Patent Publication Number: US-11391208-B2

Title: Fan blade anti-icing concept

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a divisional of U.S. application Ser. No. 16/683,552 filed Nov. 14, 2019, which claims priority to U.S. provisional patent application No. 62/925,848 filed Oct. 25, 2019, the entire contents of which are incorporated by reference herein. 
    
    
     TECHNICAL FIELD 
     The disclosure relates generally to anti-icing of fan blades in a gas turbine engine. 
     BACKGROUND 
     Ice can form and adhere to the fan blades of a gas turbine engine under certain conditions during flight. The weight of the ice buildup can result in imbalance of the fan and can be detrimental to efficient air flow. When ice breaks away and is released, the fan can be imbalanced, orbiting increases, vibration occurs and impact from ice particles can cause foreign object damage. 
     SUMMARY 
     The disclosure describes a fan blade anti-icing system for a gas turbine engine comprising: a fan hub mounted for rotation about an axis; and a fan blade extending radially outwardly from the fan hub, the fan blade having a base and an airfoil extending radially outwardly from the fan base, the airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge, the fan blade further having a radial passage extending from a blade air inlet in the blade base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet upstream of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil. 
     In accordance with another aspect, there is provided a fan blade for a gas turbine engine comprising: a fan blade having a blade base and an airfoil with a radially outward axis, the airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge, the fan blade further having a radial passage extending from a blade air inlet in the blade base for communication with a source of heated air, a rearwardly directed passage in communication with the radial passage and having a blade air outlet upstream of the trailing edge and oriented tangentially to the convex side surface or the concave side surface of the airfoil. 
     In accordance with a still further general aspect, there is provided a method of impeding icing on an airfoil surface of a fan blade of an aircraft engine, the method comprising: receiving heated pressurized air inside the fan blade; and directing the heated pressurized air exiting the fan blade to flow in a downstream direction over the airfoil surface of the fan blade. 
     Further details of these and other aspects of the subject matter of this application will be apparent from the detailed description included below and the drawings. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
         FIG. 1  shows an axial cross-section view of a turbo-fan gas turbine engine. 
         FIG. 2  is an isometric transparent view of a section through an integrally bladed fan rotor in accordance with the present description. 
         FIG. 3  is an enlarged view of  FIG. 2  showing an internal radial passage and multiple rearwardly directed passages with heated air outlets to create a Coanda effect heated air film over the airfoil surface of a fan blade. 
         FIG. 4  is a further isometric transparent view of the fan blade as shown in  FIG. 3  with a radial passage and having multiple rearwardly directed passages. 
         FIG. 5  is an isometric solid view of an alternative fan blade having a recessed pocket and cover defining an air plenum, the cover including multiple outlets to provide for the formation of a film of anti-icing air over a surface of the airfoil. 
         FIG. 6  is an isometric solid view along line  6 - 6  of  FIG. 5  showing the drilled radial passage. 
         FIG. 7  is a view of the recessed pocket of  FIG. 5  with the cover removed. 
         FIG. 8  is a radial section along line  8 - 8  of  FIG. 7  showing the intersection of the radial passage with the recessed pocket. 
         FIG. 9  is a further alternative to  FIG. 5  with the multiple outlets in a radial alignment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  shows an axial cross-section through an aircraft engine. According to the illustrated embodiment, the aircraft engine is a turbo-fan gas turbine engine. It is understood that the aircraft engine could adopt various forms others than the illustrated example. Air intake into the engine passes over fan blades  1  in a fan case  2  and is then split into an outer annular flow through the bypass duct  3  and an inner flow through the low-pressure axial compressor  4  and high-pressure centrifugal compressor  5 . Compressed air exits the compressor through a diffuser  6  and is contained within a plenum  7  that surrounds the combustor  8 . Fuel is supplied to the combustor  8  through fuel tubes  9  and fuel is mixed with air from the plenum  7  when sprayed through nozzles into the combustor  8  as a fuel air mixture that is ignited. A portion of the compressed air within the plenum  7  is admitted into the combustor  8  through orifices in the side walls to create a cooling air curtain along the combustor walls or is used for cooling to eventually mix with the hot gases from the combustor and pass over the nozzle guide vane  10  and turbines  11  before exiting the tail of the engine as exhaust. 
     The present description and drawings relate to anti-icing features of the fan blades  1 . The compressors  4 ,  5  and combustor  8  create pressurized air having a temperature greater than ambient and at least above the freezing temperature of water at flight altitude. Heated pressurized air can be bled from the compressors  4 ,  5  and combustor  8  and directed through the engine to the fan hub  12  via ducts within the hollow central engine shafts for example. 
     With reference to  FIGS. 2 and 3 , there is shown an integrally bladed fan rotor having a fan hub  12  with a hub air inlet in communication with a source of heated air, such as the compressors  4 ,  5  and combustor  8 . The exemplified fan hub  12  is generally cylindrical or conical with a radially outer surface with multiple blades spaced about the circumference. Each blade has a blade base  14 . The fan hub  1  has a plurality of hub air outlets in communication with the hub air inlet to distribute heated pressurized air to each blade. 
     The fan blade has an airfoil  15  with a radially outward axis  16  generally normal to the direction of air flow into the engine (see arrow). The airfoil  15  has a leading edge  17 , a trailing edge  18 , a convex side surface  19  between the leading and trailing edge  17 ,  18  and a concave side surface  20  between the leading and trailing edge  17 ,  18 . 
