Forward-swept wing aircraft with boundary layer inhalation and distributed electric propulsion system

Abstract

A forward-swept wing aircraft with a boundary layer inhalation and a distributed electric propulsion system. An aircraft is disclosed, comprising: a fuselage extending along a longitudinal axis; a forward-swept wing extending from the fuselage; at least one horizontal stabilizer secured to the fuselage; and a distributed electric propulsion system operatively connected to an electric power source. The distributed electric propulsion system has an air inlet located rearward of the intersection between the forward-swept wing and the fuselage and opening into a boundary layer region on a surface of the fuselage.

Description

Technical Field

[0001] The present application relates generally to aircraft and, more particularly, to improved high performance aircraft. Background Art

[0002] Continuous efforts have been made to improve the design of civil and commercial aircraft in order to achieve minor improvements in range, fuel consumption, performance, etc. However, the overall conventional design of these civil and commercial aircraft has not changed significantly over the past few decades. As a result, aircraft manufacturers have begun to focus on developing new aircraft designs that can provide better aerodynamic performance and reduce fuel consumption. In particular, new aircraft designs are sought that can provide improved high-speed and low-speed performance while increasing range (and / or reducing fuel consumption). Summary of the Invention

[0003] Thus, according to a first aspect, there is provided an aircraft comprising: a fuselage extending along a longitudinal axis; a forward-swept wing extending from the fuselage; at least one horizontal stabilizer secured to the fuselage; and a distributed electric propulsion system operatively connected to an electrical power source, the distributed electric propulsion system having an air inlet located rearward of an intersection between the forward-swept wing and the fuselage and opening into a boundary layer region on a surface of the fuselage.

[0004] The aircraft of the first aspect defined above may further comprise, in whole or in part and in any combination, one or more of the following additional elements and / or definitions of existing elements:

[0005] at least one combustion engine drivingly engaged with at least one generator, the source of electrical power being said at least one generator;

[0006] said at least one combustion engine being a gas turbine engine;

[0007] the fuselage defining an air inlet fluidly connected to an inlet of the gas turbine engine, the air inlet opening into the boundary layer region;

[0008] The air inlet extends circumferentially around the fuselage;

[0009] The sweep angle of the forward-swept wing extends from the wing to a lateral axis perpendicular to the longitudinal axis, with the sweep angle ranging from 10 to 45 degrees;

[0010] The dihedral angle of the forward-swept wing extends from a plane containing both the longitudinal axis and a lateral axis perpendicular to the longitudinal axis to a mid-plane between the pressure and suction sides of the wing, the dihedral angle being in the range of -3 to 5 degrees;

[0011] A cross section of the fuselage taken on a plane normal to the longitudinal axis is elliptical;

[0012] The air inlet of the distributed electric propulsion system has a substantially rectangular shape;

[0013] The distributed electric propulsion system includes multiple electric fans placed side by side;

[0014] The plurality of electric fans includes at least four electric fans.

[0015] The at least one combustion engine and the at least one generator include two combustion engines and two generators, each generator being drivingly engaged with a respective one of the two combustion engines; and

[0016] The at least one horizontal stabilizer is a canard-type horizontal stabilizer located in front of the forward-swept wing.

[0017] According to a second aspect, an aircraft is also provided, comprising: a fuselage extending along a longitudinal axis, the fuselage defining a payload section and an engine section; a forward-swept wing extending from the fuselage, the forward-swept wing having a wing root behind its wingtip; at least one horizontal stabilizer fixed to the fuselage; vertical and horizontal stabilizer assemblies fixed to the fuselage and located at its rear end; and a plurality of electric fans arranged side by side and operatively connected to an electric power source, the inlets of the plurality of electric fans being fluidly connected to an air inlet located behind the forward-swept wing and receiving air from a boundary layer region of the aircraft.

