To address the limitation of continuum/rarefied-flow coupling methods in handling multiple grid types, a three-dimensional Navier-Stokes (N-S)/Direct Simulation Monte Carlo (DSMC) coupling framework for arbitrary grid types was developed. A linked-list data structure was adopted to uniformly describe different grid topologies and face-adjacency relations in the continuum and rarefied regions. Based on the face-cell adjacency information, a molecular cell-to-cell migration search algorithm was developed by combining an oriented-volume criterion based on the divergence theorem with a line segment-triangle intersection method. The local Knudsen number was employed as the continuum breakdown criterion to automatically partition the N-S and DSMC regions and generate complex three-dimensional coupling interfaces. Bidirectional information transfer was realized by generating one or multiple layers of information-transfer cells on both sides of the interface and applying the boundary-state method. Hypersonic flow over a sphere was simulated using tetrahedral, hexahedral, and hybrid grids. The results showed that, compared with full-domain DSMC simulations, the computational time was reduced by 45.6% for tetrahedral grids, 39.8% for hybrid grids, and 33.7% for hexahedral grids. The wall pressure coefficient and heat flux obtained with the three grid types agreed well with the reference DSMC results. These results indicate that the proposed method has good computational efficiency and adaptability to different grid types.

Using simultaneous OH planar laser induced fluorescence (PLIF) and stereoscopic particle image velocimetry (S-PIV), this study investigates the stability boundaries and flow-flame interaction mechanisms of hydrogen-enriched methane/air premixed flames in a low-swirl burner. Two turbulence generators (TGs) featuring nearly identical blockage ratios but different orifice diameters were compared. The macroscopic flame morphologies, blow-off limits, and the underlying stabilization mechanisms were examined across a range of hydrogen fractions, equivalence ratios, and bulk flow velocities. Results show that both TG configurations produce three characteristic flame shapes: bowl, W, and crown. However, TG2, designed with larger and more sparsely distributed orifices, demonstrated a significantly broader stability range than TG1. Detailed flow field analysis revealed that TG2 generates a high-speed central jet core flanked by low-speed outer regions at the burner exit. This specific velocity profile induces an additional shear layer (ASL) that promotes rapid flow deceleration downstream. Consequently, the flame can anchor further upstream, which effectively extends the blow-off limits. Building upon these mechanistic insights, a predictive model based on the Damköhler (Da) number was developed. By correlating the local flow and chemical time scales, this model can accurately predict flame blow-off boundaries under diverse operating conditions. Ultimately, this study highlights the critical role of turbulence generator geometry in the design of low-swirl combustor injectors, providing valuable guidelines for flame stability optimization in practical hydrogen-enriched combustion systems.

To clarify the effects of relative swirl direction on the pure-hydrogen dual-swirl combustion dynamics, atmospheric tests were conducted to compare counter-rotating (CR) and co-rotating (CO) configurations with identical fuel orifices, flow passages, and nominal swirl-number magnitudes. The air volume flow rates were 2.20, 3.30, and 4.40×10⁻³ m³/s, and the equivalence ratios ranged from 0.05 to 1.00, covering 120 conditions in total. Dynamic pressure and high-speed OH* chemiluminescence measurements were used to characterize the pressure amplitude-frequency and flame projection morphology. Pressure fluctuations varied non-monotonically with equivalence ratio. CR exhibited narrower high-response regions with a maximum pressure root mean square of 0.28 kPa; CO produced broader high-response regions at medium and high flow rates, reaching 0.55 kPa. At the highest flow rate and equivalence ratios of 0.80 to 1.00, the CO value was 2.6 to 3.0 times that of CR. The dominant frequencies ranged from 0.70 to 1.14 kHz. CO exhibited continuous narrowband spectral ridges at medium-to-high flow rates and high equivalence ratios, whereas CR showed stronger broadband and multifrequency features during strong low-flow responses. The projected area increased rapidly to (2.5 to 2.8)×10³ mm² at equivalence ratios of 0.20 to 0.50 and then approached saturation. CR formed a tall, narrow envelope earlier, while CO retained a low, wide cup-shaped structure over a broader range. The relative swirl direction changed the pressure-response windows, spectral organization, and projected flame morphology.

This study investigated the impact of casing deformation on the Rotor 37 transonic compressor rotor using validated numerical simulations at design speed. The results showed that under the ideal (undeformed) casing condition, the breakdown of the rotor’s tip-leakage vortex served as the direct trigger for compressor stall. As elliptical deformation τ increased, the compressor’s stability margin steadily decreased. At maximum elliptical deformation (τ = 0.9), stall margin dropped by 4.67% relative to the ideal casing, and flow instability first appeared in passages adjacent to the smallest tip clearance. Simultaneously, casing deformation induced pronounced circumferential variability in the flow field and blockage-zone extent within the blade-tip region. On one hand, circumferential non-uniformity in leading-edge tip clearance and local flow incidence controlled the leading-edge leakage strength in each passage, shaping blockage-zone extent. Blockage was most severe near maximum tip clearance, triggering instability in that passage. On the other hand, circumferential non-uniformity in trailing-edge tip clearance and suction-side separation location controlled the trailing-edge leakage intensity and separation-blockage severity. These factors maximize blockage near minimum tip clearance, causing the earliest instability in that passage.

To enhance computational efficiency and accuracy in reliability analysis for implicit and highly nonlinear structures, this study develops an improved adaptive Kriging-based reliability analysis method. in the infill strategy of the Kriging model, a sign function and a distance-based weight function were incorporated to construct an Extended Expected Improvement learning function, ensuring that sample updates are concentrated near the limit state surface. An adaptive reliability analysis framework was established by integrating Monte Carlo simulation, which improved computational stability and reduced the number of additional samples required. Validation through classical mathematical examples and engineering problems demonstrated that, compared with Monte Carlo simulation, the relative error in failure probability was within 0.5%, and the number of new samples was significantly reduced compared to existing reliability analysis methods. Finally, a finite element case study on cavitation reliability in an aviation fuel centrifugal pump showed that the error in failure probability was only 1.06% relative to Monte Carlo simulation, while the computational cost was reduced by 51.06% compared to existing reliability analysis approaches. These results verify that the proposed method achieves both high precision and high efficiency in reliability analysis of implicit nonlinear structures.

To evaluate the performance of engine inlet hot-air anti-icing systems under different structural designs, a mesh-independent point-cloud slice attention network (PSANet) was developed for rapid pointwise prediction of the surface temperature and runback water distribution over the inlet. In this method, the anti-icing surface was represented as a point cloud, and six-dimensional structural design parameters of the piccolo tube were used as inputs. Multi-band Fourier feature mapping, conditional modulation, slice attention, and vector-gated reconstruction were introduced, and a multi-task learning strategy was adopted to jointly predict the temperature distribution and the mass flow rate distribution of runback water. Based on an in-house numerical simulation method for anti-icing, 5 000 cases of anti-icing surface temperature and runback water mass flow rate under different piccolo-tube structural designs were generated. These cases were randomly divided into training, validation, and test sets, with each case containing 23 958 surface points. The test results show that the mean absolute error and root mean square error for temperature prediction were 0.503 K and 0.841 K, respectively; the mean absolute error and root mean square error for runback water mass flow rate prediction were 0.011 mg/s and 0.018 mg/s, respectively. The prediction time for a single case was 38.57 ms, representing an acceleration of approximately 6, 880 times compared with the anti-icing numerical simulation method. Overall, the PSANet method achieves high prediction accuracy and computational efficiency, and can be applied to sensitivity analysis of structural design parameters and iterative design of hot-air anti-icing systems.

