Abstract:
A method of protecting a gas turbine engine according an exemplary aspect of the present disclosure includes, among other things, the steps of determining at least one flight condition of an aircraft and comparing the at least one flight condition to a programmed condition. The method further includes the steps of moving a plurality of inlet vanes of a low pressure compressor from a first position to a second position if the step of comparing the at least one flight condition to the programmed flight condition determines the programmed flight condition are met and deflecting any foreign objects with the plurality of inlet vanes.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 61/593,305 which was filed on Jan. 31, 2012. 
    
    
     BACKGROUND OF THE INVENTION 
     Gas turbine engines can have multiple low pressure compressor stages close coupled with a fan. When a geared architecture is employed in the gas turbine engine, a tip speed of a rotor of a low pressure compressor can be increased, and the stage count of the low pressure compressor can be reduced. Variable vanes may then be needed to improve operability of the low pressure compressor. Variable vanes increase operability by moving an angle of inlet air flowing into a first stage compressor to improve operation of the low pressure compressor. 
     An aircraft can be exposed to inclement weather that could affect the operation of the gas turbine engine. For example, the inclement weather can produce hail or ice that could affect operation. 
     SUMMARY OF THE INVENTION 
     A method of protecting a gas turbine engine according an exemplary aspect of the present disclosure includes, among other things, the steps of determining a flight condition of an aircraft and comparing the flight condition to a programmed condition. The method further includes the steps of moving a plurality of inlet vanes of a low pressure compressor from a first position to a second position if the step of comparing the flight condition to the programmed flight condition determines that the programmed flight condition are met and thus deflecting any foreign objects with the plurality of inlet vanes. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include a plurality of inlet vanes having a first position that is a non-diversion position and a second position that is a diversion position. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include a plurality of inlet vanes that move about 15° to about 50°, respectively, from the non-diversion position to the diversion position. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include a flight condition that is an altitude. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include a flight condition that is a power setting. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include flight conditions that are an altitude and a power setting. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include the steps of generating a signal in response to the step of comparing a flight condition to a programmed flight condition, and moving a plurality of inlet vanes from a first position to a second position in response to the signal. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include a programmed flight condition where an altitude is 10,000 to 25,000 feet and a power condition is idle. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include a low pressure compressor includes a plurality of rotors, a flowpath that exists between an outer diameter of the plurality of rotors and an outer casing of a gas turbine engine, and the step of deflecting any foreign objects with a plurality of inlet vanes that directs the foreign objects in the flowpath for extraction with a downstream bleed. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include foreign objects that are at least one of hail, ice or dirt. 
     A method of protecting another gas turbine engine according an exemplary aspect of the present disclosure includes, among other things, the steps of determining flight conditions of an aircraft, where the flight conditions are altitude and a power setting, and comparing the flight conditions to programmed flight conditions. The method also includes the steps of generating a signal in response to the step of comparing the flight conditions to the programmed flight conditions if the programmed flight conditions are met and moving a plurality of inlet vanes of a low pressure compressor from a non-diversion position to a diversion position in response to the signal. The method also includes the step of deflecting any foreign objects with the plurality of inlet vanes in a flowpath for extraction with a downstream bleed, where the flowpath is defined by an outer diameter of a plurality of rotors of the low pressure compressor and an outer casing of a gas turbine engine. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include a plurality of inlet vanes that move about 15° to about 50° from the non-diversion position to the diversion position. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include programmed flight conditions where an altitude is 10,000 to 25,000 feet and a power condition is idle. 
     In a further non-limited embodiment of any of the forgoing method embodiments, the method may include foreign objects that are at least one of hail, ice or dirt. 
     A gas turbine engine according an exemplary aspect of the present disclosure includes, among other things, a controller that determines a flight condition of an aircraft, compares the flight condition to a programmed flight condition, determines if the programmed flight condition is met, and generates a signal if the programmed flight condition is met. The gas turbine engine also includes a low pressure compressor, where the low pressure compressor includes a plurality of inlet vanes, and the plurality of inlet vanes move from a first position to a second position in response to the signal from the controller to deflect any foreign objects. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a plurality of inlet vanes having a first position that is a non-diversion position and a second position that is a diversion position. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a plurality of inlet vanes that move about 15° to about 50° from the non-diversion position to the diversion position. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a flight condition that is an altitude. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a flight condition that is a power setting. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a flight condition that is an altitude and a power setting. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a programmed flight condition where an altitude is 10,000 to 25,000 feet and a power condition is idle. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a low pressure compressor includes a plurality of rotors, a flowpath that exists between an outer diameter of the plurality of rotors and an outer casing of the gas turbine engine, and the step of deflecting any foreign objects with a plurality of inlet vanes that directs the foreign objects in the flowpath for extraction with a downstream bleed. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include a downstream bleed that exits the gas turbine engine through a duct. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include an actuator that moves the plurality of variable vanes between the first position and the second position. 
