In view of the problem of creep-fatigue cracking caused by thin-walled structure with dense holes in gas turbine engine, the DZ125 thin-walled plate specimens with holes were designed, and the creep-fatigue test was carried out at 850 ℃. Based on the elastic-plastic finite element analysis results of double-hole thin plate, the maximum stress path between two holes was defined as the critical area, and the equivalent stress between holes describing the complex stress state of the plate with holes was proposed. The finite element analysis and test results illustrated that the non-uniform stress between holes is the key factor to determine the cycle life; the cycle life decreased with the increase of the equivalent stress between holes. In addition, the tensile stress along the loading direction is the dominant stress of creep-fatigue failure of thin plate with holes; and the crack was originated from the high stress area of edge holes. Finally, when the the ratio of distance between adjacent holes to the diameter was about 4.22, the equivalent stress changed significantly, so the parameter should be controlled above this critical value in the design.

In order to adapt to the range of non-orthogonal parameters caused by the self-constrained parametric modeling method, a mapping method was proposed to map the sampling points of the traditional orthogonal design space to the non-orthogonal space. A feasible domain boundary identification method based on machine learning logistic regression was constructed to judge the compliance of the sample space response with relevant criteria. This method was applied to a high pressure turbine two-/ three-tooth tenon structure, finding that an engine preliminary design scheme may fail to meet the design requirements in the detailed design phase of tenon structure, and tenon structure parameters may be advanced to the preliminary design stage, together with the superior parameter design, to prevent the design difficulties in the detailed design stage. Through the boundary identification method, it was found that when the speed was reduced to 15 575 r /min or the outer diameter of the wheel was reduced to 303 mm, the two-tooth tenon structure can be designed. When the speed was reduced to 15 584 r /min or the outer diameter of the wheel was reduced to 301 mm, the three-tooth tenon structure can be designed. This method can be used as an iterative feedback information among various disciplines in the initial engine design stage. The proposed method can be generalized to other details of the engine structure design.

In order to suppress the flow-induced coupling vibration of rotor blades/bladed drum and stator caused by the close modal frequencies with the same nodal diameter between the rotor and stator, an ad-hoc forward design method based on displacement and strain energy density distribution was proposed to tailor the frequency margin, providing a solid theoretical basis for designing the gap of the frequency; and the stiffened area can be preliminarily determined through a single modal analysis, which significantly shortened the design cycles. The method was then applied to an industrial rotor/stator finite element model to improve the resonance margin. The influence of structural parameters of stiffeners on the modal characteristics of rotor and stator was investigated. The resonance margin of the dangerous mode increased from 3.47% to 10.56% by stiffening. The generality of this method is good, making it suitable for other types of engines.

A high-pressure turbine blade of a high thrust-to-weight ratio turbofan adopted edge plate damper design to reduce the vibration stress of the blade. During the core machine ground bench test, the frequency-domain characteristics of the strain test data showed drift phenomenon, and the vibration energy presented random feature in a narrow frequency band. The phenomenon was discussed and analyzed at first. Then the equivalent model of a single blade considering the influence of the edge plate damper was established according to test data. Based on the time integration method and the nonlinear modal theory, the response characteristics, friction characteristics, resonant frequency characteristics and friction damping characteristics of the system were calculated under different rotation speeds. The simulation results showed that the rotation speed fluctuation affected the frequency domain distribution characteristics of the vibration response and the resonant frequency range distribution of the system: at 11 713 r/min, the nonuple excitation order induced the resonance frequency band of 1 756—1 952 Hz, at 13 500 r/min, the octuple excitation order induced the resonance frequency around 1 800 Hz, and at 13 687 r/min, the octuple excitation order induced the resonance frequency band of 1 596—1 824 Hz. Under stable rotational speed, the dry friction force jumped alternatively between the maximum and minimum values, which reflected a friction damping effect and brought about unstable changes in the additional stiffness of the system. The nonlinear modal theory results showed that the modal frequency of the damped blade changed with the vibration response amplitude. The decrease of the blade stiffness or the increase of the friction force amplitude can enlarge the modal frequency range.

