A model-based robustness analysis method for aero-engine nozzle control schedule

Abstract

The application discloses a model-based aero-engine nozzle control plan robustness analysis method, and designs a nozzle pressure ratio control plan, a falling pressure ratio control plan and a slip control plan based on aero-engine component-level models respectively; three kinds of nozzle control plans are respectively used for simulation at a design point and in a flight envelope, control effects of the three control plans are tested; different disturbances faced by the aero-engine nozzle control are simulated, full-envelope flight tests are respectively conducted on the three control plans under the different disturbances, and robustness analysis is conducted. The application solves the evaluation problem of robustness differences among different nozzle control plans under the influence of various uncertain factors, can be used for robustness evaluation of various nozzle control plans, and further obtains robustness of different nozzle control plans, thereby providing a basis for selection of the nozzle control plan, accelerating the control plan research and development process, and being favorable for shortening the research and development period and reducing the research and development cost.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to the field of aero-engine control, in particular to a model-based aero-engine nozzle control schedule robustness analysis method. BACKGROUND

[0002] The tail nozzle is one of the important components of an aero-engine, and its main function is to make the gas discharged by the gas generator continue to expand, convert the available work of the gas into kinetic energy, and be sprayed backward at high speed to generate reaction thrust for the engine. The nozzle is the throat area of the tail nozzle. For aero-engines with afterburner and hypersonic speed, the nozzle must be controlled. The former is because the afterburner connection will impact the low-pressure turbine and cause fan surge. The latter needs to control the nozzle regardless of whether there is an afterburner or not, because the thrust loss of incomplete expansion of the gas is very large at hypersonic speed, and the nozzle must be controlled to ensure full expansion of the gas. Therefore, reasonable control of the nozzle is crucial to the safety and performance of the engine, and the control of the nozzle is also one of the important components of the engine control system.

[0003] Currently, aero-engines mainly use nozzle closed-loop control to adapt to the open-loop control of afterburner fuel. The control variable is the nozzle area, and the selection of the control parameter can be diverse. Some use the aero-engine pressure ratio before the nozzle as a constant, so that the core engine working state remains unchanged when the engine is in afterburner state. Some use the high and low pressure turbine total pressure ratio as a constant to control the nozzle. The nozzle control schedule is designed according to the working state of the engine to design the target value of the nozzle control parameter. The engine can both consider the aerodynamic stability of the fan and exert the performance of the engine. However, the design of the nozzle control schedule is not for a certain individual engine, but for engines of the same model with average performance. Therefore, under the individual differences of the engine, the effect of the nozzle control schedule will be affected. In addition, the aero-engine provides a high-temperature, high-pressure, and strong alternating working environment for the sensors of the control system, which is easy to cause the sensor to drift and other faults that are not easy to detect. The nozzle control schedule will also be affected due to the inaccurate sensor information used, and even more dangerous situations such as surge may occur. Therefore, there is an urgent need in the art for a model-based aero-engine nozzle control schedule robustness analysis method. SUMMARY

[0004] In order to solve the above technical problems, the purpose of the present application is to provide a model-based aero-engine nozzle control schedule robustness analysis method. The following technical solutions are adopted.

[0005] A model-based aero-engine nozzle control schedule robustness analysis method, comprising the following steps:

[0006] Step S100: based on the fan component level model known fan component characteristics, obtain the fan common working line;

[0007] Step S200: based on the engine component level model, build the nozzle control system;

[0008] Step S300: control the engine from the throttle lever PLA=15 degree idle state to the throttle lever PLA=65 degree intermediate state at the design point (ground), obtain the nozzle pressure ratio control plan, the supercharging ratio control plan, and the differential speed control plan;

[0009] Step S400: traverse the throttle lever from the throttle lever PLA=15 degree idle state to the maximum PLA=115 degree at the design point, and test the three nozzle control plans respectively;

[0010] Step S500: according to the flight envelope, randomly sample the flight envelope points, and control the engine to fly to each envelope point by using the three nozzle control plans in turn, traverse the throttle lever, obtain the fan actual pressure ratio parameter and the fan common working line pressure ratio parameter, and analyze the control effect of the nozzle control plan in the full envelope range;

[0011] Step S600: simulate the disturbance faced by the engine nozzle control, and perform full envelope flight test on the three nozzle control plans respectively under different disturbances;

[0012] Step S700: simulate one disturbance condition of step S600 each time, perform full envelope simulation according to step S500, output the thrust, fan pressure ratio, and common working line pressure ratio of the intermediate state and the maximum state, and perform robustness analysis.

[0013] Optionally, in step S100, the fan component characteristics refer to the flow pressure ratio characteristic curve of the fan; the common working line reflects the matching relationship between the non-design point engine components, is obtained by solving the common working equation under the conditions of meeting the air flow balance, pressure balance, coaxial work balance, and speed balance, and the tail nozzle and turbine guide vane being in a critical state; each fan flow pressure ratio characteristic line is divided into 100 parts, each division point is defined as point[i] (i=0…100, indicating 101 points), and then the common working point is located at the 50th point, i.e., point

[50] .

[0014] Optionally, in step S200, the building of the nozzle control system adopts closed-loop control, specifically, first taking point

[50] as the control target, and then using a PI controller to obtain the nozzle control given value.

[0015] Optionally, in step S300, the throttle lever is increased by 1 step from the throttle lever PLA=15 degree idle state to the throttle lever PLA=65 degree intermediate state.

[0016] Optionally, wherein each throttle lever down engine operation step is not less than 2000 steps, to ensure the engine is in a steady state, point to reach the common work line point

[50] at the output nozzle control plan required parameters.

[0017] Optionally, wherein the fall pressure ratio control plan, supercharging ratio control plan is shown as follows:

[0018] π tAim = f(n 2cor_T2 )

[0019] π cAim = f(n 1cor_T2 )

[0020] n 1corAim = f(n 2cor )

[0021]

[0022]

[0023]

[0024]

[0025] π tAim is the target fall pressure ratio, π cAim is the target supercharging ratio, n 1corAim is the target fan speed, n 2cor is the high pressure compressor speed, n 1cor is the fan speed; n2 is the high pressure compressor physical speed, n 2std is the high pressure compressor reference physical speed, T2 is the engine intake total temperature, n1 is the fan physical speed, n 1std is the fan reference physical speed; P s31 is the compressor outlet static pressure, P t6 is the mixed chamber inlet total pressure, P t2 is the fan inlet total pressure, π t is the fall pressure ratio, π c is the supercharging ratio.

