GB2595482A

GB2595482A - Aircraft propulsor

Info

Publication number: GB2595482A
Application number: GB2007955.4A
Authority: GB
Inventors: Macmanus David; Sheaf Christopher
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2020-05-28
Filing date: 2020-05-28
Publication date: 2021-12-01
Also published as: GB202007955D0

Abstract

An aircraft propulsor 200 includes a nacelle 208, a fan (116, fig 1) and a core plug 218. The nacelle includes a leading edge 210 and a trailing edge 212 downstream of the leading edge defining a duct 214 therein. The fan is received within the duct and is configured to rotate about a rotational axis Y-Y’ and generate a primary fluid flow. The core plug is at least partially received within the nacelle and includes a boundary wall 220 radially inward of the nacelle so that the primary fluid flow from the fan passes between the nacelle and the boundary wall 220. The core plug extends axially beyond the trailing edge of the nacelle relative to the rotational axis (Y-Y’) of the fan. The boundary wall of the core plug includes a flow control feature 226 located downstream of the trailing edge of the nacelle for reducing boundary layer separation of the primary fluid flow. The flow control feature may be slots (228, figure 3), vortex generators (240, figure 5) or jet vortex generators (244, figure 9) located on the boundary wall.

Description

AIRCRAFT PROPULSOR

FIELD OF THE DISCLOSURE

The present disclosure generally relates to an aircraft propulsor, and in 5 particular, to a flow control feature of an aircraft propulsor.

BACKGROUND

Exhaust systems of aircraft engines typically include a central core plug that extends downstream of an exhaust casing. Exhaust core plugs are generally designed to increase the efficiency of an exhaust flow discharged from the engine. For some engine configurations, the core plug has an increased length in order to provide a desired efficiency of the exhaust flow. This increase in length may be associated with an increase the mass of the core plug, and consequently the mass of the engine. An overall length of the engine may also increase.

SUMMARY

According to a first aspect there is provided an aircraft propulsor. The aircraft propulsor includes a nacelle including a leading edge and a trailing edge downstream of the leading edge. The nacelle defines a duct therein. The aircraft propulsor further includes a fan received within the duct and configured to rotate about a rotational axis and generate a primary fluid flow. The aircraft propulsor further includes a core plug at least partially received within the nacelle. The core plug includes a boundary wall radially inward of the nacelle so that the primary fluid flow from the fan passes between the nacelle and the boundary wall. The core plug extends axially beyond the trailing edge of the nacelle relative to the rotational axis of the fan. The boundary wall of the core plug includes a flow control feature located downstream of the trailing edge of the nacelle for reducing boundary layer separation of the primary fluid flow. The flow control feature may therefore improve a flow efficiency of the primary fluid flow.

The flow control feature may be configured to change the momentum of and/or introduce streamwise vorticity to a boundary layer of the primary fluid flow to reduce boundary layer separation of the primary fluid flow.

The flow control feature may be configured to output a secondary fluid flow to reduce boundary layer separation of the primary fluid flow.

The flow control feature may avoid premature and undesirable boundary layer separation of the primary fluid flow on the boundary wall of the core plug of the aircraft propulsor. The flow control feature may advantageously add momentum and/or introduce streamwise vorticity to the boundary layer of the primary fluid flow to avoid separation of the boundary layer from the boundary wall.

The flow control feature may allow for a reduction in a length and a mass of the 15 core plug due to the reduction in boundary layer separation of the primary fluid flow. Therefore, a compact and a lightweight core plug may be used in certain aircraft propulsor configurations.

The flow control feature may include one or more slots in the boundary wall.

Each of the one or more slots receives a secondary fluid flow therethrough. The one or more slots may increase the momentum of the boundary layer of the primary fluid flow at the boundary wall of the core plug, thereby preventing separation of the boundary layer of the primary fluid flow from the boundary wall.

The one or more slots may include at least one of an annular slot having an angular extent of 360 degrees relative to the rotational axis, a partially annular slot having an angular extent less than 360 degrees, and a non-axisymmetric slot.

At least one of the one or more slots may be defined between a first edge of an outer surface of the boundary wall and a second edge of the outer surface of the boundary wall that is radially inward and downstream of the first edge.

At least one of the one or more slots may be configured to output its secondary fluid flow at an angle inclined relative to an outer surface of the boundary wall. This may allow improved control of the secondary fluid flow to effectively increase the momentum of the boundary layer of the primary fluid flow based on a type and an operation of the aircraft propulsor.

