Method and system for obtaining thermodynamic cycle parameters of a low-cost configuration turbine engine

Abstract

The present invention belongs to the field of turbine engines and provides a method and system for obtaining thermodynamic cycle parameters for a low-cost turbine engine. The method comprises: determining an engine configuration based on a cost design target for the turbine engine; obtaining a feasible domain of thermodynamic cycle parameters based on the engine configuration, using overall performance indicators and fan inlet size thresholds as constraints. The overall performance indicators include thrust and fuel consumption; and determining thermodynamic cycle parameters within the feasible domain of thermodynamic cycle parameters based on turbine engine performance requirements and the turbine engine's resistance to intake distortion. The method of the present invention rapidly obtains thermodynamic cycle parameters without requiring repeated iterations, significantly shortening the design cycle.

Description

Technical Field

[0001] The present invention belongs to the field of turbine engines and relates to turbine engine performance design technology, and in particular to a method and system for obtaining thermodynamic cycle parameters of a low-cost turbine engine. Background Art

[0002] Thermodynamic cycle parameters of turbine engines include overall cycle parameters (including fan pressure ratio π f , total pressure ratio ∑π or compressor pressure ratio π c , total temperature at combustion chamber outlet T4, bypass ratio B, fan inlet flow rate W2 or compressor inlet flow rate W 25 ), component performance parameters (including efficiency, total pressure recovery coefficient, etc.), and air system bleed volume. The selection of thermodynamic cycle parameters is the most important step in the overall engine performance design, determining the quality of engine performance.

[0003] Currently, the existing analysis method for thermodynamic cycle parameters is to first determine the thermodynamic cycle parameters and the aerodynamic parameters of each component based on thrust, fuel consumption, size, and other requirements. The engine configuration is then determined based on the aerodynamic parameters of each component. This method allows for a wide range of selectable thermodynamic cycle parameters, potentially leading to a large number of turbine stages and a complex engine configuration. However, due to the high cost requirements of short-life turbine engines, this method may not meet these requirements. Repeated iterations are required to determine an overall performance solution and engine configuration that meets these requirements, significantly increasing the design cycle.

[0004] Therefore, it is necessary to design an efficient method for obtaining the thermodynamic cycle parameters of short-life turbine engines that meets low-cost requirements. Summary of the Invention

[0005] In order to solve the technical problems of existing thermodynamic cycle parameter analysis methods, such as failure to meet the low-cost requirements of engines, slow iterative convergence, and long iteration cycles, the present invention discloses a method for obtaining thermodynamic cycle parameters of a low-cost turbine engine, the method comprising the following steps:

[0006] S1. Determine the engine configuration based on the cost design target of the turbine engine;

[0007] S2. Obtaining a feasible region of thermodynamic cycle parameters based on the engine configuration and taking overall performance indicators and a fan inlet size threshold as constraints, wherein the overall performance indicators include a thrust indicator and a fuel consumption rate indicator;

[0008] S3. Determine the thermodynamic cycle parameters within the feasible domain of the thermodynamic cycle parameters based on the turbine engine performance requirements and the turbine engine's resistance to intake distortion.

[0009] Furthermore, in the above step S1, determining the engine configuration based on the cost design target of the turbine engine includes:

[0010] S11. Based on the user's cost design target for the turbine engine, a turbine type, a compressor bleed type, and a compressor guide vane type are obtained. The compressor bleed type includes bleed and non-bleed, the compressor guide vane type includes adjustable guide vanes and non-adjustable guide vanes, and the turbine type includes an integrally cast turbine and a welded turbine.

[0011] S12. Using an engine cost model, according to the cost design target, the compressor bleed type, the compressor guide vane type, and the turbine type, obtain the number of fan stages, the number of compressor stages, and the number of turbine stages.

