To improve the accuracy of recognizing knock intensity in aviation kerosene piston engines, a dedicated test bench was built, combustion tests were conducted, and cylinder pressure signals under various operating conditions were collected to extract knock information. Wavelet packet transform was applied to decompose the pressure signals, and energy entropy analysis was used to identify sub-band components effectively characterizing knock levels. The knock feature sub-band for this engine was found to be concentrated within the 7.5—18.75 kHz range. Knock information within this band was reconstructed, and 22 knock feature indicators were extracted from the time and frequency domains, forming a bi-domain knock feature image. To accurately identify and analyze knock intensity, both a multilayer perceptron (MLP) model and a convolutional neural network (CNN) model were developed. These models used the dual-domain knock feature indicators and images as input parameters and were validated under four different operating conditions. The results demonstrated high recognition accuracy for both models, with the CNN model achieving an average accuracy of 93.46%, surpassing the MLP model’s 88.50% by 4.96%. This highlighted the CNN model’s superior accuracy and generalization capability for knock intensity recognition in aviation kerosene engines.

Due to the particularity of the working fluid and extreme aviation conditions, dissolved gases （DG） are important factors that cannot be ignored in fuel cavitation. To reveal the independent mechanism of DG in the cavitation process, the cavitating flow characteristics around a two-dimensional Clark Y-11.7% hydrofoil in room-temperature water were investigated. Particular attention was paid to the convection-diffusion behavior of dissolved oxygen in liquid water as well as the dissolution/degassing processes. Using the Clark Y-11.7% hydrofoil as the research object, an in-depth exploration of the influence mechanism of DG on the fundamental cavitation of a hydrofoil was performed. Within the framework of the Schnerr-Sauer （SS） cavitation model, the convection-diffusion effect and dissolution/degassing processes of DG were coupled to establish the SS-DG cavitation model. Numerical simulations of hydrofoil cavitation were then conducted using this model. Combined with experimental results, the effects of DG on cavitation structures and hydrodynamic characteristics were analyzed in detail under two angles of attack (8° and 20°) and three different dissolved oxygen concentrations. The results showed that different concentrations of dissolved oxygen in the liquid altered the cavity structure near the leading edge of the hydrofoil. Under both angles of attack, the average lift and drag coefficients were only slightly affected by dissolved oxygen concentration, but its presence modified the unsteady characteristics. In addition, due to large-scale flow separation, the time-averaged dissolved oxygen concentration in the cavitation region at a 20° angle of attack was significantly higher than that at 8° under the same conditions.

In order to investigate the combustion mechanism of oblique detonation with hydrocarbon fuel internal injection, numerical simulations and direct-connected combustion heating experiments were conducted on an oblique detonation combustor simulating a flight Mach number of 8. A small-scale oblique detonation experimental piece was designed and the aviation kerosene was used as the fuel. Fuel injection and atomization was organized through 4×Φ0.3 mm holes on a diamond-shaped strut plate positioned in the center of the flow channel. An R2 mm bump was utilized to initialize detonation and stationary combustion on a 20° wedge surface. Numerical simulations of the fuel injection mixing process and engine combustion process were performed using a 10-step, 11-component chemical reaction mechanism, and the Reynolds-averaged Navier-Stokes (RANS) equations. The results indicated that the liquid fuel evaporated within a distance of 274 mm and partially mixed with the high-speed incoming flow. Stable oblique detonation flow fields were formed on both sides of the wedge surface. The flame position and morphological characteristics were observed through optical measurements. The combustion zone remained stable within an effective experimental window of 2.2 s. The experimental results are in good agreement with the numerical simulations in terms of pressure along the flow path and characteristics of the chemical reaction zone. The research findings demonstrate the technical feasibility of internal-injection hydrocarbon fuel-based oblique detonation engines.

The species distribution of reacting flow was investigated in the staged swirl spray flames under near-realistic conditions for advanced aero-engine applications. Simultaneous planar laser-induced fluorescence (PLIF) measurements of OH radicals and fuel (kerosene) were conducted in a centrally staged combustor at an inlet pressure of 0.5 MPa and an inlet temperature of 600 K. To address the difficulties and challenges in PLIF measurements of complex spray flames under engine-relevant conditions, a series of optimization measures were taken in the experimental procedure, optical layout, image acquisition, and post-processing. The simultaneous OH/fuel PLIF results revealed four typical flame structures in the combustor. Under constant inlet temperature, pressure, and total fuel supply, a reduction in the fuel staging ratio weakened the coupling between the pilot and main flames, leading to diverse variations in flame structures, and demonstrating the critical role of the pilot stage in flame stabilization.