The flow control feature may further include a plurality of vortex generators disposed on the boundary wall downstream of the one or more slots. The vortex generators may be configured to receive at least a part of the secondary fluid flow from the one or more slots. The arrangement of the one or more slots and the vortex generators may more effectively prevent separation of the boundary layer of the primary fluid flow from the boundary wall.

The flow control feature may include a plurality of vortex generators disposed on the boundary wall downstream of the trailing edge of the nacelle. The vortex generators may introduce streamwise vorticity to the boundary layer of the primary fluid flow to avoid separation of the boundary layer from the boundary wall Each vortex generator may include an upper surface inclined with respect to the outer surface of the boundary wall.

Each vortex generator may protrude a perpendicular distance HvG from the outer surface of the boundary wall and extends a length L3vG along the outer surface 25 of the boundary wall. L3vG may be about 3 to about 5 times HvG.

At least one vortex generator may be inclined azimuthally with respect to an outer surface of the boundary wall. An azimuthal angle between the at least one vortex generator and the outer surface of the boundary wall may be from about 30 10 degrees to about 15 degrees.

The flow control feature may include a plurality of jet vortex generators. Each jet vortex generator may include an aperture in the boundary wall for outputting a secondary fluid flow. The jet vortex generators may increase the momentum of the boundary layer of the primary fluid flow at the boundary wall of the core plug, thereby avoiding separation of the boundary layer of the primary fluid flow from the boundary wall.

Each jet vortex generator may be configured to output its secondary fluid flow at an angle inclined relative to an outer surface of the boundary wall. This may allow improved control of the secondary fluid flow to effectively increase the momentum of the boundary layer based on the type and the operation of the aircraft propulsor.

At least one jet vortex generator may be configured to output its secondary fluid flow at an angle inclined azimuthally with respect to an outer surface of the boundary wall. The azimuthal angle may be from about 30 degrees to about 60 degrees.

The propulsor may further include a plurality of angularly spaced apart jet vortex generators An angle between the adjacent angularly spaced jet vortex generators may be from about 10 degrees to about 20 degrees relative to the rotational axis. The propulsor may further include a plurality of jet vortex 20 generators that are axially spaced from each other relative to the rotational axis.

At least a portion of the boundary wall may have a frustoconical shape. The portion of the boundary wall may have a half angle from about 10 degrees to about 25 degrees.

The propulsor may further include a guide vane connecting the nacelle to the core plug.

A space between the trailing edge of the nacelle and the core plug may define a 30 sole exhaust of the propulsor.

According to a second aspect there is provided an aircraft propulsion system. The aircraft propulsion system includes a prime mover and the propulsor of the first aspect. The fan of the propulsor is drivably couplable to the prime mover.

S

The prime mover may be at least one of an electric motor and a gas turbine engine.

The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of an aircraft propulsion system having an aircraft propulsor Figure 2 is a partial sectional side view of the aircraft propulsor having a core plug with a flow control feature; Figure 3 is a partial sectional side view of the aircraft propulsor with the flow control feature including one or more slots; Figures 4A-4C are partial rear views of the core plug of the aircraft 20 propulsor illustrating different configurations of the one or more slots of Figure 3; Figure 5 is a partial sectional side view of the aircraft propulsor with the flow control feature including a plurality of vortex generators; Figure 6 is a partial rear view of the core plug of the aircraft propulsor illustrating the vortex generators of Figure 5; Figure 7 is a partial sectional side view of the aircraft propulsor with the flow control feature including a plurality of vortex generators disposed downstream of the one or more slots; Figure 8 is a partial rear view of the core plug of the aircraft propulsor illustrating the vortex generators and the one or more slots of Figure 7; Figure 9 is a partial sectional side view of the aircraft propulsor with the flow control feature including a plurality of jet vortex generators; and Figures 10A-10B are partial rear views of the core plug of the aircraft propulsor illustrating the jet vortex generators of Figure 9.

DETAILED DESCRIPTION

Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will 5 be apparent to those skilled in the art.

In the following disclosure, the following definitions are adopted. The terms "upstream" and "downstream" are considered to be relative to an air flow or gas flow through an aircraft propulsor. The terms "axial" and "axially" are considered 10 to relate to the direction of a rotational axis X-X' of an aircraft propulsor.

