A method for engineering analysis of booster interference aerodynamic heating

Abstract

The application discloses a method for analyzing interference aerodynamic heat of a booster, comprising the following steps: obtaining air flow parameters at different time according to flight trajectory calculation results; calculating shock wave angles of a booster nose at different time according to the air flow parameters at different time; determining a shock wave direct impact area, a first area and a second area according to the calculated shock wave angles of the booster nose; simplifying the first area into a nose stagnation point and performing boundary layer outer edge parameter calculation and cold wall heat flux density calculation; simplifying the second area into a conical surface and performing boundary layer outer edge parameter calculation and cold wall heat flux density calculation; and forming a bundled launch vehicle conical booster interference aerodynamic heating environment according to the boundary layer outer edge parameter calculation results and the cold wall heat flux density calculation results. The application can quickly provide a round of interference aerodynamic heating calculation results in the launch vehicle demonstration, scheme initial stage or multi-trajectory iteration process, so as to carry out missile body heat protection design and evaluate load and strength problems.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of carrier rockets, and particularly relates to a booster interference aerodynamic heat engineering analysis method. BACKGROUND

[0002] When a carrier rocket flies at a high speed in the air, friction with the atmosphere occurs, and aerodynamic heating phenomenon is generated. The cold wall heat flux of the rocket body surface changing with time obtained through aerodynamic heating calculation can provide a basis for the structural design and heat protection design of the model. For a single core stage carrier rocket, there are relatively mature engineering calculation methods for stagnation points, spherical surfaces, conical surfaces, cylindrical surfaces and the like. However, for a bundled carrier rocket, during high-speed flight, the booster and the core stage interfere with each other, and the interference shock wave formed has a secondary flushing effect on the rocket body, which can seriously affect the aerodynamic heating state of the shock wave interference region, and the interference has different shock wave angles at different times in the entire ascent stage. The existing aerodynamic heating for bundled interference generally adopts a numerical analysis method, which needs to establish a large number of grid models, and a large amount of computing resources and time are required for each working condition, which cannot meet the needs of rapid iteration of configurations and trajectories, and cannot analyze different times of the full trajectory. SUMMARY

[0003] The technical problem of the present application is to overcome the shortcomings of the prior art and provide a booster interference aerodynamic heat engineering analysis method, which can quickly provide a round of interference aerodynamic heating calculation in the carrier rocket demonstration, early scheme or multi-trajectory iteration process, for developing rocket body heat protection design and evaluating load and strength problems.

[0004] In order to solve the above technical problems, the present application discloses a booster interference aerodynamic heat engineering analysis method, comprising:

[0005] According to the flight trajectory calculation results, the air inflow parameters at different times are obtained;

[0006] According to the air inflow parameters at different times, the shock wave angle of the booster spherical head at different times is calculated;

[0007] According to the calculated shock wave angle of the booster spherical head, the shock wave direct impact region, the first region and the second region are determined; wherein the first region is a region 500mm above and below the shock wave direct impact region; the second region is a region 500mm-1000mm above and below the shock wave direct impact region;

[0008] The first region is simplified as a spherical head stagnation point, and the boundary layer outer edge parameter calculation and cold wall heat flux density calculation are performed;

[0009] The second region is simplified as a conical surface, and the boundary layer outer edge parameter calculation and cold wall heat flux density calculation are performed;

[0010] According to the calculation results of the boundary layer outer edge parameters and the calculation results of the cold wall heat flux density, the interference aerodynamic heating environment of the bundled launch vehicle conical booster is formed.

[0011] In the above-mentioned booster interference aerodynamic heating engineering analysis method, the air inflow parameters include: the time-varying inflow gas Mach number M ∞ , the inflow gas temperature T ∞ , the inflow gas pressure P ∞ , and the inflow gas density ρ ∞ .

[0012] In the above-mentioned booster interference aerodynamic heating engineering analysis method, the shock angle θ s of the booster nose at different times is calculated by the following formula:

[0013] θ s =θ c {1.093+[0.06(7-θ c ·M ∞ ) / (θ c ·M ∞ ) 3 / 2 ]}

[0014] Wherein, θ c represents the half-cone angle of the booster nose.

[0015] In the above-mentioned booster interference aerodynamic heating engineering analysis method, when the first region is simplified as the ball head stagnation point, the boundary layer outer edge parameters are calculated, and the calculated boundary layer outer edge parameters include: the boundary layer gas temperature Ts, the boundary layer gas pressure Ps, the boundary layer gas density ρ S , the boundary layer gas viscosity μ S and the boundary layer gas velocity V s .

