For the anisotropic thermophoresis of a charged spheroidal colloidal particle in aqueous media, based on the boundary layer approximation and the extremely thin electric double layer assumption, the analytical formulations of thermophoretic velocity, thermodiffusion coefficient and thermophoretic force under the thermal conductivity effect were derived. The results showed that the thermal conductivity effect not only changed the values of thermodiffusion coefficient and thermophoretic force of anisotropic thermophoresis of spheroidal particles, but also affected greatly the directions of thermodiffusion coefficient and thermophoretic force. The values of thermodiffusion coefficient and thermophoretic force of polymer spheroidal particles were always higher than those of “normal” particles (i.e. particles with the same thermal conductivity as the fluid), while the thermodiffusion coefficient and thermophoretic force of metal oxide spheroidal particles were lower than those of “normal” particles. Compared with the “normal” particles, the anisotropic thermophoresis of polymer spheroidal particles was weakened, and that of metal oxide spheroidal particles was enhanced. The influence of thermal conductivity on the anisotropic thermophoresis of spheroidal colloidal particles was systematically studied, which can shed a light on the engineering design of microscale thermophoresis turbine for the micro-aerial vehicles.

The non-axisymmetric endwall can effectively reduce the circumferential pressure difference and control the complex secondary flow inside the passage. Consequently, this enhanced the aerodynamic performance of the airfoil. The non-axisymmetric endwall significantly impacted the outlet area of the film holes and the trajectory of the coolant. The concave-convex changes along the circumferential and axial directions on a flat plate were studied to explore the effects of different shaping positions on the film cooling characteristics. A row of discrete film holes at positions of 10%C_ax, 40%C_ax, and 70%C_ax inside the non-axisymmetric vane cascade were positioned. The numerical method of solving the time-averaged Reynolds equations was used to investigate the influence of the endwall shape on the cooling characteristics of the film holes. The results were as follows. Only when there were concave-convex changes along the circumferential direction on the flat plate, excluding the most concave and convex points, could other positions effectively reduce the blow-off of the coolant jet at a high blowing ratio. In the case of concave or convex changes along the axial direction, the exit area of the holes significantly increased behind the most concave point and in front of the most convex point. Compared with the flat plate, the maximum outlet area of the holes can increase by 51.4%, resulting in a substantial expansion of the high adiabatic effectiveness region. The non-axisymmetric endwall profiling based on passage pressure difference method can reduce the circumferential pressure gradient and the intensity of the horse-shoe vortex, improving aerodynamic performance. The endwall shaping in the film hole arrangement can improve the coolant attachment. The overall cooling performance was better under a higher blowing ratio. When M=1.0 and M=1.5, the area of high adiabatic effectiveness region with η>0.2 can be increased by 5.58% and 5.51%, respectively, compared with the flat endwall.

Liquid ammonia is used as a new nitrogen-based coolant. Considering the real operating temperature and pressure of the scramjet engine, a three-dimensional numerical model was established to analyze the flow and heat transfer characteristics of liquid ammonia. The thermal properties were calculated by PR equation and Chung method. The phase change of ammonia was calculated by Lee evaporation model. Pyrolysis of ammonia was also considered. The flow heat transfer characteristics of liquid ammonia under different temperatures and pressures were studied by taking into account the process of phase change and pyrolysis. The cooling feature of ammonia was compared with traditional hydrocarbon fuel. The heat sink of liquid ammonia and hydrocarbon fuel was compared and analyzed under the same condition. The results showed that the heat transfer capacity of liquid ammonia increased with the rise of pressure. When pressure increased from 3 MPa to 17 MPa, heat transfer coefficient increased by 8.02%. Under the same mass flow rate, heat transfer coefficient of ammonia was higher and the wall temperature decreased by 36.3% in non-pyrolysis zone and 9.1% in pyrolysis zone.

The swirl lobe mixer can promote efficient mixing of air and fuel-rich gas in the combustion chamber of the air-turbo-rocket (ATR) and improve the combustion efficiency of the engine combustion chamber. Based on combustion chamber inlet parameters under engine ground conditions, the effects of four lobe rotation angles (0°, 3°, 5°, 7°) on the vortex and temperature distribution at the tail edge of combustion chamber lobes were compared by numerical simulation, and the engine performance under different lobe rotation angles was analyzed from three aspects: combustion efficiency, pressure loss and hot mixing efficiency. The results showed that lobe rotation can enhance the flow vortex intensity and promote the thermal mixing efficiency. When the lobe rotation angle was 5°, the thermal mixing efficiency increased by 1.2%. The combustion efficiency was the highest when the lobe rotation angle was 5°, and the combustion efficiency at the outlet was 3.35% higher than that of the conventional lobe mixer. The total pressure loss of the combustion chamber was mainly caused by the combustion reaction, and the lobe rotation could enhance the combustion effect and increase the pressure loss to a certain extent.

The research focused on the flow heat transfer characteristics and efficient calculation methods in compact heat exchangers. To achieve the research objectives, a fast calculation method for the temperature distribution of a compact heat exchanger based on the porous media theory was established. The heat transfer models of the staggered tube bundle of the pre-cooler configuration for the reference ATREX engine and the involute spiral tube bundle pre-cooler configuration for the SABRE engine were analyzed for performance, respectively. The corresponding methodology and procedure for rapid calculation of heat transfer performance were established. The results showed that the prediction results were consistent with the FLUENT 3D tube bundle model simulation results but the time consumed was less than 1% of CFD simulation. The advantages of the fast prediction method in terms of calculation cost and speed were demonstrated.

