Patent Publication Number: US-6983608-B2

Title: Methods and apparatus for assembling gas turbine engines

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH &amp; DEVELOPMENT 
     The government may have rights in this invention pursuant to government contract number N00019-01-C-0147. 
    
    
     BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engines and more particularly, to methods and apparatus for assembling gas turbine engines. 
     Known gas turbine engines include at least one rotor shaft supported by bearings which are in turn supported by annular frames. At least some known turbine frames include an annular casing that is spaced radially outwardly from an annular hub. A plurality of circumferentially-spaced apart struts extend between the annular casing and the hub. More specifically, within at least some known turbine engines, the struts, casing, and hub are integrally-formed together. In other known turbine engines, multi-piece frames are used in which only the struts and casing are integrally formed together. 
     Because at least some of the struts extend through a flow path defined within the engine, at least some of the struts are surrounded by, and extend through, a fairing that facilitates shielding the struts from hot combustion gases flowing through the flow path. More specifically, to facilitate increasing the structural integrity of fairings positioned in the flowpath, at least some known fairings are fabricated as a single-piece casting that includes at least one internal serpentine cooling passage. However, airflow and structural design requirements of such fairings may complicate the assembly of the struts to the engine frame. For example, because such fairings are unitary, the fairings may only be utilized with multi-piece frames. More specifically, each unitary strut is positioned around an inner end of each strut, slid radially outward towards a cantilevered end of each strut, and is coupled in position using a plurality of precisely-machined fastening/coupling hardware. Accordingly, because of the additional assembly and coupling hardware associated with multi-piece frames, and because of the tolerances that may be necessary to meet structural requirements, manufacturing and assembly costs of such frames may be more costly and time-consuming than associated with other known frames. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, a method for assembling a gas turbine engine is provided. The method comprises providing an engine frame including an integrally formed outer band, an inner band, and a plurality of circumferentially-spaced apart struts extending radially therebetween, and providing at least one fairing that is formed as an integral single piece casting and includes a first sidewall and a second sidewall connected at a leading edge and a trailing edge such that at least one cooling chamber is defined therebetween. The method also comprises coupling the at least one fairing around at least one strut such that the strut extends through the fairing at least one cooling chamber and such that during the coupling process the fairing is only transitioned axially around the strut rather being slid radially along the strut. 
     In another aspect, a fairing for use with a gas turbine frame strut is provided. The fairing is cast as an integral single piece and includes a first sidewall and a second sidewall connected together at a leading edge and a trailing edge such that at least one cooling chamber is defined therebetween. The fairing includes at least one partition and at least one parting line. The at least one partition is formed integrally with, and extends between, the first and second sidewalls. The at least one parting line divides the fairing into a forward portion and a separate aft portion that are removably coupled together. 
     In a further aspect, a gas turbine engine is provided. The engine includes an engine frame and at least one fairing. The engine frame includes an outer band, an inner band, and a plurality of circumferentially-spaced apart struts extending radially therebetween. The plurality of struts are formed integrally with the outer and inner bands. The at least one fairing is configured to be coupled around one of the plurality of struts such that a respective strut extends through the at least one fairing. The fairing is formed as an integral single piece and includes a first sidewall and a second sidewall connected together at a leading edge and a trailing edge such that at least one cooling chamber is defined therebetween. The fairing further includes at least one partition and at least one parting line. The at least one partition extends between the first and second sidewalls. The at least one parting line separates the fairing into a forward portion and a separate aft portion that are removably coupled together. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic illustration of an exemplary gas turbine engine; 
         FIG. 2  is an aft-facing-forward view of an exemplary turbine frame that may be used with the turbine engine shown in  FIG. 1 ; 
         FIG. 3  is an partial cross-sectional side view of the turbine engine shown in  FIG. 1  and including the turbine frame shown in  FIG. 2 ; 
         FIG. 4  is a cross-sectional view of an exemplary fairing that may be used with the turbine frame shown in  FIG. 3 ; and 
         FIG. 5  is an enlarged view of a portion of the fairing shown in  FIG. 4  and taken along area  5 — 5 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a schematic illustration of a gas turbine engine  10  including a fan assembly  12  and a core engine  13  including a high pressure compressor  14 , and a combustor  16 . Engine  10  also includes a high pressure turbine  18 , a low pressure turbine  20 , and a booster  22 . Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disc  26 . Engine  10  has an intake side  28  and an exhaust side  30 . In one embodiment, the gas turbine engine is a GE90 available from General Electric Company, Cincinnati, Ohio. Fan assembly  12  and turbine  20  are coupled by a first rotor shaft  31 , and compressor  14  and turbine  18  are coupled by a second rotor shaft  32 . 