[0026] Optionally, wherein in step S600, the different disturbances include: sensor noise, sensor drift, engine guide vane drift, individual differences, performance degradation.

[0027] Optionally, wherein in step S600, the nozzle fall pressure ratio control plan, supercharging ratio control plan, slip control plan are respectively tested by full envelope flight under different disturbances, specifically comprising the following steps:

[0028] S601: Inject sensor noise, for drop pressure ratio control, sensor parameters involve n2, T2, P s31 t6 , take the speed signal n2, sensor parameters T2, P s31 t6 Inject noise, T2 noise is set to ±0.6% of the range, P s31 t6 are set to ±0.36% of the range; for boost pressure ratio control, sensor parameters involve n1, T2, P t2 t6 , take the speed signal n1, T2 noise is set to ±0.6% of the range, P t2 t6 are set to ±0.36% of the range, P t2 noise is set to ±0.36% of the range; for differential speed control, sensor parameters involve n1, n2, T2, take the speed signals n1, n2, set T2 noise to ±0.6% of the range;

[0029] For drop pressure ratio control, because n2 is a speed signal, there is no drift, so take the speed signal n2, sensor parameters T2, P s31 t6 Inject noise, T2 measurement usually uses a platinum resistance temperature sensor with an accuracy of ±0.5%, and the temperature at the location of T2 is not high, so there is no sensor high-temperature calibration error, in addition to the measurement process there is a slow drift noise, about 20% of the accuracy value per 1000h, in summary, T2 noise is set to ±0.6% of the range. P s31 t6 is a pressure sensor parameter, its accuracy is ±0.3% of the range, and there is a slow drift noise in the measurement process, with a maximum drift of 20% of the accuracy, in summary, P s31 t6 noise is set to ±0.36% of the range; for boost pressure ratio control, sensor parameters involve n1, T2, P t2 t6 , as with drop pressure ratio control, take the speed signal n1, the same sensor noise injection settings are the same as the drop pressure ratio control plan, pressure sensor P t2 noise is set to the same as above, ±0.36% of the range.

[0030] S602: For sensor drift, only for P s31 t6 , P t2 , set the drift drift from ±1% to ±6%;

[0031] ​​​​​​​​​​The drifts described here are undetectable small drift faults, and the drift amount will not be large, so the drift amount is uniformly set from ±1% to ±6%.

[0032] S603: For the guide vane drift, set the drift of the high-pressure guide vane angle, and the drift range is ±1 degree;

[0033] According to the principle of the engine, the drift of the high-pressure guide vane angle has a great influence on the nozzle control, so only the drift of the high-pressure guide vane angle is set. The guide vane drift is caused by thermal expansion and contraction, and the gap between the intermediate mechanisms, so the drift angle will not be large, so the drift angle is set to be within ±1 degree.

[0034] S604: For individual differences, the engine individual performance differences are represented by adjusting the engine model component efficiency or area difference parameters. The standard of the maximum difference is that the thrust fluctuation is not greater than 2.5% of the original model thrust in the ground intermediate state, and the thrust fluctuation in the air cruising state is also not greater than 2.5% of the original model thrust;

[0035] S605: For performance degradation, the performance degradation of the rotating part is represented by pulling the flow and efficiency of the rotating part in the model, and the maximum degradation boundary before overhaul is set to be that the thrust in the ground intermediate state is unchanged, and the turbine rear temperature rises by 30K.

[0036] Optionally, in step S700, the robustness analysis, the key performance indicators of the engine are the engine thrust and the fan pressure ratio, and the different disturbances in step S600 are Δ 噪声 , Δ 传感器 , Δ 导叶 , Δ 个体 , Δ 衰退 , respectively. The robustness quantitative evaluation index of the key performance indicators of the engine is calculated.

[0037] Optionally, the dimensionless relative fluctuation of the key performance indicators of the engine is calculated according to different disturbances, which is used as the robustness evaluation index of the key performance indicators of the engine. The larger the value is, the worse the robustness is. The dimensionless relative fluctuation formula is as follows:

[0038]

[0039]

[0040] In the formula, j represents the performance index simulation output under different disturbances, normal represents the normal output of the performance index without disturbance, F represents the thrust, π Lc represents the fan pressure ratio.

[0041] The present disclosure has the following beneficial technical effects: in the development process of the engine control system, the method can simulate several typical uncertain factors faced by the control system in the real working process, pre-analyze the robustness characteristics of various designed nozzle control plans, provide a basis for the selection of nozzle control plans, speed up the control plan development process, and is conducive to shortening the development cycle and reducing the development cost. BRIEF DESCRIPTION OF DRAWINGS

[0042] The accompanying drawings illustrate exemplary embodiments of the present disclosure and together with the general description of the disclosure given above, and the detailed description of the embodiments below, serve to explain the principles of the present disclosure. These drawings are included herewith and constitute a part of this specification.

[0043] Figure 1 is a flow chart of a model-based aero-engine nozzle control plan robustness analysis method in an embodiment of the present disclosure;

[0044] Figure 2 is a fan common operating line obtained when designing a nozzle control plan in an embodiment of the present disclosure;

[0045] Figure 3 is a schematic diagram of a nozzle control system when generating a nozzle control plan in an embodiment of the present disclosure;

[0046] Figure 4 are three nozzle control plans designed and generated in an embodiment of the present disclosure;

[0047] Figure 5 is a schematic diagram of three nozzle control plans in an embodiment of the present disclosure;

[0048] Figure 6 is a ground verification result of three nozzle control plans in an embodiment of the present disclosure;

[0049] Figure 7 is a 1000-point graph randomly sampled in a flight envelope in an embodiment of the present disclosure;

[0050] Figure 8 is a relative change of an engine fan pressure ratio under no-fault traversal of a full envelope in an embodiment of the present disclosure;

[0051] Figures 9 to 13 is a robustness analysis graph under various disturbances of a full envelope in an embodiment of the present disclosure. DETAILED DESCRIPTION

[0052] The following will be described in conjunction with the accompanying drawings Figures 1 to 13Embodiments and the like make further detailed description of the present disclosure. It can be understood that the specific embodiments described herein are only for the purpose of explanation and are not limitations of the present disclosure. In addition, it should be noted that only parts related to the present disclosure are shown in the drawings for ease of description.

