FAN NACELLE TRAILING EDGE

Description

BACKGROUND

[0001] A turbofan engine is a gas turbine engine that includes a fan arranged at the front of the gas turbine engine. Some of the air passing through the fan is directed into a core of the gas turbine engine. The remainder of the air bypasses the core through a fan duct that surrounds the core. The fan duct includes a location with a minimum cross-sectional area, referred to as a nozzle. During certain flight conditions, a pressure ratio across the nozzle may be insufficient, causing a momentary fan stall.

[0002] EP 3 023 623 A1, in accordance with its abstract, states a gas turbine engine having a forward thrust mode and a reverse thrust mode. The gas turbine engine includes a variable pitch fan configured for generating forward thrust in the forward thrust mode of the engine and reverse thrust in the reverse thrust mode of the engine. The engine also includes a fan cowl surrounding the variable pitch fan, wherein the fan cowl forms a bypass duct for airflow generated by the fan. The fan cowl includes an aft edge that defines a physical flow area of the bypass duct, and a deflection device configured for deflecting airflow near the aft edge, wherein the deflection device is configured for operation in the reverse thrust mode of the engine. The physical flow area of the bypass duct at the aft edge remains the same in the forward thrust mode of the engine and in the reverse thrust mode of the engine.

[0003] GB 1 106 077 A, in accordance with its abstract, states a gas turbine jet power plant having a passage for air by-passing the air used for combustion, the passage leading to a discharge nozzle, an inflatable device partly defining the outer wall of the passage or nozzle and being capable of being expanded against a rigid outer shroud radially inwards to vary the gross sectional area of the discharge nozzle.

[0004] US 5 806 303 A, in accordance with its abstract, states a multiple bypass turbofan engine including a core engine assembly has a fan bypass duct radially outward of the core engine assembly and has first and second inlets disposed between forward and aft fans driven by a low pressure turbine and a core engine turbine respectively. An inlet duct having an annular duct wall is disposed radially inward of the bypass duct and connects the second inlet to the bypass duct and has disposed within a supercharger means for compressing air which is drivingly connected to the core turbine. The engine has an duct with an afterburner in an upstream portion and at a downstream end it has an exhaust nozzle with a fixed geometry throat. The nozzle may also have a fixed geometry exhaust nozzle outlet with or without a means for blowing at the throat. One embodiment of the aft fan may have radially inner and outer rows of aft fan vane airfoils separated by a non-rotatable portion of the annular duct wall such that the outer row of aft fan vane airfoils are disposed in the inlet duct and the aft fan vane airfoils are independently variable. Radially inner and outer rows of aft fan rotor blade airfoils are separated by a rotatable portion of the annular duct wall such that the outer row of aft fan rotor blade airfoils are disposed in the inlet duct adjacent to and longitudinally aft of radially inner and outer rows of aft fan vane airfoils, respectively thus providing the supercharger means. Radially inner and outer rows of aft fan rotor blade airfoils are separated by a rotatable portion of the annular duct wall such that the outer row of aft fan rotor blade airfoils are disposed in the inlet duct adjacent to and longitudinally aft of radially inner and outer rows of aft fan vane airfoils, respectively thus providing the supercharger means.

[0005] US 5 305 599 A, in accordance with its abstract, states a control of a gas turbine engine, suitable for powering an aircraft, accomplished by varying outlet nozzle cross section and/or a mixing of bypass air with exhaust gas at an inlet to the nozzle. The engine includes a core engine having a combustion chamber and high and low pressure turbines followed by an exhaust plenum. High and low pressure compressors driven by the turbine assembly direct air into the combustion chamber and into a bypass duct. The bypass duct extends from an outlet of the low pressure compressor. Pressure of fluid (air or exhaust) is measured at an inlet to the high pressure compressor, in the bypass duct, and in the exhaust plenum. Ratios of the pressure of the duct pressure or of the pressure at the inlet of the high-pressure compressor, to the plenum pressure are employed in a feedback loop for control of nozzle size, and also of mixing valve position in an alternate embodiment of the invention. The feedback loop includes computation of a desired pressure ratio based on desired power and inlet temperature to the low-pressure compressor.

SUMMARY

[0006] According to one aspect, a gas turbine engine nacelle assembly is provided in accordance with claim 1.