Figure 1 illustrates a sectional side view of an aircraft propulsion system 100. The aircraft propulsion system 100 includes an aircraft propulsor 102 (hereinafter referred to as "the propulsor 102"). In the illustrated example, the propulsor 102 is attached to a wing 104 of the aircraft (not shown) through a pylon 106. The propulsor 102 includes a nacelle 108. The nacelle 108 includes a leading edge 110 and a trailing edge 112 downstream of the leading edge 110. The nacelle 108 defines a duct 114 between the leading edge 110 and the trailing edge 112. The propulsor 102 further includes a fan 116 received within the duct 114. The fan 116 is configured to rotate about the rotational axis X-X' of the fan 116 and generate a primary fluid flow as shown by an arrow A. The propulsor 102 further includes a core plug 118 at least partially received within the nacelle 108. The core plug 118 includes a boundary wall 120 radially inward of the nacelle 108 so that the primary fluid flow from the fan 116 passes between the nacelle 108 and the boundary wall 120. The core plug 118 extends axially beyond the trailing edge 112 of the nacelle 108 relative to the rotational axis X-X' of the fan 116. At least a portion of the boundary wall 120 may have a frustoconical shape. For example, the portion of the boundary wall 120 in Figure 1 has a half angle from about 10 degrees to about 25 degrees. The propulsor 102 may further include a secondary fluid flow, which may be provided by a secondary fluid source 122.

In some examples, the secondary fluid source 122 is a local compressor. In some examples, the secondary fluid source 122 is part of an auxiliary power unit (APU). In some examples, the secondary fluid flow is be a bleed airflow that is diverted from the duct 114. It should be understood that the secondary air flow may be generated using any other source in fluid communication with the propulsor 102.

The aircraft propulsion system 100 further includes a prime mover 124. The fan 116 is drivably coupled to the prime mover 124. The prime mover 124 generates power that is used to drive the fan 116. The prime mover 124 is at least one of an electric motor, a gas turbine engine, or a combination thereof. For example, the prime mover 124 may be an engine, such as a gas turbine engine, or a combination of an engine and an electric motor, such as a hybrid propulsion system.

In the case of hybrid propulsion systems, the engine and/or the electric motor may be used to drive the fan 116. For example, the electric motor may be used to supply rotational power to the fan 116 during periods of operation in which the electric motor can provide greater efficiency than the engine and/or in which a short-term or a low level of propulsive power is required, whereas the engine may be operated during periods of operation in which a long-term or a more constant, sustained level of propulsive power is required. In some examples, the aircraft propulsion system 100 includes multiple fans or electric motors or engines, such as a distributed propulsion system. It will be appreciated that the arrangement shown in Figure 1 is by way of example only, and various alternatives are within the scope of the present disclosure.

As shown in Figure 1, the prime mover 124 is coupled to the fan 116 via a speed reduction device 126. In some examples, the speed reduction device 126 is an epicyclic gearbox. In some examples, the epicyclic gearbox is a planetary gear system. However, the speed reduction device 126 may be any type of speed reduction device. For example, the speed reduction device 126 may be a layshaft-based gear system. The use of the speed reduction device 126 may allow the prime mover 124 to operate at a higher speed, and the fan 116 to operate at a lower speed as compared to the speed of the prime mover 124. Further, a larger fan may be utilized due to the reduction of tip speed.

In some examples, the prime mover 124 is a gas turbine engine. Gas turbine 5 engines typically include one or more compressors, a combustor, and one or more turbines. The fan 116 may provide an engine air flow to the gas turbine engine. The engine air flow may then be compressed by the one or more compressors. The compressed air exhausted from the one or more compressors is directed into the combustor where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive the one or more turbines before being exhausted through the core plug 118. The one or more turbines may drive the one or more compressors by suitable interconnecting shafts.

Figure 2 illustrates a partial sectional side view of an arrangement of an aircraft propulsor 200 (hereinafter referred to as "the propulsor 200") of the aircraft propulsion system 100. The propulsor 200 includes a single primary fluid flow as shown by an arrow B. The propulsor 200 includes a nacelle 208. The nacelle 208 includes a leading edge 210 and a trailing edge 212 downstream of the leading edge 210. The nacelle 208 defines a duct 214 between the leading edge 210 and the trailing edge 212. The propulsor 200 further includes the fan 116 (shown in Figure 1) received within the duct 214 and configured to rotate about a rotational axis Y-Y' of the fan 116 and generate the primary fluid flow.

