Abstract:
A method of controlling a variable vane assembly includes the steps of sensing a first angular deflection of a first array of variable vanes about a first vane axis, and a second angular deflection of a second array of variable vanes about a second vane axis, the first array of variable vanes axially spaced from the second array of variable vanes, and adjusting the angular deflection of one of the first and second arrays of variable vanes, based on the sensed angular deflections from the other of the first and second arrays of variable vanes. A compressor including the variable vane assembly and a method of operating the variable vane assembly for a compressor are also disclosed.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a variable vane drive system for a gas turbine engine. 
         [0002]    A gas turbine engine typically includes a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive at least the compressor. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. Some gas turbine engines include a fan section driven by the turbine section. 
         [0003]    Some areas of the engine may include variable vanes. The compressor, for example, may include multiple stages of variable vanes. In some compressor designs, vanes can only be scheduled at first and second positions. The first and second positions are determined by on-ground testing. During flight, the vanes are moved between the first and second positions based on the engine&#39;s mode of operation. The vanes may experience wear on primary air-contacting faces in the first and second positions and may need to be replaced if the wear interferes with the engine performance. 
       SUMMARY 
       [0004]    A method of controlling a variable vane assembly according to an exemplary aspect of the present disclosure includes, among other things, sensing a first angular deflection of a first array of variable vanes about a first vane axis, and a second angular deflection of a second array of variable vanes about a second vane axis, the first array of variable vanes axially spaced from the second array of variable vanes, and adjusting the angular deflection of one of the first and second arrays of variable vanes, based on the sensed angular deflections from the other of the first and second arrays of variable vanes. 
         [0005]    In a further non-limiting embodiment of the foregoing method, the second array of variable vanes comprises inlet guide vanes. 
         [0006]    In a further non-limiting embodiment of any of the foregoing methods, the method further includes the step of measuring an updated angular deflection of the first array of variable vanes with respect to the first vane axis subsequent to performing the adjusting step. 
         [0007]    In a further non-limiting embodiment of any of the foregoing methods, the method includes measuring using rotary variable differential transformers. 
         [0008]    In a further non-limiting embodiment of any of the foregoing methods, the method further includes the step of calculating whether the first array of variable vanes needs to be adjusted. 
         [0009]    In a further non-limiting embodiment of any of the foregoing methods, the method includes calculating and adjusting steps using a full authority digital engine controller. 
         [0010]    In a further non-limiting embodiment of any of the foregoing methods, the method includes adjusting with a bellcrank. 
         [0011]    In a further non-limiting embodiment of any of the foregoing methods, the method further includes the step of inputting a desired angular deflection of the first array of variable vanes. 
         [0012]    A compressor for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a plurality of stages, each stage including a plurality of vane arms, wherein the vane arms each include a portion that engages a variable vane and are secured to at least one movable annular ring, at least one sensor arranged at each stage, the at least one sensor configured to measure an angular position of the variable vanes, and a controller configured adjust the at least one movable annular ring based on the angular position measured by the sensor. 
         [0013]    In a further non-limiting embodiment of the foregoing compressor, the at least one movable annular ring includes first and second movable annular rings and the vane arms include first and second ends configured to be secured to the first and second annular rings, respectively. 
         [0014]    In a further non-limiting embodiment of any of the foregoing compressors, the compressor further includes an actuator arranged at each stage, the actuator configured to move the at least one annular ring, thereby moving the variable vanes. 
         [0015]    In a further non-limiting embodiment of any of the foregoing compressors, the controller controls the actuator of each stage independently. 
         [0016]    In a further non-limiting embodiment of any of the foregoing compressors, the compressor further includes a plurality of inlet guide vanes arranged upstream from the foremost stage of the plurality of stages. 
         [0017]    In a further non-limiting embodiment of any of the foregoing compressors, at least one of the plurality of inlet guide vanes includes an inlet guide vane sensor configured to determine an angular position of the at least one inlet guide vane. 
         [0018]    In a further non-limiting embodiment of any of the foregoing compressors, the controller is configured to control the actuators based on information from the inlet guide vane sensor. 
