Patent Publication Number: US-10323536-B2

Title: Active clearance control for axial rotor systems

Description:
FIELD 
     The present disclosure relates generally to axial rotor systems of a gas turbine engine and, more particularly, to a stator assembly capable of moving forward and aft relative to a rotor assembly. 
     BACKGROUND 
     Gas turbine engines typically include compressors having multiple rows, or stages, of rotating blades and multiple stages of stators. The rotating blades rotate about an axis while the stators are fixed such that they do not rotate about the axis. A gap can exist between an outer diameter edge of the rotors and an outer diameter edge of the stators. The size of this gap affects the efficiency of the compressor as the smaller the gap is, the less the pressure loss occurs. However, elimination of this gap would be detrimental because the compressor is occasionally subjected to external forces, such as aerodynamic maneuvers, unbalanced loads of the rotors, thermal expansion of the rotors or the stators or the like. 
     SUMMARY 
     What is described is a system for increasing efficiency of a gas turbine engine. The system includes a stator assembly including at least one stator airfoil. The system also includes a rotor assembly including at least one rotor airfoil configured to rotate about an axis. The system also includes an actuator coupled to the stator assembly and configured to actuate the stator assembly in an axial direction relative to the rotor assembly, creating an axial movement such that a clearance between the at least one rotor airfoil and the stator assembly varies based on an axial position of the stator assembly. 
     Also described is a system for increasing efficiency of a compressor section of a gas turbine engine. The system includes a rotor assembly including a rotor outer diameter edge, and at least one rotor airfoil configured to rotate about an axis and to compress a fluid. The system also includes a stator assembly including a stator outer diameter edge, and a stator airfoil configured to condition the fluid, such that the rotor outer diameter edge and the stator outer diameter edge define a conic shape. The system also includes an actuator coupled to the stator assembly and configured to actuate the stator assembly in an axial direction relative to the rotor assembly, creating an axial movement such that a clearance between the at least one rotor airfoil and the stator assembly varies based on an axial position of the stator assembly. 
     Also described is a method for increasing efficiency of a compressor. The method includes receiving, by a controller, an input indicating an amount of force to be applied to the compressor. The method also includes determining, by the controller, a determined direction and a determined amount to move a stator assembly in an axial direction relative to a rotor assembly based on the input. The method also includes instructing, by the controller, an actuator coupled to the stator assembly to actuate the stator assembly the determined amount in the determined direction. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, is best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements. 
         FIG. 1  illustrates cross-sectional view of an exemplary gas turbine engine, in accordance with various embodiments; 
         FIG. 2  illustrates a cross-sectional view of a low pressure compressor section of the gas turbine engine of  FIG. 1 , in accordance with various embodiments; 
         FIG. 3  illustrates a cross-sectional view of two axial positions of a stator assembly relative to an outer diameter edge of a rotor, in accordance with various embodiments; 
         FIG. 4  illustrates a controller coupled to an actuator of the low pressure compressor section of  FIG. 2 , in accordance with various embodiments; and 
         FIG. 5  illustrates flowchart corresponding to a method to be performed by the controller of  FIG. 4 , in accordance with various embodiments. 
     
    
    
     DETAILED DESCRIPTION 
     With reference to  FIG. 1 , a gas turbine engine  20  is provided. An A-R-C axis illustrated in each of the figures illustrates the axial (A), radial (R) and circumferential (C) directions. As used herein, “aft” refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine engine. As used herein, “forward” refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion. As utilized herein, radially inward refers to the negative R direction and radially outward refers to the R direction. 
     Gas turbine engine  20  can be a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines include an augmentor section among other systems or features. In operation, fan section  22  drives coolant along a bypass flow-path B while compressor section  24  drives coolant along a core flow-path C for compression and communication into combustor section  26  then expansion through turbine section  28 . Although depicted as a turbofan gas turbine engine  20  herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings can be applied to other types of turbine engines including three-spool architectures. 
