A high lift-to-drag ratio, long-endurance, low-speed airfoil

Abstract

This invention discloses a high lift-to-drag ratio, long-endurance, low-speed airfoil, revealing its maximum relative camber, maximum relative thickness, and the curve equations satisfied by its upper and lower edges. The origin of the coordinate system containing the airfoil is defined as the leading edge of the chord. The X-axis coincides with the chord line, pointing from the leading edge to the trailing edge of the airfoil. The Y-axis is perpendicular to the X-axis, pointing in the direction of the airfoil's curvature. The chord length is defined as C; therefore, the maximum relative camber is 6%C, and the maximum relative thickness is 4.5%C. The airfoil of this invention exhibits a higher lift-to-drag ratio and a higher lift coefficient while maintaining a lower drag coefficient. It reduces the weight of the wing structure, and its thin, curved shape results in a smaller internal wing cavity, optimizing aircraft weight while concentrating stress on the structure.

Description

Technical Field

[0001] This invention relates to the field of unmanned aerial vehicle (UAV) airfoil design technology, specifically to a high lift-to-drag ratio, long-endurance, low-speed airfoil. Background Technology

[0002] The wing plays a crucial role in the shape design of an aircraft, and its aerodynamic and mechanical properties are primarily influenced by the airfoil. To improve the aerodynamic and mechanical performance of an aircraft, high-performance airfoils are required. Compared to conventional aircraft, low-altitude, low-speed aircraft have unique challenges in aerodynamic design and analysis. Different design methods exist for low-altitude, low-speed aircraft compared to medium- and high Reynolds number aircraft. Low-altitude, low-speed aircraft should employ high lift-to-drag ratio laminar flow airfoil design technology to enhance cruise performance. Lightweight low-altitude, low-speed aircraft require wings that combine excellent lift-to-drag aerodynamic characteristics with a relatively light structural weight to successfully complete long-endurance reconnaissance and detection missions. Therefore, while ensuring good structural strength, these types of UAVs generally use thin airfoils with a high lift-to-drag ratio of 9% to 11%. While pursuing excellent lift characteristics, the nose-down moment of the airfoil must also be controlled within a suitable range to reduce overall trim drag.

[0003] Currently, the high-lift airfoils used in domestic low-speed UAVs generally fall into two categories: one with greater thickness, smaller camber, and a lower lift-to-drag ratio, which is advantageous for wing structural mechanics design but detrimental to improving the UAV's cruise, climb, and other aerodynamic performance; the other with smaller thickness, larger camber, and a higher lift-to-drag ratio, which is disadvantageous for wing structural mechanics design but beneficial to improving the UAV's cruise, climb, and other aerodynamic performance. Currently, there is no airfoil that combines both excellent wing structural mechanics design and good aerodynamic performance, thus meeting the operational requirements of low-speed, long-endurance UAVs. Summary of the Invention

[0004] The purpose of this invention is to provide a high lift-to-drag ratio, long-endurance, low-speed airfoil, specifically a long-endurance, low-speed airfoil. This airfoil has a higher lift-to-drag ratio at lower Reynolds numbers, possessing a higher lift coefficient while having a lower drag coefficient; it can reduce the weight of the wing structure, and its thin, curved shape results in a smaller wing cavity, optimizing aircraft weight while concentrating the stress structure.

[0005] To achieve the above objectives, the present invention is specifically implemented through the following technical solutions:

[0006] A high lift-to-drag ratio, long-endurance, low-speed airfoil has a horizontal line passing through the left end point of the airfoil as the X-axis, with the X-axis pointing from the leading edge of the airfoil to the trailing edge and coinciding with the chord line of the airfoil; and a vertical line passing through the left end point of the airfoil as the Y-axis, with the Y-axis perpendicular to the X-axis and pointing in the direction of the curvature of the upper edge of the airfoil.

