Aircraft Wing Systems

Abstract

A wing system for an aircraft is provided, the wing system having a stowed configuration, a pre-deployed configuration, and a deployed configuration. The wing system includes two wings, each having an airfoil shape and deployable to pivot about a respective pivot axis between a pre-deployed configuration and a deployed configuration. In the stowed configuration, the two wings are in a first substantially overlapping spatial relationship with each other and are retractable within an envelope having an envelope cross-sectional shape and a corresponding envelope cross-sectional area. In the pre-deployed configuration, the two wings are in a second substantially overlapping spatial relationship with each other and are deployable to the deployed configuration. In the deployed configuration, each of the wings is capable of generating aerodynamic lift in an airflow. Each airfoil shape of each wing is a slotted airfoil having a primary element, a secondary element, and a chord, the secondary element being pivotable relative to the primary element and spaced apart by a gap. Each airfoil shape has a respective maximum thickness and a respective maximum absolute thickness. In the stowed configuration, each secondary element of each airfoil on one wing is set at a different flap angle compared to each secondary element of each airfoil on the other wing.

Description

[Technical Field]

[0001] The present invention relates to wing systems for aircraft, and more particularly to high-lift airfoils and wings therefor.

[0002] Some types of aircraft, for example, some types of unmanned aerial vehicles (UAVs), are deployed as subsonic aircraft. In some cases, the aircraft is launched with its lift-producing wings folded close to the fuselage, and the wings are subsequently deployed to enable aerodynamic flight.

[0003] As a non-limiting example, US 2017 / 369150 describes, in one embodiment, a wing for an unmanned aerial vehicle. The unmanned aerial vehicle includes a wing first body having a first end proximate to an airframe. A second end is opposite the first end. A first connector is at the first end of the wing first body. A connector rotatably connects the wing to the airframe. A second connector is at the second end of the airframe. The wing second body is rotatably connected to the first body via the second connector.

[0004] Also, as a non-limiting example, CN107380402 discloses an unmanned aerial vehicle with folding wings, which includes a fuselage, wings, a horizontal stabilizer, a vertical stabilizer, a folding wing release mechanism, and a fixing base, in which positioning holes are formed in the wings, horizontal stabilizer, and vertical stabilizer, and positioning clamp posts are disposed on the fixing base, and the positioning clamp posts are aligned with the positioning holes. Summary of the Invention

[0005] According to a first aspect of the presently disclosed subject matter, there is provided a wing system for an aircraft having a stowed configuration, a pre-deployed configuration, and a deployed configuration, the wing system comprising: two wings, each having an airfoil shape and deployable to pivot about a respective pivot axis between the pre-deployed configuration and the deployed configuration, wherein: In the stowed configuration, the two wings are in a first substantially overlapping spatial relationship with respect to one another and are retractable or arranged to be retractable within an envelope having an envelope cross-sectional shape and a corresponding envelope cross-section; In the pre-deployment configuration, the two wings are in a second substantially overlapping spatial relationship relative to one another and are deployable to the deployed configuration; and In the deployed configuration, each of the wings is capable of generating aerodynamic lift in an airflow; each airfoil shape is a slotted airfoil having a primary element, a secondary element, and a chord, said secondary element being pivotable relative to said primary element and spaced apart by a gap, each said airfoil shape having a respective maximum thickness and a respective maximum absolute thickness; Here, in the stowed configuration, the secondary elements of each of the airfoil shapes of one of the wings are set at a different flap angle compared to the secondary elements of each of the airfoil shapes of the other of the wings. For example, each of the wings has a span and an airfoil shape in at least one respective cross-section perpendicular to the respective span.

[0006] Additionally or alternatively, for example, in a stowed configuration, the secondary elements of each of the airfoil shapes of one of the wings are set at a different flap angle compared to the secondary elements of each of the airfoil shapes of the other of the wings.

[0007] Additionally or alternatively, for example, the two wings may include a first wing having the first airfoil shape, the first airfoil shape including a first leading edge and a first trailing edge, and a second wing having the first airfoil shape, the second airfoil shape including a second leading edge and a second trailing edge, e.g., in the pre-deployment configuration, the first leading edge is substantially overlapped with the second trailing edge.

[0008] Additionally or alternatively, for example, the ratio of the respective maximum absolute thickness to the chord of each of the wings is such that the wings can be deployed from the pre-deployed configuration to the deployed configuration without interfering with each other.

[0009] Additionally or alternatively, for example, each secondary element of each airfoil shape of each wing has a hinge axis to allow pivoting of the respective secondary element relative to the respective primary element. For example, the hinge axis faces the pressure surface of the respective secondary element. Additionally or alternatively, for example, for each airfoil shape, each hinge axis is spaced apart from the pressure surface of the respective secondary element. For example, the spacing is less than 50% of the maximum thickness of the secondary element. For example, the spacing may be 40%, 30%, 20%, 10%, 5%, 1%, or 0% of the maximum thickness of the secondary element.

[0010] Additionally or alternatively, for example, in the pre-deployment configuration, a respective secondary element of each airfoil shape of each wing is set to a respective first negative flap angle.

[0011] Additionally or alternatively, for example, in the pre-deployment configuration, each secondary element of each airfoil shape of each wing is set to the same flap angle.

[0012] Additionally or alternatively, for example, in the pre-deployment configuration, each secondary element of each airfoil shape of each wing is set to a respective first flap angle, so that no portion of each second flap intersects an imaginary line tangent to the leading edge of the respective airfoil shape and intersecting the respective hinge axis.

[0013] Additionally or alternatively, for example, each of the airfoil shapes may be arranged at its respective minimum absolute thickness and maximum chord, such that in the stowed configuration, the airfoil shapes of each of the two wings in the first substantially overlapping spatial relationship can be stowed within the envelope cross-sectional shape while simultaneously maximizing the occupancy of the envelope cross-section.

[0014] Additionally or alternatively, for example, the wing system is configured to be attached to the underside of an aircraft fuselage. For example, in the stowed configuration, the secondary elements of the airfoil shapes of the wing are set to negative flap angles. Additionally or alternatively, for example, in the stowed configuration, the secondary elements of the airfoil shapes of the uppermost wing are set to negative flap angles in a range of −25° to −35°, and the secondary elements of the airfoil shapes of the lowermost wing are set to negative flap angles in a range of −55° to −65°. Additionally or alternatively, for example, in the stowed configuration, the secondary elements of the airfoil shapes of the wing are set to second flap angles, such that the trailing edges and hinge axes of the secondary elements are on opposite sides of the chord lines of the airfoil shapes.

[0015] Additionally or alternatively, for example, the wing system is configured to be mounted above an aircraft fuselage. For example, in the stowed configuration, each secondary element of the airfoil shape of each wing is set to a respective positive flap angle. Additionally or alternatively, for example, in the stowed configuration, each secondary element of the airfoil shape of the uppermost wing is set to a respective positive flap angle within a range of +15° to +25°, and each secondary element of the airfoil shape of the lowermost wing is set to a respective positive flap angle within a range of +1° to +10°.

[0016] Additionally or alternatively, for example, each of the airfoil shapes has a first maximum spacing of its respective pressure surface from the chord line and a second maximum spacing of its respective hinge axis from the chord line, where the first maximum spacing and the second maximum spacing are similar to, for example, within 10% of each other.

[0017] Additionally or alternatively, for example, for each of the airfoil shapes, the respective maximum thickness-to-chord ratios and the respective maximum absolute thickness-to-chord ratios are within 10% of each other, such as 1% of the chord.

[0018] Additionally or alternatively, for example, for each of the airfoil shapes, the respective maximum thickness to chord ratio is in the range of 0.10 to 0.14.

[0019] Additionally or alternatively, for example, for each of the airfoil shapes, the respective maximum absolute thickness to chord ratio is in the range of 0.10 to 0.15.

[0020] Additionally or alternatively, for example, each of the airfoil shapes has its respective maximum thickness at a location less than 25% of the chord from the leading edge of the airfoil shape.

[0021] Additionally or alternatively, for example, each of the airfoil shapes has its respective maximum thickness at about 20% of the chord from the leading edge of the airfoil shape.

[0022] Additionally or alternatively, for example, each of the airfoil shapes has a camber shape with a maximum camber of less than or equal to about 6% of the chord.

[0023] Additionally or alternatively, for example, each of the airfoil shapes includes at least one of the following: each said gap varies between about 2.3% and about 1.2% of said chord as each said flap angle increases from 0° to 30°; Each of the gaps varies between about 2.3% and about 2.5% of the chord as each of the flap angles decreases from 0° to −20°.

[0024] Additionally or alternatively, for example, for each of the airfoil shapes, the respective leading edge radius is about 2% of the chord.

[0025] Additionally or alternatively, for example, for each airfoil shape, each primary element has a primary element trailing edge and each secondary element has a secondary element leading edge; the primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord; the primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord, and each longitudinal overlap varies between about 5% and about 3% of the chord as each flap angle increases from 0° to 30°; and the primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord, and for each airfoil shape, the respective longitudinal overlap varies between about 5% and about 5.5% of the chord as the respective flap angle decreases from 0° to −20°.

[0026] Additionally or alternatively, for example, each of the airfoil shapes may be configured as a high-lift airfoil shape. For example, the airfoil shapes may be configured to have a maximum lift coefficient of at least 2.5 at a flap angle of zero, or the airfoil shapes may be configured to have a maximum lift coefficient of greater than 2.3 at a flap angle of zero. Additionally or alternatively, for example, the maximum lift coefficient corresponds to an angle of attack of about 15° to about 16°.

[0027] Additionally or alternatively, for example, the airfoil shape may be arranged to have a lift coefficient of 1.0, which corresponds to an angle of attack of approximately 0°.

[0028] Additionally or alternatively, for example, the airfoil shapes of the two wings have a total cross-sectional area corresponding to an occupancy rate of 40% to 30% of the envelope cross-sectional area.

[0029] Additionally or alternatively, for example, the envelope cross-sectional shape may be in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion. For example, for each of the airfoil shapes, the respective first ratio of chord to diameter is in the range of 70% to 87%. For example, for each of the airfoil shapes, the respective first ratio is in the range of 75% to 85%. Additionally or alternatively, for example, for each of the airfoil shapes, the respective second ratio of maximum thickness to diameter is in the range of 7% to 12%. Additionally or alternatively, for example, for each of the airfoil shapes, the respective second ratio of maximum absolute thickness to diameter is in the range of 8% to 14%.

[0030] Additionally or alternatively, for example, the wings may be arranged in a non-tapered manner, where each wing has a uniform airfoil shape along its respective span, or, for example, the wings may be arranged in a tapered manner, where each wing has a uniform airfoil shape along its respective span.

[0031] Additionally or alternatively, for example, the wing system may be arranged to selectively allow the wings to transition from the stowed configuration to the pre-deployed configuration under predetermined conditions.

