A Design Method for Wide-Speed-Range Airfoil by Integrating Waverider Profile and Low-Speed Airfoil

Abstract

The present invention discloses a wide-speed-range airfoil design method integrating a waverider profile and a low-speed airfoil, which relates to the field of aircraft airfoil design and includes: S1. Based on the lift generation mechanism of the waverider compression characteristics in the high-speed stage and the flow-around characteristics in the low-speed stage, the front section of the airfoil is taken as the waverider profile of the high-speed airfoil, and the rear section of the airfoil is taken as the rear half of the low-speed airfoil; S2. According to the mission characteristics of the aircraft, the nonlinear weight method is used to calculate the distribution ratio of the waverider profile and the low-speed airfoil to obtain the basic airfoil; S3. The grid parameterization deformation is performed on the predetermined areas before and after the maximum thickness position of the basic airfoil to complete the wide-speed-range airfoil optimization design. The present invention provides a wide-speed-range airfoil design method integrating a waverider profile and a low-speed airfoil, which fully combines the high lift-to-drag ratio of the high-speed waverider airfoil and the high lift coefficient of the low-speed airfoil, and meets the airfoil requirements for the design of wide-speed-range aircraft by reasonably selecting the ratios of the waverider airfoil and the low-speed airfoil.

Description

Technical Field

[0001] The present invention relates to the field of aircraft airfoil design, and more specifically, to a wide-speed range airfoil design method in which a waverider section is integrated with a low-speed airfoil. Background Art

[0002] An airfoil usually refers to the cross-sectional shape of an aircraft wing, tail wing, missile wing surface, helicopter rotor blade, and propeller blade that is parallel to the aircraft's symmetry plane or perpendicular to the leading edge (or the line connecting the 1 / 4 chord length points), also known as a wing section or blade section.

[0003] Wide-speed range aircraft are highly integrated and fused with aerospace technology and multiple disciplines, and have the characteristics of high flight speed, short response time, long range and short flight time, etc., and advanced aerodynamic layout design is the forerunner of aircraft design and the key core technology of wide-speed range aircraft design. Specifically, aerodynamic performance is an important factor affecting the range and a key link in the overall design of the aircraft. Lift-to-drag ratio is an important parameter for achieving the range index of the aircraft. How to improve the lift-to-drag ratio characteristics is a key issue faced in aircraft design. Therefore, the advanced aerodynamic layout design of high-speed aircraft faces huge difficulties and there is a "lift-to-drag ratio barrier".

[0004] The waverider configuration is an effective attempt to pursue a high lift-to-drag ratio and break through the "lift-to-drag ratio barrier". Its principle is to limit the high-pressure airflow after the shock wave to the lower surface of the aircraft, and not allow it to bypass the leading edge and leak to the upper surface of the aircraft, so as to achieve pressure closure on the upper and lower surfaces, thereby obtaining a much higher lift-to-drag ratio than ordinary shapes under the design state. Wide-speed range aircraft must not only meet a larger lift coefficient at low speed, but also have a larger lift-to-drag ratio when the lift-to-weight balance is achieved. Smaller flight resistance can not only reduce the size and weight of the engine, but also help the aircraft accelerate the climb at low speed, reducing fuel consumption and flight time during the climb.

[0005] However, the use of waveriders faces the following two major problems when flying at low speeds:

[0006] First, the lift-drag ratio at low speed is relatively low, and the aircraft has a large flight resistance (although the lift coefficient is larger at low speed);

[0007] Second, the stall angle of attack is small during low-speed flight, and high-speed aircraft are prone to stall during low-speed flight. The excellent aerodynamic performance of low-speed aircraft design usually requires a large aspect ratio wing and a suitable low-speed airfoil. Therefore, the requirements of high-speed aircraft for aerodynamic layout are in great conflict with those of low-speed aircraft, which is a technical bottleneck for realizing the design of wide-speed aircraft. How to achieve an aerodynamic layout that takes into account both high and low speeds has always been a cutting-edge technology in aircraft design, but no one has yet proposed a wing aerodynamic layout design that can enable the aircraft to take into account both high-speed and low-speed flight needs. Summary of the invention

[0008] One object of the present invention is to solve at least the above problems and / or deficiencies and provide at least the advantages described hereinafter.

[0009] To achieve these objects and other advantages of the present invention, a wide-speed airfoil design method integrating a waverider profile and a low-speed airfoil is provided, including:

[0010] S1. Based on the lift generation mechanism of the waverider compression characteristics in the high-speed stage and the flow-around characteristics in the low-speed stage, take the waverider profile of the high-speed airfoil for the front section of the airfoil and the latter half of the low-speed airfoil for the rear section of the airfoil;

[0011] S2. According to the mission characteristics of the aircraft, use the non-linear weight method to calculate the allocation ratio of the waverider profile and the low-speed airfoil to obtain the basic airfoil;

[0012] S3. Select a predetermined area before and after the maximum thickness position of the basic airfoil for grid parametric deformation to complete the wide-speed airfoil optimization design for different wide-speed flight requirements.

[0013] Preferably, in S1, the design process of the front section of the airfoil includes:

[0014] S10. Extract the windward surface of the waverider streamline under the predetermined working conditions;

[0015] S11. Obtain the conical streamline on the windward surface through streamline tracking, and use the conical streamline as the lower surface of the front section of the airfoil;

[0016] S12. Use the von Kármán curve as the upper surface of the front section of the airfoil.

[0017] Preferably, in S2, the non-linear weight method is to decompose the influence of controlling the high-speed airfoil and the low-speed airfoil by constructing the following airfoil ratio objective function to decompose the influence of controlling the high-speed airfoil and the low-speed airfoil:

[0018]

[0019] In the above formula, e z is the reference function for evaluating the high-speed airfoil, the reference function for evaluating the low-speed airfoil, z is the independent variable, f ( z ) is the weight coefficient, and f ( z ) designs the allocation ratio of the high-speed airfoil and the low-speed airfoil through the following non-linear weighting function :

[0020]

[0021] In the above formula, α 、β is the shape parameter for controlling the distribution pattern, and .

