CN114330080B - Prediction method for aircraft surface-symmetric cross-basin flow field - Google Patents
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Abstract
The invention discloses a method for predicting a cross-basin flow field with symmetrical aircraft surface, which relates to the field of cross-basin flow field simulation and comprises the following steps: performing flow field solution based on a first physical space grid corresponding to the aircraft and the inflow conditions to obtain speed and temperature information of the flow field; generating a three-dimensional speed space grid, setting the grid range of the three-dimensional speed space grid based on an outer boundary to obtain a three-dimensional hemispherical area, obtaining a hemispherical encryption area and a spherical encryption area, generating a hemispherical speed space grid based on grid spacing distribution information, the hemispherical encryption area and the spherical encryption area, and symmetrically copying the hemispherical speed space grid to obtain a spherical speed space grid; performing iterative solution on the basis of the first physical space grid and the spherical three-dimensional velocity space grid to obtain a three-dimensional physical space flow field of the aircraft; the method realizes the rapid prediction of the aircraft surface-symmetric cross-basin flow field.
Description
Technical Field
The invention relates to the field of cross-basin flow field simulation, in particular to a prediction method of an aircraft surface-symmetric cross-basin flow field.
Background
With the development of adjacent space aircrafts, cross-flow-area flow fields with continuous flow and thin flow coexisting are more and more common, which are generally solved by NS/DSMC zone overlapping or Boltzmann equation at present, but the selection of continuity failure parameters and the judgment process of interfaces are more complicated in the former, and the latter is rapidly developed in recent years due to the fact that the same control equation is adopted in a plurality of flow areas. Xu et al propose a Unified Gas dynamics Scheme (UGKS) that couples particle collision and migration processes and achieves automatic transition of equations between continuous and dilute flows by adjusting relaxation time, and this method is widely used in full-flow-domain full-speed-domain flow prediction, such as multiple physical flows, turbulence simulation, etc. However, the method brings huge calculation cost due to the fact that the method is dispersed in a physical space and a speed space simultaneously.
To improve computational efficiency, many optimization techniques have emerged for velocity space grids, such as cartesian grid adaptation, unstructured grids, and the like. In comparison, the unstructured grid program is simpler to implement, the far-field boundary can be any shape, and the encryption of a designated area is easy to perform, so that the unstructured grid program is favored by many scholars. A common aircraft model is generally a plane-symmetric shape (the plane of symmetry is typically the z =0 plane), and the spatial flow is also plane-symmetric when the incoming flow does not sideslip. In the numerical simulation process, the flow field of the half-mode aircraft is calculated only by adopting a symmetric boundary condition on a symmetric plane, and then the result of the full flow field is obtained through the symmetric boundary. The half-mold calculation method can reduce about half of calculated amount and memory, and is a common method for predicting the plane symmetric flow field of the aircraft with the complex plane symmetric appearance in engineering. However, in the UGKS algorithm, the unstructured velocity space grid generated by the conventional method often cannot ensure the plane symmetry of the grid, so that the processing of distribution functions in symmetric boundaries is very complex, and interpolation processing is required in most cases, which affects the calculation accuracy. In order to avoid complex symmetrical boundary processing and ensure calculation accuracy, even if the surface-symmetrical flow field simulation with a surface-symmetrical appearance is adopted, a plurality of researchers directly carry out cross-basin flow field simulation on the full-mode aircraft, and the waste of calculation resources is caused.
Disclosure of Invention
The invention aims to realize rapid prediction of a plane-symmetric cross-basin flow field of an aircraft.
