Method for determining surge margin loss of compression part caused by performance degradation of engine

Abstract

The invention provides a method for determining surge margin loss of a compression part caused by performance degradation of an engine. The method comprises the steps that a first blade tip gap and a first labyrinth gap are obtained through measurement; determining an engine overall performance model; measuring a first performance parameter of the engine; obtaining a first modified engine overall performance model; the engine is in a wear state, a second performance parameter is measured, and a second corrected engine overall performance model is obtained; the engine is decomposed, and a second blade tip gap and a second labyrinth gap are measured again; obtaining a performance degradation relation equation; obtaining a third blade tip gap and a third labyrinth gap during total recession; obtaining a third modified engine overall performance model; and determining a surge boundary, obtaining a common working line of compression parts of the engine, and calculating the surge margin loss during total decline of the engine through a difference value between the common working line and the surge boundary. According to the method, the surge margin loss of the compression part caused by performance degradation can be determined according to the long-term test run result of the engine.

Description

Technical Field

[0001] This invention relates to the field of aero-engines, and more specifically to the field of methods for determining surge margin loss of compression components caused by engine performance degradation. Background Technology

[0002] A gas turbine engine is a thermodynamic machine that uses continuously flowing air as its working medium. Through the rotation of an impeller, it compresses and expands the working fluid, converting the chemical energy of the fuel into useful work. However, after prolonged operation, gas turbine engines experience problems such as component wear, surface roughness, corrosion, thermal fatigue, and thermal deformation, leading to performance degradation and a decline in engine performance.

[0003] The pattern of engine performance degradation varies depending on the engine's cycle parameters and model characteristics. A key challenge in engine aerodynamic stability design is obtaining a relatively reasonable surge margin loss in the compression components due to performance degradation in the early stages of a newly developed engine. This, in turn, provides reliable reference and input for various design stages, type approval, and airworthiness certification. Summary of the Invention

[0004] One object of the present invention is to provide a method for determining the loss of surge margin of compression components due to engine performance degradation.

[0005] The method to achieve the above objective includes the following steps: S1. During engine assembly, the first blade tip clearance and the first tooth clearance are measured and obtained; the overall engine performance model is determined; S2. The engine is run, and the first engine performance parameter is measured; S3. The overall engine performance model is corrected using the first performance parameter obtained in S2 to obtain a first corrected overall engine performance model; S4. A long-term engine test is conducted, putting the engine in a wear state, and a second performance parameter is measured. The overall engine performance model is corrected based on the second performance parameter to obtain a second corrected overall engine performance model; S5. The engine is disassembled, and the second blade tip clearance and the second tooth clearance are measured again; S6. Based on the second blade tip clearance, the second tooth clearance, and the first... S7. Using the difference between the blade tip clearance and the first tooth clearance, and the test run time, obtain the performance degradation relationship equation; S8. Use the second modified engine overall performance model and the performance degradation relationship equation to obtain the third blade tip clearance and the third tooth clearance during full degradation; Use the third blade tip clearance and the third tooth clearance to modify the engine overall performance model to obtain the third modified engine overall performance model; S9. Use the third blade tip clearance and the third tooth clearance obtained in S7 to determine the surge boundary, use the third blade tip clearance and the third tooth clearance obtained in S7 and the third modified engine overall performance model to obtain the common working line of the engine compression components, and calculate the surge margin loss during full degradation of the engine by the difference between the common working line and the surge boundary.

[0006] In one or more embodiments, in S1, the first tip clearance and the first tooth clearance are first determined according to different engine types.

[0007] In one or more embodiments, the engine type is a dual-shaft direct-drive high-bypass turbofan engine, and the first grate clearance includes the CDP grate clearance, the high-pressure turbine front grate clearance, the high-pressure turbine rear grate clearance, the low-pressure turbine front grate clearance, and the low-pressure turbine rear grate clearance; the first blade tip clearance includes the rotor blade tip clearance of the fan, the booster stage, the high-pressure compressor, the high-pressure turbine, and the low-pressure turbine.

[0008] In one or more embodiments, the engine type is a single-rotor turbojet engine, the first tip clearance includes the compressor and turbine blade clearance; the first grate clearance is the CDP grate clearance, the turbine front grate clearance, and the turbine rear grate clearance.

[0009] In one or more embodiments, in steps S2 and S4, performance parameters are measured on the same engine measurement section, which includes the fan inlet section, the high-pressure compressor inlet section, the high-pressure compressor outlet section, the fan outlet section, the high-pressure turbine outlet section, and the low-pressure turbine outlet section.

