High-lift two-element airfoil for general aviation and uavs

Abstract

The application relates to a high-lift two-section airfoil suitable for general aircraft and unmanned aerial vehicles, which reaches the best use aerodynamic efficiency at a lift coefficient of 1.6, and the airfoil has a slow stall performance, and the maximum use lift coefficient is more than 2.0. The main wing airfoil has a front edge radius of 1.19% C, a maximum thickness of 13.80% C, a maximum thickness position of 36.32% C, a maximum camber of 6.12% C and a maximum camber position of 82.28% C. The rudder surface type has a front edge radius of 0.32% C, a maximum thickness of 2.50% C, a maximum thickness position of 101.81% C, a maximum camber of 3.12% C, a maximum camber position of 79.58% C and a rotation shaft position of (0.86, -0.04), wherein C is the base airfoil chord length before segmentation. When the rudder surface lift coefficient increases, the main wing lift coefficient also increases under the action of the circulation effect, and the climbing rate of the aircraft is improved, which is very important for flight safety.

Description

Technical Field

[0001] This application relates to the field of airfoil design technology for aircraft, and in particular to a high-lift two-section airfoil suitable for general aviation aircraft and unmanned aerial vehicles. Background Technology

[0002] The potential for performance improvement in single-segment airfoils is relatively limited. In practice, the maximum lift coefficient of currently used single-segment airfoils typically does not exceed 2.0, and to ensure flight safety margins, the lift coefficient is kept below 1.4. When applied to UAVs, due to three-dimensional effects, the lift coefficient generally does not exceed 1.2. This limits further improvements in UAV airfoil performance. A certain company, through research, discovered that increased payload leads to a sharp increase in drag for UAVs. Therefore, to improve the UAV's loiter capability, its lift coefficient needs to be further increased. By extensively arranging a designed two-segment airfoil on the wing, they increased the UAV's lift coefficient to over 1.5, ultimately improving the cruise factor by over 30%, enabling the Heron UAV to achieve an ultra-long loiter time of over 54 hours. Figure 1 As shown, (a) represents the cruise factor and lift coefficient distribution of a medium aspect ratio wing; (b) represents the cruise factor and lift coefficient distribution of a long-endurance high-lift wing.

[0003] Another advantage of two-segment airfoils is their relatively simple structure and light weight, which minimizes additional load when applied to UAVs. Simultaneously, two-segment airfoils can achieve multiple operating states, including acceleration, cruise, braking, and landing, giving the wing good adaptability to different operating conditions. This meets the growing multi-mission requirements of modern UAVs. Currently, UAVs using slotted flaps, such as... Figure 2 As shown, (a) is a first-class drone, (b) is a second-class drone, and (c) is a third-class drone. The first two types of drones are known for their strong loitering ability.

[0004] The principle behind the increased aerodynamic efficiency of a two-section airfoil is quite complex. AMO Smith once summarized the principle, believing that the slotted airfoil produces five main aerodynamic effects: 1) The leading-edge slot weakens the pressure peak of the main wing, making it less prone to flow separation, thus improving stall capability; 2) The circulation-induced effect of the trailing-edge control surface can increase the lift of the main wing; 3) The pressure in the flow direction region at the trailing edge of the main wing is reduced due to the presence of the trailing-edge control surface, thus reducing the adverse pressure gradient and making separation less likely; 4) The flow at the trailing edge of the main wing can achieve effective pressure recovery in the slot, which also makes the flow less prone to separation; 5) The external flow passing through the slot forms a new boundary layer on the trailing-edge control surface, which can delay flow separation on the control surface. Summary of the Invention

[0005] Therefore, it is necessary to provide a high-lift two-section airfoil suitable for general aircraft and unmanned aerial vehicles in view of the above technical problems.

[0006] A high-lift two-section airfoil suitable for general aircraft and unmanned aerial vehicles, the high-lift two-section airfoil comprising a main wing airfoil and a control surface; the leading edge radius of the main wing airfoil is 1.19%C, the maximum thickness of the main wing airfoil is 13.80%C, the maximum thickness position is 36.32%C, the maximum camber of the main wing airfoil is 6.12%C, and the maximum camber position is 82.28%C; the leading edge radius of the control surface is 0.32%C, the maximum thickness of the control surface is 2.50%C, the maximum thickness position is 101.81%C, the maximum camber of the control surface is 3.12%C, the maximum camber position is 79.58%C, and the control surface rotation axis position is (0.86, -0.04); wherein C is the basic airfoil chord length before segmentation.

[0007] The high-lift two-section airfoil suitable for general aircraft and unmanned aerial vehicles has the optimal use aerodynamic efficiency at a lift coefficient of 1.6, has a slow stall performance, and has a maximum use lift coefficient of more than 2.0. The leading edge radius of the main wing airfoil is 1.19%C, the maximum thickness is 13.80%C, the maximum thickness position is 36.32%C, the maximum camber is 6.12%C, and the maximum camber position is 82.28%C. The leading edge radius of the control surface is 0.32%C, the maximum thickness is 2.50%C, the maximum thickness position is 101.81%C, the maximum camber is 3.12%C, the maximum camber position is 79.58%C, and the rotation axis position is (0.86, -0.04). C is the basic airfoil chord length before segmentation. The lift coefficient of the control surface increases, and the lift coefficient of the main wing also increases under the action of the circulation effect, thereby improving the climb rate of the aircraft, which is very important for flight safety. BRIEF DESCRIPTION OF DRAWINGS

[0008] Figure 1 Fig. 1 is a cruise factor and lift coefficient distribution for a medium aspect ratio wing, wherein (a) is the cruise factor and lift coefficient distribution of a medium aspect ratio wing, and (b) is the cruise factor and lift coefficient distribution of a long endurance high-lift wing;

[0009] Figure 2 Fig. 2 is a schematic diagram of a UAV using a slotted flap, wherein (a) is a first type of UAV, (b) is a second type of UAV, and (c) is a third type of UAV;

[0010] Figure 3 Fig. 3 is a configuration of a high-lift two-section airfoil suitable for general aircraft and unmanned aerial vehicles in one embodiment;

[0011] Figure 4 Fig. 4 is a comparison of the airfoil before and after segmentation in another embodiment;

[0012] Figure 5 Fig. 5 is a comparison of the airfoil before and after segmentation optimization in another embodiment;

