Vortex lift waverider aircraft design method based on combined flow field and molded line

Abstract

The invention belongs to waverider aircrafts, and particularly relates to a vortex lift waverider aircraft design method based on a combined flow field and a molded line, which comprises the following steps: firstly, determining the combined flow field design of an aircraft, dividing the whole flow field of the aircraft into two parts for design, the first part is a fuselage middle flow field and a fuselage-to-wing transition section flow field, and the second part is a fuselage-to-wing transition section flow field; the second part is an osculating cone reference flow field of the wing part; then, a combined design molded line is determined, in the flow field design of the fuselage and the transition section, an upper surface bottom molded line is determined and generated in a given horizontal projection molded line mode, and meanwhile, it is ensured that the connection transition problem of a flow line at a flow field interface is properly handled when the design molded line is converted; then streamline tracking and lofting are carried out, wherein under the condition that a horizontal projection molded line of the osculating cone flow field is known, a waverider lower surface is generated through streamline tracking; and finally, performing streamline tracking and streamline lofting by taking the design molded line as an input condition to obtain the vortex lift waverider aircraft.

Description

Technical Field

[0001] This invention relates to waverider aircraft, specifically a design method for vortex lift waverider aircraft based on combined flow fields and profiles. Background Technology

[0002] The waverider adopts a hypersonic aerodynamic layout with a waverider body, blended wing-body, V-tail, integrated design of the forebody and propulsion system, two parallel turbine-based combined cycle engines mounted in the rear, and an integrated design of the rear body and nozzle. In order to achieve hypersonic flight, when developing air-breathing hypersonic vehicles, it is generally desirable for the aircraft body to have a high lift-to-drag ratio. At the same time, it is desirable for the air intake area to provide as much effective air source as possible to the combustion chamber with minimal airflow energy loss, so as to reduce flight drag.

[0003] However, waveriders designed for hypersonic conditions exhibit poor low-speed performance and require coupling with vortex lift to enhance their low-speed flight performance. Waveriders have two main design elements: shock wave flow field and input profile. Currently, most designs couple the vortex lift mechanism by customizing the profile shape, while the shock wave flow field is less considered. How to comprehensively utilize both profile and shock wave flow field design elements to improve the wide-speed-range aerodynamic performance of waveriders has considerable application value.

[0004] To address the aforementioned technical shortcomings, a design solution for vortex-lift waverider aircraft based on combined flow field and profile is proposed. Summary of the Invention

[0005] To address the above problems, the present invention provides the following technical solution:

[0006] A design method for a vortex-lift waverider aircraft based on combined flow field and profile is applicable to the design principle of waveriders with vortex lift characteristics. The steps of the vortex-lift waverider aircraft design method are as follows:

[0007] (1) Determine the combined flow field design of the aircraft. The entire flow field of the aircraft is divided into two parts for design. The first part is the flow field in the middle of the fuselage and the flow field in the transition section from the fuselage to the wing. The second part is the reference flow field of the wing section.

[0008] The airflow field design in the middle of the fuselage is designed as a two-dimensional curved shock wave. The airflow parameters of the planar shock wave flow field are more uniform, which is beneficial to the design of the air intake.

[0009] The transition section flow field design is as follows: the spanwise length of the transition section flow field is 0.84m, and the shock wave angles at the starting and ending points of the shock wave profile are linearly reduced from 14° to 8° along the spanwise direction, thus achieving a transition to a flow field with small shock wave angles.

[0010] The wing section flow field design uses a scissor-cone flow field design with a shock wave angle of 8°.

[0011] (2) Determine the combined design profile. In the flow field design of the fuselage and transition section, determine the bottom profile of the upper surface. The planar shape of the wing is designed as a large swept delta wing with a sweep angle of 62°. It is generated using the given horizontal projection profile method. At the same time, when the design profile is converted, ensure that the connection and transition of the streamline at the flow field interface is properly handled.

[0012] (3) Streamline tracing and lofting: Given the horizontal projection profile of the tangential cone flow field, the wave-riding lower surface is generated by streamline tracing;

[0013] (4) Generate wave-riding configuration: Using the above design profile as input conditions, perform streamline tracing and streamline lofting to obtain the vortex lift wave-riding aircraft.

