GAS TURBINE ENGINE, NACELLE THEREOF, AND ASSOCIATED METHOD OF OPERATING A GAS TURBINE ENGINE
The nacelle can have an inlet fluidly connecting a main gas path of a gas turbine engine core, the inlet having an inlet edge connecting an external skin to an internal duct wall, and a step formed in a surface of at least one of the skin and the duct wall, the step delimiting a first portion of the surface from a second portion of the surface, the second portion of the surface being recessed relative to the first portion of the surface, the second portion of the surface extending away from both the step and the inlet edge, whereas the first portion of the surface extends between the inlet edge and the step.
The application relates generally to gas turbine engines and, more particularly, to ice mitigation systems therefor.
BACKGROUND OF THE ARTThis engine nacelle skins, which are exposed to the environment, may be subject to ice accumulation. In the case of engine nacelles, ice accumulation in the vicinity of the inlet can be particularly undesirable as accumulating ice can eventually separate from the surface and represent a source of foreign object damage (FOD). To mitigate ice accumulation to the inlet portion of engine nacelles, it was known to provide heating within the nacelle, such as via hotter air bled from the compressor for instance. Although known systems were satisfactory to a certain extent, there always remains room for improvement, such as in reducing the amount of heat required to achieve the intended purpose for instance.
SUMMARYIn one aspect, there is provided an aircraft engine nacelle comprising an inlet fluidly connecting a main gas path of a gas turbine engine core, the inlet having an inlet edge connecting an external skin to an internal duct wall, and a step formed in a surface of at least one of the skin and the duct wall, the step delimiting a first portion of the surface from a second portion of the surface, the second portion of the surface being recessed relative to the first portion of the surface, the second portion of the surface extending away from both the step and the inlet edge, whereas the first portion of the surface extends between the inlet edge and the step.
In another aspect, there is provided a method of operating a gas turbine engine, the method including a flow of air circulating along a surface of an inlet portion of the gas turbine engine, the flow of air drawing water droplets along the surface until the water droplets reach an edge of a step leading to a recessed portion of the surface, the flow of air separating the water droplets from the surface at the edge of the step.
In another aspect, there is provided an aircraft engine nacelle comprising an inlet fluidly connecting a main gas path of a gas turbine engine core, the inlet having an inlet edge connecting an external skin to an internal duct wall, and means for a flow of air circulating along the internal duct wall to draw water droplets along the surface until the water droplets reach an edge, and to separate the water droplets from the surface at the edge of the step.
In another aspect, there is provided an aircraft engine nacelle comprising an inlet fluidly connecting a main gas path of a gas turbine engine core, the inlet having an inlet edge connecting an external skin to an internal duct wall, and means for a flow of air circulating along the external skin to draw water droplets along the surface until the water droplets reach an edge, and to separate the water droplets from the surface at the edge of the step.
In a further aspect, there is provided an aircraft engine comprising a gas turbine engine core having a main gas path extending, in serial flow communication, across a compressor section, a combustor, and a turbine section, the gas turbine engine core housed within a nacelle, the nacelle having an inlet fluidly connecting the main gas path, the inlet having an inlet edge connecting an external skin to an internal duct wall, and a step formed in a surface of at least one of the skin and the duct wall, the step delimiting a first portion of the surface from a second portion of the surface, the second portion of the surface being recessed relative to the first portion of the surface, the second portion of the surface extending away from both the step and the inlet edge.
Reference is now made to the accompanying figures in which:
Gas turbine engines 10 typically have a main gas path extending through the compressor section 14, combustor 16 and turbine section 18. Turbofan engines, in particular, have a bypass path formed around the core engine, within the nacelle 22. In any case, the inlet fluidly communicates with the main gas path, and in this case, it also communicates with the bypass path. The shape of the nacelle 22 depends on the type of engine and is typically selected in a manner to accommodate the specifics of the engine.
