Method for evaluating thermodynamic cycle parameters of turboshaft engine based on cost and task requirements

Abstract

The invention discloses a turboshaft engine thermodynamic cycle parameter evaluation method based on cost and task requirements, and the method achieves the evaluation of the thermodynamic cycle parameters of a turboshaft engine on the premise of comprehensively considering the coupling influence of cycle parameters, all parts and air system parameters. According to the method, the relation between the thermodynamic cycle parameters and all parts and the whole machine cost is established, and the distribution condition of the engine cost and the engine and fuel oil weight coefficient about the cycle parameters is obtained by combining the balance between the aircraft task and the weight, so that a reliable and scientific basis is provided for thermodynamic cycle parameter evaluation and further optimization; the method can be used for evaluating the applicability of different thermodynamic cycle parameter combinations to the flight platform, so that the cost of an engine is more advantageous, the minimum residual weight coefficient of the flight platform is still within an acceptable range, and the voyage of the flight platform is further improved.

Description

Technical Field

[0001] The present application relates to the field of aviation engine technology, and in particular, to a method for evaluating thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements. Background Art

[0002] As a power-output engine, turboshaft engines are the primary propulsion systems for platforms such as helicopters and tiltrotor aircraft. In the civilian market, the economic efficiency of an aircraft platform has become a crucial factor in determining its market competitiveness. The manufacturing and periodic overhaul costs of turboshaft engines, as well as fuel costs, are the primary components of the direct operating costs of an aircraft platform, collectively accounting for over 60%. Therefore, reducing these related costs is one of the primary ways to improve the economic efficiency of an aircraft platform and has become a crucial consideration in turboshaft engine design.

[0003] Thermodynamic cycle parameters of a turboshaft engine include combustion chamber outlet temperature and compressor pressure ratio. While meeting the target flight platform power, increasing cycle parameters such as combustion chamber outlet temperature and pressure ratio can reduce engine weight and improve fuel consumption. However, this increases material and process 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, combined with current industrial technology and cost conditions, can effectively control engine manufacturing and periodic overhaul costs, as well as fuel costs, and provide a sound foundation for the specific design of engine components. However, currently available thermodynamic cycle parameter design methods primarily focus on optimizing performance indicators, with limited consideration of cost and mission requirements. The impact of thermodynamic cycle parameters on engine cost is not evaluated. Furthermore, the selection of thermodynamic cycle parameters during the design process rarely considers the weight distribution and power spectrum under different flight platform conditions. Consequently, in order to fully balance engine weight and fuel consumption, over- or under-design may occur, failing to meet the flight platform's economic requirements. Summary of the Invention

[0004] The present application provides a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, which is used to solve the technical problem that the existing thermodynamic cycle parameter design has insufficient comprehensive consideration of cost and mission requirements and does not meet the economic requirements of the flight platform.

[0005] This application is implemented through the following scheme: A method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements includes the following steps: S1. Establish an engine benchmark scheme and its design point performance model, and determine the cost ratio of each component; S2. Determine the compressor structure schemes suitable for different pressure ratios, establish cost calculation formulas for the compressor rotor blade disk and compressor stator blade assembly, and then establish a calculation formula for the total compressor cost coefficient based on the cost ratio relationship of the engine baseline scheme; S3. Based on the differences in process cost, material cost, and yield rate of gas turbine structural schemes in different gas temperature ranges, a calculation formula for the total cost coefficient of turbines in different gas temperature ranges is established; S4. Establish a calculation formula for the total cost coefficient of the combustion chamber based on the materials, cooling structure and coating scheme used for 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; S5. Establish a calculation formula for the total cost coefficient of the intake and exhaust devices based on the material cost ratio of the intake and exhaust devices of the engine baseline solution and the converted compressor inlet flow rate; S6. Establish a calculation formula for the engine manufacturing and periodic overhaul cost coefficient based on the total compressor cost coefficient, the engine baseline compressor cost ratio, the gas turbine total cost coefficient, the engine baseline gas turbine cost ratio, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust device total cost coefficients; S7. Establish a formula for calculating the engine cost coefficient based on the engine manufacturing and periodic overhaul cost coefficients and fuel consumption rate, combined with the relationship between fuel cost and engine cost; S8. Based on the weight distribution and power spectrum of the flight platform under different mission conditions, calculate the residual weight, engine fuel weight, and minimum residual weight coefficient under different mission conditions, and complete the suitability assessment of the thermodynamic cycle parameters for 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 combustion chamber outlet temperature and pressure ratio, and perform iterative calculations to obtain the distribution of the engine use cost coefficient and the flight platform minimum residual weight coefficient on the thermodynamic cycle parameter analysis diagram.

[0006] Compared with the existing technology, this application has the following beneficial effects: This application proposes a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements. The method establishes the relationship between the thermodynamic cycle parameters and the costs of each component and the entire aircraft, taking into account the coupling effects of the cycle parameters with the parameters of each component and the air system. Furthermore, the method combines the balance between the mission and weight of the aircraft to obtain the distribution of the engine cost, engine, and fuel weight coefficients with respect to the cycle parameters. This method provides a feasible approach for evaluating and optimizing the thermodynamic cycle parameters, specifically including: 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. It also establishes a relationship model between thermodynamic cycle parameters and component costs. The application 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. 2) This application establishes a minimum residual weight coefficient for the flight platform based on the weight distribution and power spectrum of the aircraft under different conditions, combined with the impact of thermodynamic cycle parameters on engine weight and fuel consumption rate. This is to characterize the impact of different thermodynamic cycle parameters on the mission capability margin of the flight platform, and is used to evaluate the applicability of different thermodynamic cycle parameter combinations to the flight platform, making the engine cost more advantageous while keeping the minimum residual weight coefficient of the flight platform within an acceptable range, further improving the range of the flight platform.

[0007] In addition to the above-described purposes, features and advantages, the present application also has other purposes, features and advantages. The present application will be further described in detail below with reference to the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0008] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the present application and, together with the description, serve to explain the principles of the present application.

[0009] In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, a person skilled in the art can derive other drawings based on these drawings without inventive work, among which: Figure 1 It is a schematic diagram of the existing thermodynamic cycle parameter analysis; Figure 2 This is a schematic diagram of a typical turboshaft engine layout; Figure 3 This is a flow chart of a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements according to a preferred embodiment of the present application; Figure 4 This is a schematic diagram of the compressor inlet outer diameter and the compressor inlet converted flow rate; Figure 5 It is a schematic diagram of the relationship between the turbine bleed air ratio and the combustion chamber outlet temperature; Figure 6 It is a diagram showing the relationship between the engine weight coefficient and the compressor converted flow rate; Figure 7 It is a schematic diagram of the distribution of the engine use cost coefficient on the thermodynamic cycle parameter analysis diagram; Figure 8It is a schematic diagram of the distribution of the minimum residual weight coefficient of the flight platform on the thermodynamic cycle parameter analysis diagram; Figure 9 This is a schematic diagram of a module of a turboshaft engine thermodynamic cycle parameter evaluation device based on cost and mission requirements in a preferred embodiment of the present application; Figure 10 This is a schematic block diagram of an electronic device according to a preferred embodiment of the present application; Figure 11 It is a diagram of the internal structure of a computer device according to a preferred embodiment of the present application.