[0013] Figure 6 For another embodiment, the wing profile is segmented, where (a) is the slot profile generation, and (b) is the slot parameter;

[0014] Figure 7 For another embodiment, the optimized cruise configuration is compared with the original wing profile FX63-137 aerodynamic performance, where (a) is the lift-drag characteristic curve, (b) is the lift coefficient and lift-drag ratio curve, and (c) is the lift coefficient and cruise factor curve;

[0015] Figure 8 For another embodiment, the pressure cloud map before and after segmentation of the original wing profile FX63-137 is compared, where (a) is the pressure cloud map before segmentation, and (b) is the pressure cloud map after segmentation;

[0016] Figure 9 For another embodiment, the pressure distribution before and after segmentation at 3° angle of attack is compared;

[0017] Figure 10 For another embodiment, the pressure cloud map before and after segmentation at 12° angle of attack is compared, where (a) is the pressure cloud map before segmentation, and (b) is the pressure cloud map after segmentation;

[0018] Figure 11 For another embodiment, the pressure distribution before and after segmentation at 12° angle of attack is compared;

[0019] Figure 12 For another embodiment, the high-lift configuration is compared with the original wing aerodynamic performance, where (a) is the lift coefficient curve with angle of attack, and (b) is the lift coefficient and lift-drag ratio curve;

[0020] Figure 13 For another embodiment, the pressure cloud map before and after segmentation is compared, where (a) is the pressure cloud map before segmentation, and (b) is the pressure cloud map after segmentation;

[0021] Figure 14 For another embodiment, the pressure distribution before and after segmentation is compared. DETAILED DESCRIPTION

[0022] In order to make the purpose, technical scheme and advantages of the present application more clear, the present application is further described in detail below in combination with the drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and do not limit the present application.

[0023] The high-lift two-segment wing profile suitable for general aircraft and unmanned aerial vehicles proposed in the present application is referred to as GYW-SA-138 wing profile.

[0024] In one embodiment, a high-lift two-element airfoil suitable for general aviation and unmanned aerial vehicles is provided, the high-lift two-element airfoil comprising a main wing airfoil and a control surface; the main wing airfoil has a leading edge radius of 1.19%C, a maximum thickness of 13.80%C, a maximum thickness location of 36.32%C, a maximum camber of 6.12%C, and a maximum camber location of 82.28%C; the control surface has a leading edge radius of 0.32%C, a maximum thickness of 2.50%C, a maximum thickness location of 101.81%C, a maximum camber of 3.12%C, a maximum camber location of 79.58%C, and a control surface pivot location of (0.86, -0.04); wherein C is the chord length of the base airfoil before segmentation.

[0025] The design Reynolds number of the GYW-SA-138 airfoil is 8x10 5 . Three states are set, namely cruising state, accelerating state and taking-off and landing state, and the control surface deflection angles are 6°, 16° and -14° respectively. The configuration of the high-lift two-element airfoil suitable for general aviation and unmanned aerial vehicles is shown in Figure 3

[0026] Specifically, the high-lift two-element airfoil suitable for general aviation and unmanned aerial vehicles is based on the FX63-137 airfoil, segmented, and the slot tip is selected at the position of 0.82 times the chord length according to experience, the control surface pivot location is (0.86, -0.04), and a group of optimal parameters is selected by trying various segmentation parameters to obtain the basic two-element airfoil. The comparison of the airfoil shapes before and after segmentation is shown in Figure 4 . Then, the two-element airfoil is taken as the basis, and the computational fluid dynamics method and Kriging surrogate model optimization are used to optimize the flight conditions with Reynolds numbers of 3x10 5 -1x10 6 . The lift coefficient of 1.6 reaches the best use of aerodynamic efficiency, the angle of attack is set to 3°, the control surface deflection is 6°, and the maximum lift-drag ratio is K α＝3°,δ＝6° . Figure 5 Comparison of airfoils before and after segmentation optimization.

[0027] In the present application, the Reynolds number of 3x10 5 -1x10 6 is defined as a lower Reynolds number.

[0028] ​The high-lift two-element airfoil suitable for general aircraft and unmanned aerial vehicles has an optimal use aerodynamic efficiency at a lift coefficient of 1.6, and has a slow stall performance, and a maximum use lift coefficient of more than 2.0. The main wing airfoil has a front edge radius of 1.19% C, a maximum thickness of 13.80% C, a maximum thickness position of 36.32% C, a maximum camber of 6.12% C, and a maximum camber position of 82.28% C. The rudder surface has a front edge radius of 0.32% C, a maximum thickness of 2.50% C, a maximum thickness position of 101.81% C, a maximum camber of 3.12% C, a maximum camber position of 79.58% C, and a rotation shaft position of (0.86, -0.04), wherein C is the base airfoil chord length before segmentation. As the rudder surface lift coefficient increases, the main wing lift coefficient also increases under the action of the circulation effect, which improves the climb rate of the aircraft, which is very important for flight safety.

[0029] In one embodiment, the high-lift two-element airfoil design process includes: obtaining a reference airfoil; determining a basic two-element airfoil based on the reference airfoil using a segmentation strategy; the basic two-element airfoil includes a main wing airfoil and a rudder surface; setting the range of optimization parameters of the basic two-element airfoil; the optimization parameters include rudder surface setback width, drop height, wing leading edge shape control point, and main wing trailing edge shape control point; obtaining sample points using an optimal hypercube design method according to the range of optimization parameters; establishing a Kriging surrogate model according to the sample points; setting optimization objectives; the optimization objectives include: a first optimization objective and a second optimization objective; the first optimization objective is to maximize the maximum lift-drag ratio when the angle of attack is 3° and the rudder surface deflection is 6°; the second optimization objective is to maximize the lift coefficient when the angle of attack is 12° and the rudder surface rotation is 16°; optimizing the airfoil using a Kriging surrogate model optimization method and an SST-IDDES calculation method according to the optimization objectives to determine the high-lift two-element airfoil.

