Wing-body integrated wide-speed-range aircraft layout design method considering afterbody effect

Abstract

The present invention proposes a layout design method for a wing-body blended wide-speed range aircraft based on consideration of the rearbody effect. The method first generates a tapered waverider forebody for supersonic wide-speed range aircraft. Then, based on the basic airfoil, the basic wing and strake airfoils of the wide-speed range aircraft are aerodynamically optimized to obtain a multi-section wide-speed range waverider airfoil. The length of the waverider rearbody is then calculated and the wide-speed range aircraft symmetry plane is designed. Finally, based on the osculating cone method and transonic strake constraints, the waverider forebody parameters and the waverider rearbody length are used to obtain three-dimensional layout parameters. The optimized basic wing airfoil, strake optimized airfoil, and symmetry plane are assembled to generate the wide-speed range aircraft layout. The present invention couples the tapered waverider, wide-speed range airfoil optimization, and osculating cone double-swept waverider generation method, taking into account the rearbody aerodynamic design, thereby improving the aerodynamic performance of the wide-speed range aircraft at subsonic, transonic, and supersonic speeds.

Description

Technical Field

[0001] The present invention relates to the field of aircraft design, and in particular to a wide-speed range aircraft layout design method for wing-body fusion based on consideration of rear body effect, which is used in the design of a new generation of supersonic passenger aircraft. Background Art

[0002] A wide-speed-range aircraft is a reusable aircraft that can take off and land horizontally and has supersonic cruise capability. Its flight speeds cover subsonic, transsonic, and supersonic speeds. The cruising speed is increased from Ma=3 to Ma=5, and the maximum flight speed even reaches Ma=8.

[0003] Currently, wide-range aircraft designs generally incorporate waveriders. While waveriders can improve the high-speed performance of wide-range aircraft, they struggle to balance subsonic and transonic lift-drag characteristics with stability. To address this issue, one approach, as seen in aircraft such as the X-15, Sanger, and Talon-A, involves horizontal takeoff and landing from subsonic or supersonic platforms, air-launching a supersonic aircraft, thereby avoiding low-speed conditions for the second-stage supersonic aircraft. This approach is clearly unsuitable for supersonic passenger aircraft design. Another approach is to enhance the subsonic, transsonic, and / or supersonic performance of the waverider, enabling single-stage wide-range flight. Methods for improving the low-speed performance of the waverider include integrating it with the wing to form a waverider-wing configuration, and utilizing vortex lift to enhance lift. Waverider-wing configurations are categorized into two types: a double-swept integral waverider configuration and a spliced, non-full-aircraft waverider configuration. The former adopts an improved waverider generation method to construct the waveriding effect of the whole aircraft, but the airfoil of its wing is generally a simple wedge shape, the upper surface of the fuselage is generally flat, and there is still room for improvement in the low-speed flight performance; the latter starts from the wide-speed range wing, fixes the shape of the waverider forebody, splices the wide-speed range fuselage and wing behind the waverider forebody, and designs the winglet / basic wing plane layout through experience or proxy model optimization to generate a wide-speed range aircraft that is not a 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 disadvantages under higher supersonic speed conditions.

[0004] It is worth noting that the existing wide-speed range aircraft design has significant rear body design defects: the current technical route generally focuses on the optimization of the forebody waveriding configuration, retaining only the vertical end surface or simply docking the power system in the rear body area, and has not yet effectively developed the transonic efficiency enhancement potential of the rear body aerodynamic surface. Summary of the Invention

[0005] In response to the problems existing in the prior art, the present invention provides a layout design method for a wing-body fusion wide-speed range aircraft based on taking into account the rearbody effect. The method couples the cone-guided waverider, wide-speed range airfoil optimization and osculating cone double-swept waverider generation method, and considers the aerodynamic design of the rearbody, thereby improving the aerodynamic performance of the wide-speed range aircraft at subsonic, transsupersonic and supersonic speeds.

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

[0007] The method for designing a layout of a wing-body fusion wide-speed range aircraft based on taking into account the rear body effect comprises the following steps:

[0008] Step 1: Generate a cone-guided waverider precursor for a wide-speed range aircraft for supersonic flight;

[0009] Step 2: Based on the basic airfoil, aerodynamic optimization is performed on the basic airfoil and the strake airfoil of the wide-speed range aircraft to obtain a multi-section wide-speed range waverider airfoil:

[0010] Step 2.1: Determine the basic airfoil of the basic wing and the basic airfoil of the winglet; perform parameterization on the basic airfoil to obtain the optimization parameters, and determine the design space of the optimization parameters;

[0011] Step 2.2: Set the design conditions and optimization models for the optimized basic wing and winglet airfoils; 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 in 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 winglet;

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

[0014] Step 3.1: First, according to the flow field shock wave angle β, the incoming flow Mach number Ma, the specific heat ratio γ and the object 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 post-wave Mach number The angle between the left characteristic line family and the center line ;

[0017] Step 3.2: According to the formula

[0018]

[0019] Calculate the length of the waverider body ;in is the proportion of the waverider precursor in the total length of the machine, is the waverider precursor length;

[0020] Step 4: Design the 2D shape of the wide-speed range aircraft symmetry surface:

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

[0022]

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

[0024] Step 5: Based on the osculating cone method and transonic winglet constraints, the three-dimensional layout parameters are obtained using the waverider front body parameters and the waverider rear body length. The basic wing optimized airfoil, the winglet optimized airfoil and the symmetry surface are assembled to generate a wide-speed range aircraft layout.

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

[0026] Furthermore, the specific process of generating the cone-guided waverider precursor for a wide-speed range supersonic flight vehicle in step 1 is as follows:

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

[0028] Step 1.2: Specify the bottom shape of the cone and find a surface that passes through the bottom shape and is parallel to the incoming flow direction. The intersection of this surface and the shock wave surface is the leading edge of the waverider. The surface formed by the leading edge and the bottom shape is the upper surface of the waverider.

