CN113742845B - Method for correcting transition model by hypersonic flow field temperature - Google Patents
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Abstract
The invention relates to a method for correcting transition model by hypersonic flow field temperature, which comprises the following steps: determining the input conditions of a test model and a flow field, and solving an NS equation and gamma-Re to be corrected θ Transition to a model to obtain predicted flow field parameters; calculating the ratio of the reference temperature to the boundary layer outer edge temperature to obtain a temperature correction factor; gamma-Re to be corrected θ Correcting the transition critical momentum thickness Reynolds number in the transition model to obtain corrected gamma-Re θ Transition model; solving NS equation and modified gamma-Re θ Transition to the model to obtain corrected flow field parameters; judging whether the corrected flow field parameters are converged or not; if the calculation is converged, stopping the calculation to finish the correction, otherwise, returning to the iterative calculation. The invention can solve the problem of original gamma-Re θ When the transition model is used for hypersonic boundary layer transition prediction, the transition region length prediction precision is low.
Description
Technical Field
The invention relates to the technical field of hypersonic aircraft design, in particular to a method for correcting a transition model by adopting hypersonic flow field temperature, an aircraft transition prediction method, computer equipment and a computer readable storage medium.
Background
The accuracy of transition starting position and transition zone length prediction directly determines hypersonic speed flightThe advancement of the aerodynamic profile design of the vehicle, and the reliability of engine intake tract start. At present, the Reynolds averaging method combined with transition mode theory is widely applied to the engineering design application of the actual hypersonic aircraft, wherein the typical representative is gamma-Re θ The transition model can effectively predict flow field parameters. However, gamma-Re in the prior art θ The transition model is developed according to a low-speed boundary layer transition mechanism, and is used for hypersonic boundary layer transition prediction analysis, the problem that the transition zone length is longer than the real transition process often occurs, and the prediction accuracy is low.
Disclosure of Invention
The invention aims to provide a transition model correction technology for a hypersonic flow field aiming at least one part of the defects, so as to solve the problem of longer prediction result of a transition region in the prior art.
In order to achieve the above purpose, the invention provides a method for correcting a transition model by adopting hypersonic flow field temperature, which comprises the following steps:
s1, determining input conditions of a test model and a flow field, and solving an NS equation and gamma-Re to be corrected θ Transition to a model to obtain predicted flow field parameters;
s2, calculating the ratio of the reference temperature to the boundary layer outer edge temperature according to the input conditions and the predicted flow field parameters;
s3, obtaining a temperature correction factor based on the ratio of the reference temperature to the boundary layer outer edge temperature;
s4, based on the temperature correction factor, gamma-Re to be corrected θ Correcting the transition critical momentum thickness Reynolds number in the transition model to obtain corrected gamma-Re θ Transition model;
s5, solving an NS equation and corrected gamma-Re θ Transition to the model to obtain corrected flow field parameters;
s6, comparing the corrected flow field parameter obtained in the step S5 with the predicted flow field parameter obtained in the step S1, and judging an NS equation and corrected gamma-Re θ Whether the transition model calculates convergence or not; if the calculation is converged, the calculation is stopped to complete the correction,otherwise, the corrected flow field parameter is used as the predicted flow field parameter, and the step S2 is returned to perform iterative calculation.
Optionally, in the step S2, a reference temperature T and a boundary layer outer edge temperature T are calculated e Is a ratio of the following formula:
wherein T is w Indicating the wall temperature, T e Represents the temperature of the outer edge of the boundary layer,
T s represents the total temperature of incoming flow, gamma is the specific heat ratio, M e Indicating the boundary layer outer edge mach number.
Optionally, the boundary layer outer edge Mach number M e The value of (2) is determined as follows:
if the test model has a single structure and no shock wave interference exists in the flow field, the Mach number M of the incoming flow is obtained according to the input condition of the flow field ∞ Future stream Mach number M ∞ Is assigned to the boundary layer peripheral Mach number M e ；
Otherwise, according to the predicted flow field parameters, adopting boundary layer searching operation to search the Mach number M at the outer edge of the boundary layer e Is a numerical value of (2).
Optionally, the boundary layer searching operation is adopted to search the Mach number M at the outer edge of the boundary layer according to the predicted flow field parameters e Comprises the following numerical values:
according to the predicted flow field parameters, the total temperature T of the incoming flow, which is 0.995 times of the total temperature of the flow field in the section of the boundary layer, is positioned along the normal direction of the wall surface of the flow field s The position of the boundary layer is the corresponding position of the outer edge of the boundary layer;
searching Mach number at corresponding position of boundary layer outer edge as boundary layer outer edge Mach number M e Is a numerical value of (2).