In order to achieve the above object, the present invention provides a method for predicting a cross-basin flow field with plane symmetry of an aircraft, where the method includes:
step 1: the method comprises the steps of obtaining overall dimension information of an aircraft, wherein the aircraft is a plane-symmetric aircraft, constructing a first model corresponding to the aircraft based on the overall dimension information of the aircraft, cutting half of the first model to obtain a second model, and generating a first physical space grid based on the second model;
step 2: performing flow field solution based on the first physical space grid and the inflow conditions to obtain the speed and temperature information of the flow field, determining a first position corresponding to a temperature lowest point according to the flow field temperature information, and marking the speeds of the first position in the x direction, the y direction and the z direction as follows:、and;
and step 3: generating a three-dimensional velocity space grid, generating an outer boundary of the three-dimensional velocity space grid based on the incoming flow condition, and aligning the three-dimensional velocity space grid based on the outer boundary、Andthe grid ranges in three directions are arranged to obtain a three-dimensional hemispherical area,、andthe grid spacing increases in equal proportion from the origin of coordinates in three directions. With origin of coordinates (,,) Macroscopic velocity corresponding to the lowest point of temperature: (,,) The part is the sphere center, a hemispherical encryption area and a spherical encryption area are respectively arranged, the radius of the hemisphere and the radius of the sphere are determined by the size of the local temperature, and the radius is larger when the temperature is lower. And generating a hemispherical speed space grid based on the constraint conditions. Relating the hemispherical grid toSymmetrically copying the plane to obtain a spherical speed space grid;
and 4, step 4: and carrying out iterative solution on the basis of the first physical space grid and the spherical three-dimensional velocity space grid to obtain the three-dimensional physical space flow field of the aircraft.
The method comprises the following steps: according to the method, the velocity space grid with the plane symmetry characteristic is generated through the local encryption and plane symmetry methods of the velocity space grid, and then the distribution function symmetric boundary conditions are combined, and the half-mode aircraft is adopted for calculation, so that the rapid prediction of the plane symmetry cross-basin flow field is realized.
Preferably, the first model is symmetrical about a z =0 plane in a three-dimensional physical space in which the x-axis, the y-axis and the z-axis are located.
Preferably, in step 2 of the method, an NS solver is used to solve the flow field. The NS solver can be used for realizing rapid and accurate solution of the flow field.
Preferably, the incoming flow conditions in the method are as follows: density of incoming flowSpeed of incoming flow in x directionY-direction upward flow velocityVelocity of incoming flow in z-directionAnd pressure of incoming flow。
Preferably, in the method, the three-dimensional velocity space grid is based on the outer boundary、Andthe grid ranges in three directions are set, including: gridding the three-dimensional velocity space inThe grid range of the direction is set to be greater than or equal toAnd is less than or equal toWherein, in the step (A),is a mode of the incoming flow velocity,the grid range of the direction is set to be greater than or equal toAnd is less than or equal to 0.
Preferably, the iterative solution is performed by using a unified gas dynamics method in step 4 of the method. The method adopts a unified gas dynamics method to carry out iterative solution, can quickly and accurately realize the flow field simulation of the half-mode aircraft, and further can quickly and accurately calculate to obtain the three-dimensional space flow field.
Preferably, the step 4 in the method specifically includes: and carrying out iterative solution on the three-dimensional physical space grid and points in the three-dimensional velocity space grid based on the first physical space grid and the spherical three-dimensional velocity space grid, and processing related boundaries to obtain the three-dimensional physical space flow field of the aircraft.
Preferably, step 1 of the method further comprises: symmetric boundary conditions are specified, and the processing of the relevant boundary in step 4 comprises processing the symmetric boundary.
Preferably, the processing the symmetric boundary in the method includes: setting the macroscopic quantity of the first virtual point and the second virtual point in the symmetric boundary condition and setting the distribution function of the first virtual point and the second virtual point;
wherein, the macroscopic quantity of the first virtual point and the second virtual point in the symmetric boundary condition is set as:
,,,,,,,,,;is the density at the first virtual point and,is the density at the first interior point,is the density at the second virtual point and,is the density at the second interior point,is the velocity in the x direction at the first virtual point,is the velocity in the x-direction at the first interior point,is the speed in the x direction at the second virtual point,is the velocity in the x-direction at the second interior point,is the velocity in the y direction at the first virtual point,is the velocity in the y-direction at the first interior point,is the velocity in the y direction at the second virtual point,is the velocity in the y-direction at the second interior point,is the velocity in the z direction at the first virtual point,is the velocity in the z direction at the first interior point,is the velocity in the z direction at the second virtual point,is the velocity in the z direction at the second interior point,is the pressure at the first virtual point and,is the pressure at the first interior point,is the pressure at the second virtual point and,is the pressure at the second interior point;
the distribution function of the first virtual point and the second virtual point is set as:
wherein f is a distribution function defined in a three-dimensional physical space and a three-dimensional velocity space in the unified gas dynamics method, the first virtual point is symmetrical with the first interior point about the symmetrical boundary, and the second virtual point is symmetrical with the second interior point about the symmetrical boundary.