The swirl effects on the performance of lobed exhaust system from a helicopter engine are studied based on computational fluid dynamics. Firstly, using the exhaust system of a civil helicopter engine as a verification example, the calculated pressure loss is in good agreement with the flight data, which proves the reliability of simulation method. Then, the aerodynamic characteristics of the exhaust system under non-swirl are analyzed. The ejection ratio of the lobed nozzle with a lobe height of 0.55R is 71% higher than the cylindrical nozzle, while the total pressure loss coefficient only increases by 17%. Finally, the effect of swirl on the performance of a lobed exhaust system is studied. The lobed nozzle has a counteract rotation effect, and the exhaust swirl enhances the ejection ratio and total pressure loss coefficient. When the swirl angle is less than 10°, the total pressure loss coefficient and ejection flow rate increase slowly. While when the swirl angle is larger than 15°, both the total pressure loss coefficient and ejection flow rate increase rapidly. Compared with non-swirl flow, the total pressure loss coefficient and ejection flow ratio of lobed exhaust system increase by 84% and 46%, respectively, at 25° swirl angle.

To comprehensively and accurately characterize the performance characteristics of axial compressors with a variety of pressure ratios across the full rotational speed range, based on previous research, this study further investigated the parametric relationships at the peak efficiency points of compressors with all pressure-ratio using a symbolic regression which is a machine learning approach. A modeling framework leveraging symbolic regression for complex, multi-source, and multi-parameter systems is proposed. Using the PySR symbolic regression library, data points are fitted individually and subjected to multi-round screening and validation. From the characteristic data of 47 compressors, general correlations on the peak efficiency points are derived: both the total pressure ratio–rotational speed and mass flow rate–rotational speed relationships conform to a hyperbolic tangent function, whereas the efficiency–rotational speed relationship exhibits a piecewise quadratic and exponential behavior. By integrating these derived correlations on peak efficiency points with previously established equations for the near-surge and near-choke regions of the characteristic curves, a nonlinear compressor performance model was developed. The proposed model demonstrates the capability to predict the performance characteristics of a broad range of compressors over nearly the entire speed range with high fidelity.

Recent research progress on shock-induced droplet deformation and breakup in supersonic flows was reviewed from the perspectives of droplet breakup dynamics and two-phase physical property effects, near-field shock structures around droplets, turbulence evolution, internal droplet flow, and numerical simulation methods. Existing relevant studies mainly focused on the influences of two-phase physical parameters on breakup modes, the complex shock-wave systems formed during shock–droplet interactions and their characteristics, and the regulatory effects of turbulence evolution on interfacial instability and droplet breakup processes. It is generally recognized that shock-induced droplet breakup is a complex multiscale process jointly governed by aerodynamic forces, interfacial instabilities, and turbulence. However, current studies on the effects of shock waves, turbulence, and related factors on droplet breakup remain largely at a qualitative stage. Future research should be based on cross-scale high-fidelity numerical methods to further strengthen the quantitative characterization of the effects of shock waves and turbulence on droplet breakup and to achieve the decoupling analysis of their respective influences on the breakup process.

To investigate the atomization field of liquid-liquid coaxial centrifugal dual injectors, a cross-scale computational framework based on the VOF-to-DPM （volume of fluid to discrete phase model） method was constructed and validated through comparative analysis with cold-flow experiments. The simulation model achieved a spray cone angle error of within 0.85% and the Sauter mean diameter error of no more than 20%. The study found that for the liquid-liquid coaxial centrifugal injector, an increase in the flow rates of both the inner and outer injectors led to a gradual increase in the spray cone angle, while the droplet size simultaneously decreased in both the interaction and non-interaction zones. The influence of inter-injector distance on the atomization characteristics was also examined. Simulation results for different injectors spacings indicated that as the distance between the spray plumes increased, the spray cone angle gradually increased. The variation pattern of the Sauter mean diameter of the droplets was related to the state of liquid film interaction: when the interaction point lied within a continuous and stable liquid film, a larger injector spacing resulted in a smaller Sauter mean diameter; conversely, when the interaction occurred within an unstable, breaking liquid film, the trend was reversed.

The adaptive cycle engine (ACE) achieves comprehensive performance advantages for various flight missions through the collaborative regulation of variable mechanisms. However, the numerous variable mechanisms and intense changes in the operating states of ACE results in significant shortcomings in the convergence and computational efficiency of the existing performance simulation model. These shortcomings make it difficult to meet the demands for large-scale performance optimization design of ACE in the aircraft-engine co-design. A performance simulation acceleration method driven by both principles and data is proposed. This method constructs a general global initial guess prediction model based on principle analysis and a data-driven self-expanding initial guess variable fidelity surrogate model, forming a principle-data joint-driven engine initial guess prediction framework. This framework addresses the problems of poor convergence and low computational efficiency caused by the difficulty in reasonably selecting initial guesses for the performance simulation models under various configurations and design schemes. Numerical simulation verification is conducted to optimize the ACE control laws at equal-inlet flow throttle state, and the results demonstrate that the proposed method achieves more than a 38% improvement in convergence capability and a 45% increase in simulation computation speed compared to traditional methods. This method effectively supports large-scale simulation optimization of ACE design parameters and control laws in aircraft-engine co-design, holding significant engineering application value.

To address the challenges of dynamic optimization, flow-rate disturbances, and safety-operation constraints in the exhaust environment pressure control system during high-altitude chamber flight-environment simulation tests, a composite model predictive control (MPC) strategy incorporating an extended state observer (ESO) and a control barrier function (CBF), referred to as MPC-ESO-CBF, was proposed. The integrated characteristics and dynamic processes of the core equipment in the high-altitude chamber exhaust environment simulation system were analyzed. To meet the demand for dynamic tracking control of exhaust environment pressure, an MPC controller was designed to achieve optimal control performance. Considering unmodeled dynamics and external disturbances in the system, an ESO was constructed to enable real-time estimation and compensation. To ensure safe system operation, a CBF was introduced to strictly confine key state variables within a prescribed safe region. A high-fidelity digital simulation platform for exhaust pressure control in high-altitude chambers was developed, and the proposed control strategy was systematically validated and compared with a linear active disturbance rejection controller (LADRC). The validation results showed that, compared with LADRC, the MPC-ESO-CBF strategy exhibited superior dynamic performance under transient thrust-variation conditions: the adjustment time of exhaust-environment pressure was reduced by approximately 38.7%, the transient peak disturbance magnitude decreased by about 33.7%, and the steady-state average error was reduced by roughly 49.1%.

For the LIPS-300S ion thruster, which operates in multiple operating modes for deep space exploration missions, a belief reliability modeling, experiment design and verification framework was established. This framework encompassed reliability modeling, experimental design, and model updating and verification. A reliability model for the multi-mode ion thruster was constructed based on belief reliability theory considering sputtering erosion-induced electron backstreaming failure. An experimental design method was proposed with the aim of accurately verifying the degradation laws of the ion thruster. Subsequently, equivalence updating was applied to interdisciplinary equations based on the experimental controllable and measurable capabilities, mitigating model biases caused by unverified detailed physical derivations. Additionally, a linear correction factor was introduced to calibrate the degradation equations, addressing discrepancies between simulations and actual performance degradation. The effectiveness of the framework was validated using 12,000 h reliability experimental data from the LIPS-300S ion thruster. The results demonstrate that the updated and calibrated model can accurately characterize the performance and degradation patterns of the ion thruster across different operating modes, with 80% of the prediction errors within 5%. This work supports credible reliability evaluation for multi-mode ion thrusters, and effectively shortens the experimental time for reliability verification.