[0029] Step 1.3: Starting from the waverider leading edge line, according to the Taylor-Maccoll equation, the lower surface of the waverider is obtained by streamline tracing, and a cone-guided waverider front is generated.

[0030] Furthermore, in step 2.1, for the basic wing, the NACA64A204 airfoil is selected as the basic airfoil; for the wingtip, an inverted wedge with a flat bottom, a maximum thickness at 45% of the chord length, a maximum thickness of 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 operating conditions are selected as: Ma=0.4, 2° angle of attack, 10km altitude; Ma=1, 2° angle of attack, 10km altitude; Ma=5, 5° angle of attack, 30km altitude; corresponding to subsonic, transonic and supersonic operating conditions respectively.

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

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

[0034]

[0035] in is the basic wing optimization objective, is the basic wing objective function, is the weighted coefficient of the basic wing penalty function, is the penalty function of the basic wing lift coefficient, is the penalty function of the basic wing drag coefficient, is the basic wing lift-to-drag ratio penalty function, is the basic wing thickness penalty function; represents the constraints, They are the lift-to-drag ratios of the optimized airfoil of the basic wing at supersonic, transonic and subsonic conditions respectively. They are the lift-to-drag ratios of the basic airfoil at supersonic, transonic and subsonic conditions, respectively. are the lift coefficients of the optimized airfoil of the basic wing at subsonic, transonic and supersonic conditions respectively, are the lift coefficients of the basic airfoil at subsonic, transonic and supersonic conditions respectively, They are the drag coefficients of the optimized airfoil of the basic wing at subsonic, transonic and supersonic conditions respectively. are the drag coefficients of the basic airfoil at subsonic, transonic and supersonic conditions respectively. The maximum thickness of the optimized airfoil for the basic wing, is the maximum thickness of the basic airfoil;

[0036] Basic wing objective function for:

[0037]

[0038] in They are respectively the subsonic lift coefficient weighting coefficient, transonic drag coefficient weighting coefficient and supersonic lift-to-drag ratio weighting coefficient of the normalized basic wing optimized airfoil;

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

[0040]

[0041] in 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 supersonic state;

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

[0043]

[0044] in 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 supersonic state;

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

[0046]

[0047] in 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 supersonic state;

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

[0049]

[0050] in is the normalized weight of basic wing thickness;

[0051] For the wingtip, the optimization model is:

[0052]

[0053] in is the winglet optimization target, is the winglet objective function, is the weighted coefficient of the wing penalty function, is the penalty function of the winglet lift coefficient, is the penalty function of the winglet drag coefficient, is the winglet lift-to-drag ratio penalty function, is the winglet thickness penalty function; represents the constraints, The lift-to-drag ratios of the optimized winglet airfoil at supersonic, transonic and subsonic conditions are respectively: They are the lift-to-drag ratios of the basic winglet airfoil at supersonic, transonic and subsonic conditions respectively. The lift coefficients of the optimized winglet airfoil at subsonic, transonic and supersonic conditions are respectively: are the lift coefficients of the basic winglet airfoil at subsonic, transonic and supersonic conditions respectively. The drag coefficients of the optimized winglet airfoil at subsonic, transonic and supersonic conditions are respectively, are the drag coefficients of the basic winglet airfoil at subsonic, transonic and supersonic conditions respectively. The maximum thickness of the airfoil optimized for the winglets, is the maximum thickness of the basic airfoil of the wingtip;

[0054] The Basefit2 expression is:

[0055]

[0056] in They are the subsonic lift-to-drag ratio weighting coefficient, transonic lift-to-drag ratio weighting coefficient and supersonic lift-to-drag ratio weighting coefficient of the normalized winglet optimized airfoil respectively;

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

[0058]

[0059] in is the normalized weight of the winglet lift coefficient under subsonic conditions, is the normalized weight of the winglet lift coefficient under transonic conditions, is the normalized weight of the winglet lift coefficient under supersonic conditions;

[0060] The penalty function expression of the drag coefficient of the wing strip is:

[0061]

[0062] in is the normalized weight of the drag coefficient of the wingtip in the subsonic state, is the normalized weight of the wingtip drag coefficient under transonic conditions, is the normalized weight of the drag coefficient of the wingtip in supersonic state;

[0063] The lift-to-drag ratio penalty function expression of the winglet is:

[0064]

[0065] in is the normalized weight of the lift-to-drag ratio of the wingtip in the subsonic state, is the normalized weight of the winglet lift-to-drag ratio in the transonic state, is the normalized weight of the winglet lift-to-drag ratio under supersonic conditions;

[0066] The penalty function for winglet thickness is:

[0067]

[0068] in Normalized weight for winglet thickness.

[0069] Furthermore, in step 2.2, the NACA64A204 airfoil is used as the base airfoil of the backup airfoil. The backup airfoil is optimized under three operating conditions: Ma = 4, 30 km altitude, 5° angle of attack; Ma = 5, 30 km altitude, 5° angle of attack; and Ma = 6, 30 km altitude, 5° angle of attack, with the goal of improving the lift-to-drag ratio. The established optimization model is:

[0070]

[0071] in is the optimization target of the spare airfoil, is the standby airfoil objective function, is the weighted coefficient of the penalty function for the spare airfoil, is the penalty function of the lift coefficient of the spare airfoil, is the penalty function for the drag coefficient of the alternative airfoil, is the basic wing lift-to-drag ratio penalty function, is the basic wing thickness penalty function; represents the constraints, The lift-to-drag ratios of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6 are respectively: These are the lift-to-drag ratios of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6, respectively. They are respectively the lift coefficients of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6. They are respectively the lift coefficients of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6. The drag coefficients of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6 are respectively, These are the drag coefficients of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6, respectively. The maximum thickness of the airfoil optimized for standby, The maximum thickness of the base airfoil for the reserve airfoil;

[0072] The Basefit3 expression is:

[0073]

