CN115935852A - Aerodynamic thermal engineering calculation method and system based on aircraft ballistic data - Google Patents
Aerodynamic thermal engineering calculation method and system based on aircraft ballistic data Download PDFInfo
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Abstract
The application discloses an aerodynamic thermal engineering calculation method and system based on aircraft ballistic data, wherein the aerodynamic thermal engineering calculation method based on the aircraft ballistic data specifically comprises the following steps: inputting the ballistic data; calculating parameters of the atmospheric free flow; judging whether shock waves are generated or not; if the shock wave is generated, setting the shape of the shock wave; calculating wave-rear airflow parameters corresponding to the shape of the shock wave; judging whether the gas is high-temperature gas or not according to the wave-rear gas flow parameters; if the gas is high-temperature gas, correcting the high-temperature gas; and calculating the heat flux density of the surface of the aircraft. The method and the device rely on the ballistic data, can provide the development condition of the aerodynamic heat in the whole flight state, and are convenient for searching aerodynamic heat severe points and performing aerodynamic heat analysis on the whole flight process; in the method, an atmospheric model is used for calculating atmospheric parameters, so that the data volume needing to be input is reduced; the shock wave shape is considered, so that the applicable working condition is wider, and the gas high-temperature effect is considered, so that the applicable working condition is wider.
Description
Technical Field
The application relates to the field of rockets, in particular to a pneumatic thermal engineering calculation method and a pneumatic thermal engineering calculation system based on aircraft ballistic data.
Background
When the aircraft flies at a high speed, high-speed airflow passes through the surface, the surface temperature of the aircraft is obviously improved due to the blocking effect of the viscosity of the air, and the aerodynamic heat problem becomes a problem which must be considered in the aerodynamic and structural design of the aircraft. In the initial stage of the development of the hypersonic aircraft, the pneumatic thermal engineering algorithm plays a great role because the engineering algorithm has high calculation efficiency and certain guarantee on precision. A flow chart of a conventional computing method is shown in fig. 1. In the input working condition, incoming flow atmospheric parameters, flight altitude, flight Mach number, simplified appearance parameters of the aircraft and the like need to be input, and the requirement is complex. Therefore, in the traditional pneumatic heat calculation method, more input parameters are needed, and the pretreatment is complicated; in the calculation process, the calculation working condition is single, batch processing cannot be carried out, and the calculation workload is large; when the airflow is calculated in a traditional mode, the change of a state equation of the air caused by the temperature change is generally not considered, so that only the working condition with a small Mach number can be calculated, and only the condition that the flight attack angle is equal to zero can be considered, and in the actual condition, the flight attack angle can cause the hot current density of the windward side to be larger; and the mode is not combined with a flight trajectory, so that the maximum aerodynamic heat flow and the overall heat flow development condition in the whole flight are difficult to express.
Therefore, how to provide a pneumatic thermal engineering calculation method which enables the flight calculation result to be more accurate is a problem which needs to be solved urgently by the technical personnel in the field.
Disclosure of Invention
The application provides an aerodynamic heat engineering calculation method based on aircraft ballistic data, which specifically comprises the following steps: inputting the ballistic data; responding to the input ballistic data, and calculating the parameters of the free flow of the atmosphere; judging whether shock waves are generated or not in response to the completion of the calculation of the atmospheric free flow parameters; if the shock wave is generated, setting the shape of the shock wave; calculating wave-rear airflow parameters corresponding to the shape of the shock wave; judging whether the gas is high-temperature gas or not according to the wave-rear gas flow parameters; if the gas is high-temperature gas, correcting the high-temperature gas; and calculating the heat flux density on the surface of the aircraft after finishing the correction of the high-temperature gas.
As above, if no shock wave is generated, it is directly determined that the gas is not a high-temperature gas.
The method is characterized in that if the gas is non-high-temperature gas, the heat flow on the surface of the aircraft is directly calculated.
As above, among others, the atmospheric free stream parameter calculation was performed using the Yang Bingwei atmospheric model.
As above, wherein setting the shape of the shock wave includes setting the shock wave shape to a normal shock wave if the aircraft is a large blunt and short aircraft; if the aircraft is a sharp or slender aircraft, the shock shape is set to oblique shock.
