Turboshaft turboprop engine overall performance scheme design method meeting multiple working conditions

Abstract

According to the turboshaft turboprop engine overall performance scheme design method meeting the multiple working conditions, an engine nonlinear equation set containing a plurality of design points is constructed through an aerodynamic thermal process equation and engine complete machine matching conditions, and cycle parameters of aerodynamic design points and working points and rotating speeds under other working conditions are iterated; the invention relates to a process and a method for enabling an engine to simultaneously meet the requirements of component matching, performance requirements and load limitation at a plurality of design points (20-30 points), namely a multi-design-point model modeling and performance matching method meeting multi-working-condition-point calculation. In the process of designing engine cycle parameters and control laws under a new working condition combination by adopting a multi-design-point model, whether physical rotating speeds, conversion rotating speeds and turbine front temperatures of residual working condition points exceed limits or not under the existing cycle parameters and control laws is checked in sequence; and in the process of replacing the original working condition in the multi-design-point model with the working condition if the working condition exceeds the limit, a three-item replacement principle method is adopted for the working condition points of the engine.

Description

Technical Field

[0001] The present application relates to the field of engine performance design, and in particular to a method for designing an overall performance scheme for a turboshaft and turboprop engine that meets multiple operating conditions. Background Art

[0002] In order to successfully complete the flight mission, the aircraft often has several critical operating points within its flight envelope. The aircraft will make requirements on the engine performance at these critical operating points. These critical operating points are the engine's performance constraint points. Helicopters or transport aircraft need to operate continuously in a variety of environments. The turboshaft / turboprop engines they use have many performance requirements within the full envelope, usually reaching 20 to 30 points. These performance constraint points usually include typical operating conditions such as high speed, plateau, and plains. When designing the overall performance scheme of the engine, it is necessary to ensure that the maximum turbine inlet temperature does not exceed the temperature when the engine is operating under all flight conditions, and the physical / converted speed does not exceed the limit. These constraints ensure that the engine operates stably within the design margin and meets the engine performance constraints.

[0003] Traditional engine overall performance calculation models typically select a single point as the engine's design point, then assign a control law to calculate the performance of other non-design points from that point. Key constraints are then identified from these non-design points, and the performance of each constraint is determined to meet the aircraft's requirements. If not, the engine's design point parameters and control laws must be adjusted through continuous iteration.

[0004] The effect is not good when applied to turboshaft / turboprop engines for the following reasons:

[0005] 1. Turboshaft / turboprop engines have many performance requirements within the full envelope, 20 to 30 points, resulting in a long design cycle and a large number of modifications.

[0006] The existing method has low computational efficiency and can only simultaneously meet the requirements of a small number of operating points such as takeoff and high-temperature and high-altitude takeoff. The multi-design point model has limitations in terms of iterative variables, number of equilibrium equations, and calculated operating points. It is unable to comprehensively consider the performance requirements of multiple operating points to determine the overall performance plan of the engine.

[0007] 2. The overall engine performance design process requires continuous iteration and calculation or testing of engine cycle parameters and control laws based on various performance requirements. This large number of performance requirements results in high labor costs and engine R&D costs, resulting in low efficiency. Existing technical solutions can ensure more accurate and efficient aircraft engine thermal cycle analysis, but they cannot design control laws for the engine's maximum state based on dozens of operating points.

[0008] 3. The existing model can only calculate a limited number of parameters, and a large number of manual iterations are required when dealing with multiple operating points. The accuracy and range of the overall performance parameter design may not be accurate, which may lead to excessive engine thrust / power and an incomplete overall engine performance solution.

[0009] Therefore, there is an urgent need for a calculation method that can simultaneously calculate and determine engine cycle parameters and control laws at multiple operating points to improve the efficiency of engine overall performance scheme design, optimize the design process, and achieve the goal of replacing manual iteration with computers.

[0010] To address this issue, a multi-design-point thermodynamic cycle analysis method has been proposed for thermodynamic cycle analysis involving multiple performance requirements and multiple constraints in aeroengine design. The construction and solution process of the multi-design-point thermodynamic cycle analysis method are detailed. A single-spool gas generator turboshaft engine with a free-style power turbine is analyzed using both the single-design-point and multi-design-point methods. The resulting design domains are analyzed using both methods. Results show that certain regions within the design domain obtained using the traditional single-design-point method are infeasible due to conflicts between performance requirements and operational constraints. These infeasible regions may contain the optimal design point, rendering the optimal design point infeasible. In the multi-design-point analysis, multiple (or all) operating conditions with performance requirements and operational constraints are used as design points, allowing the optimal design point to be found within the appropriate operating conditions.