[0074] in They are respectively the normalized lift-to-drag ratio weighted coefficient of the spare optimized airfoil under the working condition of Ma=4, the lift-to-drag ratio weighted coefficient under the working condition of Ma=5, and the lift-to-drag ratio weighted coefficient under the working condition of Ma=6;

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

[0076]

[0077] in is the normalized weight of the lift coefficient of the spare airfoil under the condition of Ma=4, is the normalized weight of the lift coefficient of the spare airfoil under the condition of Ma=5, is the normalized weight of the lift coefficient of the spare airfoil under the condition of Ma=6;

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

[0079]

[0080] in is the normalized weight of the drag coefficient of the spare airfoil under the condition of Ma=4, is the normalized weight of the drag coefficient of the spare airfoil under the condition of Ma=5, is the normalized weight of the drag coefficient of the spare airfoil under the Ma=6 condition;

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

[0082]

[0083] in is the normalized weight of the lift-to-drag ratio of the spare airfoil under the condition of Ma=4, is the normalized weight of the lift-to-drag ratio of the spare airfoil under the condition of Ma=5, is the normalized weight of the lift-to-drag ratio of the spare airfoil under the Ma=6 condition;

[0084] The penalty function for the spare airfoil thickness is:

[0085]

[0086] in is the normalized weight of the spare airfoil thickness.

[0087] Furthermore, in step 5, the three-dimensional layout parameters include the first sweep angle of the wing slat , wingspan of winglets , the length of the basic wing minus the wing span of the wing strip , mathematically represented as:

[0088]

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

[0090] Beneficial effects:

[0091] Compared with existing technologies, the present invention has the following beneficial technical effects: By combining a tapered waverider with an osculating tapered double-swept waverider design method and transonic winglet / basic wing design constraints, this method achieves an integrated waverider design for a wide-speed range aircraft. This suppresses pressure leakage from the fuselage's lower surface at supersonic speeds. Simultaneously, winglet vortices and leading-edge vortices are generated at low speeds, improving the lift coefficient of the wide-speed range aircraft. The intersection of the shock wave and the left-hand characteristic line determines the proportion of the waverider's leading body in the overall aircraft length and the length of the waverider's rear body, achieving supersonic drag reduction and improved lift-to-drag ratio. Symmetrical surface design suppresses airflow separation at the trailing edge, generating a large low-pressure region on the upper surface of the waverider's rear body at supersonic speeds. Using a surrogate model, the multi-airfoil cross-section is optimized, significantly improving the airfoil's aerodynamic performance at subsonic, transsonic, and supersonic speeds, meeting the assembly requirements of wide-speed range aircraft. The aforementioned multiple design techniques are integrated into the wide-speed range aircraft design method, ultimately improving the aerodynamic performance of the wide-speed range aircraft at subsonic, transsonic, and supersonic speeds.

[0092] Additional aspects and advantages of the present invention will be set forth in part in the description which follows and, in part, will be obvious from the description which follows, or may be learned by 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 apparent and readily understood from the following description of the embodiments with reference to the accompanying drawings, in which:

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

[0095] Figure 2 Schematic diagram of parameters involved in the generation of cone-guided waverider;

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

[0097] Figure 4 : Comparison of the winglet airfoil before and after optimization;

[0098] Figure 5 :Calculation of wide-speed range aircraft layout geometric parameters by cone-guided waverider and osculating cone method;

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

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

[0101] Figure 8 : Comparison of lift-to-drag ratio between the wide-speed range flat trailing edge configuration designed in this example and the control layout at different supersonic attack angles. DETAILED DESCRIPTION

[0102] The following describes in detail embodiments of the present invention. The embodiments are exemplary and intended to explain the present invention, but are not to be construed as limiting the present invention.

[0103] See Figure 1 The layout design method for a wing-body fusion wide-speed range aircraft based on taking into account the rear body effect proposed in this embodiment includes the following steps:

[0104] Step 1: Use the Conical Flowfield Method to generate the waverider tip for a wide-speed range supersonic flight vehicle, i.e., to generate the conical waverider precursor for the wide-speed range vehicle. The Conical Flowfield Method is a classic waverider design method. Its core concept is based on the conical shock wave theory, and it generates the waverider shape by tracing the streamlines behind the shock wave. The specific process is as follows:

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

[0106] Step 1.2: Specify the bottom shape of the cone and find a surface that passes through the bottom shape of the cone and is parallel to the incoming flow direction. The intersection of the surface and the shock wave surface is the leading edge line of the waverider. The surface formed by the leading edge line and the bottom shape line is the upper surface of the waverider. Figure 2 As shown, the cone bottom shape line of the upper surface of the waverider specified in this embodiment is a quartic equation, which is expressed as:

[0107]

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

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

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

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

[0112] In the conical flow field length When the above three geometric parameters are determined, the coordinates of the vertex of the quartic function and the tangent equations of the two endpoints can be determined, thereby calculating the coefficients of the quartic equation. For the specific calculation method, please 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 waverider leading edge line, according to the Taylor-Maccoll equation, the lower surface of the waverider is obtained by streamline tracing, and the cone-guided waverider front is generated. For the specific calculation method, please refer to the paper (Wang Xiaoyan. Research on the Design Method of Waverider Based on Three-Dimensional Leading Edge [D]. Changsha: National University of Defense Technology, 2018.). The length of the cone-guided waverider front obtained in this example is .