A pneumatic thermal engineering computing system based on aircraft ballistic data specifically comprises: the device comprises an input unit, an atmospheric free flow parameter calculation unit, a shock wave judgment unit, a shock wave shape setting unit, a backward airflow parameter calculation unit, a high-temperature gas judgment unit, a correction unit and a calculation unit; the input unit is used for inputting the ballistic data; the atmospheric free flow parameter calculation unit is used for responding to the input ballistic data and calculating the atmospheric free flow parameter; the shock wave judging unit is used for responding to the completion of the calculation of the atmospheric free flow parameters and judging whether shock waves are generated or not; the shock wave shape setting unit is used for setting the shape of the shock wave if the shock wave is generated; the wave-backward airflow parameter calculating unit 350 is configured to calculate a wave-backward airflow parameter corresponding to the shape of the shock wave; the high-temperature gas judging unit is used for judging whether the gas is high-temperature gas or not according to the wave-rear gas flow parameters; a correction unit for performing on the high-temperature gas if the gas is the high-temperature gas; and the calculating unit is used for calculating the surface heat flow.
As described above, if the shock wave determining unit determines that no shock wave is generated, the shock wave determining unit directly determines that the gas is not a high-temperature gas.
In the above, if the high-temperature gas judgment unit judges that the high-temperature gas is not the high-temperature gas, the calculation unit directly calculates the heat flow on the surface of the aircraft.
As above, the atmospheric free flow parameter calculation unit performs the atmospheric free flow parameter calculation using the Yang Bingwei atmospheric model.
As above, wherein the shock wave shape setting unit sets the shape of the shock wave to include that if the aircraft is a large blunt and short aircraft, the shock wave shape is set to a normal shock wave; if the aircraft is a sharp or slender aircraft, the shock shape is set to oblique shock.
The application has the following beneficial effects:
the method and the system for calculating the aerodynamic thermal engineering based on the aircraft ballistic data rely on the ballistic data, can provide the aerodynamic thermal development condition in the whole flight state, and are convenient for searching aerodynamic thermal severe points and performing aerodynamic thermal analysis on the whole flight process; in the method, an atmospheric model is used for calculating atmospheric parameters, so that the data volume needing to be input is reduced; the shock wave shape is considered, so that the applicable working condition is wider, and the gas high-temperature effect is considered, so that the applicable working condition is wider.
Drawings
In order to more clearly illustrate the embodiments of the present application 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, it is obvious that the drawings in the description below are only some embodiments described in the present application, and other drawings can be obtained by those skilled in the art according to these drawings.
FIG. 1 is a flow chart of a prior art method of performing pneumatic thermal engineering calculations;
FIG. 2 is a flow chart of a method of aerodynamic thermal engineering calculations based on aircraft ballistic data provided in accordance with an embodiment of the present application;
FIG. 3 is an internal block diagram of a pneumatic thermal engineering computing system based on aircraft ballistic data provided in accordance with an embodiment of the present application;
FIG. 4 is a diagram of a case of generating normal shock waves according to an embodiment of the present application;
fig. 5 is a case of generating oblique shock waves according to an embodiment of the present application.
Detailed Description
The technical solutions in the embodiments of the present application are clearly and completely described below with reference to the drawings in the embodiments of the present application, and it is obvious that the described embodiments are some, but not all, embodiments of the present application. All other embodiments obtained by a person skilled in the art based on the embodiments in the present application without making any creative effort belong to the protection scope of the present application.
According to the method, through improvement of an engineering calculation method, the missile path data of the aircraft are input as initial data, so that the pneumatic heat flow of the characteristic points and the characteristic positions of the aircraft in the whole flight can be calculated at the same time, the development conditions of the heat flow along with the time, the flight altitude and other parameters can be clearly shown, and the worst pneumatic heat condition can be found and discussed conveniently. Meanwhile, the high-temperature condition is considered, and when the flow field gas flow parameters are calculated, the high-temperature gas effect is introduced, the thermodynamic property and the transport property of the high-temperature gas effect are corrected, so that the calculation result is more accurate and is suitable for higher flight speed.
Example one
As shown in fig. 1, a method for calculating aerodynamic thermal engineering based on aircraft ballistic data according to an embodiment of the present application specifically includes the following steps:
step S210: and inputting the ballistic data.
Specifically, the trajectory data specifically includes flight time, flight altitude, and flight mach number (or flight speed).
The method is based on the ballistic data, can provide the aerodynamic heat development condition in the whole flight state, and is convenient for searching aerodynamic heat severe points and performing aerodynamic heat analysis on the whole flight process.
Step S220: in response to the input ballistic data, atmospheric free-flow parameter calculations are performed.
The atmosphere free flow parameters are specifically calculated through a Yang Bingwei atmosphere model, and the Yang Bingwei atmosphere model can obtain the atmosphere free flow parameters at the height through height calculation.
