A method for designing a non-uniform inflow precursor / inner turning inlet integrated design

Abstract

The application discloses a non-uniform inflow precursor / inner rotating inlet integrated design method, and relates to the technical field of precursor inlet integrated design. The method comprises the following steps: keeping the performance advantages of a vehicle precursor; selecting a non-uniform flow parameter at an inlet position according to a preselected inlet position of an inlet; generating a three-dimensional inner convergent inlet profile; selecting a turbojet channel inlet position on the compression profile of the three-dimensional inner convergent inlet; designing a turbojet channel; and finally completing the non-uniform inflow precursor combined power inlet integrated design, thereby providing a new idea for the inner and outer flow integrated design of a wide-speed-range vehicle under a non-uniform condition.

Description

Technical Field

[0001] This invention relates to the field of integrated design technology for forebody inlet ducts, and more specifically to an integrated design method for a non-uniform inlet flow forebody / internal turning inlet duct. Background Technology

[0002] Combined propulsion technology is a key technology for achieving wide-speed-range flight of aircraft. Air-breathing propulsion provides high specific impulse, and the key to its success lies in the integrated design of the aircraft fuselage and the air intake, i.e., the integrated design of internal and external flow. The most challenging aspect is the integrated design of the aircraft forebody and the air intake. Existing integrated technologies include: the integration of the aircraft fuselage and the two-dimensional air intake based on wedge-guided waverider theory, which is developed by analyzing the coupling problem between the wedge-guided waverider and the air intake under planar flow conditions; and the integration of the external flow tangential waverider and the internal flow two-dimensional air intake based on the tangential conical waverider theory. In China, You Yancheng et al. have conducted in-depth research on waverider theory and applied it to inward contraction flow, proposing a dual-waverider theory method applicable to both the aircraft fuselage and the three-dimensional inward contraction air intake, solving the integration problem of hypersonic external and internal flow. However, current integrated methods have not conducted in-depth research on the problem of non-uniform incoming flow generated by the aircraft forebody. Due to the influence of forebody pre-compression, the airflow entering the intake will have strong non-uniform distortion characteristics. This characteristic will seriously affect the distortion performance of the turbojet channel of the combined power intake and the anti-back pressure capability of the ramjet channel. It is proposed that the integrated design technology of forebody / intake considering non-uniform incoming flow is crucial to the successful realization of wide speed range flight. Therefore, this patent proposes to use the three-dimensional characteristic line theory to realize the integrated design of non-uniform incoming flow forebody combined power intake. Summary of the Invention

[0003] The problem this invention aims to solve is to provide an integrated design method for a non-uniform inlet forebody / internal converging inlet. While maintaining the performance advantages of the aircraft forebody, it proposes non-uniform flow parameters at the inlet position based on a pre-selected inlet position, generates a three-dimensional internal converging inlet profile, selects the turbojet channel inlet position on the compression profile of the three-dimensional internal converging inlet, designs the turbojet channel, and finally completes the integrated design of the non-uniform inlet forebody combined power inlet. This provides a new approach for the integrated design of internal and external flow of wide-speed-range aircraft under non-uniform conditions.

[0004] The technical solution provided by this invention to solve the above problems is as follows:

[0005] A method for integrated design of non-uniform incoming flow forebody / internal inlet duct, the method comprising the following steps:

[0006] (1) Generate the aircraft forebody based on the incoming flow conditions and geometric parameter requirements;

[0007] (2) Based on the aircraft forebody generated in step (1), the flow field parameters of the lower surface of the aircraft forebody are obtained by numerical simulation.

[0008] (3) Select the inlet shape of the combined power inlet, and interpolate the non-uniform flow field parameters at the location of the inlet profile of the combined power inlet based on the flow field parameters of the lower surface of the aircraft forebody obtained in step (2).

[0009] (4) Based on the combined power intake inlet shape and non-uniform flow field parameters obtained in step (3), the three-dimensional feature line method is used to generate the combined power intake compression profile, and the combined power intake compression profile outlet is stretched backward in a straight line to obtain the combined power intake stamping channel.

[0010] (5) Select the inlet shape of the turbojet channel on the surface of the combined power intake ramjet channel obtained in step (4), take the inlet shape as the starting profile and the inlet cross-sectional shape of the turbojet engine as the ending line, rely on cubic spline transition, use the mesh surface method to generate the combined power intake turbojet channel, and finally complete the integrated design method of non-uniform inflow forebody / internal turning intake through geometric modification.

[0011] Preferably, the incoming flow conditions in step (1) are the Mach number at the transition point of the wide-speed-range aircraft and the flight altitude; the geometric parameters include the aspect ratio, width constraint, height constraint, etc.; based on the above two input parameters, the aircraft forebody is generated using waveriding theory, and then the lower surface of the aircraft forebody is obtained.

