A numerical simulation method and system for reverse jet flow of reusable launch vehicle

Abstract

The present invention discloses a numerical simulation method and system for a reverse jet of a reusable carrier rocket. The method comprises: establishing a nozzle-related grid model; using the nozzle throat as the inlet boundary for jet calculation and determining other relevant boundary conditions of the nozzle; controlling the nozzle throat to give a given mass percentage according to the actual fuel gas components in the engine, and controlling the incoming air flow to give a given air component percentage according to the actual incoming flow encountered; and controlling the incoming flow Mach number and the static pressure value of the nozzle throat used for the reverse jet of the carrier rocket to gradually change during the calculation process, so as to perform simulation calculations on the reverse jet of the carrier rocket.

Description

Technical Field

[0001] The present application relates to the field of launch vehicle aerodynamic calculations, and in particular to a numerical simulation method and system for reverse jet flow of a reusable launch vehicle. Background Art

[0002] In order to reduce the cost of space launches and increase the frequency of launches, the development of reusable launch vehicles is of great significance to the development of my country's space industry.

[0003] When a reusable launch vehicle's first stage returns, it relies on engine ignition for deceleration and braking. The high-temperature gas ejected from the nozzle is in the opposite direction of the incoming airflow, a typical example of a reverse jet flow. Due to the mutual interference between the free incoming airflow and the engine jet, especially during the initial re-entry phase, when the altitude is high and the atmospheric pressure is very low, the gas at the nozzle exit expands dramatically. At this time, the incoming airflow Mach number is high, and there is strong mutual interference between the incoming airflow and the jet flow. The flow field structure is extremely complex, making it very difficult to accurately simulate this flow field structure.

[0004] At present, domestic research on reverse jets mainly focuses on jet drag reduction and heat reduction. The research content is mostly on jets with small apertures at the head. The total temperature and total pressure of the jets are relatively low, and the aerodynamic interference with the incoming flow is not strong, so the simulation difficulty is relatively low. For the reverse jet simulation of large-scale rocket engines (such as launch vehicle engines), China is still in a technical blank stage. Summary of the Invention

[0005] In order to accurately simulate the reverse jet flow field of a reusable carrier rocket during its return, and thus obtain the aerodynamic performance of a sub-stage, the present invention provides a numerical simulation method and system for the reverse jet of a reusable carrier rocket. By referring to the actual operating environment of the nozzle and simulating the actual incoming flow, and controlling the incoming flow Mach number and the static pressure value at the nozzle throat used for the reverse jet of the carrier rocket to gradually change during the calculation process, accurate flow field structure simulation is provided for the calculation of the aerodynamic performance of subsequent sub-stages.

[0006] To solve the above technical problems, the first aspect of the present invention discloses a numerical simulation method for reverse jet flow of a reusable launch vehicle, the method comprising:

[0007] Establish nozzle related mesh model;

[0008] The nozzle throat is used as the inlet boundary of the jet calculation, and other relevant boundary conditions of the nozzle are determined;

[0009] Control the nozzle throat to give a given mass percentage according to the actual gas composition in the engine, and control the incoming air flow to give a given air composition percentage according to the actual incoming air flow encountered;

[0010] The incoming flow Mach number and the static pressure value at the nozzle throat used to control the reverse jet of the carrier rocket are gradually changed during the calculation process to simulate the reverse jet of the carrier rocket.

[0011] Optionally, after establishing the nozzle-related grid model, the method further includes:

[0012] The grid of the gas expansion area downstream of the jet is refined.

[0013] Optionally, the method further includes:

[0014] The pressure at the far-field inlet boundary is set to the local atmospheric pressure calculated according to the actual flight altitude.

[0015] Optionally, before controlling the nozzle throat to set a given mass percentage according to the actual fuel gas composition in the engine, the method further comprises:

[0016] Set the nozzle throat temperature to the true value.

[0017] Optionally, the incoming flow Mach number and the nozzle throat static pressure value used for controlling the reverse jet of the carrier rocket are gradually changed during the calculation process, specifically including:

[0018] Step A, setting an initial value of the incoming flow Mach number and an initial value of the nozzle throat static pressure to perform simulation calculation;

[0019] Step B: After the calculation converges, keep the initial value of the incoming flow Mach number unchanged and increase the static pressure value at the nozzle throat. Use the convergence result as the initial condition for this calculation.