[0011] Therefore, how to improve the efficiency and comprehensiveness of the overall performance design of turboshaft and turboprop engines is a problem that needs to be solved. Summary of the Invention

[0012] The purpose of this application is to provide a method for designing an overall performance scheme of a turboshaft-turboprop engine that meets multiple working conditions, so as to solve the problems of low efficiency and incompleteness in the overall performance scheme design of a turboshaft-turboprop engine in the prior art.

[0013] The technical solution of this application is: a method for designing an overall performance scheme of a turboshaft turboprop engine that meets multiple operating conditions, comprising:

[0014] Sort all current operating points, use all current operating points as design points to establish a multi-design point model, and store it in a specific location of the system; use the multi-design point model to design engine cycle parameters and control laws that meet the performance requirements of the previous specific number of operating points;

[0015] Design a replacement control principle, call a non-design point model of the same model engine within the system, input the engine cycle parameters and control law into the non-design point model, the non-design point calculation model selects the next operating point according to the operating point ranking as the operating condition to be checked, operates the operating point in combination with the current engine cycle parameters and control law, obtains the performance parameters of the operating point, and determines whether there are any performance parameters that exceed the limit, then selects the next operating point and proceeds to the next step;

[0016] Determine whether all operating points have been checked; if not, rerun the non-design point calculation model to obtain performance parameters of the newly added operating point and determine whether any performance parameters exceed the limit. If so, replace one of the operating points in the current multi-design point model with the newly added operating point according to the replacement control principle, input the new operating point into the multi-design point model to obtain new engine cycle parameters and control laws, and repeat this step;

[0017] If all operating points are checked, the engine cycle parameters and control rules that meet all operating point combinations are formed based on the current operating point.

[0018] Preferably, the multi-design point model design method includes:

[0019] Set the iteration step size and convergence threshold, select multiple operating points as design points, obtain the component parameters and speed of each design point, select one design point as the aerodynamic design point, obtain the cycle parameters of the aerodynamic design point and the speed and component operating points of the remaining design points, obtain the coupling characteristic diagram of the aerodynamic design point, and calculate the performance residual, component matching residual, and load limit residual of the design point through the residual equation; calculate the performance residual, component matching residual, and load limit residual of the corresponding design point through the residual equation for the remaining design points;

[0020] Obtain the performance requirements, component matching, and load limits for each design point; determine whether the performance residuals, component matching residuals, and load limit residuals of each design point meet the performance requirements, component matching, and load limits, respectively. If so, obtain the performance residuals, component matching residuals, and load limit residuals of each operating point, compare them with the performance residuals, component matching residuals, and load limit residuals of the previous cycle based on the convergence threshold, and determine whether convergence has occurred. If not, modify the engine flow and free turbine output power based on the step size, and iterate using the multivariate Newton method until convergence;

[0021] If so, the multi-design point model is completed.

[0022] Preferably, the performance requirement is determined according to the power requirement of the engine; and the load limit is determined according to the limits of the thermal load, mechanical load, and aerodynamic load of the engine.

[0023] Preferably, when determining whether there are any performance parameters that exceed the limit:

[0024] If so, replace the operating point with the previous specified number of operating points according to the replacement control principle, input them into the multi-design point model to obtain new engine cycle parameters and control laws, and input them into the non-design point calculation model again. Select the next operating point again and input it into the non-design point model to execute the next step;

[0025] If not, the non-design point calculation model selects the next operating point in sequence and inputs it into the non-design point model and executes the next step.

[0026] Preferably, each operating point includes different engine total compression ratios and turbine inlet temperature limit values. By changing the total compression ratios and turbine inlet temperature limit values, different overall performance schemes and corresponding maximum state control laws can be obtained; by changing the engine total compression ratios and turbine inlet temperature limit values, a series of overall performance schemes and corresponding maximum state control laws can be obtained, and the relationship between the cruise fuel consumption rate and the design converted flow rate can be obtained.

[0027] Preferably, after the engine cycle parameters and control laws are designed, the parameter characteristics within the first specific number of operating points can be obtained, including whether it is a throttling operating condition, whether there is a parameter reaching the limit value and the type of parameter reaching the limit value; the parameter type includes the high / low pressure rotor converted speed, the high / low pressure rotor physical speed and the turbine front temperature.

