Simulation method and simulation system for thermal environment of ground equipment of carrier rocket launching site

Abstract

The invention provides a thermal environment simulation method and system for ground equipment of a carrier rocket launching site. The method comprises the steps that rocket engine jet pipe gas flow parameters are acquired; constructing simulation models of the rocket engine jet pipe, the ground equipment and the diversion trench; constructing a body grid in the simulation model, and importing the rocket engine jet pipe gas flow parameters and the simulation model into fluid simulation software for simulation; the simulation result is analyzed, and gas flow field data near the ground equipment are obtained. According to the method, the influence area range and strength of the gas flow on the ground equipment are obtained through simulation calculation, the area boundary where the high-speed and high-temperature gas flow possibly generates ablation and related parameters of ablation are obtained, and therefore a design reference basis is provided for subsequent thermal protection design.

Description

Technical Field

[0001] The present application relates to the field of thermal environment simulation technology, and in particular to a thermal environment simulation method and simulation system for ground equipment at a launch site for a carrier rocket. Background Art

[0002] In the prior art, the thermal environment simulation of the ground equipment of the launch vehicle launch site calculates the heat flux distribution at different positions of the ground gas flow field through the radiation angle coefficient, and it is impossible to accurately obtain the data of the ground heat flux. The existing patent publication number is CN117874918A, which discloses a method and system for simulating and analyzing the thermal environment of the multi-engine nozzle of the launch vehicle in parallel. The method includes: obtaining the simulation model of the rocket, the test bench and the single-sided guide groove, and meshing the simulation model; simulating and calculating the simulation model according to the jet parameters to obtain the simulation calculation results; loading the surrounding fluid material parameters to simulate and analyze the simulation calculation results to obtain the simulation analysis results, and determining the thermal environment data of the rocket, the single-sided guide groove and the test bench according to the simulation analysis results; judging whether the thermal environment data of the rocket, the single-sided guide groove and the test bench meet the design requirements. If they do not meet the design requirements, the thickness, width and height of the test bench, the profile of the single-sided guide groove, and the thermal protection scheme of the rocket are redesigned, and the values ​​are assigned to the simulation model for re-simulation calculation. The parallel thermal environment simulation analysis method and system for the engine nozzle mainly involve the process of obtaining thermal environment parameters, with little description of the specific methods. There is little description of the fluid mechanics professional analysis methods and software simulation operation procedures used in the simulation process, and factors such as takeoff drift and engine nozzle swing angle that need to be considered during the rocket launch are not included in the simulation system.

[0003] Therefore, the technical problem that needs to be solved urgently is: how to provide a simulation method and system for the thermal environment of ground equipment at a launch site for a carrier rocket, obtain the impact area and intensity of the gas flow on the ground equipment through simulation calculation, and obtain the area boundary and related parameters of ablation (temperature, flushing speed, pressure, heat flow) of the high-speed and high-temperature gas flow that may cause ablation. This provides a design reference for subsequent thermal protection design. Summary of the invention

[0004] The purpose of this application is to provide a method and system for simulating the thermal environment of ground equipment at a launch site for a carrier rocket, obtain the impact area and intensity of the gas flow on the ground equipment through simulation calculation, obtain the area boundary and related parameters of the ablation (such as temperature, flushing speed, pressure, heat flow, etc.) where high-speed and high-temperature gas flow may cause ablation, and thereby provide a design reference for subsequent thermal protection design.

[0005] To achieve the above-mentioned purpose, as the first aspect of the present application, the present application provides a method for simulating the thermal environment of ground equipment at a carrier rocket launch site, the method comprising: obtaining the gas flow parameters of a rocket engine nozzle; constructing a simulation model of the rocket engine nozzle, ground equipment and guide groove; constructing a volume grid in the simulation model, and importing the rocket engine nozzle gas flow parameters and the simulation model into a fluid simulation software for simulation; analyzing the simulation results to obtain the gas flow field data near the ground equipment.

