A method, apparatus and storage medium for determining a pneumatic stability boundary

Abstract

A method, apparatus, and storage medium for determining aerodynamic stability boundaries are disclosed. The method includes: establishing a three-dimensional computational fluid dynamics (CFD) simulation model of all components of the engine; determining the corresponding fuel step parameters for the engine, and operating the compression system (fan or compressor) in the three-dimensional CFD simulation model under unstable conditions based on the fuel step parameters; obtaining the flow field parameters of the compression system under unstable conditions, and determining the aerodynamic stability boundary of the compression system based on the flow field parameters. This solution combines a full-component three-dimensional CFD simulation model with the fuel step method to achieve a high-precision aerodynamic stability boundary determination method that more closely approximates physical reality. It solves the problem of quickly and accurately obtaining the instability boundary of the compression system in the early stages of engine design without relying on physical prototype manufacturing and testing.

Description

Technical Field

[0001] This disclosure relates to, but is not limited to, the field of compressor technology, and particularly to a method, apparatus, and storage medium for determining aerodynamic stability boundaries. Background Technology

[0002] Compressors are widely used in the power and energy industries. With technological advancements and increasing application demands, compressors need to operate at higher loads to achieve higher performance. High-load compressors face more severe stability issues. Compressor performance is typically represented by characteristic curves. The left boundary of the characteristic curve is the compressor's instability boundary. When the compressor's operating conditions exceed this instability boundary, flow instability phenomena occur, such as surge. During surge, the airflow within the compressor oscillates significantly, ultimately leading to decreased compressor performance and structural damage, and in severe cases, endangering the safety of the driven vehicle or generator set. Therefore, quickly and accurately assessing the compressor's instability boundary during the compressor design phase is crucial for reducing compressor development costs and improving compressor performance and reliability.

[0003] Current engineering applications typically employ low-dimensional prediction methods and multi-dimensional coupled prediction methods. Low-dimensional prediction methods require extensive data for correction, exhibit poor universality across different computational cases, and fail to obtain detailed flow field information. While multi-dimensional coupled prediction methods utilize high-dimensional CFD (Computational Fluid Dynamics) modeling in the compression system portion, enabling the acquisition of relatively detailed flow field information, their modeling of boundary conditions leading to compression system instability often relies on empirical models or low-dimensional simulation models. This results in a coarse capture of downstream boundary conditions of the compression system, failing to fully simulate the overall machine environment in which the compression system operates, and ultimately failing to obtain high-precision stable boundaries of the compression system under the overall machine environment. Summary of the Invention

[0004] The following is an overview of the subject matter described in detail herein. This overview is not intended to limit the scope of the claims.

[0005] This disclosure provides a method, apparatus, and storage medium for determining aerodynamic stability boundaries, which can solve the problem of not being able to quickly and accurately obtain the instability boundaries of the compression system in the early stages of engine design without relying on physical prototype manufacturing and testing.

[0006] One embodiment of this disclosure provides a method for determining aerodynamic stability boundaries, including:

[0007] A three-dimensional computational fluid dynamics (CFD) simulation model of all components of the engine is established based on the engine's constituent parts.

[0008] Determine the fuel step parameters corresponding to the engine, and make the compression system in the three-dimensional CFD simulation model work under unstable conditions according to the fuel step parameters;

[0009] The flow field parameters of the compression system (fan and compressor) under unstable conditions are obtained, and the aerodynamic stability boundary of the compression system is determined based on the flow field parameters.

[0010] An embodiment of this disclosure also provides an aerodynamic stability boundary determination device, including: a memory and a processor;

[0011] The memory is used to store the program for determining the aerodynamic stability boundary;

[0012] The processor is configured to read the program for determining aerodynamic stability boundaries and execute the aerodynamic stability boundary determination method as described in any embodiment of this disclosure.

[0013] An embodiment of this disclosure also provides a non-transient computer-readable storage medium storing a computer program, wherein the computer program, when executed by a processor, is capable of implementing the aerodynamic stability boundary determination method as described in any embodiment of this disclosure.

[0014] Compared with related technologies, the aerodynamic stability boundary determination method, apparatus, and storage medium provided in this disclosure achieve high-dimensional and refined modeling by establishing a three-dimensional CFD simulation model of all main flow path components of the engine. This allows for the simulation of flow field details, thereby improving the accuracy of the obtained aerodynamic stability boundary. By determining the corresponding fuel step parameters of the engine and employing a fuel step method for pressure simulation, transient and continuous aerodynamic boundary changes can be achieved. This makes the simulation of the process of the compression system approaching the aerodynamic instability boundary more closely resemble physical reality, further improving the accuracy of the obtained aerodynamic stability boundary. This solution enables the simulation of the engine's overall instability boundary before whole-engine testing and even before engine prototype manufacturing, solving the problem of quickly and accurately obtaining the compression system instability boundary in the early stages of engine design without relying on physical prototype manufacturing and testing.

[0015] Other features and advantages of the embodiments of this application will be set forth in the following description, and will be apparent in part from the description, or may be learned by practicing the embodiments of this application. The objects and other advantages of the embodiments of this application may be realized and obtained by means of the structures particularly pointed out in the description, claims and drawings. Attached Figure Description

[0016] The accompanying drawings are used to provide an understanding of the technical solutions of this disclosure and form part of the specification. They are used together with the embodiments of this disclosure to explain the technical solutions of this disclosure and do not constitute a limitation on the technical solutions of this disclosure.