[0012] Preferably, the flow field parameters of the lower surface of the aircraft forebody in step (2) are obtained by numerical simulation based on the lower surface of the aircraft forebody generated in step (1). The numerical simulation is based on a full three-dimensional structured grid and is solved using the Reynolds average method. After the solution is completed, the flow field parameter information of the whole field, including Mach number, static pressure, static temperature, etc., is output and stored in the form of a data table to provide data support for step (3).

[0013] Preferably, in step (3), according to geometric requirements, a combined power intake inlet is set on the lower surface of the aircraft forebody. The inlet shape is selected as a rounded rhombus. The rounded rhombus is discretized, and the flow field parameter data of the lower surface of the aircraft forebody obtained in step (2) is interpolated to obtain the non-uniform flow field parameters at the location of the combined power intake inlet profile.

[0014] Preferably, in step (4), based on the combined power intake inlet shape and non-uniform flow field parameters obtained in step (3), the three-dimensional characteristic line method is used with the combined power intake inlet as the starting point and the corresponding non-uniform flow field parameters as the initial conditions. The compression profile of the combined power intake is solved one by one downstream in a step-by-step manner. A series of compression profiles are arranged in the circumferential direction to finally obtain the compression surface of the combined power intake. The outlet profile of the compression surface is used as the starting profile and is stretched backward in a straight line to obtain the ramming channel of the combined power intake.

[0015] Preferably, step (5) involves selecting the inlet shape and position of the turbojet channel on the surface of the obtained combined power intake ramjet channel. The turbojet channel inlet shape is a rounded rectangle, arranged on the side wall of the ramjet channel. The rounded rectangle is used as the starting profile, and the turbojet engine inlet cross-section shape is used as the ending line. The turbojet channel is generated by using a convergence from both sides to the middle, relying on cubic splines to complete the mesh surface. Finally, the integrated design method of non-uniform inflow forebody / internal turning intake is completed by geometric modification.

[0016] The beneficial effects of this invention are: the integrated configuration of the non-uniform inflow forebody combined with the propulsion inlet generated by this design method introduces a new design concept for the integration of non-uniform inflow forebody combined with the propulsion inlet. This design method enables the design of propulsion systems under non-uniform conditions, effectively improving the operational capabilities of wide-range aircraft. Attached Figure Description

[0017] The accompanying drawings, which are provided to further illustrate the invention and constitute a part of this invention, are illustrative embodiments of the invention and their descriptions are used to explain the invention and do not constitute an undue limitation of the invention.

[0018] Figure 1 This is a geometric diagram showing the positions of the aircraft's forebody and the air intake.

[0019] Figure 2 This is a schematic diagram of the aircraft's forebody, combined power intake compression profile, and ramjet channel.

[0020] Figure 3 This is an overall diagram of the integrated configuration of the non-uniform inlet forebody / internal inlet.

[0021] Figure 4 It is the lower surface of the non-uniform inlet flow precursor / internal intake integrated configuration.

[0022] The markings in the diagram are as follows: 1 represents the lower surface of the aircraft forebody, 2 represents the combined propulsion air intake, 3 represents the combined propulsion air intake compression profile, 4 represents the ramjet channel, 5 represents the turbojet channel, and 6 represents the turbojet engine inlet. Detailed Implementation

[0023] The following will describe in detail the implementation of the present invention with reference to the accompanying drawings and embodiments, so that the process of how the present invention uses technical means to solve technical problems and achieve technical effects can be fully understood and implemented accordingly.

[0024] The main implementation steps of the integrated design method for non-uniform inlet forebody / inner turn intake include:

[0025] (1) This implementation plan is designed for a Mach 4 wide-speed-range vehicle. The design Mach number is taken as the transition point Mach number (Mach 3), and the flight altitude is taken as 20km. At this flight altitude, the corresponding incoming static pressure is 5529Pa, the incoming static temperature is 216.65K, and the incoming density is 0.088908kg / m³. 3 The above incoming flow parameters are used as input conditions. The geometric constraint slenderness ratio is selected as 1.8:1. The design process adopts an inviscid flow field. Under inviscid conditions, supersonic airflow has similarity. Therefore, after the slenderness ratio is determined, the specific geometric dimensions can be dimensionless. In this implementation scheme, the width constraint is set to 1, so the length direction is 1.8, and the height constraint is 0.3. According to the above incoming flow parameters and geometric parameter conventions, the incident wave angle is selected as 15° under Mach 3 conditions. The wave-riding theory and streamline tracing method are used to generate the lower surface 1 of the aircraft forebody. The designed forebody geometry is thin in the middle and rear sides, with a large sweep angle to ensure aerodynamic performance in all flight conditions.