[0022] Preferably, in S3, the processing flow of the grid parameterization deformation includes:

[0023] S30. Select the 15% area before and after the maximum thickness position of the basic airfoil, construct a control area using non-uniform FFD control nodes, and realize airfoil parameterization through the change of control nodes;

[0024] S31. Define the change range of the control area nodes, extract samples of the control nodes using the DOE method, and in the CFD simulation software, change the control nodes in the way of FFD parameterization, so as to realize the deformation design of the grid area.

[0025] Preferably, in S31, in the FFD parameterization, the B-spline function as shown in the following formula is adopted for grid deformation parameterization:

[0026]

[0027] In the above formula, C(s,t) is the airfoil surface coordinate, i is the control node serial number in the x direction, l is the number of control nodes in the x direction, j is the control node serial number in the y direction, m is the number of control nodes in the y direction, P i,j is the FFD initial airfoil coordinate, N i,p is x the basis function in the N j,q is y the basis function in the s , t ) is the calculated local coordinate of the control node, and , p , q are respectively x , y the corresponding orders in the

[0028] Preferably, in S3, after the wide-speed domain airfoil optimization design, the parameters of the optimized airfoil are as follows:

[0029] The maximum thickness is 8.59%, and the maximum thickness is located at 61.09% of the airfoil;

[0030] The airfoil camber is 0.995%, and the airfoil leading edge radius is 3 mm.

[0031] Preferably, in S3, after the wide-speed range airfoil is optimized and designed, the parameters of the optimized airfoil are as follows:

[0032] The maximum thickness is 4.64%, and the maximum thickness is located at 64.1% of the airfoil;

[0033] The maximum camber is 1.01%, and the maximum camber is located at 49.1% of the airfoil;

[0034] The leading edge blunt radius gradually increases along with the leading edge curve, and the radius of the airfoil stagnation point is 1.908 mm. The curvature radius of other places in the airfoil leading edge area except the stagnation point is about 3 mm.

[0035] Preferably, in S3, after the wide-speed range airfoil is optimized and designed, the parameters of the optimized airfoil are as follows:

[0036] The maximum thickness is 4.0%, and the maximum thickness is located at 63.4% of the airfoil;

[0037] The maximum camber is 1.01%, and the maximum camber is located at 48.6% of the airfoil;

[0038] The leading edge blunt radius of the airfoil gradually increases along with the leading edge curve. The radius of the airfoil stagnation point is 1.6 mm. The curvature radius of other places in the airfoil leading edge area except the stagnation point is about 3 mm.

[0039] The present invention has at least the following beneficial effects: The wide-speed range airfoil design method proposed by the present invention fully combines the high lift-to-drag ratio of the high-speed waverider airfoil and the high lift coefficient of the low-speed airfoil, and forms an airfoil that meets the design requirements of the wide-speed range aircraft by reasonably selecting the ratio of the waverider airfoil and the ratio of the low-speed airfoil, providing a new design method for the optimization design of the wide-speed range airfoil.

[0040] Furthermore, the airfoil design method of the present invention has an obvious design mechanism. It does not blindly search for optimization through parameterization. It can adjust the ratio of the two basic airfoils according to the specific requirements of wide-speed range flight, and locally optimize the airfoil thickness, camber, etc. by adjusting the ratio. It can also directly optimize to obtain the camber and thickness of the airfoil. The designed airfoil has good high-speed and low-speed aerodynamic performance and can meet the design requirements of wide-speed range flight.

[0041] Other advantages, objectives and features of the present invention will be partially reflected by the following description, and partially will be understood by those skilled in the art through the research and practice of the present invention. Brief Description of the Drawings

[0042] Figure 1 Schematic diagram of the airfoil optimization design based on the FFD method of the present invention;

[0043] Figure 2Schematic diagram of the front and rear transition of the two-segment airfoil on the lower surface of the present invention;

[0044] Figure 3 Schematic diagram of the leading-edge blunting of the airfoil of the present invention;

[0045] Figure 4 Basic airfoil of the integrated design of the low-speed airfoil and the waverider profile in Example 1 of the present invention ( h = 8.59%);

[0046] Figure 5 Lift coefficient curve at Mach number 0.3 in Example 1 of the present invention ( h = 8.59%);

[0047] Figure 6 Lift-to-drag ratio curve at Mach number 0.3 in Example 1 of the present invention ( h = 8.59%);

[0048] Figure 7 Lift coefficient curve at Mach number 6.0 in Example 1 of the present invention ( h = 8.59%);

[0049] Figure 8 Lift-to-drag ratio curve at Mach number 6.0 in Example 1 of the present invention ( h = 8.59%);

[0050] Figure 9 Comparison diagram of the airfoil before and after optimization in Example 2 of the present invention;

[0051] Figure 10 Relative thickness distribution of the airfoil before and after optimization in Example 2 of the present invention;

[0052] Figure 11 Wide-speed airfoil after optimization in Example 2 of the present invention ( h = 4.86%);

[0053] Figure 12 Airfoil section after optimization in Example 2 of the present invention ( h = 4.86%);

[0054] Figure 13 Variation of lift coefficient with angle of attack at Mach number 0.3 in Example 2 of the present invention ( h = 4.86%);

[0055] Figure 14 Variation of moment coefficient with angle of attack at Mach number 0.3 in Example 2 of the present invention ( h = 4.86%);

[0056] Figure 15 Variation of lift-to-drag ratio with angle of attack at Mach number 0.3 in Example 2 of the present invention ( h = 4.86%);

[0057] Figure 16 The lift coefficient of Example 2 of the present invention at Mach number 6.0 varies with the angle of attack ( h = 4.86%);