[0010] In the figure: 1. Air intake device; 2. Multi-stage axial flow; 3. Centrifugal impeller; 4. Combustion chamber; 5. Gas turbine; 6. Power turbine; 7. Exhaust device. DETAILED DESCRIPTION

[0011] It should be understood that the specific embodiments described herein are merely used to explain the technical solutions of the present application and are not intended to limit the present application.

[0012] In order to better understand the technical solution of the present application, a detailed description will be given below in conjunction with the accompanying drawings and specific implementation methods.

[0013] Existing turboshaft engine packages usually have components such as an air intake device, a compressor, a combustion chamber, a gas turbine, a power turbine and an exhaust device. The gas turbine and the compressor are connected to the same drive shaft, and the combustion chamber is arranged between the gas turbine and the compressor. Air is continuously sucked in by the compressor through the air intake device, and after compression, it enters the combustion chamber and burns with the injected fuel to become high-temperature combustion gas, and then enters the gas turbine to expand and perform work. At this time, the expansion work is transmitted to the compressor through the drive shaft to realize continuous operation of the engine, and the gas at the outlet of the gas turbine drives the power turbine to continue to expand and perform work to drive other equipment to work, such as helicopter rotors, generators, etc. The gas at the outlet of the power turbine enters the atmosphere through the exhaust device.

[0014] Parameters that affect engine performance include the total pressure loss coefficient of the given intake system, the converted compressor flow rate, compressor efficiency, total pressure loss in the combustion chamber, gas turbine inlet temperature, gas turbine efficiency, power turbine efficiency, total pressure loss in the transition section, exhaust loss, and the air system bleed ratio. In conventional overall performance design, facing a certain power target and taking into account the performance level of components, parameter analysis of the combustion chamber outlet temperature and compressor pressure ratio is carried out, and a "carpet map" is drawn. Based on the constraints of performance, structural strength, etc., the feasible region is determined to obtain appropriate thermodynamic cycle parameters. See Figure 1 Among them, fuel consumption rate reflects the fuel consumption level of the engine; unit power refers to the power corresponding to unit air flow, which can represent the weight and size of the engine.

[0015] To improve the rationality of turboshaft engine performance design, existing technologies have leveraged the coupled relationships between various components and the air system to more precisely select thermodynamic cycle parameters. However, these technologies focus on optimizing performance metrics and pay little attention to engine cost. The compressor, combustion chamber, gas turbine, power turbine, and intake and exhaust systems are the primary 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 mainstream gas, and their combined cost accounts 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 variations in the structure, materials, and processes of these components, directly impacting engine manufacturing and scheduled overhaul costs. Given the technological development stage and overhaul cycle, scheduled overhaul costs can be considered directly related to manufacturing costs. Furthermore, due to the coupled relationships between various components and the air system parameters, excessive pursuit of low engine weight when selecting thermodynamic cycle parameters can lead to increased fuel consumption, negatively impacting the mission capability of the flight platform. Therefore, achieving a balance between engine weight and fuel consumption is an effective means of enhancing flight platform capabilities.

[0016] In addition to the thermodynamic cycle parameters, the diversity of component configurations will also affect the cost of the engine. However, with the advancement of technology, engines of the same power level have gradually converged in terms of overall layout, component configuration selection, etc. For example, considering the power expansion potential of the engine, turboshaft engines in the power range of 1000kW to 2000kW mostly use a "multi-stage axial flow + centrifugal impeller" combined compressor configuration and a two-stage gas turbine configuration (see Figure 2 ) includes: an air intake device 1, a multi-stage axial flow 2, a centrifugal impeller 3, a combustion chamber 4, a gas turbine 5, a power turbine 6, and an exhaust device 7, which provides favorable conditions for cost evaluation.

[0017] It should be noted that the execution entity of this embodiment can be a computing service device with data processing, network communication, and program execution capabilities, such as a tablet computer, personal computer, or mobile phone, or a turboshaft engine thermal cycle parameter evaluation device based on cost and mission requirements that can achieve the above functions. The following describes this embodiment and the following embodiments using a turboshaft engine thermal cycle parameter evaluation device based on cost and mission requirements as the execution entity.

[0018] like Figure 1 As shown, the preferred embodiment of the present application provides a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, comprising the steps of: S1. Establish an engine benchmark scheme and its design point performance model, and determine the cost ratio of each component; S2. Determine the compressor structure schemes suitable for different pressure ratios, establish cost calculation formulas for the compressor rotor blade disk and compressor stator blade assembly, and then establish a calculation formula for the total compressor cost coefficient based on the cost ratio relationship of the engine baseline scheme; S3. Based on the differences in process cost, material cost, and yield rate of gas turbine structural schemes in different gas temperature ranges, a calculation formula for the total cost coefficient of turbines in different gas temperature ranges is established; S4. Establish a calculation formula for the total cost coefficient of the combustion chamber based on the materials, cooling structure and coating scheme used for 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; S5. Establish a calculation formula for the total cost coefficient of the intake and exhaust devices based on the material cost ratio of the intake and exhaust devices of the engine baseline solution and the converted compressor inlet flow rate; S6. Establish a calculation formula for the engine manufacturing and periodic overhaul cost coefficient based on the total compressor cost coefficient, the engine baseline compressor cost ratio, the gas turbine total cost coefficient, the engine baseline gas turbine cost ratio, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust device total cost coefficients; S7. Establish a formula for calculating the engine cost coefficient based on the engine manufacturing and periodic overhaul cost coefficients and fuel consumption rate, combined with the relationship between fuel cost and engine cost; S8. Based on the weight distribution and power spectrum of the flight platform under different mission conditions, calculate the residual weight, engine fuel weight, and minimum residual weight coefficient under different mission conditions, and complete the suitability assessment of the thermodynamic cycle parameters for 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 combustion chamber outlet temperature and pressure ratio, and perform iterative calculations to obtain the distribution of the engine use cost coefficient and the flight platform minimum residual weight coefficient on the thermodynamic cycle parameter analysis diagram.