[0007] According to one aspect, an aircraft is provided in accordance with claim 9.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

[0008] 

Figure 1A is a schematic side view of a gas turbine engine;

Figure 1B is a side view of a trailing edge of a fan nacelle for the gas turbine engine of Figure 1A according to an example not forming part of the claimed subject matter;

Figure 2A is a side view of a trailing edge of a fan nacelle for the gas turbine engine of Figure 1A according to another example not forming part of the claimed subject matter;

Figure 2B is a side view of a trailing edge of a fan nacelle for the gas turbine engine of Figure 1A according to the claimed subject matter;

Figures 3A, 3B, 4A, 4B are examples which do not relate to the claimed subject matter;

Figure 3A is a side view of a trailing edge of a fan nacelle for the gas turbine engine of Figure 1A, wherein the trailing edge includes a movable member illustrated in a stowed position;

Figure 3B is a side view of the trailing edge of Figure 3A, wherein the movable member is illustrated in a deployed position;

Figure 4A is a side view of a trailing edge of a fan nacelle the gas turbine engine of Figure 1A, wherein the trailing edge includes an inflatable member illustrated in an uninflated state;

Figure 4B is a side view of the trailing edge of Figure 4A, wherein the inflatable member is illustrated in an inflated state; and

Figure 5 is a top view of an aircraft comprising turbofan engines according to at least one aspect.

DETAILED DESCRIPTION

[0009] In the following, reference is made to aspects presented in this disclosure. However, the scope of the present disclosure is not limited to specific described aspects. Although aspects disclosed herein may achieve advantages over other possible solutions or over the prior art, whether or not a particular advantage is achieved by a given aspect is not limiting of the scope of the present disclosure. Thus, the following aspects, features, aspects and advantages are merely illustrative and are not considered elements or limitations of the appended claims except where explicitly recited in a claim(s). Likewise, reference to "the invention" shall not be construed as a generalization of any inventive subject matter disclosed herein and shall not be considered to be an element or limitation of the appended claims except where explicitly recited in a claim(s).

[0010] In aspects described herein, a trailing edge of a fan nacelle for a gas turbine engine includes a stationary or movable exterior surface that diverts air flowing over an exterior of the fan nacelle away from a region downstream of a fan duct of the gas turbine engine. Diverting this air decreases air pressure in the region downstream of the fan duct. The reduced pressure of this region results in an increased fan pressure ratio across a nozzle of the fan duct, thereby improving airflow through the fan duct and reducing the likelihood that the fan of the turbofan engine experiences a momentary fan stall.

[0011] Figure 1A is a side view of a gas turbine engine 100 according to one aspect, wherein certain features that would be blocked from view are illustrated in broken line. The gas turbine engine 100 includes an engine core 102 surrounded by a core nacelle 118. Components of the engine core 102 rotate about a longitudinal axis 130 to power a fan 116 arranged in front of the engine core 102. The fan 116 is surrounded by a fan nacelle 104. The core nacelle 118 and the fan nacelle 104 comprise a gas turbine engine nacelle assembly 190. A portion of the air passing the fan 116 enters the engine core 102 (as indicated by arrows A) and the remainder of the air passing the fan 116 enters a fan duct 112 (as indicated by arrows B). The fan 116, during operation, rotates to drive air into the fan duct 112, thereby increasing the pressure of the air in the fan duct 112 over ambient air pressure, such as the air pressure in the region 120 aft of the fan duct 112. The increased pressure of the air in the fan duct 112 divided by the pressure of the air in the region 120 aft of the fan duct 112 defines a fan pressure ratio. Airflow through the fan duct 112 is a function of the fan pressure ratio. As the fan pressure ratio decreases toward the value of 1, airflow through the fan duct 112 decreases and the likelihood of a fan stall increases. A fan stall is a momentary cessation or reversal of airflow through the fan.

[0012] The fan nacelle 104 includes a leading edge 106 and a trailing edge 108 the fan nacelle 104 also includes a fan duct skin 110 arranged in the fan duct 112. The fan nacelle 104 also includes an exterior skin 114 arranged to direct airflow around an exterior of the gas turbine engine 100 (indicated by arrows C). The fan duct skin 110 and the exterior skin 114 terminate at the trailing edge 108.

[0013] Figure 1B is a detail side view of a trailing edge 108 of a portion of the fan nacelle 104. As shown in Figure 1B, the fan duct 112 is defined by the core nacelle 118 and the fan duct skin 110 of the fan nacelle 104. In the aspect illustrated in Figure 1B, the fan duct 112 defines a nozzle 122 that is aligned with the trailing edge 108 of the fan nacelle 104. In various other aspects, the nozzle 122 of the fan duct 112 maybe located upstream of the trailing edge 108 of the fan nacelle 104. Again, the nozzle 122 defines a narrowest flow area through the fan duct 112. A pressure P_FAN upstream of the nozzle 122 divided by a pressure P_AMB in the region 120 aft of the fan duct 112 and downstream of the nozzle 122 defines the fan pressure ratio.