[0026] (2) Based on the lower surface 1 of the aircraft forebody obtained in step (1), numerical simulation calculations are carried out at Mach 3 and 20km according to the inflow conditions at the design point to obtain the flow field information of the lower surface of the aircraft forebody and extract the corresponding flow field parameters to provide data support for the subsequent air intake design. The required flow field parameters include x-coordinate, y-coordinate, z-coordinate, Mach number, density, static pressure, static temperature, x-direction velocity, y-direction velocity and z-direction velocity. The numerical simulation is based on a full three-dimensional structured grid with no less than 5 million structured grid cells. The Reynolds average method is used for the solution. The turbulence model is selected as the k-omega SST model, double precision, second-order upwind scheme. During the solution process, the lift coefficient, drag coefficient and moment coefficient of the lower surface 1 of the aircraft forebody are monitored. When each monitored quantity no longer changes, convergence is determined. After the solution is completed, the flow parameter information of the whole field is extracted and a data table is made. The data table is in the form of horizontal arrangement of variables and vertical arrangement of data, with a total of 10 columns of data. This data table will provide non-uniform inflow support for step (3).

[0027] (3) Select the inlet shape 2 of the combined propulsion inlet, and based on the flow field parameters of the lower surface of the aircraft forebody obtained in step (2), use the inverse distance weighted interpolation method to obtain the non-uniform flow parameters of the combined propulsion inlet. The inverse distance weighted interpolation method assumes that the surrounding discrete points have a local influence on the interpolation points, and that this influence weakens with increasing distance. The core of this method lies in solving the weight function corresponding to the surrounding discrete points. The weight of each discrete point can be expressed as:

[0028]

[0029]

[0030] In the formula, subscript 0 represents the interpolation point, subscript i represents the surrounding discrete points, and n represents the number of discrete points closest to the interpolation point. The arbitrary parameters at the interpolation point can be obtained based on the weights of each discrete point.

[0031]

[0032] Design a combined power inlet 2 on the lower surface 1 of the aircraft forebody. The inlet shape is selected as a rounded rhombus. Discretize the rounded rhombus to obtain 100 discrete points on the inlet shape. The shape of the incident shock wave of the inlet is selected as an inward contraction shock wave that fully covers the inlet of the inlet. Then, interpolate in the flow field data table of the lower surface 1 of the aircraft forebody obtained in step (2) to obtain the non-uniform flow field parameters at the location of the profile of the combined power inlet 2.

[0033] (4) Based on the discrete points of the combined power inlet 2 obtained in step (3) and the non-uniform flow field parameters of the lower surface of the aircraft forebody 1 at the Mach number of Mach 3 at the transition point, the non-coaxial characteristic line method is used to solve the compression profile of the combined power inlet 2 downstream in a step-by-step manner, with the discrete points generated by the combined power inlet 2 as the starting point and the corresponding non-uniform flow field parameters as the initial conditions. The non-coaxial characteristic line method introduces the change of the radius of curvature in the traditional characteristic line theory. The local radius of curvature of the corresponding incident shock surface is calculated based on the non-uniform parameters of the inlet flow and the inlet 2. Then, the characteristic line equation and compatibility equation under this condition are solved. Finally, the discrete compression profile of the combined power inlet compression surface 3 is obtained. The combined power inlet compression surface 3 has the characteristic of three-dimensional inward contraction, with a total contraction ratio of 4 and an internal contraction ratio of 1.4. The combined power inlet ram air passage 4 is obtained by stretching the outlet profile of the combined power inlet compression surface 3 backward in a straight line as the starting profile.

[0034] (5) The step (5) is to select the inlet shape and inlet position of the turbojet channel on the surface of the obtained combined power intake ramjet channel 4. The inlet shape of the turbojet channel is a rounded rectangle with an inlet length-to-width ratio of 2:1. The width dimension is the same as the width of the ramjet channel sidewall. The turbine channel is arranged on the sidewall of the ramjet channel. The turbine engine adopts a single-engine scheme, and the ramjet engine adopts a dual-engine scheme. Therefore, the same geometric turbine channel inlet is opened on the left and right ramjet channels at the same time. Taking the two turbine channel outlets on the sidewall of the ramjet channel as the starting point and the turbojet engine inlet cross-sectional shape 6 as the ending line, the combined power intake turbojet channel 5 is generated by the grid surface completed by the cubic spline in the form of converging from both sides to the middle. The control parameters of the cubic spline are determined by four parameters: starting point coordinates, starting point slope, ending point coordinates, and ending point slope. The starting point coordinates are located at the inlet center of the combined power intake turbojet channel 5, and the ending point coordinates are located at the turbojet engine inlet cross-sectional shape 6. In this embodiment, the starting point slope is tan(30°) and the outlet slope is tan(0°). The integrated design of the non-uniform inlet forebody combined with the power intake was completed through five steps, culminating in geometric modification. The overall layout and geometry are shown in the attached figure. Figure 1 As shown.