[0019] Compared with the prior art, this embodiment has the following beneficial effects: This embodiment proposes a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements. By comprehensively considering the coupling effects of the cycle parameters with the various components and air system parameters, the relationship between the thermodynamic cycle parameters and the costs of the various components and the entire aircraft is established. Furthermore, by balancing the mission and weight of the aircraft, the distribution of engine cost, engine and fuel weight coefficients with respect to the cycle parameters is obtained. This provides a feasible method for evaluating and optimizing the thermodynamic cycle parameters, specifically including: 1) This example utilizes the component technology 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 design of major aerodynamic components. A model is then established to determine the relationship between thermodynamic cycle parameters and component costs. This analysis analyzes the combined impact of changes in thermodynamic cycle parameters on manufacturing and scheduled overhaul costs, as well as fuel costs, providing a basis for further optimization of thermodynamic cycle parameters. 2) This embodiment establishes a minimum residual weight coefficient for the flight platform based on the weight distribution and power spectrum of the aircraft under different conditions, combined with the impact of thermodynamic cycle parameters on engine weight and fuel consumption rate. This coefficient characterizes the impact of different thermodynamic cycle parameters on the flight platform's mission capability margin. This is used to evaluate the applicability of different thermodynamic cycle parameter combinations to the flight platform, making the engine cost more competitive while keeping the minimum residual weight coefficient of the flight platform within an acceptable range, further improving the flight platform's range.

[0020] Preferably, step S1 specifically includes the steps of: S11. Use existing methods to perform optimization analysis of the thermodynamic cycle parameters of the engine design point, so that the power and fuel consumption rate of the engine design point meet the index requirements, and establish an engine benchmark scheme and its design point performance model; S12. Based on the existing engine baseline plan and cost situation, combined with the supplier inquiry results, determine the proportion of manufacturing costs of each component. See Table 1 for details.

[0021] Table 1 Proportion of manufacturing cost of each component ; 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 given configuration and technology level, the compressor stage pressure ratio is limited.

[0022] In order to adapt to the improvement of the compressor pressure ratio, the number of axial flow stages of the combined compressor must be increased, which will also increase the temperature of the axial flow final stage or centrifugal stage, and then the centrifugal stage will use high-temperature alloys, which will lead to increased costs. In addition to the pressure ratio, the engine flow rate and efficiency are also factors that affect the cost. The increase in flow rate will increase the outer diameter of the compressor, and the decrease in efficiency will cause the temperature of the axial flow final stage or centrifugal stage to increase accordingly. In addition, for the remaining parts of the compressor, including the outer casing of the axial flow part, the outer casing of the centrifugal part, etc., their corresponding manufacturing costs are positively correlated with the compressor flow path parts. Therefore, by obtaining the cost changes of the compressor rotor blade disk and the compressor stator blade assembly, the change in the total cost of the compressor can be basically determined. Therefore, preferably, step S2 specifically includes the steps of: S21. By statistically analyzing the compressor data of existing engines, it is possible to obtain compressor configurations applicable to different pressure ratio ranges and determine compressor structural schemes applicable to different pressure ratios. Table 2 shows combined compressor schemes applicable to 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, when the size and material remain unchanged, the total cost and quantity of blade disks are proportional. In order to improve the reliability and economy of the product, in addition to the first-stage blade disk and centrifugal impeller, the intermediate axial flow stage uses 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 corresponding to different pressure ratio ranges can be obtained; Table 2 Combined compressor structure schemes applicable to different pressure ratio ranges ; S22. The compressor inlet outer diameter determines the compressor inlet frontal area and 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 the existing engine data (see Figure 4 ): D com =a1×Wac2 2 +b1×Wac2+c1; Among them, D com is the compressor inlet outer diameter; Wac2 is the compressor inlet converted flow rate; a1, b1, c1 are coefficients, determined according to the curve fitting results; S23. For the compressor rotor blade disk, a calculation formula for the compressor blade disk cost coefficient is established based on the changes in the compressor outer diameter, the number of blade disks, and the blade disk material. This will yield the change in the compressor blade disk of the new solution relative to the engine baseline solution: ; Among them: comp,disc is the cost coefficient of the compressor rotor blade disk, which is the cost ratio of the new compressor blade disk to the engine benchmark solution; α comp,dis,mat,0 The material cost ratio of the compressor rotor blade disk of the engine baseline solution is generally 20% to 35%; D com is the compressor inlet outer diameter; D com,0 is the compressor inlet outer diameter of the engine base scheme; N comp,disc is the number of compressor rotor blades; N comp,disc,0 is the number of compressor rotor blades in the engine baseline solution; N comp,dis,GH is the number of compressor high-temperature alloy rotor blade disks; N comp,dis,GH,0 C is the number of high-temperature alloy rotor blades in the engine baseline scheme compressor; mat,GH / C mat,Ti The cost ratio of high-temperature alloys to titanium alloys takes into account the difference in material density and unit price; S24. For the compressor stator blade assembly, the cost of the intermediate stage stators is relatively stable, while the increase in the number of stages will increase the number of intermediate stage stators accordingly. Therefore, based on the influence of the number of stages and the change in the outer diameter of the compressor stator blade assembly, the calculation formula for the compressor stator blade assembly cost coefficient is established: ζ comp,stator =(1-α comp,stator,mat,0 +α 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 ); Among them: comp,stator is the cost coefficient of the compressor stator blade assembly, that is, the cost ratio of the new solution to the engine benchmark solution; α comp,stator,mat,0 D is the material cost ratio of the compressor stator blade assembly of the engine baseline solution; com is the compressor inlet outer diameter; D com,0 is the compressor inlet outer diameter of the engine base scheme; N comp is the number of compressor stages; N comp,0 is the number of compressor stages of the engine benchmark scheme; C comp,sator,mid,0 / C comp,sator, 0 The ratio of the cost of a single intermediate stage stator 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, combined with the cost ratio of the engine baseline solution, establish a calculation formula for the total compressor cost coefficient: 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 ; Among them: comp is the total cost coefficient of the compressor; C comp,disc,0 is the cost of the compressor rotor blade disk of the engine baseline solution; C comp,sator, 0 is the cost of the compressor stator blade assembly of the engine baseline solution; comp,discis the cost coefficient of the compressor rotor blade disk ζ comp,stato is the cost coefficient of the compressor stator blade assembly.

