Wave-rider profile and low-speed airfoil profile fused wide-speed-range airfoil profile design method

Abstract

The invention discloses a waverider profile and low-speed airfoil profile fused wide-speed-range airfoil profile design method, and relates to the field of aircraft airfoil profile design, and the method comprises the following steps: S1, based on a lift force generation mechanism of high-speed stage waverider compression characteristics and low-speed stage streaming characteristics, taking a high-speed airfoil profile waverider profile from an airfoil profile front section, and taking a low-speed airfoil profile rear half part from an airfoil profile rear section; s2, according to the task characteristics of the aircraft, a nonlinear weight method is adopted to calculate the distribution proportion of the waverider profile and the low-speed airfoil profile to obtain a basic airfoil profile; and S3, selecting predetermined areas in front of and behind the maximum thickness position of the basic airfoil to perform grid parameterized deformation so as to complete the wide-speed-range airfoil optimization design. According to the wide-speed-range airfoil profile design method integrating the waverider profile and the low-speed airfoil profile, the high lift-drag ratio of the high-speed waverider airfoil profile and the high lift coefficient of the low-speed airfoil profile are fully combined, and the airfoil profile requirement of wide-speed-range aircraft design is met by reasonably selecting the proportion of the waverider airfoil profile and the proportion of the low-speed airfoil profile.

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: 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); 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

[0006] An object of the present invention is to solve at least the above problems and / or disadvantages and to provide at least the advantages which will be described hereinafter.

[0007] In order to achieve these purposes and other advantages of the present invention, a method for designing a wide-speed airfoil that integrates a waverider profile with a low-speed airfoil is provided, comprising: 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. Select the predetermined area before and after the maximum thickness position of the basic airfoil for grid parameter deformation to complete the wide-speed range airfoil optimization design according to different wide-speed range flight requirements.

[0008] Preferably, 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. The von Karman curve is used as the upper surface of the front section of the airfoil.

[0009] Preferably, in S2, the nonlinear weighting method is by constructing the following airfoil proportion objective function: To decompose the effects of controlling high-speed 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 .

[0010] Preferably, 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.

[0011] Preferably, in S31, in FFD parameterization, the following B-spline function is used to perform 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.

[0012] Preferably, 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.

[0013] Preferably, 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. The curvature radius of other places in the leading edge area of ​​the airfoil except the stagnation point is about 3 mm.

[0014] Preferably, 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 about 3 mm.

[0015] The present invention includes at least the following beneficial effects: the wide-speed range airfoil design method proposed in 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 reasonably selects the proportion of the waverider airfoil and the proportion of the low-speed airfoil to form an airfoil that meets the design requirements of a wide-speed range aircraft, thereby providing a new design method for the optimization design of wide-speed range airfoils.

[0016] Furthermore, the airfoil design method of the present invention has an obvious design mechanism. It is not a blind optimization through parametric optimization. The ratio of the two basic airfoils can be adjusted according to the specific needs of wide-speed range flight. The airfoil thickness, curvature, etc. can be locally optimized by adjusting the ratio. The curvature and thickness of the airfoil can also be directly obtained through optimization. The designed airfoil has good high-speed and low-speed aerodynamic performance, and can take into account the design requirements of wide-speed range flight.