The coupling simulation method of flow and transient heat transfer with asynchronous steps was studied. CFX was used to simulate the flow and temperature distribution during 20 h after engine shut-down by using a civil high-bypass-ratio turbofan engine model. Result showed asynchronous steps ac-celeration method provided great accuracy and simulation cost. After shutdown, the circumferential temperature difference increased first under the influence of natural convection, and then decreased gradually. The temperature rise of positions like rotor disc bottom and bearing was significant because of heat soak back. To a certain extent, blowing cold air and barring gear was efficient in reducing non-uniformity of circumferential temperature and peak temperature in shut-down process.

In order to obtain the influence of inlet temperature on the pulse detonation combustion characteristics, the combustion state, peak pressure, pressure wave and flame propagation velocity in detonation tube were studied under the conditions of detonation ignition frequency of 15—25 Hz and inlet temperature of 300—450 K by means of experimental and theoretical analysis. The results showed that: from the experimental and theoretical analysis results, the peak pressure was inversely proportional to the inlet temperature; when the inlet temperature was 300 K, the overall flame propagation velocity was low, and the fastest reached 2000 m/s at 350 K and 400 K, but the flame propagation velocity of 2000 m/s cannot be maintained at 400 K. Increasing the inlet temperature had a great influence on the cyclic initiation performance of pulse detonation combustion. Specifically, at the inlet temperature of 300 K, stable initiation can be achieved under 15, 20 Hz and 25 Hz. When the inlet temperature was increased to 350 K, stable initiation can be achieved under 15 Hz, a frame drop phenomenon was observed under 20 Hz, and continuous combustion was achieved under 25 Hz. On the whole, the increase of inlet temperature could reduce the peak pressure of detonation and significantly increase the probability of continuous combustion.

Under the trend of rapid development of compound coaxial helicopter, simulation analysis of dynamic electromagnetic scattering characteristics of compound coaxial helicopter was carried out systematically to explore its electromagnetic scattering law for stealth design and target recognition. A method of dynamic time-varying electromagnetic computing mesh that can characterize the coupled rotational motions of coaxial-rotor and tail rotor was constructed. The dynamic radar cross section (RCS) of the coaxial helicopter was calculated by the high frequency methods, and the contribution and influence mechanism of the coaxial-rotor and the tail rotor to the dynamic RCS of the whole helicopter were revealed, which layed a foundation for the sequence division of the RCS reduction of the dynamic and static components in the key threat azimuth of the helicopter. The change law and correlation of micro-Doppler of rotating parts were summarized, and an idea of identifying and judging the maneuvering intention of compound coaxial helicopter was proposed. The analyses showed that, in the existence of coaxial-rotor and tail rotor, the RCS of the whole aircraft exhibited dynamic periodic oscillation characteristics, and the oscillation amplitude was related to the static RCS of non-rotating components. The backward-swept blade tip could cause a micro-Doppler scintillation band that lagged behind the inner section of the blade in time, and the lag time depended on the blade rotation speed and the sweep angle of the blade tip. At any radar incident angle, the maximum micro-Doppler frequency shift of the coaxial-rotor and the tail rotor was represented by a set of determined values. By analyzing the micro-Doppler trend of these two, the flight attitude change of the helicopter can be accurately identified, and then its maneuvering intention can be inferred.

Circular graphite seal structures with spiral groove, triangle groove and V-groove rotor micro-texture were designed. Numerical solution models for leakage flow characteristics of circular graphite seal with rotor micro-texture were established based on two-phase flow theory. On the basis of verifying the accuracy of numerical methods, the flow field characteristics and leakage characteristics of circular graphite seal with three rotor micro-texture were analyzed under different inlet pressures and rotor speeds. The influence of rotor micro-texture on the dynamic pressure effect and cavitation phenomenon of the circular graphite seal was studied, and the influence mechanism of rotor micro-texture on the oil leakage flow characteristics of the circular graphite seal was revealed. The results showed that with the increase of inlet pressure, the peak pressure and leakage of the rotor micro-texture circumference graphite seal increased linearly, and the gas phase volume fraction decreased. The peak pressure of the spiral groove structure was the largest, which increased by 21.37% compared with the V-groove structure. The rotor speed had great influence on the pressure peaks of the three micro-textures, but little influence on the gas phase volume fraction, with the maximum increase of 12.63%. Spiral groove structure had the best sealing performance and dynamic pressure effect, which can effectively reduce the friction and wear of graphite ring. The leakage flow characteristics of circular graphite seals can be analyzed more accurately by using the two-phase flow theory method because of the dynamic pressure effect and cavitation in the seal gap.

There are higher requirements for the seal performance and service life of the sealing structure between counter-rotating shafts of the aero-engine. The structure of carbon fiber rotating brush seal between rotating shafts was proposed. Based on the theory of bristles centrifugal effect mechanics and porous media theory, the method of bristles shape variable correction porosity and resistance coefficient was used to establish a solution model for the leakage flow characteristics of carbon fiber rotating brush seal between rotating shafts. The influence of bristle lay angle, bristle diameter and rotating speed on the mechanical deformation characteristics of carbon fiber bristles was analyzed, and the leakage flow characteristics of brush seal between rotating shafts were studied. The results showed that the deformation and rotation angle of the bristle could increase by increasing the speed of the inner rotor, reducing the diameter of the bristle and increasing the bristle lay angle. Considering the centrifugal effect of the bristles, when the inner rotor speed was 3230 r/min, the deformation of the T650 carbon fiber bristles in the radial direction of the shaft was 16.43% of the deformation of the Haynes25 bristles. When the inner rotor speed was 10000 r/min, the bristle lay angle increased from 30° to 45°, and the radial deformation of the bristles on the shaft was increased by 104.02%. Under the same working conditions, the fine-diameter bristles sealing effect was better. When the contact gap between the bristles and the outer rotor was 0 mm, the bristle diameter was reduced from 0.07 mm to 0.05 mm, and the leakage was reduced by 12.42%.

