Numerical Study On Heat Transfer Mechanism Of Cooling Structure Of High Temperature Rotor Of Heavy Duty Gas Turbine

Heavy-duty gas turbine collects the most advanced scientific and technological achievements of human social engineering technology.Its development level shows a country's industrial production capacity.It plays a vital role in the efficient,clean and safe use of energy.The high temperature rotor blade of heavy-duty gas turbine not only has to bear the high temperature higher than the allowable temperature of its manufacturing materials,but also is always in a high-speed rotation state in operation.Under the action of large centrifugal force,Coriolis force and its derivative buoyancy force,the working environment of the rotor blade of heavy-duty gas turbine is the worst.Therefore,the research of high temperature rotor blade and its cooling structure is the commanding point technology in the development of the flow passage components of heavy-duty gas turbine.In this thesis,the following preparations were made for the study of high temperature rotor blades.Referring to the S2 flow surface calculation software developed by Harbin Institute of technology,one-dimensional design software,quasi three-dimensional design software and one-dimensional heat flow pipe network calculation software for multi-stage gas-cooled turbine of heavy-duty gas turbine were developed.Aiming at the integrated design and numerical simulation of internal and external flow and heat coupling of heavy-duty gas turbine,the Mark II blade with experimental data and C3 X blade with multiple exhaust film holes were used as calculation models to verify the numerical method adopted in this thesis.By using one-dimensional and quasi three-dimensional design software,the advanced design experience of 9FA was digested and absorbed,the flow passage of F-class hesvy duty turbine with medium and low calorific value fuel was designed independently,and the flow passage characteristics were compared with those of9 FA.The results show that the enthalpy drop of each stage of these two three-stage turbines was reduced step by step,and the gradient of the second stage was smaller.The distribution of the enthalpy drop of the flow passage can not only improve the load of the upstream stage,but also reduce the exit loss of the flow passage.Due to the combination of meridional convergence and control vortex in the flow path design of the turbine with medium and low calorific value,the distribution of aerodynamic parameters along the blade height in the axial clearance of the cascade was relatively uniform,and the upstream cascade provides a good inlet flow field for the downstream cascade,while reducing the radial mixing loss.However,due to the low initial temperature and large flow of cold air,the flow efficiency of the medium and low calorific value turbine was 1.0% lower than that of the 9FA turbine.In addition,using one-dimensional heat flow pipe network calculation software,the effects of changing the cooling channel of mark ? cascade prototype into "slub" channel with sudden expansion and contraction and the high-speed rotation of high-temperature rotor on the internal cooling were rapidly predicted.It was found that strengthening the disturbance of the internal cooling wall will strengthen the heat transfer effect between the cold air and the blade wall,which lays a theoretical foundation for the modification design of the internal cooling structure of high-temperature rotor.Taking the 9FA turbine rotor as the master type,under the condition of lack of original geometric data,through repeated theoretical deliberation and conjugate heat transfer numerical simulation,the advanced design experience of 9FA turbine was explored.Based on the characteristics of fuel supply in China,the initial scheme of high temperature rotor and its internal cooling structure was designed.Then,the flow characteristics and heat transfer characteristics of the initial scheme were obtained by means of conjugate heat transfer numerical simulation,and the feasibility and existing problems of the scheme were analyzed theoretically.Then,three modifications were carried out to obtain the prototype scheme of high temperature rotor.In the modification design,the corner form of the cooling channel,the size of the baffle impact hole,the width of the trailing edge outlet channel,the position and angle of the filling hole were mainly adjusted.The numerical analysis of the prototype schemes hows that the flow and heat transfer characteristics of the blade were good,and the expected design goal was basically achieved.Finally,the prototype scheme was further modified: the form of the spoiler rib in the cooling chamber was improved,60° V-shaped rib was adopted,and three through slots were added along the blade height between the cooling channels VI and VII.The numerical results show that: the temperature of the outer surface of the blade decreases,the local heat transfer enhancement effect was better;the inlet parameters of the cold air become more reasonable;the mass flow rate of the cold air increases,the spanwise distribution was more uniform;the film injection and covering ability were enhanced.Obviously,the expected improvement of cold air flow field and cooling effect has been partially realized.The modified blade,that was,the final cooling structure,has important reference value for engineering application.In order to reduce the temperature of the metal surface of the blade as much as possible and prolong the service life of the blade,the effect of thermal barrier coating on the heat transfer performance of the blade was numerically studied.Compared with the uncoated high temperature rotor,the temperature difference between the coated blade surface and the gas film decreases,which reduces the heat flux through the blade,resulting in a more uniform temperature distribution on the metal surface of the blade.Except for the special tip area,the surface temperature of the coated blade was 80 K lower than that of the uncoated blade.In general,the internal cooling structure was equipped with supplementary air holes.In this thesis,the effect of supplementary air holes was understood by comparing the flow and heat transfer performance of the internal cooling structure with and without supplementary air holes under the condition of the same cold air flow.The numerical simulation results of the internal cooling structure of high temperature rotor with and without supplementary air holes show that the cooling effect of the internal cooling structure without supplementary air holes was significantly higher than that of the internal cooling structure with holes.However,it has no engineering application value,due to the long and tortuous process in the internal cooling structure of cold air,the flow resistance was greatly increased,and the required inlet pressure of cold air has far exceeded the maximum pressure of the compressor air supply.Comparing the internal cooling structure with supplementary air holes and no supplementary air holes,the cooling efficiency was lower,and the blade surface temperature and temperature gradient were higher.However,when the blade surface temperature and its gradient do not exceed the standard,the required inlet pressure of cold was in the range of air pressure compressor supply.Whether there were supplementary air holes in the rear cavity of the internal cooling structure,the results show that not only the temperature distribution on the blade surface was the same,but also the inlet pressure of the cold air was almost the same,so which can be cancelled.In view of the difficulties in the experimental study of the cooling effect of high temperature rotor,the influence of high-speed rotation on the cooling system was further studied by using the conjugate heat transfer calculation method.The calculation results show,in the rotating state,the turbulence intensity of the cold air increases significantly,the flow resistance and heat transfer ability were enhanced,and the blade surface temperature meets the design requirements in a static state,and must meet the design requirements in a rotating state.The numerical results show that it was feasible to replace the dynamic test with the static test for the high temperature rotor cooling structure of gas turbine.In subsonic wind tunnel,the aerodynamic characteristics of high temperature rotor profile of F-class heavy-duty gas turbine with medium and low calorific value fuel were tested.The test results show that the blade profile loss curve of high temperature rotor with incidence angle was relatively flat,which indicates that the blade profile selected for F-class heavy-duty gas turbine with medium and low calorific value fuel has good characteristics of variable incidence angle.