[0030] Specifically, the airfoil segmentation method is as shown in Figure 6 , wherein (a) is a slot profile generation, and (b) is a slot parameter. The main geometric parameters of the two-element airfoil are the slot parameter and the rudder surface deflection angle. Generally, the slot shape can be divided into L-type, S-type and C-type, as shown in Figure 6 , and the main characteristic parameters of the slot are Gap and OL values. The segmentation strategy is to first determine the slot entry point and the slot tip point, and according to experience, the position of the entry point is selected at 0.82 chord length position. A plurality of control points are set at the slot position, and a B-spline curve is used to connect the slot entry, tip and control points to generate the slot profile. Then, the rudder surface is set to have a setback dx and a rotation β around the rudder surface rotation shaft (x, y), and in order to control the flow separation of the rudder surface, a relatively full leading edge is designed for the rudder surface, which is generally considered to be beneficial to the flow field of the rudder surface not to produce flow separation at large rudder surface deflection angles.

[0031] Kriging surrogate model, which is a kind of surface interpolation and response surface approximation unbiased optimal estimation method, is widely used in engineering optimization field after its proposal.

[0032] Kriging model is composed of global model f(x) and local deviation model Z(x):

[0033] y(x) = f(x) + Z(x)

[0034] Where the global model is replaced by a constant β, and the local deviation is a static random process with a mean of 0, and the covariance matrix represents the degree of local deviation: Where, is the correlation matrix, symmetric along the diagonal, R(x (i) ,y (j) ) is the correlation function between sample points x (i) and x (j) , generally using Gaussian function:

[0035] The correlation parameter vector θ composed of n correlation parameters θ t is the key to the construction of Kriging surrogate model. After the correlation function is determined, the approximate response model of y(x) can be established

[0036]

[0037] Where Y is the response value containing sample data, is the estimated coefficient vector, and f is the unit column vector.

[0038] r(x) = [R(x,x 1 ), R(x,x 2 ), … R(x,x n )]

[0039] Where r(x) is the correlation vector between the unknown vector and the known vector.

[0040] The estimate of variance is:

[0041]

[0042] σ 2 and are functions of θ, then the correlation function θ can be obtained by maximizing the following relationship:

[0043]

[0044] The Kriging surrogate model can be obtained by solving the above problem. The model can achieve good fitting effect on nonlinear problems, and is very suitable for aerodynamic optimization problems.

[0045] A large amount of calculation data is needed to establish a surrogate model, which requires designing a scientific experimental method to cover as many experimental conditions as possible with fewer experimental points. The common experimental methods at present are orthogonal array method, central composite design method, Latin hypercube design method, and the method used in the embodiment is optimal hypercube design method, which improves the uniformity of the original Latin hypercube design and makes it have better space filling effect, so that the surrogate model fitted by the results of the same number of sample points is more true and accurate.

[0046] The optimization design objectives are as follows: one is to achieve the best use of aerodynamic efficiency with a lift coefficient of 1.6, set the angle of attack to 3° and the rudder deflection to 6°, and the optimization objective is to maximize the lift-drag ratio K α＝3o,δ＝6o . The other is to have a slow speed performance of the airfoil, with a maximum use of lift coefficient of more than 2.0, set the angle of attack to 12° and the rudder rotation to 16°, and the optimization objective is to maximize the lift coefficient Cl α＝12o,δ＝16o . The SST-IDDES calculation method is used, the Reynolds number is 8×10 5 , and the optimization objective can be described as follows: finding a set of variables x1 to make:

[0047] Objective maxCl α＝12o,δ＝16o (x1)

[0048] maxK α＝3o,δ＝6o (x1)

[0049] Wherein, x1 is the up-down fluctuation amplitude of the airfoil.

[0050] In one embodiment, the rudder setback width and the drop height range are [0.0064-0.0096] and [0.008-0.012] respectively; the horizontal and vertical coordinates of the wing leading edge shape control points range are [0.072, 0.108] and [0.024-0.036] respectively; and the horizontal and vertical coordinates of the main wing trailing edge shape control points range are [0.08-0.12] and [0.012-0.024] respectively.

[0051] Specifically, the optimization parameters are set to 6, which are the rudder setback width Width, the drop height Height, the wing leading edge shape control points (F x , F y ), and the main wing trailing edge shape control points (M x , M y ), and the parameter ranges are shown in Table 1. The optimal Latin hypercube method is used to select 500 groups of data to construct experimental data as shown in Table 2.

[0052] Table 1 Parameter Range

[0053] Parameter F x ]]> F y ]]> Height Range 0.072～0.108 0.024～0.036 0.008～0.012 Parameter M x ]]> M y ]]> Width Range 0.08～0.12 0.012～0.024 0.0064～0.0096

[0054] Table 2 500 Set Data Construction Experiment Data Table

[0055]

[0056]

[0057]

[0058] In one of the embodiments, starting from the trailing edge, counterclockwise rotation takes the coordinate values of the upper surface of the 115 main wing airfoils, the lower surface of the 136 main wing airfoils, the 116 rudder surface coordinate values, and the 1 rudder surface rotation axis position coordinate values; the coordinates corresponding to the airfoil upper and lower surfaces, the rudder surface, and the rotation axis position are as follows: X and Y represent the discrete point coordinate values of the airfoil upper and lower surfaces, the rudder surface, and the rotation axis position in the two-dimensional coordinate system, respectively;

