique into the gas turbine engine and in contact with the tension bolt of the gas turbine engine. 
     In another embodiment of the gas turbine engine rotor disc, the second axial side is adapted for engaging with a drive shaft of the gas turbine engine, for example via a Hirth coupling. This provides a simple construct for positioning and integrating the rotor disc of the present technique into the gas turbine engine and in contact with the drive shaft of the gas turbine engine. 
     In another aspect of the present technique, a gas turbine rotor disc assembly is presented. The gas turbine rotor disc assembly includes a gas turbine rotor disc and a plurality of turbine blades. The gas turbine rotor disc is according to the aforementioned aspect of the present technique. The turbine blades are arranged circumferentially at the blade retention arrangement of the rotor disc. The turbine blades extend radially outwards from the blade retention arrangement of the rotor disc. In the gas turbine rotor disc assembly of the present technique, the stress profile within the hub of the rotor disc is optimized and thereby stress concentration at the edge of the hub is reduced or obviated. The rotor disc experiences reduction in dishing. Due to the present rotor disc, the gas turbine of the present technique may be constructed with thinner than conventional rotor discs. Furthermore, the location of the blades of the gas turbine rotor disc assembly is free from or subjected to reduced effect from consequences of dishing of the rotor disc. 
     In yet another aspect of the present technique, a gas turbine engine is presented. The gas turbine engine includes a gas turbine rotor disc assembly. The gas turbine rotor disc assembly is according to the aforementioned aspect of the present technique. In the gas turbine engine of the present technique, the stress profile within the hub of the rotor disc is optimized and thereby stress concentration at the edge of the hub is reduced or obviated. The rotor disc experiences reduction in dishing. Due to the present rotor disc, the gas turbine of the present technique may be constructed with thinner than conventional rotor discs. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The above mentioned attributes and other features and advantages of the present technique and the manner of attaining them will become more apparent and the present technique itself will be better understood by reference to the following description of embodiments of the present technique taken in conjunction with the accompanying drawings, wherein: 
         FIG. 1  shows part of a gas turbine engine in a sectional view and in which a gas turbine rotor disc of the present technique is incorporated or the gas turbine rotor disc assembly of the present technique is incorporated; 
         FIG. 2  schematically illustrates a conventionally known rotor disc; 
         FIG. 3  schematically illustrates the conventionally known rotor disc as arranged within the gas turbine; 
         FIG. 4  schematically illustrates an exemplary embodiment of the gas turbine rotor disc of the present technique; 
         FIG. 5  schematically illustrates the gas turbine rotor disc of the present technique as arranged within the gas turbine; 
         FIG. 6  schematically illustrates the gas turbine rotor disc of the present technique as viewed along a rotational axis of the gas turbine rotor disc of the present technique; 
         FIG. 7  schematically illustrates a way of determining a first and a second axial extent in the hub of the gas turbine rotor disc; 
         FIG. 8  schematically illustrates another way of determining the first and the second axial extent in the hub of the gas turbine rotor disc; 
         FIG. 9  schematically illustrates yet another way of determining the first and the second axial extent in the hub of the gas turbine rotor disc; 
         FIG. 10  schematically illustrates a stress profile in a hub of the conventionally known rotor disc of  FIGS. 2 and 3 ; and 
         FIG. 11  schematically illustrates a stress profile in a hub of the gas turbine rotor disc of the present technique of  FIGS. 4 and 5 . 
     
    
    
     DETAILED DESCRIPTION OF INVENTION 
     Hereinafter, above-mentioned and other features of the present technique are described in details. Various embodiments are described with reference to the drawing, wherein like reference numerals are used to refer to like elements throughout. In the following description, for the purpose of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be noted that the illustrated embodiments are intended to explain, and not to limit the invention. It may be evident that such embodiments may be practiced without these specific details. 
     It may be noted that in the present disclosure, the terms “first”, “second”, etc. are used herein only to facilitate discussion, and carry no particular temporal or chronological significance unless otherwise indicated. 
       FIG. 1  shows an example of a gas turbine engine  10  in a sectional view. The gas turbine engine  10  comprises, in flow series, an inlet  12 , a compressor or compressor section  14 , a combustor section  16  and a turbine section  18  which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis  20 . The gas turbine engine  10  further comprises a shaft  22  which is rotatable about the rotational axis  20  and which extends longitudinally through the gas turbine engine  10 . The shaft  22  drivingly connects the turbine section  18  to the compressor section  14 . 
