An airfoil assembly for a variant aircraft

Abstract

The application provides a wing profile combination for a variable aircraft, comprising a transonic flight advantage wing profile, a supersonic flight advantage wing profile and a hypersonic flight advantage wing profile. The three wing profiles have different flight advantages, and the lift-drag characteristics of each cruise flight state are better than those of a reference wing profile NACA64A-204. The lift-drag ratio characteristics and the lift coefficient at different angles of attack in each cruise flight state are improved, and the three wing profiles can be changed to each other through a wing profile deformation mechanism, so that the flight performance can be optimized in different speed domains according to different flight states. Therefore, the wing profile combination for the variable aircraft provided by the application can be used for trans-speed domain flight of the variable aircraft, and is a wing profile variable solution for trans-speed domain flight of the variable aircraft.

Description

Technical Field

[0001] This invention belongs to the field of aerodynamics technology, specifically relating to an airfoil combination for morphing aircraft. Background Technology

[0002] In both military and civilian fields, the demand for multi-mission aircraft is growing, especially those capable of multi-speed cruise. Traditional aircraft do not always possess optimal aerodynamic characteristics across their entire flight speed range. Airfoil selection and design are crucial aspects of aircraft design, influencing cruise speed, takeoff and landing performance, stall speed, handling performance (especially near stall), and aerodynamic efficiency across all flight phases. Aerodynamic design experience shows that for the same configuration, the flow mechanisms for lift enhancement and drag reduction differ across speed ranges, resulting in significantly different airfoil requirements. For low-speed aircraft, relatively thicker airfoils are preferred; for supersonic aircraft, relatively thinner quadrilateral, hexagonal, or double-arc airfoils can be used. Therefore, the different cruise speeds of aircraft create a conflict in the requirements for airfoil shapes at high and low speeds. Summary of the Invention

[0003] To address the shortcomings of existing technologies, this invention provides an airfoil combination for morphing aircraft, which can effectively solve the aforementioned problems.

[0004] The technical solution adopted in this invention is as follows:

[0005] This invention provides an airfoil combination for a morphing aircraft, the airfoil combination including a transonic flight-dominant airfoil, a supersonic flight-dominant airfoil, and a hypersonic flight-dominant airfoil;

[0006] The transonic flight-advantage airfoil has a maximum thickness of 4.0%C, a maximum thickness location of 42.5%C, a maximum camber of 1.63%C, and a maximum camber location of 66.3%C.

[0007] The supersonic flight-advantage airfoil has a maximum thickness of 4%C, a maximum thickness location of 49.4%C, a maximum camber of 0.83%C, and a maximum camber location of 74.5%C.

[0008] The hypersonic flight-advantage airfoil has a maximum thickness of 4%C, a maximum thickness location of 53.4%C, a maximum camber of 1.12%C, and a maximum camber location of 73.9%C; where C is the airfoil chord length.

[0009] Preferably, the geometric coordinate expressions for the upper and lower surfaces of the airfoil in the airfoil assembly are as follows:

[0010]

[0011]

[0012] in:

[0013] y up (x) represents the ordinate of the upper surface of the airfoil;

[0014] y low (x) represents the ordinate of the lower surface of the airfoil;

[0015] A upi The coefficients of the expression representing the geometric coordinates of the upper surface of the airfoil; i = 0, 1, 2, ..., 8, representing a total of 9 coefficients;

[0016] A lowi The coefficients of the expression representing the geometric coordinates of the lower surface of the airfoil;

[0017] x represents the x-coordinate of the surface of a unit airfoil.

[0018] The preferred geometric coordinate expression coefficients for the transonic flight-dominant airfoil are:

[0019] A up0 ]] A up1 ]] A up2 ]]> A up3 ]] A up4 ]] A up5 ]] A up6 ]] <![CDATA[A up7 ]]> <![CDATA[A up8 ]]> 0.04374 0.06455 0.07530 0.50443 0.15629 0.01812 0.20727 0.12926 0.12797 <![CDATA[A low0 ]]> <![CDATA[A low1 ]]> <![CDATA[A low2 ]]> <![CDATA[A low3 ]]> <![CDATA[A low4 ]]> <![CDATA[A low5 ]]> <![CDATA[A low6 ]]> <![CDATA[A low7 ]]> <![CDATA[A low8 ]]> -0.02256 -0.01800 -0.03036 -0.01166 -0.04051 0.01478 -0.01151 0.05886 -0.01185

