Wing-body fusion wide-speed-range aircraft layout design method based on consideration of afterbody effect

Abstract

The invention provides a wing-body fusion wide-speed-range aircraft layout design method based on consideration of an afterbody effect. The method comprises the following steps: firstly, generating a cone-guided waverider precursor of a wide-speed-range aircraft for supersonic flight; then, based on the basic airfoil profile, performing pneumatic optimization on the basic wing profile and the strake wing profile of the wide-speed-domain aircraft to obtain a multi-section wide-speed-domain waverider airfoil profile; then calculating the length of a waverider afterbody and designing a symmetry plane of the wide-speed-range aircraft; and finally, based on an osculating cone method and transonic speed strake wing constraints, obtaining three-dimensional layout parameters by using the waverider precursor parameters and the waverider afterbody length, and assembling the basic wing optimized airfoil profile, the strake wing optimized airfoil profile and the symmetry plane to generate the layout of the wide-speed-domain aircraft. According to the method, the cone-guided waverider generation method, the wide-speed-range airfoil optimization method and the osculating cone double-sweepback waverider generation method are coupled, and the aerodynamic design of the afterbody is considered, so that the aerodynamic performance of the wide-speed-range aircraft under the subsonic / transonic / supersonic speed is improved.

Description

Technical Field

[0001] The present invention relates to the field of aircraft design, and particularly to a layout design method of a wing-body integrated wide-speed-range aircraft based on considering afterbody effects, which is used for the design of a new generation of supersonic airliners. Background Art

[0002] A wide-speed-range aircraft is an aircraft that can take off and land horizontally, is reusable, and has supersonic cruise capabilities. Its flight speed covers subsonic, transonic, and supersonic speeds, with the cruise speed increasing from Ma = 3 to Ma = 5, and the maximum flight speed even reaching Ma = 8.

[0003] Currently, the design of wide-speed-range aircraft generally combines the waverider design. Although the waverider can improve the high-speed performance of wide-speed-range aircraft, it is difficult to balance the subsonic / transonic lift-drag characteristics and stability. To solve this problem, one method is that aircraft such as the X-15, Sanger, and Talon-A take off and land horizontally on a subsonic / supersonic platform and launch a supersonic aircraft in the air, enabling the two-stage supersonic aircraft to avoid low-speed operating conditions. Obviously, this method cannot be used in the design of supersonic airliners; another method is to improve the subsonic / transonic / supersonic performance of the waverider to achieve single-stage wide-speed-range flight. Methods for improving the low-speed performance of the waverider include using a configuration that combines with a wing to form a waverider-wing layout, and a waverider-vortex wave layout that uses vortex lift augmentation. The waverider-wing layout is specifically divided into two categories: a double-swept integral waverider layout and a spliced non-full-aircraft waverider layout. The former uses an improved waverider generation method to construct the full-aircraft waverider effect, but the airfoil of its wing is generally a simple wedge shape, and the upper surface of the fuselage is generally flat, so there is still room for improvement in low-speed flight performance; the latter starts from a wide-speed-range wing, fixes the shape of the waverider forebody, splices a wide-speed-range fuselage and wing behind the waverider forebody, and designs the leading-edge flap / basic wing planform layout through empirical or surrogate model optimization to generate a wide-speed-range aircraft with a non-full-aircraft waverider. This method ignores the role of the aircraft projection surface shape in maintaining the high-pressure area on the lower surface and has drawbacks in the case of relatively high supersonic speeds.

[0004] It should be noted that there are significant afterbody design defects in the existing wide-speed-range aircraft design: the current technical route generally focuses on optimizing the forebody waverider configuration, only retains a vertical end face or simply docks the power system in the afterbody area, and has not effectively developed the transonic efficiency enhancement potential of the afterbody aerodynamic surface. Summary of the Invention

[0005] Aiming at the problems existing in the prior art, the present invention provides a layout design method of a wing-body integrated wide-speed-range aircraft based on considering afterbody effects. This method couples a cone-guided waverider, wide-speed-range airfoil optimization, and a method for generating a close-cone double-swept waverider, and considers afterbody aerodynamic design, thereby being able to improve the aerodynamic performance of wide-speed-range aircraft at subsonic / transonic / supersonic speeds.

[0006] The technical solution of the present invention is as follows:

[0007] The layout design method of a wing-body fusion wide-speed domain aircraft based on considering the afterbody effect includes the following steps:

[0008] Step 1: Generate a conical-guided waverider forebody for a wide-speed domain aircraft for supersonic flight;

[0009] Step 2: Based on the basic airfoil, aerodynamically optimize the basic wing airfoil and strake wing airfoil of the wide-speed domain aircraft to obtain a multi-section wide-speed domain waverider airfoil:

[0010] Step 2.1: Respectively determine the basic airfoil of the basic wing and the basic airfoil of the strake wing; perform parametric processing on the basic airfoil to obtain optimization parameters and determine the design space of the optimization parameters;

[0011] Step 2.2: Set the design conditions and optimization models for optimizing the basic wing airfoil and the strake wing airfoil; the design conditions are divided into subsonic, transonic, and supersonic conditions;

[0012] Step 2.3: For the optimization model established in Step 2.2, sample the optimization parameters within the design space, use the surrogate model, and solve the optimization model through the optimization algorithm to achieve multi-objective optimization and obtain the optimized airfoil of the basic wing and the optimized airfoil of the strake wing;

[0013] Step 3: Calculate the length of the waverider afterbody:

[0014] Step 3.1: First, according to the shock wave angle β of the flow field, the incoming flow Mach number Ma, the specific heat ratio γ, and the surface angle θ, according to the formula

[0015] M 2 = 1 + [ ( γ − 1 ) / 2 ] M a 2 s i n 2 β γ M a 2 s i n 2 β − ( γ − 1 ) / 2 1 s i n ( β − θ ) α = s i n − 1 ( 1 / M 2 )

[0016] calculate the Mach number behind the wave and the included angle between the left-running characteristic line family and the center line ;

[0017] Step 3.2: According to the formula

[0018]

[0019] calculate the length of the waverider afterbody ; where is the proportion of the waverider forebody in the overall aircraft length, is the length of the waverider forebody;

[0020] Step 4: Design the two-dimensional shape of the symmetry plane of the wide-speed domain aircraft;

[0021] Keep the symmetry plane of the waverider forebody unchanged, set the original shape of the symmetry plane of the waverider afterbody to be all straight lines on the upper and lower sides. On the basis of the original shape, keep one side straight and adjust the other side; calculate the aerodynamic parameters of the symmetry plane of the adjusted wide-speed-range aircraft under subsonic, transonic, and supersonic conditions, and calculate the aerodynamic parameters according to the formula

[0022]

[0023] Perform weighted calculation to obtain the aerodynamic performance of the symmetry plane of the adjusted wide-speed-range aircraft; where represents the weighting function of the symmetry plane, 、 and respectively represent the subsonic lift coefficient, transonic drag coefficient, and supersonic lift-to-drag ratio of the symmetry plane in sequence; 、 and respectively represent the weighting coefficient of the subsonic lift coefficient of the symmetry plane, the weighting coefficient of the transonic drag coefficient of the symmetry plane, and the weighting coefficient of the supersonic lift-to-drag ratio of the symmetry plane in sequence; the smaller the calculation result of the weighting function, the better the aerodynamic performance of the symmetry plane; obtain the symmetry plane of the wide-speed-range aircraft with the best aerodynamic performance;

[0024] Step 5: Based on the osculating cone method and the transonic strake wing constraint, use the waverider forebody parameters and the length of the waverider afterbody to obtain the three-dimensional layout parameters, and assemble the basic wing optimized airfoil, the strake wing optimized airfoil, and the symmetry plane to generate the layout of the wide-speed-range aircraft.

[0025] Furthermore, in step 1, the osculating cone method is used to generate the osculating cone waverider forebody for the wide-speed-range aircraft for supersonic flight.

