Aero-engine key component life analysis reference cycle determination method and system

Abstract

The invention relates to the technical field of aero-engines, and discloses an aero-engine key component life analysis reference cycle determination method and system, and the method comprises the steps: comprehensively considering multi-dimensional factors based on a plurality of typical mission profiles of an aircraft, and determining the actual power of each flight state, calculating performance parameters corresponding to each flight state of each typical mission profile engine by using the engine performance simulation model; constructing an engine reference cycle to carry out transient temperature field analysis and transient stress analysis on the engine key component; determining the service life assessment key position of the key component according to the stress analysis result, and calculating the low-cycle fatigue life of the key component in combination with the material fatigue performance data; and carrying out statistical analysis on the engine load spectrum to obtain the total frequency and cycle number consumption of each section. According to the method, the low-cycle fatigue damage calculation precision of the engine structure under the service condition can be effectively improved, advanced failure or overuse of the structure caused by extensive damage evaluation is avoided, and economical efficiency and safety are both considered.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engines, and in particular to a method and system for determining a reference cycle for life analysis of key parts of an aero-engine. Background Art

[0002] The use of aircraft engine life-limited parts limits the maximum number of flight cycles allowed for each engine life-limited part. Based on verified analysis, testing, or operational experience, the combination of loads, material properties, environmental impacts, and operating conditions, including the effects of parts that affect these parameters, is fully understood or predicted. This ensures that each engine life-limited part reaches its approved service life and is removed from service before any harmful engine consequences occur. To ensure the reliability of the safe life analysis, the reliability of the predicted loads and material properties must be guaranteed. If the conditions or operating cycles used for the safe life assessment do not cover the majority of conditions during normal use, the safe life assessed and verified through testing will not be sufficiently reliable, posing a risk to safety during service.

[0003] Existing methods for calculating the lifespan of critical aircraft engine components typically use field data and design experience from similar aircraft models to determine a relatively conservative steady-state lifespan analysis, such as a 35-degree takeoff temperature. These methods then conduct safe lifespan assessments and test verification using steady-state heat transfer and stress analysis. During engine service, actual low-cycle consumption is calculated based on steady-state operating cycles, without considering the increased damage caused by transient thermal effects. This results in low reliability in the lifespan assessment of critical engine components. Summary of the Invention

[0004] In view of this, the present invention provides a method and system for determining the reference cycle for life analysis of key parts of an aircraft engine, so as to solve the problem that the life assessment of key parts of an engine in the prior art is not accurate enough, resulting in low accuracy in calculating the low-cycle fatigue damage of key parts of the engine under service conditions.

[0005] In a first aspect, the present invention provides a method for determining a reference cycle for life analysis of key components of an aircraft engine, comprising:

[0006] Based on multiple typical mission profiles of the aircraft, the actual power of the engine in each flight state is determined by comprehensively considering equipment power extraction, bleed air loss and performance degradation factors. The engine performance simulation model is then used to calculate the performance parameters corresponding to each flight state of the engine in each typical mission profile.

[0007] By integrating typical mission profiles, the state change process of the typical engine working process is formed and the engine reference cycle is constructed;

[0008] Based on the engine performance parameters and air system analysis results of each state of the engine reference cycle, transient temperature field analysis of key components is performed to obtain temperature field analysis results;

[0009] Perform transient stress analysis on key components based on the reference cycle and temperature field analysis results to obtain stress analysis results;

[0010] Based on the stress analysis results under the reference cycle, the key positions for life assessment of key components are determined, and the safe life of each key component of the engine under the reference cycle is calculated in combination with the material fatigue performance data;

[0011] The engine load spectrum is statistically analyzed to obtain the total frequency and cycle consumption of each section during the service life.

[0012] The method for determining a reference cycle for life analysis of key parts of an aircraft engine provided by an embodiment of the present invention accurately determines the flight state parameters of the engine through comprehensive consideration of multiple factors, constructs a reference cycle that can cover extreme operating conditions, and combines thermal-mechanical coupling analysis to achieve accurate calculation of the response of key parts; the safe life obtained based on this fully considers uncertain factors, and can also quantify the load management of the entire life cycle and is compatible with multiple models; ultimately, it achieves precise life assessment, improves the safety of aircraft engines, reduces maintenance costs, and enhances the economy and reliability of design.

[0013] In an optional embodiment, a typical engine reference cycle includes: start-idle-takeoff-cruise-idle-shutdown. For a turbofan engine or a turboprop engine, the reference cycle includes: start-idle-takeoff-cruise-idle-reverse thrust / reverse propeller-shutdown.

[0014] The embodiment of the present invention explicitly includes "start-slow-take-off-cruise-slow-stop" into the reference cycle, which completely covers the whole process of engine working conditions from start preparation, low-power preheating, high-power take-off, stable cruising to deceleration and shutdown. The speed, temperature, load and other parameters of the engine under these working conditions are significantly different. The reference cycle covers these states, which can simulate the various physical changes that the engine undergoes in actual operation, provide real and effective boundary conditions for the stress and temperature analysis of key components, and make the life assessment more in line with actual use. For turbofan engines or turboprop engines, additional reverse thrust / reverse propeller states are added to accurately adapt to the special working characteristics of such models that use reverse thrust devices to slow down during the landing phase, avoiding deviations in the life assessment of key components due to the lack of working conditions, effectively improving the calculation accuracy of low-cycle fatigue damage of engine structures under service conditions, avoiding premature failure or excessive use of structures due to rough damage assessment, and taking into account both economy and safety.