[0020] The coefficients of the geometric coordinate expression for the supersonic flight-dominant airfoil are:

[0021] <![CDATA[A up0 ]]> <![CDATA[A up1 ]]> <![CDATA[A up2 ]]> <![CDATA[A up3 ]]> <![CDATA[A up4 ]]> <![CDATA[A up5 ]]> <![CDATA[A up6 ]]> <![CDATA[A up7 ]]> <![CDATA[A up8 ]]> 0.01650 0.02955 0.06200 0.05185 0.09169 0.06921 0.11533 0.05914 0.15379 <![CDATA[A low0 ]]> <![CDATA[A low1 ]]> <![CDATA[A low2 ]]> <![CDATA[A lww3 ]]> <![CDATA[A low4 ]]> <![CDATA[A low5 ]]> <![CDATA[A low6 ]]> <![CDATA[A low7 ]]> <![CDATA[A lww8 ]]> -0.00395 -0.00280 -0.00980 -0.00663 -0.10912 0.01217 -0.08213 0.07150 -0.00690

[0022] The coefficients of the geometric coordinate expression for the dominant airfoil for hypersonic flight are:

[0023] <![CDATA[A up0 ]]> <![CDATA[A up1 ]]> <![CDATA[A up2 ]]> <![CDATA[A up3 ]]> <![CDATA[A up4 ]]> <![CDATA[A up5 ]]> <![CDATA[A up6 ]]> <![CDATA[A up7 ]]> <![CDATA[A up8 ]]> 0.01363 0.03596 0.05054 0.03718 0.10299 0.05475 0.17636 0.07315 0.16849 <![CDATA[A low0 ]]> <![CDATA[A low1 ]]> <![CDATA[A low2 ]]> <![CDATA[A low3 ]]> <![CDATA[A low4 ]]> <![CDATA[A low5 ]]> <![CDATA[A low6 ]]> <![CDATA[A low7 ]]> <![CDATA[A low8 ]]> -0.00530 -0.00198 -0.00808 -0.00677 -0.09657 0.00608 -0.08054 0.09001 -0.00685

[0024] in:

[0025] A up0 A up1 A up2 A up3 A up4 A up5 A up6 A up7 A up8 , representing the 0th, 1st, 2nd, 3rd, 4th, 5th, 6th, 7th, and 8th expression coefficients, respectively.

[0026] Preferably, the three airfoils included in the airfoil combination are adjusted to change with each other through an airfoil deformation mechanism.

[0027] Preferably, the airfoil deformation mechanism includes a flexible skin (5), an I-beam (6), and a telescopic link (7);

[0028] The flexible skin (5) has an airfoil shape. The I-beam (6) is installed inside the flexible skin (5). The I-beam (6) is movable, thereby changing the maximum thickness position of the airfoil. One end of the telescopic link (7) is connected to the flexible skin (5), and the other end of the telescopic link (7) is connected to the I-beam (6). The telescopic link (7) moves in extension and retraction under the action of the actuator, thereby causing the flexible skin (5) to deform and change the shape of the airfoil represented by the flexible skin (5).

[0029] The airfoil combination for a morphing aircraft provided by this invention has the following advantages:

[0030] This invention provides an airfoil combination for morphing aircraft, comprising a transonic flight-dominant airfoil, a supersonic flight-dominant airfoil, and a hypersonic flight-dominant airfoil. These three airfoils, each with different flight advantages, exhibit superior lift-drag characteristics compared to the baseline airfoil NACA64A-204 in their respective cruise flight states. Furthermore, they show improvements in lift-drag ratio and lift coefficient at different angles of attack during their respective cruise flight states. Moreover, the three airfoils can be interchanged via an airfoil deformation mechanism. Therefore, this invention provides an airfoil combination for morphing aircraft, suitable for multi-speed-range flight. Attached Figure Description

[0031] Figure 1 A comparison of the geometric shapes of the airfoil combination designed for this invention and the NACA64A-204 airfoil;

[0032] Figure 2 This is a schematic diagram of the airfoil deformation mechanism;

[0033] Figure 3 This is a comparison diagram of the pressure distribution curves of the transonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil under transonic design conditions;

[0034] Figure 4 A comparison diagram of the lift curves of the transonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil under transonic conditions;

[0035] Figure 5 A comparison of lift-drag characteristic curves of the transonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil under transonic conditions;

[0036] Figure 6 A comparison of lift-to-drag ratio curves of the transonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil at transonic speeds;

[0037] Figure 7 A comparison of the moment characteristic curves of the transonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil under transonic conditions;

[0038] Figure 8 This is a comparison diagram of the pressure distribution curves of the supersonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil under supersonic design conditions.