Experiments were conducted to investigate the primary breakup characteristics of liquid jets injected into a crossflow with linear distribution of incoming velocity. Jet breakup modes, column breakup point, jet surface wave and surface velocity were extracted and analyzed by high-speed camera combined with backlighting method, and the deformation and penetration laws of the jet were described. Experiments were conducted on five structures at atmospheric pressure and 320 K environment with selection of average momentum ratio 20—80 and gas Weber number 5.6—40. The results showed that the positive gradient delayed the breakup modes and the negative gradient advanced it; when the incoming flow was non-uniform, the jet deformation, penetration, and surface wave became complicated, which also delayed or advanced the location of the column breakup point. Phenomenological analysis can effectively explain and correlate the measurements of the primary breakup characteristics of liquid jets with non-uniform incoming flow, and the predictive expressions for jet deformation and column breakup height adapted to the conditions of this experiment were presented.

For multi-stage brush seal, stage number has a great influence on the characteristics of leakage flow and heat transfer. The theoretical formula, which explains how the stage number affects the characteristics of leakage flow and heat transfer in multi-stage brush seal, was derived. The solution model of multi-stage brush seal was established and the numerical simulation was carried out based on the fluid-solid-thermal coupling calculation method. The leakage and pressure distribution measurement experiment of multi-stage brush seal was designed. The experimental device of leakage and pressure distribution characteristics of multi-stage brush seal was built. The experimental results were compared with the numerical results to verify the accuracy of the numerical simulation. The distribution characteristics of flow field and temperature field between different series brush seals were analyzed and compared. The influences of series on the leakage characteristics, flow characteristics and heat transfer characteristics of brush seal were studied. The results showed that under the same pressure ratio, from one to three stages brush seals, for each additional level, the pressure drop value borne by each stage of brush tows decreased, and the maximum pressure drop of brush tows decreased by 30.0%—36.3%. The maximum velocity of the flow field appeared below the rear baffle of the last stage brush tow, and the maximum velocity of the flow field decreased with the increase of stage number. Under the same pressure ratio, with the increase of stage number, the sealing performance of the brush seal was enhanced. For each additional stage, the leakage was reduced by 15.7%—22.0%. Under the same pressure ratio, the maximum temperature increased with the increase of stage number, and the maximum temperature increased by 3.7%—9.1% for each additional stage.

To study the flow demand of nitrogen-enriched air under the effect of dynamic change of gas phase space volume of multi-bay fuel tanks, a dynamic distribution inerting model of multi-bay fuel tank flow was established according to the equations of conservation of components, conservation of mass, and gas state equations, and a dynamic distribution method based on the change of gas phase space volume of compartments was proposed for the flow of nitrogen-enriched air; a modeling simulation was carried out for the multi-bay fuel tank of a certain type of aircraft as an example, and combined with the AMEsim software, the modeling simulation was carried out to compare the inerting effects of dynamic distribution, average distribution, and proportional distribution. The results showed that: The proposed dynamic distribution method better realizes the oxygen volume fraction control of multi-bay tanks, the nitrogen-enriched air flow required the shortest inerting time according to the dynamic distribution method, and the dynamic distribution method had a higher degree of gas-phase spatial inerting compared with the average and proportional distribution methods; in terms of volumetric tank exchange, the dynamic distribution method can effectively reduce the number of volume exchanges from 3.5% to 7.6%, and the volumetric tank exchange of the 3 allocation methods are, in descending order: average distribution, proportional distribution, and dynamic distribution.

A coupled computational model of a transonic axial compressor and a rotating detonation combustor Stage 37 was constructed to investigate the impact of forward-propagating pressure waves generated by rotating backpressure on the internal flow field of the compressor. Furthermore, the effect of backpressure velocity on the compressor's performance was analyzed. It was observed that the forward-propagating wave interacted with both the stator and rotor blades, resulting in a slight upstream shift in the attached shockwave at the leading edge of the rotor blades due to the influence of the rotating backpressure. Moreover, the flow field structure of the stator was more significantly affected by rotational backpressure, and correspondingly, the forward pressure wave was relatively more strongly suppressed by the stator. In comparison with the peak efficiency point, the compressor’s performance was found that CJ velocity backpressure (CJ was the detonation propagation velocity in the Chapman-Jouguet theory) had a positive effect on the compressor’s efficiency, with an increase of 1% and wider operating range.

In order to efficiently design and analyze the turbine blade cooling structures, a rapid thermal analysis method for film cooling coupled with thermal barrier coatings (TBC) was established and basically verified. Based on the morphologies of the conductive heat flow streamlines in the metal region, the heat transfer processes were separated into three parts of mutual adiabatic, so as to reduce the difficulties of equation construction. Topologically transforming methods were applied, thus the complicated 3D heat transfer processes were translated into 1D ones like flat plates or fan-shaped segments with uniform wall thickness. A set of 40 direct calculation formulas were derived accordingly, which achieves a rapid solution of the weighted average temperature at the objective walls. Many effects on the wall temperature could be analyzed by this method, such as the flow and heat transfer conditions, the geometries of both film cooling hole and TBC, and also the material properties of them. The accuracy of the method was investigated by comparing its results with the fiducial ones obtained from 3D simulations. The relative error (δ) of these wall temperatures were within ±1.25% considering various TBC coverage schemes and blowing ratios (M) operating conditions, which met the requirements of engineering applications.