[0028] Preferably, the specific number of operating points is 5 operating points; the performance parameters of the turboshaft / turboprop engine of the dual-rotor gas generator are 5, namely, the physical speed of the high / low pressure rotor, the converted speed of the high / low pressure rotor, and the temperature before the turbine; the performance parameters of the turboshaft / turboprop engine of the single-rotor gas generator are three, namely, the physical speed of the rotor, the converted speed of the rotor, and the temperature before the turbine.

[0029] The overall performance scheme design method of a turboshaft-turboprop engine that meets multiple operating conditions in this application has the following advantages:

[0030] 1. A nonlinear engine equation group containing multiple design points is constructed through aerodynamic and thermodynamic process equations and engine overall matching conditions. The cycle parameters of the aerodynamic design points and the operating points and speeds under other operating conditions are iterated to make the engine meet the requirements of component matching, performance requirements and load limits at multiple design points (20 to 30 points) at the same time. In other words, it is a method for multi-design point modeling and performance matching that meets the requirements of multi-operating point calculation.

[0031] 2. In the process of using the multi-design point model to design the engine cycle parameters and control laws under the new operating condition combination, check in turn whether the physical speed, converted speed and turbine pre-temperature of the remaining operating points under the existing cycle parameters and control laws exceed the limit. If they exceed the limit, use the operating condition to replace the original operating condition in the multi-design point model. In this process, the three replacement principles of the engine operating points are used.

[0032] 3. The process and method for obtaining a series of overall performance schemes and corresponding maximum state control laws by changing the engine total pressure ratio and turbine pre-temperature limit values ​​in multiple design models, that is, a method for multi-design point performance matching design control laws for turboshaft / turboprop engines that meets multi-operating point calculations. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] In order to more clearly illustrate the technical solutions provided by this application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of this application.

[0034] Figure 1 This is a schematic diagram of the overall process of this application;

[0035] Figure 2 This is a schematic diagram of the structure of the dual-rotor gas generator engine of this application;

[0036] Figure 3 This is a schematic diagram of the structure of a single-rotor gas generator engine for this application;

[0037] Figure 4 This is a flow chart of the multi-design point model design method for this application. DETAILED DESCRIPTION

[0038] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. 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 making creative efforts are within the scope of protection of the present invention.

[0039] A method for designing an overall performance scheme of a turboshaft and turboprop engine that meets multiple working conditions is disclosed. Taking a turboprop engine as an example, the invention method is introduced.

[0040] like Figure 1 , including the following steps:

[0041] Step S100: sort all current operating points, establish a multi-design point model using all current operating points as design points, and store the model in a specific location of the system; use the multi-design point model to design engine cycle parameters and control laws that meet the performance requirements of the specified number of operating points;

[0042] Preferably, after the engine cycle parameters and control laws are designed, the parameter characteristics within the first specific number of operating points can be obtained, including whether the engine is in a throttling condition, whether any parameters have reached their limits, and the types of parameters that have reached their limits. A throttling condition is defined as an operating condition in which all limiting parameters (high / low pressure rotor converted speed, high / low pressure rotor physical speed, and turbine inlet temperature) have not reached their limits.

[0043] Preferably, the specific number of operating points is 5 operating points; the performance parameters of the turboshaft / turboprop engine of the dual-spool gas generator are 5, namely, the physical speed of the high / low pressure rotor, the converted speed of the high / low pressure rotor, and the temperature before the turbine; the performance parameters of the turboshaft / turboprop engine of the single-spool gas generator are three, namely, the physical speed of the rotor, the converted speed of the rotor, and the temperature before the turbine. The structure of the dual-spool gas generator is as follows: Figure 2 ; The structure of the single rotor gas generator is as follows Figure 3 .

[0044] like Figure 4 Preferably, the multi-design point model design method includes:

[0045] Set the iteration step size and convergence threshold, select multiple operating points as design points, obtain the component parameters and speed of each design point, select one design point as the aerodynamic design point, obtain the cycle parameters of the aerodynamic design point and the speed and component operating points of the remaining design points, obtain the coupling characteristic diagram of the aerodynamic design point, and calculate the performance residual, component matching residual and load limit residual of the design point through the residual equation; for the remaining design points, calculate the performance residual, component matching residual and load limit residual of the corresponding design point through the residual equation.

