阅读说明：本技术 一种航空发动机飞行包线控制方法 (Aeroengine flight envelope control method ) 是由 陈伟博 姜繁生 于明 邴连喜 于 2021-03-23 设计创作，主要内容包括：本申请属于飞机发动机控制领域,涉及一种航空发动机飞行包线控制方法,所述方法包括：确定标准大气条件下发动机飞行包线右边界总体性能参数；计算出涡轮转子基体温度T-(基体)；确定飞机在热天条件下所处最高温度T-(2max),将标准天飞机最高进气温度设为T-(2min),在T-(2min)与T-(2max)之间按设定步长补入多个控制点；给出一个初始值n1′,计算出飞机在热天条件下的涡轮转子基体温度T′-(基体)；确定T′-(基体)与T-(基体)之间的关系,调整初始值n1′,直至T′-(基体)与T-(基体)相等,确定各控制点的低压转子转速n1、高压转子转速n2以及低压涡轮出口温度T-6随涡轮转子进口温度T-2的之间的关系。本申请有助于保持航空发动机使用寿命,减小因强度、结构完整性等带来的安全性问题。(The application belongs to the field of control over aircraft engines, and relates to a method for controlling an aircraft engine flight envelope, which comprises the following steps: determining the overall performance parameters of the right boundary of the engine flight envelope under the standard atmospheric condition; calculating the temperature T of the turbine rotor base Base body (ii) a Determining the highest temperature T of the airplane under the condition of hot weather 2max Setting the maximum inlet air temperature of the standard spacecraft to T 2min At T 2min And T 2max A plurality of control points are added according to a set step length; given an initial value n1 ', the turbine rotor matrix temperature T ' of the aircraft under hot weather conditions is calculated ' Base body (ii) a Determining T' Base body And T Base body The relationship between the two, adjust the initial value n1 'to T' Base body And T Base body The low-pressure rotor speed n1, the high-pressure rotor speed n2 and the low-pressure turbine outlet temperature T of each control point are determined to be equal 6 Dependent on turbine rotor inlet temperature T 2 The relationship between (a) and (b). The application helps to maintain the service life of the aircraft engine and reduces the safety problem caused by strength, structural integrity and the like.)

1. An aeroengine flight envelope control method is characterized by comprising the following steps:

step S1, determining the overall performance parameters of the engine flight envelope right boundary under the standard atmospheric condition;

step S2, calculating the temperature T of the turbine rotor base body according to the three-dimensional simulation calculation program of the turbine rotor blade_{Base body}；

Step S3, determining the highest temperature T of the airplane under the hot weather condition_2maxSetting the maximum inlet air temperature of the standard spacecraft to T_2minAt T_2minAnd T_2maxA plurality of control points are added according to a set step length;

step S4, an initial value n1 ' is given according to the value of the low-pressure rotor speed n1 under the condition of lower than the standard day, and the turbine rotor matrix temperature T ' of the airplane under the hot day condition is calculated '_{Base body}；

Step S5 of determining T'_{Base body}And T_{Base body}Adjusting initial value n1 'to T'_{Base body}And T_{Base body}The low-pressure rotor speed n1, the high-pressure rotor speed n2 and the low-pressure turbine outlet temperature T of each control point are determined to be equal₆Dependent on turbine rotor inlet temperature T₂The relationship between (1);

step S6, low pressure rotor speed n1, high pressure rotor speed n2 and low pressure turbine outlet temperature T at each control point₆Dependent on turbine rotor inlet temperature T₂The relationship between them is used as the control plan of the right boundary of the hot sky flight envelope to control the engine.

2. The aircraft engine flight envelope control method of claim 1, wherein the engine flight envelope right boundary overall performance parameter comprises compressor inlet flow W₂₅Compressor outlet temperature T₃Compressor outlet temperature P₃Temperature T of front combustion gas of main runner turbine rotor₄₁Flow rate W₄₁Pressure P₄₁。

3. The aircraft engine flight envelope control method of claim 1, wherein in step S1, the engine flight envelope right boundary overall performance parameter under the standard atmospheric temperature condition is determined according to an engine control plan under the standard atmospheric temperature condition, wherein the engine control plan under the standard atmospheric temperature condition includes a relationship between a low pressure rotor speed, a high pressure rotor speed, and a low pressure turbine outlet temperature with a turbine rotor inlet temperature.