[0039] Figure 9 A comparison diagram of the lift curves of the supersonic flight-advantage airfoil of the present invention and the NACA64A-204 airfoil under supersonic conditions;

[0040] Figure 10 A comparison of lift-drag characteristic curves of the supersonic flight-advantage airfoil of the present invention and the NACA64A-204 airfoil under supersonic conditions;

[0041] Figure 11 This is a comparison of the lift-to-drag ratio curves of the supersonic flight-advantage airfoil of the present invention and the NACA64A-204 airfoil at supersonic speeds.

[0042] Figure 12 A comparison of the moment characteristic curves of the supersonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil under supersonic conditions;

[0043] Figure 13 This is a comparison diagram of the pressure distribution curves of the hypersonic flight advantage airfoil of the present invention and the NACA64A-204 airfoil under hypersonic design conditions.

[0044] Figure 14 A comparison diagram of the lift curves of the hypersonic flight-advantage airfoil of the present invention and the NACA64A-204 airfoil in hypersonic conditions;

[0045] Figure 15 This is a comparison of the lift-drag characteristics of the hypersonic flight-advantage airfoil of the present invention and the NACA64A-204 airfoil at hypersonic speeds.

[0046] Figure 16 This is a comparison chart of the lift-to-drag ratio curves of the hypersonic flight-advantage airfoil of the present invention and the NACA64A-204 airfoil at hypersonic speeds.

[0047] Figure 17 A comparison of the moment characteristic curves of the hypersonic flight-advantage airfoil of the present invention and the NACA64A-204 airfoil at hypersonic speeds.

[0048] in:

[0049] 1 represents the aerodynamic characteristic curve of the transonic flight advantage airfoil of the present invention;

[0050] 2 represents the aerodynamic characteristic curve of the supersonic flight advantage airfoil of the present invention;

[0051] 3 represents the aerodynamic characteristic curve of the hypersonic flight advantage airfoil of the present invention;

[0052] 4 represents the aerodynamic characteristic curve of the NACA64A-204 airfoil used for comparison;

[0053] The components are named as follows: 5. Flexible skin, 6. I-beam, and 7. Telescopic connecting rod. Detailed Implementation

[0054] To make the technical problems solved, the technical solutions, and the beneficial effects of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and are not intended to limit the invention.

[0055] This invention provides an airfoil combination for morphing aircraft, comprising a transonic flight-dominant airfoil, a supersonic flight-dominant airfoil, and a hypersonic flight-dominant airfoil. These three airfoils, each with different flight dominance, exhibit superior lift-drag characteristics compared to the baseline airfoil NACA64A-204 in their respective cruise flight states. Furthermore, they offer improvements in lift-drag ratio and lift coefficient at different angles of attack during their respective cruise flight states. Moreover, the airfoil combination can be adapted to each other via an airfoil deformation mechanism. Therefore, this invention provides an airfoil combination for morphing aircraft, suitable for multi-speed-range flight.

[0056] Specifically, this invention addresses the conflicting requirements for airfoil shapes at high and low speeds due to varying aircraft cruise speeds. Based on the NACA64A-204 baseline airfoil, it designs an airfoil combination for variant aircraft, comprising a transonic flight-dominant airfoil, a supersonic flight-dominant airfoil, and a hypersonic flight-dominant airfoil. The airfoil design specifications are as follows:

[0057] 1. For transonic flight-advantageous airfoils: H = 9 km, Ma = 0.8, Re = 7.61 × 10⁻⁶ 6 Alpha = 1.5°, with a lift-to-drag ratio of not less than 80 and a lift coefficient of not less than 0.6 at the transonic design angle of attack; where H represents cruise altitude, Ma represents Mach number, Re represents Reynolds number, and Alpha represents angle of attack.

[0058] 2. For supersonic flight-advantageous airfoils: H = 10 km, Ma = 2, Re = 1.70 × 10⁻⁶ 7Alpha = 4°, which provides a high lift-to-drag ratio under the premise of improved lift coefficient at the supersonic design angle of attack;

[0059] 3. For hypersonic flight, the dominant airfoil is: H = 26 km, Ma = 6, Re = 4.23 × 10⁻⁶ km. 6 Alpha = 5°, which provides a high lift-to-drag ratio under the premise of improved lift coefficient at hypersonic design angle of attack;

[0060] 4. Airfoil thickness is in the range of 4%.