[0014] Compared with the prior art, the beneficial effects of the present invention are:

[0015] 1. In the design method of vortex lift waverider aircraft based on combined flow field and profile of the present invention, firstly, by designing the flow field of the vortex lift waverider aircraft, the aircraft can effectively utilize vortex lift during subsonic and transonic flight, thereby enhancing the lift characteristics of the aircraft. Especially in high Mach number flight environments, the use of appropriate shock wave configuration and flow field design helps to reduce airflow separation and improve lift-to-drag ratio.

[0016] 2. In the design method of vortex lift waverider aircraft based on combined flow field and profile of the present invention, the present invention makes the airflow parameters more uniform by fine design of the flow field in the middle of the fuselage and the flow field in the transition section, especially by the application of two-dimensional curved surface shock wave and planar shock wave flow field. This design feature is beneficial to the stability of the air intake and reduces the negative impact of non-uniform airflow on the air intake, thereby improving the overall aerodynamic efficiency of the aircraft and the stability of the engine working state.

[0017] 3. In the design method of vortex lift waverider aircraft based on combined flow field and profile of the present invention, by accurately designing the spanwise length and shock wave angle change of the flow field in the transition section, the flow field of the fuselage and wing can achieve a smooth transition between different parts. The shock wave angle is linearly reduced to a small value along the spanwise direction, avoiding flow loss caused by excessively strong shock waves. This flow field transition design greatly reduces the overflow of high-pressure airflow on the lower surface and improves the overall aerodynamic efficiency of the aircraft.

[0018] 4. In the design method of vortex lift waverider aircraft based on combined flow field and profile of the present invention, the present invention adopts a large swept delta wing design and generates the planar shape of the wing part by a given horizontal projection profile, which effectively reduces the complexity in traditional design. By ensuring that the connection and transition of streamlines at the flow field interface is properly handled, the seamless connection of fuselage, transition section and wing part is ensured, thereby reducing airflow instability and further improving the aerodynamic performance of the aircraft.

[0019] 5. In the design method of vortex lift waverider aircraft based on combined flow field and profile of the present invention, the flow field is solved based on the inverse characteristic line method, which is different from the conventional aircraft design which is a reverse design method of flow field first and shape later, thus reducing the complex calculation and experimental process in the traditional design method. Attached Figure Description

[0020] To facilitate understanding by those skilled in the art, the present invention will be further described below with reference to the accompanying drawings;

[0021] Figure 1 This invention relates to a design method for a vortex lift waverider aircraft based on a combined flow field and profile, which includes a shock wave design profile for the flow field in the mid-fuselage of the aircraft.

[0022] Figure 2 This invention relates to the design flow field of the fuselage and the transition section from fuselage to wing in a vortex lift waverider design method based on combined flow field and profile.

[0023] Figure 3 This is a schematic diagram of the shear flow cone field of the wing section in the vortex lift waverider design method based on combined flow field and profile of the present invention.

[0024] Figure 4 This is a schematic diagram of the waverider design profile of the vortex lift waverider design method based on combined flow field and profile of the present invention.

[0025] Figure 5 This is a schematic diagram showing the positional relationship between the leading edge point in the tangent plane and its horizontal projection point in the design method of a vortex lift waverider aircraft based on combined flow field and profile according to the present invention.

[0026] Figure 6 This is a waverider configuration diagram of a vortex lift waverider design method based on combined flow field and profile according to the present invention.

[0027] Figure 7 This is a diagram showing the distribution of characteristic lines in a vortex-driven supersonic flow, which is part of the design method for a vortex-lift waverider aircraft based on a combined flow field and profile according to the present invention.

[0028] Figure 8This is a schematic diagram of the verification flow field geometry for the design method of a vortex lift waverider aircraft based on combined flow field and profile of the present invention.

[0029] Figure 9 The flow field characteristic line calculation results and pressure and Mach number cloud maps calculated by CFD software are presented in Case 1 of the present invention, which is a design method for a vortex lift waverider aircraft based on a combined flow field and profile.

[0030] Figure 10 The results of Case 2 flow field characteristic line calculation and pressure and Mach number cloud map calculated by CFD software are presented in the present invention, which is a design method for a vortex lift waverider aircraft based on combined flow field and profile.

[0031] Figure 11 The Case 3 flow field characteristic line calculation results and pressure and Mach number cloud maps calculated by CFD software are presented for the design method of a vortex lift waverider aircraft based on combined flow field and profile of the present invention.