[0010] In one or more embodiments, the performance parameters include temperature, pressure, flow rate, engine thrust, fuel consumption, and rotor speed.

[0011] In one or more embodiments, the engine testing includes a 150-hour endurance test, an AMT acceleration test, and an IMI mission test.

[0012] In one or more embodiments, the overall engine performance model is a variable specific heat model based on the characteristics of engine components.

[0013] In one or more embodiments, in S7, the engine overall performance model is in a full decay state when the thrust reaches the full decay thrust limit, the engine overall performance model speed reaches the full decay speed limit, or the engine overall performance model exhaust temperature reaches the full decay exhaust temperature limit.

[0014] In one or more embodiments, the performance degradation equation is T = K × D + T0, where D is the difference between the second tip clearance and the first tip clearance or the difference between the second tooth clearance and the first tooth clearance, K is a proportionality coefficient, T0 is the first tip clearance or the first tooth clearance, and T is the third tip clearance or the third tooth clearance at full degradation.

[0015] The method described above for determining the surge margin loss of compression components caused by engine performance degradation can determine the surge margin loss of compression components caused by performance degradation based on the results of long-term engine testing. This provides a more detailed reference and input for the engine design stage and solves the problem of how to obtain the surge margin loss of compression components under full degradation in newly developed engines. Attached Figure Description

[0016] The above and other features, properties and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings and embodiments, wherein:

[0017] Figure 1 This is a flowchart of a method for determining the surge margin loss of compression components caused by engine performance degradation;

[0018] Figure 2 This is a flowchart of a specific embodiment of the method;

[0019] Figure 3 It is a working curve graph. Detailed Implementation

[0020] The present invention will be further described below with reference to specific embodiments and accompanying drawings. More details are set forth in the following description in order to provide a full understanding of the present invention. However, the present invention can obviously be implemented in many other ways different from those described herein. Those skilled in the art can make similar extensions and derivations based on actual application situations without departing from the spirit of the present invention. Therefore, the scope of protection of the present invention should not be limited by the content of this specific embodiment.

[0021] It should be noted that these and other accompanying drawings are merely examples and are not drawn to scale, and should not be construed as limiting the scope of protection of the present invention.

[0022] Surge is a common phenomenon in gas turbine engines. When surge occurs, the boost ratio of the engine's compression components fluctuates significantly, the flow rate decreases dramatically, or backflow occurs. Surge can lead to damage to engine components, a significant increase in exhaust temperature, or engine shutdown. In general engine design, specific operating conditions are defined as surge limits, serving as an important reference for designing engine performance.

[0023] Surge margin is a physical quantity that characterizes the relative distance between the operating point of the compressor component and the surge boundary. A smaller surge margin means that the operating point of the compressor component is closer to the surge boundary. A surge margin of 0 or a negative value means that the compressor component will surge and cannot operate normally. Surge margin can be used to measure the safe distance between the compressor's normal operating conditions and the occurrence of surge.

[0024] Surge margin loss refers to the reduction in the surge margin of a compressor under specific conditions. Mechanical wear leads to a decrease in surge margin. This reduction means that the safety margin between the compressor's normal operating conditions and the occurrence of surge is reduced, making the compressor more susceptible to surge and affecting the stability and safety of the engine.

[0025] The method described in this invention is used to calculate the surge margin loss of compression components caused by the performance degradation of a gas turbine engine.

[0026] Generally, engine performance degradation is mainly due to the increased tip clearance of rotating components (compressor, turbine), which leads to a decline in component performance. This is mainly reflected in the changes in geometric parameters after the engine has been used for a long time. The geometric parameters related to performance degradation are identified as tip clearance and tooth clearance.

[0027] The main steps of this invention include:

[0028] S1. During engine assembly, the first blade tip clearance and the first tooth clearance are measured and obtained; the overall engine performance model is determined.

[0029] S2. Run the engine and measure its first performance parameters;

[0030] S3. Use the first performance parameters obtained in S2 to correct the overall engine performance model, and obtain the first corrected overall engine performance model;

[0031] S4. Conduct long-term engine test to put the engine in a wear state and measure the second performance parameter. Based on the second performance parameter, modify the overall engine performance model to obtain the second modified overall engine performance model.