[0012] Furthermore, in the above step S2, according to the engine configuration, the feasible region of thermodynamic cycle parameters is obtained with the overall performance index and the fan inlet size threshold as constraints, including:

[0013] S21. Designing a feasible region of fan pressure ratio and fan efficiency according to the number of fan stages of the engine configuration;

[0014] S22. Designing a compressor pressure ratio and a compressor efficiency feasible region based on the fan pressure ratio, the number of compressor stages, the compressor bleed type, and the compressor guide vane type of the engine configuration;

[0015] S23. Designing a feasible region for the total temperature at the combustion chamber outlet based on the turbine type of the engine configuration and subjecting the total temperature at the combustion chamber outlet to a constraint of being less than the processing temperature of the turbine type;

[0016] S24. Obtaining a preliminary air intake volume of the air system using the feasible region of the total temperature at the combustion chamber outlet and the number of turbine stages of the engine configuration;

[0017] S25, calculating the total pressure ratio according to the fan pressure ratio, the compressor pressure ratio, and the total pressure recovery coefficient of the intermediate casing, and obtaining a feasible region of the bypass ratio according to the total pressure ratio and using the fuel consumption index as a constraint;

[0018] S26. Obtain a fan inlet flow rate feasible region based on the fan pressure ratio feasible region, the compressor pressure ratio feasible region, the combustion chamber outlet total temperature feasible region, and the bypass ratio feasible region, with the thrust index and the fan inlet size threshold as constraints.

[0019] Furthermore, in the above steps S24 and S26, the initial air intake volume of the air system is proportional to the feasible range of the total temperature at the combustion chamber outlet and the number of turbine stages; the fan inlet flow rate is proportional to the thrust index and inversely proportional to the fan inlet size.

[0020] Furthermore, in the above step S3, the thermodynamic cycle parameters are determined within the feasible region of the thermodynamic cycle parameters based on the turbine engine performance requirements and the turbine engine's resistance to intake distortion, including:

[0021] S31. Obtaining an overall performance index design margin based on the turbine engine performance requirement, and obtaining a compression component surge margin using the turbine engine's anti-intake distortion method;

[0022] S32. Based on the feasible domain of the thermodynamic cycle parameters, using the engine overall performance design model, obtain the ratio of the total pressure at the mixer's outer inlet to the total pressure at the inner inlet;

[0023] S33. Determine thermodynamic cycle parameters within the thermodynamic cycle parameter feasible domain according to the compression component surge margin, the overall performance index design margin, and the ratio.

[0024] Furthermore, in the above step S31, obtaining the overall performance index design margin according to the turbine engine performance requirement analysis result includes:

[0025] S311. Analyze turbine engine performance requirements and determine the priority order of thrust and fuel consumption.

[0026] S312. When the priority of thrust is higher than the priority of fuel consumption rate, obtaining a thrust design margin with the fuel consumption rate being equal to the fuel consumption rate index as a constraint and the thrust being equal to a given multiple of the thrust index as a goal;

[0027] S313. When the priority of fuel consumption rate is higher than the priority of thrust, a fuel consumption rate design margin is obtained with the thrust being equal to the thrust index as a constraint and the fuel consumption rate being equal to a given multiple of the fuel consumption rate index as a goal.

[0028] Preferably, in the above steps S312 and S313, the given multiple of the thrust index is 1.05 to 1.10 times, and the given multiple of the fuel consumption rate index is 0.90 to 0.95 times.

[0029] An embodiment of the present invention also provides a thermodynamic cycle parameter acquisition system for a low-cost turbine engine, comprising an engine configuration building module, a feasible domain design module, and a thermodynamic cycle parameter determination module.

[0030] The engine configuration building module is used to determine the engine configuration according to the cost design target of the turbine engine;

[0031] The feasible region design module is used to obtain a feasible region of thermodynamic cycle parameters according to the engine configuration, with the overall performance index and the fan inlet size threshold as constraints, wherein the overall performance index includes a thrust index and a fuel consumption index;

[0032] The thermodynamic cycle parameter determination module is used to determine the thermodynamic cycle parameters within the thermodynamic cycle parameter feasible domain according to the turbine engine performance requirements and the turbine engine anti-intake distortion.