[0017] Figure 1 This is a flowchart of an embodiment of the aerodynamic stability boundary determination method disclosed herein;

[0018] Figure 2 This is a flowchart illustrating a specific method for determining aerodynamic stability boundaries according to an embodiment of this disclosure;

[0019] Figure 3 This is a schematic diagram of flow rate changes during a surge process according to an embodiment of the present disclosure;

[0020] Figure 4 This is a schematic diagram illustrating the slope change process of the flow rate change curve during surge according to an embodiment of this disclosure;

[0021] Figure 5 This is a schematic diagram of a surge dynamic process according to an embodiment of the present disclosure;

[0022] Figure 6 This is a schematic diagram of a full-engine simulation model according to an embodiment of the present disclosure;

[0023] Figure 7 This is a schematic diagram of an aerodynamic stability boundary determination device according to an embodiment of the present disclosure. Detailed Implementation

[0024] This disclosure describes several embodiments, but these descriptions are exemplary and not limiting, and it will be apparent to those skilled in the art that many more embodiments and implementations are possible within the scope of the embodiments described herein. Although many possible combinations of features are shown in the drawings and discussed in the detailed description, many other combinations of the disclosed features are also possible. Unless specifically limited, any feature or element of any embodiment may be used in combination with, or may replace, any feature or element of any other embodiment.

[0025] This disclosure includes and contemplates combinations of features and elements known to those skilled in the art. The embodiments, features, and elements disclosed in this disclosure may also be combined with any conventional features or elements to form a unique inventive scheme as defined by the claims. Any feature or element of any embodiment may also be combined with features or elements from other inventive schemes to form another unique inventive scheme as defined by the claims. Therefore, it should be understood that any feature shown and / or discussed in this disclosure may be implemented individually or in any suitable combination. Therefore, the embodiments are not limited except by the limitations imposed by the appended claims and their equivalents. Furthermore, various modifications and changes may be made within the scope of the appended claims.

[0026] In some engineering applications, there are low-dimensional prediction methods and high-dimensional prediction methods for predicting the stability boundary of compressors.

[0027] Low-dimensional prediction methods primarily establish appropriate physical or empirical parameter models to construct the relationship between compressor geometric characteristics, aerodynamic performance, and stability boundaries, thereby enabling rapid prediction of stability boundaries during the design phase. While this method is fast, it requires extensive data to refine the model parameters in the early stages, resulting in poor generalizability across different computational cases. Furthermore, the model construction process utilizes numerous adjustable parameters, most of which are empirical. These parameters require adjustment and refinement by the user, demanding significant engineering experience and a deep understanding of physics, making it difficult to learn. Additionally, low-dimensional methods cannot obtain detailed flow field information, making it challenging to provide detailed improvement suggestions for the specific design of the compression system based on the prediction results, thus limiting their guiding value in the design cycle.

[0028] While multidimensional coupled prediction methods utilize high-dimensional CFD modeling in the compression system to obtain relatively detailed flow field information, their modeling of boundary conditions leading to compression system instability often relies on empirical models or low-dimensional simulation models. This results in a coarse capture of boundary condition characteristics, and the results still require correction through actual system testing. In the application of multidimensional coupled prediction methods, the method of moving the compression system operating point to the instability position often involves repeatedly modifying the valve function of the downstream low-dimensional program of the compression system to approximate the actual compression system operating conditions. However, this approximation method is intermittent and cannot fully simulate the continuous approach to the instability point caused by the fuel step method in real tests, thus leading to a decrease in the accuracy of the prediction results.

[0029] One embodiment of this disclosure provides a method for determining aerodynamic stability boundaries, such as... Figure 1 As shown, the following steps may be included:

[0030] Step S110: Establish a three-dimensional computational fluid dynamics (CFD) simulation model of all components of the engine based on the engine's constituent parts;

[0031] Step S120: Determine the fuel step parameters corresponding to the engine, and make the compression system in the three-dimensional CFD simulation model work in an unstable condition according to the fuel step parameters;

[0032] Step S130: Obtain the flow field parameters of the compressor under unstable conditions, and determine the aerodynamic stability boundary of the compression system based on the flow field parameters.

[0033] For example, the engine may be an aircraft engine.

[0034] This embodiment provides a method for determining the aerodynamic stability boundary of a compression system under whole-machine conditions. By establishing a three-dimensional CFD simulation model of all engine components, high-dimensional and refined modeling is achieved, which can simulate flow field details and thus improve the accuracy of the obtained aerodynamic stability boundary. By determining the corresponding fuel step parameters of the engine and using the fuel step method for breath suppression, transient and continuous aerodynamic boundary changes can be realized. This makes the simulation of the compression system approaching the aerodynamic instability boundary more closely resemble physical reality, thereby further improving the accuracy of the obtained aerodynamic stability boundary. In summary, this embodiment achieves a high-precision method for determining aerodynamic stability boundaries by combining a full-component three-dimensional CFD simulation model and the fuel step method.