[0032] S5. Disassemble the engine and measure the second blade tip clearance and the second comb tooth clearance again;

[0033] S6. Based on the differences between the second blade tip clearance, the second tooth clearance and the first blade tip clearance, the first tooth clearance, the performance degradation relationship equation is obtained;

[0034] S7. Use the second modified engine overall performance model and performance degradation relationship equation to obtain the third tip clearance and third tooth clearance under full degradation; use the third tip clearance and third tooth clearance to modify the engine overall performance model and obtain the third modified engine overall performance model.

[0035] S8. The surge boundary of the compression component is determined by using the third blade tip clearance and the third grate tooth clearance obtained in S7. The common working line of the engine compression component is obtained by using the third blade tip clearance and the third grate tooth clearance obtained in S7 and the third modified engine overall performance model. The surge margin loss during full engine decay is calculated by the difference between the common working line and the surge boundary.

[0036] Specifically, in S1, the first tip clearance and the first grate clearance are first determined according to different engine types. For example, for a twin-shaft direct-drive high-bypass turbofan engine, the first grate clearance includes the CDP grate clearance, the high-pressure turbine front grate clearance, the high-pressure turbine rear grate clearance, the low-pressure turbine front grate clearance, and the low-pressure turbine rear grate clearance; the first tip clearance includes the rotor tip clearances of the fan, booster stage, high-pressure compressor, high-pressure turbine, and low-pressure turbine. As another example, for a single-rotor turbojet engine, the first tip clearance includes the compressor and turbine blade inter-blade tip clearances; the first grate clearance includes the CDP grate clearance, the turbine front grate clearance, and the turbine rear grate clearance.

[0037] Subsequently, during engine assembly, the first blade tip clearance and the first tooth clearance were measured. It can be understood that the first blade tip clearance and the first tooth clearance are the initial clearance values ​​of the engine.

[0038] In step S1, it is also necessary to determine the overall engine performance model. The overall engine performance model is used to analyze and predict the operating performance of complex power units such as aero engines and gas turbines under different operating conditions. In a specific embodiment, the overall engine performance model is a variable specific heat model based on the characteristics of engine components.

[0039] After the engine is assembled, step S2 is performed to conduct engine performance testing, measuring and obtaining the engine's first characteristic parameters under 3 to 5 engine conditions. In some embodiments, the engine measurement sections include the fan inlet section, the high-pressure compressor inlet section, the high-pressure compressor outlet section, the fan outlet section, the high-pressure turbine outlet section, and the low-pressure turbine outlet section. The first performance parameters include the first temperature, the first pressure, the first flow rate, the first engine thrust, the first fuel consumption, and the first engine speed.

[0040] To ensure sufficient engine stability, the engine should be held in each state for a sufficiently long time; a minimum dwell time of 15 minutes is recommended. The engine should be measured across as many cross-sections as possible.

[0041] After obtaining the first characteristic parameters, the overall engine performance model is modified so that the calculation results can reproduce the test results, thus obtaining the first modified overall engine performance model.

[0042] Subsequently, long-term engine testing is conducted, such as 150-hour endurance tests, AMT acceleration tests, and IMI mission tests. The main purpose of long-term testing is to allow the engine to degrade rapidly. The test cycle should be arranged as closely as possible to the actual operating cycle of the engine to reproduce the engine's real-world usage scenarios to the greatest extent possible.

[0043] After a long-term mission test, the engine is in a worn state. Performance parameters at the same cross-section of the engine under this worn state are measured again to obtain second performance parameters, such as second temperature, second pressure, second flow rate, second engine thrust, second fuel consumption, and second speed at the same cross-section.

[0044] Based on the second performance parameters obtained above, the overall engine performance model is modified to obtain a second modified overall engine performance model, so that the calculation results of the second modified overall engine performance model can reproduce the test results in step S4.

[0045] In step S5, the engine is disassembled and disassembled to measure the second blade tip clearance and the second tooth clearance. Due to the long-term test run, the engine has experienced wear, therefore, there is a difference between the second blade tip clearance and the second tooth clearance and the first blade tip clearance and the first tooth clearance.

[0046] Based on this difference, the differences in each geometric parameter are adjusted by the same proportional coefficient to obtain the performance degradation relationship equation. Overall engine performance calculations are then performed until the engine reaches the thrust, speed, or exhaust temperature limits determined during full decay at the time of design, obtaining the relevant geometric parameter values ​​for full decay.

