A method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements.

Abstract

This application discloses a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements. This method, considering the coupling effects of cycle parameters with the parameters of various components and the air system, establishes the relationship between thermodynamic cycle parameters and the cost of each component and the entire engine. Furthermore, by balancing the aircraft's mission and weight, it obtains the distribution of engine cost, engine and fuel weight coefficients with respect to cycle parameters. This provides a reliable scientific basis for evaluating and further optimizing thermodynamic cycle parameters. It can be used to assess the applicability of different combinations of thermodynamic cycle parameters to the flight platform, making the engine cost more advantageous while maintaining the minimum residual weight coefficient of the flight platform within an acceptable range, further improving the flight platform's range.

Description

Technical Field

[0001] This application relates to the field of aero-engine technology, and in particular to a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements. Background Technology

[0002] As a power-output engine, the turboshaft engine is the primary power plant for platforms such as helicopters and tiltrotor aircraft. In the civilian market, the economics of an aircraft platform have become a crucial factor determining its market competitiveness. The manufacturing and periodic overhaul costs of turboshaft engines, along with fuel costs, constitute the main components of the direct operating costs of an aircraft platform, accounting for over 60% of the total. Therefore, reducing these related costs of turboshaft engines is one of the main ways to improve the economics of aircraft platforms and has become a requirement that must be considered in the design of turboshaft engines.

[0003] The thermodynamic cycle parameters of a turboshaft engine include the combustion chamber outlet temperature and the compressor pressure ratio. While meeting the target flight platform's power requirements, increasing cycle parameters such as combustion chamber outlet temperature and pressure ratio can reduce engine weight and improve fuel efficiency, but this increases material and manufacturing costs, and may result in over- or under-design relative to the platform's mission requirements. Selecting appropriate thermodynamic cycle parameters based on the flight platform's mission requirements, current industrial technology, and cost status can effectively control engine manufacturing and periodic overhaul costs, as well as fuel costs, and can also provide a good foundation for the specific design of various engine components. However, currently published thermodynamic cycle parameter design methods mainly focus on performance optimization, with less comprehensive consideration of cost and mission requirements. They fail to assess the impact of thermodynamic cycle parameters on engine cost, and the selection of thermodynamic cycle parameters during the design process rarely considers weight distribution and power spectrum under different flight platform conditions. To fully consider the balance between engine weight and fuel consumption, over- or under-design may occur, failing to meet the economic requirements of the flight platform. Summary of the Invention

[0004] This application provides a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, which addresses the technical problem that existing thermodynamic cycle parameter designs do not adequately consider cost and mission requirements, and therefore do not meet the economic requirements of flight platforms.

[0005] This application is achieved through the following solution:

[0006] A method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, including the following steps:

[0007] S1. Establish the engine benchmark scheme and its design point performance model, and determine the cost ratio of each component;

[0008] S2. Determine the compressor structure scheme applicable to different pressure ratios, establish the cost calculation formula for compressor rotor blade disk and compressor stator blade assembly, and then establish the compressor total cost coefficient calculation formula by combining the cost ratio relationship of the engine benchmark scheme.

[0009] S3. To address the differences in process cost, material cost, and yield rate among gas turbine structural schemes for different gas temperature ranges, establish a calculation formula for the total cost coefficient of turbines for different gas temperature ranges.

[0010] S4. Based on the materials, cooling structure, and coating scheme of the combustion chamber flame tube, the outer diameter and wall thickness of the combustion chamber casing, the number of nozzles, and the inlet conversion flow rate, establish the formula for calculating the total cost coefficient of the combustion chamber;

[0011] S5. Establish a formula for calculating the total cost coefficient of the intake and exhaust systems based on the material cost ratio of the intake and exhaust systems in the engine baseline design and the converted flow rate of the compressor inlet.

[0012] S6. Establish a formula for calculating the engine manufacturing and periodic overhaul cost coefficient based on the compressor total cost coefficient, the compressor cost ratio of the engine benchmark scheme, the gas turbine total cost coefficient, the gas turbine cost ratio of the engine benchmark scheme, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust device total cost coefficient.

[0013] S7. Based on the engine manufacturing and periodic overhaul cost coefficient and fuel consumption rate, and combined with the relationship between fuel cost and engine usage cost, establish the engine usage cost coefficient calculation formula.

[0014] S8. Based on the weight distribution and power spectrum of the flight platform under different mission conditions, calculate the remaining weight, engine fuel requirement weight, and minimum remaining weight coefficient under different mission conditions, and complete the applicability assessment of the thermodynamic cycle parameters relative to the flight platform mission.

[0015] S9. Substitute the coefficients or formulas obtained in steps S2 to S8 into the design point performance model, give the range and step size of the changes in combustion chamber outlet temperature and pressure ratio, and carry out iterative calculations to obtain the distribution of engine operating cost coefficient and flight platform minimum remaining weight coefficient on the thermodynamic cycle parameter analysis diagram.

[0016] Compared with the prior art, this application has the following advantages:

[0017] This application proposes a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements. This method, considering the coupled influence of cycle parameters with the parameters of various components and the air system, establishes the relationship between thermodynamic cycle parameters and the cost of each component and the entire engine. Furthermore, by balancing the aircraft's mission and weight, it obtains the distribution of engine cost and engine and fuel weight coefficients with respect to cycle parameters. This provides a feasible method for evaluating and optimizing thermodynamic cycle parameters, specifically including:

[0018] 1) This application utilizes the component technology solutions and cost levels of existing engine products to determine the relationship between parameters such as combustion chamber outlet temperature, pressure ratio, and flow rate and the main aerodynamic component solutions, establishes a model of the relationship between thermodynamic cycle parameters and component costs, and analyzes the comprehensive impact of changes in thermodynamic cycle parameters on manufacturing and periodic overhaul costs and fuel costs, providing a basis for further optimization of thermodynamic cycle parameters;

[0019] 2) Based on the weight distribution and power spectrum of the aircraft under different conditions, this application establishes the minimum residual weight coefficient of the flight platform in combination with the influence of thermodynamic cycle parameters on engine weight and fuel consumption rate. This coefficient is used to characterize the influence of different thermodynamic cycle parameters on the mission capability margin of the flight platform and to evaluate the applicability of different combinations of thermodynamic cycle parameters to the flight platform. This makes the engine cost more advantageous, and the minimum residual weight coefficient of the flight platform is still within an acceptable range, further improving the range of the flight platform.

[0020] In addition to the purposes, features, and advantages described above, this application has other purposes, features, and advantages. A further detailed description of this application will be provided below with reference to the figures. Attached Figure Description

[0021] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments consistent with this application and, together with the description, serve to explain the principles of this application.

[0022] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, for those skilled in the art, other drawings can be obtained based on these drawings without creative effort, wherein:

[0023] Figure 1 This is a schematic diagram of the existing thermodynamic cycle parameter analysis;

[0024] Figure 2 This is a schematic diagram of a typical turboshaft engine layout;

[0025] Figure 3This is a schematic diagram of the process for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, according to a preferred embodiment of this application.

[0026] Figure 4 This is a schematic diagram showing the conversion between the compressor inlet outer diameter and the compressor inlet flow rate;

[0027] Figure 5 This is a schematic diagram showing the relationship between the turbine bleed air ratio and the combustion chamber outlet temperature.

[0028] Figure 6 This is a schematic diagram showing the relationship between the engine weight coefficient and the compressor equivalent flow rate;

[0029] Figure 7 This is a schematic diagram showing the distribution of the engine operating cost coefficient on the thermodynamic cycle parameter analysis diagram;

[0030] Figure 8 This is a schematic diagram of the distribution of the minimum residual weight coefficient of the flight platform on the thermodynamic cycle parameter analysis diagram;

[0031] Figure 9 This is a schematic diagram of a turboshaft engine thermodynamic cycle parameter evaluation device module based on cost and mission requirements according to a preferred embodiment of this application;

[0032] Figure 10 This is a schematic block diagram of an electronic device according to a preferred embodiment of this application;

[0033] Figure 11 This is an internal structural diagram of a computer device according to a preferred embodiment of this application.

