Over-temperature protection control method for takeoff of aero-engine

Abstract

The invention provides an aero-engine take-off overtemperature protection control method, and belongs to the field of aero-engine control, and the method comprises the steps: sequentially judging whether an engine is in a take-off state, whether the air inlet temperature of the engine is higher than the critical air inlet temperature, and whether the take-off height is lower than the critical height; when the engine is in the take-off state, the air inlet temperature of the engine is higher than the critical air inlet temperature, and the take-off height is lower than the critical height, the engine enters an overtemperature protection state, and the low-pressure rotating speed control values of the engine in the middle and stress application states are corrected; the duration of the take-off state of the engine in the take-off process is obtained, and when the duration of the take-off state reaches a duration threshold value or the engine exits from the take-off state, the low-pressure rotating speed control values of the engine in the middle state and the stress application state are restored to normal control. The problem of overtemperature in the take-off process of the engine is solved, the risk of overtemperature in the take-off process of the engine is reduced, and the reliability of the engine is improved.

Description

Technical Field

[0001] This application belongs to the field of aero-engine control, and particularly relates to a control method for over-temperature protection during takeoff of an aero-engine. Background Technique

[0002] With the continuous development of aero-engine technology, the temperature in front of the turbine is getting higher and higher. For a certain generation of engines, the temperature load they can withstand is limited, and the higher the temperature level, the more sensitive the impact on the life of engine components. When an aero-engine is installed and used, the outlet temperature T6 of the low-pressure turbine is generally used as the main fuel control parameter. At present, the temperature sensor of the turbine outlet temperature T6 can accurately measure the steady-state exhaust temperature of the engine. However, when the engine is in a transient process, especially during takeoff in hot weather, the warm-up degree of the engine is very insufficient, the component efficiency is low. Due to the lag in the dynamic response of the temperature sensor of the turbine outlet temperature T6, there is a "bulge" phenomenon in parameters such as the engine state speed and thrust, which first rise and then decrease, indicating that the measurement result of the turbine outlet temperature T6 is very different from the actual temperature of the engine. If it is directly used for controlling the transient temperature of the engine, it may cause the engine to overheat, increase the turbine thermal load and vibration stress, and even damage the structural integrity of the engine, directly threatening flight safety.

[0003] At present, some engines directly use the measured value of the temperature sensor of the turbine outlet temperature T6 for transient temperature control. Although it ensures sufficient takeoff thrust in hot weather, due to the large difference between the measurement result and the actual temperature of the engine, the engine will overheat, and even exceed the maximum allowable value of the entire envelope during takeoff, which has an extremely serious impact on the life of the turbine blades (analysis shows that the life of the turbine blades will be reduced by about 10 times), greatly increasing the engine life damage and use and maintenance costs; if the lead correction method of the temperature sensor of the turbine outlet temperature T6 is adopted, since the time constant of the temperature sensor of the turbine outlet temperature T6 is related to the material and structure of the sensor itself, and also related to the working environment parameters such as temperature, pressure, and flow rate at the measurement position of the sensor, a large number of sensor wind tunnel tests and engine debugging work need to be carried out, with high technical difficulty and high research, development and use costs. Summary of the Invention

[0004] The purpose of this application is to provide a control method for over-temperature protection during takeoff of an aero-engine to solve or mitigate at least one problem in the background technique.

[0005] The technical solution of this application is: A control method for over-temperature protection during takeoff of an aero-engine, including:

[0006] Sequentially determine whether the engine is in a takeoff state, whether the engine intake temperature is higher than the critical intake temperature, and whether the takeoff altitude is lower than the critical altitude;

[0007] When the engine is in the takeoff state, the engine intake air temperature is higher than the critical intake air temperature, and the takeoff altitude is lower than the critical altitude, it enters the over-temperature protection state, and corrects the low-pressure speed control value n in the intermediate and afterburner states of the engine 1max to make the low-pressure speed control values in the intermediate and afterburner states of the engine satisfy n 1max = n 1max - n 1-XZ , where n 1-XZ is the low-pressure speed correction value in the intermediate and afterburner states of the engine;

[0008] Obtain the duration of the engine takeoff state during the takeoff process. When the takeoff state duration reaches the duration threshold or the engine exits the takeoff state, make the low-pressure speed control value n in the intermediate and afterburner states of the engine 1max return to normal control.

[0009] Preferably, according to the characteristic parameters of the aircraft and the engine during takeoff, it is judged whether the engine is in the takeoff state. The characteristic parameters of the aircraft and the engine include the engine throttle lever angle PLA, the basic state of the engine, the state of the aircraft landing gear, and the ground speed V of the aircraft D .

[0010] Preferably, the critical temperature is determined according to the engine exhaust gas temperature margin and the warm-up degree of the engine before takeoff.

[0011] Preferably, the critical altitude is two or more to adapt to plain airports and plateau airports at different altitudes.