[0018] The aircraft of the second aspect defined above may also include, in whole or in part and in any combination, one or more of the following additional elements and / or definitions of existing elements:

[0019] at least one combustion engine located in the engine section of the fuselage and drivingly engaged with at least one generator, the source of electrical power being the at least one generator;

[0020] The at least one combustion engine is at least one gas turbine engine;

[0021] the fuselage defining an air inlet fluidly connected to an inlet of the gas turbine engine, the air inlet opening into the boundary layer region;

[0022] The air inlet extends circumferentially around the fuselage;

[0023] The sweep angle of the forward-swept wing extends from the wing to a lateral axis perpendicular to the longitudinal axis, with the sweep angle ranging from 10 to 45 degrees;

[0024] The dihedral angle of the forward-swept wing extends from a plane containing both the longitudinal axis and a lateral axis perpendicular to the longitudinal axis to a mid-plane between the pressure side and the suction side of the wing, the dihedral angle being in the range of -3 to 5 degrees;

[0025] The air inlet of the distributed electric propulsion system has a substantially rectangular shape; and

[0026] The at least one horizontal stabilizer is a canard-type horizontal stabilizer located forward of a wing root of the forward-swept wing. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Reference is now made to the accompanying drawings, in which:

[0028] Figure 1 is a schematic perspective view of an aircraft according to one embodiment of the present disclosure;

[0029] Figure 2 It is taken along a plane perpendicular to the longitudinal axis of the aircraft Figure 1 a schematic cross-section of the aircraft;

[0030] Figure 3 yes Figure 1 a schematic cross-sectional view of a wingtip section and a wing root section of one of the wings of an aircraft; and

[0031] Figure 4 yes Figure 1 Another schematic perspective view of an aircraft illustrating airflow over the aircraft. DETAILED DESCRIPTION

[0032] With reference to the drawings, and more particularly with reference to Figure 1 An aircraft is shown generally at 10 and is generally described in this disclosure to illustrate some components for reference. The aircraft 10 has a fuselage 12 having a front end 12a at which a cockpit is located and a rear end 12b supporting an empennage or tail assembly 14. The fuselage 12 extends along a longitudinal axis L. The aircraft 10 also has wings 16 extending from the fuselage 12. A cross-section of the fuselage 16 taken on a plane normal to the longitudinal axis L may be elliptical.

[0033] The fuselage 12 defines a fuselage air inlet 12c. Figure 1 As shown, the fuselage air inlet 12c extends around at least a portion of the circumference of the fuselage 12. The fuselage air inlet 12c opens into a first boundary layer region B1 of the surface of the fuselage 12. The fuselage air inlet 12c may be annular and may be contained in a plane orthogonal to the longitudinal axis L. The fuselage air inlet 12c is oriented toward the front end 12a of the fuselage 12.

[0034] In the illustrated embodiment, the fuselage 12 defines a payload section 12d and an engine section 12e. The payload section 12d may be configured to receive, for example, passengers and / or cargo. The engine section 12e may be located rearward of the payload section 12d. In the illustrated embodiment, the engine section 12e is located between the payload section 12d and the tail assembly 14 of the aircraft 10. Figure 1 As shown, fuselage air inlet 12c is located between engine section 12d and payload section 12e. Fuselage air inlet 12c may be located forward of engine section 12c. Fuselage air inlet 12c is configured to allow air from first boundary layer region B1 of fuselage 12 to penetrate into engine section 12e. More information on this aspect is provided below.

[0035] The wing 16 has a leading edge 16a and a trailing edge 16b, the trailing edge 16b being behind and downstream of the leading edge 16a. The wing 16 has a pressure side 16c ( Figure 2 ) and suction side 16d( Figure 2 ), the suction side 16d is opposite the pressure side 16c and extends from the leading edge 16a to the trailing edge 16b. Each wing 16 is defined by a plurality of airfoil segments 16e distributed from its wing root 16f to the wing tip 16g. The chord of the airfoil segment 16e extends from the leading edge 16a to the trailing edge 16b and may decrease from the wing root 16f to the wing tip 16g. The taper ratio of the wing 16 is defined as the chord of the wing 16 at the wing tip 16g divided by the chord of the wing at the wing root 16f. In the illustrated embodiment, the taper ratio is in the range of 0.2 to 0.5. In a specific embodiment, the taper ratio is in the range of 0.29 to 0.31.

