CN116341421B - Hypersonic flow field numerical simulation method, hypersonic flow field numerical simulation system, electronic equipment and storage medium - Google Patents
Hypersonic flow field numerical simulation method, hypersonic flow field numerical simulation system, electronic equipment and storage medium Download PDFInfo
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Abstract
The application discloses a hypersonic flow field numerical simulation method, a hypersonic flow field numerical simulation system, electronic equipment and a storage medium, and belongs to the technical field of aerodynamics and numerical simulation technologies. The hypersonic flow field numerical simulation method comprises the following steps: building a hypersonic flow field numerical simulation frame of the aircraft; acquiring a grid micro-element interface of the hypersonic flow field numerical simulation framework; judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method; and determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity. The hypersonic flow field numerical simulation method can improve the accuracy of hypersonic flow field numerical simulation.
Description
Technical Field
The application relates to the field of aerodynamics and numerical simulation, in particular to a hypersonic flow field numerical simulation method, a hypersonic flow field numerical simulation system, electronic equipment and a storage medium.
Background
The hypersonic flow research can lay a foundation for the research and development of hypersonic aircrafts such as aeroplanes, carriers and the like. Compared with subsonic, transonic and supersonic flow, hypersonic flow has a plurality of sharp characteristics, strong shock wave interruption exists in a flow field, complex physical phenomena such as chemical reaction, thermodynamic excitation and the like occur in high-temperature gas after a wave, a high-temperature gas unbalance effect occurs, and the aerodynamic characteristics of a hypersonic aircraft are seriously influenced. Numerical simulation is one of the most dominant research approaches due to ground test conditions and equipment limitations, and the recurrence of many high-flow complex physical phenomena.
The numerical format has a great influence on the stability and accuracy of the hypersonic aircraft flow field numerical simulation. The AUSM format gives consideration to the high precision of flux differential FDS and the robustness of flux split FVS, has the advantages of small value dissipation, high contact break resolution, strong shock wave capturing performance and the like, and is widely applied to hypersonic flow field value simulation. In the AUSM class format, the interface sound speedDirectly affecting the resolution of the physical discontinuity capture.
Therefore, how to improve the accuracy of hypersonic flow field numerical simulation is a technical problem that a person skilled in the art needs to solve at present.
Disclosure of Invention
The application aims to provide a hypersonic flow field numerical simulation method, a hypersonic flow field numerical simulation system, electronic equipment and a storage medium, which can improve the accuracy of hypersonic flow field numerical simulation.
In order to solve the technical problems, the application provides a hypersonic flow field numerical simulation method, which comprises the following steps:
building a hypersonic flow field numerical simulation frame of the aircraft;
acquiring a grid micro-element interface of the hypersonic flow field numerical simulation framework;
judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method;
and determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity.
Optionally, the hypersonic flow field numerical simulation framework of the building aircraft comprises:
determining a numerical calculation grid of the hypersonic flow field according to the appearance of the aircraft;
and on the numerical calculation grid, adopting an AUSM type format discrete flow control equation set to obtain the hypersonic flow field numerical simulation framework.
Optionally, determining whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity includes:
judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity according to the left and right limit parameters of the grid infinitesimal interface;
the left and right limit parameters comprise a left and right limit value of the normal direction speed of the interface, a left and right limit value of the pressure of the interface and a left and right limit value of the sound speed of the interface.
Optionally, judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity according to the left and right limit parameters of the grid infinitesimal interface includes:
judging whether the left and right limit parameters of the grid infinitesimal interface accord with a first constraint condition and a second constraint condition;
if yes, judging the flow type of the grid infinitesimal interface to be the Jiang Jibo discontinuity;
if not, judging the flow type of the grid infinitesimal interface to be a relatively gentle area;
wherein the first constraint condition is thatThe method comprises the steps of carrying out a first treatment on the surface of the The second constraint condition is that，/>Indicating the left limit value of the interface normal direction speed, +.>Right limit value of speed representing interface normal direction, +.>Indicating the left limit value of the interface pressure +.>Indicating the right limit value of the interface pressure +.>Represents the interface sound speed left limit,/" >Represents the right limit value of interface sound velocity,/>Indicating a threshold mach number.
Optionally, before calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface, the method further includes:
calculating critical sound velocity of the grid infinitesimal interface by using a first formula;
wherein the first formula is，/>Represents critical sound velocity, ++>Indicating the left limit value of the interface normal direction speed, +.>Indicating the right limit value of the interface normal direction velocity.
Optionally, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface includes:
calculating the interface sound velocity of the grid infinitesimal interface by using a second formula;
wherein the second formula is；/>Represents interface sound speed, ++>Represents critical sound velocity, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
Optionally, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method includes:
calculating the interface sound velocity of the grid infinitesimal interface by using a third formula;
wherein the third formula is；/>Represents interface sound speed, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
The application also provides a hypersonic flow field numerical simulation system, which comprises:
The frame building module is used for building a hypersonic flow field numerical simulation frame of the aircraft;
the interface acquisition module is used for acquiring a grid infinitesimal interface of the hypersonic flow field numerical simulation framework;
the type judging module is used for judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method;
and the numerical simulation module is used for determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity.
