CN113792508A - Aerodynamic heat calculation method considering surface quality injection effect - Google Patents
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Abstract
The invention discloses a pneumatic thermal calculation method considering surface quality injection effect, which comprises the following steps: s1, acquiring the geometric shape of the hypersonic aircraft; s2, performing grid division on the geometric shape of the obtained hypersonic aircraft; s3, acquiring hypersonic incoming flow speed data, hypersonic incoming flow temperature data, hypersonic incoming flow density data and hypersonic incoming flow pressure data, and inputting surface quality injection gas mass flow rate data and surface quality injection gas temperature data into a computer processor; s4, calculating surface mass injection gas density data, surface mass injection gas velocity data, surface mass injection gas pressure data and surface mass injection gas temperature data; s5, calculating wall surface heat flow data of the hypersonic flight vehicle, and expressing the aerodynamic heating environment of the hypersonic flight vehicle through the wall surface heat flow data; the method can be used for predicting the aerodynamic thermal environment of the hypersonic aircraft with surface quality injection more accurately.
Description
Technical Field
The invention relates to the field of aerodynamic heat and thermal protection, in particular to an aerodynamic heat calculation method considering surface quality injection effect.
Background
Hypersonic aircrafts are faced with severe aerodynamic thermal environments, and researchers adopt various thermal protection means to break through thermal barriers. Among various thermal protection means, ablation heat protection technology has been widely used for decades; the sweating cooling technology is hopeful to solve the problem of thermal protection of the hypersonic aircraft in the near space due to the unique advantages. Both of the above-mentioned two thermal protection techniques involve surface quality injection: such as pyrolysis, evaporation/sublimation and various chemical reactions in ablation heat-proof technology, which all include the phenomenon that gas is extracted from the surface of a material and enters a flow field, and the phenomenon that coolant in sweating cooling technology is injected into the flow field from the surface of a porous medium.
The existing aerodynamic heat calculation method considering the surface quality injection effect mainly has the following defects:
1. the calculation method of engineering experience is summarized, although the calculation speed is high, the detailed influence of surface quality injection on the hypersonic aircraft on the aerodynamic thermal environment cannot be explored;
2. only the influence of vertical surface jet flow in the surface quality ejection phenomenon of the hypersonic aircraft is considered, and the influence brought by the flow viscosity of parallel surfaces is not considered;
3. when model surface data is processed in the calculation process, only the normal velocity of surface quality injection gas is used for simulating vertical surface jet flow in surface quality injection, the pressure gradient between a wall surface grid point and a first layer grid point adjacent to the wall surface is zero, and the surface quality injection effect and the like cannot be truly reflected.
In conclusion, the existing aerodynamic heat calculation method considering the surface quality injection effect has the problems of single considered phenomenon and inaccurate and imprecise prediction phenomenon.
Disclosure of Invention
The invention aims to overcome the defects of the prior art, provides the aerodynamic heat calculation method considering the surface quality ejection effect, and can accurately predict the influence of the hypersonic aircraft surface quality ejection effect on the aerodynamic heat environment.
The purpose of the invention is realized by the following scheme:
a method for calculating aerodynamic heat by considering surface quality injection effect comprises the following steps:
s1, acquiring the geometric shape of the hypersonic aircraft;
s2, performing grid division on the geometric shape of the obtained hypersonic aircraft;
s3, acquiring hypersonic incoming flow speed data, hypersonic incoming flow temperature data, hypersonic incoming flow density data and hypersonic incoming flow pressure data, and inputting surface quality injection gas mass flow rate data and surface quality injection gas temperature data into a computer processor;
s4, calculating surface mass injection gas density data, surface mass injection gas velocity data, surface mass injection gas pressure data and surface mass injection gas temperature data in the computer processor based on the data in the step S3 and the surface mass injection gas constant data;
and S5, in the computer processor, substituting the data in the steps S1-S4 into an N-S equation to calculate to obtain wall surface heat flow data of the hypersonic flight, expressing the pneumatic thermal environment of the hypersonic flight vehicle through the wall surface heat flow data, and then presenting the pneumatic thermal environment based on a computer display unit.
Further, in step S2, the following sub-steps are included: and gradually drawing the grids corresponding to the geometric appearances of the hypersonic aircraft according to the sequence of point-to-line, line-to-surface and surface-to-body aiming at the appearance of the hypersonic aircraft in combination with working conditions, and then finishing grid division.