[0058] Figure 17 The moment coefficient of Example 2 of the present invention at Mach number 6.0 varies with the angle of attack ( h = 4.86%);

[0059] Figure 18 The lift-to-drag ratio of Example 2 of the present invention at Mach number 6.0 varies with the angle of attack ( h = 4.86%);

[0060] Figure 19 The pressure contour of Example 2 of the present invention at Mach number 0.3 and AOA = 6.0 ( h = 4.86%);

[0061] Figure 20 The pressure contour of Example 2 of the present invention at Mach number 6.0 and AOA = 2.0 ( h = 4.86%);

[0062] Figure 21 The optimized wide-speed-range airfoil of Example 3 of the present invention ( h = 4.0%);

[0063] Figure 22 The airfoil with optimized thickness of Example 3 of the present invention ( h = 4.0%);

[0064] Figure 23 The lift coefficient of Example 3 of the present invention at Mach number 0.3 varies with the angle of attack ( h = 4.0%);

[0065] Figure 24 The lift-to-drag ratio of Example 3 of the present invention at Mach number 0.3 varies with the angle of attack ( h = 4.0%);

[0066] Figure 25 The lift coefficient of Example 3 of the present invention at Mach number 6.0 varies with the angle of attack ( h = 4.0%);

[0067] Figure 26 The lift-to-drag ratio of Example 3 of the present invention at Mach number 6.0 varies with the angle of attack ( h = 4.0%);

[0068] Figure 27 Schematic diagram of the weight ratio for designing the high- and low-speed airfoils of the present invention using a non-linear weighting function;

[0069] Figure 28 Schematic diagram of the weighted sum result of the high- and low-speed airfoils after the design of the present invention based on the weight ratio. Detailed implementation manners

[0070] The present invention will be further described in detail below with reference to the accompanying drawings, so that those skilled in the art can implement it according to the description in the specification.

[0071] As the basis for the aerodynamic layout design of an aircraft, the airfoil plays a key role in the aerodynamic performance of the aircraft. Conducting innovative design of wide-speed airfoils is one of the key problems faced by aircraft design. Therefore, in response to the requirements of the aerodynamic performance of wide-speed aircraft, the present invention provides a wide-speed airfoil design method based on the coupling design of wave-riding characteristics and low-speed airfoils, and obtains a wide-speed airfoil with a high lift-to-drag ratio and a high lift coefficient at the low-speed takeoff and landing stage and a high lift-to-drag ratio at high speed. The wide-speed airfoil design process will be described in detail below:

[0072] Generally speaking, the wing loading of a wide-speed aircraft is usually between 350-600 kg / m 2 , and the lift-weight balance that the aircraft needs to meet during the entire flight stage is as follows:

[0073]

[0074] Among them, M is the mass of the aircraft, g is the acceleration due to gravity, ρ is the atmospheric density, v is the flight speed of the aircraft, S is the wing area, C L is the lift coefficient.

[0075] Assume that the Mach number at high-speed flight is 6.0, the altitude is 28 km, the Mach number at horizontal takeoff is 0.3, the air density at the ground (altitude 1 km) is taken as 1.1 kg / m3, the speed of sound is taken as 336 m / s, the density at an altitude of 28 km is taken as 0.02507 kg / m3, the speed of sound is taken as 300 m / s, and the mass change during takeoff is not considered. Then the required lift coefficient ratio of the aircraft is:

[0076]

[0077] In the above formula, C L，1km is the lift coefficient at an altitude of 1 km, C L，28km is the lift coefficient at an altitude of 28 km, ρ 28km is the atmospheric density at an altitude of 28 km, ρ 1km is the atmospheric density at an altitude of 1 km, v28km is the flight speed of the aircraft at an altitude of 28 km, v 1km is the flight speed of the aircraft at an altitude of 1 km.

[0078] To improve the cruise performance of the aircraft, usually the lift coefficient during high-speed flight is taken at the flight state corresponding to the maximum lift-to-drag ratio, and the angle of attack corresponding to the maximum lift-to-drag ratio is usually between 3° and 6°. To ensure the safety of takeoff and landing, the angle of attack during low-speed flight does not exceed 15°. If for the same aircraft, assuming the ratio of the angles of attack corresponding to high and low speeds is 1 / 3, and the lift curve slope during the low-speed stage is 0.055, then the lift curve slope during the high-speed stage is 0.0227. Through analysis, it can be seen that the same aircraft needs to satisfy a large change in the lift coefficient in the high-speed and low-speed stages, and needs to satisfy a certain stall angle of attack at low speed. Then, when designing a wide-speed-range aircraft, it is necessary to first complete the design of the wide-speed-range airfoil. The camber and thickness of the airfoil are important design parameters that affect the lift coefficient, lift-to-drag ratio, and stall angle of attack of the airfoil. When the airfoil is designed with a certain camber, it will reduce the size of the separated area on the leeward side at the same angle of attack and increase the lift coefficient at a small angle of attack. However, due to the increase in camber, it will also cause an increase in the frontal area, and the lift-to-drag ratio of the airfoil will decrease significantly during the high-speed stage. Therefore, how to achieve a large lift coefficient and a high lift-to-drag ratio at a large angle of attack in the low-speed stage of the airfoil, while improving the lift-to-drag ratio of the airfoil at a small angle of attack in the high-speed stage is the difficulty in the design of the wide-speed-range airfoil.

