The computational model for two-phase flow including air and supercooled water droplet was established under nanosecond pulse dielectric barrier discharge （NS-DBD） actuation. The compu-tational fluid dynamics method was adopted to simulate the water impingement characteristics of airfoil under NS-DBD actuation. The influences of angle of attack, liquid water content, mean volume diameter and intensity of actuation on the water collection coefficient were studied. Finally, the aerodynamic effect and thermal effect during the process of plasma discharge were coupled to analyze the anti-icing mechanism of plasma actuation. It showed that, the thermal and mechanic effects coupled resulting from the NS-DBD plasma discharge can reduce the peak value of water collection coefficient, but can’t change the water impingement limit of airfoil. The best effect of actuation appeared at the time of 10—15 μs. The higher the strength of actuation increased, the lower the water collection coefficient decreased and the slower it rose again. The actuation had a weaker influence on the larger mean volume diameter. The mechanics and thermal effect from plasma actuation can generate a layer of wavy high temperature air film on the surface of airfoil to decrease the water collection coefficient distribution.

The study of concentration polarization, and the analysis of its influencing factors and mechanism are theoretical basis for improving the performance of hollow fiber membrane and the separation efficiency. By establishing a mathematical model of hollow fiber membrane separation of oxygen and nitrogen, the influences of parameters, such as pressure difference between the inside and outside of the filament, initial oxygen concentration of intake air, intake air flow rate and back pressure outside the filament on the polarization coefficient, were analyzed to explore the influence degree of operating parameters on the concentration polarization, and a technical scheme was put forward to reduce the concentration polarization. Factors such as pressure difference between the inside and outside of the filament, initial oxygen concentration of intake air, intake air flow rate, and back pressure outside the filament had a certain effect on the polarization coefficient, among them, the polarization coefficient was positively correlated with pressure difference inside and outside of the filament, initial oxygen concentration of intake air, and back pressure outside the filament, and was negatively correlated with intake air flow. Negative pressure suction on the permeation side can be used to increase the membrane permeation flow rate and reduce the concentration polarization. The influence of concentration polarization on membrane performance was reduced by controlling various factors.

The fluid-solid-thermal coupling theory of the split floating ring seal was analyzed, and a numerical solution model of fluid-solid-thermal coupling of split floating ring seals with/without grooves was established. On the basis of verifying the accuracy of the numerical method, the flow field characteristics, mechanical characteristics and leakage characteristics of the segmented floating ring seal without shallow groove, rectangular groove and triangular groove under the coupling effect of high temperature and high pressure with different inlet pressure, temperature and speed were studied. The results showed that the fluid-solid-thermal coupling numerical solution model of three kinds of shallow-groove structure floating ring seal can accurately solve the flow field characteristics, mechanical characteristics and leakage characteristics. The fluid moved along the circumferential direction under the action of high-speed shear force, and the local high-pressure zone was generated in the shallow groove structure due to extrusion, which formed the dynamic pressure effect. Under the same inlet pressure condition, the split floating ring seal with triangular groove structure had the largest opening force. When the inlet pressure was 600 kPa, it increased by 9.6% compared with the structure without shallow groove, and the leakage amount was kept relatively small. The temperature had little effect on the opening force of the floating ring, and the higher temperature indicated the smaller leakage amount, the maximum reduction was 6%. The rotating speed had a great influence on the opening force of the floating ring of rectangular groove and triangular groove structure, and the maximum increase was 3.2%, but had little influence on the closing force and leakage.

Experimental study of U-shaped crater holes with depth and expansion angle on a flat plate using pressure sensitive paint technique was conducted to measure and compare the adiabatic film cooling efficiency of cylindrical holes and crater holes. The results showed that the cooling performance of the cratered hole structure was better than that of the cylindrical hole structure, while the cooling performance of the cratered hole structure with expansion angle was better than that of the depth crater hole structure. The average film cooling efficiency of the optimal depth crater hole (crater depth d=0.5D, blowing ratio M=2.0, D meant diameter of cylindrical hole) compared with the cylindrical was improved by nearly 800% in the region of X/D<5 (X meant the position in mainstream direction) and nearly 700% in the region of 5<X/D<14; the average film cooling efficiency of the optimal expansion angle crater hole (expansion angles α=30°, M=2.0) compared with the non-expansion crater hole was improved by nearly 180% in the region of X/D<5 and nearly 170% in the region of 5<X/D<14. The increased depth of the crater hole facilitated the expansion of the coolant on the wall surface; the expansion angle crater hole structure facilitated the expansion of the cooling jet coverage in the spreading and flow direction. The magnitude of the enhancement of the crater depth and expansion angle on the film cooling performance of the crater hole structure increased with the increasing blowing ratio.

When the temperature and flow velocity of the freestream were 50 ℃ and 10 m/s, respectively, ground experimental studies on frosting and defrosting of the microtubule heat exchanger were carried out, respectively, under the conditions of two different freestream humidity values (1.8 g/kg and 6.4 g/kg). The ground experimental results showed that the freestream humidity value had significant influence on the frosting and defrosting performances of the microtubule heat exchanger. In the frosting experiments, the amount of frost layer increased with the increase of the humidity value of the freestream, but the relationship between them was not proportional. The greater freestream humidity value indicated the lower wall surface temperature of the heat exchanger tube bundle, and the higher growth rate of the pressure loss coefficient of freestream indicated the lower heat transfer rate of heat exchanger. However, in the defrosting experiments, the amount of frost layer and the pressure loss coefficient of freestream did not increase with the increase of freestream humidity value. And with the increase of freestream humidity value, the wall surface temperature of heat exchanger tube bundle increased, and the heat transfer rate of the heat exchanger decreased.

