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A turbofan 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

A turbofan 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.

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1. A turbofan engine, the engine comprising: a propulsive fan;an inlet wall surrounding an inlet flow path, the inlet wall extending axially from an upstream end to a downstream end adjacent the fan, the inlet wall at the downstream end surrounding an annular portion of the inlet flow path bordered

1. A turbofan engine, the engine comprising: a propulsive fan;an inlet wall surrounding an inlet flow path, the inlet wall extending axially from an upstream end to a downstream end adjacent the 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 inner wall adjacent the fan defining a downstream height of the inlet flow path; anda plurality of vanes circumferentially spaced around the inlet and extending radially inwardly from the inlet wall, a maximum radial distance between a tip of the vanes and the inlet wall defining a maximum height of the vanes, the maximum height of the vanes being at most 50% of the downstream height of the flow path, the annular flow path having an annular outer portion containing the vanes, the vanes being oriented to swirl a flow upstream of the fan within the annular outer portion of the inlet flow path such that an exit angle induced in the flow by the vanes reduces a relative Mach number of the flow at blade tips of the fan, wherein the height of the vanes varies between a leading edge and a trailing edge of the vanes. 2. The engine as defined in claim 1, wherein the maximum height of the vanes is at most 25% of the downstream height of the flow path. 3. The engine as defined in claim 1, wherein a circumferential spacing between adjacent ones of the vanes is irregular. 4. The engine as defined in claim 1, wherein of at least some of the vanes have different stagger angles from one another. 5. The engine as defined in claim 1, wherein the vanes have a stagger angle of from 10 to 15 degrees adjacent the inlet wall. 6. The engine as defined in claim 1, wherein the vanes are pivotally retained to the inlet wall such that the vanes have a variable stagger angle. 7. The engine as defined in claim 1, further comprising a heating system in heat exchange relationship with the vanes. 8. The engine as defined in claim 7, 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. 9. The engine as defined in claim 1, wherein the vanes have a chord extending between 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. 10. The engine as defined in claim 1, wherein the exit angle induced in the flow by the vanes reduces the relative Mach number of the flow at the blade tips to a value of Mach 1.3 or less. 11. The engine as defined in claim 10, wherein the exit angle induced in the flow by the vanes reduces the relative Mach number of the flow at the blade tips to a value of at least Mach 1.2. 12. A gas turbine engine comprising: a propulsive fan including an array of circumferentially spaced blades configured for rotation and extending radially between a root and a tip with a maximum radial dimension between the root and the tip defining a maximum blade span;an annular inlet including:an axially extending wall, the wall having an upstream wall portion extending axially upstream from the fan blades, the upstream wall portion defining an inlet flow path for directing air to the fan, anda plurality of vanes circumferentially spaced around the annular inlet and extending radially inwardly from the upstream wall portion, a maximum radial distance between a tip of the vanes and the upstream wall portion defining a maximum height of the vanes, the maximum height of the vanes being at most 50% of the maximum blade span, the vanes being oriented to swirl a flow upstream of the tip of the blades of the fan such that an exit angle induced in the flow by the vanes reduces a relative Mach number of the flow at the tip of the blades, wherein the height of the vanes varies between a leading edge and a trailing edge of the vanes. 13. The engine as defined in claim 12, wherein the maximum height of the vanes is 25% or less than the maximum blade span. 14. The engine as defined in claim 12, wherein a circumferential spacing between adjacent ones of the vanes is irregular. 15. The engine as defined in claim 12, wherein of at least some of the vanes have different stagger angles from one another. 16. The engine as defined in claim 12, further comprising a heating system in heat exchange relationship with the vanes. 17. The engine as defined in claim 12, wherein the vanes has a chord extending between 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. 18. The engine as defined in claim 12, wherein the exit angle induced in the flow by the vanes reduces the relative Mach number of the flow at the tip of the blades to a value of Mach 1.3 or less. 19. The engine as defined in claim 18, wherein the exit angle induced in the flow by the vanes reduces the relative Mach number of the flow at the tip of the blades to a value of at least Mach 1.2. 20. A method of reducing a relative Mach number at tips of fan blades of a gas turbine engine, the method comprising: directing a flow of air to the fan blades through an inlet flow path having a downstream radial height adjacent the fan blades, including:swirling the flow of air upstream of the blade tips within an annular outer portion of the inlet flow path to change an angle of the flow within the annular outer portion such that the relative Mach number at the blade tips is reduced, the annular outer portion including a circumferential array of vanes extending a radial distance from a wall surrounding the inlet flow path, the radial distance defining a height of the vanes, the height being at most 50% of the downstream radial height, wherein the height of the vanes varies between a leading edge and a trailing edge of the vanes; andallowing the flow of air to flow freely within a remaining central portion of the inlet flow path. 21. The method as defined in claim 20, wherein the radial distance of the annular outer portion is at most 25% of the downstream radial height. 22. The method as defined in claim 20, wherein swirling the flow of air is performed at different angles along a circumference of the inlet flow path. 23. The method as defined in claim 20, wherein swirling the flow is performed at an angle of from 10 to 15 degrees adjacent the wall with respect to a central longitudinal axis of the wall. 24. The method as defined in claim 20, wherein swirling the flow of air upstream of the blade tips to change the angle of the flow within the annular outer portion is performed such that the relative Mach number at the blade tips is reduced to a value of Mach 1.3 or less. 25. The method as defined in claim 24, wherein swirling the flow of air upstream of the blade tips to change the angle of the flow within the annular outer portion is performed such that the relative Mach number at the blade tips is reduced to a value of at least Mach 1.2.