[0006] The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket as described above, wherein the method further includes: determining whether the gas flow field data near the ground equipment meets the current design requirements; if not, modifying the ground equipment to improve the thermal protection capability; otherwise, there is no need to modify the ground equipment.

[0007] The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket as described above, wherein the method further comprises: reconstructing the volume grid of the modified ground equipment model and re-simulating.

[0008] The method for simulating the thermal environment of ground equipment at a carrier rocket launch site as described above, wherein the rocket engine nozzle gas flow parameters include combustion chamber pressure, total gas temperature, gas flow rate, gas specific heat capacity, gas thermal conductivity and / or gas viscosity.

[0009] In the method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket as described above, when constructing a volume grid in a simulation model, the density of the grid is automatically adjusted according to the gradient change of the gas flow field, and the grid is further refined in key areas.

[0010] In the method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket as described above, factors of takeoff drift and / or engine nozzle swing angle are incorporated into the simulation software of the simulation system.

[0011] In the method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket as described above, the gas flow field data includes flow velocity, temperature, pressure and / or heat flow data.

[0012] The method for simulating the thermal environment of ground equipment at a carrier rocket launch site as described above, wherein the ground equipment includes a fixed launch platform, a mobile launch platform and / or an erection frame.

[0013] As a second aspect of the present application, the present application provides a thermal environment simulation system for ground equipment at a launch site for a carrier rocket, which executes the thermal environment simulation method for ground equipment at a launch site for a carrier rocket, and the system includes:

[0014] An acquisition module is used to obtain the rocket engine nozzle gas flow parameters;

[0015] Construction module, used to construct simulation models of rocket engine nozzles, ground equipment and guide troughs;

[0016] The simulation module is used to construct a volume grid in the simulation model and import the rocket engine nozzle gas flow parameters and simulation model into the fluid simulation software for simulation;

[0017] The analysis module is used to analyze the simulation results and obtain the gas flow field data near the ground equipment.

[0018] The above-mentioned carrier rocket launch site ground equipment thermal environment simulation system, wherein the system also includes:

[0019] The judgment module is used to judge whether the gas flow field data near the ground equipment meets the current design requirements. If not, the ground equipment is modified to improve the thermal protection capability. Otherwise, there is no need to modify the ground equipment.

[0020] The beneficial effects achieved by this application are as follows:

[0021] (1) The method for simulating the gas flow field of ground equipment for rocket launch in this application takes into account a relatively comprehensive range of influencing factors and can more accurately predict the ground gas flow field during rocket launch. It not only provides guidance for the design of underground diversion troughs, making it safer and more efficient to conduct high-temperature gas flow, but also can test the thermal protection design of ground launch equipment to protect the safety of ground equipment. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in the present application. For those skilled in the art, other drawings can also be obtained based on these drawings.

[0023] Figure 1 This is a method flow chart of a method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to an embodiment of the present application.

[0024] Figure 2 This is a heat flux distribution diagram of the ground launch pad and rocket erection frame of an embodiment of the present application.

[0025] Figure 3 This is a total temperature cloud diagram of the ground equipment profile according to an embodiment of the present application.

[0026] Figure 4 It is a schematic diagram of the gas flow streamlines of the underground guide trough nozzle of an embodiment of the present application.

[0027] Figure 5 This is a cloud diagram of the total pressure on the wall of the underground diversion trough in an embodiment of the present application.

[0028] Figure 6 This is a schematic diagram of the air flow velocity of the nozzle gas flow impacting the ground launch platform in an embodiment of the present application.

[0029] Figure 7 This is a cloud diagram of the heat flux intensity on the surface of the ground launch platform according to an embodiment of the present application.

[0030] Figure 8 This is a schematic diagram of the structure of a thermal environment simulation system for ground equipment at a launch site for a carrier rocket according to an embodiment of the present application. DETAILED DESCRIPTION

[0031] The following is a clear and complete description of the technical solutions in the embodiments of the present application in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present application.

[0032] Embodiment 1

[0033] like Figure 1 As shown, the present application provides a method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket, the method comprising the following steps:

[0034] Step S1, obtaining the rocket engine nozzle gas flow parameters.