     During operation, air flows through fan assembly  12 , in a direction that is substantially parallel to a central axis  34  extending through engine  10 , and compressed air is supplied to high pressure compressor  14 . The highly compressed air is delivered to combustor  16 . Airflow (not shown in  FIG. 1 ) from combustor  16  drives turbines  18  and  20 , and turbine  20  drives fan assembly  12  by way of shaft  31 . 
       FIG. 2  is an aft-facing-forward view of an exemplary turbine frame  40  that may be used with gas turbine engine  10 .  FIG. 3  is an partial exemplary cross-sectional side view of engine  10 , including turbine frame  40 . Engine  10  includes a row of rotor blades  42  coupled to a rotor disk  44 . Frame  40  and disk  44  are positioned substantially co-axially about a longitudinal or axial centerline axis  46  extending through engine  10 , and as such, are in flow communication with hot combustion gases  48  discharged from a combustor (not shown in  FIG. 2  or  3 ), such as combustor  16 . 
     Turbine frame  40  includes a plurality of circumferentially-spaced apart, and radially-extending support struts  50 . Each strut  50  extends between a radially outer ring or band  52  and a radially inner hub or band  54 . In the exemplary embodiment, frame  40  is cast integrally with struts  50  and bands  52  and  54 . In the exemplary embodiment, outer band  52  is securely coupled to an annular casing  56  of engine  10 , and inner band  54  is securely coupled to an annular bearing support  58 . Struts  50  and bearing support  58  provide a relatively rigid assembly for transferring rotor loads induced during engine operation. 
     Each strut  50  extends through a fairing  60  which, as described in more detail below, facilitates shielding each strut  50  from combustion gases flowing through engine  10 . In the exemplary embodiment, each fairing  60  is fabricated from a high temperature cast alloy. Moreover, cooling fluid is channeled into an internal cooling chamber (not shown in  FIG. 2  or  3 ) defined within each strut  50  to facilitate reducing an operating temperature of each strut  50  and fairing  60 . 
     Fairings  60  are coupled at respective radially outer and inner ends  62  and  64  to corresponding annular outer and inner liners  66  and  68 . Liners  66  and  68  confine a flow of the combustion gases  48  therebetween, and are therefore correspondingly heated by combustion gases  48  during engine operation. Fairings  60  and liners  66  and  68  are supported by respective bands  52  and  54  to accommodate substantially unrestrained differential thermal movement therewith. 
     In the exemplary embodiment, turbine frame  40  also includes a plurality of vanes  70  coupled to, and extending between, outer and inner liners  66  and  68 , respectively, such that each vane  70  is positioned between adjacent circumferentially-spaced fairings  60 . Accordingly, in the exemplary embodiment, engine frame  40  includes nine fairings  60  and struts  50  spaced apart substantially uniformly around a perimeter of frame  40 , and nine vanes  70  spaced substantially equally between each respective pair of circumferentially-spaced struts  50 . Vanes  70  are substantially identical in configuration to fairings  60 , except that no strut  50  extends radially therethrough. In an alternative embodiment, frame  40  does not include any vanes  70 . 
       FIG. 4  is a cross-sectional view of fairing  60 .  FIG. 5  is an enlarged view of a portion of fairing  60  and taken along area  5 — 5 . Each fairing  60  includes a first sidewall  80  and a second sidewall  82  that is spaced apart from first sidewall  80 . First sidewall  80  extends longitudinally between fairing ends  62  and  64  (shown in  FIGS. 2 and 3 ) and defines a pressure side of fairing  60 . Second sidewall  82  also extends longitudinally between fairing ends  62  and  64  and defines a suction side of fairing  60 . Sidewalls  80  and  82  are joined at a leading edge  84  and at an axially-spaced trailing edge  86  of fairing  60 , such that a cooling chamber  88  is defined within fairing  60 . More specifically, each sidewall  80  and  82  has an inner surface  90  and an opposite outer surface  92 . Outer surface  92  defines a gas flowpath surface. Cooling chamber  88  is defined by inner surface  90  and is bounded between sidewalls  80  and  82 . 
