A method for designing load spectrum of aviation turboshaft engine based on thermal damage number

Abstract

The present invention relates to the technical field of aviation turboshaft engines, and provides a method for designing a load spectrum of an aviation turboshaft engine based on a thermal damage number, comprising: according to information of multiple typical mission profiles of the engine, using an engine performance model, analyzing the power demand in each typical mission profile, obtaining the engine state and its duration in each mission profile under a first application scenario, i.e., the engine load spectrum parameters in the first application scenario; fitting the relationship between the cumulative thermal damage and temperature of the working blades of the engine gas turbine; calculating the cumulative thermal damage of the engine under different temperature durations in each state under the first application scenario based on the relationship; applying the engine to a second application scenario, and keeping the cumulative thermal damage unchanged, obtaining the duration and time proportion of each state of the engine in the engine load spectrum, i.e., the engine load spectrum parameters in the second application scenario. The present invention can effectively solve the rationality and accuracy problems of conversion when performing equivalent conversion of each state for a newly developed engine.

Description

Technical Field

[0001] The invention relates to the technical field of aviation turboshaft engines, and in particular to a load spectrum design method for aviation turboshaft engines based on thermal damage numbers. Background Art

[0002] The engine design load spectrum depends on the time history of the engine working state parameters and flight state parameters. The engine working state parameters mainly include power state, speed, temperature, etc., and the flight state parameters mainly include flight altitude, flight speed, atmospheric temperature, etc. The design load spectrum is compiled based on the above parameters and duration.

[0003] The load spectrum design method of the prior art is mainly based on the flight mission profile or the existing engine field use data. Through the engine performance model calculation or rain flow counting statistical method, the engine gas generator speed range and turbine temperature related time matrix or the power turbine speed range and turbine temperature related time matrix is ​​obtained, and then compared with the speed or temperature of the engine in each state, the time proportion of each state of the engine is obtained. However, there are large differences in the use of engines in different regions. For example, the power of engines used in plateau areas is greatly reduced compared to that in plain areas, resulting in a larger maximum state proportion required for the use of engines in plateaus. Under the condition that the engine capacity remains unchanged, the time proportion of other states must be reduced accordingly. However, how to adjust the time proportion of each state of the engine, the prior art lacks conversion formulas, or uses the empirical coefficients of the relationship between each state obtained from the use of basic engines for equivalence. However, the empirical method is only applicable to the load spectrum design of improved engines. Due to the lack of historical data support, the compilation of the load spectrum of the newly developed engine cannot be correctly converted.

[0004] The existing empirical method for equivalent conversion of various states of engine load spectrum is limited to the field of load spectrum design related to improved engines supported by empirical data of basic engines. When the newly developed engines need to perform equivalent conversion of various states, it is unable to solve the problems of rationality and accuracy of the conversion, which makes the existing technology have great limitations in actual engineering applications and cannot meet the current design needs of newly developed dynamic load spectrum. Summary of the invention

[0005] The purpose of the present invention is to solve at least one technical problem in the background technology and to provide a method for designing the load spectrum of an aviation turboshaft engine based on thermal damage number.

[0006] To achieve the above object, the present invention provides a method for designing a load spectrum of an aviation turboshaft engine based on a thermal damage number, comprising:

[0007] According to information of multiple typical mission profiles of the engine, the power demand in each typical mission profile is analyzed by using an engine performance model, and the engine state and duration of each mission profile under the first application scenario are obtained, that is, the engine load spectrum parameters under the first application scenario;

[0008] Fitting engine gas turbine blade cumulative thermal damage and The relationship between the temperature and the engine state in the first application scenario is calculated based on the relationship Cumulative thermal damage under temperature duration;

[0009] The engine is applied to the second application scenario, and the accumulated thermal damage is kept unchanged, and the duration and time proportion of each state of the engine in the engine load spectrum are obtained, that is, the engine load spectrum parameters in the second application scenario.