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

[0115] The wing of a wide-speed range aircraft is a wing-body blended wing with slats, which consists of two parts: the slats, which are used to trim the transition between the wing and the fuselage and have an arched leading edge for generating vortices at transonic / subsonic speeds; and the wing part with a small leading edge sweep angle and relatively straight, called the basic wing. For its definition, see (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 ensure that both the winglets and the basic wing have good aerodynamic performance in a wide speed range, different optimized airfoils need to be assembled on the basic wing and the winglets to meet the aerodynamic performance requirements of different wing sections. The specific process is as follows:

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

[0118] The basic airfoil is parameterized. Here, the conventional 5th-order CST parameterization method in this field is used to parameterize the upper and lower surfaces of the basic airfoil, with a total of 12 design variables (i.e., optimization parameters). Its mathematical expression is:

[0119]

[0120] in is the abscissa of the airfoil, is the vertical coordinate of the upper surface of the airfoil, is the ordinate of the lower surface of the airfoil, is a class function, is a type function, Subscript and represent the upper and lower surfaces respectively, and They represent the upper surface trailing edge coordinates and the lower surface trailing edge coordinates respectively. N1 and N2 are constants, which are 0.5 and 1 respectively. N is the order, which is 5 in this embodiment. and is the undetermined coefficient, i.e., the optimization parameter. For the basic airfoil, the design space for the optimization parameters of the upper and lower surfaces is set to [-0.03, 0.03]. For the slat airfoil, the design space for the optimization parameters of the upper surface is set to [-0.03, 0.03], while no design space is set for the lower surface, thereby maintaining the straight line characteristics of the lower surface of the slat inverted wedge airfoil.

[0121] Step 2.2: Set the design conditions and optimization model for the optimized basic wing airfoil and winglet airfoil.

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

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

[0124]

[0125] in is the basic wing optimization objective, is the basic wing objective function, is the weighted coefficient of the basic wing penalty function, is the penalty function of the basic wing lift coefficient, is the penalty function of the basic wing drag coefficient, is the basic wing lift-to-drag ratio penalty function, is the basic wing thickness penalty function; represents the constraints, They are the lift-to-drag ratios of the optimized airfoil of the basic wing at supersonic, transonic and subsonic conditions respectively. They are the lift-to-drag ratios of the basic airfoil at supersonic, transonic and subsonic conditions, respectively. are the lift coefficients of the optimized airfoil of the basic wing at subsonic, transonic and supersonic conditions respectively, are the lift coefficients of the basic airfoil at subsonic, transonic and supersonic conditions respectively, They are the drag coefficients of the optimized airfoil of the basic wing at subsonic, transonic and supersonic conditions respectively. are the drag coefficients of the basic airfoil at subsonic, transonic and supersonic conditions respectively. The maximum thickness of the optimized airfoil for the basic wing, is the maximum thickness of the basic wing airfoil.

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

[0127]

[0128] in They are the subsonic lift coefficient weighting coefficient, transonic drag coefficient weighting coefficient and supersonic lift-to-drag ratio weighting coefficient of the normalized basic wing optimized airfoil, respectively.

[0129]

[0130] Normalized to get are all set weighting coefficients.

[0131] The penalty function is obtained by weighting the evaluation of the aerodynamic parameters under the three working conditions. The evaluation process of the aerodynamic parameters will determine whether the aerodynamic parameters of the optimization result are lower than the initial value. If the aerodynamic parameters are higher than the initial value, the penalty function is set to 0. If they are lower than the initial value, 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] in 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 supersonic state.

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

[0136]

[0137] in 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 supersonic state.

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

[0139]

[0140] in 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 supersonic state.

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

[0142]

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

[0144] For the wingtip, the optimization model is:

[0145]

[0146] in is the winglet optimization target, is the winglet objective function, is the weighted coefficient of the wing penalty function, is the penalty function of the winglet lift coefficient, is the penalty function of the winglet drag coefficient, is the winglet lift-to-drag ratio penalty function, is the winglet thickness penalty function; represents the constraints, The lift-to-drag ratios of the optimized winglet airfoil at supersonic, transonic and subsonic conditions are respectively: They are the lift-to-drag ratios of the basic winglet airfoil at supersonic, transonic and subsonic conditions respectively. The lift coefficients of the optimized winglet airfoil at subsonic, transonic and supersonic conditions are respectively: are the lift coefficients of the basic winglet airfoil at subsonic, transonic and supersonic conditions respectively. The drag coefficients of the optimized winglet airfoil at subsonic, transonic and supersonic conditions are respectively, are the drag coefficients of the basic winglet airfoil at subsonic, transonic and supersonic conditions respectively. Optimize the maximum thickness of the airfoil for the wing slats, is the maximum thickness of the basic airfoil of the wingtip.

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

[0148]

[0149] in They are the subsonic lift-to-drag ratio weighted coefficient, transonic lift-to-drag ratio weighted coefficient and supersonic lift-to-drag ratio weighted coefficient of the normalized slat optimized airfoil, respectively.

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

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

[0152]

[0153] in is the normalized weight of the winglet lift coefficient under subsonic conditions, is the normalized weight of the winglet lift coefficient under transonic conditions, is the normalized weight of the winglet lift coefficient under supersonic conditions.

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

[0155]

[0156] in is the normalized weight of the drag coefficient of the wingtip in the subsonic state, is the normalized weight of the wingtip drag coefficient under transonic conditions, is the normalized weight of the wingtip drag coefficient under supersonic conditions.

[0157] The lift-to-drag ratio penalty function expression of the winglet is:

[0158]

[0159] in is the normalized weight of the lift-to-drag ratio of the wingtip in the subsonic state, is the normalized weight of the winglet lift-to-drag ratio in the transonic state, is the normalized weight of the winglet's lift-to-drag ratio under supersonic conditions.

[0160] The penalty function for winglet thickness is:

[0161]

[0162] in Normalized weight for winglet thickness.

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

[0164]

[0165] in is the optimization target of the spare airfoil, is the standby airfoil objective function, is the weighted coefficient of the penalty function for the spare airfoil, is the penalty function of the lift coefficient of the spare airfoil, is the penalty function for the drag coefficient of the spare airfoil, is the basic wing lift-to-drag ratio penalty function, is the basic wing thickness penalty function; represents the constraints, The lift-to-drag ratios of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6 are respectively: These are the lift-to-drag ratios of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6, respectively. They are respectively the lift coefficients of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6. They are respectively the lift coefficients of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6. The drag coefficients of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6 are respectively, These are the drag coefficients of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6, respectively. The maximum thickness of the airfoil optimized for standby, The maximum thickness of the base airfoil for the spare airfoil.