[0053] It should be noted that the embodiments and features in the embodiments in the present disclosure can be combined with each other without conflict. The technical solutions of the present disclosure will be described in detail below with reference to the drawings and in combination with the embodiments.

[0054] Unless otherwise specified, the exemplary embodiments / examples shown will be understood as providing exemplary features of various details that can implement the technical concepts of the present disclosure in practice. Therefore, unless otherwise specified, the features of various embodiments / examples can be additionally combined, separated, interchanged and / or rearranged without departing from the technical concepts of the present disclosure.

[0055] In the drawings, cross-hatching and / or shading are generally used to make the boundaries of adjacent components clear. As such, unless otherwise specified, the presence of cross-hatching or shading is not meant to imply that a particular material, material property, dimension, ratio, commonality of components between illustrations, and / or any other characteristic, attribute, property, etc. of the components is preferred, required, or should be inferred. In addition, the dimensions and the relative dimensions of the components in the drawings are intended to be exemplary only and can not be to scale unless otherwise specified. When example embodiments can be implemented differently, a specific process sequence can be performed in a different order than described. For example, two consecutively described processes can be performed substantially simultaneously or in an order opposite to that described. In addition, the same reference numbers represent the same components.

[0056] When a component is referred to as being "on" or "over" another component, "connected to" or "coupled to" another component, it can be directly on, directly connected to, or directly coupled to the other component, or there can be intervening components. However, when a component is referred to as being "directly on", "directly connected to", or "directly coupled to" another component, there are no intervening components. For this reason, the term "connected" can refer to a physical connection, an electrical connection, etc., with or without intervening components.

[0057] For descriptive purposes, the present disclosure can use spatial or relative terms, such as "below," "lower," "bottom," "on," "above," "upper," "top," "over," and "side" (e.g., as in "sidewall"), to describe the relative relationship between one element and another element as illustrated in the figures. The spatial or relative terms are intended to encompass different orientations of the device in use, operation, and / or manufacture in addition to the orientation depicted in the figures. For example, if the device in the figures is turned over, elements described as "below" or "under" other elements or features would then be oriented "above" the other elements or features. Thus, the exemplary term "below" can encompass both an orientation of above and below. The device can be otherwise oriented (e.g., rotated 90 degrees or at other orientations) and the spatial or relative descriptors used herein interpreted accordingly.

[0058] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting. As used herein, the singular forms "a," "an," and "the" are intended to include the plural forms as well, unless the context clearly indicates otherwise. Furthermore, to the extent that the terms "including," "includes," "having," "has," "with," or "contains" or variants thereof are used in either the detailed description and / or the claims, such terms are intended to be inclusive in a manner similar to the term "comprising." It is also noted that the terms "substantial," "approximately," and other similar terms, as used herein, are used in their normal, ordinary sense and are not used in an absolute sense.

[0059] The present application provides a model-based robustness analysis method for an aero-engine nozzle control plan, comprising the following steps:

[0060] Step S100: Obtain a fan common operating line based on known fan component characteristics of a fan component level model;

[0061] Step S200: Build a nozzle control system based on an engine component level model;

[0062] Step S300: Control the engine from a throttle lever PLA=15 degree idle state to a throttle lever PLA=65 degree intermediate state at a design point (ground) to obtain a nozzle pressure ratio control plan, a pressure ratio control plan, and a speed difference control plan;

[0063] Step S400: Traverse the throttle lever PLA=15 degree idle state to the maximum PLA=115 degree at the design point to test the three nozzle control plans, respectively.

[0064] Step S500: According to the flight envelope, randomly sample the flight envelope points, control the engine to fly to each envelope point by using the three nozzle control plans in turn, traverse the throttle lever, obtain the fan actual pressure ratio parameter and the fan common working line pressure ratio parameter, and analyze the control effect of the nozzle control plan in the full envelope range;

[0065] Step S600: Simulate the disturbance faced by the engine nozzle control, and perform full envelope flight test on the three nozzle control plans under different disturbances respectively;

[0066] Step S700: Simulate a disturbance condition of step S600 each time, perform full envelope simulation according to step S500, output the thrust, fan pressure ratio, and common working line pressure ratio of the intermediate state and maximum state, and perform robustness analysis.

[0067] Optionally, in step S100, the fan component characteristic refers to the fan flow pressure ratio characteristic curve; the common working line reflects the matching relationship between the non-design point engine components, which is obtained by solving the common working equation under the conditions of meeting the air flow balance, pressure balance, coaxial work balance, and speed balance, and the tail nozzle and turbine guide vane being in a critical state; each fan flow pressure ratio characteristic line is divided into 100 parts, each divided point is defined as point[i] (i=0…100, indicating 101 points), and then the common working point is located at the 50th point, i.e., point

[50] .

[0068] Optionally, in step S200, the nozzle control system is built by using closed-loop control, specifically, point

[50] is used as the control target, and then a PI controller is used to obtain the nozzle control given value.

[0069] Optionally, in step S300, the throttle lever is traversed from the idle state of the throttle lever PLA=15 degrees to the intermediate state of the throttle lever PLA=65 degrees by increasing by 1 step each time.

[0070] Optionally, the engine runs for no less than 2000 steps under each throttle lever to ensure that the engine is in a steady state, point reaches the common working line point

[50] , and the required parameters of the nozzle control plan are output.