[0020] Step C, repeating step B until the static pressure value at the nozzle throat reaches the corresponding actual value;

[0021] Step D, performing simulation calculation based on the initial value of the flow Mach number and the actual value corresponding to the static pressure value at the nozzle throat;

[0022] Step E: After the calculation converges, the static pressure value at the nozzle throat is maintained at the actual value, and the incoming flow Mach value is increased to perform simulation calculation again;

[0023] Step F, repeating step E until the incoming flow Mach number reaches the corresponding actual value.

[0024] Optionally, after step F, the method further includes:

[0025] The incoming flow Mach number reaches a corresponding actual value as the initial value of a higher Mach number state.

[0026] Optionally, the method further includes:

[0027] The next angle of attack state used to control the reverse jet of the carrier rocket is initialized using the convergence result of the previous angle of attack calculation.

[0028] A second aspect of the present invention discloses a numerical simulation system for reverse jet flow of a reusable launch vehicle, the system comprising:

[0029] A grid model building unit is used to build nozzle-related grid models;

[0030] The boundary condition setting unit is used to use the nozzle throat as the inlet boundary for jet calculation and determine the relevant boundary conditions of the nozzle; the far field inlet is the inlet boundary, the far field outlet is the opening boundary, and the nozzle-related solid wall is set as the no-slip wall boundary;

[0031] The first control unit is used to control the nozzle throat to set a given mass percentage according to the actual gas composition in the engine, and to control the incoming air flow to set a given air composition percentage according to the actual incoming air flow;

[0032] The second control unit is used to control the incoming flow Mach number and the static pressure value of the nozzle throat used for the reverse jet of the carrier rocket to gradually change during the calculation process, so as to simulate the reverse jet of the carrier rocket.

[0033] According to a third aspect of the present invention, a computer-readable storage medium is disclosed, on which a computer program is stored. When the program is executed by a processor, the steps of the above method are implemented.

[0034] A fourth aspect of the present invention discloses a computer device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor implements the steps of the above method when executing the program.

[0035] Through one or more technical solutions of the present invention, the present invention has the following beneficial effects or advantages:

[0036] The present invention discloses a numerical simulation method and system for the reverse jet of a reusable launch vehicle. By referring to the actual operating environment of the nozzle and simulating the actual incoming flow, the incoming flow Mach number and the static pressure value at the nozzle throat used for the reverse jet of the launch vehicle are controlled to gradually change during the calculation process, thereby providing accurate flow field structure simulation for the calculation of the aerodynamic performance of subsequent sub-stages.

[0037] The above description is only an overview of the technical solution of the present invention. In order to more clearly understand the technical means of the present invention, it can be implemented in accordance with the contents of the specification. In order to make the above and other purposes, features and advantages of the present invention more obvious and easy to understand, the specific implementation methods of the present invention are specifically listed below. BRIEF DESCRIPTION OF THE DRAWINGS

[0038] Various other advantages and benefits will become apparent to those skilled in the art upon reading the detailed description of the preferred embodiment below. The accompanying drawings are for illustration purposes only and are not to be construed as limiting the present invention. The same reference numerals are used throughout the accompanying drawings to denote the same components.

[0039] In the attached figure:

[0040] Figure 1 A flow chart of a numerical simulation method for reverse jet flow of a reusable launch vehicle according to an embodiment of the present invention is shown;

[0041] Figure 2 A diagram showing a specific implementation process of step-by-step calculation according to an embodiment of the present invention is shown;

[0042] Figure 3-Figure 4 Shows a verification schematic diagram according to the present invention;

[0043] Figure 5 A schematic diagram of a low-altitude, low-inflow Mach reverse jet flow field of a reusable carrier rocket calculated using the method of the present invention is shown;

[0044] Figure 6 A schematic diagram of a low-altitude, low-inflow Mach reverse jet flow field of a reusable carrier rocket calculated using the method of the present invention is shown;

[0045] Figure 7 A schematic diagram of a numerical simulation system for reverse jet flow of a reusable launch vehicle according to an embodiment of the present invention is shown. DETAILED DESCRIPTION

[0046] Exemplary embodiments of the present invention will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present invention are shown in the accompanying drawings, it should be understood that the present invention can be implemented in various forms and should not be limited by the embodiments set forth herein. Rather, these embodiments are provided to enable a more thorough understanding of the present invention and to fully convey the scope of the present invention to those skilled in the art.

[0047] The embodiment of the present invention discloses a numerical simulation method for the reverse jet of a reusable launch vehicle. The method of the present invention is applicable to the simulation of reverse jet states such as high-altitude high-inflow Mach reverse jet, low-altitude low Mach reverse jet, single-nozzle jet, and multi-nozzle jet in the return phase of a reusable launch vehicle. Figure 1 , the method comprises the following steps:

[0048] Step 101: Establish a nozzle-related grid model.