The density, speed, pressure and distribution functions on two sides of the symmetric boundary can meet the physical fact of mirror symmetry through the arrangement, and the iterative solution of the step 4 can obtain a correct space flow field result.
Preferably, the processing of the relevant boundary in step 4 of the method further includes processing a fixed wall boundary, an incoming flow boundary, and an outgoing flow boundary.
One or more technical schemes provided by the invention at least have the following technical effects or advantages:
the invention encrypts the designated area of the velocity space unstructured grid based on the flow field calculation result, then generates the plane symmetric velocity space grid, and combines the symmetric boundary processing of the distribution function to realize the rapid prediction of the aircraft cross-flow field plane symmetric flow field. According to the method, local grid encryption is performed, so that the grid quantity is reduced, the calculation time is reduced, and the prediction of the method is quicker; the plane symmetric shape is calculated by adopting an aircraft half-mold, so that the calculated amount is reduced by half; the invention adopts the symmetric processing of the velocity space grid, facilitates the search of the velocity space symmetric points in the symmetric boundary and avoids the complex and fussy interpolation calculation process.
Drawings
The accompanying drawings, which are included to provide a further understanding of the embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention;
FIG. 1 is a schematic flow chart of a prediction method of an aircraft surface-symmetric cross-basin flow field;
FIG. 2 is a schematic view of a scaled model of an X38-like aircraft;
FIG. 3 is a grid schematic of the entire velocity space;
fig. 7 is a schematic diagram of symmetric boundary processing.
Detailed Description
In order that the above objects, features and advantages of the present invention can be more clearly understood, a more particular description of the invention will be rendered by reference to the appended drawings. It should be noted that the embodiments of the present invention and features of the embodiments may be combined with each other without conflicting with each other.
In the following description, numerous specific details are set forth in order to provide a thorough understanding of the present invention, however, the present invention may be practiced otherwise than as specifically described and thus the scope of the present invention is not limited by the specific embodiments disclosed below.
Example one
Referring to fig. 1, fig. 1 is a schematic flow chart of a method for predicting a planar symmetric cross-basin flow field of an aircraft, an embodiment of the present invention provides a method for predicting a planar symmetric cross-basin flow field of an aircraft, where the method includes:
step 1: the method comprises the steps of obtaining overall dimension information of an aircraft, wherein the aircraft is a plane-symmetric aircraft, constructing a first model corresponding to the aircraft based on the overall dimension information of the aircraft, cutting half of the first model to obtain a second model, and generating a first physical space grid based on the second model; for a certain plane-symmetric shape aircraft (assuming that an aircraft model is symmetric about a z =0 plane, a coordinate system of a three-dimensional physical space is composed of three coordinate axes (x, y and z) which are perpendicular to each other, and z is a third coordinate axis), a physical space grid is generated for the half-model shape of the aircraft by adopting commercial software, wherein the commercial software can be Gridgen software, ICEM software or other software capable of generating the physical space grid.
And 2, step: based on the physical space grid, a specified incoming flow condition is adopted (the incoming flow condition comprises the following parameters of incoming flow densitySpeed of incoming flow in x directionY-direction upward flow velocityVelocity of incoming flow in z-directionAnd pressure of incoming flow) And solving the flow field by using an NS solver to obtain macroscopic conservation quantities such as the speed, the temperature and the like of the flow field. Determining the position of the lowest temperature point according to the temperature distribution condition and marking the macroscopic velocity value of the position (,,). The NS solver calculation method can be found in the following documents: method for super-Yan, computing hydrodynamics and application]Beijing: beijing university of aerospace publisher, 2006. Other solvers may also be used in this embodiment, and the specific name and type of the solver are not limited in the embodiment of the present invention.
