An aero-engine whole machine three-dimensional simulation debugging method and device and a storage medium

Abstract

This invention discloses a method, equipment, and storage medium for three-dimensional simulation and debugging of aero-engines, belonging to the field of aero-engine aerodynamic-thermal simulation. The core of this method lies in systematically solving the simulation convergence problem through a two-level strategy: gradient pre-convergence for individual components and dynamic adaptive coupling for the entire engine. At the component level, robust pre-convergence of the flow field is achieved through progressive physical loading and stability control for the compressor, combustion chamber, and turbine. At the engine level, dynamic calculation parameter adjustment and flow field stability control based on real-time flow field stability ensure a fast and stable coupling process. This invention also expands the functions of geometric mesh standardization, post-processing verification, variable operating condition debugging, and real-time monitoring and anomaly handling. This invention significantly improves the convergence efficiency, success rate, and result reliability of simulations, reducing the computational interruption rate by more than 75%, and is suitable for the engineering design and optimization of aero-engines.

Description

Technical Field

[0001] This invention relates to the field of aerodynamic and thermodynamic simulation technology for aero-engines, specifically to a three-dimensional simulation and debugging method, equipment, and storage medium for aero-engines. Background Technology

[0002] Three-dimensional simulation calculation of the entire aero-engine is a core technical means for the research and development of modern aero-engine power plants. It can predict the complex flow field characteristics inside the engine during the design stage through numerical simulation, thereby reducing the cost of physical testing and shortening the research and development cycle. However, the existing simulation debugging methods have the following technical bottlenecks: (1) Component debugging stage: unreasonable initial boundary conditions of the compressor can easily lead to flow field oscillation under the reverse pressure gradient; direct coupling of combustion reaction in the combustion chamber can cause violent fluctuations in the flow field; instantaneous activation of the turbine cooling system can cause flow field convergence difficulties. (2) Integrated stage: due to the difference in flow field parameters of each component, direct coupling can easily cause sudden changes in interface parameters; insufficient mesh resolution in the region of violent physical field gradient can lead to amplification of calculation errors; fixed settings of relaxation factor and Coulomb number cannot adapt to the dynamic change process of flow field. (3) Extended scenario stage: flow field instability is easy to occur during the simulation of the transition process of variable operating conditions; there is a lack of effective iteration strategies when multiple physical fields are coupled (such as thermo-solid coupling); the lack of abnormal situation handling mechanism during the simulation process leads to a high calculation interruption rate.

[0003] The aforementioned problems result in low convergence efficiency and insufficient reliability of simulation results, making it difficult to meet the needs of engineering applications. Summary of the Invention

[0004] The purpose of this invention is to overcome the shortcomings of the prior art and provide a method, equipment and storage medium for debugging three-dimensional simulation of aero-engines, which starts from the robustness of the internal flow field of components and ensures that the three-dimensional simulation of the whole aero-engine can be completed efficiently and reliably through hierarchical convergence and active control.

[0005] To solve the above-mentioned technical problems, the present invention adopts the following technical solution:

[0006] Firstly, this invention provides a three-dimensional simulation and debugging method for an aero-engine, the core of which lies in systematically solving the simulation convergence problem through a two-level strategy of component gradient pre-convergence and whole-engine dynamic adaptive coupling. It mainly includes the following steps:

[0007] Component gradient pre-convergence steps: Before coupling the entire machine, the compressor, combustion chamber, and turbine are individually debugged. For the compressor, an initial flow field with a pressure gradient is established, and a progressive synchronous loading mechanism of back pressure and speed is adopted, while stall warning and adjustment are performed based on flow field monitoring; for the combustion chamber, a staged activation strategy is adopted, sequentially solving the cold flow field, calculating the fuel mixing process, and activating the combustion reaction; for the turbine, a step-by-step activation mechanism for the cold air system and a progressive loading mechanism for operating parameters are adopted.

[0008] The whole machine dynamic adaptive coupling steps are as follows: the whole machine coupling is carried out on the basis of component flow field pre-convergence, and a strategy of dynamically adjusting the relaxation factor and Coulomb number based on the stability of inlet flow is adopted. A velocity monitoring and relaxation factor adjustment mechanism for local backflow is also established.

[0009] In a second aspect, the present invention provides an electronic device, including a memory and a processor, wherein the memory stores a computer program, and the processor executes the computer program to implement the steps of the above-described method.

[0010] Thirdly, the present invention provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, performs the steps of the above-described method.

