Abstract:
A method enables a rotor assembly for a gas turbine engine to be fabricated. The method includes forming a blade including an airfoil extending from an integral dovetail used to mount the blade within the rotor assembly, and extending a projection from at least a portion of the blade, such that the stresses induced within at least a portion of the blade are facilitated to be maintained below a predetermined failure threshold for the blade to facilitate preventing failure of the blade.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engine blades, and more specifically to methods and apparatus for facilitating preventing failure of gas turbine engine blades. 
     At least some known gas turbine engines include a core engine having, in serial flow arrangement, a fan assembly and a high pressure compressor which compress airflow entering the engine. A combustor ignites a fuel-air mixture which is then channeled through a turbine nozzle assembly towards low and high pressure turbines which each include a plurality of rotor blades that extract rotational energy from airflow exiting the combustor. 
     Failure of a component within a system may significantly damage the system and/or other components within the system, and may also require system operation be suspended while the failed component is replaced or repaired. More particularly, when the component is a turbofan gas turbine engine fan blade, a blade-out may cause damage to a blade that is downstream from the released blade. More specifically, depending upon the severity of the damage to the downstream blade, other blades downstream from the released blade or the damaged trailing blade may also be damaged. Damage to the trailing blade may cause the trailing blade to fail, thereby possibly requiring operation of the turbofan gas turbine engine be suspended, and/or damage to other fan blades and/or other components within the turbofan gas turbine engine. 
     For example, at least some known turbofan gas turbine engines include a fan base having a plurality of fan blades extending radially outwardly therefrom. The impact of a released blade upon a trailing blade may cause the trailing blade to rock about an axis tangential to rotation of the fan. The trailing blade initially rocks about the tangential axis toward a forward-section of the trailing blade such that the trailing blade may be dislodged radially outwardly away from its disk slot. The motion of the trailing blade about the tangential axis then reverses due to rotation of the fan, causing the trailing blade to rock backwards toward an aft end of the trailing blade. The rocking of the blade may induce compressive and tensile stresses in the blade. The magnitude of these tensile and compressive stresses in the trailing blade may exceed the failure threshold of the blade material causing the trailing blade to fail. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, a method is provided for fabricating a fan assembly for a gas turbine engine. The method includes forming a blade including an airfoil extending from an integral dovetail used to mount the blade within the rotor assembly, and extending a projection from at least a portion of the blade, such that the stresses induced within at least a portion of the blade are facilitated to be maintained below a predetermined failure threshold for the blade to facilitate preventing failure of the blade. 
     In another aspect, a gas turbine engine blade is provided that includes an airfoil, a dovetail formed integrally with said airfoil, and a projection that extends outwardly from at least one of the airfoil and the dovetail. The projection is configured to facilitate at least partially restricting movement of the blade to facilitate preventing failure of the blade. 
     In yet another aspect, a fan assembly for a gas turbine engine is provided. The fan assembly includes a fan hub, and at least one fan blade that extends radially outwardly from the fan hub. The fan blade includes a dovetail, an airfoil extending outwardly from the dovetail, and a projection that extends outwardly from the dovetail for maintaining stress induced within at least one of the dovetail and the airfoil below a predetermined failure threshold for the fan blade. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a schematic illustration of an exemplary turbofan gas turbine engine; 
     FIG. 2 is a perspective view of a portion an exemplary fan blade that may be included in the turbofan gas turbine engine shown in FIG. 1; 
     FIG. 3 is a cross-sectional view of a portion of the fan assembly shown in FIG.  1  and taken along line  3 — 3  of FIG. 2; and 
     FIG. 4 is a cross-sectional view of a portion of the fan assembly shown in FIG.  3  and taken along line  4 — 4  of FIG.  3 . 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     As used herein, the terms “failure” and “fail” may include any damage or other condition that at least partially impairs a component from functioning properly, such as, for example, any damage or other condition that at least partially impairs a component from functioning properly may include, but is not limited to, complete breakage of the component, partial breakage of the component, a change in the shape of the component, and a change in the properties of the component. The above examples are intended as exemplary only, and thus are not intended to limit in any way the definition and/or meaning of the terms “failure” and “fail”. In addition, although the invention is described herein in association with a turbofan gas turbine engine, and more specifically for use with a fan blade within a turbofan gas turbine engine, it should be understood that the present invention may be applicable to any component. Accordingly, practice of the present invention is not limited to fan blades or other components of turbofan gas turbine engines. 