     In operation of the gas turbine engine  10 , air  24 , which is taken in through the air inlet  12  is compressed by the compressor section  14  and delivered to the combustion section or burner section  16 . The burner section  16  comprises a longitudinal axis  35  of the burner, a burner plenum  26 , one or more combustion chambers  28  and at least one burner  30  fixed to each combustion chamber  28 . The combustion chambers  28  and the burners  30  are located inside the burner plenum  26 . The compressed air passing through the compressor  14  enters a diffuser  32  and is discharged from the diffuser  32  into the burner plenum  26  from where a portion of the air enters the burner  30  and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas  34  or working gas from the combustion is channelled through the combustion chamber  28  to the turbine section  18  via a transition duct  17 . 
     This exemplary gas turbine engine  10  has a cannular combustor section arrangement  16 , which is constituted by an annular array of combustor cans  19  each having the burner  30  and the combustion chamber  28 , the transition duct  17  has a generally circular inlet that interfaces with the combustor chamber  28  and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine  18 . 
     The turbine section  18  comprises a number of blade carrying discs  36  attached to the shaft  22 . In the present example, two discs  36  each carry an annular array of turbine blades  38 . However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, guiding vanes  40 , which are fixed to a stator  42  of the gas turbine engine  10 , are disposed between the stages of annular arrays of turbine blades  38 . Between the exit of the combustion chamber  28  and the leading turbine blades  38  inlet guiding vanes  44  are provided and turn the flow of working gas onto the turbine blades  38 . 
     The combustion gas from the combustion chamber  28  enters the turbine section  18  and drives the turbine blades  38  which in turn rotates the shaft  22 . The guiding vanes  40 ,  44  serve to optimise the angle of the combustion or working gas on the turbine blades  38 . 
     The turbine section  18  drives the compressor section  14 . The compressor section  14  comprises an axial series of vane stages  46  and rotor blade stages  48 . The rotor blade stages  48  comprise a rotor disc supporting an annular array of blades. The compressor section  14  also comprises a casing  50  that surrounds the rotor stages and supports the vane stages  48 . The guide vane stages include an annular array of radially extending vanes that are mounted to the casing  50 . The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operational conditions. 
     The casing  50  defines a radially outer surface  52  of the passage  56  of the compressor  14 . A radially inner surface  54  of the passage  56  is at least partly defined by a rotor drum  53  of the rotor which is partly defined by the annular array of blades  48 . 
     The present technique is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present technique is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications. 
     The terms axial, radial and circumferential are made with reference to the rotational axis  20  of the engine, unless otherwise stated. 
       FIG. 4  schematically illustrates an exemplary embodiment of a turbine engine rotor disc  1 , and  FIG. 5  schematically illustrates the turbine engine rotor disc  1  of  FIG. 4  when incorporated with the gas turbine engine  10  of  FIG. 1  and contacted with a tension bolt  4  on one side and with a drive shaft  3  on the other side of the rotor disc  1 . A rotor disc retention assembly  100  of the gas turbine engine  10 . The rotor disc retention assembly  100  comprises the tension bolt  4 , the rotor disc  1  and the rotational axis  15 . The tension bolt  4  and the rotor disc  1  are arranged around the rotational axis ( 15 ). The turbine engine rotor disc  1 , hereinafter also referred to as the rotor disc  1 , is one of the rotor discs  36  depicted in  FIG. 1 , particularly the rotor disc  1  is that rotor disc  36  that is contacted with the tension bolt  4 . It may be noted that although only one rotor disc  1  is depicted between the tension bolt  4  and the drive shaft  3  in  FIG. 5 , there may be additional rotor discs  36  between the rotor disc  1  of  FIG. 5  and the drive shaft  3  of  FIG. 5 . In such an arrangement with one or more rotor discs  36  in addition to the rotor disc  1  of the present technique, the rotor disc  1  of the present technique is contacted with an adjacent rotor disc  36  via a Hirth coupling  2 , which may then be contacted with a subsequent adjacent rotor disc  36  via another Hirth coupling, and which in turn may be contacted to the drive shaft  3  via yet another Hirth coupling  2 . In the aforementioned arrangement with one or more rotor discs  36  in addition to the rotor disc  1  of the present technique, the rotor disc  1  is that rotor disc that is directly contacted or connected to the tension bolt  4 . 
