CN117852435A - Method for calculating transonic flow field of aircraft - Google Patents
Method for calculating transonic flow field of aircraft Download PDFInfo
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- CN117852435A CN117852435A CN202311747697.4A CN202311747697A CN117852435A CN 117852435 A CN117852435 A CN 117852435A CN 202311747697 A CN202311747697 A CN 202311747697A CN 117852435 A CN117852435 A CN 117852435A
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Abstract
The invention discloses a method for calculating a transonic flow field of an aircraft, which comprises the following steps: performing unstructured grid division on the aircraft model; introducing an unstructured grid division file into fluid mechanics simulation software, setting a simulation type to be stable, creating a calculation domain, setting a starting condition and a boundary condition of simulation, and selecting a local time step control time scale, wherein the method comprises the following steps: when the flow speed in the calculation area is high or the grid density is low, selecting a smaller time scale, otherwise, selecting a larger time scale; and setting a solving monitoring point, observing a calculation result curve, and if the result curve tends to be stable or uniformly vibrate in a small amplitude along with the time step, considering that the result is converged. The invention realizes high-efficiency, stable and rapid convergence in transonic flow field calculation.
Description
Technical Field
The invention relates to the technical field of aircraft flow field calculation, in particular to an aircraft transonic flow field calculation method.
Background
In the flight process of the aircraft, the resistance encountered in the flight is larger and larger along with the increase of the flight speed, and when the flight Mach number is close to 1, the phenomena of rapid increase of the resistance and loss can occur. This is due to the local shock wave that occurs at the supersonic region in the transonic flow field, which causes the flow separation due to the interaction of the shock wave with the boundary layer. The air flow causes viscous stress, heat conduction and the like due to the generation of abrupt compression and expansion and internal friction, so that the resistance, lift force, stability and maneuverability of the aircraft are changed sharply and irregularly and complexly. With further increases in Mach numbers, the range of local shock waves continues to expand. In practice, as the flight speed increases further to supersonic speeds, the shock wave is converted to a disjunct shock wave that occurs in front of the aircraft. Because the shock wave is far away from the aircraft body, the supersonic shock wave generally does not interfere with the flow in the boundary layer of the body, and the boundary layer separation on the surface of the body is not caused. Thus, the supersonic flow regime is much more stable than transonic. It is quite difficult to handle the transonic flow theoretically: first, the linearization theory is not applicable and the nonlinear equation must be solved; secondly, there are two flow patterns of sub-supersonic speed in the flow field at the same time, and the boundary between them is unknown; again, since the shock wave is severely disturbed by the boundary layer, the viscosity effect and the like must be considered. The method can be used for analyzing and calculating only under individual conditions, solves the problem of transonic flow and mainly depends on experiments.
The small guided projectile is used as a calculation object, and the problems of low convergence speed, poor calculation robustness, easy divergence and the like exist in the calculation process of the transonic flow field. The general CFD solver formula is completely implicit, the time scale is generally estimated based on L/U, where L is the characteristic length in the computational domain, and U is the flow rate, and in order to increase the convergence speed, a larger time scale is usually selected, but when the time scale is too large, the convergence will oscillate or even diverge, and if the time scale is too small, the convergence speed will be slow.
Disclosure of Invention
The invention aims to: the invention aims to overcome the defects of the prior art, and provides a method for calculating a transonic flow field of an aircraft, which adopts a local time scale, can select different time scales in different areas of a calculation domain, and can obtain a higher convergence rate.
The technical scheme is as follows: the invention relates to a method for calculating a transonic flow field of an aircraft, which comprises the following steps:
s1: performing unstructured grid division on the aircraft model;
s2: introducing an unstructured grid division file into fluid mechanics simulation software, setting a simulation type to be stable, creating a calculation domain, setting a starting condition and a boundary condition of simulation, and selecting a local time step control time scale, wherein the calculation method comprises the following steps of: time scale = local time scale x local time factor;
s3: setting a solving monitoring point, observing a calculated result curve, and if the result curve advances along with the time step, the result curve tends to vibrate smoothly or uniformly in a small amplitude, and the vibration interval is not more than 10 -5 The result is considered to have converged.
Further perfecting the technical scheme, the dividing rule of the local time scale is as follows: selecting different time scales in different areas of the calculation domain, when the flow speed in the calculation domain is large or the grid size is small or the aspect ratio is large, selecting a smaller time scale, otherwise, selecting a larger time scale; the local time factor is dependent on the mach number and decreases exponentially as the mach number increases.