[0011] The beneficial effects of this invention include:

[0012] This invention employs a component-based convergence strategy, using single-component flow field pre-convergence to avoid global failures caused by local divergence in whole-machine simulation. It improves computational stability through progressive loading, and significantly enhances simulation robustness under complex operating conditions through methods such as simultaneous loading of compressor back pressure and speed, and step-by-step activation of turbine cooling. The whole-machine co-simulation utilizes a dynamic relaxation factor strategy to reduce iterations and improve computational efficiency. This invention features modular scalability, supports rapid replacement of component models, and can adapt to simulation requirements of different engine configurations. Attached Figure Description

[0013] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0014] Figure 1 This is a flowchart illustrating the overall process of a three-dimensional simulation and debugging method for an aero-engine according to an embodiment of the present invention. Detailed Implementation

[0015] The embodiments of this application will now be described in detail with reference to the accompanying drawings.

[0016] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0017] It should be noted that various aspects of embodiments within the scope of the appended claims are described below. It will be apparent that the aspects described herein can be embodied in a wide variety of forms, and any particular structure and / or function described herein is merely illustrative. Based on this application, those skilled in the art will understand that one aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number of aspects set forth herein can be used to implement the device and / or practice the method. Additionally, this device and / or method can be implemented using structures and / or functionalities other than one or more of the aspects set forth herein.

[0018] It should also be noted that the illustrations provided in the following embodiments are only schematic representations of the basic concept of this application. The illustrations only show the components related to this application and are not drawn according to the number, shape and size of the components in actual implementation. In actual implementation, the form, quantity and proportion of each component can be arbitrarily changed, and the layout of the components may also be more complex.

[0019] Firstly, this invention provides a three-dimensional simulation and debugging method for an aero-engine. This embodiment uses the simulation of a certain type of aero-engine (E3) as an example, but the scope of protection of this invention is not limited to this specific model. Figure 1 As shown, the method mainly includes the following steps:

[0020] S101: Geometric Modeling and Mesh Preprocessing

[0021] First, establish geometric simplification standards: non-aerodynamic feature dimensions <5% of the main flow diameter can be ignored, and sharp corner transitions are adopted when the fillet radius is <1mm. For the E3 engine's overall geometric model, perform a cleanup operation: remove non-aerodynamic features (such as process holes and lifting structures), for example, lifting lugs with a diameter of 8mm (less than 5% of the main flow diameter φ300mm, i.e., 15mm) and process holes with a radius of 0.5mm (using sharp corner transitions). Retain core aerodynamic areas such as the compressor blade passage (including stationary and moving blades), combustion chamber flame tube (including fuel nozzle mounting base), and turbine blade passage (high-pressure turbine, low-pressure turbine). Control the component docking error to ensure that the flange misalignment is ≤0.1mm and the coaxiality of the axis is ≤0.05mm / m; for example, the measured misalignment of the flange face between the compressor outlet and the combustion chamber inlet of the E3 engine is 0.08mm (≤0.1mm), and the coaxiality of the turbine and the combustion chamber axis is 0.04mm / m (≤0.05mm / m), which meets the design requirements.

[0022] Subsequently, a computational mesh for the entire machine was generated: a structured mesh was used for the compressor, with a mesh growth rate of 1.1–1.2 on the blade surface and the tip clearance mesh was refined to the first layer height. ≈1. For the combustion chamber's flame tube region, an unstructured tetrahedral mesh is used, with 5-7 layers of prismatic meshes in the wall boundary layer, resulting in a total mesh count ≥5 million. For the turbine's moving and stationary blade passages, a hybrid mesh (structured near the blades and unstructured within the blade cooling passages) is used. The cooling hole mesh uses a Cartesian-cut mesh, ensuring the mesh orthogonality within the holes is >0.8. Mesh quality acceptance criteria are established, requiring a minimum volumetric distortion rate >0.3, a maximum aspect ratio <20, and orthogonality >0.6. For example, the compressor uses a structured hexahedral mesh, with a blade surface mesh growth rate of 1.15. The first layer mesh height at the tip clearance (0.5mm) is 0.02mm, ensuring the dimensionless value of the first mesh height. The total mesh size is approximately 1, with a total mesh count of 8 million. The combustion chamber flame tube region uses an unstructured tetrahedral mesh, with 6 layers of prism meshes on the walls (layer height gradually changing from 0.05 to 0.3 mm), resulting in a total mesh count of 6 million (≥5 million). The turbine's moving and stationary blade passages use a hybrid mesh (structured mesh on the blade surface, unstructured mesh inside the blade cooling passages). The cooling holes (1.2 mm in diameter) use Cartesian-cut meshes with an orthogonality of 0.85 (>0.8). Mesh quality was checked, and the minimum volumetric distortion rate across the entire flow domain was found to be 0.35 (>0.3), the maximum aspect ratio 18 (<20), and the orthogonality 0.65 (>0.6), meeting the acceptance criteria. This step lays a solid foundation for subsequent high-precision simulations.