The application also provides a storage medium, on which a computer program is stored, which when executed realizes the steps executed by the hypersonic flow field numerical simulation method.
The application also provides electronic equipment, which comprises a memory and a processor, wherein the memory stores a computer program, and the processor realizes the steps executed by the hypersonic flow field numerical simulation method when calling the computer program in the memory.
The application provides a hypersonic flow field numerical simulation method, which comprises the following steps: building a hypersonic flow field numerical simulation frame of the aircraft; acquiring a grid micro-element interface of the hypersonic flow field numerical simulation framework; judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method; and determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity.
According to the method, the grid infinitesimal interface is obtained from the hypersonic flow field numerical simulation framework of the aircraft, and the flow type of the grid infinitesimal interface is judged. If the flow type of the grid infinitesimal interface is a strong shock wave discontinuity, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface, wherein the shock wave discontinuity without numerical value dissipation can be accurately captured in one grid in the mode, the resolution is high, and the simulation accuracy is ensured; if the flow type of the grid infinitesimal interface is not a strong shock wave discontinuity, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method, wherein the interface sound velocity is built by the arithmetic average method, the difference between left and right limit values is balanced, and the method is simple and convenient to calculate and has good robustness and grid applicability; and finally, determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity. According to the method, the corresponding interface sound velocity determining method is selected according to the flow type of the grid infinitesimal interface, and the accuracy of hypersonic flow field numerical simulation can be improved. The application also provides a hypersonic flow field numerical simulation system, a storage medium and an electronic device, which have the beneficial effects and are not repeated here.
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For a clearer description of embodiments of the present application, the drawings that are required to be used in the embodiments will be briefly described, it being apparent that the drawings in the following description are only some embodiments of the present application, and other drawings may be obtained according to the drawings without inventive effort for those skilled in the art.
FIG. 1 is a flow chart of a hypersonic flow field numerical simulation method for interface sound velocity distinguishing processing provided by an embodiment of the application;
FIG. 2 is a schematic diagram illustrating an exploded view of the oblique shock plane velocity according to an embodiment of the present application;
FIG. 3 is a flow chart of a robust and accurate simulation of hypersonic flow provided by an embodiment of the present application;
FIG. 4 is a graph showing a comparison of the pressure distribution of an electric blunt cone flow field according to an embodiment of the present application;
FIG. 5 is a schematic diagram showing a converging curve of heat flow at a stagnation point of an electric blunt cone according to an embodiment of the present application;
fig. 6 is a schematic structural diagram of a hypersonic flow field numerical simulation system for interface sound velocity distinguishing processing according to an embodiment of the present application.
Detailed Description
For the purpose of making the objects, technical solutions and advantages of the embodiments of the present application more apparent, the technical solutions of the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present application, and it is apparent that the described embodiments are some embodiments of the present application, but not all embodiments of the present application. All other embodiments, which can be made by those skilled in the art based on the embodiments of the application without making any inventive effort, are intended to be within the scope of the application.
Referring to fig. 1, fig. 1 is a flowchart of a hypersonic flow field numerical simulation method for interface sound velocity distinguishing processing according to an embodiment of the present application.
The specific steps may include:
s101: building a hypersonic flow field numerical simulation frame of the aircraft;
the embodiment can be applied to electronic equipment with an aerodynamic analysis function, and a hypersonic flow field numerical simulation frame of an aircraft can be built in the following manner: and determining a numerical calculation grid of the hypersonic flow field according to the appearance of the aircraft, and obtaining the hypersonic flow field numerical simulation framework by adopting an AUSM (Advection Upstream Splitting Method ) type format discrete flow control equation set on the numerical calculation grid.
S102: acquiring a grid micro-element interface of the hypersonic flow field numerical simulation framework;
in the process of numerical simulation, the calculated flow field needs to be divided into small units, called grid units or grid micro-elements, and then the interface between each micro-element is called a grid micro-element interface. A plurality of grid microcell interfaces exist in the hypersonic flow field numerical simulation framework.
S103: judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method;
in the high-speed flight process of the hypersonic aircraft, a shock wave phenomenon occurs in a surrounding flow field of the aircraft, gas state parameters such as flow field pressure, density, temperature and the like are discontinuous when crossing shock waves, jump change exists, and the shock wave can be generally called shock wave interruption, and if the shock wave is strong, the shock wave can be called strong shock wave interruption; in the numerical calculation process, if a grid infinitesimal interface simulates a strong shock wave discontinuous phenomenon in a flow field, the embodiment refers to the grid infinitesimal interface as a strong shock wave discontinuous surface. The embodiment can set the grid infinitesimal interface with the left and right limit parameters meeting the preset constraint conditions as a strong shock wave discontinuity.
In this embodiment, the grid infinitesimal interface is divided into a strong shock wave discontinuity and a relatively gentle region according to the flow type, and all grid infinitesimal interfaces that are not strong shock wave discontinuities are relatively gentle regions.
If the grid infinitesimal interface is a strong shock wave discontinuity, calculating the critical sound velocity of the grid infinitesimal interface, and calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity. If the grid infinitesimal interface is a relatively gentle region, calculating the interface sound velocity of the grid infinitesimal interface through an arithmetic average method.