[0014] In the aspect illustrated in Figure 1B not forming part of the claimed subject matter, the fan duct skin 110 includes a trailing edge portion 160 that converges toward the core nacelle 118 in the aft direction. The exterior skin 114 includes a trailing edge portion 150 that includes a surface 154 that is parallel with respect to the trailing edge portion 160 of the fan duct skin 110. As a result, the trailing edge 108 of the fan nacelle 104 includes a blunt end 152. The parallel surface 154 of the trailing edge portion 150 of the exterior skin 114 diverts airflow away from the region 120 aft of the fan duct 112 compared to a trailing edge portion 150 that converges with the fan duct skin 110 all the way to the trailing edge 108. By diverting the airflow away from the region 120 aft of the fan duct 112, the pressure P_AMB in the region 120 and downstream of the nozzle 122 is reduced. Further, the resulting blunt end 152 creates a low-pressure region immediately aft of the blunt end 152 and outboard of the nozzle 122. This low pressure outboard of the nozzle 122 also decreases the pressure P_AMB in the region 120 aft of the fan duct 112.

[0015] Figure 2A is a detail side view of a trailing edge 108 of a fan nacelle 104 according to another aspect not forming part of the claimed subject matter. The fan nacelle 104 includes an exterior skin 114 with a trailing edge portion 208 that includes a surface 209 that diverges from a trailing edge portion 160 of the fan duct skin 110 in the aft direction. The exterior skin 114 includes an inflection point 206. At locations upstream of the inflection point 206, the exterior skin 114 converges toward the fan duct skin 110 in the aft direction. At locations downstream of the inflection point 206, the exterior skin 114 diverges from the fan duct skin 110 in the aft direction. Relative to the aspect shown in Figure 1B in which the trailing edge portions 150 and 160 of the exterior skin 114 and fan duct skin 110 are parallel with respect to each other, the diverging surface 209 of the trailing edge portion 208 of the exterior skin 114 diverts airflow away from the region 120 after the fan duct 112 to a greater degree, as indicated by arrow D. Furthermore, the diverging surface 209 of the trailing edge portion 208 of the exterior skin 114 results in a larger blunt end 212, resulting in a larger low-pressure region aft of the trailing edge 108. As a result of the larger low-pressure region and the greater diversion of airflow away from the region 120 aft of the fan duct 112, the diverging surface 209 results in a lower pressure P_AMB in the region 120 than the parallel trailing edge portions 150 and 160 illustrated in Figure 1B. However, the fan nacelle 104 with the diverging surface 209 illustrated in Figure 2A may result in increased drag relative to the fan nacelle 104 with the parallel trailing edge portions 150 and 160 illustrated in Figure 1B.

[0016] In the aspect illustrated in Figure 2A not forming part of the claimed subject matter, the diverging surface 209 of the trailing edge portion 208 of the exterior skin 114 diverges from the trailing edge portion 160 of the fan duct skin 110 in a linear manner. Figure 2B illustrates an aspect according to the claimed subject matter in which a trailing edge portion 208' includes a diverging surface 209' that diverges in a nonlinear manner. Specifically, the diverging surface 209' diverges at greater angles at greater distances from the inflection point 206 in the aft direction. The diverging surface 209' results in even greater diversion of airflow away from the region 120 aft of the fan duct 112 (indicated by arrow E) and an even larger low-pressure region 212' aft of the trailing edge 108. As a result of the even larger low-pressure region and the greater diversion of airflow, the diverging surface 209' may result in an even lower pressure P_AMB in the region 120 than the diverging surface 209 illustrated in Figure 2A. However, the fan nacelle 104 with the diverging surface 209' may result in even more increased drag relative to the fan nacelle 104 with the linear diverging surface 209 illustrated in Figure 2A or the fan nacelle 104 with the parallel trailing edge portions 150 and 160 illustrated in Figure 1B.

[0017] In the above-described exemplary fan nacelles 104 with parallel or diverging trailing edge portions 150, the pressure P_AMB in the region 120 aft of the fan duct 112 is decreased. Assuming the pressure P_FAN in the fan duct 112 remains the same, the resulting fan pressure ratio decreases, resulting in improved fan stall margin.

