The current research status of blind cavity bolt tightening technology for aircraft engine rotors was summarized. Guided by the technical requirements for tightening bolts in rotor blind cavities, three aspects were discussed: tightening process, tightening equipment, and tightening inspection. Firstly, focusing on the two stages of rotor rabbet fit and bolt connection, optimized processes such as rabbet press-fitting process, torque-rotation angle method, and multi-axis tightening method were summarized. Secondly, focusing on equipment configuration design and accuracy control methods, research results such as torque transmission methods of tightening equipment, mechanism deployment methods, overall configuration development, as well as bolt positioning and tightening accuracy control methods were summarized. Finally, focusing on bolt tightening inspection methods, current detection technologies for tightening torque, rotation angle, and pre-tightening force were summarized. The research results indicate that to fully overcome the limitations of the narrow space in rotor blind cavities and enhance rotor assembly quality, the blind cavity bolt tightening technology will develop towards an integrated process of press-fitting, tightening, and measurement, intelligent robotic tightening, and multi-modal fusion measurement.

To investigate the ignition, extinction and flame crossover characteristics of a newly designed dual-vortex-controlled radial staged combustor, an experimental study was conducted to examine the ignition and flame propagation mechanisms in the concave cavity and mainstream zone under different operating modes, as well as stable combustion patterns and the ignition/extinction characteristics of the combustor under various operating conditions. The results indicated that the combustor could be successfully ignited when fuel was independently supplied to either the outer-ring concave cavity or the mainstream zone, flame propagation and stable combustion were also achievable between the outer-ring concave cavity and the mainstream zone. Under both flame propagation modes, the stabilized combustion flame exhibited consistent pattern, and distributed within the concave cavity, its downstream region and the downstream of the swirl cup. The outer-ring concave cavity achieved a minimum fuel-air ratio of 0.0094 for ignition and 0.0034 for extinction under room temperature and atmospheric pressure conditions, demonstrating significantly superior ignition/extinction performance compared with the mainstream zone.

Rotating seal ring are used in aero engines to block low-temperature bleed air in the bearing cavity from high-temperature bleed air in the turbine. In order to obtain the influence of rotation speed and inlet and outlet pressure ratio on the sealing effect, an experimental system with high pressure ratio and high speed was built. Through the measurement of multi-section parameters, combined with the comparative experiments of sealing clearance and contact surface lubrication, the evolution law of leakage characteristics was analyzed. The results show that the leakage is reduced by 36.4% when the rotation speed is increased from 0 to 10000 r/min (pressure ratio 2.5), and the leakage increases by 91.8% when the pressure ratio is increased from 1.1 to 2.5 (10000 r/min). The assembly clearance affects less than 5% within the 0.5mm tolerance, and the addition of lubricating coatings reduces leakage by 15-60%. Numerical simulations show that the average contact stress increases by up to 1.24 MPa under the experimental conditions, which improves the sealing effect, and the increase of friction coefficient reduces the average contact stress by more than 1.35 MPa, which is not conducive to sealing. The quantitative relationship between rotational speed and seal performance and the efficiency mechanism of lubrication are revealed, which provides a basis for the design of rotary seals for aero engines.

To investigate the mixing characteristics of fuel and supersonic inflow, a numerical simulation study was conducted based on the annular-slot-orifice injection structure in the combustion chamber of an open-barrel-type rotating detonation ramjet engine. The study explored the effects of fuel injection angle, injection method, and axial position of the injection ring on the mixing characteristics of fuel and supersonic inflow. The results showed that increasing the fuel injection angle improves the uniformity of fuel mixing, reduces total pressure loss, and leads to more intense collisions between transverse jets. When the fuel is injected in a non-colliding manner, the mixing uniformity increases, and the collision area between the supersonic inflow and the fuel enlarges. Moving the injection ring upstream enhances the uniformity of fuel mixing, and the mixing distance between the transverse jet and the supersonic inflow becomes longer. Analysis of the flow field structure revealed that the supersonic inflow and ethylene transverse jets undergo cylindrical wake flow within the expansion section, forming a fan-shaped continuous expansion region extending downstream. Upon entering the combustion chamber, the primary mixing flow forms a series of fragmented small vortex structures, thereby promoting the mixing effect of ethylene and air.