[0032] Figure 12 The present invention provides the Case 4 flow field characteristic line calculation results and pressure and Mach number cloud maps calculated by CFD software for a vortex lift waverider design method based on combined flow field and profile.

[0033] Figure 13 This is a diagram showing the distribution of flow field pressure along the surface ACD in the design method of a vortex lift waverider based on combined flow field and profile of the present invention.

[0034] Figure 14 This is a comparison diagram of the streamlines obtained by program solution and CFD solution in the design method of vortex lift waverider aircraft based on combined flow field and profile of the present invention.

[0035] Figure 15 This is a slice cloud diagram of a vortex lift waverider design method based on combined flow field and profile, according to the present invention.

[0036] Figure 16 This is a diagram showing the lift-to-drag ratio calculation results of a vortex lift-wave riding aircraft design method based on combined flow field and profile according to the present invention.

[0037] Figure 17 This is a diagram showing the lift coefficient calculation results of a vortex lift waverider design method based on combined flow field and profile according to the present invention.

[0038] Figure 18 This invention presents a vortex system structure diagram for a vortex lift waverider design method based on combined flow field and profile. Detailed Implementation

[0039] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0040] Example

[0041] like Figure 1 - Figure 18 As shown, a design method for a vortex-lift waverider aircraft based on combined flow field and profile is applicable to the design principle of waverider aircraft with vortex lift characteristics. The steps of the vortex-lift waverider aircraft design method are as follows:

[0042] (1) Determine the combined flow field design of the aircraft. The entire flow field of the aircraft is divided into two parts for design. The first part is the flow field in the middle of the fuselage and the flow field in the transition section from the fuselage to the wing. The second part is the reference flow field of the wing section. The design condition of the incoming flow is Ma=8, and the corresponding Mach angle is:

[0043]

[0044] Therefore, the shock angle cannot be lower than this value. In order to make the wing design as thin as possible, the shock angle of the wing section's tangential flow field is set to 8°.

[0045] Mid-fuselage flow field design: Designed as a two-dimensional curved shock wave. The airflow parameters of the planar shock wave flow field are more uniform, which is beneficial to the air intake design. Its shock wave profile is as follows: Figure 1 The solid red line in the middle is a quadratic curve. The curve equation is determined by three parameters: the shock angle at the starting point, the shock angle at the ending point, and the flow field length. The surface of the wall influence zone is still defined as a straight line and is tangent to the surface of the shock-dependent zone at the intersection of the two. The shock angle at the starting point of the shock curve is set to 8°, and the shock angle at the ending point is 14°. The length of the entire reference flow field is 10m.

[0046] Transition section flow field design: The spanwise length of the transition section flow field is 0.84m. The shock wave angles at the start and end points of the shock wave profile linearly decrease from 14° to 8° along the spanwise direction, achieving a transition to a flow field with small shock wave angles. Because the entire spacecraft is symmetrical about the XOY plane, therefore... Figure 2The flow field diagram for the fuselage and the transition section from the fuselage to the wing only shows half of the flow field symmetrical about the XOY plane. The spanwise length of the curved surface shock wave flow field in the middle of the fuselage is 1.12m, and the spanwise length of the flow field in the transition section is 0.84m (with a section on each side). The transition flow field is imagined to be composed of a series of parallel two-dimensional curved surface flow field slices distributed along the spanwise direction. The vertical plane slices in the diagram are ignored here for the first time. The influence of the transverse flow between the slice flow fields is ignored. These two-dimensional curved surface flow fields are used to approximate the three-dimensional flow field. The shock wave profile of this series of two-dimensional curved surface flow fields is similar to the shock wave profile of the flow field in the middle of the fuselage. The control parameters are the same for the initial shock wave angle and the abscissa of the shock wave termination point. Only the shock wave angle θ at the termination point decreases linearly from 14° to 8° along the spanwise direction, and finally turns into a plane straight shock wave. At the same time, the transition to a flow field with a small shock wave angle is also achieved.