[0046] Obtain the performance requirements, component matching, and load limits of each design point; determine whether the performance residuals, component matching residuals, and load limit residuals of each design point meet the performance requirements, component matching, and load limits, respectively. If so, obtain the performance residuals, component matching residuals, and load limit residuals of each operating point, and compare them with the performance residuals, component matching residuals, and load limit residuals of the previous cycle based on the convergence threshold to determine whether convergence has occurred. If not, modify the engine flow and free turbine output power according to the step size, and use the multivariate Newton method to iterate until convergence.

[0047] If so, the multi-design point model is completed.

[0048] Among them, the multi-design point model of the turboprop engine is shown in the following table:

[0049] Table 1 Multi-design point model of turboprop engine

[0050]

[0051] 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 low-pressure compressor pressure ratio; 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 residuals; n 1cor,c is the calculated value of low pressure converted speed, n 2cor,c This is the calculated value of high pressure converted speed.

[0052] The residual equations E1 and E2 for the aerodynamic design point represent the power residual and exhaust velocity residual, respectively, and the iteration variables are the engine flow rate and free turbine output power. The residual equations and iteration variables for the remaining design points are the same as those for the traditional off-design point model. In this model, the engine's total pressure ratio is an input parameter and does not participate in the iteration. The fuel consumption rate at each design point is an output parameter and does not serve as a performance constraint. The residual equations and iteration variables for the remaining design points are the same as those for the traditional off-design point model.

[0053] The multi-design point model contains multiple design points, which can fully consider the distribution of all limiting parameters, so that the design range of the overall performance and control law is large enough, while keeping the number of equations contained in the model as small as possible to improve convergence (too many design points will increase the dimension of the equations and deteriorate the convergence of the model).

[0054] Preferably, the performance requirement is determined according to the power requirement of the engine; and the load limit is determined according to the limits of the thermal load, mechanical load, and aerodynamic load of the engine.

[0055] Step S200: Design a replacement control principle and call a non-design point model for the same engine model within the same system. The non-design point model is the existing model. The engine cycle parameters and control law are input into the non-design point model. The non-design point calculation model selects the next operating point according to the operating point ranking as the operating condition to be checked. The non-design point calculation model operates at this operating point in combination with the current engine cycle parameters and control law, obtains the performance parameters of this operating point, and determines whether any performance parameters exceed the limit.

[0056] If so, replace the operating point with the previous specified number of operating points according to the replacement control principle, input them into the multi-design point model to obtain new engine cycle parameters and control laws, and input them into the non-design point calculation model again. Select the next operating point again and input it into the non-design point model to execute the next step;

[0057] If not, the non-design point calculation model selects the next operating point in sequence and inputs it into the non-design point model and executes the next step.

[0058] Preferably, the replacement control principles include:

[0059] Obtaining the throttling operating conditions in a specific number of operating points in the multi-design point model, and replacing the operating conditions in the throttling state in the specific number of operating points in the multi-design point model with the checked operating conditions;

[0060] If a throttling condition does not exist, the checked condition replaces the condition in which the corresponding over-limit parameter has reached the limit in the specified number of operating points in the multi-design point model. For example, if the turbine inlet temperature in the checked condition exceeds the limit, the condition in which the turbine inlet temperature has reached the limit in the original multi-design point model is replaced.

[0061] If multiple parameters exceed the limits for the operating condition being examined, the parameter with the largest excess is selected and replaced with the corresponding existing operating condition in the multi-design point model. For example, if the turbine inlet temperature exceeds the limit by 2% and the high-pressure speed exceeds the limit by 1%, the existing operating condition where the turbine inlet temperature reaches the limit is replaced.

[0062] Step S300, determine whether all operating points have been checked; if not, run the non-design point calculation model again, obtain the performance parameters of the newly added operating point and determine whether there are any over-limit performance parameters. If so, replace one of the operating points in the current multi-design point model with the newly added operating point according to the replacement control principle, input it into the multi-design point model to obtain new engine cycle parameters and control laws, and repeat this step.

[0063] Step S400: After all operating points have been checked, engine cycle parameters and control rules that meet all operating point combinations are generated according to the current operating point.

[0064] Preferably, each operating point includes different engine overall compression ratio and turbine inlet temperature limits. These are inherent parameters of the operating point and do not participate in the iteration. By varying the overall compression ratio and turbine inlet temperature limits, different overall performance scenarios and corresponding maximum state control laws can be obtained. By referring to the concept of single-design-point cyclic analysis, varying the engine overall compression ratio and turbine inlet temperature limits can yield a range of overall performance scenarios and corresponding maximum state control laws, deriving the relationship between cruise fuel consumption and design-converted flow, and obtaining the results of a multi-design-point cyclic analysis.