And step 3: an unstructured velocity space grid is generated using commercial software. The commercial software may be gridggen software, ICEM software, or other software capable of generating a physical space grid, and the type of software for generating the physical space grid is not particularly limited by the invention. The three directions of the three-dimensional velocity space grid are respectively、And. To facilitate symmetric boundary processing (see below for details), a velocity space grid is set with respect toPlane symmetry. According to the characteristics of distribution functionWill be a velocity spaceThe grid range of the direction is set as. WhereinIs the mode of the incoming flow velocity.The direction is set as,0). Forming a three-dimensional hemispherical region.、Andthe grid spacing increases in equal proportion from the origin of coordinates in three directions. With origin of coordinates (,,) Macroscopic velocity corresponding to the lowest point of temperature: (,,) The part is the sphere center, a hemispherical encryption area and a spherical encryption area are respectively arranged, the radius of the hemisphere and the radius of the sphere are determined by the size of the local temperature, and the radius is larger when the temperature is lower. And generating a hemispherical speed space grid based on the constraint conditions. Relating the hemispherical grid toAnd symmetrically copying the plane to obtain the spherical speed space grid.
Wherein, the velocity space grid is determined by two aspects: the size of the outer boundary of the mesh and the spatial distribution of the mesh. The outer grid boundaries are generally determined by the flow velocity (given in step 2). The spatial distribution of the unstructured velocity spatial grid is not uniform, and encryption processing is generally performed near a macroscopic velocity corresponding to a coordinate origin and a temperature minimum point to ensure the accuracy and the effectiveness of a calculation result.
And 4, step 4: and (3) based on the half-mode physical space grid in the step (1) and the spherical velocity space grid in the step (3), performing iterative solution by adopting a unified gas dynamics method, realizing flow field simulation of the half-mode aircraft, and calculating to obtain a three-dimensional space flow field. The specific solving process of the unified gas dynamics method refers to the following documents: jiang Ding Wu, a research on gas dynamics algorithm based on analytical solution of model equation, a doctor paper of the Chinese aerodynamic research and development center, 2016 (6 months).
Step 1 in the method also comprises the step of specifying symmetrical boundaries, and in the calculation process of step 4, various boundary processing is involved besides iterative calculation of points in the grid. Such as a solid wall boundary, an incoming flow boundary, an outgoing flow boundary, a symmetric boundary. Wherein the solid wall boundary refers to the boundary of the contact part of the aircraft surface and the external gas, and is used for identifying the solid area; the inflow boundary refers to a boundary flowing into the calculation region and is used for controlling the gas inflow characteristic; outflow boundary refers to the boundary of the outflow calculation region for controlling the gas outflow characteristics; the symmetric boundary is a boundary set for a flow having symmetry, and its normal velocity is generally 0 in order to avoid repetition of a large number of calculations. The fixed wall boundary, the incoming flow boundary and the outgoing flow boundary all adopt conventional processing methods, and the specific implementation process can refer to documents: jiang Ding Wu, a research on gas dynamics algorithm based on analytical solution of model equation, a doctor paper of the Chinese aerodynamic research and development center, 2016 (6 months).
