Acceleration rule design method based on turboprop engine acceleration performance calculation model

Abstract

The invention relates to the technical field of aero-engine overall performance and control, in particular to an acceleration rule design method based on a turboprop engine acceleration performance calculation model. According to the method, in the process of solving a performance calculation model of the turboprop engine, fuel flow iterative solving is carried out for limiting parameter errors, a pitch angle optimization method is matched, and implementation steps and a design process of turboprop engine acceleration law design are achieved. The calculation amount generated by frequently calling a turboprop engine performance calculation model is reduced, and the design efficiency of the turboprop engine acceleration law is remarkably improved. Meanwhile, by optimizing the pitch angle, it is ensured that the turboprop engine keeps safe and stable operation in the acceleration process, and the maximum performance potential of the turboprop engine is fully exerted.

Description

Technical Field

[0001] This invention relates to the field of overall performance and control technology of aero-engines, specifically a method for designing acceleration laws based on a turboprop engine acceleration performance calculation model. Background Technology

[0002] A gas turbine propeller engine, or turboprop engine for short, is a power unit that uses high-temperature gas generated by a gas generator to drive a free turbine, which in turn drives a propeller to generate thrust. The turboprop engine involved in this invention, according to the airflow sequence, consists of an air intake, a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, a free turbine, and an exhaust nozzle. The low-pressure compressor and low-pressure turbine are coaxial, as are the high-pressure compressor and high-pressure turbine. The free turbine drives the propeller through a reduction gear. Turboprop engines are widely used in military transport aircraft, regional jets, and general aviation due to their high fuel economy, high power-to-weight ratio, and stable and reliable performance.

[0003] Generally speaking, during takeoff, climb, and acceleration of an aircraft equipped with a turboprop engine, the turboprop engine undergoes an acceleration process. During acceleration, the turboprop engine's workload changes from low to high, specifically including an increase in engine thrust, speed, and temperature, and a decrease in compressor surge margin. These changes cause the engine's operating point to approach the safe operating boundary. At the same time, because fuel flow and pitch angle need to be continuously adjusted during acceleration, the fuel flow adjustment rate and pitch angle adjustment rate may also approach the maximum adjustment capability.

[0004] For the acceleration process of a turboprop engine, careful design of the variations in fuel flow and pitch angle can ensure safe and stable engine operation during acceleration, maximizing the engine's performance potential and thus improving aircraft maneuverability. The process of designing the variations in fuel flow and pitch angle based on performance and safety requirements is called acceleration law design. However, conventional acceleration law design methods for turboprop engines still have shortcomings, specifically: Conventional acceleration law design methods primarily rely on turboprop engine performance calculation models. However, because these models lack the ability to directly solve for performance potential boundaries, designers often need to repeatedly call and iterate through these models to verify whether the designed control laws meet safety and physical constraints and ensure that the designed acceleration laws can realize the engine's maximum acceleration performance potential. Therefore, in engineering practice, using conventional acceleration law design methods is not only time-consuming and labor-intensive but also computationally inefficient, making it difficult to meet the urgent needs of next-generation turboprop engine acceleration law design.

[0005] Given the aforementioned limitations of conventional acceleration law design methods, it is necessary to explore a method that can fully consider various safety and physical constraints during acceleration, so as to ensure that the engine's performance potential is fully realized under the premise of safety and stability during engine acceleration. Summary of the Invention

[0006] The purpose of this invention is to overcome the shortcomings of the prior art and provide a method for iteratively solving the fuel flow rate in the process of solving the performance calculation model of a turboprop engine, thereby enabling the component-level model of the turboprop engine to solve the performance potential boundary of the engine. In addition, the gradient method is used to optimize the pitch angle, thereby realizing the design of the acceleration law of the turboprop engine.

[0007] To achieve the above objectives, the technical solution adopted by this invention is: a method for designing acceleration laws based on a turboprop engine acceleration performance calculation model, comprising the following steps: Step 1: Given flight conditions, using the known quantity vector and independent variable vector of the turboprop engine as input parameters, based on the turboprop engine performance calculation model, obtain the turboprop engine's high and low pressure compressor surge margin, engine cross-section parameters, engine common working equation parameters, and engine thrust. Furthermore, based on the relationship that turboprop engines with the same cross-section have equal converted flow rate and pressure, as well as the power balance of coaxial rotors, the error vector of the common working equation of the turboprop engine is calculated. Among them, the known vectors include fuel flow rate and propeller pitch angle; Step 2: Based on the surge margins of the high and low pressure compressors, engine cross-sectional parameters, and the given minimum surge margins and maximum turbine inlet temperatures of the high and low pressure compressors, obtain the relative errors of the surge margins of the high and low pressure compressors, the relative errors of the total temperature of the combustion chamber outlet and the high-pressure turbine inlet cross-section, and thus determine the calculation mode and limiting parameter errors of the turboprop engine acceleration process. Step 3: Based on the convergence of the error vector and the constraint parameter error of the engine common working equation, and combined with Newton's iteration formula, iteratively calculate the updated independent variable vector. Meanwhile, based on the surge margins of the high and low pressure compressors, engine cross-sectional parameters, calculation mode, limiting parameter errors, and known quantity vectors, the fuel flow gradient is obtained using the difference quotient method. Then, combined with Newton's iterative formula, the updated fuel flow is calculated iteratively to ensure that the surge margins of the high and low pressure compressors are not lower than the minimum surge margins of the high and low pressure compressors, and that the total temperature of the combustion chamber outlet and the high-pressure turbine inlet cross-section is not higher than the maximum turbine inlet temperature. Furthermore, based on the given maximum fuel flow rate adjustment rate, the updated fuel flow rate, and the known quantity vector, the limited fuel flow rate after the maximum fuel flow rate adjustment rate is obtained; Step 4: Based on the given flight conditions, update the independent variable vector, and limit the fuel flow, use the gradient optimization method to iteratively calculate the updated propeller pitch angle based on the turboprop engine performance calculation model. Based on the given maximum propeller pitch angle adjustment rate, the updated propeller pitch angle, and the known quantity vector, the restricted propeller pitch angle after the maximum propeller pitch angle adjustment rate is limited is obtained. Then, based on the limited fuel flow and the limited propeller pitch angle, the updated known quantity vector is obtained. Step 5: Based on the given flight conditions, update the independent variable vector and the known quantity vector. Based on the turboprop engine performance calculation model, iteratively calculate and obtain the updated engine cross-section parameters, updated engine thrust, high-pressure compressor surge margin, and updated low-pressure compressor surge margin. That is, the engine acceleration law design result is obtained by the change law of fuel flow and propeller pitch angle during the engine acceleration process.