[0036] Aircraft 10 is known as a forward-swept wing aircraft. In this type of aircraft, a wing 16 extends from a wing root 16f located on fuselage 12 to a wingtip 16g located away from fuselage 12; wingtip 16g is forward of wing root 16f. In other words, wing 16 extends forward from its wing root 16f to its wingtip 16g toward the front end 12a of fuselage 12. This differs from conventional aircraft in which the wing extends toward the tail assembly. A sweep angle θ of wing 16 is defined between wing 16 and a lateral axis A of aircraft 10, which is perpendicular to longitudinal axis L. More specifically, sweep angle θ extends from lateral axis A to a quarter-chord line 16h of wing 16. Quarter-chord line 16h is defined by a line extending through a series of points, each located at one-quarter of the chord line from the leading edge 16a of each wing segment 16e. The chord line of each wing segment 16e extends from the leading edge 16a to the trailing edge 16b of wing 16. The sweep angle θ extends from the quarter chord line 16h to the lateral axis A of the aircraft 10. In the illustrated embodiment, the sweep angle θ is in the range of 10 to 45 degrees. In a particular embodiment, the sweep angle θ is in the range of 25 to 30 degrees.

[0037] Now refer to Figure 2, the wing 16 may define an angle of dihedral φ. Angle of dihedral φ corresponds to the angle from a plane P containing both the longitudinal axis L and the lateral axis A of the aircraft 10 to the wing 16. When the aircraft 10 is supported by the ground, plane P may be parallel to the ground. More specifically, angle of dihedral φ extends from plane P containing the longitudinal axis L and the lateral axis A to a mid-plane 16i of the wing 16 located halfway between its pressure side 16c and suction side 16d. In the depicted embodiment, angle of dihedral φ ranges from -3 to 5 degrees. In a specific embodiment, angle of dihedral φ ranges from 3 to 5 degrees. In the illustrated embodiment, the aircraft 10 has positive dihedral, meaning that the wing 16 points upward and extends or tilts away from the ground from its root to its tip. This is in contrast to negative dihedral, in which the wing tends to tilt toward the ground.

[0038] Now refer to Figure 3 , for one of the wings 16 having a practical relative orientation, airfoil sections of one wing are shown at the wing root 16f and the wing tip 16g. Each airfoil section has a chord line extending from the leading edge 16a of the wing 16 to the trailing edge 16b. For illustrative purposes, the chord lines of each of the wing root section 16f and the wing tip section 16g are shown in dashed lines and are extended beyond the leading edge 16a and the trailing edge 16b. As illustrated, the chord lines of the two airfoil sections are not parallel to each other. This means that the wing 16 has a twist angle γ defined between the chord lines of the airfoil sections at the wing root 16f and the wing tip 16g. In the illustrated embodiment, the twist angle γ ranges from 0 to 6 degrees. In a specific embodiment, the twist angle γ ranges from 3 to 5 degrees. In a specific embodiment, the wing has a twist angle of 0 degrees.

[0039] Return Reference Figure 1 The tail assembly 14 includes a vertical stabilizer 14a and a horizontal stabilizer 14b. More specifically, the tail assembly 14 includes two vertical stabilizers 14a positioned at opposite lateral ends of the tail assembly 14, and a horizontal stabilizer 14b extending from one of the vertical stabilizers 14a to the other. In the illustrated embodiment, the horizontal stabilizer 14b is secured to the vertical stabilizer 14a between its wing root and its wing tip.

[0040] Tail assembly 14 includes an upper wall 14c, a lower wall 14d opposite and spaced apart from upper wall 14c, and side walls 14e extending from upper wall 14c to lower wall 14d. In the illustrated embodiment, side walls 14e are defined by vertical stabilizer 14a. In other words, both upper wall 14c and lower wall 14d are connected to vertical stabilizer 14a.

[0041] The upper wall 14c, the lower wall 14d, and the side walls 14e surround the engine compartment 18, which has an inlet 18a and an outlet 18b. As shown, the inlet 18a faces the front end 12a of the fuselage 12, while the outlet 18b is oriented away from the front end 12a of the fuselage 12. The inlet 18a and the outlet 18b can be substantially rectangular in shape. In other words, the width of both the inlet 18a and the outlet 18b of the engine compartment 18 taken along the lateral axis A is greater than the width taken along the vertical axis V (which is perpendicular to both the longitudinal axis L and the lateral axis A). Figure 2 ) intercepted height.