Wherein the atmospheric free stream parameters comprise of atmospheric free stream: density, viscosity coefficient, sound velocity, pressure, aircraft flight mach number, total temperature, total pressure, enthalpy, entropy, dynamic pressure, and the like.
The atmospheric free flow parameter calculated by the method is compared with an international standard atmospheric table, the error is within five thousandths, and the calculation of the atmospheric free flow parameter is more accurate.
The Yang Bingwei atmospheric model is used for calculating the atmospheric free flow parameters, and the data volume needing to be input is reduced.
Step S230: and judging whether shock waves are generated or not according to the atmospheric free flow parameters.
Specifically, the atmospheric free flow parameters include the flight mach number of the aircraft, so whether shock waves are generated or not is judged according to the flight mach number of the aircraft.
The shock wave generation standard is set to generate shock waves if the flying Mach number of the aircraft is greater than 1; and if the flight Mach number of the aircraft is less than 1, no shock wave is generated.
If the shock wave is generated, step S240 is performed. If no shock wave is generated, it is directly determined that the gas is not a high temperature gas, and step S280 is directly performed.
Step S240: the shape of the shock wave is set.
The determination of the shock wave shape depends on a simplified model of the aircraft and needs to be set by self.
Specifically, for a large blunt and very short aircraft, the shock shape is set to normal shock; for sharp or slender aircraft, the shock shape is set to oblique shock.
After the shock wave shape is set, step S250 is performed.
Step S250: and calculating the wave-rear airflow parameters corresponding to the shape of the shock wave.
Wherein, the calculation of the wave-rear parameters is carried out according to the set normal shock wave or oblique shock wave. The calculation of the wave-rear parameters comprises the calculation of the front-rear pressure ratio, the density ratio, the temperature ratio and the total enthalpy value of the normal shock wave and the oblique shock wave.
Wherein subscript 1 represents incoming flow, subscript 2 represents wave-rear, P is pressure, ρ is density, T is temperature, V is velocity, ma 1 The Mach number of incoming flow is shown, gamma represents specific heat ratio, and the value of common air is 1.4.
The calculation of the wave-rear parameters when a normal laser is: when hypersonic incoming flow passes through the bluff body, a normal shock wave (a shock wave perpendicular to the direction of the airflow) is generated at the aircraft head, and the shape of the shock wave is shown in fig. 4. .
Thus according to a continuous methodEquation of stroke, momentum and energy, and the pressure ratio before and after normal shock wave can be derived
Density ratio->
Temperature ratio>
And so on.
The calculation of the wave-rear parameters when being a ramp laser is: if the direction of the airflow after the wave is not vertical to the shock wave surface, the shock wave is called oblique shock wave. The shape of the oblique shock wave is shown in fig. 5, and the oblique shock wave can be seen as a plane shock wave formed by superposing a flow field of a normal shock wave, a flow field parallel to a shock wave surface and a straight uniform current field parallel to the shock wave surface. Shock angle β: the included angle formed by the shock wave surface and the incoming flow airflow direction; airflow deflection angle θ: the included angle between the air flow direction of the wave back and the incoming flow direction of the wave front.
Similarly, the pressure intensity ratio before and after normal shock wave can be derived according to a continuous equation, a momentum equation and an energy equation
Density ratio->
Temperature ratio>
When the deflection angle theta of the airflow and the Mach number Ma of the incoming flow 1 When known, the shock angle β can be determined by the following equation:
wherein the total enthalpy value h can be obtained by the formula: h =1008.4t + 0.5V, where T denotes the temperature of the wave front (or wave back) and V denotes the speed.
Step S260: and judging whether the gas is high-temperature gas or not according to the wave-backward gas flow parameters.
Whether the gas is high-temperature gas or not is judged according to the wave-back parameters obtained by calculation according to the generated corresponding shock wave shapes, specifically, whether the gas is high-temperature gas or not is judged according to the total enthalpy value in the wave-back parameters, and when the total enthalpy value h is larger than a certain characteristic value (the characteristic value can be set to 167500), the gas is judged to be high-temperature gas.
If it is a high temperature gas, step S270 is performed. If not, step S280 is performed.
Step S270: and correcting the high-temperature gas.
Specifically, for the sake of simple calculation and accuracy assurance, the high-temperature gas is corrected in the following manner: according to a high-temperature air thermodynamic function table of the Soviet Union in the prior art, a fitting formula with the error smaller than 10% of the original table data is calculated, and the fitting formula is suitable for correcting the thermodynamic characteristics and the transportation mode of high-temperature gas.
