Turboshaft turboprop engine design margin determination method considering performance degradation

Abstract

The invention belongs to the field of engine design, and particularly relates to a turboshaft turboprop engine design margin determination method considering performance attenuation, which comprises the following steps: setting attenuation factors of each component on a trained multi-design point model to form a turboprop engine model considering performance attenuation, keeping power requirements and load limits of each design point unchanged, and determining the design margin of the turboshaft turboprop engine. Performing performance attenuation and design margin correlation analysis on the turboprop engine model; the multi-design-point model is used for calculation, the engine can meet the requirements of component matching, performance requirements and load limitation at multiple design points at the same time, the engine performance attenuation model is established based on the multi-design-point model, and the design parameters of the engine under the performance attenuation condition can be rapidly and accurately evaluated. A multi-design-point model is used for calculation under the performance attenuation condition and the performance non-attenuation condition, the obtained design parameters are subtracted, and the design margin of the engine can be rapidly, accurately and quantitatively evaluated.

Description

Technical Field

[0001] This application belongs to the field of engine design, and particularly relates to a method for determining the design margin of a turboshaft and turboprop engine considering performance degradation. Background Art

[0002] During the entire life cycle of a turboshaft and turboprop engine, due to the influence of factors such as engine carbon deposition, corrosion, and deformation caused by the operating environment, the performance of components will be reduced, resulting in a decrease in the output shaft power of the engine and an increase in the fuel consumption rate. Therefore, it is necessary to leave a margin for design parameters at the beginning of the design. Due to the limitations of aerodynamic load, thermal load, and mechanical load, parameters such as corrected flow rate, rotor speed, corrected rotor speed, compressor pressure ratio, and turbine inlet temperature need to be restricted during engine design. Moreover, the performance degradation of each component of the engine will also affect these parameters. Therefore, it is necessary to consider the influence of performance degradation on the restricted parameters when designing the engine. The influence of the performance degradation of each component of the engine on the restricted parameters cannot be generalized, and in some cases, the performance degradation is beneficial to the restriction of certain parameters, and in this case, there is no need to leave a margin. The influence of engine performance degradation on the restricted parameters is very complex, and the non-exceedance of the restricted parameters is an important indicator to ensure the normal operation of the engine. Therefore, it is very important to study the correlation between performance degradation and engine design margin.

[0003] When using the conventional method to analyze the correlation between turboshaft and turboprop performance degradation and engine design margin, the trial-and-error method is mainly used. First, a certain key operating condition point is taken as the design point, and the engine is analyzed cyclically starting from the performance requirements at the design point to determine the cyclic parameters at the design point. The second step is to design the corresponding control law according to the limitations of aerodynamic load, thermal load, and mechanical load, and consider the influence of the performance degradation of each component to leave a design margin to obtain the performance of the engine at off-design points. If the performance calculation results at off-design points cannot meet the performance requirements of the aircraft for the engine, it is necessary to redesign the cyclic parameters, control law, and design margin of the engine design point, and finally achieve the goal proposed by the aircraft through repeated iteration. This method requires multiple manual iterations, has a long design cycle, and the accuracy of the iteration results is limited.

[0004] Therefore, how to efficiently obtain the correlation between performance degradation and engine design margin is a problem that needs to be solved. Summary of the Invention

[0005] The purpose of this application is to provide a method for determining the design margin of a turboshaft and turboprop engine considering performance degradation, so as to solve the problems of long design cycle and limited accuracy of iteration results in the existing design method for the correlation between performance degradation and engine design margin.

[0006] The technical solution of this application is: A method for determining the design margin of a turboshaft and turboprop engine considering performance degradation includes:

[0007] Obtain the overall performance matching model of the turboprop engine, and add a balance equation for load limitation to the overall performance matching model of the turboprop engine; sort all the current operating points, and establish a multi-design point model with all the current operating points as design points;

[0008] Obtain the cycle parameters, rotational speeds, and component operating points of multiple design points in sequence, input them into the multi-design point model, set the performance requirement residuals, component matching residuals, and iteration variables for each design point, and iteratively calculate the cycle parameters of the design points and the operating points and rotational speeds under other operating conditions in the multi-design point model until the engine simultaneously meets the requirements of component matching, performance requirements, and load limitation at multiple design points, and complete the calculation of the multi-design point model;

[0009] Set the attenuation factors of each component on the trained multi-design point model to form a turboprop engine model considering performance attenuation, keep the power requirements and load limitations of each design point unchanged, conduct a correlation analysis of performance attenuation and design margin for the turboprop engine model, and obtain the engine scheme and design parameters considering performance attenuation; obtain the engine design parameters without considering performance attenuation, and take the difference from the engine design parameters considering performance attenuation to obtain the engine design margin under performance attenuation conditions.