[0035] Among them, the rocket engine nozzle gas flow parameters include thermophysical properties such as combustion chamber pressure, total gas temperature, gas flow rate, gas specific heat capacity, gas thermal conductivity and / or gas viscosity.

[0036] Step S2, constructing simulation models of the rocket engine nozzle, ground equipment and guide grooves.

[0037] Among them, the ground equipment includes a fixed launch platform, a mobile launch platform and / or an erection frame, etc.

[0038] Step S3, constructing a volume grid in the simulation model, and importing the rocket engine nozzle gas flow parameters and the simulation model into the fluid simulation software for simulation.

[0039] As a specific embodiment of the present invention, when constructing a volume grid in a simulation model, an adaptive grid technology can be used to automatically adjust the density of the grid according to the gradient change of the gas flow field, and further refine the grid in key areas, so as to ensure both calculation accuracy and improve calculation efficiency. Key areas are, for example, the direct action area of ​​gas impacting ground equipment.

[0040] As a specific embodiment of the present invention, this application adopts a relatively high-resolution simulation grid to obtain sufficiently fine gas flow field data.

[0041] As a specific embodiment of the present invention, factors such as takeoff drift and engine nozzle swing angle are incorporated into the simulation software of the simulation system. In view of the takeoff drift factor, the drift trajectory and attitude change law of the rocket under different wind directions and wind speeds are studied. A comprehensive analysis system combining wind tunnel tests and numerical simulations is established to obtain the influence of takeoff drift on the distribution of the gas flow field. In view of the engine nozzle swing angle factor, the dynamic influence of the swing angle change rate on the gas flow field is analyzed, and the real-time changes in the gas injection direction and intensity of the engine nozzle under different swing angle motion states are simulated. The engine nozzle swing angle parameters are accurately integrated into the thermal environment, so that the simulation results more realistically reflect the complex thermal environment of the actual launch site.

[0042] As a specific embodiment of the present invention, the calculation in the simulation process needs to be based on the Reynolds-averaged Navier-Stokes equation, and the turbulent motion is regarded as the superposition of the time-averaged and instantaneous pulsating flows:

[0043]

[0044] Where, ρ represents density; t represents time; It represents the rate of change of density with time; u = (u, v, w) represents the fluid velocity vector, u, v, w represent the components of velocity in the x, y, z directions respectively; div(ρu) is the divergence term, which describes the net outflow rate of mass per unit volume.

[0045]

[0046] Wherein, μ represents dynamic viscosity; p represents pressure; represents the pressure gradient in the x direction; u′, v′, w′ represent the pulsating components of velocity (used to consider turbulence, etc.); S u represents the momentum source term in the x-direction; it may include the effects of external forces such as gravity in the x-direction; gradu represents the gradient of velocity u, which represents the diffusion effect of viscous force on the momentum in the x-direction.

[0047]

[0048] The meaning of each symbol is similar to the momentum equation in the x direction. v is the momentum source term in the y direction; represents the pressure gradient in the y direction. gradv represents the gradient of velocity v, which represents the diffusion effect of viscous force on momentum in the y direction.

[0049]

[0050] The meaning of each symbol is similar to the momentum equation in the x direction. w is the momentum source term in the z direction; represents the pressure gradient in the z direction. gradw represents the gradient of velocity w, which represents the diffusion effect of viscosity on the momentum in the z direction.

[0051] There are four main turbulence models used in simulation software, namely: k-Epsilon (i.e. k-ε model, which is a two-equation model in turbulence pattern theory, where k represents turbulent kinetic energy and ε represents the turbulent kinetic energy dissipation rate), k-Omega (i.e. k-ε model, which is a two-equation model in turbulence pattern theory, where k represents turbulent kinetic energy and ε represents the turbulent kinetic energy dissipation rate), Spalart-Allmaras (i.e. SA model, mainly used to solve the simulated transport equation of motion eddy viscosity) and Reynolds (also known as Reynolds equation, which is a partial differential equation obtained by analyzing the momentum balance of fluid flow. It describes the relationship between the shear stress, pressure gradient and fluid velocity to which the fluid is subjected during the flow process. By solving this equation, important information such as velocity distribution and pressure distribution of the fluid flow can be obtained. It is applicable to various fluid flow conditions such as laminar flow and turbulent flow) stress equation.