[0033] The present invention's thermodynamic cycle parameter acquisition method first analyzes low-cost requirements to determine an engine configuration. This engine configuration is then combined with overall performance indicators (thrust and fuel consumption) and a fan inlet size threshold to determine the feasible domain (i.e., range) of the thermodynamic cycle parameters. Finally, the final thermodynamic cycle parameters are derived from this feasible domain, taking into account turbine engine performance requirements and the turbine engine's resistance to inlet distortion. This method rapidly obtains thermodynamic cycle parameters without requiring repeated iterations, significantly shortening the design cycle. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without any creative work.

[0035] Figure 1 This is a flow chart of a method for obtaining thermodynamic cycle parameters of a low-cost turbine engine disclosed in an embodiment of the present invention;

[0036] Figure 2 This is a schematic diagram of the relationship between the air intake volume of the air system disclosed in an embodiment of the present invention, the total temperature at the combustion chamber outlet, and the number of turbine stages;

[0037] Figure 3 A schematic diagram illustrating the effects of compressor pressure ratio and combustion chamber outlet total temperature on thrust and fuel consumption according to an embodiment of the present invention;

[0038] Figure 4 This is an architectural diagram of a thermodynamic cycle parameter acquisition system for a low-cost turbine engine according to an embodiment of the present invention;

[0039] Among them, 401 is an engine configuration construction module; 402 is a feasible domain design module; and 403 is a thermodynamic cycle parameter determination module. DETAILED DESCRIPTION

[0040] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0041] The following describes the embodiments of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that, in the absence of conflict, the features of the following embodiments and embodiments can be combined with each other. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without making creative work are within the scope of protection of this application.

[0042] The embodiment of the present invention discloses a method for obtaining thermodynamic cycle parameters of a low-cost turbine engine. Figure 1 As shown, the method includes the following steps:

[0043] S1. Determine the engine configuration based on the cost design target of the turbine engine;

[0044] S2. Obtaining a feasible region of thermodynamic cycle parameters based on the engine configuration and taking overall performance indicators and a fan inlet size threshold as constraints, wherein the overall performance indicators include a thrust indicator and a fuel consumption rate indicator;

[0045] S3. Determine the thermodynamic cycle parameters within the feasible domain of the thermodynamic cycle parameters based on the turbine engine performance requirements and the turbine engine's resistance to intake distortion.

[0046] Furthermore, in the above step S1, determining the engine configuration based on the cost design target of the turbine engine includes:

[0047] S11. Based on the user's cost design objectives for the turbine engine, the turbine type, compressor bleed type, and compressor guide vane type are obtained. The compressor bleed type includes bleed and non-bleed, and the compressor guide vane types include adjustable guide vanes and non-adjustable guide vanes. During design, to reduce the cost of short-life turbine engines, a compressor with no bleed and non-adjustable guide vanes is preferably selected. This also helps reduce engine weight, and the lack of adjustable guide vanes improves operational reliability. Turbine types include integrally cast turbines and welded turbines. Integrally cast turbines are manufactured through an integrated molding process, which significantly reduces turbine cost.

[0048] S12. Using the engine cost model, the number of fan stages, compressor stages, and turbine stages is determined based on the cost design target, the compressor bleed type, the compressor guide vane type, and the turbine type. During design, to minimize engine cost, the target cost is decomposed and the number of impeller stages (including fan, compressor, and turbine) is determined based on the principle of minimizing the number of stages. Determining the number of stages significantly narrows the range of options for parameters such as the fan and compressor pressure ratios, facilitating rapid convergence on a low-cost overall performance solution.

