Abstract:
A turbofan engine deicing system includes a core nacelle ( 12 ) housing a turbine. A turbofan ( 20 ) is arranged upstream from the core nacelle. A controller ( 50 ) manipulates the turbofan in response to detecting an icing condition for avoiding undesired ice buildup on the turbofan engine ( 10 ) and nacelle parts. In one example, a variable area nozzle ( 40 ) is actuated to generate pressure pulses or a surge condition to break up any ice buildup. The icing condition can be determined by at least one sensor ( 52 ) and/or predicted based upon icing conditions schedules.

Description:
This application claims priority to PCT Application Serial No. PCT/US2006/039946, filed on Oct. 12, 2006. 
     BACKGROUND OF THE INVENTION 
     This invention relates to a deicing system for use with, for example, a turbofan engine. 
     During icy conditions, ice may build up on static and rotating components of the aircraft engine unless preventative measures are taken. Ice buildup causes propulsion system operability, safety and performance difficulties. Different techniques, such as anti-icing and deicing systems, can cause efficiency losses as well as a need for larger safety and operability margins to keep the propulsion system operational during icing conditions. These techniques result in an increase in system weight and fuel burn. 
     One example anti-icing system relies upon hot compressor bleed air, which can result in up to several percent fuel burn debit during its brief usage. Electrical heaters have also been employed, but are undesirable because they extract power from the engine. Specialty coatings have also been used to prevent icing, but are typically costly and deteriorate over time. Deicing systems are employed after ice formation and typically require an undesirably large amount of energy to actuate and break off ice accumulation. 
     What is needed is an ice preventative system that does not increase the weight of the engine or result in increased fuel burn and reduced operability of the engine. 
     SUMMARY OF THE INVENTION 
     A turbofan engine deicing system includes a core nacelle housing a turbine. A turbofan is arranged upstream from the core nacelle. A controller manipulates the turbofan in response to detecting an icing condition for avoiding undesired ice buildup on the turbofan engine and nacelle parts. In one example, a control device is commanded by the controller in response to the icing condition. In one example, the control device includes a variable area nozzle that is actuated to generate pressure pulses or a surge condition to break up any accumulated ice. The icing condition can be determined by at least one sensor and/or predicted based upon icing conditions schedules. 
     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  is a cross-sectional view of an example geared turbofan engine. 
         FIG. 2  is a partially broken perspective view of the geared turbofan engine shown in  FIG. 1  and including a deicing system. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     A geared turbofan engine  10  is shown in  FIG. 1 . A pylon  38  secures the engine  10  to an aircraft. The engine  10  includes a core nacelle  12  that houses a low spool  14  and high spool  24  rotatable about an axis A. The low spool  14  supports a low pressure compressor  16  and low pressure turbine  18 . In the example, the low spool  14  drives a turbofan  20  through a gear train  22 . The high spool  24  supports a high pressure compressor  26  and high pressure turbine  28 . A combustor  30  is arranged between the high pressure compressor  26  and high pressure turbine  28 . Compressed air from compressors  16 ,  26  mixes with fuel from the combustor  30  and is expanded in turbines  18 ,  28 . 
     In the examples shown, the engine  10  is a high bypass turbofan arrangement. In one example, the bypass ratio is greater than 10:1, and the turbofan diameter is substantially larger than the diameter of the low pressure compressor  16 . The low pressure turbine  18  has a pressure ratio that is greater than 5:1, in one example. The gear train  22  is an epicycle gear train, for example, a star gear train, providing a gear reduction ratio of greater than 2.5:1. It should be understood, however, that the above parameters are only exemplary of a contemplated geared turbofan engine. That is, the invention is applicable to other engines including direct drive turbofans. 
     Airflow enters a fan nacelle  34 , which surrounds the core nacelle  12  and turbofan  20 . The turbofan  20  directs air into the core nacelle  12 , which is used to drive the turbines  18 ,  28 , as is known in the art. Turbine exhaust E exits the core nacelle  12  once it has been expanded in the turbines  18 ,  28 , in a passage provided between the core nacelle and a tail cone  32 . 