Self-pulsation in fuel injection is an important cause of combustion instability, and the "gas-like" nature of supercritical fuel makes its self-pulsation yield a more significant effect on oil-gas mixing and combustion. The self-pulsation of supercritical kerosene jets in the rotating flow field was investigated to find out the causes of the self-pulsation and the ways to suppress it. The supercritical kerosene injection processes under different companion flow conditions, recess ratios, and gas-liquid area ratios were investigated by using a high-speed schlieren technique using the proper orthogonal decomposition. The results showed that the oscillation modes of self-pulsation were represented by axial and radial oscillations, and the accompanying gas affected the self-pulsation region. When the gas-liquid area ratio was 674, the cause of self-pulsation was oil-associated gas interaction, and the self-pulsation decreased with the decrease of the recess ratio; when the gas-liquid area ratio was 416 and 178, the cause of self-pulsation was associated gas-environment interaction, and the self-pulsation increased with the decrease of the recess ratio.

To explore the phase distribution characteristics of the oil-gas two-phase flow in scavenge pipe of aero-engine lubrication system, and achieve accurate flow pattern identification, a flow pattern identification model with image processing was proposed based on the typical flow patterns images obtained from a horizontal pipe under the working condition of aero-engine. Four typical flow patterns emerged in this experiment: slug, stratified, wavy, and annular flow. By image processing technologies such as bilateral filter, binarization, and wavelet decomposition, feature parameters were extracted from the images and used as input to the model. The identification model based on Elman Recurrent Neural Networks was established through training and verification, and it can successfully identify four different flow patterns. The identification accuracy of the model was 93.06%, and the robustness index macro-F1 was 97.60% on the verification set.

A well-defined visualized single-head swirl combustor was designed to study nanosecond pulsed plasma ignition and flame propagation characteristics of ethanol spray in combination with experimental investigation and numerical simulation. High-speed flame shadow and CH^* chemiluminescence images were taken simultaneously to explore the time scale of the flame propagation process at a wide range of ignition locations. The successful ignition process from different discharge positions can be eventually divided into four phases: flame propagation, combustion of wall-attached fuel, flame morphology transformation, and flame morphology stabilization. The ignition delay time was analyzed by a statistical method, revealing that the optimal ignition location in the combustor was at the inner boundary of the central recirculation zone, and the optimal ignition location near the wall must correspond to the widest part of the central recirculation zone.

The method of fixing gas mass flow rate and adjusting gas flow temperature is generally used to achieve aerodynamic thermal equivalent simulation in high temperature gas flow thermal tests. A method was proposed to quickly determine the test parameters of the gas flow by using the numerical calculation results of the high temperature gas flow and establishing a heat flux surrogate model, thus forming a closed-loop control method with the gas flow temperature as the control target. Kriging surrogate model method was used to establish the cold wall heat flux surrogate model under different flow rates and gas flow temperatures, and the hot wall heat flux surrogate model under fixed gas flow rates, and different gas flow temperatures and specimen surface temperatures. The average relative errors of the predicted heat flux at the test operating points were 0.27% and 0.48%, respectively, indicating high prediction accuracy. Further, the mass flow rate and temperature of gas flow were determined according to the aerodynamic thermal simulation requirements of 9 typical status points of a flight envelope. The relevant numerical calculation verification showed that the maximum relative error between the high-temperature gas flow heating hot wall heat flux and the aerodynamic thermal simulation requirements of each status point was 3.04%, and the average relative error was 0.87%, realizing the full-time sequence high-precision simulation function.

Artificial hail is mostly used in aircraft shell and engine testing, and its mechanical properties have a direct impact on the test results. Considering the storage problem of artificial hail after production, the influences of storage parameters (storage temperature and time) on the mechanical properties of artificial hails were explored with the help of experimental research methods. In terms of storage temperature, the results showed that the estimated maximum compressive strength values of pure ice hails and cotton-containing hails increased to a certain extent at −10, −15 ℃ and −25 ℃ storage temperatures, and the maximum increment was 73% and 80%. In terms of storage time, the results showed that after storage for 5, 10 h and 20 h at −20 ℃, the estimated maximum compressive strength values of pure ice hails and cotton-containing hails showed a trend of first decrease and then increase, and the maximum value was obtained under the condition of 20 hours of storage, which increased by 113% and 54%, respectively, compared with the unstored hails. Therefore, the influences of storage parameters on the mechanical properties of hail cannot be ignored and need to be considered during airworthiness testing.

In order to study the effect of spark energy and phase on the ignition success rate of aeroengine combustion chamber, large eddy simulation （LES） method was used to simulate the ignition process of axial radial cyclone combustion chamber. The simulation results showed that the ignition was still random after increasing the spark energy to 10 J. Not all single pulse spark ignitions can establish a stable flame under the same combustible condition, and the influence of spark phase on ignition success rate was not improved after increasing the spark energy. The single spark energy release time of 0.2 ms was far less than the flow field pulsation period of 8.76 ms, mainly accounting for the serious influence of spark phase against the ignition result. Under the premise of the same total ignition energy, the multi-pulse ignition was carried out in the spark phase where the single pulse ignition failed. It was proved that using multiple spark pulses instead of single spark pulse ignition is an effective way to reduce the influence of randomness on ignition success rate in a flow field pulsation period.

