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] The present application belongs to the field of engine design, and in particular relates to a method for determining a design margin of a turboshaft turboprop engine taking performance degradation into consideration. Background Art

[0002] During the life cycle of turboshaft turboprop engines, the effects of carbon deposition, corrosion, deformation, etc. caused by environmental factors will reduce the performance of components, resulting in a decrease in engine output shaft power and an increase in fuel consumption. It is necessary to leave margins for design parameters at the beginning of the design. Due to the limitations of aerodynamic loads, thermal loads, and mechanical loads, the engine needs to limit parameters such as converted flow, rotor speed, rotor converted speed, compressor pressure ratio, and turbine pre-turbine temperature during design. The performance degradation of various engine components will also affect these parameters, so it is necessary to consider the impact of performance degradation on limiting parameters when designing the engine. The impact of performance degradation of various engine components on limiting parameters cannot be generalized, and performance degradation is beneficial to the limitation of certain parameters, so there is no need to leave margins at this time. The impact of engine performance degradation on limiting parameters is very complex, and limiting parameters are not exceeded, which 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 conventional methods 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 point is used as the design point, and the engine is cyclically analyzed based on the performance requirements of the design point to determine the cycle parameters of 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 impact of the performance degradation of each component to leave a design margin to obtain the performance of the engine at the non-design point. If the performance calculation results of the non-design point cannot meet the performance requirements of the aircraft for the engine, it is necessary to redesign the engine design point cycle parameters, control laws, and design margins, and finally achieve the goals proposed by the aircraft through repeated iterations. This method requires multiple manual iterations, a long design cycle, and limited accuracy of the iterative results.

[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 turboprop engine taking into account performance degradation, so as to solve the problems of long design cycle and limited accuracy of iterative results of existing design methods for the correlation between performance degradation and engine design margin.

[0006] The technical solution of the present application is: a method for determining the design margin of a turboshaft turboprop engine taking into account performance degradation, comprising:

[0007] Obtaining the overall performance matching model of the turboprop engine, adding a load limit balance equation to the overall performance matching model of the turboprop engine; sorting all current operating points, and establishing a multi-design point model using all current operating points as design points;

[0008] The cycle parameters, speeds and component operating points of multiple design points are obtained in sequence, and input into the multi-design point model, and the performance requirement residual, component matching residual and iteration variable of each design point are set, and the cycle parameters of the design point and the operating points and speeds under other working conditions are iteratively calculated in the multi-design point model until the engine satisfies the requirements of component matching, performance requirements and load limit at multiple design points at the same time, and the multi-design point model calculation is completed;

[0009] The attenuation factor of each component is set on the trained multi-design point model to form a turboprop engine model that takes performance attenuation into account. The power demand and load limit of each design point are kept unchanged, and the correlation between performance attenuation and design margin is analyzed for the turboprop engine model to obtain the engine scheme and design parameters when performance attenuation is considered. The engine design parameters when performance attenuation is not considered are obtained and subtracted from the engine design parameters when performance attenuation is considered to obtain the engine design margin under performance attenuation conditions.

[0010] Preferably, the specific method for analyzing the correlation between performance degradation and design margin of a turboprop engine model is as follows:

[0011] The flight conditions and power requirements of each design point are obtained and input into the multi-design point model to obtain the engine scheme and design parameters when performance degradation is not considered; then, the design parameters such as the rotation speed, turbine inlet temperature and flow rate when performance degradation is not considered are obtained according to the engine scheme when performance degradation is not considered;

[0012] The attenuation factor of each component is obtained from the performance attenuation database and is sequentially input into the multi-design point model to form a turboprop engine model that takes performance attenuation into account; the load limit of each design point is kept unchanged, and the multi-design point model is used for calculation to obtain the engine solution when performance attenuation is taken into account;

[0013] Keep the power requirements of each design point unchanged, perform performance degradation calculations in the turboprop engine model, obtain an engine solution with margin, and then obtain the engine design parameters when considering performance degradation based on the engine solution with margin;

[0014] Difference between the engine design parameters when performance degradation is not considered and the engine design parameters when performance degradation is considered, to obtain the engine design margin and the characteristic diagram of each component under the performance degradation condition;

[0015] The component attenuation correction factor is set, and the characteristic diagram of each component is interpolated and multiplied 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 1.0, and the calculation formulas of 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 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 diagram of each component, and C a is the attenuation correction factor.