[0071] Optionally, the drop pressure ratio control plan and the boost pressure ratio control plan are respectively as follows:

[0072] π tAim =f(n 2cor_T2 )

[0073] π cAim =f(n 1cor_T2 )

[0074] n 1corAim = f(n 2cor )

[0075]

[0076]

[0077]

[0078]

[0079] π tAim is target pressure ratio, π cAim is target pressure ratio, n 1corAim is target fan speed, n 2cor is target high pressure compressor speed, n 1cor is fan speed; n2 is high pressure compressor physical speed, n 2std is high pressure compressor reference physical speed, T2 is engine inlet total temperature, n1 is fan physical speed, n 1std is fan reference physical speed; P s31 is compressor exit static pressure, P t6 is mixer inlet total pressure, P t2 is fan inlet total pressure, π t is pressure ratio, π c is pressure ratio.

[0080] Optionally, in step S600, the different disturbances include: sensor noise, sensor drift, engine guide vane drift, individual difference, performance degradation.

[0081] Optionally, in step S600, the nozzle pressure ratio control plan, the pressure ratio control plan, and the speed difference control plan are respectively subjected to full envelope flight test under different disturbances, and specifically comprising the following steps:

[0082] S601: inject sensor noise, for pressure ratio control, sensor parameters involve n2, T2, P s31 , P t6 , take the speed signal n2, inject noise to the sensor parameters T2, P s31 , P t6 , T2 noise is set to ±0.6% of the range, P s31 , P t6 are all set to ±0.36% of the range; for pressure ratio control, sensor parameters involve n1, T2, P t2 , P t6 , take the speed signal n1, T2 noise is set to ±0.6% of the range, P t2 , Pt6 All set to ±0.36% of the range, P t2 Noise set to ±0.36% of the range; for slip control, sensor parameters involve n1, n2, T2, take the speed signal n1, n2, T2 noise set to ±0.6% of the range;

[0083] In step S601, for the drop pressure ratio control, because n2 is the speed signal, there is no drift, so take the speed signal n2, sensor parameters T2, P s31 , P t6 Inject noise, T2 measurement usually uses platinum resistance temperature sensor, the accuracy is ±0.5%, and the temperature at the location of T2 is not high, so no sensor high temperature calibration error is set, in addition to the measurement process also exists slow drift noise, about 20% of the accuracy of the drift per 1000h, in summary, T2 noise set to ±0.6% of the range, P s31 , P t6 For pressure sensor parameters, the accuracy is ±0.3% of the range, and there is a slow drift noise in the measurement process, the maximum drift is 20% of the accuracy, in summary, P s31 , P t6 Noise is set to ±0.36% of the range; for the boost pressure ratio control, sensor parameters involve n1, T2, P t2 , P t6 , as the drop pressure ratio control, take the speed signal n1, the same sensor noise injection setting is the same as the drop pressure ratio control plan, pressure sensor P t2 Noise is set to ±0.36% of the range as above.

[0084] S602: for sensor drift, only for P s31 , P t6 , P t2 , set the drift drift from ±1% to ±6%;

[0085] In step S602, the drift is a small drift fault that cannot be detected, and the drift amount will not be large, so the drift amount is uniformly set from ±1% to ±6%.

[0086] S603: for guide vane drift, set the drift of the high pressure guide vane angle, the drift range is ±1 degree;

[0087] In step S603, according to the principle of the engine, the high pressure guide vane angle drift has a great influence on the nozzle control, so only the high pressure guide vane angle is set to drift, and the guide vane drift is due to thermal expansion and contraction, and the influence of intermediate mechanism gap and other factors, so the drift angle will not be large, so the drift angle is set to within ±1 degree.

[0088] S604: For individual differences, the engine individual performance differences are characterized by adjusting the engine model component efficiency or area difference parameters, and the standard of the maximum difference is that the thrust fluctuation is not greater than 2.5% of the original model thrust in the ground intermediate state, and the thrust fluctuation in the air cruising state is also not greater than 2.5% of the original model thrust;

[0089] S605: For performance degradation, the performance degradation of the rotating component is characterized by pulling the flow and efficiency of the rotating component in the model, and the maximum degradation boundary before overhaul is set as that the thrust in the ground intermediate state is unchanged, and the temperature after the turbine rises by 30K.

[0090] Optionally, in step S700, the robustness analysis, the key performance indicators of the engine are engine thrust and fan pressure ratio, and the different disturbances in step S600 are Δ 噪声 , Δ 传感器 , Δ 导叶 , Δ 个体 , Δ 衰退 , respectively, and the robustness quantitative evaluation indicators of the key performance indicators of the engine under the nozzle pressure ratio control plan, the supercharging ratio control plan and the slip control plan are calculated.

[0091] Optionally, the dimensionless relative fluctuation of the key performance indicators of the engine under different disturbances is calculated as the robustness evaluation indicator of the key performance indicators of the engine, and the greater the value is, the worse the robustness is, and the dimensionless relative fluctuation formula is as follows:

[0092]

[0093]

[0094] In the formula, j represents the performance indicator simulation output under different disturbances, normal represents the normal output of the performance indicator without disturbance, F represents thrust, π Lc represents fan pressure ratio.

[0095] In one embodiment, with reference to Figure 1 The present disclosure provides a model-based nozzle control plan robustness analysis method for an aero-engine, which comprises:

[0096] Step S100: obtaining a fan common working line based on fan component characteristics of a fan component level model;

[0097] Wherein, the fan component characteristic refers to the flow pressure ratio characteristic curve of the fan; the common working line reflects the matching relationship between the engine components at the non-design point, and is obtained by solving the common working equation under the conditions of meeting the flow balance, pressure balance, coaxial work balance and rotating speed balance, and the tail nozzle and turbine guide vane being in a critical state; each fan flow pressure ratio characteristic line is divided into 100 parts, each division point is defined as point[i] (i=0…100, indicating 101 points), and then the common working point is located at the 50th point, i.e. point

[50] .

[0098] As shown in Figure 2 , wherein the data is normalized, and the horizontal distribution line is the common working line, i.e. the position of the fan common working point on each rotating speed line is point

[50] .

[0099] Step S200: Building a nozzle control system based on an engine component level model; wherein the building of the nozzle control system adopts closed-loop control, specifically, taking point

[50] as the control target, and then using a PI controller to obtain the nozzle control given value.

