Reverse design method of supersonic inlet based on complex three-dimensional internal flow field

Abstract

The present invention discloses a method for reverse design of a supersonic inlet based on a complex three-dimensional internal flow field, which relates to the field of hypersonic inlet design, including: specifying the incident shock wave and the reflected shock wave of the complex three-dimensional internal flow field according to the design requirements, and solving the zero-order and first-order flow parameters after the three-dimensional non-axisymmetric incident shock wave; discretizing the three-dimensional internal flow field composed of the incident shock wave and the reflected shock wave into a series of coaxial two-dimensional reference surfaces; decomposing the zero-order and first-order flow parameters into directions along the two-dimensional reference surface and perpendicular to the two-dimensional reference surface, and solving the zero-order and first-order flow parameter initial values ​​in the internal flow field in each reference surface, and performing lateral iteration on the initial values ​​of the parameters to reconstruct the entire three-dimensional internal flow field by reverse solution. The present invention provides a method for reverse design of a supersonic inlet based on a complex three-dimensional internal flow field, wherein the design of the inlet theoretically considers the lateral flow, and realizes the reverse design of the full three-dimensional flow of a known shock wave.

Description

Technical Field

[0001] The present invention relates to the field of near-space hypersonic inlet design, and more specifically, to a supersonic inlet reverse design method based on a complex three-dimensional internal flow field. Background Art

[0002] The air intake is a key component that connects the aircraft and the engine, and is an important part of the aircraft propulsion system. During the flight of the aircraft, air flows through the air intake into the engine, providing the engine with the required air flow. In this process, the air intake can convert the kinetic energy of the high-speed airflow into potential energy, decelerate and increase the pressure of the air, thereby effectively improving the quality of the air entering the engine. Therefore, the performance of the air intake is an important factor that restricts the engine performance and even the flight performance of the entire aircraft. Especially for supersonic or hypersonic aircraft, the impact of the air intake on the overall performance of the aircraft is more significant. Studies have shown that for a first-stage hydrocarbon fuel aircraft with a Mach number of 5 to 7, the specific impulse of the propulsion system can be increased by 3% to 5% for every 1% increase in the compression efficiency of the air intake (Heiser WH, Pratt DT. Hypersonic Airbreathing Propulsion [C]. Washington DC, USA: AIAA Inc, 1994.). In addition, the quality of the matching between the air inlet and the engine is directly related to flight safety; and the air inlet is one of the three strongest forward scattering sources of the aircraft, which is directly related to the stealth performance of the aircraft (Zhao Zhenshan, Li Yanliang, Dong Jun. Numerical Simulation of the Internal and External Flow Field of the Bump Air Inlet [C]. The 15th National Conference on Computational Fluid Dynamics. 2012.). It can be seen that improving the performance of the air inlet is of great significance to the improvement of the overall performance of the engine and the overall performance of the aircraft. Therefore, all aviation powers attach great importance to the design of the air inlet.

[0003] Although significant progress has been made in the field of hypersonic inlet research, and component performance is constantly improving. However, to date, hypersonic inlets still use a two-dimensional axisymmetric reference flow field, and then use streamline tracking technology to obtain the corresponding profile. However, the inlet obtained based on the two-dimensional axisymmetric reference flow field does not have a lateral flow, but only a directional flow, which will greatly reduce the compression efficiency of the inlet. Studies have shown that in the case of inlet design based on a two-dimensional axisymmetric flow field, the influence of the lateral pressure gradient cannot be ignored (Chauffour, ML, and Lewis, MJ, "Corrected Waverider Design for Inlet Applications," 40th AIAA / ASME / SAE / ASEE Joint Propulsion Conference and Exhibit, AIAA, Fort Lauderdale, Florida, 2004, p.3405.doi:10.2514 / 6.2004-3405.), especially in the case of hypersonic speed, the non-uniform flow field caused by the lateral pressure gradient greatly reduces the performance of the propulsion system.

[0004] Although significant progress has been made in various studies in the field of hypersonic inlet research, and component performance is constantly improving. However, to date, even the inlet designed using the kissing method still uses an axisymmetric reference flow field, and then uses streamline tracking technology to obtain the corresponding profile. However, the inlet obtained based on the axisymmetric reference flow field does not have lateral flow, only directional flow. At present, researchers have not found an effective method to obtain a full three-dimensional reference flow field, and then obtain an inlet with full three-dimensional flow. The lateral flow plays a vital role in improving the performance of the inlet. At the same time, researchers generally use the traditional characteristic line method for the inverse design of the reference flow field, which is not only complex to program, but also has poor stability, which limits the selection range of the basic flow field, thereby reducing the geometric construction range of the inlet. It can be seen that one of the problems restricting the design of hypersonic inlets is the lack of a supersonic inlet inverse design method based on a complex three-dimensional internal flow field. Summary of the invention

[0005] An object of the present invention is to solve at least the above problems and / or disadvantages and to provide at least the advantages which will be described hereinafter.