[0047] As an example of this explanation, assuming a twin-shaft direct-drive high-bypass turbofan engine, the first tip clearances of the fan, booster stage, high-pressure compressor, high-pressure turbine, and low-pressure turbine determined in step S1 are T, respectively. fan,1 T bst,1 T hpc,1 T hpt,1 T lpt,1 The determined clearances of the first toothed teeth at the CDP, before the high-pressure turbine, after the high-pressure turbine, before the low-pressure turbine, and after the low-pressure turbine are L, respectively. cdp,1 L hptf,1 L hptr,1 L lptf,1 L lptr,1 .

[0048] The second tip clearances of the fan, booster stage, high-pressure compressor, high-pressure turbine, and low-pressure turbine, as determined in step S5, are T, respectively. fan,2 T bst,2 T hpc,2 T hpt,2 T lpt,2 The determined clearances of the second toothed sections at CDP, before the high-pressure turbine, after the high-pressure turbine, before the low-pressure turbine, and after the low-pressure turbine are L, respectively. cdp,2 L hptf,2 L hptr,2 L lptf,2 L lptr,2 The difference in blade tip clearance between each component is:

[0049] DT fan =T fan,2 -T fan,1

[0050] DT bst =T bst,2 -T bst,1

[0051] DT hpc =T hpc,2 -T hpc,1

[0052] DT hpt =T hpt,2 -T hpt,1

[0053] DT lpt =T lpt,2 -T lpt,1

[0054] The difference in tooth clearance between each component is:

[0055] DL cdp =L cdp,2 -L cdp,1

[0056] DL hptf =L hptf,2 -L hptf,1

[0057] DL hptr =L hptr,2 -L hptr,1

[0058] DL lptf =L lptf,2 -L lptf,1

[0059] DL lptr =L lptr,2 -L lptr,1

[0060] Assuming a given tip clearance ratio k, the tip clearance expressions for the fan, booster stage, high-pressure compressor, high-pressure turbine, and low-pressure turbine in calculating the overall engine performance are as follows:

[0061] T fan =k×DT fan +T fan,1

[0062] T bst =k×DT bst +T bst,1

[0063] Thpc =k×DT hpc +T hpc,1

[0064] T hpt =k×DT hpt +T hpt,1

[0065] T lpt =k×DT lpt +T lpt,1

[0066] Assuming a given tooth clearance ratio C, the expressions for the tooth clearance at the engine CDP, before the high-pressure turbine, after the high-pressure turbine, before the low-pressure turbine, and after the low-pressure turbine are as follows:

[0067] L cdp =C×DL cdp +L cdp,1

[0068] L hptf =C×DL hptf +L hptf,1

[0069] L hptr =C×DL hptr +L hptr,1

[0070] L lptf =C×DL lptf +L lptf,1

[0071] L lptr =C×DL lptr +L lptr,1

[0072] In other words, the performance degradation relationship equation can be expressed as T=K×D+T0, where D is the difference between the second tip clearance and the first tip clearance, or the difference between the second tooth clearance and the first tooth clearance, K is a proportionality coefficient, such as k or C, T0 is the first tip clearance or the first tooth clearance, and T is the third tip clearance or the third tooth clearance during full degradation.

[0073] Based on the above performance degradation relationship equation, execute step S7, try to give different K values ​​in the degradation relationship equation, and modify the overall engine performance model. When the speed of the overall engine performance model reaches the full degradation speed limit, or the thrust of the overall engine performance model reaches the full degradation thrust limit, or the exhaust temperature of the overall engine performance model reaches the full degradation exhaust temperature limit, it is defined as the full degradation state, and the third modified overall engine performance model, as well as the third tooth clearance value and the third blade tip clearance value under the full degradation state are obtained.

[0074] Since the test run time is limited in step S4, the performance degradation equation should be the degradation relationship under the test run time condition. The degradation relationship under the limited test run time can be used to know the degradation relationship under the full degradation state.

[0075] Based on the obtained values ​​of the third tooth clearance and the third blade tip clearance when the engine is in full decay state, the surge boundary of each compression component can be determined.

[0076] The common operating line of each compression component is obtained by calculating using the third modified engine performance model, such as... Figure 3 As shown. Figure 3 The horizontal axis represents the converted airflow rate, and the vertical axis represents the boost ratio. The difference between the common operating line of each compression component and the surge boundary under full decay conditions is calculated, and this difference is the surge margin loss.

[0077] T' represents the baseline surge boundary line, which is the surge boundary line of the new engine under normal operating conditions and non-full fade. △T represents the common operating line margin loss.