[0035] The aerodynamic stability boundary determination method in this embodiment can simulate the engine's overall instability boundary before whole-engine testing and even before engine prototype manufacturing. This solves the problem of quickly and accurately obtaining the compression system's instability boundary in the early stages of compression system design without relying on physical prototype manufacturing and testing. This is of great significance for engine development. For example, it allows for earlier design improvements, reducing development costs and increasing efficiency.

[0036] In an exemplary embodiment of this disclosure, the unstable condition can also be referred to as a surge condition, which is characterized by physical periodic oscillations. For example, when physical quantities (such as pressure, flow rate, etc.) used to characterize the performance of the compression system exhibit large-amplitude periodic oscillations, it indicates that the compression system has entered a surge operating state.

[0037] In an exemplary embodiment of this disclosure, the three-dimensional CFD simulation model may include the following sub-models: engine inlet extension model, engine intake duct model, compressor model, combustion chamber model, high-pressure turbine model, low-pressure turbine model, exhaust nozzle model, bypass duct model, and engine outlet extension model.

[0038] The three-dimensional CFD simulation model may also include some or all of the following sub-models: fan model, mixer and ejector model;

[0039] All sub-models are three-dimensional models, and all sub-models are suitable for URANS (Unsteady Reynolds-Averaged Navier-Stokes) simulation.

[0040] The aerodynamic stability boundary determination method in this embodiment achieves high-dimensional and refined modeling by performing full-component 3D CFD modeling of all main components of the real engine. This modeling can simulate the real 3D fluid flow of this type of engine, resulting in higher accuracy of the final aerodynamic stability boundary.

[0041] It should be noted that the model division in this embodiment is not limited to the above. As long as the established three-dimensional CFD simulation model includes models of all engine components, it is acceptable. For example, the fan model and the compressor model can also be combined into a compression system model.

[0042] For example, the simulation time in URANS is not required to capture the observed flow instability. The mesh scale used in the simulation should also be adjusted according to general requirements to capture the flow instability. The simulation can use a closed Reynolds stress turbulence model. The closed Reynolds stress turbulence model used is generally the SST (Shear Stress Transport) model, which has a strong ability to capture flow separation, or other turbulence models with the ability to capture flow separation.

[0043] In one example of this embodiment, the length of the engine inlet extension section of the engine inlet extension section model is greater than twice the diameter at the inlet and greater than once the length of the intake duct.

[0044] In the engine inlet extension section model, all outer boundaries of the engine inlet extension section are set to be open, except for the boundary connected to the engine intake duct, so that gas can freely enter and exit.

[0045] In one example of this embodiment, the engine intake duct model includes the main intake duct, possible branch ducts, and adjustment mechanisms corresponding to the branch ducts;

[0046] The intake of the air intake is the outlet of the extended section of the engine inlet;

[0047] For engines without a fan, the intake outlet is the compressor model inlet; for engines with a fan, the intake outlet is the fan model inlet.

[0048] In one example of this embodiment, the fan model can be based on the actual geometry of the engine fan components. For example, the fan model should perfectly match the actual geometry of the engine fan components, especially the fan blade geometry and tip clearance, which should be consistent with the actual operating state of a real engine.

[0049] For example, if the fan blades and blade clearance of an engine vary significantly under different operating conditions, then separate models should be created for each operating condition.

[0050] For example, if a fan has a multi-duct design, then all of its ducts should be included.

[0051] The inlet of the fan model is the outlet of the engine intake duct model; the outlet of the fan component is the inlet of the compressor model, or for a fan model with multi-bypass design features, its outlet can also be connected to the corresponding bypass duct model.

[0052] In one example of this embodiment, the compressor model can be built according to the actual geometry of an engine compressor. For example, the built compressor model can be completely based on the actual geometry of an engine compressor, especially the compressor blade geometry and tip clearance should be consistent with the actual engine operating conditions.

[0053] For example, if the compressor blade clearance varies significantly under different operating conditions, separate models should be built for each operating condition.

[0054] For example, if the compressor has a multi-duct design, it should include all its ducted compressor inlets corresponding to the engine intake duct model outlet or fan model outlet, with the compressor model outlet connected to the combustion chamber model inlet.

[0055] For example, if the fan and compressor components contain devices with variable geometry or involving changes in flow channels, they should be modeled separately according to their different operating conditions. It can also be understood that the three-dimensional CFD simulation model may include models of variable geometry devices and / or flow channel changing devices.

[0056] In one example of this embodiment, the combustion chamber model includes a combustion chamber shell and a flame tube structure; the outer cavity volume of the combustion chamber model is consistent with the outer cavity volume of the actual combustion chamber of the engine, and the inner volume of the combustion chamber flame tube of the combustion chamber model is consistent with the inner volume of the actual combustion chamber flame tube of the engine.

[0057] Wherein, the outer cavity volume of the combustion chamber model refers to the cavity volume formed by the outer shell of the combustion chamber; the inner volume of the combustion chamber flame tube of the combustion chamber model refers to the volume enclosed by the flame tube structure.

[0058] The shape and location of the holes on the flame tube should not be simplified; they should be modeled strictly according to the actual geometry.

[0059] The complex piping structure inside the combustion chamber can be simplified.

[0060] In this example, the combustion chamber model retains its outer shell and flame tube structure. The volume of the outer cavity of the combustion chamber and the volume of the flame tube inside the combustion chamber must be consistent with the physical reality so that the simulated combustion reaction is closer to the combustion reaction of a real engine.