[0023] S102: Systematically acquire key section parameters

[0024] Based on the overall performance parameters of the engine under design conditions, the core section parameters are determined according to the following formula using aerodynamic and thermodynamic relationships.

[0025] Compressor outlet pressure:

[0026] in, This refers to the static pressure at the compressor inlet. This refers to the static pressure at the compressor outlet. Design the pressure ratio for the compressor.

[0027] Compressor outlet temperature:

[0028] in, This refers to the total temperature at the compressor inlet. This refers to the total temperature at the compressor outlet. (Specific heat ratio of air) This refers to the compressor's adiabatic efficiency.

[0029] Compressor inlet airflow angle:

[0030] in The compressor inlet airflow angle. , These are the axial and circumferential velocity components at the compressor inlet, respectively.

[0031] Combustion chamber outlet pressure:

[0032] in This is the combustion chamber pressure loss rate (usually taken as 3% to 5%). The combustion chamber outlet pressure, This is the static pressure at the compressor outlet.

[0033] Combustion chamber outlet temperature:

[0034] in, The combustion chamber outlet temperature, This refers to the total temperature at the compressor outlet. The air-fuel ratio is LHV, and the lower heating value of the fuel is LHV. For combustion efficiency, .

[0035] Turbine outlet pressure:

[0036] in, For turbine outlet pressure, The combustion chamber outlet pressure, This is the turbine expansion ratio.

[0037] Turbine outlet temperature:

[0038] in, Turbine outlet temperature, This refers to the turbine adiabatic efficiency.

[0039] Turbine inlet airflow angle: ,in The turbine inlet airflow angle, The radial velocity at the turbine inlet. The turbine inlet velocity is the combined velocity.

[0040] For example, based on the design operating parameters of the E3 engine, the core section parameters are calculated, and the operating parameters are: compressor inlet static pressure. =101325Pa, compressor inlet total temperature =288K, compressor design pressure ratio =25, compressor adiabatic efficiency =0.88, compressor inlet axial component =150m / s, compressor circumferential velocity component =30m / s; Combustion chamber pressure loss rate =4%, fuel-air ratio f a =0.015, fuel lower calorific value LHV=43MJ / kg, combustion efficiency =0.95; Turbine expansion ratio =5, turbine adiabatic efficiency =0.9, turbine inlet radial velocity =50m / s, turbine inlet velocity =300m / s.

[0041] Therefore, the compressor outlet of the E3 engine:

[0042] static pressure =101325Pa×25=2533125Pa;

[0043] Total temperature =288K×[1+(250.286-1) / 0.88]=825K;

[0044] Inlet airflow angle =arctan (30 / 150)=11.3°.

[0045] Combustion chamber outlet:

[0046] pressure =2533125Pa × 0.96 = 2431800Pa;

[0047] temperature =825K+(0.015×43× ×0.95) / (1005×1.015)=1450K.

[0048] Turbine outlet:

[0049] pressure =2431800Pa / 5=486360Pa;

[0050] temperature ;

[0051] Inlet airflow angle = =9.6°.

[0052] S103: Gradient Debugging of Sub-components

[0053] Next, the component gradient pre-convergence step is initiated: before overall engine coupling, independent flow field adjustments are performed on the compressor, combustion chamber, and turbine. A progressive physical loading strategy is used to smoothly transition the flow field of each component from a stable initial state to the design conditions, and stability control is implemented based on real-time flow field monitoring results during the loading process. Specifically,

[0054] Compressor commissioning:

[0055] Initial boundary settings: Inlet total pressure, total temperature, and airflow angle are given; outlet back pressure is set to 80%–90% of the inlet total pressure to establish a pressure gradient; hub and casing wall surfaces are set as adiabatic, non-slip boundaries, with roughness... For example, the initial boundaries of the E3 engine compressor are set as follows: inlet given total pressure 101325 Pa, total temperature 288 K, airflow angle 11.3°, and outlet back pressure set to 85% of the inlet total pressure (86126 Pa) to establish a pressure gradient initial flow field.