[0079] To achieve the above invention purpose, the present invention draws on the lift generation mechanisms of wave-riding compression and low-speed flow around the body. By extracting the windward surface of the typical wave-riding streamline (it should be noted that the typical streamline here refers to the wave-riding streamline obtained under the flight conditions of a Mach number of 6, an angle of attack of 0°, and a shock wave angle of 12°, and the shock wave angle can be adjusted according to design requirements) and the rear half of the low-speed airfoil. Among them, as Figures 1 - 2 shown (it should be noted that Figure 1 the deformed airfoil is represented by the orange line, the reference airfoil is represented by the blue line, and various colored dots are used to represent the FFD control points, Figure 2 the lower surface of the airfoil after arc transition is represented by the purple line, the tail of the wave-riding profile is represented by the light blue line, and the front end of the low-speed airfoil is represented by the blue line), the front half of the airfoil is designed using the wave-riding concept. The lower surface is the conical streamline obtained through streamline tracking to reduce shock leakage during the high-speed stage. The upper surface of the wave-riding uses the von Kármán curve to achieve the effect of expanding the volume and improve the low-speed performance of the airfoil. The front section of the airfoil designed in this way is through the integrated design of the wave-riding profile and the von Kármán curve, as Figure 3 shown ( Figure 3The medium purple line represents the upper surface of the leading edge of the blunted airfoil, the light blue line represents the lower surface of the leading edge of the blunted airfoil, and the red line represents the blunted leading edge). Its leading edge radius can be adjusted according to the flight mission, and different leading edge radii can be satisfied to balance the influence of aerodynamic heat and aerodynamic drag;

[0080] The second half of the airfoil adopts the second half of the low-speed airfoil. On the one hand, it increases the lift of the airfoil in the low subsonic stage, and on the other hand, it can reduce the bottom drag of the airfoil during high-speed flight. That is, a fusion design method that fully combines the high performance of the waverider in the high-speed stage and the high performance of the low-speed airfoil is proposed. The front section of the airfoil is taken as the high-speed waverider profile, and the rear section of the airfoil is taken as the second half of the low-speed airfoil, so that the airfoil makes full use of the compression characteristics of the waverider in the high-speed stage to meet the high lift-to-drag ratio requirements; in the low-speed stage, it makes full use of the flow-around characteristics of the airfoil to relieve the separation at large angles of attack and meet the high-lift and high-lift-to-drag ratio requirements in the low-speed stage, so as to balance the high-speed flight and low-speed flight performance.

[0081] Furthermore, in order to enable the airfoil to adjust the proportion of the waverider profile and the low-speed airfoil segment according to the mission characteristics of the aircraft, so as to adapt to different aircraft design requirements, the present invention also designs a non-linear weight calculation method for allocating the proportion of the waverider profile and the low-speed airfoil to determine the allocation ratio of the high- and low-speed airfoils according to the usage range of the airfoil. Specifically, the present invention proposes two basic functions to decompose and control the influence of the high-speed airfoil and the low-speed airfoil, and constructs the following design airfoil ratio objective function :

[0082]

[0083] In the above formula, e z is the reference function for evaluating the high-speed airfoil, is the reference function for evaluating the low-speed airfoil, z is the independent variable, f ( z ) is the weight coefficient. Considering that the maximum thickness of a general airfoil is in the range of 40% - 70% of the airfoil, the present patent constructs the following non-linear weighting function :

[0084]

[0085] In the above formula, α , β are the shape parameters for controlling the distribution pattern, and , where, α = 4, β = 3 airfoils are more inclined to high-speed flight, α = 3, β = 4 airfoils are more inclined to low-speed flight.

[0086] The characteristic of the above objective function is that the independent variablez The value range of is 0 - 1, which is mainly used to control the ratio of the high-speed airfoil along the chord length. Since the high-speed waverider airfoil increases with the increase of the chord length, that is, the maximum thickness of the airfoil is positively correlated with the ratio of the high-speed airfoil, the objective function of the design The maximum value of appears in the range of 50% - 80% of the independent variable, meeting the design requirements of the maximum thickness of the airfoil in engineering. This method realizes the fusion of two flight modes through non-linear weights and is applicable to the trade-off analysis of multi-stage tasks.

[0087] Furthermore, on the basis of the above basic airfoil, considering the waverider characteristics of the airfoil leading edge and the load design of the airfoil trailing edge, further optimization is carried out on the basic airfoil. Mainly, a 15% area before and after the position of the maximum thickness of the airfoil is selected for grid parameterization, FFD control nodes are set to control the deformation of the airfoil, the DOE method is used to extract samples of design parameters, and then the CFD method combined with the FFD method is used to complete the data calculation of the samples, including low-speed performance and high-speed performance.

[0088] Generally, there are three types of basis functions for FFD parameterization: Bernstein polynomials, B-spline functions, and NURBS functions. Among them, Bernstein polynomials are a global influence method, whose characteristic is that when changing one control node, all coordinates will change, and it does not have local support. B-spline functions and NURBS functions have local support and local deformation ability. According to the design idea of this patent, the B-spline function is selected here for grid deformation parameterization.

[0089] The specific implementation process of the B-spline function is as follows:

[0090] First, divide the control nodes for the area to be deformed. By changing the control nodes, the area near them can be deformed. The specific affected area depends on the order of the basis function. The implementation theory of the FFD parameters is as follows:

[0091]

[0092] Among them, i is the serial number of the control node in the x direction, l is the number of control nodes in the x direction, j is the serial number of the control node in the y direction, m is the number of control nodes in the y direction, P i,j is the initial non-uniform control node coordinate of the FFD, N i,p is x the basis function in the N j,q is y the basis function in the s , t) is the calculated local coordinates of the control node, and , p 、 q are respectively x 、 y the orders corresponding to the directions.

[0093] To construct the B-spline function, it is necessary to construct the knot vector in the x direction U , and its general formula is:

[0094]

[0095]

[0096] where i is the FFD control node number, p is the order, r is the number of vector nodes, n + 1 is the number of FFD control points, u is the local coordinate, u i is the value of the i-th vector node, u i+p is the value of the (i + p)-th vector node, and R represents an r + 1-dimensional real number set. Here, 0 / 0 = 0 is defined for program solving. The construction method of the knot vector in the y direction is the same as that in the x direction, which will not be described here.