In order to give full play to the low pollution emission characteristics of the low-swirl combustion technology, the influence of oxygen concentration change on low-swirl combustion was investigated experimentally. Based on the low swirl burner, using propane as fuel, by changing the CO₂ mass flow rate through the system to adjust the premixed oxygen concentration and by controlling the air flow rate out of the nozzle, gas ratio and premixed oxygen concentration, the impact of oxygen concentration on low-swirl premixed combustion in flame form, combustion flame temperature and combustion emission product NOx was explored. Under the conditions of the test system, the lowest oxygen concentration can be achieved at the flow rate of 3.0 m/s at the nozzle, the gas ratio of 0.65, and the lowest critical oxygen concentration of 17.43%. It was found that the limit of fuel flameout increased linearly with the increase of air velocity at nozzle. With the decrease of oxygen concentration, the central incompleteness of low-swirl combustion became smaller and the flame stability of combustion nozzle became stronger. The overall temperature of the flame showed a decreasing trend. The emission of NOx from combustion was gradually reduced.

In order to improve the aerodynamic efficiency of the ducted propeller, high machining and assembly accuracy were required to maintain a small clearance between the blade tip and the duct, resulting in a significant increase in the processing, operation and maintenance costs. To solve this problem, the design philosophy of improving hovering efficiency with large blade clearance was studied based on the grooved duct configuration. Calculation was carried out based on Reynolds averaged NS equation and multiple reference frame model. Firstly, based on the baseline design of a conventional ducted propeller, it was found that when the blade tip clearance ratio increased to four times the reference tip clearance ratio, the hovering efficiency decreased by more than 25%. Therefore, the increase of the blade tip clearance led to a significant decrease in the hovering efficiency. Secondly, based on parametric analysis, it was found that tip position and groove shape are key variables determining the hovering efficiency. By properly extending the blade tip into the groove, the hovering efficiency can be equivalent to that of the baseline design when the blade tip clearance ratio reached 6.7 times of the reference tip clearance ratio. If the circular groove was used, the hovering efficiency can be increased by 5.2% when the blade tip clearance ratio was 2 times of the reference tip clearance ratio. The results of numerical simulation and experiments indicated that the grooved duct configuration can improve the aerodynamic efficiency and reduce costs with a large tip clearance ratio.

To analyze the impact of rotor speed changes on handling quality, a helicopter flight dynamics model coupled with engine dynamic characteristics was established. Then, an engine aerodynamic thermodynamic model and a rotor flight mechanics model were established. On this basis, a helicopter flight dynamics model includeing the dynamic characteristics of the engine, aerodynamic characteristics of various components of the helicopter, and aerodynamic interference between components was established. Taking the UH-60A helicopter as an example, the impact of rotor speed changes caused by engine power turbine output on helicopter handling quality was studied. Research showed that when the rotor speed decreased, the power required by the helicopter decreased, while the rotor counter torque and total pitch increased. The torque characteristics, altitude characteristics, and total pitch yaw coupled with the handling quality all deteriorated with the decrease of rotor speed. The bandwidth of the three channels of helicopter roll, pitch, and yaw increased as the rotor speed decreased.

In order to accurately obtain the yaw, pitch and main thrust performance of an engine, a four-fulcrum and six-component engine testing platform was developed on the basis of previous studies. This platform integrated thrust test and static performance calibration of an engine, and realized thrust measurement of a certain engine. The measuring system and static performance calibration method were introduced, and the static calibration test of force measuring unit was carried out. The test results showed that the designed dynamic six-component force tester had high measuring accuracy, with the overall relative error of 0.31%.

In order to clarify the effect of area ratio of the first bend on the flow heat transfer and structural response of serpentine nozzle, the effect of the thermal-solid coupling response under different area ratios of the first bend was investigated using the serial two-way loose coupling method based on the multi-physics coupling analysis software MPCCI (mesh-based parallel code coupling interface). Results were obtained as follows: the overall heatflux distribution of the nozzle under the action of large curvature and multi-bend structure was non-uniform, while the heatflux distribution of nozzles was similar; the strongest heat transfer occurred at the first bend of different nozzles, and with the increase of the outlet area ratio of the first bend, the heat flux at the first bend decreased gradually; the vortex structure was generated at the exit isometric section, making the heat transfer blocked. In structural stress response, the maximum stresses of all nozzles appeared at the exit end of the upper wall. Serpentine nozzle with the first bend outlet area ratio of 0.8 at t=39.92 s time first exhibited the maximum stress, with the decrease of area ratio of the first bend, the maximum stress value of each nozzle delayed the emergence of the maximum stress value was gradually reduced, compared with area ratio of the first bend of 0.5 ahead of the 7.12 s, the maximum value of the stress was reduced by 7.4%.