[0019] Preferably, the performance requirement residual and component matching residual include: the performance requirement residual and component matching residual of the aerodynamic design point, and the performance requirement residual and component matching residual of the remaining design points; the iteration variables of the aerodynamic design point are the air flow Wa, the total pressure ratio OPR and the power turbine extraction power P; in the twin-rotor gas generator turboprop engine, the iteration variables of the remaining design points are the low-pressure compressor pressure ratio Z F , high pressure compressor pressure ratio Z C , turbine pre-temperature T4, high-pressure turbine drop pressure ratio π HT 、Low-pressure turbine pressure drop ratio π LT 、Power turbine pressure drop ratio π FT , low speed n1 and high speed n2; in the single-spool turboshaft turboprop engine, the iteration variables of the other design points are the compressor pressure ratio Z C , turbine pre-temperature T4, turbine pressure drop ratio π T 、Power turbine pressure drop ratio π FT and speed n.

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

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

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

[0023] Fuel consumption rate residual equation:

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

[0025] The residual equation of the total pressure at the tail nozzle outlet is:

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

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

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

[0029] The residual equation of power turbine output power and propeller power requirement is:

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

[0031] The residual equations of low-pressure compressor and low-pressure turbine power are:

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

[0033] High-pressure compressor and high-pressure turbine power residual equations:

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

[0035] The flow residual equation at the low-pressure compressor outlet and the high-pressure compressor inlet is:

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

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

[0038]

[0039] The residual equation of the high-pressure turbine outlet and low-pressure turbine inlet flow rate is:

[0040]

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

[0042]

[0043] The residual equation of power turbine outlet and tail nozzle flow rate:

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

[0045] Where P is the output power, P' is the first-order derivative of P, and 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 converted flow rate of the low-pressure compressor inlet section, W cor22 is the converted flow rate of the high pressure compressor inlet section, W cor4 is the converted flow rate of the high-pressure turbine inlet section, W4 is the flow rate of the high-pressure turbine inlet section, T4 is the total temperature of the high-pressure turbine inlet section, P4 is the total pressure of the high-pressure turbine inlet section; W cor4B is the converted flow rate of the low-pressure turbine inlet section, W 4B Low-pressure turbine inlet cross-sectional flow rate, T 4B Total temperature of low-pressure turbine inlet section, P 4B Total pressure at the low-pressure turbine inlet section; W cor4C is the converted flow rate of the power turbine inlet cross section, W 4C Power turbine inlet cross-sectional flow rate, T 4C Total temperature of power turbine inlet section, P 4C Total pressure at the power turbine inlet section; W cor7 W is the converted flow rate of the tail nozzle inlet cross section, cor9 Convert flow rate to tail nozzle outlet cross section.

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

[0047] The residual equation of power turbine output power and propeller power requirement is:

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

[0049] Compressor and turbine power residual equations:

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

[0051] Residual equation of combustion chamber outlet and turbine inlet flow rate:

[0052]

[0053] The residual equation of flow rate at turbine outlet and power turbine inlet is:

[0054]

[0055] The residual equation of power turbine outlet and tail nozzle flow rate:

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

[0057] Where P is the output power, P' is the first-order derivative of P, and L T is the turbine power, L C is the compressor power, W cor4 is the converted flow rate of the turbine inlet section, W4 is the flow rate of the turbine inlet section, T4 is the total temperature of the turbine inlet section, P4 is the total pressure of the turbine inlet section; W cor5 is the converted flow rate of the turbine outlet section, W5 is the flow rate of the turbine outlet section, T5 is the total temperature of the turbine outlet section, P5 is the total pressure of the turbine outlet section; W cor7 is the converted flow rate of the power turbine outlet cross section, W cor9 Convert flow rate to tail nozzle outlet cross section.

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

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

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

[0061] Preferably, in the multi-design point model, multiple design points are iteratively calculated simultaneously, and among the multiple design points 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, 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 rotation speed and the component operating point, and the output is the performance residual and the component matching residual; at the same time, the residuals of the load limits of all design points are also used as output.

