Abstract:
An example gas turbine engine compressor includes a first compressor section. The first compressor section includes rotating stage that includes rotating blades and a stationary stage upstream thereof that includes stationary guide vanes. The stationary vanes controllably pivot about respective pivot axises for providing flow control into the rotating stage.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/708,076, which was filed on 1 Oct. 2012 and is incorporated herein by reference. 
     
    
     BACKGROUND 
       [0002]    This disclosure relates generally to a compressor section of a gas turbine engine and, more particularly, to variable vanes that influence flow to a low pressure compressor of the gas turbine engine. 
         [0003]    A gas turbine engine typically includes a fan section, 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 the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
         [0004]    The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine, and fan section rotate at a common speed in a common direction. 
         [0005]    A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds. 
         [0006]    Although geared architectures have improved propulsive efficiency, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer, and propulsive efficiencies. 
       SUMMARY 
       [0007]    A gas turbine engine compressor according to an exemplary aspect of the present disclosure includes, among other things, a first compressor section, the first compressor section including at least one rotating stage that includes rotating blades and at least one stationary stage upstream thereof that includes stationary guide vanes, which controllably pivot about respective pivot axises for providing flow control into the rotating stage. 
         [0008]    In a further non-limiting embodiment of the foregoing gas turbine engine compressor, the first compressor section is a low pressure compressor section and the gas turbine engine compressor further comprises a second compressor section that is a high pressure section. The low pressure compressor section may experience lower pressures than the high pressure compressor section during operation. 
         [0009]    In a further non-limiting embodiment of either of the foregoing gas turbine engine compressors, the first compressor section may be an axially forwardmost compressor section of the gas turbine engine relative to a direction of flow through the gas turbine engine. 
         [0010]    In a further non-limiting embodiment of any of the foregoing gas turbine engine compressors, the stationary vane stage may be the axially forwardmost vane stage of the first compressor section. 
         [0011]    In a further non-limiting embodiment of any of the foregoing gas turbine engine compressors, a first stage of the first compressor section may be the stationary stage. 
         [0012]    In a further non-limiting embodiment of any of the foregoing gas turbine engine compressors, the first compressor section may be operatively coupled to a fan drive shaft of the gas turbine engine. 
         [0013]    In a further non-limiting embodiment of any of the foregoing gas turbine engine compressors, the fan drive shaft may be operatively coupled to a geared architecture configured to drive a fan of the gas turbine engine at a different rotational speed than a rotational speed of the fan drive shaft. 
         [0014]    In a further non-limiting embodiment of any of the foregoing gas turbine engine compressors, the low pressure compressor may be positioned axially between a fan of the gas turbine engine and a high pressure compressor of the gas turbine engine. 
         [0015]    In a further non-limiting embodiment of any of the foregoing gas turbine engine compressors, the pivotable vanes are inlet guide vanes. 
         [0016]    A method of controlling flow into a compressor of a gas turbine engine, wherein the compressor has a first compressor section, at least one rotating stage that includes rotating blades and at least one stationary stage upstream thereof that includes stationary guide vanes, which controllably pivot about respective pivot axises for providing flow control into the rotation stage. according to another exemplary aspect of the present disclosure includes, among other things, pivoting the guide vanes to influence flow to the rotating blades. 
         [0017]    In a further non-limiting embodiment of the foregoing method of controlling flow, the stationary vanes may form a portion of a first stage of the compressor. 
         [0018]    In a further non-limiting embodiment of either of the foregoing methods of controlling flow, the method includes pivoting the stationary vanes from a first position to a second position to influence the flow, the first position defining a first throat area in the compressor, the second position corresponding to a second throat area in the compressor that may be between 62 percent and 65 percent of the first throat area. 
         [0019]    A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan including a plurality of fan blades rotatable about an axis; a compressor section including a first compressor section; a combustor in fluid communication with the compressor section; a turbine section in fluid communication with the combustor; a geared architecture driven by the turbine section for rotating the fan about the axis; and the first compressor section, and at least one rotating stage that includes rotating blades and at least one stationary stage upstream thereof that includes stationary guide vanes, which controllably pivot about respective pivot axises for providing flow control into the rotation stage. 
         [0020]    In a further non-limiting embodiment of the foregoing gas turbine engine, the first compressor section may be a low pressure section and the engine further comprises a second compressor section that may be a high pressure section, the low pressure compressor section experiences lower pressures than the high pressure compressor section during operation. 
         [0021]    In a further non-limiting embodiment of either of the foregoing gas turbine engines, the stationary vane stage may be the forwardmost stage of the low pressure compressor relative to a direction of flow through the engine. 
         [0022]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the stationary vanes may be configured to move from a first position to a second position to influence the flow, the first position corresponding to a first compressor throat area, the second position corresponding to a second compressor throat area that may be between 62 percent and 65 percent of the first throat area. 
         [0023]    Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     
    
     
       DESCRIPTION OF THE FIGURES 
         [0024]    The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the detailed description. The figures that accompany the detailed description can be briefly described as follows: 
           [0025]      FIG. 1  shows a section view of an example gas turbine engine. 
           [0026]      FIG. 2  shows a close up section view of a low pressure compressor of the gas turbine engine of  FIG. 1 . 
           [0027]      FIG. 3  shows a variable vane assembly from the low pressure compressor of  FIG. 2 . 
           [0028]      FIG. 4  shows a section view of variable vanes of the variable vane assembly of  FIG. 3  in a first position. 
           [0029]      FIG. 5  shows a section view of variable vanes of the variable vane assembly of  FIG. 3  in a second position that restricts more flow into the low pressure compressor than the first position. 
       