[0017] Other advantages, objectives and features of the present invention will be embodied in part through the following description, and in part will be understood by those skilled in the art through study and practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] Figure 1 Schematic diagram of airfoil optimization design based on FFD method of the present invention; Figure 2 Schematic diagram of the front and rear transition of the two sections of the airfoil on the lower surface of the present invention; Figure 3 Schematic diagram of the airfoil leading edge passivation of the present invention; Figure 4 Example 1 of the present invention: Basic airfoil designed by integrating low-speed airfoil and waverider profile ( h =8.59%)); Figure 5 Lift coefficient curve of Mach number 0.3 in Example 1 of the present invention ( h =8.59%)); Figure 6 Lift-to-drag ratio curve of Mach number 0.3 in Example 1 of the present invention ( h =8.59%)); Figure 7Lift coefficient curve of Mach number 6.0 in Example 1 of the present invention ( h =8.59%)); Figure 8 Lift-to-drag ratio curve of Mach number 6.0 of Example 1 of the present invention ( h =8.59%)); Fig. 9 Comparison diagram of the airfoil before and after optimization of Example 2 of the present invention; Fig.10 Relative thickness distribution of the airfoil before and after optimization in Example 2 of the present invention; Fig.11 The wide speed range airfoil optimized in Example 2 of the present invention ( h =4.86%)); Fig.12 The optimized airfoil section of Example 2 of the present invention ( h =4.86%)); Fig.13 The lift coefficient of embodiment 2 of the present invention at Mach number 0.3 varies with the angle of attack ( h =4.86%)); Fig.14 The moment coefficient at Mach number 0.3 in Example 2 of the present invention varies with the angle of attack ( h =4.86%)); Fig.15 The lift-to-drag ratio of Example 2 of the present invention at Mach number 0.3 varies with the angle of attack ( h =4.86%)); Fig.16 The lift coefficient of embodiment 2 of the present invention at Mach number 6.0 varies with the angle of attack ( h =4.86%)); Fig.17 The moment coefficient at Mach number 6.0 of embodiment 2 of the present invention varies with the angle of attack ( h =4.86%)); Fig.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%)); Fig.19 Pressure cloud diagram of Mach number 0.3 and AOA=6.0 in Example 2 of the present invention ( h =4.86%)); Fig. 20 Pressure cloud diagram of Mach number 6.0 and AOA=2.0 in Example 2 of the present invention ( h =4.86%)); Fig.21 The wide speed range airfoil optimized in Example 3 of the present invention ( h =4.0%)); Fig. 22The airfoil after thickness optimization in Example 3 of the present invention ( h =4.0%)); Fig.23 The lift coefficient of embodiment 3 of the present invention at Mach number 0.3 varies with the angle of attack ( h =4.0%)); Fig.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%)); Fig.25 The lift coefficient of embodiment 3 of the present invention at Mach number 6.0 varies with the angle of attack ( h =4.0%)); Fig.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%)); Fig. 27 A schematic diagram of the weight ratio of the high and low speed airfoils of the present invention designed using a nonlinear weighting function; Fig.28 It is a schematic diagram of the weighted sum result of the high and low speed airfoils after the present invention is designed based on the weight ratio. DETAILED DESCRIPTION

[0019] The present invention is further described in detail below in conjunction with the accompanying drawings so that those skilled in the art can implement the invention with reference to the description.

[0020] As the basis of aircraft aerodynamic layout design, airfoil plays a key role in the aerodynamic performance of aircraft. Carrying out innovative design of wide-speed airfoil is one of the key problems faced by aircraft design. Therefore, in response to the demand for aerodynamic performance of wide-speed aircraft, the present invention provides a wide-speed airfoil design method based on waveriding characteristics and low-speed airfoil coupling design, and obtains a wide-speed airfoil with a high lift-to-drag ratio and a high lift coefficient in the low-speed takeoff and landing stage and a high lift-to-drag ratio at high speed. The wide-speed airfoil design process is described below: Generally speaking, the wing loading of wide-speed aircraft is usually 350-600kg / m 2 The lift-weight balance that the aircraft needs to meet during the entire aircraft phase is as follows: in, 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.

[0021] Assuming the high-speed flight Mach number is 6.0, the altitude is 28km, the horizontal takeoff Mach number is 0.3, the air density on the ground (1km above sea level) is 1.1kg / m3, the speed of sound is 336m / s, the air density at 28km is 0.02507kg / m3, the speed of sound is 300m / s, and the mass change during takeoff is not considered, then the required lift coefficient ratio of the aircraft is: 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 1 km above sea level, v 28km 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.

