Gas turbine engine with partial inlet vane

Abstract

Am aircraft engine including an axially extending inlet wall surrounding an inlet flow path. A radial distance between the inlet wall and the inner wall adjacent the fan defines a downstream height of the inlet flow path. A plurality of vanes are circumferentially spaced around the inlet, each of the vanes extending radially inwardly from the inlet wall, a maximum radial distance between a tip of each of the vanes and the inlet wall defining a maximum height of the vane. The maximum height of the vane is at most 50% of the downstream height of the flow path. In another embodiment, the maximum height of the vane is at most 50% of the maximum fan blade span. A method of reducing a relative Mach number at fan blade tips is also discussed.

1 . An aircraft engine comprising:
 a propulsive fan mounted for rotation about an axis, the propulsive fan having fan blades;   an inlet for directing ambient air to the propulsive fan, the inlet having an inlet wall surrounding an inlet flow path, the inlet wall extending axially from an upstream end to a downstream end adjacent the propulsive fan, the inlet wall at the downstream end surrounding an annular portion of the inlet flow path bordered radially inwardly by an annular inner wall, a radial distance between the inlet wall and the annular inner wall adjacent the propulsive fan defining a downstream height of the inlet flow path; and   a plurality of vanes circumferentially spaced around the inlet, each of the vanes extending radially inwardly from the inlet wall, each of the plurality of vanes having an airfoil cross-section including a pressure wall and a suction wall extending chordwise between a leading edge and a trailing edge, a maximum radial distance between a tip of each of the vanes and the inlet wall defining a maximum height of each of the plurality of vanes, the maximum height being at most  50 % of the downstream height of the inlet flow path, and wherein the plurality of vanes are oriented to induce a swirl to an incoming flow about the axis of the propulsive fan in a direction of rotation of the propulsive fan.   
     
     
         2 . The aircraft engine as defined in  claim 1 , wherein the maximum height of each of the vanes is at most 25% of the downstream height of the flow path. 
     
     
         3 . The aircraft engine as defined in  claim 1 , wherein the height of each of the vanes varies between the leading edge and the trailing edge of the vane. 
     
     
         4 . The aircraft engine as defined in  claim 1 , wherein a circumferential spacing between adjacent ones of the vanes is irregular. 
     
     
         5 . The aircraft engine as defined in  claim 1 , wherein at least some of the vanes have different stagger angles from one another. 
     
     
         6 . The aircraft engine as defined in  claim 1 , wherein the vanes each have a stagger angle of 20 degrees or less adjacent the inlet wall. 
     
     
         7 . The aircraft engine as defined in  claim 1 , wherein the vanes are pivotally retained to the inlet wall such that the vanes each have a variable stagger angle. 
     
     
         8 . The aircraft engine as defined in  claim 1 , further comprising a heating system in heat exchange relationship with the vanes. 
     
     
         9 . The aircraft engine as defined in  claim 8 , wherein the heating system includes a heated fluid circulated around the inlet wall radially outwardly of the inlet flow path and in heat exchange relationship with the vanes. 
     
     
         10 . The aircraft engine as defined in  claim 1 , wherein each of the vanes has a chord extending between the leading and trailing edges of the vane, the chord defining a chord length, and a ratio of the maximum height over the chord length is 0.5 or lower. 
     
     
         11 . The aircraft engine as defined in  claim 1 , wherein the vanes each have a stagger angle selected so that the exit angle of the swirl induced in the incoming flow by the vanes reduces a relative Mach number at a tip of the fan blade tips to a value of Mach 1.3 or less. 
     
     
         12 . The aircraft engine defined in  claim 1 , wherein the vanes each have a stagger angle selected so that the exit angle of the swirl induced in the incoming flow by the vanes reduces a relative Mach number at a tip of the fan blade to a value of at least Mach 1.2 and at most Mach 1.3 
     
     
         13 . The aircraft engine defined in  claim 1 , wherein the vanes each have a vane exit angle less than 20 degrees. 
     
     
         14 . The aircraft engine defined in  claim 13 , wherein the vane exit angle is comprised between 10 degrees and 15 degrees.