[0049] The nozzle is a crucial component of the engine. Proper nozzle structural design is crucial for various launch vehicle controls and is essential for accurately simulating the flow field structure. During the modeling process, actual nozzle structural designs can be referenced. For example, the nozzle-related mesh model is designed by establishing boundaries such as the external far field, the rocket body wall, the nozzle wall, and the nozzle throat. The mesh is then refined locally within the nozzle, including the downstream gas expansion region of the jet.

[0050] Step 102: Use the nozzle throat as the inlet boundary for jet flow calculation, and determine other relevant boundary conditions of the nozzle.

[0051] This example uses conventional fluid simulation software, such as CFX. Of course, after using the nozzle throat as the inlet boundary for the jet calculation, other boundary conditions are also set. For example, the far-field inlet is used as the inlet boundary, the far-field outlet is used as the opening boundary, and the nozzle-related solid walls are set as no-slip wall boundaries.

[0052] In order to accurately simulate the complex flow field structure, the parameters used in this embodiment are basically actual values. For example, the pressure at the far field inlet boundary is set to the local atmospheric pressure calculated based on the actual flight altitude.

[0053] Step 103, controlling the nozzle throat to set a given mass percentage according to the actual fuel gas components in the engine, and controlling the incoming air flow to set a given air component percentage according to the actual incoming air flow.

[0054] To accurately simulate complex flow field structures, the parameters used in this embodiment are essentially real-world values. Specifically, the nozzle throat temperature and gas composition parameters are set to realistic values ​​based on actual engine parameters. Taking gas composition as an example, Table 1 lists the real-world composition parameters used in calculations under certain typical conditions. Furthermore, the nozzle throat is controlled to have a given mass percentage based on the actual gas composition in the engine. Furthermore, the free stream is also assigned the corresponding air composition percentage based on the actual incoming flow.

[0055] Table 1

[0056] Components Gas Free flow CO 0.3 0 <![CDATA[CO2]]> 0.2 0 <![CDATA[H2O]]> 0.27 0 <![CDATA[H2]]> 0.23 0 <![CDATA[O2]]> 0 0.22 <![CDATA[N2]]> 0 0.78

[0057] Step 104 , controlling the incoming flow Mach number and the static pressure value at the nozzle throat used for the reverse jet of the carrier rocket to gradually change during the calculation process, so as to perform simulation calculation on the reverse jet of the carrier rocket.

[0058] In this embodiment, to avoid computational divergence, the incoming Mach number and nozzle throat static pressure are configured to change gradually during the calculation process. Specifically, the incoming Mach number and nozzle throat static pressure are initialized for calculation, and after the calculation converges, the previous convergence result is used as the initial condition for the subsequent calculation. The initial value of the incoming Mach number remains unchanged, and the nozzle throat static pressure is gradually changed until convergence occurs again. Once the nozzle throat static pressure meets the conditions, it remains unchanged, and the incoming Mach value is gradually changed to participate in the calculation. Each change must be calculated until convergence before the next calculation can be performed, until the incoming Mach value reaches the actual value.

[0059] For the convenience of describing and explaining this embodiment, see Figure 2 , is a diagram of the specific implementation process of step-by-step calculation, which specifically includes the following steps:

[0060] Step A: Set an initial value of the incoming flow Mach number and an initial value of the nozzle throat static pressure to perform simulation calculations.

[0061] Step B: After the calculation converges, keep the initial value of the incoming flow Mach number unchanged and increase the static pressure value at the nozzle throat. Use the convergence result as the initial condition for this calculation.

[0062] Step C: Repeat step B until the static pressure value at the nozzle throat reaches the corresponding actual value.

[0063] Step D: performing simulation calculation based on the initial value of the flow Mach number and the actual value corresponding to the static pressure value at the nozzle throat.

[0064] Step E: After the calculation converges, the static pressure value at the nozzle throat is maintained at the actual value, and the incoming flow Mach value is increased to perform simulation calculation again.

[0065] Step F, repeating step E until the incoming flow Mach number reaches the corresponding actual value.

[0066] For example, if you need to calculate a high inflow Mach reverse jet with an inflow of 5.0 Mach, first set the inflow to 1.2 Mach and the throat static pressure value to 1MPa as the initial values ​​to start the calculation (the actual value corresponding to the throat static pressure value of a certain engine is 5MPa). After the calculation is stable, increase the throat pressure to 2MPa, use the result of the previous convergence as the initial condition, calculate this time until convergence, and then use the same method to increase the throat pressure to the actual value of 5MPa (the final throat pressure), thereby gradually increasing the throat pressure to the actual value. After the above calculation converges, gradually increase the inflow Mach number. For example, increase the inflow from 1.2 Mach to 2.0 Mach, and it will basically converge after 1000 iterative calculation steps. Then increase the inflow parameters to 5.0 Mach, and it will converge after about 1000 more calculation steps. It should be noted that during this calculation process, the Timescale Control setting cannot use the default option. You need to select the Local Timescale Factor local time scale coefficient, and the TimescaleFactor time step coefficient must remain at 1.