[0008] Furthermore, in step one, the flight conditions... for: , In the formula, For flight altitude, The flight Mach number; The known quantity vector for: , In the formula, For fuel flow rate, The propeller pitch angle; The independent variable vector for: , In the formula, For low-pressure compressor pressure ratio, This refers to the pressure ratio of the high-pressure compressor. For low-pressure compressor speed, This refers to the speed of the high-pressure compressor. For free turbine speed, Calculate the flow rate for the high-pressure turbine. Calculate the flow rate for the low-pressure turbine. Calculate the flow rate for the free turbine; Based on the turboprop engine performance calculation model for: , In the formula, These are the engine cross-sectional parameters. For high-pressure compressor surge margin, For low-pressure compressor surge margin, These are the parameters for the common working equations of the engine. For engine thrust, For flight conditions, Given a vector of quantities, A vector of independent variables; The engine cross-sectional parameters mentioned above Including section number n Engine section parameters corresponding to time , is represented as: , In the formula, For section numbering, The section number is The corresponding static temperature at that time The section number is The corresponding total temperature at that time The section number is The corresponding static pressure at that time The section number is The corresponding total pressure at that time The section number is The corresponding mass flow rate at that time; Among them, the section number This corresponds to the undisturbed far-foreground section and section number. This corresponds to the inlet section and section number of the air intake. This corresponds to the inlet / outlet and low-pressure compressor inlet sections and section numbers. This corresponds to the low-pressure compressor outlet and high-pressure compressor inlet sections, and section numbers. This corresponds to the high-pressure compressor outlet and combustion chamber inlet cross-sections and their numbers. This corresponds to the combustion chamber outlet and high-pressure turbine inlet sections, and section numbers. This corresponds to the high-pressure turbine outlet and low-pressure turbine inlet sections, and section numbers. This corresponds to the low-pressure turbine outlet and free turbine inlet sections, and section numbers. , This corresponds to the free turbine outlet and tailpipe inlet section, and the tailpipe outlet section; The parameters of the common working equation for the engine include: low-pressure compressor power. High-pressure compressor power High-pressure turbine power Low-pressure turbine power Free Turbine Power propeller power Tail nozzle inlet reference pressure ; Therefore, the relationships between the equal flow rates and pressures of turboprop engines with the same cross-section, as well as the power balance of the coaxial rotors, as described in step one, specifically include: The independent variable vector Engine cross-sectional parameters In the parameters of the engine common working equation, the equivalent flow rate corresponding to section number 2 is equal to the equivalent flow rate corresponding to section number 25; High-pressure turbine converted flow rate The converted flow rate is equal to that corresponding to section number 4, low-pressure turbine converted flow rate. The converted flow rate is equal to that corresponding to section number 5, free turbine converted flow rate. It is equal to the converted flow rate corresponding to section number 6; High-pressure turbine power and high-pressure compressor power Balanced, low-pressure turbine power With low-pressure compressor power Balanced, free turbine power With propeller power balance; Total pressure and nozzle inlet reference pressure corresponding to section number 7 Equal to each other, the error vector of the engine common working equation is calculated. Among them, the section number is n The corresponding converted flow rate From engine cross-section parameters calculate: , in, The section number is The corresponding mass flow rate at that time; The section number is The corresponding total temperature at that time; The section number is Corresponding total pressure; Furthermore, the error vector of the calculated engine common working equations is... This includes the difference between the converted flow rate corresponding to section number 2 and the converted flow rate corresponding to section number 25; High-pressure turbine converted flow rate The difference between the converted flow rate and the converted flow rate corresponding to section number 4, the converted flow rate of the low-pressure turbine. The difference between the converted flow rate and the converted flow rate corresponding to section number 5, the converted flow rate of the free turbine. The difference between the converted flow rate and the flow rate corresponding to section number 6; High-pressure turbine power and high-pressure compressor power The difference is in the power of the low-pressure turbine. With low-pressure compressor power The difference is in the power of the free turbine. With propeller power difference; Total pressure and nozzle inlet reference pressure corresponding to section number 7 difference.

[0009] Furthermore, step two specifically includes: Step 21: Based on the surge margin of the high-pressure compressor Minimum surge margin of high-pressure compressor Calculate the relative error of surge margin of high-pressure compressor for: , Step 22: Based on the surge margin of the low-pressure compressor Minimum surge margin of low-pressure compressor Calculate the relative error of surge margin of low-pressure compressor. for: , Step 23: Based on the engine cross-sectional parameters corresponding to the engine combustion chamber outlet and high-pressure turbine inlet cross-sections. Maximum turbine inlet temperature Calculate the relative error of total temperature at the combustion chamber outlet and the high-pressure turbine inlet. for: , In the formula, This refers to the total temperature corresponding to section number 4; Step 24: Based on the relative error of the low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet , Determined calculation model Specifically, it includes: Given a computational pattern Initial value: ; when , , hour, ; when , , hour, ; when , , hour, ; Step 25: According to the calculation mode Relative error of low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet Determined limiting parameter error Specifically, it includes: when hour, ; when hour, ; when hour, ; when hour, .

[0010] Furthermore, the minimum surge margin for the high-pressure compressor mentioned in step two is 5%~15%, the minimum surge margin for the low-pressure compressor is 5%~15%, and the maximum turbine inlet temperature is 1500°C. K ~1650 K .

[0011] Furthermore, step three specifically includes: Step 31: Based on the error vector of the common working equation and limiting parameter error The convergence conditions include: when or When this happens, iterative solutions for independent variables, iterative solutions for fuel flow rate, and fuel flow rate change rate constraints are performed. when and At that time, the fuel flow rate will be updated. Update the vector of independent variables This enables the limitation on the rate of change of fuel flow. Among them, the iterative solution of independent variables is based on the error vector of the engine's common working equation. Therefore, by combining Newton's iteration formula, the updated independent variable vector can be obtained through iterative calculation. : , In the formula, A vector of independent variables, This represents the error vector of the engine's common operating equations. For the Jacobian matrix: , In the formula, Represents the error vector of the engine's common working equations For independent variable vectors The gradient operator; The specific iterative solution for fuel flow rate is as follows: First, according to the calculation mode Known vector Compressor surge margin, engine section parameters corresponding to section number 4 The gradient of fuel flow rate is obtained using the difference quotient method. ,include: when hour: ; when hour: , In the formula, For low-pressure compressor surge margin, Fuel flow rate, of which This represents the symbol for calculating the parameter difference in the difference quotient method; when hour: , In the formula, For high-pressure compressor surge margin, Fuel flow rate; when hour: , In the formula, This represents the total temperature corresponding to section number 4. Fuel flow rate; Then, according to the calculation mode Fuel flow gradient Limiting parameter error Known vector By combining Newton's iterative formula, the updated fuel flow rate is obtained through iterative calculation. for: ; In the formula, Fuel flow rate; Step 32: Based on the known quantity vector And the updated fuel flow obtained in step 31 With less than 1.33 Maximum fuel flow rate adjustment rate for updating fuel flow The maximum fuel flow rate regulation rate is then limited to obtain the restricted fuel flow rate. : , In the formula, Fuel flow rate, 0.025 is the calculation step size, unit is fuel flow rate. s .