[0061] For example, the combustion chamber model uses a vortex dissipation model, and the reaction rate is obtained by the following equation:

[0062]

[0063] Where: R i Here, A is the chemical reaction rate, ∈ is the model constant, k is the turbulent dissipation rate, and Y is the turbulent kinetic energy. F It is the mass fraction of fuel, Y O is the mass fraction of the oxidant, and s is the stoichiometric ratio.

[0064] For example, the combustion reaction can be simulated using a simplified multi-step reaction model or a probability function model.

[0065] For example, the type and composition of the working fluid in the engine model should be selected based on the chosen combustion model. The composition of the working fluid should change during the combustion reaction.

[0066] For example, the fuel injection quantity in the combustion chamber is the input parameter. Fuel is injected into the combustion chamber through a separate fuel inlet boundary set at the actual physical fuel nozzle location.

[0067] For example, the inlet of the combustion chamber model is connected to the outlet of the compressor model; the outlet of the combustion chamber model is connected to the inlet of the high-pressure turbine model.

[0068] The combustion chamber model in this example can directly simulate the phase change process of fuel oil and is applicable to all mainstream fuels.

[0069] In summary, compared to traditional single-component simulation models that only use an equivalent heat source model for simplified simulation, the combustion chamber model in this embodiment can cooperate with other models in the entire component to directly simulate the combustion reaction, making the final aerodynamic stability boundary closer to physical reality.

[0070] In one example of this embodiment, the established high-pressure turbine model and low-pressure turbine model should be completely based on the actual geometry of the engine.

[0071] For example, the throat area of ​​each row of blades in the model is given entirely according to the actual geometry and is not used as a separate input parameter. In particular, the turbine blade geometry and tip clearance should be consistent with those of a real engine under operating conditions.

[0072] For example, if the turbine blade clearance varies significantly under different engine operating conditions, separate models should be developed for each operating condition.

[0073] For example, if there are devices in the turbine components that have variable geometry or involve changes in the flow path, they should also be modeled separately according to their different operating conditions.

[0074] For example, the working fluid in the turbine component should be a variable-component working fluid, the composition of which should be automatically determined based on the progress of the combustion reaction in the combustion chamber.

[0075] For example, the turbine component inlet is connected to the combustion chamber component outlet; for engines without a mixer and ejector component, the turbine component outlet is connected to a mixer and ejector model; for engines without a mixer and ejector component, the turbine component outlet is connected to a tailpipe model.

[0076] In one example of this embodiment, the length and cross-sectional area of ​​the established bypass duct model, mixer and ejector model should be consistent with the actual geometry. As for the engine accessories and pipeline geometry in the bypass duct model, mixer and ejector model, they can be simplified accordingly.

[0077] For example, the mixer and ejector model is located between the turbine model and the tailpipe model. In addition to connecting to the turbine model, its inlet section will also connect to different bypass duct models depending on the engine type.

[0078] For example, depending on the engine type, the inlet of the bypass duct model is connected to the outlet of the fan bypass duct model or the outlet of the fan model, while the outlet of the compressor bypass duct model is connected to the mixer and ejector model or a separate tail nozzle model.

[0079] In one example of this embodiment, the established nozzle model should be completely based on the actual geometry of the engine. For nozzle models with adjustable capabilities, separate nozzle models should be established according to their different operating states.

[0080] For example, the nozzle model inlet is connected to the mixer and ejector model, the bypass duct model, or the turbine component model (high-pressure turbine model, low-pressure turbine model), and the nozzle model outlet is connected to the engine outlet extension.

[0081] In one example of this embodiment, the inlet position of the established engine outlet extension model is the outlet position of the tail nozzle model.

[0082] For example, in engine configurations with multiple nozzle components, a separate extension section should be established for each nozzle, or an integrated extension section component with multiple inlets should be established. The length of the extension section component should be at least three times the diameter of all nozzles. Except for the inlets, all other boundaries of the extension section should be designed to allow free flow of the working gas.

[0083] In one example of this embodiment, if the fan blade clearance of the engine changes beyond a preset fan blade clearance threshold under different operating conditions, then the fan model includes fan models under different operating conditions; if the compressor blade clearance of the engine changes beyond a preset compressor blade clearance threshold under different operating conditions, then the compressor model includes compressor models under different operating conditions; if the high-pressure turbine blade clearance of the engine changes beyond a preset high-pressure turbine blade clearance threshold under different operating conditions, then the high-pressure turbine model includes high-pressure turbine models under different operating conditions; if the low-pressure turbine blade clearance of the engine changes beyond a preset low-pressure turbine blade clearance threshold under different operating conditions, then the low-pressure turbine model includes low-pressure turbine models under different operating conditions.

[0084] For example, for all established component models, during the model calculation process, the flow information at the interface between component models, namely flow field parameters such as pressure, temperature, and velocity, is transmitted in real time.

[0085] The aerodynamic stability boundary determination method in this embodiment forms a complete 3D engine model for overall engine stability simulation by performing 3D CFD modeling on all main flow components of the engine. Since there are no similar engine design test results available for correction in the early design stages, the advantages of the 3D CFD simulation model in this embodiment are more pronounced, offering higher accuracy compared to existing simulation models. This makes the aerodynamic stability boundary determination method in this embodiment particularly suitable for the early design stages, enabling early detection and timely correction of problems, saving R&D costs and improving R&D efficiency.