         [0019]    A method of controlling a variable vane assembly for a compressor, according to an exemplary aspect of the present invention includes, among other things, securing a first plurality of variable vanes to a plurality of vane arms, the vane arms secured to a first movable annular ring at a first end and a second movable annular ring at a second end, measuring an angular deflection of the first plurality of variable vanes with respect to a vane axis, and moving the first and second annular rings in response to the measured angular deflection of the first plurality of variable vanes. 
         [0020]    In a further non-limiting embodiment of the foregoing method, the method further includes the step of measuring an angular deflection of a second plurality of variable vanes. 
         [0021]    In a further non-limiting embodiment of any of the foregoing methods, the moving step is in response to the measured angular deflection of the second plurality of variable vanes. 
         [0022]    In a further non-limiting embodiment of any of the foregoing methods, the second plurality of variable vanes are inlet guide vanes. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0023]      FIG. 1  schematically illustrates an example gas turbine engine with variable vane control system. 
           [0024]      FIG. 2  schematically illustrates an example compressor with variable vane control system. 
           [0025]      FIG. 3  illustrates a flowchart of method for controlling variable vanes. 
           [0026]      FIG. 4  illustrates a portion of alternate example compressor. 
       
    
    
     DETAILED DESCRIPTION 
       [0027]      FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26 , and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
         [0028]    Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
         [0029]    The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
         [0030]    The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
         [0031]    A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0032]    The example low pressure turbine  46  has a pressure ratio that is greater than about  5 . The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
         [0033]    A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
         [0034]    The core airflow flowpath C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes vanes  60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  58 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
         [0035]    The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6:1), with an example embodiment being greater than about ten (10:1). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
         [0036]    In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
         [0037]    A significant amount of thrust is provided by air in the bypass flowpath B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
         [0038]    “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment, the low fan pressure ratio is less than about 1.45. 
         [0039]    “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7 ° R)]̂0.5. The “Low corrected fan tip speed,” as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
         [0040]    The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about twenty-six (26) fan blades. In another non-limiting embodiment, the fan section  22  includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about six (6) turbine rotors schematically indicated at  34 . In another non-limiting example embodiment, the low pressure turbine  46  includes about three (3) turbine rotors. A ratio between the number of fan blades and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
         [0041]    Referring to  FIG. 2  with continuing reference to  FIG. 1 , the high pressure compressor  52  can include a compressor case  68  surrounding one or more stages. In the example shown in  FIG. 2 , the high pressure compressor  52  includes first, second, and third stages  62 ,  64 ,  66 . The example high pressure compressor  52  includes inlet guide vanes (IGVs)  98 , located at the compressor  52  inlet and extending radially relative to the engine axis A along an IGV axis I. The IGVs  98  help direct air into the high pressure compressor  52 . The IGVs  98  are arranged upstream from the first compressor stage  62 . In one example, the example IGVs  98  are a first set of variable vanes, pivotable along the IGV axis I, which extends radially out of the engine core axis A, by an actuator  99 . 
         [0042]    The high pressure compressor  52  includes second set of variable vanes  70  aft of the IGVs  98  extending radially relative to the engine axis A along a variable vane axis V. The variable vane axis V extends into the compressor parallel to IGV axis I. The variable vanes each have a vane arm  72  secured to a synchronizing ring (sync-ring)  74 . The vane arm  72  is connected to a trunnion (not shown) which is in turn connected to an airfoil (not shown). The trunnion and airfoil extend radially into the compressor  52 . As the sync-ring  72  rotates, the vane arm  72  is actuated such that it pivots about an axis of the trunnion. This in turn causes the trunnion and thus the airfoil to rotate about the same trunnion axis. Each stage  62 ,  64 ,  66  includes an actuator (A)  90 , which is connected to a bellcrank  92 . The bellcrank  92  is actuated by the actuator  90  to rotate the sync-ring  74 , thus rotating the variable vanes  92 . 
         [0043]    In the disclosed embodiment, each of the stages  62 ,  64 ,  66  includes a sensor  100  on one or more of the variable vanes  70  in that stage. The sensors  100  measure the angular displacement of the variable vane  70  with respect to vane axis V. The IGVs  98  may also include a sensor  100  to measure the displacement of the IGVs  98  with respect to the IGV axis I. The sensors  100  are, for example, rotary variable differential transformers (RVDTs). 