Optionally, the temperature correction factor f is obtained in the step S3 T The following formula is adopted:
where c is a correction constant and Pr is a Planet number.
Optionally, in the step S4, γ -Re to be corrected θ Temperature correction is carried out on transition critical momentum thickness Reynolds number of the transition model, and the following formula is adopted:
Re θcnew ＝Re θc f T
wherein Re is θc Representing gamma-Re to be corrected θ Transition critical momentum thickness Reynolds number, re of transition model θcnew Represents corrected gamma-Re θ Transition critical momentum thickness Reynolds number of transition model.
Optionally, the method further comprises the following steps:
and S7, outputting the corrected flow field parameters to finish hypersonic flow field transition prediction.
The invention also provides an aircraft transition prediction method, which comprises the following steps:
constructing a test model grid based on the aircraft;
correcting gamma-Re by adopting the method for correcting transition model by adopting hypersonic flow field temperature according to any one of the above methods θ Transition model;
gamma-Re based on completion of correction θ And (5) transition model, and performing transition prediction of the boundary layer of the aircraft.
The invention also provides computer equipment, which comprises a memory and a processor, wherein the memory stores a computer program, and the processor realizes the steps of the method adopting the hypersonic flow field temperature correction transition model according to any one of the above steps when executing the computer program.
The invention also provides a computer readable storage medium, on which a computer program is stored, which when being executed by a processor, implements the steps of the method for adopting hypersonic flow field temperature correction transition model described in any one of the above.
The technical scheme of the invention has the following advantages: the invention provides a method for correcting transition model by hypersonic flow field temperature, an aircraft transition prediction method, computer equipment and a computer readable storage medium θ The hypersonic flow field transition model correction is carried out by adopting the transition model construction theory, the influence of the object plane load coefficient along with the local Mach number, the wall temperature, the incoming flow static temperature and the wall temperature and incoming flow static temperature ratio is eliminated to a certain extent by adopting the reference temperature, the problem of the hypersonic flow field transition region prediction bias length can be solved, and the hypersonic aircraft transition region length prediction precision is improved.
Drawings
Fig. 1 shows a change curve of transition critical momentum thickness reynolds number with transition occurrence momentum thickness reynolds number;
FIG. 2 is a schematic diagram of steps of a method for modifying a transition model by using hypersonic flow field temperature in an embodiment of the invention;
FIG. 3 shows an incoming flow Mach number M of a 93-10 skirt cone model calculated by a hypersonic flow field temperature correction transition model in an embodiment of the invention ∞ The object plane dimensionless pressure distribution at=5.91 is compared with the measured wind tunnel test data;
FIG. 4 shows an incoming flow Mach number M of a 93-10 skirt cone model calculated by a hypersonic flow field temperature correction transition model in an embodiment of the invention ∞ Object plane dimensionless temperature at=5.91 is compared with measured wind tunnel test data.
Detailed Description
For the purpose of making the objects, technical solutions and advantages of the embodiments of the present invention more apparent, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention, and it is apparent that the described embodiments are some embodiments of the present invention, but not all embodiments of the present invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
As previously described, the original gamma-Re θ The transition model is developed according to a low-speed boundary layer transition mechanism, and in the hypersonic boundary layer transition prediction analysis, the problem that the transition zone length is longer than the real transition process often occurs. Therefore, the hypersonic flow characteristics, hypersonic boundary layer transition mechanism and gamma-Re are necessarily combined θ And correcting the transition mode construction theory to improve the prediction precision of the transition zone length of the actual hypersonic speed complex-appearance aircraft.
The hypersonic flow is mainly characterized in that airflow is strongly compressed due to a high Mach flow field, and further, as the temperature of the flow field is rapidly increased after a shock wave formed by strong compressibility and a large temperature gradient in a boundary layer is caused by strong shearing, corresponding correction is required to be carried out on the influence of temperature effect on a transition mode. In the original gamma-Re θ In the transition model, a main relation F for controlling the length of the transition zone length The temperature related item does not exist, and the characteristic of severe temperature change in the hypersonic flow field is ignored. At the same time F length When the calculated transition zone length is increased, the transition starting position is advanced, and the transition occurrence momentum thickness Reynolds number is increased
And transition critical momentum thickness Reynolds number Re θc Monotonically strong correlations as shown in fig. 1. If the transition zone length is to be simulated as accurately as possible, the Reynolds number of the thickness of the momentum relative to the transition occurrence is +.>
Transition critical momentum thickness Reynolds number Re θc At low->
The region is as large as possible and is at high +.>
The area is then suitably reduced. In view of this, the present invention refers to the reference temperature T and the boundary layer outer edge temperature T e Is introduced into transition critical momentum thickness Reynolds number Re θc The influence of the object plane load coefficient on the local Mach number, the wall temperature, the incoming flow static temperature and the incoming flow static temperature ratio can be eliminated to a certain extent by adopting the reference temperature.