[0067] As an optional embodiment, for a high Mach number state (e.g., Mach 6, Mach 7), the incoming Mach value can be used as the initial value of the higher Mach number state. For example, the result of Mach 5 is used for initialization.

[0068] As an optional embodiment, for other numerical values ​​used in the rocket's reverse jet, such as angle of attack calculation, the next angle of attack state used in the rocket's reverse jet is initialized using the convergence result of the previous angle of attack calculation. For example, a 3° angle of attack can be initialized using the result of a 0° angle of attack, a 6° angle of attack can be initialized using the result of a 3° angle of attack, and so on.

[0069] In order to verify the method of the present invention, the method established by the present invention was used to calibrate a head reverse jet test. Figure 3 Schematic diagram of the calculated reverse jet flow field (CFD) and the experimental schlieren comparison (EXP). Figure 3 It can be seen that the calculated positions of the bow shock wave at the front end of the blunt body, the barrel shock wave at the edge of the jet, and the Mach disk at the end of the jet are basically consistent with the shock wave positions in the experimental schlieren diagram.

[0070] Figure 4 The figure shows the comparison between the calculated and experimental results of the static pressure curve of the reverse jet head. Figure 4It can be seen that the head surface pressure changes with the central angle in both the calculation and the test are basically consistent. In the non-jet state, the head surface pressure decreases with the increase of the central angle. The deviation between the calculation (NoJet-CFD) and the test (NoJet-Exp) in the non-jet state is basically within 3%. In the case of reverse jet, the surface pressure decreases as a whole. According to PN=P / P of The ratio of the static pressure on the head surface to the total pressure after the shock wave is calculated using the formula, where P represents the static pressure on the head surface, and P of It represents the total pressure after the shock wave, PN is the ratio of the two, and the maximum PN does not exceed 0.3. The deviations between the calculation with reverse jet (Jet-CFD) and the experiment with reverse jet (Jet-Exp) are basically within 6%, and the pressure peak occurs between the central angle of 50° and 60°.

[0071] Figure 5 The figure is a schematic diagram of the low-altitude, low-inflow Mach reverse jet flow field of a reusable carrier rocket calculated using the method of the present invention. Here, mach number represents the Mach number.

[0072] Figure 6 The figure shows the Mach reverse jet flow field of a reusable carrier rocket at high altitude and high incoming flow calculated by the method of the present invention.

[0073] Based on the same inventive concept as in the above embodiment, the embodiment of the present invention also discloses a numerical simulation system for reverse jet flow of a reusable launch vehicle, see Figure 7 , the system comprising:

[0074] The grid model building unit 701 is used to build a nozzle-related grid model.

[0075] The boundary condition setting unit 702 is used to use the nozzle throat as the inlet boundary for the jet calculation and determine the relevant boundary conditions of the nozzle. The far-field inlet is the inlet boundary, the far-field outlet is the opening boundary, and the relevant solid wall of the nozzle is set as the no-slip wall boundary.

[0076] The first control unit 703 is used to control the nozzle throat to give a given mass percentage according to the actual fuel gas composition in the engine, and to control the incoming air flow to give a given air composition percentage according to the actual incoming air flow.

[0077] The second control unit 704 is used to control the incoming flow Mach number and the static pressure value at the nozzle throat used for the reverse jet of the carrier rocket to gradually change during the calculation process, so as to simulate the reverse jet of the carrier rocket.

[0078] Based on the same inventive concept as in the aforementioned embodiments, an embodiment of the present invention further discloses a computer-readable storage medium having a computer program stored thereon, which implements the steps of any of the aforementioned methods when executed by a processor.

[0079] Based on the same inventive concept as in the aforementioned embodiments, an embodiment of the present invention further discloses a computer device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein when the processor executes the program, the steps of any of the aforementioned methods are implemented.

[0080] Through one or more embodiments of the present invention, the present invention has the following beneficial effects or advantages:

[0081] The present invention discloses a numerical simulation method and system for the reverse jet of a reusable launch vehicle. By referring to the actual operating environment of the nozzle and simulating the actual incoming flow, the incoming flow Mach number and the static pressure value at the nozzle throat used for the reverse jet of the launch vehicle are controlled to gradually change during the calculation process, thereby providing accurate flow field structure simulation for the calculation of the aerodynamic performance of subsequent sub-stages.