     FIG. 1 is a schematic illustration of a turbofan gas turbine engine  10  including a fan assembly  12 , 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 a fan hub  24  having a plurality of disk slots (not shown in FIG. 1) therein and spaced circumferentially about fan hub  24 . Fan assembly  12  also includes an array of fan blades  30  that extend radially outward from the disk slots and fan hub  24  to a fan blade airfoil tip  32 . Fan assembly  12  rotates about an axis of rotation  40 . Engine  10  has an intake side  42  and an exhaust side  44 . In one embodiment, engine  10  is a GE-90 engine commercially available from General Electric Aircraft Engines, Cincinnati, Ohio. 
     In operation, air flows through fan assembly  12  and compressed air is supplied to high pressure compressor  14 . The highly compressed air is delivered to combustor  16  where it is mixed with fuel and ignited. The combustion gases are channeled from combustor  16  and used to drive turbines  18  and  20 , and turbine  20  drives fan assembly  12 . 
     FIG. 2 is a perspective view of a portion an exemplary fan blade  30  that may be used with fan assembly  12  (shown in FIG.  1 ). Each blade  30  includes a hollow airfoil  50  and an integral dovetail  52  that is used for mounting airfoil  50  to fan hub  24  in a known manner. Each airfoil  50  includes a first contoured sidewall  54  and a second contoured sidewall  56 . First sidewall  54  is convex and defines a suction side of airfoil  50 , and second sidewall  56  is concave and defines a pressure side of airfoil  50 . Sidewalls  54  and  56  are joined at a leading edge  58  and at an axially-spaced trailing edge  60  of airfoil  50 . More specifically, airfoil trailing edge  60  is spaced chordwise and downstream from airfoil leading edge  58 . First and second sidewalls  54  and  56 , respectively, extend longitudinally or radially outward in span from a blade root  62  positioned adjacent dovetail  52 , to airfoil tip  32  (shown in FIG.  1 ). Fan blade  30  extends a length 64 from a forward end  66  to an aft end  68 . Dovetail  52  includes a first pressure face contact surface  70  and a second pressure face contact surface  72 . 
     FIG. 3 is a cross-sectional view of a portion of fan assembly  12  taken along line  3 — 3  of FIG.  2 . FIG. 4 is a cross-sectional view of a portion of fan assembly  12  taken along line  4 — 4  of FIG.  3 . Specifically, within FIGS. 3 and 4, fan blade  30  is coupled within fan hub  24 . More specifically, fan blade  30  is received and secured, also referred to herein as seated, within a disk slot  74  defined in fan hub  24 . In one embodiment, fan hub  24  includes a plurality of disk slots  74  defined therein and spaced circumferentially about fan hub  24 . 
     Each disk slot  74  extends at least length 64 such that each dovetail  52  is completely received therein. When each fan blade dovetail  52  is seated within a respective disk slot  74 , each fan blade  30  extends radially outward from fan hub  24 . Disk slot  74  includes a radially inner surface  76 , and a portion  78  of disk slot  74  is shaped complimentary to a portion of dovetail  52 , such that when dovetail  52  is seated within disk slot  74 , first pressure face contact surface  70  is adjacent a first disk slot pressure surface  80 , and second pressure face contact surface  72  contacts a second disk slot pressure surface  82 . 
     In the exemplary embodiment, dovetail  52  includes a blade spacer  84  that extends outwardly from a radially inner surface  86  of dovetail  52 . Alternatively, dovetail  52  does not include spacer  84 . More specifically, spacer  84  extends radially inwardly towards fan hub  24  and disk slot radially inner surface  76 . When fan blade  30  is seated within disk slot  74 , blade spacer  84  extends a distance  88  from dovetail radially inner surface  86  such that a nominal blade/disk radial gap  90  is defined between a radially inner surface  92  of spacer  84  and disk slot radially inner surface  76 . In the exemplary embodiment, blade spacer  84  extends substantially across fan blade length 64. Alternatively, in another embodiment blade spacer  84  extends across only a portion of fan blade length 64. In the exemplary embodiment, blade spacer  84  is a separate component coupled dovetail  52 . In an alternative embodiment, blade spacer  84  is formed integrally with fan blade dovetail  52 . 
     Fan blade dovetail  52  also includes a projection  94  that extends outwardly from blade spacer  84 . More specifically, projection  94  extends from dovetail  52  and radially inwardly towards axis  40 , fan hub  24 , and disk slot radially inner surface  76 . When fan blade  30  is seated within disk slot  74 , projection  94  is positioned a distance  96  from blade spacer radially inner surface  92  such that a projection/disk slot radial gap  98  is defined between disk slot radially inner surface  76  and a radially inner surface  100  of projection  94 . In one embodiment, gap  90  is approximately equal 0.190 inches, and gap  98  is approximately equal 0.040 inches. 