[0052] The k-Epsilon model includes two equations: turbulent kinetic energy and turbulent kinetic energy dissipation rate.

[0053] The turbulent kinetic energy equation is:

[0054]

[0055] The equation for the turbulent kinetic energy dissipation rate is:

[0056]

[0057] Where ρ represents the fluid density; k represents the turbulent kinetic energy (the intensity of the turbulent pulsation); t represents the time; u i Represents the velocity component in the i direction; i = 1, 2, 3; corresponding to the x, y, and z directions respectively; x i Indicates that the coordinate is in the i direction; x j Indicates that the coordinate is in the j direction; E ij The component representing the deformation rate (used to measure the deformation rate of fluid microgroups), μ t is the turbulence model (eddy) viscosity; σ k The turbulent Prandtl number, which represents the turbulent kinetic energy k, is an empirical constant used to correct the turbulent diffusion term; ε represents the dissipation rate of turbulent kinetic energy. σ ε The turbulent Prandtl number represents the rate of dissipation of turbulent kinetic energy ε. 1ε and C 2ε is an empirical constant used to adjust the importance of each term in the equation.

[0058] Specifically, the turbulent kinetic energy and the turbulent kinetic energy dissipation rate are calculated according to the turbulent kinetic energy equation and the turbulent kinetic energy dissipation rate equation.

[0059] Step S4, analyzing the simulation results to obtain gas flow field data near the ground equipment.

[0060] The gas flow field data includes flow velocity, temperature, pressure and / or heat flow data.

[0061] As another specific embodiment of the present invention, the simulation results are analyzed. In addition to obtaining flow rate, pressure and heat flow data, the dynamic change process of the gas flow field at different times is analyzed, and the transient characteristics of the thermal environment are analyzed. For example, the evolution of the thermal environment of ground equipment is observed at the initial stage of rocket takeoff, the acceleration stage and the end of engine operation.

[0062] Step S5, judging whether the gas flow field data near the ground equipment meets the current design requirements, if not, modifying the ground equipment to improve the thermal protection capability; reconstructing the volume mesh of the modified ground equipment model, and re-simulating.

[0063] Specifically, it is determined whether the gas flow field data of the ground equipment meets the current design requirements. If not, the ground equipment is modified to improve the thermal protection capability; the volume mesh of the modified ground equipment model is reconstructed and the simulation is performed again. If the gas flow field data of the ground equipment meets the current design requirements, there is no need to modify the ground equipment.

[0064] like Figure 2 The heat flux distribution diagram of the ground launch pad and the rocket erector is shown in Figure 1. The heat flux at the position of the erector close to the rocket nozzle is the largest and gradually decreases towards the surroundings. The heat flux at the ground launch pad is relatively small.

[0065] like Figure 3 The figure shows the total temperature cloud diagram of the ground equipment profile. After the rocket is ignited, the surface temperature of the ground equipment close to the engine rises sharply, and the temperature of the ground equipment corresponding to the engine nozzle is the highest, and the temperature gradually decreases towards the surrounding area.

[0066] like Figure 4 The figure shows the schematic diagram of the gas flow line of the underground guide groove nozzle. The high-temperature and high-speed gas ejected from the engine nozzle is introduced into the guide groove from the inlet of the guide groove. In the guide groove, the gas flow flows along the wall of the guide groove due to the restriction and guidance of the guide groove wall.

[0067] like Figure 5The figure shows the total pressure cloud diagram of the underground diversion trough wall. The high-temperature and high-speed gas ejected from the engine nozzle is introduced into the diversion trough from the inlet. The pressure at the diversion trough wall facing the direction of the high-temperature and high-speed gas ejected from the engine nozzle is the largest and gradually decreases toward the surrounding area.