[0015] In an optional embodiment, the performance parameters include overall cross-sectional parameters and rotor speed; when determining the takeoff state of the life reference cycle, based on the calculation results of the overall cross-sectional parameters of the engine for each typical mission profile, the takeoff state corresponding to the most severe use environment of the takeoff state in all typical mission profiles is selected as the takeoff state of the reference cycle, wherein the most severe use environment is determined by the speed and the gas turbine rotor front temperature;

[0016] When determining the impact of engine performance degradation on overall cross-sectional parameters, the engine gas temperature at which the engine life is half degraded and the flight hours are minimum is used to represent the most severe operating environment in which the engine performance degradation affects the overall cross-sectional parameters;

[0017] Based on actual usage, determine the take-off state duration that is not less than the duration required by the typical mission profile, as well as the cruise state duration that maintains the heat conduction of the wheel to reach stability.

[0018] The present invention specifies that performance parameters encompass both overall cross-sectional parameters and rotor speed, providing comprehensive data support for engine operating condition analysis. Overall cross-sectional parameters reflect the engine's internal flow field characteristics, while rotor speed is directly related to power output. The combination of these two parameters accurately characterizes the engine's operating characteristics under different operating conditions, making subsequent life analysis based on these parameters more aligned with actual operating conditions and avoiding analysis biases caused by missing or incomplete parameters. By selecting the most severe takeoff operating environment from all typical mission profiles and using the engine gas temperature at half the life decay and minimum flight hours to represent the most severe performance decay scenario when determining the impact of engine performance degradation on overall cross-sectional parameters, this ensures that the reference cycle covers the extreme operating conditions the engine may face, avoiding the limitations of conventional operating condition analysis. This makes the life assessment of key components under extreme stresses and temperatures more conservative, identifies potential failure risks in advance, and effectively ensures the safety and reliability of the engine in complex and changing environments. A reasonable takeoff time setting aligns with actual flight operational requirements, avoiding load calculation errors caused by taking too short or too long a time. Stable cruise heat conduction ensures the accuracy of key component temperature field analysis, thereby improving the reliability of stress analysis.

[0019] In an optional embodiment, the transient temperature field analysis of key components based on the engine performance parameters and air system analysis results of each state of the engine reference cycle to obtain the temperature field analysis results includes:

[0020] Determine the overall performance parameters of the engine at each state point in the reference cycle according to the overall performance program, including flow, total temperature, and total pressure at each cross-sectional position of the engine;

[0021] Determine the engine chamber temperature and pressure corresponding to each state point of the reference cycle based on the overall performance parameters and the engine air system design results;

[0022] Determine the engine flow path gas temperature, pressure distribution, and blade surface fluid temperature and pressure based on the engine component performance analysis results;

[0023] Based on the performance analysis results of the air system and engine components, determine the boundary conditions for the temperature field analysis of key components, including the fluid temperature and heat transfer coefficient of key components;

[0024] Determine the temperature analysis load step length based on the entire time history of the reference cycle, and determine the boundaries of each load step by linear interpolation based on the thermal analysis boundary conditions at each state point of the reference cycle;

[0025] Perform transition state temperature field analysis to obtain the temperature distribution of key components at each time point in the reference cycle and the temperature change history of each position over time.

[0026] The embodiment of the present invention systematically integrates the performance parameters of the entire engine cross-section, collaboratively analyzes the air system and flow path loads, and accurately determines the thermal boundary conditions; dynamically sets the load step size based on the reference cycle to capture the transition temperature changes of key components; the high-precision temperature field analysis results thus obtained can accurately assess the thermal fatigue risk of key components, thereby providing a scientific basis for design optimization and maintenance strategy formulation, and significantly improving engine reliability and economy.

[0027] In an optional embodiment, performing transient stress analysis of key components based on the reference cycle and temperature field analysis results to obtain stress analysis results includes:

[0028] The centrifugal load corresponding to each load step is obtained by linear interpolation based on the load step length determined by temperature analysis and the speed value of each state point;

[0029] Centrifugal loads and temperature loads are applied to key components to perform transition state stress analysis, and the stress distribution of key components at each time point in the reference cycle and the stress change history of each position with time are obtained.

[0030] The embodiment of the present invention realizes the dynamic coupling of centrifugal and temperature loads by accurately interpolating the centrifugal load based on the temperature analysis load step; captures the transient stress changes of key components with synchronized load steps, and accurately characterizes the stress history of each position of the blade disk; the high-precision stress analysis results thus obtained can greatly improve the accuracy of fatigue damage assessment.

[0031] In an optional embodiment, determining the key positions for life assessment of key components based on stress analysis results under a reference cycle, and calculating the safe life of each key component of the engine under the reference cycle in combination with material fatigue performance data, includes:

[0032] Determine the stress history and temperature history at each position of the blade according to the stress analysis results;

[0033] Based on the stress analysis results, determine the maximum transition state equivalent stress value at the key position of the key component life assessment;

[0034] Based on the maximum transition state equivalent stress value, temperature field and fatigue performance data of key component materials, fatigue life analysis is carried out to obtain the low cycle fatigue life of key components. Combined with low cycle fatigue test verification, the verified safe life is obtained.

[0035] The embodiment of the present invention integrates stress and temperature history data to accurately locate high-risk positions of key components, with the maximum transition state equivalent stress as the core indicator; combined with the fatigue performance of the material, the fatigue life is converted into a safe life through safety factor adjustment, fully considering uncertain factors; the results can be directly used for maintenance cycle formulation and design improvement, effectively balancing safety and economy, and can not only accurately control the failure risk of key components, but also avoid the cost waste caused by excessive maintenance or premature failure.