A single-slotted Fowler flap high-lift device for a hybrid wing-body (HWB) configured amphibious aircraft was designed, incorporating blowing/suction zones on the upper surface of the main wing. It systematically validated the lift-enhancement effectiveness of the co-flow jet (CFJ) technology on this novel aerodynamic configuration and elucidated the mechanisms behind its aerodynamic performance improvements. Numerical simulation results demonstrated that: when the flap deflection angle reached 40°, the maximum lift coefficient of the high-lift configuration attained 2.7; after implementing CFJ, the flow velocity over the upper wing surface increased, accompanied by a significant rise in negative pressure. The initial flow separation on the wing surface was effectively controlled, with the maximum lift coefficient increasing to 3.2, stall angle of attack extended by 4°, drag coefficient reduced by over 20%, and available range of pitching moment expanded. Additionally, CFJ enhanced the surface flow velocity on the flaps and suppressed the potential local separation under high angle-of-attack conditions, increasing the leading-edge negative pressure peak by up to 30% and significantly improving the flap lift-enhancement efficiency.

To address the significant contradiction between the optimal blade shapes of tiltrotor aircraft in hover and cruise modes, a mission scenario-driven aerodynamic optimization design method for tiltrotor blades was developed. This approach comprehensively considered the total energy consumption across different mission scenarios to ensure overall improvements in flight performance and efficiency. By incorporating typical mission phases including vertical takeoff, cruise, and hover and their respective durations, the method transformed the multi-point multi-objective optimization problem into a single-objective optimization problem with the total energy consumption of the entire mission scenario as the objective. The shape differences of the optimized blades under cruise and hover-focused mission scenarios were analyzed. Results indicated that the developed method achieved high aerodynamic performance of tiltrotor blades across multiple operating conditions, with optimization time reduced by 69.67% compared with multi-objective optimization.

For the cooling design for the complex and compact structure of the swirl-cup combustor head, three cooling design schemes were proposed. Numerical simulations were conducted to calculate the flow field, temperature field, and wall temperature distribution for different schemes. The results showed that, without cooling, corner vortices near the wall in the head led to the formation of high-temperature stagnation vortices by drawing in burning gases, causing a wide range of high temperatures on the surface of the flame tube head with a maximum wall temperature of 1178 K. The head cooling designs involved introducing cooling air between the inner and outer fairings to form a protective air film and disperse the near-wall corner vortices, effectively reducing the head wall temperature to varying degrees. The use of inclined multi-hole in the thermal shield was limited by geometrical dimensions, resulting in uneven development and poor adhesion of the air film, and leading to ineffective cooling. By designing a guide ring structure, it is possible to control the diversion of cooling in a compact space and to adjust the flow path design to increase the cooling airflow rate and velocity, forming a uniform air film adhering to the wall at the thermal shield and the swirler housing, with a wall temperature below 870 K. Under high temperature and high pressure conditions, the optimized scheme was validated through combustor component tests. The results demonstrated that the combustor head wall temperature did not exceed the material's long-term permissible operating temperature, proving that the head cooling design could effectively control the wall temperature.

An experimental system was built to study the effects of RP-3 and RP-5 fuels at different fuel temperatures on the flow rate characteristics, cone angle, Sauter mean diameter (SMD) and droplet size distribution. The results showed that under the same fuel pressure, as the fuel temperature decreased from 20 ℃ to −30 ℃, the density, kinematic viscosity, and surface tension of RP-3 and RP-5 fuels increased, and the flow rate also increased. The flow rate of RP-5 fuel was 2% to 4% higher than that of RP-3; the cone angle decreased by 14% for RP-3 fuel and 23% for RP-5 fuel; as SMD increased, the droplet size distribution index first decreased and then increased, indicating that the decrease in fuel temperature led to an increase in the average droplet size, but the unevenness of droplet size decreased and the droplet size distribution became more concentrated.

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.

To address the thermal protection requirements of combustor liners under high-temperature and elevated heat loads, the single-tube combustor model incorporating tangential effusion cooling technology was developed. A conjugate heat transfer analysis was employed to numerically investigate the influence of varying inlet temperature, inlet pressure, and fuel-air ratio on liner wall temperature and cooling performance. The results show that the higher inlet temperatures reduced cooling film coverage effectiveness, weakened heat transfer intensity within effusion holes and on the outer liner wall, decreased average integrated cooling efficiency, and increased both wall temperature and thermal gradients. Under the condition of this paper, variations in inlet pressure exhibited minimal influence on near-wall flow and heat transfer, with negligible differences in effusion cooling effectiveness. Under extreme conditions that the proportion of cooling air is only about 19%, fuel-air ratio of 0.042, and combustor temperature rise approaching 1300 K), the liner maintained a high average integrated cooling efficiency of 70%—90%, demonstrating the significant potential of optimized tangential effusion cooling designs for next-generation ultra-high-temperature combustors.