S104: and determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity.
The numerical simulation results generated according to the interface sound velocity include, but are not limited to, time or space distribution of hypersonic flow parameters, influence rules of hypersonic flow phenomena, aerodynamic characteristics of hypersonic aircrafts, aerodynamic heat environment, flight trajectories, control strategies, system design schemes and the like, so that aerodynamic characteristics of the aircrafts can be evaluated according to the numerical simulation results.
In the embodiment, a grid infinitesimal interface is obtained from a hypersonic flow field numerical simulation frame of the aircraft, and the flow type of the grid infinitesimal interface is judged. If the flow type of the grid infinitesimal interface is a strong shock wave discontinuity, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface, wherein the shock wave discontinuity without numerical value dissipation can be accurately captured in one grid in the mode, the resolution is high, and the simulation accuracy is ensured; if the flow type of the grid infinitesimal interface is not a strong shock wave discontinuity, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method, wherein the interface sound velocity is built by the arithmetic average method, the difference between left and right limit values is balanced, and the method is simple and convenient to calculate and has good robustness and grid applicability; and finally, determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity. According to the method, the corresponding interface sound velocity determining method is selected according to the flow type of the grid infinitesimal interface, and the accuracy of hypersonic flow field numerical simulation can be improved.
As a further introduction to the corresponding embodiment of FIG. 1, it may be determined whether the flow type of the grid-hogel interface is a strong shock discontinuity in the following manner: judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity according to the left and right limit parameters of the grid infinitesimal interface; the left and right limit parameters comprise a left and right limit value of the normal direction speed of the interface, a left and right limit value of the pressure of the interface and a left and right limit value of the sound speed of the interface.
Specifically, the process of judging the flow type of the grid infinitesimal interface according to the left and right limit parameters of the grid infinitesimal interface includes: judging whether the left and right limit parameters of the grid infinitesimal interface accord with a first constraint condition and a second constraint condition; if yes, judging the flow type of the grid infinitesimal interface to be the Jiang Jibo discontinuity; if not, judging the flow type of the grid infinitesimal interface to be a relatively gentle area; wherein the first constraint condition is thatThe method comprises the steps of carrying out a first treatment on the surface of the Said second constraint is +.>，/>Indicating the left limit value of the interface normal direction speed, +.>Right limit value of speed representing interface normal direction, +.>Indicating the left limit value of the interface pressure,indicating the right limit value of the interface pressure +. >Represents the interface sound speed left limit,/">Represents the right limit value of interface sound velocity,/>Indicating a threshold mach number.
As a further introduction to the corresponding embodiment of fig. 1, the critical sound velocity of the mesh infinitesimal interface may also be calculated using a first formula before calculating the interface sound velocity of the mesh infinitesimal interface from the critical sound velocity of the mesh infinitesimal interface; wherein the first formula is，/>Represents critical sound velocity, ++>Indicating the left limit value of the interface normal direction speed, +.>Indicating the right limit value of the interface normal direction velocity.
Further, if the flow type of the grid infinitesimal interface is a strong shock wave discontinuity, the process of calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface includes: calculating the interface sound velocity of the grid infinitesimal interface by using a second formula;
wherein the second formula is；/>Represents interface sound speed, ++>Represents critical sound velocity, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
Further, if the flow type of the grid infinitesimal interface is a relatively gentle region, the process of calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method includes: calculating the interface sound velocity of the grid infinitesimal interface by using a third formula;
Wherein the third formula is；/>Represents interface sound speed, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
The flow described in the above embodiment is explained below by way of an embodiment in practical application.
Numerical format vs. high supernumerical simulation stabilityAnd the influence of the accuracy is large. The AUSM format gives consideration to the high precision of flux differential FDS and the robustness of flux split FVS, has the advantages of small value dissipation, high contact break resolution, strong shock wave capturing performance and the like, and is widely applied to hypersonic flow field value simulation. In the AUSM class format, the interface sound speedDirectly affecting the resolution of the physical discontinuity capture. The three common treatments are as follows:
1) Arithmetic mean method. The stability is good, the format has better robustness and grid applicability, and particularly, the method can balance the difference of the left limit value and the right limit value for a gentle region with smaller parameter gradient, and is widely applicable to various gas models such as complete gas, chemical unbalanced gas, thermochemical unbalanced gas and the like. However, for a strong shock intermittent interface, the value dissipation is relatively large, the shock wave cannot be captured in a grid, and the resolution is relatively low.
2) And (5) a maximum value method. A high-precision, stable forward shock can be captured in one cell, but a ramp shock without numerical dissipation cannot be captured.