[0016] In the above-mentioned booster interference aerodynamic heating engineering analysis method, the solving formulas of the boundary layer gas temperature Ts, the boundary layer gas pressure Ps, the boundary layer gas density ρ S , the boundary layer gas viscosity μ S and the boundary layer gas velocity V s are as follows:

[0017]

[0018]

[0019]

[0020]

[0021]

[0022] wherein R represents air gas constant, R0 represents core cylinder segment radius.

[0023] In the above-mentioned booster interference aerodynamic heat engineering analysis method, when the first region is simplified as a spherical head stagnation point, and the cold wall heat flux density is calculated, the calculation formula of the cold wall heat flux density q r is as follows:

[0024]

[0025] wherein Pr represents gas Prandtl number, C p represents other constant pressure specific heat capacity.

[0026] In the above-mentioned booster interference aerodynamic heat engineering analysis method, when the second region is simplified as a conical surface, and the boundary layer outer edge parameters are calculated, the calculated boundary layer outer edge parameters include: conical surface boundary layer gas temperature T e , conical surface boundary layer gas density ρ e , conical surface boundary layer gas velocity V e and conical surface boundary layer gas pressure P e .

[0027] In the above-mentioned booster interference aerodynamic heat engineering analysis method, the calculation formula of the conical surface boundary layer gas temperature T e , the conical surface boundary layer gas density ρ e , the conical surface boundary layer gas velocity V e and the conical surface boundary layer gas pressure P e is as follows:

[0028]

[0029]

[0030]

[0031]

[0032] wherein γ represents gas specific heat ratio, P2 represents post-shock pressure, C p represents other constant pressure specific heat capacity.

[0033] In the above-mentioned booster interference aerodynamic heat engineering analysis method, the calculation formula of the post-shock pressure P2 is as follows:

[0034]

[0035] In the above-mentioned booster interference aerodynamic heat engineering analysis method, when the second region is simplified as a conical surface, and the cold wall heat flux density is calculated, the calculation formula of the cold wall heat flux density q e is as follows:

[0036]

[0037] wherein c f represents the wall friction coefficient.

[0038] The present application has the following advantages:

[0039] The application discloses a booster interference aerodynamic heat engineering analysis method. BRIEF DESCRIPTION OF DRAWINGS

[0040] Figure 1 is a step flow chart of a booster interference aerodynamic heat engineering analysis method in the embodiment of the present application;

[0041] Figure 2 is a layout diagram of booster shock wave interference of a certain bundled rocket in the embodiment of the present application. DETAILED DESCRIPTION

[0042] In order to make the purpose, technical scheme and advantages of the present application clearer, the following will further describe the disclosed embodiments of the present application in combination with the drawings.

[0043] As Figure 1 In the embodiment, the booster interference aerodynamic heat engineering analysis method comprises:

[0044] Step 1: obtaining air inflow parameters at different time according to flight trajectory calculation results.

[0045] In the embodiment, the air inflow parameters include but are not limited to: inflow gas Mach number M ∞ , inflow gas temperature T ∞ , inflow gas pressure P ∞ and inflow gas density ρ ∞ , etc.

[0046] Step 2: calculating shock wave angle of the booster ball head at different time according to the air inflow parameters at different time.

[0047] In the embodiment, the shock wave angle θ s of the booster ball head at different time can be calculated through the following formula:

[0048] θ c = θ c {1.093 + [0.06 (7- θ ∞ · M c ) / (θc M ∞ ) 3 / 2 ]}

[0049] wherein, θ c represents the half-cone angle of the boost nose.

[0050] As shown in Table 1, the Mach number M ∞ of the incoming airflow at different time is given, and the shock angle θ s of the boost nose at different time is calculated.

[0051] Time (s) Mach number Shock angle 61.9 1.5 48.1471° 72.1 2.0 36.97° 80.2 2.5 31.715° 87.4 3.0 28.71° 93.7 3.5 26.805° 100.4 4.0 25.52° 107.6 4.5 24.61° 119.3 5.0 23.941°

[0052] Table 1, the change of the shock angle θ s with the Mach number M ∞ of the incoming airflow is shown.

[0053] Step 3, according to the calculated shock angle of the boost nose, the directly impacted region, the first region and the second region are determined.

[0054] In this embodiment, after the directly impacted region is determined, the region 500mm above and below the directly impacted region is regarded as the first region, and the region 500mm-1000mm above and below the directly impacted region is regarded as the second region. As shown in Figure 2 , region A is the directly impacted region (also called the stagnation point region), region B is the first region, and region C is the second region.

[0055] Step 4, the first region is simplified as the stagnation point of the nose, and the calculation of the outer edge parameters of the boundary layer and the calculation of the cold wall heat flux density are carried out.