[0010] Preferably, the specific method for conducting a correlation analysis of performance attenuation and design margin for the turboprop engine model is as follows:

[0011] Obtain the flight conditions and power requirements of each design point, input them into the multi-design point model, and obtain the engine scheme and design parameters without considering performance attenuation; then obtain the design parameters such as rotational speed, turbine inlet temperature, and flow rate without considering performance attenuation according to the engine scheme without considering performance attenuation;

[0012] Obtain the attenuation factors of each component from the performance attenuation database, and input them into the multi-design point model in sequence to form a turboprop engine model considering performance attenuation; keep the load limitations of each design point unchanged, and use the multi-design point model for calculation to obtain the engine scheme considering performance attenuation;

[0013] Keep the power requirements of each design point unchanged, conduct a calculation to eliminate performance attenuation in the turboprop engine model, obtain an engine scheme with a margin, and then obtain the engine design parameters considering performance attenuation according to the engine scheme with a margin;

[0014] Take the difference between the engine design parameters without considering performance attenuation and the engine design parameters considering performance attenuation to obtain the engine design margin under performance attenuation conditions and the characteristic diagrams of each component;

[0015] Set the component attenuation correction factor, interpolate the characteristic diagrams of each component, and multiply them by the component attenuation correction factor to obtain the flow rate, pressure ratio, and efficiency parameters of each component after performance attenuation.

[0016] Preferably, the reference of the component attenuation correction factor is taken as 1.0, and the calculation formulas for the flow rate, pressure ratio, and efficiency parameters of each component after performance attenuation are:

[0017] a act = C a ·a;

[0018] In the formula, a act is the flow rate, pressure ratio, and efficiency parameters of each component after performance attenuation, a is the flow rate, pressure ratio, and efficiency parameters obtained by interpolation in the characteristic diagrams of each component, and C a is the attenuation correction factor.

[0019] Preferably, the performance requirement residuals and component matching residuals include: the performance requirement residuals and component matching residuals at the aerodynamic design point, and the performance requirement residuals and component matching residuals at the remaining design points; the iterative variables at the aerodynamic design point are the air flow rate Wa, the total pressure ratio OPR, and the power extracted by the power turbine P; in a dual-rotor gas generator turboprop engine, the iterative variables at the remaining design points are the low-pressure compressor pressure ratio Z F , the high-pressure compressor pressure ratio Z C , the turbine inlet temperature T4, the high-pressure turbine pressure drop ratio π HT , the low-pressure turbine pressure drop ratio π LT , the power turbine pressure drop ratio π FT , the low-speed rotation speed n1 and the high-speed rotation speed n2; in a single-rotor turboshaft turboprop engine, the iterative variables at the remaining design points are the compressor pressure ratio Z C , the turbine inlet temperature T4, the turbine pressure drop ratio π T , the power turbine pressure drop ratio π FT and the rotation speed n.

[0020] Preferably, the performance requirement residuals and component matching residuals at the aerodynamic design point include:

[0021] The residual equation of the power turbine output power and the propeller power requirement:

[0022] E1 = (P - P') / P

[0023] The residual equation of the specific fuel consumption:

[0024] E2 = (SFC - SFC') / SFC'

[0025] The residual equation of the total pressure at the nozzle exit:

[0026] E3 = (P9 - P9') / P9';

[0027] Wherein, P is the output power, P' is the first derivative of P, SFC is the specific fuel consumption, and SFC' is the first derivative of SFC; P9 is the total pressure at the outlet section of the tail nozzle, and P9' is the first derivative of P9.

[0028] Preferably, in the dual-rotor gas generator turboprop engine, the remaining design point performance requirement residuals and component matching residuals include:

[0029] Residual equation of the output power of the power turbine and the power requirement of the propeller:

[0030] e1 = (P - P') / P

[0031] Residual equation of the power of the low-pressure compressor and the low-pressure turbine:

[0032] e2 = (L LT -L F ) / L F

[0033] Residual equation of the power of the high-pressure compressor and the high-pressure turbine:

[0034] e3 = (L HT -L HC ) / L HC

[0035] Residual equation of the flow rate at the outlet of the low-pressure compressor and the inlet of the high-pressure compressor:

[0036] e4 = (W cor2 -W cor22 ) / W cor22

[0037] Residual equation of the flow rate at the outlet of the combustion chamber and the inlet of the high-pressure turbine:

[0038]

[0039] Residual equation of the flow rate at the outlet of the high-pressure turbine and the inlet of the low-pressure turbine:

[0040]

[0041] Residual equation of the flow rate at the inlet of the low-pressure turbine and the inlet of the power turbine:

[0042]

