CN113742845A - Method for modifying forward transition model by adopting hypersonic velocity flow field temperature - Google Patents

Method for modifying forward transition model by adopting hypersonic velocity flow field temperature Download PDF

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CN113742845A
CN113742845A CN202111047248.XA CN202111047248A CN113742845A CN 113742845 A CN113742845 A CN 113742845A CN 202111047248 A CN202111047248 A CN 202111047248A CN 113742845 A CN113742845 A CN 113742845A
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赵金山
石义雷
张志刚
陈挺
肖雨
粟斯尧
余嘉
廖军好
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Ultra High Speed Aerodynamics Institute China Aerodynamics Research and Development Center
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Abstract

The invention relates to a method for modifying a model by adopting hypersonic velocity flow field temperature, which comprises the following steps: determining a test model and flow field input conditions, solving an NS equation and gamma-Re to be correctedθA transition model is used for obtaining a predicted flow field parameter; calculating the ratio of the reference temperature to the boundary layer outer edge temperature to obtain a temperature correction factor; gamma-Re to be correctedθModifying the transition critical momentum thickness Reynolds number of the transition model to obtain the modified gamma-ReθTransition to the model; solving NS equation and corrected gamma-ReθThe transition model obtains the corrected flow field parameters; judging whether the corrected flow field parameters are converged; if the convergence is achieved, stopping calculation to finish correction, and otherwise, returning to the iterative calculation. The invention can solve the problem of original gamma-ReθTransition modelWhen the method is used for predicting the transition of the hypersonic speed boundary layer, the prediction accuracy of the transition region length is low.

Description

Method for modifying forward transition model by adopting hypersonic velocity flow field temperature
Technical Field
The invention relates to the technical field of hypersonic velocity aircraft design, in particular to a method for adopting a hypersonic velocity flow field temperature modification transition model, an aircraft transition prediction method, computer equipment and a computer readable storage medium.
Background
The accuracy of the transition starting position and the transition region length prediction directly determines the advancement of the aerodynamic appearance design of the hypersonic aircraft and the reliability of the starting of the engine air inlet channel. Currently, the Reynolds average method combined with the transition mode theory is widely applied to the engineering design application of the actual hypersonic aircraft, wherein a typical representative example is gamma-ReθAnd the transition model can effectively predict flow field parameters. However, gamma-Re of the prior artθThe transition model is developed according to a low-speed boundary layer transition mechanism, and is used for the problem that the transition region length is longer than the real transition process in the prediction analysis of the hypersonic boundary layer transition, and the prediction precision is low.
Disclosure of Invention
The invention aims to provide a transition model correction technology aiming at a hypersonic velocity flow field to solve the problem that a transition region prediction result is long in the prior art, aiming at least part of defects.
In order to achieve the purpose, the invention provides a method for modifying a transition model by adopting hypersonic velocity flow field temperature, which comprises the following steps:
s1, determining the test model and the flow field input conditions, solving the NS equation and the gamma-Re to be correctedθA transition model is used for obtaining a predicted flow field parameter;
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 gamma-Re to be corrected based on the temperature correction factorθModifying the transition critical momentum thickness Reynolds number of the transition model to obtain the modified gamma-ReθTransition to the model;
s5, solving NS equation and corrected gamma-ReθThe transition model obtains 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 the NS equation and the corrected gamma-ReθWhether the transition model calculates convergence; if the flow field parameter is converged, stopping calculation to finish correction, otherwise, taking the corrected flow field parameter as the predicted flow field parameter, and returning to the step S2 for iterative calculation.
Optionally, in the step S2, the reference temperature T and the boundary layer outer edge temperature T are calculatedeThe following formula is adopted for the ratio of (A):
Figure BDA0003251381110000021
wherein, TwIndicating wall temperature, TeThe temperature of the outer edge of the boundary layer is shown,
Figure BDA0003251381110000022
Figure BDA0003251381110000023
Tsrepresenting the total temperature of the incoming flow, gamma is the specific heat ratio, MeIndicating the boundary layer outer edge mach number.
Optionally, the boundary layer outer edge mach number MeThe value of (c) is determined by:
if the test model has a single structure and no shock wave interference in the flow field, obtaining the Mach number M of the incoming flow according to the input conditions of the flow field∞Mach number M of future stream∞The value of (A) is assigned to the boundary layer outer edge Mach number Me；
Otherwise, searching the Mach number M of the outer edge of the boundary layer by adopting boundary layer searching operation according to the predicted flow field parameterseThe numerical value of (c).