[0047] Flow field design of the wing section, such as Figure 3 As shown: A herringbone cone flow field design is used, with a shock wave angle of 8° and a shock wave exit profile of y = 0.09(z - 1.4). 3 +0.04(z-1.4) 2 -10tan8°, z∈[1.4,3.0], the semi-span of the aircraft is designed to be 3m, so the shock wave exit profile ends at z=3. It should be noted that at the z=1.4m plane, that is, at the junction of the planar two-dimensional flow field and the tangential cone flow field, the local tangential plane curvature radius of the tangential cone flow field at this point is 12.5m, which is large enough. At the same time, the designed shock wave angle is small enough, so the streamlines emanating from the same point of these two flow fields are not significantly different in spatial position. Therefore, the streamlines near this interface will not experience surface generation algorithm failure or surface distortion during the lofting process.

[0048] (2) Determine the combined design profile. In the flow field design of the fuselage and transition section, determine the bottom profile of the upper surface. The planar shape of the wing is designed as a large swept delta wing with a sweep angle of 62°. It is generated using the given horizontal projection profile method. At the same time, when the design profile is converted, ensure that the connection and transition of the streamline at the flow field interface is properly handled.

[0049] As the loading component of an aircraft, the shape of the fuselage projected along the flow direction affects its volumetric efficiency to some extent. Therefore, when designing the fuselage and transition sections, the bottom profile of the upper surface of the waverider is given in the corresponding flow field, and its design profile is as follows:

[0050] y = -0.35714z 2 -0.5z - 0.5, z ∈ [0, 0.56]

[0051] y = -0.125z 2 +0.14z-0.9312, z∈[0.56,1.4]

[0052] The planform shape of the wing is crucial for wide-speed-range design. Therefore, the wing section is generated using a given horizontal projection profile in its corresponding tangential conical flow field. Here, the wing planform is designed as a large-sweep delta wing with a sweep angle of 62°. The horizontal projection profile equation is: x = -1.8908z + 5.6727, z ∈ [1.4, 3.0]. It should be noted that while combined profile design increases the diversity of waverider design methods, attention must be paid to the transition of streamlines at the flow field interface during profile conversion. To ensure the smoothness of the surface generated by the lofting of nearby streamlines during profile conversion, the horizontal projection profile of the wing section must be determined based on the previously designed upper surface profile. Figure 2 For example, in the diagram, the ΔDEF region represents the planar straight shock wave flow field section ignoring the influence of transverse flow. BH is the streamline of this flow field section, and point B is the starting point of this streamline and also the leading edge point of the aircraft. For the combined profiles to be connected, point B must satisfy both the requirement that its projection point C in the XOZ plane is the starting point of the horizontal projection profile, and its projection point A in the YOZ plane is the ending point of the upper surface profile. From these constraints, it is easy to derive ΔAEB~ΔDEF, and thus... Given the coordinates of point A on the upper surface profile, the coordinates of point C can be obtained using the similarity relationship described above, thus yielding the expression for the horizontal projection profile:

[0053] x=-1.8908z+5.6727,z∈[1.4,3.0]

[0054] (3) Streamline tracing and lofting: Given the horizontal projection profile of the herringbone cone flow field, the wave-riding lower surface is generated through streamline tracing, combined with... Figure 5 To explain, point B is a discrete point located on the shock wave exit profile. The coordinates of point B are known. The ΔABO1 region is the tangent plane corresponding to point B. Point A is the vertex of the tangent cone. O1 is the center of the circle of curvature. Knowing the coordinates of point B and the expression for the shock wave exit profile, we can determine the length of the radius of curvature O1B and the angle θ between the local tangent plane and the vertical plane. From these two parameters, we can deduce the coordinates of the center O1. The vertex A of the tangent cone and the center O1... 1，Having the same y-axis and z-axis coordinates, the length of O1A can be determined by the tangent relationship in the right angle ΔABO1: O1B=O1Atan8°. Thus, the three-dimensional spatial coordinates of point A are also determined. Point A′ is the projection point of point A onto the XOZ plane, sharing the same x-axis and z-axis coordinates as point A. Point B′ is the projection point of point B onto the XOZ plane, similarly sharing the same x-axis and z-axis coordinates as point B. Therefore, A′B′ is the horizontal projection line of the shock wave profile AB within the local tangent plane onto the XOZ plane, and CG is the leading edge horizontal projection profile. A′B′ and... The points intersect at point P′. Point P′ can be determined by simultaneously solving the equations of the straight line A′B′ and the equations of the horizontal projection profile of CG. Point P′ is the projection of the leading edge point P on the local tangent plane onto the XOZ plane. Point P and point P′ share the same x-axis and z-axis coordinates. Therefore, the distance PD from point P to the XOZ plane can be determined. Determining PD is equivalent to locating the leading edge point in the tangent cone flow field ABO1. Streamlines are traced in this two-dimensional plane to obtain the expression of the streamlines in this plane coordinate system. Then, the streamlines are transformed into a three-dimensional rectangular coordinate system as shown in the figure using the trigonometric relationships in the figure.