[0078] The common operating line of any compression component during full engine decay is obtained by calculating using the third modified engine performance model, such as... Figure 3 As shown by the S-line, S' represents the common operating line of the new engine under normal operating conditions and non-full decay. △S represents the common operating line margin loss.

[0079] The sum of △S and △T is the surge margin loss N when the engine is in full decline, that is, N = △S + △T.

[0080] The above method can determine the surge margin loss of the compression component caused by performance degradation based on the results of long-term engine testing, providing a more detailed reference and input for the engine design stage, effectively improving design quality, and solving the problem of how to obtain the surge margin loss of the compression component under full degradation in newly developed engines.

[0081] A specific embodiment of this method is provided below.

[0082] (1) Determine the geometric parameters related to performance degradation based on the characteristics of the engine model. Generally, engine performance degradation is mainly caused by the increase in the tip clearance of rotating components (compressor, turbine), leading to a decrease in component performance, and the increase in the grating clearance (compressor, turbine), leading to an increase in cooling air leakage. Therefore, these two types of data should be selected as geometric parameters related to performance degradation. The specificity of these two types of data is also related to the engine type. For example, for a twin-shaft direct-drive high-bypass turbofan engine, the tip clearance includes the tip clearance of five components: fan, booster stage, high-pressure compressor, high-pressure turbine, and low-pressure turbine, while the grating clearance includes the CDP grating clearance, the high-pressure turbine front grating clearance, the high-pressure turbine rear grating clearance, the low-pressure turbine front grating clearance, and the low-pressure turbine rear grating clearance. For another example, for a single-rotor turbojet engine, the tip clearance includes the compressor and turbine tip clearances, while the grating clearance includes the turbine front grating clearance and the turbine rear grating clearance.

[0083] (2) During engine assembly, the relevant geometric parameter values ​​determined in step (1) are measured and obtained;

[0084] (3) Conduct engine performance testing, measuring engine characteristic section temperature, pressure, and flow parameters, as well as engine thrust, fuel consumption, and other performance parameters, at 3-5 engine stable states. To ensure sufficient engine stability, the dwell time in each state should be long enough, with a recommended dwell time of not less than 15 minutes. The engine measurement sections should be as numerous as possible. Recommended measurement sections and parameters for twin-shaft direct-drive high-bypass turbofan engines include: engine inlet flow rate, temperature, and pressure; high-pressure compressor inlet temperature and pressure; fan outlet temperature and pressure; high-pressure turbine outlet temperature and pressure; and low-pressure turbine outlet temperature and pressure.

[0085] (4) Based on the test and cross-sectional parameter measurement results in step (3), the engine reference overall performance model is corrected. This correction process mainly corrects the relevant geometric parameters in step (1) so that the calculation results can reproduce the test results, thereby confirming the relevant geometric parameter values ​​measured in step (1).

[0086] (5) Conduct long-term engine tests, such as 150-hour endurance tests, AMT acceleration tests, and IMI mission tests. The main purpose of these long-term tests is to allow the engine to degrade rapidly. The test cycle should be arranged as closely as possible to the actual operating cycle of the engine to reproduce the engine's real-world operating scenarios to the greatest extent possible.

[0087] (6) After the long-term mission test is completed, the engine performance admission test is carried out. The engine characteristic section temperature, pressure, flow parameters, engine thrust, fuel consumption and other performance parameters are measured under the same engine conditions as in step (3).

[0088] (7) Disassemble the engine and measure the relevant parameter values ​​determined in step (1);

[0089] (8) Based on the test results of step (6), only the geometric parameters related to step (1) are corrected, and the engine reference overall performance model is corrected again so that the calculation results can reproduce the test results and confirm the relevant geometric parameter values ​​measured in step (7).

[0090] (9) According to the differences of each geometric parameter determined in steps (8) and (4), the differences of each geometric parameter are adjusted in the same proportion for the blade tip clearance and the tooth classification, and the overall performance calculation of the engine is carried out until the engine reaches the speed or exhaust temperature limit determined at the time of full decay during the design, and the relevant geometric parameter values ​​determined in step (1) at the time of full decay are obtained.

[0091] (10) Determine the surge boundary of each compression component based on the geometric parameters determined in step (9);

[0092] (11) Based on the geometric parameters determined in step (9), the common working line of each compression component is calculated by the overall performance calculation model;

[0093] (12) Calculate the surge margin loss of each compression component when the performance deteriorates based on steps (10) and (11).