After the correction of the high temperature gas is completed, step S280 is performed.
Step S280: and calculating the heat flux density of the surface of the aircraft.
The surface heat flux density comprises the aircraft stagnation surface heat flux density, an oblique cone with an attack angle and a flat plate surface heat flux density.
In particular, where the stagnation point resides is the point at which the forward most end of the aircraft is in a state of stagnation. Stagnation point surface heat flux q ws The concrete expression is as follows:
subscript w represents a wall parameter; subscript s denotes post wave parameters, pr and Le denote dimensionless parameters, where Pr =0.71, le =1.0; h is d For average air enthalpy, the value is 15419.98KJ/kg; h is s Is the stagnation point enthalpy. ρ w represents the wall surface density.
Gradient of stagnation point velocity
Can be obtained from the modified newton equation:
wherein R is 0 Is a hemispherical radius; p S The total pressure after wave; rho S Is the density of the stationary points after the wave; p ∞ Is the incoming hydrostatic pressure.
Wherein the heat flux q of the stagnation point surface ws It can be specifically expressed as:
where ρ is sl ＝1.225kg/m Incense stick ；V e =7900m/s, ∞ represents the incoming flow, and V ∞ represents the incoming flow speed.
The heat flux density q of the surface of the stationary point can be calculated by any one of the above modes ws 。
Specifically, for the calculation of the surface heat flux density of an oblique cone, a flat plate and the like with an attack angle, an axisymmetric analogy method and an equivalent cone method can be adopted.
Generally, the equivalent cone method works well for heat flow distribution on the windward bus. Therefore, in this embodiment, an equivalent cone method is used to calculate the surface heat flux density of the oblique cone, the flat plate, and the like with the angle of attack.
The principle of the equivalent cone method is as follows: the flow around the object with the cone of attack is replaced by a flow around the object with an equivalent cone of zero angle of attack flow, the equivalent cone being a function of the angle of attack and the circumferential angle.
The equivalent taper angle θ TC is a function of the angle of attack and the angle of the cone, and is calculated as follows:
wherein α is the angle of attack; theta.theta. c Is a conical angle;
the meridian angle.
The equivalent cone angle is introduced, so that the heat flow with the attack angle can be changed into the heat flow with the cone surface with the zero attack angle.
Example two
As shown in fig. 3, the system for calculating aerodynamic thermal engineering based on aircraft ballistic data according to the embodiment of the present application specifically includes: input means 310, atmospheric free flow parameter calculation means 320, shock wave determination means 330, shock wave shape setting means 340, post-wave airflow parameter calculation means 350, high-temperature gas determination means 360, correction means 370, and calculation means 380.
Wherein the input unit 310 is used for inputting the ballistic data.
The method is based on the ballistic data, can provide the aerodynamic heat development condition in the whole flight state, and is convenient for searching aerodynamic heat severe points and performing aerodynamic heat analysis on the whole flight process.
The atmospheric free flow parameter calculation unit 320 is configured to perform atmospheric free flow parameter calculation in response to the input ballistic data.
The atmosphere free flow parameters are specifically calculated through a Yang Bingwei atmosphere model, and the Yang Bingwei atmosphere model can obtain the atmosphere free flow parameters at the height through height calculation.
The atmospheric free flow parameter calculated by the method is compared with an international standard atmospheric table, the error is within five thousandths, and the calculation of the atmospheric free flow parameter is more accurate.
The Yang Bingwei atmospheric model is used for calculating the atmospheric free flow parameters, and the data volume needing to be input is reduced.
The shock wave determining unit 330 is configured to determine whether a shock wave is generated in response to completion of calculation of the atmospheric free flow parameter.
Specifically, the shock wave generation standard is set such that if the mach number is greater than 1, a shock wave is generated; if the mach number is less than 1, no shock wave is generated, and the high temperature gas determination unit 350 is executed.
The shock wave shape setting unit 340 is used for setting the shape of the shock wave if the shock wave is generated.
The judgment of the shock wave shape depends on a simplified model of the aircraft and needs to be set by self.
Specifically, for a large blunt and very short aircraft, the shock shape is set to a normal shock; for sharp or slender aircraft, the shock shape is set to oblique shock.
The wave-backward airflow parameter calculating unit 350 is configured to calculate a wave-backward airflow parameter corresponding to the shape of the shock wave.