[0022] In order to improve the cruising performance of the aircraft, the lift coefficient during high-speed flight is usually taken as 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°. In order to ensure takeoff and landing safety, the angle of attack of low-speed flight should not exceed 15°. If for the same aircraft, it is assumed that the angle of attack ratio corresponding to high-speed and low-speed flight is 1 / 3, if the lift line slope in the low-speed stage is 0.055, then the lift line slope in the high-speed stage is 0.0227. Through analysis, it can be seen that the same aircraft needs to meet the larger lift coefficient changes in the high-speed and low-speed stages, and needs to meet a certain stall angle of attack at low speed. Therefore, when completing the design of 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, the area of ​​leeward separation at the same angle of attack will be reduced, and the lift coefficient at small angles of attack will be increased. However, since the increase in camber will also lead to an increase in the frontal area, the lift-to-drag ratio of the airfoil at high speeds will drop significantly. Therefore, how to achieve a high lift coefficient and a high lift-to-drag ratio at low speeds and high angles of attack of the airfoil, and at the same time improve the lift-to-drag ratio of the airfoil at small angles of attack at high speeds, is the difficulty in designing airfoils in a wide speed range.

[0023] To achieve the above-mentioned purpose, the present invention draws on the lift generation mechanism of waverider compression and low-speed flow, extracts the windward surface of the typical waverider streamline (it should be noted that the typical streamline here refers to the waverider streamline obtained under the flight conditions of Mach number 6, angle of attack 0°, shock wave angle 12°, and the shock wave angle can be adjusted according to design requirements) and the rear half of the low-speed airfoil, wherein, as Figure 1-Figure 2 As shown (it should be noted that Figure 1 The orange line represents the deformed airfoil, the blue line represents the reference airfoil, and dots of various colors are used to represent FFD control points. Figure 2 The purple line in the middle represents the lower surface airfoil after the arc transition, the light blue line represents the tail of the waverider section, and the blue line represents the front end of the low-speed airfoil). The front half of the airfoil is designed with the waverider concept. The lower surface is a conical streamline obtained by streamline tracing to reduce shock wave leakage at the high-speed stage. The upper surface of the waverider adopts the von Karman 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 designed by integrating the waverider section and the von Karman curve, as shown in the figure. Figure 3 As shown ( Figure 3 The purple line indicates the upper surface of the airfoil leading edge after passivation, the light blue line indicates the lower surface of the airfoil leading edge after passivation, and the red line indicates the passivated leading edge). The leading edge radius can be adjusted according to the flight mission to meet different leading edge radii and balance the effects of aerodynamic heat and aerodynamic drag. The rear half of the airfoil adopts the rear half of the low-speed airfoil, which, on the one hand, increases the lift of the airfoil in the low subsonic stage, and on the other hand, reduces 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 takes the high-speed waverider section, and the rear section of the airfoil takes the rear half of the low-speed airfoil. This allows the airfoil to fully utilize 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 fully utilizes the flow characteristics of the airfoil to alleviate separation at large angles of attack, meet the high lift and high lift-to-drag ratio requirements in the low-speed stage, and take into account both high-speed and low-speed flight performance.

[0024] Furthermore, in order to make the airfoil adjust the proportion of the waverider section and the low-speed airfoil section according to the mission characteristics of the aircraft, so as to adapt to different aircraft design requirements, the present invention also designs a nonlinear weight calculation method for allocating the proportion of the waverider section and the low-speed airfoil, so as to determine the allocation proportion of the high-speed and low-speed airfoils according to the use 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 proportion objective function: : In the above formula, e z As a reference function for evaluating high-speed airfoils, is the reference function for evaluating low-speed airfoils, 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% to 70%, this patent constructs the following nonlinear weighting function: : In the above formula, α , β is the shape parameter that controls the distribution morphology, and ,in, α =4, β =3 airfoil is more inclined to high-speed flight. α =3, β =4 airfoil is more suitable for slow speed flight.