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

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

[0064] 2. Use the multi-design point model to perform calculations under conditions of performance degradation and non-degradation, and make differences in the obtained design parameters to quickly, accurately and quantitatively evaluate the design margin of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0065] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

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

[0067] Figure 2 A multi-design point model of a twin-spool gas generator turboprop engine for this application;

[0068] Figure 3 A multi-design point model of a single-spool gas generator turboprop engine for this application;

[0069] Figure 4 This is a flow chart of the performance degradation and design margin correlation analysis method for this application. DETAILED DESCRIPTION

[0070] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0071] A method for determining the design margin of a turboshaft turboprop engine considering performance degradation is proposed. The scheme for 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 twin-rotor gas generator turboprop engine.

[0072] Some parameters of the twin-rotor gas generator turboprop engine are shown in Table 1.

[0073] Table 2 Parameters of twin-rotor gas generator turboprop engine

[0074]

[0075]

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

[0077]

[0078] The steps include:

[0079] Step S100, obtaining the overall performance matching model of the turboprop engine, adding a load limit balance equation to the overall performance matching model of the turboprop engine; sorting all current operating points, and establishing a multi-design point model using 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, multiple design points are iteratively calculated at the same time, and one of the multiple design points iterated at the same time 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 speed and the component operating point, and the output is the performance residual and the component matching residual; at the same time, the residual of the load limit of all design points is also used as the output.

[0081] In the multi-design point model, the multi-design point model of the twin-spool gas generator turboprop engine is as follows: Figure 2 The multi-design point model of a single-rotor gas generator turboprop engine is shown in Figure 3 shown.

[0082] Step S200, sequentially obtain cycle parameters, rotation speeds and component operating points of multiple design points, wherein the number of the multiple design points is preferably 5; input them into a multiple design point model, set the performance requirement residual, component matching residual and iteration variable of each design point, iteratively calculate the cycle parameters of the design point and the operating points and rotation speeds under other working conditions in the multiple 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 multiple design point model.

[0083] For the aerodynamic design point, the performance requirements are that the output power of the power turbine is equal to the propeller power requirement and the fuel consumption rate is equal to the target fuel consumption rate. In addition, in order to ensure that the exhaust speed is approximately equal to the target exhaust speed, the residual equation of the total pressure at the tail nozzle outlet is added.

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

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

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

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

[0088] 2) Fuel consumption rate residual equation:

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

[0090] 3) Residual equation of total pressure at tail nozzle outlet:

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

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

[0093] Preferably, in a twin-rotor gas generator turboprop engine, for the remaining design points, the component matching conditions are the same as the non-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 of power turbine output power and propeller power requirement:

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

[0096] 2) Low-pressure compressor and low-pressure turbine power residual equation:

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

[0098] 3) High-pressure compressor and high-pressure turbine power residual equation:

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

[0100] 4) Flow residual equation at the low-pressure compressor outlet and the high-pressure compressor inlet:

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

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

[0103]

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

[0105]

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

[0107]

[0108] 8) Residual equation of power turbine outlet and tail nozzle flow rate:

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

[0110] Where P is the output power, P' is the first-order derivative of P, and 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 converted flow rate of the low-pressure compressor inlet section, W cor22 is the converted flow rate of the high pressure compressor inlet section, W cor4 is the converted flow rate of the high-pressure turbine inlet section, W4 is the flow rate of the high-pressure turbine inlet section, T4 is the total temperature of the high-pressure turbine inlet section, P4 is the total pressure of the high-pressure turbine inlet section; W cor4B is the converted flow rate of the low-pressure turbine inlet section, W 4B Low-pressure turbine inlet cross-sectional flow rate, T 4B Total temperature of low-pressure turbine inlet section, P 4B Total pressure at the low-pressure turbine inlet section; W cor4C is the converted flow rate of the power turbine inlet cross section, W 4C Power turbine inlet cross-sectional flow rate, T 4C Total temperature of power turbine inlet section, P 4C Total pressure at the power turbine inlet section; W cor7 W is the converted flow rate of the tail nozzle inlet cross section, cor9 Convert flow rate to tail nozzle outlet cross section.