[0026] Furthermore, the specific process of generating the osculating cone waverider forebody for the wide-speed-range aircraft for supersonic flight in step 1 is as follows:

[0027] Step 1.1: Determine the half-cone angle of the cone generating the conical shock wave flow field and the incoming flow Mach number Ma, and specify the length of the conical flow field ;

[0028] Step 1.2: Specify the conical bottom profile, find the surface passing through the conical bottom profile and parallel to the incoming flow direction. The intersection line of the surface and the shock wave surface is the leading edge line of the waverider, and the surface formed by the leading edge line and the bottom profile is the upper surface of the waverider;

[0029] Step 1.3: Starting from the leading edge line of the waverider, according to the Taylor-Maccoll equation, obtain the lower surface of the waverider through streamline tracing to generate the osculating cone waverider forebody.

[0030] Furthermore, in step 2.1, for the basic wing, the NACA64A204 airfoil is selected as the basic airfoil; for the strake wing, an inverted wedge with a flat bottom, the maximum thickness at 45% of the chord length, the maximum thickness being 3.5% of the chord length, and a leading edge radius of 10 mm is selected as the basic airfoil.

[0031] Furthermore, in step 2.2, the design conditions are selected as: Ma = 0.4, angle of attack of 2°, altitude of 10 km; Ma = 1, angle of attack of 2°, altitude of 10 km; Ma = 5, angle of attack of 5°, altitude of 30 km; corresponding to subsonic, transonic, and supersonic conditions respectively.

[0032] Furthermore, in step 2.2, the optimization models are respectively:

[0033] For the basic wing airfoil, the optimization model is:

[0034]

[0035] where is the optimization objective of the basic wing, is the objective function of the basic wing, is the penalty function weighting coefficient of the basic wing, is the lift coefficient penalty function of the basic wing, is the drag coefficient penalty function of the basic wing, is the lift-to-drag ratio penalty function of the basic wing, is the thickness penalty function of the basic wing; represents the constraint condition, are respectively the lift-to-drag ratios of the optimized airfoil of the basic wing under supersonic, transonic, and subsonic conditions in sequence, are respectively the lift-to-drag ratios of the basic airfoil of the basic wing under supersonic, transonic, and subsonic conditions in sequence, are respectively the lift coefficients of the optimized airfoil of the basic wing under subsonic, transonic, and supersonic conditions in sequence, are respectively the lift coefficients of the basic airfoil of the basic wing under subsonic, transonic, and supersonic conditions in sequence, are respectively the drag coefficients of the optimized airfoil of the basic wing under subsonic, transonic, and supersonic conditions in sequence, are respectively the drag coefficients of the basic airfoil of the basic wing under subsonic, transonic, and supersonic conditions in sequence, is the maximum thickness of the optimized airfoil of the basic wing, is the maximum thickness of the basic airfoil of the basic wing;

[0036] The objective function of the basic wing is:

[0037]

[0038] where They are, in sequence, the weighted coefficient of the subsonic lift coefficient, the weighted coefficient of the transonic drag coefficient, and the weighted coefficient of the supersonic lift-drag ratio of the optimized airfoil of the basic wing after normalization;

[0039] The penalty function expression of the lift coefficient of the basic wing is:

[0040]

[0041] where is the normalization weight of the lift coefficient of the basic wing in the subsonic state, is the normalization weight of the lift coefficient of the basic wing in the transonic state, is the normalization weight of the lift coefficient of the basic wing in the supersonic state;

[0042] The penalty function expression of the drag coefficient of the basic wing is:

[0043]

[0044] where is the normalization weight of the drag coefficient of the basic wing in the subsonic state, is the normalization weight of the drag coefficient of the basic wing in the transonic state, is the normalization weight of the drag coefficient of the basic wing in the supersonic state;

[0045] The penalty function expression of the lift-drag ratio of the basic wing is:

[0046]

[0047] where is the normalization weight of the lift-drag ratio of the basic wing in the subsonic state, is the normalization weight of the lift-drag ratio of the basic wing in the transonic state, is the normalization weight of the lift-drag ratio of the basic wing in the supersonic state;

[0048] The penalty function expression of the thickness of the basic wing is:

[0049]

[0050] where is the normalization weight of the thickness of the basic wing;

[0051] For the strake wing, the optimization model is:

[0052]

[0053] where is the optimization target of the strake wing, is the objective function of the strake wing, is the penalty function weighted coefficient of the strake wing, is the penalty function of the side-edge wing lift coefficient, is the penalty function of the side-edge wing drag coefficient, is the penalty function of the side-edge wing lift-drag ratio, is the penalty function of the side-edge wing thickness; represents the constraint condition, are respectively the lift-drag ratios of the optimized airfoil of the side-edge wing under supersonic, transonic, and subsonic conditions in sequence, are respectively the lift-drag ratios of the basic airfoil of the side-edge wing under supersonic, transonic, and subsonic conditions in sequence, are respectively the lift coefficients of the optimized airfoil of the side-edge wing under subsonic, transonic, and supersonic conditions in sequence, are respectively the lift coefficients of the basic airfoil of the side-edge wing under subsonic, transonic, and supersonic conditions in sequence, are respectively the drag coefficients of the optimized airfoil of the side-edge wing under subsonic, transonic, and supersonic conditions in sequence, are respectively the drag coefficients of the basic airfoil of the side-edge wing under subsonic, transonic, and supersonic conditions in sequence, is the maximum thickness of the optimized airfoil of the side-edge wing, is the maximum thickness of the basic airfoil of the side-edge wing;

[0054] The expression of Basefit2 is:

[0055]

[0056] where are respectively the weighted coefficients of the subsonic lift-drag ratio, transonic lift-drag ratio, and supersonic lift-drag ratio of the optimized airfoil of the side-edge wing after normalization;

[0057] The expression of the penalty function of the side-edge wing lift coefficient is:

[0058]

[0059] where is the normalization weight of the side-edge wing lift coefficient in the subsonic state, is the normalization weight of the side-edge wing lift coefficient in the transonic state, is the normalization weight of the side-edge wing lift coefficient in the supersonic state;

[0060] The expression of the penalty function of the side-edge wing drag coefficient is:

[0061]

[0062] where is the normalization weight of the side-edge wing drag coefficient in the subsonic state, is the normalization weight of the side-edge wing drag coefficient in the transonic state, is the normalized weight of the side-edge wing drag coefficient in the supersonic state;

[0063] The penalty function expression of the side-edge wing lift-drag ratio is:

[0064]

[0065] where is the normalized weight of the side-edge wing lift-drag ratio in the subsonic state, is the normalized weight of the side-edge wing lift-drag ratio in the transonic state, is the normalized weight of the side-edge wing lift-drag ratio in the supersonic state;

[0066] The penalty function of the side-edge wing thickness is:

[0067]

[0068] where is the normalized weight of the side-edge wing thickness.

[0069] Furthermore, in step 2.2, the NACA64A204 airfoil is also used as the basic airfoil of the backup airfoil, and the optimization of the backup airfoil aiming to improve the lift-drag ratio is carried out under three working conditions: Ma = 4, altitude of 30 km, angle of attack of 5°, and Ma = 5, altitude of 30 km, angle of attack of 5°, and Ma = 6, altitude of 30 km, angle of attack of 5°. The established optimization model is:

[0070]

[0071] where is the optimization objective of the backup airfoil, is the objective function of the backup airfoil, is the weighted coefficient of the penalty function of the backup airfoil, is the lift coefficient penalty function of the backup airfoil, is the drag coefficient penalty function of the backup airfoil, is the lift-drag ratio penalty function of the basic airfoil, is the thickness penalty function of the basic airfoil; represents the constraint condition, are the lift-drag ratios of the backup optimized airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the lift-drag ratios of the basic airfoil of the backup airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the lift coefficients of the backup optimized airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the lift coefficients of the basic airfoil of the backup airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the drag coefficients of the backup optimized airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, They are respectively the drag coefficients of the basic airfoil of the standby airfoil under three working conditions of Ma = 4, 5, and 6. is the maximum thickness of the standby optimized airfoil. is the maximum thickness of the basic airfoil of the standby airfoil.