[0086] In an exemplary embodiment of this disclosure, the step of causing the compression system in the three-dimensional CFD simulation model to operate under unstable conditions according to the fuel step parameters includes:

[0087] The fuel step parameters are confirmed based on the engine configuration;

[0088] The corresponding parameters of the three-dimensional CFD simulation model are set according to the fuel step parameters so that the compression system in the three-dimensional CFD simulation model operates under unstable conditions.

[0089] In one example of this embodiment, the fuel step parameters include fuel injection quantity increase time, fuel injection quantity increase factor, constant fuel injection quantity duration, and fuel injection quantity decrease time. The step of setting the corresponding parameters of the three-dimensional CFD simulation model according to the fuel step parameters, so that the compression system in the three-dimensional CFD simulation model operates under unstable conditions, includes:

[0090] The fuel injection quantity in the combustion chamber of the three-dimensional CFD simulation model is adjusted so that the fuel injection quantity of the three-dimensional CFD simulation model increases from the initial injection quantity to the fuel injection quantity increase multiple within the fuel injection quantity increase time.

[0091] After the fuel injection quantity in the three-dimensional CFD simulation model reaches the fuel injection quantity increase factor, the current fuel injection quantity is maintained until the constant fuel injection quantity is reached for a certain duration.

[0092] Reduce the fuel injection quantity of the three-dimensional CFD simulation model so that the fuel injection quantity of the three-dimensional CFD simulation model drops to the initial injection quantity during the fuel injection quantity fallback time.

[0093] In this embodiment, the fuel injection quantity in the combustion chamber model is increased by using the fuel step method, thereby realizing the continuous change of the aerodynamic boundary between the compressor and the combustion chamber, and then realizing data acquisition and calculation of the aerodynamic stability boundary.

[0094] In an exemplary embodiment of this disclosure, obtaining the flow field parameters of the compression system under unstable conditions and determining the aerodynamic stability boundary of the compression system based on the flow field parameters includes:

[0095] Numerical calculations are performed based on the fuel step parameters, and the changes in the flow field parameters are recorded throughout the entire instability condition.

[0096] Based on the change process of the flow field parameters, the moment when the aerodynamic stability boundary condition is satisfied is determined, and the value of the key flow field parameter at that moment is taken as the aerodynamic stability boundary.

[0097] For example, the flow field parameters may include parameters of one or more of the following physical quantities: temperature (total temperature, static temperature), pressure (total pressure, static pressure), density, and velocity; the key flow field may refer to a certain flow field parameter at a key cross section, such as the area-averaged static pressure at the outlet boundary of the compression system.

[0098] The recording of the changes in flow field parameters such as pressure and temperature at different cross-sectional locations can be done using measurement and recording methods commonly used in this field.

[0099] In one example of this embodiment, determining the moment when the aerodynamic stability boundary condition is satisfied based on the change process of the flow field parameters, and using the value of the flow field parameters at that moment as the aerodynamic stability boundary, may include:

[0100] The area-average static pressure at the outlet interface of the compression system is selected as P;

[0101] Determine the curve of the physical quantity P as a function of time, and calculate the slope S of the curve at each time t using the slope calculation formula. P (t); where S P (t) represents the slope of the physical quantity P at time t, where t represents the time of the curve;

[0102] Determine the moment when the stability boundary conditions are met, and use the physical quantity P corresponding to that moment as the stability boundary of the compressor.

[0103] For example, the slope calculation formula is:

[0104]

[0105] Among them, S P (t) represents the slope of the curve of the physical quantity P, W represents the window length for calculating the slope and is a positive integer; ΔP represents the change of the physical quantity P within the time length of W; f represents the sampling frequency;

[0106] For example, the stability boundary condition is:

[0107]

[0108] Among them, S P (N) represents the slope of physical quantity P at time N, where A represents the threshold; m c This represents the compressor flow rate. The physical meaning of this formula is that when the rate of change of the physical quantity P reaches A, and the compressor flow rate is positive at this point, the surge boundary is considered to have been reached. This operating condition is used as the compressor surge boundary point. The value of A is related to the different compressor configurations (such as axial compressors, centrifugal compressors, and combined axial-centrifugal compressors) of the aero-engine being calculated. In practical applications, it is usually selected based on existing calculation examples of similar engine configurations.

[0109] It should be noted that the method for determining the aerodynamic stability boundary is not limited to the method in this example that determines the surge boundary based on the change of the physical quantity P over time. Other methods can also be used, which will not be explained in detail here.

[0110] In summary, the aerodynamic stability boundary determination method disclosed herein, based entirely on three-dimensional modeling and employing a fuel step method to induce suffocation, can simulate real high-dimensional fluid flow and preserve flow details, and can directly simulate combustion reactions in the simulation model of combustion chamber components, thereby obtaining a more accurate and physically realistic start-up stability boundary.

[0111] The aerodynamic stability boundary determination method described in this disclosure will be described in its entirety below with a complete embodiment, such as... Figure 2 As shown, it may include the following steps:

[0112] Step S210: Establish a three-dimensional CFD simulation model of all components of the main channel of the aero-engine.