As shown in fig. 2, the method for correcting the transition model by adopting the hypersonic flow field temperature provided by the embodiment of the invention comprises the following steps:
s1, determining input conditions of a test model and a flow field, and solving an NS equation and gamma-Re to be corrected θ And (5) transferring the model to obtain predicted flow field parameters. Wherein, solving NS equation and gamma-Re θ The transition model can be solved through numerical discrete; the flow field parameters include Mach number, pressure, temperature, density, speed in three directions, viscosity coefficient, heat conduction coefficient, object plane heat flux, friction coefficient, etc. of each grid point.
S2, calculating a reference temperature T and a boundary layer outer edge temperature T according to the input conditions and the predicted flow field parameters e Is a ratio of (2).
S3, based on the reference temperature T and the boundary layer outer edge temperature T e To obtain the temperature correction factor f T 。
S4, based on the temperature correction factor, gamma-Re to be corrected θ Correcting the transition critical momentum thickness Reynolds number in the transition model to obtain corrected gamma-Re θ And (5) transition to a model.
S5, solving an NS equation and corrected gamma-Re θ And (5) transferring the model to obtain the corrected flow field parameters.
S6, comparing the corrected flow field parameter obtained in the step S5 with the predicted flow field parameter obtained in the step S1, and judging an NS equation and corrected gamma-Re θ Whether the transition model calculates convergence or not; if the flow field parameters are converged, stopping calculation to finish correction, otherwise, taking the corrected flow field parameters as predicted flow field parameters, and returning to the step S2 to perform iterative calculation. The flow field parameters tend to be stable, namely convergence, through iterative calculationJudging NS equation and modified gamma-Re θ When the transition model is converged, the preset threshold value can be passed, and if the deviation exceeds the preset threshold value, the corrected flow field parameter obtained in the step S5 is compared with the predicted flow field parameter obtained in the step S1, so that the convergence is not considered.
The technical proposal provided by the invention aims at the original gamma-Re θ When the transition model is used for hypersonic boundary layer transition prediction, the limitation of severe flow field temperature change characteristics is not fully considered, temperature correction is carried out by introducing a concept of reference temperature, and reference temperature T and boundary layer outer edge temperature T are utilized e Is used for correcting the original gamma-Re θ Transition critical momentum thickness Reynolds number Re of transition model θc The temperature effect in hypersonic flow characteristics can be introduced into a transition model, so that the original gamma-Re is effectively solved θ When the transition model is used for hypersonic boundary layer transition prediction, the problem that the transition zone length is longer and the prediction precision is lower is solved.
Preferably, in step S2, a reference temperature T and a boundary layer outer edge temperature T are calculated e Can be represented by the following formula:
wherein T is w Representing the wall temperature, determining according to the flow field input condition, T e Represents the temperature of the outer edge of the boundary layer,
T s representing the total temperature of the incoming flow, and is obtained by calculating the input condition, wherein the common calculation mode is that
T ∞ Representing the incoming flow temperature, M ∞ Represents Mach number of incoming flow, gamma is specific heat ratio from input condition, and for calorimetric complete air, 1.4, M is usually taken e Indicating the boundary layer outer edge mach number.
Further, boundary layer peripheral edge horsesHertz number M e The value of (2) may be determined as follows:
if the test model has a single structure (such as a simple shape of a flat plate, a pointed cone and the like) and no shock wave interference exists in the flow field, the incoming flow Mach number M is obtained according to the flow field input condition ∞ Future stream Mach number M ∞ Is assigned to the boundary layer peripheral Mach number M e To perform subsequent calculations;
otherwise, according to the predicted flow field parameters, adopting boundary layer searching operation to search the Mach number M at the outer edge of the boundary layer e Is a numerical value of (2).
If the future stream Mach number M ∞ Is assigned to the boundary layer peripheral Mach number M e The actual numerical value difference between the two should not be too large, and is preferably calculated
Delta. Is approximately 0.4. For a test model of relatively simple shape, the future stream Mach number M ∞ Is assigned to the boundary layer peripheral Mach number M e The subsequent calculation can be carried out to quickly and simply determine the Mach number M at the outer edge of the boundary layer e Is a numerical value of (2).