         [0044]    Displacement information for the variable vanes  70  and IGVs  98 , measured from the sensors  100 , is communicated to a controller  102 . The controller  102  is, for example, an engine  20  controller, such as a full authority digital engine controller (FADEC). The controller  102  signals the actuators  90  to actuate the bellcranks  92 , and adjusts the variable vanes  70  in response to the angular displacement information. 
         [0045]    Referring to  FIG. 3  with continued reference to  FIG. 2 , a method for controlling variable vanes  70  within a compressor  52  is shown. In step  200 , a displacement angle of IGVs  98  in the compressor  52  is measured. In step  202 , a displacement angle of the variable vanes  70  aft of the IGVs  98  is measured. The displacement angle of the variable vanes  70  may be measured in selected compressor stages  62 ,  64 ,  66  or in every compressor stage  62 ,  64 ,  66 . 
         [0046]    In step  204 , the controller  102  calculates whether the variable vanes  70  need to be adjusted. The variable vanes  70  may be adjusted to improve engine  20  efficiency. For example, the certain modes of operation for the engine  20 , such as takeoff, landing, cruise, etc. may require certain variable vane  70  scheduling. Alternatively, variable vanes  70  may be rotated to provide a certain face for primary impingement of the air in the core air flowpath C through the compressor  52 . For example, variable vanes  70  may experience wear on a first face and may then be rotated to provide a second face for impingement of the core air. Variable vanes  70  can wear for a variety of reasons, such as foreign object damage, erosion, and debris accumulation. A slight amount of wearing may occur from these various issues, though within acceptable guidelines, but which may impact efficiency; the impact of such wear can be lessened by rotating the variable vanes  70 . This may increase the lifetime of the variable vanes  70 . 
         [0047]    If the variable vanes  70  do not need to be adjusted, as indicated by the “No” logic step  204   a,  the method returns to step  202 . If the variable vanes  70  do need to be adjusted, as indicated by the “Yes” logic step  204   b,  then the method proceeds to steps  206 ,  208 , and  210 . In step  206  the variable vanes  70  in the first stage  62  are adjusted. In step  208 , the variable vanes  70  in the second stage  64  are adjusted. In step  210 , the variable vanes  70  in the third stage  66  are adjusted. The variable vanes  70  in each stage  62 ,  64 ,  66  may be adjusted independently of one another by the actuators  90 . After adjustment, the method returns to step  200 . This closed-loop feedback control method allows for infinite incremental adjustments of the variable vane  70  angles. That is, the variable vanes  70  can be adjusted to any position with the range of motion allowed by the bellcrank  92 . Also, the control method allows for real time variable vane  70  adjustment during flight in response a user input such as a change in flight conditions. Alternatively, the variable vanes  70  may be automatically positioned at a certain angle based on the difference between the current variable vane  70  angle and a desired variable vane  70  angle. 
         [0048]      FIG. 4  shows an alternate compressor  52 ′. In the example of  FIG. 4 , each variable vane  70 ′ includes a vane arm  72 ′ which is connected to first and second sync-rings  74   a,    74   b  at first and second ends of the vane arm  72 ′, respectively. A bellcrank  92 ′ is connected to the first and second sync-rings  74   a,    74   b.  In operation, the sync-rings  74   a,    74   b  are rotated circumferentially about the engine axis A ( FIG. 1 ) by the actuator  90  via the bellcrank  92 ′. The sync-rings  74   a,    74   b  are rotated in opposite directions. This provides circumferential forces to first and second ends of the vane arm  72 ′, respectively. Applying these forces causes the vane arm  72 ′ to pivot about a vane axis V. This in turn rotates the variable vane  70 ′. The method of  FIG. 3  may be applied to the alternate compressor  52 ′ as well. 
         [0049]    While the variable vane actuation systems are described herein in the context of the high pressure compressor  52 ,  52 ′, it should be understood that the variable vane actuation system may be used in other parts of the engine which include variable vanes, for example, the high or low pressure turbines  46 ,  54 . 
         [0050]    Although embodiments of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.