The research results for improving the efficiency of failure probability calculation, including the efficient sampling-based algorithms and the efficient integration-based algorithms, were summarized. Among them, based on the traditional Monte Carlo simulation method, the importance sampling algorithm generated samples in the failure domain. The optimal sampling technology optimized the zone sample size. The zone refinement technology reduced the number of zones, thus reducing the Monte Carlo simulation sample size. In addition, the direct integration was realized by establishing the mapping relationship of failure domain at N flight cycles and the initial (N = 0) flight cycles based on the probability density theory. When the relative error with Monte Carlo was less than 5%, the calculation time cost was reduced by at least tens of times.

The concept of assembly mechanics was applied to improve the traditional horizontal assembly technology of low pressure turbine unit of commercial high bypass ratio aero-engine. An advanced horizontal assembly technology was developed based on deformation control by mechanical simulation analysis of the horizontal docking process of low pressure turbine unit. The result demonstrated that this technology realized the installation of low pressure turbine unit with high precision, high efficiency and high reliability. In addition, compared with traditional experience assembly, the guided tool structure was more concise by contributing to around 70% reduction of the weight, and the docking difficulty was greatly reduced by contributing to more than 60% increase of the assembly efficiency and 100% increase of the success rate of one-time docking.

Important research results of non-synchronous vibration were summarized through its experiment, non-synchronous vibration signal obtaining, and mechanism of non-synchronous vibration. The basic research of non-synchronous vibration, and research on blade tip timing technology combined strain gauge and PIV (particle image velocimetry) to obtain the mechanism of non-synchronous vibration, compressed sensing blade tip timing to capture non-synchronous vibration signal and rotor blade vibration active control to research the coupling mechanism. Blade tip timing technology and strain gauge, PIV and active control blade vibration technology provide advanced brace for non-synchronous vibration phenomenon research. In a long run, multi-physics experimental research will still be a mainstream research method for the non-synchronous vibration.

A one-dimensional convection-thermal conductivity model was established, and the thermal insulation effectiveness criterion of the thermal barrier coating was obtained through theoretical deduction and analysis. When the thermal resistance of the coating was greater than the heat transfer resistance of the high-temperature gas side without coating, the coating could generate a positive thermal insulation effect. On the contrary, the heat transfer coefficient for the external gas side after spraying the thermal barrier coating (TBC) had a maximum critical value. At this time, only when the external gas side heat transfer coefficient was smaller than this critical value after the TBC was sprayed, the TBC could have a thermal insulation effect, otherwise the thermal barrier coating couldn't work, even worsen the heat transfer of the blades. The temperature-drop of the TBC itself was proportional to the temperature-drop of the outer surface of metal blade with or without TBC. And a new evaluation standard of the thermal insulation effect of the coating based on the outer surface temperature of the blade metal substrate was proposed.

A time-varying model considering crack growth degradation process was established and applied to the probabilistic life analysis of crack growth in the turbine disk mortise. Firstly, a time-varying crack growth model of GH4720Li superalloy was established by introducing the Wiener stochastic process with double time scale functions. And it was verified by crack growth test of compact tensile specimens. Then, taking the root of the turbine disk’s mortise tooth as the key object, the weight function method for solving the stress intensity factor of the crack in the root of the turbine disk’s mortise tooth was established and compared with the finite element analysis results of the crack growth in the turbine disk mortise. Finally, combining the weight function and the time-varying crack growth model, the reliability analysis method for fatigue crack growth degradation of the turbine disk mortise was established. Analysis results showed that the crack growth degradation life of the turbine disk’s mortise presented a large dispersion, with an average value of 14177 cycles and a standard deviation of 1090.09 cycles. The prediction of crack growth life was 10312 cycles when the reliability was 99.87%.

A numerical simulation method of laser shock peening of curved structure based on the idea of discretization was proposed. The precise definition of the impacted area and pressure was realized through the spatial geometric relationship and the principle of energy conservation, and the numerical simulation of laser shock peening of any curved surface and any angle can be realized. Then the target mesh size was determined according to the requirements of mesh independence. Using this method, the residual stress distribution law of the characteristic for the turbine mortise structure after laser shock peening was explored. Compared with the experimental results, the prediction error was less than 20%. The research showed that the residual compressive stress in a certain depth range was introduced into the mortise of the turbine disk after laser shock peening, but due to the process accessibility caused by the structural characteristics of the curved surface, the residual stress value was lower than the plane structure at the same process level, and there were differences in different directions.