[0097] Usually, the knot vector is given in the following way:

[0098]

[0099] where u 0 = u 1… = u p = 0, u n+1 = … = u n+p+1 = 1, and u p+1 、…、 u n The control nodes are set to increase in an arithmetic progression, and at the same time, it is necessary to satisfy u p+1 > u p , u n < u n+1 .

[0100] Since s is known, then Bi,0 is determined B 0,0 , B 1,0 , B 2,0 , B 3,0 , B 4,0 … are known values B 0,1 = s1 × B 0,0 + s2 × B 1,0 , B 1,1 = s1 × B 1,0 + s2 × B 2,0 and then B i,1 are all known B 0,2 = s1 × B 0,1 + s2 × B 1,2 , B 1,2 = s1 × B 1,1 + s2 × B 2,1 and then B i,2 are all known. It can be seen that starting from order 0, the basis functions of all control nodes can be obtained sequentially

[0101] Since the node distribution on the airfoil surface is in the form of a curve, it is usually difficult to accurately control local design parameters using uniformly distributed control nodes. Here, a non-uniform control node distribution form is adopted for mesh parameterization. Assuming the initial non-uniform control node coordinates are P i,j , and the calculated local coordinates are ([[]] s , t ), then on the premise that the control nodes, node vectors, and weighting factors remain unchanged, the corresponding airfoil coordinates become

[0102]

[0103] With the node vectors and weighting factors remaining unchanged, after changing the non-uniform control node coordinates to , the airfoil coordinates become

[0104]

[0105] Through the above design method, the wide-speed-range airfoil optimization design can be carried out to complete different airfoil designs for different wide-speed-range flight requirements.

[0106] Example 1:

[0107] Considering that the cruise state of a wide-speed-range aircraft is mainly in the high-speed stage, when integrating the waverider airfoil and the low-speed airfoil in the case, it is more inclined to the high-speed flight mode. The values of the shape parameters α and β are set ( α =4, β =3), and a non-linear weighting function is obtained. The objective function is calculated with respect to the independent variable value u to obtain the optimized allocation ratio. As shown in Figures 27 - 28 , it is selected that the proportion of the high-speed waverider airfoil is 64.5% and the proportion of the low-speed airfoil is 35.5%. Among them, the waverider airfoil is obtained by solving the Taylor-Maccoll conical flow field equation, and the low-speed airfoil is selected from a specific airfoil (such as: NACA 64(1)-212) and trimmed. In order to achieve the expansion design of the airfoil, the upper surface of the airfoil is designed in the form of a von Kármán curve.

[0108] Among them, the waverider airfoil selects the conical flow field with a free-stream Mach number of 6 and a shock wave angle of 12°. The benchmark flow field information is obtained by solving the Taylor-Maccoll equation, and then the streamline coordinates obtained by streamline tracing are used. The specific Taylor-Maccoll equation is as follows:

[0109]

[0110] Among them, is the critical sound speed of the free stream, c is the sound speed, is the dimensionless radial velocity of the critical sound degree, is the dimensionless tangential velocity of the critical sound degree, θ is the polar angle.

[0111]

[0112]

[0113] In the above formula, T t is the total temperature of the free stream, R 0 is the gas constant of air, k is the specific heat ratio of air, T is the static temperature of the free stream, is the free-stream Mach number.

[0114] The second half of the airfoil is the modified low-speed airfoil NACA 64(1)-212. The maximum thickness of this airfoil is 9.97%, and the position of the maximum thickness is at 40% of the airfoil chord length.

[0115] When considering viscosity for the waverider, the state at which the maximum lift-to-drag ratio occurs is no longer the design state. Generally, at a positive angle of attack, as the angle of attack increases, a leeward region will form on the back of the aircraft, and this region is a region of expansion waves. However, due to the relatively high flight altitude of the transonic aircraft and the relatively low ambient pressure, the pressure reduction range of the expansion waves is limited, so the contribution to increasing the lift of the aircraft is not obvious. Therefore, from the perspectives of improving the aerodynamic performance of the aircraft, increasing the volume, and strengthening the structural safety, etc., a volume increase design is carried out on the upper surface of the waverider, and the design method used is the von Kármán curve. The specific profile design equation is as follows:

[0116]

[0117] Among them, H refers to the distance from the end coordinate of the given von Kármán curve to the axis, L represents the length of the von Kármán curve in the x direction, h x is the distance from the curve in the x direction to the axis. H =0 indicates that the upper surface of the waverider is parallel to the oncoming flow.

[0118] The next step is the specific design process of the airfoil:

[0119] First, the length of the waverider airfoil is selected as 1000 mm. The upper surface is designed with the von Kármán curve. The designed exit thickness is 121.6 mm. The second half uses the design idea of the transonic airfoil. The trailing edge of the airfoil uses a rear-loaded design. The upper surface of the airfoil uses an arc design method. The length of the second half of the airfoil is intercepted as 548.7 mm. The trailing edge thickness of the airfoil is 4.673 mm. The initial airfoil designed by this method has a length of 1548.7 mm, a maximum thickness of 121.6 mm, and a relative thickness of 7.85%. In order to ensure the smooth continuity of the lower surface airfoil, 98% of the position of the front half of the waverider airfoil on the lower surface and 3% of the position of the low-speed airfoil on the lower half of the lower surface are cut, and then an arc transition is carried out through a spline curve.