[0056] Implement a progressive synchronous loading mechanism: back pressure loading is performed according to the following formula:

[0057]

[0058] Among them, the back pressure increment at each step , Let n be the initial back pressure for iteration step n. The initial back pressure for iteration step n-1, For target back pressure, This represents the number of iteration steps. For example, the target back pressure of the E3 compressor is... =2533125Pa, loaded according to the aforementioned formula, for example, in step 1 ΔP=0.03×(2533125-86126) / 10≈7341Pa, (1) = 86126 + 7341 = 93467 Pa; Step 2: ΔP = 0.03 × (2533125 - 93467) / 10 ≈ 7319 Pa, (2) = 93467 + 7319 = 100786 Pa; The loading is done in 10 steps to reach the target value.

[0059] The rotational speed is applied synchronously, and the rotational speed application is based on the following formula:

[0060]

[0061] Among them, the speed increment at each step , The number of iterations. Let n be the rotational speed at iteration step n. For the rotational speed of iteration step n-1, The target rotational speed. For example, the target rotational speed is... =15000r / min, ΔN ≤ 0.03×( -N (n-1)) / 10, a total of 10 steps to complete the loading.

[0062] Establish a stall warning and adjustment mechanism: During loading, monitor in real time. When the area of ​​the backflow region in a certain blade channel exceeds a set threshold (e.g., 5%), load loading is paused and the back pressure is automatically adjusted, such as reducing the current back pressure by 5%. Loading is then resumed after stabilization. For example, if the area of ​​the backflow region in the third-stage blade channel reaches 6.2% (>5%) in step 6, load loading is immediately paused and the current back pressure is reduced by 5% (from 1,520,000 Pa to 1,444,000 Pa). Loading is resumed after stabilization for 30 steps.

[0063] Convergence Criteria: The following convergence criteria are established: In 50 consecutive iterations, the fluctuation of the total pressure recovery coefficient is <0.5%, and the efficiency fluctuation is <0.3%. For example, if the total pressure recovery coefficient fluctuates by 0.3% (<0.5%) and the efficiency fluctuates by 0.2% (<0.3%) in 50 consecutive iterations, the compressor flow field is considered converged. This process effectively avoids rotational stall and ensures a smooth transition of the compressor flow field.

[0064] Combustion chamber adjustment:

[0065] In this embodiment, the combustion chamber commissioning adopts a three-stage step-by-step activation strategy, which executes the following stages sequentially:

[0066] Phase 1 (Cold Flow Field Solution): The combustion model is disabled. A standard k-ε turbulence model and enhanced wall treatment are used to solve the pure air flow field until the residuals are below [value missing]. For example, after 800 iterations, the residual decreases to (< In this stage, without interference from chemical reactions, the flow field is first stabilized in terms of physical structure, establishing a reliable flow basis for subsequent fuel introduction and combustion, thus avoiding the initial instability caused by the coupling of flow and chemical reactions.

[0067] The second stage (fuel mixing process calculation): Based on the obtained stable flow, this stage introduces a fuel injection model to calculate fuel atomization and the fuel-air mixing process. A pressure swirl nozzle model is used for fuel injection, and the atomization angle and injection pressure are set; a fuel concentration gradient threshold is then established. To achieve the preset mixing uniformity standard and meet the requirements. ( (This is the fuel concentration gradient threshold). For example, the fuel injection model is a pressure swirling nozzle model; after 500 iterations, the fuel concentration gradient... =0.04 ( This stage of commissioning ensures that the fuel and oxidant are fully mixed in space before the chemical reaction is activated, completing pre-combustion fuel preparation without severe exothermic interference, and eliminating the risk of localized violent reactions or flameout caused by uneven mixing.

[0068] The third stage (combustion reaction activation): After confirming that the mixing homogeneity standard is met, the combustion reaction model is activated by gradually increasing the reaction rate coefficient. Where k is the reaction rate coefficient, n is the number of iterations, and I is the total number of transition steps. This is the initial reaction rate coefficient. The initial reaction rate coefficient is set at 10% of the target reaction rate coefficient, and the coefficient is gradually increased to reach the target coefficient. For example, the combustion model used is the EDC (eddy dissipation concept) combustion model, and the chemical reaction mechanism adopts the two-step reaction mechanism of aviation kerosene. The initial reaction rate coefficient is then gradually increased according to the above formula. =0.1× ( (Target reaction rate), total transition steps I=20, according to Gradually increase (n=1~20). This gentle activation method effectively avoids drastic fluctuations in the flow field. This stage does not apply the full chemical reaction intensity all at once, but rather introduces the chemical reaction gently, allowing the flow field time to gradually adapt to the exothermic process, thereby ultimately achieving a smooth and stable establishment of the entire combustion field and effectively avoiding flow field collapse caused by sudden changes in reaction rate.