[0043] Residual equation of the flow rate at the outlet of the power turbine and the tail nozzle:

[0044] e8 = (W cor7 -W cor9 ) / W cor9 ;

[0045] Wherein, P is the output power, P' is the first derivative of P, LLT is the low-pressure turbine power, L F is the low-pressure compressor power, L HT is the high-pressure turbine power, L HC is the high-pressure compressor power, W cor2 is the corrected flow rate at the low-pressure compressor inlet section, W cor22 is the corrected flow rate at the high-pressure compressor inlet section, W cor4 is the corrected flow rate at the high-pressure turbine inlet section, W4 flow rate at the high-pressure turbine inlet section, T4 total temperature at the high-pressure turbine inlet section, P4 total pressure at the high-pressure turbine inlet section; W cor4B is the corrected flow rate at the low-pressure turbine inlet section, W 4B flow rate at the low-pressure turbine inlet section, T 4B total temperature at the low-pressure turbine inlet section, P 4B total pressure at the low-pressure turbine inlet section; W cor4C is the corrected flow rate at the power turbine inlet section, W 4C flow rate at the power turbine inlet section, T 4C total temperature at the power turbine inlet section, P 4C total pressure at the power turbine inlet section; W cor7 is the corrected flow rate at the nozzle inlet section, W cor9 is the corrected flow rate at the nozzle outlet section.

[0046] Preferably, in the single-rotor turboshaft and turboprop engine, the performance requirement residual equations and component matching residual equations at other design points are as follows:

[0047] Residual equation for the output power of the power turbine and the power requirement of the propeller:

[0048] e1 = (P - P') / P

[0049] Residual equation for the power of the compressor and the turbine:

[0050] e2 = (L T - L C ) / L C

[0051] Residual equation for the flow rate at the combustor outlet and the turbine inlet:

[0052]

[0053] Residual equation for the flow rate at the turbine outlet and the power turbine inlet:

[0054]

[0055] Residual equation for the flow rate at the power turbine outlet and the nozzle:

[0056] e5 = (W cor7 - Wcor9 ) / W cor9 ;

[0057] Wherein, P is the output power, P' is the first derivative of P, L T is the turbine power, L C is the compressor power, W cor4 is the corrected flow rate at the turbine inlet section, W4 is the flow rate at the turbine inlet section, T4 is the total temperature at the turbine inlet section, P4 is the total pressure at the turbine inlet section; W cor5 is the corrected flow rate at the turbine outlet section, W5 is the flow rate at the turbine outlet section, T5 is the total temperature at the turbine outlet section, P5 is the total pressure at the turbine outlet section; W cor7 is the corrected flow rate at the power turbine outlet section, W cor9 is the corrected flow rate at the nozzle outlet section.

[0058] Preferably, the load limit residual in the multi-design point model is that the maximum value of the limit parameter among all design points is equal to the limit value, which is:

[0059] max(a1, a2,..., a n ) - a c = 0

[0060] Wherein a i is the value of the load a at the i-th design point, a c is the limit value of the load a; for a dual-rotor gas generator turboprop engine with a power turbine, a includes the turbine inlet temperature T4, the corrected rotational speed n 1cor of the low-pressure rotor, the corrected rotational speed n 2cor of the high-pressure rotor, the physical rotational speed n1 of the low-pressure rotor and the physical rotational speed n2 of the high-pressure rotor.

[0061] Preferably, in the multi-design point model, iterative calculations are performed on multiple design points simultaneously, and among the multiple design points that are iterated simultaneously, one of the design points is used as the aerodynamic design point. The input of the aerodynamic design point is the cycle parameter, and the coupling characteristic diagram of the aerodynamic design point is obtained, and the output is the performance residual and the component matching residual; the input of the remaining design points is the rotational speed and the component operating point, and the output is the performance residual and the component matching residual; meanwhile, the residual of the load limit of all design points is also used as the output.

[0062] The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation in this application has the following advantages:

[0063] 1. Using the multi-design point model for calculation can enable the engine to simultaneously meet the requirements of component matching, performance requirements, and load limit at multiple design points. Based on the multi-design point model, an engine performance degradation model is established, which can quickly and accurately evaluate the design parameters of the engine under performance degradation conditions.

[0064] 2. Calculate using the multi - design - point model under two conditions of performance degradation and non - degradation respectively, and take the difference of the obtained design parameters, which can quickly, accurately and quantitatively evaluate the design margin of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0065] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of this application.