    
    
     DETAILED DESCRIPTION 
       [0030]      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 . 
         [0031]    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. 
         [0032]    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. 
         [0033]    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. 
         [0034]    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. 
         [0035]    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. 
         [0036]    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 . 
         [0037]    The core airflow 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. 
         [0038]    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), with an example embodiment being greater than about ten (10). 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. 
         [0039]    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. 
         [0040]    A significant amount of thrust is provided by the bypass flow 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 (&#39;TSFC&#39;)”—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. 
         [0041]    “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. 
         [0042]    “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. 
         [0043]    The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section  22  includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about 6 turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about 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. 
         [0044]    Referring to  FIGS. 2 and 3  with continuing reference to  FIG. 1 , the example low pressure compressor  44  includes a variable vane assembly  62  having a plurality of radially extending variable vanes  68 . 
         [0045]    The low pressure compressor  44  is considered a low pressure compressor of the engine  20  as it experiences lower pressures during operation than the high pressure compressor  52  of the engine  20 . The example low pressure compressor  44  is positioned axially between the fan  42  of the engine  20  and the high pressure compressor  52  of the engine  20 . 
         [0046]    Notably, the low pressure compressor  44  is driven by the low speed spool  30 , which is operably coupled to the geared architecture  48  of the engine  20 . The low speed spool  30  thus includes portions that function as a fan drive shaft as the low speed spool  30  rotates the geared architecture  48  to drive the fan  42 . 
         [0047]    In this example, the variable vane assembly  62  provides the axially forwardmost stage of the low pressure compressor  44 . More specifically, in this example, the vanes  68  are inlet guide vanes and the forwardmost vanes of the low pressure compressor  44 . 
         [0048]    Each of the vanes  68  is rotatable about a respective radially extending axis, such as the axis R, to influence flow into the low pressure compressor  44 . The axis R extends radially from the axis A. Each of the vanes  68  may be rotated about its axis R between positions that permit more flow and positions that permit less flow to tailor flow into the low pressure compressor  44  to balance system operability and enhance performance. 
         [0049]    The example vanes  68  are pivoted via a pivoting mechanism that has an arm  76 . An actuator  78  moves the arm  76  to rotate the vanes  68  about their respective axises. A Full Authority Digital Engine Control (FADEC) is schematically illustrated at  80 . The FADEC  80  controls the actuator  78  to control pivoting of the vanes  68 . 
         [0050]    In this example, the positioning of the vanes  68  is controlled as a function of corrected low pressure compressor speed. In some examples, at low power settings, the vanes  68  are moved to a more closed position. At higher rotational speeds, the vanes  68  are rotated to a more open position. The more closed position permits less flow through the low pressure compressor  44  than the more open position. 
         [0051]    Referring now to  FIGS. 4 and 5  with continuing reference to  FIGS. 2 and 3 , a top view cutaway of an example embodiment of the variable vane assembly  62  includes adjacent variable vanes  68   a,    68   b  and  68   c.  The vanes  68   a - 68   c  are attached to a stationary portion of the gas turbine engine  20 , such as a case structure (not shown). The vanes  68   a - 68   c  have a suction surface  90  and a pressure surface  94 . During operation of the engine  20 , flow moving along the core flow path C moves into the low pressure compressor  44  between adjacent ones of the vanes  68   a - 68   c.  The adjacent vanes define a throat area T, which represents the minimal area between adjacent ones of the vanes  68   a - 68   c.  Flow moves into the low pressure compressor  44  through the throat area T. 
         [0052]    Various factors can influence the location and size of the throat area T. For example, the shape of the vanes  68   a - 68   c,  the stagger angle of the vanes  68   a - 68   c  relative to the core flow path C, and the orientation of the vanes  68   a - 68   c  are all possible factors that can influence the size of the throat area T. 
         [0053]      FIG. 4  shows the vanes  68   a - 68   c  when the low pressure compressor  44  is operating at a relatively high rotational speed.  FIG. 5  shows the vanes  68   a - 68   c  when the low pressure compressor  44  is operating at a relatively low rotational speed. The vanes  68   a - 68   c  are shown in a more open position in  FIG. 4  than in  FIG. 5 . The more open position corresponds to the low pressure compressor  44  operating at the relatively high rotational speed. The more closed position corresponds to the low pressure compressor  44  operating at the relatively low rotational speed. When the vanes  68   a - 68   c  are in a more open position, the throat area T is greater than when the vanes  68   a - 68   c  are in a more closed position. 
         [0054]    The shapes of the vanes  68   a - 68   c  is an illustration of one possible embodiment. The shape of the vanes  68   a - 68   c  may vary depending on, for example, the components of the low pressure compressor  44  to which the vanes  68   a - 68   c  are attached, the location of the vanes  68   a - 68   c  within the low pressure compressor  44 , gas path flow velocities, desired design characteristics of the engine  20 , and materials used in fabricating the gas turbine engine  20 . 
         [0055]    In this example,  FIG. 4  represents the vanes  68   a - 68   c  when they are in their maximum open position.  FIG. 5 , by contrast, represents the vanes  68   a - 68   c  in the maximum closed position. The throat area T between the vanes  68   a - 68   c  in the maximum closed position is between 62 percent and 65 percent of the throat area when the vanes are in the maximum open position of  FIG. 4 . The amount of rotation between the maximum closed position and the maximum open position is from −37 degrees to +18 degrees in this example. 
         [0056]    Geared gas turbine engines are unique in that the low pressure compressor  44  rotates at an increased speed compared to a low pressure compressor of prior art direct drive turbine engines. The increased rotational speed of the low pressure compressor  44  leads to different compressor behavior and operation than the low pressure compressors of direct drive engines. 
         [0057]    The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.