Further, performing unstructured meshing in S1 includes: the method is characterized in that the flow complex area of the aircraft model is refined, meanwhile, in order to better calculate the flow field characteristics in the boundary layer, the calculation accuracy is improved, meanwhile, the height of the first layer of boundary layer grid is calculated and y+ is determined, the growth rate is set to be 1.1, and 20 layers of boundary layer grids are drawn on the surface of the aircraft model. Because transonic sections avoid large nearby variable gradients, if no boundary layer is added, the flux on inaccurate walls is calculated, which can be ameliorated by plotting the boundary layer mesh.
Further, the boundary condition of the calculation domain is set as a pressure far field, the surface of the projectile body is set as a slip-free boundary condition, an SST turbulence model is selected, the incoming flow Ma number is 1.2, the attack angle is 0-16 degrees, the calculation step number is 2000, and the local time factor is set to be 0.1-0.5. And the convergence speed is increased while the convergence of the calculation result is ensured.
Further, the monitoring points include an axial force coefficient ca, a normal force coefficient cn and a pitching moment coefficient mz of the monitoring model.
Further, the aircraft model is a small guided projectile model.
Further, non-structural meshing of the small guided projectile model aircraft model includes: and refining the grids of the head, control surface and tail area of the tail wing.
Further, the local time factor used for transonic segments is 0.1.
The beneficial effects are that: compared with the prior art, the invention has the advantages that: according to the method for calculating the transonic flow field of the aircraft, provided by the invention, local time scale control is adopted, and high-efficiency, stable and rapid convergence in transonic flow field calculation is realized through unstructured grid division and simulation setting. The calculation method has the advantages of short time, good stability and high matching degree between the calculation result and the test result; compared with fluent software, the calculation speed can be increased by more than one time.
Drawings
FIG. 1 is a flow chart of a calculation method of the present invention;
FIG. 2 is a schematic diagram showing the comparison of experimental and simulated calculated axial force coefficients in the present invention;
FIG. 3 is a schematic diagram showing the comparison of test and simulation calculation normal force coefficients in the present invention;
FIG. 4 is a graph showing the comparison of the pitch moment coefficients calculated by the test and simulation in the present invention.
Detailed Description
The technical scheme of the invention is described in detail below through the drawings, but the protection scope of the invention is not limited to the embodiments.
Example 1: the method for calculating the transonic flow field of the aircraft shown in fig. 1 comprises the following steps:
s1: performing unstructured grid division on the aircraft model;
s2: introducing an unstructured grid division file into fluid mechanics simulation software, setting a simulation type to be stable, creating a calculation domain, setting a simulation starting condition and a boundary condition, selecting different time scales in different areas of the calculation domain, and selecting a smaller time scale when the flow speed in the calculation domain is large or the grid size is smaller or the length-width ratio is larger, otherwise, selecting a larger time scale;
specifically, the local time scale employs the following rule:
wherein: the basic time scale is an initial time scale set according to global simulation requirements; the flow field velocity gradient is a velocity gradient calculated on each calculation unit; the speed gradient threshold is a set speed gradient criterion, and when the speed gradient of the flow field exceeds the threshold, a local time scale is adopted; the local mesh size is a local mesh size of each computing unit; the grid size threshold is a set grid size criterion, and when the grid size of the computing unit is smaller than this threshold, a local time scale is used.
When the local time scale is adopted, the calculation method is as follows: time scale = local time scale x local time factor;
wherein C is a local time factor, Δt i Is the time scale of the ith grid cell, ΔL i For the length of the ith grid cell, U i Speed for the ith grid cell;
the local time factor C increases with mach number (Ma number, i.e., fluid velocity to sonic velocity ratio), the time factor decreases exponentially,
C＝C initial ×e -a·Ma
wherein C is initial Is an initial time factor, which represents an initial time factor value in the case of a low mach number, a is a positive constant for adjusting the slope of the index, and Ma is a mach number. Preferably, the local time factor is set to between 0.1 and 0.5, wherein the local time factor used for the transonic section is selected to be 0.1.
S3: setting a solving monitoring point, observing a calculated result curve, and if the result curve advances along with the time step, the result curve tends to vibrate smoothly or uniformly in a small amplitude, and the vibration interval is not more than 10 -5 The result is considered to have converged.