     The core nacelle  12  is supported within the fan nacelle  34  by structure  36 , which are commonly referred to as upper and lower bifurcations. A generally annular bypass flow path  39  is arranged between the core and fan nacelles  12 ,  34 . The example illustrated in  FIG. 1  depicts a high bypass flow arrangement in which approximately eighty percent of the airflow entering the fan nacelle  34  bypasses the core nacelle  12 . The bypass flow B within the bypass flow path  39  exits the fan nacelle  34  through a nozzle exit area  40 . 
     For the engine  10  shown in  FIG. 1 , a significant amount of thrust may be provided by the bypass flow B due to the high bypass ratio. Thrust is a function of density, velocity and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B. In one example, the engine  10  includes a structure associated with the nozzle exit area  40  to change the physical area and geometry to manipulate the thrust provided by the bypass flow B. However, it should be understood that the nozzle exit area may be effectively altered by other than structural changes, for example, by altering the boundary layer, which changes the flow velocity. Furthermore, it should be understood that any device used to effectively change the nozzle exit area is not limited to physical locations near the exit of the fan nacelle  34 , but rather, includes altering the bypass flow B at any suitable location. 
     The engine  10  has a flow control device  41  that is used to effectively change the nozzle exit area. In one example, the flow control device  41  provides the fan nozzle exit area  40  for discharging axially the bypass flow B pressurized by the upstream turbofan  20  of the engine  10 . A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The turbofan  20  of the engine  10  is typically designed for a particular flight condition, typically cruise at 0.8M and 35,000 feet. The turbofan  20  is designed at a particular fixed stagger angle for an efficient cruise condition. The flow control device  41  is operated to vary the nozzle exit area  40  to adjust fan bypass air flow such that the angle of attack or incidence on the fan blade is maintained close to design incidence at other flight conditions, such as landing and takeoff. This enables desired engine operation over a range of flight condition with respect to performance and other operational parameters such as noise levels. In one example, the flow control device  41  defines a nominal converged position for the nozzle exit area  40  at cruise and climb conditions, and radially opens relative thereto to define a diverged position for other flight conditions. The flow control device  41  provides an approximately 20% change in the nozzle exit area  40 . 
     In one example, the flow control device  41  includes multiple hinged flaps  42  arranged circumferentially about the rear of the fan nacelle  34 . The hinged flaps  42  can be actuated independently and/or in groups using segments  44 . In one example, the segments  44  and each hinged flap  42  can be moved angularly using actuators  46 . The segments  44  are guided by tracks  48  in one example. In the example shown, the hinged flaps  42  may be manipulated to change the amount and/or direction of thrust. 
     A deicing system  54  includes a controller  50  that communicates with the actuators  46  to manipulate the flow control device  41  thereby changing the effective nozzle exit area. The controller  50  commands the control device  41  to manipulate the turbofan to avoid ice build up. In the example shown in  FIG. 2 , the flow control device  41  physically changes the nozzle exit area  40  by moving the hinged flaps  42 . 
     More specifically, effectively changing the nozzle exit area  40  can introduce pressure pulses capable of breaking up any formed ice by modulating back pressure. This can be achieved by rapidly opening and closing the hinged flaps  42  several times. A controlled surge condition introducing mechanical vibrations can also be initiated by the flow control device  41  to break up any ice on the engine  10 . 
     An ice detection sensor  52  in communication with the controller  50  can be used to detect the actual presence of ice in a desired location on the engine  10 . The controller  50  initiates a deicing procedure with the control device  41  in response to any detected ice. 
     Alternatively or additionally, the controller  50  can periodically actuate the flow control device  41  based upon a schedule or conditions that are typically favorable to ice formation. An aircraft icing sensor  60  can be used to provide information to the controller  50  for use with the icing conditions schedules. The aircraft icing sensor  60  includes atmospheric temperature and pressure sensors  56  and  58 , in one example. A deicing procedure can also be initiated manually by the pilot using a switch  59 , for example. 
     Although example embodiments of this invention have 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.