Further, for a test model with a complex appearance, according to predicted flow field parameters, boundary layer searching operation is adopted to search the Mach number M at the outer edge of the boundary layer e Comprises the following numerical values:
according to the predicted flow field parameters, the total temperature T of the incoming flow, which is 0.995 times of the total temperature of the flow field in the section of the boundary layer, is positioned along the normal direction of the wall surface of the flow field s The position of the boundary layer is the corresponding position of the outer edge of the boundary layer;
searching Mach number at corresponding position of boundary layer outer edge as boundary layer outer edge Mach number M e Is a numerical value of (2).
For a test model with a complex appearance, the Mach number M of the outer edge of the boundary layer can be accurately obtained by adopting the technical scheme e And the correction precision of the hypersonic flow field transition model is improved.
Preferably, the temperature correction factor f is obtained in step S3 T The following formula is adopted:
wherein c is a correction constant, the value range is 0.9-1, a large amount of hypersonic boundary layer transition ground wind tunnel test data can be matched in the range, pr is the Plandter number, and the value of the hypersonic boundary layer transition ground wind tunnel test data is 0.72.
Preferably, gamma-Re to be corrected θ And correcting the transition critical momentum thickness Reynolds number of the transition model by adopting the following formula:
Re θcnew ＝Re θc f T
wherein Re is θc Representing gamma-Re to be corrected θ Transition critical momentum thickness Reynolds number, re of transition model θcnew Represents corrected gamma-Re θ Transition critical momentum thickness Reynolds number of transition model. When the correction is performed, the correction is required for each grid point.
Preferably, the method for correcting the transition model by adopting the hypersonic flow field temperature further comprises the following steps:
and S7, outputting the corrected flow field parameters to finish hypersonic flow field transition prediction.
And according to the output flow field parameters, performing transition prediction on the determined test model and the flow field input conditions.
In a specific embodiment, the invention uses 93-10 skirt cone model as test model, and verifies the performance of the method for correcting transition model by hypersonic flow field temperature (the correction method is short for the invention), and the flow field input conditions comprise: mach number M of incoming stream ∞ =5.91, incoming flow unit reynolds number Re ∞ /L＝9.348×10 6 /m, incoming flow temperature T ∞ =56.2k, incoming flow turbulence pulsation intensity Tu ∞ The angle of attack alpha=0.1%, and the object plane boundary is a non-slip heat insulation wall, as shown in fig. 3 and 4, the object plane pressure distribution and the temperature distribution obtained by the correction of the correction method of the invention are in good agreement with the actually measured wind tunnel test result, 93And (5) the most area of the surface of the skirt cone model is basically coincident with the wind tunnel test result, and only the tail part is different, so that the rationality and feasibility of the invention are proved.
In some embodiments, the present invention further provides an aircraft transition prediction method, including:
constructing a test model grid based on the aircraft;
the gamma-Re is corrected by adopting the method of correcting the transition model by adopting the hypersonic flow field temperature according to any one of the embodiments θ Transition model;
gamma-Re based on completion of correction θ And (5) transition model, and performing transition prediction of the boundary layer of the aircraft.
In particular, in some preferred embodiments of the present invention, there is further provided a computer device, including a memory and a processor, where the memory stores a computer program, and the processor implements the steps of the method for using the hypersonic flow field temperature correction transition model in any of the above embodiments when the processor executes the computer program.
In other preferred embodiments of the present invention, there is also provided a computer readable storage medium having stored thereon a computer program which, when executed by a processor, implements the steps of the method for using the hypersonic flow field temperature correction transition model described in any of the above embodiments.
Those skilled in the art will appreciate that implementing all or part of the above-described embodiment of the method may be accomplished by a computer program to instruct related hardware, where the computer program may be stored in a non-volatile computer readable storage medium, and the computer program may include the above-described embodiment of the method using the hypersonic flow field temperature correction transition model when executed, and the description thereof will not be repeated here.
Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present invention, and are not limiting; although the invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit and scope of the technical solutions of the embodiments of the present invention.