4. The aircraft engine flight envelope control method of claim 1, wherein step S2 further comprises:

step S21, calculating the temperature T of the turbine rotor cooling air led out of the compressor according to the overall performance parameters, the size of the engine air system flow path and the flow path loss in the step S1_{Cooling down}Flow rate W_{Cooling down}Pressure P_{Cooling down}；

Step S22, combining the temperature T of the gas before the main runner turbine rotor₄₁Flow rate W₄₁Pressure P₄₁Calculating the temperature T of the base body of the turbine rotor according to the three-dimensional simulation calculation program of the turbine rotor blade_{Base body}。

5. The aircraft engine flight envelope control method according to claim 1, wherein in step S3, the step size is set to 10 ℃.

6. The aircraft engine on-flight envelope control method according to claim 5, wherein the adjusting the initial value n 1' in step S5 includes:

if T'_{Base body}＞T_{Base body}This indicates that given a higher initial value of n1 ', the turbine rotor blades are subjected to a higher temperature load, which requires a reduction of n1 ' and, conversely, an increase of n1 '.

Technical Field

The application belongs to the field of aircraft engine control, and particularly relates to a flight envelope control method for an aircraft engine.

Background

In the design process of a control plan of the right boundary of the flight envelope, the aircraft engine usually needs to develop comprehensive performance requirements and strength and service life requirements, the two requirements are in a mutual restriction relationship, and an excessively high engine thrust requirement means that the strength and service life of the engine need to be reduced, and vice versa. In the development requirements of domestic aeroengines, corresponding performance indexes are usually specified under standard atmospheric conditions, and under the hot weather conditions exceeding the standard atmospheric temperature, no clear index requirements are usually required in the air.

At present, when a control plan of a flight envelope of a domestic aeroengine is formulated, the situations of engine service life loss and the like under the condition of a hot day (the temperature is higher than the standard atmospheric temperature) are not considered, when the right boundary of the flight envelope exceeds the standard day thermal load, the engine is still controlled according to higher engine exhaust temperature, and meanwhile, the use envelope is not limited, so that the quick consumption of the service life of the engine and potential safety risks can be caused.

In the conventional regulation plan for controlling the engine according to the same engine exhaust temperature when the air intake temperature at the right boundary of the flight envelope changes, although the temperature of the gas before the turbine does not change greatly, the outlet temperature of the compressor increases with the increase of the air intake temperature of the engine, and the following technical defects exist:

a) as the outlet temperature of the compressor increases, the strength and life reserve of the compressor components will decrease;

b) along with the increase of the outlet temperature of the compressor, the temperature of the cooling air of the turbine blade led out from the compressor is increased, so that the heat load of the working blade of the turbine is increased, and the strength and the service life reserve of the turbine part are reduced.

Disclosure of Invention

The invention provides an engine control method based on control of turbine component cooling air and turbine rotor inlet temperature under the condition of hot-day flight envelope right boundary flight, aiming at the problem that the service life of an engine is reduced by the existing hot-day flight envelope right boundary control plan, and solves the problems of quick life consumption and potential safety of an aircraft engine reduced by the existing control plan, so that the aircraft engine has the same working life in the use process of the hot-day flight envelope right boundary as that in the use process under the condition of standard atmospheric temperature, the safety, the economy and the service life of the engine are improved, and the failure rate of the engine is reduced.

The application provides an aeroengine flight envelope control method, which comprises the following steps:

step S1, determining the overall performance parameters of the engine flight envelope right boundary under the standard atmospheric condition;

step S2, calculating the temperature T of the turbine rotor base body according to the three-dimensional simulation calculation program of the turbine rotor blade_{Base body}；

Step S3, determining the highest temperature T of the airplane under the hot weather condition_2maxSetting the maximum inlet air temperature of the standard spacecraft to T_2minAt T_2minAnd T_2maxA plurality of control points are added according to a set step length;

step S4, an initial value n1 ' is given according to the value of the low-pressure rotor speed n1 under the condition of lower than the standard day, and the turbine rotor matrix temperature T ' of the airplane under the hot day condition is calculated '_{Base body}；

Step S5 of determining T'_{Base body}And T_{Base body}Adjusting initial value n1 'to T'_{Base body}And T_{Base body}The low-pressure rotor speed n1, the high-pressure rotor speed n2 and the low-pressure turbine outlet temperature T of each control point are determined to be equal₆Dependent on turbine rotor inlet temperature T₂The relationship between (1);

step S6, low pressure rotor speed n1, high pressure rotor speed n2 and low pressure turbine outlet temperature T at each control point₆Dependent on turbine rotor inlet temperature T₂The relationship between them is used as the control plan of the right boundary of the hot sky flight envelope to control the engine.