[0059] The upper surface X, Y of the main wing airfoil has the following values: (0.8228, 0.0566), (0.8122, 0.0592), (0.8016, 0.0618), (0.7910, 0.0644), (0.7804, 0.0669), (0.7699, 0.0694), (0.7593, 0.0719), (0.7488, 0.0743), (0.7383, 0.0767), (0.7278, 0.0791), (0.7173, 0.0813), (0.7069, 0.0836), (0.6964, 0.0858), (0.6860, 0.0879), (0.6756, 0.0900), (0.6652, 0.0920), (0.6549, 0.0940), (0.6445, 0.0959), (0.6342, 0.0978), (0.6238, 0.0996), (0.6135, 0.1013), (0.6032, 0.1030), (0.5929, 0.1046), (0.5826, 0.1062), (0.5723, 0.1077), (0.5621, 0.1091), (0.5518, 0.1104), (0.5415, 0.1117), (0.5313, 0.1129), (0.5211, 0.1140), (0.5108, 0.1151), (0.5006, 0.1161), (0.4904, 0.1170), (0.4802, 0.1179), (0.4700, 0.1187), (0.4598, 0.1194), (0.4496, 0.1200), (0.4395, 0.1205), (0.4293, 0.1210), (0.4192, 0.1214), (0.4090, 0.1217), (0.3989, 0.1220), (0.3888, 0.1221), (0.3788, 0.1222), (0.3687, 0.1222), (0.3587, 0.1221), (0.3487, 0.1219), (0.3387, 0.1216), (0.3287, 0.1213), (0.3188, 0.1208), (0.3089, 0.1203), (0.2990, 0.1197), (0.2892, 0.1190), (0.2794, 0.1182), (0.2697, 0.1173), (0.2599, 0.1163), (0.2503, 0.1152), (0.2407, 0.1140), (0.2312, 0.1127), (0.2217, 0.1114), (0.2123, 0.1099), (0.2029,0.1083)，(0.1937,0.1066)，(0.1845,0.1049)，(0.1754,0.1030)，(0.1665,0.1010)，(0.1576,0.0989)，(0.1488,0.0967)，(0.1402,0.0944)，(0.1317,0.0920)，(0.1234,0.0895)，(0.1152,0.0869)，(0.1073,0.0842)，(0.0995,0.0815)，(0.0920,0.0787)，(0.0848,0.0758)，(0.0779,0.0730)，(0.0714,0.0702)，(0.0653,0.0674)，(0.0596,0.0647)，(0.0543,0.0620)，(0.0494,0.0594)，(0.0448,0.0568)，(0.0407,0.0543)，(0.0369,0.0519)，(0.0333,0.0495)，(0.0301,0.0471)，(0.0272,0.0448)，(0.0245,0.0426)，(0.0220,0.0404)，(0.0197,0.0382)，(0.0176,0.0361)，(0.0157,0.0341)，(0.0140,0.0321)，(0.0124,0.0301)，(0.0109,0.0282)，(0.0096,0.0264)，(0.0084,0.0246)，(0.0072,0.0229)，(0.0062,0.0212)，(0.0053,0.0195)，(0.0045,0.0179)，(0.0037,0.0164)，(0.0030,0.0149)，(0.0024,0.0134)，(0.0019,0.0120)，(0.0014,0.0106)，(0.0010,0.0092)，(0.0007,0.0079)，(0.0004,0.0066)，(0.0002,0.0054)，(0.0001,0.0041)，(0.0000,0.0029)，(0.0000,0.0017)，(0.0001,0.0005)；.

[0060] The values of the lower surface X, Y of the main wing airfoil are as follows: (0.0003,-0.0007), (0.0007,-0.0019), (0.0011,-0.0030), (0.0017,-0.0042), (0.0024,-0.0053), (0.0033,-0.0063), (0.0043,-0.0073), (0.0054,-0.0082), (0.0066,-0.0090), (0.0079,-0.0098), (0.0094,-0.0105), (0.0109,-0.0112), (0.0125,-0.0119), (0.0143,-0.0125), (0.0161,-0.0131), (0.0181,-0.0137), (0.0202,-0.0143), (0.0225,-0.0148), (0.0250,-0.0154), (0.0276,-0.0159), (0.0305,-0.0165), (0.0336,-0.0170), (0.0370,-0.0176), (0.0407,-0.0182), (0.0447,-0.0187), (0.0492,-0.0192), (0.0541,-0.0198), (0.0594,-0.0203), (0.0654,-0.0208), (0.0719,-0.0213), (0.0790,-0.0217), (0.0866,-0.0221), (0.0949,-0.0225), (0.1036,-0.0228), (0.1128,-0.0230), (0.1224,-0.0233), (0.1322,-0.0234), (0.1422,-0.0235), (0.1524,-0.0235), (0.1627,-0.0235), (0.1730,-0.0235), (0.1834,-0.0233), (0.1939,-0.0231), (0.2044,-0.0229), (0.2148,-0.0226), (0.2254,-0.0223), (0.2359,-0.0219), (0.2464,-0.0214), (0.2569,-0.0209), (0.2674,-0.0203), (0.2779,-0.0197), (0.2884,-0.0190), (0.2990,-0.0183), (0.3095,-0.0175), (0.3201,-0.0167), (0.3307,-0.0158), (0.3414,-0.0148), (0.3521,-0.0138)，(0.3628,-0.0128)，(0.3736,-0.0117)，(0.3844,-0.0106)，(0.3953,-0.0095)，(0.4062,-0.0083)，(0.4171,-0.0071)，(0.4281,-0.0058)，(0.4391,-0.0046)，(0.4502,-0.0033)，(0.4613,-0.0020)，(0.4724,-0.0007)，(0.4836,0.0006)，(0.4947,0.0019)，(0.5059,0.0032)，(0.5170,0.0045)，(0.5281,0.0057)，(0.5392,0.0070)，(0.5429,0.0074)，(0.5466,0.0079)，(0.5503,0.0084)，(0.5541,0.0089)，(0.5580,0.0094)，(0.5619,0.0100)，(0.5659,0.0106)，(0.5700,0.0112)，(0.5741,0.0118)，(0.5783,0.0125)，(0.5826,0.0133)，(0.5870,0.0140)，(0.5914,0.0148)，(0.5959,0.0156)，(0.6005,0.0165)，(0.6051,0.0173)，(0.6098,0.0182)，(0.6146,0.0192)，(0.6195,0.0202)，(0.6244,0.0212)，(0.6294,0.0222)，(0.6345,0.0233)，(0.6396,0.0243)，(0.6448,0.0255)，(0.6500,0.0266)，(0.6553,0.0278)，(0.6606,0.0290)，(0.6660,0.0302)，(0.6714,0.0314)，(0.6768,0.0327)，(0.6822,0.0339)，(0.6877,0.0352)，(0.6932,0.0365)，(0.6987,0.0378)，(0.7042,0.0391)，(0.7096,0.0404)，(0.7151,0.0417)，(0.7206,0.0430)，(0.7260,0.0442)，(0.7314,0.0455)，(0.7368,0.0467)，(0.7421,0.0479)，(0.7474,0.0491)，(0.7526,0.0502)，(0.7577,0.0513)，(0.7628,0.0523)，(0.7678,0.0532)，(0.7727,0.0541)，(0.7775,0.0549)，(0.7823,0.0556)，(0.7869,0.0562)，(0.7914,0.0568)，(0.7958,0.0572)，(0.8001,0.0574)，(0.8042,0.0576)，(0.8082,0.0575)，(0.8121,0.0574)，(0.8158,0.0570)，(0.8194,0.0565)，(0.8228,0.0558)，(0.8228,0.0566)；.