[0113] For example, the three-dimensional CFD simulation model includes the following sub-models: engine inlet extension section model, engine intake duct model, compressor model, combustion chamber model, high-pressure turbine model, low-pressure turbine model, exhaust nozzle model, bypass duct model, and engine outlet extension section model; the three-dimensional CFD simulation model also includes some or all of the following sub-models: fan model, mixer and ejector model; wherein, all sub-models are three-dimensional models;

[0114] In this example, the three-dimensional CFD model created for all components of the aero-engine is sufficient for URANS simulation.

[0115] It should be noted that the model division is not limited to the above; as long as the established 3D CFD simulation model includes models of all engine components, it is acceptable. For example, it can also be divided according to... Figure 6 The schematic diagram shown establishes the following models in the order of the main flow from upstream to downstream: engine intake duct model, compression system model (including fan and compressor), combustion chamber model, turbine model, exhaust nozzle model, other pipeline structure models, and exhaust nozzle external fluid domain extension section model.

[0116] Step S220: Use the fuel step method to make the compression system of the aero-engine operate in an unstable condition (i.e., surge condition).

[0117] In the constructed 3D CFD simulation model, the fuel injection quantity in the combustion chamber of the aero-engine is regulated so that the fuel injection quantity rapidly increases from the initial injection quantity to FS (the fuel injection quantity increase factor) times within time t1 (i.e., fuel injection quantity increase time), remains at a constant fuel injection quantity for time t2 (i.e., constant fuel injection quantity duration), and then rapidly decreases back to the initial injection quantity within time t3 (fuel injection quantity decrease time), thus achieving a fuel injection step. The values ​​of FS, t1, t2, and t3 can be determined based on the engine configuration. FS should be selected to ensure that the engine enters a surge operating state during the fuel injection step process.

[0118] For example, in practical applications, it can be determined based on data from existing aero-engines with similar configurations.

[0119] This step, by directly simulating combustion in the combustion chamber, obtains a more realistic "continuous change of the aerodynamic boundary between the compressor and the combustion chamber," thus making the surge process of the engine's compression system more realistic.

[0120] Step S230: Perform numerical calculations and record the surge dynamic process.

[0121] For example, this step may include: using the confirmed values ​​of FS, t1, t2, and t3, performing numerical calculations to record the changes in flow field parameters such as flow rate, pressure at different cross-sections of the engine, and temperature throughout the surge process.

[0122] Step S240: Analyze the surge dynamic process and extract the surge boundary.

[0123] Numerical calculations can be used to obtain the parameter changes of various physical quantities during surge. Then, the surge boundary of the compression system can be determined according to the following steps:

[0124] Step S241: First, obtain the time-varying process data of the physical quantity. For example... Figure 3 As shown, taking flow rate as an example, the change process of this physical quantity during surge is illustrated, which is characterized by periodic fluctuations accompanied by alternating positive and negative values.

[0125] Step S242: Calculate the slope S of the change process of physical quantity P. P The process of change of (t) is calculated using the following formula:

[0126]

[0127] Among them, S P (t) represents the slope of the curve of the physical quantity P, that is, how fast the physical quantity P changes; W represents the length of the window for calculating the slope, which is a positive integer; AP represents the amount of change of the physical quantity P within the time length of W; f represents the sampling frequency.

[0128] Step S243: Determine whether the stability boundary conditions are met. The determination conditions are as follows:

[0129]

[0130] Among them, S P (N) represents the slope of physical quantity P at time N, where A represents the threshold; m c This indicates the compressor flow rate.

[0131] Step S250: Change the model parameters and conduct tests at other speeds or under different operating conditions.

[0132] After completing the above parameter change process, you can continue to modify the rotational speed parameters in the model to complete the identification of stability boundaries at other rotational speeds.

[0133] This embodiment, based on a three-dimensional modeling method for determining aerodynamic stability boundaries, is the first to combine a full-component modeling method for aero-engines. This results in a high-dimensional, minimally simplified three-dimensional CFD simulation model capable of simulating realistic high-dimensional fluid flow without sacrificing flow field details, achieving a more refined simulation that facilitates design improvements. Furthermore, the subsequent use of a fuel step method for breath suppression more closely approximates physical reality, further enhancing accuracy. This aerodynamic stability boundary determination method allows for the determination of the engine's overall instability boundaries before full-engine testing and even before prototype fabrication, which is of significant importance for engine development. For example, it enables early design improvements, reducing development costs and increasing efficiency.

[0134] The following description uses a small turbofan engine equipped with a first-stage axial fan and a first-stage centrifugal compressor as an example to illustrate the complete embodiment described above, which may include the following steps:

[0135] Step S310. First, establish a three-dimensional CFD simulation model of the turbofan engine. For example, all main duct components of the engine, including the intake, fan, compressor, bypass duct, combustion chamber, turbine, and exhaust nozzle, can be modeled using three-dimensional URANS.

[0136] Step S320. Use the fuel step method to bring the turbofan engine into surge condition.

[0137] For example, based on the specific configuration and size of the small turbofan engine, FS = 3, t1 = 0.1s, t2 = 0.1s, and t3 = 0.1s were selected. Using the selected values, a fuel step simulation was performed, and it was confirmed, as described above, that the engine's compression system entered a surge condition during the fuel step process.