[0006] In order to achieve these purposes and other advantages of the present invention, a method for inverse design of a supersonic inlet based on a complex three-dimensional internal flow field is provided, comprising:

[0007] S1. Specify the incident shock wave and reflected shock wave of the complex three-dimensional internal flow field according to the design requirements, and solve the zero-order and first-order flow parameters after the three-dimensional non-axisymmetric incident shock wave;

[0008] S2, discretizing the three-dimensional internal flow field formed by the incident shock wave and the reflected shock wave into a series of coaxial two-dimensional reference surfaces;

[0009] S3, decomposing the zero-order and first-order flow parameters obtained in S1 into quantities along the two-dimensional reference surface and perpendicular to the two-dimensional reference surface;

[0010] S4, based on the quantities obtained in S3, solving the initial values ​​of the zero-order and first-order flow parameters in the internal flow field in each reference surface, performing lateral iteration on the initial values ​​of the parameters in each two-dimensional reference surface, so as to reconstruct the entire three-dimensional internal flow field by reverse solution;

[0011] S5. Specify the leading edge profile FCT of the air inlet on the incident shock wave surface, obtain the three-dimensional compression profile of the air inlet by streamline tracing, and stretch the tail end of the three-dimensional compression profile along the incoming flow direction to obtain the isolation section 11 of the air inlet.

[0012] Preferably, in S1, the incident shock wave is directly given by a parametric surface or a Bezier surface function, and the reflected shock wave is given by a shock wave angular growth rate of the reflected shock wave.

[0013] Preferably, in S1, when solving the zero-order and first-order flow parameters after the three-dimensional non-axisymmetric incident shock wave, the flow parameters of the vertical shock wave surface are solved by the flow field relationship, and the flow parameters of the extended shock wave surface are solved by the shock wave relationship.

[0014] Preferably, the control equations in the flow field relationship include:

[0015] Mass equation I:

[0016] Momentum equation I:

[0017] Energy equation I:

[0018]

[0019] Equation of state:

[0020]

[0021] In the above formula, p For pressure, ρ is the density, T is the temperature, m and l are two orthogonal directions along the shock wave surface, n is the normal direction of the shock wave surface, i To refer to the parameters, V m and Vl is the velocity along the shock wave surface, V n is the velocity perpendicular to the shock wave surface, γ is the specific heat ratio, R is the gas constant, K nm The shock wave surface along the curve m The normal curvature of the direction, K nl The shock wave surface along the curve l The normal curvature of the direction, K gl is the curve on the shock wave surface l The geodesic curvature of K gm is the curve on the shock wave surface m The geodesic curvature of τ gm The curve on the shock wave surface m The geodesic torsion.

[0022] Preferably, the control equation in the shock wave relationship includes:

[0023] Mass equation II:

[0024] Momentum equation II:

[0025]

[0026] Velocity equation II:

[0027]

[0028] Energy equation II:

[0029] In the above formula, p 1 is the pressure before the shock wave, p 2 is the pressure after the shock wave, ρ 1 is the shock front density, ρ 2 is the density after the shock wave, m and l are two orthogonal directions along the shock wave surface, n is the normal direction of the shock wave surface, i To refer to the parameters, V 1m Shock front m The speed of the direction, V 2m After shock wave m The speed of the direction, V 1n Shock front n The speed of the direction, V2n The shock wave trailing edge n The speed of the direction, γ is the specific heat ratio.

[0030] Preferably, in S4, the control equations for solving the initial values ​​of the parameters and the lateral iteration include:

[0031]

[0032]

[0033]

[0034] In the above formula, s and l are the characteristic line directions in the two-dimensional reference surface, φ for , P and D is the pressure gradient in the two-dimensional reference surface, W is the pressure gradient perpendicular to the two-dimensional reference surface, V is the velocity vector in the two-dimensional reference plane, w is the velocity vector perpendicular to the two-dimensional reference surface, p For pressure, ρ is the density, δ is the airflow deflection angle, μ is the Mach angle.

[0035] The present invention has at least the following beneficial effects:

[0036] Firstly, the design method of the present invention theoretically considers the lateral flow in the air inlet, realizes the inverse design of the full three-dimensional flow of the known shock wave, and at the same time, the air inlet is a full three-dimensional air inlet with lateral flow, which further improves the compression efficiency; and the lateral flow can help the low-energy flow move to both sides and the high-energy flow converge to the center, thereby improving the air inlet performance.