Preferably, the engine flight envelope right boundary generalitiesEnergy parameters include compressor inlet flow W₂₅Compressor outlet temperature T₃Compressor outlet temperature P₃Temperature T of front combustion gas of main runner turbine rotor₄₁Flow rate W₄₁Pressure P₄₁。

Preferably, in step S1, the engine flight envelope right boundary overall performance parameter under the standard atmospheric temperature condition is determined according to an engine control plan under the standard atmospheric temperature condition, wherein the engine control plan under the standard atmospheric temperature condition includes a relationship between a low pressure rotor speed, a high pressure rotor speed, and a low pressure turbine outlet temperature as a function of a turbine rotor inlet temperature.

Preferably, the step S2 further includes:

step S21, calculating the temperature T of the turbine rotor cooling air led out of the compressor according to the overall performance parameters, the size of the engine air system flow path and the flow path loss in the step S1_{Cooling down}Flow rate W_{Cooling down}Pressure P_{Cooling down}；

Step S22, combining the temperature T of the gas before the main runner turbine rotor₄₁Flow rate W₄₁Pressure P₄₁Calculating the temperature T of the base body of the turbine rotor according to the three-dimensional simulation calculation program of the turbine rotor blade_{Base body}。

Preferably, in step S3, the step size is set to 10 ℃.

Preferably, in step S5, the adjusting the initial value n 1' includes:

if T'_{Base body}＞T_{Base body}This indicates that given a higher initial value of n1 ', the turbine rotor blades are subjected to a higher temperature load, which requires a reduction of n1 ' and, conversely, an increase of n1 '.

Compared with the existing hot-day flight envelope right boundary control plan, the method and the device have the advantages that the service life of the aero-engine is kept, the safety problem caused by strength, structural integrity and the like is reduced, the operation of a pilot is facilitated, and the condition that the engine cannot be returned to a factory in advance due to high-temperature use is guaranteed.

Drawings

FIG. 1 is a flow chart of a preferred embodiment of the aircraft engine in-flight envelope control method of the present application.

FIG. 2 is a standard day engine control plan with T to the right₂The direction of temperature increase.

FIG. 3 is a hot day engine control plan, shown in dashed lines, with T to the right₂The direction of temperature increase.

FIG. 4 is turbine bucket cooling air temperature as a function of engine intake air temperature T₂Schematic diagram of the variation relationship.

FIG. 5 is turbine blade firing temperature T₄₁Dependent on the engine inlet temperature T₂Schematic diagram of the variation relationship.

FIG. 6 is a graph of turbine blade base temperature versus engine intake temperature T₂Schematic diagram of the variation relationship.

Detailed Description

In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the accompanying drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all embodiments of the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application, and should not be construed as limiting the present application. All other embodiments obtained by a person of ordinary skill in the art without any inventive work based on the embodiments in the present application are within the scope of protection of the present application. Embodiments of the present application will be described in detail below with reference to the drawings.

The application provides an aeroengine flight envelope control method, as shown in fig. 1, which mainly comprises the following steps:

step S1, determining the overall performance parameters of the engine flight envelope right boundary under the standard atmospheric condition;

step S2, calculating the temperature T of the turbine rotor base body according to the three-dimensional simulation calculation program of the turbine rotor blade_{Base body}；

Step S3, determining airplaneThe highest temperature T under hot weather conditions_2maxSetting the maximum inlet air temperature of the standard spacecraft to T_2minAt T_2minAnd T_2maxA plurality of control points are added according to a set step length;

step S4, an initial value n1 ' is given according to the value of the low-pressure rotor speed n1 under the condition of lower than the standard day, and the turbine rotor matrix temperature T ' of the airplane under the hot day condition is calculated '_{Base body}；

Step S5 of determining T'_{Base body}And T_{Base body}Adjusting initial value n1 'to T'_{Base body}And T_{Base body}The low-pressure rotor speed n1, the high-pressure rotor speed n2 and the low-pressure turbine outlet temperature T of each control point are determined to be equal₆Dependent on turbine rotor inlet temperature T₂The relationship between (1);

step S6, low pressure rotor speed n1, high pressure rotor speed n2 and low pressure turbine outlet temperature T at each control point₆Dependent on turbine rotor inlet temperature T₂The relationship between them is used as the control plan of the right boundary of the hot sky flight envelope to control the engine.