Further, in step S4, the method includes the sub-steps of:
s40, inputting the mass flow rate of the surface mass injection gas
Surface quality injection gas temperature
Surface mass ejection gas constant
And obtaining the pressure on the grid point of the first layer close to the wall surface by flow field calculation by using the hypersonic speed incoming flow speed data, the hypersonic speed incoming flow temperature data, the hypersonic speed incoming flow density data and the hypersonic speed incoming flow pressure data which are obtained in the step S3
Density on first grid points adjacent to wall
Normal velocity at a first grid point adjacent to the wall
；
S41, calculating the surface quality injection gas density data according to the following formula
：
S42, surface quality injection gas density data obtained by calculation in the step S2 is utilized
And the surface mass injection gas mass flow rate input in step S3
The normal velocity of the surface-quality injected vertical jet gas is calculated according to the following formula
：
S43, injecting gas density data by using the obtained surface quality
And the surface quality injection gas temperature input in the step S3
Surface mass ejection gas constant
The surface quality injection gas pressure is calculated according to the following formula
：
Further, the hypersonic aerial vehicle comprises a spacecraft reentry capsule.
Further, the hypersonic aerocraft comprises a blunt hypersonic reentry missile.
Further, the hypersonic aerial vehicle comprises a near space hypersonic aerial vehicle.
Further, the hypersonic aerocraft comprises a Mars and other planetary advancers.
Further, in step S3, according to the trajectory of the known hypersonic vehicle, that is, altitude data and mach number data of the hypersonic vehicle at different times, a gas parameter table is looked up to obtain hypersonic incoming flow data, hypersonic incoming flow temperature data, hypersonic incoming flow density data and hypersonic incoming flow pressure data; obtaining the coming current speed data of the hypersonic aerocraft according to the known hypersonic Mach number data and the hypersonic coming current speed data obtained by table lookup,
， 
is the incoming flow speed of the hypersonic aerocraft,
The Mach number of the hypersonic aircraft,
Is a hypersonic speed of incoming flow.
Further, in step S4, the surface quality injection gas constant is obtained from a table of physical property parameters of the gas.
The beneficial effects of the invention include:
the embodiment of the invention provides a pneumatic thermal calculation method considering surface quality injection effect, which is closer to the surface quality injection effect than the prior art. Specifically, the method considers the influence of vertical jet flow of the wall surface of the hypersonic aircraft containing the mass injection area and the influence of very high viscosity of gas flowing parallel to the wall surface, adopts the momentum conservation equation with non-zero pressure gradient in the direction of the vertical wall surface to calculate and process according to the characteristic of surface mass injection, is closer to the physical phenomenon compared with the scheme of simulating the vertical jet flow of the wall surface in the direction of velocity in the prior art, and realizes more accurate prediction of the influence of the surface mass injection effect of the hypersonic aircraft on the pneumatic thermal environment.
The input end of the calculation method provided by the embodiment of the invention is the surface quality injection gas mass flow rate, the surface quality injection gas temperature and the surface quality injection gas constant data, the surface quality injection gas mass flow rate is easier to directly obtain in practical engineering practice and experiments, and the influence and influence rule of the surface quality injection effect of the hypersonic aircraft on the pneumatic thermal environment, the flow characteristics and the like can be predicted more accurately.