[0066] Figure 1 It is a schematic diagram of the overall process of this application;

[0067] Figure 2 It is a multi - design - point model of a dual - rotor gas - generator turboprop engine of this application;

[0068] Figure 3 It is a multi - design - point model of a single - rotor gas - generator turboprop engine of this application;

[0069] Figure 4 It is a flowchart of the method for analyzing the correlation between performance degradation and design margin of this application. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0070] The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0071] A method for determining the design margin of a turboshaft - turboprop engine considering performance degradation. The solution of establishing a performance - degradation model based on a multi - design - point model can be migrated and applied to a single - rotor turboshaft - turboprop engine and a dual - rotor gas - generator turboprop engine.

[0072] Some parameter descriptions of the dual - rotor gas - generator turboprop engine are shown in Table 1.

[0073] Table 2 Parameter Descriptions of Dual - Rotor Gas - Generator Turboprop Engine

[0074]

[0075]

[0076] Some parameter descriptions of the single - rotor turboshaft - turboprop engine are shown in Table 2.

[0077]

[0078] It includes the following steps:

[0079] Step S100: Obtain the overall performance matching model of the turboprop engine, add a balance equation for load limit to the overall performance matching model of the turboprop engine; sort all current operating points, and establish a multi-design point model with all current operating points as design points; the overall performance matching model of the turboprop engine adopts an existing model.

[0080] In the multi-design point model, iterative calculations are performed on multiple design points simultaneously. Among the multiple design points that are iterated simultaneously, one of the design points is used as the aerodynamic design point. The input of the aerodynamic design point is the cycle parameter, and the coupling characteristic diagram of the aerodynamic design point is obtained, and the output is the performance residual and the component matching residual; the inputs of the remaining design points are the rotational speed and the component operating point, and the output is the performance residual and the component matching residual; the residuals of the load limit for all design points are also used as outputs.

[0081] In the multi-design point model, the multi-design point model of the dual-rotor gas generator turboprop engine is as Figure 2 shown. The multi-design point model of the single-rotor gas generator turboprop engine is as Figure 3 shown.

[0082] Step S200: Obtain the cycle parameters, rotational speeds, and component operating points of multiple design points in sequence. The number of multiple design points is preferably 5; input them into the multi-design point model, set the performance requirement residuals, component matching residuals, and iterative variables of each design point, and iteratively calculate the cycle parameters of the design points and the operating points and rotational speeds under other operating conditions in the multi-design point model until the engine simultaneously meets the requirements of component matching, performance requirements, and load limit at multiple design points, and complete the calculation of the multi-design point model.

[0083] For the aerodynamic design point, the performance requirement is that the output power of the power turbine is equal to the propeller power requirement, and the specific fuel consumption is equal to the target specific fuel consumption. In addition, in order to ensure that the exhaust velocity is approximately equal to the target exhaust velocity, a residual equation for the total pressure at the outlet of the tailpipe is added.

[0084] Preferably, the performance requirement residuals and component matching residuals include: the performance requirement residuals and component matching residuals of the aerodynamic design point, and the performance requirement residuals and component matching residuals of the remaining design points.

[0085] The performance requirement residuals and component matching residuals of the aerodynamic design point include:

[0086] 1) Residual equation for the output power of the power turbine and the propeller power requirement:

[0087] E1 = (P - P') / P (1)

[0088] 2) Residual equation for the specific fuel consumption:

[0089] E2 = (SFC - SFC') / SFC' (2)

[0090] 3) Total pressure residual equation at the nozzle exit:

[0091] E3 = (P9 - P9') / P9' (3)

[0092] The iterative variables at the aerodynamic design point are the air flow rate Wa, the total pressure ratio OPR, and the power extracted by the power turbine P.

[0093] Preferably, in a dual-rotor gas generator turboprop engine, for the remaining design points, the component matching conditions are the same as the off-design point performance matching model of a conventional engine, and the performance requirement is that the output power of the power turbine is equal to the propeller power requirement. The remaining design point performance requirement residuals and component matching residuals include:

[0094] 1) Residual equation for the difference between the output power of the power turbine and the propeller power requirement:

[0095] e1 = (P - P') / P (4)

[0096] 2) Residual equation for the power of the low-pressure compressor and the low-pressure turbine:

[0097] e2 = (L LT -L F ) / L F (5)

[0098] 3) Residual equation for the power of the high-pressure compressor and the high-pressure turbine:

[0099] e3 = (L HT -L HC ) / L HC (6)

[0100] 4) Residual equation for the flow rate between the outlet of the low-pressure compressor and the inlet of the high-pressure compressor:

[0101] e4 = (W cor2 -W cor22 ) / W cor22 (7)

[0102] 5) Residual equation for the flow rate between the outlet of the combustion chamber and the inlet of the high-pressure turbine:

[0103]

[0104] 6) Residual equation for the flow rate between the outlet of the high-pressure turbine and the inlet of the low-pressure turbine:

[0105]

[0106] 7) Residual equation for the flow rate between the inlet of the low-pressure turbine and the inlet of the power turbine:

[0107]

[0108] 8) Residual flow equation for the outlet of the power turbine and the nozzle:

[0109] e8 = (W cor7 - W cor9 ) / W cor9 (11)

[0110] Where P is the output power, P' is the first derivative of P, L LT is the low-pressure turbine power, L F is the low-pressure compressor power, L HT is the high-pressure turbine power, L HC is the high-pressure compressor power, W cor2 is the corrected flow rate at the inlet section of the low-pressure compressor, W cor22 is the corrected flow rate at the inlet section of the high-pressure compressor, W cor4 is the corrected flow rate at the inlet section of the high-pressure turbine, W4 is the flow rate at the inlet section of the high-pressure turbine, T4 is the total temperature at the inlet section of the high-pressure turbine, P4 is the total pressure at the inlet section of the high-pressure turbine; W cor4B is the corrected flow rate at the inlet section of the low-pressure turbine, W 4B is the flow rate at the inlet section of the low-pressure turbine, T 4B is the total temperature at the inlet section of the low-pressure turbine, P 4B is the total pressure at the inlet section of the low-pressure turbine; W cor4C is the corrected flow rate at the inlet section of the power turbine, W 4C is the flow rate at the inlet section of the power turbine, T 4C is the total temperature at the inlet section of the power turbine, P 4C is the total pressure at the inlet section of the power turbine; W cor7 is the corrected flow rate at the inlet section of the nozzle, W cor9 is the corrected flow rate at the outlet section of the nozzle.

[0111] The iterative variables for the remaining design points are the low-pressure compressor pressure ratio Z F , the high-pressure compressor pressure ratio Z C , the turbine inlet temperature T4 (or the fuel-air ratio FAR4 in the combustion chamber), the high-pressure turbine pressure ratio π HT , the low-pressure turbine pressure ratio π LT , the power turbine pressure ratio π FT , the low-speed rotation speed n1 and the high-speed rotation speed n2.

[0112] Preferably, in a single-rotor turboshaft turboprop engine, for the remaining design points, the component matching conditions are the same as the non-design point performance matching model of a traditional engine, and the performance requirement is that the output power of the power turbine is equal to the propeller power requirement. The performance requirement residual equation and the component matching residual equation for the remaining design points are as follows:

[0113] 1) Residual equation of power turbine output power and propeller power demand:

[0114] e1 = (P - P') / P (13)

[0115] 2) Residual equation of compressor and turbine power:

[0116] e2 = (L T - L C ) / L C (14)

[0117] 3) Residual equation of flow rate at combustor outlet and turbine inlet:

[0118]

[0119] 4) Residual equation of flow rate at turbine outlet and power turbine inlet:

[0120]

[0121] 5) Residual equation of flow rate at power turbine outlet and nozzle:

[0122] e5 = (W cor7 - W cor9 ) / W cor9 (17)

[0123] Wherein, P is the output power, P' is the first derivative of P, L T is the turbine power, L C is the compressor power, W cor4 is the converted flow rate at the turbine inlet section, W4 is the flow rate at the turbine inlet section, T4 is the total temperature at the turbine inlet section, P4 is the total pressure at the turbine inlet section; W cor5 is the converted flow rate at the turbine outlet section, W5 is the flow rate at the turbine outlet section, T5 is the total temperature at the turbine outlet section, P5 is the total pressure at the turbine outlet section; W cor7 is the converted flow rate at the power turbine outlet section, W cor9 is the converted flow rate at the nozzle outlet section.

[0124] The iterative variables at the remaining design points are the compressor pressure ratio Z C , the turbine inlet temperature T4 (or the fuel - air ratio FAR4 in the combustor), the turbine pressure ratio π T , the power turbine pressure ratio π FT and the rotational speed n.

[0125] Preferably, the load - limit residual in the multi - design - point model is that the maximum value of the limit parameter among all design points is equal to the limit value, and the formula is as follows:

[0126] max(a1, a2,..., a n ) - ac = 0 (18)

[0127] where a i is the value of the load a at the i-th design point, and a c is the limit value of the load a. For a twin-rotor gas generator turboprop engine with a power turbine, a includes the turbine inlet temperature T4, the corrected speed n 1cor of the low-pressure rotor, the corrected speed n 2cor of the high-pressure rotor, the physical speed n1 of the low-pressure rotor, and the physical speed n2 of the high-pressure rotor.

[0128] The multi-design point model of the twin-rotor gas generator turboprop engine is shown in Table 3.