     In the exemplary embodiment, cooling chamber  88  includes a plurality of inner ribs or partitions  94  which partition cooling cavity  88  into a plurality of cooling chambers  88 . Specifically, in the exemplary embodiment, fairing  60  is a single piece casting that is formed integrally with sidewalls  80  and  82 , and inner walls  94 . More specifically, airfoil  42  includes a leading edge cooling chamber  100 , a trailing edge cooling chamber  102 , and at least one intermediate cooling chamber  104 . In one embodiment, leading edge cooling chamber  100  is in flow communication with trailing edge and intermediate cooling chambers  102  and  104 , respectively. In the exemplary embodiment, at least a portion of chambers  88  is configured as a serpentine cooling passageway. 
     Leading edge cooling chamber  100  extends longitudinally or radially through fairing  60 , and is bordered by sidewalls  80  and  82 , and by fairing leading edge  84 . Each intermediate cooling chamber  104  is between leading edge cooling chamber  100  and trailing edge cooling chamber  102 , and is bordered by bordered by sidewalls  80  and  82  and by a leading edge partition  110  and an intermediate partition  112 . In the exemplary embodiment, intermediate partition  112  is slightly aft of a mid-chord (not shown) of fairing  60 . Trailing edge cooling chamber  102  extends longitudinally or radially through fairing  60 , and is bordered by sidewalls  80  and  82 , and by fairing trailing edge  86 . 
     Leading edge partition  110  and intermediate partition  112  extend between sidewalls  80  and  82 . More specifically, intermediate partition  112  is formed integrally with a pair of outer end portions  114  and  116 , and a body portion  118  extending therebetween. In the exemplary embodiment, a thickness T 1  of body portion  118  is substantially constant between ends  114  and  116 , and each end  114  and  116  has a thickness T 2  that is thicker than body thickness T 1 . In one embodiment, end thickness T 2  is created by the coupling additional material  120  to partition  112  through a known process, such as, but not limited to a known welding process. In another embodiment, partition thickness T 2  is formed integrally with partition  112  during the casting process. More specifically, in such a process, material  120  may be coupled to an existing fairing partition to modify the existing engine fairing, or alternatively, may be cast as an integral portion of a partition during fabrication of the engine frame fairing. 
     Moreover, although ends  114  and  116  are illustrated as having a generally rectangular cross-sectional profile, it should be noted that ends  114  and  116  are not limited to having a generally rectangular cross-sectional profile. For example, in another embodiment, ends  114  and  116  are chamfered and have a generally triangular cross-sectional profile. 
     In the exemplary embodiment, additional material  120  is added only to an aft side  130  of partition  112  adjacent ends  114  and  116 , such that material  120  extends from partition  118  and from sidewall inner surfaces  90 . In an alternative embodiment, additional material  120  is added to a forward side  132  of partition  112  adjacent ends  114  and  116 . In a further alternative embodiment, additional material  120  is added to respective forward and/or aft sides  132  and  130  of partition  112  adjacent ends  114  and  116 . In one embodiment, partition  118  does not extend fully longitudinally through fairing  60  between fairing ends  62  and  64 , but additional material  120  is added longitudinally through fairing  60  and along sidewall inner surface  90 , such that a cross-sectional profile of material  120  is substantially constant longitudinally through fairing  60  between ends  62  and  64 . 
     Fairing  60  is also formed with a parting line  140  such that a two-piece fairing is produced from a single casting which, as described in more detail below, facilitates coupling fairing  60  around each respective strut  50 . Specifically, parting line  140  extends from sidewall  80  to sidewall  82  through intermediate cooling chamber  104 , and divides fairing  60  into a forward portion  144  and an aft portion  146 . More specifically, part line  140  extends through intermediate cooling chamber  104  immediately upstream from intermediate partition  112 . 
     In the exemplary embodiment, parting line  104  includes a pair of cut lines  150  and  152  that are mirrored-images of each other. Specifically, cut line  150  extends between sidewall inner and outer surfaces  90  and  92 , respectively, through sidewall  80 , and similarly, cut line  152  extends between sidewall inner and outer surfaces  90  and  92 , respectively, through sidewall  82 . More specifically, in the exemplary embodiment, each cut line  150  and  152  extends at least partially through additional material  120 . 
