A steady-state performance calculation method for heavy-duty gas turbines based on multi-stage turbine cooling

Abstract

This application belongs to the field of heavy-duty gas turbine design technology, specifically involving a method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling. The method includes: designing a calculation method for compressor bleed air cooling of a multi-stage turbine: the four-stage turbine is cooled by bleed air from the compressor at stages 5, 7, 8, 10, and 15 respectively, with the bleed air volume expressed in matrix form. The product of the bleed air volume and the bleed air enthalpy is used as the mixed air energy to participate in the cooling of the main airflow; integrating the compressor bleed air cooling multi-stage turbine air system into the overall performance system, constructing a refined overall performance model for solution, and calculating the steady-state performance of the heavy-duty gas turbine; comparing with experimental data, and applying a correction coefficient matrix to correct the calculated steady-state performance of the heavy-duty gas turbine.

Description

Technical Field

[0001] This application belongs to the field of heavy-duty gas turbine design technology, specifically relating to a method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling. Background Technology

[0002] Heavy-duty gas turbines are core power equipment for efficient energy utilization, characterized by high technological intensity and significant design and manufacturing challenges. Steady-state performance calculations for heavy-duty gas turbines are fundamental to overall performance design and provide guidance for overall scheme formulation and specialized technical research and development directions.

[0003] Currently, steady-state performance calculations for heavy-duty gas turbines primarily utilize commercial software such as GASTURB and GATECYCLE. However, these methods lack high accuracy and their applicability differs from actual needs. While GASTURB performs well in calculating the performance of turbofan gas engines, its model architecture is not detailed enough for single-rotor heavy-duty gas turbines, resulting in lower calculation accuracy. GATECYCLE, mainly designed for performance schemes of gas turbines and combined cycle systems, offers somewhat coarse-grained performance calculations for gas turbines. Considering economic factors and software encapsulation models, using commercial software to calculate the steady-state performance of heavy-duty gas turbines limits designers' ability to create adaptive solutions, leading to high costs, poor economic efficiency, and low development and usage efficiency. Therefore, a more refined design for the steady-state performance calculation of heavy-duty gas turbines is needed to improve its applicability and practicality.

[0004] This application is made in view of the aforementioned technical deficiencies. Summary of the Invention

[0005] The purpose of this application is to provide a method for calculating the steady-state performance of heavy-duty gas turbines based on multi-stage turbine cooling, so as to overcome or mitigate at least one of the known technical defects.

[0006] The technical solution of this application is:

[0007] A method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling includes:

[0008] Calculation method for designing a multi-stage turbine with bleed air cooling for a compressor:

[0009] The four turbines are cooled by bleed air from the compressor in stages 5, 7, 8, 10 and 15, respectively. The bleed air volume is expressed in matrix form. The product of the bleed air volume and the bleed air enthalpy is used as the mixed air energy to participate in the cooling of the main airflow.

[0010] The compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system, and a refined overall performance model is constructed and solved to obtain the steady-state performance of the heavy-duty gas turbine.

[0011] By comparing experimental data, a correction coefficient matrix was applied to correct and calculate the steady-state performance of the heavy-duty gas turbine.

[0012] According to at least one embodiment of this application, in the above-described method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling, the compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system, and a refined overall performance model is constructed and solved to calculate the steady-state performance of the heavy-duty gas turbine, specifically as follows:

[0013] Given the gas turbine power, the compressor characteristic beta value, the characteristic beta value, and the combustion chamber outlet temperature, thermodynamic cycle calculations are performed according to the gas flow sequence to obtain the steady-state performance of the heavy-duty gas turbine.

[0014] According to at least one embodiment of this application, in the above-described method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling, the compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system, and a refined overall performance model is constructed for solution to calculate the steady-state performance of the heavy-duty gas turbine. During the process, thermodynamic cycle calculations are carried out in the order of airflow from the compressor, combustion chamber, compressor bleed air cooling multi-stage turbine, and exhaust device.

[0015] According to at least one embodiment of this application, in the above-described method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling, the compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system, and a refined overall performance model is constructed for solution to calculate the steady-state performance of the heavy-duty gas turbine. During the process, thermodynamic cycle calculations are performed according to the gas flow sequence to obtain rotor inlet flow deviation, power deviation, and turbine outlet pressure deviation. After forming a residual matrix, the NEWTON-RAPHSON algorithm is applied to solve the nonlinear equation set to calculate the steady-state performance of the heavy-duty gas turbine.

[0016] According to at least one embodiment of this application, in the above-described method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling, the steady-state performance of the heavy-duty gas turbine is calculated by comparing experimental data and applying a correction coefficient matrix. This includes multiplying the pressure ratio, flow rate, and efficiency of the compressor and turbine characteristic points by the correction coefficient matrix, respectively, and participating in the calculation of the overall thermodynamic performance of the turbine.