[0025] The above objective function is characterized by the independent variable z The value range of is 0-1, which is mainly used to control the proportion of high-speed airfoil along the chord length. Since the high-speed waverider profile airfoil increases with the chord length, that is, the maximum thickness of the airfoil is positively correlated with the proportion of the high-speed airfoil, the design objective function The maximum value of appears in the range of 50%~80% of the independent variable, which meets the design requirements of the maximum thickness of the airfoil in engineering. This method realizes the fusion of two flight modes through nonlinear weights and is suitable for multi-stage mission trade-off analysis.

[0026] Furthermore, based on the above basic airfoil, taking into account the wave-riding characteristics of the leading edge of the airfoil and the loading design of the trailing edge of the airfoil, further optimization was carried out on the basic airfoil. The area 15% before and after the maximum thickness of the airfoil was mainly selected for grid parameterization, and the FFD control node was set to control the deformation of the airfoil. The DOE method was used to extract samples of the design parameters, and then the CFD method was combined with the FFD method to complete the data calculation of the samples, including low-speed performance and high-speed performance.

[0027] There are usually three basis functions for FFD parameterization: Bernstein polynomials, B-spline functions, and NURBS functions. Among them, Bernstein polynomials are a global influence method, which is characterized by changing one control node, all coordinates will change, and it has no local support. B-spline functions and NURBS functions have local support and local deformation capabilities. According to the design concept of this patent, B-spline functions are selected here for mesh deformation parameterization.

[0028] The specific implementation process of the B-spline function is as follows: First, the control nodes are divided into the areas that need to be deformed. By changing the control nodes, the areas near them can be deformed. The specific affected area depends on the order of the basis function. The implementation theory of FFD parameters is as follows: in, 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 non-uniform control node 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.

[0029] Constructing the B-spline function requires constructing the node vector in the x direction U , the general formula is: in, i is the FFD control node number, p For the order, r is the number of vector nodes, n+1 is the number of FFD control points, u are local coordinates, u i is the i-th vector node value, u i+p is the i+pth vector node value, and R represents an r+1-dimensional real number set. Here, 0 / 0=0 is defined for program solving. The node vector construction method in the y direction is the same as that in the x direction, which will not be described here.

[0030] Usually the node vector is given in the following way: in, u 0 = u 1… = up =0, u n+1 = … = u n+p+1 =1, and u p+1 , …, u n The control nodes are set up in an arithmetic progression with increasing distribution, and they must also meet u p+1 > u p , u n < u n+1 .

[0031] Since s is known, B i,0 It is certain. 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 All are 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 All are known, and it can be seen that, starting from order 0, the basis functions of all control nodes can be found in turn.

[0032] Since the node distribution on the airfoil surface is in the form of a curve, it is usually difficult to accurately control the local design parameters using uniformly distributed control nodes. Here, a non-uniform control node distribution is used for mesh parameterization. Assume that the initial non-uniform control node coordinates are P i,j , the calculated local coordinates are ( s , t ), then under the premise that the control node, node vector and weighting factor remain unchanged, the corresponding airfoil coordinates become: The node vector and weighting factor remain unchanged, and the non-uniform control node coordinates are changed to After that, the airfoil coordinates become: Through the above design method, the wide-speed range airfoil optimization design can be carried out, and different airfoil designs can be completed according to different wide-speed range flight requirements.

[0033] Embodiment 1: Considering that the cruising state of a wide-speed range aircraft is mainly in the high-speed stage, the high-speed flight mode is preferred when integrating the waverider airfoil and the low-speed airfoil in this case, and the values ​​of the shape parameters α and β are set ( α =4, β =3), obtain the nonlinear weighting function and calculate the objective function With the value of the independent variable u The changing law of , and the optimized allocation ratio is obtained, such as Figure 27-Figure 28 As shown in the figure, the high-speed waverider airfoil accounts for 64.5% and the low-speed airfoil accounts for 35.5%. The waverider airfoil is obtained by solving the Taylor-Maccoll conical flow field equation, and the low-speed airfoil is cut by a specific airfoil (such as NACA 64(1)-212). In order to achieve the expansion design of the airfoil, the upper surface of the airfoil is designed by the von Karman curve.