[0100] Specifically, as shown in Figure 3 , the nozzle control system adopts closed-loop control, specifically, taking point

[50] as the control target (the position of the fan common working point), and then using a PI controller to obtain the nozzle control area and then transmitting it to the actuator circuit to control the engine operation until the fan is stabilized at point

[50] on the flow pressure ratio characteristic curve.

[0101] Step S300: Controlling the engine at the design point (ground) to step through from the idle state of the throttle lever PLA=15 degrees to the intermediate state of the throttle lever PLA=65 degrees, and controlling the nozzle to make the fan work at the common working line under each throttle lever. After the engine is stabilized, output the required parameters of the nozzle control plan, obtain the nozzle pressure ratio control plan, the supercharging ratio control plan and the differential speed control plan.

[0102] The relationship between the engine working state and the throttle lever angle is shown in Table 1:

[0103] Table 1 Relationship between engine working state and throttle lever angle

[0104]

[0105] Wherein, the design point is that the controller input height H = 0, Mach number Mahe = 0; the throttle lever is single step increased by 1 from the throttle lever PLA = 15 degrees slow car state to the throttle lever PLA = 65 degrees intermediate state, wherein the engine operating step under each throttle lever is not less than 2000 steps, ensuring that the engine is in a steady state, point reaches the common working line point

[50] , outputs the required parameters of the nozzle control plan, the fall pressure ratio control plan, the supercharging ratio control plan, and the differential speed control plan, which are respectively shown as the following formula (1) (2) (3), and the specific data is shown in Table 2 (data normalization processing):

[0106] π tAim =f(n 2cor_T2 ) (1)

[0107] π cAim =f(n 1cor_T2 ) (2)

[0108] n 1corAim =f(n 2cor ) (3)

[0109]

[0110]

[0111]

[0112]

[0113] π tAim is the target fall pressure ratio, π cAim is the target supercharging ratio, n 2cor_T2 is the high-pressure compressor conversion speed, n 1cor_T2 is the fan conversion speed; n2 is the high-pressure compressor physical speed, n 2std is the high-pressure compressor reference physical speed, T2 is the engine inlet total temperature, n1 is the fan physical speed, n 1std is the fan reference physical speed; P s31 is the compressor outlet static pressure, P t6 is the mixed chamber inlet total pressure, P t2 is the fan inlet total pressure, π t is the fall pressure ratio, π c is the supercharging ratio.

[0114] Table 2 control plan

[0115]

[0116]

[0117]

[0118] Figure 4 The figures show a visual representation of three control plans (data normalized). Figure (a) shows the pressure ratio control plan, with the horizontal axis representing the equivalent speed of the high-pressure compressor and the vertical axis representing the pressure ratio. The curve in the figure represents the change in the target pressure ratio as the equivalent speed of the high-pressure compressor changes. The horizontal part of the curve is used for nozzle control in idle state to ensure that the nozzle is at its maximum position in idle state. Figure (b) shows the pressure ratio control plan, with the horizontal axis representing the equivalent speed of the fan and the vertical axis representing the pressure ratio. The curve in the figure represents the change in the target pressure ratio as the equivalent speed of the fan changes. The horizontal part of the curve is also used for nozzle control in idle state, which can ensure that the nozzle is at its maximum position in idle state. Figure (c) shows the slip control plan, with the horizontal axis representing the equivalent speed of the high-pressure compressor and the vertical axis representing the equivalent speed of the fan. The curve in the figure represents the change in the target equivalent speed of the fan as the equivalent speed of the high-pressure compressor changes. The horizontal part of the curve is also used for nozzle control in idle state to ensure that the nozzle is at its maximum position in idle state.

[0119] Step S400: Test three nozzle control schemes by traversing the throttle lever from PLA = 15 degrees to the maximum PLA = 115 degrees at the design point, such as... Figure 5 As shown in Figure (a), the nozzle control system for the pressure ratio control plan uses sensors in the aero-engine control system to collect the actual pressure ratio π of the engine in this state. t The target pressure ratio π of the engine at this time is calculated according to the nozzle control plan. tAim The difference between the target pressure ratio and the actual engine pressure ratio is calculated and then fed into the PI controller to calculate the target nozzle area A. 82 , and the open-loop nozzle area A 81 To maximize the nozzle area, the nozzle control system needs to limit the nozzle area, ensuring it does not exceed the mechanical limit area of ​​the nozzle. Therefore, the final nozzle area is A. 8Aim The actuator circuit is fed into the nozzle, which controls the desired nozzle area by adjusting the hydraulic flow rate in the rodless and rod chambers of the nozzle actuator cylinder. The larger-than-expected logic in the open-loop and closed-loop control schemes prevents the nozzle area from becoming too small. The sensors used in the pressure ratio control scheme are T2, n2, and P. s31 P t6 Figure (b) shows the nozzle control system of the boost ratio control plan. Its architecture differs from Figure (a) in that the left side uses Tt2 and n1 for the "nozzle boost ratio control plan". Figure (c) shows the nozzle control system of the slip control plan. Its architecture differs from Figure (a) in that the left side uses Tt2 and n2 for the "nozzle slip control plan". The three control systems have the same structure, differing only in the state parameters and target parameters required by the control plan, as described in detail in step S300. Figure 6The results of testing the three control schemes at the design point show that the actual fan pressure ratio at the same fan conversion speed is the same as the pressure ratio on the common working line, indicating that all three control schemes enable the low-pressure compressor to operate stably on the common working line at the design point.

[0120] Step S500: Randomly sample 1000 points along the flight envelope, and sequentially use the pressure ratio control to control the boost ratio to fly to each envelope point, traversing the throttle lever. Then, under steady-state conditions, output the actual fan pressure ratio parameters and the fan common working line pressure ratio parameters, and perform a comparative analysis of their relative changes. The flight envelope is shown below. Figure 7 As shown, the statistical results of the relative change in the fan voltage ratio of the full-envelope line are as follows: Figure 8 The results show that under pressure ratio control, the relative variation range between the low-pressure compressor pressure ratio and the common operating line pressure ratio is -0.21% to 0.79%; under pressure ratio control, the relative variation range is -0.40% to 0.65%; and under slip control, the relative variation range is -1.55% to 1.24%. It can be seen that all three control schemes can ensure the engine operates well around the low-pressure compressor's common operating line across the entire envelope, but slip control deviates slightly more from the other two schemes.