To deeply understand the mechanisms of flow and heat transfer of hydrocarbon fuel for the active cooling applications, numerical investigation of convection heat transfer characteristics and entropy generation behaviors of supercritical pressure RP-3 in a vertical upward tube with inner diameter of 2 mm was carried out. The radial distributions of local entropy generation at different cross-sections along the flow direction under both normal and deteriorated heat transfer modes were obtained. The effects of wall heat flux on Nusselt number and bulk entropy generation were studied. The results indicated that the bulk entropy generation and Nusselt number had opposite trends. The increase of Nusselt number corresponded to lower irreversibility and decreased entropy generation. Along the flow direction, the dissipation entropy generation was gradually replaced by heat transfer entropy generation as the dominant mechanism in the inlet section. Along the radial direction, the laminar entropy generation played a dominant role in the viscous sublayer, and the turbulent entropy generation became the dominant mechanism in the buffer layer. Under the deteriorated heat transfer regime, the bulk entropy generation reached a minimum as the heat transfer coefficient peaked. In the section where heat transfer deterioration occurred, the ratio of heat transfer entropy generation to total entropy generation exceeded 0.99.

Temperature and salinity are critical factors influencing ice adhesion strength. Utilizing a programmable high-low temperature test chamber, static freshwater ice and saline ice were prepared under varying temperature and salinity conditions, with tangential ice adhesion forces measured to investigate the temperature-dependent trends and salinity-influenced mechanisms. The results demonstrated that the quasi-liquid layer (QLL) at the ice-substrate interface played a decisive role in the initial ice adhesion process for both ice types. For freshwater ice, the tangential adhesion strength increased linearly as temperatures decreased from 0 ℃ to −11 ℃. Beyond this critical temperature threshold ( −11 ℃), the growth rate gradually decelerated with further cooling, stabilizing substantially below −20 ℃. Saline ice required lower temperatures to form effective interfacial adhesion, exhibiting weaker tangential adhesion compared with freshwater ice. Within the salinity range of 1%—2.5%, ice adhesion strength was predominantly governed by the interfacial ice-salt distribution. This study could provide theoretical and technical references for elucidating ice formation mechanisms and developing anti-/de-icing strategies under diverse environmental conditions, particularly in marine cryogenic environments.

Based on a 2.5D woven Ceramic Matrix Composite (CMC) plate, the microscopic geometric features were identified, and the probability distribution functions were calculated. A parameterized full-scale model reflecting the dispersion of the material’s microscopic structure was established. The Monte Carlo simulation method was employed to investigate the influences of the geometric parameters’ randomness of the woven CMC material’s microscopic structure on the probability distribution of its anisotropic thermal conductivity. The results were validated through orthogonal experiments. The findings revealed that, due to the thermal conductivity difference between the matrix and fiber bundles, the temperature and heat flux distributions exhibited non-uniformity. There existed a noticeable temperature gradient at the interface between the matrix and fiber bundles. The heat flux density was higher in the matrix and oblique fiber bundles, while it was lower in the Y-direction (weft) and X-direction (warp) fiber sections. Under the influence of geometric parameter variations, changes in the volume fraction of the matrix and fiber bundles were primary factors affecting the fluctuation of the effective thermal conductivity. The impacts of the six geometric features on the probability distribution of thermal conductivity were ranked as follows: weaving angle, warp length, warp spacing, warp width, weft width and weft length.

In order to obtain the unstable growth characteristics and influence laws of surface waves during the process of swirling liquid sheet fragmentation, experimental studies on different geometric structures of closed swirl atomizers were conducted. Based on high-speed images, the growth rate of surface wave amplitude during the development of liquid sheet was extracted, and the influence of pressure drop was summarized. The research results indicated that the amplitude of the dominant liquid sheet surface wave increased exponentially along the flow direction of the liquid sheet surface, and the growth rate of the surface wave amplitude value exhibited nonlinear characteristics; the development of liquid sheet was divided into fast growth zone, slow growth zone, and fragmentation zone. A function was established for the variation of fluctuation amplitude with travel distance in different regions, and the dominant fluctuation growth rate corresponding to different development regions was extracted, reflecting the nonlinear and unstable growth laws of the liquid sheet. At the same time, the position of the turning point of the function was analyzed, and a functional relationship between the breakup length and pressure drop was established. It was found that the breakup length showed an exponential decline trend with the increase of pressure drop, and the ultimate breakup length of the atomizer liquid sheet can be predicted through the functional relationship. The ultimate breakup length increased with the increase of the atomizer outlet diameter.

In order to investigate the effect of fuel spray spatial distribution on ignition characteristics, large eddy simulation method combined with discrete phase model and dynamic thickening flame model was used to simulate the flow and combustion process in the axial-radial swirler combustion chamber. The results showed that the local droplet concentration and distribution range of the high concentration in the ignition area determined the formation of the initial flame kernel and the propagation process of the flame. Especially in the early stage of flame development, the hollow conical fuel spray flame mainly propagated in circumferential and axial directions, while the solid conical fuel spray flame propagated in circumferential, axial and radial directions at the same time, and the fuel spray spatial distribution of different injection modes corresponded to the flame morphology after stability. The large eddy simulation method was used for the first time in reproducing the phenomenon that the fuel spray distribution determined the flame distribution observed in other scholars’ experiments, proving that Sauter mean diameter (SMD) alone cannot represent all the effects of atomization characteristics on the fuel spray combustion process.

In response to the problems of large parameters, high network complexity, and low portability of aero-engine surface defect detection algorithms, a lightweight defect detection algorithm WGS-YOLO combined with attention mechanism based on YOLOv5s (you only look once version 5 small) was proposed. The algorithm utilized the ShuffleNet V2 network to construct the unit-reconstructed backbone network, reducing network complexity. The presented AW-CBAM attention mechanism (CBAM attention mechanism with adaptive adjustment of weights) was capable of simultaneously extracting channel attention features and spatial attention features, and dynamically adjusting the weight proportion of the two features through adaptive weight coefficients, which enabled the network to pay more attention to the global information of the input feature map, and apply it into the backbone network to enhance the network’s representation and generalization capabilities. A lightweight GS-C3 module was designed in the neck network by introducing depthwise separable convolution and Ghost convolution to achieve feature condensation of the input feature map, efficiently capturing important information in input features. The experimental results showed that the recognition precision of the WGS-YOLO algorithm reached 92.0%, 2.1% higher than the baseline. Meanwhile, the network parameters were reduced by 55.3%, and the computational load decreased by 57%. Therefore, the proposed algorithm can meet the design requirements of lightweight network and effectively detect major surface defects in aero-engines.