[0023] The main aerodynamic components of the turbine mainly include turbine guide vanes, turbine blades and turbine discs, which account for more than 75% of the turbine cost. Under a certain configuration and technical level, to increase the gas temperature, it is necessary to use more advanced high-temperature alloy materials and a more complex cooling structure, or increase the amount of bleed air for cooling. At the same time, changes in the combustion chamber outlet temperature, compressor pressure ratio, component efficiency and air system will affect the gas turbine rotor inlet converted flow rate. The gas turbine rotor converted flow rate directly reflects the gas turbine's flow capacity and flow area requirements, and thus determines the basic size of the turbine. However, for small and medium-sized turboshaft engines, the reduction in turbine size will increase the process difficulty of the cooling structure, and there is a certain cost offset. Therefore, when analyzing the turbine cost, the influence of the gas temperature can be mainly considered. Therefore, preferably, step S3 specifically includes the steps of: S31. Using the existing engine technology, establish the relationship between the ratio of the bleed air flow rate of each part of the turbine to the engine inlet flow rate and the combustion chamber outlet temperature, including the combustion-pilot air ratio, the combustion-driven bleed air ratio, and the power turbine bleed air ratio. Figure 5 The design results of the combustion-pilot air ratio, the combustion-dynamic bleed air ratio, and the power turbine bleed air ratio are respectively given as a function of the combustion chamber outlet temperature; the combustion-pilot air ratio refers to the bleed air ratio at the first-stage guide vane of the gas turbine, the combustion-dynamic bleed air ratio refers to the bleed air ratio at the gas turbine rotor blades (including the second-stage guide vane of the gas turbine), and the power turbine bleed air ratio refers to the bleed air ratio at the power turbine; specifically, the calculation formula for the combustion-pilot air ratio is: CO NGV = a2×T4 3 +b2×T4 2 +c2×T4+d2; Among them, CO NGV is the fuel-pilot gas ratio; T4 is the combustion chamber outlet temperature; a2, b2, c2, d2 are coefficients, determined according to the curve fitting results; The calculation formula for the combustion and bleed air ratio is: CO HPTR = a3×T4 3 +b3×T4 2 +c3×T4+d3; Among them, CO HPTR is the combustion air ratio; T4 is the combustion chamber outlet temperature; a3, b3, c3, d3 are coefficients, determined according to the curve fitting results; The calculation formula for the power turbine bleed air ratio is: CO LPTR= a4×T4 3 +b4×T4 2 +c4×T4+d4; Among them, CO LPTR is the ratio of power turbine bleed air; T4 is the combustion chamber outlet temperature; a4, b4, c4, d4 are coefficients, determined according to the curve fitting results; S32. With reference to existing engines, establish the relationship between the gas temperature range and the material level and cooling structure of the main aerodynamic components of the turbine, and determine the structural scheme applicable to different gas temperatures at the inlet of the gas turbine and the power turbine. The inlet gas temperature of the gas turbine refers to the combustion chamber outlet temperature, and the inlet gas temperature of the power turbine refers to the gas turbine outlet temperature. For gas turbines, inlet gas temperature refers to the combustion chamber outlet temperature; for power turbines, it refers to the turbine outlet temperature. Turbine inlet gas temperature has a significant impact on the material quality and cooling structure of the turbine's key aerodynamic components. While maintaining the same strength and lifespan requirements, increasing turbine inlet gas temperature necessitates the selection of materials with enhanced high-temperature and corrosion resistance. For example, the high-temperature alloys used in turbine blades can be upgraded from directionally solidified to first-, second-, or third-generation single crystals. Furthermore, adjustments to the blade cooling structure may be necessary, such as upgrading solid blades to hollow blades and further adopting hollow blades with more complex internal cooling channels. For example, a simple straight cooling channel convection heat exchange structure could be upgraded to a multi-channel structure with turbulent flow. Furthermore, cooling structures such as impingement, film cooling, and slit cooling could be added, or even more complex laminate cooling could be employed. Thermal barrier coatings could also be considered. Whether using materials with enhanced high-temperature and corrosion resistance or using blade skins with more complex cooling structures, engine cost will increase. Using existing engines, a relationship between the gas temperature range and the material quality and cooling structure of the turbine's key aerodynamic components is established (see Table 3). Table 3 lists gas turbine structural options for different gas temperature ranges. The relationship between power turbine structural options and gas temperature can also be established based on this. In Table 3, Temperature Range 1 < Temperature Range 2 < Temperature Range 3 < Temperature Range 4 < Temperature Range 5. Temperature Range 1 is used as the baseline. Within this baseline, the material level and cooling structure complexity of various components are denoted by A, while B, C, and D represent higher levels of material level and cooling structure complexity.

[0024] Table 3 Gas turbine structure schemes in different gas temperature ranges ; S33. Based on the differences in process cost, material cost, and yield rate among different turbine solutions, a calculation formula for the total cost coefficient of the turbine in different gas temperature ranges is established: 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 × C HPT,mat, k / C HPT,mat, 0,k )×Q HPT,k / Q HPT,k,0 ); Among them: HPT is the total cost coefficient of the gas turbine, that is, the cost ratio of the new solution to the engine baseline solution; α HPT,k,0 is the cost ratio of each part of the gas turbine of the engine baseline solution; α HPT, mat,k,0 C is the material cost ratio of the corresponding part of the gas turbine of the engine baseline solution; HPT,process,k / C HPT,process,0,k is the ratio of the gas turbine process cost to the engine benchmark solution, which can be determined based on statistical data; C HPT,mat, k / C HPT,mat, 0,k is the ratio of gas turbine material cost to engine base case, which can be determined based on statistical data; Q HPT,k / Q HPT,k,0 The gas turbine yield rate is the ratio of the engine benchmark solution, which can be determined based on statistical data; Power turbines generally use uncooled blades, which significantly impacts material costs. Calculations reveal the total cost coefficients for turbines in different gas temperature ranges, as shown in Tables 4 and 5.

[0025] Table 4 Total cost coefficient of gas turbine in different gas temperature ranges

[0026] Table 5 Total cost coefficient of power turbine in different gas temperature ranges

[0027] Preferably, similar to gas turbines, the materials, cooling structure, and coating scheme used in the combustion chamber flame liner are closely related to the combustion chamber outlet temperature; similar to compressors, the outer diameter and wall thickness of the combustion chamber casing, and the number of nozzles are affected by the inlet conversion flow rate. The combustion chamber casing and flame liner 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: S41. Based on the materials, cooling structure, and coating scheme used for 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, a calculation formula for the total cost coefficient of the combustion chamber is established: ζ comb =f(T4,α comb,k,0 , αcomb, mat,k,0 , C comb,process,k / C comb,process,0,k , C comb,mat, k / C comb,mat, 0,k ,Wac 31 / Wac 31,0 ,PI,Q comb,k / Q comb,k,0 ); Among them: comb is the total cost coefficient of the combustion chamber, that is, the cost ratio of the new solution to the engine baseline solution; T4 is the combustion chamber outlet temperature; α comb,k,0 is the cost ratio of each part of the combustion chamber of the engine baseline solution; α comb,mat,k,0 C is the material cost ratio of each part of the combustion chamber of the engine baseline solution; comb,process,k / C comb,process,0,k is the ratio of the process cost of each part of the combustion chamber to the engine benchmark solution; C comb,mat,k / C comb,mat,0,k is the ratio of the material cost of each part of the combustion chamber to the engine baseline solution; Wac 31 Wac is the converted flow rate at the combustion chamber inlet; 31,0 is the combustion chamber inlet conversion flow rate of the engine benchmark scheme PI is the compressor pressure ratio; Q comb,k / Q comb,k,0 It is the ratio of the yield rate of each part of the combustion chamber to the engine benchmark solution.

[0028] Preferably, the cost of the intake and exhaust devices is mainly affected by the size and is closely related to the compressor inlet conversion. Referring to the method for the compressor, step S5 specifically includes the following steps: S51. Based on the material cost ratio of the intake and exhaust devices of the engine baseline solution and the converted compressor inlet flow rate, establish the calculation formula for the total cost coefficient of the intake and exhaust devices: ζ inout =f(α inout,mat , Wac2); Among them: inout is the total cost coefficient of the combustion chamber, that is, the cost ratio of the new scheme to the engine benchmark scheme; α inout,mat is the material cost ratio of the intake and exhaust devices of the engine baseline solution; Wac2 is the converted flow rate at the compressor inlet.