[0166] The Basefit3 expression is:

[0167]

[0168] in They are respectively the lift-to-drag ratio weighted coefficient of the normalized spare optimized airfoil under the working condition of Ma=4, the lift-to-drag ratio weighted coefficient under the working condition of Ma=5, and the lift-to-drag ratio weighted coefficient under the working condition of Ma=6.

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

[0170]

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

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

[0173]

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

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

[0176]

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

[0178] The penalty function for the spare airfoil thickness is:

[0179]

[0180] in is the normalized weight of the spare airfoil thickness.

[0181] Step 2.3: For the optimization model established in step 2.2, the conventional Latin hypercube sampling method in this field is used for sampling within the design space. The grid is automatically generated using a Tcl script. The aerodynamic parameters of the samples are calculated using the Reynolds mean stress (RANS) method as the data input of the proxy model. The optimization model is solved using the conventional particle swarm algorithm in this field to achieve multi-objective optimization, and finally the optimized airfoils of the basic wing, the winglet optimized airfoil, and the spare airfoil are obtained.

[0182] The surrogate model used is the Kriging surrogate model, a common surrogate model in the field. The Kriging surrogate model is an interpolation model whose interpolation result is a linear weighting of the sample function response values, with the weighting coefficient calculated using a Gaussian exponential model. Since the optimization objective is an optimization function related to the lift coefficient, drag coefficient, and lift-to-drag ratio, this process requires establishing surrogate models for each of the 12 design variables, namely the lift coefficient, drag coefficient, and lift-to-drag ratio.

[0183] In the particle swarm optimization algorithm, each particle first obtains an initial velocity and position. An optimization function is used to evaluate the results of each particle, finding the individual best position pbest and the global best position gbest. An inertia weight is added to the initial velocity, and learning factors are added to the individual and global bests, respectively. After initialization, particle velocity and position updates begin. This process represents the particle's movement based on the inertia weight and two learning rates, respectively. The inertia weight depends on the particle's initial velocity, and the two learning rates determine the weight of movement toward the individual and global best positions in the overall velocity. The movement speed toward the individual and global best positions is a random value within the range of half the design space. The movement direction is the vector difference between the individual and global best positions and the current position. After the movement speed is determined, the particle position is updated. If the particle's position is outside the design space during the position update, the coordinate outside the design space is set to the maximum allowed value for the movement direction. After the particle position is updated, the new position serves as the initial value for the next round of particle swarm optimization. The next round of optimization continues until the maximum number of iterations is reached or the particle error falls below a threshold. The optimization result obtained at this point is the optimization result obtained by the surrogate model optimization for that round.

[0184] The final basic wing optimization results are as follows Figure 3 As shown in the figure, the optimization results of the winglets are as follows: Figure 4 As shown in Figure 2, the aerodynamic parameters of the optimized airfoil 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 feature. As shown in 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 is increased by 2.2×10 -4 The improvement in aerodynamic parameters comes from the expansion of the low-pressure area on the subsonic / transonic upper surface, and the optimization of supersonic performance comes from the double S-shaped features on the lower surface, which creates two stronger high-pressure areas and reduces the pressure coefficient on the upper surface.

[0187] After the optimization of the winglets, the maximum thickness of the airfoil is moved back to 52%, and the maximum thickness is reduced to 2.6% of the chord length. The lift-to-drag ratio in all working conditions is improved. This improvement is due to the more uniform low-pressure distribution on the subsonic upper surface and the presence of a more obvious suction peak. The leading edge pressure coefficient of the transonic upper surface is significantly reduced, and a strong compression wave is maintained on the supersonic lower surface.

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

[0189]

[0190] Step 3: Calculate the length of the waverider rear body, specifically including the following steps:

[0191] Step 3.1: The waverider rear body is the fuselage area behind the waverider body. Since the disturbance of the flow field under supersonic speed 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-moving characteristic line passing through the point, the aerodynamic characteristics of the rear body are related to the left-moving characteristic line passing through the trailing edge of the waverider front body and the leading edge of the waverider rear body.

[0192] In order to determine the length of the waverider rear body, in this embodiment, the waverider rear Mach number is first calculated based on the flow field shock wave angle β, the incoming flow Mach number Ma, the specific heat ratio γ and the object plane angle θ. The angle between the left characteristic line family and the center line , mathematically expressed as:

[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-moving characteristic line start from the trailing edge of the waverider forebody, find the intersection point of the left-moving characteristic line and the shock wave, and calculate the proportion of the waverider forebody in the total length of the aircraft. , when the length of the waverider precursor is known In the case of , the specific data is expressed as:

[0195]

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

[0197] Step 4: Design the two-dimensional shape of the symmetry surface of the wide-speed range aircraft; in the present invention, when designing the two-dimensional shape of the symmetry surface, the symmetry surface of the waverider front body is kept unchanged, and the original shape of the symmetry surface of the waverider rear body is set to be all straight lines on the upper and lower sides. Then, based on the original shape, one side is first kept as a straight line and the other side is adjusted; the aerodynamic parameters of the adjusted symmetry surface under subsonic, transonic and supersonic conditions are calculated.

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

[0199]

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

[0201] The aerodynamic parameters of each symmetry plane are evaluated using a weighted function to evaluate its aerodynamic performance. The mathematical expression of the weighted function is:

[0202]

[0203] in represents the weighting function of the symmetry plane, 、 and They represent the subsonic lift coefficient, transonic drag coefficient and supersonic lift-to-drag ratio of the symmetry plane respectively; 、 and These represent the weighted coefficients for the subsonic lift coefficient, transonic drag coefficient, and supersonic lift-to-drag ratio of the symmetric surface, respectively. The weighted coefficients are normalized using the subsonic lift coefficient, transonic drag coefficient, and supersonic lift-to-drag ratio of the symmetric surface's original shape. The smaller the calculated weighting function, the better the aerodynamic performance of the symmetric surface.