[0055] (4) Generate the waverider configuration. Using the above-mentioned design profile as input, perform streamline tracing and streamline lofting to obtain the vortex-lift waverider aircraft. That is, solve the leading edge using the above method, and perform streamline tracing and streamline lofting from the leading edge to obtain... Figure 6 The wave-riding configuration shown.

[0056] Flow field solution and streamline tracing principle: This design scheme uses the inverse characteristic line method to solve the flow field for a given shock wave shape. The inverse characteristic line method uses a predictor-corrector calculation with second-order accuracy. The governing equations of the supersonic inviscid flow field exhibit hyperbolic characteristics. Hyperbolic partial differential equations can be solved using the characteristic line method, where the characteristic lines are a family of space curves such as... Figure 7 The partial differential equations can be simplified by reducing their dimension by one along the characteristic lines. The differential relations satisfied by the physical quantities after dimension reduction are called compatibility relations. For a two-dimensional inviscid and rotatable supersonic flow field, there are three families of characteristic lines in the flow field, namely left / right Mach lines (C+ / C-) and streamlines (C0). The left / right Mach lines can also be called left / right characteristic lines. In this paper, we will use left / right characteristic lines uniformly.

[0057] Equations of left / right line characteristics:

[0058]

[0059] λ is the slope of the characteristic line, and ± represent the left-hand and right-hand characteristic lines, respectively. θ is the airflow velocity vector angle, tanθ = v / u, where u and v are the velocity components in the two directions, and μ is the Mach angle, μ = arcsin(1 / Ma).

[0060] The flow parameters along the left / right characteristic lines satisfy the compatibility relation:

[0061]

[0062] In the formula, ± represents the compatibility relationship satisfied by the left and right characteristic lines respectively, δ is a coefficient, when δ=0 is a two-dimensional planar flow field; when δ=1 is an axisymmetric flow field, y is the radial distance from the point on the characteristic line to the axis of symmetry.

[0063] Equations whose characteristic lines are streamlines:

[0064]

[0065] There are two compatibility equations along the streamlines:

[0066] ρVdV+dp=0 (4)

[0067] dp-c 2 dp = 0 (5)

[0068] The system of equations consisting of characteristic line equations (1) and (3) and compatibility equations (2), (4), and (5) can replace the system of partial differential equations consisting of continuity equations, momentum equations, and energy equations to describe the two-dimensional steady isentropic supersonic flow field. Furthermore, these equations are ordinary differential equations, which simplify and decouple the partial differential equations, thus improving the solution efficiency. In actual solution, the differential terms in equations (1)-(5) need to be replaced with difference terms, and combined with the known upstream boundary conditions during the unit process, a program is developed to advance the solution of the downstream flow field until the complete information of the reference flow field is obtained. Streamline tracing is based on the aforementioned characteristic line difference grid, using interpolation to obtain a series of discrete points on the streamlines, and then fitting them with spline curves.

[0069] Flow field solution and streamline tracing program verification:

[0070] To verify the correctness of the flow field solver program, four axisymmetric flow fields with different shock wave shapes were designed. The characteristic line program was used to calculate these four reference flow fields and the results were compared with those calculated by CFD software. Figure 8 This is a schematic diagram of the reference flow field. Curve AB is the generatrix of the shock cone, which is designed as a smooth quadratic curve in this paper, as shown below. Figure 8 As shown, the shock angle at point A, where the apex A of the shock cone coincides with the origin O, is β. a The x-coordinate of the shock wave termination point B is x. b Shock angle is β bThe expression for the shock cone generatrix AB (6) can be constrained by the three parameters mentioned above. The curve ACD is the surface corresponding to the shock wave, which can be divided into two segments. One segment is the AC segment of the shock wave dependent region ABC enclosed by AB and BC, and the other segment is the CD segment of the wall influence region BCD enclosed by BC and BD. Since the given shock wave shape can only solve the flow field of the shock wave dependent region ABC, it is necessary to artificially give the wall profile CD to complete the flow field. The principle for giving the wall profile CD here is: let the profile CD be a straight line, and the slope of the straight line is equal to the tangent of the flow angle at point C (the flow angle at point C is obtained by solving the shock wave dependent region ABC). This ensures that the entire curve ACD is continuously differentiable at point C, and minimizes the situation where the same family of characteristic lines intersect too early due to unreasonable wall profile CD, and the flow field cannot be solved.