[0029] Preferably, step S6 specifically includes the steps of: S61. Based on the total compressor cost coefficient, the compressor cost percentage of the engine baseline solution, the total gas turbine cost coefficient, the gas turbine cost percentage of the engine baseline solution, the total turbine cost coefficient, the total combustion chamber cost coefficient, and the total intake and exhaust device cost coefficients, establish the formula for calculating the engine manufacturing and periodic overhaul cost coefficient: 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 + (1-α comp,0 -α HPT,0 -α LPT,0 -α comb,0 -α inout,0 ); Among them: manu is the cost coefficient for engine manufacturing and periodic overhaul; comp is the total cost coefficient of the compressor; α comp,0 is the proportion of compressor cost of the engine baseline solution; HPT is the total cost coefficient of the gas turbine; α HPT,0 is the gas turbine cost ratio of the engine baseline solution; LPT is the total cost coefficient of the power turbine; α LPT,0 is the proportion of the power turbine cost of the engine baseline solution; comb is the total cost coefficient of the combustion chamber; α comb,0 is the combustion chamber cost ratio of the engine baseline solution; inout is the total cost coefficient of the intake and exhaust devices; α inout,0 The cost ratio of the intake and exhaust devices in the engine baseline solution.

[0030] Preferably, step S7 specifically includes the steps of: S71. Based on the engine manufacturing and periodic overhaul cost coefficients and fuel consumption rate, and the relationship between fuel cost and engine operating cost, establish the formula for calculating the engine operating cost coefficient: ζ eng =ζ manu ×α manu +SFC / SFC0×α SFC +(1-α manu -α SFC ); Among them: eng is the engine use cost coefficient; manu is the engine manufacturing and periodic overhaul cost coefficient; α manu is the ratio of manufacturing and periodic overhaul costs to engine operating costs, generally 40% to 50%; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline design point fuel consumption rate; α SFCIt is the ratio of fuel cost to engine operating cost, generally 45% to 55%.

[0031] Preferably, step S8 specifically includes the steps of: S81. Considering that the takeoff weight, payload, and power spectrum vary under different mission conditions of the flight platform, the weight distribution and power spectrum under different mission conditions of the flight platform 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 is shown in Table 6.

[0032] Table 6 Example of power spectrum provided by the flight platform (power normalized) ; S82. Based on the weight distribution of the flight platform under different mission conditions, calculate the weight of the flight platform that can be used to load the engine and fuel under different mission conditions as the remaining weight: m res =m fule +m engine =m to -m st -m pay ; Where: m res is the remaining weight of the flight platform; m fule is the fuel weight of the flight platform; m engine is the engine weight; m to is the take-off weight of the flight platform; m st is the weight of the flight platform body structure; m pay is the commercial load of the flight platform; S83. Based on the design point performance model of the baseline engine solution, using the component characteristic diagrams generated by the preliminary design, establish an engine off-design point performance model. Based on the power spectrum under different mission conditions of the flight platform, input altitude and power into the off-design point performance model, perform off-design point calculations, and obtain the engine fuel consumption rate and fuel flow rate for each flight phase. Combined with the duration of each flight phase, calculate the required fuel weight: m fule,de,0 =∑(SFC j ×P dn,j ×t j ); Where: m fule,de,0 Fuel weight required for the engine base case; SFC j is the fuel consumption rate of the engine benchmark scheme in each flight phase; P dn,j is the power of 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 solution, and the required fuel weight, and taking into account a certain residual fuel coefficient, 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 coefficient: δ res = m res / ( m engine,0+ m fule,de,0 ×β); Where: res is the residual weight coefficient of the flight platform; m res is the remaining weight of the flight platform; m engine,0 is the engine base solution weight; m fule,de,0 is the required fuel weight of the engine baseline solution; β is the residual fuel coefficient, which ranges from 10% to 20%. The minimum value of the residual weight coefficient is selected to represent the minimum ratio of the engine baseline solution weight to the required fuel weight of the flight platform residual weight, and the ratio of the engine baseline solution weight to the flight platform residual weight under this condition is recorded. S85. Although factors such as the number of compressor stages and selected materials may affect engine weight, the converted compressor 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 solution to the weight of the base engine solution: δ engine =m engine / m engine,0 ; Where: engine is the engine weight coefficient; m engine is the engine weight; m engine,0 is the engine base solution weight; Based on the existing engine data, the relationship between the engine weight coefficient and the compressor conversion flow rate is established (see Figure 6 ): δ engine = a5×Wac2 2 +b5×Wac2 1 +c5; Where: engine is the engine weight coefficient; Wac2 is the compressor inlet conversion flow; a5, b5, c5 are coefficients, determined according to the curve fitting results; S86. Based on the changes in engine weight and fuel consumption, calculate the minimum remaining weight coefficient corresponding to the new engine solution: δ res,min =(δ engine ×γ+SFC / SFC0×(1-γ))×δ res,min,0 ; Where:res,min is the minimum residual weight coefficient of the flight platform; δ engine is the engine weight coefficient; γ is the ratio of the engine weight to the residual weight under the conditions of the engine baseline solution and the minimum residual weight coefficient of the flight platform; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline solution design point fuel consumption rate; δ res,min,0 is the minimum residual weight coefficient of the flight platform corresponding to the engine baseline solution.

[0033] In the above embodiment, step S9 substitutes the coefficients or formulas obtained in steps S2 to S8 into the design point performance model, gives the range and step size of the combustion chamber outlet temperature and pressure ratio, and performs iterative calculation to obtain the distribution of the engine use cost coefficient and the minimum residual weight coefficient of the flight platform on the thermodynamic cycle parameter analysis diagram, as shown in detail. 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 the engine use cost, and the ,minimum residual weight coefficient of the flight platform is still within ,an acceptable range.

[0034] The method proposed in this application has been applied to 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 improved indicators such as specific power and fuel consumption, but also made the engine cost more competitive and further increased the range of the flight platform.