[0204] The weighted calculation results in this embodiment show that the weighted aerodynamic efficiency of the symmetry surface is the best when the lower side is a straight line and the upper side adopts the upper side curve of the 20° contraction shape, and this symmetry surface also has the largest area among the non-original shapes, so the upper side of the symmetry surface is determined to be the upper side curve of the 20° contraction shape.

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

[0206]

[0207] A weighted function was applied to the aerodynamic parameters of each symmetry plane to evaluate its performance. The weighted analysis revealed that using an airfoil on the lower surface was less effective than maintaining a flat surface. Therefore, in this example, a straight line was used on the lower fuselage surface, resulting in the final two-dimensional shape of the symmetry plane. The improvement in the aerodynamic parameters of the symmetry plane is attributed to suppressing airflow separation at the trailing edge, which expands the leading edge low-pressure region at subsonic speeds. At transonic and supersonic speeds, multiple weak expansion waves appear on the upper surface, creating a large trailing edge low-pressure region.

[0208] Step 5: Based on the osculating cone method and transonic winglet constraints, the three-dimensional layout parameters are obtained using the waverider front body parameters and the waverider rear body length. The basic wing optimized airfoil, the winglet optimized airfoil and the symmetry surface are assembled to generate a wide-speed range aircraft layout.

[0209] In this embodiment, Figure 5 As shown, it is assumed that the conical shock wave flow field extends backward, and the extension length is equal to , the cone shock radius R extending to the trailing edge of the waverider is obtained by using the flow field shock wave angle β and the trigonometric function relationship. Substitute the parameters of the cone-guided waverider front into the osculating cone method: k is equal to the offset distance , the spanwise width ψ is equal to the central angle , while preserving the reverse angle on the lower surface of the waverider in the osculating cone method For auxiliary calculation, the actual basic wing dihedral angle is still kept at 0°. The first sweep angle of the wingtip is finally derived , wingspan of winglets , the length of the basic wing minus the wing span of the wing strip The mathematical representation of this process is:

[0210]

[0211] in 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 wing slat , wingspan of winglets The chord length of the leading edge strip is 5592.384mm, the total wingspan is 6935.446mm, the chord length of the basic wing wingtip is 1917.302mm, the root chord length of the basic wing is 7477.479mm, and the root-tip ratio is 3.9.

[0213] Finally, aerodynamic calculations were performed on the obtained wide-speed range aircraft layout to obtain aerodynamic parameters, and then compared with a certain control layout:

[0214]

[0215] The results show that the flat trailing edge configuration proposed in this embodiment is superior to the control layout in terms of target aerodynamic parameters. Compared with the control layout, the flat trailing edge configuration sacrifices the proportion of the wave-riding area of the fuselage lower surface to the total length of the aircraft, resulting in a lower lift coefficient at supersonic speed than the control configuration, but it achieves an improvement in lift-to-drag ratio by significantly reducing the drag coefficient. Figures 6 to 8 Under design conditions, the flat trailing edge configuration of this embodiment outperforms the control layout in subsonic lift coefficient and supersonic lift-to-drag ratio at varying angles of attack. At low angles of attack, the transonic drag coefficient is superior to the control layout. At typical flight altitudes and angles of attack, this embodiment exhibits superior performance across a wide speed range.

[0216] The improved aerodynamic performance of this example under supersonic design conditions stems from the effective suppression of pressure leakage at the winglets. Numerical simulation results show that the lift-to-drag ratio of the winglets in the control configuration is 6.51, while the lift-to-drag ratio of the winglets in this example configuration is 10.31, further demonstrating that pressure leakage at the winglets is suppressed. The localized high-pressure areas in the flow field on the upper surface are more likely to be caused by the compression waves of the detached body due to the passivation treatment and the pressure distribution characteristics of the airfoil. The rear body contraction design utilizes the vortex wave effect at supersonic speeds, generating expansion waves and vortices on the upper surface, reducing the pressure coefficient on the upper surface.

[0217] The improvement of aerodynamic performance under the transonic design condition of this example is due to the reduction of transonic drag by the rear body design. The drag of the waverider rear body of this example configuration is reduced to 9.064×10 -3 The drag coefficient of the entire aircraft was reduced from 44.68% to 34.73%.

[0218] The improved subsonic performance of this example is primarily due to the shedding vortices generated by the waverider's leading edge, slats, and base wings, which reduce the pressure coefficient on the vehicle's upper surface, and the improved lift-to-drag characteristics of the waverider's aft body design. The vortex lift enhancement effect is dominated by the waverider's leading edge at low angles of attack, while at high angles of attack, the vortices on the waverider's leading edge, slats, and base wings collectively reduce the pressure coefficient on the vehicle's upper surface. The waverider's aft body design with a flat trailing edge provides a 40.32% increase in lift coefficient and a 37.70% decrease in drag coefficient compared to the control aft body, significantly improving the lift-to-drag ratio.

[0219] Different from previous wide-speed range aircraft design methods, the present invention combines the cone-guided waverider, the osculating cone method, and the transonic winglet / basic wing design method, and combines it with the aerodynamic design of the waverider rear body to design a wide-speed range aircraft layout with a higher lift coefficient at subsonic speeds, reduced drag at transonic speeds, and a high lift-to-drag ratio at supersonic speeds. Compared with traditional wide-speed range aircraft design methods, the wide-speed range aircraft designed by this method has advantages in aerodynamic parameters. There is obvious vortex lift at subsonic speeds, surface pressure leakage at supersonic speeds is suppressed, and the upper surface utilizes the vortex wave effect. In addition, the aerodynamic rear body design is comprehensively considered during the design process, which can provide a reference for subsequent wide-speed range aircraft that are decoupled from the power system.

[0220] Although the embodiments of the present invention have been shown and described above, it will be understood that the above embodiments are illustrative and are not to be construed as limitations on the present invention. A person skilled in the art may change, modify, replace and modify the above embodiments within the scope of the present invention without departing from the principles and purpose of the present invention.