[0037] Secondly, the design method of the present invention broadens the design range of the air intake duct by reasonably designing the lateral flow, thereby further improving the working performance of the air intake duct.

[0038] Third, compared with the traditional flow field solving method, the design method of the present invention has significantly improved calculation efficiency and accuracy. The obtained flow field is no longer limited to the traditional two-dimensional axisymmetric internal flow field. More complex three-dimensional internal flow fields can be obtained through more free shock wave selection, and more different forms of air inlets can be designed to meet aerodynamic requirements.

[0039] Fourthly, the design method of the present invention can obtain richer flow information in the flow field (including zero-order and first-order flow parameters), which cannot be achieved by other methods. The rich flow information facilitates a deeper understanding of the nature of the flow.

[0040] Other advantages, objectives and features of the present invention will be embodied in part through the following description, and in part will be understood by those skilled in the art through study and practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] Figure 1 It is a schematic diagram of the complex three-dimensional internal flow field in the present invention;

[0042] Figure 2 It is a schematic diagram of solving the flow field within the reference surface in the present invention;

[0043] Figure 3 It is a schematic diagram of solving the flow field between reference surfaces in the present invention;

[0044] Figure 4 It is a schematic diagram of the inlet flow field tracking process based on the complex three-dimensional internal flow field of the present invention;

[0045] Figure 5 It is a schematic diagram of the overall structure of the air intake duct based on the complex three-dimensional internal flow field of the present invention;

[0046] Among them, 1-incident shock wave of complex three-dimensional internal flow, 2-reflected shock wave of complex three-dimensional internal flow, 3-two-dimensional reference surface, 4-shock wave reflection position, 5-incident shock wave in the reference surface, 6-pseudo streamline in the reference surface, 7-pseudo characteristic line in the reference surface, 8-leading edge profile of the inlet, 9-streamline tracking process line, 10-compression profile of the inlet, 11-isolation section of the inlet. DETAILED DESCRIPTION

[0047] The present invention is further described in detail below in conjunction with the accompanying drawings so that those skilled in the art can implement the invention with reference to the description.

[0048] The present invention obtains an inlet compression profile designed by a supersonic inlet inverse design method based on a complex three-dimensional internal flow field, and the inlet compression profile mainly includes: a leading edge profile, a compression profile and an isolation section.

[0049] Specifically, the inverse design method of the supersonic inlet based on the complex three-dimensional internal flow field includes the following steps:

[0050] Step 1: Specify the complex three-dimensional inflow incident shock wave according to the design requirements. The incident shock wave can be directly given by the function, and the angular growth rate of the reflected shock wave is specified.

[0051] Step 2: Solve the zero-order and first-order flow parameters after the complex three-dimensional inflow incident shock wave.

[0052] Step 3: discretize the complex three-dimensional internal flow field composed of the three-dimensional incident shock wave and the reflected shock wave into a series of coaxial two-dimensional reference surfaces, and decompose the zero-order and first-order flow parameters after the incident shock wave into quantities along the reference surface and quantities perpendicular to the reference surface.

[0053] Step 4: Solve the zero-order and first-order flow information of each reference surface along the reference surface direction and perpendicular to the reference surface direction.

[0054] Step 5: Reconstruct the entire three-dimensional flow field.

[0055] Step 6: Specify the leading edge profile FCT of the inlet on the incident shock wave surface, and use streamline tracing to obtain the three-dimensional compression profile of the inlet.

[0056] In step 2 and step 4, the three-dimensional curved shock wave theory and the three-dimensional curved characteristic line method can solve the first-order flow parameters in the flow field, and on the basis of realizing the inverse solution, accelerate the solution efficiency of the entire three-dimensional flow field.

[0057] Example:

[0058] like Figure 5 As shown in the figure, the main content of the supersonic inlet inverse design method based on complex three-dimensional internal flow field is to design the inlet compression surface through the incident shock wave and reflected shock wave of the complex three-dimensional internal flow, and the inlet compression surface here mainly includes: leading edge profile, compression surface and isolation section.

[0059] The reverse design method of supersonic inlet based on complex three-dimensional internal flow field mainly includes the following steps:

[0060] Step 1: Figure 1 As shown, an incident shock wave 1 and a reflected shock wave 2 of a complex three-dimensional internal flow are specified according to design requirements. The incident shock wave 1 of the complex three-dimensional internal flow is specified by using functions such as parametric surfaces or Bezier surfaces. At the shock wave reflection position 4, a shock wave growth rate of the reflected shock wave 2 of the complex three-dimensional internal flow along the incoming flow direction is specified.

