Rapid prediction method for jet flow thermal environment

Abstract

The invention discloses a rapid prediction method for a jet flow thermal environment. The method comprises the following steps: firstly, calculating by utilizing a chemical equilibrium model to obtain characteristic parameters such as total enthalpy, equivalent gas constant and the like of fuel gas of a nozzle outlet based on fuel gas components and contents of the nozzle outlet of an engine; secondly, on the basis of the assumed condition that the total enthalpy of the nozzle outlet and the total enthalpy of the nozzle throat are equal, gas characteristic parameters of the nozzle throat part are obtained through the obtained equivalent gas characteristic parameters of the nozzle outlet; then, carrying out surface grid division on the cross section of the nozzle throat, and setting jet flow boundary conditions at the position; then, a far-field computational domain is selected, a closed space is formed by the far-field computational domain, a spray pipe and equipment, and a volume grid of a mixed structure is generated; and finally, solving the N-S equation by using a CFD calculation program to simulate a low-backpressure jet flow field in the space environment simulation equipment and obtain a thermal environment of key equipment, thereby providing input conditions for rapid evaluation of a heat protection system in a design stage, and reducing the test cost of the space environment simulation equipment.

Description

Technical Field

[0001] The present invention relates to the technical field of aerodynamic design of high-altitude and ultra-high-speed aircraft, and particularly to a rapid prediction method for jet thermal environment. Background Art

[0002] The vacuum thermal environment in which a spacecraft operates in orbit is complex and variable. In order to ensure that each component on the spacecraft can withstand the influence of this environment and work properly throughout the operation process, it is necessary to conduct thermal tests to assess the thermal structure design of the spacecraft and verify the thermal characteristics of the components. The space environment simulation test equipment is an important ground test equipment in the spacecraft environment simulation project. Its main function is to simulate orbital space environments such as vacuum, cold black, and solar radiation, expose problems in spacecraft launch and in-orbit operation in advance, and provide important guarantees for the development and in-orbit stable operation of the spacecraft. However, the thermal vacuum test has some limitations: ① The thermal vacuum test requires a long period and high cost; ② The number of test temperature measurement points is limited, and the sensor layout position is restricted by the surface shape and size of the test piece. Usually, it is impossible to measure the temperature at the concerned points inside the parts.

[0003] In recent years, with the rapid development of numerical simulation technology, researchers at home and abroad have widely used the DSMC method to calculate the engine jet flow field and study the thermal effect of the engine wake on sensitive equipment. The research results show that the DSMC method can accurately simulate the engine vacuum plume flow field, but this calculation method has high requirements for computer capabilities, calculation costs, etc. Summary of the Invention

[0004] The present invention provides a rapid prediction method for jet thermal environment, which can quickly evaluate the thermal effect of the engine jet on sensitive equipment in the space environment simulation equipment and reduce the test cost.

[0005] In a first aspect, a rapid prediction method for jet thermal environment is provided, including:

[0006] Based on the gas components and their contents at the engine nozzle exit, the total enthalpy, total pressure, and equivalent gas characteristic parameters at the nozzle exit are calculated using the chemical equilibrium model;

[0007] Based on the assumption that the total enthalpy and total pressure of the gas at the nozzle exit and the nozzle throat are equal, the gas characteristic parameters at the nozzle throat part are obtained using the equivalent gas characteristic parameters at the nozzle exit as the jet boundary condition for flow field calculation;

[0008] The cross-section of the nozzle throat is meshed, and the jet boundary condition is set at this position;

[0009] A far-field calculation domain is selected to form a closed space with the nozzle and the equipment, and surface meshes are generated;

[0010] Set the surface grid density, control the grid growth rate, and generate a volume grid with a hybrid structure;

[0011] Use a CFD calculation program to solve the Navier-Stokes equations to simulate the low-backpressure jet flow field in the space environment simulation device and obtain the thermal environment of the key equipment.