[0069] Flame stability control: When the flame front fluctuation frequency is >500Hz, the combustion model time constant is increased to 1.5 times the original value. For example, when the flame front fluctuation frequency is detected to be 550Hz (>500Hz), the combustion model time constant is increased from 0.01s to 0.015s (1.5 times), the fluctuation frequency drops to 420Hz, and stability is restored.

[0070] Turbo tuning:

[0071] A phased activation mechanism for the cooling system is implemented: the cooling air volume is gradually increased to the turbine blades, with the cooling air inlet set as the mass flow boundary. The inlet cooling air volume is calculated as follows:

[0072]

[0073] in, , The number of iterations. For the amount of cooling air in step n, For the (n-1)th step of cooling air volume, For the target airflow, To adjust the number of times For example, the target airflow rate of the E3 turbocharger. =0.5kg / s, number of adjustments =10, then =0.05kg / s. According to the formula for imported air volume, starting from 0, increase by 0.05kg / s each step until the target value is reached.

[0074] Progressive loading of operating parameters: The turbine is gradually loaded from a low state to a high state using differential pressure and rotational speed. The formula for loading the turbine inlet and outlet differential pressure is:

[0075]

[0076] in, For the pressure difference at step n, For the pressure difference at step n-1, For the target pressure difference, To adjust the number of times.

[0077] The formula for the turbine rotor speed loading is:

[0078]

[0079] in, For the nth step rotation speed, For the (n-1)th step rotational speed, For the target speed, To adjust the number of times.

[0080] For example, target pressure difference =1945440Pa ( - ), target speed =12000 r / min, =10, and loading is performed according to the turbine inlet and outlet pressure difference formula and turbine rotor speed loading formula mentioned above, in 10 steps from low state to high state. This step smoothly introduces the cooling effect and aerodynamic load.

[0081] S104: Whole-system co-simulation optimization

[0082] After debugging the compressor, combustion chamber, and turbine of the aforementioned components, the next step is to perform a dynamic adaptive coupling step for the entire machine. This involves optimizing the joint simulation of the entire machine, performing whole-machine coupling based on the pre-convergence of the component flow fields, employing a dynamically adjusted calculation parameter strategy, and establishing an active intervention mechanism for local flow field instability. This includes the following steps:

[0083] Interface parameter coupling:

[0084] First, model connection is performed to establish interfaces between components such as compressor, combustion chamber, and turbine, connecting the aero-engine into a whole. Through model connection, coupling interfaces are established between compressor outlet, combustion chamber inlet (annular interface, area 0.2m²), combustion chamber outlet, and turbine inlet (annular interface, area 0.18m²), and the "frozen rotor" method is used to handle dynamic and static interference.

[0085] Then comes the initial coupling phase: To maintain coupling stability, very conservative (i.e., small) relaxation factors and Courant numbers must be used to suppress violent oscillations in the flow field parameters, ensuring that the solver can stably establish the basic flow field structure in the initial coupling stage and avoiding computational divergence caused by excessively large initial parameters. Specifically, the solver will use relaxation factors for pressure, velocity, and energy. Set the quorum to 0.01±50% (conservative value), the Courrard number (CFL) to 1±50% (conservative value), and the iteration step to 1000 steps for initial flow field coupling simulation. For example, set the relaxation factor. =0.01, Courant number CFL=1, the flow field is initially stable (residuals remain at...). (On the order of magnitude). Those skilled in the art will understand that these values ​​are far lower than typical values ​​in conventional single-component simulations, representing typical low-value settings used to maintain coupling stability. Under these conservative parameters, with a fixed number of initial iterations, the flow field initially stabilizes, and the residuals are maintained at... The magnitude of the data verified that the initial parameter settings effectively maintained the coupling stability.