The propulsor 200 further includes a core plug 218 at least partially received within the nacelle 208. The core plug 218 includes a boundary wall 220 radially inward of the nacelle 208 so that the primary fluid flow from the fan 116 passes between the nacelle 208 and the boundary wall 220. The boundary wall 220 extends circumferentially about the rotational axis Y-Y' of the fan 116. The core plug 218 extends axially beyond the trailing edge 212 of the nacelle 208 relative to the rotational axis Y-Y' of the fan 116. The core plug 218 includes a trailing edge 224 at a rear end. The trailing edge 224 of the core plug 218 is downstream of the trailing edge 212 of the nacelle 208.

The core plug 218 generally tapers to the trailing edge 224. As shown in Figure 2, the boundary wall 220 generally tapers downstream of the nacelle 208. At least a portion of the boundary wall 220 has a frustoconical shape. For example, the portion of the boundary wall 220 has a half angle HA from about 10 degrees to about 25 degrees.

The propulsor 200 further includes a guide vane 222 connecting the nacelle 208 to the core plug 218. In some cases, for example where the prime mover 124 is not a gas turbine engine, a space between the trailing edge 212 of the nacelle 208 and the core plug 218 defines a sole exhaust of the propulsor 200. However, in other cases there may be multiple exhaust fluid flows from the propulsor 200.

The boundary wall 220 of the core plug 218 includes a flow control feature 226 located downstream of the trailing edge 212 of the nacelle 208. The flow control feature 226 is configured to reduce boundary layer separation of the primary fluid flow, for example by changing the momentum of and/or introducing streamwise vorticity to a boundary layer of the primary fluid flow to reduce boundary layer separation of the primary fluid flow. In some examples, the flow control feature 226 outputs a secondary fluid flow to reduce boundary layer separation of the primary fluid flow. Various examples of the flow control feature 226 will be described hereinafter with reference to Figures 3-10B.

In the example of Figure 3 the flow control feature 226 includes one or more slots 228 in the boundary wall 220. The one or more slots 228 receive a secondary fluid flow (shown by an arrow C) therethrough. At least one of the one or more slots 228 is defined between a first edge 230 of an outer surface 231 of the boundary wall 220 and a second edge 232 of the outer surface 231 of the boundary wall 220 that is radially inward and downstream of the first edge 230. The second edge 232 is radially inward of the first edge 230 relative to the rotational axis Y-Y'. The secondary fluid flow from at least one of the one or more slots 228 may be used to reenergise the boundary layer of the primary fluid flow. This may modify the momentum deficit of the primary fluid flow close to the boundary wall 220, thereby reducing boundary layer separation of the primary fluid flow.

The effectiveness of the one or more slots 228 may depend upon their distances from the trailing edge 212 of the nacelle 208 and the trailing edge 224 of the core plug 218, as well as upon the slot height. The slot height can be defined as Hs, the perpendicular distance between the first edge 230 and the second edge 232. The distance from the trailing edge 212 of the nacelle 208 can be defined as L1s, the perpendicular distance between a first line that passes through the slot 228 and is perpendicular to the boundary wall 220 and a second line that passes through the trailing edge 212 of the nacelle 208 and is parallel to the first line. The distance from the trailing edge 224 can be defined as L25, the distance between the slot 228 and the trailing edge 224 of the core plug 218 measured parallel to the boundary wall 220 of the core plug 218. The location of the one or more slots 228 may then be defined by: Lis 0.05 < < 0.6 Lls + L2s The value of this ratio may, in some examples, be between 0.2 and 0.5, between 0.25 and 0.45 or perhaps between 0.3 and 0.4.

The level of flow control from the one or more slots 228 may depend upon a slot massflow Ms, an associated slot velocity Vs, and a momentum added relative to a local flow velocity at the boundary wall 220. In some examples, a size (e.g., the slot height Hs) of the one or more slots 228 is selected based on a predicted or measured size of a local boundary layer. For example, the slot height Hs of at least one of the one or more slots 228 may be selected to be less than half of the size of the local boundary layer. Typical boundary layer thicknesses may be about 10-40mm, so the slot height Hs may be about 5-20mm, although smaller slot heights Hs of 5-10mm may also be appropriate, especially for smaller engines.