Optionally, the mach number M of the outer edge of the boundary layer is searched by adopting boundary layer searching operation according to the predicted flow field parameterseThe values of (a) include:
according to the predicted flow field parameters, positioning the total temperature T of the flow field in the boundary layer section to reach 0.995 times of the total temperature T of the incoming flow for the first time along the wall surface normal direction of the flow fieldsThe position of the boundary layer is the corresponding position of the outer edge of the boundary layer;
searching the Mach number of the corresponding position of the outer edge of the boundary layer as the Mach number M of the outer edge of the boundary layereThe numerical value of (c).
Optionally, the temperature correction factor f is obtained in step S3TThe following formula is adopted:
Figure BDA0003251381110000031
wherein c is a correction constant, and Pr is a Plantet number.
Optionally, in the step S4, γ — Re to be correctedθThe transition critical momentum thickness Reynolds number of the transition model is subjected to temperature correction, and the following formula is adopted:
Reθcnew＝ReθcfT
wherein, ReθcDenotes the gamma-Re to be correctedθTransition critical momentum thickness Reynolds number Re of transition modelθcnewRepresents the corrected gamma-ReθTransition critical momentum thickness Reynolds number in transition model.
Optionally, the method further comprises the following steps:
and S7, outputting the corrected flow field parameters to finish the transition prediction of the hypersonic flow field.
The invention also provides a transition prediction method of the aircraft, which comprises the following steps:
constructing a test model grid based on the aircraft;
the method for correcting the gamma-Re by adopting the model for correcting the hypersonic velocity flow field temperatureθTransition to the model;
based on corrected gamma-ReθAnd (4) a transition model is used for predicting transition 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 any one of the steps of the method for adopting the hypersonic flow field temperature transition model when executing the computer program.
The invention further provides a computer-readable storage medium, on which a computer program is stored, where the computer program, when executed by a processor, implements any of the above steps of the method for modifying a transition model using hypersonic flow field temperature.
The technical scheme of the invention has the following advantages: the invention provides a method for adopting a hypersonic velocity flow field temperature transition model, an aircraft transition prediction method, computer equipment and a computer readable storage mediumθThe model structure theory of transition carries out the correction of the transition model of the hypersonic velocity flow field, and the influence of the object surface load coefficient along with the local Mach number, the wall temperature, the incoming flow static temperature and the ratio of the wall temperature to the incoming flow static temperature is eliminated to a certain extent by adopting the reference temperature.
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FIG. 1 is a graph showing the Reynolds number of the transition critical momentum thickness with the Reynolds number of the transition occurrence momentum;
FIG. 2 is a schematic diagram illustrating a method for transition model by using hypersonic flow field temperature modification in an embodiment of the present invention;
FIG. 3 shows a transition module using hypersonic flow field temperature modification in an embodiment of the present invention93-10 skirt cone model obtained by calculation of model method at incoming flow Mach number M∞Comparing the object surface dimensionless pressure distribution with the actually measured wind tunnel test data when the object surface dimensionless pressure distribution is 5.91;
FIG. 4 shows an incoming flow Mach number M of a 93-10 skirt cone model calculated by a method for modifying a transition model by using hypersonic velocity flow field temperature in an embodiment of the invention∞And comparing the dimensionless temperature of the object plane at 5.91 with the actually measured wind tunnel test data.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
As previously described, the original γ -ReθThe transition model is developed according to a low-speed boundary layer transition mechanism, and in the prediction analysis for the transition of the hypersonic boundary layer, a problem that the length of the transition region is longer than that of a real transition process often occurs. Therefore, it is necessary to combine the hypersonic flow characteristic, the hypersonic boundary layer transition mechanism and γ -ReθThe transition mode structure theory is modified to improve the prediction precision of the transition zone length of the actual hypersonic complex-shape aircraft.