[0072] The Basefit3 expression is:

[0073]

[0074] Among them They are respectively the lift-drag ratio weighting coefficients of the standby optimized airfoil under the working condition of Ma = 4, the lift-drag ratio weighting coefficient under the working condition of Ma = 5, and the lift-drag ratio weighting coefficient under the working condition of Ma = 6 after normalization.

[0075] The penalty function expression of the lift coefficient of the standby airfoil is:

[0076]

[0077] Among them is the normalization weight of the lift coefficient of the standby airfoil under the working condition of Ma = 4. is the normalization weight of the lift coefficient of the standby airfoil under the working condition of Ma = 5. is the normalization weight of the lift coefficient of the standby airfoil under the working condition of Ma = 6.

[0078] The penalty function expression of the drag coefficient of the standby airfoil is:

[0079]

[0080] Among them is the normalization weight of the drag coefficient of the standby airfoil under the working condition of Ma = 4. is the normalization weight of the drag coefficient of the standby airfoil under the working condition of Ma = 5. is the normalization weight of the drag coefficient of the standby airfoil under the working condition of Ma = 6.

[0081] The penalty function expression of the lift-drag ratio of the standby airfoil is:

[0082]

[0083] Among them is the normalization weight of the lift-drag ratio of the standby airfoil under the working condition of Ma = 4. is the normalization weight of the lift-drag ratio of the standby airfoil under the working condition of Ma = 5. is the normalization weight of the lift-drag ratio of the standby airfoil under the working condition of Ma = 6.

[0084] The penalty function of the thickness of the standby airfoil is:

[0085]

[0086] Wherein is the normalized weight of the backup airfoil thickness.

[0087] Furthermore, in step 5, the three-dimensional layout parameters include the first sweep angle of the strake wing , the span of the strake wing , the length obtained by subtracting the span of the strake wing from the span of the basic wing , and the mathematical representation is:

[0088]

[0089] Wherein is the second sweep angle of the basic wing determined according to the transonic constraint, is the dihedral angle of the lower surface of the waverider, is the central angle, R is the radius of the conical shock wave extending to the trailing edge position of the waverider afterbody, is the offset distance.

[0090] Advantageous effects:

[0091] Compared with the prior art, the present invention has the following advantageous technical effects: By combining the cone-guided waverider with the closely-coned double-swept waverider design method and the transonic strake wing / basic wing design constraints, the overall waverider design of the waverider wide-speed domain aircraft is realized, and the pressure leakage on the lower surface of the fuselage at supersonic speed is suppressed; at the same time, strake vortices and leading-edge vortices are generated at low speed, and the lift coefficient of the wide-speed domain aircraft at low speed is improved. By determining the intersection point of the shock wave and the left-running characteristic line, the proportion of the waverider forebody in the overall length of the aircraft and the length of the waverider afterbody are determined, realizing supersonic drag reduction and an increase in the lift-to-drag ratio. By designing the shape of the symmetry plane to suppress the separation of the airflow at the trailing edge, a large-scale low-pressure area is generated on the upper surface of the waverider afterbody at supersonic speed. Using the surrogate model to optimize the multi-airfoil section greatly improves the aerodynamic performance of the airfoil in the subsonic / transonic / supersonic speed range and can meet the assembly requirements of the wide-speed domain aircraft. The above-mentioned various design means are comprehensively applied in the design method of the wide-speed domain aircraft, ultimately improving the aerodynamic performance of the wide-speed domain aircraft in the subsonic / transonic / supersonic speed range.

[0092] The additional aspects and advantages of the present invention will be partly given in the following description, partly will become obvious from the following description, or be understood through the practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0093] The above and / or additional aspects and advantages of the present invention will become obvious and easy to understand from the description of the embodiments in conjunction with the following drawings, wherein:

[0094] Figure 1 is the design framework flow chart of the present invention;

[0095] Figure 2 Generate a schematic diagram of the parameters involved in the conical leading-edge waverider;

[0096] Figure 3 : Comparison diagram of the NACA64A204 airfoil and the optimized airfoil of the basic airfoil;

[0097] Figure 4 : Comparison diagram of the airfoils of the leading-edge strips before and after optimization;

[0098] Figure 5 : Geometric parameters of the wide-speed-range aircraft layout calculated by the conical leading-edge waverider and the osculating cone method;

[0099] Figure 6 : Comparison of the lift coefficients of the wide-speed-range flat trailing-edge configuration designed in this example and the reference layout at different angles of attack in the subsonic regime;

[0100] Figure 7 : Comparison of the drag coefficients of the wide-speed-range flat trailing-edge configuration designed in this example and the reference layout at different angles of attack in the transonic regime;

[0101] Figure 8 : Comparison of the lift-to-drag ratios of the wide-speed-range flat trailing-edge configuration designed in this example and the reference layout at different angles of attack in the supersonic regime. Specific implementation manners

[0102] The embodiments of the present invention will be described in detail below. The embodiments are exemplary and are intended to explain the present invention and should not be construed as limiting the present invention.

[0103] Refer to Figure 1 , the wing-body integrated wide-speed-range aircraft layout design method based on considering the afterbody effect proposed in this embodiment includes the following steps:

[0104] Step 1: Use the conical leading-edge method to generate the waverider nose for the wide-speed-range aircraft for supersonic flight, that is, generate the conical leading-edge forebody of the wide-speed-range aircraft. In the field of waverider design, the conical leading-edge method (Conical Flowfield Method) is a classic waverider design method. Its core idea is based on the conical shock wave theory, and the waverider shape is generated by tracking the streamlines behind the shock wave. The specific process is given below:

[0105] Step 1.1: Determine the cone half-angle of the conical shock wave flow field and the incoming flow Mach number. In this example, the cone half-angle , the incoming flow Mach number Ma = 5, and specify the length of the conical flow field .

[0106] Step 1.2: Specify the conical bottom profile line, find the surface passing through the conical bottom profile line and parallel to the oncoming flow direction. The intersection line of the surface and the shock wave surface is the leading edge line of the waverider, and the surface formed by the leading edge line and the bottom profile line is the upper surface of the waverider; as Figure 2 shown, in this embodiment, the conical bottom profile line of the upper surface of the waverider is a quartic equation, expressed as:

[0107]

[0108] To determine the three parameters a, b, and c, three geometric parameters are selected:

[0109] The vertical ratio between the upper surface and the lower surface of the waverider : , where c is the ordinate of the upper surface of the waverider, is the ordinate of the intersection point of the lower surface edge of the waverider and the shock wave (points A and B in Figure 2 );

[0110] The spanwise width : is the absolute value of the angle between the line connecting point A or B and the origin and the Z-axis;

[0111] The inclination angle of the lower surface edge : is the angle corresponding to the tangent slope of the lower surface of the waverider at point A or B.

[0112] In the case where the length of the conical flow field is determined, the vertex coordinates of the quartic function and the tangent equations at both ends can be determined through the above three geometric parameters, so as to calculate the coefficients of the quartic equation. For the specific calculation method, refer to the paper (Miao Meng, Shi Jiatong, Yan Chao, et al. Optimization Design and Analysis of Hypersonic Cone-Guided Waverider [C]. The 15th National Conference on Computational Fluid Dynamics, Yantai, August 4-7, 2012). In this embodiment , , .

[0113] Step 1.3: Starting from the leading edge line of the waverider, according to the Taylor-Maccoll equation, the lower surface of the waverider is obtained through streamline tracing to generate a cone-guided waverider forebody. For the specific calculation method, refer to the paper (Wang Xiaoyan. Research on the Design Method of Waverider Based on 3D Leading Edge Line [D]. Changsha: National University of Defense Technology, 2018). The length of the cone-guided waverider forebody obtained in this embodiment .