[0012] Furthermore, step four specifically includes: Step 41: Based on the given flight conditions and the updated independent variable vector obtained in step three. Limit fuel flow Based on the turboprop engine performance calculation model, the engine thrust during the acceleration process is... To maximize the optimization objective, a gradient optimization method is used to calculate and update the propeller pitch angle. : , In the formula, This represents the symbol for calculating the parameter difference in gradient optimization methods. The propeller pitch angle is 0.1, the gradient optimization step size is 0.1, and the engine thrust is 0.1. Performance calculation model of turboprop engine Find: , In the formula, This is a placeholder; Step 42: Based on the known quantity vector The updated propeller pitch angle obtained in step 41 With less than 2.77 Maximum propeller pitch angle adjustment rate for updating propeller pitch angle By imposing restrictions, the maximum propeller pitch angle adjustment rate is obtained, resulting in the restricted propeller pitch angle. : , In the formula, The propeller pitch angle is 0.025, and the calculation step size is 0.025. s ; Step 43: Based on the limited fuel flow Limit propeller pitch angle This yields an updated vector of known quantities based on constraints on fuel flow rate and propeller pitch angle. for: .

[0013] Furthermore, step five specifically includes: Based on the given flight conditions Update the vector of independent variables Update the known vector Based on turboprop engine performance calculation model The updated engine cross-section parameters are obtained through iterative calculation. Upgrade engine thrust Update the surge margin of the high-pressure compressor. Update the surge margin of the low-pressure compressor. This refers to the design results of the engine acceleration law obtained by observing the changes in fuel flow and propeller pitch angle during engine acceleration. , In the formula, To update the engine cross-section parameters, To update the surge margin of the high-pressure compressor, To update the surge margin of the low-pressure compressor, To upgrade engine thrust, This is a placeholder.

[0014] The beneficial effects of this invention are as follows: The method provided by this invention, by iteratively solving for fuel flow rate errors in response to limiting parameter errors during the solution process of the turboprop engine performance calculation model, realizes the implementation steps and design process of turboprop engine acceleration law design; the method described in this invention avoids the problem of repeatedly calling the turboprop engine performance calculation model to solve turboprop engine performance parameters in the conventional turboprop engine acceleration law design process, because the turboprop engine performance calculation model does not have the ability to solve the over-acceleration potential boundary; by iteratively solving for fuel flow rate errors in response to limiting parameter errors, combined with the optimization of the pitch angle, the computational load of turboprop engine acceleration law design is reduced, and the design efficiency is improved; Based on the turboprop engine performance calculation model, this method improves the design approach to directly solve the engine performance potential boundary, reduces unnecessary calls to the turboprop engine performance calculation model, improves the efficiency of acceleration law design, overcomes the limitations of conventional turboprop engine acceleration law design methods, and provides strong support for turboprop engine acceleration law design. Attached Figure Description

[0015] Figure 1 This is a schematic diagram of the turboprop engine structure during the implementation of this invention; Figure 2 This is a comparison between the fuel flow variation pattern obtained during the implementation of this invention and the calculation results obtained using conventional acceleration pattern design methods; Figure 3 This is a comparison between the propeller pitch angle variation law obtained during the implementation of this invention and the calculation results obtained using conventional acceleration law design methods; Figure 4This is a comparison between the total thrust variation law obtained during the implementation of this invention and the calculation results obtained using conventional acceleration law design methods; Figure 5 This is a comparison between the surge margin variation law of the low-pressure compressor obtained during the implementation of this invention and the calculation results obtained using conventional acceleration law design methods; Figure 6 This is a comparison between the surge margin variation law of the high-pressure compressor obtained during the implementation of this invention and the calculation results obtained using conventional acceleration law design methods; Figure 7 This is a comparison between the total temperature variation law of the combustion chamber outlet and high-pressure turbine inlet section obtained during the implementation of this invention and the calculation results obtained using conventional acceleration law design methods; Figure 8 This invention is based on Figure 2 The obtained fuel flow rate regulation curve; Figure 9 This invention is based on Figure 3 The obtained propeller pitch angle adjustment rate curve; Figure 10 This is a comparison of the number of times the turboprop engine performance calculation model was called during the implementation of this invention with the number of times the conventional acceleration law design method was used. Detailed Implementation

[0016] The principles and features of the present invention are described below with reference to the accompanying drawings. The examples given are only for explaining the present invention and are not intended to limit the scope of the present invention.

[0017] As described in the background section, since the fuel flow rate and propeller pitch angle need to be continuously adjusted during acceleration, the fuel flow rate adjustment rate and propeller pitch angle adjustment rate may approach the maximum adjustment capability. Therefore, the method provided by this invention is to optimize the variation law of fuel flow rate and propeller pitch angle during acceleration so that the turboprop engine can achieve its maximum acceleration performance potential while meeting safety and physical constraints. The safety constraints specifically include that the surge margin of the high-pressure and low-pressure compressors is not lower than the minimum surge margin of the high-pressure and low-pressure compressors, and the total temperature of the combustion chamber outlet and high-pressure turbine inlet section is not higher than the maximum turbine inlet temperature. The physical constraints include the maximum fuel flow rate adjustment rate and the maximum propeller pitch angle adjustment rate.