[0012] Combined with the first aspect, in some implementation manners of the first aspect, the equivalent gas characteristic parameters include the mixture gas constant R and the equivalent specific heat ratio γ; where:

[0013]

[0014] R u is the universal gas constant, M i is the molecular weight of each gas component at the nozzle exit, C i is the mass fraction of each gas component at the nozzle exit, γ i is the specific heat ratio of each gas component at the nozzle exit.

[0015] Combined with the first aspect, in some implementation manners of the first aspect, the method uses the Mach number Ma = 1 at the nozzle throat, the mixture gas constant R, and the equivalent specific heat ratio γ to calculate the gas characteristic parameters at the nozzle throat part as the jet boundary condition for the flow field calculation:

[0016] h 0jet = h 0exit

[0017] p 0jet = p 0exit

[0018] p jet = p 0jet / (1 + 0.5×(γ - 1)) γ / γ-1

[0019] T jet = (h 0jet / Cp) / (1 + 0.5×(γ - 1))

[0020] v jet = sqrt(γRT jet )

[0021] where, h 0exit is the total enthalpy at the nozzle exit, h 0jet is the total enthalpy at the nozzle throat, p 0exit is the total pressure at the nozzle exit, p 0jet is the total pressure at the nozzle throat, p jet is the static pressure of the nozzle throat cross-section, T jet is the static temperature of the nozzle throat cross-section, Cp is the mixture gas specific heat at constant pressure, v jetis the cross-sectional velocity of the nozzle throat.

[0022] Combined with the first aspect, in some implementations of the first aspect, a hybrid grid is divided for the cross-section of the nozzle throat.

[0023] Combined with the first aspect, in some implementations of the first aspect, the grid generation rate of the cross-section of the nozzle throat is controlled to be 1.1 - 1.2, the grid scale at the junction of the jet surface and the nozzle side is 1E-2m - 1E-5m; the axial grid density is 21 - 41.

[0024] Combined with the first aspect, in some implementations of the first aspect, after the cross-sectional grid of the nozzle throat is generated, the boundary conditions of the pressure, density, and temperature of the jet are set at the cross-sectional grid nodes as the input conditions for the flow field calculation.

[0025] Combined with the first aspect, in some implementations of the first aspect, the method selects a vacuum tank as the far-field calculation domain, forms a closed space with the nozzle and the equipment, and generates a surface grid; in the vacuum test environment, the surface of the vacuum tank is set as the symmetry plane boundary condition.

[0026] Combined with the first aspect, in some implementations of the first aspect, when generating a spatial hybrid structure grid in the closed body, a hexahedral grid is generated in the space near the surface.

[0027] Combined with the first aspect, in some implementations of the first aspect, at least 2 layers of hexahedral grids are filled in all surface boundary layers, and non-hexahedral grids are filled outside the hexahedral grids.

[0028] Combined with the first aspect, in some implementations of the first aspect, the method uses the test space environment to simulate the equipment internal environment as the initial calculation field. The space environment simulation equipment belongs to the continuous flow region in the jet core area. The Navier-Stokes equation is solved numerically for the engine jet flow field to obtain the thermal environment of the key equipment.

[0029] Compared with the prior art, the solution provided by the present invention at least includes the following beneficial technical effects:

[0030] (1) A rapid prediction method for the jet thermal environment provided by the present invention adopts a hybrid grid generation technology, which can ensure the number of grids and meet the solution accuracy;

[0031] (2) A rapid prediction method for the jet thermal environment provided by the present invention uses the chemical equilibrium method to obtain the nozzle outlet parameters as the calculation conditions, which can greatly improve the calculation efficiency;

[0032] (3) The rapid prediction method for jet heat environment provided by the present invention can numerically simulate the low back-pressure full flow field of the environmental simulation device, effectively reducing the ground test cost, having great economic value, and providing a new idea and strategy for the future ground test verification of aircraft. Description of the Drawings

[0033] Figure 1 It is a calculation flow chart.

[0034] Figure 2 It is a schematic diagram of the environmental simulation device.