[0121] Step S600: Simulate the five types of uncertainties: sensor noise, sensor drift, engine guide vane drift, individual differences, and performance degradation.

[0122] Specifically, step S600 includes the following steps:

[0123] S601: Sensor noise injection. For pressure ratio control, the sensor parameters involve n2, T2, and P. s31 ,P t6 Among them, the speed signal has no drift, sudden jumps are easily filtered out, and the sensor has redundant arrangement, so the speed sensor signal n2 is accepted. Only T2 and P are given. s31 ,P t6 Injected noise. T2 is typically measured using a platinum resistance temperature sensor with an accuracy of ±0.5%. Since the temperature at the location of T2 is not high, no high-temperature calibration error is set for the sensor. In addition, there is slowly varying drift noise during the measurement process, drifting approximately 20% of the accuracy value every 1000 hours. Therefore, the noise level for T2 is set to ±0.6% of the measurement range. s31 ,P t6 The parameters of the pressure sensor are as follows: its accuracy is ±0.3% of the measurement range; it exhibits slowly varying drift noise during measurement, with a maximum drift of 20% of the accuracy; therefore, P... s31 ,P t6 Noise levels are set to ±0.36% of the measurement range. For boost ratio control, the sensor parameters involve n1, T2, and P. t2 ,P t6, With the control of the drop pressure ratio, take the speed signal n1, the same sensor noise injection settings with the drop pressure ratio control plan, pressure sensor P t2 Noise settings are the same as the range of ± 0.36%.

[0124] S602: sensor drift fault, only for P s31 ,P t6 , P t2 , the drift described here is a small drift fault that cannot be detected, and the drift amount will not be large, so the drift amount is uniformly set from ± 1% to ± 6%.

[0125] S603: guide vane drift, according to the principle of the engine, the high pressure guide vane angle drift has a great influence on the nozzle control, so only the high pressure guide vane angle is set to drift, and the guide vane drift is due to thermal expansion and contraction, and the influence of factors such as intermediate mechanism gap, so the drift angle will not be large, so the drift angle is set to within ± 1 degree.

[0126] S604: individual difference, during the manufacturing process of the aero-engine, due to various reasons such as machine tools, environment, etc., there are differences in manufacturing precision and assembly precision of parts, which is inevitable, which makes the same model engine have differences in performance, therefore, the robustness of the nozzle drop pressure ratio control plan to the individual difference of the engine needs to be analyzed. The individual difference of the engine is reflected in the model as the difference in the efficiency or area of the model components, and the boundary of the difference is that in the ground intermediate state, the thrust fluctuation is not greater than 2.5% of the original model thrust 58800N, and in the air cruising state, the thrust fluctuation is also not greater than 2.5% of the original model thrust. According to this boundary constraint, adjust the performance coefficient of the engine, simulate the larger individual difference of the engine, and according to the boundary condition, the adjustment range of the performance coefficient obtained by debugging is shown in the following table.

[0127] Table 3 individual difference pull table

[0128]

[0129] S605: performance degradation, the aero-engine is a typical mechanical product, and because its working environment is relatively harsh, during use, there will be aging, wear and tear, etc., which will cause its performance to decline, and its main performance is the decline of the performance of rotating parts. The decline is characterized by pulling the flow, efficiency and other parameters of the rotating parts in the model, and according to the related literature, the maximum decline boundary before overhaul is set as: the ground intermediate state thrust is unchanged, and the turbine rear temperature rises by 30K. Through multiple tests, one set of selected set pull coefficient is shown in the following table:

[0130] Table 4 performance degradation pull table

[0131]

[0132] Step S700: According to the simulation of step S600, simulate each case, and perform step S500 to perform full envelope simulation, output the intermediate state and maximum state thrust, fan pressure ratio, and pressure ratio at the common operating line, and perform robustness analysis such as Figures 9 to 13 .

[0133] Among them, the robustness analysis engine performance indicators are engine thrust and fan pressure ratio, and the disturbances of various uncertain factors are respectively Δ = {Δ 噪声 , Δ 传感器 , Δ 导叶 , Δ 个体 , Δ 衰退}, the robustness quantitative evaluation indicators of the key performance indicators of the two nozzle control plans can be calculated respectively.

[0134] According to each disturbance, the dimensionless relative fluctuation of the performance indicator is calculated as its robustness evaluation indicator, and the greater the value, the worse the robustness, as shown in equations (8) and (9).

[0135]

[0136]

[0137] In the formula, j represents the performance indicator simulation output under different disturbances, normal represents the normal output of the performance indicator without disturbance, F represents thrust, and π Lc represents the fan pressure ratio.

[0138] Figure 9 The engine performance indicator changes after injecting noise under the three control plans. Figure (a) is the relative change of the intermediate state thrust, and figure (b) is the relative change of the maximum state thrust. It can be seen that under the two states, the relative change of the thrust of the differential speed control is the smallest, the pressure ratio control is second, and the pressure ratio control is the largest. But the relative change range of thrust under the three control plans in the full envelope range is within 1%, which is relatively small. It shows that the robustness of the thrust of the three control plans to sensor noise is good. Figure (c) is the relative change of the intermediate state pressure ratio at the common operating line. Figure (d) is the relative change of the maximum state pressure ratio at the common operating line. It can be seen that under the intermediate state differential speed control, the maximum relative change of the low-pressure compressor pressure ratio reaches 1.5%, while the maximum relative change of the low-pressure compressor pressure ratio under the pressure ratio control and the pressure ratio control is within 1%. In the maximum state, it is almost the same as the intermediate state. In summary, for the change of the low-pressure compressor pressure ratio, the robustness of the pressure ratio control and the pressure ratio control is better than that of the differential speed control, but the three control plans can make the low-pressure compressor work around the common operating line.