[0032] Additionally or alternatively, for example, the wing system may be arranged to maintain the deflection angle of each airfoil shape of each wing negative during deployment from the pre-deployed configuration to the deployed configuration.

[0033] According to a first aspect of the presently disclosed subject matter, there is provided a wing system for an aircraft having at least a stowed configuration, a pre-deployed configuration, and a deployed configuration, the wing system comprising: two wings, each deployable to pivot about a respective pivot axis between the pre-deployed configuration and the deployed configuration, wherein: In the stowed configuration, the two wings are in a first substantially overlapping spatial relationship with respect to one another and are retractable within an envelope having an envelope cross-sectional shape and a corresponding envelope cross-section; In the pre-deployment configuration, the two wings are in a second substantially overlapping spatial relationship relative to one another and are deployable to the deployed configuration; and In the deployed configuration, each of the wings is capable of generating aerodynamic lift in an airflow; Each of the wings has a wing span and an airfoil shape in at least one cross section perpendicular to the respective wing span; each said airfoil shape is a slotted airfoil having a primary element, a secondary element, and a chord, said secondary element being pivotable relative to said primary element and spaced apart by a gap, each said airfoil shape having a respective maximum thickness and a respective maximum absolute thickness; The airfoil shapes are arranged at their respective minimum absolute thicknesses and maximum chords such that, in the stowed configuration, the airfoil shapes of the two wings in the first substantially overlapping spatial relationship can be stowed within the envelope cross-sectional shape while simultaneously maximizing the occupancy of the envelope cross-sectional area.

[0034] For example, in the stowed configuration, the corresponding secondary elements of the airfoil shapes of one of the wings are set at a different flap angle compared to the respective secondary elements of each airfoil shape of the other wings.

[0035] Additionally or alternatively, for example, the two wings may comprise the first wing having the first airfoil shape (the first airfoil shape including a first leading edge and a first trailing edge) and the second wing having the first airfoil shape (the second airfoil shape including a second leading edge and a second trailing edge), e.g., in the pre-deployed configuration, the first leading edge is generally overlapped with the second trailing edge.

[0036] Additionally or alternatively, for example, the ratio of each maximum absolute thickness to chord of each wing is such that it allows the wing to be deployed from a pre-deployed configuration without interference to the deployed configuration.

[0037] Additionally or alternatively, for example, each secondary element of each airfoil shape of each wing has a hinge axis to allow rotation of each secondary element relative to its respective primary element. For example, the hinge axis faces the pressure surface of each secondary element. Additionally or alternatively, for example, for each airfoil shape, each hinge axis is spaced apart from the pressure surface of each secondary element. For example, the spacing is less than 50% of the maximum thickness of the secondary element. For example, the spacing may be 40%, 30%, 20%, 10%, 5%, 1%, or 0% of the maximum thickness of the secondary element.

[0038] Additionally or alternatively, for example, in the pre-deployment configuration, a respective secondary element of each airfoil shape of each wing is set to a respective first negative flap angle.

[0039] Additionally or alternatively, for example, in the pre-deployment configuration, each secondary element of each airfoil shape of each wing is set to the same flap angle.

[0040] Additionally or alternatively, for example, in the pre-deployment configuration, each secondary element of each airfoil shape of each wing is set to a respective first flap angle such that no portion of each second flap is tangent to the leading edge of the respective airfoil shape and intersects an imaginary line intersecting the respective hinge axis.

[0041] Additionally or alternatively, for example, the wing system is configured to be attached to the underside of an aircraft fuselage. For example, in the stowed configuration, the secondary elements of the airfoil shapes of the wing are set to negative flap angles. Additionally or alternatively, for example, in the stowed configuration, the secondary elements of the airfoil shapes of the uppermost wing are set to negative flap angles in a range of −25° to −35°, and the secondary elements of the airfoil shapes of the lowermost wing are set to negative flap angles in a range of −55° to −65°. Additionally or alternatively, for example, in the stowed configuration, the secondary elements of the airfoil shapes of the wing are set to second flap angles, such that the trailing edges and hinge axes of the secondary elements are on opposite sides of the chord lines of the airfoil shapes.

[0042] Additionally or alternatively, for example, the wing system is configured to be mounted above an aircraft fuselage. For example, in the stowed configuration, the secondary elements of the airfoil shapes of the respective wing are set to respective positive flap angles. Additionally or alternatively, for example, in the stowed configuration, the secondary elements of the airfoil shapes of the uppermost wing are set to respective positive flap angles within a range of +15° to +25°, and the secondary elements of the airfoil shapes of the lowermost wing are set to respective positive flap angles within a range of +1° to +10°.

[0043] Additionally or alternatively, for example, each of the airfoil shapes may have a first maximum spacing of each pressure surface from the chord line and a second maximum spacing of each hinge axis from the chord line, where the first maximum spacing and the second maximum spacing are similar to, for example, within 10% of each other.

[0044] Additionally or alternatively, for example, for each of the airfoil shapes, the respective maximum thickness-to-chord ratios and the respective maximum absolute thickness-to-chord ratios are within 10% of each other, such as 1% of the chord.

[0045] Additionally or alternatively, for example, for each of the airfoil shapes, the respective maximum thickness to chord ratio is in the range of 0.10 to 0.14.

[0046] Additionally or alternatively, for example, for each of the airfoil shapes, the respective maximum absolute thickness to chord ratio is in the range of 0.10 to 0.15.

[0047] Additionally or alternatively, for example, each of the airfoil shapes has a respective maximum thickness at a location less than 25% of the chord from the leading edge of the airfoil shape.

[0048] Additionally or alternatively, for example, each of the airfoil shapes has a respective maximum thickness at about 20% of the chord from the leading edge of the airfoil shape.

[0049] Additionally or alternatively, for example, each of the airfoil shapes has a camber shape with a maximum camber of less than or equal to about 6% of the chord.

[0050] Additionally or alternatively, for example, each of the airfoil shapes may include at least one of the following: Each of the gaps varies between about 2.3% and about 1.2% of the chord as the respective flap angle increases from 0° to 30°. Each of the gaps varies between about 2.3% and about 2.5% of the chord as the respective flap angle decreases from 0° to -20°.

[0051] Additionally or alternatively, for example, for each of the airfoil shapes, the respective leading edge radius is about 2% of the chord.

[0052] Additionally or alternatively, for example, for each said airfoil shape, each said primary element has a primary element trailing edge and each said secondary element has a secondary element leading edge that includes any of the following: The primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord. The primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord, and each longitudinal overlap varies between about 5% and about 3% of the chord as the respective flap angle increases from 0° to 30°. the primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord, and for each of the airfoil shapes, the respective longitudinal overlap varies between about 5% and about 5.5% of the chord as the respective flap angle decreases from 0° to -20°.

[0053] Additionally or alternatively, for example, each of the airfoil shapes may be configured as a high-lift airfoil shape. For example, the airfoil shapes may be configured to have a maximum lift coefficient of at least 2.5 at a flap angle of zero, or the airfoil shapes may be configured to have a maximum lift coefficient of greater than 2.3 at a flap angle of zero. Additionally or alternatively, for example, the maximum lift coefficient corresponds to an angle of attack between about 15° and about 16°.

[0054] Additionally or alternatively, for example, the airfoil shape may be arranged to have a lift coefficient of 1.0, which corresponds to an angle of attack of approximately 0°.

[0055] Additionally or alternatively, for example, the airfoil shapes of the two wings have a total cross-sectional area that occupies between 40% and 30% of the envelope cross-sectional area.

[0056] Additionally or alternatively, for example, the envelope cross-sectional shape may be in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion. For example, for each of the airfoil shapes, the respective first ratio of the chord to the diameter is in the range of 70% to 87%. For example, for each of the airfoil shapes, the respective first ratio is in the range of 75% to 85%. Additionally or alternatively, for example, for each of the airfoil shapes, the respective second ratio of the maximum thickness to the diameter is in the range of 7% to 12%. Additionally or alternatively, for example, for each of the airfoil shapes, the respective second ratio of the maximum absolute thickness to the diameter is in the range of 8% to 14%.

[0057] Additionally or alternatively, for example, each of the wings may be arranged with a zero taper, where each of the wings has a uniform airfoil shape along its respective span, or, for example, each of the wings may be arranged with a non-zero taper, where each of the wings has a uniform airfoil shape along its respective span.

[0058] Additionally or alternatively, for example, the wing system may be arranged to selectively allow the wings to transition from a stowed configuration to a pre-deployed configuration under predetermined conditions.

[0059] Additionally or alternatively, for example, the wing system may be arranged to maintain a negative deflection angle of the respective airfoil shape of each wing during deployment from the pre-deployed configuration to the deployed configuration.

[0060] According to a second aspect of the presently disclosed subject matter, there is provided an aircraft comprising a wing system as defined above with respect to the first aspect of the presently disclosed subject matter, for example, the aircraft being a UAV.

[0061] According to a third aspect of the presently disclosed subject matter, there is provided a launch system including an aircraft as defined above with respect to the second aspect of the presently disclosed subject matter, housed within a fairing having an internal chamber defining said envelope, said envelope cross-sectional shape, and said envelope cross-section.

[0062] For example, the launch system is configured to selectively release the aircraft from the fairing. For example, the launch system is configured to selectively transition the wing system from the stowed configuration to the pre-deployed configuration after releasing the aircraft from the fairing. For example, the launch system is configured to selectively transition the wing system from the pre-deployed configuration to the deployed configuration and enable aerodynamic flight of the aircraft.

[0063] In at least one example feature of the presently disclosed subject matter, a deployable wing of an aircraft is provided with a two-element airfoil shape, whereby, in the stowed configuration, the secondary element of the airfoil shape of one wing is set at a different flap angle compared to each secondary element of each airfoil shape of the other wing, thereby maximizing the occupancy of an envelope cross-sectional area within which it is desired to stow the wing system in the stowed configuration. Such a feature may be particularly useful in applications of the presently disclosed subject matter where such envelope cross-section is semicircular or where the two overlapping airfoil shapes in the stowed configuration are in the form of segments or trapezoids that can have different widths for storage.

[0064] As another feature of at least one example of the presently disclosed subject matter, a deployable wing of an aircraft is provided with a two-element airfoil shape, wherein a secondary element of the airfoil shape of the one wing and each secondary element of each airfoil shape of the other wing must be pivoted to a flap angle different from a flap angle corresponding to the stowed configuration so that the wing can be deployed from the pre-deployed configuration to the deployed configuration.

[0065] As another feature of at least one example of the presently disclosed subject matter, a deployable wing of an aircraft is provided with a two-element airfoil shape that allows for a high degree of flexibility in the design of the two-element airfoil shape (including maximum absolute thickness, maximum airfoil thickness, leading edge radius, camber, flap hinge point location, and flap angle) to allow efficient packaging within the envelope provided by the fairing while providing high lift and endurance, and mission adaptability via flap deflection settings such as takeoff and landing, cruise, loiter, etc.