     As best seen in  FIGS. 3 and 4 , the blade has a radial passage  21  for receiving and distributing heated pressurized air. The radial passage  21  extends from a blade air inlet  22  in the blade base  14  that is in communication with the air outlet in the fan hub  12 . The radial passage  21  is provided adjacent to the leading edge  17 . In the example illustrated, six rearwardly directed passages  23  branch off in fluid communication with the radial passage  21 . Each passage  23  extends towards the trailing edge  18  and terminates in a blade air outlet  24  in the concave side surface  20 , the convex side surface  20  or both. 
     The blade air outlets  24  are disposed upstream of the trailing edge  18  typically within the upstream half of the airfoil and oriented predominantly tangential to the airfoil surface to emit heated pressurized air substantially parallel to the incoming air that passes over the airfoil  15 . The rearward passages  23  and blade air outlets  24  are oriented in a rearward or downstream direction substantially parallel to the incoming air direction to provide for the formation of a heated air film over the airfoil surface. In the radially inward area of the airfoil  15  adjacent to the blade platform  25  and fillet  26 , the incoming air direction is directed to be parallel to the blade platform  25  (see arrow in  FIG. 2 ). 
     The heated pressurized air from the blade air outlets  24  is ejected in streams or jets that merge smoothly with the incoming air. As a result, the streams of heated air are subjected to the Coanda effect and flow downstream attached to the concave side surface  20 . The Coanda effect is the tendency of a fluid jet to stay attached to an adjacent surface, named after Romanian inventor Henri Coanda. The effect is the tendency of a jet of fluid emerging from an orifice to follow an adjacent flat or curved surface and to draw in or entrain fluid from the surroundings so that a region of lower pressure develops. The lower pressure region between the jet and adjacent surface draws the jets towards the adjacent surface to flow parallel to or “attach” to the surface. Eventually the jet and ambient air flow tend to mix downstream due to turbulence and the Coanda effect dissipates. 
     Therefore, the jet of heated pressurized air from the blade air outlets  24  creates a film of heated air flowing parallel to and closely attached to the airfoil surface. The heated air flowing through the radial passage  21  and the multiple rearward passages  23  will heat the metal of the airfoil  15  through convection. In addition, the rearward orientation of the rearward passages  23  and blade air outlets  24  will emit jets of heated air that will flow close to and parallel to the concave side surface  20 , the convex side surface  19  or both resulting from the Coanda effect. The areas of the airfoil surface over which the heated jets of air flow will be locally heated to impede formation of ice and melt ice that has been deposited. 
     In the illustrated example, the multiple rearwardly directed passages  23  and blade air outlets  24  are radially spaced apart and are radially aligned on an imaginary line that is transverse to the blade platform  25 . Various alternative patterns of location the rearwardly directed passages  23  and blade air outlets  24  can be adopted depending on the nature of blade anti-icing required. In the examples illustrated, as seen in  FIG. 4  the blade air outlets  24  are disposed in an area of the concave side surface  20  that is an offset by a distance ‘x’ from the blade platform  25 . 
       FIGS. 5 to 9  show an alternative fan blade anti-icing arrangement with a single rearwardly directed passage formed as a recessed pocket  27  (see  FIG. 7 ) in the concave side surface  28 . As best seen in  FIG. 7 , the recessed pocket  27  intersects with and is in fluid communication with the radial passage  29 .  FIG. 6  shows the blade air inlet  30  and radial passage  29  that have been drilled to connect with the recessed pocket  27 . 
     As seen in  FIG. 5 , the recessed pocket  27  is covered with a cover  31  to define an internal plenum within the airfoil  32 . The plenum between the cover  31  and the recessed pocket  27  receives heated pressurized air from the radial passage  29 . The cover  31  and adjacent areas of the airfoil  32  are heated by convection to locally impede icing. Heated air is emitted through the blade air outlets  33  that fluidly communicate with the internal plenum beneath the cover  31 . Heated air from the blade air outlets  33  is emitted rearwardly in streams or jets that are substantially parallel to the blade platform  34  and substantially parallel to the rearward incoming air flow direction (see arrow). 
       FIGS. 5 and 9  show blade air outlets  33  and blade air outlets  35  drilled into the cover  31  and cover  36 . The blade air outlets  33 ,  35  can be drilled through the cover  31  and  36 , or may be formed as a slot in an edge of the cover  31 ,  36  or may be formed as a slot in an edge of the recessed pocket  27 . The cover  31 ,  36  may be sealed to the edges of the recessed pocket  27  by adhesive bonding, welding, or diffusion bonding for example. The corners of the recessed pocket  27  and cover  31  are rounded to reduce stress concentration issues. The cover  31 ,  36  could be formed as a round disk or other shapes if desired. The shape and location of the recessed pocket  27 , and the arrangement of blade air outlets  33 ,  35  can be selected to target the specific icing experienced by any blade configuration. The above described system can be applied to bladed and integrally bladed fans, both solid and hollow blades, and both metal and composite blades. 
     The above description is meant to be exemplary only, and one skilled in the relevant arts will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. The present disclosure may be embodied in other specific forms without departing from the subject matter of the claims. The present disclosure is intended to cover and embrace all suitable changes in technology. Modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims. Also, the scope of the claims should not be limited by the preferred embodiments set forth in the examples, but should be given the broadest interpretation consistent with the description as a whole.