[0139] Figure 10To analyze the engine performance index change under the control of pressure ratio control plan and boost ratio control plan after the pressure sensor drift, Fig. (a) and Fig. (b) respectively show the thrust change and the fan pressure ratio change under the control of pressure ratio control plan in P s31 drift, the thrust change and the fan pressure ratio change. It can be seen that in the middle state P s31 drift is greater than 5%, the thrust change of the engine tends to be horizontal, and the fan pressure ratio change is no longer coincident with the maximum state. This is because for P s31 positive drift will make the measured pressure ratio greater than the actual pressure ratio, and the pressure ratio control plan will make the nozzle area decrease. However, the nozzle has a mechanically minimum area, and P s31 drift is greater than 5%, the nozzle is reduced to the minimum mechanical area, and at this time P s31 positive drift increases, the nozzle area in the middle state no longer changes, the nozzle cannot make the pressure ratio control plan reach the target pressure ratio, and the pressure ratio control plan fails. In the maximum state, because the afterburner will increase the pressure after the low-pressure turbine, the nozzle area will normally increase to reach the target pressure ratio, and P s31 positive drift will make the nozzle area decrease, but it has not reached the minimum nozzle area, so after 5%, the fan pressure ratio change in the middle state and the maximum state is separated. In summary, in P s31 positive drift is greater than 5%, the pressure ratio control plan fails, and at this time the engine switches from the middle state to the afterburning state, and the engine core will have a relatively large fluctuation, which may cause a dangerous failure such as surge.

[0140] Fig. (c) and Fig. (d) respectively show the thrust change and the fan pressure ratio change under the control of boost ratio control plan in P t2 drift. It can be seen that in P t2 drift reaches 6%, the relative change of the thrust in the middle state is still not flat, and the relative change value is less than the value in Fig. (a) with the same drift, and the relative change of the pressure ratio in the middle state and the maximum state is coincident, and no separation occurs, indicating that the nozzle control plan has not touched the minimum mechanical area of the nozzle.

[0141] Fig. (e) and Fig. (f) respectively show the thrust change and the fan pressure ratio change under the control of pressure ratio control plan in P t6 drift. Fig. (g) and Fig. (h) respectively show the thrust change and the fan pressure ratio change under the control of boost ratio control plan in P t6 drift. Similar to the analysis above, it can be seen that under the same drift, the relative change of the thrust and the relative change of the fan pressure ratio of the boost ratio control plan are less than those of the pressure ratio control plan.

[0142] In summary, under the same sensor P t6With the same amount of drift, the robustness of the boost pressure ratio control schedule is better than the drop pressure ratio control schedule. s31 , P t2 With the same amount of drift, the robustness of the boost pressure ratio control schedule is better than the drop pressure ratio control schedule.

[0143] Figure 11 The effect of high pressure vane drift on the three control schedules. Fig. (a) and (b) show the relative thrust change in the intermediate and maximum states, respectively. Fig. (c) shows the relative fan pressure ratio change in the intermediate and maximum states. The relative thrust change in the intermediate state is greater than 3% for the slip control schedule, and the relative thrust change in the maximum state is greater than 2% for the slip control schedule. The relative fan pressure ratio change is less than 1% for all three control schedules. The robustness of the boost pressure ratio and drop pressure ratio control schedules is better than the slip control schedule.

[0144] Figure 12 The effect of engine individuality on the three control schedules. Fig. (a) and (b) show the relative thrust change in the intermediate and maximum states, respectively. Fig. (c) shows the relative fan pressure ratio change in the intermediate and maximum states. The relative thrust change is greater than 1% for all three control schedules. The relative thrust change for the slip control schedule is greater than the boost pressure ratio and drop pressure ratio control schedules. The relative fan pressure ratio change is less than 1% for all three control schedules.

[0145] Figure 13Fig. (a) and (b) are the relative changes of the intermediate and maximum thrust, respectively. Fig. (c) shows the relative changes of the fan pressure ratio in the intermediate and maximum states. Since the intermediate state does not touch the minimum area of the nozzle, the relative changes of the pressure ratio in the intermediate and maximum states are the same, so they are shown in the same figure. It can be seen that with the increase of the performance degradation, the relative changes of the engine thrust under the three nozzle control plans all increase, but in the intermediate state, the relative change of the thrust under the slip control plan is obviously the largest, and the relative changes of the thrust under the drop pressure ratio control and the boost pressure ratio control are almost the same. In the maximum state, the relative changes of the thrust under the three control plans are not much different, but the relative change of the thrust under the slip control is the largest, and the relative change of the thrust under the drop pressure ratio control is the smallest. For the change of the fan pressure ratio, the slip control plan also has the largest change. In summary, for the engine performance degradation, the robustness of the slip control plan is the worst, and the robustness of the drop pressure ratio control plan is slightly better than that of the boost pressure ratio control plan.

[0146] The following conclusions are drawn from the foregoing analysis:

[0147] (1) The three nozzle control plans have good robustness to sensor noise, which helps to improve the robustness of nozzle control, facilitates the rapid and effective design of a nozzle control plan that meets the requirements of control system development, avoids the problem of insufficient robustness of a single control plan that requires repeated iteration, and thus shortens the development cycle.

[0148] (2) For the drift of the shared sensor P t6 , the boost pressure ratio control and the drop pressure ratio control have the same trend of influence, but the robustness of the boost pressure ratio control plan is better than that of the drop pressure ratio. For different sensors P s31 , P t2 drifts at the same probability, the robustness of the boost pressure ratio control is also better than that of the drop pressure ratio.

[0149] (3) For the guide vane drift, the robustness of the boost pressure ratio control and the drop pressure ratio control to the guide vane drift is better than that of the slip control plan.

[0150] (4) The robustness of the three control plans to the large individual differences of the engine is that the boost pressure ratio control is better than the drop pressure ratio control, and the drop pressure ratio control is better than the slip control.

[0151] Therefore, the combined use of the three control plans can solve the problem of insufficient robustness when the engine has a large degradation and the sensor noise is affected, provide a basis for the selection of nozzle control plans, avoid the problem of insufficient robustness of a single control plan that requires repeated iteration, speed up the development process of the control plan, and be conducive to shortening the development cycle and reducing the development cost.