[0049] Furthermore, in the above step S2, according to the engine configuration, the feasible region of thermodynamic cycle parameters is obtained with the overall performance index and the fan inlet size threshold as constraints, including:

[0050] S21. Design the fan pressure ratio and fan efficiency feasible domain according to the number of fan stages of the engine configuration. During implementation, the fan pressure ratio and fan efficiency feasible domain can be determined according to the number of fan stages and the existing technical level of the fan. The functional relationship between the fan pressure ratio and the number of fan stages can be expressed as π f =f1(J f ), where π f is the fan pressure ratio (its value range is π f1 π f2 π f3 ...π fn ), J f is the number of fan stages. The functional relationship between fan efficiency, fan stages and fan pressure ratio can be expressed as η f =f2(J f ,π f ), where η f is the fan efficiency, and its value range can be expressed as [η f1 η f2 η f3 ...η fn ].

[0051] S22. Design the compressor pressure ratio and compressor efficiency feasible domain based on the fan pressure ratio, the number of compressor stages, the compressor bleed type, and the compressor guide vane type of the engine configuration. During implementation, determine the preliminary range of the compressor pressure ratio and efficiency based on the fan pressure ratio, the number of compressor stages, and whether the guide vanes are adjustable and bleed. The functional relationship between the compressor pressure ratio, the compressor bleed type, the compressor guide vane type, and the number of compressor stages can be expressed as π c =f3(J c ,π f , KT, FQ), where π c is the compressor pressure ratio (the feasible range is [π c1 π c2 π c3 ...π cn]), J c is the number of compressor stages, KT is the type of compressor guide vanes, and FQ is the type of compressor bleed air. The functional relationship between compressor efficiency, compressor pressure ratio, and number of compressor stages can be expressed as η c =f4(J c ,π c ), where η c is the compressor efficiency (the feasible range is [η c1， η c2， η c3， ...η cn ]).

[0052] S23. Based on the turbine type of the engine configuration, and with the total temperature at the combustion chamber outlet being less than the processing temperature of the turbine type as a constraint, design the feasible region of the total temperature at the combustion chamber outlet. The feasible region of the total temperature T4 at the combustion chamber outlet can be expressed as [T 41 T 42 T 43 ...T 4n ].

[0053] S24, using the feasible range of the total temperature at the combustion chamber outlet and the turbine stage number of the engine configuration, obtain the preliminary air intake volume of the air system. Figure 2 As shown in the figure, the air system bleed air volume is determined according to the feasible region of the total temperature at the combustion chamber outlet and the number of turbine stages. The functional relationship among the air system bleed air volume, the total temperature at the combustion chamber outlet and the number of turbine stages can be expressed as YQ=f5(T4, J T ), where T4 is the total temperature at the combustion chamber outlet, J T is the number of turbine stages, and YQ is the air bleed volume of the air system.

[0054] S25. Calculate the total pressure ratio based on the fan pressure ratio, the compressor pressure ratio, and the total pressure recovery coefficient within the intermediate casing. Obtain a feasible region for the bypass ratio based on the total pressure ratio and the fuel consumption index as a constraint. During implementation, the feasible region for the bypass ratio is determined based on the total pressure ratio and the fuel consumption index. The functional relationship among the total pressure ratio, the fuel consumption index, and the bypass ratio can be expressed as: B = f6(∑π, sfc), where sfc is the fuel consumption, ∑π is the total pressure ratio, and B is the bypass ratio, with a value range of [B1B2B3…B n ]. The total pressure ratio can be expressed as ∑π=π f *π c *σ, where σ is the total pressure recovery coefficient of the intermediate casing.

[0055] S26. Based on the feasible domain of the fan pressure ratio, the feasible domain of the compressor pressure ratio, the feasible domain of the total temperature at the combustion chamber outlet, and the feasible domain of the bypass ratio, the feasible domain of the fan inlet flow rate is obtained with the thrust index and the fan inlet size threshold as constraints. During implementation, the feasible domain of the fan inlet air flow rate is determined based on the feasible domain of the thermodynamic cycle parameters, the thrust index, and the fan inlet size threshold. The functional relationship between the thermodynamic cycle parameters, the thrust, and the fan inlet size calculated in steps S21 to S25 can be expressed as: W2=f7(π f ,π c , T4, B, Fn, φ2), where Fn is the thrust, φ2 is the fan inlet size threshold, and W2 is the fan inlet air flow rate, and its value can be expressed as [W 21 W 22 W 23 ...W 2n ]).