The most important characteristic of hypersonic flow is strong compression of air flow caused by a high mach flow field, and further, due to the fact that the temperature of the flow field is sharply increased after a shock wave formed by strong compression and large temperature gradient in a boundary layer is caused by strong shearing, corresponding correction needs to be carried out on the influence of a temperature effect on a transition mode. In the original gamma-ReθIn transition model, the main relationship formula F for controlling the transition region lengthlengthThe characteristic of severe temperature change in the hypersonic velocity flow field is ignored because temperature-related terms do not exist in the process. At the same time, FlengthWhen the transition region length is increased, the calculated transition region length is increased,the transition starting position will be advanced and the transition occurrence momentum thickness Reynolds number
Figure BDA0003251381110000051
Critical momentum thickness Reynolds number Re of transitionθcMonotone strong correlation, as shown in fig. 1. If the transition region length is to be simulated as accurately as possible, the transition occurs with a momentum thickness Reynolds number
Figure BDA0003251381110000052
Transition critical momentum thickness Reynolds number ReθcAt low level
Figure BDA0003251381110000053
The zone is as large as possible and at high
Figure BDA0003251381110000054
The area is reduced appropriately. In view of the above, the present invention provides a reference temperature T and a boundary layer outer edge temperature TeThe ratio of (a) to (b) introduces a transition critical momentum thickness Reynolds number ReθcAnd (4) performing correction, wherein the influence of the object plane load coefficient along with the local Mach number, the wall temperature, the incoming flow static temperature and the ratio of the wall temperature to the incoming flow static temperature can be eliminated to a certain extent by adopting the reference temperature.
As shown in fig. 2, a method for modifying a transition model by using a hypersonic flow field temperature according to an embodiment of the present invention includes the following steps:
s1, determining the test model and the flow field input conditions, solving the NS equation and the gamma-Re to be correctedθAnd (5) obtaining a predicted flow field parameter by the transition model. Wherein the NS equation and gamma-Re are solvedθThe transition model can be solved through numerical value dispersion; the flow field parameters comprise Mach number of each grid point, pressure, temperature, density, speed in three directions, viscosity coefficient, heat conduction coefficient, object surface heat flux, friction coefficient and the like.
S2, calculating reference temperature T and boundary layer outer edge temperature T according to input conditions and predicted flow field parameterseThe ratio of (a) to (b).
S3, based on the reference temperature T and the boundary layer outer edge temperature TeThe ratio of (a) to (b),obtaining a temperature correction factor fT。
S4 gamma-Re to be corrected based on the temperature correction factorθModifying the transition critical momentum thickness Reynolds number of the transition model to obtain the modified gamma-ReθA transition model.
S5, solving NS equation and corrected gamma-ReθAnd (5) transitioning 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 the NS equation and the corrected gamma-ReθWhether the transition model calculates convergence; if the flow field parameter is converged, stopping calculation to finish correction, otherwise, taking the corrected flow field parameter as the predicted flow field parameter, and returning to the step S2 for iterative calculation. The flow field parameters tend to be stable, namely converged, and the NS equation and the corrected gamma-Re can be judged through iterative calculationθIf the corrected flow field parameter obtained in step S5 is compared with the predicted flow field parameter obtained in step S1, and the deviation exceeds the preset threshold, it may be determined that the flow field parameter is not converged.
The technical scheme provided by the invention aims at the original gamma-ReθWhen the transition model is used for predicting transition of the hypersonic speed boundary layer, the limitation of the violent characteristic of the temperature change of the flow field is not fully considered, temperature correction is carried out by introducing a concept of reference temperature, and the reference temperature T and the temperature T of the outer edge of the boundary layer are utilizedeCorrected original gamma-ReθTransition critical momentum thickness Reynolds number Re of transition modelθcThe temperature effect in the hypersonic flow characteristic can be introduced into the transition model, so that the problem of the original gamma-Re is effectively solvedθWhen the transition model is used for transition prediction of a high supersonic velocity boundary layer, the transition region is long in length and low in prediction accuracy.
Preferably, in step S2, the reference temperature T and the boundary layer outer edge temperature T are calculatedeThe following formula can be adopted for the ratio of (A):
Figure BDA0003251381110000061
wherein, TwRepresenting wall temperature, determined by flow field input conditions, TeThe temperature of the outer edge of the boundary layer is shown,
Figure BDA0003251381110000062
Tsthe representative total temperature of the incoming flow is obtained by calculating the input conditions, and the general calculation mode is
Figure BDA0003251381110000063
T∞Representing the incoming flow temperature, M∞Representing the incoming flow Mach number, gamma being the specific heat ratio from the input conditions, and for calorimetric complete air, it is generally 1.4, MeIndicating the boundary layer outer edge mach number.
Further, the Mach number M of the outer edge of the boundary layereThe value of (c) can be determined by:
if the test model has a single structure (such as a simple shape like a flat plate, a pointed cone and the like) and the flow field has no shock wave interference, the Mach number M of the incoming flow is obtained according to the input conditions of the flow field∞Mach number M of future stream∞The value of (A) is assigned to the boundary layer outer edge Mach number MeTo perform subsequent calculations;
otherwise, searching the Mach number M of the outer edge of the boundary layer by adopting boundary layer searching operation according to the predicted flow field parameterseThe numerical value of (c).