[0114] Step 2: Based on the basic airfoil, aerodynamic optimization is carried out on the basic airfoil and the strake airfoil of the wide-speed domain aircraft to obtain a multi-section wide-speed domain waverider airfoil.

[0115] The wing of the wide speed range aircraft is a strake-wing and fuselage blended wing, which consists of two parts: a strake wing used to modify the transition between the wing and the fuselage, with an arched leading edge and used to generate vortices at transonic / supersonic speeds; and a wing part with a relatively small leading edge sweep angle and relatively straight, called the basic wing, the definition of which can be found in (Editorial Committee of the Aircraft Design Manual. Aircraft Design Manual, Volume 6: Aerodynamic Design [M]. Beijing: Aviation Industry Press, 2002: 145-155.).

[0116] In order to enable both the strake wing and the basic wing to have good wide speed range aerodynamic performance, different optimized airfoils need to be assembled on the basic wing and the strake wing to meet the aerodynamic performance requirements of different wing segments. The specific process is as follows:

[0117] Step 2.1: Determine the basic airfoil; in this embodiment, as Figure 3 shown, for the basic wing, the NACA64A204 airfoil is selected as the basic airfoil; as Figure 4 shown, for the strake wing, an inverted wedge with a straight bottom, the maximum thickness at 45% of the chord length, the maximum thickness being 3.5% of the chord length, and a leading edge radius of 10 mm is selected as the basic airfoil.

[0118] Perform parametric processing on the basic airfoil. Here, the conventional 5th-order CST parametric method in the art is used to perform parametric processing on the upper and lower surfaces of the basic airfoil respectively, with a total of 12 design variables (i.e., optimization parameters). Its mathematical expression is:

[0119]

[0120] where is the abscissa of the airfoil, is the ordinate of the upper surface of the airfoil, is the ordinate of the lower surface of the airfoil, is the class function, is the shape function, the subscript of and represent the upper surface and the lower surface respectively, and represent the trailing edge coordinates of the upper surface and the lower surface in sequence. The class function The N1 and N2 in are constants, taking 0.5 and 1 respectively. The N in the shape function is the order, which is taken as 5 in this embodiment, and are undetermined coefficients, i.e., optimization parameters. For the basic wing airfoil, the design space of the upper and lower surface optimization parameters is set to [-0.03, 0.03]. For the strake wing airfoil, the design space of the upper surface optimization parameter is set to [-0.03, 0.03], while no design space is set for the lower surface, so as to maintain the straight line feature of the lower surface of the strake wing inverted wedge airfoil.

[0121] Step 2.2: Set the design conditions and optimization models for optimizing the basic wing airfoil and the strake wing airfoil.

[0122] The design conditions selected are: Ma = 0.4, angle of attack = 2°, altitude = 10 km; Ma = 1, angle of attack = 2°, altitude = 10 km; Ma = 5, angle of attack = 5°, altitude = 30 km; corresponding to subsonic, transonic and supersonic conditions respectively.

[0123] For the basic wing airfoil, the optimization model is:

[0124]

[0125] where is the optimization objective of the basic wing, is the objective function of the basic wing, is the penalty function weighting coefficient of the basic wing, is the lift coefficient penalty function of the basic wing, is the drag coefficient penalty function of the basic wing, is the lift-to-drag ratio penalty function of the basic wing, is the thickness penalty function of the basic wing; represents the constraint condition, are the lift-to-drag ratios of the optimized airfoil of the basic wing under supersonic, transonic and subsonic conditions respectively, are the lift-to-drag ratios of the basic airfoil of the basic wing under supersonic, transonic and subsonic conditions respectively, are the lift coefficients of the optimized airfoil of the basic wing under subsonic, transonic and supersonic conditions respectively, are the lift coefficients of the basic airfoil of the basic wing under subsonic, transonic and supersonic conditions respectively, are the drag coefficients of the optimized airfoil of the basic wing under subsonic, transonic and supersonic conditions respectively, are the drag coefficients of the basic airfoil of the basic wing under subsonic, transonic and supersonic conditions respectively, is the maximum thickness of the optimized airfoil of the basic wing, is the maximum thickness of the basic airfoil of the basic wing.

[0126] Since the contribution of a small sweep angle to subsonic lift is dominant, the goal of optimization under subsonic conditions is to maximize the lift of a wide-speed aircraft. Transonic drag reduction is the main means to improve the transonic performance of current wide-speed aircraft, and the lift-to-drag ratio is the most important factor during supersonic cruise. Therefore, in this embodiment, the objective function of the basic wing is as follows:

[0127]

[0128] where are respectively the weighted coefficients of the subsonic lift coefficient, the transonic drag coefficient, and the supersonic lift-to-drag ratio of the optimized airfoil of the basic wing after normalization, and are obtained through the formula

[0129]

[0130] where are all set weighted coefficients.

[0131] The penalty function is obtained by evaluating and weighting the aerodynamic parameters under the three conditions. The evaluation process of the aerodynamic parameters will judge whether the aerodynamic parameters of the optimization result are lower than the initial values. If the aerodynamic parameters are higher than the initial values, the penalty function is set to 0. If they are lower than the initial values, the penalty function returns the difference between the original aerodynamic parameters and the optimization result. Specifically:

[0132] The penalty function expression of the basic wing lift coefficient is:

[0133]

[0134] where is the normalization weight of the basic wing lift coefficient in the subsonic state, is the normalization weight of the basic wing lift coefficient in the transonic state, is the normalization weight of the basic wing lift coefficient in the supersonic state.

[0135] The penalty function expression of the basic wing drag coefficient is:

[0136]

[0137] where is the normalization weight of the basic wing drag coefficient in the subsonic state, is the normalization weight of the basic wing drag coefficient in the transonic state, is the normalization weight of the basic wing drag coefficient in the supersonic state.

[0138] The penalty function expression of the basic wing lift-to-drag ratio is:

[0139]

[0140] where is the normalized weight of the basic wing lift - to - drag ratio in the subsonic state, is the normalized weight of the basic wing lift - to - drag ratio in the transonic state, is the normalized weight of the basic wing lift - to - drag ratio in the supersonic state.

[0141] The expression of the basic wing thickness penalty function is:

[0142]

[0143] where is the normalized weight of the basic wing thickness.

[0144] For the strake wing, the optimization model is:

[0145]

[0146] where is the strake wing optimization objective, is the strake wing objective function, is the strake wing penalty function weighting coefficient, is the strake wing lift coefficient penalty function, is the strake wing drag coefficient penalty function, is the strake wing lift - to - drag ratio penalty function, is the strake wing thickness penalty function; represents the constraint condition, are respectively the lift - to - drag ratios of the optimized strake wing airfoil in supersonic, transonic and subsonic working conditions in sequence, are respectively the lift - to - drag ratios of the basic strake wing airfoil in supersonic, transonic and subsonic working conditions in sequence, are respectively the lift coefficients of the optimized strake wing airfoil in subsonic, transonic and supersonic working conditions in sequence, are respectively the lift coefficients of the basic strake wing airfoil in subsonic, transonic and supersonic working conditions in sequence, are respectively the drag coefficients of the optimized strake wing airfoil in subsonic, transonic and supersonic working conditions in sequence, are respectively the drag coefficients of the basic strake wing airfoil in subsonic, transonic and supersonic working conditions in sequence, is the maximum thickness of the optimized strake wing airfoil, is the maximum thickness of the basic strake wing airfoil.

[0147] Since for the strake wing, the lift - to - drag ratio is the most important factor, the Basefit2 expression is:

[0148]

[0149] where They are respectively the subsonic lift-drag ratio weighting coefficient, the transonic lift-drag ratio weighting coefficient, and the supersonic lift-drag ratio weighting coefficient of the optimized airfoil of the strake wing after normalization.