Leading edge slat can inhibit the flow separation at large angles of attack and effectively improve the aerodynamic performance. Determining the geometry parameters of slat is vital to improve the aerodynamic performance. Therefore, the aerodynamic performance of leading edge slat was optimized based on multi-objective genetic algorithm. NACA 0012 was selected to design the geometry parameters of slats, and the input parameters such as slat width, position and curves degree were defined. The optimized Latin-hypercube method was used for sampling. The sample data were used to establish the input parameters and the surrogate model with lift and drag coefficient as the target parameters to improve the optimization efficiency, and the multi-objective genetic algorithm was used to optimize the target parameters in the glob-al scope. Results indicated the lift coefficient increased obviously with the large curves degree of slat. Leading-edge slat improved the aerodynamic performance by increasing the turbulent kinetic energy of upper wing sur-face and reducing the pressure coefficient of the front edge of the upper wing. After optimization, the maximum lift coefficient of the airfoil with slat was increased by 12.5% and the stall angle of attack was delayed by 5°. It could provide theoretical support for the optimization design of airfoil aerodynamic performance with slat.

TNO (Netherlands Organization for Applied Scientific Research) model, which is built based on the modeling of the wave number-frequency spectrum of turbulent boundary layer fluctuation pressure, is a rapid prediction method for trailing edge scattering noise. However, the wave number-frequency spectrum used in the TNO model omitted the TT (turbulence-turbulence) term from the fluctuating pressure source terms, leading to significant deviations in noise prediction at mid-to-high frequencies. To address this, the flow field input of the model was optimized, and the TT term of the Chase Ⅰ model was introduced to improve the accuracy of the model. Based on the far-field noise data of NACA0018 airfoil obtained from wind tunnel experiments, the effectiveness of the improved model was verified. The results showed that compared with the original model using XFOIL method to calculate the flow field as the input of wave number frequency spectrum, using RANS （Reynolds-averaged Navier-Stokes） method to calculate the flow field input was more accurate; the introduction of TT term had a significant impact on the prediction of wave number frequency spectrum amplitude in high wave number regions; the improved fast prediction method for trailing edge scattering noise significantly improved the prediction accuracy of high-frequency noise; the original TNO model had an issue that noise prediction deviations increased with the angle of attack, but this was substantially mitigated in the improved model.

In order to improve the aerodynamic performance of dragonfly-inspired tandem flapping wings, the influences of three parameters, i.e. pitching amplitude, wing spacing and phase difference between forewing and hindwing, on the lifting efficiency of dragonfly-inspired tandem flapping wing were systematically analyzed by the combination of the neural network and CFD. The results showed that pitching amplitude, wing spacing and phase difference between forewing and hindwing had important influence on the aerodynamic performance of dragonfly-inspired tandem flapping wings. In the analysis parameter range, compared with the flapping wing with worst parameter combination, the lifting efficiency of the flapping wing with the best parameter combinations increased 90.33% by using neural network optimization. Furthermore, through the analysis of flow field of dragonfly-inspired tandem flapping wing with different parameter combinations, it was found that the trailing vortices shedding from the forewing can be reattached to the surface of the hindwing for the tandem flapping wing with the optimal combination of parameter, which can weaken the vortex intensity during the upstroke of the hindwing and reduce the energy consumption of the flapping wing, so that the flapping wing can generate better aerodynamic performance.

A method of increasing stability of suspension by combining active jet device and sling feedback was proposed. The coupling model of CH-47 helicopter/elastic sling/rigid body hanging object was established. Targeting the swing and unstable yaw motion of the hanging object, an active jet device was designed to generate damping force and torque to restrain the swing and yaw of the hanging object through the thrust of the ducted fan. Considering the driving force limitation of the jet device, sling feedback was introduced to further inhibit the swing of the hanging object. The attitude control of the helicopter adopted active disturbance rejection control to suppress the disturbance caused by the movement of the hanging object. The effects of key parameters such as the sling feedback gain compensation and the maximum driving force on the stability of the slung and the energy consumption of the jet were studied. The results showed that the introduction of active jet significantly inhibited the swing and yaw of the hanging object; under reasonable parameter configuration, the combined stability augmentation method had lower energy consumption and better stability augmentation performance than active jet device alone.

Based on Hamilton’s principle and the multi-body dynamics method, a coordinated system was established to accurately describe the spatial position and motion relationship of the rotor, wing, pylon, and other moving parts. Various offsets and the coupling effects of aerodynamics, structural deformations, and inertial forces were fully considered. The influence of the coupling effects resulting from the elastic deformation of components was retained. For the necessary analytical framework, mass, stiffness, and damping matrices were derived in non-rotating coordinates. Subsequently, a sophisticated rotor/wing multi-mode coupled aeroelastic stability analysis model for tiltrotor aircraft was formulated. Utilizing this model, the impact of various parameters on the multi-mode coupling aeroelastic stability of tiltrotor aircraft was analyzed. These parameters included vertical bending, chord bending, and torsional stiffness of the wing, as well as the wing forward sweep, mast height, hub pre-cone angle, rotor flapping stiffness, and pitch-flap coupling coefficient. The results indicated that the system’s stability was most sensitive to the torsional stiffness in three directions of wing stiffness. Furthermore, the wing forward sweep and the pitch-flap coupling coefficient did not effectively suppress whirl flutter. Increasing the mast height and the wing torsional stiffness can keep the whirl flutter boundary unchanged. Additionally, placing the tilting hinge closer to the rear of the elastic shaft and increasing the rotor flapping stiffness can enhance the whirl flutter stability margin. Moreover, comprehensive consideration of the hub pre-cone angle and suspension height can increase the whirl flutter stability margin in a more efficient manner.