[0018] To achieve the above objectives, the present invention provides the following specific embodiments: Example 1: A method for designing acceleration laws based on a turboprop engine acceleration performance calculation model, comprising the following steps: S01, Given Flight Conditions for: , In the formula, For flight altitude, The flight Mach number; Known vector for: , In the formula, For fuel flow rate, The propeller pitch angle; Independent variable vector for: , In the formula, For low-pressure compressor pressure ratio, This refers to the pressure ratio of the high-pressure compressor. For low-pressure compressor speed, This refers to the speed of the high-pressure compressor. For free turbine speed, Calculate the flow rate for the high-pressure turbine. Calculate the flow rate for the low-pressure turbine. Calculate the flow rate for the free turbine; Based on turboprop engine performance calculation model : , In the formula, These are the engine cross-sectional parameters. For high-pressure compressor surge margin, For low-pressure compressor surge margin, These are the parameters for the common working equations of the engine. For engine thrust, For flight conditions, Given a vector of quantities, A vector of independent variables; The surge margins of the high- and low-pressure compressors and the engine cross-sectional parameters of the turboprop engine were obtained. Engine common working equation parameters, engine thrust; Among them, engine cross-sectional parameters Including section number n Engine section parameters corresponding to time , is represented as: , In the formula, For section numbering, The section number is The corresponding static temperature at that time The section number is The corresponding total temperature at that time The section number is The corresponding static pressure at that time The section number is The corresponding total pressure at that time The section number is The corresponding mass flow rate at that time; Among them, the section number This corresponds to the undisturbed far-foreground section and section number. This corresponds to the inlet section and section number of the air intake. This corresponds to the inlet / outlet and low-pressure compressor inlet sections and section numbers. This corresponds to the low-pressure compressor outlet and high-pressure compressor inlet sections, and section numbers. This corresponds to the high-pressure compressor outlet and combustion chamber inlet cross-sections and their numbers. This corresponds to the combustion chamber outlet and high-pressure turbine inlet sections, and section numbers. This corresponds to the high-pressure turbine outlet and low-pressure turbine inlet sections, and section numbers. This corresponds to the low-pressure turbine outlet and free turbine inlet sections, and section numbers. , This corresponds to the free turbine outlet and tailpipe inlet section, and the tailpipe outlet section; The obtained parameters for the common operating equations of the engine include: low-pressure compressor power. High-pressure compressor power High-pressure turbine power Low-pressure turbine power Free Turbine Power propeller power Tail nozzle inlet reference pressure .

[0019] S02. Based on the relationship of equal flow rate and pressure for turboprop engines with the same cross-section, and the power balance of coaxial rotors, Then we have the independent variable vector of S01. Engine cross-sectional parameters In the parameters of the engine common working equation, the equivalent flow rate corresponding to section number 2 is equal to the equivalent flow rate corresponding to section number 25; High-pressure turbine converted flow rate The converted flow rate is equal to that corresponding to section number 4, low-pressure turbine converted flow rate. The converted flow rate is equal to that corresponding to section number 5, free turbine converted flow rate. It is equal to the converted flow rate corresponding to section number 6; High-pressure turbine power and high-pressure compressor power Balanced, low-pressure turbine power With low-pressure compressor power Balanced, free turbine power With propeller power balance; Total pressure and nozzle inlet reference pressure corresponding to section number 7 Equal to each other, the error vector of the engine common working equation is calculated. Among them, the section number is n The corresponding converted flow rate From engine cross-section parameters calculate: , in, The section number is The corresponding mass flow rate at that time; The section number is The corresponding total temperature at that time; The section number is Corresponding total pressure; Furthermore, the error vector of the calculated engine common working equations is... include: The difference between the converted flow rate corresponding to section number 2 and the converted flow rate corresponding to section number 25; High-pressure turbine converted flow rate The difference between the converted flow rate and the converted flow rate corresponding to section number 4, the converted flow rate of the low-pressure turbine. The difference between the converted flow rate and the converted flow rate corresponding to section number 5, the converted flow rate of the free turbine. The difference between the converted flow rate and the flow rate corresponding to section number 6; High-pressure turbine power and high-pressure compressor power The difference is in the power of the low-pressure turbine. With low-pressure compressor power The difference is in the power of the free turbine. With propeller power difference; Total pressure and nozzle inlet reference pressure corresponding to section number 7 difference.

[0020] S03, Based on the surge margin of the high-pressure compressor Minimum surge margin of high-pressure compressor Calculate the relative error of surge margin of high-pressure compressor for: .

[0021] S04, Based on the surge margin of low-pressure compressors Minimum surge margin of low-pressure compressor Calculate the relative error of surge margin of low-pressure compressor. for: .

[0022] S05. Engine section parameters based on the combustion chamber outlet and high-pressure turbine inlet sections corresponding to section number 4. Maximum turbine inlet temperature Calculate the relative error of total temperature at the combustion chamber outlet and the high-pressure turbine inlet. for: , In the formula, This is the total temperature corresponding to section number 4.

[0023] S06, Based on the relative error of low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet , Determined calculation model Specifically, it includes: Given a computational pattern Initial value: ; when , , hour, ; when , , hour, ; when , , hour, .

[0024] S07, According to the calculation mode Relative error of low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet Determined limiting parameter error Specifically, it includes: when hour, ; when hour, ; when hour, ; when hour, .

[0025] S08, Error vector based on the engine common working equation and limiting parameter error The convergence situation is as follows or When this happens, iterative solutions for independent variables, iterative solutions for fuel flow rate, and fuel flow rate change rate constraints are performed. The iterative solution of independent variables involves using the error vector of the engine's common working equations, combined with Newton's iteration formula, to iteratively calculate and update the independent variable vector. : , In the formula, A vector of independent variables, This represents the error vector of the engine's common operating equations. For the Jacobian matrix: , In the formula, Represents the error vector of the engine's common working equations For independent variable vectors The gradient operator.

[0026] S09, According to the calculation mode Known vector Compressor surge margin, engine section parameters corresponding to section number 4 The gradient of fuel flow rate is obtained using the difference quotient method. ,include: when hour: ; when hour: , In the formula, For low-pressure compressor surge margin, Fuel flow rate, of which This represents the symbol for calculating the parameter difference in the difference quotient method; when hour: , In the formula, For high-pressure compressor surge margin, Fuel flow rate; when hour: , In the formula, This represents the total temperature corresponding to section number 4. This refers to fuel flow rate.

[0027] S10, According to the calculation mode Fuel flow gradient Limiting parameter error Known vector By combining Newton's iterative formula, the updated fuel flow rate is obtained through iterative calculation. for: , In the formula, This refers to fuel flow rate.

[0028] S11. Based on the known quantity vector The updated fuel flow obtained from S12 With less than 1.33 Maximum fuel flow rate adjustment rate for updating fuel flow The maximum fuel flow rate regulation rate is then limited to obtain the restricted fuel flow rate. : , In the formula, Fuel flow rate, 0.025 is the calculation step size, unit is fuel flow rate. s .

[0029] S12, based on the given flight conditions and the updated independent variable vector obtained in step three. Limit fuel flow Based on the turboprop engine performance calculation model, the engine thrust during the acceleration process is... To maximize the optimization objective, a gradient optimization method is used to calculate and update the propeller pitch angle. : , In the formula, This represents the symbol for calculating the parameter difference in gradient optimization methods. The propeller pitch angle is 0.1, the gradient optimization step size is 0.1, and the engine thrust is 0.1. Performance calculation model of turboprop engine Find: , In the formula, This is a placeholder.

[0030] S13. Based on the known quantity vector The updated propeller pitch angle obtained in step 41 With less than 2.77 Maximum propeller pitch angle adjustment rate for updating propeller pitch angle By imposing restrictions, the maximum propeller pitch angle adjustment rate is obtained, resulting in the restricted propeller pitch angle. : , In the formula, The propeller pitch angle is 0.025, and the calculation step size is 0.025. s .