[0152] In the description of the specification, the description of the terms "one embodiment / way", "some embodiments / ways", "an example", "a specific example", or "some examples" and the like means that the specific features, structures, materials or characteristics described in connection with the embodiment / way or example are included in at least one embodiment / way or example of the present application. In the present specification, the illustrative description of the above terms does not necessarily refer to the same embodiment / way or example. Also, the specific features, structures, materials or characteristics described can be combined in any appropriate manner in any one or more embodiments / ways or examples. In addition, the person skilled in the art can combine and combine the different embodiments / ways or examples described in the specification and the features of the different embodiments / ways or examples, without contradiction.

[0153] In addition, the terms "first", "second" are only for the purpose of description, and cannot be understood as indicating or implying relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined with "first", "second" can explicitly or implicitly include at least one of the features. In the description of the present application, the meaning of "a plurality of" is at least two, for example, two, three, etc., unless otherwise specifically limited.

[0154] The person skilled in the art should understand that the above-mentioned embodiments are only for the purpose of clearly illustrating the present disclosure, and are not intended to limit the scope of the present disclosure. Based on the above disclosure, other changes or modifications can also be made by those skilled in the art, and these changes or modifications are still within the scope of the present disclosure.

Claims

1. A model-based robustness analysis method for aero-engine nozzle control planning, wherein, The method includes the following steps: Step S100: Obtain the common working line of the fan based on the characteristics of the fan component according to the fan component-level model; Step S200: Build the nozzle control system based on the engine component-level model; Step S300: Based on the nozzle control system, control the engine to traverse from the idle state to the intermediate state at the design point, and obtain the nozzle pressure ratio control plan, boost pressure ratio control plan, and slip control plan respectively. Step S400: Traverse the throttle lever from idle state to maximum state at the design point, and test the nozzle pressure ratio control plan, boost ratio control plan, and slip control plan respectively; Step S500: Randomly sample flight envelope points based on the flight envelope to perform full envelope simulation. Sequentially use the nozzle pressure ratio control plan, boost ratio control plan, and slip control plan to fly to each envelope point, traverse the throttle lever, obtain the actual fan pressure ratio parameters and the fan common working line pressure ratio parameters, and compare and analyze the control effects of the nozzle pressure ratio control plan, boost ratio control plan, and slip control plan. Step S600: Simulate different disturbances faced by the engine nozzle control, and conduct full envelope flight tests on the nozzle pressure ratio control plan, pressure ratio control plan, and slip control plan under different disturbances; Step S700: Whenever a disturbance situation described in step S600 is simulated, a full envelope simulation is performed according to step S500, and the thrust, fan pressure ratio, and pressure ratio at the common working line of the intermediate and maximum states are output respectively. Robustness analysis is then performed to provide a basis for the selection of the nozzle control plan.

2. The method according to claim 1, wherein, In step S100, the characteristics of the fan components refer to the flow rate-pressure ratio characteristic curve of the fan; the common working line of the fan is obtained by solving the common working equation when the airflow balance, pressure balance, coaxial work balance and speed balance are satisfied, and the tail nozzle and turbine guide are in a critical state.

3. The method according to claim 2, wherein, In step S100, the flow-pressure ratio characteristic curve of each fan is divided into 100 parts, and each dividing point is defined as point[i], i=0…100, so that the common working point is at point 50, i.e. point[50].

4. The method according to claim 3, wherein, In step S200, a closed-loop control system is built for the nozzle control system, with point[50] as the control target and a PI controller is used to obtain the nozzle control setpoint.

5. The method according to any one of claims 2-4, wherein, Step S300: The throttle lever is incremented by 1 step from the slow state where the throttle lever PLA=15 degrees to the intermediate state where the throttle lever PLA=65 degrees.

6. The method according to claim 5, wherein, The engine running step length under each throttle lever shall not be less than 2000 steps to ensure that the engine is in a steady state. When the point reaches the common working line point[50], the required parameters for the nozzle pressure ratio control plan, boost pressure ratio control plan and slip control plan shall be output.

7. The method according to claim 1, wherein, The nozzle pressure ratio control plan, boost ratio control plan, and slip control plan are respectively shown in the following formulas: ， ， ， ， ， ， ， π tAim It is the target drop ratio. π cAim It is the target boost ratio. n 1corAim Calculate the speed for the target fan. n 2cor To convert the speed of the high-pressure compressor. n 1cor Calculate the fan speed; n 2 represents the physical speed of the high-pressure compressor. n 2std This is the reference physical speed of the high-pressure compressor. T 2 represents the engine intake air temperature. n 1 represents the physical speed of the fan. n 1std This is the fan's reference physical speed; P s31 It is the static pressure at the compressor outlet. P t6 It is the total pressure at the inlet of the mixing chamber. P t2 It is the total pressure at the fan inlet. π t It is the pressure drop ratio. π c It's the boost ratio.

8. The method according to claim 7, wherein, In step S600, the different disturbances include: sensor noise, sensor drift, engine guide vane drift, individual differences, and performance degradation.

9. The method according to claim 1, wherein, Step S700 robustness analysis: The key engine performance indicators are engine thrust and fan pressure ratio. Let the different disturbances in step S600 be denoted as Δ = {Δ 噪声 Δ 传感器 Δ 导叶 Δ 个体 Δ 衰退 }, respectively calculate the robustness quantitative evaluation indicators of the key performance indicators of the engine for the nozzle pressure ratio control plan, boost ratio control plan, and slip control plan.

10. The method according to claim 9, wherein, The dimensionless relative fluctuation of the engine's key performance indicators is calculated based on different disturbances and used as a robustness evaluation index for the engine's key performance indicators. The larger the value, the worse the robustness. The formula for the dimensionless relative fluctuation is as follows: ， ， In the formula: j This represents the simulated performance output under different disturbances. "normal" represents the normal performance output without disturbance. "F" represents thrust, and "π" represents the thrust. Lc This indicates the fan pressure ratio.

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