[0065] The following is an example to illustrate:

[0066] In a specific embodiment, there are a total of 29 points for the power demand of the full envelope, including ground take-off state, high-altitude take-off state, hot-day take-off, high-altitude high-speed, high-altitude low-speed and other states. The power demand is the demand for the output power of the power turbine, the power extraction is the power extraction of the high-pressure rotor, the bleed air flow rate is the bleed air flow rate of the aircraft, and the bleed air position is the outlet of the high-pressure compressor.

[0067] The design pressure ratio of the low-pressure compressor is 5.0 and the efficiency is 0.84; the design pressure ratio of the high-pressure compressor is 8.0 and the efficiency is 0.85; the combustion efficiency is 0.995 and the total pressure loss is 0.055; the limit of the turbine inlet temperature is taken as 1900K, the efficiency of the high-pressure turbine is 0.88, the efficiency of the low-pressure turbine is 0.89, and the limit of the physical speed and corrected speed of the high / low-pressure compressors is 100%. According to the process shown in Table 4, the multi-design point matching method is used to design the engine cycle parameters and the maximum state control law. The calculation results are shown in Table 4. The bold numbers in the table are the points where the corrected speed and physical speed of the high and low-pressure rotors and the turbine inlet temperature reach the limit.

[0068] Table 4 Calculation Results of the Full Envelope Steady-State Performance of a Certain Type of Engine

[0069]

[0070]

[0071] The exhaust velocity of the turboshaft / turboprop engine is lower than the speed of sound, and the tail nozzle is not critical. Therefore, at different flight Mach numbers, the engines are not similar and the control laws will bifurcate. The bifurcation of the low-pressure physical speed, low-pressure corrected speed and turbine inlet temperature of the engine is more obvious. When the Mach number is relatively large, the turbine inlet temperature is on the low side and the low-pressure rotor speed is on the high side. With the increase of the engine inlet temperature, the control parameters of the maximum state control law are as follows:

[0072] When T2 < 263.7K, control n 1cor = 1.0;

[0073] When 263.7K < T2 < α, control n2 = 100.0%;

[0074] When α < T2 < 331K, control T4 = 1900K.

[0075] α is between 290K and 305K, which is the engine inlet temperature for switching from controlling the maximum of n2 to controlling the maximum of T4. This temperature varies with the flight Mach number, being 290K when the flight Mach number is 0.0 and 305K when the Mach number is 0.7.

[0076] In summary, the present application has the following advantages:

[0077] 1. A nonlinear engine equation group containing multiple design points is constructed through aerodynamic and thermodynamic process equations and engine overall matching conditions. The cycle parameters of the aerodynamic design points and the operating points and speeds under other operating conditions are iterated to make the engine meet the requirements of component matching, performance requirements and load limits at multiple design points (20 to 30 points) at the same time. In other words, it is a method for multi-design point modeling and performance matching that meets the requirements of multi-operating point calculation.

[0078] 2. In the process of using the multi-design point model to design the engine cycle parameters and control laws under the new operating condition combination, check in turn whether the physical speed, converted speed and turbine pre-temperature of the remaining operating points under the existing cycle parameters and control laws exceed the limit. If they exceed the limit, use the operating condition to replace the original operating condition in the multi-design point model. In this process, the three replacement principles of the engine operating points are used.

[0079] 3. The process and method for obtaining a series of overall performance schemes and corresponding maximum state control laws by changing the engine total pressure ratio and turbine pre-temperature limit values ​​in multiple design models, that is, a method for multi-design point performance matching design control laws for turboshaft / turboprop engines that meets multi-operating point calculations.

[0080] 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. Other structures can refer to common designs. In the absence of conflicts, the same embodiment and different embodiments of the present invention can be combined with each other.

[0081] 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 scope of protection of the present invention.