[0035] Figure 3 It is a schematic diagram of the grid structure.

[0036] Figure 4 It is a heat cloud diagram of the surface of the key equipment in the environmental simulation device.

[0037] Figure 5 It is a schematic diagram of the distribution of measuring points on the surface of the key equipment in the environmental simulation device.

[0038] Figure 6 It is a schematic diagram of the comparison between the heat flux on the surface of the key equipment in the environmental simulation device and the test results. Detailed Embodiment

[0039] The present invention will be further described in detail below with reference to the drawings and specific embodiments.

[0040] As Figure 1 shown, the present invention provides a rapid prediction method for jet heat environment, and the specific steps are as follows.

[0041] Step (1), based on the gas components and their contents at the engine nozzle exit, use the chemical equilibrium model to calculate the total enthalpy, total pressure and equivalent gas characteristic parameters at the nozzle exit.

[0042] The equivalent gas characteristic parameters include the mixture gas constant R and the equivalent specific heat ratio γ; based on the gas component contents at the engine nozzle exit, parameters such as the total enthalpy, total pressure, mixture gas constant R and equivalent specific heat ratio γ at the nozzle exit are obtained based on the chemical equilibrium method; where:

[0043] γ = ∑C i γ i

[0044] R u is the universal gas constant (8314 J / kmol / K), M i is the molecular weight of each gas component at the nozzle exit, C i is the mass fraction of each gas component at the nozzle exit, γ i is the specific heat ratio of each gas component at the nozzle exit.

[0045] Step (2), based on the assumption that the total enthalpy and total pressure of the gas at the nozzle exit and the nozzle throat are equal, the gas characteristic parameters at the nozzle throat are obtained by using the equivalent gas characteristic parameters at the nozzle exit obtained above, and used as the jet boundary condition for the flow field calculation.

[0046] Based on the condition that the total enthalpy and total pressure of the nozzle exit and the nozzle throat are equal, using Ma = 1 at the nozzle throat, the mixed gas constant R and the equivalent specific heat ratio γ, the gas pressure, temperature, velocity and other characteristic parameters of the nozzle throat cross-section are calculated as the jet boundary condition for the flow field calculation:

[0047] h 0jet = h 0exit

[0048] p 0jet = p 0exit

[0049] p jet = p 0jet / (1 + 0.5×(γ - 1)) γ / γ-1

[0050] T jet = (h 0jet / Cp) / (1 + 0.5×(γ - 1))

[0051] v jet = sqrt(γRT jet )

[0052] Among them, h 0exit is the total enthalpy at the nozzle exit, h 0jet is the total enthalpy at the nozzle throat, p 0exit is the total pressure at the nozzle exit, p 0jet is the total pressure at the nozzle throat, p jet is the static pressure of the nozzle throat cross-section, T jet is the static temperature of the nozzle throat cross-section, Cp is the specific heat at constant pressure of the mixed gas, v jet is the velocity of the nozzle throat cross-section.

[0053] Step (3), perform mesh division on the nozzle throat cross-section and set the jet boundary condition at this position.

[0054] Perform a hybrid mesh division on the nozzle throat cross-section, encrypt it near the nozzle wall surface, and at the same time, it is necessary to ensure the quality of the spatial mesh and the calculation efficiency. In one embodiment, control the surface mesh generation rate to be 1.1 - 1.2, the mesh scale at the junction of the jet surface and the nozzle side surface is 1E-2 m - 1E-5 m; the axial mesh density is 21 - 41.

[0055] After the surface mesh is generated, as the surface jet condition, boundary conditions such as the pressure, density, and temperature of the jet are set at the cross-section mesh nodes as the input conditions for the flow field calculation.

[0056] Step (4): Select the far-field calculation domain, form a closed space with the nozzle and the equipment, and generate the spatial surface mesh.

[0057] Select the vacuum tank as the far-field calculation domain, form a closed space with the nozzle and the equipment, and generate the spatial surface mesh. In the vacuum test environment, set the surface of the vacuum tank as the symmetric plane boundary condition.