[0150] Similarly, the penalty function is obtained by evaluating and weighting the aerodynamic parameters under three working conditions. The evaluation process of the aerodynamic parameters will judge whether the aerodynamic parameters of the optimization result are lower than the initial values. If the aerodynamic parameters are higher than the initial values, the penalty function is set to 0. If they are lower than the initial values, the penalty function returns the difference between the original aerodynamic parameters and the optimization result. Specifically:

[0151] The penalty function expression of the strake wing lift coefficient is:

[0152]

[0153] where is the normalization weight of the strake wing lift coefficient in the subsonic state, is the normalization weight of the strake wing lift coefficient in the transonic state, is the normalization weight of the strake wing lift coefficient in the supersonic state.

[0154] The penalty function expression of the strake wing drag coefficient is:

[0155]

[0156] where is the normalization weight of the strake wing drag coefficient in the subsonic state, is the normalization weight of the strake wing drag coefficient in the transonic state, is the normalization weight of the strake wing drag coefficient in the supersonic state.

[0157] The penalty function expression of the strake wing lift-drag ratio is:

[0158]

[0159] where is the normalization weight of the strake wing lift-drag ratio in the subsonic state, is the normalization weight of the strake wing lift-drag ratio in the transonic state, is the normalization weight of the strake wing lift-drag ratio in the supersonic state.

[0160] The penalty function of the strake wing thickness is:

[0161]

[0162] where is the normalization weight of the strake wing thickness.

[0163] In addition, in this embodiment, the NACA64A204 airfoil is also used as the basic airfoil of the backup airfoil, and the optimization of the backup airfoil aiming to improve the lift-drag ratio is carried out under three working conditions: Ma = 4, altitude of 30 km, angle of attack of 5°, Ma = 5, altitude of 30 km, angle of attack of 5°, and Ma = 6, altitude of 30 km, angle of attack of 5°. The established optimization model is:

[0164]

[0165] where is the optimization target of the backup airfoil, is the objective function of the backup airfoil, is the penalty function weighting coefficient of the backup airfoil, is the lift coefficient penalty function of the backup airfoil, is the drag coefficient penalty function of the backup airfoil, is the lift-drag ratio penalty function of the basic airfoil, is the thickness penalty function of the basic airfoil; represents the constraint condition, are the lift-drag ratios of the backup optimized airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the lift-drag ratios of the basic airfoil of the backup airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the lift coefficients of the backup optimized airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the lift coefficients of the basic airfoil of the backup airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the drag coefficients of the backup optimized airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, are the drag coefficients of the basic airfoil of the backup airfoil under the three working conditions of Ma = 4, 5, and 6 in sequence, is the maximum thickness of the backup optimized airfoil, is the maximum thickness of the basic airfoil of the backup airfoil.

[0166] The expression of Basefit3 is:

[0167]

[0168] where are the lift-drag ratio weighting coefficients of the backup optimized airfoil under the Ma = 4 working condition, the lift-drag ratio weighting coefficient under the Ma = 5 working condition, and the lift-drag ratio weighting coefficient under the Ma = 6 working condition after normalization in sequence.

[0169] The expression of the lift coefficient penalty function of the backup airfoil is:

[0170]

[0171] Among them is the normalized weight of the lift coefficient of the backup airfoil under the condition of Ma = 4, is the normalized weight of the lift coefficient of the backup airfoil under the condition of Ma = 5, is the normalized weight of the lift coefficient of the backup airfoil under the condition of Ma = 6.

[0172] The penalty function expression of the drag coefficient of the backup airfoil is:

[0173]

[0174] Among them is the normalized weight of the drag coefficient of the backup airfoil under the condition of Ma = 4, is the normalized weight of the drag coefficient of the backup airfoil under the condition of Ma = 5, is the normalized weight of the drag coefficient of the backup airfoil under the condition of Ma = 6.

[0175] The penalty function expression of the lift-to-drag ratio of the backup airfoil is:

[0176]

[0177] Among them is the normalized weight of the lift-to-drag ratio of the backup airfoil under the condition of Ma = 4, is the normalized weight of the lift-to-drag ratio of the backup airfoil under the condition of Ma = 5, is the normalized weight of the lift-to-drag ratio of the backup airfoil under the condition of Ma = 6.

[0178] The penalty function of the thickness of the backup airfoil is:

[0179]

[0180] Among them is the normalized weight of the thickness of the backup airfoil.

[0181] Step 2.3: For the optimization model established in Step 2.2, within the design space, use the conventional Latin hypercube sampling method in the field to conduct sampling, use Tcl script to automatically generate grids, use the Reynolds-averaged Navier-Stokes (RANS) method to calculate the aerodynamic parameters of the samples as the data input of the surrogate model, and use the conventional particle swarm optimization algorithm in the field to solve the optimization model to achieve multi-objective optimization, and finally obtain the optimized airfoils of the basic airfoil, the strake airfoil, and the backup airfoil.

[0182] Among them, the surrogate model adopts the conventional Kriging surrogate model in this field. The Kriging surrogate model is an interpolation model, and the interpolation result is a linear weighting of the sample function response values, and the weighting coefficients are calculated through a Gaussian exponential model. Since the optimization objective is an optimization function related to the lift coefficient, drag coefficient, and lift-to-drag ratio, in this process, it is necessary to establish surrogate models corresponding to the lift coefficient, drag coefficient, and lift-to-drag ratio for 12 design variables respectively.

[0183] In the particle swarm algorithm, each particle first obtains an initial velocity and position, evaluates the results of each particle using the optimization function, finds the individual best position pbest and the global best position gbest, adds an inertia weight to the initial velocity, and adds learning factors to the individual best and global best respectively. After completion of the initialization, the particle velocity and particle position are updated. This process indicates that the particles move according to the inertia weight and two learning rates respectively. The inertia weight depends on the initial velocity of the particle, and the two learning rates determine the weights of the movements towards the individual best and global best positions in the total movement velocity. The movement velocities towards the individual best and global best positions are random values, and their velocity ranges are half of the design space. The movement direction is the vector difference between the individual best and global best and the current position coordinates. After obtaining the movement velocity, the particle position is updated. If it is found that the particle exceeds the design space when updating the particle position, the coordinates exceeding the design space are commanded to be the maximum value allowed by the movement direction. After updating the particle position, the new particle position is used as the initial value for the next round of particle swarm optimization, and the next round of optimization is carried out until the maximum number of iterations is reached or the error of the particle is less than the threshold. The optimization result obtained at this time is the optimization result obtained by using the surrogate model in this round.

[0184] Finally, the optimization result of the basic wing is as Figure 3 shown, and the optimization result of the strake wing is as Figure 4 shown. The aerodynamic parameters after airfoil optimization are as follows:

[0185]

[0186] After the basic wing is optimized, the maximum thickness of the airfoil moves backward, the leading edge becomes a sharp leading edge, and the lower surface presents a double-S characteristic. As can be seen from the table, the subsonic lift coefficient and supersonic lift-to-drag ratio of the optimized airfoil of the basic wing are optimized, and the transonic drag coefficient increases by 2.2×10 -4 , but the transonic lift-to-drag ratio is improved. The improvement of the aerodynamic parameters is due to the expansion of the low-pressure area on the upper surface in the subsonic / transonic range, and the optimization of the supersonic performance is due to the double-S characteristic on the lower surface generating two stronger high-pressure areas, and the pressure coefficient on the upper surface is reduced.

[0187] After optimization, the maximum thickness of the strake wing airfoil is shifted to 52%, and the maximum thickness is reduced to 2.6% of the chord length. The lift-drag ratio under each working condition is improved. This improvement is due to the more uniform low-pressure distribution on the upper surface at subsonic speeds and the more obvious suction peak. At transonic speeds, the pressure coefficient at the leading edge of the upper surface is significantly reduced, and at supersonic speeds, a strong compression wave is maintained on the lower surface.