The calculation method provided by the embodiment of the invention can lay a good foundation for the fluid-solid-heat coupling calculation research of the ablation technology and the sweating cooling technology in the thermal protection technology by more accurately and truly simulating the surface quality ejection effect.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
FIG. 1 is a schematic diagram of direction settings in an implementation of a computation method in an embodiment of the invention; wherein the content of the first and second substances,
is the direction of the normal direction of the device,
， 
are two directions on a tangent plane;
FIG. 2 is a schematic diagram of model surface data processing in an aerodynamic thermal calculation method taking into account surface quality injection effect according to an embodiment of the present invention;
FIG. 3 is a flowchart of model surface data processing in an aerodynamic thermal calculation method with consideration of surface mass ejection effects according to an embodiment of the present invention;
FIG. 4 is a diagram of a model used in a prior art theoretical calculation and experimental model; wherein, the model is a cone with a half cone angle of 5 degrees, wherein, the front 3.75 inches L from the head of the model along the surface is a non-surface quality injection area (non-shadow part), namely a normal solid wall surface; starting at 3.75 inches and moving back 16.25 inches, i.e., up to the end of the mold, the surface quality shot region (shaded);
FIG. 5 is a comparison graph of the prior theoretical calculation results and experimental data with the wall surface heat flow results calculated by the aerodynamic heat calculation method considering the surface quality injection effect of the present invention; wherein the abscissa is the non-dimensionalized distance (s/L =95.25mm along the model surface length. the ordinate is the non-dimensionalized wall heat flow (wall heat flow)
Reference heat flow
). In the upper right diagram of fig. 5, cap represents the result calculated by the aerodynamic heat calculation method considering the surface mass ejection effect of the present invention in a cap calculation platform, Theory represents the result calculated by a theoretical method, and Experiment represents the result. m =0 for an injection rate of 0%, m =0.04 for an injection rate of 0.04%, and m =0.08 for an injection rate of 0.08%, and lines and symbols corresponding to the respective results are shown graphically in the upper right corner of fig. 5. The injection rate is as follows: surface mass injection gas mass flow rate ratioAn incoming mass flow rate;
FIG. 6 is a flowchart illustrating steps of an embodiment of a method of the present invention.
Detailed Description
All features disclosed in all embodiments in this specification, or all methods or process steps implicitly disclosed, may be combined and/or expanded, or substituted, in any way, except for mutually exclusive features and/or steps.
Example 1
As shown in fig. 1 to 6, a method for calculating aerodynamic heat considering surface mass ejection effect includes the steps of:
s1, acquiring the geometric shape of the hypersonic aircraft;
s2, performing grid division on the geometric shape of the obtained hypersonic aircraft;
s3, acquiring hypersonic incoming flow speed data, hypersonic incoming flow temperature data, hypersonic incoming flow density data and hypersonic incoming flow pressure data, and inputting surface quality injection gas mass flow rate data and surface quality injection gas temperature data into a computer processor;
s4, calculating surface mass injection gas density data, surface mass injection gas velocity data, surface mass injection gas pressure data and surface mass injection gas temperature data in the computer processor based on the data in the step S3 and the surface mass injection gas constant data;
and S5, in the computer processor, substituting the data in the steps S1-S4 into an N-S equation to calculate to obtain wall surface heat flow data of the hypersonic flight, expressing the pneumatic thermal environment of the hypersonic flight vehicle through the wall surface heat flow data, and then presenting the pneumatic thermal environment based on a computer display unit.
Example 2
On the basis of embodiment 1, in step S2, the following sub-steps are included: and gradually drawing the grids corresponding to the geometric appearances of the hypersonic aircraft according to the sequence of point-to-line, line-to-surface and surface-to-body aiming at the appearance of the hypersonic aircraft in combination with working conditions, and then finishing grid division.
Example 3
On the basis of embodiment 1, in step S4, the method includes the sub-steps of:
s40, inputting the mass flow rate of the surface mass injection gas
Surface quality injection gas temperature
Surface mass ejection gas constant
And obtaining the pressure on the grid point of the first layer close to the wall surface by flow field calculation by using the hypersonic speed incoming flow speed data, the hypersonic speed incoming flow temperature data, the hypersonic speed incoming flow density data and the hypersonic speed incoming flow pressure data which are obtained in the step S3
Density on first grid points adjacent to wall
Normal velocity at a first grid point adjacent to the wall
；
S41, calculating the surface quality injection gas density data according to the following formula
：
S42, surface quality injection gas density data obtained by calculation in the step S2 is utilized
And step S3 input surface mass injection gas mass flow rate
The normal velocity of the surface-quality injected vertical jet gas is calculated according to the following formula
：
S43, injecting gas density data by using the obtained surface quality
And the surface quality injection gas temperature input in the step S3
Surface mass ejection gas constant
The surface quality injection gas pressure is calculated according to the following formula
：
In other embodiments of the invention based on example 1, it is noted that the hypersonic aircraft comprises a spacecraft re-entry capsule.
In other embodiments of the present invention based on example 1, it should be noted that the hypersonic aircraft includes a blunt hypersonic re-entry missile.