[0034] Among them, the waverider airfoil selects the conical flow field when the incoming flow Mach number is 6 and the shock wave angle is 12°. The reference flow field information is obtained by solving the Taylor-Maccoll equation, and then the streamline coordinates are obtained by streamline tracing. The specific Taylor-Maccoll equation is as follows: in, is the critical sound speed of the free stream, c is the speed of sound, is the dimensionless radial velocity of the critical acoustic intensity, is the dimensionless tangential velocity of the critical acoustic intensity, θ is the polar angle.

[0035] In the above formula, T t It is to flow the total temperature, R 0 is the gas constant of air, k is the specific heat ratio of air, T It is to flow with tranquility and warmth, It is the incoming flow Mach number.

[0036] The rear part of the airfoil is a modified low-speed airfoil NACA 64(1)-212, which has a maximum thickness of 9.97% and a maximum thickness position of 40% of the airfoil chord length.

[0037] When considering the viscosity of the waverider, the state of the maximum lift-to-drag ratio is no longer the design state. Usually, at a positive angle of attack, as the angle of attack increases, a leeward area will form on the back of the aircraft, which is an expansion wave area. However, due to the high flight altitude of the cross-speed domain aircraft and the low ambient pressure, the decompression range of the expansion wave is limited, so the contribution to the lift of the aircraft is not obvious. Therefore, from the perspective of improving the aerodynamic performance of the aircraft, increasing the volume and strengthening the structural safety, the upper surface of the waverider is designed to increase the volume. The design method used is the von Karman curve, and the specific line design equation is as follows: in, H It refers to the distance between the given von Karman curve end point coordinate and the axis. L represents the length of the von Karman curve along the x direction, h x is the distance of the curve from the axis along the x direction. H =0 means that the upper surface of the waverider is parallel to the incoming flow.

[0038] The next step is the specific design process of the airfoil: First, the length of the waverider airfoil is 1000mm, the upper surface is designed with the von Karman curve, and the designed outlet thickness is 121.6mm. The rear half adopts the transonic airfoil design idea, the airfoil trailing edge adopts the rear loading design, and the upper surface of the airfoil adopts the arc design method. The length of the rear half of the airfoil is 548.7mm, and the thickness of the airfoil trailing edge is 4.673mm. The initial airfoil length designed by this method is 1548.7mm, the maximum thickness is 121.6mm, and the relative thickness is 7.85%. In order to ensure the smoothness and continuity of the lower surface airfoil, the 98% position of the waverider airfoil in the front half of the lower surface and the 3% position of the low-speed airfoil in the rear half of the lower surface are cut, and then the arc transition is performed through the spline curve.

[0039] Considering that the leading edge of the waverider airfoil is pointed, it cannot meet the processing design requirements regardless of high-speed or low-speed flight. At the same time, the heat protection requirements must also be considered at the high-speed stage. Generally, the larger the leading edge radius, the better the heat protection effect. However, 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. This patent performs a passivation treatment on the leading edge of the airfoil, and also performs an arc transition at the junction of the leading and trailing edges of the lower surface of the airfoil, such as Figure 4 As shown ( Figure 4 The x-axis is the chord coordinate of the airfoil, and the y-axis is the curvilinear coordinate of the airfoil perpendicular to the chord. The designed airfoil length is 1410 mm, and the maximum absolute thickness is 121.05 mm. The maximum thickness of the airfoil is obtained ( h is the maximum thickness) is 8.59%, the maximum thickness position is 61.09%, the airfoil camber is 0.995%, the normalized leading edge radius is 3.5mm, the tail angle is 14.82°, and the designed airfoil is named HyAerofoil0859.