Lightweight thermal protection materials are a key solution for large-area thermal protection systems in modern aerospace vehicles. A thermal response model coupling aerodynamic heat transfer and thermochemical processes was developed to comparatively investigate the thermal protection performance of nanoporous resin-based low-ablation materials and non-ablative ceramic tiles under two typical of thermal environments. It revealed the energy dissipation pathways and their respective contribution weights. Results showed that, compared with ceramic tiles, low-ablation materials achieved an overall energy dissipation of approximately 10%—15% through a synergistic combination of mechanisms including thermal blockage, heat sink effects, pyrolysis gas release, and pyrolysis reactions, demonstrating superior multi-path heat dissipation capabilities. Radiative heat dissipation was identified as a core mechanism shared by both materials, with its contribution significantly increasing under intensified thermal loads. Moreover, the nanoporous structure of the low-ablation material presented low thermal conductivity, along with higher specific heat capacity and density, enabling more effective heat confinement in the upper layer and significantly reducing heat transfer to the interior, thus resulting in a lower backside temperature compared with ceramic tiles. The study confirmed that resin-based materials can dynamically regulate the proportion of energy dissipation mechanisms and the distribution of heat sinks along the thickness direction, achieving efficient thermal protection across a wide range of thermal environments.

The wall temperature non-uniformity issue in tangential effusion cooling for a swirl combustor flame tube was addressed. Numerical simulations were employed to investigate the influence of effusion hole arrangement on cooling performance. A zoned refinement strategy was proposed based on near-wall flow characteristics: establishing a dense zone in the high-temperature hotspot region (axial 1/4—2/3 position) with sparse and transition zones at both ends. While maintaining constant total coolant flow rate, the hole spacing in the intensified cooling zone was reduced from 11 times the hole diameter in the baseline configuration to 8 times the hole diameter in the optimized configuration, with corresponding adjustments made to the hole spacing distribution in both the sparse and transition zones. Results demonstrated that the reduced hole spacing in the dense zone significantly decreased the peak wall temperature (up to 50 K reduction) and the high-temperature area, while the tapered hole spacing in transition zones enhanced film-cooling continuity, lowering wall temperature gradient below 44 K/cm to meet thermal protection requirements. The optimized configuration achieved 74% overall cooling efficiency, providing both theoretical foundation and technical support for swirl combustor liner cooling design.

To improve the cycle thermal efficiency and output power of the swashplate engine, a simulation model of a swashplate engine with pre-chamber hydrogen jet igniting ammonia was established. The influences of key structural parameters of the pre-chamber on the thermal jet in the pre-chamber and the engine performance parameters were analyzed. The results showed that appropriately increasing the cross-sectional ratio can advance the start time of the thermal jet and increase the penetration length of the thermal jet; when the cross-sectional ratio reached 0.225, the flame propagation speed in the cylinder was faster, and the comprehensive effect was the best. With the increase of the pre-chamber volume, the penetration distance and distribution area of the jet flame first increased and then decreased; when the volume ratio was 1%, the jet flame had a larger penetration distance and a larger distribution area, enabling the engine to achieve optimal balance between power output and economy under this working condition. The influence of the orifice diameter on the performance of the Swashplate engine also showed a trend of first increasing and then decreasing with the increase of the orifice diameter; when the orifice diameter was 1mm, the flame propagation speed in the main combustion chamber was faster, the distribution area was larger, and the comprehensive performance was better. This could provide theoretical reference and technical support for the optimal design of the pre-chamber structure of the swashplate engine and the efficient application of hydrogen-ammonia fuel.