At least one of the one or more slots 228 is configured to output its secondary fluid flow at an angle inclined relative to the boundary wall 220. The angle of the secondary fluid flow relative to the boundary wall 220 can be defined as the angle as formed between the ejected massflow and the boundary wall 220. It should be understood that appropriate values for the slot parameters defined above can be chosen based on desired application attributes.

Figure 4A-4C are partial rear views of the core plug 218 of the aircraft propulsor 200 illustrating different configurations of the one or more slots 228 of Figure 3.

As shown in Figure 4A-4C, the one or more slots 228 may include at least one of an annular slot 234 having an angular extent of 360 degrees, a partially annular slot 236 having an angular extent less than 360 degrees, and a nonaxisymmetric slot 238. Figure 4A illustrates the annular slot 234 having an angular extent of 360 degrees relative to the rotational axis Y-Y'. Figure 4B illustrates the partially annular slot 236 having an angular extent less than 360 degrees. In some examples, multiple partially annular slots 236 are provided angularly spaced from each other. An angular extent of each partially annular slot 236 may be equal to or less than 180 degrees, 90 degrees, 60 degrees, or 30 degrees. In some examples, the multiple partially annular slots 236 are uniformly distributed about the rotational axis Y-Y'. In some other examples, the multiple partially annular slots 236 are non-uniformly distributed about the rotational axis Y-Y'. Figure 4C illustrates the non-axisymmetric slot 238 relative to the rotational axis Y-Y'.

The slot massflow Ms and the slot velocity Vs may differ between the annular slot 234, the partially annular slot 236 and the non-axisymmetric slot 238. It should be understood that any combination of the one or more annular slots as described above can be utilized and various alternatives are within the scope of the present disclosure.

In the example of Figure 5 the flow control feature 226 includes a plurality of vortex generators 240 disposed on the boundary wall 220 downstream of the trailing edge 212 of the nacelle 208. The vortex generators 240 introduce streamwise vorticity and increase momentum of the boundary layer of the primary fluid flow. This reduces separation of the boundary layer of the primary fluid flow from the boundary wall 220.

Each vortex generator 240 includes an upper surface 242 inclined with respect to the boundary wall 220. The location of the vortex generators 240 can be defined in terms of their distances from the trailing edge 212 of the nacelle 208 and the trailing edge 224 of the core plug 218. The distance from the trailing edge 212 of the nacelle 208 can be defined as distance L1vG, which is the perpendicular distance between a first line that passes through the centre of the vortex generator 240 and is perpendicular to the boundary wall 220 and a second line that passes through the trailing edge 212 of the nacelle 208 and is parallel to the first line. The distance from the trailing edge 224 of the core plug 218 can be defined as distance L2vG, which is the distance between the centre of the vortex generator 240 and the trailing edge 224 of the core plug 218 when measured parallel to the boundary wall 220 of the core plug 218. The length of the vortex generators 240 can be defined as distance L3vG, which is the distance along the boundary wall 220 the vortex generator 240 extends.

The locations of the vortex generators 240 may then be defined by: LlvG 0.05 < < 0.6 LlvG + L2vG The value of this ratio may, in some examples, be between 0.2 and 0.5, between 0.25 and 0.45 or perhaps between 0.3 and 0.4. Circumferentially, adjacent vortex generators 240 may be spaced apart by between about 5 and 20 25 degrees.

Each vortex generator 240, which may be arranged to produce co-rotating or counter-rotating vortices, protrudes a perpendicular distance HvG from the outer surface 231 of the boundary wall 220. In some examples, HvG is selected to be between 0.5 and 1.5 times the predicted or measured height of a local boundary layer. In some examples, L3vG is about 3 to about 5 times HvG. As noted above, typical boundary layer thicknesses may be about 10-40mm. Thus, HvG may be about 5-60mm, although smaller slot heights HvG of 5-30mm may also be appropriate, especially for smaller engines. L3vG may then be about 30-200mm, for example 30-120mm. Smaller values of L3vG, for example, 15-150mm or 15-90mm may also be appropriate, especially for smaller engines.

At least one of the vortex generators 240 may be inclined azimuthally with respect to the boundary wall 220. For example, at least one of the vortex generators 240 may be inclined relative to the primary fluid flow to generate streamwise vortices at an angle. The azimuthal angle can be defined as avG (shown in the front-on view of the rear of the core plug 218 in Figure 6). In some examples, the azimuthal angle avG between at least one of the vortex generators 240 and the boundary wall 220 is from about 10 degrees to about 15 degrees.