[0061] The values of the rudder surfaces X, Y are as follows: (1.0181, -0.0291), (1.0149, -0.0269), (1.0097, -0.0237), (1.0042, -0.0204), (0.9983, -0.0171), (0.9920, -0.0138), (0.9852, -0.0104), (0.9781, -0.0070), (0.9705, -0.0036), (0.9625, -0.0001), (0.9542, 0.0035), (0.9455, 0.0070), (0.9364, 0.0107), (0.9271, 0.0144), (0.9176, 0.0181), (0.9078, 0.0218), (0.8979, 0.0256), (0.8879, 0.0294), (0.8737, 0.0340), (0.8723, 0.0344), (0.8708, 0.0349), (0.8692, 0.0353), (0.8676, 0.0357), (0.8659, 0.0361), (0.8642, 0.0365), (0.8623, 0.0369), (0.8604, 0.0374), (0.8585, 0.0378), (0.8565, 0.0381), (0.8544, 0.0385), (0.8523, 0.0389), (0.8501, 0.0392), (0.8479, 0.0395), (0.8456, 0.0398), (0.8433, 0.0401), (0.8410, 0.0403), (0.8387, 0.0405), (0.8363, 0.0407), (0.8339, 0.0408), (0.8316, 0.0409), (0.8292, 0.0409), (0.8268, 0.0409), (0.8245, 0.0409), (0.8222, 0.0408), (0.8199, 0.0406), (0.8176, 0.0404), (0.8155, 0.0402), (0.8133, 0.0399), (0.8113, 0.0396), (0.8093, 0.0392), (0.8075, 0.0388), (0.8057, 0.0384), (0.8040, 0.0379), (0.8025, 0.0373), (0.8011, 0.0368), (0.7998, 0.0362), (0.7987, 0.0355), (0.7977, 0.0349), (0.7969, 0.0342), (0.7963, 0.0335), (0.7958, 0.0327), (0.7955, 0.0320), (0.7954, 0.0312), (0.7955, 0.0305), (0.7958, 0.0297), (0.7963, 0.0289), (0.7970, 0.0281), (0.7979, 0.0274), (0.7990, 0.0266), (0.8002, 0.0258), (0.8017, 0.0251), (0.8034, 0.0243), (0.8052, 0.0236), (0.8073, 0.0229), (0.8095, 0.0222), (0.8119, 0.0215), (0.8144, 0.0208), (0.8171, 0.0202), (0.8199, 0.0195), (0.8229, 0.0189), (0.8260, 0.0182), (0.8292, 0.0176), (0.8325, 0.0170), (0.8359, 0.0164), (0.8393, 0.0158), (0.8428, 0.0152), (0.8464, 0.0146), (0.8500, 0.0140), (0.8537, 0.0134), (0.8573, 0.0128), (0.8610, 0.0122), (0.8647, 0.0116), (0.8684, 0.0110), (0.8721, 0.0104), (0.8757, 0.0098), (0.8794, 0.0092), (0.8830, 0.0087), (0.8923, 0.0071), (0.9014, 0.0055), (0.9104, 0.0037), (0.9192, 0.0018), (0.9278, -0.0001), (0.9363, -0.0022), (0.9445, -0.0043), (0.9526, -0.0066), (0.9604, -0.0089), (0.9680, -0.0113), (0.9754, -0.0138), (0.9825, -0.0164), (0.9892, -0.0190), (0.9957, -0.0217), (1.0019, -0.0244), (1.0077, -0.0271), (1.0131, -0.0298), (1.0166, -0.0316), (1.0181, -0.0291).

[0062] The values of the rudder pivot position X, Y are as follows: (0.86, -0.04).

[0063] In the verification embodiment, the FX63-137 airfoil is segmented, the slot tip is selected at 0.82 times the chord length position, and the rudder rotation axis position is (0.86, -0.04) according to experience, and a group of optimal parameters is selected by trying various segmentation parameters to obtain a basic two-segment airfoil; then, the two-segment airfoil is used as a basis, and a computational fluid dynamics method and a Kriging surrogate model are used for optimization to obtain a high-lift two-segment airfoil suitable for general aircraft and unmanned aerial vehicles. The optimization results of the Kriging surrogate model are shown in Table 3, and it can be found that the overall gain is less than 10%, but the performance of the segmented airfoil is at a relatively optimal level because the slot parameters have been continuously adjusted before optimization.

[0064] Table 3 Comparison of airfoil aerodynamic performance before and after optimization

[0065] Parameter K α＝3°,δ＝6° ]]> Cl α＝12°,δ＝6° ]] Before optimization 110 2.1 After optimization 118 2.2 Gain 7.3％ 4.8％

[0066] The aerodynamic performance of the optimized cruise configuration and the original airfoil FX63-137 is compared, as shown in Figure 7 , wherein (a) is the lift-drag characteristic curve, (b) is the lift coefficient and lift-drag ratio curve, and (c) is the lift coefficient and cruise factor curve. From Figure 7 , it can be seen that the lift coefficient of the airfoil after deflection of 6° is significantly increased, and the segmented airfoil has a certain slow-speed performance. Considering the safety margin, the lift coefficient of 1.62 is selected as the cruise lift coefficient. As can be seen from the middle graph, after the lift coefficient is greater than 1.4, the drag coefficient of the segmented airfoil is significantly smaller than that of the original airfoil, and from the right graph, it can be seen that the cruise factor of the segmented airfoil is much larger than that of the original airfoil after the lift coefficient is 1.5. The final obtained airfoil has a cruise factor that is 25% higher than the original airfoil.

[0067] Further analysis of the flow field information of the cruise configuration before and after optimization at 3° angle of attack, the pressure contour comparison of the original airfoil FX63-137 before and after segmentation is shown in Figure 8 , wherein (a) is the pressure contour before segmentation, and (b) is the pressure contour comparison after segmentation. From Figure 8 , it can be seen that the suction range of the upper surface of the segmented airfoil is expanded, and the pressure loading of the lower surface, especially near the segmentation position, is increased, so that the lift coefficient of the cruise configuration is increased.