[0086] Finally, dynamic parameter adjustments are performed: when the engine inlet flow fluctuation is monitored to be stable within the preset range, the relaxation factor and Coulomb number are activated and linearly increased. Specifically, when the engine inlet (compressor inlet) flow curve shows a fluctuation amplitude of <5% for 50 consecutive steps, the following steps are performed: The relaxation factor is linearly increased, where n is the number of iterations. Let n be the relaxation factor in step n. The initial relaxation factor, The maximum relaxation factor is I, and the total number of transition steps is denoted by Courant number. Linear improvement, where Let be the Courland number at step n. For example, if the engine inlet flow rate curve shows a fluctuation of 3% (<5%) for 50 consecutive steps, the starting parameters are increased, and the relaxation factor... =0.01, =0.7, adjust the total number of steps I=500, and increase the relaxation factor according to the relaxation factor increase formula mentioned above. The Courant number (CFL) is gradually increased from 0.01 to 0.7; the Courant number (CFL) is gradually increased from 1 to 5 according to the Courant number enhancement formula described above. This strategy significantly accelerates convergence while ensuring stability.

[0087] Flow field stability control:

[0088] Backflow monitoring: Establish a backflow monitoring mechanism. When a backflow velocity exceeds a threshold in a specific local area (typically at the turbine outlet), for example, the threshold is set to... , i.e., reflux rate Furthermore, if the relaxation factor in that local area continues for more than 100 steps, it will automatically decrease to 50% of the current value. The mainstream speed is the average speed at the engine's current position. For example, =30m / s, a backflow velocity was detected in the turbine outlet region. =35m / s (>0.1) =30m / s) and continue for 120 steps, the system automatically reduces the relaxation factor of the local area (turbine outlet area) from 0.6 to 0.3, and the backflow disappears after 50 steps. This mechanism effectively prevents the spread of local instability.

[0089] Pressure fluctuation suppression: Establish pressure fluctuation suppression. When the total outlet pressure fluctuation exceeds 2% of the design value, activate the global damping coefficient adjustment. Increase damping. Among these, The damping coefficient is... The initial damping coefficient, This represents a percentage fluctuation. For example, the total pressure fluctuation at the nozzle exit of the E3 engine reaches 2.5% (>2% × design value), calculated as... ( =0.1, By adjusting the damping coefficient from 0.1 to 0.1025 (=2.5%), the fluctuation amount was reduced to 1.8%.

[0090] S105: Post-processing and Verification of Simulation Results

[0091] This step extracts key performance parameters from the simulation results and calculates each parameter index using the following formula.

[0092] Overall thrust calculation: Where F is the thrust, The mass flow rate of air passing through the engine. The average airflow velocity at the engine nozzle exit section is given. The average airflow velocity at the inlet cross-section of the engine intake duct. Due to environmental pressures, This refers to the static pressure at the engine nozzle exit cross-section. This represents the area of ​​the engine nozzle exit cross-section.

[0093] Fuel consumption rate calculation: SFC represents fuel consumption rate. This refers to the fuel flow rate.

[0094] For example, thrust calculation, substituting parameters from a specific embodiment: =50kg / s, =500m / s, =150m / s, =486360Pa, =101325Pa, =0.2m², =0.75kg / s. The calculated thrust F=94507N and fuel consumption rate SFC=0.008 kg / (N·h) were compared with the design value of E3 engine (0.0082kg / (N·h)) and the deviation was 2.4%, which verified the accuracy of the simulation results.

[0095] S106: Variable Operating Condition Simulation Debugging

[0096] Variable operating condition commissioning: Interpolate the operating condition parameters, select the maximum thrust condition and the idle condition as feature points, and establish a parameter interpolation model: ,in For the interpolation target operating condition, For parameters of adjacent feature points, These are characteristic parameters for slow-speed operation. Here, k represents the characteristic parameter for maximum thrust operation, and k is the current rotational speed. For the characteristic speed of the idle operation, The characteristic speed under maximum thrust conditions. Variable operating condition transient rate control: speed change rate ≤ 500 r / min / s, fuel flow change rate ≤ 0.05 kg / s².

[0097] For dynamic response characteristic tuning, a segmented loading strategy is adopted for acceleration: in the low-speed range (<50% of rated speed), the speed increase per step is ≤3%, and in the high-speed range (≥50% of rated speed), the increase per step is ≤1.5%. For deceleration, back pressure and speed are reduced synchronously, with the back pressure reduction rate lagging behind the speed reduction rate by 5-10 steps to avoid excessive expansion. For example, interpolation is performed based on the design point and maximum thrust point to control the speed change rate at 500 r / min / s and the fuel change rate at 0.005 kg / s². During acceleration, the speed increase per step is 2% in the low-speed range (<7500 r / min) and 1% in the high-speed range (≥7500 r / min), ensuring dynamic stability during the transition process.