[0007] The equation for the upper edge curve of the airfoil is as follows:

[0008] When x∈[0, 2.5], y U =1.329400x + -0.309680x 2 + 0.032832x 3 ;

[0009] When x∈(2.5, 7.5], y U =0.543000 + 0.603067x + -0.025920x 2 +0.000789x 3 ;

[0010] When x∈(7.5, 15], y U =0.506800 + 0.563793x + -0.015268x 2 +0.000153x 3 ;

[0011] When x∈(15, 25], y U =1.412000 + 0.430667x + -0.009040x 2 +0.000061x 3 ;

[0012] When x∈(25, 40], y U =3.061000 + 0.266890x + -0.003699x 2 +0.000004x 3 ;

[0013] When x∈(40, 60], y U =5.323000 + 0.163367x + -0.002490x 2 +0.000003x 3 ;

[0014] When x∈(60, 80], y U =9.026000 + 0.015850x + -0.000590x 2 + -0.000005x 3 ;

[0015] When x∈(80, 100], y U = (-9.56666666666675e-5)x 3+0.0243450000000001x 2 - 2.24478333333334x + 76.9700000000007;

[0016] The equations for the lower edge curve of the airfoil are as follows:

[0017] When x∈[0, 2.5], y L =-0.684000x + 0.305120x 2 + -0.033664x 3 ;

[0018] When x∈(2.5, 7.5], y L =-0.405000 + -0.037267x + 0.030480x 2 + -0.001365x 3 ;

[0019] When x∈(7.5, 15], y L =-0.982200 + 0.195093x + -0.000008x 2 + -0.000063x 3 ;

[0020] When x∈(15,25], y L =-1.127000 + 0.232467x + -0.002880x 2 +0.000005x 3 ;

[0021] When x∈(25,40], y L =-1.307000 + 0.256167x + -0.003890x 2 +0.000019x 3 ;

[0022] When x∈(40,60], y L =-2.397000 + 0.283883x + -0.003475x 2 +0.000009x 3 ;

[0023] When x∈(60,80], y L =-6.695000 + 0.461717x + -0.005885x 2 +0.000019x3 ;

[0024] When x∈(80,100], y L =-120.215000 + 4.291650x + -0.048795x 2 +0.000179x 3 ;

[0025] Where x is the x-coordinate, x∈[0, 100], y U Let y be a point on the upper edge curve of the airfoil. L Let C be a point on the lower edge curve; define the chord length as C, then the maximum relative camber of the airfoil is 6%C; the maximum relative thickness of the airfoil is 4.5%C.

[0026] The beneficial effects of this invention are as follows:

[0027] (1) The airfoil of the present invention has a higher lift-to-drag ratio at a lower Reynolds number, and has a higher lift coefficient while having a lower drag coefficient;

[0028] (2) The airfoil of the present invention can reduce the weight of the wing structure. The thin and curved shape makes the inner cavity of the wing smaller, which can optimize the weight of the aircraft, but the stress structure is more concentrated.

[0029] (3) The airfoil of the present invention has a very smooth leading edge, which can generate a suitable turbulent boundary layer in the airfoil and delay the occurrence of airflow separation; even if boundary layer separation occurs, it often starts from the trailing edge and gradually develops forward, the stall process is relatively gentle, the safety is high, and it has good stall performance.

[0030] (4) The airfoil of the present invention has an extremely high maximum lift coefficient. The large camber (more curved upper surface) makes the airflow acceleration more significant and generates a stronger pressure difference (greater suction on the upper surface), thereby generating greater lift at the same speed and wing angle of attack Alfa.

[0031] (5) The airfoil of the present invention can reduce the takeoff / landing speed of the aircraft and generate sufficient lift at low airspeed, and is also suitable for short takeoff and landing aircraft;

[0032] (6) The airfoil of the present invention can improve the maneuverability of the aircraft. High lift means that a smaller turning radius and stronger instantaneous turning ability can be achieved;

[0033] (7) The airfoil of the present invention has very good gliding performance and also has excellent climb performance, especially for improving the aircraft's loiter time record. Attached Figure Description

[0034] Figure 1 This is a schematic diagram of the overall structure of the present invention;

[0035] Figure 2 This is a comparison of the lift coefficient and drag coefficient curves of the airfoil of the present invention with those of the conventional airfoils CLARK Y and E374;

[0036] Figure 3 A comparison diagram of the lift-to-drag ratio curves and pitching moment curves of the airfoil of the present invention with those of the conventional airfoils CLARK Y and E374;

[0037] Figure 4 The lift coefficient and drag coefficient curves of the CLARK Y airfoil at different Reynolds numbers;

[0038] Figure 5 The curves show the lift coefficient and drag coefficient of the airfoil E374 at different Reynolds numbers;

[0039] Figure 6 The curves show the lift coefficient and drag coefficient of the airfoil (WY4560) of this invention at different Reynolds numbers. Detailed Implementation

[0040] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments.