It should be noted that the stream Mach number M is obtained in the future∞The value of (A) is assigned to the boundary layer outer edge Mach number MeSubsequent calculation is carried out, and the actual numerical difference between the two values is not too large, preferably
Figure BDA0003251381110000071
δ ≈ 0.4. For a test model with a relatively simple shape, the future stream Mach number M∞The value of (A) is assigned to the boundary layer outer edge Mach number MeSubsequent calculation is carried out, and the Mach number M of the outer edge of the boundary layer can be quickly and simply determinedeThe numerical value of (c).
Further, for complex-shaped test models, based on predictionsThe flow field parameters of the method are obtained by searching the Mach number M of the outer edge of the boundary layer by adopting the boundary layer searching operationeThe values of (a) include:
according to the predicted flow field parameters, positioning the total temperature T of the flow field in the boundary layer section to reach 0.995 times of the total temperature T of the incoming flow for the first time along the wall surface normal direction of the flow fieldsThe position of the boundary layer is the corresponding position of the outer edge of the boundary layer;
searching the Mach number of the corresponding position of the outer edge of the boundary layer as the Mach number M of the outer edge of the boundary layereThe numerical value of (c).
For a test model with a complex shape, by adopting the technical scheme, the Mach number M of the outer edge of the boundary layer can be accurately obtainedeThe value of (2) improves the correction precision of the hypersonic flow field transition model.
Preferably, the temperature correction factor f is obtained in step S3TThe following formula is adopted:
Figure BDA0003251381110000072
the c is a correction constant, the value range is 0.9-1, a large amount of ground wind tunnel transition test data of the hypersonic speed boundary layer can be matched in the range, Pr is a Plantt number, and the value of the method is 0.72.
Preferably, gamma-Re to be correctedθModifying the transition critical momentum thickness Reynolds number of the transition model by adopting the following formula:
Reθcnew＝ReθcfT
wherein, ReθcDenotes the gamma-Re to be correctedθTransition critical momentum thickness Reynolds number Re of transition modelθcnewRepresents the corrected gamma-ReθTransition critical momentum thickness Reynolds number in transition model. When the correction is performed, the correction is required for each grid point.
Preferably, the method for modifying the forward rotation model by using the hypersonic flow field temperature further includes the following steps:
and S7, outputting the corrected flow field parameters to finish the transition prediction of the hypersonic flow field.
And transition prediction can be carried out on the determined test model and the flow field input condition according to the output flow field parameters.
In a specific embodiment, the invention takes a 93-10 skirt cone model as a test model, and verifies the performance of the method for modifying the model by using the hypersonic flow field temperature (the modification method of the invention for short) provided by the invention, wherein the flow field input conditions include: mach number M of incoming flow∞(5.91) Reynolds number Re of incoming flow∞/L＝9.348×106M, temperature T of incoming flow∞56.2K, turbulence pulsation intensity Tu of incoming flow∞The object plane pressure distribution and the temperature distribution obtained by the correction method of the invention are better in accordance with the actually measured wind tunnel test result, the large part area of the 93-10 skirt cone model surface is basically superposed with the wind tunnel test result, and only the tail part has difference, thereby proving the rationality and the feasibility of the invention, as shown in fig. 3 and fig. 4.
In some embodiments, the invention further provides an aircraft transition prediction method, which includes:
constructing a test model grid based on the aircraft;
the method for correcting the gamma-Re by adopting the hypersonic velocity flow field temperature correction transition model in any one of the above embodimentsθTransition to the model;
based on corrected gamma-ReθAnd (4) a transition model is used for predicting transition of the boundary layer of the aircraft.
Particularly, 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, when executing the computer program, implements the steps of the method for using the hypersonic flow field temperature modification transition model in any one of the above embodiments.
In other preferred embodiments of the present invention, a computer-readable storage medium is further provided, where a computer program is stored on the computer-readable storage medium, and when being executed by a processor, the computer program implements the steps of the method for using the hypersonic flow field temperature modification transition model in any one of the above embodiments.
It will be understood by those skilled in the art that all or part of the processes of the method of the embodiments described above may be implemented by a computer program, and the computer program may be stored in a non-volatile computer-readable storage medium, and when executed, the computer program may include the processes of the method embodiments of the model for transitioning through hypersonic flow field temperature modification, and will not be described again here.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention, but not to limit it; although the present 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 solutions described in the foregoing embodiments may still be modified, or some technical features may be equivalently replaced; and such modifications or substitutions do not depart from the spirit and scope of the corresponding technical solutions of the embodiments of the present invention.