[0188] In this example, after optimizing the NACA64A204 airfoil under the working conditions of Ma = 4 / 5 / 6, the aerodynamic parameters of the alternative optimized airfoil are as follows in the table:

[0189]

[0190] Step 3: Calculate the length of the afterbody of the waverider, which specifically includes the following steps:

[0191] Step 3.1: The afterbody of the waverider is the fuselage area behind the waverider body. Since the disturbance of the downstream flow field at supersonic speeds can only propagate backward, and according to the characteristic line theory, the flow field characteristics of a point are determined by the flow field in front of the left-running characteristic line passing through this point, the aerodynamic characteristics of the afterbody are related to the left-running characteristic line passing through the trailing edge of the forebody of the waverider and the leading edge of the afterbody of the waverider.

[0192] In order to determine the length of the afterbody of the waverider, in this embodiment, first, according to the shock wave angle β of the flow field, the incoming flow Mach number Ma, the specific heat ratio γ, and the surface angle θ, the Mach number behind the wave and the angle between the left-running characteristic line family and the center line are calculated. The mathematical expression is:

[0193] M 2 = 1 + [ ( γ − 1 ) / 2 ] M a 2 s i n 2 β γ M a 2 s i n 2 β − ( γ − 1 ) / 2 1 s i n ( β − θ ) α = s i n − 1 ( 1 / M 2 )

[0194] Step 3.2: Let the left-running characteristic line start from the trailing edge of the forebody of the waverider, find the intersection point of the left-running characteristic line and the shock wave, so as to calculate the proportion of the forebody of the waverider in the total length of the whole aircraft , and in the case of knowing the length of the forebody of the waverider, the length of the afterbody of the waverider can be deduced. The specific data expression is:

[0195]

[0196] The length of the afterbody of the waverider calculated in this embodiment.

[0197] Step 4: Design the two-dimensional shape of the symmetry plane of the wide-speed-range aircraft; in the present invention, when designing the two-dimensional shape of the symmetry plane, the symmetry plane of the forebody of the waverider remains unchanged, and the original shape of the symmetry plane of the afterbody of the waverider is set to be all straight lines on the upper and lower sides. Then, on the basis of the original shape, first keep one side as a straight line and adjust the other side; calculate the aerodynamic parameters of the adjusted symmetry plane under subsonic, transonic, and supersonic working conditions.

[0198] In this embodiment, the following five symmetry planes are selected. Specifically, for the lower side of the symmetry plane being a straight line, the upper sides respectively adopt the upper curve of the basic wing optimized airfoil, the upper curve of the alternative optimized airfoil, and the upper curve of the 20° contraction shape. For the upper side of the symmetry plane being a straight line, the lower side adopts the lower curve of the basic wing optimized airfoil and the lower curve of the alternative optimized airfoil. Aerodynamic calculations are performed, and the aerodynamic evaluation results of each symmetry plane are as follows:

[0199]

[0200] The 20° contraction shape uses an upper surface shape based on the design of the aircraft afterbody, and its shape can be referred to the paper (MARTENS R E. F-15 Nozzle / afterbody integration[J]. Journal of Aircraft.1976, 13(5): 327-333.).

[0201] The aerodynamic performance of the aerodynamic parameters of each symmetry plane is evaluated using a weighting function. The mathematical expression of the weighting function is:

[0202]

[0203] where represents the weighting function of the symmetry plane, , and respectively represent the subsonic lift coefficient, transonic drag coefficient, and supersonic lift-to-drag ratio of the symmetry plane in sequence; , and respectively represent the weighting coefficient of the subsonic lift coefficient of the symmetry plane, the weighting coefficient of the transonic drag coefficient of the symmetry plane, and the weighting coefficient of the supersonic lift-to-drag ratio of the symmetry plane in sequence. The weighting coefficients are normalized using the subsonic lift coefficient, transonic drag coefficient, and supersonic lift-to-drag ratio of the original shape of the symmetry plane respectively. The smaller the calculation result of the weighting function, the better the aerodynamic performance of the symmetry plane.

[0204] The weighting calculation results in this embodiment show that: for the symmetry plane with a straight lower side and the upper curve of the 20° contraction shape on the upper side, the weighted aerodynamic efficiency is the best, and this symmetry plane also has the largest area among the non-original shapes. Therefore, the upper side of the symmetry plane is determined to be the upper curve of the 20° contraction shape.

[0205] Then, with the upper side of the symmetry plane fixed as the upper curve of the 20° contraction shape, the lower side of the symmetry plane respectively adopts the lower curve of the basic wing optimized airfoil and the lower curve of the alternative optimized airfoil, and the calculated aerodynamic parameters are:

[0206]

[0207] The aerodynamic parameters of each symmetry plane are evaluated using a weighting function to assess their performance. Based on the weighted results, it is analyzed that the optimization effect of using an airfoil on the lower surface is inferior to maintaining a flat plane. Therefore, in this example, a straight line is used for the lower surface of the fuselage to obtain the final two-dimensional shape of the symmetry plane. The improvement of the aerodynamic parameters of the symmetry plane is due to suppressing the airflow separation at the trailing edge, expanding the low-pressure area at the leading edge at subsonic speeds, and there are multiple weak expansion waves on the upper surface at transonic and supersonic speeds, generating a large-scale low-pressure area at the trailing edge.

[0208] Step 5: Based on the osculating cone method and the transonic strake wing constraint, use the waverider forebody parameters and the waverider afterbody length to obtain the three-dimensional layout parameters, and assemble the basic wing optimized airfoil, strake wing optimized airfoil, and symmetry plane to generate a wide-speed domain aircraft layout.

[0209] In this embodiment, as Figure 5 shown, assume that the conical shock wave flow field extends backward, and the extension length is equal to , and use the shock wave angle β of the flow field and the trigonometric function relationship to obtain the conical shock wave radius R extending to the trailing edge position of the waverider afterbody. Substitute the parameters of the conical-guided waverider forebody into the osculating cone method: k is equal to the offset distance , the spanwise width ψ is equal to the central angle , and at the same time retain the dihedral angle on the lower surface of the waverider body in the osculating cone method for auxiliary calculation, and the actual dihedral angle of the basic wing remains 0°. Finally, the first sweep angle of the strake wing, the span of the strake wing, and the length after subtracting the span of the strake wing from the span of the basic wing are derived. The mathematical representation in this process is:

[0210]

[0211] where is the second sweep angle of the basic wing determined according to the transonic constraint.

[0212] Finally, the parameters obtained in this embodiment are: the second sweep angle of the basic wing, the first sweep angle of the strake wing, the span of the strake wing, the chord length of the strake wing is 5592.384 mm, the total span is 6935.446 mm, the tip chord length of the basic wing is 1917.302 mm, the root chord length of the basic wing is 7477.479 mm, and the taper ratio is 3.9.

[0213] Finally, aerodynamic calculations are performed on the obtained wide-speed domain aircraft layout to obtain aerodynamic parameters, and a comparison is made with a certain type of reference layout:

[0214]

[0215] The results show that the flat trailing edge configuration proposed in this embodiment is comprehensively superior to the control layout in terms of the target aerodynamic parameters. The flat trailing edge configuration sacrifices the proportion of the waverider region on the lower surface of the fuselage in the full length of the aircraft compared to the control layout, resulting in a lower lift coefficient at supersonic speeds than the control configuration. However, the lift-to-drag ratio is improved by a significant reduction in the drag coefficient. Referring to Figures 6 - 8 , in this embodiment, the flat trailing edge configuration is superior to the control layout in terms of changing the angle of attack, subsonic lift coefficient, and supersonic lift-to-drag ratio under the design conditions, and the transonic drag coefficient is superior to the control layout at small angles of attack. At typical flight altitudes and angles of attack, the configuration of this embodiment has good wide-speed range performance.

[0216] The improvement in aerodynamic performance under the supersonic design conditions of this example is due to the effective suppression of the pressure leakage at the strake wing. The numerical simulation results show that the lift-to-drag ratio of the strake wing in the control layout is 6.51, and the lift-to-drag ratio of the strake wing in the configuration of this example is 10.31, further indicating that the pressure leakage at the strake wing is suppressed. The local high-pressure region in the upper surface flow field is more due to the detached compression wave caused by the blunting treatment and the pressure distribution characteristics of the airfoil. The aft body contraction design utilizes the vortex-wave effect at supersonic speeds, generating expansion waves and vortices on the upper surface, and reducing the pressure coefficient on the upper surface.