[0042] In the depicted embodiment, the inlet 18a of the engine compartment 18 opens into a second boundary layer region B2 located on the surface of the fuselage 12 of the aircraft 10. The second boundary layer region B2 is located downstream of the first boundary layer region B1 relative to the air flow F circulating around the aircraft 10. In other words, the engine compartment 18 is fluidically connected to the second boundary layer region B2. Further details regarding this aspect are provided below. The inlet 18a of the engine compartment 18 is located behind the intersection between the forward-swept wing 16 and the fuselage 12.

[0043] Still refer to Figure 1 The aircraft 10 includes a front horizontal stabilizer 20, which is a canard-type stabilizer and is hereinafter referred to as a canard. The canard 20 is disposed adjacent the front end 12a of the fuselage 12. The canard 20 is located between the front end 12a of the fuselage 12 and the wing root 16f of the wing 16. In the illustrated embodiment, the canard 20 includes two wings 20a disposed on opposite sides of the fuselage 12. Each of the two wings 20a of the canard 20 may define a sweep angle α defined from a quarter chord line 20b of the wing 20a of the canard 20 to a lateral axis A of the aircraft 10. The canard 20 may or may not have a twist angle that is not equal to zero, an anhedral angle that is not equal to zero, and / or a taper ratio that is not equal to one.

[0044] Return Reference Figure 1 To propel the aircraft, the aircraft is equipped with a distributed power system 22 housed within the engine compartment 18 of the tail assembly 14. In the illustrated embodiment, the distributed power system 22 includes electric fans 22a. One or more electric fans 22a may be used. In the illustrated embodiment, four electric fans 22a are used and positioned side by side. Each electric fan 22a is operatively connected to an electric power source S. In a particular embodiment, the distributed power system 22 includes six fans, three on each side of the fuselage and three behind the wings. The number and diameter of the fans may be a function of the overall design parameters of the aircraft 10.

[0045] The distributed electric propulsion system 22 has an air inlet 22b that corresponds to the inlet 18a of the engine compartment 18. In other words, the air inlet 22b of the distributed electric propulsion system 22 opens into the second boundary layer region B2 of the aircraft 10. The inlet of each electric fan 22a is fluidically connected to the second boundary layer region B2 via the inlet 18a of the engine compartment 18. The outlet of each electric fan 22a is fluidically connected to the outlet 18b of the engine compartment 18.

[0046] Also refer to Figure 4 In the illustrated embodiment, and in operation, the distributed electric propulsion system 22 is configured to draw air from the second boundary layer region B2 to ingest at least a portion of the boundary layer developed over the surface of the fuselage 12. In a particular embodiment, the air in the second boundary layer region B2 comes from the boundary layer developed over the surface of the wing 16. Figure 4 As shown, a portion F1 of the airflow F over the wing 16 is deflected inwardly toward the fuselage 12 due to the forward-swept nature of the wing 16 and enters a second boundary layer region B2 on the fuselage. As shown, the forward sweep of the wing 16 helps increase the amount of air that circulates through the second boundary layer region B2, thereby helping to increase the mass flow of air entering the inlet 18a of the engine compartment 18. Therefore, compared to an aircraft with a conventional swept wing, this configuration (i.e., the forward-swept wing 16 and the air inlet 22b of the distributed electric propulsion system 22 located rearward of the intersection between the forward-swept wing 16 and the fuselage 12) can provide a greater mass flow of air directed into the inlet air inlet 22b.

[0047] In particular embodiments, absorbing at least a portion of the boundary layer reduces the drag exerted by the air flow F around the aircraft 10 and increases or improves the aerodynamic performance of the aircraft 10 compared to configurations not configured to absorb the boundary layer.

[0048] In the illustrated embodiment, the forward-swept wings 16 converge toward an air intake 22 b of a distributed electric propulsion system 22 and may deflect at least a portion of the flow circulating around the aircraft 10 toward the distributed electric propulsion system 22 .