[0111] The iteration variables for the remaining design points are the low-pressure compressor pressure ratio Z F , high pressure compressor pressure ratio Z C , turbine pre-temperature T4 (or combustion chamber oil-gas ratio FAR4), high-pressure turbine drop pressure ratio π HT 、Low-pressure turbine pressure drop ratio π LT 、Power turbine pressure drop ratio π FT , low speed n1 and high 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 conventional 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 equations and component matching residual equations for the remaining design points are as follows:

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

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

[0115] 2) Compressor and turbine power residual equations:

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

[0117] 3) Residual equation of combustion chamber outlet and turbine inlet flow rate:

[0118]

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

[0120]

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

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

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

[0124] The iteration variables for the remaining design points are the compressor pressure ratio Z C , turbine pre-temperature T4 (or combustion chamber oil-gas ratio FAR4), turbine drop pressure ratio π T 、Power turbine pressure drop ratio π FT and speed n.

[0125] Preferably, the load limit residual in the multi-design point model is the maximum value of the limit parameter in all design points 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 load a at the i-th design point, a c is the limit value of load a. For a twin-rotor gas generator turboprop engine with a power turbine, a includes the turbine inlet temperature T4, the low-pressure rotor conversion speed n 1cor , high pressure rotor conversion speed n 2cor , the physical speed of the low-pressure rotor n1 and the physical speed of the high-pressure rotor n2.

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

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

[0130]

[0131] Among them, n 1cor is the low pressure rotor conversion speed, n 2cor is the converted 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 temperature before the turbine, and m is the number of operating points; Z C is the high pressure compressor pressure ratio; Z F is the pressure ratio of the low-pressure compressor; π LT is the low-pressure turbine pressure drop ratio; π HT is the high-pressure turbine pressure drop ratio, π FT is the free turbine pressure drop ratio; e1-e8 are all residuals; n 1cor,c is the calculated value of low pressure conversion speed, n 2cor,c It is the calculated value of high pressure converted speed.

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

[0133] Step S300, setting the attenuation factor of each component on the trained multi-design point model to form a turboprop engine model that takes performance attenuation into account, keeping the power demand and load limit of each design point unchanged, performing a correlation analysis between performance attenuation and design margin on the turboprop engine model, and obtaining an engine scheme and design parameters when performance attenuation is considered; obtaining the engine design parameters when performance attenuation is not considered, and subtracting them from the engine design parameters when performance attenuation is considered to obtain the engine design margin under performance attenuation conditions.

[0134] Preferably, in a twin-rotor gas generator turboprop engine, the power requirements and load limits at each design point include the turbine inlet temperature T4, the low-pressure rotor conversion speed n 1cor , high pressure rotor conversion speed n 2cor , low-pressure rotor physical speed n1 and high-pressure rotor physical speed n2; in a single-rotor turboshaft turboprop engine, the power requirements and load limits of each design point include the turbine pre-temperature T4, the rotor conversion speed n cor and the rotor physical 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] The flight conditions and power requirements of each design point are obtained and input into the multi-design point model to obtain the engine scheme and design parameters when performance degradation is not considered; then, the design parameters such as the rotation speed, turbine inlet temperature and flow rate when performance degradation is not considered are obtained according to the engine scheme when performance degradation is not considered;

[0137] The attenuation factor of each component is obtained from the performance attenuation database and is sequentially input into the multi-design point model to form a turboprop engine model that takes performance attenuation into account; the load limit of each design point is kept unchanged, and the multi-design point model is used for calculation to obtain the engine solution when performance attenuation is taken into account;

[0138] Keeping the power demand at each design point unchanged, the performance degradation calculation is performed in the turboprop engine model to obtain an engine solution with margin. Then, based on the engine solution with margin, the engine design parameters such as speed, turbine inlet temperature and flow rate are obtained when considering performance degradation.

[0139] Difference between the engine design parameters when performance degradation is not considered and the engine design parameters when performance degradation is considered, to obtain the engine design margin and the characteristic diagram of each component under the performance degradation condition;

[0140] The component attenuation correction factor is set, and the characteristic diagram of each component is interpolated and multiplied by the component attenuation correction factor to obtain the flow rate, pressure ratio, efficiency and other parameters of each component after performance attenuation.