[0061] Step 2: Solve the zero-order and first-order flow parameters after the three-dimensional non-axisymmetric incident shock wave. Specifically, the flow parameters of the vertical shock wave surface are solved through the flow field relationship, and the flow parameters of the extended shock wave surface are solved through the shock wave relationship.

[0062] Furthermore, the specific control equations for solving the flow parameters using the flow field relationship are as follows:

[0063] Mass equation:

[0064] Momentum equation:

[0065] Energy equation:

[0066]

[0067] Equation of state:

[0068]

[0069] in, p For pressure, ρ is the density, T is the temperature, m and l are two orthogonal directions along the shock wave surface, n is the normal direction of the shock wave surface, i To represent the parameters, in order to make the state equation neat, the state equations of the same form for partial derivatives in different directions are represented by the parameters i In short, V m and V l is the velocity along the shock wave surface, V n is the velocity perpendicular to the shock wave surface, γ is the specific heat ratio, R is the gas constant, K nm The shock wave surface along the curve m The normal curvature of the direction, K nl The shock wave surface along the curve l The normal curvature of the direction, K gl is the curve on the shock wave surface l The geodesic curvature of K gm is the curve on the shock wave surface m The geodesic curvature of τ gm The curve on the shock wave surface m The geodesic torsion.

[0070] The specific control equations for solving the flow parameters using the shock wave relationship are as follows:

[0071] Mass equation:

[0072] Momentum equation:

[0073] Velocity equation:

[0074]

[0075] Energy equation:

[0076]

[0077] In the above formula, p 1 is the pressure before the shock wave, p 2 is the pressure after the shock wave, ρ 1 is the shock front density, ρ 2 is the density after the shock wave, m and l are two orthogonal directions along the shock wave surface, n is the normal direction of the shock wave surface, i To represent the parameters, in order to make the above equations neat, the equations of the same form for partial derivatives in different directions are represented by the parameters i In short, V 1m Shock front m The speed of the direction, V 2m After shock wave m The speed of the direction, V 1n Shock front n The speed of the direction, V 2n The shock wave trailing edge n The speed of the direction, γ is the specific heat ratio.

[0078] Step 3: Figure 1-Figure 3 As shown, the three-dimensional internal flow field composed of the incident shock wave 1 of the complex three-dimensional internal flow and the reflected shock wave 2 of the complex three-dimensional internal flow is discretized into a series of coaxial two-dimensional reference surfaces 3, and the incident shock wave 5 in the reference surface is obtained. The first-order flow parameters and velocities obtained in step 2 are projected to the direction along the two-dimensional reference surface 3 and the direction perpendicular to the two-dimensional reference surface 3.

[0079] Step 4: Figure 1-Figure 3 As shown, in each two-dimensional reference surface 3, the initial values ​​of the zero-order and first-order flow parameters in the internal flow field are solved by using the pseudo streamlines 6 and the pseudo characteristic lines 7 in the reference surface.

[0080] Step 5: Figure 3 As shown, the initial flow parameters in each two-dimensional reference surface 3 are iterated laterally, and finally the entire three-dimensional internal flow field is reconstructed by reverse solution.

[0081] The control equations for solving the initial values ​​of the above parameters and lateral iterations are as follows:

[0082]

[0083]

[0084] In the above formula, s and l are the characteristic line directions in the two-dimensional reference surface, φ for , P and D is the pressure gradient in the two-dimensional reference surface, W is the pressure gradient perpendicular to the two-dimensional reference surface, V is the velocity vector in the two-dimensional reference plane, w is the velocity vector perpendicular to the two-dimensional reference surface, p For pressure, ρ is the density, δ is the airflow deflection angle, μ is the Mach angle. It should be noted that in the process of solving the initial value of this parameter and lateral iteration, the operation steps are to first assume that all lateral values ​​are zero, then perform the initial solution, and then perform iterative update after the initial solution.

[0085] Step 6: Figure 4 As shown, given the required leading edge profile 8 of the air inlet, the streamline is used to trace the process line 9 to obtain the compression profile 10 of the air inlet that meets the requirements, and the tail end of the compression profile is stretched along the incoming flow direction to obtain the isolation section 11 of the air inlet, and the following is obtained: Figure 5 The air intake is shown.