[0049] In the depicted embodiment, the aircraft 10 includes at least one combustion engine 24, which may be located in the engine section 12 e of the fuselage 12. The at least one combustion engine 24 is drivingly engaged with at least one generator 26 for generating electrical power to power the electric fan 22 a of the distributed electric propulsion system 22. In the illustrated embodiment, the electrical power source S of the aircraft 10 corresponds to the at least one generator 26. Alternatively, the electrical power source S may be a battery.

[0050] In the illustrated embodiment, the at least one combustion engine 24 includes two combustion engines 24, which can be placed side by side in the engine section 12e of the fuselage 12. Each of the two gas turbine engines 24 is drivingly coupled to a respective one of two generators 26. The combustion engines 24 can be gas turbine engines. Any other suitable engine can be used without departing from the scope of the present disclosure.

[0051] Each of the two gas turbine engines 24 has an inlet fluidically connected to the fuselage air inlet 12c and the first boundary layer region B1. In other words, the inlet of the gas turbine engine 24 is in fluidic flow communication with the first boundary layer region B1 of the aircraft 10 via the fuselage air inlet 12c. In operation, the gas turbine engine 24 is configured to draw air from the first boundary layer region B1 to absorb at least a portion of the boundary layer that develops above the surface of the fuselage 12. In certain embodiments, absorbing at least a portion of the boundary layer reduces the drag force exerted by the air flow F surrounding the aircraft 10 and increases or improves the aerodynamic performance of the aircraft 10, compared to a configuration not configured to absorb the boundary layer.

[0052] In certain embodiments, the combination of boundary layer absorption by the air inlet of the distributed electric propulsion system 22 and the forward-swept wings 16 provides benefits in aircraft performance. These benefits can include, for example, improved wing performance at low speeds, which saves power. In certain embodiments, the wings, due to their inverse span gradient, allow their highest-loaded areas to experience boundary layer absorption. At low speeds, this can reduce adverse pressure gradients across the wing. At high speeds, this can have adverse effects when the velocity of the air entering the distributed electric propulsion system's air inlet is lower than the velocity of the airflow circulating around the aircraft, potentially leading to expanded flow ducts and, consequently, backpressure that can increase impact strength. In certain embodiments, the aircraft is designed so that the inlet is not an obstructed area at the trailing edge of the wing. Aircraft design parameters can all be functions of aircraft speed relative to the speed at the fan inlet. For example, with current engines, during takeoff, the cross-sectional area of ​​the flow duct absorbed by the engine decreases towards the engine. This is known as a manifold. Manifolds occur when the aircraft's speed is lower than the velocity of the airflow drawn into the engine at maximum power. During cruise, the splitter duct is present because the aircraft speed is higher than the speed of the airflow being drawn into the engine inlet.

[0053] The forward-swept wing 16 allows a load distribution from the wingtip to the wing root, where the maximum load occurs at the wing. This can allow for improved low-speed stall characteristics compared to a configuration with a swept wing by maintaining wingtip control. The load characteristics from the wingtip to the wing root of the forward-swept wing can allow for a higher level of boundary layer absorption compared to a configuration with a swept wing. This boundary layer absorption can allow for a reduction in propulsion power requirements compared to a configuration lacking boundary layer absorption. The boundary layer absorption at the end of the forward-swept wing can be designed to act as boundary layer control and can improve the load characteristics at the wing root section of the wing. The combination of the forward-swept wing and the boundary layer absorption can provide an aerodynamic concept that is more efficient than conventional configurations, and this aerodynamic concept can utilize the combined interaction of all three concepts (i.e., boundary layer absorption, distributed electric propulsion system, and forward-swept wing) and can result in significantly improved high-speed performance and low-speed characteristics compared to conventional aircraft lacking these concepts.

[0054] To operate the aircraft, the airflow around the forward-swept wings is deflected toward the ground to maintain the aircraft's lift. A spanwise component is imparted to the airflow, causing it to move toward the aircraft's fuselage. A portion of the airflow is drawn into the boundary layer region located on the fuselage's outer surface, downstream of the intersection between the fuselage and the forward-swept wings. This drawn-in portion of the airflow is accelerated by an electric fan.

[0055] The above description is intended to be exemplary only, and those skilled in the art will recognize that changes may be made to the described embodiments without departing from the scope of the invention disclosed. Other modifications falling within the scope of the invention will be apparent to those skilled in the art upon review of this disclosure, and such modifications are intended to fall within the scope of the appended claims.