[0041] like Figures 1 to 6 As shown, this invention discloses a high lift-to-drag ratio, long-endurance, low-speed airfoil, named WY4560, specifically designed for long-endurance, low-speed applications. Its specific airfoil structure is as follows: the horizontal line passing through the left endpoint of the airfoil is the X-axis, with its direction extending from the leading edge to the trailing edge and coinciding with the chord line; the vertical line passing through the left endpoint of the airfoil is the Y-axis, perpendicular to the X-axis and pointing in the direction of the airfoil's mid-curve curvature.

[0042] The equation for the upper edge curve of the airfoil is expressed as follows:

[0043] When x∈[0, 2.5], y U =1.329400x + -0.309680x 2 + 0.032832x 3 ;

[0044] When x∈(2.5, 7.5], y U =0.543000 + 0.603067x + -0.025920x 2 +0.000789x 3 ;

[0045] When x∈(7.5, 15], y U =0.506800 + 0.563793x + -0.015268x 2 +0.000153x3 ;

[0046] When x∈(15, 25], y U =1.412000 + 0.430667x + -0.009040x 2 +0.000061x 3 ;

[0047] When x∈(25, 40], y U =3.061000 + 0.266890x + -0.003699x 2 +0.000004x 3 ;

[0048] When x∈(40, 60], y U =5.323000 + 0.163367x + -0.002490x 2 +0.000003x 3 ;

[0049] When x∈(60, 80], y U =9.026000 + 0.015850x + -0.000590x 2 + -0.000005x 3 ;

[0050] When x∈(80, 100], y U = (-9.56666666666675e-5)x 3 +0.0243450000000001x 2 - 2.24478333333334x + 76.9700000000007;

[0051] The equations for the lower edge curve of the airfoil are as follows:

[0052] When x∈[0, 2.5], y L =-0.684000x + 0.305120x 2 + -0.033664x 3 ;

[0053] When x∈(2.5, 7.5], y L =-0.405000 + -0.037267x + 0.030480x 2 + -0.001365x 3 ;

[0054] When x∈(7.5, 15], y L =-0.982200 + 0.195093x + -0.000008x 2 + -0.000063x 3 ;

[0055] When x∈(15,25], y L =-1.127000 + 0.232467x + -0.002880x 2 +0.000005x 3 ;

[0056] When x∈(25,40], y L =-1.307000 + 0.256167x + -0.003890x 2 +0.000019x 3 ;

[0057] When x∈(40,60], y L =-2.397000 + 0.283883x + -0.003475x 2 +0.000009x 3 ;

[0058] When x∈(60,80], y L =-6.695000 + 0.461717x + -0.005885x 2 +0.000019x 3 ;

[0059] When x∈(80,100], y L =-120.215000 + 4.291650x + -0.048795x 2 +0.000179x 3 ;

[0060] Where x is the x-coordinate, x∈[0, 100], y U Let y be a point on the upper edge curve of the airfoil. L Let C be a point on the lower edge curve; define the chord length as C, then the maximum relative camber of the airfoil is 6%C; the maximum relative thickness of the airfoil is 4.5%C.

[0061] In this invention, the origin of the coordinate system containing the airfoil is defined as the leading edge of the chord. The X-axis coincides with the chord line and points from the leading edge to the trailing edge of the airfoil. The Y-axis is perpendicular to the X-axis and points in the direction of the curvature of the airfoil's arc. The chord length of the airfoil is defined as 100, allowing for proportional scaling up or down designs.

[0062] In this embodiment, with the chord length x ranging from [0, 100], the dimensions of the airfoil are shown in Table 1 below.