[0017] This application has at least the following beneficial technical effects:

[0018] This paper presents a method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling. On the basis of compressor bleed air cooling of a four-stage turbine, a refined model for calculating the performance of the heavy-duty gas turbine is built, and a coefficient matrix correction method is applied to further improve the calculation accuracy of the performance model. Attached Figure Description

[0019] Figure 1This is a schematic diagram of the bleed position of the axial compressor provided in the embodiments of this application;

[0020] Figure 2 This is a schematic diagram of the cooling turbine calculation mode in the commercial software GASTURB model calculation provided in this application embodiment;

[0021] Figure 3 This is a schematic diagram of the compressor bleed air cooling four-stage turbine and bleed air ratio provided in the embodiments of this application;

[0022] Figure 4 This is a schematic diagram illustrating the integration of a compressor bleed air cooling multi-stage turbine air system into the overall performance system solution provided in this application embodiment;

[0023] Figure 5 This is a power spectrum of a heavy-duty gas turbine field test provided in an embodiment of this application;

[0024] Figure 6 This is a comparison chart of the computational accuracy between the GASTURB model and the overall performance refinement model provided in the embodiments of this application;

[0025] Figure 7 This is a schematic diagram of the compressor characteristic matrix correction process provided in the embodiments of this application;

[0026] Figure 8 This is a comparison chart of the computational accuracy between the GASTURB model provided in the embodiments of this application and the refined model for overall performance calculation using the correction coefficient matrix.

[0027] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation

[0028] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings, and other related parts can be referred to the general design.

[0029] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.

[0030] Compared to multi-rotor gas turbines, single-shaft gas turbines have a relatively simple structure. In commercial software like GASTURB, the steady-state performance calculations for single-rotor heavy-duty gas turbines involve combined cooling calculations before and after the turbine rotor, resulting in fewer variables and balance equations in the system equations. While simplifying turbine calculations, this reduces the accuracy and realism of the overall turbine performance calculations. Therefore, this application provides a method for calculating the steady-state performance of heavy-duty gas turbines based on multi-stage turbine cooling. This method cools the four-stage turbine with bleed air from the compressor, integrating the complex air cooling system into the nonlinear system of the performance model. This improves the model's refinement. Furthermore, by comparing experimental data and using a coefficient matrix correction method, the steady-state performance is corrected, further enhancing the calculation accuracy of the heavy-duty gas turbine steady-state performance model.

[0031] Calculation method for designing a multi-stage turbine with bleed air cooling for a compressor:

[0032] The single-shaft heavy-duty gas turbine's main components include a 15-stage axial-flow compressor, a combustion chamber, and a 4-stage turbine. The bleed air location of the 15-stage axial-flow compressor is also specified. Figure 1 As shown, air is drawn from levels 5, 7, 8, 10, and 15, respectively.

[0033] In the commercial software GASTURB model calculation, the cooling turbine calculation mode is as follows: Figure 2 As shown, cooling mixing calculations are performed separately before and after rotor calculations. This simplified performance calculation affects the effectiveness and accuracy of the overall machine performance calculation. Therefore, a calculation method for compressor bleed air cooling of a multi-stage turbine is designed, with a compressor bleed air cooling four-stage turbine and an bleed air ratio as shown in the example. Figure 3 As shown.

[0034] The four-stage turbine is cooled by bleed air from the compressor in stages 5, 7, 8, 10, and 15, respectively. The bleed air volume is expressed in matrix form, and the product of the bleed air volume and the bleed air enthalpy is used as the mixed air energy to participate in the cooling of the main airflow. In this way, the complex air cooling system is integrated into the nonlinear system of the performance model for solution calculation.

[0035] The compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system, and a refined overall performance model is constructed for solution:

[0036] For a single-rotor heavy-duty gas turbine used in power generation, the variable and residual settings and calculation process for the overall performance nonlinear system incorporating compressor bleed air cooling multi-stage turbine are as follows: Figure 4 As shown.

[0037] Given the gas turbine power, firstly, the compressor characteristic beta value, characteristic beta value, and combustion chamber outlet temperature are given. Then, thermodynamic cycle calculations are performed according to the gas flow sequence (compressor, combustion chamber, compressor bleed air cooling multi-stage turbine, exhaust device) to obtain rotor inlet flow deviation, power deviation, and turbine outlet pressure deviation. After forming the residual matrix, the NEWTON-RAPHSON algorithm is applied to solve the nonlinear equations to calculate the steady-state performance of the heavy-duty gas turbine.