[0056] In this embodiment, when the first region is simplified as the stagnation point of the nose, the calculated outer edge parameters of the boundary layer include but are not limited to: the boundary layer gas temperature Ts, the boundary layer gas pressure Ps, the boundary layer gas density ρ S , the boundary layer gas viscosity μ S and the boundary layer gas velocity V s .

[0057] Preferably, the solving formula of the boundary layer gas temperature Ts, the boundary layer gas pressure Ps, the boundary layer gas density ρ S , the boundary layer gas viscosity μ S and the boundary layer gas velocity V s is as follows:

[0058]

[0059]

[0060]

[0061]

[0062]

[0063] wherein, R represents air gas constant, R0 represents core cylinder segment radius.

[0064] As shown in the following table 2, the boundary layer outer edge parameters calculated by step 4 are given:

[0065] Time (s) Ts (K) Ps (Pa) S (kg / m 3 )]]> ​ 61.9 314.1 62823.1 0.6966 72.1 390.8 50159.2 0.4471 80.2 491.7 37147.6 0.2631 87.4 625.9 25018.8 0.1392 93.7 786.4 16086.1 0.0712 100.4 1018.7 8821.9 0.0301 107.6 1327.9 4182.4 0.0109 119.3 1522.1 988.7 0.0022

[0066] Table 2, the boundary layer outer edge parameters calculated by step 4 are given:

[0067] Further, when the first region is simplified as a spherical head stagnation point, the cold wall heat flux density q r is calculated, and the calculation formula of the cold wall heat flux density q p is as follows:

[0068]

[0069] wherein, Pr represents gas Prandtl number, C p represents other constant pressure specific heat capacity.

[0070] As shown in the following table 3, the cold wall heat flux density q r calculated by step 4 is given:

[0071] Time (s) q r (kW / m 2 )]]> 61.9 31.34 72.1 36.86 80.2 42.2 87.4 44.38 93.7 45.68 100.4 44.38 107.6 41.72 119.3 19.9

[0072] Table 3, the cold wall heat flux density calculated by step 4 is given:

[0073] Step 5, the second region is simplified as a conical surface, and the boundary layer outer edge parameter calculation and the cold wall heat flux density calculation are carried out.

[0074] In this embodiment, when the second region is simplified as a conical surface, the boundary layer outer edge parameters calculated by the boundary layer outer edge parameter calculation include: the conical surface boundary layer gas temperature T e , the conical surface boundary layer gas density ρ e , the conical surface boundary layer gas velocity V e and the conical surface boundary layer gas pressure P e .

[0075] Preferably, the calculation formula of the conical surface boundary layer gas temperature T e , the conical surface boundary layer gas density ρ e , the conical surface boundary layer gas velocity V e and the conical surface boundary layer gas pressure P e is as follows:

[0076]

[0077]

[0078]

[0079]

[0080] wherein γ represents the gas specific heat ratio, and P2 represents the pressure behind the shock wave.

[0081] The calculation formula of the pressure P2 behind the shock wave is as follows:

[0082]

[0083] The boundary layer outer edge parameters calculated by step 5 are shown in the following table 4:

[0084] Time (s) T e (K)]]> e (kg / m 3 )]]> ​ V e (m / s) <!-- 5 -->]]> 61.9 255.6 0.76626 343.8 72.1 271.4 0.49181 491.0 80.2 293.3 0.28941 633.0 87.4 323.9 0.15312 781.1 93.7 358.9 0.07832 929.2 100.4 416.0 0.03311 1103.3 107.6 492.2 0.01199 1299.2 119.3 519.1 0.00242 1423.3

[0085] Table 4, the cold wall heat flux density calculation table by step 5

[0086] Further, when the second region is simplified as a conical surface, the cold wall heat flux density q e is calculated, and the calculation formula is as follows:

[0087]

[0088] wherein c f represents the wall friction coefficient.

[0089] The cold wall heat flux density q e calculated by step 5 is shown in the following table 5:

[0090] Time (s) q e (kW / m 2 )]]> 61.9 43.4 72.1 43.7 80.2 41.2 87.4 37.1 93.7 31.96 100.4 25.1 107.6 17.7 119.3 6.3

[0091] Table 5, the cold wall heat flux density calculation table by step 5

[0092] Step 6, according to the calculation results of the boundary layer outer edge parameters and the cold wall heat flux density, the interference aerodynamic heating environment of the bundled launch vehicle conical booster is formed.