[0063] Table 1

[0064] X 0 1.25 2.5 5 7.5 10 15 20 25 <![CDATA[Upper arc y U > 0 1.242 1.901 3.009 3.941 4.771 6.045 6.9 7.487 <![CDATA[Lower arc y L > 0 -0.444 -0.329 0 0.454 0.905 1.73 2.413 2.968 X 30 40 50 60 70 80 90 95 100 <![CDATA[Upper arc y U > 7.852 8.087 7.683 6.881 5.701 4.214 2.393 1.407 0.275 <![CDATA[Lower arc y L > 3.399 3.953 4.193 3.998 3.42 2.477 1.285 0.587 0

[0065] The airfoil of this invention can reduce the weight of the wing structure. Its thin and curved shape results in a smaller internal cavity for the wing, optimizing the aircraft's weight, but also concentrating the stress structure. The airfoil has a very smooth leading edge, which enables the wing to generate a suitable turbulent boundary layer, delaying the occurrence of airflow separation. Even if boundary layer separation occurs, it tends to gradually progress from the trailing edge forward, resulting in a gentler stall process, high safety, and good stall performance.

[0066] The airfoil of this invention has an extremely high maximum lift coefficient. Its large camber (more curved upper surface) significantly accelerates the airflow, generating a stronger pressure differential (greater suction on the upper surface), thus producing greater lift at the same speed and wing angle of attack (Alfa). This allows for reduced takeoff / landing speeds, generating sufficient lift at low airspeeds, and is also suitable for short takeoff and landing aircraft. It improves aircraft maneuverability; high lift means a smaller turning radius and stronger instantaneous turn capability. It exhibits excellent gliding performance while also providing superior climb performance, particularly beneficial for improving aircraft loiter time records.

[0067] Comparative Example:

[0068] The airfoil WY4560 of this invention is compared with the conventional airfoils CLARK Y and E374. For example... Figure 2 As shown, at a Reynolds number Re of 67,000, within the range of wing angle of attack Alfa from -1 to 8 degrees, the lift coefficient Cl of the airfoil WY4560 of this invention is higher than the former two; while within the range of wing angle of attack Alfa from 1 to 6 degrees, the drag coefficient Cd of the WY4560 airfoil is lower than the former two.

[0069] like Figure 3As shown, also at a Reynolds number Re of 67000, in the lift-to-drag ratio curve Cl / Cd, within the range of wing angle of attack Alfa from -1 to 6 degrees, the lift-to-drag ratio curve of the WY4560 airfoil of this invention is above the other two, which means better lift-to-drag performance. In the pitch moment curve Cm / a, the WY4560 curve is consistently at the bottom. The smaller the pitch moment coefficient Cm, the weaker the "tendency" or "effectiveness" to make the aircraft pitch up or down.

[0070] The pitch moment coefficient Cm is calculated using the formula: Cm = M / (q * S * c). Here, M is the pitch moment, and c is the mean aerodynamic chord length, which is crucial for analyzing the longitudinal static stability of an aircraft. When an aircraft using the WY4560 airfoil is affected by airflow fluctuations, and the wing angle of attack changes, a negative Cm value typically indicates that the aircraft has a "nose-down" moment, contributing to flight stability. Cm directly determines the longitudinal stability and maneuverability of the aircraft.

[0071] Stability: A stable aircraft, when disturbed (such as by an updraft causing the nose to pitch up), should automatically generate a restoring moment to return it to its original attitude. This is manifested by a negative slope (dCm / d@<0) in the curve of Cm as a function of angle of attack α. The negative slope is a measure of longitudinal static stability.

[0072] Handling: Changes in Cm are the basis for pilots or flight control systems to alter the aircraft's pitch attitude by manipulating control surfaces such as elevators and all-moving horizontal stabilizers. Pulling back on the control stick causes the elevators to deflect upwards, changing the lift of the horizontal stabilizers and thus producing a negative change in Cm, causing the nose to rise.