[0071]

[0072] Table 1 lists the design parameters for four reference flow fields. Cases 1-4 are all axisymmetric flow fields, with other parameters remaining unchanged except for the shock wave angle β at the termination point. b To adjust the shock wave shape. Shock angle β from Case 1 to Case 4. b As the flow increases, the shock wave shape becomes increasingly steep. The operating conditions for all four reference flow fields are: free flow Ma ∞ =6, static pressure P0 = 2550Pa, static temperature T0 = 221.5K, corresponding to atmospheric conditions at an altitude of 25km.

[0073] Table 1

[0074]

[0075] Figure 9-12 The results of the flow field characteristic line calculations for Cases 1-4 are compared with the pressure and Mach number contour maps calculated by CFD software. In these four figures, the upper half of each figure shows the pressure and Mach number contour maps of the flow field solved by the inverse characteristic line method, while the lower half shows the corresponding results from the CFD software. Since the CFD solution requires wavefront flow field information, the figure will have an additional wavefront flow field compared to the waveback flow field solved by the inverse characteristic line method. It should be noted that since the CFD software does not have the ability to solve the inverse problem (inversely calculating the flow field given the shock wave shape), the object surface coordinate information calculated by the inverse characteristic line method is used as the input condition. The object surface geometry is reconstructed using CAD software, and then the inviscid flow field is solved using CFD software.

[0076] From the perspectives of shock wave shape, flow field pressure, and Mach number distribution, the results obtained by the inverse characteristic line method and CFD software are extremely similar. However, a qualitative comparison based solely on flow field contour maps is not entirely convincing. To better verify the accuracy of the inverse characteristic line method, such as... Figure 13As shown, the distribution law of flow field pressure along the surface ACD of the object was extracted from Case 1-Case 4 of CFD software, and quantitatively compared with the results calculated by the inverse characteristic line.

[0077] Figure 13 In the figure, the wall pressure curves for all four flow fields show inflection points. This is because although the forced straight line segment CD of the object surface ensures a geometric transition with the shock wave dependent region AC segment at point C, the pressure transition at point C is not smooth in most cases. Therefore, the inflection point of the pressure curve in the figure corresponds to the boundary point C between the shock wave dependent region and the wall-affected region on the object surface. From Case 1 to Case 4, with other parameters unchanged, β... b As the shock increases, it becomes increasingly "steep." The steeper the shock, the larger the shock-dependent region ACD is, enclosed by the shock profile AB, the counter-leftward characteristic line BC emanating from the shock's endpoint, and the object surface AC. Therefore, point C shifts backward. This also explains why the inflection points of the pressure curves from Case 1 to Case 4 in the figure shift backward.

[0078] The inverse characteristic line in the figure and the pressure value along the surface calculated by CFD software still have some deviation in the early part of the flow field, but the two match better as the flow progresses. This is because in the compatibility equation (2) of the axisymmetric flow field, the denominator includes the radial distance, while the radial distance from the discrete point near the cone tip of the conical flow field to the axis of symmetry is very small, resulting in a large number dividing the small number, which is the cause of the error.

[0079] To verify the correctness of the streamline tracing program, based on the four axisymmetric reference flow fields of Case 1-Case 4, a streamline in its meridional plane was calculated using CFD software and the streamline tracing program. The radial distance from the starting point of the streamline to the axis of symmetry is 0.2m.

[0080] like Figure 14 As shown, the solid black line with arrows is the streamline obtained by the CFD software, and the series of discrete points are the points on the streamline obtained by the feature line program. The two are highly overlapping.

[0081] Based on the above results, it can be concluded that the inverse characteristic line and streamline tracing program written in this chapter meets the requirements and can be applied to waverider design.

[0082] Configuration aerodynamic performance simulation:

[0083] To verify whether a vortex-lift waverider aircraft designed based on a combined flow field and combined profile can effectively ride waves, the viscous flow field of the aircraft was simulated when it flew at a design Mach number with an angle of attack of 0° and 2°. Figure 15 Pressure slice cloud images of the upper surface of the aircraft were extracted under these two flight angles of attack.