[0010] According to one aspect of the present invention, the cumulative thermal damage is The relationships between temperatures include:

[0011] ;

[0012] ;

[0013] Where:

[0014] It is the cumulative heat damage;

[0015] For Temperature-dependent damage per unit time function;

[0016] is the life damage correction function related to the rotor speed Ng and the atmospheric temperature T0;

[0017] for duration;

[0018] a is the slope;

[0019] b is a constant;

[0020] c is the intercept;

[0021] By calculating the engine states The stress of the gas turbine blades at different temperatures is measured by referring to the creep-life equation of the blade material to obtain the damage of the gas turbine blades under different engine conditions, thereby fitting different a, b, c coefficients at temperature.

[0022] According to one aspect of the present invention, the cumulative thermal damage is obtained by accumulating the thermal damage per unit time, and the thermal damage per unit time is obtained by linear interpolation of the damage values ​​at typical temperature points:

[0023] ;

[0024] Where:

[0025] T is a Temperature value;

[0026] For Thermal damage per unit time at temperature value;

[0027] For Bi A typical temperature point with a higher temperature value;

[0028] For Bi Typical temperature point with a lower temperature value;

[0029] For Bi Thermal damage per unit time at a typical temperature point with a higher temperature value;

[0030] For Bi Thermal damage per unit time at a typical temperature point one level lower in temperature value.

[0031] According to one aspect of the present invention, the time proportion of the duration of each engine state in the engine load spectrum is obtained as follows:

[0032] Calculate the engine status Duration at temperature And the duration The ratio of the time of each task profile in the second application scenario.

[0033] According to one aspect of the present invention, the engine states include: a maximum state, an intermediate state and a maximum continuous state.

[0034] According to the solution of the present invention, the aviation turboshaft engine can first be placed in, for example, a plateau area (the first application scenario) for use, so that the aviation turboshaft engine can perform flight missions in a relatively harsh environment. In this way, after obtaining the cumulative thermal damage in the relatively harsh environment by calculation, the calculated cumulative thermal damage is used as the extreme value, thereby forming an effective historical data support, and then achieving correct and reliable equivalent conversion in, for example, a plain area (the second application scenario), ensuring that the design of the time proportion of the duration of each engine state in the engine load spectrum is reliable and accurate.

[0035] According to the solution of the present invention, the present invention can effectively solve the problems of rationality and accuracy of conversion when performing equivalent conversion of various states for newly developed engines, thereby avoiding usage limitations in actual engineering applications and meeting the current design requirements of newly developed dynamic load spectra. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 The flowchart schematically shows a method for designing a load spectrum of an aircraft turboshaft engine based on thermal damage number according to an embodiment of the present invention. DETAILED DESCRIPTION

[0037] The present invention will now be discussed with reference to exemplary embodiments. It should be understood that the embodiments discussed are only to enable those skilled in the art to better understand and thus implement the present invention, rather than implying any limitation on the scope of the present invention.

[0038] As used herein, the term “including” and variations thereof are to be interpreted as open-ended terms meaning “including, but not limited to.” The term “based on” is to be interpreted as “based, at least in part, on.” The terms “one embodiment” and “an embodiment” are to be interpreted as “at least one embodiment.”

[0039] Figure 1 A flow chart schematically shows a method for designing a load spectrum of an aviation turboshaft engine based on thermal damage number according to an embodiment of the present invention. Figure 1 As shown, in this embodiment, the load spectrum design method of an aviation turboshaft engine based on thermal damage number includes:

[0040] According to information of multiple typical mission profiles of the engine, the power demand in each typical mission profile is analyzed by using an engine performance model, and the engine state and duration of each mission profile under the first application scenario are obtained, that is, the engine load spectrum parameters under the first application scenario;

[0041] Fitting engine gas turbine blade cumulative thermal damage and The relationship between the temperature and the engine state in the first application scenario is calculated based on the relationship The cumulative heat damage under temperature duration (i.e., the number of heat damages);

[0042] The engine is applied to the second application scenario, and the accumulated thermal damage is kept unchanged, and the duration and time proportion of each state of the engine in the engine load spectrum are obtained, that is, the engine load spectrum parameters in the second application scenario.