Taking a compressor blade made of GH2787 as the research object, an approach of coupling laser shock peening and shot peening was proposed for taking advantages of individual laser shock peening and shot peening. This led to a high level of residual stress and sensitivity inhibition of defects on component surface, so as to comprehensively improve the surface integrity of the blade. The surface integrity and fatigue strength of the strengthened blades were tested, and the coupled laser-shot peening exhibited the most significant effect on compressor blade fatigue strength improvement. The fatigue strength of the coupled laser-shot peened compressor blades was improved by 20.6%, which was further increased to 28.1% after following polishing treatment. The increase of fatigue strength was significantly higher than the individual laser shock peening and shot peening, i.e., 15.9% and 18.3%, respectively. This work provides theoretical basis and data support for the implementation of compressor blade surface strengthening technology and anti-fatigue design.

In view of the fact that the performance characteristics of rolling bearings vibration signals are inevitably affected by environmental noise and lack of Markov characteristics, a life prediction method of rolling bearings in noisy environment based on IKF-ARIMA-NARNN was proposed. Firstly, considering that the Kalman filter （KF） denoising ignored the correlation of data, a denoising method based on incremental Kalman filter （IKF） was proposed. Secondly, due to the evolution process of bearing performance characteristic parameters that fails to meet the Markov characteristics sometimes, an analysis model of the evolution process of bearing performance characteristic parameters based on the autoregressive integrated moving average （ARIMA） with nonlinear error term was established. At the same time, the dynamic nonlinear autoregressive neural network （NARNN） was used to estimate the nonlinear random error term of the degradation model, so as to realize the life prediction of rolling bearings. Finally, the effectiveness and applicability of the proposed method was verified by the case analysis of rolling bearing engineering, and the prediction accuracy was at least 26.75% and 51.25% higher than the traditional method.

To achieve the automation and intelligence of the radiographic inspection of turbine blades in aeroengines, and to effectively address the time-consuming, labor-intensive, and inefficient problems in traditional radiographic inspection methods, a research initiative was undertaken to develop a defect detection method for X-ray images of turbine blade based on unsupervised learning. A defect inspection algorithm suitable for X-ray images of aeroengine turbine blades was proposed based on an unsupervised generative adversarial network. It consisted of a generator network, a discriminator network, and an extra encoder network. Reconstruction, discrimination, encoding, and intermediate encoding loss were designed, and the weighted sum of the four losses was used to construct the objective function. Using non-defective X-ray images for model training. A defect inspection model for X-ray images of aeroengine turbine blades was established based on the trained generator network. The effects of input image size, encoding size, and type of reconstruction loss on the performance of the defect detection model were studied. Results showed that the proposed model with an input image size of 128 pixel×128 pixel, 600 encoding size, and L₂ reconstruction loss can achieve an area under curve (AUC) of 0.911. The defect inspection algorithm can realize strict technical indicators of zero missing rate for actual production, but the false detection rate (>62.1%) was relatively high. As an auxiliary detection method applied in actual production, it can improve the manual detection efficiency by 1.6 times.

Considering the aeroelastic instability problem of the labyrinth seal of aeroengine, a model for solving the damped vibration reduction of labyrinth seal was established. Based on the energy method, the influences of the structure parameters of the damping ring and its installation position on the aeroelastic stability of the labyrinth seal were studied, and the mechanism of the influence of the damping ring on the aeroelastic stability of the labyrinth seal was revealed. The results showed that the aerodynamic damping ratio of the first mode and second mode was −0.41% and −0.049%, respectively, without the damper ring, which resulted in aeroelastic instability. However after installation of the damping ring, the aerodynamic damping ratio of the second mode changed from negative to positive, and aerodynamic damping ratio of the first mode was still negative, but the damping ratio value increased, the damping ring improved the aeroelastic stability of the labyrinth seal. When the installation interference of the damping ring was larger, the opening amount was smaller, moreover, when the aerodynamic damping ratio of the labyrinth seal was larger, the aeroelastic stability was better. When the damping ring was installed at the cantilever end, the aerodynamic damping ratio was the largest, meantime the labyrinth seal was the most stable. When the damping ring was installed at the wave node, the aerodynamic damping ratio and the amplitude reduction ratio were the smallest, and then the labyrinth seal stability was poor.

In order to optimize the unbalance response feature of a mid turbine frame dual rotor system, and can make it operate stably within a large speed range under the same unbalance mass, a simplified particles system was established based on the mid turbine frame dual rotor system. It was proved that there was vibration energy transmission between the particles in the simplified particles system based on the dynamic vibration absorber theory deduction. The stiffness combination and damping conditions with the best vibration energy transmission efficiency were calculated. According to the ideas above, the support stiffness of a dual-rotor system test rig with mid turbine frame was optimized. The finite element method was used to analyze the rotor dynamics feature and its unbalance response. The result showed that the rotor system also had a similar vibration energy transmission phenomenon. After the optimization, the strain energy proportion of the power turbine shaft in the key mode decreased by more than 31.9%. Compared with the unoptimized rotor system, the unbalance response of power turbine shaft decreased by more than 73.0% and the response of the power turbine disk dedcreased by more than 55.3%, indicating that using the vibration energy transmission method to optimize a mid turbine frame dual rotor system can reduce the unbalanced response of the system effectively.