The complex flow field around aircraft engines during carrier landing and boltering was analyzed by establishing a full-scale numerical model of both carrier and aircraft. The overset grid and moving pressure measurement techniques combined with delayed detached-eddy simulation (DDES) were developed to resolve the complex vortex systems generated by the advancing carrier hull. The effects of wind speed, wind direction angle, and angle of attack on the total pressure distortion at the engine inlet were analyzed, and the dynamic evolution characteristics of the total pressure distortion were revealed. Key findings indicated that under headwind conditions, counter-rotating vortex pairs formed on both sides of the carrier hull, while crosswinds caused significantly asymmetric flow separation, with turbulence intensity enhanced on the leeward side. With the increase of wind angle, the total pressure distortion index exhibited decreased mean values but elevated standard deviations. Specifically at 90° yaw angle (relative to 0°), the mean value dropped from 0.0649 to 0.0498 (23.3% reduction), while the standard deviation surged from 1.04×10⁻³ to 4.97×10⁻³ (377.9% increase). Particularly during touchdown, the distortion transients became especially prominent due to ground effects and pitch adjustment.

The intercooled recuperated turbofan engine has more health parameters than the ordinary three-shaft turbofan engine, which leads to a larger calculation amount of fault diagnosis. To enhance the speed of gas path fault diagnosis of intercooled recuperated turbofan engine under multi-component degradation, a fast gas path fault diagnosis method for intercooled recuperated turbofan engine was developed. Based on the traditional nested iterative diagnosis architecture, a non-nested iterative architecture was designed, which considered the degradation of 6 rotating components and 2 heat exchanger components. By dividing the components, the engine component matching and fault diagnosis could be completed in the same Newton-Raphson iterative algorithm. The sensor selection optimization was carried out to improve the ability of measuring parameters to identify the degradation of different components. The results of simulation test on all conditions of engine life cycle showed that the maximum diagnostic error of this method was 0.0036%, the average error was 0.0008%, and the average calculation time was less than 0.063 seconds. Compared with the traditional diagnostic architecture, this non-nested architecture can significantly reduce the number of sub-component calls and fault diagnosis time while ensuring the diagnostic accuracy. It provides the method and theoretical support for overcoming the conflict between high-precision fault diagnosis and rapid diagnosis of multi-component faults.

To investigate the influence of main stage swirl number on NOx emissions in a centrally-staged lean premixed prevaporized combustor across wide operating conditions, three-dimensional numerical simulations were used to analyze the turbulent combustion and NO formation characteristics during the reference emissions Landing and Take-Off (LTO) cycle. And the simulation methodology was validated against experimental NOx emission data. Furthermore, the influence of main stage swirl number on NOx emissions under LTO cycle conditions was simulated and analyzed. The results revealed that the influence of main swirl number on NOx emissions varied under different operating conditions. At idle conditions, an increase in main swirl number enhanced fuel-air mixing in the pilot combustion zone, leading to higher combustion temperature and elevated NOx emissions. Under high-load conditions, as the main swirl number increased, additional airflow was entrained into the re-circulation zone, further reducing the fuel-to-air ratio in the lean primary combustion zone and thereby decreasing NOx emissions. With increasing main swirl number over the range of 0.5 to 0.9, total NOx emissions during the reference emissions landing and takeoff cycle decreased.

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.

To investigate the effect of crossflow within turbine blades on film cooling characteristics, a numerical study was conducted. An in-house GCFD code was employed to simulate the impact of the non-uniform variation in coolant flow across a limited channel on film cooling characteristics. The research focused on a 15-hole cylindrical film cooling flat plate, where the coolant flow channel is perpendicular to the mainstream flow channel in space. The coolant flow area equals the total cross-sectional area of all film cooling holes. Results show that the non-uniform crossflow and downstream blockage apply different tangential momentum proportions to the entrance of each film hole, causing changes in the flow structure of the holes. The interaction between holes further leads to different characteristics of film coverage, including windward and leeward deviations, as well as a symmetric distribution. As the blowing ratio increases from 0.2 to 1.5, the non-linear variation of parameters intensifies. The ratio of the maximum to minimum discharge coefficient decreases, reaching a minimum value of 1.04. The ratio of the maximum to minimum space-averaged cooling effectiveness increases, with a maximum value of 3.06.