[0056] Furthermore, in the above steps S24 and S26, the initial air intake volume of the air system is proportional to the feasible range of the total temperature at the combustion chamber outlet and the number of turbine stages; the fan inlet flow rate is proportional to the thrust index and inversely proportional to the fan inlet size.

[0057] Furthermore, in the above step S3, the thermodynamic cycle parameters are determined within the feasible region of the thermodynamic cycle parameters based on the turbine engine performance requirements and the turbine engine's resistance to intake distortion, including:

[0058] S31. Obtaining an overall performance index design margin based on the turbine engine performance requirements, and obtaining a compression component surge margin using the turbine engine's anti-intake distortion. Specifically, the following steps are included:

[0059] S311. Analyze turbine engine performance requirements and determine the priority order of thrust and fuel consumption.

[0060] S312. When the priority of thrust is higher than the priority of fuel consumption rate, obtaining a thrust design margin with the fuel consumption rate being equal to the fuel consumption rate index as a constraint and the thrust being equal to a given multiple of the thrust index as a goal;

[0061] S313. When the priority of fuel consumption rate is higher than the priority of thrust, a fuel consumption rate design margin is obtained with the thrust being equal to the thrust index as a constraint and the fuel consumption rate being equal to a given multiple of the fuel consumption rate index as a goal.

[0062] Preferably, in the above steps S312 and S313, the given multiple of the thrust index is 1.05 to 1.10 times, and the given multiple of the fuel consumption rate index is 0.90 to 0.95 times.

[0063] When implementing the above steps S311 to S313, the emphasis of the aircraft's engine performance requirements (thrust or fuel consumption rate) is captured to roughly determine the overall performance index design margin. For example, for supersonic engines, thrust takes precedence over fuel consumption rate; for long-endurance engines, fuel consumption rate takes precedence over thrust; during design, the design margin of priority indicators is increased as much as possible (exceeding 5% to 10%), where: thrust is mainly related to fan inlet air flow, bypass ratio, combustion chamber outlet total temperature, fan pressure ratio, and compressor pressure ratio; fuel consumption rate is mainly related to bypass ratio, combustion chamber outlet total temperature, fan pressure ratio, and compressor pressure ratio; the effects of compressor pressure ratio and combustion chamber outlet total temperature on thrust and fuel consumption rate are as follows: Figure 3 shown.

[0064] During implementation, the aircraft's requirements for engine stability are generally in the form of "engine inlet comprehensive distortion index greater than or equal to the set value". The compression component surge margin is roughly determined based on the aircraft's requirements for engine stability, the number of fan / compressor component stages and their approximate pressure distortion sensitivity coefficient. When conditions such as the number of compression component stages remain unchanged, the greater the design pressure ratio, the smaller the compression component surge margin.

[0065] S32. Based on the feasible domain of the thermodynamic cycle parameters, the engine overall performance design model is used to obtain the ratio of the total pressure at the mixer outer inlet to the total pressure at the inner inlet. For example, for a short-life turbine hybrid engine, the total pressure at the mixer outer inlet P is 16 The ratio of the total pressure of the internal inlet P6 (P 16 / P6) can be between 0.9 and 1.1, which is mainly related to the fan pressure ratio, bypass ratio and total temperature at the combustion chamber outlet.

[0066] S33. Determine thermodynamic cycle parameters within the thermodynamic cycle parameter feasible region based on the compression component surge margin, the overall performance index design margin, and the ratio. During implementation, optimal values of the thermodynamic cycle parameters can be determined from the thermodynamic cycle parameter feasible region by compromise based on factors such as the thermodynamic cycle parameter feasible region, the overall performance index design margin, the compression component surge margin, the feasibility of each component, and the ratio from step S32.