The level of flow control from each vortex generator 240 may depend upon any one or more of the distance L1 va the distance L2vG, the azimuthal angle avG, and the size of the vortex generators 240 (HvG and L3vG). The effectiveness of the vortex generators 240 may depend up on the co-rotating and counter-rotating arrangements of the vortex generators 240.

In some examples, multiple vortex generators 240 are circumferentially arranged around the core plug 218. Figure 6 illustrates a circumferential arrangement of the vortex generators 240. In some examples, the vortex generators 240 are arranged around the full circumferential extent of the core plug 218, in other words along an angular extent of 360 degrees. In another example, the vortex generators 240 are arranged around the core plug 218 along an angular extent less than 360 degrees, in other words less than the full circumferential extent of the core plug 218. In other examples, the vortex generators 240 are located around the core plug 218 in a non-axisymmetric arrangement. In some examples, the vortex generators 240 are located in discrete circumferential ranges around the core plug 218. For instance, there may be a group of vortex generators between 0 and 45 degrees, no vortex generators between 45 and 90 degrees, and another group of vortex generators between 90 and 135 degrees. In some examples, the vortex generators 240 are located in a range of axial and/or azimuthal positions around the core plug 218 with respect to the rotational axis Y-Y'. The arrangement of the vortex generators 240 shown in Figure 6 is exemplary in nature and various alternative arrangements are possible.

As shown in Figure 7, the flow control feature 226 may include both vortex generators 241 and one or more slots 228. The vortex generators 241 are disposed on the boundary wall 220 downstream of the one or more slots 228. The vortex generators 241 are configured to receive at least a part of the secondary fluid flow from the one or more slots 228. The vortex generators 241 may be similar to the vortex generators 240 of Figures 5 and 6.

The dimensions and parameters of the one or more slots 228 may be same as described above with reference to Figures 3 and 4. Similarly, the dimensions and parameters of the vortex generators 241 may be same as that of the vortex generators 240. The relative dimensions and parameters of the one or more slots 228 and the vortex generators 241 can be chosen based on desired application attributes.

A distance AsvG can be defined for the combination of a slot 228 and a vortex 20 generator 241. Specifically, AsvG can be defined as the distance from a slot 228 to the end of the downstream edge of the vortex generator 241, measured along the boundary wall 220. In some examples, the effectiveness of the flow control feature 226 may depend upon the distance Asvc. related to the slot height Hs by: In some examples AsvG is a SV G r "--^ "--^ is The value of this ratio may, in some examples, be between 1.5 and 4.5 or perhaps between 2 and 4.

The size and parameters of the one or more slots 228 can be varied as per desired application attributes. Similarly, the size and parameters of each vortex generator 241 can also be varied.

Figure 8 illustrates a partial front-on view of the rear of the core plug 218 of the aircraft propulsor 200 illustrating the flow control feature 226 of Figure 7. As shown in Figure 8, the one or more slots 228 may be an annular slot 229 having 5 a circumferential extent of 360 degrees. The vortex generators 241 are located adjacent to and downstream of the annular slot 229. The vortex generators 241 are circumferentially arranged along a circumferential extent of 360 degrees. The relative arrangement of the one or more slots 228 and the vortex generators 241 shown in Figure 8 is exemplary in nature, and various alternative 10 arrangements of the one or more slots 228 and the vortex generators 241 are possible.

In the example of Figure 9, the flow control feature 226 includes a plurality of jet vortex generators 244. Each jet vortex generator 244 includes an aperture in the boundary wall 220 for outputting the secondary fluid flow. At least one of the jet vortex generators 244 introduces streamwise vorticity and increases momentum of the boundary layer of the primary fluid flow close to the outer surface 231 of the boundary wall 220. This may reduce separation of the boundary layer of the primary fluid flow from the boundary wall 220.

The location of the jet vortex generators 244 can be defined by their distance LljvG from the trailing edge 212 of the nacelle 208 and their distance L2jvG from the trailing edge 224 of the core plug 218. The distance Ll jvG can be defined as the perpendicular distance between a first line passing through a jet vortex generator 244 perpendicular to the boundary wall 220 and a second line passing through the trailing edge 212 of the nacelle 208 and parallel to the first line. The distance L2jvG can be defined as the distance between a jet vortex generator 244 and the trailing edge 224 of the core plug 218 when measured parallel to the boundary wall 220 of the core plug 218. The locations of the jet vortex generators 244 may then be defined by: LljvG 0.05 < < 0.6 LljvG The value of this ratio may, in some examples, be between 0.2 and 0.5, between 0.25 and 0.45 or perhaps between 0.3 and 0.4.

