Numerical simulation of the acoustic mode characteristics of the combustion chamber with the dome was carried out based on the coupled computational aeroacoustic (CAA)/low-order thermoacoustic network method. The eigenfrequencies, damping rates and acoustic pressure distribution clouds of the main acoustic modes of the combustion chamber were obtained, the accuracy and efficiency of the coupling methods were verified, and the influence of the dome on the acoustic mode characteristics was compared and analyzed. The results showed that, after considering the influence of the dome which was coupled with the acoustic pressure distribution of the combustion chamber, the damping rate of the first-order transverse and longitudinal modes of the combustion chamber was reduced by 24.5% and 16%, but the effect on eigenfrequency was not obvious; compared with the CAA method, the coupling method can improve the calculation efficiency by 49% under the condition of ensuring the calculation accuracy, making it very suitable for the engineering of multi-nozzle liquid rocket engine combustion chamber.

The wing/rotor aerodynamic interference of multi-rotor tilting wing aircraft in helicopter mode was analyzed. Computational model for the aerodynamic interference analysis was established based on CFD method. The aerodynamic interference of the wing/rotor in hover and forward flight was numerically calculated, the influence of aerodynamic interference on the aerodynamic characteristics of the wing was analyzed, and the influence law of flap and aileron deflections was further studied. The results showed that, under the influence of rotor wake, the wing generated a large backward force, accounting for 5.4% of the total rotor force. The flap and aileron deflections during hovering can achieve effective longitudinal and heading maneuvers, and the flap and aileron up deflections during forward flight can reduce the backward wing force by 30%, thereby increasing the maximum forward flight speed.

The noise generated by impinging jets is a critical issue during the takeoff and landing of fighter jets and carrier-based aircraft. However, the mechanism behind the generation of tone in subsonic impinging jets remains poorly understood. An impinging jet with a nozzle pressure ratio of 1.69 and an impingement distance of three nozzle diameters was studied. Large eddy simulation was employed to capture the near-field flow, and an acoustic analogy model was used to predict the far-field noise. The extended Helmholtz decomposition and principal correlation decomposition methods were applied to analyze the modal characteristics and the correlation between the near-field flow and far-field noise. The results revealed that, due to the influence of the impingement plate, a supersonic region appeared in the flow field, forming shock cells. The noise mode of the impinging jet exhibited an axisymmetric structure. The tone in the far field was closely related to the shock waves in the near field. The associated modal structure manifested as waves that gradually propagated outward from the impingement region, and the transmission process connected the vortex structures on the flat plate with the shock structures in the free jet region. The pressure field in the near field dominated noise generation compared with the velocity field. The alternating positive and negative amplitudes of the modal amplitudes in the impingement region constituted a key factor in the generation of tone. This study could elucidate the mechanisms of impinging jet noise and provide a theoretical foundation for the optimization and control of impinging jet noise.

A two-stage counter-rotating axial-flow compressor (CRAC) was studied and the unsteady flow characteristics induced by the coupling of blade tip clearance size (TCS) and rotational speed changes were investigated. Based on numerical simulation results, the proper orthogonal decomposition (POD) method was employed to extract the POD modal distribution patterns. Further analysis on the flow characteristics of the blade TCS under near-stall conditions was conducted. Numerical and FFT results showed that as the blade TCS increased, the leakage flow and the fragmentation of blade tip leakage vortex (TLV) intensified, leading to an enlargement of the range of unsteady fluctuations and a reduction in fluctuation intensity. The frequency of blade tip leakage flow (TLF) fluctuations in the rear rotor (R2) increased. The interference between the blade TLF overflowing channel of R2 and the leading edge of adjacent blades was attributed to the sudden increase in unsteadiness at the leading edge. The POD analysis confirmed the above conclusions and revealed the reasons for the static pressure fluctuation region on the blade surface. It was also found that an increase in blade TCS resulted in a decrease in the dominant modal structure of the three-dimensional flow field in the front rotor (R1), and the higher-order modal structure of R1 migrated radially towards the blade root. As the rotational speed increased, the interference between the upstream and downstream rotors also increased.