To account for mixed uncertainties, including geometric and material factors, under the operating conditions of aero-engine turbine disks, this study developed a probabilistic low-cycle fatigue life model. A comparative analysis of various uncertainty models was conducted, and a probabilistic representation model for geometric uncertainty was established. Through the integration of a deep neural network surrogate model with quasi-Monte Carlo simulation algorithms, an interval quasi-Monte Carlo sampling approach was formulated for reliability analysis incorporating geometric and material mixed uncertainties. This enabled probabilistic fatigue life prediction and reliability analysis, thereby quantifying the upper and lower bounds of life reliability. At a reliability level of 99.87%, the lower bound of fatigue life was 1.61×10⁴ cycles and the upper bound was 2.53×10⁴ cycles, These results provide methodological support for high-reliability turbine disk design.

To address the issue that existing flight state recognition methods fail to consider the temporal sequence and regularity of flight, resulting in insufficient recognition accuracy under non-stationary flight conditions, a helicopter flight state recognition method based on temporal attention and state transition constraints was proposed. A multi-channel perceptual residual module was introduced to effectively extract key representation information from flight parameters. A temporal attention module was designed to capture the correlation of flight parameters over time. A Markov state transition constraint module was constructed, combining state transition priors to output continuous and flight-regulated state recognition results. Experiments on a flight parameter dataset collected from an actual helicopter flight indicated that, compared with mainstream methods, the proposed method improved the precision, recall, and F1 score of state recognition under non-stationary flight conditions by 2.88%, 2.73%, and 2.81%, respectively. The model achieved a single inference time of 23 ms and a computational cost of 20 Mflops, thus demonstrating significant engineering application value.

Manufacturing deviations in turbine blades inevitably cause film cooling hole misalignment relative to the cast profile, compromising cooling effectiveness. This study proposed an adaptive hole positioning optimization method based on 3D-scanned geometric deviations, and validated through conjugate heat transfer (CHT) simulations and experiments. By integrating reverse engineering with a rigid-body transformation algorithm, an optimized blade model was generated to compensate for profile distortions. Comparative analysis against the ideal design and an unadjusted reverse-engineered blade demonstrated that the proposed method significantly mitigated the adverse effects of hole drift. Under typical operating conditions, the optimized blade achieved local temperature reductions of up to 9.82 K in critical pressure-side regions and improved mid-span cooling efficiency by approximately 1.3%, particularly at high coolant mass flow rates. These findings confirm the efficacy of the adaptive strategy in enhancing cooling robustness under real-world manufacturing constraints.

To improve the bird impact resistance of shouldered aero-engine fan blades, a new design method is investigated. The bird strike on fan-blades simulation in the rotating state was verified with a comparison to the experiment. By studying bird-impact induced failures at different locations and the sensitivity of structural parameters, the risk locations and the key structural parameters affecting the bird impact resistance were identified. An new design was proposed. The results show that the 57% to 83% blade height on the leading edge is the weak location range of bird impacts on the shouldered fan blades; the thickness of the leading edge fillet and the thickness of the leading edge profile are the key structural parameters that affect the bird impact resistance of blades. Increasing the two parameters can effectively improve the bird impact resistance of the blade. The bird impact resistance improvement design scheme can be applied to engineering design.

Aiming at the circumferential equidistant rubbing fault between the rotor and the case in the development process of the engine, the basic theory of the vibration of the cylindrical case was introduced. The form and reason for circumferential equidistant rubbing between the rotor Sand the cylindrical case were analyzed. The mechanism of the circumferential equidistant rubbing between the rotor and the cylindrical case was revealed. The conditions of circumferential equidistant rubbing between the rotor and cylindrical case were proposed. A general identification method for circumferential equidistant rubbing is established. The identification method was illustrated using an example of a circumferential equidistant rubbing fault between the rotor and the seal seat in the engine. The research has shown that the circumferential equidistant rubbing between the rotor and the cylindrical case is usually the result of two vibration interactions excited by two independent excitation sources. The fluid will excite the pitch diameter travelling wave resonance of the cylindrical case. When the pitch diameter travelling wave resonance and the rotor vibration satisfy some conditions, the circumferential equidistant rubbing will occur. The number of rubbing points depends on the pitch diameter number and the ratio of the rotor speed to the pitch diameter resonance frequency of the cylindrical case. The example of a circumferential equidistant rubbing fault between the rotor and seal seat in the real aero-engine proves the effectiveness of the general identification method for circumferential equidistant rubbing. The general identification method can guide rubbing fault diagnosis and case dynamics design.

In probe-and-drogue aerial refueling, the position of the drogue is challenging to control under the bow wave effect, frequently leading to docking failure. To address this issue, a dynamic simulation platform for the probe-and-drogue aerial refueling docking process was developed to calculate the displacement response of the drogue at different pre-docking distances. The computational results indicated that the pre-docking distance significantly affected the variation in drogue sinkage. A larger pre-docking distance extended the duration of the docking process, thereby increasing the cumulative interference time of the bow wave effect, and resulting in a greater dynamic variation amplitude of sinkage throughout the process. To mitigate the excessive impact of the bow wave effect on drogue sinkage, the longitudinal initial position envelopes for successful docking were determined for different pre-docking distances. Optimization results demonstrate that, in addition to aligning with the center of the drogue, applying an upward offset of 50–70 cm based on the pre-docking distance can significantly enhance the docking success rate. Moreover, this strategy exhibits robust performance against atmospheric disturbances.

The influence of swirl intensity and hydrogen nozzle configuration on the combustion performance of a pure-hydrogen combustor was investigated. A dual-swirler design with hydrogen injection at the venturi exit was adopted. Through combined numerical and experimental analysis, the effects of inner swirl presence, outer swirl number (ranging from 0.9 to 1.5), and hydrogen orifice number (12 and 15) on mixing, flame structure, and NOx emissions were systematically examined. Results indicated that increasing the outer swirl number from 0.9 to 1.5 reduced the mixing distance from 45 mm to 15 mm and lowered NOx emissions by approximately 10 r/min at a constant equivalence ratio. While the inner swirl promoted flame stability, an inner axial-flow design dispersed the central high-temperature zone, decreasing its volume by about 30% and further suppressing NOx formation. Increasing the number of hydrogen orifices from 12 to 15 improved upstream circumferential fuel distribution, reduced local peak temperatures, and shortened flame length without altering the overall mixing rate, thereby reducing NOx emissions. The combustor also exhibited effective thermoacoustic suppression, with pressure oscillations maintained below 20 Pa, ensuring stable operation across all tested conditions. This work provides a viable design strategy for high-efficiency, low-emission pure-hydrogen combustion systems, supporting the development of clean energy technologies and carbon-neutrality objectives.

An optimization method for shot peening parameters considering real rough tooth surfaces was proposed to meet the high strength and toughness requirements of high-performance aeronautical gears. A coupled finite element method-discrete element method (FEM-DEM） model incorporating ground surface topography was developed to investigate the effects of surface coverage, shot diameter, and shot peening intensity on shot peening performance. A three-factor (surface coverage, shot flow rate, air pressure) and three-level shot peening simulation scheme was designed using response surface methodology to achieve maximum residual compressive stress and minimum roughness. Polynomial models were established to analyze the interactions among these factors and their influence on shot peening effectiveness. The MS-NM algorithm was utilized to optimize the shot peening process with constraints on surface coverage, shot flow rate, and air pressure. The proposed process was then validated through simulations on different initial roughness models. The results showed that the prediction error between the response surface model and simulation results is less than 2%, indicating that the model can be effectively used for shot peening effect prediction and process optimization.