Claims

1. A method for designing an overall performance scheme of a turboshaft and turboprop engine that meets multiple operating conditions, characterized in that: include: Sort all current operating points, use all current operating points as design points to establish a multi-design point model, and store it in a specific location of the system; Calling the multi-design point model to design the engine cycle parameters and control laws that meet the performance requirements of the first specific number of operating points; Design a replacement control principle, call a non-design point model of the same model engine within the system, input the engine cycle parameters and control law into the non-design point model, the non-design point calculation model selects the next operating point according to the operating point ranking as the operating condition to be checked, operates the operating point in combination with the current engine cycle parameters and control law, obtains the performance parameters of the operating point, and determines whether there are any performance parameters that exceed the limit, then selects the next operating point and proceeds to the next step; Determine whether all operating points have been checked; if not, rerun the non-design point calculation model to obtain performance parameters of the newly added operating point and determine whether any performance parameters exceed the limit. If so, replace one of the operating points in the current multi-design point model with the newly added operating point according to the replacement control principle, input the new operating point into the multi-design point model to obtain new engine cycle parameters and control laws, and repeat this step; If all operating points are checked, the engine cycle parameters and control rules that meet all operating point combinations are formed based on the current operating point.

2. The method for designing an overall performance scheme of a turboshaft-turboprop engine that meets multiple operating conditions according to claim 1, characterized in that: The multi-design point model design method comprises: Set the iteration step size and convergence threshold, select multiple operating points as design points, obtain the component parameters and speed of each design point, select one design point as the aerodynamic design point, obtain the cycle parameters of the aerodynamic design point and the speed and component operating points of the remaining design points, obtain the coupling characteristic diagram of the aerodynamic design point, and calculate the performance residual, component matching residual, and load limit residual of the design point through the residual equation; calculate the performance residual, component matching residual, and load limit residual of the corresponding design point through the residual equation for the remaining design points; Obtain the performance requirements, component matching, and load limits for each design point; determine whether the performance residuals, component matching residuals, and load limit residuals of each design point meet the performance requirements, component matching, and load limits, respectively. If so, obtain the performance residuals, component matching residuals, and load limit residuals of each operating point, compare them with the performance residuals, component matching residuals, and load limit residuals of the previous cycle based on the convergence threshold, and determine whether convergence has occurred. If not, modify the engine flow and free turbine output power based on the step size, and iterate using the multivariate Newton method until convergence; If so, the multi-design point model is completed.

3. The method for designing an overall performance solution for a turboshaft and turboprop engine that meets multiple operating conditions according to claim 2, wherein: The performance requirement is determined according to the power requirement of the engine; the load limit is determined according to the limits of the thermal load, mechanical load, and aerodynamic load of the engine.

4. The method for designing a turboshaft-turboprop engine overall performance scheme that meets multiple operating conditions according to claim 1, wherein: When determining whether there are performance parameters that exceed the limit: If so, replace the operating point with the previous specified number of operating points according to the replacement control principle, input them into the multi-design point model to obtain new engine cycle parameters and control laws, and input them into the non-design point calculation model again. Select the next operating point again and input it into the non-design point model to execute the next step; If not, the non-design point calculation model selects the next operating point in sequence and inputs it into the non-design point model and executes the next step.

5. The method for designing an overall performance solution for a turboshaft and turboprop engine that meets multiple operating conditions according to claim 1, wherein: Each operating point contains different engine total compression ratio and turbine inlet temperature limit values. By changing the total compression ratio and turbine inlet temperature limit values, different overall performance schemes and corresponding maximum state control laws can be obtained; by changing the engine total compression ratio and turbine inlet temperature limit values, a series of overall performance schemes and corresponding maximum state control laws can be obtained, and the relationship between the cruise fuel consumption rate and the design converted flow rate can be obtained.

6. The method for designing an overall performance scheme of a turboshaft and turboprop engine that meets multiple operating conditions according to claim 1, characterized in that: After the engine cycle parameters and control laws are designed, the parameter characteristics within the first specific number of operating points can be obtained, including whether it is a throttling operating condition, whether there is a parameter reaching the limit value and the type of parameter reaching the limit value; the parameter type includes the high / low pressure rotor converted speed, the high / low pressure rotor physical speed and the turbine front temperature.

7. The method for designing an overall performance solution for a turboshaft and turboprop engine that meets multiple operating conditions according to claim 1, wherein: The specific number of operating points is 5 operating points; the performance parameters of the turboshaft / turboprop engine of the dual-spool gas generator are 5, namely the physical speed of the high / low pressure rotor, the converted speed of the high / low pressure rotor, and the turbine front temperature; There are three performance parameters for a turboshaft / turboprop engine with a single-rotor gas generator, namely the rotor physical speed, the rotor converted speed and the turbine inlet temperature.

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