[0031] S14, Based on the limited fuel flow Limit propeller pitch angle This yields an updated vector of known quantities based on constraints on fuel flow rate and propeller pitch angle. for: .

[0032] S15. Based on the given flight conditions Update the vector of independent variables Update the known vector Based on turboprop engine performance calculation model The updated engine cross-section parameters are obtained through iterative calculation. Upgrade engine thrust Update the surge margin of the high-pressure compressor. Update the surge margin of the low-pressure compressor. This refers to the design results of the engine acceleration law obtained by observing the changes in fuel flow and propeller pitch angle during engine acceleration. , In the formula, To update the engine cross-section parameters, To update the surge margin of the high-pressure compressor, To update the surge margin of the low-pressure compressor, To upgrade engine thrust, This is a placeholder.

[0033] Example 2 is the same as Example 1, except that the minimum surge margin of the high-pressure compressor in S03 is 5%~15%, the minimum surge margin of the low-pressure compressor in S04 is 5%~15%, and the maximum turbine inlet temperature in S05 is 1500°C. K ~1650 K .

[0034] Example 3: Same as Example 1, except that: S08, Error vector based on the engine common working equation and limiting parameter error The convergence situation is as follows and At that time, the fuel flow rate will be updated. Update the vector of independent variables This leads to the implementation of the same fuel flow rate limitation as S11.

[0035] like Figure 1-10 As shown, in order to further illustrate the technical solution and technical effects of the present invention, the following is based on... Figure 1 The turboprop engine structure shown uses the acceleration law design method of the turboprop engine acceleration performance calculation model provided by this invention, and the following specific calculation example is provided: S01, Given Flight Conditions for: , In the formula, flight altitude For 8000 m Mach number of flight Ma It is 0.65; Known vector for: , In the formula, fuel flow rate for propeller pitch angle for ; Independent variable vector for: , In the formula, the low-pressure compressor pressure ratio The pressure ratio of the high-pressure compressor is 0.8. The low-pressure compressor speed is 0.6. The high-pressure compressor speed is 49%. 76%, free turbine speed The high-pressure turbine converted flow rate is 54%. The calculated flow rate of the low-pressure turbine is 377. The free turbine has a flow rate of 864. It is 1431; Given the minimum surge margin of the high-pressure compressor Minimum surge margin of low-pressure compressor Maximum turbine inlet temperature ; Based on turboprop engine performance calculation model : , In the formula, These are the engine cross-sectional parameters. For high-pressure compressor surge margin, For low-pressure compressor surge margin, These are the parameters for the common working equations of the engine. For engine thrust, For flight conditions, Given a vector of quantities, A vector of independent variables; Obtain the surge margin of the high-pressure compressor of the turboprop engine The surge margin of the low-pressure compressor is 20.8%, and the engine cross-section parameters are 23.1%. Engine thrust It is 2700.6 N Engine common working equation parameters; Among them, engine cross-sectional parameters Including section number n Engine section parameters corresponding to time , is represented as: , In the formula, For section numbering, The section number is The corresponding static temperature at that time The section number is The corresponding total temperature at that time The section number is The corresponding static pressure at that time The section number is The corresponding total pressure at that time The section number is The corresponding mass flow rate at that time; Among them, the section number This corresponds to the undisturbed far-foreground section and section number. This corresponds to the inlet section and section number of the air intake. This corresponds to the inlet / outlet and low-pressure compressor inlet sections and section numbers. This corresponds to the low-pressure compressor outlet and high-pressure compressor inlet sections, and section numbers. This corresponds to the high-pressure compressor outlet and combustion chamber inlet cross-sections and their numbers. This corresponds to the combustion chamber outlet and high-pressure turbine inlet sections, and section numbers. This corresponds to the high-pressure turbine outlet and low-pressure turbine inlet sections, and section numbers. This corresponds to the low-pressure turbine outlet and free turbine inlet sections, and section numbers. , This corresponds to the free turbine outlet and tailpipe inlet section, and the tailpipe outlet section; The obtained parameters for the common operating equations of the engine include: low-pressure compressor power. It is 135.4 kW High-pressure compressor power It is 91.1 kW High-pressure turbine power It is 98.2 kW Low-pressure turbine power It is 155.7 kW Free Turbine Power 302.8 kW propeller power It is 327.6 kW Tail nozzle inlet reference pressure It is 34283.4 Pa .

[0036] S02. Based on the relationship of equal flow rate and pressure for turboprop engines with the same cross-section, and the power balance of coaxial rotors, Then we have the vector of independent variables. Engine cross-sectional parameters In the parameters of the engine common working equation, the equivalent flow rate corresponding to section number 2 is equal to the equivalent flow rate corresponding to section number 25; High-pressure turbine converted flow rate The converted flow rate is equal to that corresponding to section number 4, low-pressure turbine converted flow rate. The converted flow rate is equal to that corresponding to section number 5, free turbine converted flow rate. It is equal to the converted flow rate corresponding to section number 6; High-pressure turbine power and high-pressure compressor power Balanced, low-pressure turbine power With low-pressure compressor power Balanced, free turbine power With propeller power balance; Total pressure and nozzle inlet reference pressure corresponding to section number 7 Equal to each other, the error vector of the engine common working equation is calculated. Among them, the section number is n The corresponding converted flow rate From engine cross-section parameters calculate: , in, The section number is The corresponding mass flow rate at that time; The section number is The corresponding total temperature at that time; The section number is Corresponding total pressure; Furthermore, the error vector of the calculated engine common working equations is... include: The difference between the converted flow rate corresponding to section number 2 and the converted flow rate corresponding to section number 25; High-pressure turbine converted flow rate The difference between the converted flow rate and that corresponding to section number 4 is -29.8, the converted flow rate of the low-pressure turbine. The difference between the converted flow rate and that corresponding to section number 5 is 54.1, which is the converted flow rate of the free turbine. The difference between the converted flow rate and that corresponding to section number 6 is 81.5; High-pressure turbine power and high-pressure compressor power The difference is 7.1 kW Low-pressure turbine power With low-pressure compressor power The difference is 20.3. kW Free Turbine Power With propeller power The difference is -24.8. kW ; Total pressure and nozzle inlet reference pressure corresponding to section number 7 The difference is 2639.5. Pa .

[0037] S03, Based on the surge margin of the high-pressure compressor Minimum surge margin of high-pressure compressor Calculate the relative error of surge margin of high-pressure compressor 1.31: .

[0038] S04, Based on the surge margin of low-pressure compressors Minimum surge margin of low-pressure compressor Calculate the relative error of surge margin of low-pressure compressor. It is 1.08: .