In other embodiments of the present invention based on example 1, it should be noted that the hypersonic flight vehicle includes a hypersonic flight vehicle in a near space.
In another embodiment of the present invention based on example 1, it should be noted that the hypersonic flight vehicle includes a mars and other planetary introducers.
In other embodiments of the present invention based on example 1, it should be noted that, in step S3, according to the trajectory of the known hypersonic vehicle, that is, altitude data of the hypersonic vehicle at different times and mach number data of the hypersonic vehicle, a gas parameter table is looked up to obtain hypersonic incoming flow data, hypersonic incoming flow temperature data, hypersonic incoming flow density data and hypersonic incoming flow pressure data; obtaining the coming current speed data of the hypersonic aerocraft according to the known hypersonic Mach number data and the hypersonic coming current speed data obtained by table lookup,
， 
is the incoming flow speed of the hypersonic aerocraft,
The Mach number of the hypersonic aircraft,
Is a hypersonic speed of incoming flow.
In another embodiment of the present invention based on example 1, in step S4, the surface mass injection gas constant is obtained from a table of physical property parameters of the gas.
The results obtained by calculation by adopting the aerodynamic heat calculation method considering the surface quality injection effect in the embodiment of the invention are compared and evaluated with theoretical calculation solutions and experimental data previously obtained by Marvin, J.G. et al (Marvin J, Akin, C M, Combined Effects of Mass Addition and No. blast on Boundary-Layer Transition, AIAA Journal, Vol.8, number 5, 1970, pp. 857-863. ss), wherein the theoretical solutions are researched by Marvin, J.G. et al through a consideration tableThe experimental data are measured by Marvin, J, G, and the like in an Ames 3.5-ft hypersonic wind tunnel. The calculation and experiment model is shown in fig. 4, and the working conditions of the calculation and experiment are shown in table 1. Fig. 5 shows the wall heat flow results calculated by this method compared to the wall heat flow results calculated and tested by Marvin, j.g. Adopting corresponding boundary conditions according to conditions during calculation, wherein the wall surface area without surface quality injection is set as the boundary condition of the viscous wall surface, and the wall temperature of the isothermal wall is constant
(ii) a The surface quality injection area adopts the model surface data processing method in the aerodynamic heat calculation method considering the surface quality injection effect, wherein the surface quality injection gas temperature
(ii) a Setting the surface mass injection gas constant as an air gas constant according to theoretical calculation and experimental corresponding conditions; the surface quality index gas mass rate is derived from a dimensionless number of injection rates, the injection rate being equal to
Wherein
The gas density is injected for the surface quality,
the surface quality is used for injecting the normal velocity of gas,
in order to obtain the density of the incoming gas flow,
in this embodiment, injection rates of 0% (i.e., no surface mass injection viscous wall surface) and 0 are selected for the incoming flow gas velocity according to experiments and calculations.04% and 0.08%. It can be seen from fig. 5 that the results are approximately the same and the trends are consistent.
TABLE 1 calculated and experimental operating conditions data
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The functionality of the present invention, if implemented in the form of software functional units and sold or used as a stand-alone product, may be stored in a computer readable storage medium. Based on such understanding, the technical solution of the present invention may be embodied in the form of a software product, which is stored in a storage medium, and all or part of the steps of the method according to the embodiments of the present invention are executed in a computer device (which may be a personal computer, a server, or a network device) and corresponding software. And the aforementioned storage medium includes: various media capable of storing program codes, such as a usb disk, a removable hard disk, or an optical disk, exist in a read-only Memory (RAM), a Random Access Memory (RAM), and the like, for performing a test or actual data in a program implementation.
Claims (9)
1. A method for calculating aerodynamic heat by considering surface quality injection effect is characterized by comprising the following steps:
s1, acquiring the geometric shape of the hypersonic aircraft;
s2, performing grid division on the geometric shape of the obtained hypersonic aircraft;
s3, acquiring hypersonic incoming flow speed data, hypersonic incoming flow temperature data, hypersonic incoming flow density data and hypersonic incoming flow pressure data, and inputting surface quality injection gas mass flow rate data and surface quality injection gas temperature data into a computer processor;
s4, calculating surface mass injection gas density data, surface mass injection gas velocity data, surface mass injection gas pressure data and surface mass injection gas temperature data in the computer processor based on the data in the step S3 and the surface mass injection gas constant data;
and S5, in the computer processor, substituting the data in the steps S1-S4 into an N-S equation to calculate to obtain wall surface heat flow data of the hypersonic flight, expressing the pneumatic thermal environment of the hypersonic flight vehicle through the wall surface heat flow data, and then presenting the pneumatic thermal environment based on a computer display unit.