The details will be described below.

In step S1, the engine flight envelope right boundary overall performance parameters including the compressor inlet flow W are calculated according to the engine control plan under the standard atmospheric temperature condition₂₅Compressor outlet temperature T₃Compressor outlet temperature P₃Temperature T of front combustion gas of main runner turbine rotor₄₁Flow rate W₄₁Pressure P₄₁。

In the present embodiment, as shown in fig. 2, the engine control plan under the standard atmospheric temperature condition includes a low-pressure rotor speed n1, a high-pressure rotor speed n2, and a low-pressure turbine outlet temperature T₆Dependent on turbine rotor inlet temperature T₂Relationship between turbine rotor inlet temperature.

In some alternative embodiments, step S2 further includes:

step S21, calculating the total performance parameters and the engine air system flow path size and flow path loss according to the step S1Temperature T of cooling gas of turbine rotor led out from gas compressor_{Cooling down}Flow rate W_{Cooling down}Pressure P_{Cooling down}；

Step S22, combining the temperature T of the gas before the main runner turbine rotor₄₁Flow rate W₄₁Pressure P₄₁Calculating the temperature T of the base body of the turbine rotor according to the three-dimensional simulation calculation program of the turbine rotor blade_{Base body}。

In some alternative embodiments, in step S3, the maximum inlet air temperature of the aircraft in hot weather is calculated by altitude and mach number according to the actual flight capability of the aircraft, and is set as T_2maxSetting the maximum inlet air temperature of the standard spacecraft to T_2minAccording to a 10 deg.C (adjustable) step size at T_2minAnd T_2maxFill in i points in between, add T_2minAnd T_2maxThe total of (i +2) points is accumulated, and calculation is performed for each point according to the following steps S4 and S5.

In step S4, an initial value n1 ' is given according to the n1 value under the standard day condition, the pneumatic parameters in step S1 and step S2 are calculated, and the turbine rotor matrix temperature T ' under the hot day condition is calculated '_{Base body}。

In step S5, if T'_{Base body}＜T_{Base body}When the initial value n1 'is lower and the engine thrust is more reduced, n 1' is increased, and the turbine rotor base temperature T is recalculated in step S4_{Base body}To T'_{Base body}＝T_{Base body}(ii) a If T'_{Base body}＞T_{Base body}Indicating that the initial value n1 'given at this time is high and the temperature load to which the turbine rotor blades are subjected is high, n 1' needs to be reduced, and the turbine rotor base temperature T is recalculated according to step S4_{Base body}To T'_{Base body}＝T_{Base body}(ii) a Determining T'_{Base body}＝T_{Base body}(i +2) points n1, n2, T under the conditions₆And engine inlet temperature T₂The relationship (2) of (c).

In step S6, (i +2) points n1, n2, T₆And engine inlet temperature T₂As a hot sky flight envelopeThe right boundary control plan indirectly ensures that the turbine matrix (metal) temperature does not exceed the standard, and the obtained control plan is shown by a dotted line in figure 3.

The trend of the temperature of cooling gas of the turbine rotor blade, the temperature of gas of the rotor before the turbine and the temperature of the outlet of the gas compressor along with the temperature of the inlet of the engine under the hot weather condition, which are calculated by the method, is shown in the figures 4-6, so that the temperature of the turbine matrix under the hot weather condition is equal to the temperature of the turbine matrix under the standard weather condition.

Compared with the existing right boundary control plan of the hot-day flight envelope, the method is beneficial to maintaining the service life of the aircraft engine, reducing the safety problems caused by strength, structural integrity and the like, facilitating the operation of a pilot and ensuring that the engine cannot return to a factory in advance due to high-temperature use.

The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.