[0066] As another feature of at least one example of the presently disclosed subject matter, an aircraft deployable wing is provided with a two-element airfoil shape, which may enable extending the flight envelope, particularly in the low speed range, by providing high lift characteristics, including, for example, by reducing the thickness-to-chord ratio.

[0067] In another aspect of at least one example of the presently disclosed subject matter, a deployable wing of an aircraft is provided with a two-element airfoil shape, which has high lift and long durability. [Brief explanation of the drawings]

[0068] For a better understanding of the subject matter disclosed herein, and to illustrate how it may be carried out in practice, embodiments will now be described, by way of non-limiting example, with reference to the accompanying drawings, in which:

[0069] 1(a) is a cross-sectional view of an aircraft according to a first example of the subject matter of the present disclosure in a stowed configuration stored in a fairing, and FIG. 1(b) is a cross-sectional view of the aircraft according to the example of FIG. 1 in a pre-deployment configuration.

[0070] FIG. 2 is a cross-sectional view of the example aircraft of FIG. 1 in a deployed configuration.

[0071] Figure 3 shows a slotted airfoil section on which the example aircraft wing of Figure 1 is based. Figure 3(a) shows a schematic of the flap deflection of the flap of the airfoil of Figure 3, and Figure 3(b) shows the relative sizes of the primary and secondary elements of the airfoil of Figure 3.

[0072] FIG. 4 shows a schematic of the geometric relationship between the maximum absolute thickness of the slotted airfoil section of FIG. 3 and the airfoil thickness.

[0073] FIG. 5(a) shows the reference slotted airfoil section, and FIG. 5(b) shows the slotted airfoil section of FIG.

[0074] FIG. 6 compares diagrammatically the thickness distribution of the airfoils of FIGS. 5(a) and 5(b) along their respective chords.

[0075] FIG. 7 provides a schematic comparison of the camber distribution of the airfoils of FIGS. 5(a) and 5(b) along their respective chords.

[0076] Figure 8 schematically compares the lift coefficient distributions of the airfoils of Figures 5(a) and 5(b) with the angle of attack when the flap angle is zero and the Reynolds number is 1,000,000. Figure 8(a) schematically compares the lift coefficient distributions of the airfoils of Figures 5(a) and 5(b) with the drag coefficient when the flap angle is zero and the Reynolds number is 1,000,000. Figure 8(b) schematically compares the hinge moment coefficient distributions of the second-order elements of the airfoils of Figures 5(a) and 5(b) with the lift coefficient when the flap angle is zero and the Reynolds number is 1,000,000. Figure 8(c) schematically compares the hinge moment coefficient distributions of the second-order elements of the airfoils of Figures 5(a) and 5(b) with the lift coefficient when the flap angle is +10° and the Reynolds number is 1,000,000.

[0077] FIG. 9 shows a schematic of the gap size and longitudinal overlap between the primary and secondary elements of the slotted airfoil section of FIG.

[0078] Figure 10(a) schematically compares the gap size distribution of the airfoils of Figures 5(a) and 5(b) with the flap deflection angle. Figure 10(b) schematically compares the longitudinal overlap distribution of the airfoils of Figures 5(a) and 5(b) with the flap deflection angle.

[0079] Figure 11(a) shows the slotted airfoil section of Figure 3 in a stowed configuration, Figure 11(b) shows the slotted airfoil section of Figure 3 in a pre-deployed configuration, and Figure 11(c) shows the slotted airfoil section of Figure 3 in a deployed configuration.

[0080] Figure 12(a) is a cross-sectional view of the aircraft of Figure 1 showing the envelope therein, and Figure 12(b) shows the cross-sectional envelope of Figure 12(a).

[0081] FIG. 13 is a cross-sectional view of an aircraft according to a variation of the example of FIG. 1 in a stowed configuration retracted into a fairing.

[0082] FIG. 14 compares the lift coefficient distribution of the example airfoil of FIG. 3 with the angle of attack when the flap angles are 0°, +10°, and −10°.

[0083] Referring to FIGS. 1, 1(a), and 2, an aircraft according to a first example of the subject matter of this disclosure is generally designated 100 and includes a wing system 200 which is novel per se.

[0084] In at least this example, aircraft 100 is a powered, fixed-wing aircraft, includes a subsonic configuration, and has a fuselage section 110 (also referred to herein as fuselage 110), wing system 200, a suitable propulsion system (not shown), and tails (not shown, including, by way of example and not limitation, horizontal and vertical stabilizers). Additionally, while the subject matter of this disclosure finds particular application in UAV aircraft, the subject matter of this disclosure may also be applied to manned aircraft, mutatis mutandis, particularly general aviation, subsonic transport, naval aviation, and the like. In alternative variations of this example and in other examples, aircraft 100 may instead be an unpowered, fixed-wing, subsonic aircraft, manned or unmanned.

[0085] Aircraft 100 further comprises a control system including a controller (not shown) arranged to operate aircraft 100 at least within its flight envelope.

[0086] As will become more apparent herein, the wing system 200 has a stowed configuration SC, a pre-deployed configuration PC, and a deployed configuration DC.

[0087] The wing system 200 includes two wings 220, specifically a left wing 220P and a right wing 220S. Accordingly, unless otherwise specified herein, reference numeral 220 also refers to each of the left wing 220P and the right wing 220S.

[0088] Each of the wings 220 is mounted to the fuselage 110 for pivoting about a respective pivot axis PA, such that each of the wings 220 can pivot about its respective axis PA between a respective pre-deployment position PC and a respective deployed position DC. In at least this example, the pivot axes PA of the two wings 220 are coaxial, although in other variations of this example and other examples, the pivot axes PA of the two wings 220 are not coaxial, e.g., parallel and spaced apart, or non-parallel and spaced apart. Examples of suitable mechanisms for providing such pivoting are known in the art.

[0089] In at least this example, the pivot axis PA of each wing 220 is parallel to the yaw axis of the aircraft 100 .

[0090] Left wing 220P and right wing 220S are substantially mirror images of each other and are otherwise substantially identical, at least in this example, except for the overlapping arrangement of portions of each wing 220 near each pivot axis PA.

[0091] In at least this example, aircraft 100 is arranged to be stored in fairing 800, while wing system 200 is in a storage configuration SC. For example, such fairing 800 may provide aerodynamic isolation for aircraft 100 as it accelerates to a required speed and altitude, and / or such fairing may protect the aircraft during storage or transport. For example, aircraft 100 may be accelerated to a required speed and altitude via an external launch system, such as a launch barrel, or via a booster rocket (e.g., as its payload), or may be delivered to a required altitude and required speed, for example, via an aircraft carrier.

[0092] Fairing 800 has an internal shape and a fairing cross-section in a transverse plane perpendicular to the longitudinal axis LAF of the fairing to accommodate a respective cross-section of aircraft 100 (i.e., a respective external shape and cross-sectional area of ​​aircraft 100). At least in this example, fairing 800 has a substantially uniform internal shape and a substantially uniform cross-section in a transverse plane perpendicular to the longitudinal axis LAF for at least a cross-section of the aircraft that includes wing 220 in stowed configuration SC. The longitudinal axis LAF of the fairing is generally defined at the centerline of the fairing.

[0093] It should be noted that, at least for some implementations of the present example in which acceleration of the aircraft 100 (while stowed in the fairing 800 of each assembly) is achieved by applying forces to the assembly (e.g., thrust from a booster rocket or thrust resulting from firing from a launch barrel), it may be advantageous to align such forces with the fairing's longitudinal axis LAF, and particularly with the fairing's centerline. In at least such cases, there may be corresponding advantages to locating the aircraft 100's center of gravity CG on the fairing's longitudinal axis LAF ( FIG. 1 ), and particularly on the fairing's centerline. For example, such an arrangement may minimize or eliminate the risk of nutation. In at least some such cases, the aircraft 100 may be positioned such that its center of gravity CG is stowed within the fuselage 110. Thus, in a cross-section along the fairing's longitudinal axis LAF, the respective cross-sections of the fuselage 110 occupy more than half of the respective interior cross-sections of the fairing 800, while the two wings 220 (in the stowed configuration SC) occupy less than half of the interior cross-sections of the fairing 800. In at least some such cases, the fuselage 110 has a greater mass than the wings 220. As will become more apparent herein, in accordance with one aspect of the subject matter of the present disclosure, the maximum absolute thickness ABS of the wings is minimized, allowing the two wings 220 to be attached to the smallest possible portion of the interior volume provided by the fairing 800, thereby allowing the cross-sectional area of ​​the fuselage to be maximized, which in turn can maximize, for example, the payload capacity or the aircraft 100.

[0094] 1, the wings 220 are below the fuselage, i.e., the suction surface of the top wing 220 (e.g., starboard wing 220S) of the wing system 200 faces the bottom 115 of the fuselage 110. However, in at least one alternative variation of the example of FIG. 1, now referring to FIG. 13, the wings 220 may be above the fuselage, i.e., the pressure surface of the bottom wing 220 (e.g., starboard wing 220P) of the wing system 200 faces the top 116 of the fuselage 110.

[0095] 3, each wing 220 is based on (i.e., includes a plurality of corresponding airfoil sections) a two-element slotted airfoil 300, also referred to interchangeably herein as an SA airfoil or SA airfoil 300, which exhibits particularly high-lift characteristics. In other words, each wing 220 includes a plurality of airfoil sections each corresponding to the two-element slotted airfoil 300.

[0096] As used herein, the term "slotted airfoil" (SA-airfoil) generally refers to a two-element wing (also referred to herein as a two-element airfoil) consisting of a first or primary element, an airfoil body, and a second or secondary element, in the form of a flap or aileron. The secondary element of the SA-airfoil is separated from the body by a slot that is substantially open to the airflow with any deflection of the flap. A feature of the slotted airfoil type structure is that it facilitates actuation of the secondary element through positive or negative deflection angles. Another feature of this type of structure is that a box-like, relatively rigid structure can be achieved. Generally, when the second airfoil element is coupled to an actuator, the SA-airfoil can provide adjustable geometric features and can have built-in options for operating to provide takeoff / landing flaps, ailerons, air brakes, and deceleration in maximum speed flight. The SA-airfoil can include a mission-adapted two-element high-lift arrangement and is inherently a high-lift airfoil, particularly suited for the development of long-endurance UAVs and other aircraft that loiter with a high lift coefficient. Such SA-airfoils can be designed for cruise / loitering flight with high lift coefficients, optionally relying on rotation of secondary elements about external hinge points to adjust the airfoil to various flight configurations.

