Aerodynamic Design Of Marine High Power Axial-compressor

High-power marine gas turbine is the key and core equipment for the development of large warships in China.Long-term tracking and imitation make us lack of fully autonomous design capability of marine gas turbine compressor.This causes the problems that design system is not complete,design level is not high,and professional design rule is missing.Therefore,it is urgent to carry out the basic research on the aerodynamics and thermodynamics of marine gas turbine compressor.Development of modern compressor is toward the trend of high average stage pressure ratio,high efficiency and large surge margin,which puts forward much higher requirement for the compressor aerodynamic design.For this reason,the investigation on aerodynamic design of high power marine axial compressor is conducted in this thesis.Through the research progress survey of high-power marine gas turbine compressor in domestic and abroad,according to the fundamental principles and methodology of compressor aerodynamic design in combination with the characteristics and actual situation of marine gas turbine,aerodynamic design system of high-power marine multistage axial compressor suitable for our country has been established.The primary purpose of one-dimensional inverse problem is to estimate the compressor efficiency and main geometric parameters according to a given flow rate and pressure ratio.The preliminary design scheme of compressor is determined mainly based on empiricalrelationship.In the one-dimensionalcalculation and optimization stage,the AXIAL software module in Concepts-NREC engineering design system is adopted,which can predict and analyze the performance of turbomachinery under variable conditions.Calculation program of S2 based on streamline curvature method is adopted for the inverse problem of S2 flow,with the main task as the design the distribution law of cascade performance parameters along the blade height for the given meridian flow size.Based on the design experience of CDA blade,in which the inlet air flow parameters is transformed into non-dimensionalone,and the relationship between blade dimensionless parameters and the blade modeling parametersis obtained.CDA blade profile modeling parameters are determined through the dimensionless parameters,and then CDA blade profile is achieved.In order to enable the designed CDA blade to various flow environments and play the role of flow control,the parameters such as the location of the middle arc,the location of the maximum thickness and the radius of the leadingtrailing edge are linked with the aerodynamic design results.The full 3D numerical simulation is carried out using NUMECA software.According to the three-dimensional numerical simulation results,the matching design between stages is performed,with the consideration the off-design performance of the compressor in the appropriate angle of attack of different areas.Based on the established aerodynamic design system for high-power multistage axial compressor,a design caseof a high-power gas turbine compressor is carried out,and the aerodynamic design of the 6 stage axial compressor is completed.The design of One dimension inverse problem,analysis of one-dimensional characteristics,S2 inverse problem calculation,blade modeling and 3D CFD calculation and analysis are carried out.The load distribution at various stages is gradually reduced from the front to the rear,and the spanwise direction is distributed in accordance with the isobaric ratio in order to avoid excessive radial mixing loss,the designof S2 does not strictly follow the traditional design rules(such as equal circulation,equal reaction degree etc.),based on one-dimensional design parameters,which are obtained throughcontinuous adjustment of design parameters.According to the CDA technical feature,a blade modeling program for engineering practiceis developed,which can realize the functions of different camber line,different thickness distributions,andbending and sweeping of stack line.The three-dimensional numerical simulation results show that with consideration of the 0.5mm clearance of the rotor blade tip,the design efficiency of the 6 stage compressor can reaches 89.06%,the surge margin of the design speed is 17.6%,and the off-design conditions is also good.Theoretical analysis of the matching between the stages is performed.the compressor is deviated from the design state at the equal speed characteristic line,and the compressor characteristics is changedin all stages.The relative value of this change will be amplified step by step,and the relative value of the exit stage from the design state is the highest.This is the step by step magnification of the speed characteristic change of the multistage compressor.When the speed is lower than the design speed,it will cause severe stall or surge at the inlet or front stage.The higher the pressure ratio of the compressor,the trend of the surge in front and choke behind is much stronger.The performance analysis of 6 stage compressor is carried out under various conditions,the weighted average calculation is performed in the circumferential direction,and the distribution of the airflow angle in the meridional flow path is then obtained.With the increase of the inlet air angle,the outlet air angle remains basically unchanged,and the static blade has a strong "adsorption" ability.When exceeding this range of adsorption,the outlet air angle will change,that is,the so-called backward angle effect.Under Low working conditions,the performance deterioration of compressor first occurs at the top area of rotor blade withlarge positive incidence angle,and the negative incidence angle of last stage compressor also causes large aerodynamic losses,this cumulative effect leads to that it is difficult to match the angle of the back stages,especially around the low speed conditions,this matching is more difficult.On one hand,the design point performance should be taken into account,and the wide range variable condition performance should be met.However,how to balance the contradiction between the cumulative effect and design requirements is the difficulty in the design progress.Through the theoretical and calculation,the scheme of turning angle under low speed conditions is presented.Surge simulation analysis of marinehigh-power multistage axial compressor is carried out.Method and program for estimating the surge boundary of compressor is proposed,which can be used to determine the stable working boundary of compressor based on the value of compressor rotor's diffuser.Compressor characteristic curve is derived by program calculation,and the key parameters such as diffuser criterion value,diffuser coefficientwD,inlet rotor diffuser 2 1A/ A and aerodynamic load criterion cp? of the rotor on stall boundary are also obtained.According to the engineering examples,calculation formula is suitable to calculate the surge boundary of new compressor.However,in order to improve the analysis,experimental datashould be accumulated.Aerodynamic performance test requirements of compressor,parameter measurement and measurement method,data processing and error analysis are carried out.Research results from this thesis can provide a solid foundation for the aerodynamic design of marine high-power axial compressor,and have important theoretical and practical significance for promoting the development of marine high-power marine gas turbines.