The symmetric boundary processing is used in combination with the plane-symmetric velocity space grid in the invention, and the symmetric boundary condition processing method in the method is as follows:
in this embodiment, the processing the symmetric boundary in the method includes: setting the macroscopic quantity of the first virtual point and the second virtual point in the symmetric boundary condition and setting the distribution function of the first virtual point and the second virtual point;
as shown in fig. 7, a virtual point 1 in fig. 7 is a first virtual point, a virtual point 2 is a second virtual point, an inner point 1 is a first inner point, and an inner point 2 is a second inner point, and macroscopic quantities of the first virtual point and the second virtual point in the symmetric boundary condition are set as:
,,,,,,,,,(ii) a The purpose of this is that if the flow is symmetric about a certain face, the face can be set as a symmetric boundary when generating the physical space grid, which can save half of the physical space grid.Is the density at the first virtual point or points,is the density at the first interior point,is the density at the second virtual point and,is the density at the second interior point,is the velocity in the x direction at the first virtual point,is the velocity in the x-direction at the first interior point,is the speed in the x direction at the second virtual point,is the velocity in the x-direction at the second interior point,is the velocity in the y direction at the first virtual point,is the velocity in the y-direction at the first interior point,is the velocity in the y direction at the second virtual point,is the velocity in the y-direction at the second interior point,is the velocity in the z direction at the first virtual point,is the velocity in the z direction at the first interior point,is the velocity in the z direction at the second virtual point,is the velocity in the z direction at the second interior point,is the pressure at the first virtual point and,is the pressure at the first interior point,is the pressure at the second virtual point and,is the pressure at the second interior point;
the distribution function of the first virtual point and the second virtual point is set as:
wherein f is a distribution function defined in a three-dimensional physical space and a three-dimensional velocity space in the unified gas dynamics method, the first virtual point is symmetrical with the first interior point about the symmetrical boundary, and the second virtual point is symmetrical with the second interior point about the symmetrical boundary.
By means of the arrangement, density, speed, pressure and distribution functions in the symmetric boundary can meet the physical fact of mirror symmetry, and by doing so, the iterative solution of the step 4 can obtain a correct space flow field result.
Due to the adoption of the method in step 3The velocity space grid with plane symmetry does not need any interpolation processing in the assignment process of the virtual point distribution function, thereby reducing the calculation amount and ensuring the calculation precision.
Specific implementation examples of the X38-like aircraft scale model flow field simulation are given below.
The incoming flow conditions are set as follows: mach number 8.0, angle of attack 20 degrees, knudsen number 0.275, reference length 0.28 m. Fig. 2 shows a schematic diagram of the appearance of a scaled model of an aircraft of class X38, X, y, and z in fig. 2 are coordinate axes of a three-dimensional physical space, and fig. 3 shows a schematic diagram of the whole velocity space grid generated in step 3. FIG. 4 is a velocity space gridA cross-sectional view. FIG. 5 is a velocity space gridA cross-sectional view. FIG. 6 is a velocity space gridA cross-sectional view. Local encryption is performed near the origin of coordinates and the lowest point of the temperature of the NS prediction solution. The time consumption of symmetric boundaries in the half-mold physical space grid calculation is negligible, so the single-step time is reduced by about 50%. Therefore, aiming at the aircraft surface-symmetric cross-basin flow field, the speed space unstructured grid optimization technology reduces the calculation time and improves the calculation efficiency, and the method is a rapid flow field prediction method.
In the embodiment, the designated area of the velocity space unstructured grid is encrypted based on the flow field calculation result of the NS solver, then the surface symmetric velocity space grid is generated, and the rapid prediction of the aircraft cross-flow-field surface symmetric flow field is realized by combining the distribution function symmetric boundary processing.
While preferred embodiments of the present invention have been described, additional variations and modifications in those embodiments may occur to those skilled in the art once they learn of the basic inventive concepts. Therefore, it is intended that the appended claims be interpreted as including preferred embodiments and all such alterations and modifications as fall within the scope of the invention.
It will be apparent to those skilled in the art that various changes and modifications may be made in the present invention without departing from the spirit and scope of the invention. Thus, if such modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalents, the present invention is also intended to include such modifications and variations.