     In the exemplary embodiment, projection  94  is a separate component coupled to, or frictionally coupled with, blade spacer  84 . In an alternative embodiment, projection  94  is formed integrally with blade spacer  84 . In one embodiment, fan blade  30  does not include blade spacer  84 , and rather projection  94  extends outwardly from dovetail radially inner surface  86  towards axis  40 , fan hub  24 , and disk slot radially inner surface  76 . In an alternative embodiment, fan blade  30  does not include blade spacer  84 , and projection  94  is either integrally formed with dovetail  52 , or is coupled to dovetail  52 . Projection  94  extends a distance  102  from fan blade aft end  68  toward fan blade forward end  66 . Although projection  94  is herein illustrated as extending distance  102  from aft end  68  toward forward end  66 , it should be understood that projection  94  may be positioned anywhere along blade spacer radially inner surface  92  to facilitate preventing failure of fan blade  30 , as described below. For example, in an alternative embodiment, projection  94  is positioned adjacent fan blade forward end  66 . 
     Fan assembly  12  includes an axis  104  that is tangential to disk slot radially inner surface  76 . Although axis  104  is herein illustrated as extending through a general center of fan blade length 64, it should be understood that axis  104  may extend through any portion of blade  30  along length 64, and tangentially to disk slot radially inner surface  76 . 
     During rotation of fan assembly  12 , when a blade mounted to fan hub  24  upstream from blade  30  fails, or is ejected from its respective disk slot, a condition herein referred to as “blade-out”, a portion of such a fan blade may impact fan blade  30 . Such contact may cause fan blade  30  to rock, or rotate about axis  104 . Specifically, initially, fan blade  30  rotates about axis  104  towards fan blade forward end  66  such that forward end  66  is forced radially inwardly towards disk slot radially inner surface  76 , and such that fan blade aft end  68  is forced radially outwardly away from disk slot radially inner surface  76 . More specifically, such impact may cause fan blade forward end  66  to partially unseat from disk slot  74 . As the stress wave, initiated by the release blade impact, is reflected and propagates through blade  30 , the rotational motion about axis  104  is reversed, thus causing fan blade  30  to rotate towards fan blade aft end  68  such that fan blade forward end  66  is forced radially outwardly away from disk slot radially inner surface  76 , and such that fan blade aft end  68  is forced radially inwardly toward disk slot radially inner surface  76 . More specifically, fan blade aft end  68  may partially unseat from disk slot  74 . 
     When fan blade aft end  68  is at least partially unseated from disk slot  74 , pressure between fan blade first pressure face contact surface  70  and first disk slot pressure surface  80 , and fan blade second pressure face contact surface  72  and second disk slot pressure surface  82 , is concentrated at fan blade forward end  66 . More specifically, a relatively high amount of compressive stress may be concentrated in fan blade aft end  68  and a relatively high amount of tensile stress may be concentrated in fan blade forward end  66 . The magnitude of these tensile and compressive stresses in fan blade  30  may exceed a predetermined failure threshold for at least a portion of fan blade  30 , thus causing fan blade  30  to partially or completely fail. However, projection  94  restricts movement of fan blade  30 , and more specifically restricts rotation of fan blade  30  about axis  104 , thus facilitating reducing tensile stresses that may be induced within fan blade forward end  66 . More specifically, as fan blade aft end  68  is unseated from disk slot  74 , projection  94  partially restricts inward radial displacement of fan blade aft end  68  such that only a limited amount of tensile stress may become concentrated in fan blade forward end  66 . Accordingly, projection  94  facilitates maintaining stress levels within fan blade  30  below a failure threshold of fan blade  30 . 
     The above-described tool is cost-effective and highly reliable for facilitating preventing failure of a component. The tool facilitates maintaining stresses induced within at least a portion of a component below a predetermined failure threshold of the component. More specifically, the tool at least partially restricts movement of a component to maintain tensile and compressive stresses within the component below a failure threshold of the component. As a result, the tool facilitates preventing failure of a component in a cost-effective and reliable manner. 
     Exemplary embodiments of blades and assemblies are described above in detail. The systems are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. Each blade and assembly component can also be used in combination with other tool and assembly components. 
     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.