Patent Publication Number: US-11391174-B2

Title: Vane arm clip for variable stator vanes

Description:
BACKGROUND 
     The present disclosure relates to a gas turbine engine and, more particularly, to a vane arm clip therefor. 
     Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. 
     Some gas turbine engines include variable stator vanes that can be pivoted about their individual axes to change an operational performance characteristic. Typically, the variable stator vanes are robustly designed to handle the stress loads that are applied to change the position of the vanes. A mechanical linkage is typically utilized to rotate the variable stator vanes. Because forces on the variable stator vanes can be relatively significant, forces transmitted through the mechanical linkage can also be relatively significant. Variable vanes are mounted about a pivot and are attached to an arm that is in turn actuated to adjust each of the vanes of a stage. A specific orientation between the arm and vane is required to assure that each vane in a stage is adjusted as desired to provide the desired engine operation. Newer compressor designs have resulted in higher compression ratios, loads and increased vane arm strength requirements that may be difficult to meet with sheet metal vane arm designs. 
     SUMMARY 
     A variable vane actuation system for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a stem section, along an axis, the stem section includes a groove; a vane arm comprising a claw section received at least partially into the groove; a clip mounted to the claw section; and a fastener fastened to the stem section along the axis to retain the clip and the vane arm to the stem section. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the clip comprises an aperture and the claw comprises an aperture along the axis. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the clip comprises a full hoop with a slot opposite the aperture, the slot fits into the groove. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the claw section comprises opposing fingers that are curved from an upper surface. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the opposing fingers each terminate at respective faces spaced from each other to define a claw opening received into the groove. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that a vane is rotatable about the axis. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the stem extends from an outer trunion of the vane. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the vane comprises an airfoil. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the clip is mounted to the claw section via an interference fit. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the clip is manufactured of sheet metal that is between 40-70 mils (1.0-1.8 mm) thick. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the stem section extends from a vane through an engine case. 
     A method of assembling a variable vane actuation system according to one disclosed non-limiting embodiment of the present disclosure includes locating a clip at least partially around a claw section of a vane arm. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that locating the clip at least partially around the claw section provides an interference fit between the clip and the claw section. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes retaining the clip to the claw section via a bolt threaded into a threaded bore of a stem section of a vane. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes threading the bolt into the threaded bore along an axis of rotation of the stem section. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes locating the claw section at least partially into a groove in the stem section prior to locating the clip at least partially around the claw section. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes locating the clip at least partially into the groove in the stem section. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that locating the clip at least partially into the groove in the stem section transverse to an axis of rotation of the stem section. 
     The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be appreciated; however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of an example gas turbine engine architecture. 
         FIG. 2  is a schematic view of a variable vane system for a gas turbine engine. 
         FIG. 3  is a partial perspective view of one stage of a variable vane system for a gas turbine engine. 
         FIG. 4  is a partial perspective view of a variable vane system for a gas turbine engine according to one disclosed non-limiting embodiment. 
         FIG. 5  is a sectional view of an interface between a one variable vane and a vane arm with the clip of  FIG. 7  according to one disclosed non-limiting embodiment. 
         FIG. 6  is a perspective view of a vane arm according to one disclosed non-limiting embodiment. 
         FIG. 7  is a perspective view of a vane arm clip according to one disclosed non-limiting embodiment. 
         FIG. 8  is a perspective view of a vane arm clip according to another disclosed non-limiting embodiment. 
         FIG. 9  is a sectional view of an interface between one variable vane and a vane arm with the clip of  FIG. 8 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool GTF (geared turbofan) that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engine architectures might include an augmentor section and exhaust duct section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion thru the turbine section  28 . Although depicted as a GTF in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with GTF as the teachings may be applied to other types of turbine engines such as a Direct-Drive-Turbofan with high, or low bypass augmented turbofan, turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an intermediate pressure compressor (“IPC”) between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an intermediate pressure turbine (“IPT”) between the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”). 
     The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing compartments  38 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  (“LPC”) and a low pressure turbine  46  (“LPT”). The inner shaft  40  drives the fan  42  directly or thru a geared architecture  48  to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. 
     The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  (“HPC”) and high pressure turbine  54  (“HPT”). A combustor  56  is arranged between the HPC  52  and the HPT  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     Core airflow is compressed by the LPC  44  then the HPC  52 , mixed with fuel and burned in the combustor  56 , then expanded over the HPT  54  and the LPT  46 . The turbines  54 ,  46  rotationally drive the respective low spool  30  and high spool  32  in response to the expansion. The main engine shafts  40 ,  50  are supported at a plurality of points by the bearing compartments  38 . It should be understood that various bearing compartments  38  at various locations may alternatively or additionally be provided. 
     In one example, the gas turbine engine  20  is a high-bypass geared aircraft engine with a bypass ratio greater than about six (6:1). The geared architecture  48  can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3:1, and in another example is greater than about 3.0:1. The geared turbofan enables operation of the low spool  30  at higher speeds which can increase the operational efficiency of the LPC  44  and LPT  46  to render increased pressure in relatively few stages. 
     A pressure ratio associated with the LPT  46  is pressure measured prior to the inlet of the LPT  46  as related to the pressure at the outlet of the LPT  46  prior to an exhaust nozzle of the gas turbine engine  20 . In one non-limiting embodiment, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the LPC  44 , and the LPT  46  has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans, where the rotational speed of the fan  42  is the same (1:1) of the LPC  44 . 