Claims (7)
1. A prediction method for an aircraft surface-symmetric cross-basin flow field is characterized by comprising the following steps:
step 1: the method comprises the steps of obtaining overall dimension information of an aircraft, wherein the aircraft is a plane-symmetric aircraft, constructing a first model corresponding to the aircraft based on the overall dimension information of the aircraft, cutting half of the first model to obtain a second model, and generating a first physical space grid based on the second model;
step 2: performing flow field solving based on the first physical space grid and the incoming flow conditions to obtain speed and temperature information of the flow field, and determining a first bit corresponding to a temperature lowest point according to the flow field temperature informationSetting and marking the speed of the first position in the x direction, the y direction and the z direction as follows:、and;
and step 3: generating a three-dimensional velocity space grid, generating an outer boundary of the three-dimensional velocity space grid based on the incoming flow condition, and aligning the three-dimensional velocity space grid based on the outer boundary、Andthe grid ranges in three directions are arranged to obtain a three-dimensional hemispherical area,、andthe position is a coordinate origin, and a hemispherical encryption area is obtained by taking the coordinate origin as a sphere center so as to、Andobtaining a spherical encryption area for the center of sphere, generating a hemispherical velocity space grid based on grid spacing distribution information, the hemispherical encryption area and the spherical encryption area, relating the hemispherical velocity space grid toSymmetrically copying the plane to obtain a spherical speed space grid;
and 4, step 4: performing iterative solution on the basis of the first physical space grid and the spherical three-dimensional velocity space grid to obtain a three-dimensional physical space flow field of the aircraft;
the step 4 specifically includes: based on the first physical space grid and the spherical three-dimensional velocity space grid, carrying out iterative solution on the three-dimensional physical space grid and points in the three-dimensional velocity space grid, and processing related boundaries to obtain a three-dimensional physical space flow field of the aircraft;
the step 1 further comprises: specifying symmetric boundary conditions, wherein the processing of the relevant boundary in step 4 comprises processing the symmetric boundary;
processing the symmetric boundary includes: setting the macroscopic quantity of the first virtual point and the second virtual point in the symmetric boundary condition and setting the distribution function of the first virtual point and the second virtual point;
wherein, the macroscopic quantity of the first virtual point and the second virtual point in the symmetric boundary condition is set as:
,,,,,,,,,;is the density at the first virtual point and,is the density at the first interior point,is the density at the second virtual point and,is the density at the second interior point,is the velocity in the x direction at the first virtual point,is the velocity in the x-direction at the first interior point,is the speed in the x direction at the second virtual point,is the velocity in the x-direction at the second interior point,is the velocity in the y direction at the first virtual point,is the velocity in the y-direction at the first interior point,is the velocity in the y direction at the second virtual point,is the velocity in the y-direction at the second interior point,is the velocity in the z direction at the first virtual point,is the velocity in the z direction at the first interior point,is the velocity in the z direction at the second virtual point,is the velocity in the z direction at the second interior point,is the pressure at the first virtual point and,is the pressure at the first interior point,is the pressure at the second virtual point and,is the pressure at the second interior point;
the distribution function of the first virtual point and the second virtual point is set as:
wherein f is a distribution function defined in a three-dimensional physical space and a three-dimensional velocity space in the unified gas dynamics method, the first virtual point is symmetrical with the first interior point about the symmetrical boundary, and the second virtual point is symmetrical with the second interior point about the symmetrical boundary.
2. The method for predicting the aircraft surface-symmetric cross-basin flow field according to claim 1, wherein the first model is symmetric about a z =0 plane in a three-dimensional physical space where an x-axis, a y-axis and a z-axis are located.
3. The method for predicting the aircraft surface-symmetric cross-basin flow field according to claim 1, wherein an NS solver is adopted to solve the flow field in step 2 of the method.
4. The method for predicting the aircraft surface-symmetric cross-basin flow field according to claim 1, wherein the inflow conditions are as follows: density of incoming flowSpeed of incoming flow in x directionY-direction upward flow velocityVelocity of incoming flow in z-directionAnd pressure of incoming flow。
5. The method for predicting the cross-basin flow field of the aircraft according to claim 4, wherein the three-dimensional velocity space grid based on the outer boundary is、Andthe grid ranges in three directions are set, including: gridding the three-dimensional velocity space inAndthe grid range of the direction is set to be greater than or equal toAnd is less than or equal toWherein, in the step (A),is a mode of the incoming flow velocity,the grid range of the direction is set to be greater than or equal toAnd is less than or equal to 0.
6. The method for predicting the aircraft surface-symmetric cross-basin flow field according to claim 1, wherein the iterative solution is performed by using a unified gas dynamics method in step 4 of the method.
7. The method for predicting the aircraft surface-symmetric cross-basin flow field according to claim 1, wherein the processing of the relevant boundaries in step 4 further comprises processing of a fixed wall boundary, an inflow boundary and an outflow boundary.
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