Claims

1. A layout design method for a wide-speed range aircraft with wing-body blending taking into account the rear body effect, characterized by: The following steps are involved: Step 1: Generate a cone-guided waverider precursor for a wide-speed range aircraft for supersonic flight; Step 2: Based on the basic airfoil, aerodynamic optimization is performed on the basic airfoil and the strake 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 winglet; perform parameterization on the basic airfoil to obtain the optimization parameters, and determine the design space of the optimization parameters; Step 2.2: Set the design conditions and optimization models for the optimized basic wing and winglet airfoils; 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 in 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 winglet; Step 3: Calculate the length of the waverider body: Step 3.1: First, according to the flow field shock wave angle β, the incoming flow Mach number Ma, the specific heat ratio γ and the object surface angle θ, according to the formula Calculate the post-wave Mach number The angle between the left characteristic line family and the center line ; Step 3.2: According to the formula Calculate the length of the waverider body ;in is the proportion of the waverider precursor in the total length of the machine, is the waverider precursor length; Step 4: Design the 2D shape of the wide-speed range aircraft symmetry surface: Keep the symmetry plane of the waverider front body unchanged, set the original shape of the symmetry plane of the waverider rear body to be straight lines on the upper and lower sides, and on the basis of the original shape, keep one side straight and adjust the other side; calculate the aerodynamic parameters of the adjusted symmetry plane of the wide-speed range aircraft under subsonic, transonic and supersonic conditions, and adjust the aerodynamic parameters according to the formula Perform weighted calculation to obtain the adjusted aerodynamic performance of the wide-speed range aircraft symmetry plane; represents the weighting function of the symmetry plane, 、 and They represent the subsonic lift coefficient, transonic drag coefficient and supersonic lift-to-drag ratio of the symmetry plane respectively; 、 and They represent the weighted coefficient of the subsonic lift coefficient of the symmetric surface, the weighted coefficient of the transonic drag coefficient of the symmetric surface, and the weighted coefficient of the supersonic lift-to-drag ratio of the symmetric surface respectively; the smaller the calculated result of the weighted function, the better the aerodynamic performance of the symmetric surface; the symmetric surface with the best aerodynamic performance in a wide speed range of the aircraft is obtained; Step 5: Based on the osculating cone method and transonic winglet constraints, the three-dimensional layout parameters are obtained using the waverider front body parameters and the waverider rear body length. The basic wing optimized airfoil, the winglet optimized airfoil and the symmetry surface are assembled to generate a wide-speed range aircraft layout.

2. The layout design method for a wing-body blended wide-speed range aircraft based on taking into account the rear body effect according to claim 1, characterized in that: In step 1, the cone-guided method is used to generate a cone-guided waverider forebody for supersonic flight in a wide speed range.

3. The layout design method for a wing-body blended wide-speed range aircraft based on consideration of the rear body effect according to claim 2, characterized in that: The specific process of step 1 to generate the cone-guided waverider precursor for a wide-speed range aircraft for supersonic flight is as follows: Step 1.1: Determine the cone semi-cone angle and the incoming flow Mach number Ma that generates the conical shock wave flow field, and specify the conical flow field length ; Step 1.2: Specify the bottom shape of the cone and find a surface that passes through the bottom shape and is parallel to the incoming flow direction. The intersection of this surface and the shock wave surface is the leading edge of the waverider. The surface formed by the leading edge and the bottom shape is the upper surface of the waverider. Step 1.3: Starting from the waverider leading edge line, according to the Taylor-Maccoll equation, the lower surface of the waverider is obtained by streamline tracing, and a cone-guided waverider front is generated.

4. The layout design method for a wing-body blended wide-speed range aircraft based on taking into account the rear body effect according to claim 1, characterized in that: In step 2.1, for the basic wing, select the NACA64A204 airfoil as the basic airfoil; for the wingtip, select an inverted wedge with a flat bottom, a maximum thickness at 45% of the chord length, a maximum thickness of 3.5% of the chord length, and a leading edge radius of 10 mm as the basic airfoil.

5. The layout design method for a wing-body blended wide-speed range aircraft based on taking into account the rear body effect according to claim 1, characterized in that: In step 2.2, the design conditions are selected as: Ma=0.4, 2° angle of attack, 10km altitude; Ma=1, 2° angle of attack, 10km altitude; Ma=5, 5° angle of attack, 30km altitude; corresponding to subsonic, transonic and supersonic conditions respectively.