     Gas turbine engine  20  generally comprise a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure  36  via several bearing systems  38 ,  38 - 1 , and  38 - 2 . It should be understood that various bearing systems  38  at various locations can alternatively or additionally be provided, including for example, bearing system  38 , bearing system  38 - 1 , and bearing system  38 - 2 . 
     Low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure (or first) compressor section  44  and a low pressure (or first) turbine section  46  inner shaft  40  is connected to fan  42  through a geared architecture  48  that can drive fan  42  at a lower speed than low speed spool  30 . Geared architecture  48  includes a gear assembly  60  enclosed within a gear housing  62 . Gear assembly  60  couples inner shaft  40  to a rotating fan structure. High speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and high pressure (or second) turbine section  54 . A combustor  56  is located between high pressure compressor  52  and high pressure turbine  54 . A mid-turbine frame  57  of engine static structure  36  is located generally between high pressure turbine  54  and low pressure turbine  46 . Mid-turbine frame  57  supports one or more bearing systems  38  in turbine section  28 . Inner shaft  40  and outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The core airflow C is compressed by low pressure compressor section  44  then high pressure compressor  52 , mixed and burned with fuel in combustor  56 , then expanded over high pressure turbine  54  and low pressure turbine  46 . Mid-turbine frame  57  includes airfoils  59  which are in the core airflow path. Turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     Gas turbine engine  20  is a high-bypass geared aircraft engine. The bypass ratio of gas turbine engine  20  can be greater than about six (6). The bypass ratio of gas turbine engine  20  can also be greater than ten (10). Geared architecture  48  can be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture  48  can have a gear reduction ratio of greater than about 2.3 and low pressure turbine  46  can have a pressure ratio that is greater than about five (5). The bypass ratio of gas turbine engine  20  can be greater than about ten (10:1). The diameter of fan  42  can be significantly larger than that of the low pressure compressor section  44 , and the low pressure turbine  46  can have a pressure ratio that is greater than about five (5:1). Low pressure turbine  46  pressure ratio is measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of low pressure turbine  46  prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of particular embodiments of a suitable geared architecture engine and that the present disclosure contemplates other turbine engines including direct drive turbofans. 
     The next generation of turbofan engines are designed for higher efficiency and use higher pressure ratios and higher temperatures in high pressure compressor  52  than are conventionally experienced. These higher operating temperatures and pressure ratios create operating environments that cause thermal loads that are higher than the thermal loads conventionally experienced, which occasionally shortens the operational life of current components. 
     With reference now to  FIG. 2 , low pressure compressor section  44  includes a rotor assembly  206  and a stator assembly  210 . Fluid flows aft into low pressure compressor section  44  as indicated by arrow  220  where it is initially conditioned by a guide vane  200 . A rotor  202  coupled to rotor assembly  206  propels the fluid aft by rotating about the A axis. After being propelled by rotor  202 , the fluid is again conditioned by a guide vane  204 . Guide vane  200  and guide vane  204  are coupled to a case  222  and are stationary relative to the rotating rotor  202 . 
     After conditioning by guide vane  204 , the fluid is propelled aft (i.e., compressed) by a rotor  208 A, conditioned by a stator  212 A, propelled aft by a rotor  208 B, conditioned by a stator  212 B, propelled aft by a rotor  208 C, conditioned by a stator  212 C, propelled aft by a rotor  208 D, conditioned by a stator  212 D and propelled aft by a rotor  208 D. In that regard, low pressure compressor section  44  includes five stages of rotors  208  separated by four stators  212 . The rotors  208  rotate about the A axis while the stators  212  do not rotate about the A axis. Case  222  circumferentially surrounds each of the rotors and stators. 
     Stator assembly  210  has an outer diameter edge  216  from which the stators  212  extend radially inward to an inner diameter edge  217  defined by the radially inner edges of stators  212 . Rotor assembly  206  includes an inner diameter edge  215  from which rotors  208  extend radially outward to an outer diameter edge  214  defined by the radially outer edges of rotors  208 . 