[0036] In a second aspect, the present invention provides a system for determining a reference cycle for life analysis of key components of an aircraft engine, the system comprising:

[0037] The typical mission profile acquisition module is used to determine the actual power of the engine in each flight state based on the aircraft's typical mission profile, taking into account equipment power extraction, bleed air loss, and performance degradation factors. It also uses the engine performance simulation model to calculate the performance parameters corresponding to each flight state of the engine in each typical mission profile.

[0038] The engine reference cycle construction module is used to integrate various typical mission profiles, form the state change process of the engine's typical working history, and construct the engine reference cycle;

[0039] The temperature field analysis module is used to perform transient temperature field analysis of key components based on the engine performance parameters and air system analysis results of each state of the engine reference cycle to obtain temperature field analysis results;

[0040] Stress analysis module, used to perform transient stress analysis of key components based on reference cycle and temperature field analysis results to obtain stress analysis results;

[0041] The low-cycle fatigue life calculation module is used to determine the key positions for life assessment of key components based on the stress analysis results under the reference cycle, and calculate the low-cycle fatigue life of each key component of the engine under the reference cycle in combination with the material fatigue performance data;

[0042] The statistical analysis module is used to perform statistical analysis on the engine load spectrum to obtain the total frequency and cycle consumption of each section during the service life cycle.

[0043] In a third aspect, the present invention provides a computer device comprising: a memory and a processor, the memory and the processor being communicatively connected to each other, the memory storing computer instructions, and the processor executing the computer instructions to execute the method for determining a reference cycle for life analysis of key parts of an aircraft engine according to the first aspect or any corresponding embodiment thereof.

[0044] In a fourth aspect, the present invention provides a computer-readable storage medium having computer instructions stored thereon, the computer instructions being used to enable a computer to execute the method for determining a reference cycle for life analysis of critical parts of an aircraft engine according to the first aspect or any corresponding embodiment thereof.

[0045] In a fifth aspect, the present invention provides a computer program product comprising computer instructions for causing a computer to execute the method for determining a reference cycle for life analysis of critical components of an aerospace engine according to the first aspect or any corresponding embodiment thereof. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] In order to more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the specific embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0047] Figure 1 1 is a flow chart of a method for determining a reference cycle for life analysis of key components of an aircraft engine according to an embodiment of the present invention;

[0048] Figure 2 is a schematic diagram of the reference cycle for life analysis;

[0049] Figure 3 It is a schematic diagram of the engine blade structure;

[0050] Figure 4 is a structural block diagram of a system for determining a reference cycle for life analysis of key components of an aircraft engine according to an embodiment of the present invention;

[0051] Figure 5 Schematic diagram of the hardware structure of a computer device according to an embodiment of the present invention. DETAILED DESCRIPTION

[0052] To make the purpose, technical solutions, and advantages of the embodiments of the present invention more clear, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without making creative efforts shall fall within the scope of protection of the present invention.

[0053] Existing technology typically uses field data and design experience from similar aircraft models to determine a relatively conservative steady-state condition for life analysis, such as a 35-degree takeoff temperature. Safe life assessments and test verification are then conducted through steady-state heat transfer and stress analysis. During engine service, actual low-cycle consumption is calculated based on steady-state operating cycles, without considering the increased damage caused by transient thermal effects. This results in poor reliability in safe life analysis and poses risks to aircraft safety during service.

[0054] Therefore, this embodiment provides a method for determining a reference cycle for life analysis of key components of an aircraft engine. Based on the typical mission profile of an aircraft, while taking into account the influence of factors such as power margin, engine manufacturing dispersion, and performance degradation, a reference cycle for life analysis of key components of an engine is comprehensively determined. The reference cycle covers the most severe operating environment encountered in normal operation. The stress and temperature history of key components are calculated through the reference cycle, and then the safe life is calculated and tested for verification. This can effectively improve the reliability of the safe life assessment and obtain a verified safe life value in units of reference cycles. Figure 1 As shown, the process includes the following steps:

[0055] Step S101, based on multiple typical mission profiles of the aircraft, the actual power of the engine in each flight state is determined by comprehensively considering equipment power extraction, bleed air loss and performance degradation factors, and the performance parameters corresponding to each flight state of the engine in each typical mission profile are calculated using the engine performance simulation model.

[0056] Specifically, the aircraft platform of the embodiment of the present invention provides a typical mission spectrum, which is a combination of aircraft status, altitude, time, and power requirements. There are generally multiple typical mission profiles, as shown in Tables 1 and 2:

[0057] Table 1

[0058]

[0059] Table 2

[0060]

[0061] Furthermore, the embodiment of the present invention determines the actual engine power values ​​required for each state under the typical mission profile based on the power requirements of the aircraft platform's typical mission profile, taking into account the power extraction of equipment such as the aircraft platform's power generation, bleed air loss, installation loss (power margin), and engine performance degradation, as shown in Table 3 below.

[0062] Table 3

[0063]

[0064] Furthermore, the performance parameters corresponding to each flight state of the engine for each typical mission profile are calculated based on the engine performance simulation model, including rotor speed and overall cross-sectional parameters, as shown in Tables 4 and 5.

[0065] Table 4

[0066] parameter Speed ​​value (r / min) Engine rotor speed ***

[0067] Table 5

[0068] parameter Flow rate (kg / s) Total temperature (K) Total pressure (kPa) Engine import *** *** *** Compressor inlet *** *** *** compressor outlet *** *** *** Combustion chamber inlet *** *** *** Combustion chamber outlet *** *** *** Gas turbine rotor front *** *** *** Gas turbine outlet *** *** *** Power turbine inlet *** *** *** Power turbine outlet *** *** *** nozzle outlet *** *** ***

[0069] Step S102 : synthesize each typical task profile to form a state change process of a typical engine working history and construct an engine reference cycle.