[0097] In this example implementation, where the wings 220 have a generally rectangular planform, the SA airfoil 300 is uniform in size and shape along each span of the wings 220.

[0098] In this example implementation, where the wings 220 are generally trapezoidal (i.e., tapered) or any other shape, the SA airfoil 300 is uniform in shape along the span of each of the wings 220, but the size of the SA airfoil 300 at each spanwise position along the span of each of the wings varies proportionally to the size of the chord C at each spanwise position.

[0099] In at least one such implementation, the respective thickness-to-chord ratios along the respective spans of the wings 220 may be uniform.

[0100] In at least one variation of the above implementation, the respective thickness-to-chord ratios along the respective spans of the wings 220 may be non-uniform. For example, the respective thickness-to-chord ratios along the respective spans may decrease in a direction from the respective root toward the respective tip. In such cases, according to one aspect of the presently disclosed subject matter, the respective maximum thickness-to-chord ratios between the root and the tip are still maintained within a range that provides high lift characteristics for the respective AS airfoil 300. For example, such a range may be 10% to 14%. In at least some such cases, there may be an advantage to having a relatively thin tip (compared to the root).

[0101] The airfoil 300 includes a first or primary element 332 (also referred to herein interchangeably as an "airfoil body" or "body") and a pivotable secondary element 334 (also referred to herein interchangeably as an "airfoil flap" or "flap"). Correspondingly, therefore, each wing 220 has a primary wing portion 222 (also referred to herein as a main wing portion) corresponding to the primary element 332 of the airfoil 300 and a secondary wing portion 224 (also referred to herein as a flap portion) corresponding to the secondary element 334 of the airfoil 300. At least in this example, the two-element structure of each wing 220, i.e., each primary wing portion 222 and each secondary wing portion 224, generally extends from the wing tip 221 to the wing root 226 ( FIG. 2 ), while the portion 227 of the wing 220 that is inboard of the wing root 226 and in overlapping relationship with the fuselage 110 in deployed configuration DC, does not require an aerodynamic shape or two-element wing structure.

[0102] The primary element 332 includes a leading edge 338 of the SA airfoil 300 that coincides with the leading edge 252 of each blade 220. The primary element 332 also includes major portions 331a and 333a of the suction surface 331 and pressure surface 333 of the SA airfoil 300. The primary element 332 also includes a trailing edge 338a, which is interchangeably referred to herein as the trailing edge of the primary element 332.

[0103] The secondary element 334 is arranged as a single continuous flap portion and includes the trailing edge 339 of the SA airfoil 300 that coincides with the trailing edge 254 of the wing 220, and the secondary element 334 also includes minor portions 331b and 333b of the suction surface 331 and the pressure surface 333 of the SA airfoil 300, respectively.

[0104] The SA airfoil 300 includes a chord C defined along a chord line CL that corresponds to a zero angle of attack α for the SA airfoil 300. In other words, the chord line CL is generally parallel to the longitudinal axis of the aircraft 100 or has a desired inclination angle relative to the longitudinal axis of the aircraft 100, and / or the chord line CL is generally parallel to the roll axis of the aircraft 100 or has a desired inclination angle about the pitch axis of the aircraft 100.

[0105] It should be noted that in at least some implementations of the above examples, the chord lines CL automatically have the desired inclination relative to the longitudinal axis of the aircraft 100 when the wings 220 are deployed from the pre-deployed position PC to the deployed position DC without the need to manipulate the wings 220 other than to pivot the wings about the pivot axis PA. In other words, in the pre-deployed position PC and in the corresponding stowed position SC, the wings 220 can be oriented relative to each other and relative to the pivot axis PA such that the wings 220 automatically provide the desired inclination relative to the longitudinal axis of the aircraft 100 in the deployed position DC.

[0106] Secondary element 334 is pivotable about hinge point 359 (also referred to herein as the hinge axis) to provide a range of positive and negative deflection angles δ of secondary element 334 relative to primary element 332 .

[0107] The deflection angle δ is referred to herein interchangeably as the flap deflection angle δ or the flap angle δ.

[0108] The deflection angle δ of the secondary element 334 is defined relative to the flap datum line FL, which is at an angle θ relative to the chord line CL, i.e., in Figure 3, the leading edge 335 is vertically higher than the trailing edge 339. In other words, a zero deflection angle δ of the secondary element 334 corresponds to the secondary element 334 being aligned with the flap datum line FL.

[0109] Referring also to FIG. 3( a), a positive deflection angle δ corresponds to a flap reference line FL that is at an angle greater than the angle θ relative to the chord line CL (i.e., the secondary element 334 pivots about the hinge point 359 in a clockwise direction in FIG. 3), while a negative deflection angle δ corresponds to a flap reference line FL that is at an angle less than the angle θ relative to the chord line CL (i.e., the secondary element 334 pivots about the hinge point 359 in a counterclockwise direction in FIG. 3).

[0110] It should be noted that in at least some implementations of the above example, the position of the chord line CL for the airfoil 300 may also be selected to correspond to a zero angle of attack α for the SA airfoil 300. This results in a lift coefficient C for the airfoil when the deflection angle δ is zero. l is 1.0, and at the same time, the secondary element 334 can be pivoted to a deflection angle δ within an "aerodynamic" range, for example, up to ±20° or up to ±25°, to enable aerodynamic operation of the secondary element 334. An "aerodynamically operating secondary element 334" is one in which, at each deflection angle δ within the aerodynamic range, attached flow is maintained on the airfoil 300 and its efficient control is provided by the airfoil 300; i.e., within the aerodynamic range, changes in the deflection angle δ increase or decrease the lift coefficient C of the airfoil 300. l It is also noted that in the above example implementation, the position of the chord line CL for the airfoil 300 reduces drag during cruise and also provides a maximum lift coefficient C at zero deviation δ, consistent with high lift performance. l , i.e., the maximum lift coefficient C lis greater than 2.2 under such conditions, for example 2.3 or greater. In yet another alternative variation of this example, the "aerodynamic" range may be, for example, up to ±30°.

[0111] In at least some implementations of the above example, the angle θ may be approximately 12°.

[0112] In yet other implementations of the above example, the angle θ can be in a range of about 8° to about 16°. In examples where the angle θ is less than 12°, e.g., less than 8° to 12°, the range of positive / negative deflection angles δ to enable aerodynamic operation of the secondary element 334 can vary from an aerodynamic range of greater than +20° to less than −20° (e.g., up to +24° / −16°) compared to ±20° in the example of FIG. 3, or from an aerodynamic range of greater than +25° to less than −25° (e.g., up to +29° / −21°) compared to ±25° in the example of FIG. 3. In examples where the angle θ is greater than 12°, for example greater than 12° to 16°, the range of positive / negative deflection angles δ to enable aerodynamic operation of the secondary element 334 may vary from an aerodynamic range of less than +20° to greater than −20° (e.g., up to +16° / −24°) compared to ±20° in the example of FIG. 3, or may vary from an aerodynamic range of less than +25° to greater than −25° (e.g., up to +21° / −29°) compared to ±25° in the example of FIG. 3.

[0113] Referring to FIG. 14, at a flap deflection angle δ of 0°, at least in this example, the SA airfoil 300 exhibits a maximum lift coefficient of about 2.5 at an angle of attack α of about 15°-16°, while the lift coefficient is about 1.0 at an angle of attack α of about 0° and about 2.1 at an angle of attack α of about 20°. l For the same airfoil 300, at a positive flap deflection angle δ of +10°, the SA airfoil 300 exhibits a maximum lift coefficient of over 2.6 at an angle of attack α of about 14°-15°, while the lift coefficient is about 1.6 at an angle of attack α of about 0° and decreases to about 2.1 at an angle of attack α of about 20°. lFor the same airfoil 300, at a negative flap deflection angle δ of −10°, the SA airfoil 300 exhibits a maximum lift coefficient of about 2.15 at an angle of attack α of about 17°, while the lift coefficient is about 0.4 at an angle of attack α of about 0° and decreases to about 2.0 at an angle of attack α of about 20°. l decreases to.

[0114] It should be noted that, at least in this example, the secondary element 334 can be pivoted to a deflection angle δ within an aerodynamic range of, for example, at least ±20° or ±25° to enable aerodynamic movement of the secondary element 334. However, at least in this example, it is also possible to pivot the secondary element 334 to a deflection angle δ over a “mechanical range” greater than the aerodynamic range of ±20° or ±25°. Such a mechanical range may be limited by mechanical constraints. For example, the example of FIG. 3 has a negative mechanical deflection angle δ limit of approximately −72°, at which the suction surface of the secondary element 334 abuts the aft end of the primary element 332. In this example, the positive mechanical deflection angle δ limit is greater than 90°, at which the aft edge of the secondary element 334 abuts the pressure surface of the primary element 332. In alternative variations of this example, the mechanical range of deflection angle δ can be any of: +90° to -90°; +90° to -80°; +90° to -70°; +90° to -60°; +80° to -90°; +90° to -80°; +80° to -70°; +80° to -60°; +70° to -90°; +70° to -80°; +70° to -70°; +70° to -60°; +60° to -90°; +60° to -80°; +60° to -70°; +60° to -60°; +50° to -90°; +50° to -80°; +50° to -70°; +50° to -60°.

[0115] It should be noted that the actuators of each secondary element 334 can be adapted to the required range of positive and negative deflection angles (including the aerodynamic range and mechanical range required for the particular wing 220).

[0116] 3(b), at least in this example, primary element 332 has a primary chord Cl that is about 77% of chord C, where primary chord Cl is defined as the distance between leading edge 338 and trailing edge 338a along a direction parallel to chord line CL. Also, at least in this example, secondary element 334 has a secondary chord C2 that is about 27.5% of chord C, where secondary chord C2 is It is defined as the distance between leading edge 335 and trailing edge 339 along a direction parallel to chord line CL when secondary element 334 is at a deflection angle δ of 0°. In alternative variations of this example and other examples, primary element 332 can have a primary chord C1 within any of the following ranges of chord C: 70%-80%, 65%-85%, or 75%-80%. In this example or other variations of this example, secondary element 334 can have a secondary chord C2 within any of the following ranges of chord C: 20%-30%, 25%-35%, or 25%-30%.

[0117] Slot 355 separates a leading end 335 (also referred to interchangeably herein as a leading edge of secondary element 334 ) of secondary element 334 from a trailing end 338 a of primary element 332 .

[0118] For example, the slot 355 may have an exit width of about 2.3% of the chord C of the airfoil in the undeflected position of the flap element, ie, at a deflection angle δ of zero.

[0119] Without being bound by theory, the inventors believe that airflow through the slots 255 provides improved aerodynamic efficiency and linearity to the airfoil, provides or enhances high lift characteristics to the airfoil, and / or provides or enhances flap efficiency characteristics to the airfoil, compared to a similar non-slotted airfoil.