[0043] In this embodiment, taking the analysis of a typical mission profile I power state (i.e., engine state) of a newly developed aviation turboshaft engine as an example, the engine typical mission profile I state is shown in Table 1:

[0044]

[0045] Further, according to one embodiment of the present invention, the cumulative thermal damage of the working blades of the gas turbine of the engine is fitted with The relationship between the temperatures is:

[0046] Through the stress and life analysis of the engine gas turbine blades, multiple groups of different corresponding engine states are obtained. Temperature and gas turbine blade thermal damage data, based on the thermal damage data fitting cumulative thermal damage and The relationship between temperature;

[0047] Further, according to one embodiment of the present invention, the cumulative thermal damage is The relationships between temperatures include:

[0048] ;

[0049] Where:

[0050] It is the cumulative heat damage;

[0051] For Temperature-dependent damage per unit time function;

[0052] is the life damage correction function related to the rotor speed Ng and the atmospheric temperature T0;

[0053] for duration;

[0054] In this embodiment, , These two correlation functions are the specific expressions of the known creep-stress relationship in the creep estimation of turbine blades of aviation turboshaft engines.

[0055] In this embodiment, the formula It can be seen that the cumulative thermal (creep) damage is related to Temperature, speed, etc., that is, related to the engine status. Temperature and operating speed will fluctuate slightly according to engine performance changes. However, for turboshaft engines, especially engines with high thermodynamic cycle parameters, small changes in speed have a lower impact on the damage of gas turbine blades than temperature. Therefore, the corresponding relationship between temperature and speed determined by the engine performance model can be used to eliminate the speed variable. The cumulative thermal (creep) damage of the turbine blades is Temperature has the following relationship:

[0056] ;

[0057] Where:

[0058] a, b and c are relationship coefficients;

[0059] a is the slope;

[0060] b is a constant;

[0061] c is the intercept;

[0062] Furthermore, by calculating the engine states The stress of the gas turbine blades at different temperatures is measured. The life and damage of the gas turbine blades under different engine conditions are obtained by referring to the creep-life equation of the blade material, so as to fit different a, b, c coefficients at temperature.

[0063] In this embodiment, in fact, the higher the engine state, the greater the thermal damage per unit time. For an aviation turboshaft engine with high thermodynamic cycle parameters, the temperature and stress of its gas turbine blades after work are high, which is the main limiting component affecting the engine life. By counting the thermal (creep) damage of the gas turbine blades, the relative relationship between the thermal damages of various engine states can be characterized. Therefore, the thermal (creep) damage count of the turbine blades is the above-mentioned cumulative thermal damage.

[0064] Further, according to an embodiment of the present invention, the above-mentioned cumulative thermal damage is obtained by accumulating the thermal damage per unit time, and the thermal damage per unit time is obtained by linear interpolation of the damage values ​​at typical temperature points:

[0065] ;

[0066] Where:

[0067] T is a Temperature value;

[0068] for Thermal damage per unit time at temperature value;

[0069] For Bi A typical temperature point with a higher temperature value;

[0070] For Bi Typical temperature point with a lower temperature value;

[0071] For Bi Thermal damage per unit time at a typical temperature point with a higher temperature value;

[0072] For Bi Thermal damage per unit time at a typical temperature point one level lower in temperature value.

[0073] Further, according to an embodiment of the present invention, the engine is applied to a second application scenario, the accumulated thermal damage is kept unchanged, and the time proportion of the duration of each state of the engine in the engine load spectrum is obtained, which includes:

[0074] Make the formula The accumulated thermal damage in remains unchanged;

[0075] Calculate the engine load spectrum for each state of the engine Duration at temperature And the duration The ratio of the time of each mission profile in the second application scenario. This setting can make the time proportion of each state of the engine in the load spectrum equivalently adjusted according to the use requirements while the cumulative thermal damage remains unchanged. The equivalent adjustment of each state of the engine design load spectrum is shown in Table 2.

[0076]

[0077] Further, according to an embodiment of the present invention, the engine states include: a maximum state, an intermediate state and a maximum continuous state.