[0070] Specifically, each typical task profile is comprehensively sorted out to form an engine state change process that can represent a typical working process, namely, the engine reference cycle, such as Figure 2 As shown in FIG, a typical reference cycle engine state history should include “start-idle-takeoff-cruise-idle-shutdown”. For turbofan engines or turboprop engines, it usually also includes reverse thrust / reverse propeller state, which can be defined as “start-idle-takeoff-cruise-idle-reverse thrust / reverse propeller-shutdown”.

[0071] The embodiments of the present invention explicitly incorporate the "start-slow-takeoff-cruise-slow-shutdown" cycle into the reference cycle, fully covering the entire engine operating process from start-up preparation, low-power preheating, high-power takeoff, stable cruising, to deceleration and shutdown. Under these operating conditions, engine parameters such as speed, temperature, and load vary significantly. The reference cycle covers these states and can simulate the various physical changes experienced by the engine in actual operation, providing realistic and effective boundary conditions for stress and temperature analysis of key components, making life assessment more consistent with actual usage. For turbofan or turboprop engines, an additional "reverse thrust / reverse propeller state" is added to accurately adapt to the special operating characteristics of these aircraft models, which use reverse thrust devices to slow down during landing. Taking a turbofan engine as an example, the reverse thrust state changes the direction of the engine's airflow, and parameters such as internal pressure and speed differ from those in the normal state. Including this state in the reference cycle avoids deviations in the life assessment of key components due to missing operating conditions, ensuring the integrity and reliability of life analysis for special aircraft models under all operating conditions.

[0072] The parameters of each state point in the typical reference cycle state used in the life analysis are shown in Table 6.

[0073] Table 6

[0074]

[0075] Specifically, when determining the maximum operating state "take-off state" of the life reference cycle, it is necessary to determine it based on the calculation results of the overall cross-sectional parameters of the engine for each typical mission profile, and select the most severe operating environment of the flight state among all typical mission profiles. The most severe environment is determined by the speed and the temperature before the gas turbine rotor. For example, the speed and temperature corresponding to the "take-off state" of typical mission profile one are N1 and T1, the speed and temperature corresponding to the "take-off state" of typical mission profile two are N2 and T2, and the speed and temperature corresponding to typical mission profile n are Nn and Tn. The take-off state corresponding to the maximum N and T needs to be selected as the take-off state of the reference cycle.

[0076] When determining the impact of engine performance degradation on overall cross-sectional parameters, the embodiments of the present invention consider the gas temperature of the engine with half life degradation (i.e., 1 / 2 of the impact of complete engine deterioration) and the minimum flight hours to represent the most severe operating environment, ensuring that the reference cycle covers the extreme operating conditions that the engine may face, avoiding the limitations of conventional operating condition analysis, making the life assessment of key components under extreme stresses and temperatures more conservative, identifying potential failure risks in advance, and effectively ensuring the safety and reliability of the engine in complex and changing environments.

[0077] The takeoff duration within the life reference cycle is determined based on actual use for a typical helicopter or fixed-wing aircraft, and should not be less than the duration required by the typical mission profile, typically defined as 5 minutes. The cruise duration within the life reference cycle is typically maintained for 5 to 15 minutes to stabilize the heat transfer from the wheel disc. This embodiment optimizes the takeoff time to meet actual flight operational requirements, avoiding load calculation errors caused by excessively short or long takeoff times. The stable cruise heat transfer state ensures the accuracy of critical component temperature field analysis, thereby improving the reliability of stress analysis.

[0078] Step S103 : Based on the engine performance parameters of each state point of the engine reference cycle and the air system analysis results, a transient temperature field analysis of key components is performed to obtain temperature field analysis results.

[0079] The embodiment of the present invention is Figure 3 Taking the reference cycle of the engine blade structure shown in the figure as an example, the transient temperature field analysis process under the reference cycle includes the following steps:

[0080] A1. Determine the overall performance parameters of the engine at each state point in the reference cycle according to the overall performance program, including flow rate, total temperature, and total pressure at each cross-sectional position of the engine;

[0081] A2, based on the overall performance parameters and the engine air system design results, determine the engine chamber temperature and pressure corresponding to each state point of the reference cycle;

[0082] A3, based on the engine component performance analysis results, determine the engine flow path gas temperature and pressure distribution and the blade surface fluid temperature and pressure;

[0083] A4. Based on the performance analysis results of the air system and engine components, determine the boundary conditions for the blade and disk temperature field analysis, including the fluid temperature and heat transfer coefficient at the blade and disk locations;

[0084] A5, based on the entire time history of the reference cycle, determine the temperature analysis load step length, and determine the boundaries of each load step by linear interpolation based on the thermal analysis boundary conditions at each state point of the reference cycle;

[0085] A6, perform transition temperature field analysis to obtain the blade disk temperature field corresponding to each load step of the reference cycle.

[0086] The embodiment of the present invention systematically integrates the performance parameters of the entire engine cross-section, collaboratively analyzes the air system and flow path loads, and accurately determines the thermal boundary conditions; dynamically sets the load step size based on the reference cycle to capture the transition temperature changes of key components; the high-precision temperature field analysis results thus obtained can be used to accurately assess the thermal fatigue risk of key components.

[0087] Step S104 , performing transient stress analysis on key components based on the reference cycle and temperature field analysis results to obtain stress analysis results.