[0138] Step S330. After establishing the three-dimensional CFD simulation model, perform numerical calculations to obtain the flow rate change process during surge (e.g., Figure 3 (As shown).

[0139] Step S340. Select flow rate as the physical parameter for extracting the steady-state boundary, and based on the flow rate change process (e.g. Figure 3 As shown in the figure, the slope S of the curve of the change of physical quantity P during this process is calculated. P The process of change of (t) (e.g.) Figure 4 (As shown).

[0140] In this example, the window length for calculating the slope is W = 200, the sampling frequency is f = 200000, and the threshold is A = -50. Based on this, the time of the steady-state boundary point B can be calculated to be 0.0027s. At this point, the compressor's steady-state boundary point flow rate is 0.39 kg / s, and the pressure ratio is 4.46. This is the steady-state boundary point of the aero-engine compression system at this speed in this example.

[0141] At this point, the calculation of the engine's overall aerodynamic instability boundary under this operating condition has been completed. The instability boundary under other operating conditions can also be calculated using the same method.

[0142] Compared to traditional methods, the aerodynamic stability boundary determination method disclosed herein provides a more accurate prediction of the instability boundary of the compression system under full-engine conditions. Existing full-engine environment boundary calculation methods typically use the operation of a throttling valve downstream of the compression system to gradually reduce the operating flow rate of the compression system towards the instability point. This process is usually intermittent, which differs significantly from the continuous process of the operating point moving into the instability position during engine testing and daily use. The aerodynamic stability boundary determination method disclosed herein can simulate this continuous process while taking into account the complex influence of upstream and downstream components of the compression system on the instability boundary in detail. Figure 5 The figure shows the overall machine instability boundary obtained by this method. The black trajectory in the figure is the whole process of the compression system under the surge condition in the whole machine environment obtained by this method simulation. The circles in the black trajectory are the instability boundary points determined according to the method of this disclosure.

[0143] One embodiment of this disclosure also provides an aerodynamic stability boundary determination device, see [link to relevant documentation]. Figure 7 This includes: memory and processor;

[0144] The memory is used to store the program for determining the aerodynamic stability boundary;

[0145] The processor is configured to read the program for determining aerodynamic stability boundaries and execute the aerodynamic stability boundary determination method as described in any embodiment of this disclosure.

[0146] The processor in the embodiments of this disclosure can be a general-purpose processor, including a central processing unit (CPU), a network processor (NP), a microprocessor, etc., or other conventional processors. The processor can also be a digital signal processor (DSP), an application-specific integrated circuit (ASIC), an off-the-shelf programmable gate array (FPGA), discrete logic or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, or other equivalent integrated or discrete logic circuits, or a combination of the above devices. That is, the processor in the embodiments described above can be any processing device or combination of devices that implements the methods, steps, and logic block diagrams disclosed in the embodiments of this disclosure. If the embodiments of this disclosure are implemented in part in software, then instructions for software can be stored in a suitable non-volatile computer-readable storage medium, and one or more processors can be used to execute the instructions in hardware to implement the methods of the embodiments of this disclosure. The term "processor" as used herein can refer to the above-described structure or any other structure suitable for implementing the techniques described herein.

[0147] An embodiment of this disclosure also provides a non-transient computer-readable storage medium storing a computer program, wherein the computer program, when executed by a processor, can implement the aerodynamic stability boundary determination method as described in any embodiment of this disclosure.

[0148] In summary, existing engine compression system stability prediction models either fail to establish separate models for the upstream and downstream components of the compression system, or simply simplify these components into ideal pipeline and cavity structures for dimensionality reduction modeling. This disclosure, however, provides a method, apparatus, and storage medium for determining aerodynamic stability boundaries, which comprehensively considers the influence of upstream and downstream components and performs refined modeling, offering the following advantages compared to existing technologies:

[0149] Compared with traditional low-dimensional prediction methods, the aerodynamic stability boundary determination method disclosed in this paper realizes high-dimensional modeling of the entire engine, with fewer empirical parameters and adjustments in the calculation model, and can obtain high prediction accuracy without the need for a large amount of experimental data correction.

[0150] Compared with existing multi-dimensional coupling prediction methods, the aerodynamic stability boundary determination method disclosed in this paper performs a more refined simulation of the aerodynamic boundary of the engine compression system. While predicting the aerodynamic instability boundary, it can more accurately restore the flow field details near the engine instability condition, which is beneficial for users to further analyze the instability mechanism and propose targeted improvement measures in practical applications.

[0151] Compared to existing methods that bring the compression system close to the aerodynamic instability boundary, this method, based on a full three-dimensional simulation model of the entire machine, uses the fuel step method to realize transient and continuous aerodynamic boundary changes. This makes the simulation of the compression system approaching the aerodynamic instability boundary more closely resemble physical reality and can obtain more accurate results.