6. The method for designing a wide-speed-range aircraft layout with wing-body fusion taking into account the rear body effect according to claim 1 or 4, characterized in that: In step 2.2, the optimization models are: For the basic airfoil, the optimization model is: in is the basic wing optimization objective, is the basic wing objective function, is the weighted coefficient of the basic wing penalty function, is the penalty function of the basic wing lift coefficient, is the penalty function of the basic wing drag coefficient, is the basic wing lift-to-drag ratio penalty function, is the basic wing thickness penalty function; represents the constraints, They are the lift-to-drag ratios of the optimized airfoil of the basic wing at supersonic, transonic and subsonic conditions respectively. They are the lift-to-drag ratios of the basic airfoil at supersonic, transonic and subsonic conditions, respectively. are the lift coefficients of the optimized airfoil of the basic wing at subsonic, transonic and supersonic conditions respectively, are the lift coefficients of the basic airfoil at subsonic, transonic and supersonic conditions respectively, They are the drag coefficients of the optimized airfoil of the basic wing at subsonic, transonic and supersonic conditions respectively. are the drag coefficients of the basic airfoil at subsonic, transonic and supersonic conditions respectively. The maximum thickness of the optimized airfoil for the basic wing, is the maximum thickness of the basic airfoil of the basic wing; Basic wing objective function for: in They are the subsonic lift coefficient weighting coefficient, transonic drag coefficient weighting coefficient and supersonic lift-to-drag ratio weighting coefficient of the normalized basic wing optimized airfoil respectively; The penalty function expression of the basic wing lift coefficient is: in 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 supersonic state; The penalty function expression of the basic wing drag coefficient is: in 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 supersonic state; The basic wing lift-to-drag ratio penalty function expression is: in 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 supersonic state; The basic wing thickness penalty function expression is: in is the normalized weight of basic wing thickness; For the wingtip, the optimization model is: in is the winglet optimization target, is the winglet objective function, is the weighted coefficient of the wing penalty function, is the penalty function of the winglet lift coefficient, is the penalty function of the winglet drag coefficient, is the winglet lift-to-drag ratio penalty function, is the winglet thickness penalty function; represents the constraints, The lift-to-drag ratios of the optimized winglet airfoil at supersonic, transonic and subsonic conditions are respectively: They are the lift-to-drag ratios of the basic winglet airfoil at supersonic, transonic and subsonic conditions respectively. The lift coefficients of the optimized winglet airfoil at subsonic, transonic and supersonic conditions are respectively: are the lift coefficients of the basic winglet airfoil at subsonic, transonic and supersonic conditions respectively. The drag coefficients of the optimized winglet airfoil at subsonic, transonic and supersonic conditions are respectively, are the drag coefficients of the basic winglet airfoil at subsonic, transonic and supersonic conditions respectively. The maximum thickness of the airfoil optimized for the winglets, is the maximum thickness of the basic airfoil of the wingtip; The Basefit2 expression is: in They are the subsonic lift-to-drag ratio weighting coefficient, transonic lift-to-drag ratio weighting coefficient and supersonic lift-to-drag ratio weighting coefficient of the normalized winglet optimized airfoil respectively; The penalty function expression of the wing lift coefficient is: in is the normalized weight of the winglet lift coefficient under subsonic conditions, is the normalized weight of the winglet lift coefficient under transonic conditions, is the normalized weight of the winglet lift coefficient under supersonic conditions; The penalty function expression of the drag coefficient of the wing strip is: in is the normalized weight of the drag coefficient of the wingtip in the subsonic state, is the normalized weight of the wingtip drag coefficient under transonic conditions, is the normalized weight of the drag coefficient of the wingtip in supersonic state; The lift-to-drag ratio penalty function expression of the winglet is: in is the normalized weight of the lift-to-drag ratio of the wingtip in the subsonic state, is the normalized weight of the winglet lift-to-drag ratio in the transonic state, is the normalized weight of the winglet lift-to-drag ratio under supersonic conditions; The penalty function for winglet thickness is: in Normalized weight for winglet thickness.

7. The layout design method for a wing-body blended wide-speed range aircraft based on consideration of the rear body effect according to claim 6, characterized in that: In step 2.2, the NACA64A204 airfoil is used as the base airfoil for the backup airfoil. Optimization of the backup airfoil is performed under three operating conditions: Ma = 4, 30 km altitude, 5° angle of attack; Ma = 5, 30 km altitude, 5° angle of attack; and Ma = 6, 30 km altitude, 5° angle of attack, with the goal of improving the lift-to-drag ratio. The established optimization model is: in is the optimization target of the spare airfoil, is the standby airfoil objective function, is the weighted coefficient of the penalty function for the spare airfoil, is the penalty function of the lift coefficient of the spare airfoil, is the penalty function for the drag coefficient of the spare airfoil, is the basic wing lift-to-drag ratio penalty function, is the basic wing thickness penalty function; represents the constraints, The lift-to-drag ratios of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6 are respectively: These are the lift-to-drag ratios of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6, respectively. They are respectively the lift coefficients of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6. They are respectively the lift coefficients of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6. The drag coefficients of the spare optimized airfoil under the three working conditions of Ma=4, 5, and 6 are respectively, These are the drag coefficients of the basic airfoil of the spare airfoil under the three working conditions of Ma=4, 5, and 6, respectively. The maximum thickness of the airfoil optimized for standby, The maximum thickness of the base airfoil for the reserve airfoil; The Basefit3 expression is: in They are respectively the normalized lift-to-drag ratio weighted coefficient of the spare optimized airfoil under the working condition of Ma=4, the lift-to-drag ratio weighted coefficient under the working condition of Ma=5, and the lift-to-drag ratio weighted coefficient under the working condition of Ma=6; The penalty function expression of the reserve airfoil lift coefficient is: in is the normalized weight of the lift coefficient of the spare airfoil under the condition of Ma=4, is the normalized weight of the lift coefficient of the spare airfoil under the condition of Ma=5, is the normalized weight of the lift coefficient of the spare airfoil under the condition of Ma=6; The penalty function expression of the reserve airfoil drag coefficient is: in is the normalized weight of the drag coefficient of the spare airfoil under the condition of Ma=4, is the normalized weight of the drag coefficient of the spare airfoil under the condition of Ma=5, is the normalized weight of the drag coefficient of the spare airfoil under the Ma=6 condition; The expression of the penalty function of the alternative airfoil lift-to-drag ratio is: in is the normalized weight of the lift-to-drag ratio of the spare airfoil under the condition of Ma=4, is the normalized weight of the lift-to-drag ratio of the spare airfoil under the condition of Ma=5, is the normalized weight of the lift-to-drag ratio of the spare airfoil under the Ma=6 condition; The penalty function for the spare airfoil thickness is: in is the normalized weight of the spare airfoil thickness.

8. The layout design method for a wing-body blended wide-speed range aircraft based on consideration of the rear body effect according to claim 7, characterized in that: In step 5, the three-dimensional layout parameters include the first sweep angle of the wing slat , wingspan of winglets , the length of the basic wing minus the wing span of the wing strip , mathematically represented as: in is the second sweep angle of the basic wing determined according to the transonic constraint, is the reverse angle on the lower surface of the waverider, is the central angle, R is the radius of the conical shock wave extending to the trailing edge of the waverider body, is the offset distance.

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