[0088] Specifically, the process of performing transient stress analysis of key components in the embodiment of the present invention includes the following steps:

[0089] B1, obtain the centrifugal load corresponding to each load step by linear interpolation based on the load step determined by temperature analysis and the speed value of each state point;

[0090] B2, apply centrifugal load and temperature load to the key parts to perform transition state stress analysis, and obtain the stress distribution of the key parts at each time point in the reference cycle and the stress change history of each position over time.

[0091] Based on the load step size (e.g., 0.5 seconds / step) of the temperature analysis, the present embodiment converts the speed at each state point (e.g., 100% rated speed at takeoff) into a continuous centrifugal load through linear interpolation. Taking the blade disk as an example, during startup, the transient change in rotor speed from 0 to the rated value is discretized into multiple load steps. The centrifugal force within each step can be accurately calculated, avoiding the neglect of speed fluctuations in traditional steady-state analysis and reducing errors in the calculation of radial stress on the blade disk. The time-varying temperature load (e.g., the blade temperature rise curve over time) obtained from the transient temperature field analysis is coupled with the centrifugal load and applied to the blade disk, simulating the combined effects of thermal expansion and centrifugal force in actual operation.

[0092] The load step size is consistent with the temperature analysis, ensuring strict synchronization of centrifugal loads, temperature loads, and time histories. This avoids undercalculation of stress fluctuations due to load step mismatch, improves the timeliness of transition state stress analysis, and accurately depicts the stress history of key components at various locations. The high-precision stress analysis results obtained in this way can significantly improve the accuracy of fatigue damage assessment.

[0093] Step S105 , based on the stress analysis results under the reference cycle, key positions for life assessment of key components are determined, and low-cycle fatigue life of key components of the engine under the reference cycle is calculated in combination with material fatigue performance data.

[0094] Specifically, the following steps are included:

[0095] C1, determine the stress history and temperature history of each position of key parts based on the stress analysis results.

[0096] C2. Based on the stress analysis results, determine the maximum transition state equivalent stress value at the key position of the key component life assessment.

[0097] C3, based on the maximum transition state equivalent stress value, temperature field and fatigue performance data of key component materials, fatigue life analysis is carried out to obtain the low cycle fatigue life of key components. Combined with low cycle fatigue test verification, the verified safe life is obtained.

[0098] Specifically, the material fatigue performance data is the stress-life curve, that is, the SN curve, which ensures that the fatigue data used has a reliability of -3σ. Low cycle fatigue refers to the material or component under high stress amplitude and low cycle number (usually <10 4 ~10 6 The safe life refers to the minimum number of cycles under which a component does not experience fatigue failure under a specified reliability level (e.g., 90% confidence level) and safety factor. The embodiments of the present invention determine the fatigue life distribution of key components under target operating conditions (e.g., stress amplitude, temperature, and ambient medium), verify whether the design safety factor meets regulatory requirements through low-cycle fatigue testing, quantify reliability using tools such as the Weibull distribution, and introduce conservative factors based on engineering safety requirements. Ultimately, a safe life value that can be used for design is obtained, ensuring that key components experience extremely low probability of fatigue failure in actual service.

[0099] Based on stress analysis results, the present invention extracts stress-time and temperature-time curves at various locations on key components, enabling spatiotemporal coupled analysis of thermal and mechanical loads. For example, by synchronously recording stress amplitude and temperature fluctuations at the tenon of a turbine blade, it is possible to quantify the impact of the decrease in material yield strength at high temperatures on fatigue life, thus avoiding the overestimation of fatigue life that would result from ignoring temperature effects in a single stress analysis.

[0100] By identifying the maximum transition-state equivalent stress value, high-risk areas (such as the edges of the wheel hub holes and the bolt holes on the blade edge plates) can be precisely located. For example, in the takeoff state of a certain engine's high-pressure turbine wheel, the equivalent stress at the center of the hub plate reaches 480MPa, far exceeding that at the edge. This location is identified as a key location for life assessment, allowing subsequent life calculations to focus on the true weak links and improve analysis efficiency.

[0101] The safe life results obtained through analysis can be directly used to formulate maintenance cycles (such as replacing blades every 5,000 reference cycles) or design improvements (such as increasing the thickness of key positions), effectively balancing safety and economy, accurately controlling the risk of failure of key components, and avoiding cost waste caused by excessive maintenance or premature failure.

[0102] Step S106 , statistically analyzing the engine load spectrum to obtain the total frequency and cycle consumption of each section within the service life period.

[0103] The embodiment of the present invention performs a statistical analysis on the engine load spectrum as shown in Table 7, based on the duration of each typical mission profile and the frequency ratio within the service life cycle (such as 10,000 hours), and the total frequency of each profile within the service life cycle.

[0104] Table 7

[0105]

[0106] The embodiments of the present invention collect engine mission profile data in real time (such as recording the load history of each flight through sensors) and compare it with the total frequency model in the design phase. This can dynamically evaluate the engine health status and identify abnormal load accumulation in advance (for example, if a single profile load exceeds the standard due to a malfunction on a certain aircraft, the risk can be quickly located through total frequency statistics). Total frequency analysis can assist in optimizing task scheduling. For example, while ensuring the task completion rate, the task allocation of different load profiles can be adjusted to balance the damage and consumption of various engine components and extend the overall service life.

[0107] The embodiment of the present invention further uses the rain flow method to count the number of various cycles for each typical mission profile within the service life cycle (Table 8), and converts the equivalent fatigue damage consumption under each cycle (in reference cycles) based on the stress and life analysis results of each cycle to obtain the cycle consumption C (in reference cycles) within the service life cycle. See Table 9 for details.