[0217] The improvement in aerodynamic performance under the transonic design conditions of this example is due to the aft body design reducing the transonic drag. The wave-rider aft body drag in the configuration of this example is reduced to 9.064×10 -3 , and the proportion of the drag coefficient in the whole aircraft is reduced from 44.68% to 34.73%.

[0218] The improvement in subsonic performance of this example is mainly due to the detached vortices generated by the leading edge of the waverider, the strake wing, and the basic wing reducing the pressure coefficient on the upper surface of the aircraft, and the improvement of the lift-to-drag characteristics by the wave-rider aft body design. In the lift augmentation effect of vortex lift, the vortices of the wave-rider forebody dominate at small angles of attack, and the vortices on the wave-rider forebody, strake wing, and basic wing jointly reduce the pressure coefficient on the upper surface of the aircraft at large angles of attack. In the wave-rider aft body design, the lift coefficient provided by the aft body of the flat trailing edge layout is increased by 40.32% compared to the aft body of the control layout, and the drag coefficient is reduced by 37.70%, achieving a significant improvement in the lift-to-drag ratio.

[0219] Different from the previous wide-speed range aircraft design methods, the present invention combines the cone-guided waverider, the osculating cone method, and the transonic strake wing / basic wing design method, and combines the aerodynamic design of the wave-rider aft body to design a wide-speed range aircraft layout with a higher lift coefficient at subsonic speeds, drag reduction at transonic speeds, and a high lift-to-drag ratio at supersonic speeds. Compared with the traditional wide-speed range aircraft design methods, the wide-speed range aircraft designed by this method has advantages in aerodynamic parameters, with obvious vortex lift at subsonic speeds, suppression of surface pressure leakage at supersonic speeds, and utilization of the vortex-wave effect on the upper surface. And in the design process, the aerodynamic aft body design is comprehensively considered, which can provide a reference for the subsequent wide-speed range aircraft decoupled from the power system.

[0220] Although the embodiments of the present invention have been shown and described above, it can be understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those of ordinary skill in the art can make changes, modifications, substitutions, and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. A layout design method for a wing-body blended wide-speed-range aircraft considering the afterbody effect, characterized in that: It includes the following steps: Step 1: Generate a conical-guided waverider forebody for a wide-speed-range aircraft for supersonic flight; Step 2: Based on the basic airfoil, conduct aerodynamic optimization on the basic wing airfoil and strake wing airfoil of the wide-speed-range aircraft to obtain a multi-section wide-speed-range waverider airfoil: Step 2.1: Determine the basic airfoil of the basic wing and the basic airfoil of the strake wing respectively; perform parametric processing on the basic airfoil to obtain optimization parameters and determine the design space of the optimization parameters; Step 2.2: Set the design conditions and optimization models for optimizing the basic wing airfoil and the strake wing airfoil; the design conditions are divided into subsonic, transonic, and supersonic conditions; Step 2.3: For the optimization model established in Step 2.2, sample the optimization parameters within the design space, use the surrogate model, and solve the optimization model through the optimization algorithm to achieve multi-objective optimization and obtain the optimized airfoil of the basic wing and the optimized airfoil of the strake wing; Step 3: Calculate the length of the waverider afterbody: Step 3.1: First, according to the shock wave angle β of the flow field, the incoming flow Mach number Ma, the specific heat ratio γ, and the surface angle θ, according to the formula Calculated post-shock Mach number and the angle between the left-running characteristic family and the centerline ; Step 3.2: According to the formula The length of the afterbody of the waverider is calculated ; where is the proportion of the forebody of the waverider in the overall length of the aircraft, is the length of the forebody of the waverider; Step 4: Design the two-dimensional shape of the symmetric plane of the wide-speed-range aircraft: Keep the symmetric plane of the waverider forebody unchanged, set the original shape of the symmetric plane of the waverider afterbody to be all straight lines on the upper and lower sides, on the basis of the original shape, keep one side straight and adjust the other side; calculate the aerodynamic parameters of the adjusted symmetric plane of the wide-speed-range aircraft under subsonic, transonic, and supersonic conditions, and according to the formula for the aerodynamic parameters Perform weighted calculation to obtain the aerodynamic performance of the symmetry plane of the wide-speed-range aircraft after adjustment; among them represents the weighting function of the symmetry plane, , and respectively represent the subsonic lift coefficient, transonic drag coefficient and supersonic lift-drag ratio of the symmetry plane in sequence; , and respectively represent the weighting coefficient of the subsonic lift coefficient of the symmetry plane, the weighting coefficient of the transonic drag coefficient of the symmetry plane and the weighting coefficient of the supersonic lift-drag ratio of the symmetry plane in sequence; the smaller the calculation result of the weighting function, the better the aerodynamic performance of the symmetry plane; obtain the symmetry plane of the wide-speed-range aircraft with the best aerodynamic performance Step 5: Based on the osculating cone method and the transonic strake wing constraint, use the waverider forebody parameters and the length of the waverider afterbody to obtain the three-dimensional layout parameters, and assemble the optimized airfoil of the basic wing, the optimized airfoil of the strake wing, and the symmetric plane to generate the layout of the wide-speed-range aircraft.

2. The layout design method of a wing-body fusion wide-speed-range aircraft based on considering afterbody effect according to claim 1, characterized in that: In Step 1, the conical-guided waverider forebody for a wide-speed-range aircraft for supersonic flight is generated using the conical guidance method.

3. The layout design method of a wing-body fusion wide-speed-range aircraft based on considering the afterbody effect according to claim 2, wherein: The specific process of generating the conical-guided waverider forebody for a wide-speed-range aircraft for supersonic flight in Step 1 is as follows: Step 1.1: Determine the cone half-cone angle and the incoming flow Mach number Ma that generate the conical shock wave flow field, and specify the length of the conical flow field ; Step 1.2: Specify the conical bottom profile line, find the surface passing through the conical bottom profile line and parallel to the incoming flow direction, and the intersection line of the surface and the shock wave surface is the leading edge line of the waverider, and the surface formed by the leading edge line and the bottom profile line is the upper surface of the waverider; Step 1.3: Starting from the leading edge line of the waverider, according to the Taylor-Maccoll equation, obtain the lower surface of the waverider through streamline tracing to generate the conical-guided waverider forebody.

4. A layout design method for a wing-body blended wide-speed-range aircraft based on considering afterbody effects as claimed in claim 1, characterized in that: In Step 2.1, for the basic wing, select the NACA64A204 airfoil as the basic airfoil; for the strake wing, select an inverted wedge with a flat bottom, the maximum thickness at 45% of the chord length, the maximum thickness being 3.5% of the chord length, and a leading edge radius of 10 mm as the basic airfoil.

5. The layout design method of a wing-body fusion wide-speed-range aircraft based on considering the afterbody effect according to claim 1, characterized in that: In Step 2.2, select the design conditions as: Ma = 0.4, 2° angle of attack, 10 km altitude; Ma = 1, 2° angle of attack, 10 km altitude; Ma = 5, 5° angle of attack, 30 km altitude; corresponding to subsonic, transonic, and supersonic conditions respectively.