[0034] In the diagram: 1. Intake device; 2. Multi-stage axial flow; 3. Centrifugal impeller; 4. Combustion chamber; 5. Gas turbine; 6. Power turbine; 7. Exhaust device. Detailed Implementation

[0035] It should be understood that the specific embodiments described herein are merely illustrative of the technical solutions of this application and are not intended to limit this application.

[0036] To better understand the technical solution of this application, a detailed description will be provided below in conjunction with the accompanying drawings and specific implementation methods.

[0037] Existing turboshaft engine packages typically include components such as an intake system, compressor, combustion chamber, gas turbine, power turbine, and exhaust system. The gas turbine and compressor are connected on the same drive shaft, and the combustion chamber is located between the gas turbine and compressor. Air is continuously drawn in by the compressor through the intake system, compressed, and then enters the combustion chamber to burn with injected fuel to become high-temperature gas. This gas then enters the gas turbine, expands, and does work. At this time, the expansion work is transmitted to the compressor through the drive shaft, thus enabling the engine to operate continuously. The gas exiting the gas turbine drives the power turbine to continue expanding and doing work to drive other equipment, such as helicopter rotors and generators. The gas exiting the power turbine enters the atmosphere through the exhaust system.

[0038] Parameters affecting engine performance include the total pressure loss coefficient of the given intake system, compressor equivalent flow rate, compressor efficiency, total pressure loss of the combustion chamber, gas turbine inlet temperature, gas turbine efficiency, power turbine efficiency, total pressure loss in the transition section, exhaust losses, and bleed air ratio in the air system. In conventional overall performance design, considering a defined power target and the performance levels of components, parameter analysis of the combustion chamber outlet temperature and compressor pressure ratio is conducted to create a "carpet map." Based on constraints related to performance and structural strength, the feasible region is determined to obtain suitable thermodynamic cycle parameters. (See...) Figure 1 Among them, fuel consumption rate reflects the engine's fuel consumption level; power per unit volume refers to the power corresponding to a unit airflow, which can characterize the engine's weight and size.

[0039] To improve the rationality of turboshaft engine performance design, existing technologies have explored the coupling relationships between parameters of various components and the air system to allow for more precise selection of thermodynamic cycle parameters. However, these technologies focus primarily on performance optimization, with less consideration for engine cost. The compressor, combustion chamber, gas turbine, power turbine, and intake and exhaust systems are the main aerodynamic components of a turboshaft engine. Their specific design is closely related to parameters such as the temperature, pressure, and flow rate of the engine's main gas stream, accounting for over 65% of the engine's total cost. While maintaining engine power and lifespan, the combustion chamber outlet temperature and pressure ratio can lead to differences in the structure, materials, and manufacturing processes of these components, directly impacting the engine's manufacturing and periodic overhaul costs. In the context of technological development and a fixed overhaul cycle, periodic overhaul costs can be considered directly related to manufacturing costs. Furthermore, due to the coupling relationships between the parameters of various components and the air system, excessively pursuing lower engine weight when selecting thermodynamic cycle parameters may lead to increased fuel consumption, negatively impacting the flight platform's mission capabilities. Therefore, achieving a balance between engine weight and fuel consumption is an effective means of improving flight platform capabilities.

[0040] Besides thermodynamic cycle parameters, the diversity of component configurations also affects engine costs. However, with technological advancements, engines of the same power rating have gradually converged in terms of overall layout and component configuration selection. For example, considering the engine's power expansion potential, turboshaft engines in the 1000kW–2000kW power range often adopt a combined compressor configuration of "multi-stage axial flow + centrifugal impeller" and component configurations such as a two-stage gas turbine (see...). Figure 2 The system includes: intake device 1, multi-stage axial flow 2, centrifugal impeller 3, combustion chamber 4, gas turbine 5, power turbine 6, and exhaust device 7, which provides favorable conditions for cost assessment.

[0041] It should be noted that the executing entity in this embodiment can be a computing service device with data processing, network communication, and program execution functions, such as a tablet computer, personal computer, or mobile phone, or a turboshaft engine thermodynamic cycle parameter evaluation device based on cost and task requirements capable of performing the above functions. The following description uses a turboshaft engine thermodynamic cycle parameter evaluation device based on cost and task requirements as an example to illustrate this embodiment and the subsequent embodiments.

[0042] like Figure 1 As shown, a preferred embodiment of this application provides a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, including the following steps:

[0043] S1. Establish the engine benchmark scheme and its design point performance model, and determine the cost ratio of each component;

[0044] S2. Determine the compressor structure scheme applicable to different pressure ratios, establish the cost calculation formula for compressor rotor blade disk and compressor stator blade assembly, and then establish the compressor total cost coefficient calculation formula by combining the cost ratio relationship of the engine benchmark scheme.

[0045] S3. To address the differences in process cost, material cost, and yield rate among gas turbine structural schemes for different gas temperature ranges, establish a calculation formula for the total cost coefficient of turbines for different gas temperature ranges.

[0046] S4. Based on the materials, cooling structure, and coating scheme of the combustion chamber flame tube, the outer diameter and wall thickness of the combustion chamber casing, the number of nozzles, and the inlet conversion flow rate, establish the formula for calculating the total cost coefficient of the combustion chamber;

[0047] S5. Establish a formula for calculating the total cost coefficient of the intake and exhaust systems based on the material cost ratio of the intake and exhaust systems in the engine baseline design and the converted flow rate of the compressor inlet.

[0048] S6. Establish a formula for calculating the engine manufacturing and periodic overhaul cost coefficient based on the compressor total cost coefficient, the compressor cost ratio of the engine benchmark scheme, the gas turbine total cost coefficient, the gas turbine cost ratio of the engine benchmark scheme, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust device total cost coefficient.

[0049] S7. Based on the engine manufacturing and periodic overhaul cost coefficient and fuel consumption rate, and combined with the relationship between fuel cost and engine usage cost, establish the engine usage cost coefficient calculation formula.

[0050] S8. Based on the weight distribution and power spectrum of the flight platform under different mission conditions, calculate the remaining weight, engine fuel requirement weight, and minimum remaining weight coefficient under different mission conditions, and complete the applicability assessment of the thermodynamic cycle parameters relative to the flight platform mission.

[0051] S9. Substitute the coefficients or formulas obtained in steps S2 to S8 into the design point performance model, give the range and step size of the changes in combustion chamber outlet temperature and pressure ratio, and carry out iterative calculations to obtain the distribution of engine operating cost coefficient and flight platform minimum remaining weight coefficient on the thermodynamic cycle parameter analysis diagram.

[0052] Compared with the prior art, this embodiment has the following beneficial effects:

[0053] This embodiment proposes a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements. Considering the coupled influence of cycle parameters with the parameters of various components and the air system, it establishes the relationship between thermodynamic cycle parameters and the cost of each component and the entire engine. Furthermore, by balancing the aircraft's mission and weight, it obtains the distribution of engine cost, engine and fuel weight coefficients with respect to cycle parameters. This provides a feasible method for evaluating and optimizing thermodynamic cycle parameters, specifically including:

[0054] 1) This embodiment utilizes the component technology solutions and cost levels of existing engine products to determine the relationship between parameters such as combustion chamber outlet temperature, pressure ratio, and flow rate and the main aerodynamic component solutions. It establishes a model relating thermodynamic cycle parameters to component costs and analyzes the comprehensive impact of changes in thermodynamic cycle parameters on manufacturing and periodic overhaul costs, as well as fuel costs, providing a basis for further optimization of thermodynamic cycle parameters.