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.

Infrared detector technologies pose a serious threat to the survivability of modern military aircraft with wide use of turbofan engines. Based on the air tab mixing enhancement technique with active flow control, experimental and numerical study was carried out on the application and mixing efficiency of air tab in multi-channel turbofan nozzle, and explored the use and control of mixing to effectively reduce the temperature of the jet to inhibit infrared radiation and meet the demand of the aircraft stealth performance. Experimental results of the model of mixing enhancement with air tab were analyzed, and verified the feasibility of the study method by comparing the experimental and numerical results. The verified numerical results show that the air tab can improve the thermal mixing efficiency of the nozzle from about 0.17 to over 0.3, and the high temperature jet core area downstream of the nozzle is reduced by more than 30%. Well mixing performance is achieved with almost no total pressure and thrust loss from mixing.

A synergistic design method combining geometric configuration and flow control was proposed to enhance the low-Mach-number self-starting capability and mitigate vortex region in hypersonic inward-turning inlets. The method combined geometric reconstruction of “side wall segmentation, forward sweep, and outward translation” with a controlled side overflow strategy, significantly reducing the effective internal contraction ratio and achieving coordinated suppression and removal of separated flow and boundary layer. Three-dimensional numerical results show that, compared with the conventional rear-cut-lip configuration, the novel inlet achieves significant performance improvements over a wide Mach number range. The self-starting Mach number decreases from 3.9 to 3.2 and the starting Mach number decreases from 3.3 to 3.1. When the inflow Mach number is 4.0 and 6.0, the actual captured mass flow increases by 9.74% and 6.54%, respectively, and the exit total pressure recovery coefficient increases by 9.91% and 4.45%. Additionally, the exit distortion index is reduced by at least 13%, and at an inflow Mach number of 4.0, the vortex region is essentially eliminated.

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.

To investigate the load on the pintle in a coaxial thrust variable solid rocket motor, this study numerically analyzes the force acting on pintles with different profiles under various typical operating conditions. The results show that during takeoff stage, the maximum load on pintles matches the engine thrust, with no significant differences observed between profiles. In cruise and final acceleration phases of the sustenance stage, concave-profile pintles bear 63% and 51% of the load compared to convex-profile counterparts, while conical-profile pintles experience intermediate loads. When actively adjusting pintle positions, the load increases linearly with rising combustion chamber pressure. Pressure fluctuations within the combustion chamber also cause proportional linear variations in pintle loads. The force on the pintle can be significantly reduced by optimizing its shape. After adding a groove to the tail of the convex surface pintle, the maximum load can be reduced by 95%.

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°.

Due to the strong nonlinearity of shroud contact interfaces and enormous computational cost associated with multimodal damping analysis, conventional methods fail to achieve comprehensive, efficient, and accurate solutions. To address this challenge, an improved whale optimization algorithm-extreme learning machine (IWOA-ELM) model was proposed for multimodal damping characteristic prediction, enabling fast and highly accurate mapping from extremely limited inputs to massive outputs under very short training times and small datasets. A lightweight training dataset was constructed using the energy method combined with finite-element modal analysis. An improved WOA (IWOA) was developed by incorporating a history-memory-based group collaborative strategy and Sobol reverse initialization, and its superior global optimization capability was verified using the CEC2017 benchmark suite. The IWOA was further employed to optimize the weights and biases of the ELM, yielding a powerful network capable of realizing extremely low-input and ultra-high-output prediction. Experimental validation was conducted on a shrouded bladed-disk structure. The input consisted of damping ratios at five stress points under the first bending mode, while the output corresponded to the complete damping characteristics (2000 damping values) of the first 20 vibration modes over a stress range of 0—100 MPa. The results showed that the proposed IWOA-ELM achieved a mean squared error of 3.74×10⁻⁶, which was reduced by 87% compared with the conventional ELM, demonstrating its outstanding prediction accuracy. Moreover, the prediction time was only 0.51 s, improving computational efficiency by nearly 2000 times, which enabled rapid evaluation of large-scale multimodal shroud damping characteristics. The proposed IWOA-ELM damping prediction model made it possible for fast and comprehensive consideration of multimodal vibration-reduction performance, effectively reducing resonance risks and exhibiting strong potential for practical engineering applications.