This study examined the ignition performance and its influencing factors for a triple-header combustor employing torch ignition under low-temperature fuel conditions. A specialized low-temperature fuel delivery system was designed, and high-speed imaging techniques were utilized to capture the dynamic flame propagation process during ignition. Experimental investigations were conducted to analyze the effects of total pressure loss and air intake orifice size of the igniter on ignition characteristics. The results demonstrated that the ignition performance of the combustor deteriorated significantly under low-temperature fuel conditions. Under condition 1 and condition 3, successful ignition could not be achieved when the fuel supply rate of the starting nozzle reached 9 kg/h. Enlarging the air intake orifice area of the torch igniter and increasing the total pressure loss coefficient were found to enhance the ignition performance and extend the operational envelope of the engine. Furthermore, −35 ℃ low-temperature fuel substantially affected the flame propagation process to sectors distant from the igniter, markedly increasing the ignition delay time. The longest ignition completion time was 806 ms, while under normal-temperature fuel conditions, it could be reduced to 433 ms.

For evaluating the safety of Turboshaft engine power turbine over-speed protective device, the simulation test for over-speed was carried out on a specific type of Turboshaft engine under prescribed conditions and the fuel decay law of engine over-speed was obtained. Furthermore, a method for predicting maximum rotor speed after losing load was proposed. According to the relationship between the fuel flow decay law and output power, pneumatic torque, power turbine speed and acceleration, the power turbine speed at different moments during over-speed process was calculated by using sectional iterative calculation method. The maximum power turbine speed when shaft fracture occurs was obtained. Finally, a difference analysis was conducted on the highest speed post-shaft breakage of the over-speed protection devices equipped with different fuel cutoff timings. The results indicate that the duration of the stable decline phase during the fuel cutoff process plays a decisive role in the maximum speed of the power turbine speed of the power turbine. After adjusting the fuel cutoff time from 65 ms to 185 ms, the maximum power turbine speed rosed from 116.1% to 123.2%. To ensure the safety of the turbine after losing load, it is advisable to control the fuel cutoff time within 162 ms.

To resolve the conflict between high hover energy consumption and long endurance in compound-wing unmanned aerial vehicle (UAV), a high-energy-density hydrogen-electric extended-range hybrid power system is proposed. A cross-modal power scheduling strategy dynamically allocates output between the fuel cell and lithium battery, maintaining a 95% state of charge (SOC) during cruise to ensure hovering power redundancy and optimize global hydrogen consumption. Validated through simulations and flight tests, the prototype achieved a 28 min hover and a 230 km cruise. Notably, it reduced hydrogen consumption by approximately 8.3% compared to the rule-based equivalent consumption minimization strategy (RB-ECMS) strategy, providing an efficient and clean hybrid power solution.

To investigate the interference of a high-speed ground effect vehicle’s airframe on the nacelle flow field and installation effect on the net propulsive thrust, 3D numerical simulations were conducted for both isolated and installed nacelle configurations. Using thrust-drag bookkeeping method, thrust and drag components were extracted to analyse thrust changes between the two configurations in cruise conditions. Additionally, the effects of axial nacelle installation positions on net propulsive thrust were studied. Results showed that the wing’s entrainment and acceleration of the engine jet resulted in an increase of 0.4% in intrinsic net force and 1.5% in internal nacelle force, while the airframe’s influence on the nacelle cowl surface pressure distribution significantly raised external nacelle force in 18.24%, leading to net propulsive thrust reduction. Moving the nacelle closer to the wing enhanced jet entrainment and reduced external nacelle force by weakening the body interference, thus improving net propulsive thrust. The net propulsive thrust at the position closest to the wing was 1.46% higher than that at the position farthest from the wing.

The osculating cone waverider is studied in order to improve its wide area performance, five parameters, including the leading edge curve function coefficients a and l, shock wave curve function coefficients b and c, and the semi-cone Angle δ_c of the reference flow field, were selected as the design variables, and the sensitivity analysis of the design parameters was carried out in the range of Ma=4～7. The second generation of non-dominated sorting genetic algorithm is used to optimize the volume ratio, design point Ma6 and non-design point Ma=4 lift-drag ratio as the target variables. The results show a significant nonlinear relationship between the design parameters and the performance parameters, wherein the semi-cone angle of the reference flow field plays a crucial role in determining both the geometric and aerodynamic performance of the waverider. Compared with the reference model, the volume ratio of the optimized waverider is increased by 7.1%. At the design point of Ma=6 and the non-design point of Ma=4, the lift-drag ratio is increased by 3.4% and 10.7%. The optimized waverider over a wide Mach number range shows higher aerodynamic efficiency and lower total pressure loss, with a lift to drag ratio of 4.88 at Ma=7, which can be maintained at 3.90 even after leading edge bluntness. The optimized design significantly enhances the aerodynamic performance of both the design and non-design points of the waverider, and possesses good wide range adaptability.