[0120] Considering that the leading edge of the waverider airfoil is pointed, it cannot meet the processing and design requirements for both high-speed flight and low-speed flight. At the same time, for high-speed flight, the heat protection requirements also need to be considered. Generally, the larger the leading edge radius, the better the heat protection effect, but the larger the leading edge radius, the greater the resistance. For low-speed flight, the design of the leading edge radius will also affect the stall characteristics of the airfoil. The larger the radius, the better the stall performance. In this patent, the leading edge of the airfoil is bluntly processed, and an arc transition is also carried out at the connection of the leading and trailing edges of the lower surface of the airfoil, asFigure 4 as shown in Figure 4 where the x-axis is the chordwise coordinate of the airfoil and the y-axis is the airfoil curve coordinate perpendicular to the chordwise direction. After design, the length of the airfoil is 1410 mm, the maximum absolute thickness is 121.05 mm, and the maximum thickness of the airfoil ( h is the maximum thickness) is 8.59%, the position of the maximum thickness is 61.09%, the camber of the airfoil is 0.995%, the normalized leading edge radius is 3.5 mm, the trailing edge angle is 14.82°, and the designed airfoil is named HyAerofoil0859.

[0121] To study the influence of the leading edge radius of the airfoil on the aerodynamic performance of the airfoil, this patent conducts a comparative analysis on the blunt radii of 1 mm, 3 mm, and 5 mm at the leading edge, and the analysis states correspond to Mach numbers of 0.3 and 6.0, and the curves of the lift coefficient and the lift-to-drag ratio under different design conditions are obtained.

[0122] Figure 5 The comparative curve of the lift coefficient varying with the angle of attack under different leading edge radii at Mach number 0.3 (where Figure 5 the red solid line is the lift coefficient curve when the leading edge radius R = 1 mm, the purple dashed line is the lift coefficient curve when the leading edge radius R = 3 mm, and the green dashed line is the lift coefficient curve when the leading edge radius R = 5 mm). At Mach number 0.3 and an angle of attack of 10°, the lift coefficients corresponding to the leading edge radii of 1 mm, 3 mm, and 5 mm are 1.017, 0.930, and 0.928 respectively. It can be seen from the lift coefficient curve that the lift line slope of this airfoil in the low-speed section becomes smaller as the angle of attack increases. This may be because the relatively small leading edge radius and thickness cause the separation region on the airfoil surface to become larger when the angle of attack increases.

[0123] Figure 6 Shown is the comparative curve of the lift-to-drag ratio varying with the angle of attack under different leading edge radii at Mach number 0.3. The maximum lift-to-drag ratios corresponding to the leading edge radii of 1 mm, 3 mm, and 5 mm are 20.53, 23.47, and 22.62 respectively, and the corresponding lift coefficients are 0.408, 0.403, and 0.404 respectively, and the angle of attack is 0°.

[0124] Figure 7 Shown is the comparative curve of the lift coefficient varying with the angle of attack under different leading edge radii at Mach number 6.0. At Mach number 6.0, the lift coefficient increases as the angle of attack increases, and the lift line slope increases slightly as the angle of attack increases. This may also be related to the increase in the separation region on the back part to a certain extent, but the influence of the blunt radius size on the lift coefficient is not as large as that at low speeds.

[0125] Figure 8The figure shows the comparison curves of the lift-drag ratio varying with the angle of attack at Mach number 6.0 for different leading-edge radii. The maximum lift-drag ratios corresponding to leading-edge radii of 1 mm, 3 mm, and 5 mm are 4.287, 3.664, and 3.19 respectively, and the corresponding angles of attack are 4°, 6°, and 6° respectively. Through analysis, in the low-speed stage, when the blunting radius is greater than 3 mm, the change in the maximum lift-drag ratio becomes gentle. In the high-speed stage, in order to alleviate the aerodynamic heat load, the leading edge of the aircraft must meet a certain blunting radius. After weighing, the leading-edge radius in the design of a wide-speed-range airfoil is also preferably selected to be about 3 mm, which can take into account the aerodynamic requirements of both low speed and high speed. That is, when the leading-edge radius is 3 mm, at Mach number 0.3, the lift coefficient corresponding to an angle of attack of 10° is 0.93, the maximum lift-drag ratio is 23.47, and the corresponding drag coefficient is 0.4; at Mach number 6.0, the maximum lift-drag ratio is 3.664, and the corresponding angle of attack is 6°.

[0126] Example 2:

[0127] Considering that the flight speed of a wide-speed-range aircraft has to face low-speed, transonic, and high-speed stages, the aircraft puts forward higher requirements for lift augmentation and drag reduction throughout the full-speed range, and also puts forward design requirements for the leading-edge radius, airfoil thickness, and airfoil camber of the airfoil. Low-speed airfoils usually adopt a larger leading-edge radius, a larger airfoil thickness, and an appropriate airfoil camber to ensure a high lift coefficient, a high lift-drag ratio, and a large stall angle of attack at low speed, and the airfoil thickness is usually above 10%. The airfoil thickness of a fighter is usually about 4%, and the camber is small. The thickness of a waverider increases with the increase in the size of the aircraft, but the airfoil thickness usually adopted is controlled within 5%. The lift-drag ratio of a high-speed airfoil is also closely related to the leading-edge radius, maximum thickness, and area of the airfoil. Generally speaking, the wide-speed-range airfoil design should also choose an airfoil with a small thickness, a small leading-edge radius, and a small camber.

[0128] As Figures 11 - 12 shown, for the requirements of wide-speed-range flight, the thickness distribution and maximum thickness of the airfoil are optimized. After optimization using the FFD method, the maximum thickness of the airfoil is 4.86%, the position of the maximum thickness is at 64.1%, the maximum camber is 1.01%, the position of the maximum camber is 49.1%, and the airfoil area is 0.03041. The leading-edge blunting radius gradually increases along with the leading-edge curve. The radius at the stagnation point is 1.908 mm, and the curvature radius at other places in the leading-edge region is basically about 3 mm. The designed airfoil is named HyAerofoil0486.