6. The layout design method of a wing-body fusion wide-speed regime aircraft based on considering the afterbody effect according to claim 1 or 4, characterized in that: In Step 2.2, the optimization models are respectively: For the basic wing airfoil, the optimization model is: Among them is the basic wing optimization objective is the basic wing objective function is the weighting coefficient of the basic wing penalty function is the lift coefficient penalty function of the basic wing is the drag coefficient penalty function of the basic wing is the lift-drag ratio penalty function of the basic wing is the thickness penalty function of the basic wing; represents the constraint condition are respectively the lift-drag ratios of the optimized airfoil of the basic wing under supersonic, transonic and subsonic conditions are respectively the lift-drag ratios of the basic airfoil of the basic wing under supersonic, transonic and subsonic conditions are respectively the lift coefficients of the optimized airfoil of the basic wing under subsonic, transonic and supersonic conditions are respectively the lift coefficients of the basic airfoil of the basic wing under subsonic, transonic and supersonic conditions are respectively the drag coefficients of the optimized airfoil of the basic wing under subsonic, transonic and supersonic conditions are respectively the drag coefficients of the basic airfoil of the basic wing under subsonic, transonic and supersonic conditions is the maximum thickness of the optimized airfoil of the basic wing is the maximum thickness of the basic airfoil of the basic wing; Basic wing objective function is as follows: wherein are respectively the weighted coefficient of the subsonic lift coefficient, the weighted coefficient of the transonic drag coefficient, and the weighted coefficient of the supersonic lift-drag ratio of the optimized airfoil of the basic airfoil after normalization; The penalty function expression of the basic wing lift coefficient is: Among them is the normalized weight of the basic wing lift coefficient in the subsonic state, is the normalized weight of the basic wing lift coefficient in the transonic state, is the normalized weight of the basic wing lift coefficient in the supersonic state; The penalty function expression of the basic wing drag coefficient is: Among them is the normalized weight of the basic wing drag coefficient in the subsonic state, is the normalized weight of the basic wing drag coefficient in the transonic state, is the normalized weight of the basic wing drag coefficient in the supersonic state; The penalty function expression of the basic wing lift-drag ratio is as follows: wherein is the normalized weight of the basic wing lift-drag ratio in the subsonic state, is the normalized weight of the basic wing lift-drag ratio in the transonic state, is the normalized weight of the basic wing lift-drag ratio in the supersonic state; The penalty function expression of the basic wing thickness is as follows: Among them is the normalized weight of the basic wing thickness For the strake wing, the optimization model is: Among them is the optimization objective of the strake wing, is the objective function of the strake wing, is the weighting coefficient of the penalty function of the strake wing, is the penalty function of the lift coefficient of the strake wing, is the penalty function of the drag coefficient of the strake wing, is the penalty function of the lift-to-drag ratio of the strake wing, is the penalty function of the thickness of the strake wing; represents the constraint condition, are respectively the lift-to-drag ratios of the optimized airfoil of the strake wing under supersonic, transonic and subsonic conditions in sequence, are respectively the lift-to-drag ratios of the basic airfoil of the strake wing under supersonic, transonic and subsonic conditions in sequence, are respectively the lift coefficients of the optimized airfoil of the strake wing under subsonic, transonic and supersonic conditions in sequence, are respectively the lift coefficients of the basic airfoil of the strake wing under subsonic, transonic and supersonic conditions in sequence, are respectively the drag coefficients of the optimized airfoil of the strake wing under subsonic, transonic and supersonic conditions in sequence, are respectively the drag coefficients of the basic airfoil of the strake wing under subsonic, transonic and supersonic conditions in sequence, is the maximum thickness of the optimized airfoil of the strake wing, is the maximum thickness of the basic airfoil of the strake wing; The Basefit2 expression is: wherein are respectively the subsonic lift-drag ratio weighting coefficient, the transonic lift-drag ratio weighting coefficient, and the supersonic lift-drag ratio weighting coefficient of the optimized airfoil of the normalized edge strip wing in sequence; The penalty function expression of the strake wing lift coefficient is: wherein is the normalized weight of the leading-edge wing lift coefficient in the subsonic state, is the normalized weight of the leading-edge wing lift coefficient in the transonic state, is the normalized weight of the leading-edge wing lift coefficient in the supersonic state; The penalty function expression of the strake wing drag coefficient is: where is the normalized weight of the strake wing drag coefficient in the subsonic state, is the normalized weight of the strake wing drag coefficient in the transonic state, is the normalized weight of the strake wing drag coefficient in the supersonic state; The penalty function expression of the strake wing lift-drag ratio is: where is the normalized weight of the lift-to-drag ratio of the strake wing in the subsonic state, is the normalized weight of the lift-to-drag ratio of the strake wing in the transonic state, is the normalized weight of the lift-to-drag ratio of the strake wing in the supersonic state; The penalty function of the strake wing thickness is: Among them is the normalized weight of the strake wing thickness 7. A layout design method for a blended wing-body wide-speed-range aircraft based on considering the afterbody effect as claimed in claim 6, characterized in that: In step 2.2, the NACA64A204 airfoil is also used as the basic airfoil of the backup airfoil, and the optimization of the backup airfoil aiming to improve the lift-drag ratio is carried out under three working conditions: Ma = 4, altitude of 30 km, angle of attack of 5°, Ma = 5, altitude of 30 km, angle of attack of 5°, and Ma = 6, altitude of 30 km, angle of attack of 5°. The established optimization model is: Among them is the optimization target of the backup airfoil is the objective function of the backup airfoil is the penalty function weighting coefficient of the backup airfoil is the lift coefficient penalty function of the backup airfoil is the drag coefficient penalty function of the backup airfoil is the lift-to-drag ratio penalty function of the basic airfoil is the thickness penalty function of the basic airfoil represents the constraint condition are respectively the lift-to-drag ratios of the backup optimized airfoil under three working conditions of Ma = 4, 5, and 6 are respectively the lift-to-drag ratios of the basic airfoil of the backup airfoil under three working conditions of Ma = 4, 5, and 6 are respectively the lift coefficients of the backup optimized airfoil under three working conditions of Ma = 4, 5, and 6 are respectively the lift coefficients of the basic airfoil of the backup airfoil under three working conditions of Ma = 4, 5, and 6 are respectively the drag coefficients of the backup optimized airfoil under three working conditions of Ma = 4, 5, and 6 are respectively the drag coefficients of the basic airfoil of the backup airfoil under three working conditions of Ma = 4, 5, and 6 is the maximum thickness of the backup optimized airfoil is the maximum thickness of the basic airfoil of the backup airfoil Among them, the Basefit3 expression is: Among them are respectively the lift-drag ratio weighting coefficients of the normalized standby optimized airfoil under the Ma = 4 condition, the lift-drag ratio weighting coefficient under the Ma = 5 condition, and the lift-drag ratio weighting coefficient under the Ma = 6 condition; The penalty function expression of the backup airfoil lift coefficient is: where is the normalized weight of the spare airfoil lift coefficient under the condition of Ma = 4, is the normalized weight of the spare airfoil lift coefficient under the condition of Ma = 5, is the normalized weight of the spare airfoil lift coefficient under the condition of Ma = 6; The penalty function expression of the backup airfoil drag coefficient is: Among them is the normalized weight of the spare airfoil drag coefficient under the condition of Ma = 4, is the normalized weight of the spare airfoil drag coefficient under the condition of Ma = 5, is the normalized weight of the spare airfoil drag coefficient under the condition of Ma = 6; The penalty function expression of the backup airfoil lift-drag ratio is: where is the normalized weight of the lift-drag ratio of the backup airfoil under the condition of Ma = 4, is the normalized weight of the lift-drag ratio of the backup airfoil under the condition of Ma = 5, is the normalized weight of the lift-drag ratio of the backup airfoil under the condition of Ma = 6; The penalty function of the backup airfoil thickness is: Among them is the normalized weight of the backup airfoil thickness 8. A layout design method for a wing-body fusion wide-speed range aircraft based on considering afterbody effects as claimed in claim 7, characterized in that: In step 5, the three-dimensional layout parameters include the first sweep angle of the strake wing , the span of the strake wing , the length obtained by subtracting the span of the strake wing from the span of the basic wing , which is mathematically represented as: Wherein is the second sweep angle of the basic wing determined according to the transonic constraint, is the dihedral angle of the lower surface of the waverider, is the central angle, and R is the radius of the conical shock wave extending to the trailing edge position of the afterbody of the waverider, is the offset distance.

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