[0120] 3(a), the deflection angle δ of the secondary element 334 generally depends on the particular mode of operation of the wing 220. For example, in cruise / loitering mode, the deflection angle δ is 0°, in takeoff / landing mode TLM, the deflection angle δ is +20° to +25°, in aileron mode AM, the deflection angle δ is ±20° to ±25°, and in maximum speed deflection release mode DM to negative flap deflection, the deflection angle δ is −10° to −15°.

[0121] The exact shape and size of the slot 355 may generally vary with the deflection angle δ of the secondary element 334 , and thus will depend largely on the particular mode of operation of the airfoil 220 .

[0122] Furthermore, at least in this example, in the storage configuration SC, the deflection angle δ can be between +20° and −60°.

[0123] For example, in at least one implementation of the above example, referring again to FIG. 1 , one wing 220 (e.g., right wing 220S) can have a deflection angle δ (annotated in FIG. 1 as angle δ1) of approximately −30°, e.g., −32°, while the other wing 220 (e.g., left wing 220P) can have a deflection angle δ (annotated in FIG. 1 as angle δ2) of approximately −60°. For example, in at least one other implementation of the above example, referring again to FIG. 13 , one wing 220 (e.g., right wing 220S) can have a deflection angle δ (annotated in FIG. 13 as angle δ1′) of approximately +20°, while the other wing 220 (e.g., left wing 220P) can have a deflection angle δ of approximately +5° (annotated in FIG. 13 as angle δ2′).

[0124] Furthermore, at least in this example, with reference to FIG. 1(a), in the pre-deployment positioned PCs, the deflection angles δ may be identical to one another, for example, between −10° and −15°.

[0125] Referring again to FIG. 3 , the hinge point 359 of the secondary element 334 is spatially fixed relative to the SA airfoil 300. In particular, the hinge point 359 of the secondary element 334 is spatially fixed at a position displaced outward relative to the lower (pressure) surface of the secondary element 334. In at least this example, the spacing FS is minimized to enable minimizing the value of the maximum absolute thickness ABT of the airfoil 300, and therefore the hinge point 359 is as close to the secondary element 334 as possible. In other variations and examples of this example, the hinge point 359 can be enclosed within the secondary element 334. Also, in other variations and examples of this example, the hinge point 359 can face the suction surface of the secondary element 334, i.e., each small portion 331b of the suction surface 331.

[0126] Preferably, the hinge point 359 is positioned further relative to the secondary element 334 so that the secondary element 334 is aerodynamically stable relative to the hinge point 359 (i.e., the hinge point is preferably forward of the aerodynamic center of the secondary element 334) and also minimizes actuator loads where possible.

[0127] For example, the spacing FS is less than 50% of the maximum thickness of the secondary element 334. For example, the spacing FS can be 40%, 30%, 20%, 10%, 5%, 1%, or 0% of the maximum thickness of the secondary element 334.

[0128] A suitable actuation mechanism (not shown) is provided for actuating secondary element 334 to adopt a range of desired positive and negative deflection angles, including the aerodynamic range and the mechanical range.

[0129] In an alternative variation of this example, secondary element 334 may instead be formed as a slotted flap portion rather than a continuous flap portion. In such an example, the slotted flap portion may have a slotted flap structure that includes two or more sub-elements rigidly connected to each other and separated by a minor slot such that the secondary element still rotates about hinge point 359 as a rigid body.

[0130] Note that leading edge 338 is also rounded and has leading edge radius 315 to promote high lift characteristics. For example, leading edge radius 315 may be greater than 1.7% of chord C. In this example, leading edge radius 315 is approximately 2% of chord C.

[0131] It should also be noted that the leading edge 335 is also rounded to allow for a wide range of deflection angles δ while maintaining attached flow on the secondary element 334 .

[0132] It should be noted, with particular reference to FIG. 4, which shows the airfoil 300 in a deployed PC, that the SA airfoil 300 also has a minimum value of maximum absolute thickness ABT that is the same magnitude as the maximum thickness t of the SA airfoil 300.

[0133] "Maximum absolute thickness" means the maximum distance between any point on the bottom of the SA airfoil 300 and any point on the top of the SA airfoil 300, including the hinge point 359, along a direction parallel to the respective pivot axes PA.

[0134] By pivoting the secondary elements 334 about the hinge points 359 to a suitable negative deflection angle δ, the spacing of any portion of the secondary elements 334 from the chord line CL can be minimized in a direction parallel to the respective pivot axis PA (and extending downward in the view of FIG. 3 ) to provide a minimum value for such maximum absolute thickness ABT for a given shape of the SA airfoil 300. For example, such a suitable negative deflection angle δ can be similar to the value of the angle θ, for example. For example, such a suitable negative deflection angle δ can be in the range of, for example, −10° to −15°.

[0135] At least in this example, when the deflection angle δ of the secondary element 334 is zero, the trailing edge 339 is below the hinge point 359, while when the deflection angle δ is negative so as to provide a minimum value for the maximum absolute thickness ABT, the trailing edge 339 is above the hinge point 359.

[0136] At least in this example, where hinge point 359 is spaced a distance FS from the pressure surface of secondary element 334, the appropriate negative deflection angle δ is the angle that causes all portions of secondary element 334 to be closer to chord line CL than hinge point 359. At least at this location, the position of secondary element 334 is referenced as a pre-deployment position of secondary element 334, which corresponds to pre-deployment configuration PC.

[0137] This minimum value of maximum absolute thickness ABT can be further minimized for a given chord C of the SA airfoil 300 by controlling one or more of the camber, the spacing of the lowest point 301 of the pressure "weld" 310 of the airfoil 300 from the chord line CL, and the maximum airfoil thickness t.

[0138] For example, by reducing the camber of the airfoil 300, the minimum value of the maximum absolute thickness ABT can be reduced.

[0139] For example, decreasing the spacing of the lowest point 301 of the pressure "welt" 310 of the airfoil 300 from the chord line CL can decrease the minimum value of the maximum absolute thickness ABT.

[0140] For example, by reducing the maximum thickness of the airfoil 300, the minimum value of the maximum absolute thickness ABT can be reduced.

[0141] On the other hand, reducing the camber of the airfoil 300 and / or reducing the spacing of the lowest point 301 of the pressure "weld" 310 of the airfoil 300 from the chord line CL and / or reducing the maximum thickness of the airfoil 300 may lead to a reduction in the leading edge radius, which is contrary to general principles when designing high-lift two-element airfoils.

[0142] Referring to FIG. 3, at least in this example, the spacing BS of the pressure pressure pressure minimum point 301 from the chord line CL of the airfoil 300 and the spacing BS of the flap hinge point 359 from the chord line CL of the airfoil 300 are similar in magnitude.

[0143] At least for each airfoil shape 300, the respective maximum thickness-to-chord ratios and the respective minimum maximum absolute thickness-to-chord ratios are within 10% of each other. For example, at least in this example, the respective maximum thickness-to-chord ratios and the respective minimum maximum absolute thickness-to-chord ratios are within 1% of the chord C.

[0144] In at least this example, for each airfoil shape 300, the respective maximum thickness to chord ratio is, for example, in the range of 0.10 to 0.14.

[0145] In at least this example, for each airfoil shape 300, the respective maximum absolute thickness to chord ratio is, for example, in the range of 0.10 to 0.15.

[0146] In at least this example, each airfoil shape 300, and in particular its primary elements 332, has a maximum camber of about 6% of chord C. Each airfoil shape 300, and in particular its primary elements 332, can be determined relative to a Cartesian coordinate system as Δy / chord C, where for each spacing x from the leading edge 338 (i.e., each x coordinate), the value of Δy is determined from the following equation: Δy = (y upper (x) + y lower (x)) / 2 where y upper (x) is the y coordinate of the upper (suction) surface of the primary element 332, and y lower (x) is the y-coordinate of the lower (pressure) surface of the primary element 332. Other ways of defining camber are of course possible. For example, the camber measured in this way may change when a tilt angle is applied to the chord line CL.

[0147] 3 and 4, at least in this example, the minimum value of the maximum absolute thickness ABT of the SA airfoil 300 is actually the minimum height that the airfoil 300 can adopt along a direction parallel to the pivot axis PA. Thus, the minimum value of the maximum absolute thickness ABT is the distance between a first imaginary line IM1 tangentially contacting the top of the airfoil 300 and a second imaginary line IM2 tangentially contacting the bottom of the airfoil 300, which includes the hinge point 359. At least in this example, the bottom of the airfoil 300 is the hinge point 359 (when the secondary elements 334 are in their pre-deployment configurations), although in alternative variations of this example, the hinge point 359 could be closer to the chord line CL, and thus the bottom of the airfoil 300 could instead be the lowest point 301 of the pressure "weld" 310 of the airfoil 300. The pressure antinode 310 may be defined as the portion or point of the primary element 332 that is most spaced below the chord line CL as seen in FIG.

[0148] Thus, the minimum value of the maximum absolute thickness ABT of the SA airfoil 300 is the minimum spacing (in a direction parallel to the pivot axis PA) required to be free of obstacles such that each wing 220 and airfoil 300 can pivot about its respective pivot axis PA without colliding with such obstacles, e.g., with obstacles such as other wings 220, during deployment from the pre-deployment position PC to the deployment position DC.

[0149] Referring also to FIGS. 5(a) to 10(b), the SA shape 300 can be compared with the reference airfoil shape DA, for example, as follows.

[0150] The reference airfoil DA is a well-known high-lift, long-endurance two-element airfoil, such as the Heron tip airfoil as disclosed in the following publications: "IAI Long-Endurance UAV Aerodynamic Concepts," by Misha Shepshelovich, February 25-26, 2004. The Heron 1 tip-slotted airfoil is shown in Figure 4 of the publication. It has a maximum thickness-to-chord ratio of 17.5%, providing high lift characteristics (Figure 3). "Aerodynamics of Contaminated UAV Wings," by B. Marcus, M. Sarid, and Misha Shepshelovich, February 25-26, 2005. The Heron 1 tip-slotted airfoil is shown in Figure 1 of the publication. It has a maximum thickness-to-chord ratio of 17.5%, providing high lift characteristics (Figure 9). "Icing on Two-Element High-Lift UAV Wings," by Danny Abramov and Misha Shepshelovich, March 6-7, 2013. The Heron 1 tip-slotted airfoil, shown in Figure 4 of the publication, referenced as the SAT-41 airfoil, and discussed on page 5 thereof, has high-lift characteristics (Figure 7).

[0151] According to the three publications, the Heron 1 tip slotted airfoil provides durability and lift characteristics with a maximum lift coefficient of approximately 2.3 to 2.5 at zero flap deflection.