3) A calculation method based on critical sound velocity. The shock wave discontinuous pluronic relation is satisfied, and the numerical format accurately captures shock wave discontinuities without numerical dissipation in a grid, whether the shock waves are normal shock waves or oblique shock waves, so that the resolution is higher. However, the critical sound velocity of the method is calculated by adopting the gas enthalpy value and the gas specific heat ratio, is derived from the complete gas assumption, and cannot be directly applied to Gao Chaogao temperature unbalanced flow simulation. For unbalanced flow, not only the gas components with the left and right limits and the thermodynamic temperature (translational, rotational, vibration, electron and other temperature modes) are different, but also the specific heat ratio of the gas is different, and the formation enthalpy of the gas components may be negative, so that the interface gas enthalpy value may be negative, and therefore, the critical sound velocity calculation method based on the complete gas assumption is not applicable any more. For example, the existing interface sound velocity calculation method for hypersonic high temperature unbalanced flow simulation is easy to operate the negative number, so that calculation failure is caused, and calculation stability and applicability are affected.
In order to further develop a more robust and accurate hypersonic flow numerical simulation method based on an AUSM type format, the application provides a hypersonic flow robust and accurate simulation scheme, which is used for constructing a hypersonic flow simulation framework from an AUSM type numerical discrete format and judging a strong shock break area and a relatively gentle area according to interface parameter characteristics; in a relatively gentle region, an arithmetic average method is adopted to construct interface sound velocity, balance left and right limit value differences, and the method is simple and convenient to calculate and has better robustness and grid applicability; in a strong shock wave discontinuous region, an interface sound velocity is fitted by adopting a method based on a critical sound velocity, so that shock wave discontinuous without numerical value dissipation can be accurately captured in a grid, the resolution is high, and the simulation accuracy is ensured; the critical sound velocity can be calculated based on the shock wave discontinuous pluronic relation, so that the calculation is simple and convenient, and the problems of ' difference in limit value between left and right specific heat ratios ', ' negative component enthalpy ', negative component evolution ' and the like possibly occurring in the traditional method (the method for calculating the gas enthalpy value and the gas specific heat ratio) during the high-unbalance flow simulation are avoided, so that the method has higher stability and universality.
The specific implementation manner of the embodiment is as follows:
step 1: and (5) constructing a hypersonic flow numerical simulation framework. And importing a numerical calculation grid, and constructing a numerical iteration solution basic framework by adopting an AUSM type format discrete flow control equation set.
The numerical iteration solution flow control equation set is one of the most common methods for hypersonic flow numerical simulation, and the basic principle and the implementation method of the iteration solution framework are described in detail in a plurality of publications, and are not repeated here, but only the technical content related to the invention is described.
The computational mesh involved in this embodiment includes, but is not limited to, a common hypersonic flow simulation mesh such as a structural mesh, an unstructured mesh, or a structure-unstructured hybrid mesh. The flow control equation set is a conservation type N-S equation or Euler equation, which is the most commonly used control equation form for hypersonic flow simulation. The flow control equation sets are discrete in an AUSM class format, where the AUSM class format includes, but is not limited to, AUSM, AUSMD, AUSMDV, AUSM +, AUSM+ -up, AUSMPW, AUSMPW +, M-AUSMPW+, and the like, which employ numerical discrete formats of interfacial sound velocity. The AUSM format has the advantages of low numerical dissipation, high contact break resolution, high shock wave capturing performance and the like, and is widely applied to hypersonic flow field numerical simulation because the AUSM format has the advantages of high precision of flux differential FDS and robustness of flux split FVS.
Step 2: and judging the flow type of the grid micro-element interface. In the iterative process, based on the left and right limit parameters on the grid infinitesimal interface, judging whether the interface is a strong shock wave discontinuity. If the shock wave is a strong shock wave discontinuity, executing the step 3; if the shock wave is not a strong shock wave discontinuity, the method is a relatively gentle area, and step 4 is executed.
There are many ways to determine the strong shock discontinuities, and one of the more efficient ways is described herein as an example: based on the interface normal speed and the left and right limiting values of the pressure intensity, judging:
if it isAnd->Step 3 is executed if the shock wave is a strong shock wave discontinuity; otherwise, the step 4 is executed if the shock wave discontinuities are not shock wave discontinuities, i.e. relatively gentle areas.
Here, theAnd->Limiting value of interface normal direction speed, +.>And->Limiting the interface pressure to a left limit value and a right limit value; />Is a threshold Mach number. In the high-speed flight process of the hypersonic aircraft, a shock wave phenomenon can occur in a surrounding flow field around the aircraft, flow field parameters such as flow field pressure, density, temperature, speed and the like are discontinuous when crossing shock waves, jump change exists, and the shock wave can be generally called shock wave interruption, and if the shock wave is strong, the shock wave can be called strong shock wave interruption; since the strong shock interruption does not have strict judgment standard at present, the embodiment adopts +. >As a threshold parameter for strong shock discontinuities, when the normal incoming flow Mach number at the shock discontinuity is higher than +>The discontinuities are, in turn, strong shock discontinuities. />The value of (2) is not fixed and can be properly adjusted according to the actual working condition, wherein->。
According to the normal shock theory, hypersonic flow is obtained before shock wave, and subsonic flow is obtained after shock wave. For oblique shock waves, the speed of the oblique shock waves is decomposed on the shock wave surface, and the speed jumps in the normal direction of an interface. Referring to fig. 2, fig. 2 is a schematic diagram illustrating a ramp surface velocity decomposition according to an embodiment of the present application, where a ramp (lattice voxel interface) in fig. 2 represents an Oblique clock (Cell-interface),is the interface left limit speed vector,/>Is the interface right limit speed vector,/>For the left limit speed vector +.>Interface normal component, ++>For the left limit speed vector +.>Is used to determine the tangential component of (a),for right limit speed vector +.>Interface normal component, ++>For right limit speed vector +.>Tangential component of>Representing the vector included angle. Normal speed->And->The characteristics of forward shock Mach number change are met, supersonic flow is carried out before shock wave, subsonic flow is carried out after shock wave, and therefore:
；
here, theAnd- >The limiting sound velocity is about the interface, but this condition is only a necessary condition for shock break, and not a sufficient condition. In the non-shock intermittent flow field region, progressive crossing between supersonic velocity and subsonic velocity, such as Laval nozzle throat, can also occurAnd the flow of the ultrasonic velocity, etc., so that the judgment is assisted by matching with other conditions.