The model vibration in wind tunnel test may easily lead to the coupling of oil flow velocity, while the traditional optical flow method cannot identify or eliminate the vibration. Therefore, an optical flow decoupling algorithm based on resection was proposed. The low resolution optical flow value solved by the interpolated image correlation method was taken as the iterative initial value of the optical flow. Then, based on the resection, a calculation model for vibration angle and displacement under vibration conditions was derived, and then the vibration displacement was eliminated to achieve speed decoupling. The simulation test showed that the AEEs （average endpoint errors） of the traditional optical flow method and the optical flow decoupling algorithm were 0.20, 0.98, 1.44, 1.63 pixel/s and 0.19, 0.20, 0.22, 0.27 pixel/s, respectively, at the simulated vibration angles of 0°, 5°, 10° and 15°, and the measurement error of vibration angle was less than 0.1°; the fluorescence oil flow test further indicated that the velocity field streamline obtained by the optical flow decoupling algorithm based on rear intersection was more accurate, smooth and clearer, which can effectively solve the velocity coupling problem. This method has certain practical engineering application value.

For the continuum/rarefied hybrid flow field formed by hypersonic aircraft in near-space, the DSMC (direct simulation Monte Carlo) method requires significant computational resources, while the CFD (computational fluid dynamics) method cannot accurately simulate rarefaction effects. Based on in-depth research on the coupling algorithm for continuous/rarefied flows, an unstructured grid coupled CFD/DSMC method was proposed. This method had the characteristics of strong universality and good adaptability, and there was no need to smooth the irregular complex interface when coupling calculation was carried out. Simulations of hypersonic flow around a circular cylinder and blunt cone were performed, and the numerical results demonstrated that the method is highly applicable to irregular interfaces and complex geometries. The effectiveness and computational efficiency of the method were validated by comparing the results with those obtained using overlapping grids. Compared with traditional DSMC methods, the computational efficiency of the two test cases was improved by factors of 2.3 and 3.16, respectively.

The progresses made since the late 20th century in addressing bird strike issues on aircraft engine fan blades, covering aspects such as the geometric modeling of birds, transient dynamic numerical analysis methods, fan blade response and damage from bird strikes, as well as key bird strike testing techniques, were reviewed. The results indicated that foreign research on bird strikes on aircraft engine fan blades entered a deep application phase, while domestic research still faced challenges such as unclear research frameworks, undefined systems, and insufficient support. It’s also suggested that future research should focus on high-precision bird modeling, bird strikes on new material blades, bird strike testing techniques for blades in rotating states, and methods for analyzing the dynamic response of the entire engine post-strike.

In view of the difficulty in extracting weak fault signals from rolling bearings in aircraft engines due to the influence of environmental noise, an aircraft engine main bearing fault feature extraction method based on threshold parameter decision screening was proposed. To adaptively select the parameters in variational mode decomposition (VMD), the particle swarm optimization (PSO) algorithm was used to optimize the parameters in the VMD algorithm, which were used as preparameters to process the raw vibration signals collected by sensors, bringing about K modal components. Furthermore, a new parameter harmonization formula was proposed. This formula balanced kurtosis and correlation coefficient and combines them into a single parameter P. Then, based on threshold parameter criteria, high signal-to-noise ratio signals were selected and integrated to generate new vibration signals. Finally, weak fault features of the bearing were extracted by using envelope spectrum analysis. The results showed that parameter harmonic formula and threshold parameter decision method can balance the relationship between kurtosis and correlation coefficient, and filter out the components with higher kurtosis value but less effective information, so the method can effectively extract fault features of rolling bearings in both simple and complex transmission paths, providing an effective means for complex signal processing and diagnosis of main bearing faults of aircraft engines.

Based on the finite element method, a ball-plane three-dimensional simulation model with the contact of the third body particles was established. The influence of the third body on fretting wear under partial slip condition was mainly studied. The contact behavior between the third body and the joint surface under the action of linear elastic material and elastic-plastic material was analyzed. The influences of different sizes on the contact behavior of the third body particles were studied. The effect of third body particles on fretting wear under different loading conditions was studied. The results showed that the elastic-plastic material can better reflect the contact characteristics of the third body. The third body particles were plastically deformed by the larger contact pressure, which reduced the contact pressure between the contact surfaces. The contact pressure of the third body particles with a diameter of 0.8 μm was the largest, and the contact pressure was the smallest at 0.2 μm, and the plastic deformation increased with the increase of diameter. In the initial stage of fretting, the existence of the third body particles could reduce the wear. For partial slip fretting wear, a smaller contact width or a larger displacement amplitude could lead to a decrease in the contact pressure of the third body particles, and an increase in friction dissipation energy and wear.

Considering the influence of material performance differences in different organizational zones, the burst speed analysis of FGH99 alloy dual property turbine disk was conducted based on the average stress method, ultimate strain method, energy method, and plastic instability method. The turbine disk burst speed tests were conducted at room temperature and high temperature of 441 ℃, and radial fracture occurred in all cases. Comparison between the analysis and the test results showed that: the predicted result was larger than test data when using the average stress method with the error of 9.42%, and the prediction of fracture mode was not accurate. The correction factor 0.80 should be considered when predicting the radial burst speed; the maximum error between the predicted burst speed using the limit strain method and the energy method and the experimental results was 0.98%, and the predicted fracture starting position was consistent with the experimental results; the plastic instability method accurately predicted the fracture mode, and the maximum error in burst speed was 2.27%.