[0035] like Figure 9 As shown, another preferred embodiment of the present application further provides a method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements, comprising: Cost ratio determination module, used to establish the engine benchmark scheme and its design point performance model, and determine the cost ratio of each component; The compressor total cost coefficient calculation module is used to determine the compressor structure scheme suitable for different pressure ratios, establish the cost calculation formula for the compressor rotor blade disk and compressor stator blade assembly, and then establish the compressor total cost coefficient calculation formula based on the cost ratio relationship of the engine benchmark scheme; The turbine total cost coefficient calculation module is used to establish a calculation formula for the turbine total cost coefficient in different gas temperature ranges based on the differences in process cost, material cost and yield rate of gas turbine structural schemes in different gas temperature ranges; The combustion chamber total cost coefficient calculation module is used to establish a calculation formula for the combustion chamber total cost coefficient based on the material, cooling structure and coating scheme used in 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; The module for calculating the total cost coefficient of the intake and exhaust devices is used to establish a calculation formula for the total cost coefficient of the intake and exhaust devices based on the material cost ratio of the intake and exhaust devices of the engine baseline scheme and the compressor inlet conversion flow rate; The engine manufacturing and periodic overhaul cost coefficient calculation module is used to establish a calculation formula for the engine manufacturing and periodic overhaul cost coefficient based on the total compressor cost coefficient, the compressor cost ratio of the engine baseline solution, the total gas turbine cost coefficient, the gas turbine cost ratio of the engine baseline solution, the total turbine cost coefficient, the total combustion chamber cost coefficient, and the total cost coefficients of the intake and exhaust devices; The engine cost coefficient calculation module is used to establish an engine 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 cost; The applicability assessment module is used to calculate the residual weight, engine fuel weight, and minimum residual weight coefficient under different mission conditions based on the weight distribution and power spectrum of the flight platform under different mission conditions, and complete the applicability assessment of the thermodynamic cycle parameters relative to the flight platform mission; 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 combustion chamber outlet temperature and pressure ratio, and carry out iterative calculation to obtain the distribution of the engine usage cost coefficient and the minimum residual weight coefficient of the flight platform on the thermodynamic cycle parameter analysis diagram.

[0036] The turboshaft engine thermodynamic cycle parameter evaluation device based on cost and mission requirements provided in this application adopts the turboshaft engine thermodynamic cycle parameter evaluation method based on cost and mission requirements in the above-mentioned embodiment, which can solve the technical problem that the existing thermodynamic cycle parameter design does not give sufficient comprehensive consideration to cost and mission requirements and does not meet the economic requirements of the flight platform. Compared with the existing technology, the beneficial effects of the turboshaft engine thermodynamic cycle parameter evaluation device based on cost and mission requirements provided in this application are the same as the beneficial effects of the turboshaft engine thermodynamic cycle parameter evaluation method based on cost and mission requirements provided in the above-mentioned embodiment, and the other technical features of the turboshaft engine thermodynamic cycle parameter evaluation device based on cost and mission requirements are the same as the features disclosed in the above-mentioned embodiment method, and are not further described here.

[0037] like Figure 10 As shown, a preferred embodiment of the present application also provides an electronic device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor 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 embodiment when executing the computer program.

[0038] The electronic device provided in the application utilizes the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method described in the aforementioned embodiment. This device can address the technical issue that existing thermodynamic cycle parameter designs fail to comprehensively consider cost and mission requirements, thus failing to meet the economic requirements of flight platforms. Compared to the prior art, the beneficial effects of the electronic device provided in this application are the same as those of the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method described in the aforementioned embodiment. Other technical features of the electronic device are the same as those disclosed in the aforementioned embodiment and are not further elaborated here.

[0039] like Figure 11 As shown, the preferred embodiment of the present application further provides a computer device, which can be a terminal or a liveness detection server, and its internal structure diagram can be as shown in FIG. Figure 11 As shown. The computer device includes a processor, a memory and a network interface connected via a system bus. The processor of the computer device is used to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system and a computer program. The internal memory provides an environment for the operation of the operating system and the computer program in the non-volatile storage medium. The network interface of the computer device is used to communicate with other external computer devices via a network connection. When the computer program is executed by the processor, the steps of the above-mentioned method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements are implemented.

[0040] Those skilled in the art will understand that Figure 11 The structure shown in the figure is only a block diagram of a part of the structure related to the solution of the present application, and does not constitute a limitation on the computer device to which the solution of the present application is applied. The specific computer device may include more or fewer components than shown in the figure, or combine certain components, or have a different component arrangement.

[0041] The computer device provided in this application utilizes the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method described in the aforementioned embodiment, resolving the technical issue that existing thermodynamic cycle parameter designs fail to comprehensively consider cost and mission requirements, thus failing to meet the economic requirements of flight platforms. Compared to the prior art, the beneficial effects of the computer device provided in this application are the same as those of the cost- and mission-requirement-based turboshaft engine thermodynamic cycle parameter evaluation method described in the aforementioned embodiment. Other technical features of the electronic device are the same as those disclosed in the aforementioned embodiment and are not further elaborated here.

[0042] A preferred embodiment of the present application also provides a storage medium, which includes a stored program, and when the program is running, controls the device where the storage medium is located to execute the steps of the turboshaft engine thermodynamic cycle parameter evaluation method based on cost and mission requirements in the above embodiment.

[0043] It should be noted that the steps shown in the flowcharts of the accompanying drawings can be executed in a computer system such as a set of computer-executable instructions, and that, although a logical order is shown in the flowcharts, in some cases, the steps shown or described can be executed in an order different from that shown here.

[0044] If the functions described in the method of this embodiment are implemented in the form of a software functional unit and sold or used as an independent product, they can be stored in a storage medium readable by one or more computing devices. Based on this understanding, the part of the embodiment of this application that contributes to the prior art or the part of the technical solution can be embodied in the form of a software product, which is stored in a storage medium and includes a number of instructions for enabling a computing device (which can be a personal computer, server, mobile computing device or network device, etc.) to execute all or part of the steps of the method described in each embodiment of this application. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (ROM, Read-Only Memory), random access memory (RAM, Random Access Memory), disk or optical disk, and other media that can store program code.

[0045] Those skilled in the art will appreciate that the embodiments of the present application may be provided as methods, systems, or computer program products. Therefore, the present application may take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware. Furthermore, the present application may 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 the present application may be implemented in various computer languages, for example, the object-oriented programming language C++ and the embedded programming language C.

[0046] The present application is described with reference to the flowcharts and / or block diagrams of the methods, devices (systems), and computer program products according to the embodiments of the present application. It should be understood that each process and / or block in the flowchart and / or block diagram, as well as the combination of processes and / or blocks in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 a process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0047] These computer program instructions may also be stored in a computer readable memory that can direct a computer or other programmable data processing device to work in a specific manner, so that the instructions stored in the computer readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 a process or multiple processes and / or boxes Figure 1 The function specified in one or more boxes.

[0048] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operational steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing the instructions executed on the computer or other programmable device for implementing the process. Figure 1 a process or multiple processes and / or boxes Figure 1 A step that specifies a function in one or more boxes.

[0049] The present application also provides a computer program product, comprising a computer program, which, when executed by a processor, implements the steps of the method for evaluating the thermodynamic cycle parameters of a turboshaft engine based on cost and mission requirements as described above.

[0050] The computer program product provided in this application can address 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 to existing technologies, the beneficial effects of the computer program product provided in this application are similar to those of the cost- and mission-based turboshaft engine thermodynamic cycle parameter evaluation method provided in the aforementioned embodiment, and are not further elaborated here.

[0051] Although the preferred embodiments of the present application have been described, those skilled in the art may make additional changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present application.