According to the development requirements of efficient cooling and safety and reliability of turbine blades, based on the surrogate model, the sensitivity analysis of the typical structural parameters of the impingement/effusion cooling structure such as injection angle, film plate thickness, impingement distance, spacing distance and diameter ratio of the film and impingement holes, affecting overall cooling effectiveness and maximum thermal stress was performed. Two optimization schemes, maximizing overall cooling effectiveness and simultaneously increasing overall cooling effectiveness and reducing maximum thermal stress, were realized. Results showed that the high thermal stress region appeared near the film cooling holes; the injection angle is the main factor affecting the cooling effectiveness and maximum thermal stress. There was a competitive relationship between the two optimization objectives. The overall cooling effectiveness was increased by 2.9%, and the maximum thermal stress was reduced by 12.5% through multi-objective optimization method.

The effects of different chemical reaction mechanisms on the flame surface database and simulation results based on the steady laminar flamelet model (SLFM) were analyzed by using one-dimensional flame analysis and constructing a flamelet database. Based on the large eddy simulation (LES) program AECSC (aero engine combustor simulation code) software, the SLFM model was combined with DRG (direct relation graph) method simplified mechanism, Smooke mechanism, and GRI 3.0 detailed mechanism to simulate Flame D, E, and F jet flames, among which the average temperature and pulsation values of GRI 3.0 mechanism were the closest to the experimental data. The LES-SLFM model was faster and had comparable overall accuracy compared with the LES and probability density functional transport equation turbulent combustion (TPDF) models. The chemical mechanism affected the flamelet database and thus the time and accuracy of the simulation. The LES-SLFM model combined with the detailed mechanism was fast and had suitable accuracy, which should be further tested in the combusor simulation, presenting potential and value for future applications.

The essence of dynamic combustion was explained. Based on the requirements of both military and civil aircraft engines, the importance of conducting dynamic combustion technologies and supporting the optimization design of aero-engine combustors was highlighted. The research progress on dynamic combustion technologies was summarized and compared from domestic and foreign aspects in terms of fundamental research, model combustor research, numerical simulation, and engineering application. The challenges in developing dynamic combustion technologies were analyzed, four key technologies to be developed in China were presented, and their connotation and technical pathways were expounded. Based on the design requirements of advanced engines along with current dynamic combustion progress, suggestions including establishing database, developing design methods, building theoretical system and training research teams, were put forward for developing dynamic combustion technologies in China.

A numerical investigation was conducted to study the effect of swirl numbers of the main stage and the pilot stage on the flow characteristics of a concentric staged lean direct injection (LDI) combustor. The swirl numbers at different reference cross-sections were used to measure the swirling intensity of the two stages respectively. The results showed that the width and reverse flow rate of the central recirculation zone (CRZ) increased with the increment of S_v (the pilot stage outlet swirl number). The S_v constituting a threshold for significantly enhancing the reverse flow intensity of the CRZ was between 0.60 and 0.64. Upstream the CRZ vortex center, the main stage swirling flow brought about the radial expansion of the CRZ, but downstream it suppressed the CRZ and reduced the reverse flow rate. The flow characteristics changed significantly when S_n (the main stage swirl number at the flow passage outlet) increased from 0.46 to 0.60. S_v and S_n provided a better measure of the swirling intensity levels of the two stages respectively, relative to the swirl numbers defined at the outlet of the swirler vanes.