[0082] Although the preferred embodiments of the present application have been described, those skilled in the art may make additional changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present application.

[0083] Obviously, those skilled in the art may make various changes and modifications to this application without departing from the spirit and scope of this application. Thus, if these modifications and variations of this application fall within the scope of the claims of this application and their equivalents, this application is intended to include these modifications and variations.

Claims

1. A numerical simulation method for reverse jet flow of a reusable launch vehicle, characterized in that: The method comprises: Establish nozzle related mesh model; The nozzle throat is used as the inlet boundary of the jet calculation, and other relevant boundary conditions of the nozzle are determined; Control the nozzle throat to give a given mass percentage according to the actual gas composition in the engine, and control the incoming air flow to give a given air composition percentage according to the actual incoming air flow encountered; The inflow Mach number and the nozzle throat static pressure value used to control the reverse jet flow of the carrier rocket are gradually changed during the calculation process to simulate the reverse jet flow of the carrier rocket, specifically including: Step A, setting an initial value of the incoming flow Mach number and an initial value of the nozzle throat static pressure to perform simulation calculation; Step B: After the calculation converges, keep the initial value of the incoming flow Mach number unchanged and increase the static pressure value at the nozzle throat. Use the convergence result as the initial condition for this calculation. Step C, repeating step B until the static pressure value at the nozzle throat reaches the corresponding actual value; Step D, performing simulation calculation based on the initial value of the incoming flow Mach number and the actual value corresponding to the nozzle throat static pressure value; Step E: After the calculation converges, the static pressure value at the nozzle throat is maintained at the actual value, and the incoming flow Mach value is increased to perform simulation calculation again; Step F, repeating step E until the incoming flow Mach value reaches the corresponding actual value.

2. The method according to claim 1, wherein After establishing the nozzle-related grid model, the method further includes: The grid of the gas expansion area downstream of the jet is refined.

3. The method according to claim 1, wherein The method further comprises: The pressure at the far-field inlet boundary is set to the local atmospheric pressure calculated from the actual flight altitude.

4. The method according to claim 1, wherein Before controlling the nozzle throat to set a mass percentage according to the actual fuel gas composition in the engine, the method further includes: Set the nozzle throat temperature to the true value.

5. The method according to claim 1, wherein After step F, the method further comprises: The incoming flow Mach number reaches a corresponding actual value as the initial value of a higher Mach number state.

6. The method according to claim 1, wherein The method further comprises: The next angle of attack state used to control the reverse jet of the carrier rocket is initialized using the convergence result of the previous angle of attack calculation.

7. A numerical simulation system for reverse jet flow of a reusable launch vehicle, characterized in that: The system comprises: A grid model building unit is used to build nozzle-related grid models; The boundary condition setting unit is used to use the nozzle throat as the inlet boundary for jet calculation and determine the relevant boundary conditions of the nozzle; the far field inlet is the inlet boundary, the far field outlet is the opening boundary, and the nozzle-related solid wall is set as the no-slip wall boundary; The first control unit is used to control the nozzle throat to set a given mass percentage according to the actual gas composition in the engine, and to control the incoming air flow to set a given air composition percentage according to the actual incoming air flow; The second control unit is configured to control the incoming flow Mach number and the nozzle throat static pressure value used for the reverse jet of the carrier rocket to gradually change during the calculation process, so as to simulate the reverse jet of the carrier rocket, specifically comprising: Step A, setting an initial value of the incoming flow Mach number and an initial value of the nozzle throat static pressure to perform simulation calculation; Step B: After the calculation converges, keep the initial value of the incoming flow Mach number unchanged and increase the static pressure value at the nozzle throat. Use the convergence result as the initial condition for this calculation. Step C, repeating step B until the static pressure value at the nozzle throat reaches the corresponding actual value; Step D, performing simulation calculation based on the initial value of the incoming flow Mach number and the actual value corresponding to the nozzle throat static pressure value; Step E: After the calculation converges, the static pressure value at the nozzle throat is maintained at the actual value, and the incoming flow Mach value is increased to perform simulation calculation again; Step F, repeating step E until the incoming flow Mach value reaches the corresponding actual value.

8. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the program is executed by a processor, the steps of the method according to any one of claims 1 to 6 are implemented.

9. A computer device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein: When the processor executes the program, the steps of the method according to any one of claims 1 to 6 are implemented.

Images