In order to understand the vibration characteristics of the blade-disc system and to check whether the damping effect meets the technical specifications, it is necessary to design and carry out dry friction damping blade vibration characteristics experiment. Real blade vibration characteristics experiments encounter more challenges in the design process than simulated blade experiments. A certain type of turbine blade with underplatform damper was taken as an example to design an experimental programme for the vibration damping characteristics of real turbine blades. The frequency response function curves of excited blades under different positive pressures were obtained by carrying out vibration damping characteristic experiments of turbine blades. The data showed that underplatform damping device within the design frequency range of the blade group vibration had a significant inhibition effect, the best vibration damping effect can reduce the excited blade co-vibration amplitude up to 60%.

Based on the macroscopic phenomenological creep model, the finite element subroutine was compiled to simulate and analyze the creep deformation behavior of real turbine blades. In the engineering practice, through simulation of creep deformation in service environment, the important assessment positions of turbine blades were determined. Considering the effect of creep, the stress relaxation effects on the concentration positions were analyzed, and the prediction endurance life of the verifying positions with/without stress relaxation was compared. By changing the deviation angle between the crystal axis and the blade height direction, the allowable deviation angle in the engineering application was simulated. These results indicated that the important assessment positions of turbine blades were generally located at the root of the spoiler columns and the corner of the cooling channels. The stress relaxation phenomenon occurred at these positions with the creep progresses, affecting the prediction endurance life of turbine blades. The tolerance of deviation angle in engineering was designed as 10°, beyond which the creep performance of turbine blade material decreased significantly.

An experimental investigation of the influence of seal cavity leakage on compressor was carried out on two compressors with variable inlet throttle ratio and cavity pressure respectively. The investigation results showed that seal cavity leakage flow rate increased with the decrease of inlet throttle ratio or the increase of cavity pressure. As the seal cavity leakage flow rate increased, the outlet pressure of main flow decreased and the blade wake area expanded near compressor hub, reducing the compressor performance. In a compressor experiment with multiple seal cavities, compared with the front seal cavity, the influence of leakage flow rate at rear seal cavity on compressor efficiency was greater, and the higher rotating speed indicated the greater influence.

In order to study the influence of structural parameters on the hydraulic and acoustic performance of a pump-jet propelled submarine in cruising condition, the DES (detached eddy simulation) was used to calculate the unsteady flow in the whole basin. The pressure fluctuation or velocity fluctuation of the flow field was output as the sound source, and the acoustic FEM (finite element method) was used to predict the flow noise. The experimental study was carried out on the circulating water channel of Lanzhou University of Technology. The noise and thrust information were collected by hydrophone and thrust data acquisition system, and the numerical results were verified. The results showed that fewer impeller and guide vane number and less area ratio of nozzle import and export can improve the internal flow condition of pump-jet. The total thrust of pump-jet propulsor decreased with the increase of the number of impeller blades or the number of guide vane blades, but increased with the increase of the area ratio of nozzle import and export. The open water efficiency decreased with the increase of three structural parameters. The increase of impeller blades number caused all kinds of noise to show an overall upward trend, while the increase of different kinds of guide vane number and area ratio of nozzle import and export showed different influence laws on different kinds of noise. When the structural parameters changed, the directivity distribution of various flow noises was basically the same.

To address the performance degradation caused by eroded leading edge of aero-engine supersonic fan blade, a Kriging model between the curvature control point of fan blade leading edge and three optimization objectives of total pressure loss coefficient, total pressure recovery coefficient and material removal rate was constructed, and the second Non-dominated Sorting Genetic Algorithm (NSGA-Ⅱ) was used to optimize the eroded leading edge. The results showed that the material removal rate was negatively correlated with the total pressure loss coefficient, and positively correlated with the total pressure recovery coefficient. Compared with the eroded cascade, the total pressure loss coefficient of the blade with trimmed leading edge decreased by 8.74%, the total pressure recovery coefficient increased by 2.31%, and the material removal rate was 1.94%. Optimized cascade was able to improve the flow around the leading edge, and reduce the intensity of the lip shock and the area of the subsonic region. The bow shock of the eroded cascade intersected the passage normal shock on the suction side, which intensified the interaction with the boundary layer. The optimized cascade made the passage shock move forward, and the shock waves met in advance within the passage, thus restraining the development of the wake.

It is difficult to conduct the full-scale testing for J-class heavy-duty gas turbine transonic compressors (with an inlet design mass-flow rate of roughly 1000 kg/s), the diameter of the first stage compressor was nearly 3 m, the turbomachinery designers scaled-down the geometry to meet thepower limitation, which is a convenient method based on similarity principle. This work performed comparative investigations of scaling effects on the aerodynamic performance and loss mechanisms based on the numerical simulation. The results showed that the larger model operated at a higher isentropic efficiency condition. The Casey’s prediction models had a deviation of less than 0.5%. The shock/tip leakage vortex loss and shock-boundary layer interaction loss of the larger model were stronger than those in the prototype model, while the profile boundary loss and wake mixing loss of the prototype model were stronger than the larger model. The profile losses (boundary layer and wake mixing loss) were the main loss source causing the difference in the isentropic efficiency between the prototype and scaled-up compressor rotor, meanwhile, the difference caused by tip secondary flow loss was not the main source. This investigation is intended to provide suggestions for industrial preliminary design stages of the heavy-duty gas turbine compressors.