In order to investigate the effects of atomization characteristics on the lean blow-out (LBO) performance, an experimental study was carried out on the spray and the combustion. The LBO limits and outlet temperatures near the LBO condition of different atomizers were measured and the combustion efficiency was estimated in a single dome rectangular model combustor with a dual-radial and a dual-axial swirl cup, respectively. The results showed that, （1） the LBO performance of the combustors with different structures and atomizers was very different. The LBO performance of the dual-radial combustors was generally better than that of the dual-axial combustors. （2）There was no obvious correlation between spray Sauter mean diameter （SMD）and LBO limit, and SMD alone cannot represent the influence of all atomization characteristics on LBO performance. （3）The hollow spray kept the flame base away from the recirculation zone, and the solid spray kept the flame base concentrated in the recirculation zone. （4）The combustion efficiency and outlet temperature of combustors with different atomization characteristics were very different near the LBO. The outlet temperature at LBO represented the ideal LBO limit. The combustion efficiency reflected the difference between the actual LBO limit and the ideal LBO limit.

To study the atomization performance of the centrifugal nozzle in the initial atomization stage, a traditional phase Doppler particle analyzer (PDPA) and a digital off-axis holography (DOH) were used. Using aviation kerosene RP-3 as the working fluid and maintaining the fuel pressure at 0.8 MPa, the atomization performance of the centrifugal nozzle at different fuel temperatures (240—300 K) was studied. The experimental results showed that: the spatial distribution of the Sauter mean diameter (SMD) in the initial atomization stage presented a “single peak” distribution, and with the increase of the axial distance, the peak value of the SMD increased, and the peak position of the SMD moved to the outside; for the same axial position of the initial atomization stage, because the change of fuel temperature could affect the length of the liquid film and the process of droplets breaking, the fuel temperature had no obvious regularity to the SMD distribution; the algorithm of droplets identification could identify the overlapping droplets as liquid filaments or irregular droplets. And the algorithm might exclude overlapping droplets from the statistical range, so the SMD and the peak position of the differential distribution of droplets size and number measured by DOH was smaller than that of DOH; DOH can directly observe the breaking process of the liquid film and the distribution of droplets, helping to analyze the experimental results.

In order to investigate the main characteristics of the liquid film formed by impinging jets on solid walls, the shape and thickness of the liquid film were investigated experimentally based on ultraviolet light emitting diode induced fluorescence （LEDIF）and a high-speed camera. The experiment results showed that the length and width of the liquid films on the curved wall and flat wall increased as the jet velocity increased. As the airflow velocity increased, the lengths of the liquid films on both the flat and curved walls increased, while the widths decreased. The width of the liquid film increased slightly, but the length of the liquid film increased obviously as the radius of curvature increased. The thickness of the liquid film on the flat and curved walls decreased gradually as the jet velocity increased. The transition occurred when the jet velocity reached the critical value, and the thickness of the liquid film increased rapidly. The critical velocity of the liquid film on the curved wall was 19.10—25.08 m/s, while that of the liquid film on the flat wall was 25.08—35.92 m/s, approximately. As the airflow velocity increased, the thickness of the liquid film on the flat wall decreased gradually, while the thickness of the liquid film on the curved wall increased when x=0—55 mm, and decreased when x>55 mm. For the different radius of curvature, the liquid film thickness had a shape of “W” along the circumferential direction Ψ₁. And the “W” was flattened gradually as the radius of curvature increased, but the thickness in the middle ( Ψ₁ =0°) was kept unchanged.

By solving the thickness and velocity distribution of the liquid sheet formed by an oblique liquid jet impinging onto a wall, in combination with the energy equation and empirical approximation, a semi-empirical model was established to predict the liquid sheet boundary. The model can directly describe the influence of various factors without numerical iteration of the complex equations. To verify the accuracy of the model, experimental studies were carried out, and the influences of the jet velocity, impact angle, viscosity and surface tension were analyzed. Then, the model results were compared with the experimental results. It was found that the model had a good prediction ability for the liquid sheet shape. The correlation coefficients of the experimental and model boundary curves under all working conditions were larger than 0.99. The model also had a high prediction accuracy even for the downstream complex flow area, with the error about 1%.