Wherein, the calculation of the wave-rear parameters is carried out according to the set normal shock wave or oblique shock wave. The calculation of the wave-rear parameters comprises the calculation of the front-rear pressure ratio, the density ratio, the temperature ratio and the total enthalpy value of the normal shock wave and the oblique shock wave.
Wherein subscript 1 represents incoming flow, subscript 2 represents wave-rear, P is pressure, ρ is density, T is temperature, V is velocity, ma 1 The Mach number of incoming flow is shown, gamma represents specific heat ratio, and the value of common air is 1.4.
The calculation of the wave-rear parameters when a normal laser is: when hypersonic incoming flow passes through the bluff body, a normal shock wave (a shock wave perpendicular to the direction of the airflow) is generated at the aircraft head, and the shape of the shock wave is shown in fig. 4. .
Therefore, according to a continuous equation, a momentum equation and an energy equation, the pressure intensity ratio before and after the normal shock wave can be deduced
Density ratio->
Temperature ratio>
And so on.
The calculation of the wave-rear parameters when being a ramp laser is: if the direction of the airflow after the wave is not vertical to the shock wave surface, the shock wave is called oblique shock wave. The oblique shock wave can be seen as a planar shock wave formed by superposing a flow field of a normal shock wave, a flow field parallel to a shock wave surface and a direct uniform flow field parallel to the shock wave surface, as shown in fig. 5. Shock angle β: the included angle formed by the shock wave surface and the incoming flow airflow direction; airflow deflection angle θ: the included angle between the air flow direction of the wave back and the incoming flow direction of the wave front.
Similarly, the pressure intensity ratio before and after normal shock wave can be derived according to a continuous equation, a momentum equation and an energy equation
Density ratio (X;) device>
Temperature ratio>
When the deflection angle theta of the airflow and the Mach number Ma of the incoming flow 1 When known, the shock angle β can be determined by:
wherein the total enthalpy value h can be obtained by the formula: h =1008.4t + 0.5V, where T denotes the temperature of the wave front (or wave back) and V denotes the speed.
The high temperature gas determination unit 360 is configured to determine whether the gas is a high temperature gas according to the atmospheric free flow parameter.
If the gas is not a high temperature gas, the calculation unit 380 is executed.
The correcting unit 370 corrects the high-temperature gas if the gas is a high-temperature gas.
Specifically, for the sake of simple calculation and accuracy assurance, the high-temperature gas is corrected in the following manner: according to a high-temperature air thermodynamic function table of the Soviet Union, calculating a fitting formula with the error smaller than 10% from the original table data, and correcting the thermodynamic characteristics and the transportation mode of the high-temperature gas by using the fitting formula.
The calculation unit 380 is used to calculate the surface heat flux density.
Parked is the point at which the forward most end of the aircraft is in a stagnation state. Stagnation point surface heat flux q ws The concrete expression is as follows:
subscript w represents a wall parameter; subscript s denotes post wave parameters, pr and Le denote dimensionless parameters, where Pr =0.71, le =1.0; h is d For average air enthalpy, the value is 15419.98KJ/kg; h is s Is the stagnation point enthalpy. ρ w represents the wall surface density.
Gradient of stagnation point velocity
Can be obtained from the modified newton equation:
wherein R is 0 Is a hemispherical radius; p S The total pressure after wave; rho S Is the density of the stationary points after the wave; p ∞ Is the incoming hydrostatic pressure.
Wherein the stagnation point surface heat flux q ws It can be specifically expressed as:
where ρ is sl ＝1.225kg/m Incense stick ；V e =7900m/s, ∞ represents the incoming flow, and V ∞ represents the incoming flow velocity.
The heat flux density q of the surface of the stationary point can be calculated by any one of the above modes ws 。
Specifically, for the calculation of the surface heat flux density of an oblique cone, a flat plate and the like with an attack angle, an axisymmetric analogy method and an equivalent cone method can be adopted.
Generally, the equivalent cone method works well for heat flow distribution on the windward bus. Therefore, in this embodiment, the equivalent cone method is used to calculate the surface heat flux density of an oblique cone with an attack angle, a flat plate, and the like.
The principle of the equivalent cone method is as follows: the flow around the object with the cone of attack is replaced by a flow around the object with an equivalent cone of zero angle of attack flow, the equivalent cone being a function of the angle of attack and the circumferential angle.
The equivalent cone angle θ TC is a function of the angle of attack and the cone angle, and is calculated as follows:
wherein α is the angle of attack; theta.theta. c Is a conical angle;
the meridian angle.