[0068] Further analysis of the surface pressure, the pressure distribution comparison before and after segmentation at 3° angle of attack is shown in Figure 9 , from Figure 9 , it can be seen that the rudder lift coefficient is increased, and the main wing lift coefficient is also increased under the action of the circulation effect.

[0069] Further analysis of the pressure contour of the cruise configuration before and after segmentation at 12° angle of attack, the pressure contour comparison before and after segmentation at 12° angle of attack is shown in Figure 10As shown, (a) is the pressure contour map before segmentation, and (b) is the pressure contour map after segmentation. From Figure 10 As can be seen at high angles of attack, the suction on the upper surface and the pressure on the lower surface increase after segmentation, and the flow separation range of the airfoil increases slightly. However, considering that the cruise configuration has sufficient safety margin while ensuring the lift coefficient, this does not affect the use of the airfoil.

[0070] Further analysis of its surface pressure, and a comparison of pressure distribution before and after optimization at a 12° angle of attack, are shown below. Figure 11 As shown. From Figure 11 It can be seen that the lift coefficient of the airfoil is still greater than that of the original airfoil under the condition of 12° angle of attack.

[0071] Further analysis of the aerodynamic performance of the high-lift configuration and the original airfoil (segmented canard) is provided, and a comparison of their aerodynamic performance is shown below. Figure 12 As shown, (a) is the lift coefficient as a function of angle of attack, and (b) is the lift coefficient versus lift-to-drag ratio curve. From Figure 12 As can be seen, with a 3° angle of attack and a 16° control surface deflection, the airfoil lift coefficient reaches 2.1. At this point, the tail-strike angle constraint is minimal, expanding the fuselage design range. Simultaneously, it can be observed that the airfoil still possesses a significant lift coefficient margin under this condition, resulting in strong takeoff and landing safety. Comparing the lift-to-drag ratio curves, it can be seen that the segmented high-lift configuration at this lift coefficient has a maximum lift-to-drag ratio of approximately 120, which is over 100% higher than the lift-to-drag ratio near the maximum lift coefficient of the original airfoil. This significantly improves the aircraft's climb rate, which is crucial for flight safety.

[0072] Further analysis of the flow field information of the high-lift configuration before and after the segmentation at a 3° angle of attack, compared with the original airfoil, and the pressure contour maps before and after the segmentation are shown below. Figure 13 As shown, (a) is the pressure contour map before segmentation, and (b) is the pressure contour map after segmentation. From Figure 13 As can be seen, the suction force on the upper surface increases significantly after segmentation, while the loading pressure on the lower surface increases significantly.

[0073] Further analysis of its surface pressure, and comparison of pressure distribution before and after segmentation, for example... Figure 14 As shown, from Figure 14 As can be seen, the lift coefficient increases significantly, while no flow separation occurs on the upper surface.

[0074] The accelerated configuration selection 2° angle of attack verification, rudder surface deflection-14°, the lift coefficient at this time is 0.38, compared with the cruise state lift coefficient 1.62, reduced four times, the cruise speed can be increased by 2 times. Compared with the large lift take-off and landing configuration 3° angle of attack condition using lift coefficient 2.0 reduced five times, the speed can be increased by 2.3 times. It can be seen that for the unmanned aerial vehicle using segmented high lift wing, the speed range is greatly changed under the condition of unchanged angle of attack, which plays an important role in the execution of various tasks of unmanned aerial vehicle.

[0075] Note: Figure 7 , Figure 9 , Figure 11 , Figure 12 , Figure 14 , the basic wing profile is FX63-137 wing profile, and the GYW-SA-138 wing profile is a high-lift two-section wing profile suitable for general aircraft and unmanned aerial vehicles.

[0076] The technical features of the above embodiments can be combined arbitrarily, and in order to make the description simple, not all possible combinations of the technical features in the above embodiments are described, however, as long as the combination of the technical features does not exist contradictory, it should be considered as the scope of the present application.

[0077] The above embodiments only express several embodiments of the present application, and the description is more specific and detailed, but it should not be understood as a limitation on the scope of the patent. It should be pointed out that for ordinary skilled in the art, without departing from the concept of the present application, a number of modifications and improvements can be made, which are within the scope of the present application. Therefore, the protection scope of the present application should be subject to the appended claims.

Claims

1. A high-lift two-element airfoil suitable for general aviation and unmanned aerial vehicles, characterized in that, The high-lift two-element airfoil comprises a main wing airfoil and a rudder surface; a leading edge radius of the main wing airfoil is 1.19%C, a maximum thickness of the main wing airfoil is 13.80%C, a maximum thickness position of the main wing airfoil is 36.32%C, a maximum camber of the main wing airfoil is 6.12%C, and a maximum camber position of the main wing airfoil is 82.28%C; a leading edge radius of the rudder surface is 0.32%C, a maximum thickness of the rudder surface is 2.50%C, a maximum thickness position of the rudder surface is 101.81%C, a maximum camber of the rudder surface is 3.12%C, and a maximum camber position of the rudder surface is 79.58%C, and a rudder rotation axis position is (0.86, -0.04); wherein C is a basic airfoil chord length before segmentation.

2. The high-lift two-element airfoil suitable for general aviation and drones of claim 1, wherein, The high-lift two-element airfoil design process comprises: obtaining a reference airfoil; determining a basic two-element airfoil based on the reference airfoil by using a segmentation strategy; the basic two-element airfoil comprises a main wing airfoil and a rudder surface; setting a range of optimization parameters of the basic two-element airfoil; the optimization parameters comprise a rudder surface setback width, a drop height, a main wing leading edge shape control point, and a main wing trailing edge shape control point; obtaining sample points by using an optimal hypercube design method according to the range of the optimization parameters; establishing a Kriging surrogate model according to the sample points; setting optimization objectives; the optimization objectives comprise a first optimization objective and a second optimization objective; the first optimization objective is that a maximum lift-drag ratio is maximum when an attack angle is 3° and a rudder deflection is 6°; the second optimization objective is that a lift coefficient is a maximum design lift coefficient when the attack angle is 12° and the rudder rotation is 16°; and optimizing the airfoil by using a Kriging surrogate model optimization method and an SST-IDDES calculation method according to the optimization objectives to determine a high-lift two-element airfoil.