[0068] like Figure 6 The figure shows the velocity of the gas flow from the nozzle impacting the ground launch platform. The maximum velocity of the gas flow at the tail of the nozzle is 2800m / s, which gradually decreases towards the surroundings.

[0069] like Figure 7 The figure shows the heat flux intensity cloud map on the surface of the ground launch platform. The surface of the ground launch platform forms multiple points with the highest heat flux density (Heat Flux) of 11 (W / m 2 ), and the heat flux density increases from the highest heat flux density point to 11 (W / m 2 ) gradually decreases towards the surrounding area.

[0070] Embodiment 2

[0071] like Figure 8 As shown, the present application provides a thermal environment simulation system 100 for ground equipment at a launch site for a carrier rocket, the system comprising:

[0072] The acquisition module 10 is used to obtain the rocket engine nozzle gas flow parameters.

[0073] The construction module 20 is used to construct simulation models of the rocket engine nozzle, ground equipment and guide grooves.

[0074] The simulation module 30 is used to construct a volume grid in the simulation model and import the rocket engine nozzle gas flow parameters and the simulation model into the fluid simulation software for simulation.

[0075] The analysis module 40 is used to analyze the simulation results and obtain the gas flow field data near the ground equipment.

[0076] The judgment module 50 is used to judge whether the gas flow field data near the ground equipment meets the current design requirements. If not, the ground equipment is modified to improve the thermal protection capability. Otherwise, there is no need to modify the ground equipment.

[0077] As other embodiments of the present invention. The present application also provides a computer storage medium, wherein the computer storage medium stores computer instructions, and when the computer instructions are called, they are used to execute the address mapping method of the large-capacity solid-state hard disk. The computer storage medium contains one or more program instructions, and the one or more program instructions are used by the processor to execute a method for simulating the thermal environment of ground equipment at a launch site of a carrier rocket.

[0078] The embodiments disclosed in the present invention provide a computer-readable storage medium, in which computer program instructions are stored. When the computer program instructions are executed on a computer, the computer executes the above-mentioned method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket.

[0079] An embodiment of the present invention provides a processor for processing the above-mentioned method for simulating the thermal environment of ground equipment at a launch site of a carrier rocket.

[0080] In the embodiment of the present invention, the processor may be an integrated circuit chip having the ability to process signals. The processor may be a general-purpose processor, a digital signal processor (DSP), an application-specific integrated circuit (ASIC), a field programmable gate array (FPGA), or other programmable logic devices, discrete gate or transistor logic devices, or discrete hardware components.

[0081] The methods, steps and logic block diagrams disclosed in the embodiments of the present invention can be implemented or executed. The general processor can be a microprocessor or the processor can also be any conventional processor, etc. The steps of the method disclosed in the embodiments of the present invention can be directly embodied as a hardware decoding processor for execution, or can be executed by a combination of hardware and software modules in the decoding processor. The software module can be located in a mature storage medium in the field such as a random access memory, a flash memory, a read-only memory, a programmable read-only memory or an electrically erasable programmable memory, a register, etc. The processor reads the information in the storage medium and completes the steps of the above method in combination with its hardware.

[0082] The storage medium may be a memory, which may be, for example, a volatile memory or a nonvolatile memory, or may include both volatile and nonvolatile memory.

[0083] Among them, the non-volatile memory can be a read-only memory (ROM), a programmable read-only memory (PROM), an erasable programmable read-only memory (EPROM), an electrically erasable programmable read-only memory (EEPROM), or a flash memory. The volatile memory can be a random access memory (RAM), which is used as an external cache. By way of example but not limitation, many forms of RAM are available, such as static random access memory (SRAM), dynamic random access memory (DRAM), synchronous dynamic random access memory (SDRAM), double data rate synchronous dynamic random access memory (DDRSDRAM), enhanced synchronous dynamic random access memory (ESDRAM), synchronous link dynamic random access memory (SLDRAM) and direct RAM bus random access memory (DRRAM).