Some or all of the jet vortex generators 244 may be configured to output their 5 secondary fluid flow at an angle ajvG inclined relative to the boundary wall 220. Some or all of the jet vortex generators 244 may also be configured to output their secondary fluid flow at an angle 13.Jvc; (shown in Figure WA) inclined azimuthally with respect to the boundary wall 220. In some examples, the azimuthal angle is from about 30 degrees to about 60 degrees. Each of the jet 10 vortex generators 244 may be arranged to produce co-rotating or counter-rotating vortices.

The level of flow control from each jet vortex generator 244 may depend upon one or more of the distance LING, the distance LZNG, the angle ajvG, and the angle 8NG. The level of flow control from each jet vortex generator 244 may further depend upon a massflow MJVG, an associated velocity VJVG, and a momentum added relative to a local flow velocity at the boundary wall 220. The effectiveness of each jet vortex generator 244 may depend up on the co-rotating and counter-rotating arrangements of the jet vortex generators 244. In some examples, a ratio of an exit velocity to a local velocity of each jet vortex generator 244 is about 1.5 to 2.5.

The propulsor 200 may include a plurality of angularly spaced apart jet vortex generators 244, as shown in Figure 10A. In some examples, an angle between adjacent angularly spaced jet vortex generators 244 is from about 10 degrees to about 20 degrees relative to the rotational axis Y-Y'. In some examples, as shown in Figure 103, the propulsor 200 further includes a plurality of jet vortex generators 244 that are axially spaced from each other along the rotational axis Y-Y'.

The jet vortex generators 244 may be arranged around the full circumferential extent of the core plug 218, in other words along an angular extent of 360 degrees. In another example, the jet vortex generators 244 are arranged around the core plug 218 along an angular extent less than 360 degrees, in other words less than the full circumferential extent of the core plug 218. In other examples, the jet vortex generators 244 are located around the core plug 218 in a nonaxisym metric arrangement. In some examples, the jet vortex generators 244 are located in discrete circumferential ranges around the core plug 218. For instance, there may be a group of vortex generators between 0 and 45 degrees, no vortex generators between 45 and 90 degrees, and another group of vortex generators between 90 and 135 degrees.

The arrangement of the jet vortex generators 244 shown in Figures 10A and 10B are exemplary in nature and various alternative arrangements are possible. In the illustrated examples of Figures 10A and 10B, each jet vortex generator 244 has a circular shape. In some other examples, the jet vortex generators 244 include apertures having different alternative shapes, such as square, rectangular, oval, elliptical, polygonal, and the like, based on application requirements.

The use of flow control features 226 as described above with reference to Figures 3-10B may reduce a required length and mass of the core plug 218 while maintaining flow efficiency of the primary fluid flow.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