The theory and engineering approach for on-site balancing of fan rotors in high-bypass ratio aero-engines were studied to address crucial aspects such as identifying variable amplitude speed impulse signals, establishing the correlation between vibration signals and the angular location of rotor imbalance, and expeditious balancing screw arrangement. Additionally, the investigation scrutinized fan rotor vibration in high-bypass ratio aero-engines, and a method was proposed to ascertain trial weight quantity and angle. The research also examined the trajectory of imbalance variation with rotor speed. Following a single on-site balance, fan rotor vibration experienced a remarkable reduction of over 80%, while low-pressure turbine rotor vibration decreased by more than 70%. This successful outcome validated both the balance theory and engineering methodology. Furthermore, it underscored the advantages of minimizing fan rotor vibration in mitigating vibrations in the low-pressure turbine rotor.

A full-scale rotor system test rig was designed to investigate the vibration response of the aero-engine rotor under sudden base impact load during the process of landing, based on the dynamic similarity principle and structural feature similarity principle. An experiment method for the vibration characteristics of aero-engine rotor under sudden base impact load driven by high-speed motor on the vibration and shock table was proposed. The vibration response characteristics of the rotor under axial and vertical impact loads were obtained. The influences of impact parameters, such as impact speed, impact load size, impact direction, and impact pulse width, on rotor vibration response were studied through experiments. Experimental results showed that when the rotor system received sudden base impact load, the vibration response increased significantly, and then returned to the stable state before impact. The frequency components of instantaneous vibration response were complex, and there were irregular frequency components other than the fundamental frequency. In addition, the impact parameters had significant impact on the vibration response of the rotor, and special attention should be paid to the vertical vibration response. The study can provide a reference for the vibration characteristics experiment and structural safety design of aero-engine rotor under sudden base impact load thanks to its important engineering application value.

In order to establish a life prediction method for welded joints under multiaxial tensile and shear stress conditions, a butterfly shaped test piece of GH4169 electron beam welded joints was designed and fatigue tests were conducted at different loading angles (0°, 15 °, 30°, and 45°) to obtain F-N curves respectively. The influences of different shear/tensile stress ratios γ on the fatigue performance of welded joints were explored. The critical distance method was applied to determine the characteristic size of the notch root of the welded pieces, and the effective stress parameters on the weld plane were obtained. Furthermore, relevant correction coefficients $ k(\gamma ) $ were introduced to modify the material parameters in the Findley model, and a unified multiaxial fatigue life prediction model suitable for different loading angles was established. By comparing with test results, it was verified that the proposed model had high prediction accuracy, with an error within ±2 times dispersion band.

A method for detection and diagnosis of early weak faults in rolling bearings was proposed. It addressed the challenge of timely detection using existing methods. The method utilized the Gini index to extract feature indicators from the complete life cycle data of rolling bearings, enabling timely detection of early weak faults. Additionally, it applied enhanced singular spectral decomposition and the honey badger algorithm to optimize the maximum correlated kurtosis deconvolution method for effectively decomposing early weak fault signals in rolling bearings, maximizing noise reduction and highlighting fault impact, and enabling effective diagnosis of early weak faults in rolling bearings. Tests using the Cincinnati rolling bearing lifespan dataset demonstrated that the proposed method can detect and diagnose early weak faults in rolling bearings with a lead time of 1700 minutes and 30 minutes, respectively.

Considering the issue of hot corrosion in the TC4 alloy compressor blades of aircraft engines, a non-destructive testing method based on nonlinear surface waves was proposed. Firstly, the propagation behavior of nonlinear ultrasonic surface waves with different sizes of pits was studied through numerical simulation, and the relationship between the sizes of elliptical-pitting and nonlinear coefficients was investigated. Subsequently, by establishing a model of the ultrasonic testing probe wedge block, the impact of wedge characteristic dimensions on the surface wave conversion efficiency was examined. Finally, a nonlinear ultrasonic testing system was designed and constructed to experimentally validate the numerical results. The findings indicated that the length and angle of the detection probe wedge block affected the surface wave conversion efficiency. Through numerical simulation and experimental validation, it was observed that, under the same depth-to-width ratio, the corrosion defect size and nonlinear coefficient exhibited a negative correlation, suggesting that the nonlinear ultrasonic coefficient can be used to characterize the extent of material corrosion damage.

In view of the problem that the weak fault characteristics of rolling bearings in the early stage affected by background environmental noise are difficult to be extracted, a rolling bearing vibration feature extraction and characterization method was proposed based on maximum correlated kurtosis deconvolution optimized by gray wolf optimization (GWO) algorithm. The complete ensemble empirical model decomposition with adaptive noise (CEEMDAN) was used to decompose the weak fault vibration signal disturbed by strong background environmental noise into several signal components, and the signal components were screened and reconstructed according to the kurtosis and correlation coefficient as the screening index in this method. The maximum correlated kurtosis deconvolution (MCKD) optimized by the GWO algorithm filtered out the noise components in the reconstructed signal, enhanced the weak fault feature components and performed envelope demodulation to extract the weak fault features. A comprehensive verification of the effectiveness of the vibration signal fault feature extraction and characterization method was carried out based on the rolling bearing test bench data and the real whole machine data of the aero-engine. The results showed that this method can effectively filter out the strong background environmental noise part in the weak fault vibration signal and enhance the weak fault characteristics, indicating that the peak factor of the denoising signal processed by this method increased by 2.43 compared with the original vibration signal in the turbofan engine experiment, so it effectively enhanced the shock component in the vibration signal. The method proposed can be used as one of the effective methods for fault diagnosis of aero-engine.