[0094] It should be noted that the use of terms such as "first," "second," and "third" to define the components and performance models is merely for the purpose of distinguishing the different stages of the corresponding components and performance models. Unless otherwise stated, the above terms have no special meaning and therefore should not be construed as limiting the scope of protection of this application.

[0095] Furthermore, this application uses specific terms to describe embodiments of the application. For example, "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic related to at least one embodiment of the application. Therefore, it should be emphasized and noted that "an embodiment," "one embodiment," or "an alternative embodiment" mentioned twice or more in different locations in this specification do not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0096] While the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the invention. Any variations and modifications can be made by those skilled in the art without departing from the spirit and scope of the invention. Therefore, any modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the scope of the invention, fall within the protection scope defined by the claims of the present invention.

Claims

1. A method for determining the surge margin loss of compression components due to engine performance degradation, characterized in that, Includes the following steps: S1. During engine assembly, the first blade tip clearance and the first tooth clearance are measured and obtained; the overall engine performance model is determined. S2. Run the engine and measure its first performance parameters; S3. Use the first performance parameters obtained in S2 to correct the overall engine performance model, and obtain a first corrected overall engine performance model; S4. Conduct long-term engine test to put the engine in a wear state, measure the second performance parameter, and modify the overall engine performance model based on the second performance parameter to obtain the second modified overall engine performance model. S5. Disassemble the engine and measure the second blade tip clearance and the second comb tooth clearance again; S6. Based on the difference between the second blade tip gap, the second tooth gap and the first blade tip gap, the first tooth gap, the performance degradation relationship equation is obtained; S7. Using the second modified engine overall performance model and the performance degradation relationship equation, obtain the third tip clearance and the third tooth clearance during full decay; Using the third blade tip clearance and the third tooth clearance, the overall engine performance model is corrected to obtain the third corrected overall engine performance model; S8. The surge boundary of the compression component is determined using the third blade tip clearance and the third grate tooth clearance obtained in S7. The common working line of the engine compression component is obtained using the third blade tip clearance and the third grate tooth clearance obtained in S7 and the third modified engine overall performance model. The surge margin loss during full engine decay is calculated by the difference between the common working line and the surge boundary.

2. The method as described in claim 1, characterized in that, In S1, the first blade tip clearance and the first tooth clearance are determined according to different engine types.

3. The method as described in claim 2, characterized in that, The engine type is a dual-shaft direct-drive high-bypass turbofan engine. The first grate clearance includes the CDP grate clearance, the high-pressure turbine front grate clearance, the high-pressure turbine rear grate clearance, the low-pressure turbine front grate clearance, and the low-pressure turbine rear grate clearance. The first blade tip clearance includes the rotor blade tip clearance of the fan, booster stage, high-pressure compressor, high-pressure turbine, and low-pressure turbine.

4. The method as described in claim 2, characterized in that, The engine type is a single-rotor turbojet engine, and the first blade tip clearance includes the compressor blade tip clearance and the turbine blade tip clearance; the first grate clearance includes the CDP grate clearance, the turbine front grate clearance, and the turbine rear grate clearance.

5. The method as described in claim 1, characterized in that, In steps S2 and S4, performance parameters are measured on the same engine measurement section, which includes the fan inlet section, the high-pressure compressor inlet section, the high-pressure compressor outlet section, the fan outlet section, the high-pressure turbine outlet section, and the low-pressure turbine outlet section.

6. The method as described in claim 1, characterized in that, The performance parameters include temperature, pressure, flow rate, engine thrust, fuel consumption, and rotor speed.

7. The method as described in claim 1, characterized in that, The engine testing includes a 150-hour endurance test, an AMT acceleration test, and an IMI mission test.

8. The method as described in claim 1, characterized in that, The overall engine performance model is a variable specific heat model based on the characteristics of engine components.

9. The method as described in claim 1, characterized in that, In S7, the engine is in a full decay state when the thrust of the overall engine performance model reaches the full decay thrust limit, the engine speed of the overall engine performance model reaches the full decay speed limit, or the exhaust temperature of the overall engine performance model reaches the full decay exhaust temperature limit.

10. The method as described in claim 1, characterized in that, The performance degradation relationship equation is T=K×D+T0, where D is the difference between the second blade tip clearance and the first blade tip clearance or the difference between the second tooth clearance and the first tooth clearance, K is a proportionality coefficient, T0 is the first blade tip clearance or the first tooth clearance, and T is the third blade tip clearance or the third tooth clearance during full degradation.

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