     It is desirable for a distance  260  between outer diameter edge  216  of stator assembly  210  and outer diameter edge  214  of rotor assembly  206  to be small. As fluid is propelled aft, pressure builds between each stage of low pressure compressor section  44 . As distance  260  increases, more air leaks forward between each stage. However, it is preferable for distance  260  to be greater than zero as it is desirable to include room for tolerances. As gas turbine engine  20  is in use and being maneuvered, loads, or forces, are applied to rotor assembly  206  that cause rotor assembly  206  to move in the radial direction. These loads include maneuver loads, the normal pulling of rotors  208  as it rotates due to non-centered weights, differential thermal growth between rotor assembly  206  and stator assembly  210  and the like. Accordingly, distance  260  is selected so that rotor assembly  206  and stator assembly  210  are unlikely to make contact during normal operating conditions. 
     A tie shaft  205  holds rotor  202  and rotors  208  together axially so they do not separate in the axial direction. A bearing  218  is coupled to case  222  and resists radial force of rotor assembly  206  to reduce the likelihood of rotor assembly  206  changing position radially relative to case  222 . A ball bearing resists radial force of rotor assembly  206  to further reduce the likelihood of rotor assembly  206  changing position radially relative to case  222 . The ball bearing allows rotor assembly  206  to expand in the aft direction due to thermal and pressure forces. 
     A forward end  266  of stator assembly  210  is coupled to an actuator  228 . A forward sliding seal  232  allows stator assembly  210  to move forward and aft while forming a seal with case  222 . Similarly, an aft end  268  of stator assembly  210  is coupled to case  222  via an aft sliding seal  230  that allows stator assembly  210  to move in the axial direction relative to case  222  while forming a seal with case  222 . Actuator  228  can include any actuator capable of changing the position of stator assembly  210  relative to case  222  and, thus, rotor assembly  206 . As illustrated, actuator  228  utilizes a roller cam actuation system. In another embodiment, an actuator is positioned at the aft end of stator assembly  210  instead of or in addition to actuator  228  positioned at the forward end of stator assembly  210 . 
     As illustrated, outer diameter edge  216  of stator assembly  210  and inner diameter edge  217  of rotor assembly  206  form a conic shape such that the larger plane surface of the conic shape is forward and the radius of the conic shape decreases towards the vertex of the conic shape in the aft direction. Accordingly, by actuating stator assembly  210  in the forward direction, the radius of the conic shape is reduced, thus reducing distance  260  and increasing the efficiency of low pressure compressor section  44  by reducing the amount of fluid leaking between stages. 
     A forward flange  262  of stator assembly  210  is coupled to a forward end  270  of a linear guide rail  226  and an aft flange  264  of stator assembly  210  is coupled to an aft end  272  of linear guide rail  226 . A carriage  224  is coupled to case  222  and slidably coupled to linear guide rail  226 . Accordingly, linear guide rail  226  can move forward and aft relative to carriage  224  and thus case  222 . Carriage  224  and linear guide rail  226  are designed such that linear guide rail  226  and carriage  224  resist radial motion relative to case  222 . Stated another way, carriage  224  and linear guide rail  226  resist a radial force of stator assembly  210  and carriage  224  and linear guide rail  226  allows axial movement of stator assembly  210 . 
     With reference now to  FIG. 3 , a portion  308  of outer diameter edge  216  of stator assembly  210  is shown in a first position  302  and a second position  300  relative to rotor  208 A. First position  302  of portion  308  is positioned aft of second position  300  of portion  308 . When outer diameter edge  216  is in first position  302 , a distance  306  exists between portion  308  and rotor  208 A. As outer diameter edge  216  moves forward relative to rotor  208 A to second position  300 , a new distance  304  exists between portion  308  and rotor  208 A. Because of the conic shape defined by stator assembly  210  and rotor assembly  206 , distance  304  is smaller than distance  306 . 