To improve the ignition performance of the aero-engine combustor, different configurations of two-stage axial swirlers were designed based on the swirler design parameters. Ground and high-altitude ignition tests were carried out to obtain the ignition fuel-air-ratios for a model combustor equipped with the designed swirlers. The effects of outer swirl number, the air split ratio between the inner and outer swirlers, and Venturi throat size on combustor ignition performance were analyzed. The results showed that under the same total air flow rate and inner swirl number, when the outer swirl number increased from 1.77 to 2.15, the combustor ignition performance first decreased and then increased, and the ignition characteristic was better when the outer swirl number was smaller. The inner and outer swirler flow ratios had a significant impact on the combustor ignition performance under high altitude and low pressure conditions, and the ignition characteristic was the best when the ratio was 3∶7. If other structural parameters were kept unchanged, the Venturi throat radius had a significant impact on the high altitude ignition performance of the combustor, and it had good ignition characteristics when the throat radius was 8 mm.

In order to investigate the properties of various physical parameters in the thermoacoustic oscillation in the Rijke Tube, the flame temperature was measured by single-end tunable diode laser absorption spectroscopy (TDLAS), at the TDLAS measurement frequency of 5 kHz, which effectively revealed the properties of temperature in thermoacoustic oscillation. Result showed that, the temperature varied regularly at approximate 230 Hz, which coincided with the eigen frequency of the Rijke tube. The static pressure at the exit of the Rijke tube fluctuated at the same frequency. Besides, the flame chemiluminescence intensity was captured by the high-speed camera. The result showed that the flame chemiluminescence intensity area also fluctuated periodically due to the variation of static pressure. The frequencies of temperature, static pressure and chemiluminescence intensity were consistent with each other.

According to the flow mechanism and organization method of squealer tip, the analysis of aerodynamic and heat transfer characteristics of squealer tip considering cooling, and the modeling of leakage flow of squealer tip were summarized. The results showed that the flow inside the cavity had an obvious effect on the heat transfer. There was a strong interaction between the cooling gas in the cavity and the leakage flow. Reasonable cavity shape and jet hole position can effectively improve the aerodynamic performance and reduce the heat load of the blade tip; The uncertainty of blade tip machining and aerodynamic parameters could significantly affect the performance of squealer tip. By modeling the vortex structure in the cavity, the performance prediction model of the squealer tip considering the compressibility of air jet was verified by experiment and numerical simulation results. The model can effectively evaluate the performance of the squealer tip and provide a reference for engineering design.

The three-dimensional flow field of an aeroengine flow pipe measured by area-integrated method was numerically simulated, and the effects of radial distribution, axial position, rake size, total pressure rake, total temperature rake and static pressure measurement position on the flow measurement by area-integrated method were analyzed under different Reynolds numbers. The results showed that, with the area-integrated method of high precision measurement flow, the measurement layout can completely and accurately reflect the flow field information on the whole measurement section, especially in the boundary layer of the flow pipe ring wall, and the measurement points should be arranged as much as possible. Too many measuring points may result in excessive number and oversize of measuring harrows, causing great interference to the flow field in the flow tube and reducing the flow measurement accuracy. On the basis of the equal torus measuring point layout scheme, the flow measurement error was not be changed if the measuring points in the main flow area were reduced appropriately, thus shortening the size of the measuring harrow, reducing the clogging ratio of the measuring harrow and the aerodynamic force of the measuring harrow, and significantly saving the test cost. However, it is necessary to check the applicability of the equal torus truncation scheme and the axial measurement section position in the whole working range of the flow pipe. The closer the measuring section to the rake body indicated the greater the relative error of flow measurement. The length and height of rake had little influence on the relative error of the flow measurement, but the width of the rake had great influence on the relative error of the flow measurement. In order to reduce the clogging ratio of the measuring rake of flow pipe, it is recommended that the total pressure rake should be arranged in the upstream of total temperature rake, and the measuring section of static pressure and total pressure should be arranged in the same axial position. The radial layout and number of total temperature or total pressure measuring points, the axial position of total temperature rake or total pressure rake and static pressure measuring section directly affected the error of measuring flow by area-integrated method. Special attention should be paid to the test layout of each rake when flow testing scheme was made.