[0040] In order to study the influence of the leading edge radius of the airfoil on the aerodynamic performance of the airfoil, this patent conducted a comparative analysis on the leading edge blunting radii of 1mm, 3mm and 5mm. The analyzed states corresponded to Mach numbers 0.3 and Mach numbers 6.0, and the curves of lift coefficient and lift-to-drag ratio under different design conditions were obtained.

[0041] Figure 5 Comparison curve of lift coefficient with angle of attack at 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=1mm, the purple dotted line is the lift coefficient curve when the leading edge radius R=3mm, and the green dotted line is the lift coefficient curve when the leading edge radius R=5mm). At Mach number 0.3 and 10° angle of attack, the lift coefficients corresponding to the leading edge radii of 1mm, 3mm, and 5mm are 1.017, 0.930, and 0.928, respectively. From the lift coefficient curve, it can be seen that the slope of the lift line of this airfoil in the low-speed section becomes smaller and smaller with the angle of attack. It may be that the smaller leading edge radius and thickness cause the separation area of ​​the airfoil surface to become larger when the angle of attack increases.

[0042] Figure 6 Shown is a comparison curve of lift-to-drag ratio versus angle of attack at different leading edge radii at Mach number 0.3. The maximum lift-to-drag ratios corresponding to leading edge radii of 1mm, 3mm, and 5mm are 20.53, 23.47, and 22.62, respectively. The corresponding lift coefficients are 0.408, 0.403, and 0.404, respectively. The angle of attack is 0°.

[0043] Figure 7Shown is a comparison curve of the lift coefficient changing with the angle of attack at different leading edge radii at Mach number 6.0. At Mach number 6.0, the lift coefficient increases with the angle of attack, and the slope of the lift line increases slightly with the increase in the angle of attack, which may also be related to the increase in the back separation area. However, the effect of the blunting radius on the lift coefficient is not as great as at low speed.

[0044] Figure 8 The figure shows the lift-to-drag ratio curve with different leading edge radii at Mach number 6.0. The maximum lift-to-drag ratios corresponding to the leading edge radii of 1mm, 3mm, and 5mm are 4.287, 3.664, and 3.19, respectively, and the corresponding angles of attack are 4°, 6°, and 6°, respectively. Through analysis, when the passivation radius is greater than 3mm at low speed, the change of the maximum lift-to-drag ratio becomes gentle. In order to alleviate the aerodynamic heat load at high speed, the leading edge of the aircraft must meet a certain passivation radius. After weighing, the leading edge radius of about 3mm is also selected when designing a wide speed range airfoil, which can take into account the aerodynamic requirements of low speed and high speed at the same time. That is, when the leading edge radius is 3mm, the lift coefficient corresponding to the Mach number of 0.3 and the 10° angle of attack is 0.93, the maximum lift-to-drag ratio is 23.47, and the corresponding lift coefficient is 0.4; the maximum lift-to-drag ratio is 3.664 at Mach number 6.0, and the corresponding angle of attack is 6°.

[0045] Embodiment 2: Considering that the flight speed of wide-speed range aircraft has to face low-speed, transonic and high-speed stages, the aircraft has higher requirements for lift increase and drag reduction in the full speed range, and has put forward design requirements for the leading edge radius, airfoil thickness and airfoil camber of the airfoil. Low-speed airfoils usually adopt larger leading edge radius, larger airfoil thickness and appropriate airfoil camber to ensure low-speed high lift coefficient, high lift-to-drag ratio and large stall angle of attack. The airfoil thickness is usually above 10%. The airfoil thickness of fighter jets is usually around 4% and the camber is small. The thickness of the waverider increases with the increase of the aircraft size, but the airfoil thickness is usually controlled within 5%. The lift-to-drag ratio of the high-speed airfoil is also closely related to the leading edge radius, maximum thickness and area of ​​the airfoil. On the whole, the wide-speed range airfoil design should also choose an airfoil with small thickness, small leading edge radius and small camber.