[0078] According to the above scheme of the present invention, the aviation turboshaft engine can first be placed in, for example, a plateau area (the first application scenario) for use, so that the aviation turboshaft engine can perform flight missions in a relatively harsh environment. In this way, after obtaining the cumulative thermal damage in the relatively harsh environment by calculation, the calculated cumulative thermal damage is used as the extreme value, thereby forming an effective historical data support, and then realizing correct and reliable equivalent conversion in, for example, a plain area (the second application scenario), ensuring that the design of the time proportion of the duration of each engine state in the engine load spectrum is reliable and accurate.

[0079] According to the above scheme of the present invention, the present invention can effectively solve the rationality and accuracy problems of conversion when performing equivalent conversion of various states for newly developed engines, thereby avoiding usage limitations in actual engineering applications and meeting the current design requirements of newly developed dynamic load spectra.

[0080] Those skilled in the art will appreciate that the modules and algorithm steps described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, or a combination of computer software and electronic hardware. Whether these functions are performed in hardware or software depends on the specific application and design constraints of the technical solution. Professional and technical personnel can use different methods to implement the described functions for each specific application, but such implementation should not be considered to be beyond the scope of the present invention.

[0081] The above description is only a preferred embodiment of the present application and an explanation of the technical principles used. Those skilled in the art should understand that the scope of the invention involved in the present application is not limited to the technical solution formed by a specific combination of the above technical features, but should also cover other technical solutions formed by any combination of the above technical features or their equivalent features without departing from the inventive concept. For example, the above features are replaced with the technical features with similar functions disclosed in this application (but not limited to) by each other.

[0082] It should be understood that the size of the serial numbers of each step in the content of the invention and the implementation methods of the present invention does not absolutely mean the order of execution. The execution order of each process should be determined by its function and internal logic, and should not constitute any limitation on the implementation process of the implementation methods of the present invention.

Claims

1. A method for designing load spectrum of aviation turboshaft engine based on thermal damage number, characterized in that: include: According to information of multiple typical mission profiles of the engine, the power demand in each typical mission profile is analyzed by using an engine performance model, and the engine state and duration of each mission profile in the first application scenario are obtained; Fitting engine gas turbine blade cumulative thermal damage and The relationship between the temperature and the engine state in the first application scenario is calculated based on the relationship Cumulative thermal damage under temperature duration; Applying the engine to the second application scenario, keeping the accumulated thermal damage unchanged, and obtaining the duration and time proportion of each state of the engine in the engine load spectrum; The cumulative thermal damage The relationships between temperatures include: ； ； Where: It is the cumulative heat damage; For Temperature-dependent damage per unit time function; is the life damage correction function related to the rotor speed Ng and the atmospheric temperature T0; for duration; a is the slope; b is a constant; c is the intercept; By calculating the engine states The stress of the gas turbine blades at different temperatures is measured by referring to the creep-life equation of the blade material to obtain the damage of the gas turbine blades under different engine conditions, thereby fitting different a, b, c coefficients at temperature.

2. The method for designing the load spectrum of an aviation turboshaft engine based on thermal damage number according to claim 1, characterized in that: The cumulative thermal damage is obtained by accumulating the thermal damage per unit time, and the thermal damage per unit time is obtained by linear interpolation of the damage values ​​at typical temperature points: ； Where: T is a Temperature value; for Thermal damage per unit time at temperature value; For Bi A typical temperature point with a higher temperature value; For Bi Typical temperature point with a lower temperature value; For Bi Thermal damage per unit time at a typical temperature point with a higher temperature value; For Bi Thermal damage per unit time at a typical temperature point one level lower in temperature value.

3. The method for designing the load spectrum of an aviation turboshaft engine based on thermal damage number according to claim 1, characterized in that: The duration and time proportion of each engine state in the engine load spectrum are obtained as follows: Calculate the engine status Duration at temperature And the duration The ratio of the time of each task profile in the second application scenario.

4. The method for designing the load spectrum of an aviation turboshaft engine based on thermal damage number according to any one of claims 1 to 3, characterized in that: The engine states include: a maximum state, an intermediate state and a maximum continuous state.

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