A bolt hole design optimization method based on Non-Uniform Rational B-Splines (NURBS) curves was proposed. The three-dimensional model of the bolt hole was constructed using a two-segment NURBS curve with pole-based control, with the optimization objective of minimizing the maximum equivalent stress. The pole coordinates served as design variables, while constraints included the maximum radial displacement of the baffle, maximum circumferential stress of the bolt hole, and bolt clamping area. A mathematical optimization model for the bolt hole was established, and the enhanced learning differential evolution (QLDE) algorithm was employed for optimization. The results showed that the optimized bolt hole achieved a 19.0% reduction in maximum equivalent stress, a 12.7% decrease in maximum circumferential stress, a 0.3% reduction in maximum radial displacement of the baffle, and a 0.5% increase in bolt clamping area, while the low-cycle fatigue life of the baffle improved by 122%. Experimental validation confirmed the effectiveness of the proposed method, with significantly enhanced low-cycle fatigue life. This addressed the failure issue of premature crack initiation in baffle bolt holes due to inadequate low-cycle fatigue life.

To mitigate severe corner separation in large-turning-angle diffuser cascades, this study systematically investigated the control effects of endwall synthetic jets on a cascade with a geometric turning angle of 68°. The jets were arranged at four distinct positions: the leading edge, mid-section (encompassing the locations anterior and posterior to the separation reattachment point), and trailing edge of corner separation. By coupling three core control parameters—jet angle, momentum coefficient, and excitation frequency, this study elucidated the modulation mechanisms of synthetic jets on the endwall boundary layer, separation bubble, and complex secondary flow structures. The results indicated that end-wall synthetic jets can effectively weaken corner separation. Specifically, the end-wall synthetic jet arranged at the middle region of the separation (before the reattachment point) with a jet angle of 30° can effectively act on the core area of the separation bubble, break up the corner separation vortex, and block the boundary layer migration. This configuration achieved a maximum reduction of 20.27% in total pressure loss coefficient and an increase of 13.07% in static pressure rise coefficient. In contrast, the end-wall synthetic jet placed at the leading edge of the separation with the same jet angle of 30° inhibited the separation expansion from the source, resulting in a 14.69% reduction in total pressure loss coefficient and a 9.66% increase in static pressure rise coefficient. However, the end-wall synthetic jets arranged at the middle region of the separation (after the reattachment point) and the trailing edge exhibited limited control effects. The study proposed a control strategy covering the entire evolution stage of corner separation, and revealed the adaptation mechanism between synthetic jets and separation stages, offering an approach to optimize the aerodynamic performance of large-turning-angle diffuser cascades.

Tandem hypersonic vehicles suffer from nonlinear multibody interference and attitude instability during booster separation, which seriously affect flight safety and reliability. Conventional separation simulation methods have difficulty in capturing this complex dynamic process accurately. A Numerical Virtual Flight method for tandem hypersonic vehicle stage separation was proposed, and an aerodynamics-motion-control coupled simulation platform was established. Different separation schemes were designed and evaluated at Mach 10 by considering the initial separation attitude, separation-mechanism actuation, and separation control strategy. The results showed that an initial angle of attack of 4° brought the two stages close to trim and improved the initial separation stability. Compared with free separation, the hydraulic strut separation mechanism increased the relative axial distance between the two stages by 0.49 m at 0.2 s and reduced the relative pitch angle by about 52%. The nonlinear dynamic inversion control system further improved attitude stability during separation; under high-gain control parameters, the pitch-angle fluctuation amplitude of the main stage was reduced by about 90% compared with the uncontrolled case. The final scheme satisfies the safe-separation criteria and provides reliable data support for two-stage separation scheme design.

Thermal loads inside and outside the aircraft pose a non-negligible challenge for horizontal takeoff and landing, reusable high-Mach aircrafts. To meet the thrust requirements of the vehicle and utilize the thermal loads efficiently, four thermodynamic cycle layouts of propulsion and energy management system were proposed based on a precooled engine with a closed helium cycle. These layouts differed in the strategies for heat absorption and utilization. Optimizations were conducted with the power output of closed cycles, specific impulse, and comprehensive performance as the respective objectives, while maintaining the same thrust among the cycle layouts. To comprehensively assess the performance of cycles, a multi-attribute evaluation method for propulsion and energy management systems was proposed, and the Grey Relational Analysis was adapted. The results indicated that the multi-branch staged cooled cycle could achieve a maximum output power of 18.23 MW; however, its system-level performance was not optimal. In contrast, the dual-branch partially cooled cycle exhibited the best comprehensive performance because of its balanced performance under different optimization objectives and application scenarios. The research results provide insights for multi-metric integration design of propulsion and energy management systems in high-Mach vehicles.

Special tasks such as reconnaissance and eavesdropping poses higher requirements for the miniaturization and controllable flight of flapping wing aircraft. A single axis double-crank double-rocker coaxial four-wing flapping mechanism is designed based on a single-crank double-rocker flapping mechanism. According to the kinematics of the mechanism, the symmetry of the flapping angles of the left and right wings during one flapping cycle is analyzed. This flapping mechanism causes a certain instantaneous symmetry difference between the left and right flapping rods. However, the difference is very small, with a maximum value of about 2.7 degrees. The overall flapping angle amplitudes of the left and right sides are 36.6 degrees, and the average flapping angles of the left and right sides are 15 degrees. This flapping mechanism has good overall symmetry and lateral stability. A modular assembly coaxial four-wing flapping wing prototype is developed and flight tests are conducted. The prototype can perform maneuvers such as hand throwing or “hovering” taking-off, forward flight, climbing, turning, and circle flight. Research results provide important design reference for developing flapping wing aircraft with multiple pairs of wings to achieve more concealment and miniaturization.

By adjusting three parameters—bird impact speed, impact position, and engine speed, the transient bird-strike response of the fan blades of the CFM56-7B high-bypass-ratio civil turbofan engine was analyzed by explicit-dynamics simulation. Results showed that for impact speeds of 60, 80, and 130 m/s, the peak Mises stress occurred at about 4.2—4.3 ms. Longer impact duration and smaller axial-relative velocity caused more severe cumulative damage. Impact position changed the peak moment and the stress distribution. At 30%, 50%, and 80% of blade length the stress peaks occurred at about 2.6, 4.1, and 4.2 ms, respectively. An impact at 50% blade length produced stress concentration at the blade root. Engine speed was positively correlated with the damage; peak stress increased markedly at high speeds. Under the conditions of a mass of 1.85 kg, a speed of 130 m/s, and an engine speed of 5175 r/min, the damage from a bird striking the outer casing was much greater than that from a strike on the inlet fairing. This damage presented as circumferential stress concentration and local plastic damage at the blade tip.

Two-stroke piston engines often employ turbocharging coupled with exhaust resonance technology to address air intake challenges at high altitudes. However, for four-cylinder engines, exhaust gas blocking and pressure interference caused by in-phase exhaust in the exhaust resonance system lead to performance degradation at high altitudes. Therefore, a dual resonance tube exhaust system was designed and optimized based on the exhaust pressure wave propagation theory and a genetic algorithm. Three-dimensional fluid dynamics simulations were conducted to verify that the dual resonance tube exhaust system solved the problem of exhaust gas blocking and pressure interference while effectively utilizing pressure waves to optimize the exhaust backpressure distribution. The simulation and optimization results showed that the engine equipped with the dual exhaust resonance system maintained the original power and achieved a slight increase under conditions up to an altitude of 5000 m. Under most high-altitude conditions above 5000 m, the optimized power increased by more than 5% compared with the original engine, and the improvement increased with altitude. At a high altitude of 10 000 m, the maximum power increase reached 34.7%. The study confirmed that the dual resonance tube exhaust system can effectively solve the problems of exhaust gas blocking and pressure interference caused by in-phase exhaust in a four-cylinder engine.