[0035] This integrated design method for the non-uniform inflow forebody combined power inlet maintains the advantages of both the aircraft forebody and the combined power inlet. Based on the flow parameters of the lower surface of the aircraft forebody, it conducts inlet compression profile design, which can effectively improve the aerodynamic performance of the combined power inlet and broaden the working range of the integrated configuration. This introduces a new approach to the integrated design of internal and external flow for wide-speed-range aircraft.

[0036] Compared with existing technologies, the advantages of this invention are: the integrated configuration of the non-uniform inflow forebody combined propulsion inlet generated by this design method can introduce new design ideas for the integration of non-uniform inflow forebody combined propulsion inlets. This design method enables the design of propulsion systems under non-uniform conditions, effectively improving the operational capabilities of hypersonic vehicles.

Claims

1. A method for integrated design of a non-uniform incoming flow forebody / internal inlet duct, characterized in that, The method includes the following steps: (1) Generate the aircraft forebody based on the incoming flow conditions and geometric parameter requirements; (2) Based on the aircraft forebody generated in step (1), the flow field parameters of the lower surface of the aircraft forebody are obtained by numerical simulation. (3) Select the inlet shape of the combined power inlet, and interpolate the non-uniform flow field parameters at the location of the inlet profile of the combined power inlet based on the flow field parameters of the lower surface of the aircraft forebody obtained in step (2). (4) Based on the combined power intake inlet shape and non-uniform flow field parameters obtained in step (3), the three-dimensional feature line method is used to generate the combined power intake compression profile, and the combined power intake compression profile outlet is stretched backward in a straight line to obtain the combined power intake stamping channel. (5) Select the inlet shape of the turbojet channel on the surface of the combined power intake ramjet channel obtained in step (4), take the inlet shape as the starting profile and the turbojet engine inlet cross section shape as the ending line, rely on cubic spline transition, use the mesh surface method to generate the combined power intake turbojet channel, and finally complete the integrated design method of non-uniform inflow forebody / inner turn intake through geometric modification. In step (4), based on the combined power intake inlet shape and non-uniform flow field parameters obtained in step (3), the three-dimensional characteristic line method is used with the combined power intake inlet as the starting point and the corresponding non-uniform flow field parameters as the initial conditions. The compression profile of the combined power intake is solved one by one downstream in a step-by-step manner. A series of compression profiles are arranged in the circumferential direction to finally obtain the compression surface of the combined power intake. The outlet profile of the compression surface is used as the starting profile and stretched backward in a straight line to obtain the ramming channel of the combined power intake.

2. The integrated design method for a non-uniform incoming flow forebody / internal inlet duct according to claim 1, characterized in that, The incoming flow conditions in step (1) are the Mach number of the transition point of the wide-speed-range aircraft and the flight altitude; the geometric parameters include the aspect ratio, width constraint, and height constraint; based on the above two input parameters, the aircraft forebody is generated using waveriding theory, and then the lower surface of the aircraft forebody is obtained.

3. The integrated design method for a non-uniform incoming flow forebody / internal inlet duct according to claim 1, characterized in that, The flow field parameters on the lower surface of the aircraft forebody in step (2) are obtained by numerical simulation based on the lower surface of the aircraft forebody generated in step (1). The numerical simulation is based on a fully three-dimensional structured grid and is solved using the Reynolds average method. After the solution is completed, the flow field parameters, including Mach number, static pressure, and static temperature, are output and stored in the form of a data table to provide data support for step (3).

4. The integrated design method for a non-uniform incoming flow forebody / internal inlet duct according to claim 1, characterized in that, In step (3), according to geometric requirements, a combined power intake inlet is set on the lower surface of the aircraft forebody. The inlet shape is selected as a rounded rhombus. The rounded rhombus is discretized, and the flow field parameter data of the lower surface of the aircraft forebody obtained in step (2) is interpolated to obtain the non-uniform flow field parameters at the location of the combined power intake inlet profile.

5. The integrated design method for a non-uniform incoming flow forebody / internal inlet duct according to claim 1, characterized in that, Step (5) involves selecting the inlet shape and position of the turbojet channel on the surface of the obtained combined power intake ramjet channel. The turbojet channel inlet shape is a rounded rectangle, arranged on the side wall of the ramjet channel. The rounded rectangle is used as the starting profile, and the turbojet engine inlet cross-section shape is used as the ending line. The turbojet channel is generated by using cubic splines to complete the mesh surface of the combined power intake turbojet channel. Finally, the non-uniform inflow forebody / internal turning intake integrated design method is completed by geometric modification.

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