[0108] Table 8

[0109] Number of cycles (times) 10,000-hour life cycle Category I cycle (0-takeoff-0) S1 Category II cycle (slow-maximum-slow) S2 Category III cycle (cruise-max-cruise) S3

[0110] Table 9

[0111]

[0112] The embodiment of the present invention performs damage analysis on a typical mission profile and uses the rain flow counting method to decompose the irregular load time history (such as engine speed fluctuations and temperature fluctuations) into several stress cycles, thus solving the problem of quantifying the non-periodicity of the actual load. The main cycle and secondary cycle parameters are determined and converted into reference cycle consumption, which improves the accuracy of fatigue damage calculation. The total fatigue damage consumption is measured in "reference cycle numbers" to intuitively reflect the life consumption progress of key engine parts. For example, when C approaches the design safe life (such as 6000 reference cycles), a maintenance warning is automatically triggered to avoid premature failure or overuse due to ambiguous damage assessment.

[0113] This embodiment also provides a system for determining reference cycles for life analysis of key aircraft engine components. This system is used to implement the above-mentioned embodiments and preferred implementations, and details already described will not be repeated. As used below, the term "module" may refer to a combination of software and / or hardware that implements a predetermined function. Although the systems described in the following embodiments are preferably implemented in software, implementation using hardware, or a combination of software and hardware, is also possible and contemplated.

[0114] This embodiment provides a system for determining a reference cycle for life analysis of key parts of an aircraft engine. Figure 4 Shown, including:

[0115] The typical mission profile acquisition module 401 is used to determine the actual power of the engine in each flight state based on the aircraft's typical mission profile, taking into account equipment power extraction, bleed air loss, and performance degradation factors, and calculate the performance parameters corresponding to each flight state of the engine in each typical mission profile using the engine performance simulation model;

[0116] The engine reference cycle construction module 402 is used to synthesize various typical mission profiles to form a state change process of a typical engine working history and construct an engine reference cycle;

[0117] The temperature field analysis module 403 is used to perform transient temperature field analysis of key components based on the engine performance parameters of each state of the engine reference cycle and the air system analysis results to obtain temperature field analysis results;

[0118] The stress analysis module 404 is used to perform transient stress analysis of key components based on the reference cycle and temperature field analysis results to obtain stress analysis results;

[0119] The low-cycle fatigue life calculation module 405 is used to determine the key positions for life assessment of key components based on the stress analysis results under the reference cycle, and calculate the low-cycle fatigue life of each key component of the engine under the reference cycle in combination with the material fatigue performance data;

[0120] The statistical analysis module 406 is used to perform statistical analysis on the engine load spectrum to obtain the total frequency and cycle consumption of each section within the service life period.

[0121] In some optional embodiments, a typical engine reference cycle includes: start-idle-takeoff-cruise-idle-shutdown. For a turbofan engine or a turboprop engine, the reference cycle includes: start-idle-takeoff-cruise-idle-reverse thrust / reverse propeller-shutdown.

[0122] In some optional embodiments, the performance parameters include overall cross-sectional parameters and rotor speed. When determining the takeoff state of the life reference cycle, based on the calculation results of the overall cross-sectional parameters of the engine for each typical mission profile, the takeoff state corresponding to the most severe use environment of the takeoff state in all typical mission profiles is selected as the takeoff state of the reference cycle, wherein the most severe use environment is determined by the speed and the gas turbine rotor front temperature. When determining the impact of engine performance degradation on the overall cross-sectional parameters, the engine gas temperature at half the life degradation and the minimum flight hours is used to represent the most severe use environment in which the engine performance degradation affects the overall cross-sectional parameters.

[0123] Based on actual usage, determine the take-off state duration that is not less than the duration required by the typical mission profile, as well as the cruise state duration that maintains the heat conduction of the wheel to reach stability.

[0124] In some optional implementations, the temperature field analysis module 403 includes:

[0125] An overall performance parameter acquisition unit, used to determine the overall performance parameters of each state point of the engine reference cycle according to the overall performance program, including the flow rate, total temperature and total pressure at each cross-sectional position of the engine;

[0126] An engine chamber temperature and pressure acquisition unit, configured to determine the engine chamber temperature and pressure corresponding to each state point of a reference cycle based on overall performance parameters combined with engine air system design results;

[0127] A performance analysis unit, used to determine the engine flow path gas temperature and pressure distribution and the blade surface fluid temperature and pressure based on the engine component performance analysis results;

[0128] The temperature field analysis boundary acquisition unit is used to determine the temperature field analysis boundary conditions of key components based on the performance analysis results of the air system and engine components, including the fluid temperature and heat transfer coefficient of the key components;

[0129] The load step boundary acquisition unit is used to determine the temperature analysis load step length based on the entire time history of the reference cycle, and to determine the boundaries of each load step by linear interpolation based on the thermal analysis boundary conditions of each state point of the reference cycle;

[0130] The key component temperature field analysis unit is used to perform transition state temperature field analysis to obtain the temperature distribution of key components at each time point in the reference cycle and the temperature change history of each position over time.

[0131] In some optional implementations, the stress analysis module 404 includes:

[0132] A centrifugal load acquisition unit is used to obtain the centrifugal load corresponding to each load step by linear interpolation based on the load step determined by temperature analysis and the speed value of each state point;

[0133] The stress analysis unit is used to apply centrifugal loads and temperature loads to key components to perform transitional stress analysis, thereby obtaining the stress distribution of key components at each time point in the reference cycle and the stress change history of each position over time.