[0055] 2) Based on the weight distribution and power spectrum of the aircraft under different conditions, this embodiment combines the influence of thermodynamic cycle parameters on engine weight and fuel consumption rate to establish the minimum residual weight coefficient of the flight platform. This coefficient characterizes the influence of different thermodynamic cycle parameters on the mission capability margin of the flight platform and is used to evaluate the applicability of different combinations of thermodynamic cycle parameters to the flight platform. This makes the engine cost more advantageous, and the minimum residual weight coefficient of the flight platform is still within an acceptable range, further improving the range of the flight platform.

[0056] Preferably, step S1 specifically includes the following steps:

[0057] S11. Use existing methods to complete the thermodynamic cycle parameter optimization analysis of the engine design point, so that the power, fuel consumption rate and other indicators of the engine design point meet the index requirements, and establish the engine benchmark scheme and its design point performance model.

[0058] S12. Based on the existing engine benchmark scheme and cost situation, and in conjunction with the supplier inquiry results, determine the proportion of manufacturing cost of each component, as shown in Table 1.

[0059] Table 1. Proportion of Manufacturing Cost for Each Component

[0060] ;

[0061] The main aerodynamic components of a compressor include the compressor rotor blade disk and the compressor stator blade assembly, which together account for more than 65% of the compressor cost. Given a specific configuration and technological level, the stage pressure ratio of a compressor is limited.

[0062] To accommodate the increased compressor pressure ratio, the number of axial stages in the combined compressor must be increased. This increases the temperature of the final axial stage or centrifugal stage, necessitating the use of high-temperature alloys in the centrifugal stage, thus increasing costs. Besides pressure ratio, engine flow rate and efficiency also affect cost. Increased flow rate leads to an increase in compressor outer diameter, while decreased efficiency results in a corresponding increase in temperature of the final axial stage or centrifugal stage. Furthermore, the manufacturing costs of other compressor components, including the outer casing of the axial and centrifugal sections, are positively correlated with the compressor flow path components. Therefore, by obtaining the cost changes of the compressor rotor blade disk and compressor stator blade assembly, the overall cost change of the compressor can be largely determined. Therefore, preferably, step S2 specifically includes the following steps:

[0063] S21. By statistically analyzing the compressor data of existing engines, the compressor configurations suitable for different pressure ratio ranges can be obtained, and the compressor structure schemes suitable for different pressure ratios can be determined. Table 2 shows the combined compressor schemes suitable for different pressure ratio ranges. In the table, pressure ratio range 1 < pressure ratio range 2 < pressure ratio range 3 < pressure ratio range 4. Due to similar processes, the total cost of blade disks is directly proportional to the number of blade disks while keeping the size and materials unchanged. To improve the reliability and economy of the product, except for the first-stage blade disk and centrifugal impeller, the intermediate axial flow stages should use two-stage integral blade disks as much as possible to reduce the number of blade disks. Based on the statistical results of existing engines, the relationship between the number of compressor stages, the number of blade disks, the number of blade disks made of different materials, the number of stator stages, and the pressure ratio can be obtained for different pressure ratio ranges.

[0064] Table 2. Combination compressor structural schemes applicable to different pressure ratio ranges

[0065] ;

[0066] S22. The compressor inlet outer diameter determines the compressor inlet frontal area, which is a key factor affecting the compressor inlet converted flow rate. Under the same technical level, the relationship between the compressor inlet outer diameter and the compressor inlet converted flow rate is established based on existing engine data (see...). Figure 4 ):

[0067] D com =a1×Wac2 2 +b1×Wac2+c1;

[0068] Among them, D com Where is the compressor inlet outer diameter; Wac2 is the compressor inlet converted flow rate; a1, b1, and c1 are coefficients determined based on curve fitting results;

[0069] S23. For compressor rotor blade disks, based on changes in compressor outer diameter, number of blade disks, and blade disk material, establish a formula for calculating the compressor blade disk cost coefficient, thereby obtaining the changes in the new compressor blade disk scheme relative to the engine baseline scheme:

[0070] ;

[0071] Where: ζ comp,disc The compressor rotor blade disk cost coefficient represents the cost ratio of the new compressor blade disk design relative to the engine baseline design; α comp,dis,mat,0 The material cost ratio of the compressor rotor blade disk in the benchmark engine design is generally 20% to 35%; D com D is the outer diameter of the compressor inlet. com,0 The compressor inlet outer diameter is the reference diameter for the engine design; N comp,disc N represents the number of compressor rotor blade disks. comp,disc,0 The number of compressor rotor blade disks in the engine baseline design; N comp,dis,GH N represents the number of high-temperature alloy rotor blade disks in the compressor. comp,dis,GH,0 The number of high-temperature alloy rotor blade disks for the compressor in the engine baseline design; C mat,GH / C mat,Ti The cost ratio of high-temperature alloys to titanium alloys needs to take into account the differences in material density and unit price;

[0072] S24. For compressor stator blade assemblies, the cost of intermediate stage stators is relatively stable, while increasing the number of stages will correspondingly increase the number of intermediate stage stators. Therefore, based on the influence of changes in the number of stages and outer diameter of compressor stator blade assemblies, a formula for calculating the cost coefficient of compressor stator blade assemblies is established:

[0073] ζcomp,stator =(1-α comp,stator,mat,0 +α com,stator,mat,0 ×D com 2 / D com,0 2 )×

[0074] (1+ (N comp - N comp,0 )×C comp,sator,mid,0 / C comp,sator, 0 );

[0075] Where: ζ comp,stator α represents the cost coefficient of the compressor stator blade assembly, i.e., the cost ratio of the new design to the engine baseline design; comp,stator,mat,0 The material cost ratio of the compressor stator blade assembly in the engine baseline design; D com D is the outer diameter of the compressor inlet. com,0 The compressor inlet outer diameter is the reference diameter for the engine design; N comp N represents the number of compressor stages. comp,0 The number of compressor stages for the engine baseline design; C comp,sator,mid,0 / C comp,sator, 0 This represents the ratio of the cost of a single intermediate stage stator to the total cost of the stator blade assembly, and is typically between 15% and 25%.

[0076] S25. Using the cost coefficients of the compressor rotor blade disk and compressor stator blade assembly, and combining the cost ratio of the engine benchmark scheme, establish a formula for calculating the total cost coefficient of the compressor:

[0077] ζ comp = C comp,disc,0 / (C comp,sator, 0 + C comp,disc, 0 )×ζ comp,disc + C comp,sator,0 /

[0078] (C comp,sator, 0 + C comp,disc, 0 )×ζ comp,stato ;

[0079] Where: ζ comp C is the total cost coefficient for the compressor. comp,disc,0 Cost of compressor rotor blade disk for engine baseline design; C comp,sator, 0 Cost of compressor stator blade assembly for engine baseline design; ζ comp,disc Cost coefficient ζ for compressor rotor blade disk comp,stato This is the cost coefficient for the compressor stator blade assembly.

[0080] The main aerodynamic components of a turbine include the turbine guide vane, turbine blades, and turbine disk, accounting for over 75% of the turbine cost. Given a specific configuration and technological level, increasing the gas combustion temperature requires the use of more advanced high-temperature alloy materials and more complex cooling structures, or increasing the bleed air volume for cooling. Simultaneously, changes in combustion chamber outlet temperature, compressor pressure ratio, component efficiency, and the air system will affect the converted flow rate at the gas turbine rotor inlet. The converted flow rate directly reflects the gas turbine's flow capacity and flow channel area requirements, thus determining the turbine's basic dimensions. However, for small and medium-sized turboshaft engines, reducing the turbine size increases the manufacturing difficulty of the cooling structure, resulting in some cost offsetting. Therefore, when analyzing turbine costs, the influence of gas combustion temperature can be primarily considered. Therefore, preferably, step S3 specifically includes the following steps:

[0081] S31. Using existing engine technology, establish the relationship between the proportion of bleed air flow in each part of the turbine to the engine inlet flow and the combustion chamber outlet temperature, including the proportion of bleed air from the combustion engine to the turbine, the proportion of bleed air from the combustion engine to the turbine, and the proportion of bleed air from the power turbine. Figure 5 The design results for the first-stage guide air ratio, the combustion turbine bleed air ratio, and the power turbine bleed air ratio are given as a function of the combustion chamber outlet temperature. The first-stage guide air ratio refers to the bleed air ratio at the first-stage guide vane of the gas turbine, the combustion turbine bleed air ratio refers to the bleed air ratio at the working blades of the gas turbine (including the second-stage guide vane), and the power turbine bleed air ratio refers to the bleed air ratio at the power turbine. Specifically, the formula for calculating the first-stage guide air ratio is:

[0082] CO NGV = a2×T4 3 +b2×T4 2 +c2×T4+d2;

[0083] Among them, CO NGV T1 is the ratio of induced draft gas to induced draft gas; T4 is the combustion chamber outlet temperature; a2, b2, c2, and d2 are coefficients determined based on curve fitting results.