Aiming at the axisymmetric inlet designed for an axial supersonic through-flow fan, a design method suitable for this kind of inlet is proposed. This kind of inlet has no expansion section structure, the outlet is the throat, and the whole flow field is supersonic. Firstly, the initial inlet is designed with Ma₀=2.5 as the design point, and the influence of typical parameters such as half cone angle on the inlet is explored. The results show that with the increase of half cone angle, the total pressure recovery coefficient σ at the outlet of the inlet increases and the length decreases. After analysis, shortening the inlet length is more effective for improving the σ. The central body surface of the internal pressure section is designed in the form of curved surface to shorten the length of the inlet and the test model is designed for wind tunnel test. The length of the inlet is shortened by about 23.3%. The numerical simulation results show that the σ of the inlet increases at different angles of attack at Ma₀=2.5 and 2.0. The curve of pressure ratio along the center of the numerical simulation and the schlieren of the external pressure section are in good agreement with the wind tunnel test, which verifies the correctness of the numerical simulation method and the inlet design method in this paper.

Taking a certain gas turbine as the research object, a component-based overall performance computational model was developed. Sensitivity analysis was conducted to select appropriate correction coefficients for component characteristics. Test data was then coupled to design a coordinated operating equation set, and multi-point optimization was applied to refine the component characteristics. This methodology ultimately yielded a high-precision overall performance simulation model for the gas turbine. Based on the data obtained from the modified overall performance simulation model, a gas turbine performance prediction model was established using a multilayer perceptron (MLP). The results showed that, compared with the test data, the prediction model based on the modified gas turbine overall performance simulation data had an error of less than 1% in the performance evaluation of each steady-state operating point, which met the engineering accuracy requirements. It was also 74.7% faster than the corrected model, significantly reducing the computing time, and providing technical support for real-time prediction of the overall performance of gas turbine.

The influence of fretting wear on crack initiation of dovetail joint structure was studied. The ball/plane finite element model was established, and the method of the fretting wear simulation and Smith-Watson-Topper (SWT) parameter calculation was given and verified by comparing with existing literature. A two-dimensional finite element analysis model of dovetail structure was established, and SWT parameter values of contact surface were calculated. It was found that fretting wear made stress distribution uniform and the SWT parameter values decline, so the risk of crack initiation was reduced. Because the wear on the upper edge was much greater than that on the lower edge, the location of SWT maximum value transferred from the upper edge of initial stage (no wear) to the lower edge, then, fretting wear made crack initiate on the lower edge. Considering fretting wear, the influences of friction coefficient, base angle and centrifugal force on SWT parameters of dovetail were discussed, and the fretting behavior between contact surfaces of dovetail structure was studied. The results showed that with the decrease of friction coefficient, the surface contact state changed from partial slip to complete slip, and the increase of surface wear resulted in the decline of SWT parameter. When the dovetail angle was 45°, the SWT parameters value was the smallest and the structure was the most reliable. With the enlargement of centrifugal force, the contact surface deformation and stress increased, so the SWT parameter could gradually augment.

Aiming at the problem that the vibration signal fault features of rolling bearing are weak when the fault occurs, a rolling bearing fault diagnosis method based on grey wolf algorithm optimized feature mode decomposition was proposed. Firstly, the signal was decomposed by feature mode decomposition to obtain several modal components. Then, the kurtosis-correlation coefficient selection criterion was used to filter and classify the modal components, and the high noise signal and low noise signal were output. Secondly, the high noise signal was decomposed by wavelet packet, and the signal component was reconstructed by weighted fusion index composed of kurtosis, skewness and information entropy to complete signal noise reduction and fault feature enhancement. At the same time, in order to solve the disadvantage that the key parameters of feature mode decomposition and the wavelet packet basis of wavelet packet transform need to be set artificially, the information entropy of the reconstructed signal divided by the kurtosis was selected as the objective function, and the grey wolf algorithm was used to optimize in a certain range and substituted. Finally, the signal was envelope demodulated to extract fault features. The simulation signals, the data set of Western Reserve University, the data of the main bearing test bed of turbofan engine and the fault test data of deep groove ball bearing were used to verify the proposed method. The results showed that the fault features of the processed signal were obvious, and the noise reduction effect of the method was good, the signal-to-noise ratio of the simulated noisy signal increased by 10.76, and the kurtosis value of the aircraft main bearing noisy signal increased by 1.94.