[0134] In some optional implementations, the low cycle fatigue life calculation module 405 includes:

[0135] Determine the stress history and temperature history of each position of key parts based on the stress analysis results;

[0136] Based on the stress analysis results, determine the maximum transition state equivalent stress value at the key position of the key component life assessment;

[0137] Based on the maximum transition state equivalent stress value, temperature field and fatigue performance data of key component materials, fatigue life analysis is carried out to obtain the low cycle fatigue life of key components. Combined with low cycle fatigue test verification, the verified safe life is obtained.

[0138] The further functional description of each of the above modules and units is the same as that of the above corresponding embodiments and will not be repeated here.

[0139] The aircraft engine key parts life analysis reference cycle determination system in this embodiment is presented in the form of a functional unit, where the unit refers to an ASIC (Application Specific Integrated Circuit) circuit, a processor and memory that executes one or more software or fixed programs, and / or other devices that can provide the above functions.

[0140] The embodiment of the present invention also provides a computer device having the above Figure 4 The reference cycle determination system for life analysis of key parts of an aircraft engine is shown.

[0141] See also Figure 5 , Figure 5 is a schematic diagram of the structure of a computer device provided by an optional embodiment of the present invention, such as Figure 5 As shown, the computer device includes: one or more processors 10, memory 20, and interfaces for connecting various components, including high-speed interfaces and low-speed interfaces. Various components utilize different buses to communicate with each other and can be installed on a common mainboard or installed in other ways as needed. The processor can process the instructions executed in the computer device, including instructions stored in the memory or on the memory to display the graphical information of the GUI on an external input / output device (such as, a display device coupled to the interface). In some optional embodiments, if necessary, multiple processors and / or multiple buses can be used together with multiple memories and multiple memories. Equally, multiple computer devices can be connected, and each device provides part of the necessary operations (for example, as a server array, a group of blade servers, or a multi-processor system). Figure 5 A processor 10 is taken as an example.

[0142] The processor 10 may be a central processing unit, a network processor, or a combination thereof. The processor 10 may further include a hardware chip. The hardware chip may be an application-specific integrated circuit, a programmable logic device, or a combination thereof. The programmable logic device may be a complex programmable logic device, a field programmable gate array, a general purpose array logic, or any combination thereof.

[0143] The memory 20 stores instructions that can be executed by at least one processor 10, so as to enable at least one processor 10 to execute the method shown in the above embodiment.

[0144] The memory 20 may include a program storage area and a data storage area, wherein the program storage area may store an operating system and application programs required for at least one function; the data storage area may store data created based on the use of the computer device, etc. In addition, the memory 20 may include a high-speed random access memory, and may also include a non-transient memory, such as at least one disk storage device, a flash memory device, or other non-transient solid-state storage device. In some optional embodiments, the memory 20 may optionally include a memory remotely located relative to the processor 10, and these remote memories may be connected to the computer device via a network. Examples of the above-mentioned network include, but are not limited to, the Internet, an intranet, a local area network, a mobile communication network, and combinations thereof.

[0145] The memory 20 may include a volatile memory, such as a random access memory; the memory may also include a non-volatile memory, such as a flash memory, a hard disk or a solid-state drive; the memory 20 may also include a combination of the above types of memory.

[0146] The computer device further includes a communication interface 30 for the computer device to communicate with other devices or a communication network.

[0147] The embodiment of the present invention also provides a computer-readable storage medium. The above-mentioned method according to the embodiment of the present invention can be implemented in hardware, firmware, or implemented as a computer code that can be recorded in a storage medium, or implemented as a computer code that is originally stored in a remote storage medium or a non-temporary machine-readable storage medium and downloaded through a network and will be stored in a local storage medium, so that the method described herein can be stored in such software processing on a storage medium using a general-purpose computer, a dedicated processor, or programmable or dedicated hardware. Among them, the storage medium can be a magnetic disk, an optical disk, a read-only storage memory, a random access memory, a flash memory, a hard disk or a solid-state drive, etc.; further, the storage medium can also include a combination of the above-mentioned types of memory. It can be understood that a computer, a processor, a microprocessor controller or programmable hardware includes a storage component that can store or receive software or computer code. When the software or computer code is accessed and executed by a computer, a processor or hardware, the method shown in the above embodiment is implemented.

[0148] A portion of the present invention may be applied as a computer program product, such as a computer program instruction, which, when executed by a computer, can call or provide the method and / or technical solution according to the present invention through the operation of the computer. Those skilled in the art should understand that the form in which the computer program instruction exists in a computer-readable medium includes, but is not limited to, a source file, an executable file, an installation package file, etc. Accordingly, the way in which the computer program instruction is executed by the computer includes, but is not limited to: the computer directly executes the instruction, or the computer compiles the instruction and then executes the corresponding compiled program, or the computer reads and executes the instruction, or the computer reads and installs the instruction and then executes the corresponding installed program. Here, the computer-readable medium may be any available computer-readable storage medium or communication medium that can be accessed by the computer.

[0149] Although the embodiments of the present invention have been described with reference to the accompanying drawings, those skilled in the art may make various modifications and variations without departing from the spirit and scope of the present invention. Such modifications and variations are all within the scope defined by the appended claims.

Claims

1. A method for determining a reference cycle for life analysis of key parts of an aircraft engine, characterized in that: include: Based on multiple typical mission profiles of the aircraft, the actual power of the engine in each flight state is determined by comprehensively considering equipment power extraction, bleed air loss and performance degradation factors. The engine performance simulation model is then used to calculate the performance parameters corresponding to each flight state of the engine in each typical mission profile. By integrating typical mission profiles, the state change process of the typical engine working process is formed and the engine reference cycle is constructed; Based on the engine performance parameters and air system analysis results at each state point of the engine reference cycle, transient temperature field analysis of key components is performed to obtain temperature field analysis results; Perform transient stress analysis on key components based on the reference cycle and temperature field analysis results to obtain stress analysis results; Based on the stress analysis results under the reference cycle, the key positions for the life assessment of key components are determined, and the low-cycle fatigue life of each key component of the engine under the reference cycle is calculated in combination with the material fatigue performance data; The engine load spectrum is statistically analyzed to obtain the total frequency and cycle consumption of each section during the service life.