[0018] In the above-described exemplary aspects, the trailing edge 108 of the fan nacelle 104 is modified in a manner that reduces pressure in the region 120 behind the fan duct 112 but also increases aerodynamic drag of the fan nacelle 104. Figures 3A and 3B and Figures 4A and 4B illustrate aspects not forming part of the claimed subject matter in which the fan nacelle 104 includes actuatable and inflatable portions, respectively, that decrease pressure in the region 120 aft of the fan duct 112 during phases of flight when the fan pressure ratio is relatively low or below a fan pressure ratio threshold but do not decrease pressure in the region 120 (and therefore do not increase drag) when the fan pressure ratio is not low or is below a fan pressure ratio threshold.

[0019] Figures 3A and 3B are detail side views of a trailing edge 108 of a fan nacelle 104 that includes a flap 306 in the trailing edge portion 308 of the exterior skin 114. The flap 306 is illustrated in a stowed position in Figure 3A and the flap 306' is illustrated in a fully deployed position in Figure 3B. In the stowed position, the flap 306 is flush with portions of the exterior skin 114 upstream of the trailing edge portion 308. As shown in Figure 3A, air flowing over the exterior skin 114 is directed toward the region 120 aft of the fan duct 112 when the flap 306 is in the stowed position, as indicated by arrow F. When deployed, the flap 306' in the trailing edge portion 308' diverges from the trailing edge portion 160 of the fan duct skin 110. Further, the air flowing over the exterior skin 114 is diverted away from the region 120 aft of the fan duct 112, as indicated by arrow F'. Additionally, a low-pressure region is generated aft of the deployed flap 306', which also reduces the pressure P_AMB in the region 120 aft of the fan duct 112.

[0020] The fan nacelle 104 generally has a cylindrical shape or a nearly cylindrical shape. Consequently, the flap 306 may comprise a plurality of flap segments arranged about a circumference of the trailing edge portion 308 of the fan nacelle. In certain applications, the flap segments may be spaced at particular intervals around the fan nacelle 104.

[0021] The flap 306 (or flap segments) is typically attached to the fan nacelle 104 via hinges or the like. The flap 306 (or flap segments) may be actuated between the stowed position and the deployed position via an actuator, such as a hydraulic piston, a pneumatic piston, a solenoid, an electric motor or a geometry-changing material, such as electroactive polymers. As noted above, the flap 306 may be manipulated between at least two positions - a fully stowed position as shown in Figure 3A and a fully deployed position as shown in Figure 3B. In such aspects, the flap 306 may be moved to the deployed position during phases of flight and/or conditions during which flight testing has demonstrated the fan pressure ratio to be low. For example, on certain aircraft, the fan pressure ratio is lowest during takeoff and climb phases of flight, and the flaps 306 may be moved to the deployed position during these phases of flight. In other aspects, the flap 306 may include many positions between a fully stowed position and a fully deployed position. In such aspects, the flap 306 may be moved to a particular position that achieves a predetermined minimum fan pressure ratio. In such aspects, the gas turbine engine 100 and/or an aircraft using the gas turbine engine 100 may be equipped with pressure sensors that enable calculation of the fan pressure ratio. Referring again to Figure 1A, the gas turbine engine 100 includes a first pressure sensor 180 arranged in the fan duct 112 at a location upstream of the fan nozzle 122 that is operable to measure the pressure in the fan duct P_FAN. The gas turbine engine 100 also includes a second pressure sensor 182 arranged along an exterior of the fan nacelle 104 that measures ambient pressure P_AMB. A controller 184 for the gas turbine engine 100 is operable to receive data from the first pressure sensor 180 and the second pressure sensor 182 to calculate the fan pressure ratio. In the event the calculated fan pressure ratio drops below a predetermined threshold value, the controller 184 actuates the flap 306 (or flap segments) to lower the pressure P_AMB in the region 120 aft of the fan duct 112 to increase the fan pressure ratio. In one aspect, the controller 184 can move the flap 306 between the fully stowed position shown in Figure 3A and the fully deployed position shown in Figure 3B to increase the fan pressure ratio. In another aspect, the controller 184 can move the flap 306 from the fully stowed position shown in Figure 3A to a partially-deployed position to achieve a fan pressure ratio equal to or higher than the predetermined threshold value.