An inlet portion 20 of an example nacelle 22 is shown enlarged in
The example inlet portion 20 is provided with a heater internal to the rounded portion 108. The heater can be a heating air conduit 110 having a plurality of apertures dissipating hotter air bled from the compressor, for instance, or any suitable heater. The heater can be used to heat water, in solid or liquid phase, which comes into contact with the inlet portion 20, to avoid it forming and accumulating a layer of ice, which could eventually dislodge and represent a potential FOD. The power directed to the heater can be modulated as a function of the amount of power expected to be required to achieve this purpose, for instance.
However, the heater has a limited range, and even if, for a given power, it can avoid ice accumulation within its range, liquid water circulating along the surface can eventually exit its range and form an ice accumulation downstream of its range. To avoid this, the range of the heater can be extended, to a certain extent, by supplying additional power (hotter water will travel farther before freezing, especially if it runs along a warmer surface), but this is done at the cost of the additional power, which is typically undesired. Moreover, some embodiments may have practical limitations to the amount of extension of heater range achievable by added power.
More specifically, in the example presented in
The height of the step 112A can vary greatly depending on the size of the engine and the specifics of the embodiment. However, for the purpose of providing an order of magnitude, it can be said here that the height of the step 112A can be expected to be between 0.010″ and 0.200″ in most practical applications. Greater heights may represent a flow distortion judged as being too large, while not providing sufficient compensating advantages, whereas a height smaller than 0.010″ may not be sufficient to cause ejection of the water droplets 113. The exact height for a specific application can be determined based on simulation or testing, for instance. Similarly, the sharpness of the step, i.e. the dimension of the fillet radius of the edge of the step, can vary greatly from one embodiment to another and can be chosen in view of optimizing the efficiency of a specific embodiment. Typically, the ratio of the fillet radius to the height of the step can be between 0 and 1, and the fillet radius can thus be less than 0.200″, for example.
In the specific embodiment illustrated in
In the case of a turbofan engine, providing a step 112A along the bypass duct wall, as opposed to along the nacelle skin, can be particularly useful in avoiding ice accumulation which would be likely to otherwise form a potential source of FOD.
Indeed, ice accumulating on the nacelle skin 106 will typically not represent a potential source of FOD during flight, however it can still be undesired for other reasons. It will be noted that the flow dynamics during takeoff are very different than during flight, hence if ice has accumulated on a nacelle skin 106 on a previous flight and remains in the vicinity of the inlet, a sufficient velocity of air may be drawn forwardly along the skin, towards the inlet, during the next takeoff, causing detachment of the ice accumulation and a FOD. Accordingly, several reasons may motivate the use of a backward facing step 112B on the external skin 106 of the nacelle 22 in addition to, or perhaps even instead of, a step 112A on the internal duct wall 104. This backward facing step 112B on the external skin 106 can be used to eject water running along the surface during flight, prevent the ejected water from freezing and forming an ice accumulation during flight, and thus prevent eventual aspiration of such an ice accumulation by a reverse flow occurring during takeoff, for instance.
It will be understood that in the specific case of a turbofan gas turbine engine, the inlet 20 extends annularly around the engine's main axis 11, and therefore the inlet edge 102, skin 106, and duct wall 104 can be axisymmetric around the main axis 11. In such a context, the step can be designed in a manner to extend around the entire circumference of the inlet 20, for instance. However, in some embodiments, it may be determined that one or more targeted circumferential portions of the inlet 20 are more prone to ice accumulation, and the step can be designed to extend only partially around the circumference, in coincidence with the one or more circumferential portions more prone to ice accumulation. In the case of a turbofan gas turbine engine, the duct wall 104 can be a an outer bypass duct wall for instance. It will be noted, however, that in alternate embodiments, the step can be provided on nacelle inlets 20 of gas turbine engines 10 having other geometries, such as different types of gas turbine engines 10, and the step can thus be adapted accordingly.