In accordance with the velocity and pressure distribution features of the offset paired vortex distortion at the exit of a typical large S-bend inlet, a method for defining the offset paired vortex boundary conditions was established by using a vortex model. The coupling effect mechanism between offset paired vortex distortion and fan was explored by carrying out numerical simulation research. From the research results, it indicated that the upstream distorted flow mixing was strengthened, and the tangential velocity increased under the suction effect of the fan. The condition of the offset paired vortex inlet caused the deterioration of fan performance, and a serious loss of stability margin. Among them, the loss of stability margin was 7.1% under the condition of mutual exclusion vortex with 35 m²/s vortex circulation, and 7.9% under the condition of mutual attraction vortex. As the strength of offset paired vortex increased, the fan performance deteriorated more, and the loss of stability margin was also greater. Under the condition of the offset paired vortex embodied in the large area of attack angle was caused by the axial speed loss and tangential pre-rotation of the fan blade tip. This caused a relatively high total pressure ratio in some blade channels of the fan, while the remaining positions of the fan were still at a relatively low total pressure ratio level, resulting in a decrease in the fan overall total pressure ratio near the stall point of the fan and an advance in the stall point.

To investigate the impact of the variation in the widest position of the tip winglets on the leakage flow at the blade tip of the compressor, the widest position of the pressure surface tip winglets in the circumferential direction with an addition of 1.5 times the blade width was studied. The widest positions of the tip winglets were located at 10% to 90% of the axial chord length of the blade, with intervals of 20% of the axial chord length. Subsequently, the effects of different widest position tip winglets schemes on the clearance flow field at the compressor blade tip were analyzed. The results showed that: the change of the widest position of the tip winglet on the axial chord length had a direct influence on the flow field regulation effect. PW0.1c scheme with the widest position of the winglet at 10% axial chord length increased the total pressure loss by 2.39%, while PW0.7c scheme with the widest position at 70% axial chord length had the most obvious improvement effect on the flow field, reducing the total pressure loss by 1.56%.

The effects of main design parameters of the ejector nozzle, including the length of the main nozzle, the throat area of the ejector nozzle, the position of the throat and the area of the expansion section, on the internal flow-field and thrust performance in the design state were studied through numerical simulations. The results showed that, in order to improve the thrust coefficient of the ejector nozzle, it is advantageous to reduce the length of the main nozzle, appropriately increase the throat area of the ejector nozzle, move the position of the ejector nozzle throat backward and increase the cross-sectional area of the expansion section in the ejector. The influence of various design parameters on the flow mechanism was mainly embodied in the degree of the mainstream expansion and the characteristics of the shear layer. It was verified that the ejector nozzle could achieve good aerodynamic performance within the full operating range (Mach number of 0 to 4). In the design state, the thrust coefficient could reach 0.96.

In order to investigate the influence of angle of attack on boundary layer transition and aerodynamic heating distribution, a novel wing-body configuration featuring a large swept wing was proposed, and the improved k-ω-γ transition model was used to calculate the boundary layer transition under different angles of attack at Ma=6. The research results indicated that variations in the angle of attack significantly affected boundary layer transition and aerodynamic heating of the wing-body configuration, and the boundary layer transition on the surface initiated from the wing-body junction and gradually expanded downstream along the connection, exhibiting a widening trend on both sides. Notably, as the angle of attack increased from −10° to 10°, the transition range on the upper surface of the fuselage and wing initially decreased before increase. On the lower surface, the transition range of the fuselage increased first and then decreased, while on the wing, it decreased first and then increased. As for aerodynamic heating, on the windward side of the wing-body configuration, the heat flux distribution was primarily influenced by the boundary layer transition due to the relatively weak flow structure. On the leeward side, where the flow structure was stronger, the heat flux distribution was affected by both flow structure and boundary layer transition. Additionally, streamwise hot streak structures were identified on the upper surface of the wing at 6° and 10° angles of attack. The comparison between the flow structure and heat flux distribution revealed that the formation of hot streaks was related to the evolution of streamwise vortices and corner stream-wise vortices.

During taking off of carrier-based aircraft, the engine exhaust impinged on the jet blast deflector generates high-intensity noise, which endangers the hearing health of deck staff and the safety of equipment structure. Under the circumstance of subsonic cold jet impinging on a inclined plate, by taking advantage of chevron nozzles and micro structured inclined plates, the suppression effects of two noise suppression strategies on impinging jet noise were studied. The effects of the chevron nozzles on impinging jet noise characteristics, such as chevron number, chevron penetration and chevron length, were analyzed carefully. The effects on the directivity and intensity of impinging jet noise of the micro structured impinged plates, such as triangular streamwise grooves, pyramid bulges, embedded circular holes, and triangular crossing grooves, were explored. Investigation was carried out to assess the noise-suppressing superposition effects of multiple suppression strategies, which consisted of chevron nozzles and micro structured inclined plates. The experimental results showed that chevron nozzles efficiently suppressed impinging jet noise, and increasing chevron penetration can further improve the ability of noise reduction. Micro structured impinged plates can reduce noise in the low frequencies, and the triangular crossing groove plate had the best suppression effect, achieving a maximum reduction in the overall sound pressure levels of up to 3.68 dB at the downstream. The combined strategies of noise reduction significantly suppressed downstream noise levels, manifesting a distinct nonlinear superposition mechanism. The study conclusions had engineering guiding significance for suppressing impinging jet noise of carrier-based aircraft.

The mechanism of slotted stator blade controlling boundary layer separations was studied, and the transonic single-stage axial flow compressor NASA Stage 35 was selected as the research object. Based on three front slot outlet positions and two rear slot outlet positions, six double-slot schemes were designed. The numerical calculation results indicated that under small and medium mass flow rate conditions, all the schemes improved the compressor total performance. When the outlets of front slot and rear slot were located at 13%C_a (C_a was the blade tip axial chord length) and 60%C_a, the compressor total performance was the best. And without significantly reducing the design point compressor efficiency, the scheme exhibited the highest compressor efficiency improvement of 1.3% under the near stall condition. The internal flow field analysis showed that at 99% blade span, higher jet momentum and the jet direction closer to the separation vortex core better eliminated the low-energy fluids generated by the boundary layer separation. However, under the interaction of radial pressure gradient and reverse flow in the separation zone, the jet with lower velocity near the slot undersurface migrated to lower blade span and the leading edge of the blade, resulting in an increase in the thickness of the boundary layer at the lower blade span.