A dynamic absorber with nonlinear stiffness characteristics (nonlinear energy sink, NES) was introduced into a three-degree-of-freedom system to study the suppression of multi-frequency vibrations in the main system, and then extended to the real casing structure of an aero-engine with continuous medium characteristics. In terms of theoretical analysis, based on the complex variable averaging method, the intrinsic mechanism and necessary conditions for the nonlinear energy sink to achieve transient resonance capture and targeted energy transfer in a three-degree-of-freedom nonlinear system were revealed, and the dynamic response characteristics of the nonlinear energy sink in a multi-degree-of-freedom system were obtained. In terms of regularity analysis, the influences of different nonlinear stiffness and damping parameters on the energy dissipation characteristics of the nonlinear energy sink in a three-degree-of-freedom nonlinear system were studied, and the optimal vibration dissipation characteristics of the nonlinear energy sink in a multi-degree-of-freedom system were obtained. Based on the vibration energy visualization analysis method, the design method of the nonlinear energy sink was extended to the real engine combined casing structure, and the process of vibration energy transfer from the casing vibration source to the nonlinear energy sink was successfully visualized. From the perspective of vibration energy, the targeted energy absorption and dissipation characteristics of the nonlinear energy sink on the casing structure were revealed. A single nonlinear energy sink effectively suppressed the first three modal vibrations of the casing, reducing the amplitudes by 25.1%, 25.3%, and 25.2%, respectively, achieving broadband vibration suppression of the thin-walled casing structure by the nonlinear energy sink.

To optimize the vibration phenomenon of the swash plate axial piston pump, a finite element model of the piston/cylinder pair of the swash plate axial piston pump was established, and the model was verified to be highly accurate through tests, and the influence of working conditions parameters such as load pressure at the oil discharge port, cylinder rotational speed, and inclination angle of the swash plate on the pressure and cavitation characteristics of the oil film of plunger micro-textured were analyzed by using the isometric test method, and the influence of the three types of plunger micro-textured structures on the oil film characteristics was compared based on the analysis of the pressure and cavitation characteristics. Based on the pressure characteristics and cavitation characteristics, the influence of three types of plunger micro-textured on the oil film characteristics is compared. The results show that: when the plunger is in the phase of transition between the suction area and discharge area, the pressure gradient of the piston/cylinder pair film increases with the increase of load pressure at the discharge port, cylinder rotational speed, and swash-plate inclination, the positive pressure of the oil film increases and the degree of cavitation increases, and the degree of cavitation decreases with the rise of suction port pressure; the type of the plunger micro-textured is directly related to the characteristics of the oil film, and the three types of the micro-textured with rectangle grooves are the best, When the groove types are rectangular, trapezoidal, and inverted trapezoidal, the peak gas phase volume fractions are 7.143%, 7.216%, and 7.293%, respectively, and the peak positive pressures are 187.16 N, 211.5 N, and 219.8 N, respectively, which provide references for the optimal design of the vibration phenomenon of the piston pump.

The time series of the full-cycle friction torque of the bearing were taken as the analysis object, the fractal theory was used as the main analysis method, the optimal delay time was determined through the mutual information method, and the optimal embedding dimension was determined through the observation method, finally the correlation dimension was obtained. The results showed that the friction torque performance of the bearing had chaotic characteristics; the embedding dimension was positively correlated with the nonlinearity and chaos degree of the system, while the correlation dimension was negatively correlated. The correlation dimension can be used as an indicator for measuring the performance of bearings. Bearings A and B can predict bearing failure 120 min and 160 min in advance respectively. Based on the principle of maximum entropy, it was verified that the failure time of bearing A differed from that predicted by the fractal theory by 12 min, accounting for only 10% of the predicted time of 120 min. The deviation error was at a relatively small level, fully verifying the effectiveness and reliability of the fractal theory in characterizing the performance evolution of bearings.