     In one example, a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10668 meters). This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The relatively low Fan Pressure Ratio according to one example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (“T”/518.7) 0.5  in which “T” represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
     With reference to  FIG. 2 , one or more stages of the LPC  44  and/or the HPC  52  include a variable vane system  100  that can be rotated to change an operational performance characteristic of the gas turbine engine  20  for different operating conditions. The variable vane system  100  may include one or more variable vane stages. 
     The variable vane system  100  may include a plurality of variable stator vanes  102  (also shown in  FIG. 3 ) circumferentially arranged around the engine central axis A. The variable stator vanes  102  each include a variable vane body that has an airfoil portion such that one side of the airfoil portion generally operates as a suction side and the opposing side of the airfoil portion generally operates as a pressure side. Each of the variable stator vanes  102  generally spans between an inner diameter and an outer diameter relative to the engine central axis A. 
     Each of the variable stator vanes  102  includes an inner trunion  104  that is receivable into a corresponding socket and an outer trunion  106  mounted through an outer engine case  108  such that each of the variable stator vanes  102  can pivot about a vane axis V ( FIG. 3 ). 
     The variable vane system  100  further includes a synchronizing ring assembly  110  to which, in one disclosed non-limiting embodiment, each of the outer trunions  106  are attached through a vane arm  112  along a respective axis D. It should be appreciated that although a particular vane arm  112  is disclosed in this embodiment, various linkages of various geometries may be utilized. 
     The variable vane system  100  is driven by an actuator system  118  with an actuator  120 , a drive  122  and an actuator arm  124  (also shown in  FIG. 4 ). Although particular components are separately described, it should be appreciated that alternative or additional components may be provided. 
     With reference to  FIG. 4 , the vane arm  112  links each outer trunion  106  to the synchronizing ring assembly  110 . Rotation of the synchronizing ring assembly  110  about the engine axis A ( FIG. 1 ) drives the vane arm  112  to rotate the outer trunion  106  of each of the variable stator vanes  102 . 
     Each vane arm  112  interfaces with the synchronizing ring assembly  110  via a pin  130 . The pin  130  is swaged to an end section  140  of the vane arm  112  within an aperture  142 . A collar  144  of the pin  130  may be utilized to locate the pin  130  at an appropriate depth prior to swaging. The pin  130  is received within a bushing  146  that fits within a sleeve  148  in the synchronizing ring assembly  110 . The bushing  146  permits the pin  130  and the vane arm  112  to rotate together relative to the synchronizing ring assembly  110 . 
     With reference to  FIG. 5 , each vane arm  112  ( FIG. 6 ) also interfaces with its respective stator vane  102  via a claw section  160  opposite the end section  140  to rotate the outer trunion  106 . The claw section  160  includes opposing fingers  162 ,  164  that are curved from an upper surface  166  and terminate at respective opposed faces  168 ,  170 . The upper surface  166  extends from an arm  174  and includes an aperture  172 . Each vane arm  112  may be manufactured of a sheet metal material that is between 40-70 mils (1.0-1.8 mm) thick. The opposed faces  168 ,  170  are spaced to provide an opening  176  that interfaces with the outer trunion  106 . 
     The outer trunion  106  includes a stem section  180  that forms a groove  182 . An outer diameter of the stem section  180  is greater than the diameter of the groove  182 . A threaded bore  186  is defined through the outer trunion  106  along the axis T. 
     A clip  190  ( FIG. 7 ) includes opposing arms  192 ,  194  that are curved from a clip upper surface  196  to be essentially perpendicular thereto. The clip upper surface  196  includes a clip aperture  198  that corresponds with the aperture  172 . That, is the clip  190  is essentially bracket shaped to fit over the claw section  160  ( FIG. 7 ). The clip  190  may be manufactured of nickel sheet metal that is between 40-70 mils (1.0-1.8 mm) thick. 
     The threaded bore  186  receives a bolt  188  that retains the clip  190  and the vane arm  112  to the outer trunion  106  such that opposed faces  168 ,  170  engage the groove  182 . That is, the bolt  188  extends through the clip aperture  198  and the aperture  172 , then threads into the threaded bore  186  along axis T. 
     The clip  190  may be fitted to the claw section  160  via an interference fit  200  across the opposing fingers  162 ,  164  to provide strength during all operating conditions or fitted with small clearance fit to provide additional strength only in a high load situation such as a surge condition. The clip  190  provides strength much like a leaf spring. 
     With reference to  FIG. 8 , a clip  190 A according to another embodiment is a full hoop. The clip  190 A includes a clip aperture  198  in an upper surface  196  and a slot  197  in a lower surface  199 . The upper surface  196  and the lower surface  199  include radiused arms  210 ,  212  therebetween that fit around the opposing fingers  162 ,  164  ( FIG. 9 ). 
     The clip aperture  198  receives the bolt  188  and the slot fits into the groove  182 . In this embodiment, the clip  190 A is installed transverse to the axis T. The clip  190 A thereby provides full circumferential support to the claw section  160 . A spacer  310  may be provided for clearance. 
     The clip  190 ,  190 A provides strength to the claw section  160  much like a leaf spring. The clip  190 ,  190 A is inexpensive and relatively easy to manufacture that significantly supplements the strength of existing sheet metal vane arm design. In addition, as the clip is assembled to the vane arm, the clip  190 ,  190 A provides resistance to rotation caused by run on torque assembly forces. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason, the appended claims should be studied to determine true scope and content.