[0067] Specifically, for each value in the feasible domain of each thermodynamic cycle parameter (fan pressure ratio, compressor pressure ratio, combustion chamber outlet total temperature, bypass ratio, fan inlet flow) determined in step S2, the engine overall performance calculation model can be used for traversal calculation (for example, when all five parameters have six values, the calculation results have a total of five values). 6= 15625 kinds); among the traversal calculation results, the results that simultaneously meet the overall performance indicators and the constraints (compression component surge margin, ratio of the mixer outer inlet total pressure to the inner inlet total pressure) are screened out. This result is the final optional range of the thermodynamic cycle parameters. Within the final optional range, the result with the largest margin for the overall performance focus indicator (thrust or fuel consumption rate) can be selected as the optimal result of the thermodynamic cycle parameters.

[0068] The present invention's thermodynamic cycle parameter acquisition method first analyzes low-cost requirements to determine an engine configuration. This engine configuration is then combined with overall performance indicators (thrust and fuel consumption) and a fan inlet size threshold to determine the feasible domain (i.e., range) of the thermodynamic cycle parameters. Finally, the final thermodynamic cycle parameters are derived from this feasible domain, taking into account turbine engine performance requirements and the turbine engine's resistance to inlet distortion. This method rapidly obtains thermodynamic cycle parameters without requiring repeated iterations, significantly shortening the design cycle.

[0069] Based on the same inventive concept, an embodiment of the present invention also provides a system for acquiring the thermal cycle parameters of a low-cost turbine engine, as described in the following embodiments. Since the principle of solving the problem by the system for acquiring the thermal cycle parameters of a low-cost turbine engine is similar to the method for acquiring the thermal cycle parameters of a low-cost turbine engine disclosed in the above embodiments, the implementation of the system for acquiring the thermal cycle parameters of a low-cost turbine engine can refer to the implementation of the method for acquiring the thermal cycle parameters of a low-cost turbine engine, and the repeated parts will not be repeated. As used below, the terms "unit" or "module" can be a combination of software and / or hardware that implements a predetermined function. Although the system described in the following embodiments is preferably implemented in software, implementation in hardware, or a combination of software and hardware, is also possible and conceivable.

[0070] Figure 4 This is a structural block diagram of a thermodynamic cycle parameter acquisition system for a low-cost turbine engine according to an embodiment of the present invention. The system includes an engine configuration construction module 401, a feasible domain design module 402, and a thermodynamic cycle parameter determination module 403. The structure is described below.

[0071] The engine configuration building module 401 is used to determine the engine configuration according to the cost design target of the turbine engine;

[0072] The feasible region design module 402 is used to obtain a feasible region of thermodynamic cycle parameters according to the engine configuration, with the overall performance index and the fan inlet size threshold as constraints, wherein the overall performance index includes a thrust index and a fuel consumption index;

[0073] The thermodynamic cycle parameter determination module 403 is used to determine the thermodynamic cycle parameters within the thermodynamic cycle parameter feasible domain according to the turbine engine performance requirements and the turbine engine's resistance to intake distortion.

[0074] In this embodiment, a computer device is provided, 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, the method for obtaining thermodynamic cycle parameters of any low-cost turbine engine configuration described above is implemented to solve the technical problem that the existing thermodynamic cycle parameter analysis method has an excessively large selection range, which may result in a large number of turbine stages, making the engine configuration more complicated and making it difficult to design a short-life turbine engine that meets the low-cost requirements.

[0075] Specifically, the computer device may be a computer terminal, a server or a similar computing device.

[0076] In this embodiment, a computer-readable storage medium is provided, wherein the computer-readable storage medium stores a computer program for executing any of the above-mentioned methods for acquiring thermodynamic cycle parameters of a low-cost turbine engine.