[0084] The formula for calculating the combustion-driven bleed air ratio is:

[0085] CO HPTR = a3×T4 3 +b3×T4 2 +c3×T4+d3;

[0086] Among them, CO HPTR T4 is the combustion chamber outlet temperature; a3, b3, c3, and d3 are coefficients determined based on curve fitting results.

[0087] The formula for calculating the bleed air ratio of the power turbine is:

[0088] CO LPTR = a4×T4 3 +b4×T4 2 +c4×T4+d4;

[0089] Among them, CO LPTR T4 is the bleed air ratio of the power turbine; T4 is the combustion chamber outlet temperature; a4, b4, c4, and d4 are coefficients determined based on curve fitting results.

[0090] S32. Referring to existing engines, establish the relationship between the gas temperature range and the material level and cooling structure scheme of the main aerodynamic components of the turbine, and determine the structural schemes applicable to different gas inlet temperatures of the gas turbine and the power turbine. Here, the gas inlet temperature of the gas turbine refers to the combustion chamber outlet temperature, and the gas inlet temperature of the power turbine refers to the gas turbine outlet temperature.

[0091] For gas turbines, the inlet gas temperature refers to the combustion chamber outlet temperature; for power turbines, it refers to the gas turbine outlet temperature. The turbine inlet gas temperature has a crucial impact on the material quality and cooling structure of the turbine's main aerodynamic components. Under the same strength and lifespan requirements, an increase in turbine inlet gas temperature necessitates the use of materials with stronger high-temperature and corrosion resistance. For example, the high-temperature alloys used in turbine blades may be upgraded from directionally solidified to first-generation, second-generation, or third-generation single-crystal alloys. Simultaneously, adjustments to the blade cooling structure may be required, replacing solid blades with hollow blades and further employing hollow blades with more complex internal cooling channels. This could involve upgrading simple direct-cooling convection heat transfer structures to multi-channel convection heat transfer structures with turbulence design, and potentially adding impingement, film cooling, or slotted cooling structures, or even more complex laminated cooling systems, and considering the use of thermal barrier coatings. Both the use of materials with stronger high-temperature and corrosion resistance and the selection of blades with more complex cooling structures will increase engine costs. Table 3 establishes the relationship between the gas temperature range and the material quality and cooling structure of the turbine's main aerodynamic components, referencing existing engines. Table 3 presents gas turbine structural schemes for different gas temperature ranges. The relationship between power turbine structural schemes and gas temperature can also be established with reference to this table. In Table 3, temperature range 1 < temperature range 2 < temperature range 3 < temperature range 4 < temperature range 5. Taking temperature range 1 as the baseline scheme, the material level and cooling structure complexity of different parts in the baseline scheme are represented by A, while B, C, and D represent increases in material level and cooling structure complexity.

[0092] Table 3 Gas turbine structural schemes for different gas temperature ranges

[0093] ;

[0094] S33. To address the differences in process cost, material cost, and yield rate among different turbine designs, establish a formula for calculating the total turbine cost coefficient for different gas temperature ranges:

[0095] ζ HPT = ∑(α HPT,k,0 × ((1-α HPT, mat,k,0 ) ×C HPT,process,k / C HPT,process,0,k +α HPT, mat,k,0 ×

[0096] C HPT,mat, k / C HPT,mat, 0,k )×Q HPT,k / Q HPT,k,0 );

[0097] Where: ζ HPT The total cost coefficient for the gas turbine is the cost ratio of the new design to the benchmark engine design; α HPT,k,0 The cost breakdown of each component of the gas turbine in the benchmark engine design; α HPT, mat,k,0 The material cost ratio of the corresponding part of the gas turbine in the engine benchmark scheme; C HPT,process,k / C HPT,process,0,k The ratio of gas turbine manufacturing costs to the benchmark engine design can be determined based on statistical data; C HPT,mat, k / C HPT,mat, 0,k The ratio of gas turbine material costs to the benchmark engine design can be determined based on statistical data; Q HPT,k / Q HPT,k,0 The ratio of the gas turbine yield rate to the engine baseline can be determined based on statistical data;

[0098] Power turbines typically employ uncooled blades, which significantly impacts material costs. Calculations show that the total turbine cost coefficient varies across different gas temperature ranges, as illustrated in Tables 4 and 5.

[0099] Table 4. Total Cost Coefficient of Gas Turbines for Different Gas Temperature Ranges

[0100]

[0101] Table 5. Total Cost Coefficient of Power Turbines in Different Gas Temperature Ranges

[0102]

[0103] Preferably, similar to a gas turbine, the materials, cooling structure, and coating scheme used in the combustion chamber flame tube section are closely related to the combustion chamber outlet temperature; similar to a compressor, the outer diameter and wall thickness of the combustion chamber casing section, and the number of nozzles, are affected by the inlet converted flow rate. The combustion chamber casing and flame tube section account for more than 95% of the total cost of the combustion chamber; therefore, referring to the methods for compressors and gas turbines, step S4 specifically includes the following steps:

[0104] S41. Based on the materials, cooling structure, and coating scheme of the combustion chamber flame tube, the outer diameter and wall thickness of the combustion chamber casing, the number of nozzles, and the inlet converted flow rate, establish the formula for calculating the total cost coefficient of the combustion chamber:

[0105] ζ comb =f(T4,α comb,k,0 α comb, mat,k,0 C comb,process,k / C comb,process,0,k C comb,mat, k /

[0106] C comb,mat, 0,k ,Wac 31 / Wac 31,0 ,PI,Q comb,k / Q comb,k,0 );

[0107] Where: ζ comb T4 is the total cost coefficient of the combustion chamber, i.e., the cost ratio of the new scheme to the engine baseline scheme; T4 is the combustion chamber outlet temperature; α comb,k,0 The cost breakdown of each component of the combustion chamber in the engine baseline design; α comb,mat,k,0 The material cost ratio of each part of the combustion chamber in the engine baseline design; C comb,process,k / C comb,process,0,k C represents the ratio of the manufacturing costs of each part of the combustion chamber to the benchmark engine design; comb,mat,k / C comb,mat,0,k The ratio of material costs for each part of the combustion chamber to the engine baseline design; Wac 31 Calculated flow rate for combustion chamber inlet; Wac 31,0 For the engine baseline design, the converted flow rate at the combustion chamber inlet is PI, where PI represents the compressor pressure ratio; Q comb,k / Q comb,k,0 This represents the ratio of the yield rate of each part of the combustion chamber to the engine baseline.

[0108] Preferably, the cost of the intake and exhaust devices is mainly affected by their size and is closely related to the compressor inlet conversion. Referring to the compressor method, step S5 specifically includes the following steps:

[0109] S51. Based on the material cost ratio of the intake and exhaust systems in the engine baseline design and the converted flow rate of the compressor inlet, establish a formula for calculating the total cost coefficient of the intake and exhaust systems:

[0110] ζ inout =f(α inout,mat ,Wac2);

[0111] Where: ζ inout The total cost coefficient for the combustion chamber is the cost ratio of the new design relative to the benchmark engine design; α inout,mat The ratio of material costs for the intake and exhaust systems in the engine baseline design; Wac2 is the converted flow rate at the compressor inlet.