[0052] Obviously, those skilled in the art may make various changes and modifications to this application without departing from the spirit and scope of this application. Thus, if these modifications and variations of this application fall within the scope of the claims of this application and their equivalents, this application is intended to include these 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 steps: S1. Establish an engine benchmark scheme and its design point performance model, and determine the cost ratio of each component; S2. Determine the compressor structure schemes suitable for different pressure ratios, establish cost calculation formulas for the compressor rotor blade disk and compressor stator blade assembly, and then establish a calculation formula for the total compressor cost coefficient based on the cost ratio relationship of the engine baseline scheme; S3. Based on the differences in process cost, material cost, and yield rate of gas turbine structural schemes in different gas temperature ranges, a calculation formula for the total cost coefficient of turbines in different gas temperature ranges is established; S4. Establish a calculation formula for the total cost coefficient of the combustion chamber based on the materials, cooling structure and coating scheme used for 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; S5. Establish a calculation formula for the total cost coefficient of the intake and exhaust devices based on the material cost ratio of the intake and exhaust devices of the engine baseline solution and the converted compressor inlet flow rate; S6. Establish a calculation formula for the engine manufacturing and periodic overhaul cost coefficient based on the total compressor cost coefficient, the engine baseline compressor cost ratio, the gas turbine total cost coefficient, the engine baseline gas turbine cost ratio, the turbine total cost coefficient, the combustion chamber total cost coefficient, and the intake and exhaust device total cost coefficients; S7. Establish a formula for calculating the engine cost coefficient based on the engine manufacturing and periodic overhaul cost coefficients and fuel consumption rate, combined with the relationship between fuel cost and engine cost; S8. Based on the weight distribution and power spectrum of the flight platform under different mission conditions, calculate the residual weight, engine fuel weight, and minimum residual weight coefficient under different mission conditions, and complete the suitability assessment of the thermodynamic cycle parameters for 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 combustion chamber outlet temperature and pressure ratio, and perform iterative calculations to obtain the distribution of the engine use cost coefficient and the flight platform minimum residual 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 perform optimization analysis of the thermodynamic cycle parameters of the engine design point, so that the power and fuel consumption rate of the engine design point meet the index requirements, and establish an engine benchmark scheme and its design point performance model; S12. Based on the existing engine baseline plan and cost situation, combined with supplier inquiry results, determine the proportion of manufacturing costs 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, wherein step S2 specifically comprises the following steps: S21. By collecting statistics on compressor data of existing engines, compressor configurations applicable to different pressure ratio ranges can be obtained, and compressor structural solutions applicable to different pressure ratios can be determined; S22. Based on existing engine data, establish the relationship between the compressor inlet outer diameter and the compressor inlet converted flow rate: D com =a1×Wac2 2 +b1×Wac2+c1; in, D com is the compressor inlet outer diameter; Wac2 is the compressor inlet converted flow rate; a1, b1, c1 are coefficients, determined according to the curve fitting results; S23. Based on the changes in the compressor outer diameter, number of blades, and blade material, a calculation formula for the compressor blade cost coefficient is established to obtain the change in the compressor blade of the new scheme relative to the engine baseline scheme: ; Among them: comp,disc is the cost coefficient of the compressor rotor blade disk, which is the cost ratio of the new compressor blade disk to the engine benchmark solution; α comp,dis,mat,0 The material cost ratio of the compressor rotor blade disk of the engine baseline solution is generally 20% to 35%; D com is the compressor inlet outer diameter; D com,0 is the compressor inlet outer diameter of the engine base scheme; N comp,disc is the number of compressor rotor blades; N comp,disc,0 is the number of compressor rotor blades in the engine baseline solution; N comp,dis,GH is the number of compressor high-temperature alloy rotor blade disks; N comp,dis,GH,0 C is the number of high-temperature alloy rotor blades in the engine baseline scheme compressor; mat,GH / C mat,Ti The cost ratio of high-temperature alloys to titanium alloys takes into account the difference in material density and unit price; S24. Based on the influence of the number of stages and outer diameter changes of the compressor stator blade assembly, a calculation formula for the cost coefficient of the compressor stator blade assembly is established: 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 ); Among them: comp,stator is the cost coefficient of the compressor stator blade assembly, that is, the cost ratio of the new solution to the engine benchmark solution; α comp,stator,mat,0 D is the material cost ratio of the compressor stator blade assembly of the engine baseline solution; com is the compressor inlet outer diameter; D com,0 is the compressor inlet outer diameter of the engine base scheme; N comp is the number of compressor stages; N comp,0 is the number of compressor stages of the engine benchmark scheme; C comp,sator,mid,0 / C comp,sator, 0 The ratio of the cost of a single intermediate stage stator 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, combined with the cost ratio of the engine baseline solution, establish a calculation formula for the total compressor cost coefficient: 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 ; Among them: comp is the total cost coefficient of the compressor; C comp,disc,0 is the cost of the compressor rotor blade disk of the engine baseline solution; C comp,sator, 0 is the cost of the compressor stator blade assembly of the engine baseline solution; comp,disc is the cost coefficient of the compressor rotor blade disk ζ comp,stato is the cost coefficient of the compressor stator blade assembly.