To evaluate the measurement accuracy of five-hole probes, the method of evaluating measurement uncertainties based on statistical Monte Carlo method (MCM) was developed and compared with the maximum error limit method and the guide to the expression of uncertainty method (GUM). At the same time, the influence of sampling numbers M on the evaluation result of MCM was studied. The method was applicable to measurement models without mathematical expressions, and the influence of nonlinearity of the model can be considered. The probability density function (PDF) can more scientifically characterize the distribution of the input quantity, instead of being limited to the normal distribution. To verify the method, a five-hole probes was calibrated and used in a standard wind tunnel. The results showed that for the static pressure, the shortest 95% coverage interval provided by MCM was 11.1% smaller than the GUM. The difference accounted for 33.3% of the standard deviation. Furthermore, using MCM, the non-linear influence in data processing procedure can be considered compared with GUM. The MCM was applied to evaluate the measurement results of the flow field behind the cascade using the five-hole probes. The results showed that the uncertainty distribution of each parameter was similar to the error distribution in the whole measured cross section. The uncertainty was larger in blade tip leakage vortex area. However, it can avoid the trouble that the relative static pressure can not be expressed, and the influence of gross error can be removed. The statistical error of MCM itself can be solved by appropriately increasing M while considering the computer performance and time cost.

A numerical simulation method of the whole aero-engine based on the circumferentially averaged method was developed independently. Based on the Navier-Stokes equation, the governing equation of the circumferentially averaged throughflow model was deduced. In view of the shortcomings of the traditional model for the source terms of the equation, an inviscid blade force model considering the influence of airfoil, a compressor spanwise distribution loss model based on machine learning, and a circumferential non-uniformity model based on theoretical analysis were proposed. On this basis, the quasi-3D modeling of the combustion chamber was further completed, and finally the quasi-3D steady state simulation of the whole aero-engine was realized. The whole aero-engine simulation of turbojet engine WP11 was completed by using the whole aero-engine circumferentially averaged steady state simulation program CAM developed, and the simulation results of the Russian S2 program AES-S2 were compared and analyzed. The results showed that compared with the Russian S2 program AES-S2, the circumferentially aver-aged quasi-3D simulation program developed had higher simulation accuracy. In comparison of the quasi-3D calculation results of WP11’s design point, the error of turbine mass flow calculated by CAM was less than that of AES-S2 by over 8%. In terms of engine thrust, the error was reduced by more than 16%. The convergence of CAM was better. The amplitude of the turbine mass flow calculated by CAM was less than that of AES-S2 by over 10%. In terms of engine thrust, the amplitude of the calculation result was reduced by more than 20%.

To explore the reasons for fatigue failure of inlet guide vane (IGV) downstream the strut in the transition section of a gas turbine compressor, unsteady numerical simulation and one-way fluid-structure coupling methods were used to analyze the effects of strut on unsteady flow in 1.5 stages high-pressure compressor and forced response characteristics of the IGV. Furthermore, the analysis results were checked by fatigue strength experiments. The results showed that the isentropic efficiency in the design point of compressor was reduced by 3.6 percent points compared with the model without strut. Vortex shedding alternately at the trailing edge of the strut caused the inlet attack angle of the IGV to deviate from the design value, resulting in a decrease in aerodynamic performance and a significant increase in the unsteady pressure pulsation on the IGV surface. With the increase of circumferential distance between IGV and strut, the influence of strut on IGV first increased and then decreased rapidly. The perturbation frequency resulted from the shedding vortex was relatively disperse, and the low-order resonance of IGV induced by the low frequency component was the fundamental reason of the fatigue failure of IGV; the maximum vibration stress can reach 400 MPa. The numerical simulation results were consistent with the experiment, proving the reliability of the simulation.

A key technique for reducing the aerodynamic excitation of downstream rotor blades is the asymmetry layout of inlet guide vanes. The half and half layout divided can be defined in terms of both the number of blades and the vane spacing. This study derives a theoretical solution to aerodynamic excitation with half and half layout and compares the two design schemes’ differences. We compare the aerodynamic excitation properties under two asymmetry schemes using a 1.5-stage compressor as an example, examine the spectral properties of the aerodynamic forces on the rotor blade surface, investigate the relationship between the excitation properties and the vane spacing, and identify the frequency separation mechanism of the aerodynamic excitation for half and half layout. The findings demonstrate that altering the guide vane row’s vane spacing causes the aerodynamic excitation to be frequency divided, the amplitude distribution of the aerodynamic response of rotor blades is affected by the propagation law of aerodynamic excitation, and that the spatial distribution and frequency domain characteristics of the pneumatic excitation differ between the two asymmetry schemes.