[0098] S107: Simulation Process Monitoring and Anomaly Handling

[0099] Establish real-time monitoring metrics, including the following:

[0100] Monitor the slope of the flow and pressure curves. When the absolute value of the slope is >0.01 for 20 consecutive steps, issue a calculation divergence warning and reduce the relaxation factor by half.

[0101] The extreme values ​​of the physical field are detected, and upper limits for pressure (1.5 times the design value) and temperature (1.2 times the design value) are set. When the upper limit for pressure or temperature is exceeded, the relaxation factor is reduced by half.

[0102] For example, during real-time monitoring, when the iteration reaches 1200 steps, if the slope of the pressure curve is >0.01 and persists for 20 steps, the system issues a divergence warning and reduces the relaxation factor from 0.5 to 0.25. Simultaneously, if the local temperature in the combustion chamber exceeds the limit, the relaxation factor is further reduced to 0.125, and the temperature recovers.

[0103] A computational divergence recovery mechanism is established. During the computation process, the results are saved periodically, such as every 50 steps, retaining only the two most recent files. When the computation diverges, the previously saved results are retrieved, such as opening the results of the previous 50 steps, reducing the relaxation factor to 30% of the current value, and recalculating. For example, in anomaly recovery, the computation results are automatically saved every 50 steps (retaining the two most recent files). If the computation diverges at step 1500, the results of the previous 50 steps (step 1450) are retrieved, the relaxation factor is reduced to 30% of the current value (from 0.125 to 0.0375), and recalculation continues until convergence at step 2000.

[0104] Through the above specific implementation methods, the convergence efficiency of the three-dimensional simulation calculation of the E3 engine was improved by 40%, the calculation interruption rate was reduced from 35% to 8%, and the deviation between the results and experimental data was ≤5%. This fully demonstrates the effectiveness and superiority of the present invention in improving simulation convergence, robustness, and efficiency.

[0105] Secondly, embodiments of the present invention provide an electronic device, including a memory and a processor. The memory stores a computer program, and the processor executes the computer program to implement the steps of the above-described method. The specific steps are as described above and will not be repeated here.

[0106] Thirdly, embodiments of the present invention provide a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, performs the steps of the above-described method. The specific steps are as described above and will not be repeated here.

[0107] Specifically, computer-readable storage media, including both permanent and non-permanent, removable and non-removable media, can store information using any method or technology. Information can be computer-readable instructions, data structures, program modules, or other data. Examples of computer-readable storage media include, but are not limited to, phase-change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other memory technologies, CD-ROM, digital versatile optical disc (DVD) or other optical storage, magnetic tape, magnetic disk storage or other magnetic storage devices, or any other non-transferable medium that can be used to store information accessible by a computing device. As defined herein, computer-readable storage media does not include transient media, such as modulated data signals and carrier waves.