As the supersonic velocity air flow is subjected to the excitation, the air pressure, temperature and density rise, namely:
；
here, the、/>And->Is the pressure, density and temperature after shock wave, +.>、/>And->Is shock wave front pressure, density and temperature. As the gas state equation, it can be seen that:
；
and due to、/>Thus, there are:
；
；
it can be seen that at shock discontinuities, the "jump" amplitude of pressure is greater than the "jump" amplitude of density or temperature, with more significant variation. Therefore, the shock break can be judged in an auxiliary way by using the variation amplitude of the pressure. For the calorimetric complete gas, the parameters before and after shock wave meet the relationship of Rankine-Hugoniot, and the parameters can be obtained through transformation:
；
here, theIs the ratio of specific heat of gas. For normal shock wave>Is the shock wave front Mach number; for oblique shock waves>Mach number dominant for the normal velocity of the shock wave front interface. The specific heat ratio of the gas is satisfied whether the gas is completely gas or non-equilibrium gas >Thus, there are:
；
when the shock is strong, i.e. the strong shock is interruptedWhen the method is used, the following steps are included:
；
therefore, the left and right limit values for the strong shock interface pressureAnd->The following relationship needs to be satisfied:
；
the present embodiment uses the above relationship as an auxiliary judgment of the strong shock break.
Step 3: and (5) processing the sound velocity of the interface of the strong shock wave discontinuities. In the iterative process, the interface sound velocity is calculated based on the critical sound velocity method. After the step is completed, the process goes to step 5;
there are many methods for calculating the critical sound velocity, and one of them, which is more efficient and robust, is described here as an example: based on the pluronic relation on the shock wave discontinuities, calculating a critical sound velocity to obtain an interface sound velocity:
；
；
here, theIs critical sound velocity +.>And->Respectively the left and right limit sound velocity of the interface, +.>Is the interfacial sound velocity.
In the AUSM format, the critical sound velocity is calculated by a plurality of methods, for example, the critical sound velocity can be calculated based on the gas enthalpy value and the gas specific heat ratio, and the specific process is as follows:
the method comprises the following steps:
；
；
the second method is as follows:
；
here, theFor the removal of the interfacial average enthalpy after tangential kinetic energy, < >>And->Total enthalpy of left and right limit of interface, +.>Is the average specific heat ratio of gas +.>And->Is the specific heat ratio of the limiting gas at the left and right sides of the interface. Both of these methods have certain limitations: the method I is obtained by deducting the calorimetric complete gas, is only suitable for complete gas simulation, and is necessary to calculate the gas enthalpy value, and an additional gas enthalpy value calculation process is required to be introduced for solving the Euler equation set which does not contain an energy equation; the second method is theoretically suitable for unbalanced gas simulation by considering the difference of specific heat ratios of left and right interfaces, which is caused by chemical reaction equivalence, but is easy to fail in prescription operation and relatively poor in simulation stability because the problems of negative formation enthalpy possibly existing in gas components, negative value (namely negative difference) possibly occurring due to subtraction of left and right interface parameters and the like are not considered.
The embodiment provides a more efficient and robust calculation method based on the pluronic relation, and theoretical basis thereof is introduced as follows: from the shock theory, it is known that at the shock break, the interface normal velocity "jump" satisfies the pluronic relationship:
；
the critical sound speed can therefore be written directly as:
；
compared with the traditional method, the method is much simpler and more convenient, avoids the problems of ' extra calculation of enthalpy value possibly needed by Euler equation solution ', ' difference of interface specific heat ratio limit value ', ' negative formation enthalpy of component ', negative difference ' and the like, is suitable for complete gas, is suitable for unbalanced gas, and is more stable and has good applicability.
From the definition of the critical sound velocity, it is known that:
；
；
here, theAnd->Front and back velocities, +.>And->The wave front and wave back sound speeds, respectively. According to the speed characteristic of shock break, the wave front is supersonic flow, the wave back is subsonic flow, namely +.>、/>Thus, there are:
；
；
and due toThus, there are:
；
it can be seen that the critical sound speed of the shock wave should be between the wave front and wave back sound speeds. Therefore, the embodiment further limits the interface sound velocity, avoids the deviation of the interface sound velocity which is too large or too small due to numerical errors and the like, and enhances the stability of the interface sound velocity;
；
Step 4: relatively gentle region interface sound speed processing. In the iterative process, based on an arithmetic average method, the difference between the left limit value and the right limit value is balanced, and the interface sound velocity of a relatively gentle region is calculated:
；
after this step, the process proceeds to step 5.