4. The method for evaluating turboshaft engine thermodynamic cycle parameters based on cost and mission requirements according to claim 1, wherein step S3 specifically comprises the following steps: S31. Utilizing existing engine technical solutions, establish the relationship between the ratio of bleed air flow rates of various turbine components to the engine inlet flow rate and the combustion chamber outlet temperature, including the combustion-pilot air ratio, the combustion-dynamic bleed air ratio, and the power turbine bleed air ratio. The combustion-pilot air ratio refers to the bleed air ratio at the first stage guide vane of the gas turbine; the combustion-dynamic bleed air ratio refers to the bleed air ratio at the gas turbine rotor blades; and the power turbine bleed air ratio refers to the bleed air ratio at the power turbine. S32. With reference to existing engines, establish the relationship between the gas temperature range and the material level and cooling structure of the main aerodynamic components of the turbine, and determine the structural scheme applicable to different gas temperatures at the inlet of the gas turbine and power turbine, including: The inlet gas temperature of a gas turbine refers to the combustion chamber outlet temperature, and the inlet gas temperature of a power turbine refers to the gas turbine outlet temperature; S33. Based on the differences in process cost, material cost, and yield rate among different turbine solutions, a calculation formula for the total cost coefficient of the turbine in different gas temperature ranges is established: 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 ); Among them: HPT is the total cost coefficient of the gas turbine, that is, the cost ratio of the new solution to the engine baseline solution; α HPT,k,0 is the cost ratio of each part of the gas turbine of the engine baseline solution; α HPT, mat,k,0 C is the material cost ratio of the corresponding part of the gas turbine of the engine baseline solution; HPT,process,k / C HPT,process,0,k is the ratio of the gas turbine process cost to the engine benchmark solution, which can be determined based on statistical data; C HPT,mat, k / C HPT,mat, 0,k is the ratio of gas turbine material cost to engine base case, which can be determined based on statistical data; Q HPT,k / Q HPT,k,0 It is the ratio of the gas turbine yield rate to the engine benchmark solution, which 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, wherein in step S31, the formula for calculating the fuel-pilot air ratio is: CO NGV = a2×T4 3 +b2×T4 2 +c2×T4+d2; in, CO NGV is the fuel-pilot gas ratio; T4 is the combustion chamber outlet temperature; a2, b2, c2, d2 are coefficients, determined according to the curve fitting results; The calculation formula for the combustion and bleed air ratio is: WHAT HPTR = a3×T4 3 +b3×T4 2 +c3×T4+d3; Among them, CO HPTR is the combustion air ratio; T4 is the combustion chamber outlet temperature; a3, b3, c3, d3 are coefficients, determined according to the curve fitting results; The calculation formula for the power turbine bleed air ratio is: WHAT LPTR = a4×T4 3 +b4×T4 2 +c4×T4+d4; Among them, CO LPTR is the power turbine bleed air ratio; T4 is the combustion chamber outlet temperature; a4, b4, c4, d4 are coefficients, which are determined according to 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 used for 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, a calculation formula for the total cost coefficient of the combustion chamber is established: 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 ,Wac 31 / Wac 31,0 ,PI,Q comb,k / Q comb,k,0 ); Among them: comb is the total cost coefficient of the combustion chamber, that is, the cost ratio of the new solution to the engine benchmark solution; T4 is the combustion chamber outlet temperature; α comb,k,0 is the cost ratio of each part of the combustion chamber of the engine baseline solution; α comb,mat,k,0 C is the material cost ratio of each part of the combustion chamber of the engine baseline solution; comb,process,k / C comb,process,0,k is the ratio of the process cost of each part of the combustion chamber to the engine benchmark solution; C comb,mat,k / C comb,mat,0,k Wac is the ratio of the material cost of each part of the combustion chamber to the engine baseline solution; 31 Wac is the converted flow rate at the combustion chamber inlet; 31,0 is the combustion chamber inlet conversion flow rate of the engine benchmark scheme PI is the compressor pressure ratio; Q comb,k / Q comb,k,0 It is the ratio of the yield rate of each part of the combustion chamber to the engine benchmark solution.

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 devices of the engine baseline solution and the converted compressor inlet flow rate, establish the calculation formula for the total cost coefficient of the intake and exhaust devices: g inout =f(a inout,mat ,Wac2); Among them: inout is the total cost coefficient of the combustion chamber, that is, the cost ratio of the new scheme to the engine benchmark scheme; α inout,mat is the material cost ratio of the intake and exhaust devices of the engine baseline solution; 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 total compressor cost coefficient, the compressor cost percentage of the engine baseline solution, the total gas turbine cost coefficient, the gas turbine cost percentage of the engine baseline solution, the total turbine cost coefficient, the total combustion chamber cost coefficient, and the total intake and exhaust device cost coefficients, 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 )； Among them: manu is the cost coefficient for engine manufacturing and periodic overhaul; comp is the total cost coefficient of the compressor; α comp,0 is the proportion of compressor cost of the engine baseline solution; HPT is the total cost coefficient of the gas turbine; α HPT,0 is the gas turbine cost ratio of the engine baseline solution; LPT is the total cost coefficient of the power turbine; α LPT,0 is the proportion of the power turbine cost of the engine baseline solution; comb is the total cost coefficient of the combustion chamber; α comb,0 is the combustion chamber cost ratio of the engine baseline solution; inout is the total cost coefficient of the intake and exhaust devices; α inout,0 The cost ratio of the intake and exhaust devices in the engine baseline solution.

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 the relationship between fuel cost and engine operating cost, establish the formula for calculating the engine operating cost coefficient: g eng =ζ manu ×a manu +SFC / SFC0×α SFC +(1-a manu -a SFC ); Among them: eng is the engine use cost coefficient; manu is the engine manufacturing and periodic overhaul cost coefficient; α manu is the ratio of manufacturing and periodic overhaul costs to engine operating costs, generally 40% to 50%; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline design point fuel consumption rate; α SFC It is the ratio of fuel cost to engine operating cost, generally 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 takeoff, climb, cruise, descent, landing, and flight reserve phases. S82. Based on the weight distribution of the flight platform under different mission conditions, calculate the weight of the flight platform that can be used to load the engine and fuel under different mission conditions as the remaining weight: m res =m fule +m engine =m to -m st -m pay ; Where: m res is the remaining weight of the flight platform; m fule is the fuel weight of the flight platform; m engine is the engine weight; m to is the take-off weight of the flight platform; m st is the weight of the flight platform body structure; m pay is the commercial load of the flight platform; S83. Based on the design point performance model of the baseline engine solution, using the component characteristic diagrams generated by the preliminary design, establish an engine off-design point performance model. Based on the power spectrum under different mission conditions of the flight platform, input altitude and power into the off-design point performance model, perform off-design point calculations, and obtain the engine fuel consumption rate and fuel flow rate for each flight phase. Combined with the duration of each flight phase, calculate the required fuel weight: m fule,de,0 =∑(SFC j ×P dn,j ×t j ); Where: m fule,de,0 Fuel weight required for the engine base case; SFC j is the fuel consumption rate of the engine benchmark scheme in each flight phase; P dn,j is the power of 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 solution, and the required fuel weight, and taking into account a certain residual fuel coefficient, 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 coefficient: δ res = m res / ( m engine,0+ m fule,de,0 ×β); Where: res is the residual weight coefficient of the flight platform; m res is the remaining weight of the flight platform; m engine,0 is the engine base solution weight; m fule,de,0 is the fuel weight required for the engine baseline solution; β is the residual fuel coefficient, which ranges from 10% to 20%; S85. Determine the engine weight factor based on the ratio of the weight of the new engine solution to the weight of the base engine solution: d engine =m engine / m engine,0 ; Where: engine is the engine weight coefficient; m engine is the engine weight; m engine,0 is the engine base solution weight; Based on the existing engine data, the relationship between the engine weight coefficient and the compressor conversion flow is established: δ engine = a5×Wac2 2 +b5×Wac2 1 +c5; Where: engine is the engine weight coefficient; Wac2 is the compressor inlet conversion flow; a5, b5, c5 are coefficients, determined according to the curve fitting results; S86. Based on the changes in engine weight and fuel consumption, calculate the minimum remaining weight coefficient corresponding to the new engine solution: d res,min =(δ engine ×γ+SFC / SFC0×(1-γ))×δ res,min,0 ; Where: res,min is the minimum residual weight coefficient of the flight platform; δ engine is the engine weight coefficient; γ is the ratio of the engine weight to the residual weight under the conditions of the engine baseline solution and the minimum residual weight coefficient of the flight platform; SFC is the engine design point fuel consumption rate; SFC0 is the engine baseline solution design point fuel consumption rate; δ res,min,0 is the minimum residual weight coefficient of the flight platform corresponding to the engine baseline solution.

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