[0152] The SA airfoil 300 has significant geometric / aerodynamic differences relative to the reference airfoil shape DA, for example:

[0153] 3 and 7, the SA airfoil 300 includes a camber line CAM that, at least in this example, peaks at about 6% of chord C and, at least in this example, remains within a range of 5% to 6% of chord C from about 21% to about 60% of chord C from the leading edge 338. This compares to the camber line of the reference airfoil shape DA, which peaks at about 8% of each chord and remains within a range of 6% to 8% of chord from about 21% to about 60% of chord from the leading edge. In FIG. 7, the camber of the airfoil 300 and the reference airfoil DA are defined in a similar manner.

[0154] 3 and 6, the SA airfoil 300, at least in this example, has a maximum thickness t' of about 13% of chord C, and includes a thickness distribution that, at least in this example, remains between 10% and 13% of chord C from about 5% to about 50% of chord C from the leading edge 338. This compares to the thickness distribution of the reference airfoil shape DA, which has a maximum thickness t' of about 17.5% of each chord, and remains between 10% and 17.5% of chord from about 5% to about 50% of chord from the leading edge.

[0155] 5(a) and 5(b), the maximum thickness t of the airfoil 300 is reached at a chord position ct that is less than 25% of the chord C, or for example, about 20% of the chord C, from the leading edge 338 of the airfoil 300. In contrast, the maximum thickness t' of the reference airfoil DA is reached at a chord position ct' that is about 30% of the chord from its leading edge 338'. Furthermore, the leading edge radius 305 of the airfoil 300 is about 2% of the chord C from the leading edge 338 of the airfoil 300. In contrast, the leading edge radius 305' of the reference airfoil DA is about 2.3% of the chord from its leading edge 338'.

[0156] Thus, because the maximum thickness t of airfoil 300 decreases more rapidly (17.5% to 13%) relative to the maximum thickness t' of reference airfoil DA, the effective leading edge radius relative to the maximum thickness of airfoil 300 is significantly larger than the maximum thickness of reference airfoil DA, compared to the relatively small decrease in the respective leading edge radii (2.3% to 2%).

[0157] Furthermore, referring again to Figures 5(a) and 5(b), the spacing FS of the flap hinge point 359 from the secondary element 334 of the airfoil 300 is significantly smaller than the spacing FS' of the flap hinge point 359' from the secondary element of the reference airfoil DA.

[0158] 5(a) and 5(b), the spacing BS of the pressure-pressure-lowest point 301 from the chord line CL of the airfoil 300 is significantly smaller than the spacing BS' of the pressure-pressure-lowest point 301' from the chord line CL' of the reference airfoil DA. Furthermore, the spacing of the pressure-pressure-lowest point 301 from the leading edge 338 of the airfoil 300 is significantly smaller than the spacing of the pressure-pressure-lowest point 301' from the leading edge 338' of the reference airfoil DA.

[0159] 8, the SA airfoil 300 exhibits a maximum lift coefficient of about 2.5 at an angle of attack α of about 15° or about 16°. At least in this example, the lift coefficient is about 1.0 at an angle of attack α of about 0°, decreasing to about 2.1 at an angle of attack α of about 20°. This compares to a maximum lift coefficient of about 2.5 for the reference airfoil shape DA at an angle of attack α of about 13° or 14°, and a maximum lift coefficient of about At an angle of attack α of 0°, the lift coefficient is about 1.1, decreasing to about 1.95 at an angle of attack α of about 20°. Meanwhile, as can be seen in Figures 8(a), 8(b), and 8(c), the SA airfoil 300 exhibits higher drag and higher actuator loads compared to the reference airfoil shape DA.

[0160] 9 and 10(a), the secondary element 334 is spatially positioned relative to the primary element 332 to provide a characteristic relationship between the gap size M of gap 355 (between the primary element 332 and the secondary element 334) and the deflection angle δ of the secondary element 334 relative to the primary element 332. This relationship is illustrated in FIG. 10(a) and shows that the gap size M of gap 355 is a maximum of about 2.5% of chord C at a deflection angle δ of −20°, and then the gap size M decreases to about 2.3% of chord C at a deflection angle δ of 0° and continues to decrease to about 1.2% of chord C at a deflection angle δ of 30°. This compares with a maximum gap size of about 2.9% of chord at a deflection angle δ of −20° for the reference airfoil shape DA, which decreases to about 2.4% of chord at a deflection angle δ of 0° and continues to decrease to about 1.4% of chord at a deflection angle δ of 30°.

[0161] 9 and 10(b), the secondary elements 334 are spatially positioned relative to the primary elements 332 to provide a characteristic relationship between the longitudinal overlap M2 (between the primary elements 332 and the secondary elements 334) and the deflection angle δ of the secondary elements 334 relative to the primary elements 332. This relationship is illustrated in FIG. 10(b) and shows that the longitudinal overlap M2 is a maximum of about 5.7% of chord C at a deflection angle δ of −20°, and then the longitudinal overlap M2 decreases to about 5% of chord C at a deflection angle δ of 0° and continues to decrease to about 3% of chord C at a deflection angle δ of 30°. This compares with a maximum longitudinal overlap of about 8% of chord at a deflection angle δ of −20° for the reference airfoil shape DA, which decreases to about 6% of chord at a deflection angle δ of 0° and continues to decrease to about 2% of chord at a deflection angle δ of 30°.

[0162] It is contemplated that the airfoil 300 may have certain characteristics relative to the reference airfoil shape DA, including, for example: Reduction of maximum thickness. This allows for a reduction in maximum absolute thickness and / or an increase in maximum lift in the relative forward position. Reduction of maximum camber, which reduces the maximum absolute thickness and therefore the maximum lift. A relatively large leading edge radius to maximum thickness ratio (relative to maximum thickness), which allows for increased maximum lift. Hinge location close to the bottom flap surface, which reduces the minimum maximum absolute thickness. A more forward and less pronounced pressure "belly" on the pressure surface of the airfoil, which allows a reduction in the minimum maximum absolute thickness and a slight increase in maximum lift.

[0163] As described above, the wing system 200 has a stowed configuration SC, a pre-deployed configuration PC, and a deployed configuration DC.

[0164] 1, 11(a), 12(a), and 12(b), in a stowed configuration SC, the two wings 220P, 200S are in a first substantially overlapping spatial relationship relative to one another and are retractable within an envelope EV having an envelope cross-sectional shape EP and a corresponding envelope cross-section EA, in which the leading edge of one wing 220 substantially overlaps the trailing edge of the other wing 220.

[0165] Such envelope EV is defined, at least in this example, by the space available in the fairing 800 of the wing system 200, given the cross-section of the fuselage 110. At least in this example, where the fairing 800 has a circular cross-section, the envelope EV is defined as a sector of a circle having a diameter D equal to the inner diameter of the fairing 800.

[0166] Therefore, the envelope cross-sectional shape EP includes a circular arc portion EP1 and a chord portion EP2, and an envelope cross section EA is provided by the area surrounded by the cross-sectional shape EP between the circular arc portion EP1 and the chord portion EP2.

[0167] According to one aspect of the presently disclosed subject matter, the airfoil shapes 300 of the two wings 220P, 220S are arranged with their respective minimum absolute thicknesses ABT and maximum chords C, such that in the stowed configuration SC, the respective airfoil shapes 300 of the two wings 220 in their respective substantially overlapping spatial relationship can be stored within the envelope cross-section shape EP while maximizing the occupancy of the envelope cross-section EA. In other words, according to one aspect of the presently disclosed subject matter, the airfoil shapes 300 of the two wings 220P, 220S are arranged to maximize the occupancy of the envelope cross-section EA while providing high lift characteristics, taking into account the shape of the envelope cross-section shape EP into which the two overlapping airfoils 300 are required to fit. In general, the smaller the maximum thickness t, the greater the chord C that can be provided within the limits.

[0168] In at least this example, this is achieved by designing two overlapping airfoil sections 300 with wings having a minimum maximum absolute thickness ABT and simultaneously a maximum chord C, such that the two airfoils 300 of the two wings 220P, 220S are contained within the envelope cross-sectional shape EP in the first approximately overlapping spatial relationship described above. Compared to designing an airfoil of the same chord and typical high lift characteristics, this is enhanced by reducing the maximum thickness, reducing camber, and increasing the leading edge radius to thickness ratio, but it does not have to fit within the envelope EV.

[0169] In at least this example, the airfoil shapes 300 of the two wings 220 have a total cross-sectional area that corresponds to an occupancy between 40% and 30%. In other words, the airfoil shapes 300 of the two wings 220 have a total cross-sectional area that is between 40% and 30% of the envelope cross section EA.

[0170] With particular reference to FIG. 11(a), it should be noted that a feature of the slotted, two-element airfoil nature of airfoil 300 is that the pivotability of secondary element 334 allows it to be pivoted to a large negative deflection angle δ, thereby effectively reducing the magnitude of chord C to a smaller effective chord C' in the stowed configuration SC. For example, in the example airfoil 300 of FIG. 3, chord C is reduced by 13%, providing a reduced chord C' at a negative deflection angle δ of the secondary element 334 of approximately -72°. Furthermore, particularly for the airfoil 300 of the lower left wing 220P, pivoting the secondary element 334 to such a large negative deflection angle δ also allows for greater utilization of the space available in the curved portion of arch portion EP1, as best seen in FIG. 12(a).

[0171] It should also be noted that another feature of the slotted two-element airfoil nature of the airfoil 300 is that the swiveling capability of the secondary elements 334 also allows the secondary elements 334 of each of the two superimposed airfoils 220S, 220P to be swiveled at two different negative deflection angles δ, or alternatively, the secondary elements 334 of each of the two superimposed airfoils 220S, 220P to be swiveled at two different positive deflection angles δ.

[0172] For example, in the example shown in FIG. 1 , in which the wing system 200 is mounted on the underside of the fuselage 110, the upper wing 220 (e.g., the right wing 220S) can have a deflection angle δ1 in the range of −25° to −35°, e.g., approximately −30°, e.g., −32°, while the lower wing 220 (e.g., the left wing 220P) can have a deflection angle δ2 in the range of −55° to −65°, e.g., approximately −60°. Variability in the deflection angles of the wings 220, particularly to allow for different negative deflection angles δ for the two wings 220 in the stowed configuration SC, allows for the design of the airfoil 300 with a larger chord C than would otherwise be possible (e.g., larger than would be possible with a single-element airfoil). Furthermore, this variability, along with reduced camber, reduced thickness, and reduced absolute maximum thickness, also allows for maximizing the use of envelope EV and minimizing the size of such envelope EV, e.g., compared to the baseline airfoil 300. In this example, where the wings 220 are below the fuselage 110, the space available for the lower wing 220 (in a direction parallel to the chord line CL) is less than the space available for the upper wing 220, given the curvature of the arc segment EP1. However, rather than having the chord limited by the space available for the lower wing, as in the case of a single-element wing, the subject matter of the present disclosure allows a pair of two-element airfoils 300 of relatively large chords to be mounted in a "nose-to-tail" configuration, where there may be significant overlap between the two wing sections in a direction parallel to the pivot axis PA. Such overlap is removed when the wing system 200 is transitioned to the pre-deployment position PC. The "nose-to-tail" configuration means that the leading edge of each wing 220 (or corresponding airfoil 300) is in close proximity to the trailing edge of the other wing 220 (or corresponding airfoil 300).

