Abstract:
A method and system for a rotatable member of a turbine engine are provided. The rotatable member includes a substantially cylindrical shaft rotatable about a longitudinal axis, and a hub coupled to the cylindrical shaft through a conical shaft portion wherein the conical shaft portion includes a plurality of circumferentially-spaced air passages and wherein at least one of the plurality of air passages includes a non-circular cross section.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to turbine engines and, more particularly, to a method and system for maintaining cooling to internal components of turbine engines. 
         [0002]    At least some known turbine engine high pressure turbine disks include radially outer rim slots for attaching a plurality of blades to the disk using a dovetail connection. The dimensions of the slots combined with the forces exerted on the rim during various operational loadings tend to shorten the life of the disk. To strengthen the area of the rim that tends to limit the life of the disk, the dimensions of the slots may be modified. However, modification of the dovetail slot shape to increase the strength of the disk can decrease the blade cooling circuit pressure and cooling flow margins to the blades attached at the slots. 
         [0003]    In addition, improving the life of the disk by improving the rim makes particulate erosion in forward inner shaft cavity of the disk a new life limiting area. Eliminating the particulate erosion in the forward inner shaft cavity of the disk is accomplished by eliminating the deep pocket between the shaft and the disk. However, this modification results in an excessive stress concentration at the top of the shaft air hole due to the reduction in displacement attenuation between the disk hub and the area of the shaft where the hole is positioned. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0004]    In one embodiment, a rotatable member of a turbine engine includes a substantially cylindrical shaft rotatable about a longitudinal axis, and a hub coupled to the cylindrical shaft through a conical shaft portion wherein the conical shaft portion includes a plurality of circumferentially-spaced air passages and wherein at least one of the plurality of air passages includes a non-circular cross section. 
         [0005]    In another embodiment, a method of forming a turbine disk is provided. The turbine disk includes a hub coupled to a shaft portion, a radially outer rim, and a web extending therebetween. The method includes determining a first blade slot depth for receiving blades on the turbine disk, determining a second blade slot depth that facilitates reducing stress in the rim wherein the second blade slot depth is less then the first blade slot depth, forming the rim using the second slot depth, and forming the shaft portion that includes at least one air passage having a non-circular cross-section. 
         [0006]    In yet another embodiment, a turbine engine system includes a disk rotatable about a longitudinal axis. The disk includes a hub coupled to a conical shaft portion that includes a plurality of circumferentially-spaced air passages wherein at least one of the plurality of air passages includes a non-circular cross section. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]      FIGS. 1-5  show exemplary embodiments of the method and system described herein. 
           [0008]      FIG. 1  is a cross-sectional view of a high pressure turbine first stage disk assembly and a second stage disk assembly of a high pressure turbine assembly in accordance with an exemplary embodiment of the present invention; 
           [0009]      FIG. 2  is an enlarged cross-sectional view of the high pressure turbine first stage disk assembly and the second stage disk assembly shown in  FIG. 1 ; 
           [0010]      FIG. 3  is a side elevation view of the high pressure turbine first stage disk assembly in accordance with an exemplary embodiment of the present invention; 
           [0011]      FIG. 4  is a schematic view of the air passage shown in  FIG. 1  in accordance with an exemplary embodiment of the present invention; 
           [0012]      FIG. 5A  is an aftward perspective view of the conical shaft connection shown in  FIG. 1  including circular shaped holes; and 
           [0013]      FIG. 5B  is an aftward perspective view of the conical shaft connection shown in  FIG. 1  in accordance with an exemplary embodiment of the present invention. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0014]    The following detailed description illustrates embodiments of the invention by way of example and not by way of limitation. It is contemplated that the invention has general application to embodiments of turbine engine components in industrial, commercial, and residential applications. 
         [0015]    As used herein, an element or step recited in the singular and proceeded with the word “a” or “an” should be understood as not excluding plural elements or steps, unless such exclusion is explicitly recited. Furthermore, references to “one embodiment” of the present invention are not intended to be interpreted as excluding the existence of additional embodiments that also incorporate the recited features. 