3. The high-lift two-element airfoil suitable for general aviation and drones of claim 2, wherein, The ranges of the rudder surface setback width and the drop height are [0.0064~0.0096] and [0.008~0.012] respectively; the ranges of the horizontal coordinate and the vertical coordinate of the main wing leading edge shape control point are [0.072, 0.108] and [0.024~0.036] respectively; and the ranges of the horizontal coordinate and the vertical coordinate of the main wing trailing edge shape control point are [0.08~0.12] and [0.012~0.024] respectively.

4. The high-lift two-element airfoil suitable for general aviation and drones of claim 1, wherein: The coordinate values of the airfoil upper surface and lower surface, the rudder surface, and the rotation axis position are obtained by starting from the main wing trailing edge and rotating counterclockwise; X and Y represent the discrete point coordinate values of the airfoil upper surface and lower surface, the rudder surface, and the rotation axis position in a two-dimensional coordinate system respectively. wherein the upper surface X, Y of the main wing airfoil has the following values: (0.8228, 0.0566), (0.8122, 0.0592), (0.8016, 0.0618), (0.7910, 0.0644), (0.7804, 0.0669), (0.7699, 0.0694), (0.7593, 0.0719), (0.7488, 0.0743), (0.7383, 0.0767), (0.7278, 0.0791), (0.7173, 0.0813), (0.7069, 0.0836), (0.6964, 0.0858), (0.6860, 0.0879), (0.6756, 0.0900), (0.6652, 0.0920), (0.6549, 0.0940), (0.6445, 0.0959), (0.6342, 0.0978), (0.6238, 0.0996), (0.6135, 0.1013), (0.6032, 0.1030), (0.5929, 0.1046), (0.5826, 0.1062), (0.5723, 0.1077), (0.5621, 0.1091), (0.5518, 0.1104), (0.5415, 0.1117), (0.5313, 0.1129), (0.5211, 0.1140), (0.5108, 0.1151), (0.5006, 0.1161), (0.4904, 0.1170), (0.4802, 0.1179), (0.4700, 0.1187), (0.4598, 0.1194), (0.4496, 0.1200), (0.4395, 0.1205), (0.4293, 0.1210), (0.4192, 0.1214), (0.4090, 0.1217), (0.3989, 0.1220), (0.3888, 0.1221), (0.3788, 0.1222), (0.3687, 0.1222), (0.3587, 0.1221), (0.3487, 0.1219), (0.3387, 0.1216), (0.3287, 0.1213), (0.3188, 0.1208), (0.3089, 0.1203), (0.2990, 0.1197), (0.2892, 0.1190), (0.2794, 0.1182), (0.2697, 0.1173), (0.2599, 0.1163), (0.2503, 0.1152), (0.2407, 0.1140)，(0.2312, 0.1127)，(0.2217, 0.1114)，(0.2123,0.1099)，(0.2029, 0.1083)，(0.1937, 0.1066)，(0.1845, 0.1049)，(0.1754, 0.1030)，(0.1665, 0.1010)，(0.1576, 0.0989)，(0.1488, 0.0967)，(0.1402, 0.0944)，(0.1317,0.0920)，(0.1234, 0.0895)，(0.1152, 0.0869)，(0.1073, 0.0842)，(0.0995, 0.0815)，(0.0920, 0.0787)，(0.0848, 0.0758)，(0.0779, 0.0730)，(0.0714, 0.0702)，(0.0653,0.0674)，(0.0596, 0.0647)，(0.0543, 0.0620)，(0.0494, 0.0594)，(0.0448, 0.0568)，(0.0407, 0.0543)，(0.0369, 0.0519)，(0.0333, 0.0495)，(0.0301, 0.0471)，(0.0272,0.0448)，(0.0245, 0.0426)，(0.0220, 0.0404)，(0.0197, 0.0382)，(0.0176, 0.0361)，(0.0157, 0.0341)，(0.0140, 0.0321)，(0.0124, 0.0301)，(0.0109, 0.0282)，(0.0096,0.0264)，(0.0084, 0.0246)，(0.0072, 0.0229)，(0.0062, 0.0212)，(0.0053, 0.0195)，(0.0045, 0.0179)，(0.0037, 0.0164)，(0.0030, 0.0149)，(0.0024, 0.0134)，(0.0019,0.0120)，(0.0014, 0.0106)，(0.0010, 0.0092)，(0.0007, 0.0079)，(0.0004, 0.0066)，(0.0002, 0.0054)，(0.0001, 0.0041)，(0.0000, 0.0029)，(0.0000, 0.0017)，(0.0001,0.0005)；. The values of the lower surface X, Y of the main wing airfoil are as follows: (0.0003, -0.0007), (0.0007, -0.0019), (0.0011, -0.0030), (0.0017, -0.0042), (0.0024, -0.0053), (0.0033, -0.0063), (0.0043, -0.0073), (0.0054, -0.0082), (0.0066, -0.0090), (0.0079, -0.0098), (0.0094, -0.0105), (0.0109, -0.0112), (0.0125, -0.0119), (0.0143, -0.0125), (0.0161, -0.0131), (0.0181, -0.0137), (0.0202, -0.0143), (0.0225, -0.0148), (0.0250, -0.0154), (0.0276, -0.0159), (0.0305, -0.0165), (0.0336, -0.0170), (0.0370, -0.0176), (0.0407, -0.0182), (0.0447, -0.0187), (0.0492, -0.0192), (0.0541, -0.0198), (0.0594, -0.0203), (0.0654, -0.0208), (0.0719, -0.0213), (0.0790, -0.0217), (0.0866, -0.0221), (0.0949, -0.0225), (0.1036, -0.0228), (0.1128, -0.0230), (0.1224, -0.0233), (0.1322, -0.0234), (0.1422, -0.0235), (0.1524, -0.0235), (0.1627, -0.0235), (0.1730, -0.0235), (0.1834, -0.0233), (0.1939, -0.0231), (0.2044, -0.0229), (0.2148, -0.0226), (0.2254, -0.0223), (0.2359, -0.0219), (0.2464, -0.0214), (0.2569, -0.0209), (0.2674, -0.0203), (0.2779, -0.0197), (0.2884, -0.0190), (0.2990, -0.0183), (0.3095, -0.0175), (0.3201, -0.0167)，(0.3307, -0.0158)，(0.3414, -0.0148)，(0.3521, -0.0138)，(0.3628, -0.0128)，(0.3736, -0.0117)，(0.3844, -0.0106)，(0.3953, -0.0095)，(0.4062, -0.0083)，(0.4171, -0.0071)，(0.4281, -0.0058)，(0.4391, -0.0046)，(0.4502, -0.0033)，(0.4613, -0.0020)，(0.4724, -0.0007)，(0.4836, 0.0006)，(0.4947, 0.0019)，(0.5059, 0.0032)，(0.5170, 0.0045)，(0.5281, 0.0057)，(0.5392,0.0070)，(0.5429, 0.0074)，(0.5466, 0.0079)，(0.5503, 0.0084)，(0.5541, 0.0089)，(0.5580, 0.0094)，(0.5619, 0.0100)，(0.5659, 0.0106)，(0.5700, 0.0112)，(0.5741,0.0118)，(0.5783, 0.0125)，(0.5826, 0.0133)，(0.5870, 0.0140)，(0.5914, 0.0148)，(0.5959, 0.0156)，(0.6005, 0.0165)，(0.6051, 0.0173)，(0.6098, 0.0182)，(0.6146,0.0192)，(0.6195, 0.0202)，(0.6244, 0.0212)，(0.6294, 0.0222)，(0.6345, 0.0233)，(0.6396, 0.0243)，(0.6448, 0.0255)，(0.6500, 0.0266)，(0.6553, 0.0278)，(0.6606,0.0290)，(0.6660, 0.0302)，(0.6714, 0.0314)，(0.6768, 0.0327)，(0.6822, 0.0339)，(0.6877, 0.0352)，(0.6932, 0.0365)，(0.6987, 0.0378)，(0.7042, 0.0391)，(0.7096,0.0404)，(0.7151, 0.0417)，(0.7206, 0.0430)，(0.7260, 0.0442)，(0.7314, 0.0455)，(0.7368, 0.0467)，(0.7421, 0.0479)，(0.7474, 0.0491)，(0.7526, 0.0502)，(0.7577,0.0513)，(0.7628, 0.0523)，(0.7678, 0.0532)，(0.7727, 0.0541)，(0.7775, 0.0549)，(0.7823, 0.0556)，(0.7869, 0.0562)，(0.7914, 0.0568)，(0.7958, 0.0572)，(0.8001,0.0574)，(0.8042, 0.0576)，(0.8082, 0.0575)，(0.8121, 0.0574)，(0.8158, 0.0570)，(0.8194, 0.0565)，(0.8228, 0.0558)，(0.8228, 0.0566)；. The values of the rudder surface X, Y are as follows: (1.0181, -0.0291), (1.0149, -0.0269), (1.0097, -0.0237), (1.0042, -0.0204), (0.9983, -0.0171), (0.9920, -0.0138), (0.9852, -0.0104), (0.9781, -0.0070), (0.9705, -0.0036), (0.9625, -0.0001), (0.9542, 0.0035), (0.9455, 0.0070), (0.9364, 0.0107), (0.9271, 0.0144), (0.9176, 0.0181), (0.9078, 0.0218), (0.8979, 0.0256), (0.8879, 0.0294), (0.8737, 0.0340), (0.8723, 0.0344), (0.8708, 0.0349), (0.8692, 0.0353), (0.8676, 0.0357), (0.8659, 0.0361), (0.8642, 0.0365), (0.8623, 0.0369), (0.8604, 0.0374), (0.8585, 0.0378)，(0.8565, 0.0381)，(0.8544, 0.0385)，(0.8523, 0.0389)，(0.8501, 0.0392)，(0.8479, 0.0395)，(0.8456,0.0398)，(0.8433, 0.0401)，(0.8410, 0.0403)，(0.8387, 0.0405)，(0.8363, 0.0407)，(0.8339, 0.0408)，(0.8316, 0.0409)，(0.8292, 0.0409)，(0.8268, 0.0409)，(0.8245,0.0409)，(0.8222, 0.0408)，(0.8199, 0.0406)，(0.8176, 0.0404)，(0.8155, 0.0402)，(0.8133, 0.0399)，(0.8113, 0.0396)，(0.8093, 0.0392)，(0.8075, 0.0388)，(0.8057,0.0384)，(0.8040, 0.0379)，(0.8025, 0.0373)，(0.8011, 0.0368)，(0.7998, 0.0362)，(0.7987, 0.0355)，(0.7977, 0.0349)，(0.7969, 0.0342)，(0.7963, 0.0335)，(0.7958,0.0327)，(0.7955, 0.0320)，(0.7954, 0.0312)，(0.7955, 0.0305)，(0.7958, 0.0297)，(0.7963, 0.0289)，(0.7970, 0.0281)，(0.7979, 0.0274)，(0.7990, 0.0266)，(0.8002,0.0258)，(0.8017, 0.0251)，(0.8034, 0.0243)，(0.8052, 0.0236)，(0.8073, 0.0229)，(0.8095, 0.0222)，(0.8119, 0.0215)，(0.8144, 0.0208)，(0.8171, 0.0202)，(0.8199,0.0195)，(0.8229, 0.0189)，(0.8260, 0.0182)，(0.8292, 0.0176)，(0.8325, 0.0170)，(0.8359, 0.0164)，(0.8393, 0.0158)，(0.8428, 0.0152)，(0.8464, 0.0146)，(0.8500,0.0140)，(0.8537, 0.0134)，(0.8573, 0.0128)，(0.8610, 0.0122)，(0.8647, 0.0116)，(0.8684, 0.0110)，(0.8721, 0.0104)，(0.8757, 0.0098)，(0.8794, 0.0092)，(0.8830,0.0087)，(0.8923, 0.0071)，(0.9014, 0.0055)，(0.9104, 0.0037)，(0.9192, 0.0018)，(0.9278, -0.0001)，(0.9363, -0.0022)，(0.9445, -0.0043)，(0.9526, -0.0066)，(0.9604, -0.0089)，(0.9680, -0.0113)，(0.9754, -0.0138)，(0.9825, -0.0164)，(0.9892, -0.0190)，(0.9957, -0.0217)，(1.0019, -0.0244)，(1.0077, -0.0271)，(1.0131, -0.0298)，(1.0166, -0.0316)，(1.0181, -0.0291)； The values of the rudder surface rotation axis position X, Y are as follows: (0.86, -0.04).

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