     The reduction in distance between first position  302  and second position  300  reduces an amount of fluid that leaks between rotor  208 A and portion  308 . Accordingly, when portion  308  is in second position  300 , low pressure compressor section  44  is more efficient yet has less tolerance of axial movement of rotor  208 A. Thus, second position  300  is desirable when less tolerance is desired between rotor  208 A and portion  308 . When portion  308  is in first position  302 , low pressure compressor section is less efficient yet has more tolerance for axial movement of rotor  208 A. Thus, first position  302  is desirable when more tolerance is desired between portion  308  and rotor  208 A. 
     With reference to  FIGS. 2 and 4 , a controller  400  is be coupled to actuator  228 . Controller  400  can include a processor and a tangible, non-transitory memory and be capable of implementing logic. The processor can be a general purpose processor, a digital signal processor (DSP), an application specific integrated circuit (ASIC), a field programmable gate array (FPGA) or other programmable logic device, discrete gate or transistor logic, discrete hardware components, or any combination thereof. The controller  400  can receive signals generated. 
     Controller  400  receives information regarding gas turbine engine  20 , such as upcoming maneuvers, landings, takeoffs or the like; information regarding the environment, such as whether pockets of low pressure exist in the current environment; instructions from an operator of the aircraft; and/or information regarding conditions of the gas turbine engine such as rotational engine speed, temperature data, acceleration data received from accelerometers positioned in the engine, proximity of components received from proximity sensors or the like. Controller  400  determines if any loads or forces will be applied to rotor assembly  206  such as maneuver loads, thermal growth or the like based on the information. Based on the forces on rotor assembly  206 , controller  400  instructs actuator  228  to cause stator assembly  210  to be in a suitable position relative to rotor assembly  206 . When in a suitable position, low pressure compressor section  44  will function with a high efficiency while retaining a low likelihood of collision between outer diameter edge  214  of rotor assembly  206  and outer diameter edge  216  of stator assembly  210 . 
     With reference to  FIGS. 2, 4 and 5 , a method  500  is performed by controller  400  for causing actuator  228  to position stator assembly  210  in a suitable position relative to rotor assembly  206 . In block  502 , controller  400  determines that a maneuver or event is currently or is likely to change the clearance between outer diameter edge  214  and outer diameter edge  216 . Controller  400  can also or instead receive an instruction from an operator of the aircraft regarding a desired tolerance between rotor assembly  206  and stator assembly  210  and/or an indication from the operator of whether a tolerance and/or efficiency change is desired. 
     In block  504 , when a maneuver or situation is currently or can change the clearance between rotor assembly  206  and stator assembly  210  (stated differently, when an input indicates that a force will be applied to the engine), controller  400  determines an amount to actuate stator assembly  210 . As mentioned with reference to  FIG. 4 , the amount controller  400  will cause actuator  228  to actuate stator assembly  210  is an amount in which the tip clearance is sufficient to reduce the likelihood of contact between stator assembly  210  and rotor assembly  206  while providing maximum efficiency. Additionally or instead, controller  400  can receive an amount to actuate stator assembly  210  from an operator. 
     In block  506 , controller  400  instructs actuator  228  to adjust the position of stator assembly  210  relative to rotor assembly  206  the amount determined in block  504 . As discussed above, this places stator assembly  210  in an optimal position relative to rotor assembly  206  for tip clearance and efficiency of low pressure compressor section  44 . 
     The concepts disclosed herein have been described with reference to a low pressure compressor section of a gas turbine engine. However, one skilled in the art will realize that these concepts are applicable to any system including a rotor assembly having a rotor that rotates relative to an axis and a stator assembly having a stator that does not rotate relative to the axis. Additionally, the concepts have been described with reference to a stator assembly moving axially relative to a rotor assembly. However, one skilled in the art will realize that these concepts are applicable to a system in which a rotor assembly moves relative to a stator assembly. 
     Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. The scope of the disclosure, however, is provided in the appended claims.