[0022] Figures 4A and 4B are detail side views of a trailing edge 108 of a fan nacelle 104 that includes an inflatable boot 406 in the trailing edge portion 408 of the exterior skin 114. The inflatable boot 406 is illustrated in an uninflated state in Figure 4A and the inflatable boot 406' is illustrated in a fully inflated position in Figure 4B. When this inflatable boot 406 is uninflated, the inflatable boot 406 is flush with portions of the exterior skin 114 upstream of the trailing edge portion 408. As shown in Figure 4A, air flowing over the exterior skin 114 is directed toward the region 120 aft of the fan duct 112 when the inflatable boot 406 is in the uninflated state, as indicated by arrow G. When the inflatable boot 406' is inflated, a portion of the inflatable boot 406' in the trailing edge portion 408' diverges from the trailing edge portion 160 of the fan duct skin 110. When the inflatable boot 406' is inflated, the air flowing over the exterior skin 114 is diverted away from the region 120 aft of the fan duct 112, as indicated by arrow G'.

[0023] The inflatable boot 406 is attached in an airtight manner to a perimeter of the exterior skin 114 at a first location 420 and at a second location 422, creating an inflatable plenum 416 there-between. The inflatable plenum 416 communicates with a pressurized air source 412 (e.g., pressurized air from a bleed of the gas turbine engine 100) via one or more channels 414 in the fan nacelle 104. The inflatable boot 406 may be made from a rubber material or other flexible material. The inflatable boot 408 may be continuous about a perimeter of the exterior skin 114 of the fan nacelle 104. Alternatively, the inflatable boot 408 may comprise a plurality of inflatable boot segments arranged around the perimeter of the exterior skin 114 of the fan nacelle 104.

[0024] The inflatable boot 406 may include multiple inflated states to provide for varying degrees of air diversion and pressure drop in the region 120 aft of the fan duct 112. The controller 184 could measure the air pressure in the fan duct 112 using the first pressure sensor 180 and ambient air pressure using the second pressure sensor 182, and could command inflation of the inflatable boot 408 upon the fan pressure ratio dropping below the predetermined threshold value. Alternatively, the inflatable boot could be fully inflated whenever the fan pressure ratio drops below the predetermined threshold value or whenever the aircraft on which the gas turbine engine 100 is operating experiences phases of flight and/or flight conditions known to result in fan pressure ratios close to, at, or below the predetermined threshold value.

[0025] Figure 5 is a top view of an aircraft 500 that includes a fuselage 502, wings 504, and gas turbine engines 100 according to at least one aspect mounted under the wings. In various other aspects, the gas turbine engines 100 could be mounted to the fuselage 502 behind the wings 504. The aircraft 500 includes an ambient pressure sensor 506, such as a static pressure sensor that is also used by avionics on board the aircraft. The ambient pressure sensor 506 could be used in lieu of or in combination with the second pressure sensor 182, discussed above with reference to Figure 1, to calculate the fan pressure ratio.

[0026] The aircraft 500 also includes avionics 508. The avionics 508 are operable to detect flight conditions of the aircraft, including phases of flight for the aircraft. For example, the avionics 508 on modern aircraft include flight management computers that are programmed with waypoints, altitudes, and speeds for a particular flight. The avionics 508 also typically include a global positioning system (GPS) that allows for a precise determination of a location of the aircraft 500. Using this information (and possibly additional information from other avionics systems), the avionics 508 recognize a phase of flight for the aircraft 500. For example, the avionics 508 may recognize that the aircraft 500 is executing a takeoff roll along the runway, is climbing after takeoff, is cruising, is descending, or is landing. In at least one aspect, the avionics 508 can transmit a phase of flight status to the controller 184 such that the controller 184 only actuates the flap 306 to a deployed position or inflates the inflatable boot 406 in the event the fan pressure ratio drops below a predetermined threshold value and the phase of flight status indicates a phase of flight at which fan stall is more likely, such as a takeoff phase of flight or a climb phase of flight. In at least one aspect, the controller 184 may be omitted from the gas turbine engine 100, and the avionics 508 can control actuation of the flap 306 or inflation of the inflatable boot 406.

[0027] In the aspects described above, an exterior skin of a fan nacelle diverts air flowing over an exterior of the fan nacelle away from a region aft of a fan duct. Diverting the air away from the region reduces the air pressure behind the fan duct, thereby increasing a fan pressure ratio. The increased fan pressure ratio improves airflow through the fan duct and reduces the likelihood of a fan stall.

[0028] The descriptions of the various aspects have been presented for purposes of illustration, but are not intended to be exhaustive or limited to the aspects disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art. The terminology used herein was chosen to best explain the principles of the aspects, the practical application or technical improvement over technologies found in the marketplace, or to enable others of ordinary skill in the art to understand the aspects disclosed herein.