     As depicted in  FIGS. 4 and 5 , the rotor disc  1  includes a hub  60 , a web  70 , a blade retention arrangement  80 , a rotational axis  15 , a first axial side  91  and a second axial side  92 . The hub  60  is region or part of the rotor disc  99  that surrounds a central hole  11  or central bore  11 . As shown in  FIG. 5  the central bore  11  is arranged around the rotational axis  15  of the rotor disc  1  when the rotor disc  1  is positioned inside the gas turbine engine  10  of  FIG. 1 . From the hub  60  extends radially the web  70  which is section of the rotor disc  1  that connects the hub  60  to the blade retention arrangement  80 . The blade retention arrangement  80  usually comprises slots (not shown) into which roots (not shown) of a plurality of the turbine blades  38  (shown in  FIG. 1 ) are arranged or fixed. Thus the turbine blades  38  are circumferentially arranged on the rotor disc  1  and extend radially outwards, with respect to the rotational axis  15  or the rotational axis  20 , from the rotor disc  1  and particularly outwards from the blade retention arrangement  80  of the rotor disc  1 . The rotor disc  1  and the plurality of the turbine blades  38  arranged on the rotor disc  1  together form a turbine engine rotor disc assembly  100  as shown in  FIG. 1 . The rotational axis  15  of the rotor disc  1  overlaps the rotational axis  20  when the rotor disc  1  is positioned inside the gas turbine engine  10  of  FIG. 1 . 
     As shown in  FIG. 5 , a tension bolt  4  of the gas turbine engine  10  passes through the central bore  11  and is physically contacted at the first axial side  91  of the rotor disc  1 . The tension bolt  4  bears the load of the turbine engine rotor disc assembly  100 , i.e. of the rotor disc  1  and the turbine blades  38  arranged on the rotor disc  1 , when the turbine engine rotor disc assembly  100  is rotated during operation of the gas turbine engine  10 . On the second axial side  92  of the rotor disc  1 , the rotor disc  1  is contacted or coupled with the drive shaft  3  of the gas turbine engine  10  via generally a Hirth coupling  2 . The location of the Hirth coupling  2  is depicted in  FIG. 4  although  FIG. 4  does not schematically depict the Hirth coupling  2  in its entirety along with the drive shaft  3 . The drive shaft  3  rotationally couples the gas turbine engine  10  to a downstream load for example a generator (not shown). The first and the second axial sides  91  and  92  are with respect to the rotational axis  15 . The first axial side  91  is adapted for engaging the tension bolt  4  of the gas turbine engine  10 . The first axial side  91  may include a chamfered recess  13  for receiving the tension bolt  4 , as shown in  FIGS. 4 and 5 , or for receiving a nut head (not shown) connected to the tension bolt  4 .  FIG. 5  depicts the second axial side  92  connected to the drive shaft  3  via the Hirth coupling  2 , however, as aforementioned the second axial side  92  may alternatively be connected to a subsequently arranged rotor disc  36  via the Hirth coupling  2 . 
     The tension bolt  4  applies a compressive force across the disc  1  or a number of discs and to secure the disc or discs to the drive shaft  3 . The tension bolt  4  is therefore in tension. The tension bolt  4  may be attached and tightened to the drive shaft by a spline arrangement  102 . 
     The blade retention arrangement  80  has a centre of mass  82 . The centre of mass  82  may be a geometric centre of the blade retention arrangement  80  when the blade retention arrangement  80  is formed symmetrically and with a homogenous material. The blade retention arrangement  80  may be assumed to be divided by a radial plane  5  that passes through the centre of mass  82  of the blade retention arrangement  80  and is perpendicular to the rotational axis  15 .  FIGS. 4, 5 and 6  schematically depict the radial plane  5 . The radial plane  5  extends through the rotor disc  1  intersecting the central bore  11 , the hub  60 , the web  70  and the blade retention arrangement  80 . 
     As shown in  FIGS. 4 and 5 , the radial plane  5  by intersecting the hub  60  defines a first axial side portion  61  in the hub  60  towards the first axial side  91  and a second axial side portion  62  in the hub  60  towards the second axial side  92 . In the rotor disc  1 , the second axial side portion  62  axially extends between 10% and 30% more than the first axial side portion  61 .  FIGS. 7, 8 and 9  present different ways of defining the axial extension of the first axial side portion  61  and the second axial side portion  62 . 
     As shown in  FIGS. 8 and 9 , the first axial side portion  61  has an axial extent  63  and the second axial side portion  62  has an axial extent  64 . According to the present technique, in the rotor disc  1 , the axial extent  64  of the second axial side portion  62  is between 10% and 30% greater than the axial extent  63  of the first axial side portion  61 . 