The equivalent cone angle is introduced, so that the heat flow with the attack angle can be changed into the heat flow of the cone surface with the zero attack angle.
The application has the following beneficial effects:
the method and the system for calculating the aerodynamic thermal engineering based on the aircraft ballistic data rely on the ballistic data, can provide the aerodynamic thermal development condition in the whole flight state, and are convenient for searching aerodynamic thermal severe points and performing aerodynamic thermal analysis on the whole flight process; in the method, an atmospheric model is used for calculating atmospheric parameters, so that the data volume needing to be input is reduced; the shock wave shape is considered, so that the applicable working condition is wider, and the gas high-temperature effect is considered, so that the applicable working condition is wider.
Although the present application has been described with reference to examples, which are intended to be illustrative only and not to be limiting of the application, changes, additions and/or deletions may be made to the embodiments without departing from the scope of the application.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any person skilled in the art can easily conceive of the changes or substitutions within the technical scope of the present application, and shall be covered by the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (10)
1. A pneumatic thermal engineering calculation method based on aircraft ballistic data is characterized by comprising the following steps:
inputting the ballistic data;
responding to the input ballistic data, and calculating the parameters of the free flow of the atmosphere;
judging whether shock waves are generated or not in response to the completion of the calculation of the atmospheric free flow parameters;
if the shock wave is generated, setting the shape of the shock wave;
calculating the wave-rear airflow parameters corresponding to the shape of the shock wave;
judging whether the gas is high-temperature gas or not according to the wave-rear gas flow parameters;
if the gas is the high-temperature gas, correcting the high-temperature gas;
and calculating the heat flux density of the surface of the aircraft after the correction of the high-temperature gas is completed.
2. A method as claimed in claim 1, wherein if no shock wave is generated, it is directly determined as a non-high temperature gas.
3. A method as claimed in claim 2, wherein the heat flow on the surface of the aircraft is directly calculated if the gas is not a high temperature gas.
4. An aerodynamic thermal engineering calculation method based on aircraft ballistic data as claimed in claim 1 wherein the atmospheric free-flow parameter calculation is performed using a Yang Bingwei atmosphere model.
5. An aerodynamic thermal engineering calculation method based on aircraft ballistic data as claimed in claim 4 wherein setting the shape of the shock wave comprises setting the shock wave shape to a normal shock wave if the aircraft is a large blunt and short aircraft; if the aircraft is a sharp or slender aircraft, the shock shape is set to oblique shock.
6. A pneumatic thermal engineering computing system based on aircraft ballistic data is characterized by specifically comprising: the device comprises an input unit, an atmospheric free flow parameter calculation unit, a shock wave judgment unit, a shock wave shape setting unit, a backward airflow parameter calculation unit, a high-temperature gas judgment unit, a correction unit and a calculation unit;
the input unit is used for inputting the ballistic data;
the atmospheric free flow parameter calculation unit is used for responding to the input ballistic data and calculating the atmospheric free flow parameter;
the shock wave judging unit is used for responding to the completion of the calculation of the atmospheric free flow parameters and judging whether shock waves are generated or not;
the shock wave shape setting unit is used for setting the shape of the shock wave if the shock wave is generated;
the wave-rear airflow parameter calculating unit 350 is configured to calculate a wave-rear airflow parameter corresponding to the shape of the shock wave;
the high-temperature gas judging unit is used for judging whether the gas is high-temperature gas or not according to the wave-rear gas flow parameters;
a correction unit for performing on the high-temperature gas if the gas is the high-temperature gas;
and the calculating unit is used for calculating the surface heat flow.
7. An aerodynamic thermal engineering calculation system based on aircraft ballistic data as claimed in claim 6 wherein the shock wave determination unit determines directly as a non-high temperature gas if it determines that no shock wave is generated.
8. The air-thermal engineering calculation system based on aircraft ballistic data of claim 7, wherein the high-temperature gas judgment unit judges that the gas is not a high-temperature gas, and the calculation unit directly calculates the heat flow on the surface of the aircraft.
9. An aerodynamic thermal engineering calculation system based on aircraft ballistic data as claimed in claim 6 wherein the free-air flow parameter calculation unit performs free-air flow parameter calculations using a Yang Bingwei atmosphere model.
10. An aerodynamic thermal engineering calculation system based on aircraft ballistic data according to claim 6 wherein the shock shape setting unit sets the shock shape to a normal shock if the aircraft is a large blunt and short aircraft; if the aircraft is a sharp or slender aircraft, the shock shape is set to oblique shock.
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