[0112] Preferably, step S6 specifically includes the following steps:

[0113] S61. Based on the compressor total cost coefficient, the compressor cost ratio of the engine baseline design, the gas turbine total cost coefficient, the gas turbine cost ratio of the engine baseline design, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust system total cost coefficient, establish the formula for calculating the engine manufacturing and periodic overhaul cost coefficient:

[0114] ζ manu =ζ comp ×α comp,0 +ζ HPT ×α HPT,0 +ζ LPT ×α LP,0T +ζ comb ×α comb,0 +ζ inout ×α inout,0 +

[0115] (1-α comp,0 -α HPT,0 -α LPT,0 -α comb,0 -α inout,0 );

[0116] Where: ζ manu For engine manufacturing and periodic overhaul cost coefficients; ζ comp α is the total cost coefficient for the compressor; comp,0 The compressor cost percentage for the benchmark engine design; ζ HPT α is the total cost coefficient for gas turbines. HPT,0 The cost percentage of the gas turbine in the engine benchmark solution; ζ LPT α represents the total cost coefficient of the power turbine. LPT,0 The cost percentage of the power turbine in the benchmark engine design; ζ comb α is the total cost coefficient for the combustion chamber. comb,0 The combustion chamber cost percentage of the engine's benchmark design; ζ inoutα represents the total cost coefficient for the intake and exhaust systems. inout,0 The cost percentage of the intake and exhaust systems in the engine baseline design.

[0117] Preferably, step S7 specifically includes the following steps:

[0118] S71. Based on the engine manufacturing and periodic overhaul cost coefficients and fuel consumption rate, and considering the relationship between fuel costs and engine operating costs, establish a formula for calculating the engine operating cost coefficient:

[0119] ζ eng =ζ manu ×α manu +SFC / SFC0×α SFC +(1-α manu -α SFC );

[0120] Where: ζ eng For engine operating cost coefficient; ζ manu α represents the cost coefficient for engine manufacturing and periodic overhauls. manu The ratio of manufacturing and periodic overhaul costs to engine operating costs is typically 40%–50%; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline design point fuel consumption rate; α SFC This refers to the ratio of fuel costs to engine operating costs, typically ranging from 45% to 55%.

[0121] Preferably, step S8 specifically includes the following steps:

[0122] S81. Considering that the takeoff weight, payload and power spectrum of the flight platform are different under different mission conditions, the weight distribution and power spectrum of the flight platform under different mission conditions should be obtained, including the takeoff phase, climb phase, cruise phase, descent phase, landing phase and flight reserve phase. An example of the power spectrum provided by the flight platform provider is shown in Table 6.

[0123] Table 6. Example of power spectrum provided by the flight platform provider (power normalization processing)

[0124] ;

[0125] S82. Based on the weight distribution under different mission conditions of the flight platform, calculate the remaining weight of the flight platform that can be used to load the engine and fuel under different mission conditions:

[0126] m res =m fule +m engine =m to -m st -m pay ;

[0127] Where: m res The remaining weight of the flight platform; m fule For the fuel weight of the flight platform; m engine For engine weight; m to The takeoff weight of the flight platform; m st The weight of the flight platform's airframe structure; m pay For the payload capacity of the flight platform;

[0128] S83. Based on the engine baseline design point performance model, and utilizing the component characteristic diagrams generated from the preliminary design, establish an engine non-design point performance model. According to the power spectrum under different mission conditions of the flight platform, input altitude and power into the non-design point performance model to perform non-design point calculations, obtain the engine fuel consumption rate and fuel flow rate for each flight phase, and calculate the required fuel weight by combining the time of each flight phase.

[0129] m fule,de,0 =∑(SFC j ×P dn,j ×t j );

[0130] Where: m fule,de,0 Fuel weight requirements for engine baseline design; SFC j The engine's baseline fuel consumption rate for each flight phase; P dn,j Power for each flight phase; t j The time for each flight phase;

[0131] S84. Based on the remaining weight of the flight platform under different mission conditions, the engine weight of the engine baseline scheme, and the required fuel weight, and considering a certain fuel surplus factor, calculate the ratio of the remaining weight of the flight platform to the engine weight and the required fuel weight to obtain the remaining weight factor:

[0132] δ res = m res / ( m engine,0+ m fule,de,0 ×β);

[0133] Where: δ res m is the remaining weight coefficient of the flight platform. res The remaining weight of the flight platform; m engine,0 For the engine baseline design weight; m fule,de,0 β is the fuel weight required for the engine baseline scheme; β is the remaining fuel coefficient, which ranges from 10% to 20%; the minimum value of the remaining weight coefficient is selected, which can be used to characterize the minimum ratio of the remaining weight of the flight platform, the engine baseline scheme weight, and the required fuel weight, and the ratio of the engine baseline scheme weight to the remaining weight of the flight platform under this condition is recorded.

[0134] S85. Although factors such as the number of compressor stages and the materials used can affect engine weight, the compressor's converted flow rate remains the key factor causing changes in engine weight. Therefore, the engine weight coefficient is determined based on the ratio of the weight of the new engine design to the weight of the baseline engine design.

[0135] δ engine =m engine / m engine,0 ;

[0136] Where: δ engine For engine weight coefficient; m engine For engine weight; m engine,0 The weight is the benchmark weight for the engine.

[0137] Based on existing engine data, establish the relationship between the engine weight coefficient and the compressor equivalent flow rate (see...). Figure 6 ):

[0138] δ engine = a5×Wac2 2 +b5×Wac2 1 +c5;

[0139] Where: δ engine is the engine weight coefficient; Wac2 is the compressor inlet converted flow rate; a5, b5, and c5 are coefficients determined based on curve fitting results.

[0140] S86. Based on the changes in engine weight and fuel consumption rate, calculate the minimum residual weight coefficient corresponding to the new engine design:

[0141] δ res,min =(δ engine ×γ+SFC / SFC0×(1-γ))×δ res,min,0 ;

[0142] Where: δ res,min δ is the minimum residual weight coefficient for the flight platform. engine γ is the engine weight coefficient; γ is the ratio of engine weight to remaining weight under the conditions of using the engine baseline scheme and the minimum remaining weight coefficient of the flight platform; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline scheme design point fuel consumption rate; δ res,min,0 This is the minimum remaining weight coefficient of the flight platform corresponding to the engine baseline scheme.

[0143] In the above embodiment, step S9 substitutes the coefficients or formulas obtained in steps S2 to S8 into the design point performance model, giving the range and step size of the changes in combustion chamber outlet temperature and pressure ratio, and performs iterative calculations to obtain the distribution of engine operating cost coefficient and flight platform minimum residual weight coefficient on the thermodynamic cycle parameter analysis diagram, as detailed in [see...]. Figure 7 and Figure 8 As can be seen from the figure, appropriately increasing the combustion chamber outlet temperature and pressure ratio can further reduce engine operating costs, while the minimum residual weight coefficient of the flight platform remains within an acceptable range.

[0144] The method proposed in this application has been applied in the design of a certain type of turboshaft engine. Based on the method and conclusions provided in this application, appropriately increasing the combustion chamber outlet temperature and pressure ratio not only further improves indicators such as power-to-weight ratio and fuel consumption rate, but also makes the engine more cost-effective and further extends the range of the flight platform.

[0145] like Figure 9 As shown, another preferred embodiment of this application also provides a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, including:

[0146] The cost ratio determination module is used to establish the engine benchmark scheme and its design point performance model, and to determine the cost ratio of each component.

[0147] The compressor total cost coefficient calculation module is used to determine the compressor structure scheme applicable to different pressure ratios, establish the cost calculation formula for compressor rotor blade disk and compressor stator blade assembly, and then establish the compressor total cost coefficient calculation formula by combining the cost ratio relationship of the engine benchmark scheme.