[0077] Specifically, computer-readable storage media include permanent and non-permanent, removable and non-removable media that can be used to store information by any method or technology. The information can be computer-readable instructions, data structures, program modules or other data. Examples of computer-readable storage media include, but are not limited to, phase change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other memory technology, compact disc read-only memory (CD-ROM), digital versatile disc (DVD) or other optical storage, magnetic cassettes, tape disk storage or other magnetic storage devices or any other non-transmission media that can be used to store information that can be accessed by a computing device. As defined herein, computer-readable storage media does not include transitory media such as modulated data signals and carrier waves.

[0078] Obviously, those skilled in the art should understand that the various modules or steps of the above-mentioned embodiments of the present invention can be implemented using a general-purpose computing device, they can be concentrated on a single computing device, or distributed across a network composed of multiple computing devices. Alternatively, they can be implemented using program code executable by the computing device, so that they can be stored in a storage device and executed by the computing device. In some cases, the steps described herein can be performed in a different order than that shown, or they can be fabricated into separate integrated circuit modules, or multiple modules or steps can be fabricated into a single integrated circuit module for implementation. Thus, the embodiments of the present invention are not limited to any specific combination of hardware and software.

[0079] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that various modifications and variations of the present invention are possible. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention are intended to be within the scope of protection of the present invention.

Claims

1. A method for obtaining thermodynamic cycle parameters of a low-cost turbine engine, characterized in that: include: Determine the engine configuration based on the cost design target of the turbine engine; According to the engine configuration, the feasible domain of thermodynamic cycle parameters is obtained with the overall performance index and the fan inlet size threshold as constraints, wherein the overall performance index includes the thrust index and the fuel consumption rate index, including: designing the fan pressure ratio and the fan efficiency feasible domain according to the fan stage number of the engine configuration; designing the compressor pressure ratio and the compressor efficiency feasible domain by the fan pressure ratio, as well as the compressor stage number, the compressor bleed type and the compressor guide vane type of the engine configuration; according to the turbine type of the engine configuration, the total temperature at the combustion chamber outlet is less than the processing temperature of the turbine type. As a constraint, a feasible domain of the total temperature at the combustion chamber outlet is designed; the preliminary bleed air volume of the air system is obtained using the feasible domain of the total temperature at the combustion chamber outlet and the number of turbine stages of the engine configuration; the total pressure ratio is calculated based on the fan pressure ratio, the compressor pressure ratio, and the total pressure recovery coefficient of the intermediate casing, and the feasible domain of the bypass ratio is obtained based on the total pressure ratio with the fuel consumption index as a constraint; the feasible domain of the fan inlet flow rate is obtained based on the feasible domain of the fan pressure ratio, the feasible domain of the compressor pressure ratio, the feasible domain of the total temperature at the combustion chamber outlet, and the feasible domain of the bypass ratio with the thrust index and the fan inlet size threshold as constraints; The thermodynamic cycle parameters are determined within the feasible domain of the thermodynamic cycle parameters based on the turbine engine performance requirements and the turbine engine's resistance to intake distortion, including: obtaining an overall performance index design margin based on the turbine engine performance requirements, and obtaining a compression component surge margin using the turbine engine's resistance to intake distortion; obtaining a ratio of a mixer's outer duct inlet total pressure to a mixer's inner duct inlet total pressure based on the thermodynamic cycle parameter feasible domain using an engine overall performance design model; and determining the thermodynamic cycle parameters within the feasible domain of the thermodynamic cycle parameters based on the compression component surge margin, the overall performance index design margin, and the ratio.

2. The method for obtaining thermodynamic cycle parameters of a low-cost turbine engine according to claim 1, characterized in that: The cost design target of the turbine engine is used to determine the engine configuration, including: According to the user's cost design target for the turbine engine, the turbine type, the compressor bleed type, and the compressor guide vane type are obtained, wherein the compressor bleed type includes bleed and non-bleed, the compressor guide vane type includes adjustable guide vanes and non-adjustable guide vanes, and the turbine type includes an integrally cast turbine and a welded turbine; An engine cost model is adopted to obtain the number of fan stages, the number of compressor stages, and the number of turbine stages according to the cost design target, the compressor bleed air type, the compressor guide vane type, and the turbine type.