[0046] like Figure 11-Figure 12 As shown in the figure, in response to the needs of wide-speed flight, the thickness distribution and maximum thickness of the airfoil are optimized. The maximum thickness of the airfoil obtained by the FFD method is 4.86%, the maximum thickness position is 64.1%, the maximum curvature is 1.01%, the maximum curvature position is 49.1%, and the airfoil area is 0.03041. The leading edge blunting radius gradually increases with the leading edge curve, the radius of the stagnation point is 1.908mm, and the curvature radius of other places in the leading edge area is basically around 3mm. The designed airfoil is named HyAerofoil0486.

[0047] Fig. 9 The comparison of airfoil profile before and after optimization is given ( Fig. 9 In the figure, the x-axis is the chord-wise coordinate of the airfoil, the y-axis is the curve coordinate of the airfoil perpendicular to the chord, the orange solid line is the airfoil profile of Example 1, and the blue dotted line is the optimized airfoil profile of Example 2). It can be seen that since the leading edge radius and trailing edge thickness of Scheme 1 are also small, the middle position of the airfoil is mainly optimized, the maximum thickness of the airfoil is changed, and the camber does not change much.

[0048] Fig.10 The relative thickness distribution comparison before and after airfoil optimization is given ( Fig.10 In the figure, the x-axis is the chord-wise coordinate of the airfoil, the y-axis is the curvilinear coordinate of the airfoil perpendicular to the chord, the orange solid line is the relative thickness distribution of the airfoil in Example 1, and the blue dotted line is the relative thickness distribution of the airfoil after optimization in 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.

[0049] Fig.13 The lift coefficient variation curve of the optimized airfoil at Mach number 0.3 with the angle of attack 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 decreases. 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 take-off and landing of the aircraft.

[0050] Fig.14 The curve of the pitching moment coefficient of the optimized airfoil at Mach number 0.3 as a inverse of the angle of attack is given. The zero-lift moment coefficient of the airfoil is about 0.054, which means that the airfoil has a nose-up moment. As the angle of attack increases, the moment coefficient decreases first and then increases, and there is an obvious nose-up moment at large angles of attack.

[0051] Fig.15 The lift-to-drag ratio curve of the optimized airfoil at Mach number 0.3 with angle of attack is given. It can be seen that the maximum lift-to-drag ratio has increased compared with the original airfoil, and the corresponding angle of attack is 0°. Analysis shows that the reduction in thickness leads to a significant reduction in drag. Taking an angle of attack of 10° as an example, the lift-to-drag ratio of the optimized airfoil is 5.435, while 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.

[0052] Fig.16The lift coefficient curve of the optimized airfoil at Mach number 6.0 with the angle of attack is given. As the airfoil thickness decreases, the lift coefficient of the airfoil at high speed decreases, and the drag coefficient also decreases. The linearity of the lift coefficient with the angle of attack is good. However, compared with the original airfoil, the lift coefficient has decreased overall. Taking the 4° angle of attack as an example, the lift coefficient drops from 0.08 of the original airfoil to 0.0646 after optimization.

[0053] Fig.17 The curve of the pitching moment coefficient of the optimized airfoil at Mach number 6.0 as a inverse of the angle of attack is given. Unlike the low-speed stage, in the high-speed stage, the pitching moment coefficient of the airfoil gradually increases with the increase of the angle of attack, and the nose-up trend is more obvious. The reason is that the point of action of the lift is after the moment reference point. As the angle of attack increases, the lift increases, which leads to an increase in the moment.

[0054] Fig.18 The lift-to-drag ratio curve of the optimized airfoil at Mach number 6.0 is given. The maximum lift-to-drag ratio increases from 3.664 to 4.198 compared with the original airfoil. The corresponding angle of attack is 6° and the corresponding lift coefficient is 0.0932.