Taking a classic transonic turbine blade as the research object, this study adopts numerical simulation methods to investigate the influence laws of five tip clearance heights (0.23, 0.46, 0.69, 0.92 mm, and 1.38 mm) on the flow and heat transfer characteristics as well as aerodynamic performance of the turbine rotor tip and its adjacent regions. The results show that with the increase in tip clearance, the scope of the supersonic region on the blade tip expands toward the leading edge, the flow separation point shifts rearward, the interference effect between shock waves and expansion waves is enhanced, and the complexity of the flow field is aggravated. The area-averaged Nusselt Number (Nu) at a clearance height of 1.38 mm is increased by 9.9% compared with that at a clearance height of 0.23 mm. Both the strength and scale of the tip leakage vortex and the horseshoe vortex are enhanced with the increase in clearance, and their interaction is intensified. However, the area-averaged total pressure loss coefficient shows a decreasing trend with the increase in clearance, and the loss at a clearance height of 1.38 mm is reduced by 1.993% compared with that at a clearance height of 0.23 mm, which indicates that the strong confinement and concentrated mixing of the leakage flow in small clearances are the main causes of higher aerodynamic losses. In addition, the absolute leakage mass flow increases significantly with the increase in clearance, while the leakage rate per unit clearance height decreases accordingly.

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.

A decoupled optimization method is proposed for underplatform damper (UPD) design to address the lack of multi-objective mechanism insight and the high cost of iterative blade–damper modeling. We develop a parameterized equivalent beam model of the blade–UPD system that accurately reproduces the first bending-mode dynamics and the nonlinear forced response, while keeping the model size and computational cost modest. We characterize the influence of key UPD design parameters on vibration attenuation and introduce three dimensionless metrics—reduction degree, domain width, and coincidence degree—to quantify damper performance. The results show that the contact angle primarily sets the attenuation “depth” and applicability “width,” whereas the damper mass tunes the operating domain’s overlap with engine flow-field excitation. Based on these insights, we adopt a decoupled strategy: first optimize geometric parameters using a Pareto multi-objective genetic algorithm, then set the damper mass according to the excitation characteristics. For a non-symmetric UPD case, the optimized design increases the reduction degree by 13%, the domain width by 38%, and the coincidence degree by 114.6%, yielding a 43.8% decrease in the first-bending resonance amplitude at the target operating condition.

A hybrid electrically controlled rotor based on macro-fiber composite (MFC) was proposed. To explore the feasibility and control characteristics of this hybrid ECR, an aeroelastic dynamics analysis model was first established based on moderate deflection beam theory combined with the piezoelectric constitutive equations of MFC. Subsequently, using this analytical model and focusing on the MFC model rotor, simulation studies were conducted to investigate the influences of key control and design parameters on the rotor’s control response. With a design target of achieving a cyclic pitch range of ±12°, the feasibility of utilizing MFC for cyclic pitch control in the hybrid electrically controlled rotor was analyzed. The simulation results indicated that by increasing the MFC driving voltage, adding more MFC layers, and reducing blade torsional stiffness, the elastic torsional control authority of the hybrid electrically controlled rotor can be effectively enhanced. Meanwhile, the model rotor blade tip can achieve a maximum 1 Ω torsional angle of approximately ±12°, meeting the initial cyclic pitch design objective.

To improve the accuracy of dynamic modeling for aeroengine spline connected rotors, the influence mechanism of tooth surface roughness on structural dynamic characteristics was investigated. A simulation model of single-tooth-pair engagement with rough surfaces was established to reveal the evolution of normal and tangential meshing stiffness with displacement loads. A dynamic modeling method based on the virtual material method was proposed, and a finite element model of the overall structure considering surface topography was constructed. Results showed that lower roughness improved the uniformity of contact stress distribution and contact stiffness, whereas higher roughness caused contact area loss and induced tooth root slip damage. Roughness reduced the natural frequency and increased the vibration response through stiffness softening, with the first natural frequency being the most sensitive to surface conditions. The specific influence was about 7% on the first natural frequency and 11.8% on the radial vibration amplitude. In contrast, torque loading partially offset the negative effect of roughness by improving the contact state. The established cross-scale analysis framework clarified the influence of micro-topography on local contact behavior and global dynamic response, providing theoretical support for performance prediction and manufacturing process optimization of aeroengine spline connected rotors.

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.