[0058] Step (5): Set the surface mesh density, control the mesh growth rate, and generate the volume mesh with a hybrid structure.

[0059] When generating the spatial hybrid structure mesh in the closed body, to ensure the accurate simulation of the thermal environment on the jet surface, hexahedral meshes need to be generated in the space near the surface. At the same time, the spatial mesh growth rate needs to be controlled to ensure the accuracy and efficiency of the numerical calculation.

[0060] When generating the spatial hybrid structure mesh in the closed body, at least two layers of hexahedral meshes are filled in all surface boundary layers, and non-hexahedral meshes (such as tetrahedrons and pyramids) are filled outside the hexahedral meshes to reduce the number of meshes and improve the calculation efficiency.

[0061] Step (6): Use the CFD calculation program to solve the Navier-Stokes (N-S) equation to simulate the low-backpressure jet flow field in the space environment simulation equipment and obtain the thermal environment of the key equipment.

[0062] Taking the environment in the test space environment simulation equipment as the initial calculation field, numerically solve the engine jet flow field to obtain the thermal environment of the key equipment. The space environment simulation equipment belongs to the continuous flow region in the jet core area, and the thermal impact of the jet on the key equipment can be evaluated by solving the Navier-Stokes equation. This can provide input conditions for the rapid evaluation of the thermal protection system in the design stage and reduce the test cost of the space environment simulation equipment.

[0063] Embodiment:

[0064] The following combines examples to illustrate the specific implementation of the present invention.

[0065] As a specific design example of a rapid prediction method for the jet thermal environment. In this embodiment, the space simulation equipment is a large-capacity vacuum tank, and the aircraft is placed in the center of the tank, such as Figure 2As shown in the figure. The throat diameter of the engine nozzle is 11 mm, and the outlet diameter is 80 mm. The nozzle parameters are shown in Table 1 and Table 2. According to the description in this article, through chemical equilibrium calculation, it can be obtained that the equivalent gas constant of the engine nozzle throat is 401 J / kg / K, the equivalent gas specific heat ratio is 1.22, the outlet total pressure is 908388 Pa, and the outlet total temperature is 2583 K. In this way, the calculation conditions for the nozzle throat (Mach number = 1) can be obtained: static pressure 509249 Pa, static temperature 2327 K.

[0066] The space simulation device generates a hybrid grid as Figure 3 shown, and the jet surface and the device surface are meshed with increased density.

[0067] The calculation results are as Figure 4 shown. It can be seen from Figure 4 that the heat flux on the surface of the key device is concentrated in the position below the engine nozzle. The peak value of the heat flux is about 2.2 MW / m 2 , compared with the DSMC calculation results, the difference between the two is about 9.1%, as shown in Table 3, but the calculation efficiency is significantly improved.

[0068] The comparison of the heat flux on the surface of the key device with the test results is as Figure 6 (the test measurement points are as Figure 5 shown). It can be seen that the calculation results are in good agreement with the heat flux results of the test measurement points. The general deviation is about 2%, and the maximum deviation is about 20%, meeting the requirements of engineering development.

[0069] Table 1 Engine operating parameters

[0070]

[0071] Table 2 Engine outlet gas composition (mole fraction)

[0072] <![CDATA[CO 2 > <![CDATA[H 2 O]]> <![CDATA[H 2 > <![CDATA[N 2 > CO 0.080 0.31 0.23 0.33 0.05

[0073] Table 3 Comparison of the peak heat flux calculation results of the key device surface in the environmental simulation device with the DSMC results

[0074] Calculation method <![CDATA[Heat flux / (kW / m 2 )]]> Computing resource CPU The calculation method in this paper 2.2 144 CPUs, 3 days DSMC 2.4 1440 CPUs, 7 days (The calculation method in this paper - DSMC) / The calculation method in this paper % 9.1 /

[0075] Although the present invention is disclosed above with preferred embodiments, it is not used to limit the present invention. Any person skilled in the art can make possible changes and modifications without departing from the spirit and scope of the present invention. Therefore, the protection scope of the present invention should be determined by the scope defined by the claims of the present invention.