The second stage rotor blade of multi-stage compressor was investigated to efficiently calculate the response caused by aerodynamic excitation. The compressor was reduced through the inlet boundary correction, and combined with the harmonic method to improve the calculation efficiency of unsteady aerodynamic force. The relationship between frequency and response in modal space and the harmonic analysis method were used to further improve the solution efficiency. When the forced vibration response analysis under aerodynamic excitation was urgently required in engineering, the blade vibration features at resonance can be evaluated efficiently. The results showed that the excitations of the second rotor blade was mainly subject to the upstream stator passing frequencies, and the influence of the downstream stator was limited. The 15th modal shape was excited near the resonance condition and there was significant vibration stress in the blade tip area. Under the given damping and simulated working conditions, the vibration stress was about 71 MPa in the examined position and direction obtained by the transient response analysis, and it was about 92 MPa under the resonance condition evaluated by the harmonic analysis method.

In order to investigate the effects of airfoil probe head on the transonic flow field, a numerical simulation was performed in the transonic turbine cascade with airfoil probes installed at the different heights of blade’s leading edge. The variations of blade’s load performance, vortex structure, flow loss and applicability of probes with different incidence angles were analyzed. The results indicated that the airfoil probe’s head affected the loading performance of the blade, and the effect was quite sensitive to the flow incidence angles. A long streamwise vortex induced by the probe head was formed. And an attached vortex layer appeared on the suction surface of the blade at a large positive incidence angle. This attached vortex layer was the main factor for the decline of the load performance of the blade with the airfoil probes. The proportion of flow loss contributed by probes at each slice of the cascade passage decreased gradually along the flow direction. Compared with original cascade, the flow loss behind the cascade increased by 7.4% at high incidence angle. The probes installed at different positions in the spanwise direction had good applicability in the whole range of adjustable incidence angle.

In order to extend the measurement range of five-hole pressure probe (FHPP), the probe was calibrated firstly in a low-speed wind tunnel. The extending method was applied to process the calibration data, the differences of the probe calibration curves defined by different calibration coefficients were compared, and the measurement range of the FHPP in low-speed flow was determined. On this basis, the applicability study at different Mach numbers was carried out in a subsonic wind tunnel to evaluate the application effect of the extending method in compressible flow. Finally, actual measurement was carried out in the wind tunnel to analyze the error level of the extending method. The results showed that: the calibration curve had a wider monotonic range and higher calculation accuracy after using this calibration coefficient definition. At low speed, the FHPP was applicable within the measurement range of yaw angle ±70° and pitch angle ±50°. At subsonic speed, the compressible flow had a great influence on the negative measurement range of the pitch angle, but the measurement range of the yaw angle can still reach close to ±70°. The error analysis results also showed that the error level measured by the extending method was generally lower. This method for extending the angular measurement range of FHPP has a good application prospect in some turbomachinery internal flow tests with a large angle variation range.

In order to study the influence of particles on the elbow tube, the particle track model was used to calculate the gas-solid two-phase flow in the elbow tube, and the fluid-thermo-structure coupling model was applied to calculate the thermal response of the elbow tube under the two-phase flow. Finally, the influence of the particle size was studied. Results showed that solid particles were gathered on the region outside of the elbow tube near the outlet, such that the temperature and plastic strain of the inner wall at the gathering region rose about 280 K and 60%, with the corresponding local fatigue life decreasing by 48%. Particle size affected the aggregation location and concentration of solid particles, as well as the temperature and plastic strain of the wall near the aggregation location. As the particle size increased, the average volume fraction of the aggregated particles increased and then decreased, resulting in an increase and then a decrease in the internal wall temperature and plastic strain at the gathering region, with the opposite change in local fatigue life. These three quantities reached extreme values when the particle size was around 8µm, with 1042 K, 0.016697 and 244 cycle lives in that order.