[0084] The beneficial effects achieved by this application are as follows:

[0085] (1) The method for simulating the gas flow field of ground equipment for rocket launch in this application takes into account a relatively comprehensive range of influencing factors and can more accurately predict the ground gas flow field during rocket launch. It not only provides guidance for the design of underground diversion troughs, making it safer and more efficient to conduct high-temperature gas flow, but also can test the thermal protection design of ground launch equipment to protect the safety of ground equipment.

[0086] In the description of this application, the terms "first" and "second" are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the features. In the description of this application, "plurality" means two or more, unless otherwise clearly and specifically defined.

[0087] In the description of the present application, the word "for example" is used to mean "used as an example, illustration or explanation". Any embodiment described as "for example" in the present application is not necessarily to be construed as being more preferred or advantageous than other embodiments. The following description is given to enable any technician in the field to implement and use the present invention. In the following description, details are listed for the purpose of explanation. It should be understood that a person of ordinary skill in the art can recognize that the present invention can be implemented without using these specific details. In other examples, well-known structures and processes will not be elaborated in detail to avoid obscuring the description of the present invention with unnecessary details. Therefore, the present invention is not intended to be limited to the embodiments shown, but is consistent with the widest scope consistent with the principles and features disclosed in the present application.

[0088] The above description is only an embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent substitution, improvement, etc. made within the spirit and principle of the present invention should be included in the scope of the claims of the present invention.

Claims

1. A method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket, characterized in that: The method includes: Obtain rocket engine nozzle gas flow parameters; Construct simulation models of rocket engine nozzles, ground equipment, and flow channels; Construct a volume mesh in the simulation model, and import the rocket engine nozzle gas flow parameters and simulation model into the fluid simulation software for simulation; The simulation results are analyzed to obtain the gas flow field data near the ground equipment.

2. The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to claim 1, characterized in that: The method also includes: determining whether the gas flow field data near the ground equipment meets the current design requirements; if not, modifying the ground equipment to improve the thermal protection capability; otherwise, there is no need to modify the ground equipment.

3. The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to claim 2, characterized in that: The method also includes: reconstructing the volume grid of the modified ground equipment model and re-simulating.

4. The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to claim 1, characterized in that: The rocket engine nozzle gas flow parameters include combustion chamber pressure, total gas temperature, gas flow rate, gas specific heat capacity, gas thermal conductivity and / or gas viscosity.

5. The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to claim 1, characterized in that: When constructing the volume mesh in the simulation model, the density of the mesh is automatically adjusted according to the gradient change of the gas flow field, and the mesh is further refined in the key area.

6. The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to claim 1, characterized in that: The take-off drift and / or engine nozzle swing angle factors are incorporated into the simulation software of the simulation system.

7. The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to claim 1, characterized in that: The gas flow field data includes flow velocity, temperature, pressure and / or heat flow data.

8. The method for simulating the thermal environment of ground equipment at a launch site for a carrier rocket according to claim 1, characterized in that: The ground equipment includes a fixed launching platform, a mobile launching platform and / or an erecting frame.

9. A thermal environment simulation system for ground equipment at a launch site for a carrier rocket, characterized in that: The method according to any one of claims 1 to 8 is implemented, the system comprising: An acquisition module is used to obtain the rocket engine nozzle gas flow parameters; Construction module, used to construct simulation models of rocket engine nozzles, ground equipment and guide troughs; The simulation module is used to construct a volume grid in the simulation model and import the rocket engine nozzle gas flow parameters and simulation model into the fluid simulation software for simulation; The analysis module is used to analyze the simulation results and obtain the gas flow field data near the ground equipment.

10. The thermal environment simulation system for ground equipment at a launch site for a carrier rocket according to claim 9, characterized in that: The system also includes: The judgment module is used to judge whether the gas flow field data near the ground equipment meets the current design requirements. If not, the ground equipment is modified to improve the thermal protection capability. Otherwise, there is no need to modify the ground equipment.

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