[0129] Figure 9 The comparison of the airfoil profiles before and after optimization is given ( Figure 9Among them, the x-axis is the chordwise coordinate of the airfoil, the y-axis is the airfoil curve coordinate perpendicular to the chordwise direction. The orange solid line is the airfoil profile of Example 1, and the blue dashed line is the optimized airfoil profile of Example 2). It can be seen that since the leading-edge radius of the airfoil in Scheme 1 is also small and the trailing-edge thickness is also small, mainly the middle position of the airfoil is optimized, changing the maximum thickness of the airfoil, and the camber changes little.

[0130] Figure 10 The comparison of the relative thickness distribution of the airfoil before and after optimization is given ( Figure 10 Among them, the x-axis is the chordwise coordinate of the airfoil, the y-axis is the airfoil curve coordinate perpendicular to the chordwise direction. The orange solid line is the relative thickness distribution of the airfoil of Example 1, and the blue dashed line is the relative thickness distribution of the optimized airfoil of Example 2). The thickness distribution trend of the optimized airfoil is consistent with that of the original airfoil, and the thickness of the airfoil decreases significantly in the area around 64% of the maximum thickness position.

[0131] Figure 13 The curve of the lift coefficient of the optimized airfoil varying with the angle of attack at Mach number 0.3 is given. The variation law of the lift coefficient is consistent with that of the original airfoil. As the angle of attack increases, the slope of the lift line gradually becomes smaller. At an angle of attack of 10°, the lift coefficient increases, from the original 0.930 to 0.984. The increase in the lift coefficient of the optimized airfoil in the low-speed section is more conducive to the takeoff and landing of the aircraft.

[0132] Figure 14 The curve of the pitching moment coefficient of the optimized airfoil varying with the angle of attack at Mach number 0.3 is given. The zero-lift moment coefficient of the airfoil is about 0.054, that is, the airfoil has a nose-up moment by itself. As the angle of attack increases, the moment coefficient first decreases and then increases, and has an obvious nose-up moment at large angles of attack.

[0133] Figure 15 The curve of the lift-to-drag ratio of the optimized airfoil varying with the angle of attack at Mach number 0.3 is given. It can be seen that the maximum lift-to-drag ratio increases compared with the original airfoil, and the corresponding angle of attack is 0°. Analyzing, it is mainly due to the significant reduction of the resistance caused by the decrease in thickness. Taking an angle of attack of 10° as an example, the lift-to-drag ratio of the optimized airfoil is 5.435, and that of the original airfoil is 4.817. The maximum lift-to-drag ratio is 31.75, the corresponding angle of attack is 0°, and the lift coefficient is 0.338.

[0134] Figure 16 The curve of the lift coefficient of the optimized airfoil varying with the angle of attack at Mach number 6.0 is given. As the thickness of the airfoil decreases, the lift coefficient of the airfoil in the high-speed stage decreases, and the drag coefficient also decreases. The linearity of the lift coefficient varying with the angle of attack is good. However, the lift coefficient generally decreases compared with the original airfoil. Taking an angle of attack of 4° as an example, the lift coefficient decreases from 0.08 of the original airfoil to 0.0646 of the optimized airfoil.

[0135] Figure 17 The curve of pitching moment coefficient of the optimized airfoil varying with the angle of attack at Mach number 6.0 is given. Different from the low-speed stage, in the high-speed stage, as the angle of attack increases, the pitching moment coefficient of the airfoil gradually increases, and the nose-up trend becomes more obvious. The reason is that the action point of the lift is behind the moment reference point, and as the angle of attack increases, the lift increases, resulting in an increase in the moment.

[0136] Figure 18 The curve of lift-to-drag ratio of the optimized airfoil varying with the angle of attack at Mach number 6.0 is given. The maximum lift-to-drag ratio increases compared with the original airfoil, from 3.664 to 4.198. The corresponding angle of attack is 6°, and the corresponding lift coefficient is 0.0932.

[0137] Figure 19 and Figure 20 The pressure contour maps of the optimized airfoil at typical angles of attack at Mach numbers 0.3 and 6.0 are given respectively (at Mach number 0.3, the lift coefficient corresponding to the angle of attack of 10° is 0.984, the maximum lift-to-drag ratio is 31.75, and the corresponding lift coefficient is 0.338; at Mach number 6.0, the maximum lift-to-drag ratio is 4.198, and the corresponding angle of attack is 6°). It can be seen that in low-speed flow, the generation of lift of the airfoil mainly comes from the front half of the airfoil, and obvious high and low pressure regions are formed on the upper and lower surfaces of the airfoil leading edge. In the high-speed stage, the generation of lift mainly comes from the shock compression on the lower surface of the airfoil.

[0138] Example 3:

[0139] To further ensure that the designed airfoil has better adaptability, the airfoil thickness is further optimized and adjusted here. As Figures 21 - 22 shown, the maximum thickness of the airfoil is reduced from 4.64% to 4.0%. The position of the maximum thickness of the optimized airfoil is at 63.4%, the maximum camber is 1.01%, the position of the maximum camber is 48.6%, the leading-edge blunt radius gradually increases with the leading-edge curve, the radius of the stagnation point is 1.6 mm, and the curvature radius of other places in the leading-edge region is basically about 3 mm. The designed airfoil is named HyAerofoil0400.

[0140] Figure 23 The curve of lift coefficient of the airfoil of Scheme 3 varying with the angle of attack at Mach number 0.3 is given. When the airfoil thickness is further reduced to 4%, the lift coefficient corresponding to the angle of attack of 10° is 0.963, the maximum lift-to-drag ratio is 30.53, and the corresponding lift coefficient is 0.29.

[0141] Figure 24 The curve of lift-to-drag ratio of the airfoil of Scheme 3 varying with the angle of attack at Mach number 0.3 is given. The maximum lift-to-drag ratio is 30.53, the corresponding lift coefficient is 0.293, and the lift-to-drag ratio corresponding to the angle of attack of 10° is 5.46.