     In a further non-limited embodiment of any of the forgoing gas turbine engine embodiments, the gas turbine engine may include foreign objects that are at least one of hail, ice or dirt. 
     These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  shows a schematic view of a gas turbine engine; 
         FIG. 2  shows a cut away cross-sectional view of the gas turbine engine with inlet vanes blocking particles during a specific flight condition; 
         FIG. 3  shows a perspective view of a plurality of variable inlet vanes in a non-diversion position; 
         FIG. 4  shows a perspective view of a plurality of variable inlet vanes in a diversion position; and 
         FIG. 5  shows a perspective view of the movement of the plurality of variable inlet vanes between the non-diversion position and the diversion position. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as 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 might include an augmentor section (not shown) among other systems or features. 
     Although depicted as a geared turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with geared turbofans as the teachings may be applied to other types of turbine engines including three-spool or geared turbofan architectures. 
     The fan section  22  drives air along a bypass flowpath B while the compressor section  24  drives air along a core flowpath C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . 
     The 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. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The geared architecture  48  connects the low pressure compressor  44  to the fan  42 , but allows for rotation of the low pressure compressor  44  at a different speed and/or direction than the fan  42 . 
     The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and a high pressure turbine  54 . 
     A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . 
     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 . 
     The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes airfoils  60  that are in the core airflow path. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     The engine  20  is in one example a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6:1) with an example embodiment being greater than ten (10:1). The geared architecture  48  is an epicyclic gear train (such as a planetary gear system or other gear system) with a gear reduction ratio of greater than about 2.3 (2.3:1). The low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). The low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
     In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), and the fan diameter is significantly larger than that of the low pressure compressor  44 . The low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5 (2.5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     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 feet, with the engine at its best fuel consumption, also known as bucket cruise Thrust Specific Fuel Consumption (“TSFC”). TSFC is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf of thrust the engine produces at that minimum point. 
     “Fan pressure ratio” is the pressure ratio across the fan blade alone. The fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.6. 
     “Low corrected fan tip speed” is the actual fan tip speed in feet per second divided by an industry standard temperature correction of [(Tambient deg R)/518.7) 0.5 ]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 feet per second (351 meters per second). 
     Fan exit stators  74  are positioned between the fan  42  and the low pressure compressor  44 , and air travels through the fan exit stators  74  to the low pressure compressor through a front center body duct  72 . The front center body duct  72  can also have variable geometry. 
     The low pressure compressor  44  includes a plurality of vane stages. In one example, each of the stages can have variable geometry. In one example, the low pressure compressor  44  includes a plurality of variable inlet vanes  62  in a first stage, a plurality of vanes  66  in a second stage, a plurality of vanes  68  in a third stage, and a plurality of exit vanes  70  in a final stage. In one example, the plurality of exit vanes  70  have an inwardly sloping flowpath that connects the low pressure compressor  44  to the high pressure compressor  52 . In one example, only the variable inlet vanes  62  are variable. However, it is possible that vanes  66 ,  68  and  70  in the other stages can be variable. 
     The variable inlet vanes  62  are linked and moveable between a non-diversion position (shown in  FIG. 3 ) and a diversion position (shown in  FIG. 4 ).  FIG. 5  shows the movement of the vanes  62  between the non-diversion position  62   a  (shown in solid lines) and the diversion position  62   b  (shown in phantom lines). The non-diversion position  62   a  is a less obstructing position, and the diversion position  62   b  is a more obstructing position. In one example, the variable inlet vanes  62  move an angle X° between about 15° to about 50° from the non-diversion position  62   a  to the diversion position  62   b . That is, the range of movement of the variable inlet vanes  62  is about 15° to about 50°. In one example, the variable inlet vanes  62  are moved to a diversion position  62   b  to protect the health of the gas turbine engine  20 . The variable inlet vanes  62  are moved by an actuator  88 . The actuator  88  may provide for moving the vanes  62  between the diversion position  62   b  and a non-diversion position  62   a  and a plurality of intermediate positions therebetween. 
     The low pressure compressor also includes a plurality of rotors or blades  64  between each of the stages of the vanes  62 ,  66 ,  68  and  70 . An inner diameter of the plurality of rotors or blades  64  increases for successive stages of rotors  64  in the direction of airflow. In one example, an outer diameter of the rotors  64  through successive stages in the direction of airflow is nearly cylindrical. This defines a nearly annular flowpath  82  between the outer diameter of the rotors  64  and the outer casing  76 . 