CLAIMS: 1. An aircraft propulsor (200) comprising: a nacelle (208) comprising a leading edge (210) and a trailing edge (212) downstream of the leading edge (210), wherein the nacelle (208) defines a duct (214) therein; a fan (116) received within the duct (214) and configured to rotate about a rotational axis (Y-Y') and generate a primary fluid flow; a core plug (218) at least partially received within the nacelle (208) and comprising a boundary wall (220) radially inward of the nacelle (208) so that the primary fluid flow from the fan (116) passes between the nacelle (208) and the boundary wall (220), the core plug (218) extending axially beyond the trailing edge (212) of the nacelle (208) relative to the rotational axis (Y-Y') of the fan (116); and wherein the boundary wall (220) of the core plug (218) comprises a flow control feature (226) located downstream of the trailing edge (212) of the nacelle (208) for reducing boundary layer separation of the primary fluid flow.
2. The propulsor (200) of claim 1, wherein the flow control feature (226) is configured to change the momentum of and/or introduce streamwise vorticity to a boundary layer of the primary fluid flow to reduce boundary layer separation of the primary fluid flow.
3. The propulsor (200) of claim 1 or claim 2, wherein the flow control feature (226) is configured to output a secondary fluid flow to reduce boundary layer separation of the primary fluid flow.
4 The propulsor (200) of any of claims 1 to 3, wherein the flow control feature (226) comprises one or more slots (228) in the boundary wall (220), and wherein each of the one or more slots (228) receives a secondary fluid flow therethrough.
5. The propulsor (200) of claim 4, wherein the one or more slots (228) comprises at least one of: an annular slot (234) having an angular extent of 360 degrees relative to the rotational axis (Y-Y'); a partially annular slot (236) having an angular extent less than 360 degrees; and a non-axisymmetric slot (238).
6. The propulsor (200) of claim 4 or 5, wherein at least one of the one or more slots (228) is defined between a first edge (230) of an outer surface (231) of the boundary wall (220) and a second edge (232) of the outer surface (231) of the boundary wall (220) that is radially inward and downstream of the first edge (230).
The propulsor (200) of any one of claims 4 to 6, wherein at least one of the one or more slots (228) is configured to output its secondary fluid flow at an angle (as) inclined relative to an outer surface (231) of the boundary wall (220).
The propulsor (200) of any one of claims 4 to 7, wherein the flow control feature (226) further comprises a plurality of vortex generators (241) disposed on the boundary wall (220) downstream of the one or more slots (228, 229) and configured to receive at least a part of the secondary fluid flow from the one or more slots (228).
The propulsor (200) of claim 1, wherein the flow control feature (226) comprises a plurality of vortex generators (240) disposed on the boundary wall (220) downstream of the trailing edge (212) of the nacelle (208).
The propulsor (200) of claim 8 or claim 9, wherein each vortex generator (240) comprises an upper surface (242) inclined with respect to the outer surface (231) of the boundary wall (220). 7. 8. 9. 10.
11. The propulsor (200) of any one of claim 8 to 10, wherein each vortex generator (240) protrudes a perpendicular distance HvG from the outer surface (231) of the boundary wall (220) and extends a length L3vG along the outer surface (231) of the boundary wall (220), wherein L3vG is about 3 to about 5 times HVG.
12. The propulsor (200) of any one of claims 8 to 11, wherein at least one vortex generator (240) is inclined azimuthally with respect to an outer surface (231) of the boundary wall (220).
13. The propulsor (200) of claim 12, wherein an azimuthal angle (avG) between the at least one vortex generator (240) and the outer surface (231) of the boundary wall (220) is from about 10 degrees to about 15 degrees.
14. The propulsor (200) of any one of the preceding claims, wherein the flow control feature (226) comprises a plurality of jet vortex generators (244), each jet vortex generator (244) comprising an aperture in the boundary wall (220) for outputting a secondary fluid flow.
15. The propulsor (200) of claim 14, wherein each jet vortex generator (244) is configured to output its secondary fluid flow at an angle (aJvG) inclined relative to an outer surface (231) of the boundary wall (220).
16. The propulsor (200) of claim 15, wherein at least one jet vortex generator (244) is configured to output its secondary fluid flow at an angle (6.jvG) inclined azimuthally with respect to an outer surface (231) of the boundary wall (220).
17. The propulsor (200) of claim 16, wherein the azimuthal angle (6jvG) is from about 30 degrees to about 60 degrees.
18. The propulsor (200) of any one of claims 14 to 17, comprising a plurality of angularly spaced apart jet vortex generators (244), wherein an angle between adjacent angularly spaced jet vortex generators (244) is from about 10 degrees to about 20 degrees relative to the rotational axis (Y-Y').
19. The propulsor (200) of any one of claims 14 to 18, comprising a plurality of jet vortex generators (244) that are axially spaced from each other relative to the rotational axis (Y-Y').
20. The propulsor (200) of any one of the preceding claims, wherein at least a portion of the boundary wall (220) has a frustoconical shape.
21. The propulsor (200) of claim 20, wherein the portion of the boundary wall (220) has a half angle (HA) from about 10 degrees to about 25 degrees.
22. The propulsor (200) of any one of the preceding claims, further comprising a guide vane (222) connecting the nacelle (208) to the core plug (218).
23. The propulsor (200) of any one of the preceding claims, wherein a space between the trailing edge (212) of the nacelle (208) and the core plug (218) define a sole exhaust of the propulsor (200).
24. An aircraft propulsion system (100) comprising: a prime mover (124); and the propulsor (200) of any one of the preceding claims, wherein the fan (116) is drivably couplable to the prime mover (124).
25. The propulsion system (100) of claim 24, wherein the prime mover (124) is at least one of an electric motor and a gas turbine engine.