[0039] S05. Engine section parameters based on the combustion chamber outlet and high-pressure turbine inlet sections corresponding to section number 4. Maximum turbine inlet temperature Calculate the relative error of total temperature at the combustion chamber outlet and the high-pressure turbine inlet. -0.49: , In the formula, the total temperature corresponding to section number 4 is... It is 966.3 K .

[0040] S06, Based on the relative error of low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet , Determined calculation model The value is 2, specifically including: Given a computational pattern Initial value: ; when , , hour, ; when , , hour, ; when , , hour, .

[0041] S07, According to the calculation mode Relative error of low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet Determined limiting parameter error The value is 1.31, specifically including: when hour, ; when hour, ; when hour, ; when hour, .

[0042] S08, Error vector based on the engine common working equation and limiting parameter error The convergence situation is as follows or When this happens, iterative solutions for independent variables, iterative solutions for fuel flow rate, and fuel flow rate change rate constraints are performed. The iterative solution of independent variables involves using the error vector of the engine's common working equations, combined with Newton's iteration formula, to iteratively calculate and update the independent variable vector. for : , In the formula, A vector of independent variables, This represents the error vector of the engine's common operating equations. For the Jacobian matrix: , In the formula, Represents the error vector of the engine's common working equations For independent variable vectors The gradient operator.

[0043] S09, According to the calculation mode Known vector Compressor surge margin, engine section parameters corresponding to section number 4 The gradient of fuel flow rate is obtained using the difference quotient method. It is -47.99, specifically including: when hour: ; when hour: , In the formula, For low-pressure compressor surge margin, Fuel flow rate, of which This represents the symbol for calculating the parameter difference in the difference quotient method; when hour: , In the formula, For high-pressure compressor surge margin, Fuel flow rate; when hour: , In the formula, This represents the total temperature corresponding to section number 4. This refers to fuel flow rate.

[0044] S10, According to the calculation mode Fuel flow gradient Limiting parameter error Known vector By combining Newton's iterative formula, the updated fuel flow rate is obtained through iterative calculation. It is 0.064 kg / s : , In the formula, This refers to fuel flow rate.

[0045] S11. Based on the known quantity vector and update fuel flow With less than 1.33 Maximum fuel flow rate adjustment rate for updating fuel flow The maximum fuel flow rate regulation rate is then limited to obtain the restricted fuel flow rate. It is 0.064 kg / s : , In the formula, Fuel flow rate, 0.025 is the calculation step size, unit is fuel flow rate. s .

[0046] S12, based on the given flight conditions and the updated independent variable vector obtained in step three. Limit fuel flow Based on the turboprop engine performance calculation model, the engine thrust during the acceleration process is... To maximize the optimization objective, a gradient optimization method is used to calculate and update the propeller pitch angle. It is 15.12 : , In the formula, This represents the symbol for calculating the parameter difference in gradient optimization methods. The propeller pitch angle is 0.1, the gradient optimization step size is 0.1, and the engine thrust is 0.1. Performance calculation model of turboprop engine Find: , In the formula, This is a placeholder.

[0047] S13. Based on the known quantity vector The updated propeller pitch angle obtained in step 41 With less than 2.77 Maximum propeller pitch angle adjustment rate for updating propeller pitch angle By imposing restrictions, the maximum propeller pitch angle adjustment rate is obtained, resulting in the restricted propeller pitch angle. It is 15.08 : , In the formula, The propeller pitch angle is 0.025, and the calculation step size is 0.025. s .

[0048] S14, Based on the limited fuel flow Limit propeller pitch angle This yields an updated vector of known quantities based on constraints on fuel flow rate and propeller pitch angle. for: , In the formula, the fuel flow rate is limited. The speed is 0.064 kg / s, limiting the propeller pitch angle. It is 15.08 .

[0049] S15. Based on the given flight conditions Update the vector of independent variables Update the known vector Based on turboprop engine performance calculation model The updated engine cross-section parameters are obtained through iterative calculation. Upgrade engine thrust It is 3588.42 N Update the surge margin of the high-pressure compressor. The updated low-pressure compressor surge margin is 10.009%. The figure of 16.287% is the result of the engine acceleration law design, obtained by analyzing the changes in fuel flow and propeller pitch angle during engine acceleration. , In the formula, To update the engine cross-section parameters, To update the surge margin of the high-pressure compressor, To update the surge margin of the low-pressure compressor, To upgrade engine thrust, This is a placeholder.