     As schematically depicted in  FIG. 8 , measurements of the first axial extent  63  and the second axial extent  64  are limited to a region  67  of the hub  60 . In other words the measurement of the first axial extent  63  and the second axial extent  64  are performed within the region  67  of the hub  60 . The measurements of the first axial extent  63  and the second axial extent  64  are performed in a continuous straight line perpendicular to the radial plane  5 . The measurement or value of the first axial extent  63  is a measure of length or distance from the radial plane  5  to an edge of the first axial side  91  within the region  67 , i.e. a measure of length of the first axial side portion  61 . Similarly, the measurement or value of the second axial extent  64  is a measure of length or distance from the radial plane  5  to an edge of the second axial side  92  within the region  67 , i.e. a measure of length of the second axial side portion  62 . The region  67  of the hub  60  is a region or portion of the hub  60  that has geometric similarity at the first axial side  91  and the second axial side  92 . 
     The geometric similarity as used herein means that within the region  67  the first and the second axial sides  91 ,  92  both have the same shape, or one has the same shape as the mirror image of the other, mirrored across the radial plane  5 . An example of geometric similarity is when the axial sides  91 ,  92  have same or substantially similar angle of curvature at their respective edges within the region  67 . 
     As shown in  FIG. 7 , the region  67  of the hub  60  is free from an integrally formed connection  68  projecting out from the hub  60 . The integrally formed connection  68  may be adapted for contacting one or more components  7  of the gas turbine engine  10 , for example a support extending from the hub  60  and adapted to contact a subsequent rotor disc (not shown). In another words, the measurement of the axial extents  63 ,  64  do not include any such integrally formed connections  68  and are limited to a main body of the hub  60 .  FIG. 7  depicts another region  69  in the hub  60  of the rotor disc  1 . The region  69  shows the integrally formed connection  68  for example a projection  68  extending outward from the hub  60 . While determining the axial extends  63 ,  64  i.e. while measuring the first and the second axial side portions  61 ,  62  the measurements are to be performed within the region  67  or of the region  67  and not within the region  69  or of the region  69 . 
     As depicted in  FIG. 9 , the measurements of the axial extents  63 ,  64  are defined at an axial surface  88  of the hub  60 . In other words the measurement or value of the first axial extent  63  is a measure of length or distance from the radial plane  5  to an edge of the axial surface  88  of the first axial side  91 , i.e. a measure of length of the first axial side portion  61 . Similarly, the measurement or value of the second axial extent  64  is a measure of length or distance from the radial plane  5  to an edge of the axial surface  88  of the second axial side  92  i.e. a measure of length of the second axial side portion  62 . The axial surface  88  is a surface of the hub  60  that defines the central bore  11 . 
       FIG. 11  schematically illustrates a stress profile in the hub  60  of the gas turbine rotor disc  1  of the present technique, for example in the exemplary embodiment of the rotor disc  1  as depicted in  FIGS. 4 and 5 . The stress profile in the hub  60  of the rotor disc  1  may be understood comparatively with respect to the stress profile in the hub  60  of the conventionally known rotor disc  99  as depicted in  FIG. 10  for the conventionally known rotor disc  99  shown in  FIGS. 2 and 3 . 
     In the rotor disc  1  of the present technique, due to greater axial extent  64  of the second axial side portion  62 , the stress concentration is optimized and distributed differently as compared to the stress profile depicted in  FIG. 10  for the conventionally known rotor disc  99 . Due to the increased axial extent  64  of the second axial side portion  62 , the peak stress is formed substantially towards a centre of the hub  60 , instead of being formed at the edge  93  as aforementioned in case of the stress profile depicted in  FIG. 10  for the conventionally known rotor disc  99 . 
     It may be noted that the greater axial extent of the second axial side portion  62  as compared to the first axial side portion  61  results from having more material of the hub  60  at the second axial side portion  62  as compared to the first axial side portion  61  of the hub  60 , however the increase in the axial extent i.e. addition of the more material at the second axial side portion  62  as compared to the first axial side portion  61  of the hub  60  is not done as a separate component, the hub  60  including the first axial side portion  61  and the second axial side portion  62  is formed integrally as a single body along with the web  70  and the blade retention arrangement  80 . 
     While the present technique has been described in detail with reference to certain embodiments, it should be appreciated that the present technique is not limited to those precise embodiments. Rather, in view of the present disclosure which describes exemplary modes for practicing the invention, many modifications and variations would present themselves, to those skilled in the art without departing from the scope and spirit of this invention. The scope of the invention is, therefore, indicated by the following claims rather than by the foregoing description. All changes, modifications, and variations coming within the meaning and range of equivalency of the claims are to be considered within their scope.