3. The method for obtaining thermodynamic cycle parameters of a low-cost turbine engine according to claim 1, characterized in that: The initial air intake volume of the air system is proportional to the feasible range of the total temperature at the combustion chamber outlet and the number of turbine stages; the fan inlet flow rate is proportional to the thrust index and inversely proportional to the fan inlet size.

4. The method for obtaining thermodynamic cycle parameters of a low-cost turbine engine according to claim 1, characterized in that: Obtain overall performance index design margin based on turbine engine performance requirement analysis results, including: Analyze turbine engine performance requirements and determine the priority between thrust and fuel consumption; When the priority of thrust is higher than the priority of fuel consumption rate, the thrust design margin is obtained with the fuel consumption rate being equal to the fuel consumption rate index as a constraint and the thrust being equal to a given multiple of the thrust index as a goal; When the priority of the fuel consumption rate is higher than the priority of the thrust, the fuel consumption rate design margin is obtained with the thrust being equal to the thrust index as a constraint and the fuel consumption rate being equal to a given multiple of the fuel consumption rate index as a goal.

5. The method for obtaining thermodynamic cycle parameters of a low-cost turbine engine according to claim 4, characterized in that: The given multiple of the thrust index is 1.05 to 1.10 times, and the given multiple of the fuel consumption rate index is 0.90 to 0.95 times.

6. A low-cost turbine engine thermodynamic cycle parameter acquisition system, characterized in that: include: An engine configuration building module, wherein the engine configuration building module is used to determine an engine configuration according to a cost design target of a turbine engine; A feasible domain design module is used to obtain the feasible domain of thermodynamic cycle parameters based on the engine configuration, with the overall performance index and the fan inlet size threshold as constraints, the overall performance index includes the thrust index and the fuel consumption rate index, including: designing the fan pressure ratio and fan efficiency feasible domain according to the fan stage number of the engine configuration; designing the compressor pressure ratio and compressor efficiency feasible domain through the fan pressure ratio, as well as the compressor stage number, compressor bleed type and compressor guide vane type of the engine configuration; designing the compressor pressure ratio and compressor efficiency feasible domain according to the turbine type of the engine configuration, with the total temperature of the combustion chamber outlet being less than the processing The feasible domain of the total temperature at the combustion chamber outlet is designed using the processing temperature of the turbine type as a constraint; the preliminary bleed air volume of the air system is obtained using the feasible domain of the total temperature at the combustion chamber outlet and the number of turbine stages of the engine configuration; the total pressure ratio is calculated based on the fan pressure ratio, the compressor pressure ratio, and the total pressure recovery coefficient within the intermediate casing, and the feasible domain of the bypass ratio is obtained based on the total pressure ratio using the fuel consumption index as a constraint; the feasible domain of the fan inlet flow rate is obtained based on the feasible domains of the fan pressure ratio, the compressor pressure ratio, the total temperature at the combustion chamber outlet, and the bypass ratio, using the thrust index and the fan inlet size threshold as constraints; A thermodynamic cycle parameter determination module is used to determine thermodynamic cycle parameters within a feasible domain of thermodynamic cycle parameters based on turbine engine performance requirements and turbine engine anti-intake distortion, including: obtaining an overall performance index design margin based on the turbine engine performance requirements, and using the turbine engine anti-intake distortion to obtain a compression component surge margin; obtaining a ratio of a mixer outer duct inlet total pressure to a mixer inner duct inlet total pressure based on the thermodynamic cycle parameter feasible domain using an engine overall performance design model; and determining thermodynamic cycle parameters within the feasible domain of thermodynamic cycle parameters based on the compression component surge margin, the overall performance index design margin, and the ratio.

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