[0050] The calculations performed according to the steps described in this invention are compared with the calculation results obtained using conventional acceleration law design methods. Figure 2 As shown; A comparison of the propeller pitch angle variation law with the calculation results obtained using conventional acceleration law design methods. Figure 3 As shown; Comparison of the total thrust variation law with the calculation results obtained using conventional acceleration law design methods. Figure 4 As shown; A comparison of the surge margin variation law of low-pressure compressors with the calculation results obtained using conventional acceleration law design methods. Figure 5 As shown; A comparison of the surge margin variation law of high-pressure compressors with the calculation results obtained using conventional acceleration law design methods. Figure 6 As shown; A comparison of the total temperature variation patterns at the combustion chamber outlet and high-pressure turbine inlet sections with calculation results obtained using conventional acceleration law design methods. Figure 7 As shown; Figures 2-7 The solid lines with crosses represent the method described in this invention, while the dashed lines with circles represent conventional acceleration law design methods. The results of calculations using the acceleration law design method based on the turboprop engine acceleration performance calculation model disclosed in this invention are analyzed as follows: Conclusion 1: Through Figure 2 , Figure 3 It can be seen that the design results of the turboprop engine acceleration law (the variation law of fuel flow rate and propeller pitch angle during acceleration) obtained using the method described in this invention are consistent with the results obtained by conventional acceleration law design methods; through Figure 4 , Figure 5 , Figure 6 , Figure 7 It can be seen that during acceleration, the engine thrust variation law, low-pressure compressor surge margin variation law, high-pressure compressor surge margin variation law, and total temperature variation law of combustion chamber outlet and high-pressure turbine inlet section obtained by using the method described in this invention and the conventional acceleration law design method are consistent with the calculation results obtained by the conventional acceleration law design method. Conclusion 2: According to Figure 2 The obtained fuel flow rate regulation curve is as follows Figure 8 As shown; according to Figure 3 The obtained propeller pitch angle adjustment rate curve is as follows: Figure 9 As shown; pass Figure 5 , Figure 6 , Figure 7 , Figure 8 , Figure 9 It can be seen that, during acceleration, the calculation results obtained using the method described in this invention can ensure that the surge margin of the high-pressure and low-pressure compressors is not lower than the minimum surge margin of the high-pressure and low-pressure compressors, the total temperature of the combustion chamber outlet and the high-pressure turbine inlet section is not higher than the maximum turbine inlet temperature, and the fuel flow rate regulation rate and propeller pitch angle regulation rate are not higher than the given regulation rate limit. This shows that the acceleration law obtained using the method described in this invention can enable the engine to reach its maximum performance potential, proving the effectiveness of the method described in this invention. Conclusion 3: By statistically analyzing the number of times the turboprop engine performance calculation model was invoked during calculations using the method described in this invention and the conventional acceleration law design method, a comparison of the number of times the turboprop engine performance calculation model was invoked can be obtained. Figure 10 As shown; Figures 8 to 10 The solid lines with crosses represent the method described in this invention, while the dashed lines with circles represent conventional acceleration law design methods. pass Figure 10 It can be seen that the acceleration law design method based on the turboprop engine acceleration performance calculation model described in this invention calls the turboprop engine performance calculation model less often during the calculation process than the conventional acceleration law design method. This proves that the method described in this invention can reduce the amount of calculation compared to the conventional acceleration law design method, thereby improving the calculation efficiency and demonstrating the high efficiency of the method described in this invention. The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for designing acceleration laws based on a turboprop engine acceleration performance calculation model, characterized in that, Includes the following steps: Step 1: Given flight conditions, using the known quantity vector and independent variable vector of the turboprop engine as input parameters, based on the turboprop engine performance calculation model, obtain the turboprop engine's high and low pressure compressor surge margin, engine cross-section parameters, engine common working equation parameters, and engine thrust. Furthermore, based on the relationship that turboprop engines with the same cross-section have equal converted flow rate and pressure, as well as the power balance of coaxial rotors, the error vector of the common working equation of the turboprop engine is calculated. Among them, the known vectors include fuel flow rate and propeller pitch angle; Step 2: Based on the surge margins of the high and low pressure compressors, engine cross-sectional parameters, and the given minimum surge margins and maximum turbine inlet temperatures of the high and low pressure compressors, obtain the relative errors of the surge margins of the high and low pressure compressors, the relative errors of the total temperature of the combustion chamber outlet and the high-pressure turbine inlet cross-section, and thus determine the calculation mode and limiting parameter errors of the turboprop engine acceleration process. Step 3: Based on the convergence of the error vector and the constraint parameter error of the engine common working equation, and combined with Newton's iteration formula, iteratively calculate the updated independent variable vector. Meanwhile, based on the surge margins of the high and low pressure compressors, engine cross-sectional parameters, calculation mode, limiting parameter errors, and known quantity vectors, the fuel flow gradient is obtained using the difference quotient method. Then, combined with Newton's iterative formula, the updated fuel flow is calculated iteratively to ensure that the surge margins of the high and low pressure compressors are not lower than the minimum surge margins of the high and low pressure compressors, and that the total temperature of the combustion chamber outlet and the high-pressure turbine inlet cross-section is not higher than the maximum turbine inlet temperature. Furthermore, based on the given maximum fuel flow rate adjustment rate, the updated fuel flow rate, and the known quantity vector, the limited fuel flow rate after the maximum fuel flow rate adjustment rate is obtained; Step 4: Based on the given flight conditions, update the independent variable vector, and limit the fuel flow, use the gradient optimization method to iteratively calculate the updated propeller pitch angle based on the turboprop engine performance calculation model. Based on the given maximum propeller pitch angle adjustment rate, the updated propeller pitch angle, and the known quantity vector, the restricted propeller pitch angle after the maximum propeller pitch angle adjustment rate is limited is obtained. Then, based on the limited fuel flow and the limited propeller pitch angle, the updated known quantity vector is obtained. Step 5: Based on the given flight conditions, update the independent variable vector and the known quantity vector. Based on the turboprop engine performance calculation model, iteratively calculate and obtain the updated engine cross-section parameters, updated engine thrust, high-pressure compressor surge margin, and updated low-pressure compressor surge margin. That is, the engine acceleration law design result is obtained by the change law of fuel flow and propeller pitch angle during the engine acceleration process.

2. The acceleration law design method based on the turboprop engine acceleration performance calculation model as described in claim 1, characterized in that, In step one, the flight conditions mentioned for: ， In the formula, For flight altitude, The flight Mach number; The known quantity vector for: ， In the formula, For fuel flow rate, The propeller pitch angle; The independent variable vector for: ， In the formula, For low-pressure compressor pressure ratio, This refers to the pressure ratio of the high-pressure compressor. For low-pressure compressor speed, This refers to the speed of the high-pressure compressor. For free turbine speed, Calculate the flow rate for the high-pressure turbine. Calculate the flow rate for the low-pressure turbine. Calculate the flow rate for the free turbine; Based on the turboprop engine performance calculation model for: ， In the formula, These are the engine cross-sectional parameters. For high-pressure compressor surge margin, For low-pressure compressor surge margin, These are the parameters for the common working equations of the engine. For engine thrust, For flight conditions, Given a vector of quantities, A vector of independent variables; The engine cross-sectional parameters mentioned above Including section number n Engine cross-section parameters at that time , represented as: ， In the formula, For section numbering, The section number is The corresponding static temperature at that time The section number is The corresponding total temperature at that time The section number is The corresponding static pressure at that time The section number is The corresponding total pressure at that time The section number is The corresponding mass flow rate at that time; Among them, the section number This corresponds to the undisturbed far-foreground section and section number. This corresponds to the inlet section and section number of the air intake. This corresponds to the inlet / outlet and low-pressure compressor inlet sections and section numbers. This corresponds to the low-pressure compressor outlet and high-pressure compressor inlet sections, and section numbers. This corresponds to the high-pressure compressor outlet and combustion chamber inlet cross-sections and their numbers. This corresponds to the combustion chamber outlet and high-pressure turbine inlet sections, and section numbers. This corresponds to the high-pressure turbine outlet and low-pressure turbine inlet sections, and section numbers. This corresponds to the low-pressure turbine outlet and free turbine inlet sections, and section numbers. , This corresponds to the free turbine outlet and tailpipe inlet section, and the tailpipe outlet section; The parameters of the common working equation for the engine include: low-pressure compressor power. High-pressure compressor power High-pressure turbine power Low-pressure turbine power Free Turbine Power propeller power Tail nozzle inlet reference pressure ; Therefore, the relationships between the equal flow rates and pressures of turboprop engines with the same cross-section, as well as the power balance of the coaxial rotors, as described in step one, specifically include: The independent variable vector Engine cross-sectional parameters In the parameters of the engine common working equation, the equivalent flow rate corresponding to section number 2 is equal to the equivalent flow rate corresponding to section number 25; High-pressure turbine converted flow rate The converted flow rate is equal to that corresponding to section number 4, low-pressure turbine converted flow rate. The converted flow rate is equal to that corresponding to section number 5, free turbine converted flow rate. It is equal to the converted flow rate corresponding to section number 6; High-pressure turbine power and high-pressure compressor power Balanced, low-pressure turbine power With low-pressure compressor power Balanced, free turbine power With propeller power balance; Total pressure and nozzle inlet reference pressure corresponding to section number 7 Equal to each other, the error vector of the engine common working equation is calculated. Among them, the section number is n The corresponding converted flow rate From engine cross-section parameters calculate: ， in, The section number is The corresponding mass flow rate at that time; The section number is The corresponding total temperature at that time; The section number is Corresponding total pressure; Furthermore, the error vector of the calculated engine common working equations is... This includes the difference between the converted flow rate corresponding to section number 2 and the converted flow rate corresponding to section number 25; High-pressure turbine converted flow rate The difference between the converted flow rate and the converted flow rate corresponding to section number 4, the converted flow rate of the low-pressure turbine. The difference between the converted flow rate and the converted flow rate corresponding to section number 5, the converted flow rate of the free turbine. The difference between the converted flow rate and the flow rate corresponding to section number 6; High-pressure turbine power and high-pressure compressor power The difference is in the power of the low-pressure turbine. With low-pressure compressor power The difference is in the power of the free turbine. With propeller power difference; Total pressure and nozzle inlet reference pressure corresponding to section number 7 difference.