     A predetermined vane schedule  92  is programmed and stored in a controller  86  and associates vane  62  positions with predetermined flight conditions. The vane schedule  92  includes values and parameters that are pre-programmed in the controller  86  derived through analysis and testing. In one example, the programmed flight conditions are an altitude, a power setting, or both an altitude and a power setting. An altimeter  78  may determine the altitude of the gas turbine engine  20 . 
     The controller  86  uses the predetermined vane schedule  92  to position the variable inlet vanes  62  as required by real time flight condition measurements. In one example, the vane schedule  92  is associated with operating characteristics of the low pressure compressor  44 , as well as inclement weather operating parameters for the gas turbine engine  20 . In one example, the vane schedule  92  is associated with operating characteristics of the low pressure compressor  44  in view of only the inclement weather operating parameters for the gas turbine engine  20 . Inclement weather is defined as, or is a precursor to, hail, ice, rain or other environmental factors. 
     The controller  86  considers the programmed flight conditions of the vane schedule  92 . The programmed flight conditions are established when the chances of encountering hail, ice, dirt, or other foreign objects increases. At these flight conditions, the angle of the variable inlet vanes  62  will be determined to meet the health of the overall gas turbine engine  20  instead of the optimum operability of the low pressure compressor  44 . 
     In one example, inclement weather is most likely to affect engine operation during ascent and descent of the gas turbine engine  20 . In one example, the chance of foreign objects  84  entering the gas turbine engine  20  increases when the gas turbine engine  20  is substantially at idle and the gas turbine engine  20  is at an altitude of 10,000 to 25,000 feet. In one example, the gas turbine engine  20  is idle when the low spool is operating at 30% to 70% of the design speed, or bucket cruise, condition. 
     At least one flight condition is monitored, determined and compared by the controller  86  to the vane schedule  92  that considers the programmed flight conditions that are pre-programmed in the controller  86 . If the controller  86  determines that the programmed flight condition are met, the variable inlet vanes  62  are moved. 
     In one example, when the altimeter  78  detects that the altitude is between 10,000 and 25,000 feet, the gas turbine engine  20  is idle, or both the altitude is between 10,000 feet and 25,000 feet and the gas turbine engine  20  is idle, the programmed flight conditions are met. At these altitudes, the chance of hail, ice, dirt or other contaminants entering the gas turbine engine  20  increases. 
     When the programmed flight conditions are met, the controller  86  sends a signal to the actuator  88  to move the variable inlet vanes  62  to a diversion position to protect the engine  20 . As shown in  FIG. 4 , when any hail, ice, dirt, foreign objects, or other particles  84  impinge on the variable inlet vanes  62 , the flow of the direction of the particles  84  changes. The variable inlet vanes  62  deflect any hail, ice, dirt foreign objects, or other particles  84  to flow in the flowpath  82  between an inner surface  80  of an outer casing  76  of the gas turbine engine  20  and an outer diameter of the plurality rotors  64  for extraction with the downstream bleed through a duct  90 . 
     In this example, the variable inlet vanes  62  are in the diversion position when the programmed flight conditions are met, regardless of the actual weather conditions. This is important as current control systems are not fast enough to determine the presence of a storm and to react if a storm occurs. In one example, the variable inlet vanes  62  move about 15° to about 50° from the non-diversion position to the diversion position. The exact amount the variable inlet vanes  62  move depends on several factors, among them engine size, engine operability, and the conditions of inclement weather determined by the flight conditions. The non-diversion position is a less obstructing position, and the diversion position is a more obstructing position. 
       FIG. 3  shows the variable inlet vanes  62  when in a non-diversion position. As compared to  FIG. 4 , the variable inlet vanes  62  are at a different angle than shown in  FIG. 3 . Any hail, ice, dirt, foreign objects, or other particles  84  flow in the variable inlet vanes  62 . In this example, the hail, ice, dirt, foreign objects, or other particles would not impinge on the variable inlet vanes  62  and could travel inboard and into the low pressure compressor  44  and the high pressure compressor  52 . 
     In another example, the altitude is 15,000 feet during an idle descent condition at a maximum aircraft indicated airspeed. At these conditions, the chances of encountering a hailstorm which could cause engine operational issues increases. In one example, the variable inlet vanes  62  moves about 30° from the non-diversion position to the diversion position. 
     The gas turbine engine  20  can operate during flight conditions that are more likely to expose the gas turbine engine  20  to the presence of hail, ice, dirt, foreign objects, or other particles without adverse responses from the gas turbine engine  20 , optimizing the overall flight cycle. 
     Although a gas turbine engine  20  including geared architecture  48  is described, the vane schedule  92  can be employed with a gas turbine engine without a geared architecture. 
     The foregoing description is only exemplary of the principles of the invention. Many modifications and variations are possible in light of the above teachings, for instance, the inlet vanes may also have intermediate positions. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than using the example embodiments which have been specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.