[0173] Also for example, in the example shown in FIG. 13 , in which the wing system 200 is mounted on the upper side of the fuselage 110, the upper wing 220 (e.g., the right wing 220S) can have a positive deflection angle δ′ ranging from +15° to +25°, e.g., approximately +20°, while the lower wing 220 (e.g., the left wing 220P) can have a deflection angle δ′ ranging from +1° to +10°, e.g., approximately +5°. Variation in the deflection angles of the wings 220, particularly to allow for different positive deflection angles δ for the two wings 220 in the stowed configuration SC, allows for the design of the airfoil 300 with a larger chord C than would otherwise be possible (e.g., larger than would be possible with a single-element airfoil). Furthermore, this variability, along with reduced camber, reduced thickness, and reduced absolute maximum thickness, also allows for maximizing the use of envelope EV and minimizing the size of such envelope EV, e.g., compared to the baseline airfoil 300. In this example, where the wings 220 are above the fuselage 110, the space available for the lower wing 220 (in a direction parallel to the chord line CL) is less than the space available for the upper wing 220, given the curvature of the arc portions of their respective envelopes. However, rather than having the chord limited by the space available for the upper wing, as in the case of a single-element wing, the subject matter of the present disclosure allows a pair of two-element airfoils 300 of relatively large chord to be mounted in a "nose-to-tail" arrangement, with significant overlap between the two wing sections in a direction parallel to the pivot axis PA. Such overlap is eliminated when the wing system 200 is transitioned to the pre-deployment configuration PC.

[0174] It should be noted that in the stowed configuration SC, the secondary elements 334 of each of the two wings can be deflected to the maximum possible extent of their respective mechanical ranges as needed to maximize envelope utilization, e.g., the ratio of chord C to inner diameter D of the fairing 800 is between 70% and 90%, between 70% and 87%, or between 75% and 85%, respectively.

[0175] For example, the ratio of the maximum thickness t to the inner diameter D of the fairing 800 may be between 7% and 12%, between 8% and 11%, or between 9% and 10%.

[0176] For example, the ratio of the minimum value of the maximum absolute thickness ABT to the inner diameter D of the fairing 800 is 7% to 12%, 8% to 14%, or 7% to 14%, respectively.

[0177] 1(a), 4, and 11(b), in the pre-deployment configuration PC, the two wings 220P, 220S are in a second, substantially overlapping spatial relationship relative to one another and can be deployed to the deployed configuration DC, in which the leading edge of one wing 220 substantially overlaps the trailing edge of the other wing 220.

[0178] In this configuration, the two superimposed airfoils 300 pivot their respective secondary elements 334 from their respective deflection angles δ corresponding to the stowed configuration SC to a particular negative deflection angle δ, so that each airfoil 300 exhibits a lateral height equal to the minimum value of the maximum absolute thickness ABT. In this and other examples, the deflection angles δ of the two airfoils are the same, for example, between -10° and -15°.

[0179] It should be noted that, at least in this example and other examples, in the stowed configuration SC, the value of the maximum absolute thickness is greater than the minimum value of the maximum absolute thickness ABT. Thus, upon transitioning the airfoils 300 from the stowed configuration SC to the pre-deployed configuration PC, the two superimposed airfoils 300 simultaneously pivot their respective secondary elements 334 such that the maximum absolute thickness in the stowed configuration SC decreases to the minimum value of the maximum absolute thickness ABT required for the pre-deployed configuration PC.

[0180] To maximize the use of the envelope EV, the two airfoils 300 are stacked as close as possible to each other at the pivot axis PA, where the corresponding first imaginary line IM1 of the lower airfoil 300 (i.e., the left wing 220P) is spaced at or near the second imaginary line IM2 of the upper airfoil 300 (i.e., the right wing 220S). For example, in the pre-deployment position PC, the deflection angle δ can be set to an angle within the range of -10° to -15°.

[0181] The transition of wing system 200 from the stowed configuration SC to the pre-deployed configuration PC may occur in response to separating aircraft 100 from fairing 800. Such separation may occur, for example, while aircraft 100 is already in flight and traveling at a significant forward speed. When fairing 800 is detached, in the example of FIG. 1 , secondary elements 334 may pivot from a large negative deflection angle δ corresponding to the stowed configuration SC to a smaller negative deflection angle δ corresponding to the pre-deployed configuration PC, thereby allowing each airfoil 300 to minimize its lateral height and equalize its maximum absolute thickness ABT. Similarly, in the example of FIG. 13 , when fairing 800 is detached, secondary elements 334 may pivot from a respective positive deflection angle δ corresponding to the stowed configuration SC to the required negative deflection angle δ corresponding to the pre-deployed configuration PC, thereby allowing each airfoil 300 to minimize its lateral height and equalize its maximum absolute thickness ABT.

[0182] In each of the two embodiments, this allows the two stacked wings 220P, 220S to pivot about the pivot axis PA from the pre-deployment position PC to the deployed position DP without colliding or interfering with each other.

[0183] It should be noted that the maximum absolute thickness-to-chord ratio of each wing 220 is such that the wings 220 can be deployed from the pre-deployment position PC to the deployment position DC without interference therebetween.

[0184] FIG. 11(b) schematically shows an intermediate position when the wing 220 moves from the pre-deployment position PC to the deployment position DC.

[0185] For example, with particular reference to FIGS. 2 and 11(b), in the deployed configuration DC, the wings 220 are fully deployed, with the leading edges of the wings facing the airflow AS. Such airflow over the wings 220 may be the result of the aircraft 100 traveling through the air at a significant forward speed. For example, the wings 220 are pivoted about an axis approximately 90° from the pre-deployment configuration PC, with the wingspan of the wings 220 substantially parallel to the longitudinal axis of the aircraft 100. In the deployed configuration DC, each wing 220 can generate an aerodynamic lift force L in the airflow AS, thereby enabling the aircraft 100 to fly in an aerodynamic flight mode. In particular, the high lift characteristics of the airfoil 300 provide the aircraft 100 with high-lift performance. For example, in deployed configuration DC, the span of wing 220 is substantially non-parallel to the longitudinal axis of aircraft 100, e.g., 90° relative to the longitudinal axis of aircraft 100 in instances where wing 220 is a zero-sweep wing, or less than 90° in instances where wing 220 is a swept wing.

[0186] It should be noted that at the beginning of deployment DC, and thus at the end of the transition from pre-deployment PC to deployment DC, the secondary element 334 of each of the two wings 220 at deployment DC is at a relatively large negative deflection angle δ, e.g., −10° to −15°. Referring to FIG. 14 , this corresponds to a lower lift coefficient than a zero deflection angle δ, which corresponds to, for example, a cruise condition. This feature reduces the lift generated by the wing 220 when the wing is suddenly exposed to the airflow, resulting in relatively lower loading on the wing compared to providing the wing 220 at the required deflection angle δ (e.g., zero) immediately after the beginning of deployment DC. This feature, unique to aspects of the subject matter of the present disclosure, allows the wing to be designed to be lighter than it would otherwise be, due to the reduced expected impact loads upon deployment. Furthermore, this feature can be taken into account in the design process, allowing the airfoil 300 to be designed with a larger negative deflection angle δ already in the pre-deployment configuration, further reducing the initial impact loads at deployment DC. Thus, at least in this example, the secondary element 334 can be maintained at a large negative deflection angle δ, e.g., −10° to −15°, until fully deployed at the deployed configuration DC and after the initial shock load from the airflow. The secondary element 334 can be pivoted to zero or a positive deflection angle δ depending on the lift or performance requirements of the mission, such as cruise or takeoff.

[0187] It should be noted that in alternative variations of this example and in other examples, the fairing may have an internal geometry different from a segment of a circle. For example, the corresponding cross-sectional shape EP may be, for example, a polygon (e.g., rectangular, square, hexagonal, etc.), or a portion of an ellipse, or a portion of a superellipse. As with the illustrated example, mutatis mutandis, the airfoil shapes 300 of each of the two wings 220P, 220S are similarly positioned to maximize the occupancy of the respective envelope cross-section of the envelope, taking into account the envelope cross-sectional shape of the envelope into which the two overlapping airfoils 300 must fit.

[0188] It should also be noted that in still other variations of this example and in other examples, the fairings may be omitted, and instead each envelope may use a corresponding cross-sectional shape EP associated with a container in which aircraft 100 may be stored for transport or safety until required for use, in which aircraft 100 may be stored in a respective storage configuration and, once removed from the container, may be transferred to a respective pre-deployment configuration, for example, by ground crew.

[0189] It is further noted that in at least one example of the presently disclosed subject matter, in the stowed configuration, the corresponding cross-sectional shape EP may be limited by various factors, such as the required fuselage cross-section, and the presently disclosed subject matter makes it possible to provide a two-element airfoil wing that can maximize the use of this limited space.

[0190] In the method claims that follow, alphanumeric and roman numerals used to designate claim steps are provided for convenience only and do not imply any particular order for performing the steps.

[0191] For example, and finally, it should be noted that the word "including" as used throughout the appended claims should be interpreted to mean "including, but not limited to."

[0192] For example, while examples have been shown and disclosed in accordance with the subject matter of this disclosure, it should be understood that many changes can be made without departing from the scope of the subject matter of this disclosure, as claimed.