The arithmetic average method has simple and convenient calculation and good stability, can ensure that the format has better robustness and grid applicability, can balance the difference of left and right limit values especially for gentle areas with smaller parameter gradients, and is widely applicable to various gas models such as complete gas, chemical unbalanced gas, thermochemical unbalanced gas and the like. Therefore, the present embodiment directly applies the arithmetic average method in a relatively gentle region.
Step 5: and (3) completing the numerical iteration of the flow equation set, and giving out a simulation result according to the numerical simulation requirement.
The flow control equation set numerical iterative process and the numerical simulation result processing specific method are discussed or described in detail in a lot of publications or data, and are not described in detail herein, so that the present invention can be well applied to these general methods.
The numerical simulation results provided by the embodiment include, but are not limited to, time or space distribution of hypersonic flow parameters, influence rules of hypersonic flow phenomena, aerodynamic characteristics of hypersonic aircrafts, aerodynamic heat environment, flight trajectories, control strategies, system design schemes and the like, and can provide key technical support for design and evaluation of hypersonic aircrafts.
The advantages of this embodiment are: (1) The hypersonic flow simulation framework is constructed from an AUSM class numerical discrete format, and the hypersonic flow simulation framework has the advantages that the AUSM class numerical discrete format has the advantages of high precision of flux difference FDS, robustness of flux splitting FVS, small numerical dissipation, high contact break resolution, high shock wave capturing performance and the like; (2) The judging process of the grid infinitesimal interface flow characteristics comprehensively considers the speed characteristics and the pressure characteristics of shock wave discontinuities, is simple and convenient to calculate, and can rapidly and accurately capture shock wave discontinuities; (3) The embodiment not only considers the normal shock wave break-up characteristic, but also considers the speed decomposition of the oblique shock wave and the break-up characteristic thereof, so that the oblique shock wave with the same high resolution as the normal shock wave can be captured by using a common grid; (4) The interface sound velocity is fitted by comprehensively using an arithmetic average method and a critical sound velocity method, so that the advantages of the arithmetic average method and the critical sound velocity method are reserved, and the defects of the arithmetic average method and the critical sound velocity method are avoided; (5) The critical sound velocity is calculated by adopting the shock wave discontinuous pluronic relation, so that the method is simple and convenient to calculate, avoids the defects of the traditional method, and has better stability and universality; (6) the application scope of the embodiment is wide: suitable computing grids include, but are not limited to, various common hypersonic flow simulation grids such as structural grids, unstructured grids or structure-unstructured hybrid grids; suitable flowing media include, but are not limited to, common hyperflowing media such as the earth's atmosphere, mars atmosphere, various high temperature gases, and the like; suitable gas models include, but are not limited to, complete gas models, chemically unbalanced gas models, thermochemical unbalanced gas models, and other types of common high-flow gas models.
Referring to fig. 3, fig. 3 is a flow chart of robust and accurate simulation of hypersonic flow provided by the embodiment of the application, and the specific process is as follows: and (3) constructing a numerical simulation frame, judging the type of the grid infinitesimal interface, processing a strong shock wave intermittent interface, and processing a relatively gentle area interface to finish simulation output results.
Taking the numerical simulation of the super electric blunt cone model as an example, the application effect of the application is described. The present application may be used in this operating mode, but is not limited to this operating mode.
Calculating working conditions: the total length of the electric blunt cone model is 2m, the radius of the ball head is 0.175m, and the half cone angle is 4.6 degrees. The hypersonic flight conditions for numerical simulation are: mach number 13.0, flight speed 4230m/s, incoming flow density 0.0006944kg/m3, incoming flow pressure 53Pa, incoming flow temperature 265K, and flight attack angle 20 degrees.
The spatial dispersion uses the common AUSMPW+ format, which mimics thermochemical non-equilibrium flow. In order to embody the effect of the application, grids adopted in calculation are sparse near shock break, the quality is relatively poor, and the precise capture difficulty of shock break is increased. The method of this example and the conventional method are compared, respectively.
Referring to fig. 4, fig. 4 is a graph showing a comparison of the pressure distribution of an electric blunt cone flow field according to an embodiment of the present application, where (a) in fig. 4 is the pressure distribution of an electric blunt cone flow field obtained by a conventional method, and (b) is the pressure distribution of an electric blunt cone flow field obtained by the present embodiment, where pressure represents the pressure, and X and Y represent coordinate axes.
Referring to FIG. 5, FIG. 5 is a schematic diagram showing a convergence curve of heat flow at a stationary point of an electric blunt cone according to an embodiment of the present application, wherein the abscissa in FIG. 5 is Iteration number (Iteration) and heat flow intensity (Q), and kW/m 2 Units kilowatts per square meter, representing heat flow intensity, are shownThe reference value of the convergence curve of the electric blunt cone standing point heat flow is obtained by using the traditional method and the convergence curve of the electric blunt cone standing point heat flow obtained by the method.