[0029] While the foregoing is directed to aspects, other and further aspects described herein may be devised without departing from the basic scope thereof, and the scope thereof is determined by the claims that follow.

Claims

1. A gas turbine engine nacelle assembly (190), comprising:

a core nacelle (118) of a gas turbine engine (100); and

a fan nacelle (104) disposed around the core nacelle (118), wherein the fan nacelle (104) includes a leading edge (106) and a trailing edge (108), wherein the fan nacelle (104) includes a fan duct skin (110) arranged between the leading edge (106) and the trailing edge (108), wherein the core nacelle (118) and the fan duct skin (110) define a fan duct (112) there-between, the fan duct (112) including a fan nozzle (122);

wherein the fan nacelle (104) includes an exterior skin (114) arranged between the leading edge (106) and the trailing edge (108) to direct airflow around an exterior of the gas turbine engine (100), wherein the fan duct skin (110) and the exterior skin (114) terminate at the trailing edge (108), wherein a trailing edge portion (208') of the exterior skin (114) is diverging from a trailing edge portion (160) of the fan duct skin (110) in an aft direction, and wherein the trailing edge portion (208') of the exterior skin (114) diverges at increasing angles from an inflection point (206) toward the trailing edge (108).

2. The gas turbine engine nacelle assembly (190) of claim 1, wherein the trailing edge portion (208') of the exterior skin (114) diverges in a nonlinear manner.

3. The gas turbine engine nacelle assembly (190) of claim 1 or 2, wherein the fan nozzle (122) is aligned with the trailing edge (108) of the fan nacelle (104).

4. The gas turbine engine nacelle assembly (190) of any one of claims 1 to 3, wherein the fan nozzle (122) defines a narrowest flow area through the fan duct (112).

5. The gas turbine engine nacelle assembly (190) of any one of claims 1 to 4, wherein at locations upstream of the inflection point (206), the exterior skin (114) converges toward the fan duct skin (110) in the aft direction.

6. The gas turbine engine nacelle assembly (190) of any one of claims 1 to 5, wherein at locations downstream of the inflection point (206), the exterior skin (114) diverges from the fan duct skin (110) in the aft direction.

7. The gas turbine engine nacelle assembly (190) of any one of claims 1 to 6, wherein the trailing edge portion (208') of the exterior skin (114) comprises a diverging surface (209') that diverts the airflow away from a region (120) aft of the fan duct (112).

8. The gas turbine engine nacelle assembly (190) of claim 7, wherein the diverging surface (209') diverges at greater angles at greater distances from the inflection point (206) in the aft direction.

9. An aircraft (500) comprising:
a fuselage (502), a wing (504), and a gas turbine engine (100) arranged relative to the fuselage (502) and the wing (504), wherein the gas turbine engine (100) comprises a nacelle assembly (190) according to claim 1.

10. The aircraft (500) of claim 9, wherein the trailing edge portion (208') of the exterior skin (114) diverges in a nonlinear manner.

11. The aircraft (500) of claim 9 or 10, wherein the fan nozzle (122) is aligned with the trailing edge (108) of the fan nacelle (104).

12. The aircraft (500) of any one of claims 9 to 11, wherein the fan nozzle (122) defines a narrowest flow area through the fan duct (112).

13. The aircraft (500) of any one of claims 9 to 12, wherein at locations upstream of the inflection point (206), the exterior skin (114) converges toward the fan duct skin (110) in the aft direction.

14. The aircraft (500) of any one of claims 9 to 13, wherein at locations downstream of the inflection point (206), the exterior skin (114) diverges from the fan duct skin (110) in the aft direction.

15. The aircraft (500) of any one of claims 9 to 14, wherein the trailing edge portion (208') of the exterior skin (114) comprises a diverging surface (209') that diverts the airflow away from a region (120) aft of the fan duct (112).