[0142] Figure 25 The curve of the lift coefficient of the airfoil of Design Scheme III varying with the angle of attack at Mach number 6.0 is given. The lift coefficient basically varies linearly with the angle of attack, and the lift coefficient corresponding to an angle of attack of 4° is 0.0628.

[0143] Figure 26 The curve of the lift-to-drag ratio of the airfoil of Design Scheme III varying with the angle of attack at Mach number 6.0 is given. The maximum lift-to-drag ratio is 4.527, the corresponding angle of attack is 6°, and the corresponding lift coefficient is 0.0912.

[0144] The above are only some specific implementation cases of the present invention and are not used to limit all the content of the present invention. The transition characteristics of the airfoil flow are not considered in the calculation process, but the calculation methods adopted are the same for comparative selection. For those skilled in the art, the present invention can have various changes and modifications, such as the selection of the basic airfoils of the waverider airfoil and the low-speed airfoil, the ratio of the front and rear sections, the optimization of the airfoil thickness and camber, etc. The content of this invention can also be applicable to the comprehensive optimization design of airfoils flying in more than three speed ranges in the future. Any modifications, improvements, changes, etc. made within the spirit and principles of the present invention shall be included within the protection scope of the present invention.

[0145] Although the embodiments of the present invention are disclosed as above, they are not limited to the applications listed in the specification and the embodiments. It can be fully applied to various fields suitable for the present invention. For those familiar with the field, additional modifications can be easily achieved. Therefore, without departing from the general concept defined by the claims and the equivalent scope, the present invention is not limited to the specific details and the illustrated and described examples here.

Claims

1. A method for designing a wide-speed airfoil that integrates a waverider profile with a low-speed airfoil, characterized in that: include: S1. Based on the lift generation mechanism of the waverider compression characteristics at high speed stage and the flow characteristics at low speed stage, the front section of the airfoil is taken as the waverider section of the high speed airfoil, and the rear section of the airfoil is taken as the rear half of the low speed airfoil; S2. According to the mission characteristics of the aircraft, a nonlinear weight method is used to calculate the distribution ratio of the waverider profile and the low-speed airfoil to obtain a basic airfoil; S3, selecting a predetermined area before and after the maximum thickness position of the basic airfoil to perform grid parameter deformation, so as to complete the wide-speed range airfoil optimization design according to different wide-speed range flight requirements; In S1, the design process of the front section of the airfoil includes: S10, extracting the windward surface of the wave-riding streamline under predetermined working conditions; S11, obtaining a conical streamline on the windward surface by streamline tracing, and using the conical streamline as the lower surface of the front section of the airfoil; S12, using the von Karman curve as the upper surface of the front section of the waverider airfoil; In S2, the nonlinear weighting method is constructed by constructing the following airfoil proportion objective function: To decompose the effects of controlling high speed airfoils and low speed airfoils: In the above formula, e z As a reference function for evaluating high-speed airfoils, Reference function for evaluating low-speed airfoils, z is the independent variable, f ( z ) is the weight coefficient, and f ( z ) through the following nonlinear weighting function Design the distribution ratio of high-speed airfoil and low-speed airfoil: In the above formula, α , β is the shape parameter that controls the distribution morphology, and ; In S3, the process of processing the mesh parameterized deformation includes: S30, select the 15% area before and after the maximum thickness position of the basic airfoil, use the non-uniform FFD control node to construct the control area, and realize the airfoil parameterization through the change of the control node; S31. Clarify the range of changes in the nodes in the control area, use the DOE method to extract samples of the control nodes, and use the FFD parameterization method to change the control nodes in the CFD simulation software to achieve deformation design of the grid area.

2. The method for designing a wide-speed airfoil by integrating a waverider section with a low-speed airfoil as claimed in claim 1, characterized in that: In S31, in FFD parameterization, the following B-spline function is used for mesh deformation parameterization: In the above formula, C(s,t) is the coordinate of the airfoil surface, i is the control node number in the x direction, l Control the number of nodes in the x direction, j is the control node number in the y direction, m Control the number of nodes in the y direction, P i,j is the initial airfoil coordinate of FFD, N i,p for x The basis functions in the direction, N j,q for y The basis functions in the direction, ( s , t ) are the calculated local coordinates of the control node, and , p , q They are x , y The corresponding order in the direction.

3. The method for designing a wide-speed airfoil by integrating a waverider section with a low-speed airfoil as claimed in claim 1, characterized in that: In S3, after the wide speed range airfoil optimization design, the parameters of the optimized airfoil are as follows: The maximum thickness is 8.59%, and the maximum thickness is located at 61.09% of the airfoil; The airfoil camber is 0.995% and the airfoil leading edge radius is 3mm.

4. The method for designing a wide-speed airfoil by integrating a waverider section with a low-speed airfoil as claimed in claim 1, characterized in that: In S3, after the wide speed range airfoil optimization design, the parameters of the optimized airfoil are as follows: The maximum thickness is 4.64%, and the maximum thickness is located at 64.1% of the airfoil; The maximum camber is 1.01%, and the maximum camber is located at 49.1% of the airfoil; The leading edge blunting radius gradually increases with the leading edge curve, and the radius of the airfoil stagnation point is 1.908 mm, and the curvature radius of other places in the leading edge area of ​​the airfoil except the stagnation point is 3 mm.

5. The method for designing a wide-speed airfoil by integrating a waverider section with a low-speed airfoil as claimed in claim 1, characterized in that: In S3, after the wide speed range airfoil optimization design, the parameters of the optimized airfoil are as follows: The maximum thickness is 4.0%, and the maximum thickness is located at 63.4% of the airfoil; The maximum camber is 1.01%, and the maximum camber is located at 48.6% of the airfoil; The blunting radius of the leading edge of the airfoil gradually increases with the leading edge curve. The radius of the airfoil stagnation point is 1.6 mm, and the curvature radius of other places in the leading edge area of ​​the airfoil except the stagnation point is 3 mm.

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