[0012] Preferably, the low-pressure speed correction value n in the intermediate and afterburner states of the engine 1-XZ is obtained through a whole-aircraft test.

[0013] Preferably, the low-pressure speed correction value in the intermediate and afterburner states of the engine is 1% - 2% of the low-pressure speed control value n in the intermediate and afterburner states of the engine according to the engine intake air temperature. 1max

[0014] Preferably, the conversion time to return to normal control is linearly transitioned.

[0015] Preferably, the duration threshold and the conversion time are set to 25s and 3s respectively.

[0016] The aviation engine takeoff over-temperature protection control method provided by this application solves the over-temperature problem during the engine takeoff process by setting the over-temperature protection control logic for the takeoff state, reduces the risk of over-temperature during the engine takeoff process, improves the reliability of the engine, and ensures flight safety. Description of the Drawings

[0017] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of this application.

[0018] Figure 1 Schematic diagram of the over-temperature protection control method for an aero-engine takeoff in this application. Specific embodiments

[0019] To make the purpose, technical solutions, and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below in conjunction with the accompanying drawings in the embodiments of this application.

[0020] Due to the lag in the dynamic response of the temperature sensor at the turbine outlet T6, when the engine goes from idle to maximum takeoff state, there is a "bulge" characteristic where parameters such as thrust and speed first increase and then decrease. To avoid the over-temperature impact of the "bulge" characteristic on the engine, this application proposes an over-temperature protection control method for an aero-engine takeoff based on the parameter characteristics during the engine takeoff process and considering specific takeoff scenarios.

[0021] As Figure 1 shown, the over-temperature protection control method for an aero-engine takeoff provided in this application includes the following processes:

[0022] S10, determine whether the engine is in the takeoff state.

[0023] Based on the characteristic parameters of the aircraft and the engine during takeoff, determine whether the engine is in the takeoff state.

[0024] Generally, the characteristic parameters of the aircraft and the engine during takeoff include the engine power lever angle PLA, the basic state of the engine, the state of the aircraft landing gear, the ground speed V of the aircraft D etc., and thus the characteristic parameters such as the engine power lever angle PLA, the basic state of the engine, the state of the aircraft landing gear, and the ground speed V of the aircraft can be used for judgment. For example, when the power lever angle PLA of a certain engine ≥ 65°, the engine is in the intermediate or above state, the landing gear state signal is in the lowered state, and the ground speed V of the aircraft D ≥ 5.5 m / s, it can be determined that the engine is in the takeoff state.

[0025] S20, determine whether the engine intake temperature T2 is higher than the critical intake temperature.

[0026] Obtain the engine inlet air temperature T2 at takeoff, and determine whether the engine inlet air temperature T2 exceeds the critical inlet air temperature. When the engine inlet air temperature T2 at takeoff is lower than the critical inlet air temperature, the engine is generally controlled by speed and there is no over-temperature problem; when the engine inlet air temperature T2 at takeoff is higher than the critical inlet air temperature, the engine is generally controlled according to the exhaust gas temperature. At this time, by reducing the low-pressure relative physical speed n1 of the engine, over-temperature protection is achieved.

[0027] In this application, the selection of the critical inlet air temperature is related to factors such as the engine exhaust gas temperature margin and the warm-up degree of the engine before takeoff. For example, in some embodiments, the critical inlet air temperature of a certain engine is set to 283K (i.e., about 10°C).

[0028] S30, determine whether the takeoff altitude H is lower than the critical altitude.

[0029] In this application, in order to adapt to plain airports and plateau airports at different altitudes, the critical altitude can be two or more, and the same or different control laws can be set for the plain airports and plateau airports corresponding to different heights. For example, in an embodiment of this application, the critical altitude can be set to 0km, 1km, 3km, 4.5km.

[0030] S40, when it is satisfied that the engine is in the takeoff state, the inlet air temperature is higher than the critical temperature, and the takeoff altitude is lower than the critical altitude, enter the over-temperature protection state. At this time, make the low-pressure speed control value n of the engine in the intermediate and afterburner states 1max be corrected. The correction process is to reduce the low-pressure speed control value n of the engine in the intermediate and afterburner states 1max , that is, n 1max =n 1max -n 1-XZ , where n 1-XZ is the low-pressure speed correction value of the engine in the intermediate and afterburner states.

[0031] In this application, the low-pressure speed correction value n of the engine in the intermediate and afterburner states 1-XZ can be obtained through a whole-aircraft test. The test process is as follows: adjust the aircraft and engine parameters, simulate the takeoff process. When pushing from the idle speed to the maximum state, if the parameters such as the engine speed and thrust no longer show the "bulge" phenomenon of rising first and then falling or the "bulge" amount is small, the low-pressure speed correction value n of the engine in the intermediate and afterburner states can be obtained 1-XZ .