Claims (10)
1. The method for correcting the transition model by adopting the hypersonic flow field temperature is characterized by comprising the following steps of:
s1, determining input conditions of a test model and a flow field, and solving an NS equation and gamma-Re to be corrected θ Transition to a model to obtain predicted flow field parameters;
s2, calculating the ratio of the reference temperature to the boundary layer outer edge temperature according to the input conditions and the predicted flow field parameters;
s3, obtaining a temperature correction factor based on the ratio of the reference temperature to the boundary layer outer edge temperature;
s4, based on the temperature correction factor, gamma-Re to be corrected θ Correcting the transition critical momentum thickness Reynolds number in the transition model to obtain corrected gamma-Re θ Transition model;
s5, solving an NS equation and corrected gamma-Re θ Transition to the model to obtain corrected flow field parameters;
s6, comparing the corrected flow field parameter obtained in the step S5 with the predicted flow field parameter obtained in the step S1, and judging an NS equation and corrected gamma-Re θ Whether the transition model calculates convergence or not; if the flow field parameters are converged, stopping calculation to finish correction, otherwise, taking the corrected flow field parameters as predicted flow field parameters, and returning to the step S2 to perform iterative calculation.
2. The method for adopting the hypersonic flow field temperature correction transition model according to claim 1, wherein the method is characterized in that:
in the step S2, a reference temperature T and a boundary layer outer edge temperature T are calculated e Is a ratio of the following formula:
wherein T is w Indicating the wall temperature, T e Represents the temperature of the outer edge of the boundary layer,
T s represents the total temperature of incoming flow, gamma is the specific heat ratio, M e Indicating the boundary layer outer edge mach number.
3. The method for modifying transition model by hypersonic flow field temperature according to claim 2, characterized in that,
the boundary layer outer edge Mach number M e The value of (2) is determined as follows:
if the test model has a single structure and no shock wave interference exists in the flow field, the Mach number M of the incoming flow is obtained according to the input condition of the flow field ∞ Future stream Mach number M ∞ Is assigned to the boundary layer peripheral Mach number M e ；
Otherwise, according to the predicted flow field parameters, adopting boundary layer searching operation to search the Mach number M at the outer edge of the boundary layer e Is a numerical value of (2).
4. The method for adopting the hypersonic flow field temperature correction transition model according to claim 3, wherein the method is characterized in that:
the boundary layer searching operation is adopted to search the Mach number M at the outer edge of the boundary layer according to the predicted flow field parameters e Comprises the following numerical values:
according to the predicted flow field parameters, the total temperature T of the incoming flow, which is 0.995 times of the total temperature of the flow field in the section of the boundary layer, is positioned along the normal direction of the wall surface of the flow field s The position of the boundary layer is the corresponding position of the outer edge of the boundary layer;
searching Mach number at corresponding position of boundary layer outer edge as boundary layer outer edge Mach number M e Is a numerical value of (2).
5. The method for adopting the hypersonic flow field temperature correction transition model according to claim 2, wherein the method is characterized in that:
the temperature correction factor f is obtained in the step S3 T The following formula is adopted:
where c is a correction constant and Pr is a Planet number.
6. The method for adopting the hypersonic flow field temperature correction transition model according to claim 5, wherein the method is characterized in that:
in the step S4, the gamma-Re to be corrected θ Temperature correction is carried out on transition critical momentum thickness Reynolds number of the transition model, and the following formula is adopted:
Re θcnew ＝Re θc f T
wherein Re is θc Representing gamma-Re to be corrected θ Transition critical momentum thickness Reynolds number, re of transition model θcnew Represents corrected gamma-Re θ Transition critical momentum thickness Reynolds number of transition model.
7. The method for modifying a transition model by using hypersonic flow field temperature as set forth in claim 1, further comprising the steps of:
and S7, outputting the corrected flow field parameters to finish hypersonic flow field transition prediction.
8. The transition prediction method for the aircraft is characterized by comprising the following steps of:
constructing a test model grid based on the aircraft;
modifying gamma-Re by adopting the method for modifying transition model by adopting hypersonic flow field temperature according to any one of claims 1-7 θ Transition model;
gamma-Re based on completion of correction θ Transition model, enterPredicting transition of boundary layer of aircraft.
9. A computer device comprising a memory and a processor, the memory storing a computer program, characterized in that the processor, when executing the computer program, implements the steps of the method of any one of claims 1 to 7 using a hypersonic flow field temperature correction transition model.
10. A computer readable storage medium having stored thereon a computer program, wherein the computer program when executed by a processor implements the steps of the method of any one of claims 1 to 7 employing a hypersonic flow field temperature correction transition model.
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| CN116432561B (en) * | 2023-06-05 | 2024-03-12 | 中国空气动力研究与发展中心计算空气动力研究所 | Transition prediction method of hypersonic boundary layer |
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