[0038] After the overall performance refinement model was built, a comparison of the calculation accuracy between the GASTURB model and the overall performance refinement model was conducted on the field test data of the heavy-duty gas turbine. The power spectrum of the heavy-duty gas turbine field test is shown below. Figure 5 As shown, steady-state performance data are extracted and compared. The main process data for heavy-duty gas turbines include compressor outlet temperature, compressor outlet pressure, and exhaust temperature. Under equal power conditions, the calculation accuracy of the two models is compared as follows: Figure 6 As shown, compared with the GASTRUB method, the multi-stage turbine cooling method has higher calculation accuracy.

[0039] By comparing experimental data, a correction coefficient matrix was applied to correct and calculate the steady-state performance of the heavy-duty gas turbine:

[0040] To further improve computational accuracy, coefficient matrices are used to correct component characteristic points, thus refining the model's steady-state performance. In the steady-state performance calculation of the heavy-duty gas turbine, the pressure ratio, flow rate, and efficiency of the compressor and turbine characteristic points are multiplied by correction coefficient matrices and then incorporated into the overall engine thermodynamic performance calculation. The compressor characteristic matrix correction process is as follows: Figure 7 As shown, the turbine characteristic correction process is similar.

[0041] Comparison of the computational accuracy of the GASTURB model and the refined model using the applied correction coefficient matrix for overall performance calculation, for example... Figure 8 As shown, the accuracy of parameter calculation under each working condition after correction is less than 0.4%.

[0042] The above embodiments, compared with the simplified model established by the commercial software GASTURB, develop a method for calculating the steady-state performance of heavy-duty gas turbines based on multi-stage turbine cooling. After building a bleed air-cooled 4-stage turbine model, it is integrated into the overall performance equation set for calculation. Finally, correction coefficients are applied to improve the accuracy of the performance model calculation.

[0043] First, by comparing simplified turbine cooling, a calculation method for a compressor induced air cooling four-stage turbine was developed;

[0044] Then, the multi-stage turbine cooling air system is integrated into the overall performance system solution to improve the refinement of the steady-state model;

[0045] Finally, by comparing the experimental data, the steady-state performance of the gas turbine was corrected using a correction coefficient matrix, thereby further improving the calculation accuracy of the performance model.

[0046] The method for calculating the steady-state performance of heavy-duty gas turbines based on multi-stage turbine cooling provided in the above embodiments can improve the overall performance calculation accuracy of heavy-duty gas turbines, enhance the quality of performance design and performance evaluation, provide methodological support for the performance design of heavy-duty gas turbines, and also provide a new and practical solution for the steady-state performance calculation of single-rotor gas turbines containing multi-stage turbines.

[0047] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of this application.

Claims

1. A method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling, characterized in that, include: Calculation method for designing a multi-stage turbine with bleed air cooling for a compressor: The four turbines are cooled by bleed air from the compressor in stages 5, 7, 8, 10 and 15, respectively. The bleed air volume is expressed in matrix form. The product of the bleed air volume and the bleed air enthalpy is used as the mixed air energy to participate in the cooling of the main airflow. The compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system, and a refined overall performance model is constructed and solved to obtain the steady-state performance of the heavy-duty gas turbine. By comparing experimental data, a correction coefficient matrix was applied to correct and calculate the steady-state performance of the heavy-duty gas turbine.

2. The method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling according to claim 1, characterized in that, The compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system. A refined overall performance model is constructed and solved to obtain the steady-state performance of the heavy-duty gas turbine, specifically: Given the gas turbine power, the compressor characteristic beta value, the characteristic beta value, and the combustion chamber outlet temperature, thermodynamic cycle calculations are performed according to the gas flow sequence to obtain the steady-state performance of the heavy-duty gas turbine.

3. The method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling according to claim 2, characterized in that, The compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system, and a refined overall performance model is constructed for solution to obtain the steady-state performance of the heavy-duty gas turbine. In the process, thermodynamic cycle calculations are carried out in the order of airflow from the compressor, combustion chamber, compressor bleed air cooling multi-stage turbine, and exhaust device.

4. The method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling according to claim 3, characterized in that, The compressor bleed air cooling multi-stage turbine air system is integrated into the overall performance system. A refined overall performance model is constructed and solved to obtain the steady-state performance of the heavy-duty gas turbine. During the process, thermodynamic cycle calculations are carried out according to the gas flow sequence to obtain rotor inlet flow deviation, power deviation, and turbine outlet pressure deviation. After forming the residual matrix, the NEWTON-RAPHSON algorithm is applied to solve the nonlinear equations to obtain the steady-state performance of the heavy-duty gas turbine.

5. The method for calculating the steady-state performance of a heavy-duty gas turbine based on multi-stage turbine cooling according to claim 4, characterized in that, By comparing experimental data, the steady-state performance of the heavy-duty gas turbine is calculated using a correction coefficient matrix. This includes multiplying the pressure ratio, flow rate, and efficiency at the characteristic points of the compressor and turbine by the correction coefficient matrix, and then using these factors to calculate the overall thermodynamic performance of the turbine.

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