Claims (10)

1. A method for modifying a transition model by adopting hypersonic velocity flow field temperature is characterized by comprising the following steps:
s1, determining the test model and the flow field input conditions, solving the NS equation and the gamma-Re to be correctedθA transition model is used for obtaining a predicted flow field parameter;
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 gamma-Re to be corrected based on the temperature correction factorθModifying the transition critical momentum thickness Reynolds number of the transition model to obtain the modified gamma-ReθTransition to the model;
s5, solving NS equation and corrected gamma-ReθThe transition model obtains the corrected flow field parameters;
s6, stepThe corrected flow field parameters obtained in step S5 are compared with the predicted flow field parameters obtained in step S1, and the NS equation and the corrected γ -Re are determinedθWhether the transition model calculates convergence; if the flow field parameter is converged, stopping calculation to finish correction, otherwise, taking the corrected flow field parameter as the predicted flow field parameter, and returning to the step S2 for iterative calculation.
2. The method for model modification using hypersonic flow field temperature according to claim 1, characterized in that:
in step S2, the reference temperature T and the boundary layer outer edge temperature T are calculatedeThe following formula is adopted for the ratio of (A):
Figure FDA0003251381100000011
wherein, TwIndicating wall temperature, TeThe temperature of the outer edge of the boundary layer is shown,
Figure FDA0003251381100000012
Figure FDA0003251381100000013
Tsrepresenting the total temperature of the incoming flow, gamma is the specific heat ratio, MeIndicating the boundary layer outer edge mach number.
3. The method for model modification using hypersonic flow field temperature of claim 2, wherein,
mach number M of outer edge of boundary layereThe value of (c) is determined by:
if the test model has a single structure and no shock wave interference in the flow field, obtaining the Mach number M of the incoming flow according to the input conditions of the flow field∞Mach number M of future stream∞The value of (A) is assigned to the boundary layer outer edge Mach number Me；
Otherwise, according to the predicted flow field parameters, adopting boundary layer searching operation to search the boundary layerMach number of edge MeThe numerical value of (c).
4. The method for model modification using hypersonic flow field temperature according to claim 3, characterized in that:
searching the Mach number M of the outer edge of the boundary layer by adopting boundary layer searching operation according to the predicted flow field parameterseThe values of (a) include:
according to the predicted flow field parameters, positioning the total temperature T of the flow field in the boundary layer section to reach 0.995 times of the total temperature T of the incoming flow for the first time along the wall surface normal direction of the flow fieldsThe position of the boundary layer is the corresponding position of the outer edge of the boundary layer;
searching the Mach number of the corresponding position of the outer edge of the boundary layer as the Mach number M of the outer edge of the boundary layereThe numerical value of (c).
5. The method for model modification using hypersonic flow field temperature according to claim 2, characterized in that:
the temperature correction factor f is obtained in the step S3TThe following formula is adopted:
Figure FDA0003251381100000021
wherein c is a correction constant, and Pr is a Plantet number.
6. The method for model modification using hypersonic flow field temperature according to claim 5, characterized in that:
in the step S4, gamma-Re to be correctedθThe transition critical momentum thickness Reynolds number of the transition model is subjected to temperature correction, and the following formula is adopted:
Reθcnew＝ReθcfT
wherein, ReθcDenotes the gamma-Re to be correctedθTransition critical momentum thickness Reynolds number Re of transition modelθcnewRepresents the corrected gamma-ReθTransition critical momentum thickness thunder of transition modelNuo count.
7. The method for model modification using hypersonic flow field temperature according to claim 1, further comprising the steps of:
and S7, outputting the corrected flow field parameters to finish the transition prediction of the hypersonic flow field.
8. A transition prediction method for an aircraft, comprising:
constructing a test model grid based on the aircraft;
method for correcting gamma-Re by adopting hypersonic flow field temperature correction transition model according to any one of claims 1 to 7θTransition to the model;
based on corrected gamma-ReθAnd (4) a transition model is used for predicting transition of the boundary layer of the aircraft.
9. A computer device comprising a memory and a processor, wherein the memory stores a computer program, and the processor executes the computer program to implement the steps of the method for employing the hypersonic flow field temperature modification transition model according to any one of claims 1 to 7.
10. A computer-readable storage medium, on which a computer program is stored, wherein the computer program, when being executed by a processor, implements the steps of the method for employing the hypersonic flow field temperature modification transition model according to any one of claims 1 to 7.
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