[0084] Overall, the overflow intensity at 0° angle of attack is not large, while at 2° angle of attack, the overflow volume and overflow intensity are significantly weaker than at 0° angle of attack, and the wave riding effect is stronger, proving that the configuration can effectively ride waves at the design Mach number.

[0085] Under high Mach number cruise conditions, it is generally desirable for an aircraft to have a high lift-to-drag ratio. Therefore, high Mach number aerodynamic performance simulation was performed on this configuration, and the calculation conditions are shown in Table 2.

[0086] Table 2

[0087]

[0088] The calculation results of the lift-to-drag ratio are as follows: Figure 16 As shown, under the design Mach number of 8, its maximum lift-to-drag ratio is greater than that of the other two operating conditions, reaching 4.71. The maximum lift-to-drag ratio of the other two non-design operating conditions is not less than 4.5. The lift-to-drag ratio calculation here already includes the bottom resistance.

[0089] At low Mach numbers, a high lift coefficient is typically required for the aircraft. Therefore, two simulation conditions were selected: Ma = 0.4 at an altitude of 0 km and Ma = 0.8 at an altitude of 10 km. The lift coefficient calculation results are as follows: Figure 17 As shown (the reference area for calculating the lift coefficient is its horizontal projected area). The blue line in the figure represents the linear fit of the lift coefficient at small angles of attack. The lift coefficient increases non-linearly in both operating conditions. The non-linear gain of the lift coefficient is greatest when Ma = 0.4 and the angle of attack is 16°, increasing by 18.8% compared to the linear value. Figure 18 The diagram illustrates the vortex structure of the flow field under this operating condition. The figure shows a pair of strong flow-directed vortices in the leeward region of the aircraft. This is due to the gradual mixing and amplification of the cone-shaped vortex initially generated at the nose as it develops downstream, passing through the leading-edge vortices of the central wing and outer wing. Therefore, the wave-riding configuration designed using this method effectively couples the vortex lift mechanism, proving the effectiveness of the design method. Overall, under simulated low-speed conditions, the maximum lift coefficient of this configuration is no less than 1.2.

[0090] The preferred embodiments of the present invention disclosed above are merely illustrative of the invention. These preferred embodiments do not exhaustively describe all details, nor do they limit the invention to any specific implementation. Clearly, many modifications and variations can be made based on the content of this specification. This specification selects and specifically describes these embodiments to better explain the principles and practical applications of the invention, thereby enabling those skilled in the art to better understand and utilize the invention. The invention is limited only by the claims and their full scope and equivalents.

Claims

1. A design method for a vortex-lift waverider aircraft based on combined flow field and profile, characterized in that, It is adapted to the design principle of waverider aircraft with vortex lift characteristics. The design steps of the vortex lift waverider aircraft are as follows; (1) Determine the combined flow field design of the aircraft. The entire flow field of the aircraft is divided into two parts for design. The first part is the flow field in the middle of the fuselage and the flow field in the transition section from the fuselage to the wing. The second part is the reference flow field of the wing section. The airflow field design in the middle of the fuselage is designed as a two-dimensional curved shock wave. The airflow parameters of the planar shock wave flow field are more uniform, which is beneficial to the design of the air intake. The transition section flow field design is as follows: the spanwise length of the transition section flow field is 0.84m, and the shock wave angles at the starting and ending points of the shock wave profile are linearly reduced from 14° to 8° along the spanwise direction, thus achieving a transition to a flow field with small shock wave angles. The wing section flow field design uses a scissor-cone flow field design with a shock wave angle of 8°. (2) Determine the combined design profile. In the flow field design of the fuselage and transition section, determine the bottom profile of the upper surface. The planar shape of the wing is designed as a large swept delta wing with a sweep angle of 62°. It is generated using the given horizontal projection profile method. At the same time, when the design profile is converted, ensure that the connection and transition of the streamline at the flow field interface is properly handled. (3) Streamline tracing and lofting: Given the horizontal projection profile of the tangential cone flow field, the wave-riding lower surface is generated by streamline tracing; (4) Generate wave-riding configuration: Using the above design profile as input conditions, perform streamline tracing and streamline lofting to obtain the vortex lift wave-riding aircraft.

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