To address the conflict between local adaptability and global consistency in aero-engine component map correction, a dynamic adaptive correction-hybrid optimization model (DAC-HOM) partition-based collaborative correction framework was proposed. The framework employed an embedded DAC-HOM as the unified correction kernel. By coupling this kernel with a Bernstein-basis adaptive surface characterized by dual degrees of freedom in rotational speed and operating-point distribution and a covariance-adaptive bandit-enhanced hippopotamus optimization algorithm (CABE-hippo), the framework enabled adaptive optimization and dynamic boundary adjustment for smooth characteristic surfaces. Building on this, a partition-based collaborative correction strategy for dual-source heterogeneous data was introduced, in which ground test conditions and high-altitude operating conditions represented by quick access recorder (QAR) flight data were corrected separately and then fused to obtain a globally consistent characteristic map. Compared with traditional progressive correction approaches, the proposed method effectively suppressed the propagation of local errors into the global region. Numerical results showed that the framework reduced the overall mean error to 2.61%, and decreased exhaust gas temperature, high-pressure rotor speed and fuel flow errors by 83.6%, 16.5%, and 46.8%, respectively. The proposed framework significantly improved the performance matching accuracy of aero-engine mathematical models at off-design points.

This study employed the transient liquid crystal method to experimentally quantify the heat transfer on the target surface of a complex double-wall structure incorporating impingement hole, pin-fins, and slot. The investigation focuses on the effects of curvature (9° concave, 30° convex, and 75° convex) and impingement Reynolds numbers (10000—40000). The results demonstrate that as curvature increases, internal heat transfer within the double-wall structure is enhanced. At identical impingement Reynolds numbers, the average convective heat transfer coefficient for 75° convex is approximately 20 W/(m²·K) higher than that of 9° concave. For every increase of 10000 in the impingement Reynolds number, the area-averaged convective heat transfer coefficient rises by about 30 W/(m²·K) across all three curved structures. The increase in the average convective heat transfer coefficient slows down only when the impingement Reynolds number rises from 30000 to 40000. The spanwise distribution of the convective heat transfer coefficient exhibits multiple peaks, with secondary peaks surpassing the primary peak at Reynolds number of 40000. Along the flow direction, the convective heat transfer coefficient generally decreases but displays distinct distribution patterns at different longitudinal positions.

When aircraft fly through clouds containing supercooled water droplets, ice can accrete on the surfaces, posing a serious risk to flight safety. Effective anti-icing systems are therefore essential to ensure safe operation under such conditions. The micro-ejector hot-air anti-icing system, as a common anti-icing method for transport aircraft, features low hot air consumption and high anti-icing efficiency. Based on the self-developed NNW-ICE software, the computational methodologies for characterizing the micro-ejector, modeling heat transfer, and estimating pressure drop within the anti-icing channel were developed. These methodologies established a thermal calculation framework for the micro-ejector anti-icing system, yielding key parameters, including the entrainment ratio, temperature of the entrained air, external skin temperature distribution, and runback water distribution for a given geometric configuration. Based on this framework and combined with the genetic algorithm, an optimization design framework for the micro-ejector hot-air anti-icing system was established to maximize the average external skin surface temperature by adjusting the structural design parameters such as the width of the anti-icing unit, the height of the double-skin channel and the internal width of the channel. The optimized design achieved significant performance improvements: the average external surface temperature increased by 3.01 K, and the mass flow rate of runback water decreased by 39.00%. The results showed that the anti-icing optimization design framework based on thermal calculation and genetic algorithm can be applied to the micro-ejector hot-air anti-icing system with satisfactory performance.

In response to the increasingly stringent cooling requirements of turbine disc and mortise structure, an SATES (self-adaptive turbulence eddy simulation) model was employed to conduct a numerical simulation study on the “S”-shaped gap flow channels generated during the assembly of the turbine disk tenon groove and blade tenon head. The differences in flow characteristics between this channel and a conventional rectangular channel under the same operating conditions were compared and analyzed. The variation of the average wall Nusselt number with the inlet Reynolds number and the rotational Reynolds number under transient conditions in the “S”-shaped channel was investigated. Results indicated that the complex flow structure of the “S”-shaped channel led to a significantly stronger heat transfer capability compared with the rectangular channel. Under rotating conditions, the heat transfer coefficients on the two side walls of the “S”-shaped channel exhibited asymmetry. In contrast to the symmetric distribution observed in the rectangular channel, the heat transfer capability of the right-side wall in the “S”-shaped channel was significantly higher than that of the left-side wall. The average wall Nusselt number on the inner walls of both channels increased with the rise of the main flow Reynolds number and the rotational Reynolds number.

In high-altitude environment simulation for aerospace propulsion system test facilities, severe flow-rate shocks, pronounced nonlinearities, unmodeled dynamics, and complex test scenarios pose significant challenges to control performance, thereby motivating the need to overcome the limitations of fixed-parameter ADRC and experience-based trial-and-error tuning. To address this challenge, an efficient parameter self-tuning method for ADRC is proposed, based on temporal and spatial scale transformations, to improve control performance of aeropropulsion systems test facility under complex scenarios. Specifically, this method establishes quantitative mapping relationships between the system’s temporal/spatial scales and the parameters of the three core ADRC modules: the tracking differentiator (TD), the ex-tended state observer (ESO) and the nonlinear state error feedback (NLSEF). The simulation results for each module validate the effectiveness and adaptability of the proposed approach. Finally, the proposed self-tuning approach is applied to the pressure control of aeropropulsion systems under different scenarios. Experimental results show that the proposed method significantly simplifies the ADRC parameter-tuning process and improves the accuracy of parameter acquisition. Compared with the bandwidth-based linear ESO-ADRC, under the proposed parameter-tuning algorithm, the steady-state mean absolute error of the linear ESO-ADRC is reduced from 0.35 to 0.17, the maximum pressure fluctuation during the transient phase is reduced from 0.42 kPa to 0.22 kPa, and the maximum pressure recovery time is shortened from 6.5 s to 4.0 s. Furthermore, the corresponding indices of the nonlinear NESO-ADRC are further improved to 0.12, 0.13 kPa, and 2.7 s, respectively. These results indicate that the proposed method can effectively enhance pressure control performance, thereby ensuring the reliability of aeroengine flight tests.

To accurately predict the leakage characteristics of brush seals, this study developed a dynamic leakage prediction method considering hysteresis effects. The brush pack region was equivalently treated as a porous medium, and a dynamic leakage prediction model for brush seals considering hysteresis effects was established. The calculation methods for key parameters in the model, including the viscous loss coefficient, inertial loss coefficient, and porosity, were clarified, and the accuracy of the model was validated using an ultra-high-speed brush seal test rig. The proposed method was used to investigate the causes of leakage differences during acceleration and deceleration processes, as well as the effects of operating conditions and structural parameters on the hysteresis and leakage characteristics of brush seals. The results showed that the numerical predictions of leakage characteristics considering hysteresis effects agreed well with the experimental results, with a maximum error of 8.6%. Among operating parameters, an increase in inlet pressure increased the leakage and significantly intensified the hysteresis effect. Among structural parameters, increasing the bristle diameter weakened the hysteresis effect but led to higher leakage, whereas decreasing the bristle height and bristle arrangement angle simultaneously weakened the hysteresis effect and reduced the leakage. The findings of this study provide methods and a theoretical basis for the accurate prediction and performance improvement of brush seal leakage characteristics.