[0152] It will be understood by those skilled in the art that all or some of the steps, systems, or apparatuses disclosed above, and their functional modules / units, can be implemented as software, firmware, hardware, or suitable combinations thereof. In hardware implementations, the division between functional modules / units mentioned above does not necessarily correspond to the division of physical components; for example, a physical component may have multiple functions, or a function or step may be performed collaboratively by several physical components. Some or all components may be implemented as software executed by a processor, such as a digital signal processor or microprocessor, or as hardware, or as an integrated circuit, such as an application-specific integrated circuit (ASIC). Such software may be distributed on a computer-readable medium, which may include computer storage media (or non-transitory media) and communication media (or transient media). As is known to those skilled in the art, the term "computer storage medium" includes volatile and non-volatile, removable and non-removable media implemented in any method or technology for storing information (such as computer-readable instructions, data structures, program modules, or other data). Computer storage media include, but are not limited to, RAM, ROM, EEPROM, flash memory or other memory technologies, CD-ROM, digital versatile disc (DVD) or other optical disc storage, magnetic cartridges, magnetic tape, disk storage or other magnetic storage devices, or any other medium that can be used to store desired information and can be accessed by a computer. Furthermore, it is well known to those skilled in the art that communication media typically contain computer-readable instructions, data structures, program modules, or other data in modulated data signals such as carrier waves or other transmission mechanisms, and may include any information delivery medium.

Claims

1. A method of determining a pneumatic stability boundary, the method comprising: include: A three-dimensional computational fluid dynamics (CFD) simulation model of all components of the engine is established based on the engine's constituent parts. The fuel step parameters are determined based on the engine configuration. The fuel step parameters include fuel injection quantity increase time, fuel injection quantity increase factor, constant fuel injection quantity duration, and fuel injection quantity decrease time. The corresponding parameters of the three-dimensional CFD simulation model are set according to the fuel step parameters to make the compression system in the three-dimensional CFD simulation model work under unstable conditions. This includes: adjusting the fuel injection quantity in the combustion chamber of the three-dimensional CFD simulation model so that the fuel injection quantity of the three-dimensional CFD simulation model increases from the initial injection quantity to the fuel injection quantity increase multiple during the fuel injection quantity increase time; after the fuel injection quantity of the three-dimensional CFD simulation model reaches the fuel injection quantity increase multiple, maintaining the current fuel injection quantity until it reaches a constant fuel injection quantity for a certain duration; and reducing the fuel injection quantity of the three-dimensional CFD simulation model so that the fuel injection quantity of the three-dimensional CFD simulation model drops to the initial injection quantity during the fuel injection quantity fallback time. Obtaining the flow field parameters of the compression system under unstable conditions and determining the aerodynamic stability boundary of the compression system based on the flow field parameters includes: performing numerical calculations based on the fuel step parameters and recording the change process of the flow field parameters throughout the unstable conditions; determining the moment when the aerodynamic stability boundary condition is met based on the change process of the flow field parameters, and using the value of the key flow field parameter at that moment as the aerodynamic stability boundary.

2. The method for determining the aerodynamic stability boundary according to claim 1, characterized in that: The three-dimensional CFD simulation model includes the following sub-models: engine inlet extension section model, engine intake duct model, compressor model, combustion chamber model, high-pressure turbine model, low-pressure turbine model, tail nozzle model, bypass duct model, and engine outlet extension section model. The three-dimensional CFD simulation model also includes some or all of the following sub-models: fan model, mixer and ejector model; All of the sub-models are three-dimensional models.

3. The method for determining the aerodynamic stability boundary according to claim 2, characterized in that: The combustion chamber model includes a combustion chamber shell and a flame tube structure; the outer cavity volume of the combustion chamber model is consistent with the outer cavity volume of the actual combustion chamber of the engine, and the inner volume of the combustion chamber flame tube of the combustion chamber model is consistent with the inner volume of the flame tube of the actual combustion chamber of the engine. Wherein, the outer cavity volume of the combustion chamber model refers to the cavity volume formed by the outer shell of the combustion chamber; the inner volume of the combustion chamber flame tube of the combustion chamber model refers to the inner volume enclosed by the flame tube structure.

4. The method of claim 2, wherein, The combustion chamber model adopts the vortex dissipation model, and the reaction rate is obtained by the following formula: wherein, is the chemical reaction rate, is the model constant, is the turbulent dissipation rate, is the turbulent kinetic energy, is the mass fraction of fuel, is the mass fraction of oxidizer, is the stoichiometric ratio.

5. The method for determining the aerodynamic stability boundary according to claim 2, characterized in that: If the fan blade clearance of the engine changes beyond a preset fan blade clearance threshold under different operating conditions, then the fan model includes fan models under different operating conditions. If the compressor blade clearance of the engine changes beyond a preset compressor blade clearance threshold under different operating conditions, then the compressor model includes compressor models under different operating conditions. If the high-pressure turbine blade clearance of the engine changes beyond a preset high-pressure turbine blade clearance threshold under different operating conditions, then the high-pressure turbine model includes high-pressure turbine models under different operating conditions. If the low-pressure turbine blade clearance of the engine changes beyond a preset low-pressure turbine blade clearance threshold under different operating conditions, then the low-pressure turbine model includes low-pressure turbine models under different operating conditions.

6. A device for determining a pneumatic stability boundary, comprising: Memory and processor; characterized in that: The memory is used to store the program for determining the aerodynamic stability boundary; The processor is configured to read the program for determining aerodynamic stability boundaries and execute the aerodynamic stability boundary determination method as described in any one of claims 1 to 5.

7. A non-transitory computer-readable storage medium storing a computer program, wherein, When the computer program is executed by the processor, it can implement the aerodynamic stability boundary determination method as described in any one of claims 1 to 5.

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