In order to reduce the sonic boom of a supersonic passenger aircraft inlet, a rectangular low-boom supersonic inlet was designed with features of internal 0° cowl and relaxed isentropic compression, boundary layer bleed was applied to enhance the aerodynamic performance of the inlet. Through numerical simulation methods, the influences of critical inlet design parameters on aerodynamic performance and flow characteristics were studied, and the effect of inlet on boom characteristics of aircraft was also investigated. The results showed that, for a supersonic passenger cruising at Mach number 1.8, the optimal performance of the low-boom inlet counted on the bleed control, which was determined by factors such as total compression angle and bleed slot structure; at design point, the total pressure recovery coefficient of low-boom inlet achieved 0.956, the boom signature was reduced by 74% compared with conventional rectangular external compression inlet; the conventional inlet increased the boom signature of the aircraft by 136%, while the boom signature of the aircraft with low-boom inlet decreased by 4.5% on the basis of conventional one.

Considering the influence of high temperature real gas effect on boundary layer transition, the γ-Re_θ transition model modified by high Mach number was used for research. The boundary layer transition of typical cases was calculated and the results were compared with DNS （direct numerical simulation） and tests. The results showed that: for high speed and high enthalpy boundary transition, the original γ-Re_θ transition model predicted the transition location prematurely, while the transition location predicted by the transition model with high Mach number correction was more consistent with the experimental results. The high temperature real gas effect reduced the temperature and the thickness of the boundary layer, and promoted the boundary layer transition.

By using the user-defined function (UDF) programming interface of computational fluid dynamics (CFD) software FLUENT, the ignition gas and propellant gas dosing models of solid rocket motor were established. The transient process of motor ignition with different ignition cartridge locations was simulated. The results showed that: (1) The ignition location directly affected the injection process of ignition gas, thus affecting the transient temperature and pressure changes in different locations of the combustion chamber. (2) Under middle ignition condition, the ignition gas propagated to both sides of the inner passage of the propellant at the same time, resulting in the shorter time for the complete ignition of the inner surface on the propellant. (3) The ignition pressure peaks of combustion chamber under head ignition, middle ignition and tail ignition were 10.741, 9.862 MPa and 10.023 MPa, respectively. The ignition pressure peaks under head ignition were 8.9% and 7.16% higher than those under middle ignition and tail ignition, respectively. (4) Under the condition of tail ignition, the heat transfer of the ignition gas to the propellant decreased, the ignition lag period and the flame propagation period increased, and the ignition delay was as high as 12.75 ms.

The characteristics of the inductively coupled plasma (ICP) in the discharge chamber of a radio frequency ion thruster were studied as crucial factors affecting the thruster’s performance. A two-dimensional axisymmetric fluid model was developed for the radio frequency ion thruster discharge chamber. The simulation utilized a fifth-order weighted essentially non-oscillatory (WENO) scheme and the finite difference time domain (FDTD) method to solve the flow and electromagnetic field equations, respectively. Plasma characteristics inside the self-designed 3 cm diameter thruster discharge chamber were calculated. The simulation results indicated that the coupling of axial and radial induced magnetic fields, tangential induced electric fields, and tangential induced electric currents contributed mainly to maintaining ICP in the discharge chamber. In the heating mechanism of sustaining ICP, both random heating and Ohmic heating could play significant roles, and with an increase in radio frequency power, the proportion of random heating power to total heating power also increased. By increasing the radio frequency power, the beam current and propellant utilization efficiency of the radio frequency ion thruster rose, but the electrical efficiency decreased. Increasing the propellant flow rate led to an increase in thruster beam current but a gradual decrease in propellant utilization efficiency.

Considering the problem in calculating the correlation dimension using the hard threshold function that may lead to boundary disturbances, a fuzzy correlation dimension method was proposed and then used to extract the degradation features of rolling bearing. In view of the small sample prediction characteristics for the degradation stage of rolling bearing and the lack of confidence interval estimation for prediction results in traditional particle filtering, the maximum entropy particle filtering method was proposed. The experimental results showed that the fuzzy correlation dimension was consistent with the evolution trend of rolling bearing vibration signals, and can accurately extract the rolling bearing vibration performance degradation feature. Meanwhile, the average relative error of the proposed maximum entropy particle filter was only 1.31%, and the average absolute error was only 6.94, which was significantly lower than traditional particle filters and GM models. The added interval prediction function made the prediction results more diverse.

Considering the problem of unbalanced fault data in the helicopter tail drive system, a fault diagnosis method of helicopter tail drive system based on digital twin and transfer learning was proposed. A rigid-flexible coupling dynamics simulation model of the helicopter tail drive system was established to obtain high-fidelity fault simulation data truly reflecting the operating state of the helicopter tail drive system. A residual network introducing coordinate separable convolution and attention mechanism was used for fault feature extraction and classification. The domain adaptive method based on Gaussian kernel function was used to reduce the distribution difference between simulation data and experimental data in the feature space. In order to improve the robustness of the decision boundary and enhance the differentiation between categories, the cross-entry loss with margin regularization was introduced. It was experimentally verified that the fault diagnosis method based on digital twin and transfer learning can address the issue of deteriorated training effect in deep learning fault diagnosis model caused by unbalanced data. This method significantly reduced model loss and improved model accuracy by at least 2.17%, reaching the performance level of the deep learning fault diagnosis model based on normal data.