Patent Application
20210115795
The disclosure relates generally to fans for turbofan gas turbine engines, and to anti-icing of the fan blades in a gas turbine engine.
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 deposits break away and are released from the blades, the fan can be imbalanced, orbiting may increase, vibration may occur and/or impact from ice particles can cause foreign object damage to the engine.
Various anti-icing devices have been developed such as electrical heaters in surfaces prone to icing or effusion of hot air into the air flow passing over the airfoil of a blade. Hot air can quickly mix with the turbulent incoming air flow into the engine and the anti-icing effect can be dissipated. Improvement is desirable.
There is accordingly provided a rotor of a gas turbine engine comprising: a hub and a plurality of blades extending from the hub; and an anti-icing fin attached to the hub and extending radially therefrom upstream of at least one blade of the plurality of blades, the anti-icing fin including: a fin body having an airfoil shape and disposed upstream from a radially inward portion of the at least one blade; a plenum defined within the fin body in communication with a heated air source of the gas turbine engine; and a plurality of outlet orifices oriented in a downstream direction substantially parallel to a flow of incoming air over the at least one blade.
There is also provided a method of heating a fan blade of a gas turbine engine for anti-icing, the method comprising: emitting jets of heated air from a radial fin disposed upstream from a radially inward portion of a fan blade airfoil, the jets of heated air being directed by outlet orifices in a downstream direction substantially parallel to a flow of incoming air over the fan blade airfoil.
There is further provided a gas turbine engine comprising: an engine nose cone having a chamber in communication with a heated air source of the gas turbine engine; a fan having a fan hub and a circumferentially spaced apart plurality of fan blades, each fan blade having an airfoil with a leading edge and a radially inward portion of the fan blade airfoil; a circumferentially spaced apart plurality of anti-icing radial fins disposed on an annular ring, the annular ring being disposed between the engine nose cone and an associated blade being one of the plurality of fan blades, wherein each fin has a plenum in communication with the chamber and a plurality of outlet orifices oriented in a downstream direction substantially parallel to a flow of incoming air over the radially inward portion of the associated blade.
There are provided structures and/or methods for preventing ice accumulation on the fan blades of a gas turbine engine. In one embodiment, this is achieved using airfoil shaped fins upstream of a rotor, the airfoil shaped “shark fin” configured to convey heated air from the hub onto the rotor, and therefore the fan blades, to reduce or prevent ice formation on the main fan blades.
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.
The present description and drawings relate to anti-icing features of the fan 1. The compressors 4, 5 and combustor 8 create heated air having a temperature and pressure greater than ambient and at least above the freezing temperature of water at flight altitude. Heated pressurized air can originate from the compressors 4, 5 or combustor 8 and be bled off to be directed through the engine to the engine nose cone 13 via ducts within the hollow central engine shafts, for example.
With reference to
In
Each fan blade 19 has an airfoil 21 with a leading edge 29 and a trailing edge 30. A dashed line 31 in
In
Each fin 34 has a plenum 38 that is in fluid communication with the chamber 28 of the engine nose cone 27 and receives heated pressurized air via multiple inlet orifices 39. The fins 34 have one or more outlet orifices 40 on a trailing edge 41. The outlet orifices 40 can be drilled to any desired angular orientation emit jets of heated air toward the leading edge 29 of the airfoil 21, toward only the pressure side of the fan blade airfoil 21, toward only the suction side of the fan blade airfoil 21, or to both sides simultaneously. In all cases, the outlet orifices 40 are oriented in a downstream direction substantially parallel to the flow of incoming air over the radially inward portion 32 of the associated fan blade 19. As a result, the heated air jets emitted from the outlet orifices 40 merge or exit at an acute angle to the incoming airflow direction to create a boundary layer flow (i.e. the Coanda effect). The outlet orifices 40 can be located and oriented to various areas on the surface of the airfoil 21 and platform 22 that are prone to icing, which in general are located in the radially inward portion 32.
To recap the above description, an anti-icing radial fin 34 is located upstream from each fan blade 19 for heating the fan blade airfoil 21 of a gas turbine engine. The fins 34 can be supported on an annular ring 35 disposed between the engine nose cone 27 and the blade platform 22. Alternatively the fins 34 could be formed on the engine nose cone 27 as a combined element. The fin 34 has an aerodynamically shaped body with a trailing edge 41 disposed upstream from the radially inward portion 32 of the fan blade airfoil 21. The fin 34 has a plenum 38 that is in communication with a heated air source of the gas turbine engine via the internal chamber 28 of the engine nose cone 27 or through some other ducting initiating from a heated air source such as a compressor 4, 5 or combustor 8.
The fin 34 has one or more outlet orifices 40 oriented in a downstream direction substantially parallel to a flow of incoming air over the fan blade airfoil 21. In the example illustrated multiple outlet orifices 40 are disposed upstream from the leading edge 29 of the fan blade airfoil 21.
In the example illustrated, the fan blade 19 has a radial length extending from the blade platform 22 to the blade tip 33. The radially inward portion 32 of the fan blade airfoil 21 spans from the blade platform 22 to approximately one third of the radial length, defining an area generally prone to icing. In the example illustrated, the body of the fin 34 has a convex leading edge 42 axially aligned with the leading edge 29 of the radially inward portion 32 of the fan blade airfoil 21. However the trailing edge 41 of the fin 34 can axially overlap or can be circumferentially offset from the leading edge 29 of the airfoil 21 if desired.
The heated air jets can be delivered from the outlet orifices 40 of the trailing edge 41 of the fin 34 to impinge on the fan blade airfoil 21 on the pressure surface, suction surface or both. The heated air jets can merge or exit the outlet orifices 40 at an acute angle to the incoming airflow direction to create a boundary layer flow, commonly referred to as the Coanda effect. Heated air can be delivered to the blade airfoil 21 or blade platform 22 location where icing most likely to accrete, usually the radially inward ⅓ of the airfoil 21. The heated air can be delivered in sufficient quantity and pressure to prevent ice formation altogether or to limit ice formation or to lower the bond strength of the ice formed on the blade 19. These functions are designed to lower the imbalance caused by ice accumulation on the blade 19 and ice shedding from the blade 19. The fin 34 can be designed to meet all FOD (foreign object damage), structural, aerodynamic and dynamic requirements for the gas turbine engine.
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.