[0032] In the preferred embodiment of this application, the low-pressure speed correction value of the engine in the intermediate and afterburner states is usually taken as the low-pressure speed control value n of the engine in the intermediate and afterburner states 1max1% - 2% of it is adjusted according to the engine intake air temperature T2. Considering the special scenario of using at high-altitude airports, it can be not corrected. As shown in Table 1, the low-pressure speed correction value n of the engine in the intermediate and afterburning states obtained in an embodiment of the present application 1-XZ , when it is lower than the critical intake air temperature (283K), the low-pressure speed control value n of the engine in the intermediate and afterburning states 1max is not corrected. When it is 10K higher than the critical intake air temperature (i.e., 293K), the low-pressure speed correction value n of the engine in the intermediate and afterburning states 1-XZ takes 1% of the low-pressure speed control value of the engine in the intermediate and afterburning states. When it is 20K or higher than the critical intake air temperature (i.e., 303K or higher), the low-pressure speed correction value n of the engine in the intermediate and afterburning states 1-XZ takes 2% of the low-pressure speed control value of the engine in the intermediate and afterburning states.

[0033] Table 1 Low-pressure speed correction value n of the engine in the intermediate and afterburning states 1-XZ

[0034]

[0035] S50, resume normal control:

[0036] Obtain the duration X1 of the engine takeoff state during the actual takeoff process. When the duration X1 of the takeoff state reaches the duration threshold or the engine exits the takeoff state, the low-pressure speed control value n of the engine in the intermediate and afterburning states 1max resumes to normal control, and this normal control is the control without modification.

[0037] In the preferred embodiment of the present application, to ensure the smooth change of engine parameters during the recovery process, the conversion time X2 to resume to normal control transitions linearly.

[0038] In some embodiments of the present application, the duration threshold of the engine takeoff state and the conversion time can be set to 25s and 3s respectively.

[0039] The aviation engine takeoff over-temperature protection control method provided by the present application solves the over-temperature problem during the engine takeoff process by setting the over-temperature protection control logic for the takeoff state, reduces the risk of over-temperature occurrence during the engine takeoff process, improves the reliability of the engine, ensures flight safety, and at the same time provides a new method for solving similar problems in engine development, thereby ensuring the smooth progress of model development, reducing the cost of model development. In addition, this method is simple to implement, easy to improve, and has a wide adaptability.

[0040] As described above, it is only the specific implementation manner of this application, but the protection scope of this application is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed by this application should be covered within the protection scope of this application. Therefore, the protection scope of this application shall be subject to the protection scope of the claimed rights.

Claims

1. A method for controlling over-temperature protection of an aircraft engine during takeoff, characterized in that: include: Determine in sequence whether the engine is in take-off state, whether the engine intake temperature is higher than the critical intake temperature, and whether the take-off altitude is lower than the critical altitude; When the engine is in takeoff state, the engine intake temperature is higher than the critical intake temperature, and the takeoff altitude is lower than the critical altitude, the over-temperature protection state is entered, and the low-pressure speed control value n of the engine in the intermediate and afterburner states is set. 1max Correction is made so that the low pressure speed control value of the engine in the intermediate and afterburner states satisfies n 1max =n 1max -n 1-XZ , where n 1-XZ It is the low pressure speed correction value of the engine in the intermediate and afterburner states; Get the duration of the engine takeoff state during takeoff. When the takeoff state duration reaches the duration threshold or the engine exits the takeoff state, the low-pressure speed control value n of the engine intermediate and afterburner states is set. 1max Restore to normal control.

2. The aircraft engine takeoff over-temperature protection control method according to claim 1, characterized in that: According to the characteristic parameters of the aircraft and the engine at takeoff, it is judged whether the engine is in the takeoff state, wherein the characteristic parameters of the aircraft and the engine include the engine throttle lever angle PLA, the basic state of the engine, the state of the aircraft landing gear, the ground speed V D .

3. The aircraft engine takeoff over-temperature protection control method according to claim 1, characterized in that: The critical temperature is determined according to the engine exhaust temperature margin and the engine warm-up degree before takeoff.

4. The aircraft engine takeoff over-temperature protection control method according to claim 1, characterized in that: The number of critical heights is two or more to accommodate plain airports and plateau airports at different heights.

5. The aircraft engine takeoff over-temperature protection control method according to claim 1, characterized in that: The low pressure speed correction value n of the engine in the intermediate and afterburner states 1-XZ Obtained through whole machine testing.

6. The aircraft engine takeoff over-temperature protection control method according to claim 5, characterized in that: The low-pressure speed correction value of the engine in the middle and afterburner states is taken as the low-pressure speed control value n of the engine in the middle and afterburner states according to the engine intake temperature. 1max 1% to 2%.

7. The aircraft engine takeoff over-temperature protection control method according to claim 1, characterized in that: The transition time back to normal control is linear.

8. The aircraft engine takeoff over-temperature protection control method according to claim 7, characterized in that: The duration threshold and the conversion time are set to 25s and 3s respectively.

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