         [0016]      FIG. 1  is a cross-sectional view of a high pressure turbine assembly  100  in accordance with an exemplary embodiment of the present invention. In the exemplary embodiment, high pressure turbine assembly  100  includes a high pressure turbine first stage disk assembly  102  and a second stage disk assembly  104 . First stage disk assembly  102  and second stage disk assembly  104  are circumscribed about an engine centerline  106  of a gas turbine engine such as a General Electric CF6-80 aircraft gas turbine engine. First and second stage disk assemblies  102  and  104  include first and second disks  108  and  110  having slotted first and second rims  112  and  114  which receive first and second turbine blades  116  and  118 , respectively, in a dovetail fit. First and second blades  116  and  118  are axially retained within their respective first and second rims  112  and  114  by first forward and aft blade retainers  120  and second forward and aft blade retainers  122 , respectively. First and second disks  108  and  110  include first and second webs  124  and  126  extending radially inwardly from first and second rims  112  and  114 , to first and second hubs  128  and  130 , respectively. First stage disk assembly  102  includes a cooling air deswirler  132  located radially outward from a conical shaft connection  134  to a substantially cylindrical shaft  136  extending axially forwardly from first hub  128  of disk  108 . A flow of cooling air  138  is channeled from a high-pressure compressor discharge (not shown) through a cavity  140 , deswirler  132 , and through at least one of a plurality of air passages  142  that channels cooling air onboard disk assemblies  102  and  104 . At least a portion of the flow of cooling air  138  is channeled to slots  144  and  146  in first and second rims  112  and  114 . The flow of cooling air  138  is further channeled to blades  116  and  118  from slots  144  and  146 . Because slots  144  and  146  form a portion of the cooling air circuit for cooling air to blades  116  and  118 , a dimension of slots  144  and  146  is at least partially determinant of a head loss through the cooling circuit. For example, if a cross-sectional area of slots  144  and/or  146  is reduced in size, the flow of cooling air  138  to blades  116  and/or  118  may be reduced. In the exemplary embodiment, the cross-sectional area of slots  144  and/or  146  is reduced in size to facilitate reducing stress damage to first and second rims  112  and  114 . Air passages  142  also form a portion of the cooling circuit and as such a cross-sectional area of cooling passages  142  also affects the head loss in the cooling air circuit to blades  116  and/or  118 . By increasing a cross-sectional area of air passages  142 , head loss in the cooling circuit can be reduced thereby making up for the increased head loss due to reducing the size of slots  144  and  146 . However, simply increasing the diameter of air passages  142  was determined to increase stress in an area of air passages  142  and hub  128 . 
         [0017]      FIG. 2  is an enlarged cross-sectional view of high pressure turbine first stage disk assembly  102  and a second stage disk assembly  104  (shown in  FIG. 1 ). An annular cavity  148  is formed between a conical connection  134  to the cylindrical shaft  136  and the first hub  128  and is closed at an intersection of conical connection  134  and first hub  128  and open and exposed to the flow of cooling air  138  passing through the cooling air deswirler  132  at an inner diameter (ID)  152  of the first hub  128 . Dust and debris in the flow of cooling air  138  can become entrapped and build up in cavity  148  over time with continued operation of the engine. The flow of cooling air  138  has both axial and circumferential velocities relative to the rotating first hub  128 . Debris entrained in the flow of cooling air  138  can circumferentially scrub rotating internal surfaces  154  of rotating first hub  128  and, over time, cause damage to the first hub  128  and first disk  108 . Annular cavity  148  is formed as a deep pocket to provide significant attenuation from disk hub growth at the location of air passages  142 . Removing this pocket to eliminate erosion of internal surfaces  154  decreases the attenuation at the location of air passages  142  (the location of air passages  142  is fixed by the location of deswirler  132 ). Air passages  142  are shaped and oriented to maintain sufficient attenuation and to ensure workable stresses. In the exemplary embodiment, air passages  142  are positioned in alignment with an outlet of deswirler  132  to act as an extension of the diffuser/impeller, allowing the walls of the holes to put work into the flow thereby increasing pressure and reducing the swirl of the flow of cooling air  138  relative to disk  108 . 
         [0018]      FIG. 3  is a side elevation view of high pressure turbine first stage disk assembly  102  in accordance with an exemplary embodiment of the present invention. Disk assembly  102  is shown in  FIG. 3  as a first embodiment  302  in solid lines superimposed on a second embodiment  304  in dashed lines so that differences between a profile of second embodiment  304  and a profile of first embodiment  302  are more clearly apparent. First embodiment  302  includes rim slot  144  having a first depth  306 . Second embodiment  304  includes rim slot  144  that has a second depth  308 . In first embodiment  302 , first depth  306  is at least partially responsible for increased stress in slot  144 , which tends to shorten a life of disk assembly  102 . By forming disk assembly  102  using second depth  308 , which permits a larger slot bottom radius, stress in slot  144  is substantially reduced. However, a shallower depth of slot  144  also decreases a cross-sectional area of a cooling path in slot  144  and a reduced flow of cooling air  138  to blade  116  (shown in  FIG. 1 ). Because of the increased life of disk assembly  102  provided by a shallower slot  144 , the erosion in cavity  148  becomes a more life limiting area than the shallower slot  144  and presents a new problem to be solved. 