[0108] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, various modifications and variations can be made to the embodiments of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for three-dimensional simulation debugging of an aeroengine whole machine, characterized in that, The method comprises the following steps: The sub-component gradient pre-convergence step: before the whole machine coupling, the flow field of the compressor, the combustor and the turbine are adjusted independently, the flow field of each component is smoothly transitioned from the stable initial state to the design working condition through the progressive physical loading strategy, and the stability control is performed based on the real-time flow field monitoring results during the loading process; The debugging of the compressor comprises: establishing an initial flow field of a pressure gradient, and using a gradual synchronous loading mechanism of back pressure and rotating speed, and establishing a stall early warning mechanism, when the area of a cascade passage reverse flow region is greater than a set threshold, the loading is suspended and the back pressure is automatically adjusted. The back pressure is loaded in steps, and the back pressure increment of each step is , an initial back pressure of an iteration step n, an initial back pressure of an iteration step n-1, a target back pressure, a number of iteration steps; the rotating speed is loaded in steps, and the rotating speed increment of each step is , , a rotating speed of an iteration step n, a rotating speed of an iteration step n-1, a target rotating speed; the set threshold of the stall early warning mechanism is 5% of the area of the cascade passage; and the automatically adjusted back pressure is to reduce the current back pressure by 5%. The debugging of the combustion chamber comprises: sequentially performing solving a cold-state flow field under a closed combustion model, introducing a fuel injection model and calculating a mixing process of fuel and air, and activating a combustion reaction by gradually increasing a reaction rate coefficient; the reaction rate coefficient is according to a formula gradually increasing to a target value, wherein k is the reaction rate coefficient, n is the iteration step number, I is the total transition step number, is an initial coefficient of the reaction rate. The debugging of the turbine comprises: adopting a step-by-step activation mechanism for the turbine cold air system, and gradually increasing the cold air mass flow; the cold air mass flow is gradually increased according to , , is the iteration step number, is the n-step cold air amount, is the n-1-step cold air amount, is the target cold air flow, is the adjustment number, and ≥10. The whole machine dynamic adaptive coupling step: based on the pre-convergence of the component flow field, the whole machine coupling is performed, the dynamically adjusted calculation parameter strategy is adopted, and the flow field stability control is established; The dynamically adjusted calculation parameter strategy comprises: initially setting a relaxation factor and a Courant number for maintaining coupling stability for coupling; when it is monitored that the engine inlet flow fluctuation is stabilized within a preset range, starting and linearly increasing the relaxation factor and the Courant number; the initially set relaxation factor is 0.01, and the Courant number CFL is 1; the condition for starting to increase the relaxation factor and the Courant number is that the fluctuation amplitude of the engine inlet flow curve is less than 5% for 50 steps in succession; the relaxation factor is linearly increased by ; the Courant number is linearly increased by ; wherein n is the iteration step number, is the nth step relaxation factor, is the initial relaxation factor, is the maximum relaxation factor, I is the transition total step number, is the nth step Courant number; The establishing of the flow field stability control comprises: establishing a backflow monitoring mechanism, when detecting that the local backflow velocity exceeds a threshold value and lasts for a certain number of steps, reducing the relaxation factor of the local; the backflow monitoring mechanism is: when the backflow velocity and lasts for more than 100 steps, the relaxation factor of the local is reduced to 50% of the current value, wherein is the main flow velocity.

2. The method of claim 1, wherein, Before the sub-component gradient pre-convergence step, further comprising: The geometry and grid standardization step: the geometry of the engine aerodynamic core area is cleaned, non-aerodynamic features with a size less than 5% of the main flow diameter are removed, and the fillet with a radius less than 1 mm is transitioned to an acute angle; the whole machine calculation grid satisfying the minimum volume twist rate > 0.3, the maximum aspect ratio < 20, and the orthogonality > 0.6 is generated.

3. The method of claim 1, wherein, After the whole machine dynamic adaptive coupling step, further comprising: The simulation result post-processing and verification step: the whole machine thrust F and the specific fuel consumption SFC are calculated, and the calculation results are compared with the engine design values to verify the accuracy of the simulation results; The thrust ; The fuel consumption ; where F is the thrust, is the mass flow of air through the engine, is the average flow velocity of the gas flow through the engine nozzle exit section, is the average flow velocity of the gas flow through the engine inlet section, is the ambient pressure, is the static pressure of the gas flow through the engine nozzle exit section, is the area of the engine nozzle exit section, is the fuel flow.

4. The method of claim 1, wherein, After the whole machine dynamic adaptive coupling step, further comprising: The variable working condition simulation debugging and simulation process monitoring and abnormal handling step; The variable working condition simulation debugging comprises parameters based on two characteristic working condition points, and the parameters of the target working condition are obtained through an interpolation model; the variable working condition control rotation speed change rate is ≤500 r / min / s, and the fuel flow change rate is ≤0.05 kg / s²; for the acceleration process, a segmented loading strategy is adopted, in the low speed segment below 50% of the rated speed, the rotation speed increment of each step is ≤3%, and in the high speed segment above or equal to 50% of the rated speed, the rotation speed increment of each step is ≤1.5%; The simulation process monitoring and abnormal handling comprises: monitoring the slope of the flow and pressure curve, when the absolute value of the slope is greater than 0.01 for 20 steps, a calculation divergence warning is issued and the relaxation factor is reduced by half; detecting the physical field extreme value, when the pressure exceeds 1.5 times of the design value or the temperature exceeds 1.2 times of the design value, the relaxation factor is reduced by half; the calculation results are saved regularly, when the calculation diverges, the previously saved results are called, and the relaxation factor is reduced to 30% of the current value for re-computation. 5.An electronic device comprising a memory and a processor, the memory storing a computer program, wherein, The processor executes the computer program to realize the steps of the method of any one of claims 1 to 4.

6. A computer-readable storage medium having stored thereon a computer program, characterized in that, The computer program is run by the processor to perform the steps of the method of any one of claims 1 to 4.

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