Claims

1. 1. A wing system for an aircraft, the wing system having a stowed configuration, a pre-deployed configuration, and a deployed configuration; two wings, each having an airfoil shape and deployable to pivot about a respective pivot axis between the pre-deployed configuration and the deployed configuration, wherein: In the stowed configuration, the two wings are in a first substantially overlapping spatial relationship with respect to one another and are retractable within an envelope having an envelope cross-sectional shape and a corresponding envelope cross-section; In the pre-deployment configuration, the two wings are in a second substantially overlapping spatial relationship with respect to each other and can be deployed to the deployed configuration without interfering with each other; and In the deployed configuration, each of the wings is capable of generating aerodynamic lift in an airflow; each airfoil shape of each wing is a slotted airfoil having a primary element, a secondary element, and a chord, said secondary element being pivotable relative to said primary element and spaced apart by a gap, each said airfoil shape having a respective maximum thickness and a respective maximum absolute thickness; wherein in the stowed configuration, each secondary element of each airfoil of one of the wings is set at a different flap angle compared to each secondary element of each airfoil of the other of the wings; the wing system is arranged to be mounted under an aircraft fuselage, and in the stowed configuration, each secondary element of the airfoil shape of each wing is set to a respective negative flap angle, and in the stowed configuration, each secondary element of the airfoil shape of each wing is set to a respective second flap angle, whereby each trailing edge and each hinge axis of each secondary element is on opposite sides of each chord line of the airfoil shape; or the wing system is arranged to be mounted under an aircraft fuselage, and in the stowed configuration, each secondary element of the airfoil shape of an uppermost wing is set to a respective negative flap angle in the range of −25° to −35°, and each secondary element of the airfoil shape of a lowermost wing is set to a respective negative flap angle in the range of −55° to −65°, and in the stowed configuration, each secondary element of each airfoil shape of each wing is set to a respective second flap angle, whereby each trailing edge and each hinge axis of each secondary element is on opposite sides of each chord line of the airfoil shape; or the wing system is arranged to be attached to an underside of an aircraft fuselage, and in the stowed configuration, each secondary element of the airfoil shape of each wing is set to a respective negative flap angle, and in the stowed configuration, each secondary element of the airfoil shape of an uppermost wing is set to a respective negative flap angle within a range of -25° to -35° and each secondary element of the airfoil shape of a lowermost wing is set to a respective negative flap angle within a range of -55° to -65°, and in the stowed configuration, each secondary element of the airfoil shape of each wing is set to a respective second flap angle, whereby each trailing edge and each hinge axis of each secondary element are on opposite sides of each chord line of the airfoil shape. Wing systems for aircraft.

2. 10. The wing system of claim 1, comprising one of the following: - a wing system, wherein the two wings comprise a first wing having a first airfoil shape, the first airfoil shape comprising a first leading edge and a first trailing edge, and a second wing having a second airfoil shape, the second airfoil shape comprising a second leading edge and a second trailing edge; and a wing system, wherein the two wings comprise a first wing having a first airfoil shape, the first airfoil shape comprising a first leading edge and a first trailing edge, and a second wing having a second airfoil shape, the second airfoil shape comprising a second leading edge and a second trailing edge, and wherein in the pre-deployed configuration, the first leading edge is substantially overlapped with the second trailing edge.

3. 3. The wing system of claim 1, wherein the ratio of the respective maximum absolute thickness to the chord of each wing is such that the wings can be deployed from the pre-deployed configuration to the deployed configuration without interfering with each other.

4. A wing system according to any one of claims 1 to 3, comprising one of the following: - a wing system wherein each secondary element of each airfoil shape of each said wing has a hinge axis for enabling pivoting of each secondary element relative to its respective primary element; - a wing system in which each secondary element of each airfoil shape of each said wing has a hinge axis for enabling pivoting of each secondary element relative to a respective primary element, said hinge axis facing a pressure surface of each secondary element; - a wing system wherein each secondary element of each airfoil shape of each said wing has a hinge axis for permitting pivoting of the respective secondary element relative to the respective primary element, and wherein for each said airfoil shape, the respective hinge axis is spaced from the pressure surface of the respective secondary element; - a wing system in which each secondary element of each airfoil shape of each said wing has a hinge axis for enabling pivoting of the respective secondary element relative to the respective primary element, the hinge axis facing the pressure surface of the respective secondary element, and for each said airfoil shape, the respective hinge axis being spaced from the pressure surface of the respective secondary element; - a wing system wherein each secondary element of each airfoil shape of each said wing has a hinge axis for permitting pivoting of the respective secondary element relative to the respective primary element, and wherein for each said airfoil shape, the respective hinge axis is spaced from the pressure surface of the respective secondary element, and the spacing is less than 50% of the maximum thickness of the secondary element; - each secondary element of each airfoil shape of each said wing has a hinge axis to enable pivoting of the respective secondary element relative to the respective primary element, the hinge axis facing the pressure surface of the respective secondary element, and for each said airfoil shape, the respective hinge axis is spaced from the pressure surface of the respective secondary element, the spacing being less than 50% of the maximum thickness of the secondary element.

5. A wing system according to any one of claims 1 to 4, comprising one of the following: - a wing system wherein, in the pre-deployment configuration, a respective secondary element of each airfoil shape of each of the wings is set to a respective negative flap angle; - a wing system in which, in the pre-deployment configuration, each secondary element of each airfoil shape of each wing is set to a respective negative flap angle, and in the pre-deployment configuration, each secondary element of each airfoil shape of each wing is set to the same flap angle; a wing system in which, in the pre-deployment configuration, a respective secondary element of each airfoil shape of each of the wings is set to a respective negative flap angle, and in the pre-deployment configuration, a respective secondary element of each airfoil shape of each of the wings is set to a respective first flap angle, whereby no portion of a respective second flap is tangent to a leading edge of a respective airfoil shape and intersects a respective hinge axis.

6. 6. The wing system of claim 1, wherein each airfoil shape is arranged at a minimum of its respective maximum absolute thickness and a maximum of its chord, such that in the stowed configuration, the airfoil shapes of each of the two wings in the first substantially overlapping spatial relationship can be stowed within the envelope cross-sectional shape while simultaneously maximizing the occupancy of the envelope cross-section.

7. A wing system according to any one of claims 1 to 6, comprising one of the following: - a wing system, the wing system being arranged to be attached to an upper side of an aircraft fuselage; - the wing system is arranged to be mounted above an aircraft fuselage, and wherein in the stowed configuration, a respective secondary element of an airfoil shape of each wing is set to a respective positive flap angle; - the wing system is arranged to be mounted above an aircraft fuselage, and in the stowed configuration, each secondary element of the airfoil shape of the uppermost wing is set to a respective positive flap angle in a range of +15° to +25°, and each secondary element of the airfoil shape of the lowermost wing is set to a respective positive flap angle in a range of +1° to +10°; the wing system is arranged to be mounted above an aircraft fuselage, and in the stowed configuration, each secondary element of the airfoil shape of each wing is set to a respective positive flap angle, and in the stowed configuration, each secondary element of the airfoil shape of an uppermost wing is set to a respective positive flap angle in the range of +15° to +25° and each secondary element of the airfoil shape of a lowermost wing is set to a respective positive flap angle in the range of +1° to +10°.

8. A wing system according to any one of claims 1 to 7, comprising one of the following: - a wing system wherein, for each of said airfoil shapes, said respective maximum thickness-to-chord ratio is in the range of 0.10 to 0.14; - for each said airfoil shape, the respective said maximum absolute thickness to chord ratio is in the range of 0.10 to 0.

15.

9. 9. The wing system of claim 1, wherein each airfoil shape has its respective maximum thickness at a location less than 25% of the chord from the leading edge of the airfoil shape.

10. 10. The wing system of claim 1, wherein each airfoil shape has a camber shape with a maximum camber of about 6% or less of the chord.

11. 11. The wing system of claim 1, wherein for each airfoil shape, the respective leading edge radius is about 2% of the chord.

12. for each said airfoil shape, each said primary element has a primary element trailing edge and each said secondary element has a secondary element leading edge; the primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord; the primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord, and each longitudinal overlap varies between about 5% and about 3% of the chord as each flap angle increases from 0° to 30°; and 12. The wing system of claim 1, wherein the primary element trailing edge overlaps the secondary element leading edge with a longitudinal overlap in a direction parallel to the chord, and wherein for each airfoil shape, the respective longitudinal overlap varies between about 5% and about 5.5% of the chord as the respective flap angle decreases from 0° to -20°.

13. A wing system according to any one of claims 1 to 12, comprising one of the following: - a wing system, wherein each of the airfoil shapes is arranged as a high-lift airfoil shape; - a wing system wherein each said airfoil shape is arranged as a high-lift airfoil shape and wherein said airfoil shapes are arranged to have a maximum lift coefficient of at least 2.5 at a flap angle of zero; - a wing system wherein each said airfoil shape is arranged as a high-lift airfoil shape, and wherein said airfoil shapes are arranged to have a maximum lift coefficient greater than 2.3 at zero flap angle; - a wing system wherein each said airfoil shape is arranged as a high-lift airfoil shape, and wherein said airfoil shapes are arranged to have a maximum lift coefficient of at least 2.5 at a flap angle of zero, said maximum lift coefficient corresponding to an angle of attack of about 15° to about 16°; a wing system wherein each of the airfoil shapes is arranged as a high-lift airfoil shape, and wherein the airfoil shapes are arranged to have a maximum lift coefficient greater than 2.3 at a flap angle of zero, and wherein the maximum lift coefficient corresponds to an angle of attack of about 15° to about 16°.

14. A wing system according to any one of claims 1 to 13, wherein the airfoil shapes of the two wings have a total cross-sectional area corresponding to an occupancy rate of 40% to 30% of the envelope cross-sectional area.

15. A wing system according to any one of claims 1 to 14, comprising one of the following: - a wing system wherein the envelope cross-sectional shape is in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion; - the envelope cross-sectional shape is in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion, and for each of the airfoil shapes, a respective first ratio of chord to diameter is in the range of 70% to 87%; - a wing system wherein the envelope cross-sectional shape is in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion, and wherein for each of the airfoil shapes, the respective first ratio of chord to diameter is in the range of 70% to 87%, and for each of the airfoil shapes, the respective first ratio is in the range of 75% to 85%; - the envelope cross-sectional shape is in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion, and including at least one of: for each airfoil shape, the respective second ratio of maximum thickness to diameter is in the range of 7% to 12%; and for each airfoil shape, the respective second ratio of maximum absolute thickness to diameter is in the range of 8% to 14%; - the envelope cross-sectional shape is in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion, and the wing system including at least one of: for each of the airfoil shapes, a respective first ratio of the chord to the diameter is in a range of 70% to 87%, and for each of the airfoil shapes, a respective second ratio of the maximum thickness to the diameter is in a range of 7% to 12% and for each of the airfoil shapes, a respective second ratio of the maximum absolute thickness to the diameter is in a range of 8% to 14%; the envelope cross-sectional shape is in the form of a portion of a circle having a defined diameter, the portion being defined by a corresponding arc portion and a corresponding chord portion, and wherein for each of the airfoil shapes, a respective first ratio of chord to diameter is in the range of 70% to 87%, and for each of the airfoil shapes, the respective first ratio is in the range of 75% to 85%, and at least one of for each of the airfoil shapes, a respective second ratio of maximum thickness to diameter is in the range of 7% to 12%, and for each of the airfoil shapes, a respective second ratio of maximum absolute thickness to diameter is in the range of 8% to 14%.

16. 16. The wing system of claim 1, wherein the wing system is arranged to maintain a negative flap angle for each airfoil of each wing during deployment from the pre-deployed configuration to the deployed configuration.

17. An aircraft comprising a wing system as defined in any one of claims 1 to 16.

18. 20. A launch system comprising an aircraft as defined in claim 17 housed within a fairing having an internal chamber defining said envelope, said envelope cross-sectional shape, and said envelope cross-section.

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