The flow field pressure distribution calculated by the scheme and the traditional method is shown in fig. 4, and it can be seen that under the grid condition, the pressure distribution calculated by the scheme is smoother, and shock intermittent capturing is more accurate. Fig. 5 shows a convergence curve of standing point heat flow, and it can be seen that compared with the conventional method, the heat flow calculated by the method has better convergence and better accords with the reference value.
Referring to fig. 6, fig. 6 is a schematic structural diagram of a hypersonic flow field numerical simulation system according to an embodiment of the present application, where the system may include:
the frame construction module 601 is used for constructing a hypersonic flow field numerical simulation frame of the aircraft;
the interface acquisition module 602 is configured to acquire a grid infinitesimal interface of the hypersonic flow field numerical simulation framework;
A type judging module 603, configured to judge whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method;
the numerical simulation module 604 is configured to determine a numerical simulation result of the hypersonic flow field according to the interfacial sound velocity.
In the embodiment, a grid infinitesimal interface is obtained from a hypersonic flow field numerical simulation frame of the aircraft, and the flow type of the grid infinitesimal interface is judged. If the flow type of the grid infinitesimal interface is a strong shock wave discontinuity, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface, wherein the shock wave discontinuity without numerical value dissipation can be accurately captured in one grid in the mode, the resolution is high, and the simulation accuracy is ensured; if the flow type of the grid infinitesimal interface is not a strong shock wave discontinuity, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method, wherein the interface sound velocity is built by the arithmetic average method, the difference between left and right limit values is balanced, and the method is simple and convenient to calculate and has good robustness and grid applicability; and finally, determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity. According to the method, the corresponding interface sound velocity determining method is selected according to the flow type of the grid infinitesimal interface, and the accuracy of hypersonic flow field numerical simulation can be improved.
Further, the process of building the hypersonic flow field numerical simulation frame of the aircraft by the frame building module 601 includes: and determining a numerical calculation grid of the hypersonic flow field according to the appearance of the aircraft, and obtaining the hypersonic flow field numerical simulation framework by adopting an AUSM type format discrete flow control equation set on the numerical calculation grid.
Further, the process of determining whether the flow type of the grid infinitesimal interface is a strong shock discontinuity by the type determining module 603 includes: judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity according to the left and right limit parameters of the grid infinitesimal interface; the left and right limit parameters comprise a left and right limit value of the normal direction speed of the interface, a left and right limit value of the pressure of the interface and a left and right limit value of the sound speed of the interface.
Further, the process of the type judging module 603 judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity according to the left and right limit parameters of the grid infinitesimal interface includes: judging whether the left and right limit parameters of the grid infinitesimal interface accord with a first constraint condition and a second constraint condition; if yes, judging the flow type of the grid infinitesimal interface to be the Jiang Jibo discontinuity; if not, judging the flow type of the grid infinitesimal interface to be a relatively gentle area;
Wherein the first constraint condition is thatThe method comprises the steps of carrying out a first treatment on the surface of the The second constraint condition is that，/>Indicating the left limit value of the interface normal direction speed, +.>Right limit value of speed representing interface normal direction, +.>Indicating the left limit value of the interface pressure +.>Indicating the right limit value of the interface pressure +.>Represents the interface sound speed left limit,/">Represents the right limit value of interface sound velocity,/>Indicating a threshold mach number.
Further, the method further comprises the following steps:
the critical sound velocity calculation module is used for calculating the critical sound velocity of the grid infinitesimal interface by utilizing a first formula before calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface;
wherein the first formula is，/>Represents critical sound velocity, ++>Indicating the left limit value of the interface normal direction speed, +.>Indicating the right limit value of the interface normal direction velocity.
Further, the process of calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface by the type judging module 603 includes: calculating the interface sound velocity of the grid infinitesimal interface by using a second formula;
wherein the second formula is；/>Represents interface sound speed, ++>Represents critical sound velocity, ++>Represents the interface sound speed left limit,/" >Indicating the right limit value of the interface sound velocity.
Further, the process of calculating the interface sound velocity of the grid infinitesimal interface by the type judging module 603 through an arithmetic average method includes: calculating the interface sound velocity of the grid infinitesimal interface by using a third formula;
wherein the third formula is；/>Represents interface sound speed, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
Since the embodiments of the system portion and the embodiments of the method portion correspond to each other, the embodiments of the system portion refer to the description of the embodiments of the method portion, which is not repeated herein.
The present application also provides a storage medium having stored thereon a computer program which, when executed, performs the steps provided by the above embodiments. The storage medium may include: a U-disk, a removable hard disk, a Read-Only Memory (ROM), a random access Memory (Random Access Memory, RAM), a magnetic disk, or an optical disk, or other various media capable of storing program codes.
The application also provides an electronic device, which can comprise a memory and a processor, wherein the memory stores a computer program, and the processor can realize the steps provided by the embodiment when calling the computer program in the memory. Of course the electronic device may also include various network interfaces, power supplies, etc.