To investigate the effects of global profile errors on the aerodynamic performance of a transonic compressor, a five-dimensional geometric variability model characterizing blade surface global profile errors was established based on Gaussian process and principal component analysis. Combined with the non-intrusive polynomial chaos expansion based on Gaussian distribution, a surrogate mode of compressor performance was proposed. The effects of the performance were quantified using non-intrusive polynomial chaos methodology, complemented by loss source analysis to investigate the critical flow mechanisms in two extreme-performance blade profiles. Key findings indicated that the performance parameters at peak efficiency condition exhibited slight deviations from normal distributions under the effects of random profile errors, with the performance showing degradation. And the performance was most sensitive to the error at blade tip leading edges. The thinner leading edge can reduce shock intensity and weaken shock-leakage vortex interactions, with the loss decrease.

To obtain the internal flow characteristics and develop a method for identifying flutter instability in single-stage fans, dynamic pressure sensors were installed at the inlet and outlet sections, and the fan was subjected to forced flutter experimental by controlling the throttle valve opening at different speeds. The dynamic pressure data during the fan’s rotational stall and flutter were obtained. The rotational stall and flutter characteristics and the evolution process of flutter were analyzed by combining time-domain signals and frequency-domain signals. The experimental results show that when the fan experiences flutter at different speeds, the stall vortex is produced, developed, and dissipated; during the throttling flutter process, there is a transition period during which the fan transitions from rotational stall to flutter, and there is also a 1—2 cycle recovery process when exiting flutter; there are mild flutter phenomena at all speeds, which are caused by the gas compression in the flutter process that leads to a lag effect upstream and downstream, and the flutter evolution process of the fan is simplified into an equivalent force mechanical model for analysis. The mutual correlation analysis between sensors at the same circumferential and different axial positions is proposed based on the equivalent force mechanical model, and a flutter monitoring and identification method based on mutual correlation time-frequency analysis is developed. This study provides a basis for obtaining the rotational stall and flutter evolution process of the fan, as well as online monitoring and fault analysis of flutter. However, due to experimental conditions, additional dynamic pressure sensors were not installed, and more details about the rotational stall and flutter process could not be obtained.

This study employed numerical simulations to investigate the influence of controllable speed casing treatment with different rotational directions and rotational speeds on the performance of a subsonic compressor stage. The results indicated that when the rotational direction of the rotatable ring segment rotated in the same direction as the rotor, an increase in rotational speed led to a reduction in both the stable operating margin and peak efficiency. The rotatable ring segment exerted a reverse force to the tip leakage flow, thereby reducing the leakage flow velocity and exacerbating flow blockage of the flow channel. Conversely, when the rotational direction of the rotatable ring segment rotated opposite to the rotor, an increase in the rotational speed expanded the stable operating margin. The rotatable ring segment exerted a co-directional force to the tip leakage flow, thereby increasing the leakage flow velocity and improving the flow capacity of the flow channel under near stall conditions. When the rotatable ring segment rotated at 100% of the rotor speed in the opposite direction, the stable operating margin of the compressor stage improved by 16.34%, and its pressure ratio increased by 0.12%.

In fluid-structure interaction simulations of aviation gear pumps, two key issues: the distinct geometric mismatch between fluid and solid domains resulting from radially scaling gears to connect gapless tooth tip and meshing zone flow fields; and the notable risk of inaccurate load transfer due to the classical radial basis function (RBF) mapping algorithm neglecting fundamental physical conservation principles, were addressed in this research. To solve these critical problems, force and moment conservation constraints were carefully introduced between the source and target domains. The minimum norm method was effectively employed to correct the target field grid pressure, and an advanced RBF data reconstruction algorithm with significantly enhanced conservation properties was successfully proposed. Combined with detailed grid analysis and rigorous experimental verification, comprehensive fluid-structure interaction simulations were conducted for a specific type of aviation gear pump. Test results showed that the advanced RBF algorithm effectively eliminated total force and moment deviations, with an average mapping error as low as 0.0015%. When loading flow field loads onto solid grid nodes by the advanced RBF algorithm, the static maximum equivalent stress was measured at 31.61 MPa, and the dynamic meshing stress peak reached 192.18 MPa. The flow field pressure contributed 34.38% to solid stress and 29.84% to strain, respectively. These important findings clearly demonstrate that fluid-structure interaction simulation is highly necessary for accurately evaluating the actual service status of aviation gear pumps in practical operational environments.