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

[0142] a act =C a ·a (19)

[0143] In the formula, a act are the flow rate, pressure ratio, efficiency and other parameters of each component after performance degradation, a is the flow rate, pressure ratio, efficiency and other parameters interpolated from the characteristic diagram of each component, Ca is the attenuation correction factor.

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

[0145] 1. Using a multi-design point model for calculations can enable the engine to simultaneously meet the requirements of component matching, performance requirements, and load limits at multiple design points. An engine performance degradation model is established based on the multi-design point model, which can quickly and accurately evaluate the design parameters of the engine under performance degradation conditions.

[0146] 2. Use the multi-design point model to perform calculations under conditions of performance degradation and non-degradation, and make differences in the obtained design parameters to quickly, accurately and quantitatively evaluate the design margin of the engine.

[0147] Finally, it should be noted that: the drawings of the embodiments disclosed in the present invention only involve structures related to the embodiments disclosed in the present invention, and other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of the present invention can be combined with each other;

[0148] Finally: 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 in the protection scope of the present invention.

Claims

1. A method for determining the design margin of a turboshaft turboprop engine considering performance degradation, characterized in that: include: Obtaining an overall performance matching model of a turboprop engine, and adding a load-limited equilibrium equation 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; The cycle parameters, speeds and component operating points of multiple design points are obtained in sequence, and input into the multi-design point model, and the performance requirement residual, component matching residual and iteration variable of each design point are set, and the cycle parameters of the design point and the operating points and speeds under other working conditions are iteratively calculated in the multi-design point model until the engine satisfies the requirements of component matching, performance requirements and load limit at multiple design points at the same time, and the multi-design point model calculation is completed; The attenuation factor of each component is set on the trained multi-design point model to form a turboprop engine model that takes performance attenuation into account. The power demand and load limit of each design point are kept unchanged. The performance attenuation and design margin correlation analysis of the turboprop engine model is performed to obtain the engine scheme and design parameters when performance attenuation is taken into account. The engine design parameters without considering performance degradation are obtained, and the engine design parameters with considering performance degradation are subtracted to obtain the engine design margin under performance degradation conditions.

2. The method for determining the design margin of a turboshaft turboprop engine considering performance degradation according to claim 1, characterized in that: The specific method for analyzing the correlation between performance degradation and design margin of the turboprop engine model is as follows: The flight conditions and power requirements of each design point are obtained and input into the multi-design point model to obtain the engine scheme and design parameters when performance degradation is not considered; then, the design parameters such as the rotation speed, turbine inlet temperature and flow rate when performance degradation is not considered are obtained according to the engine scheme when performance degradation is not considered; The attenuation factor of each component is obtained from the performance attenuation database and is sequentially input into the multi-design point model to form a turboprop engine model that takes performance attenuation into account; the load limit of each design point is kept unchanged, and the multi-design point model is used for calculation to obtain the engine solution when performance attenuation is taken into account; Keep the power requirements of each design point unchanged, perform performance degradation calculations in the turboprop engine model, obtain an engine solution with margin, and then obtain the engine design parameters when considering performance degradation based on the engine solution with margin; Difference between the engine design parameters when performance degradation is not considered and the engine design parameters when performance degradation is considered, to obtain the engine design margin and the characteristic diagram of each component under the performance degradation condition; The component attenuation correction factor is set, and the characteristic diagram of each component is interpolated and multiplied 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 turboprop engine considering performance degradation according to claim 2, characterized in that: The reference of the component attenuation correction factor is 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; In the formula, 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 diagram of each component, and C a is the attenuation correction factor.

4. The method for determining the design margin of a turboshaft turboprop engine considering performance degradation according to claim 1, characterized in that: The performance requirement residual and component matching residual include: the performance requirement residual and component matching residual of the aerodynamic design point, and the performance requirement residual and component matching residual of the other design points; the iteration variables of the aerodynamic design point are the air flow Wa, the total pressure ratio OPR and the power turbine extraction power P; in the twin-rotor gas generator turboprop engine, the iteration variables of the other design points are the low-pressure compressor pressure ratio Z F , high pressure compressor pressure ratio Z C , turbine pre-temperature T4, high-pressure turbine drop pressure ratio π HT 、Low-pressure turbine pressure drop ratio π LT 、Power turbine pressure drop ratio π FT , low speed n1 and high speed n2; in the single-spool turboshaft turboprop engine, the iteration variables of the other design points are the compressor pressure ratio Z C , turbine pre-temperature T4, turbine pressure drop ratio π T 、Power turbine pressure drop ratio π FT and speed n.