[0086] Compared with the traditional flow field solution method, the above steps 2 to 4 can not only realize the solution of complex three-dimensional internal flow field, but also obtain the first-order flow parameters in the flow field. On the basis of realizing the reverse solution, the calculation accuracy is improved and the solution efficiency of the entire three-dimensional flow field is accelerated, so that the obtained flow field is no longer limited to the traditional two-dimensional axisymmetric internal flow field. Therefore, more and more complex three-dimensional internal flow fields can be obtained through more free shock wave selection, and more different forms of inlet ducts can be designed to meet aerodynamic requirements.

[0087] The above solution is only an illustration of a preferred embodiment, but is not limited thereto. When implementing the present invention, appropriate replacement and / or modification can be performed according to user needs.

[0088] Although the embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and the embodiments. It can be fully applied to various fields suitable for the present invention. For those familiar with the art, additional modifications can be easily realized. Therefore, without departing from the general concept defined by the claims and equivalent scope, the present invention is not limited to the specific details and the illustrations shown and described here.

Claims

1. A supersonic inlet reverse design method based on a complex three-dimensional internal flow field, characterized in that: include: S1. Specify the incident shock wave and reflected shock wave of the complex three-dimensional internal flow field according to the design requirements, and solve the zero-order and first-order flow parameters after the three-dimensional non-axisymmetric incident shock wave; S2, discretizing the three-dimensional internal flow field formed by the incident shock wave and the reflected shock wave into a series of coaxial two-dimensional reference surfaces; S3, decomposing the zero-order and first-order flow parameters obtained in S1 into quantities along the two-dimensional reference surface and perpendicular to the two-dimensional reference surface; S4, based on the quantities obtained in S3, solving the initial values ​​of the zero-order and first-order flow parameters in the internal flow field in each reference surface, performing lateral iteration on the initial values ​​of the parameters in each two-dimensional reference surface, so as to reconstruct the entire three-dimensional internal flow field by reverse solution; S5. Specify the leading edge profile FCT of the inlet on the incident shock wave surface, obtain the three-dimensional compression profile of the inlet by using streamline tracing, and stretch the tail end of the three-dimensional compression profile along the incoming flow direction to obtain the isolation section of the inlet; In S1, the incident shock wave is directly given by a parametric surface or a Bezier surface function, and the reflected shock wave is given by a shock wave angular growth rate of the reflected shock wave; In S1, when solving the zero-order and first-order flow parameters after the three-dimensional non-axisymmetric incident shock wave, the flow parameters of the vertical shock wave surface are solved by the flow field relationship, and the flow parameters of the extended shock wave surface are solved by the shock wave relationship; In S4, the governing equations for solving the initial values ​​of parameters and lateral iterations include: In the above formula, s and l are the characteristic line directions in the two-dimensional reference surface, φ for , P and D is the pressure gradient in the two-dimensional reference surface, W is the pressure gradient perpendicular to the two-dimensional reference surface, V is the velocity vector in the two-dimensional reference plane, w is the velocity vector perpendicular to the two-dimensional reference surface, p For pressure, ρ is the density, δ is the airflow deflection angle, μ is the Mach angle.

2. The method for reverse design of a supersonic inlet based on a complex three-dimensional internal flow field according to claim 1, characterized in that: The governing equations in the flow field relationship include: Mass equation I: Momentum equation I: Energy equation I: Equation of state: In the above formula, p For pressure, ρ is the density, T is the temperature, m and l are two orthogonal directions along the shock wave surface, n is the normal direction of the shock wave surface, i To refer to the parameters, V m and V l is the velocity along the shock wave surface, V n is the velocity perpendicular to the shock wave surface, γ is the specific heat ratio, R is the gas constant, K nm The shock wave surface along the curve m The normal curvature of the direction, K nl The shock wave surface along the curve l The normal curvature of the direction, K gl is the curve on the shock wave surface l The geodesic curvature of K gm is the curve on the shock wave surface m The geodesic curvature of τ gm The curve on the shock wave surface m The geodesic torsion.

3. The method for reverse design of a supersonic inlet based on a complex three-dimensional internal flow field according to claim 1, characterized in that: The governing equations in the shock wave relationship include: Mass equation II: Momentum equation II: Velocity equation II: Energy equation II: In the above formula, p 1 is the pressure before the shock wave, p 2 is the pressure after the shock wave, ρ 1 is the shock front density, ρ 2 is the density after the shock wave, m and l are two orthogonal directions along the shock wave surface, n is the normal direction of the shock wave surface, i To refer to the parameters, V 1m Shock front m The speed of the direction, V 2m After shock wave m The speed of the direction, V 1n Shock front n The speed of the direction, V 2n The shock wave trailing edge n The speed of the direction, γ is the specific heat ratio.

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