Claims

1. An aircraft comprising: a fuselage extending along a longitudinal axis, the fuselage defining an air inlet; a forward-swept wing, the forward-swept wing extending from the fuselage from a wing root to a wing tip, the air inlet being located forward of the wing root of the forward-swept wing; at least one horizontal stabilizer secured to the fuselage; a distributed electric propulsion system operatively connected to a source of electric power and having an air inlet located aft of an intersection between the forward-swept wing and the fuselage and opening into a boundary layer region on a surface of the fuselage; and At least one combustion engine is drivingly engaged with at least one generator, the source of electrical power being the at least one generator, the air inlet being fluidly connected to an inlet of the at least one combustion engine, the air inlet opening into the boundary layer region.

2. The aircraft according to claim 1, wherein: The at least one combustion engine is at least one gas turbine engine.

3. The aircraft of claim 1, wherein: The air inlet extends circumferentially around the fuselage.

4. The aircraft of claim 1, wherein: The sweep angle of the forward-swept wing extends from the forward-swept wing to a lateral axis perpendicular to the longitudinal axis, the sweep angle being in the range of from 10 to 45 degrees.

5. The aircraft of claim 1, wherein: The dihedral angle of the forward-swept wing extends from a plane containing both the longitudinal axis and a lateral axis perpendicular to the longitudinal axis to a mid-plane between the pressure side and the suction side of the forward-swept wing, the dihedral angle being in the range from -3 to 5 degrees.

6. The aircraft of claim 1, wherein: A cross-section of the fuselage taken on a plane normal to the longitudinal axis is elliptical.

7. The aircraft of claim 1, wherein: The air inlet of the distributed electric propulsion system has a substantially rectangular shape.

8. The aircraft of claim 1, wherein: The distributed electric propulsion system includes a plurality of electric fans placed side by side.

9. The aircraft of claim 1, wherein: The at least one combustion engine and the at least one generator include two combustion engines and two generators, each of the generators being drivingly engaged with a respective one of the two combustion engines.

10. The aircraft of claim 1, wherein: The at least one horizontal stabilizer is a canard-type horizontal stabilizer located in front of the forward-swept wing.

11. An aircraft comprising: a fuselage extending along a longitudinal axis, the fuselage defining a payload section and an engine section, the fuselage defining an air inlet; a forward-swept wing extending from the fuselage, the forward-swept wing having a wing root behind a wingtip thereof, the air inlet being located forward of the wing root of the forward-swept wing; at least one horizontal stabilizer secured to the fuselage; vertical and horizontal stabilizer assemblies secured to the fuselage and located at the aft end of the fuselage; a plurality of electric fans positioned side by side and operatively connected to a power source, the plurality of electric fans having inlets fluidly connected to an air intake located behind the forward-swept wing and receiving air from a boundary layer region of the aircraft; and at least one combustion engine located in the engine section of the fuselage and drivingly engaged with at least one generator, the source of electrical power being the at least one generator, the air inlet being fluidly connected to an inlet of the at least one combustion engine, the air inlet opening into the boundary layer region.

12. The aircraft according to claim 11, wherein: The at least one combustion engine is at least one gas turbine engine.

13. The aircraft according to claim 12, wherein: The air inlet extends circumferentially around the fuselage.

14. The aircraft of claim 11, wherein: The sweep angle of the forward-swept wing extends from the forward-swept wing to a lateral axis perpendicular to the longitudinal axis, the sweep angle being in the range of from 10 to 45 degrees.

15. The aircraft of claim 11, wherein: The dihedral angle of the forward-swept wing extends from a plane containing both the longitudinal axis and a lateral axis perpendicular to the longitudinal axis to a mid-plane between the pressure side and the suction side of the forward-swept wing, the dihedral angle being in the range from -3 to 5 degrees.

16. The aircraft of claim 11, wherein: The air inlet has a substantially rectangular shape.

17. The aircraft of claim 11, wherein: The at least one horizontal stabilizer is a canard-type horizontal stabilizer located forward of the wing root of the forward-swept wing.

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