[0055] Fig.19 and Fig. 20 The pressure cloud diagrams of the optimized airfoil at typical angles of attack at Mach numbers of 0.3 and 6.0 are given respectively (at Mach number 0.3, the lift coefficient corresponding to the 10° angle of attack 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 lift of the airfoil mainly comes from the front half of the airfoil, and obvious high and low pressure areas are formed on the upper and lower surfaces of the leading edge of the airfoil. In the high-speed stage, the lift is mainly generated by the shock wave compression on the lower surface of the airfoil.

[0056] Embodiment 3: In order to further ensure that the designed airfoil has good adaptability, the thickness of the airfoil is further optimized and adjusted, such as Figure 21-22 As shown in the figure, the maximum thickness of the airfoil is reduced from 4.64% to 4.0%. After optimization, the maximum thickness of the airfoil is at 63.4%, the maximum curvature is 1.01%, the maximum curvature is at 48.6%, the leading edge blunting radius gradually increases with the leading edge curve, the radius of the stagnation point is 1.6mm, and the curvature radius of other places in the leading edge area is basically around 3mm. The designed airfoil is named HyAerofoil0400.

[0057] Fig.23The lift coefficient variation curve of the three-airfoil at Mach number 0.3 with the angle of attack is given. When the airfoil thickness is further reduced to 4%, the lift coefficient corresponding to the 10° angle of attack is 0.963, the maximum lift-to-drag ratio is 30.53, and the corresponding lift coefficient is 0.29.

[0058] Fig.24 The lift-to-drag ratio curve of the three-wing scheme at Mach number 0.3 versus angle of attack 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 an angle of attack of 10° is 5.46.

[0059] Fig.25 The lift coefficient variation curve of the three-wing scheme at Mach number 6.0 with the angle of attack is given. The lift coefficient changes basically linearly with the angle of attack, and the lift coefficient corresponding to 4° angle of attack is 0.0628.

[0060] Fig.26 The lift-to-drag ratio curve of the three-wing scheme at Mach number 6.0 is given as a function of the angle of attack. The maximum lift-to-drag ratio is 4.527, the corresponding angle of attack is 6°, and the corresponding lift coefficient is 0.0912.

[0061] The above description is only a part of the specific implementation cases of the present invention, and is not intended to limit the entire content of the present invention. The calculation process does not take into account the transition characteristics of the airfoil flow, but as a comparative selection, the calculation methods used are consistent. For those skilled in the art, the present invention may have various changes and variations, such as the selection of basic airfoils of waverider airfoils and low-speed airfoils, the ratio of the front and rear sections, the optimization of the thickness and curvature of the airfoil, etc. The content of the invention may also be applied to the comprehensive optimization design of airfoils flying in more than three speed ranges. Any modifications, improvements, changes, etc. made within the ideas and principles of the present invention should be included in the protection scope of the present invention.

[0062] Although the embodiments of the present invention have been disclosed 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 art, additional modifications can be easily realized. Therefore, without departing from the general concept defined by the claims and equivalent scope, the present invention is not limited to the specific details and the illustrations shown and described 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. Select the predetermined area before and after the maximum thickness position of the basic airfoil for grid parameter deformation to complete the wide-speed range airfoil optimization design according to different wide-speed range flight requirements.

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 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. The von Karman curve is used as the upper surface of the front section of the waverider airfoil.

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 S2, the nonlinear weighting method is constructed by constructing the following airfoil proportion objective function: To decompose the effects of controlling high-speed 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 .

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, 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.

5. The method for designing a wide-speed airfoil by integrating a waverider section with a low-speed airfoil as claimed in claim 4, 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.

6. 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.

7. 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. The curvature radius of other places in the leading edge area of ​​the airfoil except the stagnation point is about 3 mm.

8. The method for designing a wide-speed airfoil that integrates 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 about 3 mm.

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