Example 2: the method provided in the embodiment 1 is applied to a small guided projectile model, and specifically comprises the following steps:
s1: performing unstructured grid division on a small guided projectile model, wherein the areas with more complex flow, such as the head part, the control surface, the combination part of the surface of the projectile body, the tail part of the tail wing and the like, are thinned, meanwhile, in order to better calculate the flow field characteristics in the boundary layer, improve the calculation accuracy, calculate y+ to determine the grid height of the first layer of the boundary layer, set the growth rate to be 1.1, draw 20 layers of the grid of the boundary layer on the surface of the projectile body, and finish grid division;
s2: introducing an unstructured grid division file into fluid mechanics simulation software, setting a simulation type to be stable, creating a calculation domain, setting a boundary condition of the calculation domain to be a pressure far field, setting a slip-free boundary condition on the surface of an elastomer, selecting an SST turbulence model, setting the number of incoming flows Ma to be 1.2, setting the attack angle to be 0-16 degrees, setting the number of calculation steps to be 2000, setting a local time factor to be 0.1, and selecting a local time step control time scale.
S3: and setting a solving monitoring point, monitoring an axial force coefficient ca, a normal force coefficient cn and a pitching moment coefficient mz of the model, observing a calculation result curve, and considering that the result is converged if the propulsion curve tends to be stable or uniformly vibrate in a small amplitude along with the time step as shown in fig. 2 to 4.
Compared with the prior art, the invention has the following technical effects: the calculation method has the advantages of short time, good stability and high matching degree between the calculation result and the test result; compared with fluent software, the calculation speed can be increased by more than one time.
As described above, although the present invention has been shown and described with reference to certain preferred embodiments, it is not to be construed as limiting the invention itself. Various changes in form and details may be made therein without departing from the spirit and scope of the invention as defined by the appended claims.
Claims (8)
1. The method for calculating the transonic flow field of the aircraft is characterized by comprising the following steps of:
s1: performing unstructured grid division on the aircraft model;
s2: introducing an unstructured grid division file into fluid mechanics simulation software, setting a simulation type to be stable, creating a calculation domain, setting a starting condition and a boundary condition of simulation, and selecting a local time step control time scale, wherein the calculation method comprises the following steps of: time scale = local time scale x local time factor;
s3: setting a solving monitoring point, observing a calculated result curve, and if the result curve advances along with the time step, the result curve tends to vibrate smoothly or uniformly in a small amplitude, and the vibration interval is not more than 10 -5 The result is considered to have converged.
2. The aircraft transonic flow field calculation method of claim 1, wherein: the dividing rule of the local time scale is as follows: selecting different time scales in different areas of the calculation domain, when the flow speed in the calculation domain is large or the grid size is small or the aspect ratio is large, selecting a smaller time scale, otherwise, selecting a larger time scale; the local time factor is dependent on the mach number and decreases exponentially as the mach number increases.
3. The aircraft transonic flow field calculation method of claim 2, further comprising: the step of performing unstructured grid division in the step S1 comprises the following steps: refining the complex flowing area of the aircraft model, determining the height of the first layer of boundary layer grid by calculating y+, setting the growth rate to be 1.1, and drawing 20 layers of boundary layer grids on the surface of the aircraft model.
4. A method of aircraft transonic flow field calculation as set forth in claim 3, characterized in that: in the step S2: the boundary condition of the calculation domain is set as a pressure far field, the surface of the projectile body is set as a slip-free boundary condition, an SST turbulence model is selected, the incoming flow Ma number is 1.2, the attack angle is 0-16 degrees, the calculation step number is 2000, and the local time factor is set to be 0.1-0.5.
5. The aircraft transonic flow field calculation method of claim 1, wherein: the monitoring points comprise an axial force coefficient ca, a normal force coefficient cn and a pitching moment coefficient mz of the monitoring model.
6. The aircraft transonic flow field calculation method of claim 1, wherein: the aircraft model is a small guided projectile model.
7. The aircraft transonic flow field calculation method of claim 5, further comprising: unstructured meshing of a small guided projectile model aircraft model includes: and refining the grids of the head, control surface and tail area of the tail wing.
8. The aircraft transonic flow field calculation method of claim 6, further comprising: the local time factor used for transonic segments is 0.1.
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