Ansprüche

1. Gasturbinentriebwerk-Gondelanordnung (190), die aufweist:

eine Kerngondel (118) eines Gasturbinentriebwerks (100); und

eine Fangondel (104), die um die Kerngondel (118) angeordnet ist, wobei die Fangondel (104) eine Vorderkante (106) und eine Hinterkante (108) aufweist, wobei die Fangondel (104) eine Fankanalhülle (110) aufweist, die zwischen der Vorderkante (106) und der Hinterkante (108) angeordnet ist, wobei die Kerngondel (118) und die Fankanalhülle (110) einen Fankanal (112) dazwischen bilden, wobei der Fankanal (112) eine Fandüse (122) aufweist;

wobei die Fangondel (104) eine Außenhülle (114) aufweist, die zwischen der Vorderkante (106) und der Hinterkante (108) angeordnet ist, um einen Luftstrom um eine Außenseite des Gasturbinentriebwerks (100) zu leiten, wobei die Fankanalhülle (110) und die Außenhülle (114) an der Hinterkante (108) enden, wobei ein Hinterkantenabschnitt (208') der Außenhülle (114) von einem Hinterkantenabschnitt (160) der Fankanalhülle (110) nach hinten divergiert und wobei der Hinterkantenabschnitt (208') der Außenhülle (114) in zunehmenden Winkeln von einem Wendepunkt (206) zur Hinterkante (108) divergiert.

2. Gasturbinentriebwerk-Gondelanordnung (190) nach Anspruch 1, wobei der Hinterkantenabschnitt (208') der Außenhülle (114) nichtlinear divergiert.

3. Gasturbinentriebwerk-Gondelanordnung (190) nach Anspruch 1 oder 2, wobei die Fandüse (122) zur Hinterkante (108) der Fangondel (104) ausgerichtet ist.

4. Gasturbinentriebwerk-Gondelanordnung (190) nach einem der Ansprüche 1 bis 3, wobei die Fandüse (122) eine schmalste Strömungsfläche durch den Fankanal (112) definiert.

5. Gasturbinentriebwerk-Gondelanordnung (190) nach einem der Ansprüche 1 bis 4, wobei an Stellen stromaufwärts vom Wendepunkt (206) die Außenhülle (114) zur Fankanalhülle (110) nach hinten konvergiert.

6. Gasturbinentriebwerk-Gondelanordnung (190) nach einem der Ansprüche 1 bis 5, wobei an Stellen stromabwärts vom Wendepunkt (206) die Außenhülle (114) von der Fankanalhülle (110) nach hinten divergiert.

7. Gasturbinentriebwerk-Gondelanordnung (190) nach einem der Ansprüche 1 bis 6, wobei der Hinterkantenabschnitt (208`) der Außenhülle (114) eine divergierende Oberfläche (209') aufweist, die den Luftstrom weg von einem Bereich (120) hinter dem Fankanal (112) ablenkt.

8. Gasturbinentriebwerk-Gondelanordnung (190) nach Anspruch 7, wobei die divergierende Oberfläche (209') in größeren Winkeln in größeren Abständen vom Wendepunkt (206) nach hinten divergiert.

9. Flugzeug (500), das aufweist:
einen Rumpf (502), eine Tragfläche (504) und ein Gasturbinentriebwerk (100), das relativ zum Rumpf (502) und zur Tragfläche (504) angeordnet ist, wobei das Gasturbinentriebwerk (100) eine Gondelanordnung (190) nach Anspruch 1 aufweist.

10. Flugzeug (500) nach Anspruch 9, wobei der Hinterkantenabschnitt (208') der Außenhülle (114) nichtlinear divergiert.

11. Flugzeug (500) nach Anspruch 9 oder 10, wobei die Fandüse (122) zur Hinterkante (108) der Fangondel (104) ausgerichtet ist.

12. Flugzeug (500) nach einem der Ansprüche 9 bis 11, wobei die Fandüse (122) eine schmalste Strömungsfläche durch den Fankanal (112) definiert.

13. Flugzeug (500) nach einem der Ansprüche 9 bis 12, wobei an Stellen stromaufwärts vom Wendepunkt (206) die Außenhülle (114) zur Fankanalhülle (110) nach hinten konvergiert.

14. Flugzeug (500) nach einem der Ansprüche 9 bis 13, wobei an Stellen stromabwärts vom Wendepunkt (206) die Außenhülle (114) von der Fankanalhülle (110) nach hinten divergiert.

15. Flugzeug (500) nach einem der Ansprüche 9 bis 14, wobei der Hinterkantenabschnitt (208') der Außenhülle (114) eine divergierende Oberfläche (209') aufweist, die den Luftstrom weg von einem Bereich (120) hinter dem Fankanal (112) ablenkt.