[0073] The smaller the Cm value, the weaker the "strength" or "effectiveness" of the aircraft's rotational tendency around its horizontal axis (whether pitching up or down). To determine whether the pitching or down tendency is weakening, the sign of Cm must be considered. To make the aircraft pitch up, a negative Cm needs to be created. If this negative Cm becomes "smaller" (e.g., from -0.15 to -0.05), it means the pitching operation is ineffective, and the pitching tendency is weakening. To make the aircraft pitch down, a positive Cm needs to be created. If this positive Cm becomes "smaller" (e.g., from +0.15 to +0.05), it means the pitching operation is ineffective, and the pitching tendency is weakening.

[0074] like Figures 4 to 6 As shown, within the Reynolds number range of 60,000 to 120,000, and within the wing angle of attack Alfa range of -3 to 4 degrees, comparing the three airfoils, the airfoil WY4560 of this invention has a larger lift coefficient Cl. It is evident that the airfoil of this invention exhibits a higher lift-to-drag ratio at lower Reynolds numbers, possessing a higher lift coefficient while maintaining a lower drag coefficient.

[0075] In the description of the embodiments of the present invention, it should be noted that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", and "outer" indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing the embodiments of the present invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the embodiments of the present invention.

[0076] The specific embodiments described in this invention are merely illustrative of the invention and are not intended to limit it. Those skilled in the art can make modifications to these embodiments without contributing any inventive step after reading this specification, but such modifications are protected by patent law as long as they fall within the scope of the claims of this invention.

Claims

1. A high lift-to-drag ratio, long-endurance, low-speed airfoil, characterized in that, The horizontal line passing through the left end point of the airfoil is taken as the X-axis, and the direction of the X-axis is from the leading edge of the airfoil to the trailing edge of the airfoil, coinciding with the chord line of the airfoil; the vertical line passing through the left end point of the airfoil is taken as the Y-axis, and the Y-axis is perpendicular to the X-axis and points in the direction of the curvature of the upper edge of the airfoil. The equation for the upper edge curve of the airfoil is as follows: When x∈[0, 2.5], y U =1.329400x + -0.309680x 2 + 0.032832x 3 ; When x∈(2.5, 7.5], y U =0.543000 + 0.603067x + -0.025920x 2 + 0.000789x 3 ; When x∈(7.5, 15], y U =0.506800 + 0.563793x + -0.015268x 2 + 0.000153x 3 ; When x∈(15, 25], y U =1.412000 + 0.430667x + -0.009040x 2 + 0.000061x 3 ; When x∈(25, 40], y U =3.061000 + 0.266890x + -0.003699x 2 + 0.000004x 3 ; When x∈(40, 60], y U =5.323000 + 0.163367x + -0.002490x 2 + 0.000003x 3 ; When x∈(60, 80], y U =9.026000 + 0.015850x + -0.000590x 2 + -0.000005x 3 ; When x∈(80, 100], y U = (-9.56666666666675e-5)x 3 + 0.0243450000000001x 2 -2.24478333333334x + 76.9700000000007; The equations for the lower edge curve of the airfoil are as follows: When x∈[0, 2.5], y L =-0.684000x + 0.305120x 2 + -0.033664x 3 ; When x∈(2.5, 7.5], y L =-0.405000 + -0.037267x + 0.030480x 2 + -0.001365x 3 ; When x∈(7.5, 15], y L =-0.982200 + 0.195093x + -0.000008x 2 + -0.000063x 3 ; When x∈(15,25], y L =-1.127000 + 0.232467x + -0.002880x 2 + 0.000005x 3 ; When x∈(25,40], y L =-1.307000 + 0.256167x + -0.003890x 2 + 0.000019x 3 ; When x∈(40,60], y L =-2.397000 + 0.283883x + -0.003475x 2 + 0.000009x 3 ; When x∈(60,80], y L =-6.695000 + 0.461717x + -0.005885x 2 + 0.000019x 3 ; When x∈(80,100], y L =-120.215000 + 4.291650x + -0.048795x 2 + 0.000179x 3 ; Where x is the x-coordinate, x∈[0, 100], y U Let y be a point on the upper edge curve of the airfoil. L Let C be a point on the lower edge curve; define the chord length as C, then the maximum relative camber of the airfoil is 6%C; the maximum relative thickness of the airfoil is 4.5%C.

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