[0061] This invention designs an airfoil combination for a morphing aircraft, compared with the NACA64A-204 reference airfoil geometry, for example... Figure 1 As shown in the table below. The transonic airfoil has a maximum thickness of 4.0%C, a maximum thickness location of 42.5%C, a maximum camber of 1.63%C, and a maximum camber location of 66.3%C; the supersonic airfoil has a maximum thickness of 4%C, a maximum thickness location of 49.4%C, a maximum camber of 0.83%C, and a maximum camber location of 74.5%C; the hypersonic airfoil has a maximum thickness of 4%C, a maximum thickness location of 53.4%C, a maximum camber of 1.12%C, and a maximum camber location of 73.9%C. Specific geometric parameters are shown in the table below. Where C is the airfoil chord length.

[0062] Airfoil name Maximum thickness Maximum thickness location Maximum curvature Location of maximum curvature Supersonic flight advantage airfoils 4％C 42.5％C 1.63％C 66.3％C Supersonic flight advantage airfoil 4％C 49.4％C 0.83％C 74.5％C Hypersonic flight advantage airfoils 4％C 53.4％C 1.12％C 73.9％C

[0063] The geometric coordinate expressions for the upper and lower surfaces of the airfoil are:

[0064]

[0065]

[0066] in:

[0067] y up (x) represents the ordinate of the upper surface of the airfoil;

[0068] y low (x) represents the ordinate of the lower surface of the airfoil;

[0069] A upi The coefficients of the expression representing the geometric coordinates of the upper surface of the airfoil; i = 0, 1, 2, ..., 8, representing a total of 9 coefficients;

[0070] A lowi The coefficients of the expression representing the geometric coordinates of the lower surface of the airfoil;

[0071] x represents the x-coordinate of the surface of a unit airfoil.

[0072] The coefficients of the expression for the geometric coordinates of the dominant airfoil for transonic flight in this invention are:

[0073] <![CDATA[A up0 ]]> <![CDATA[A up1 ]]> <![CDATA[A up2 ]]> <![CDATA[A up3 ]]> <![CDATA[A up4 ]]> <![CDATA[A up5 ]]> <![CDATA[A up6 ]]> <![CDATA[A up7 ]]> <![CDATA[A up8 ]]> 0.04374 0.06455 0.07530 0.50443 0.15629 0.01812 0.20727 0.12926 0.12797 <![CDATA[A low0 ]]> <![CDATA[A low1 ]]> <![CDATA[A low2 ]]> <![CDATA[A lww3 ]]> <![CDATA[A low4 ]]> <![CDATA[A low5 ]]> <![CDATA[A low6 ]]> <![CDATA[A low7 ]]> <![CDATA[A lww8 ]]> -0.02256 -0.01800 -0.03036 -0.01166 -0.04051 0.01478 -0.01151 0.05886 -0.01185

[0074] The coefficients of the geometric coordinate expression for the supersonic flight-dominant airfoil of this invention are:

[0075] <![CDATA[A up0 ]]> <![CDATA[A up1 ]]> <![CDATA[A up2 ]]> <![CDATA[A up3 ]]> <![CDATA[A up4 ]]> <![CDATA[A up5 ]]> <![CDATA[A up6 ]]> <![CDATA[A up7 ]]> <![CDATA[A up8 ]]> 0.01650 0.02955 0.06200 0.05185 0.09169 0.06921 0.11533 0.05914 0.15379 <![CDATA[A low0 ]]> <![CDATA[A low1 ]]> <![CDATA[A low2 ]]> <![CDATA[A low3 ]]> <![CDATA[A low4 ]]> <![CDATA[A low5 ]]> <![CDATA[A low6 ]]> <![CDATA[A low7 ]]> <![CDATA[A low8 ]]> -0.00395 -0.00280 -0.00980 -0.00663 -0.10912 0.01217 -0.08213 0.07150 -0.00690

[0076] The coefficients of the geometric coordinate expression for the hypersonic flight dominant airfoil of this invention are:

[0077] <![CDATA[A up0 ]]> <![CDATA[A up1 ]]> <![CDATA[A up2 ]]> <![CDATA[A up3 ]]> <![CDATA[A up4 ]]> <![CDATA[A up5 ]]> <![CDATA[A up6 ]]> <![CDATA[A up7 ]]> <![CDATA[A up8 ]]> 0.01363 0.03596 0.05054 0.03718 0.10299 0.05475 0.17636 0.07315 0.16849 <![CDATA[A low0 ]]> <![CDATA[A low1 ]]> <![CDATA[A low2 ]]> <![CDATA[A low3 ]]> <![CDATA[A low4 ]]> <![CDATA[A low5 ]]> <![CDATA[A low6 ]]> <![CDATA[A low7 ]]> <![CDATA[A low8 ]]> -0.00530 -0.00198 -0.00808 -0.00677 -0.09657 0.00608 -0.08054 0.09001 -0.00685

[0078] in:

[0079] A up0 A up1 A up2 A up3 A up4 A up5 A up6 A up7 A up8 , representing the 0th, 1st, 2nd, 3rd, 4th, 5th, 6th, 7th, and 8th expression coefficients, respectively.

[0080] The airfoil combination provided by this invention allows for mutual adjustment and transformation through an airfoil deformation mechanism. Therefore, during aircraft flight, the airfoil shape can be adjusted in real time according to the flight speed range, thereby obtaining an airfoil that conforms to the current flight speed range. This invention does not limit the specific structural form of the airfoil deformation mechanism; as a specific embodiment, such as... Figure 2 The diagram shown is a schematic of an airfoil deformation mechanism, which can be understood as a cross-sectional view of an airfoil. This mechanism is used to transform between different airfoil shapes, and its components include: flexible skin 5, I-beam 6, and telescopic link 7.

[0081] The flexible skin 5 has an airfoil shape, and the I-beam 6 is installed inside the flexible skin 5. The I-beam 6 is movable, thereby changing the maximum thickness position of the airfoil. One end of the telescopic link 7 is connected to the flexible skin 5, and the other end of the telescopic link 7 is connected to the I-beam 6. The telescopic link 7 moves in extension and retraction under the action of the actuator, thereby causing the flexible skin 5 to deform and change parameters such as the camber of the airfoil, so that the shape of the airfoil represented by the flexible skin 5 changes.

[0082] The following experiments will illustrate the differences in aerodynamic characteristics between the transonic, supersonic, and hypersonic airfoils designed in this invention and the baseline airfoil:

[0083] Tables 1, 2, and 3 show the main aerodynamic characteristics of the transonic flight dominant airfoil and the reference airfoil NACA64A-204 at the design point, respectively; the main aerodynamic characteristics of the supersonic flight dominant airfoil and the reference airfoil NACA64A-204 at the design point; and the main aerodynamic characteristics of the hypersonic flight dominant airfoil and the reference airfoil NACA64A-204 at the design point.

[0084] Table 1. Main aerodynamic characteristics of the transonic flight dominant airfoil and the reference airfoil NACA 64A-204 at the design point: (H = 9 km, Ma = 0.8, Re = 7.61 × 10⁻⁶) 6 (Alpha = 1.5°)

[0085] airfoil Lift coefficient drag coefficient torque coefficient Rise-to-drag ratio NACA64A-204 0.597 0.00840 -0.0663 71.1 Supersonic flight advantage airfoils 0.640 0.00779 -0.0865 82.2

[0086] Table 2. Main aerodynamic characteristics of the supersonic flight dominant airfoil and the reference airfoil NACA 64A-204 at the design point: (H = 10 km, Ma = 2, Re = 1.70 × 10⁻⁶) 7 Alpha = 4°

[0087] airfoil Lift coefficient drag coefficient torque coefficient Rise-to-drag ratio NACA64A-204 0.1518 0.02713 0.0172 5.60 Supersonic flight advantage airfoil 0.1579 0.02042 0.0091 7.73

[0088] Table 3. Main aerodynamic characteristics of the dominant and reference airfoils for hypersonic flight, NACA 64A-204, at the design point: (H = 26 km, Ma = 6, Re = 4.23 × 10⁻⁶) 6 (Alpha = 5°)

[0089] airfoil Lift coefficient drag coefficient torque coefficient Rise-to-drag ratio NACA64A-204 0.0471 0.01397 0.0031 3.36 Hypersonic flight advantage airfoils 0.0575 0.00834 -0.0009 6.89

[0090] Figure 1 A comparison diagram of the geometric shapes of the airfoil combination designed for this invention and the NACA64A-204 reference airfoil.