Revendications

1. Ensemble nacelle (190) de moteur à turbine à gaz, comprenant :

une nacelle centrale (118) d'un moteur à turbine à gaz (100) ; et

une nacelle de soufflante (104) disposée autour de la nacelle centrale (118), dans lequel la nacelle de soufflante (104) comporte un bord d'attaque (106) et un bord de fuite (108), dans lequel la nacelle de soufflante (104) comporte un revêtement de conduit de soufflante (110) agencé entre le bord d'attaque (106) et le bord de fuite (108), dans lequel la nacelle centrale (118) et le revêtement de conduit de soufflante (110) définissent un conduit de soufflante (112) entre eux, le conduit de soufflante (112) comportant une buse de soufflante (122) ;

dans lequel la nacelle de soufflante (104) comporte un revêtement extérieur (114) agencé entre le bord d'attaque (106) et le bord de fuite (108) pour diriger un écoulement d'air autour d'un extérieur du moteur à turbine à gaz (100), dans lequel le revêtement de conduit de soufflante (110) et le revêtement extérieur (114) se terminent au niveau du bord de fuite (108), dans lequel une partie de bord de fuite (208') du revêtement extérieur (114) diverge à partir d'une partie de bord de fuite (160) du revêtement de conduit de soufflante (110) dans une direction arrière, et dans lequel la partie de bord de fuite (208') du revêtement extérieur (114) diverge à des angles croissants à partir d'un point d'inflexion (206) vers le bord de fuite (108).

2. Ensemble nacelle (190) de moteur à turbine à gaz selon la revendication 1, dans lequel la partie de bord de fuite (208') du revêtement extérieur (114) diverge de manière non linéaire.

3. Ensemble nacelle (190) de moteur à turbine à gaz selon la revendication 1 ou 2, dans lequel la buse de soufflante (122) est alignée avec le bord de fuite (108) de la nacelle de soufflante (104).

4. Ensemble nacelle (190) de moteur à turbine à gaz selon l'une quelconque des revendications 1 à 3, dans lequel la buse de soufflante (122) définit une zone d'écoulement la plus étroite à travers le conduit de soufflante (112).

5. Ensemble nacelle (190) de moteur à turbine à gaz selon l'une quelconque des revendications 1 à 4, dans lequel à des emplacements en amont du point d'inflexion (206), le revêtement extérieur (114) converge vers le revêtement de conduit de soufflante (110) dans la direction arrière.

6. Ensemble nacelle (190) de moteur à turbine à gaz selon l'une quelconque des revendications 1 à 5, dans lequel à des emplacements en aval du point d'inflexion (206), le revêtement extérieur (114) diverge du revêtement de conduit de soufflante (110) dans la direction arrière.

7. Ensemble nacelle (190) de moteur à turbine à gaz selon l'une quelconque des revendications 1 à 6, dans lequel la partie de bord de fuite (208') du revêtement extérieur (114) comprend une surface divergente (209') qui détourne l'écoulement d'air d'une région (120) à l'arrière du conduit de soufflante (112).

8. Ensemble nacelle (190) de moteur à turbine à gaz selon la revendication 7, dans lequel la surface divergente (209') diverge à des angles plus grands à des distances plus grandes du point d'inflexion (206) dans la direction arrière.

9. Aéronef (500) comprenant :
un fuselage (502), une aile (504), et un moteur à turbine à gaz (100) agencé par rapport au fuselage (502) et à l'aile (504), dans lequel le moteur à turbine à gaz (100) comprend un ensemble nacelle (190) selon la revendication 1.

10. Aéronef (500) selon la revendication 9, dans lequel la partie de bord de fuite (208') du revêtement extérieur (114) diverge de manière non linéaire.

11. Aéronef (500) selon la revendication 9 ou 10, dans lequel la buse de soufflante (122) est alignée avec le bord de fuite (108) de la nacelle de soufflante (104).

12. Aéronef (500) selon l'une quelconque des revendications 9 à 11, dans lequel la buse de soufflante (122) définit une zone d'écoulement la plus étroite à travers le conduit de soufflante (112).

13. Aéronef (500) selon l'une quelconque des revendications 9 à 12, dans lequel à des emplacements en amont du point d'inflexion (206), le revêtement extérieur (114) converge vers le revêtement de conduit de soufflante (110) dans la direction arrière.

14. Aéronef (500) selon l'une quelconque des revendications 9 à 13, dans lequel à des emplacements en aval du point d'inflexion (206), le revêtement extérieur (114) diverge du revêtement de conduit de soufflante (110) dans la direction arrière.

15. Aéronef (500) selon l'une quelconque des revendications 9 à 14, dans lequel la partie de bord de fuite (208') du revêtement extérieur (114) comprend une surface divergente (209') qui détourne l'écoulement d'air d'une région (120) à l'arrière du conduit de soufflante (112).

Drawing