3. The acceleration law design method based on the turboprop engine acceleration performance calculation model as described in claim 1, characterized in that, Step two specifically includes: Step 21: Based on the surge margin of the high-pressure compressor Minimum surge margin of high-pressure compressor Calculate the relative error of surge margin of high-pressure compressor for: ； Step 22: Based on the surge margin of the low-pressure compressor Minimum surge margin of low-pressure compressor Calculate the relative error of surge margin of low-pressure compressor. for: ； Step 23: Based on the engine cross-sectional parameters corresponding to the engine combustion chamber outlet and high-pressure turbine inlet cross-sections. Maximum turbine inlet temperature Calculate the relative error of total temperature at the combustion chamber outlet and the high-pressure turbine inlet. for: ， In the formula, This refers to the total temperature corresponding to section number 4; Step 24: Based on the relative error of the low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet , Determined calculation model Specifically, it includes: Given a computational pattern Initial value: ; when , , hour, ; when , , hour, ; when , , hour, ; Step 25: According to the calculation mode Relative error of low-pressure compressor surge margin Relative error of surge margin of high-pressure compressor Relative error of total temperature at combustion chamber outlet and high-pressure turbine inlet Determined limiting parameter error Specifically, it includes: when hour, ; when hour, ; when hour, ; when hour, .

4. The acceleration law design method based on the turboprop engine acceleration performance calculation model as described in claim 1, characterized in that, The minimum surge margin for the high-pressure compressor mentioned in step two is 5%~15%, the minimum surge margin for the low-pressure compressor is 5%~15%, and the maximum turbine inlet temperature is 1500°C. K ~1650 K .

5. The acceleration law design method based on the turboprop engine acceleration performance calculation model as described in claim 1, characterized in that, Step three specifically includes: Step 31: Based on the error vector of the common working equation and limiting parameter error The convergence conditions include: when or When this happens, iterative solutions for independent variables, iterative solutions for fuel flow rate, and fuel flow rate change rate constraints are performed. when and At that time, the fuel flow rate will be updated. Update the vector of independent variables This enables the limitation on the rate of change of fuel flow. Among them, the iterative solution of independent variables is based on the error vector of the engine's common working equation. Therefore, by combining Newton's iteration formula, the updated independent variable vector can be obtained through iterative calculation. : ， In the formula, A vector of independent variables, This represents the error vector of the engine's common operating equations. For the Jacobian matrix: ， In the formula, Represents the error vector of the engine's common working equations For independent variable vectors The gradient operator; The specific iterative solution for fuel flow rate is as follows: First, according to the calculation mode Known vector Compressor surge margin, engine section parameters corresponding to section number 4 The gradient of fuel flow rate is obtained using the difference quotient method. ,include: when hour: ; when hour: ， In the formula, For low-pressure compressor surge margin, Fuel flow rate, of which This represents the symbol for calculating the parameter difference in the difference quotient method; when hour: ， In the formula, For high-pressure compressor surge margin, Fuel flow rate; when hour: ， In the formula, This represents the total temperature corresponding to section number 4. Fuel flow rate; Then, according to the calculation mode Fuel flow gradient Limiting parameter error Known vector By combining Newton's iterative formula, the updated fuel flow rate is obtained through iterative calculation. for: ； In the formula, Fuel flow rate; Step 32: Based on the known quantity vector And the updated fuel flow obtained in step 31 With less than 1.33 Maximum fuel flow rate adjustment rate for updating fuel flow The maximum fuel flow rate regulation rate is then limited to obtain the restricted fuel flow rate. : ， In the formula, Fuel flow rate, 0.025 is the calculation step size, unit is fuel flow rate. s .

6. The acceleration law design method based on the turboprop engine acceleration performance calculation model as described in claim 1, characterized in that, Step four specifically includes: Step 41: Based on the given flight conditions and the updated independent variable vector obtained in step three. Limit fuel flow Based on the turboprop engine performance calculation model, the engine thrust during the acceleration process is... To maximize the optimization objective, a gradient optimization method is used to calculate and update the propeller pitch angle. : ， In the formula, This represents the symbol for calculating the parameter difference in gradient optimization methods. The propeller pitch angle is 0.1, the gradient optimization step size is 0.1, and the engine thrust is 0.

1. Performance calculation model of turboprop engine Find: ， In the formula, This is a placeholder; Step 42: Based on the known quantity vector The updated propeller pitch angle obtained in step 41 With less than 2.77 Maximum propeller pitch angle adjustment rate for updating propeller pitch angle By imposing restrictions, the maximum propeller pitch angle adjustment rate is obtained, resulting in the restricted propeller pitch angle. : ， In the formula, The propeller pitch angle is 0.025, and the calculation step size is 0.

025. s ; Step 43: Based on the limited fuel flow Limit propeller pitch angle This yields an updated vector of known quantities based on constraints on fuel flow rate and propeller pitch angle. for: 。 7. The acceleration law design method based on the turboprop engine acceleration performance calculation model as described in any one of claims 1-6, characterized in that, Step five specifically refers to: Based on the given flight conditions Update the vector of independent variables Update the known vector Based on turboprop engine performance calculation model The updated engine cross-section parameters are obtained through iterative calculation. Upgrade engine thrust Update the surge margin of the high-pressure compressor. Update the surge margin of the low-pressure compressor. This refers to the design results of the engine acceleration law obtained by observing the changes in fuel flow and propeller pitch angle during engine acceleration. ， In the formula, To update the engine cross-section parameters, To update the surge margin of the high-pressure compressor, To update the surge margin of the low-pressure compressor, To upgrade engine thrust, This is a placeholder.

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