[0091] See Figures 3 to 7 Table 1 shows that the transonic flight advantage airfoil of the present invention improves the lift coefficient at different angles of attack in transonic conditions, and the stall angle of attack is slightly improved. At a design angle of attack of 1.5°, the lift coefficient is 0.640 and the lift-to-drag ratio is 82.2, which are better than the lift coefficient of 0.597 and the lift-to-drag ratio of 71.1 of the reference airfoil. Furthermore, it meets the design requirements of a lift coefficient greater than 0.6 and a lift-to-drag ratio greater than 80.

[0092] See Figures 8 to 12According to Table 2, the lift coefficient of the supersonic flight advantage airfoil of the present invention is improved at different angles of attack in supersonic state. At a design angle of attack of 4°, the lift coefficient is improved from 0.1518 of the reference airfoil to 0.1579, and the lift-to-drag ratio is improved from 5.60 of the reference airfoil to 7.73, thus improving the lift-to-drag ratio that meets the design specifications.

[0093] See Figures 13 to 17 As shown in Table 3, the lift coefficient of the hypersonic flight advantage airfoil of the present invention is improved at different angles of attack in hypersonic conditions. At a design angle of attack of 5°, the lift coefficient is improved from 0.0471 of the reference airfoil to 0.0575, and the lift-to-drag ratio is improved from 3.36 of the reference airfoil to 6.89, thus improving the lift-to-drag ratio that meets the design specifications.

[0094] Therefore, the three airfoils designed in this invention improve the aerodynamic performance at transonic, supersonic, and hypersonic speeds compared to the baseline airfoil NACA64A-204. The aircraft can change the airfoil shape during flight through the airfoil deformation mechanism to adapt to flight at different speeds.

[0095] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. An airfoil combination for a morphing aircraft, characterized in that, The airfoil combinations include transonic flight-dominant airfoils, supersonic flight-dominant airfoils, and hypersonic flight-dominant airfoils; The transonic flight-advantage airfoil has a maximum thickness of 4.0%C, a maximum thickness location of 42.5%C, a maximum camber of 1.63%C, and a maximum camber location of 66.3%C. The supersonic flight-advantage airfoil has a maximum thickness of 4%C, a maximum thickness location of 49.4%C, a maximum camber of 0.83%C, and a maximum camber location of 74.5%C. The hypersonic flight-advantage airfoil has a maximum thickness of 4%C, a maximum thickness location of 53.4%C, a maximum camber of 1.12%C, and a maximum camber location of 73.9%C; where C is the airfoil chord length. The airfoil combination includes three airfoils that can be adjusted to change with each other through an airfoil deformation mechanism; The airfoil deformation mechanism includes a flexible skin (5), an I-beam (6), and a telescopic link (7). The flexible skin (5) has an airfoil shape. The I-beam (6) is installed inside the flexible skin (5). The I-beam (6) is movable, thereby changing the maximum thickness position of the airfoil. One end of the telescopic link (7) is connected to the flexible skin (5), and the other end of the telescopic link (7) is connected to the I-beam (6). The telescopic link (7) moves in extension and retraction under the action of the actuator, thereby causing the flexible skin (5) to deform and change the shape of the airfoil represented by the flexible skin (5).

2. The airfoil combination for a morphing aircraft according to claim 1, characterized in that, The geometric coordinate expressions for the upper and lower surfaces of the airfoil are as follows: ； ； in: The vertical coordinate of the upper surface of the airfoil; The vertical coordinate of the lower surface of the airfoil; The coefficients of the expression representing the geometric coordinates of the upper surface of the airfoil; This indicates that there are a total of 9 expression coefficients; The coefficients of the expression representing the geometric coordinates of the lower surface of the airfoil; The x-coordinate of the surface of a unit airfoil.

3. The airfoil combination for a morphing aircraft according to claim 2, characterized in that, The coefficients of the geometric coordinate expression for the dominant airfoil for transonic flight are: ； The coefficients of the geometric coordinate expression for the supersonic flight-dominant airfoil are: ； The coefficients of the geometric coordinate expression for the dominant airfoil for hypersonic flight are: ； in: , , , , , , , , , representing the 0th, 1st, 2nd, 3rd, 4th, 5th, 6th, 7th, and 8th expression coefficients, respectively.

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