In the description, each embodiment is described in a progressive manner, and each embodiment is mainly described by the differences from other embodiments, so that the same similar parts among the embodiments are mutually referred. For the system disclosed in the embodiment, since it corresponds to the method disclosed in the embodiment, the description is relatively simple, and the relevant points refer to the description of the method section. It should be noted that it will be apparent to those skilled in the art that various modifications and adaptations of the application can be made without departing from the principles of the application and these modifications and adaptations are intended to be within the scope of the application as defined in the following claims.
It should also be noted that in this specification, relational terms such as first and second, and the like are used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Moreover, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. Without further limitation, an element defined by the phrase "comprising one … …" does not exclude the presence of other like elements in a process, method, article, or apparatus that comprises the element.
Claims (8)
1. A hypersonic flow field numerical simulation method, comprising:
building a hypersonic flow field numerical simulation frame of the aircraft;
acquiring a grid micro-element interface of the hypersonic flow field numerical simulation framework;
judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method;
determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity;
before calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface, the method further comprises the following steps: calculating critical sound velocity of the grid infinitesimal interface by using a first formula;
wherein the first formula is，/>Represents critical sound velocity, ++>Indicating the left limit value of the interface normal direction speed, +.>Indicating the right limit value of the interface normal direction speed;
the method for calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface comprises the following steps:
calculating the interface sound velocity of the grid infinitesimal interface by using a second formula;
Wherein the second formula is；/>Represents interface sound speed, ++>Represents critical sound velocity, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
2. The hypersonic flow field numerical simulation method as set forth in claim 1, wherein the constructing hypersonic flow field numerical simulation framework of the aircraft includes:
determining a numerical calculation grid of the hypersonic flow field according to the appearance of the aircraft;
and on the numerical calculation grid, adopting an AUSM type format discrete flow control equation set to obtain the hypersonic flow field numerical simulation framework.
3. The hypersonic flow field numerical simulation method of claim 1, wherein determining whether the flow type of the grid element interface is a strong shock discontinuity comprises:
judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity according to the left and right limit parameters of the grid infinitesimal interface;
the left and right limit parameters comprise a left and right limit value of the normal direction speed of the interface, a left and right limit value of the pressure of the interface and a left and right limit value of the sound speed of the interface.
4. The hypersonic flow field numerical simulation method of claim 3, wherein judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity according to the left and right limit parameters of the grid infinitesimal interface comprises:
Judging whether the left and right limit parameters of the grid infinitesimal interface accord with a first constraint condition and a second constraint condition;
if yes, judging the flow type of the grid infinitesimal interface to be the Jiang Jibo discontinuity;
if not, judging the flow type of the grid infinitesimal interface to be a relatively gentle area;
wherein the first constraint condition is thatThe method comprises the steps of carrying out a first treatment on the surface of the The second constraint condition is that，/>Indicating the left limit value of the interface normal direction speed, +.>Right limit value of speed representing interface normal direction, +.>Indicating the left limit value of the interface pressure +.>Indicating the right limit value of the interface pressure +.>Represents the interface sound speed left limit,/">Represents the right limit value of interface sound velocity,/>Indicating a threshold mach number.
5. The hypersonic flow field numerical simulation method as set forth in claim 1, wherein calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method includes:
calculating the interface sound velocity of the grid infinitesimal interface by using a third formula;
wherein the third formula is；/>Represents interface sound speed, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
6. A hypersonic flow field numerical simulation system, comprising:
The frame building module is used for building a hypersonic flow field numerical simulation frame of the aircraft;
the interface acquisition module is used for acquiring a grid infinitesimal interface of the hypersonic flow field numerical simulation framework;
the type judging module is used for judging whether the flow type of the grid infinitesimal interface is a strong shock wave discontinuity; if yes, calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface; if not, calculating the interface sound velocity of the grid infinitesimal interface by an arithmetic average method;
the numerical simulation module is used for determining a numerical simulation result of the hypersonic flow field according to the interface sound velocity;
the critical sound velocity calculation module is used for calculating the critical sound velocity of the grid infinitesimal interface by utilizing a first formula before calculating the interface sound velocity of the grid infinitesimal interface according to the critical sound velocity of the grid infinitesimal interface;
wherein the first formula is，/>Represents critical sound velocity, ++>Indicating the left limit value of the interface normal direction speed, +.>Indicating the right limit value of the interface normal direction speed;
the process of calculating the interface sound velocity of the grid infinitesimal interface by the type judging module according to the critical sound velocity of the grid infinitesimal interface comprises the following steps: calculating the interface sound velocity of the grid infinitesimal interface by using a second formula;
Wherein the second formula is；/>Represents interface sound speed, ++>Represents critical sound velocity, ++>Represents the interface sound speed left limit,/">Indicating the right limit value of the interface sound velocity.
7. An electronic device comprising a memory and a processor, wherein the memory stores a computer program, and the processor, when invoking the computer program in the memory, performs the steps of the hypersonic flow field numerical simulation method as claimed in any one of claims 1 to 5.
8. A storage medium having stored therein computer executable instructions which when loaded and executed by a processor perform the steps of the hypersonic flow field numerical simulation method as claimed in any one of claims 1 to 5.
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