[0129] Table 3 Multi-design point model of twin-rotor gas generator turboprop engine

[0130]

[0131] where n 1cor is the corrected speed of the low-pressure rotor, n 2cor is the corrected speed of the high-pressure rotor, n1 is the physical speed of the low-pressure rotor, n2 is the physical speed of the high-pressure rotor, T4 is the turbine inlet temperature, m is the number of operating points; Z C is the pressure ratio of the high-pressure compressor; Z F is the pressure ratio of the low-pressure compressor; π LT is the pressure drop ratio of the low-pressure turbine; π HT is the pressure drop ratio of the high-pressure turbine, π FT is the pressure drop ratio of the free turbine; e1 - e8 are all residuals; n 1cor,c is the calculated value of the low-pressure corrected speed, n 2cor,c is the calculated value of the high-pressure corrected speed.

[0132] For the turboprop engine of the twin-rotor gas generator, there are 3 equations for the aerodynamic design points, 8 equations for each of the remaining design points, and a total of 5 equations for the load limit residuals of all design points. The total number of design points is m, and there are a total of 8m equations.

[0133] Step S300, set the attenuation factors of each component on the trained multi-design point model to form a turboprop engine model considering performance degradation, keep the power demand and load limit of each design point unchanged, conduct an analysis on the correlation between performance degradation and design margin of the turboprop engine model, and obtain the engine scheme and design parameters considering performance degradation; obtain the engine design parameters without considering performance degradation, and take the difference from the engine design parameters considering performance degradation to obtain the engine design margin under performance degradation conditions.

[0134] Preferably, in a dual-rotor gas generator turboprop engine, the power requirements and load limits at each design point include the turbine inlet temperature T4, the corrected low-pressure rotor speed n 1cor , the corrected high-pressure rotor speed n 2cor , the physical low-pressure rotor speed n1 and the physical high-pressure rotor speed n2; in a single-rotor turboshaft turboprop engine, the power requirements and load limits at each design point include the turbine inlet temperature T4, the corrected rotor speed n cor and the physical rotor speed n.

[0135] Preferably, the specific method for analyzing the correlation between performance degradation and design margin of a turboprop engine model is as follows:

[0136] Obtain the flight conditions and power requirements at each design point, input them into the at-most design point model, and obtain the engine scheme and design parameters without considering performance degradation; then obtain the design parameters such as speed, turbine inlet temperature, and flow rate without considering performance degradation according to the engine scheme without considering performance degradation;

[0137] Obtain the degradation factors of each component from the performance degradation database and input them into the at-most design point model in sequence to form a turboprop engine model considering performance degradation; keep the load limits at each design point unchanged, and use the multi-design point model for calculation to obtain the engine scheme considering performance degradation;

[0138] Keep the power requirements at each design point unchanged, perform performance degradation elimination calculation in the turboprop engine model to obtain an engine scheme with a margin, and then obtain the engine design parameters such as speed, turbine inlet temperature, and flow rate considering performance degradation according to the engine scheme with a margin.

[0139] Take the difference between the engine design parameters without considering performance degradation and the engine design parameters considering performance degradation to obtain the engine design margin and the characteristic diagrams of each component under performance degradation conditions;

[0140] Set the component degradation correction factor, multiply the interpolated characteristic diagrams of each component by the component degradation correction factor to obtain the parameters such as flow rate, pressure ratio, and efficiency of each component after performance degradation.

[0141] Preferably, the reference of the component degradation correction factor is taken as 1.0, and the calculation formulas for the parameters such as flow rate, pressure ratio, and efficiency of each component after performance degradation are as follows:

[0142] a act = C a ·a(19)

[0143] In the formula, a act is the parameter such as flow rate, pressure ratio, and efficiency of each component after performance degradation, a is the parameter such as flow rate, pressure ratio, and efficiency interpolated from the characteristic diagrams of each component, Ca is the attenuation correction factor.

[0144] In summary, the present application has the following advantages:

[0145] 1. Using the multi-design point model for calculation can enable the engine to simultaneously meet the requirements of component matching, performance requirements, and load limitations at multiple design points. Based on the multi-design point model, an engine performance attenuation model is established, which can quickly and accurately evaluate the design parameters of the engine under the condition of performance attenuation.

[0146] 2. Using the multi-design point model to calculate under two conditions of performance attenuation and non-attenuation respectively, and taking the difference of the obtained design parameters, can quickly, accurately and quantitatively evaluate the design margin of the engine.

[0147] Finally, it should be noted that: in the accompanying drawings of the disclosed embodiments of the present invention, only the structures related to the disclosed embodiments are involved. For other structures, reference can be made to the general design. Without conflict, the same embodiment and different embodiments of the present invention can be combined with each other;

[0148] Finally: The above are only the preferred embodiments of the present invention and are not used to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principle of the present invention shall be included within the protection scope of the present invention.