Passive fluidic thrust vectoring technology has emerged as a frontier research hotspot in aircraft thrust vector control due to its advantages of requiring no external gas source, simple structure, and low observability. A supersonic axisymmetric fluidic thrust vectoring nozzle with six passive circumferentially uniform secondary flow channels was designed. By controlling the on/off combinations of these secondary flow channels, two control strategies were established. Utilizing schlieren flow visualization and air-bridge balance force measurement techniques, the jet deflection characteristics and thrust vectoring performance of the nozzle within the nozzle pressure ratio (NPR) range of 3.5 to 5.0 were systematically investigated. Experimental results demonstrated that the nozzle successfully achieved vector control in 12 circumferential directions, with maximum thrust vector angles reaching 9.8° in both the primary and secondary control directions. The highest linearity was 89.58% in the primary control direction and 97.28% in the secondary control direction. The maximum thrust vectoring efficiency of the nozzle reached 13.45°/%. This research could provide experimental evidence and design references for further engineering application of passive fluidic thrust vectoring technology in platforms requiring multi-axis control moments, such as missiles and rockets.

Numerical simulations are conducted to investigate the heat release characteristics of an RBCC engine under freestream Mach number 6 conditions. The influence of fuel injection position on flow feature, heat release distribution, combustion mode, and engine performance were analyzed with the aid of logarithmic processing and filter functions. Results showed that moving the injection position upstream shifted the shock train leading edge forward, expanded the high-pressure zone within the flow path, and significantly increased the overall pressure level. It also enlarged the combustion reaction zone, increased the total heat release, and shifted the concentrated heat release region upstream. Furthermore, upstream injection increased the peak heat release rate by reducing the corresponding cross-sectional area, although the shock induced by the fuel strut contributed more significantly to local heat release enhancement via its compression effect. All tested injection positions exhibited dominant diffusion and subsonic combustion modes. Injection upstream of the isolator cavity increased the proportion of premixed combustion, while injection of the central strut sidewall was more conducive to supersonic combustion. Within the studied range, upstream injection effectively improved combustion efficiency, thrust, and specific impulse, with injection upstream of the isolator cavity yielding the optimal engine performance.

Motivated by the urgent need to improve the propulsion motors and generators performance in aviation electrification, thermal management technologies, including loss suppression, passive cooling, and active cooling were reviewed to clarify innovation directions. Their technical characteristics and application effects were analyzed in the context of aviation scenarios. The results indicate that: the core of loss suppression lies in the coordinated optimization of material selection, structural design, and manufacturing processes; passive cooling reduces the internal temperature gradient by enhancing heat conduction paths; and active cooling exhibits a power-dependent trend, where low-to-medium power motors and generators utilize lightweight air-cooling, while medium-to-high and megawatt-scale motors and generators require high-efficiency liquid cooling architectures. Finally, critical future directions for high-power-density aviation propulsion motors and generators were identified: loss suppression via novel soft magnetic materials and precision manufacturing; advanced cooling using microchannels and triply periodic minimal surface (TPMS) structures; and next-generation thermal management exemplified by hydrogen-cooled superconducting systems.

Noise reduction tests of acoustic liners for a short nacelle inlet with a large bypass ratio were conducted on a scaled fan test rig. A sudden drop in noise reduction performance at certain frequencies was observed as rotational speed increased, and the underlying cause was analyzed. This led to the proposal of subsequent optimization directions for acoustic liner design. The study revealed that the noise reduction at twice the blade passing frequency (BPF) of the test liner experienced a sharp decline when the rotational speed increased from 88% to 91%. Modal decomposition analysis indicated that the primary mode associated with twice the BPF scattered into adjacent circumferential dominant modes at higher rotational speeds. By calculating the noise reduction differences of the circumferential modes 35, 36, and 37 across various frequencies, simulation results showed that the noise reduction performance of the 36th circumferential mode decreased at higher frequencies. This implied that as rotational speed increased, the twice the BPF frequency shifted beyond the originally optimized frequency range of the 36th circumferential mode, leading to reduced noise attenuation. In conclusion, the study suggests that in the initial design phase, acoustic liners should not only target the design operating conditions but also comprehensively consider the entire engine operating envelope, with enhanced robustness in noise reduction performance across a wider range of rotational speeds.

Aviation fluid-conveying piping systems are usually subjected to multi-point non-uniform base excitations under multi-clamp support conditions, making it difficult for traditional uniform loading models to accurately describe the actual load environment. To this end, based on the Transfer Matrix Method (TMM), this study conducts dynamic modeling and vibration analysis of pipeline systems subjected to multi-point base excitations, using typical space flow pipelines as the research subject. By introducing clamp constraints and spatial coordinate transformation matrices, the 14-equation model accounting for fluid-structure interaction (FSI) effects was extended to a multi-support spatial flow pipeline system. A method for independently applying base excitations at each clamp constraint location was proposed. By embedding excitations into the chain-solving process of the transfer matrix, it achieved the simulation of multi-point base excitations. Finally, the created model was validated by constructing an experimental system and combining it with finite element simulation. The model's prediction errors for the first six natural frequencies and the first-order resonance response amplitude were within 5.30% and 7.69%, thereby validating the reasonableness of the modeling approach. Based on this, the effects of hydraulic parameters and clamp constraint positions on the piping system were analyzed. The results indicate that fluid pressure dominates the decrease in the system's fundamental frequency, and the placement of clamps near elbows enhances system stiffness, thereby altering the vibration characteristics of the piping system. The relevant results can provide a theoretical basis for vibration reduction design and layout optimization of piping systems.

The rapid development of the aviation industry has exacerbated the environmental problems associated with traditional aviation kerosene. Developing aviation alternative fuel has become a strategic choice to address global climate change, reduce greenhouse gas emissions, ensure energy security, and promote sustainable development of the aviation industry. Three aviation alternative fuels with high application prospects, namely sustainable aviation fuel (SAF), ammonia fuel, and hydrogen fuel, were reviewed. The basic physicochemical properties, preparation pathways, and emission reduction benefits of the fuels were systematically sorted out. The application of Sustainable Aviation Fuel (SAF) in commercial flights reached a 50% mixing ratio. However, ammonia combustion faced organization challenges, and hydrogen was constrained by storage and transport. This study comprehensively analyzed the practical constraints of these three fuels regarding economy, technology maturity, infrastructure adaptability, and supply chain development. The following development priorities of aviation power were proposed: prioritizing the integration and application of SAF with existing aviation equipment, strengthening the technological research and development of hydrogen and ammonia fuels in combustion performance and storage and transportation safety, and steadily expanding their large-scale application in commercial flights. It is suggested to strengthen policy guidance, deepen cross industry collaborative innovation, and promote the integration and evolution of fuel and power technologies. This will help gradually build a green aviation energy system covering raw material acquisition, preparation and transformation, storage and transportation support, and terminal applications, thereby achieving low-carbon and high-quality development of the aviation industry.

Rapid short-range detonation initiation is key to achieving efficient ignition in fuels of different phases. This paper provides a detailed review of the latest research progress on various ignition methods and the ignition performance of fuels in different phases. It analyzes the advantages and limitations of ignition methods such as detonation tubes, shock wave focusing, laser ignition, and indirect ignition. Furthermore, it notes that in different fuel systems, gaseous fuels are highly reactive and easy to ignite, while liquid fuels possess high energy density but are difficult to ignite. Currently, significant achievements have been made in both experimental and numerical simulations of detonation waves; however, challenges remain, including difficult flame control, structural design requiring optimization, and insufficient adaptability to operating conditions. Future efforts should focus on in-depth investigation of multi-physics coupling mechanisms, the development of novel obstacles, optimization of ignition, and the expansion of their applications in new propulsion systems.

The impact of thermal loads on the dynamic characteristics of a combined-support rotor system in aero-engines was investigated. A temperature-dependent model was developed for the combined-support system, deriving the formulas for the stiffness and damping of elastic rings and squirrel cage elastic supports under varying temperatures. A 3D finite element model of the power turbine rotor was developed, incorporating thermal deformation and stress analysis based on temperature field distribution. A bidirectional coupled model integrating the combined support and rotor was proposed, enabling further analysis of thermal-induced changes in rotor dynamics. Results indicated that thermal loads caused a minor reduction in critical speeds, while vibration amplitudes exhibited significant variations only near critical speed regions. Experimental validation on a rotor test rig confirmed the theoretical analysis.