2. A method for calculating aerodynamic heat considering surface mass induction effect according to claim 1, comprising the following sub-steps in step S2: and gradually drawing the grids corresponding to the geometric appearances of the hypersonic aircraft according to the sequence of point-to-line, line-to-surface and surface-to-body aiming at the appearance of the hypersonic aircraft in combination with working conditions, and then finishing grid division.
3. A method of calculating aerodynamic heat considering surface mass induction effect according to claim 1, comprising, in step S4, the sub-steps of:
s40, inputting the mass flow rate of the surface mass injection gas
Surface quality injection gas temperature
Surface mass ejection gas constant
And using the hypersonic speed incoming flow data, the hypersonic speed incoming flow temperature data, the hypersonic speed incoming flow density data and the hypersonic speed incoming flow pressure data acquired in the step S3,obtaining the pressure on a first layer grid point adjacent to the wall surface through flow field calculation
Density on first grid points adjacent to wall
Normal velocity at a first grid point adjacent to the wall
；
S41, calculating the surface quality injection gas density data according to the following formula
：
S42, surface quality injection gas density data obtained by calculation in the step S2 is utilized
And the surface mass injection gas mass flow rate input in step S3
The normal velocity of the surface-quality injected vertical jet gas is calculated according to the following formula
：
S43, injecting gas density data by using the obtained surface quality
And the surface quality injection gas temperature input in the step S3
Surface mass ejection gas constant
The surface quality injection gas pressure is calculated according to the following formula
：
4. The method of aerodynamic heating computation taking into account surface mass ejection effects of claim 1, wherein the hypersonic aerial vehicle comprises a spacecraft re-entry capsule.
5. The method of claim 1, wherein the hypersonic aerial vehicle comprises a blunt hypersonic re-entry missile.
6. The method of claim 1, wherein the hypersonic aerial vehicle comprises a near space hypersonic aerial vehicle.
7. A method of aerodynamic heating computation taking into account surface quality ejector effects as in claim 1, wherein the hypersonic aerocraft comprises a Mars and other planetary intakes.
8. Consideration surface according to claim 1The method for calculating the aerodynamic heat of the mass ejection effect is characterized in that in step S3, according to the trajectory of the known hypersonic vehicle, namely altitude data of the hypersonic vehicle at different moments and Mach number data of the hypersonic vehicle, a gas parameter table is searched to obtain hypersonic incoming flow speed data, hypersonic incoming flow temperature data, hypersonic incoming flow density data and hypersonic incoming flow pressure data; obtaining the coming current speed data of the hypersonic aerocraft according to the known hypersonic Mach number data and the hypersonic coming current speed data obtained by table lookup,
，
is the incoming flow speed of the hypersonic aerocraft,
The Mach number of the hypersonic aircraft,
Is a hypersonic speed of incoming flow.
9. An aerodynamic thermal calculation method according to claim 1, wherein in step S4, the surface mass ejection gas constant is obtained from a table of physical property parameters of the gas.
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| CN115636078B (en) * | 2022-10-25 | 2024-06-04 | 北京理工大学 | Hypersonic elastomer surface resistance and heat reduction method based on material ablation gas injection |
| CN116542116A (en) * | 2023-07-07 | 2023-08-04 | 北京航空航天大学 | Hypersonic aircraft full flow field numerical simulation method considering micropore gaseous sweating |
| CN116542116B (en) * | 2023-07-07 | 2023-09-01 | 北京航空航天大学 | Hypersonic aircraft full flow field numerical simulation method considering micropore gaseous sweating |
| CN117972279A (en) * | 2024-03-28 | 2024-05-03 | 中国空气动力研究与发展中心超高速空气动力研究所 | Method for calculating ball head standing point heat flow based on boundary layer theory |
| CN117972279B (en) * | 2024-03-28 | 2024-05-31 | 中国空气动力研究与发展中心超高速空气动力研究所 | Method for calculating ball head standing point heat flow based on boundary layer theory |
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