Claims

1. A method for determining the design margin of a turboshaft and turboprop engine considering performance degradation, characterized in that Including: Obtain the overall performance matching model of a turboprop engine, and add a balance equation for load limitation to the overall performance matching model of the turboprop engine; Sort all current operating points, and use all current operating points as design points to establish a multi-design point model; Obtain the cycle parameters, rotational speeds, and component operating points of multiple design points in sequence, input them into the multi-design point model, set the performance requirement residuals, component matching residuals, and iterative variables for each design point, and iteratively calculate the cycle parameters of the design points and the operating points and rotational speeds under other operating conditions in the multi-design point model until the engine simultaneously meets the requirements of component matching, performance requirements, and load limitation at multiple design points, and complete the calculation of the multi-design point model; Set the attenuation factors of each component on the trained multi-design point model to form a turboprop engine model considering performance attenuation. Keep the power requirements and load limitations of each design point unchanged, and conduct a correlation analysis of performance attenuation and design margin for the turboprop engine model to obtain the engine scheme and design parameters considering performance attenuation; Obtain the engine design parameters without considering performance attenuation, subtract them from the engine design parameters considering performance attenuation, and obtain the engine design margin under performance attenuation conditions.

2. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation according to claim 1, wherein The specific method for conducting a correlation analysis of performance attenuation and design margin for the turboprop engine model is as follows: Obtain the flight conditions and power requirements of each design point, input them into the multi-design point model, and obtain the engine scheme and design parameters without considering performance attenuation; then obtain the design parameters such as rotational speed, turbine inlet temperature, and flow rate without considering performance attenuation according to the engine scheme without considering performance attenuation; Obtain the attenuation factors of each component from the performance attenuation database, and input them into the multi-design point model in sequence to form a turboprop engine model considering performance attenuation; keep the load limitations of each design point unchanged, and use the multi-design point model for calculation to obtain the engine scheme considering performance attenuation; Keep the power requirements of each design point unchanged, conduct performance attenuation elimination calculation in the turboprop engine model to obtain an engine scheme with a margin, and then obtain the engine design parameters considering performance attenuation according to the engine scheme with a margin; Subtract the engine design parameters without considering performance attenuation from the engine design parameters considering performance attenuation to obtain the engine design margin and the characteristic diagrams of each component under performance attenuation conditions; Set the component attenuation correction factor, interpolate the characteristic diagrams of each component, and multiply them by the component attenuation correction factor to obtain the flow rate, pressure ratio, and efficiency parameters of each component after performance attenuation.

3. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation as claimed in claim 2, wherein The reference of the component attenuation correction factor is taken as 1.0, and the calculation formulas for the flow rate, pressure ratio, and efficiency parameters of each component after performance attenuation are as follows: a act = C a · a; where a act are the flow rate, pressure ratio, and efficiency parameters of each component after performance degradation, a are the flow rate, pressure ratio, and efficiency parameters obtained by interpolation in the characteristic diagrams of each component, and C a is the attenuation correction factor.

4. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation according to claim 1, characterized in that The performance requirement residuals and component matching residuals include: the performance requirement residuals and component matching residuals at the aerodynamic design point, and the performance requirement residuals and component matching residuals at the remaining design points; the iterative variables at the aerodynamic design point are the air flow rate Wa, the overall pressure ratio OPR, and the power extracted by the power turbine P; in a dual-rotor gas generator turboprop engine, the iterative variables at the remaining design points are the low-pressure compressor pressure ratio Z F , the high-pressure compressor pressure ratio Z C , the turbine inlet temperature T4, the high-pressure turbine pressure ratio drop π HT , the low-pressure turbine pressure ratio drop π LT , the power turbine pressure ratio drop π FT , the low-speed rotation speed n1 and the high-speed rotation speed n2; in a single-rotor turboshaft turboprop engine, the iterative variables at the remaining design points are the compressor pressure ratio Z C , the turbine inlet temperature T4, the turbine pressure ratio drop π T , the power turbine pressure ratio drop π FT and the rotation speed n.

5. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation according to claim 4, wherein The performance requirement residuals and component matching residuals of the aerodynamic design point include: the residual equation of the power output of the power turbine and the power requirement of the propeller: E1 = (P - P') / P The residual equation of specific fuel consumption: E2 = (SFC - SFC') / SFC' The residual equation of the total pressure at the outlet of the tailpipe: E3 = (P9 - P9') / P9'; Wherein, P is the output power, P' is the first derivative of P, SFC is the specific fuel consumption, and SFC' is the first derivative of SFC; P9 is the total pressure at the outlet section of the tail nozzle, and P9' is the first derivative of P9.

6. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation as described in claim 4, characterized in that In the twin-rotor gas generator turboprop engine, the performance requirement residuals and component matching residuals at other design points include: Residual equation of the output power of the power turbine and the power requirement of the propeller: e1 = (P - P') / P Residual equation of the power of the low-pressure compressor and the low-pressure turbine: e2 = (L LT - L F ) / L F Residual equation of the power of the high-pressure compressor and the high-pressure turbine: e3 = (L HT - L HC ) / L HC Residual equation of the flow rate between the outlet of the low-pressure compressor and the inlet of the high-pressure compressor: e4 = (W cor2 - W cor22 ) / W cor22 Residual equation of the flow rate between the outlet of the combustor and the inlet of the high-pressure turbine: Residual equation of the flow rate between the outlet of the high-pressure turbine and the inlet of the low-pressure turbine: Residual equation of the flow rate between the inlet of the low-pressure turbine and the inlet of the power turbine: Residual equation of the flow rate between the outlet of the power turbine and the tail nozzle: e8 = (W cor7 - W cor9 ) / W cor9 ; Where P is the output power, P' is the first derivative of P, L LT is the low-pressure turbine power, L F is the low-pressure compressor power, L HT is the high-pressure turbine power, L HC is the high-pressure compressor power, W cor2 is the corrected flow rate at the inlet section of the low-pressure compressor, W cor22 is the corrected flow rate at the inlet section of the high-pressure compressor, W cor4 is the corrected flow rate at the inlet section of the high-pressure turbine, W4 is the flow rate at the inlet section of the high-pressure turbine, T4 is the total temperature at the inlet section of the high-pressure turbine, P4 is the total pressure at the inlet section of the high-pressure turbine; W cor4B is the corrected flow rate at the inlet section of the low-pressure turbine, W 4B is the flow rate at the inlet section of the low-pressure turbine, T 4B is the total temperature at the inlet section of the low-pressure turbine, P 4B is the total pressure at the inlet section of the low-pressure turbine; W cor4C is the corrected flow rate at the inlet section of the power turbine, W 4C is the flow rate at the inlet section of the power turbine, T 4C is the total temperature at the inlet section of the power turbine, P 4C is the total pressure at the inlet section of the power turbine; W cor7 is the corrected flow rate at the inlet section of the tail nozzle, W cor9 is the corrected flow rate at the outlet section of the tail nozzle.

7. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation as claimed in claim 4, wherein In the single-rotor turboshaft turboprop engine, the performance requirement residual equations and component matching residual equations at other design points are as follows: Residual equation of the output power of the power turbine and the power requirement of the propeller: e1 = (P - P') / P Residual equation of the power of the compressor and the turbine: e2 = (L T - L C ) / L C Residual equation of the flow rate between the outlet of the combustor and the inlet of the turbine: Residual equation of the flow rate between the outlet of the turbine and the inlet of the power turbine: Residual equation of the flow rate between the outlet of the power turbine and the tail nozzle: e5 = (W cor7 - W cor9 ) / W cor9 ; Wherein, P is the output power, P' is the first derivative of P, and L T is the turbine power, and L C is the compressor power, W cor4 is the converted flow rate at the turbine inlet section, W4 is the flow rate at the turbine inlet section, T4 is the total temperature at the turbine inlet section, and P4 is the total pressure at the turbine inlet section; W cor5 is the converted flow rate at the turbine outlet section, W5 is the flow rate at the turbine outlet section, T5 is the total temperature at the turbine outlet section, and P5 is the total pressure at the turbine outlet section; W cor7 is the corrected flow rate at the outlet section of the power turbine, W cor9 is the corrected flow rate at the outlet section of the exhaust nozzle.

8. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation according to claim 1, wherein The load limit residual in the multi-design point model is that the maximum value of the limit parameter at all design points is equal to the limit value, which is: max(a1,a2,...,a n )-a c =0 where a i is the value of the load a at the i-th design point, and a c is the limit value of the load a; for a dual-rotor gas generator turboprop engine with a power turbine, a includes the turbine inlet temperature T4, the corrected speed n 1cor of the low-pressure rotor, the corrected speed n 2cor of the high-pressure rotor, the physical speed n1 of the low-pressure rotor, and the physical speed n2 of the high-pressure rotor.

9. The method for determining the design margin of a turboshaft and turboprop engine considering performance degradation according to claim 1, characterized in that, In the multi-design point model, iterative calculations are performed on multiple design points simultaneously. Among the multiple design points that are iterated simultaneously, one of the design points is used as the aerodynamic design point. The input of the aerodynamic design point is the cycle parameter, and the coupling characteristic diagram of the aerodynamic design point is obtained. The output is the performance residual and the component matching residual; the inputs of the other design points are the rotational speed and the component operating point, and the outputs are the performance residual and the component matching residual; meanwhile, the residuals of the load limits of all design points are also used as outputs.

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