Claims

1. A rapid prediction method for jet thermal environment, characterized in that: include: Based on the gas composition and content at the engine nozzle outlet, the total enthalpy, total pressure and equivalent gas characteristic parameters at the nozzle outlet are calculated using the chemical equilibrium model. Based on the assumption that the total enthalpy and total pressure of the gas at the nozzle outlet and the nozzle throat are equal, the gas characteristic parameters at the nozzle throat are obtained using the equivalent gas characteristic parameters at the nozzle outlet as the jet boundary conditions for flow field calculation. Mesh the nozzle throat cross section and set the jet boundary conditions at this location; Select the far-field calculation domain to form a closed space with the nozzle and equipment, and generate the surface mesh; Set surface mesh density, control mesh growth rate, and generate volume meshes for hybrid structures; The CFD calculation program is used to solve the Navier-Stokes equations to simulate the low back pressure jet flow field in the space environment simulation equipment and obtain the thermal environment of key equipment.

2. The method according to claim 1, characterized in that The equivalent gas characteristic parameters include the mixed gas constant R and the equivalent specific heat ratio γ; where: R u is the universal gas constant, M i is the molecular weight of each gas component at the nozzle outlet, C i is the mass fraction of each gas component at the nozzle outlet, γ i is the specific heat ratio of each gas component at the nozzle outlet.

3. The method according to claim 2, characterized in that The method uses the nozzle throat Ma=1 and the mixed gas constant R and the equivalent specific heat ratio γ to calculate the gas characteristic parameters at the nozzle throat as the jet boundary conditions for flow field calculation: h 0jet =h 0exit p 0jet =p 0exit p jet =p 0jet / (1+0.5×(γ-1)) γ / γ- 1 T jet =(h 0jet / Cp) / (1+0.5×(γ-1)) v jet =sqrt(γRT jet ) Among them, h 0exit is the total enthalpy at the nozzle outlet, h 0jet is the total enthalpy at the nozzle throat, p 0exit is the total pressure at the nozzle outlet, p 0jet is the total pressure at the nozzle throat, p jet is the static pressure of the nozzle throat cross section, T jet is the static temperature of the nozzle throat cross section, Cp is the specific heat of the mixed gas at constant pressure, v jet is the nozzle throat cross-sectional velocity.

4. The method according to claim 1, characterized in that: Hybrid meshing is performed on the nozzle throat cross section.

5. The method according to claim 4, characterized in that The grid generation rate of the nozzle throat cross section is controlled to be 1.1-1.2, the grid scale at the junction of the jet surface and the nozzle side is 1E-2m-1E-5m; the axial grid density is 21-41.

6. The method according to claim 1, characterized in that After the nozzle throat cross-section grid is generated, the method sets the boundary conditions of the pressure, density and temperature of the jet on the cross-section grid nodes as the input conditions for flow field calculation.

7. The method according to claim 1, characterized in that The method selects a vacuum tank as a far-field calculation domain, forms a closed space with the nozzle and the equipment, and generates a spatial surface mesh; in a vacuum test environment, the surface of the vacuum tank is set as a symmetric surface boundary condition.

8. The method according to claim 1, characterized in that: When generating a spatial hybrid structured mesh in a closed volume, a hexahedral mesh is generated in the space near the surface.

9. The method according to claim 8, characterized in that All surface boundary layers are filled with at least 2 layers of hexahedral meshes, and non-hexahedral meshes are filled outside the hexahedral meshes.

10. The method according to claim 1, characterized in that The method uses the internal environment of the test space environment simulation device as the initial calculation field. The space environment simulation device belongs to the continuous flow area in the jet core area. The engine jet flow field is numerically solved by solving the Navier-Stokes equation to obtain the thermal environment of the key equipment.

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