         [0019]    A solution to the erosion problem is provided by eliminating the deep pocket of cavity  148 . However, cavity  148  serves to improve attenuation of air passages  142  from the expansion of hub  128  due to centrifugal and thermal loads. To compensate for the reduction of the attenuation and to reduce stresses to an adequate level, the shape, position, and orientation of air passages are modified and to restore adequate cooling air pressure to blades  116  and/or  118 , the area of air passages is increased. 
         [0020]      FIG. 4  is a schematic view of air passage  142  (shown in  FIG. 1 ) in accordance with an exemplary embodiment of the present invention. In the exemplary embodiment, air passage  142  includes a non-circular shape, for example, but not limited to an elliptical shape. The elliptical shape of air passage  142  includes a major axis  402  and a minor axis  404 . Air passage  142  has a width  406  across major axis and a depth  408  across minor access  404 . A circumferential line  410  circumscribes conical shaft connection  134  at an axial location through a center  412  of the elliptical shape of air passage  142 . Because conical shaft connection  134  transfers significant torque from the high pressure turbine (HPT) to a high pressure compressor (HPC), major axis of air passage  142  is canted by an angle α with respect to circumferential line  410 . In one embodiment, angle α is an angle between five degrees and twenty degrees with respect to circumferential line  410 . In another embodiment angle α is approximately fifteen degrees to maintain a highest stress peak proximate a center of major axis  402 . This results in significant stress reduction and robustness for all operations (including the stresses due to torque) by maintaining the peak stress located on the largest radius possible and in the most advantageous position on the surface of air passage  142 . The stress reduction obtained from newly shaped air passage  142  allowed for elimination of the shaft forward inner pocket cavity  148 . 
         [0021]    The elliptical shape of air passage  142  is able to achieve a greater opening area than a circular opening having an increased diameter without increasing peak hole stresses unacceptably proximate air passage  142 . The greater opening area permits an improvement in the flow circuit pressure. In combination with eliminating the deep pocket cavity  148  it would not be possible to enlarge air passage  142  as a circular hole due to a lack of space in conical shaft connection  134  proximate hub  128 . In addition, the non-circular shaped hole is sized, shaped, and oriented to act as a diffuser extension of deswirler  132  in that selecting the clocking position of a pattern of air passages  142  in relation to deswirler  132  permits control of the flow tangential mach number radially inward from air passage  142  facilitate pressure recovery. 
         [0022]      FIG. 5A  is an aftward-looking perspective view of conical shaft connection  134  (shown in  FIG. 1 ) including circular shaped holes  502 . As illustrated in  FIG. 5A , holes  502  are clocked approximately two degrees with respect to deswirler clips  504 , which are indicative of a position of vanes of the deswirler when installed.  FIG. 5B  is an aftward perspective view of conical shaft connection  134  (shown in  FIG. 1 ) in accordance with an exemplary embodiment of the present invention. In the exemplary embodiment, air passages  142  are elliptically-shaped passages that extend through conical shaft connection  134 . Major axis  402  of air passages  142  are canted approximately fifteen degrees with respect to circumferential line  410 . Air passages  142  are clocked approximately seven degrees with respect to a position of vanes in deswirler  132  when installed. Positions of a plurality of attachment clips  504  are indicative of the position of deswirler  132 . An angle β represents an amount of the clocking position of center  412  in relation to deswirler  132 . In one embodiment, angle β is between approximately three degrees and approximately fifteen degrees. In an alternative embodiment, angle β is approximately five degrees to approximately ten degrees. In the exemplary embodiment, angle β is approximately seven degrees. Setting angle β to approximately seven degrees also decreased the tangential Mach number to an acceptable value. 
         [0023]    The above-described embodiments of a method and system of forming a turbine disk provides a cost-effective and reliable means for providing cooling to components of a turbine engine and reducing stress in such components. More specifically, the methods and systems described herein facilitate increases a life of components of a high pressure turbine disk assembly such that a life of the assembly as a whole is increased. As a result, the methods and systems described herein facilitate forming and operating turbine engines in a cost-effective and reliable manner. 
         [0024]    An exemplary method and system for forming a turbine disk and maintaining are described above in detail. The apparatus illustrated is not limited to the specific embodiments described herein, but rather, components of each may be utilized independently and separately from other components described herein. Each system component can also be used in combination with other system components. 
         [0025]    While the disclosure has been described in terms of various specific embodiments, it will be recognized that the disclosure can be practiced with modification within the spirit and scope of the claims.