To gain in-depth insights into the dynamic evolution law of the tip clearance leakage vortex structure in a compressor and its impact on performance under gas-solid two-phase conditions, numerical simulations were carried out on the Rotor37 transonic compressor under both single-phase and gas-solid two-phase operating conditions using Fluent software, based on the Detached Eddy Simulation DES (DES)-Discrete Phase Model (DPM) coupling model. The Omega vortex identification criterion was adopted to identify the tip clearance leakage vortex structure, and the entropy production rate was utilized as the characteristic parameter for quantifying the compressor’s flow losses. Results demonstrated that, compared with the single-phase condition, the supersonic region in the flow passage at the 90% blade height section of the compressor blade was significantly narrowed under the gas-solid two-phase condition, the shock wave position shifted upstream, and a deceleration transition zone appeared at the leading edge of the shock wave. Moreover, identification of the tip leakage vortex structures under the two operating conditions via the Omega criterion revealed that the particle phase significantly promoted the shedding and reorganization of the vortex structures. Additionally, observations of the variations in entropy production rate at different positions on the compressor blades showed that the presence of the particle phase not only significantly increased the compressor’s flow losses, but also reduced the compressor’s pressurization capacity and isentropic efficiency.

To address the limitations of the traditional Reynolds equation method in terms of model accuracy and generalizability, a semi-analytical method based on the fundamental equations of fluid mechanics was proposed for accurate calculation of the stiffness and damping coefficients of gas foil bearings. First, the fluid-structure interaction between the gas film and the foil structure was established. Then, the eccentricity ratio was continuously varied in order to locate the static equilibrium position of the journal at a determined rotational speed. Subsequently, a first-order eccentricity perturbation was applied to obtain the four stiffness coefficients via steady-state calculations. Finally, the four damping coefficients were determined through transient calculations by imposing two simple harmonic vibrations on the journal. The results showed that as the rotational speed increased, the static equilibrium position of the journal gradually moved closer to the bearing center, the direct stiffness coefficients decreased, the cross-coupled stiffness coefficients increased, and all four damping coefficients decreased. The calculated stiffness and damping coefficients can be used in subsequent rotor dynamics analyses. The stiffness coefficients of the foil bearing were tested experimentally, and the experimental results demonstrated the feasibility of the proposed calculation method.

To ensure stable operation of the low-pressure turbine under low-Reynolds-number conditions, numerical simulation and experimental methods were adopted to analyze the operating characteristics of the low-pressure turbine cascade with respect to influencing factors including Reynolds number, load coefficient, incoming flow reduced frequency, and turbulence intensity, and further carry out optimal design adapted to the low-Reynolds-number working environment. The results showed that with the decrease in Reynolds number, the influence of reduced frequency increased; when the Reynolds number was 1.0×10⁴, the reduced frequency increased from 1.03 to 3.09, leading to a 0.9% reduction in efficiency and a 6.0% increase in loss. The optimized cascade scheme reduced the adverse pressure gradient downstream of the throat on the blade suction surface, effectively suppressed the separated flow under low-Reynolds-number conditions, and thus significantly improved the performance of the turbine cascade. Under low-Reynolds-number conditions, there was no airflow separation on the suction surface of the optimized cascade, and the pressure and energy losses were reduced by 30%—40%; no improvement was observed under high-Reynolds-number conditions. Within the range of incoming flow angle of attack from −20° to 10°, the pressure and energy losses increased by 5%—25%, while the losses remained nearly the same when the angle of attack was between −10° and 0°.

Based on the aircraft mission profile and maneuver load envelope, a transformation correlation between aircraft maneuver loads and engine maneuver loads was proposed according to the aircraft motion states, enabling quantitative prediction of blade tip clearance during maneuver flight using a full-aircraft finite element model. The axial and circumferential distribution characteristics of tip clearance variations across different engine stages were calculated and analyzed, identifying critical engine sections. Results showed that under typical maneuver conditions, the most significant tip clearance variations occurred at the high-pressure turbine and low-pressure turbine, with changes of 0.22 mm and 0.20 mm, respectively. These variations were substantially smaller than those induced by thermal and centrifugal loads. The radial tip clearance change exhibited a linear relationship with linear acceleration and gyroscopic moment, demonstrating a quadratic relationship with additional centrifugal loads. Compared with counter-rotating dual-rotor systems, the influence of gyroscopic moments on the tip clearance of each stage in co-rotating dual-rotor systems was smaller.

In order to investigate the effects of total pressure inlet distortion on the performance and stability of a counter-rotating fan, numerical simulations were conducted on the aerodynamic performance and flow field characteristics of a two-stage counter-rotating lift fan under both uniform inflow and total pressure distorted inflow conditions behind a baffle. By combining steady and unsteady calculations, a comparative analysis was performed to explore the Instability mechanism of the counter-rotating fan under total pressure distorted inflow. The results indicate that under distorted inflow conditions, the internal flow of the fan exhibits stronger unsteadiness, leading to a flow rate shift and variation in the characteristic curves obtained from steady and unsteady calculations. Therefore, the stall condition of the fan under distorted inflow should be determined based on the results from unsteady calculations. Compared with uniform inflow conditions, the distorted inflow lowers the stall boundary mass flow rate from 175.82 kg/s to 162.69 kg/s, a reduction of 7.47%, and decreases the stall margin by 9.07%. The operational range within which the counter-rotating fan can stably operate is significantly reduced, and its ability to suppress stall disturbances has decreased. Before the onset of instability, as the total pressure distortion behind the baffle propagates downstream the circumferential extent of the distorted region remains largely unchanged. Consistent with uniform inflow conditions, the instability of the fan under total pressure distorted inflow behind the baffle first occurs in the tip region of the second-stage rotor R2. After blocking the R2 passage, it propagates upstream and affects the tip region of the first-stage rotor R1, leading to global instability.

In order to quantify the discrepancies between the characteristics of the three-dimensional (3D) turbulent flow field and the two-dimensional (2D) observational results within a constant-volume combustion chamber (CVCC), a numerical study was conducted using the large-eddy simulation (LES) method based on a sliding mesh technique. Simulations were performed for the 3D transient flow fields under various combinations of initial pressure and fan rotational speed. Key parameters in the flame development region, including velocity distribution, turbulence intensity, turbulence integral length scale, flow uniformity, and isotropy, were analyzed and compared with experimental data for validation. The objective of this research is to reveal the underlying mechanisms responsible for the characteristic differences between 2D and 3D turbulent flow fields through quantitative comparison, thereby providing theoretical support for enhancing turbulence control and optimizing combustion stability within CVCCs. The results indicate significant differences between the fully 3D simulated field and the 2D projection typically available from experiments. Compared to the experimentally monitored plane, the maximum relative error in the fitted formula coefficients for turbulence intensity reached 17% on the two orthogonal monitoring planes. Similarly, the maximum relative error for the turbulence length scale coefficients reached 5%. Furthermore, the ratio of the meridional to circumferential length scales deviated substantially from the theoretical value of 2 for isotropic turbulence, indicating a worse degree of isotropy compared to the experimental plane. These significant discrepancies confirm that traditional flame observation methods based on 2D projections introduce non-negligible measurement errors, highlighting the critical importance of 3D flow field analysis in combustion diagnostics.