Returning to the illustrated example of a turbofan gas turbine engine application, the inlet edge 102 can form part of a D-duct 108 connected to a remainder of the nacelle 22, or bypass duct, as known in the art, and the heating conduit 110 can extend circumferentially within the D-duct 108, for instance. In such an embodiment, the step can coincide with, and be formed by, the junction between the D-duct 108 and adjacent sections of the nacelle 22, for instance. The heating conduit 110 can have a plurality of apertures forming heating air outlets, and be connected to a compressor to receive bleed air therefrom. An example of a possible arrangement is shown in
Returning to
The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. Still other 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.
Claims
1. An aircraft engine nacelle comprising an inlet fluidly connecting to a main gas path of a gas turbine engine core, the inlet having an inlet edge connecting an external skin to an internal duct wall, and a step formed in a surface of at least one of the external skin and the duct wall, the step delimiting a first portion of the surface from a second portion of the surface, the second portion of the surface being recessed relative to the first portion of the surface by a height of the step, the second portion of the surface extending away from both the step and the inlet edge, whereas the first portion of the surface extends between the inlet edge and the step.
2. The aircraft engine nacelle of claim 1 wherein the height of the step is of between 0.010″ and 0.200″ measured normal to the surface.
3. The aircraft engine nacelle of claim 1 wherein the step has a riser.
4. The aircraft engine nacelle of claim 1 wherein the aircraft engine is a turbofan engine, and the duct wall is an outer bypass duct wall.
5. The aircraft engine nacelle of claim 4 wherein the inlet edge is a portion of a D-duct, the D-duct connecting the skin and the duct wall.
6. The aircraft engine nacelle of claim 5 wherein the step is formed at a junction between the D-duct and the duct wall.
7. The aircraft engine nacelle of claim 5 wherein a heating air conduit is provided inside the D-duct, the heating air conduit having a plurality of heating air outlets, and being connected to a compressor bleed air source.
8. The aircraft engine nacelle of claim 4 wherein the inner duct wall has an acoustic panel, the step being located along the surface, between the inlet edge and the acoustic panel.
9. The aircraft engine nacelle of claim 1 wherein the step is formed in the duct wall.
10. The aircraft engine nacelle of claim 8 further comprising an other step formed in the skin, the other step delimiting a recessed portion of the skin, the recessed portion of the skin extending away from both the other step and the inlet edge.
11. The aircraft engine nacelle of claim 1 wherein the inlet edge, skin and duct are annular.
12. The aircraft engine nacelle of claim 11 wherein the step is backward facing.
13. A method of operating a gas turbine engine, the method including a flow of air circulating along a surface of an inlet portion of the gas turbine engine, the flow of air drawing water droplets along the surface until the water droplets reach an edge of a step leading to a recessed portion of the surface, the flow of air separating the water droplets from the surface at the edge of the step.
14. The method of claim 13 wherein the method further comprises subjecting water in solid state to heating, and thereby transforming the water in solid state into the water droplets.
15. The method of claim 13 further comprising directing the detached water droplets into one of an engine core main gas path, or a bypass duct.
16. An aircraft engine comprising a gas turbine engine core having a main gas path extending, in serial flow communication, across a compressor section, a combustor, and a turbine section, the gas turbine engine core housed within a nacelle, the nacelle having an inlet fluidly connecting the main gas path, the inlet having an inlet edge connecting an external skin to an internal duct wall, and a step formed in a surface of at least one of the skin and the duct wall, the step delimiting a first portion of the surface from a second portion of the surface, the second portion of the surface being recessed relative to the first portion of the surface, the second portion of the surface extending away from both the step and the inlet edge.
17. The aircraft engine of claim 16 wherein the step has a height of between 0.010″ and 0.200″ measured normal to the surface.
18. The aircraft engine of claim 16 wherein the aircraft engine is a turbofan engine, and the duct wall is an outer bypass duct wall.
19. The aircraft engine of claim 16 wherein the step is formed in the duct wall.
20. The aircraft engine of claim 16 wherein the inlet edge, skin and duct are annular.
Type: Application
Filed: Nov 28, 2019
Publication Date: Jun 3, 2021
Inventor: Philippe-André TETRAULT (Boucherville)
Application Number: 16/699,001