     In the exemplary embodiment, each cut line  150  and  152  defines a tongue and groove joint configuration  156  that facilitates coupling faring forward and aft portions  144  and  146 , respectively. In alternative embodiments, forward and aft portions  144  and  146  are coupled together using other joint configurations. Moreover, in another alternative embodiment, cut lines  150  and  152  are not mirrored images of each other. 
     In the exemplary embodiment, each cut line  150  and  152  extends radially inward from sidewall outer surface  92  at a location that is approximately centered with respect to each respective intermediate partition end  114  and  116 . More specifically, in the exemplary embodiment, each cut line  150  and  152  extends radially inward for a distance D 1  that is approximately equal to a thickness T 3  of each sidewall  80  and  82 . Each cut line  150  and  152  then extends aftward in a predetermined radius of curvature R 1  such that a semi-circular portion  160  is defined within partition material  120 . Each cut line  150  and  152  is then extended generally axially through partition  112  to partition forward side  132 . Accordingly, each cut line  150  and  152  defines a respective aft-facing step  164  and  166  along each gas flowpath surface  92 . 
     A retaining groove  170  is formed within each cut line  150  and  152  between each semi-circular portion  160  and partition forward side  132 . Each groove  170 , as described in ore detail below, is offset with respect to each cut line  150  and  152  to facilitate sealing along parting line  140  when fairing portions  144  and  146  are coupled together. Moreover, because each groove  170  is offset with respect to each cut line  150  and  152 , parting line  140  is divided into four sealing locations  180  spaced along line  140 . 
     During fabrication of fairings  60 , initially each fairing  60  is cast as an integrally-formed single casting. Parting line  140  is then formed within fairing  60 . Specifically, in the exemplary embodiment, each cut line  150  and  152  is formed via a primary electrical discharge machining (EDM) wire, and a secondary EDM wire is used to create grooves  170 . In addition to creating sealing locations  180 , offsetting grooves  170  with respect to each cut line  150  and  152  also facilitates compensating for wire EDM kerf. Each groove  170  is sized to receive a locking wire  174  therein which facilitates sealing between fairing portions  144  and  146 . 
     Accordingly, when parting line  140  has been formed, each fairing  60  may be coupled around each strut  50  in an axial direction rather than having to be slid radially outward from a cantilevered end of each strut  50 . More specifically, parting line  140  creates a two-piece fairing  60  that may be coupled to an integrally-formed, one-piece frame  40  such that multi-piece frame structures are not necessary. Specifically, once parting line  140  is created, fairing forward portion  144  is removably coupled to fairing aft portion  146 . Accordingly, during assembly, fairing aft portion  146  may be positioned relative to a respective strut  50  to be shielded, and such that a locking wire  174  is positioned within each sealing groove  170 . Fairing forward portion  144  is then axially coupled to aft portion  146  to complete the installation of fairing  60  such that strut  50  is shielded therein. Each locking wire  174  facilitates sealing between fairing portions  144  and  140  such that fluid leakage through each joint  156  is facilitated to be reduced. 
     Accordingly, assembly costs and times are facilitated to be reduced in comparison to those associated with multi-piece frame assemblies. Moreover, parting line  140  also enables high temperature cast alloy materials to be used to form fairings  60  without requiring more expensive multi-piece frame assemblies. 
     Moreover, fairing  60  is also reusable in that it is removable from one strut  50  and can be easily assembled on another strut  50 . Because forward and aft fairing portions  140  and  144  can assemble axially around each strut  50 , fairing  60  not only facilitates eliminating multi-piece frame structures, but also eliminates locking mechanisms and/or coupling hardware that is used with multi-piece frame assemblies. Accordingly, incorporating fairings  60  facilitate reducing design efforts from both a cost and cycle basis, along with hardware manufacturing and development cycles. 
     The above-described engine frame fairings are cost-effective and highly reliable. Each fairing is coupled axially around an integrally formed, one-piece engine frame. Accordingly, expensive coupling hardware associated with multi-piece engine frames is eliminated. Moreover, existing fairings may be modified for use as described herein. As a result, a fairing design is provided that facilitates minimizing the design efforts associated with both a cost-cycle basis, along with coupling hardware and manufacturing development cycles. 
     Exemplary embodiments of an engine frame, are described above in detail. The engine frames illustrated are not limited to the specific embodiments described herein, but rather, the fairings described herein may be utilized independently and separately from the gas turbine engine frames described herein. 
     While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.