Inspired by the flow-control capability of corrugated structures on dragonfly wings, a streamwise-arranged non-smooth surface was proposed and applied to a highly loaded diffuser cascade. Using a numerically validated simulation approach, the effects of groove location and structural parameters on the aerodynamic performance of the cascade were systematically investigated, and the underlying mechanisms were elucidated from both macroscopic flow features and near-wall flow physics. The results indicated that when the grooves were arranged within 75%—100% of axial chord length on the suction surface with an appropriate parameter combination, stable trapped vortices were formed inside the groove cavities. Through periodic interaction with the main flow, these trapped vortices introduced small-scale, amplitude-limited disturbances in the near-wall region, leading to an increase in the turbulence intermittency factor without triggering high-energy large-scale turbulent structures. Meanwhile, the local retention and redistribution of reversed-flow momentum promoted a transition of the separated region from a highly unstable, high-energy state to a more moderate and controlled turbulent state. Consequently, the separation vortex near the trailing edge on the suction side was significantly weakened, passage blockage was alleviated, and the development of the suction-side boundary layer was effectively improved. Based on the double design point loss criterion, the total pressure loss was reduced by up to 8.66% within the usable incidence angle range, and the upper boundary of the usable incidence angle was extended by approximately 0.7° toward higher values.

Facing the challenge of endoscopic particle image velocimetry (PIV) measurements within the confined spaces of aero-engine compressors and fans, a flexible endoscopic PIV method was proposed. This approach employed a flexible fiber-optic image bundle as the core image transmission component, and integrated with a self-developed endoscopic lens. It effectively overcame the limitations of traditional rigid endoscopes in vibration-prone environments, such as restricted field of view, susceptibility to motion blur, and insufficient installation reliability. Based on the test constraints of the internal flow field in a certain type of aero-engine compressor, a self-developed endoscopic imaging lens suitable for confined space detection was designed. This lens was integrated with a fiber-optic image bundle and an interframe camera to construct an endoscopic flow field acquisition system. Calibration experiments were conducted to quantify the optical loss and optical non-uniformity introduced by the fiber bundle. A pixel-by-pixel compensation method was proposed to significantly enhance the image quality of the flexible endoscopic system. Experimental validation demonstrated that this flexible endoscopic PIV system successfully captured the intended turbulent jet velocity distribution. Compared with traditional PIV measurements, the results presented an average relative error controlled within 3%, effectively verifying the system's accuracy. This method could provide a technical approach for refined measurements of complex flow fields within aero-engines.

Achieving stable and self-sustained rotating detonation under low-oxygen intake conditions is crucial for the design of rotating detonation afterburners. Through two-dimensional numerical simulations of non-premixed ethylene/air rotating detonation flow fields, this study investigated the influence of intake temperature on the critical lower limit of oxygen concentration required for rotating detonation initiation under fuel-lean conditions. The results showed that as the intake temperature of the afterburner increased from 875 K to 1025 K, the lower limit of oxygen mass fraction for stable propagation of the rotating detonation shifted from 15% to 14%. When the oxygen concentration approached the critical lower limit for initiation, increasing the intake temperature had a minor effect on the average heat release rate in the combustor flow field. However, higher intake temperatures can enhanced deflagration within the flow field, enabling self-sustained propagation of the detonation wave even when the velocity deficit exceeded 8%, thereby broadening the stable initiation boundary. Furthermore, increased intake temperature induced the emergence of complex multi-wave modes in the flow field, while reduced oxygen concentration tended to drive the rotating detonation toward propagation in a single-wave mode. Further analysis indicated that multi-wave modes significantly affected the total pressure loss and specific thrust of the afterburner, while mixed modes dominated variations in the temperature rise ratio and combustion efficiency. Increasing the intake temperature generally degrades the performance parameters of the rotating detonation afterburner; nevertheless, under appropriate intake temperature and oxygen concentration conditions, rotating detonation after burning can achieve lower total pressure loss.

Gears are crucial mechanical components to transmit power and torque, playing an irreplaceable role in aero-engines while continuously evolving towards higher speeds, heavier loads, and lighter weights. However, their dynamic performance faces substantial challenges. The issues of nodal-diameter resonance failure and modal jumping that occur within a wide operational speed range in the design optimization of an aero-engine bevel gear was addressed. A method based on the Campbell diagram was proposed to identify dangerous nodal-diameter vibration modes and optimize the structure of the gear to avoid resonance across a wide speed range. The method transformed traditional post-event, manual, experience-based judgments into a preemptive, automated, mathematically feature-driven optimization indicator, enabling separation and control of nodal-diameter vibration modes in gear vibration optimization. First, modal frequencies were distinguished by the Modal Assurance Criterion, ensuring an accurate fit to the Campbell diagram. The slopes of the frequency lines in the Campbell diagram were extracted, and based on these slopes, the presence of nodal-diameter vibration modes was identified. Their resonance frequencies were then predicted and recorded. Next, an optimization model was developed based on the dangerous nodal-diameter vibration mode identification method, to avoid resonance over a wide speed range. This aimed to minimize the gear mass while constraining the stress and the nodal-diameter resonance frequencies. The resonance-avoidance optimization process was constructed by assigning large values and the Pointer strategy. After optimization, the gear successfully avoided nodal-diameter resonance within the 75%—107% operational speed range, and its mass was reduced by 6.566%. This confirmed the effectiveness of the proposed gear nodal-diameter identification and resonance avoidance optimization method in engineering applications, offering significant support for the structural enhancement of aero-engine bevel gears.

In order to meet the thermal reduction requirements of high Mach number compression corners, numerical simulations were carried out to study the effects of gas transpiration on the flow structure and aerothermal environment of a 15° compression corner at a free stream Mach number of 11.63. The results showed that gas transpiration significantly increased the boundary layer thickness in both its active and downstream regions, reduced the velocity gradient within the boundary layer, and augmented the local pressure. When the transpiration location was positioned upstream or within the separation zone, gas transpiration led to a significant enlargement of the separation region. In contrast, when the transpiration location was positioned downstream the separation zone, gas transpiration had a negligible influence on the separation characteristics. Variations in the flow structure were found to significantly influence the wall heat flux distribution. The thickening of the boundary layer reduced the temperature gradient within it, thereby lowering the wall heat flux in both the transpiration region and the downstream region. To achieve optimal thermal reduction, transpiration cooling should cover the high heat flux region near the peak heat flux, while its effect remains confined to the downstream region to prevent the expansion of the separation zone, thereby stabilizing the peak heat flux location and effectively reducing the peak heat flux.

Turbine vane endwall with large surface area is commonly protected by zoned, multi-row film cooling arrays; however, the strong spatial variation of mainstream flow and entrainment by passage vortices within the turbine cascade flow field frequently impede uniform film coverage. Expansion-shaped film cooling holes by modulating exit momentum can promote coolant adherence to the surface. By combining numerical simulations and pressure-sensitive paint (PSP) experiments, the film cooling characteristics of shaped holes in different endwall regions were investigated, and the endwall film cooling effectiveness of a baseline cylindrical hole configuration and a combined leading-edge slot hole–passage teardrop-shaped hole configuration under varying density ratios and blowing ratios was quantified. Results showed that, at high blowing ratios, the elevated coolant momentum induced blow-off at the endwall leading edge for both configurations. The leading-edge slot holes substantially attenuated passage-vortex strength, while the passage teardrop-shaped holes demonstrated enhanced resilience against disruption by pressure side horseshoe-vortex branches; as a result, the combined configuration produced maximum increases in endwall film cooling effectiveness of 80.18% at density ratio of 1.4 and 66.16% at density ratio of 3.0. With blowing ratio increasing, endwall film cooling effectiveness exhibited a rise-then-fall behavior: cooling degradation was observed at density ratio of 1.4 when blowing ratio exceeded 3.0 and at density ratio of 3.0 when blowing ratio exceeded 4.0. The blowing ratio corresponding to peak cooling effectiveness was influenced by density ratio and the local endwall pressure gradient.