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

6. The method for determining the design margin of a turboshaft turboprop engine considering performance degradation according to claim 4, characterized in that: In the twin-spool gas generator turboprop engine, the remaining design point performance requirement residuals and component matching residuals include: The residual equation of power turbine output power and propeller power requirement is: e1=(P-P') / P The residual equations of low-pressure compressor and low-pressure turbine power are: e2=(L LT -L F ) / L F High-pressure compressor and high-pressure turbine power residual equations: e3=(L HT -L HC ) / L HC The flow residual equation at the low-pressure compressor outlet and the high-pressure compressor inlet is: e4=(W cor2 -W cor22 ) / W cor22 The residual equation of the flow rate at the combustion chamber outlet and the high-pressure turbine inlet is: The residual equation of the high-pressure turbine outlet and low-pressure turbine inlet flow rate is: The residual equation of the flow rate at the low-pressure turbine inlet and the power turbine inlet is: The residual equation of power turbine outlet and tail nozzle flow rate: e8=(W cor7 -IN cor9 ) / IN cor9 ; Where P is the output power, P' is the first-order derivative of P, and 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 converted flow rate of the low-pressure compressor inlet section, W cor22 is the converted flow rate of the high pressure compressor inlet section, W cor4 is the converted flow rate of the high-pressure turbine inlet section, W4 is the flow rate of the high-pressure turbine inlet section, T4 is the total temperature of the high-pressure turbine inlet section, P4 is the total pressure of the high-pressure turbine inlet section; W cor4B is the converted flow rate of the low-pressure turbine inlet section, W 4B Low-pressure turbine inlet cross-sectional flow rate, T 4B Total temperature of low-pressure turbine inlet section, P 4B Total pressure at the low-pressure turbine inlet section; W cor4C is the converted flow rate of the power turbine inlet section, W 4C Power turbine inlet cross-sectional flow rate, T 4C Total temperature of power turbine inlet section, P 4C Total pressure at the power turbine inlet section; W cor7 W is the converted flow rate of the tail nozzle inlet cross section, cor9 Convert flow rate to tail nozzle outlet cross section.

7. The method for determining the design margin of a turboshaft turboprop engine considering performance degradation according to claim 4, characterized in that: In the single-rotor turboshaft turboprop engine, the performance requirement residual equations and component matching residual equations of the remaining design points are as follows: The residual equation of power turbine output power and propeller power requirement is: e1=(P-P') / P Compressor and turbine power residual equations: e2=(L T -L C ) / L C Residual equation of combustion chamber outlet and turbine inlet flow rate: Residual equation of flow rate at turbine outlet and power turbine inlet: The residual equation of power turbine outlet and tail nozzle flow rate: e5=(W cor7 -IN cor9 ) / IN cor9 ; Where P is the output power, P' is the first-order derivative of P, and L T is the turbine power, L C is the compressor power, W cor4 is the converted flow rate of the turbine inlet section, W4 is the flow rate of the turbine inlet section, T4 is the total temperature of the turbine inlet section, and P4 is the total pressure of the turbine inlet section; W cor5 is the converted flow rate of the turbine outlet section, W5 is the flow rate of the turbine outlet section, T5 is the total temperature of the turbine outlet section, and P5 is the total pressure of the turbine outlet section; W cor7 is the converted flow rate of the power turbine outlet cross section, W cor9 Convert flow rate to tail nozzle outlet cross section.

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

9. The method for determining the design margin of a turboshaft turboprop engine considering performance degradation according to claim 1, characterized in that: In the multi-design point model, multiple design points are iteratively calculated simultaneously, and among the multiple design points 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, 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 rotation speed and the component operating point, and the output is the performance residual and the component matching residual; at the same time, the residuals of the load limits of all design points are also used as output.

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