[0093] In this embodiment, the interference aerodynamic heating environment of the bundled launch vehicle conical booster can be formed according to the calculation results of step 4 and step 5. For example, on the basis of table 3 and table 5, a factor of 1.2 can be considered as the interference aerodynamic heating environment condition, as shown in table 6:

[0094] Time (s) Resident point cold wall heat flow (kW / m 2 ) Conical surface cold wall heat flux (kW / m 2 )]]> 61.9 37.61 52.08 72.1 44.23 52.44 80.2 50.64 49.44 87.4 53.26 44.52 93.7 54.82 38.35 100.4 53.26 30.12 107.6 50.06 21.24 119.3 23.88 7.56

[0095] Table 6. Interfering aerothermodynamic heating environment conditions

[0096] Although the present application has been disclosed with reference to the preferred embodiments, it is not intended to limit the present application to the preferred embodiments. Any person skilled in the art, without departing from the spirit and scope of the present application, can make possible changes and modifications to the technical solutions of the present application by using the disclosed methods and technical contents. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application, without departing from the technical solutions of the present application, shall fall within the protection scope of the technical solutions of the present application.

[0097] The contents not described in detail in the specification of the present application are known to the person skilled in the art.

Claims

1. A method for analyzing the aerodynamic and thermal engineering aspects of booster interference, characterized in that, include: Based on the flight trajectory calculation results, the airflow parameters at different times are obtained; Based on the airflow parameters at different times, the shock wave angle of the booster ball head at different times is calculated. Based on the calculated shock wave angle of the booster ball head, the direct impact area of ​​the shock wave, the first area, and the second area are determined; wherein, the first area is the area within 500mm above and below the direct impact area of ​​the shock wave; the second area is the area within 500mm to 1000mm above and below the direct impact area of ​​the shock wave. The first region is simplified to a spherical head stagnation point, and boundary layer outer edge parameters and cold wall heat flux density are calculated. The second region is simplified into a cone, and boundary layer outer edge parameters and cold wall heat flux density are calculated. Based on the calculation results of the boundary layer outer edge parameters and the cold wall heat flux density, an aerodynamic heating environment is formed that interferes with the launch vehicle's conical booster.

2. The booster interference aerodynamic and thermal engineering analysis method according to claim 1, characterized in that, Airflow parameters, including: the time-varying Mach number of the incoming gas. Incoming gas temperature Incoming gas pressure and incoming gas density .

3. The booster interference aerodynamic and thermal engineering analysis method according to claim 2, characterized in that, The shock wave angle of the booster ball at different times is calculated using the following formula. : in, This indicates the semi-cone angle of the ball head that assists the push.

4. The booster interference aerodynamic and thermal engineering analysis method according to claim 3, characterized in that, When simplifying the first region to a spherical stagnation point and calculating the boundary layer outer edge parameters, the calculated boundary layer outer edge parameters include: boundary layer gas temperature. Boundary layer gas pressure Boundary layer gas density Boundary layer gas viscosity and boundary layer gas velocity .

5. The booster interference aerodynamic and thermal engineering analysis method according to claim 4, characterized in that, Boundary layer gas temperature Boundary layer gas pressure Boundary layer gas density Boundary layer gas viscosity and boundary layer gas velocity The solution formula is as follows: in, Represents the gas constant of air. This indicates the radius of the core stage section.

6. The booster interference aerodynamic and thermal engineering analysis method according to claim 5, characterized in that, When simplifying the first region to a stagnation point at the spherical head and calculating the cold wall heat flux density, the cold wall heat flux density... The calculation formula is as follows: in, Represents the Prandtl number of the gas. This indicates other isobaric specific heat capacities.

7. The booster interference aerodynamic and thermal engineering analysis method according to claim 6, characterized in that, When simplifying the second region into a cone and calculating the boundary layer outer edge parameters, the calculated boundary layer outer edge parameters include: the cone-shaped boundary layer gas temperature. Conical boundary layer gas density Conical boundary layer gas velocity and conical boundary layer gas pressure .

8. The booster interference aerodynamic and thermal engineering analysis method according to claim 7, characterized in that, Conical boundary layer gas temperature Conical boundary layer gas density Conical boundary layer gas velocity and conical boundary layer gas pressure The solution formula is as follows: in, Indicates the specific heat ratio of a gas. Indicates the pressure after the shock wave. This indicates other isobaric specific heat capacities.

9. The booster interference aerodynamic and thermal engineering analysis method according to claim 8, characterized in that, Post-shock pressure The solution formula is as follows: 。 10. The booster interference aerodynamic and thermal engineering analysis method according to claim 9, characterized in that, When simplifying the second region into a conical surface and calculating the cold wall heat flux density, the cold wall heat flux density... The calculation formula is as follows: in, This represents the coefficient of friction of the wall surface.

Images