In order to further investigate the influence of tip winglets on the stable operating margin of a transonic compressor stage, the effects of different tip winglets on NASA Stage 37 compressor stage were studied numerically by comparing the flow field characteristics of the prototype compressor stage to the compressor stage equipped with tip winglets. Additionally, the stability expansion mechanism of the pressure-side tip winglet was revealed. The results showed that with the increase of pressure-side tip winglet width, the stable operating margin increased by 7.35%, 12.27%, 19.49% and 12.34%, respectively. According to the stability expansion mechanism of the pressure-side tip winglet, the interference between the leakage vortex and the shock wave was weakened, the flow condition in the rotor area improved, and the separation on the downstream stator suction-side was reduced. Consequently, both the low-energy fluid and the flow loss in the stator area were diminished.

Inlet boundary layer (IBL) condition can make a significant effect on the flow condition in the compressors. In order to explore the effect of IBL condition on the flow characteristics at the endwall region of the highly loaded compressor tandem cascade, a steady numerical method was used to calculate the flow field of a high subsonic tandem cascade, and comparatively analyze the effects of different IBL conditions, including thickness and form (normal and skewed), on the corner separation and tip leakage flow characteristics. The results showed that: without tip clearance, increasing the IBL thickness could increase the three-dimensional corner separation range from the front blade and the circumferential range of the rear blade. The skewed IBL weakened the secondary flow in the front blade passage but enhanced the secondary flow in the rear blade passage, the total pressure loss decreased and the effect of the skewed IBL on corner flow was weakened with the increase of IBL thickness. With 1% blade height clearance, increasing the IBL thickness could increase the scale of the leakage vortex in spanwise direction and the skewed IBL could enhance the leakage vortex of the front blade but weaken the leakage vortex of the rear blade, the total pressure loss decreased but the effect of the skewed IBL on tip flow was enhanced with the increase of IBL thickness.

In order to better ensure the comprehensive aerodynamic performance of turbine air-cooled blades in the structural strength design stage, based on the traditional four-line modeling method of two-dimensional cascade, two parameters, i.e. throat width and trailing edge bending angle, were introduced to ensure the aerodynamic performance. A modeling method based on free curve for irregular cooling air inlet was proposed to alleviate the stress concentration problem at the cold air inlet and its potential value was demonstrated by optimization calculation. On this basis, the typical cooling structure inside the turbine air-cooled blade was modeled parametrically. The mass of blade and the maximal tensile stress of blade body were taken as the optimization objectives, and parameters were selected to optimize turbine air-cooled blades. After optimization, the mass of blade decreased by 0.99% and maximum tensile stress decreased by 6.55%. The optimization results showed that the relevant parameterization method can meet the design requirements of turbine blades with complex internal cooling structure, and can effectively improve turbine design efficiency.

In order to better understand the impact of pulse frequency on the flow characteristics of the twin entry turbine rotor inlet, ANSYS CFX software was applied to numerical calculation of two-channel radial turbines with different frequencies (20, 40, 80 Hz) under the condition of pulse intake, while the distribution law of rotor inlet incidence angle in one pulse period, and the distribution law of blade height and circumferential direction were obtained. The results showed that in one pulse period, there were two peaks and two troughs in the incidence angle of the rotor inlet, the incidence angle near the crest decreased with the increase of frequency, and the incidence angle near the trough increased with the increase of frequency. The incidence angle distribution along the blade height took 0.5 blade height as the dividing point, and the incidence angle of the blade root part and the blade height part presented an opposite trend with the frequency. When the inlet pressure of blade root branch was larger, the inlet flow characteristics of twin entry turbines were more stable.

In order to study the relationship between the stall inception and the flow field characteristics on rotor tip casing-wall, the throttling characteristics of a low-speed isolated rotor axial compressor was experimentally investigated. Dynamic pressure signals were measured during the throttling process by dynamic pressure sensors on the casing-wall at the rotor tip circumferential and chord-wise locations. The flow field characteristics on rotor tip casing-wall before and after the stall inception were analyzed systematically. The results showed that at the near stall condition, the tip leakage flow and the tip secondary leakage flow were in the blade passage, the tip leakage flow was stable at the rotor tip leading-edge; the tip leakage flow in the stall disturbance region swung at the rotor tip leading-edge when the spike-type stall inception occurred, the formation of spike-type stall inception was related to the instability of tip leakage flow in the local area of rotor tip as an external appearance of the tip leakage flow instability; after stall of the compressor, the tip leakage flow in the stall disturbance region spilled from the rotor tip leading-edge, the stall disturbance evolved into stall cell which rotated in circumference. During the experiment, the arrangement method of dynamic pressure sensors can effectively capture the flow field characteristics on rotor tip casing-wall when stall inception occurred.