2. The method according to claim 1, characterized in that A typical engine reference cycle includes: start-idle-takeoff-cruise-idle-shutdown. For a turbofan engine or turboprop engine, the reference cycle includes: start-idle-takeoff-cruise-idle-reverse thrust / reverse propeller-shutdown.

3. The method according to claim 2, characterized in that The performance parameters include overall cross-sectional parameters and rotor speed. When determining the takeoff state of the life reference cycle, based on the calculation results of the overall cross-sectional parameters of the engine for each typical mission profile, the takeoff state corresponding to the most severe use environment of the takeoff state in all typical mission profiles is selected as the takeoff state of the reference cycle, wherein the most severe use environment is determined by the speed and the gas turbine rotor front temperature. When determining the impact of engine performance degradation on overall cross-sectional parameters, the engine gas temperature at which the engine life is half degraded and the flight hours are minimum is used to represent the most severe operating environment in which the engine performance degradation affects the overall cross-sectional parameters; Based on actual usage, determine the take-off state duration that is not less than the duration required by the typical mission profile, as well as the cruise state duration that maintains the heat conduction of the wheel to reach stability.

4. The method according to claim 1, wherein The transient temperature field analysis of key components based on the engine performance parameters and air system analysis results at each state point of the engine reference cycle to obtain temperature field analysis results includes: Determine the overall performance parameters of the engine at each state point in the reference cycle according to the overall performance program, including flow, total temperature, and total pressure at each cross-sectional position of the engine; Determine the engine chamber temperature and pressure corresponding to each state point of the reference cycle based on the overall performance parameters and the engine air system design results; Determine the engine flow path gas temperature, pressure distribution, and blade surface fluid temperature and pressure based on the engine component performance analysis results; Based on the performance analysis results of the air system and engine components, determine the boundary conditions for the temperature field analysis of key components, including the fluid temperature and heat transfer coefficient of key components; Determine the temperature analysis load step length based on the entire time history of the reference cycle, and determine the boundaries of each load step by linear interpolation based on the thermal analysis boundary conditions at each state point of the reference cycle; Perform transition state temperature field analysis to obtain the temperature distribution of key components at each time point in the reference cycle and the temperature change history of each position over time.

5. The method according to claim 4, characterized in that The transient stress analysis of key components is performed based on the reference cycle and temperature field analysis results to obtain stress analysis results, including: The centrifugal load corresponding to each load step is obtained by linear interpolation based on the load step length determined by temperature analysis and the speed value of each state point; Centrifugal loads and temperature loads are applied to key components to perform transition state stress analysis, and the stress distribution of key components at each time point in the reference cycle and the stress change history of each position with time are obtained.

6. The method according to claim 5, characterized in that The stress analysis results under the reference cycle are used to determine the key positions for the life assessment of key components, and the low-cycle fatigue life of each key component of the engine under the reference cycle is calculated in combination with the material fatigue performance data, including: Determine the stress history and temperature history of each position of key parts based on the stress analysis results; Based on the stress analysis results, determine the maximum transition state equivalent stress value at the key position of the key component life assessment; Based on the maximum transition state equivalent stress value, temperature field and fatigue performance data of key component materials, fatigue life analysis is carried out to obtain the low cycle fatigue life of key components. Combined with low cycle fatigue test verification, the verified safe life is obtained.

7. A system for determining reference cycles for life analysis of key parts of an aircraft engine, characterized in that: include: The typical mission profile acquisition module is used to determine the actual power of the engine in each flight state based on the aircraft's typical mission profile, taking into account equipment power extraction, bleed air loss, and performance degradation factors. It also uses the engine performance simulation model to calculate the performance parameters corresponding to each flight state of the engine in each typical mission profile. The engine reference cycle construction module is used to integrate various typical mission profiles, form the state change process of the engine's typical working history, and construct the engine reference cycle; The temperature field analysis module is used to perform transient temperature field analysis of key components based on the engine performance parameters and air system analysis results of each state of the engine reference cycle to obtain temperature field analysis results; Stress analysis module, used to perform transient stress analysis of key components based on reference cycle and temperature field analysis results to obtain stress analysis results; The low-cycle fatigue life calculation module is used to determine the key positions for life assessment of key components based on the stress analysis results under the reference cycle, and calculate the low-cycle fatigue life of each key component of the engine under the reference cycle in combination with the material fatigue performance data; The statistical analysis module is used to perform statistical analysis on the engine load spectrum to obtain the total frequency and cycle consumption of each section during the service life cycle.

8. A computer device, characterized in that: include: A memory and a processor, wherein the memory and the processor are communicatively connected to each other, the memory stores computer instructions, and the processor executes the method for determining a reference cycle for life analysis of key parts of an aircraft engine according to any one of claims 1 to 6 by executing the computer instructions.

9. A computer-readable storage medium, characterized in that The computer-readable storage medium stores computer instructions, which are used to enable a computer to execute the method for determining a reference cycle for life analysis of key components of an aircraft engine according to any one of claims 1 to 6.

10. A computer program product, characterized in that The method comprises computer instructions for causing a computer to execute the method for determining a reference cycle for life analysis of a key component of an aircraft engine according to any one of claims 1 to 6.

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