The wear and degradation of sliding bearings is one of the important factors affecting the life and reliability of aviation fuel pumps. The numerical simulation model based on this object sacrifices computational efficiency while continuously improving the fitting degree with test results, which is not conducive to its application in engineering practice. This article proposes a method of parameterizing the wear amount and establishing a Gaussian process model. By parameterizing the circumferential distribution of wear amount at different speeds, eccentricities, and wear times, the input-output parameters are simplified, and a Gaussian process regression model with much higher computational efficiency than traditional simulation methods is constructed. On this basis, the error and confidence level of the model prediction results were studied under different training sets, operating conditions, and wear times, and the accuracy of the model was verified by comparing the test results with the model prediction. Research has found that selecting an appropriate training set can partially improve the predictive performance of the model, manifested as a narrowing of the 95% confidence interval; When the journal speed, eccentricity, and wear time are used as independent variables, the results provided by the prediction model are highly consistent with the simulation model, with a maximum relative error of 5.26%. Furthermore, by comparing the model predictions with test results under the same operating conditions, it was found that the relative error was less than 5%, which meets the accuracy requirements and validates the rationality of the prediction model. The rationality of the prediction model is verified; When using wear time as the independent variable, the average absolute error is the lowest, only 0.001 µm, and when using eccentricity as the independent variable, it is the highest, at 1.33 µm; Further comparative studies have found that the prediction errors of the model come from two aspects: systematic errors in the parameterization process and systematic errors in the Gaussian process regression model. The dominant factors of prediction errors under different operating conditions and wear times have been analyzed.

In substation environments, acoustic-signature-based unmanned aerial vehicle (UAV) detection systems have recently attracted significant attention because, relative to electro-optical or radio-frequency solutions, they can deliver full coverage of the surrounding airspace at substantially lower cost. In view of the complex electromagnetic noise environment and high requirements for spatial coverage in substations, a drone monitoring system based on a distributed acoustic array was proposed. The system employed dual tetrahedral microphone arrays to capture acoustic signals, and a 500 Hz high-pass filter combined with a log-Mel spectrogram extraction algorithm was designed to suppress current-induced noise. On this basis, a ResNet-18 detection network was constructed, achieving a drone detection accuracy exceeding 90% and a false alarm rate below 4% within a 200 m range. To address the instability of time delay estimation for distant acoustic signals, a time delay continuity judgment and outlier rejection strategy was introduced, along with a least-squares hyperbolic localization model, thereby improving the drone localization performance within 100m in practical scenarios. Experimental results verified the effectiveness of the proposed method.

Thrust vectoring technology has gradually become one of the important means to enhance aircraft maneuverability and fault tolerance. Effectively integrating thrust vectoring technology with traditional control systems to improve the fault tolerance of overall control strategies is a pressing research issue to be addressed. This paper establishes the model foundation for thrust vectoring control strategies by constructing an integrated aircraft/engine controlled object and actuator model. For the elevator failure scenario in thrust-vectored fixed-wing aircraft, a fault-tolerant control strategy based on switching control is proposed. This strategy combines traditional PID control with active disturbance rejection control (ADRC) algorithms, switching based on system state: PID control is used for the elevator under normal conditions, while switching to ADRC for thrust vectoring angle control during elevator failure, thereby ensuring system stability and control performance. Nonlinear simulation verification through joint testing of the control system and the integrated aircraft/engine controlled object demonstrates the effectiveness of the proposed method. The results show that, compared to the baseline controller, the designed controller improves rise time performance by approximately 15% and settling time performance by about 2% under fault conditions. Moreover, the switching control strategy maintains system stability and performance across various fault scenarios.

The effective ion heating of the ion cyclotron resonance heating (ICRH) stage is of vital importance for the variable specific impulse magnetoplasma rocket (VASIMR) to become an effective propulsion device. A multi-component fluid model in which the helicon plasma source (HPS) and the ICRH stage are connected in series was developed. The simulation was performed with different antenna length and different input frequency to analyze the ion energy coupling mechanism of the ICRH stage. The numerical results demonstrated that, the azimuthal ion current density resonated with the azimuthal electric field in the ICRH stage and ions can absorb energy from electromagnetic fields continuously due to the resonance. The resonant area got larger as the antenna length of the ICRH stage increased and ions could deposit energy from electromagnetic fields more efficiently. The input frequency of the ICRH stage had a great influence on the ion heating effect. Due to the existence of ion collisions, the optimum input frequency of the ICRH stage was lower than the ion cyclotron frequency and the heating efficiency decreased as the ion temperature increased.