Ducted fan powered aircrafts are valued for their excellent efficiency, safety and noise characteristics in the low-speed field. How to reduce the thrust loss and intake distortion at the lip in crosswind conditions needs to be further explored. In this paper, an optimization method for reducing the adverse pressure gradient was used to optimize the lip geometry. The aerodynamic characteristics and flow laws of the original and optimized lips under crosswind conditions were investigated. In order to further improve the ability of the lip to resist crosswind, a double-lip structure was proposed based on the optimized geometry, and the flow control mechanism of the double-lip structure was analyzed. The results showed that the optimization method of reducing the adverse pressure gradient allowed the lip to vary within 3% of the dimensionless thrust and total pressure distortion coefficient over the advance ratio range of 0 to 0.07. The double-lip configuration allowed a further increase in the advance ratio reduced by the lip dimensionless thrust from 0.07 to 0.44. The total thrust reduction of the ducted fan can be reduced from 18.5% to 6.6% by using the double-mouth structure when the advance ratio increased from 0 to 0.55.

To explore the effects of endwall self-adaptive injection on corner separation and loss performance of high speed compressor cascade, the different outlet positions of endwall self-adaptive injection were studied by numerical simulation method. The results showed that the application of endwall self-adaptive injection can effectively control corner separation, reduce low-energy fluid accumulation, improve cascade pressure expansion ability and reduce total pressure loss of the cascade. The maximum relative reduction of the total pressure loss coefficient of 9.17% of the cascade at attack angle i=4° was obtained. The control effect of the total pressure loss coefficient decreased and then increased with the outlet position of endwall self-adaptive injection moving towards the trailing edge. The optimal control effect was achieved when the outlet position of endwall self-adaptive injection was located near the middle of the suction separation. However, the total pressure loss significantly increased when the outlet position was too far forward. And the effect of the inflow attack angle on the loss reduction effect was also more significant while the cascade at a positive attack angle was easier to control compared with the design attack angle when the outlet position was too far forward.

The research status of composite-steel hybrid gears was reviewed. The military requirements for the research of composite-steel hybrid gears and the important research significance for improving the power density of helicopter transmission system were briefly described. The research status and development trends of composite-steel hybrid gears were described in detail from five aspects: mechanical properties characterization of composite materials for gear spokes, structural strength analysis of hybrid gears, configuration design, manufacturing and experimental verification technology. The research status of composite-steel hybrid gear application technology was summarized. The shortcomings of the application technology of composite-steel hybrid gears in China were pointed out. Besides, the future research directions including material property calibration, strength analysis, optimization method, and manufacturing and testing technology that need to be carried out urgently regarding composite-steel hybrid gears were proposed.

At present the intelligent fault diagnosis model is difficult to train due to the difficulty in obtaining the fault data of rolling bearings. A self-supervised deep one-class classification method was proposed for early fault warning of rolling bearings based on the training of normal class samples. Based on the deep one-class classification model, multi-task and self-supervision mechanisms were introduced. Input signal was extracted with the use of the depth of the deep residual network characteristics, and the proposed features were taken respectively as input of more child tasks, including support vector description (SVDD) as the output results of the classification subtasks supervision and label as the rest of the subtasks; through the joint loss function, only relying on normal class samples can complete the model of supervised learning. When the proposed method was applied to the fault warning of rolling bearings, the vibration acceleration signals of the whole life cycle were decomposed by frequency band, and the signals in different frequency bands were encoded as the input of the network. The proposed method was validated on two actual rolling bearing fault data sets. The verification results showed that the accuracy of the proposed depth classification method reached more than 99% in fault warning, which fully indicated that the proposed method has a high ability of fault warning and anomaly detection.

In order to improve the speed of rocket sled test under existing conditions, a two-stage propulsion monorail rocket sled structure was designed, and the test trajectory was designed and calculated according to the sled structure and the use of engines. Decelerating parachute and foam water tank were used to decelerate and recover the first and second stage sleds respectively, and the deceleration effect was verified by tests; the engine ignition technology and the separation technology between the stages of the multi-stage monorail sled during high-speed movement were studied, and the effect of the time controller controlling the ignition of the secondary engine and the throwing of the parachute was verified; the sled moving speed during the test was obtained by using high-speed camera and net target. It was concluded that, with the smooth completion of the test task, the structure of the two-stage sled was proved reasonable and reliable; the time controller triggered the secondary engine and deceleration device according to the set time, indicating that it can be safely and reliably applied to the rocket sled test; the deceleration parachute was thrown out in advance under high overload, so the design of the deceleration device needs to be improved; the secondary sled was successfully braked and recovered on the track, and the foam water tank had an efficient and reliable deceleration effect on the monorail sled.

As a new type of electric propulsion technology, ionic wind propulsion technology can directly convert electrical energy into thrust. In order to enhance the ionic wind propulsion effect, the causes and factors of the thrust were studied, and the wire-plane and needle-plane ionic wind propulsion models were constructed. The internal conditions of ionic wind were studied by the finite element software COMSOL Multiphysics, and the changes of ionic wind speed and volume force density were analyzed because these were proportional to the thrust. The results showed that the volume force density of the wire-plane structure was larger than that of the needle-plane structure, but the wind speed was smaller. Increasing the input voltage or decreasing the corona electrode diameter could effectively increase the wind speed and volume force density, and then increase the thrust. There was an optimal electrode spacing and a collector width to maximize the thrust. The lower ambient pressure indicated the greater thrust. The thrusts of the wire-plane structure calculated by the two methods were 0.046 N/m and 0.042 N/m, respectively, which were close to the results of related experiments. The simulation study can provide a reference for the design of ionic wind propulsion structure.