[0148] The total cost coefficient calculation module for turbines is used to establish calculation formulas for the total cost coefficient of turbines in different gas temperature ranges, taking into account the differences in process cost, material cost and yield of gas turbine structural schemes in different gas temperature ranges.

[0149] The combustion chamber total cost coefficient calculation module is used to establish the combustion chamber total cost coefficient calculation formula based on the materials, cooling structure and coating scheme of the combustion chamber flame tube, the outer diameter and wall thickness of the combustion chamber casing, the number of nozzles and the inlet conversion flow rate;

[0150] The total cost coefficient calculation module for intake and exhaust systems is used to establish a formula for calculating the total cost coefficient of intake and exhaust systems based on the material cost ratio of the intake and exhaust systems in the engine baseline design and the converted flow rate of the compressor inlet.

[0151] The engine manufacturing and periodic overhaul cost coefficient calculation module is used to establish the engine manufacturing and periodic overhaul cost coefficient calculation formula based on the compressor total cost coefficient, the compressor cost ratio of the engine benchmark scheme, the gas turbine total cost coefficient, the gas turbine cost ratio of the engine benchmark scheme, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust device total cost coefficient.

[0152] The engine operating cost coefficient calculation module is used to establish an engine operating cost coefficient calculation formula based on the engine manufacturing and periodic overhaul cost coefficients and fuel consumption rate, combined with the relationship between fuel cost and engine operating cost.

[0153] The suitability assessment module is used to calculate the remaining weight, engine fuel requirement weight, and minimum remaining weight coefficient under different mission conditions based on the weight distribution and power spectrum of the flight platform, and to complete the suitability assessment of the thermodynamic cycle parameters relative to the mission of the flight platform.

[0154] The iterative calculation module is used to substitute the obtained coefficients or formulas into the design point performance model, give the range and step size of the changes in combustion chamber outlet temperature and pressure ratio, and carry out iterative calculations to obtain the distribution of engine operating cost coefficient and flight platform minimum residual weight coefficient on the thermodynamic cycle parameter analysis diagram.

[0155] The cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation device provided in this application, employing the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method described in the above embodiments, can solve the technical problem that existing thermodynamic cycle parameter designs lack comprehensive consideration of cost and mission requirements, failing to meet the economic requirements of flight platforms. Compared with the prior art, the beneficial effects of the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation device provided in this application are the same as those of the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method provided in the above embodiments, and other technical features in the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation device are the same as those disclosed in the methods of the above embodiments, and will not be repeated here.

[0156] like Figure 10 As shown, a preferred embodiment of this application also provides an electronic device, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the steps of the method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements in the above embodiments.

[0157] The electronic device provided in this application employs the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method described in the above embodiments. This method addresses the technical problem that existing thermodynamic cycle parameter designs often lack comprehensive consideration of cost and mission requirements, failing to meet the economic requirements of flight platforms. Compared to the prior art, the electronic device provided in this application has the same beneficial effects as the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method provided in the above embodiments, and other technical features of the electronic device are the same as those disclosed in the methods of the above embodiments, and will not be repeated here.

[0158] like Figure 11 As shown, a preferred embodiment of this application also provides a computer device, which may be a terminal or a liveness detection server, and its internal structure diagram may be as follows. Figure 11 As shown, the computer device includes a processor, memory, and a network interface connected via a system bus. The processor provides computational and control capabilities. The memory includes a non-volatile storage medium and internal memory. The non-volatile storage medium stores the operating system and computer programs. The internal memory provides an environment for the operation of the operating system and computer programs stored in the non-volatile storage medium. The network interface is used to communicate with other external computer devices via a network connection. When the computer program is executed by the processor, it implements the steps of the aforementioned method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements.

[0159] Those skilled in the art will understand that Figure 11 The structure shown is merely a block diagram of a portion of the structure related to the present application and does not constitute a limitation on the computer device to which the present application is applied. Specific computer devices may include more or fewer components than those shown in the figure, or combine certain components, or have different component arrangements.

[0160] The computer equipment provided in this application employs the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method described in the above embodiments. This addresses the technical problem that existing thermodynamic cycle parameter designs often lack comprehensive consideration of cost and mission requirements, failing to meet the economic requirements of flight platforms. Compared to existing technologies, the beneficial effects of the computer equipment provided in this application are the same as those of the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method provided in the above embodiments. Furthermore, other technical features of the electronic equipment are identical to those disclosed in the methods of the above embodiments, and will not be elaborated upon here.

[0161] A preferred embodiment of this application also provides a storage medium including a stored program that, when the program is executed, controls the device where the storage medium is located to perform the steps of the method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements in the above embodiments.

[0162] It should be noted that the steps shown in the flowchart in the accompanying drawings can be executed in a computer system such as a set of computer-executable instructions, and although a logical order is shown in the flowchart, in some cases the steps shown or described may be executed in a different order than that shown here.

[0163] If the functions described in this embodiment are implemented as software functional units and sold or used as independent products, they can be stored in one or more computing device-readable storage media. Based on this understanding, the parts of this application's embodiments that contribute to the prior art or the technical solutions can be embodied in the form of a software product. This software product is stored in a storage medium and includes several instructions to cause a computing device (which may be a personal computer, server, mobile computing device, or network device, etc.) to execute all or part of the steps of the methods described in the various embodiments of this application. The aforementioned storage media include: USB flash drives, portable hard drives, read-only memory (ROM), random access memory (RAM), magnetic disks, optical disks, and other media capable of storing program code.

[0164] Those skilled in the art will understand that embodiments of this application can be provided as methods, systems, or computer program products. Therefore, this application can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, this application can take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code. The solutions in the embodiments of this application can be implemented in various computer languages, such as the object-oriented programming language C++ and the embedded programming language C.

[0165] This application is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of this application. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart... Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0166] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0167] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0168] This application also provides a computer program product, including a computer program that, when executed by a processor, implements the steps of the above-described method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements.

[0169] The computer program product provided in this application can solve the technical problem that existing thermodynamic cycle parameter designs lack comprehensive consideration of cost and mission requirements, thus failing to meet the economic requirements of flight platforms. Compared with the prior art, the beneficial effects of the computer program product provided in this application are the same as those of the turboshaft engine thermodynamic cycle parameter evaluation method based on cost and mission requirements provided in the above embodiments, and will not be repeated here.

[0170] Although preferred embodiments of this application have been described, those skilled in the art, upon learning the basic inventive concept, can make other changes and modifications to these embodiments. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments as well as all changes and modifications falling within the scope of this application.

[0171] Obviously, those skilled in the art can make various modifications and variations to this application without departing from the spirit and scope of this application. Therefore, if such modifications and variations fall within the scope of the claims of this application and their equivalents, this application also intends to include such modifications and variations.

Claims

1. A method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, characterized in that, Including the following steps: S1. Establish the engine benchmark scheme and its design point performance model, and determine the cost ratio of each component; S2. Determine the compressor structure scheme applicable to different pressure ratios, establish the cost calculation formula for compressor rotor blade disk and compressor stator blade assembly, and then establish the compressor total cost coefficient calculation formula by combining the cost ratio relationship of the engine benchmark scheme. S3. To address the differences in process cost, material cost, and yield rate among gas turbine structural schemes for different gas temperature ranges, establish a calculation formula for the total cost coefficient of turbines for different gas temperature ranges. S4. Based on the materials, cooling structure, and coating scheme of the combustion chamber flame tube, the outer diameter and wall thickness of the combustion chamber casing, the number of nozzles, and the inlet conversion flow rate, establish the formula for calculating the total cost coefficient of the combustion chamber; S5. Establish a formula for calculating the total cost coefficient of the intake and exhaust systems based on the material cost ratio of the intake and exhaust systems in the engine baseline design and the converted flow rate of the compressor inlet. S6. Establish a formula for calculating the engine manufacturing and periodic overhaul cost coefficient based on the compressor total cost coefficient, the compressor cost ratio of the engine benchmark scheme, the gas turbine total cost coefficient, the gas turbine cost ratio of the engine benchmark scheme, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust device total cost coefficient. S7. Based on the engine manufacturing and periodic overhaul cost coefficient and fuel consumption rate, and combined with the relationship between fuel cost and engine usage cost, establish the engine usage cost coefficient calculation formula. S8. Based on the weight distribution and power spectrum of the flight platform under different mission conditions, calculate the remaining weight, engine fuel requirement weight, and minimum remaining weight coefficient under different mission conditions, and complete the applicability assessment of the thermodynamic cycle parameters relative to the flight platform mission. S9. Substitute the coefficients or formulas obtained in steps S2 to S8 into the design point performance model, give the range and step size of the changes in combustion chamber outlet temperature and pressure ratio, and carry out iterative calculations to obtain the distribution of engine operating cost coefficient and flight platform minimum remaining weight coefficient on the thermodynamic cycle parameter analysis diagram.

2. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, characterized in that, Step S1 specifically includes the following steps: S11. Use existing methods to complete the thermodynamic cycle parameter optimization analysis of the engine design point, so that the power and fuel consumption rate of the engine design point meet the index requirements, and establish the engine benchmark scheme and its design point performance model. S12. Based on the existing engine benchmark scheme and cost situation, and in conjunction with the supplier inquiry results, determine the proportion of manufacturing cost of each component.

3. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, step S2 specifically includes the following steps: S21. By statistically analyzing the compressor data of existing engines, the compressor configurations applicable to different pressure ratio ranges can be obtained, and the compressor structure schemes applicable to different pressure ratios can be determined. S22. Establish the relationship between the compressor inlet outer diameter and the compressor inlet converted flow rate based on existing engine data: D com =a1×Wac2 2 +b1×Wac2+c1; in, D com Where is the compressor inlet outer diameter; Wac2 is the compressor inlet converted flow rate; a1, b1, and c1 are coefficients determined based on curve fitting results; S23. Based on the changes in compressor outer diameter, number of blade disks, and blade disk material, establish a formula for calculating the compressor blade disk cost coefficient, thereby obtaining the changes in the new compressor blade disk scheme relative to the engine baseline scheme: g comp,disc =(1-a comp,dis,mat,0 +a comp,dis,mat,0 ×D com 2 / D com,0 2 )×N comp,disc / N comp,disc,0 × ((N comp,dis -N comp,dis,GH +N comp,dis,GH,0 ) / N comp,dis +(N comp,dis,GH -N comp,dis,GH,0 ) / N comp,dis ×(1-a comp,dis,mat,0 +a comp,dis,mat,0 ×C mat,GH / C mat,Ti ); Where: ζ comp,disc The compressor rotor blade disk cost coefficient represents the cost ratio of the new compressor blade disk design relative to the engine baseline design; α comp,dis,mat,0 The material cost ratio of the compressor rotor blade disk in the engine baseline design is taken as 20% to 35%; D com D is the outer diameter of the compressor inlet. com,0 The compressor inlet outer diameter is the reference diameter for the engine design; N comp,disc N represents the number of compressor rotor blade disks. comp,disc,0 The number of compressor rotor blade disks in the engine baseline design; N comp,dis,GH N represents the number of high-temperature alloy rotor blade disks in the compressor. comp,dis,GH,0 The number of high-temperature alloy rotor blade disks for the compressor in the engine baseline design; C mat,GH / C mat,Ti The cost ratio of high-temperature alloys to titanium alloys needs to take into account the differences in material density and unit price; S24. Establish a formula for calculating the cost coefficient of the compressor stator blade assembly based on the influence of the number of stages and outer diameter variations in the compressor stator blade assembly: g comp,stator =(1-a comp,stator,mat,0 +a com,stator,mat,0 ×D com 2 / D com,0 2 )× (1+(N comp -N comp,0 )×C comp,sator,mid,0 / C comp,sator,0 ); Where: ζ comp,stator α represents the cost coefficient of the compressor stator blade assembly, i.e., the cost ratio of the new design to the engine baseline design; comp,stator,mat,0 The material cost ratio of the compressor stator blade assembly in the engine baseline design; D com D is the outer diameter of the compressor inlet. com,0 The compressor inlet outer diameter is the reference diameter for the engine design; N comp N represents the number of compressor stages. comp,0 The number of compressor stages for the engine baseline design; C comp,sator,mid,0 / C comp,sator,0 The cost of a single intermediate stage stator relative to the total cost of the stator blade assembly is 15% to 25%. S25. Using the cost coefficients of the compressor rotor blade disk and compressor stator blade assembly, and combining the cost ratio of the engine benchmark scheme, establish a formula for calculating the total cost coefficient of the compressor: ζ comp =C comp,disc,0 / (C comp,sator,0 +C comp,disc,0 )×ζ comp,disc +C comp,sator,0 / (C comp,sator,0 +C comp,disc,0 )×ζ comp,stato ; Where: ζ comp C is the total cost coefficient for the compressor. comp,disc,0 Cost of compressor rotor blade disk for engine baseline design; C comp,sator,0 Cost of compressor stator blade assembly for engine baseline design; ζ comp,disc Cost coefficient ζ for compressor rotor blade disk comp,stato This is the cost coefficient for the compressor stator blade assembly.

4. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, step S3 specifically includes the following steps: S31. Using existing engine technology, establish the relationship between the proportion of bleed air flow of each part of the turbine to the engine inlet flow and the combustion chamber outlet temperature, including the proportion of bleed air from the first stage of the gas turbine, the proportion of bleed air from the gas turbine working blades, and the proportion of bleed air from the power turbine. The proportion of bleed air from the first stage of the gas turbine refers to the proportion of bleed air from the first stage guide section of the gas turbine, the proportion of bleed air from the gas turbine working blades refers to the proportion of bleed air from the power turbine section. S32. Referring to existing engines, establish the relationship between the gas temperature range and the material levels and cooling structure schemes of the main aerodynamic components of the turbine, and determine the applicable structural schemes for different gas temperatures at the inlet of the gas turbine and the power turbine. The inlet gas temperature of a gas turbine refers to the combustion chamber outlet temperature, while the inlet gas temperature of a power turbine refers to the gas turbine outlet temperature. S33. To address the differences in process cost, material cost, and yield rate among different turbine designs, establish a formula for calculating the total turbine cost coefficient for different gas temperature ranges: g HPT =∑(α HPT,k,0 ×((1-a HPT,mat,k,0 )×C HPT,process,k / C HPT,process,0,k +a HPT,mat,k,0 ×C HPT,mat,k / C HPT,mat,0,k )×Q HPT,k / Q HPT,k,0 ); Where: ζ HPT The total cost coefficient for the gas turbine is the cost ratio of the new design to the benchmark engine design; α HPT,k,0 The cost breakdown of each component of the gas turbine in the benchmark engine design; α HPT,mat,k,0 The material cost ratio of the corresponding part of the gas turbine in the engine benchmark scheme; C HPT,process,k / C HPT,process,0,k The ratio of gas turbine manufacturing costs to the benchmark engine design can be determined based on statistical data; C HPT,mat,k / C HPT,mat,0, k represents the ratio of the gas turbine material cost to the engine baseline, which can be determined based on statistical data; Q HPT,k / Q HPT,k,0 The ratio of the gas turbine yield rate to the engine baseline can be determined based on statistical data.

5. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 4, in step S31: the calculation formula for the combustion-guide gas ratio is: CO NGV =a2×T4 3 +b2×T4 2 +c2×T4+d2; in, CO NGV T1 is the ratio of induced draft gas to induced draft gas; T4 is the combustion chamber outlet temperature; a2, b2, c2, and d2 are coefficients determined based on curve fitting results. The formula for calculating the combustion-driven bleed air ratio is: WHAT HPTR =a3×T4 3 +b3×T4 2 +c3×T4+d3; Among them, CO HPTR T4 is the combustion chamber outlet temperature; a3, b3, c3, and d3 are coefficients determined based on curve fitting results. The formula for calculating the bleed air ratio of the power turbine is: WHAT LPTR =a4×T4 3 +b4×T4 2 +c4×T4+d4; Among them, CO LPTR T4 is the bleed air ratio of the power turbine; T4 is the combustion chamber outlet temperature; a4, b4, c4, and d4 are coefficients determined based on the curve fitting results.

6. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, characterized in that, Step S4 specifically includes the following steps: S41. Based on the materials, cooling structure, and coating scheme of the combustion chamber flame tube, the outer diameter and wall thickness of the combustion chamber casing, the number of nozzles, and the inlet converted flow rate, establish the formula for calculating the total cost coefficient of the combustion chamber: ζ comb =f(T4,α comb,k,0 ,α comb,mat,k,0 ,C comb,process,k / C comb,process,0,k ,C comb,mat,k / C comb,mat,0,k ,Say 31 / Speech 31,0 ,PI,Q comb,k / Q comb,k,0 ); Where: ζ comb T4 is the total cost coefficient of the combustion chamber, i.e., the cost ratio of the new scheme to the engine baseline scheme; T4 is the combustion chamber outlet temperature; α comb,k,0 The cost breakdown of each component of the combustion chamber in the engine baseline design; α comb,mat,k,0 The material cost ratio of each part of the combustion chamber in the engine baseline design; C comb,process,k / C comb,process,0,k C represents the ratio of the manufacturing costs of each part of the combustion chamber to the benchmark engine design; comb,mat,k / Ccomb,mat,0,k represents the ratio of material costs for each part of the combustion chamber relative to the engine's baseline design; Wac 31 Calculated flow rate for combustion chamber inlet; Wac 31,0 For the engine baseline design, the converted flow rate at the combustion chamber inlet is PI, where PI represents the compressor pressure ratio; Q comb,k / Q comb,k,0 This represents the ratio of the yield rate of each part of the combustion chamber to the engine baseline.

7. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, characterized in that, Step S5 specifically includes the following steps: S51. Based on the material cost ratio of the intake and exhaust systems in the engine baseline design and the converted flow rate of the compressor inlet, establish a formula for calculating the total cost coefficient of the intake and exhaust systems: g inout =f(a inout,mat ,Wac2); Where: ζ inout The total cost coefficient for the combustion chamber is the cost ratio of the new design relative to the benchmark engine design; α inout,mat The ratio of material costs for the intake and exhaust systems in the engine baseline design; Wac2 is the converted flow rate at the compressor inlet.

8. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, characterized in that, Step S6 specifically includes the following steps: S61. Based on the compressor total cost coefficient, the compressor cost ratio of the engine baseline design, the gas turbine total cost coefficient, the gas turbine cost ratio of the engine baseline design, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust system total cost coefficient, establish the formula for calculating the engine manufacturing and periodic overhaul cost coefficient: g manu =ζ comp ×a comp,0 +g HPT ×a HPT,0 +g LPT ×a LP,0T +g comb ×a comb,0 +g inout ×a inout,0 + (1-α comp,0 -α HPT,0 -α LPT,0 -α comb,0 -α inout,0 )； Where: ζ manu For engine manufacturing and periodic overhaul cost coefficients; ζ comp α is the total cost coefficient for the compressor; comp,0 The compressor cost percentage for the benchmark engine design; ζ HPT α is the total cost coefficient for gas turbines. HPT,0 The cost percentage of the gas turbine in the engine benchmark solution; ζ LPT α represents the total cost coefficient of the power turbine. LPT,0 The cost percentage of the power turbine in the benchmark engine design; ζ comb α is the total cost coefficient for the combustion chamber. comb,0 The combustion chamber cost percentage of the engine's benchmark design; ζ inout α represents the total cost coefficient for the intake and exhaust systems. inout,0 The cost percentage of the intake and exhaust systems in the engine baseline design.

9. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, characterized in that, Step S7 specifically includes the following steps: S71. Based on the engine manufacturing and periodic overhaul cost coefficients and fuel consumption rate, and considering the relationship between fuel costs and engine operating costs, establish a formula for calculating the engine operating cost coefficient: g eng =ζ manu ×a manu +SFC / SFC0×α SFC +(1-a manu -a SFC ); Where: ζ eng For engine operating cost coefficient; ζ manu α represents the cost coefficient for engine manufacturing and periodic overhauls. manu The ratio of manufacturing and periodic overhaul costs to engine operating costs is taken as 40%–50%; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline design point fuel consumption rate; α SFC This represents the ratio of fuel cost to engine operating cost, ranging from 45% to 55%.

10. The method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to claim 1, characterized in that, Step S8 specifically includes the following steps: S81. Obtain the weight distribution and power spectrum of the flight platform under different mission conditions, including the takeoff phase, climb phase, cruise phase, descent phase, landing phase, and flight reserve phase. S82. Based on the weight distribution under different mission conditions of the flight platform, calculate the remaining weight of the flight platform that can be used to load the engine and fuel under different mission conditions: m res =m fule +m engine =m to -m st -m pay ; Where: m res The remaining weight of the flight platform; m fule For the fuel weight of the flight platform; m engine For engine weight; m to The takeoff weight of the flight platform; m st The weight of the flight platform's airframe structure; m pay For the payload capacity of the flight platform; S83. Based on the engine baseline design point performance model, and utilizing the component characteristic diagrams generated from the preliminary design, establish an engine non-design point performance model. According to the power spectrum under different mission conditions of the flight platform, input altitude and power into the non-design point performance model to perform non-design point calculations, obtain the engine fuel consumption rate and fuel flow rate for each flight phase, and calculate the required fuel weight by combining the time of each flight phase. m fule,de,0 =∑(SFC j ×P dn,j ×t j ); Where: m fule,de,0 Fuel weight requirements for engine baseline design; SFC j The engine's baseline fuel consumption rate for each flight phase; P dn,j Power for each flight phase; t j The time for each flight phase; S84. Based on the remaining weight of the flight platform under different mission conditions, the engine weight of the engine baseline scheme, and the required fuel weight, and considering a certain fuel surplus factor, calculate the ratio of the remaining weight of the flight platform to the engine weight and the required fuel weight to obtain the remaining weight factor: δ res =m res / (m engine,0+ m fule,de,0 ×β); Where: δ res m is the remaining weight coefficient of the flight platform. res The remaining weight of the flight platform; m engine,0 For the engine baseline design weight; m fule,de,0 The required fuel weight for the engine's baseline design; β is the residual fuel coefficient, ranging from 10% to 20%; S85. Determine the engine weight coefficient based on the ratio of the weight of the new engine design to the weight of the baseline engine design: d engine =m engine / m engine,0 ; Where: δ engine For engine weight coefficient; m engine For engine weight; m engine,0 The weight is the benchmark weight for the engine. Based on existing engine data, establish the relationship between the engine weight coefficient and the compressor equivalent flow rate: δ engine =a5×Wac2 2 +b5×Wac2 1 +c5; Where: δ engine is the engine weight coefficient; Wac2 is the compressor inlet converted flow rate; a5, b5, and c5 are coefficients determined based on curve fitting results. S86. Based on the changes in engine weight and fuel consumption rate, calculate the minimum residual weight coefficient corresponding to the new engine design: d res,min =(δ engine ×γ+SFC / SFC0×(1-γ))×δ res,min,0 ; Where: δ res,min δ is the minimum residual weight coefficient for the flight platform. engine γ is the engine weight coefficient; γ is the ratio of engine weight to remaining weight under the conditions of using the engine baseline scheme and the minimum remaining weight coefficient of the flight platform; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline scheme design point fuel consumption rate; δ res,min,0 This is the minimum remaining weight coefficient of the flight platform corresponding to the engine baseline scheme.

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