The permanent magnet synchronous motor is a critical component of electric propulsion systems, and its operational status is integral to the system’s safe functioning. Failure of the permanent magnet synchronous motor (PMSM) can trigger multiple physical quantity changes, making it challenging to achieve accurate fault diagnosis relying on a single signal source. To address this issue, a decision-level multi-channel information fusion fault diagnosis method was proposed, by combining convolutional neural networks and gated recurrent units (CNN-GRU) and improved Dempster-Shafer (D-S) evidence argumentation. Initially, analytical and finite element methods were employed to quantitatively analyze the vibration and current frequency domain characteristics of local demagnetization and rotor eccentricity faults in permanent magnet synchronous motors, thereby enhancing the interpretability of diagnostic results. Subsequently, a decision-level fusion diagnostic model was established, by integrating CNN-GRU and improved D-S evidence theory based on Pignistic probability distance and weighted Deng entropy. Finally, a motor fault simulation tester was constructed, and the model was validated using experimental data. The results demonstrated that multi-channel fault diagnosis is superior to single-channel diagnosis results. The decision-level fault diagnosis based on multi-source data, with the fusion of 4 channels, achieved diagnostic accuracy of 100%, 100%, and 99.3%, respectively, under three operating conditions. The proposed method accurately identified the types of permanent magnet synchronous motor faults, providing a reference for fault diagnosis in electric propulsion systems and offering potential value for engineering applications.

The intensely unsteady exhaust flow from pulse detonation combustor has a significant impact on turbine performance. The matching problem between pulse detonation combustor and turbine components is one of the difficulties restricting the development of pulse detonation turbine engine. In order to clarify the research progress on turbine performance and optimization strategies under pulse detonation gas impact, the flow characteristics and evolution process of exhaust flow field were briefly introduced with the pulse detonation combustor exhaust characteristics as the starting point. From the aspects of theoretical analysis, experimental testing and numerical simulation, the research status of key issues such as turbine performance evaluation methods, turbine operating characteristics, turbine internal flow characteristics and loss mechanism under pulse detonation gas impact was summarized. The research achievements of Northwestern Polytechnical University in the field of pulse detonation combustor and turbine component matching were reviewed, and the key problems to be solved in this field were prospected. Optimization strategies for turbine performance under pulse detonation gas impact were proposed, including stabilized pressure device, shock attenuation device, and turbine blade optimization design, etc. This indicates that the organic integration of various optimization strategies is one of the future research directions.

The coupled fluid-solid-thermal finite element model of the active cooling channels in a scramjet combustion chamber was established using the substructure method. The effects of the number of struts on the heat transfer characteristics of strut-type lattice cooling channels under identical volume fractions and loading conditions were investigated, and a multi-objective optimization of the strut cross-sectional shape was conducted. The results showed that, compared with the rectangular smooth channel, the average Nusselt number of the strut-type lattice channel increased by at least 64.2%, significantly enhancing the heat transfer performance of the cooling channel, but the flow resistance increased by more than 5.8 times; the heat transfer performance of the strut-type lattice channel first increased and then decreased with the increase in the number of struts, with the two-strut lattice channel exhibiting the best heat transfer performance. The multi-objective optimization results of the lattice channel, which aimed to maximize the average Nusselt number, and minimize the pressure drop with the highest temperature, showed that under the same volume fraction, the more number of struts meant the lesser impact of the cross-sectional shape on the flow and heat transfer performance of the lattice channel. The circular cross-section strut served as a suitable solution for balancing heat transfer and pressure drop requirements for two or more struts-type lattice channels. The optimal solutions of the single-strut lattice channel exhibited the largest improvement in the target performance, of which the average Nussle number of the compromise solution increased by 20.2%, the maximum temperature decreased by 4.4%, and the pressure drop increased by 55.1%, demonstrating better overall performance. The results could provide a reference for the design of lattice channels.

The internal flow field of a stealthy inlet with an inclined-cut entrance under the ground state exhibited complex three-dimensional separation structures, which greatly reduced the aerodynamic performance of the inlet and cannot meet the requirements of inlet/engine matching. A lip self-supplying air-jet for flow control method of the ground state of a triangular inclined cut entrance inlet was proposed. Numerical simulation results showed that, at the ground state, the original vortex on the entrance of the triangular inclined cut entrance inlet was divided into two small and one large vortices by the lip self-supplying air-jet, the scale of the vortex near the symmetry plane was significantly reduced, and the shedding of the large-scale vortex near the symmetry plane on the upper wall was delayed, improving the circumferential uniformity of the low total pressure area distribution at the exit section of the inlet, and greatly reducing the distortion index. Furthermore, the effects of the transverse spacing of the jet slot, transverse position of the jet slot, jet angle, and jet area on the aerodynamic performance and flow field structure of the stealth inlet under ground state were studied, and the optimal parameter configuration was obtained. Compared with the baseline, the total pressure recovery coefficient of the control method decreased slightly, and the circumferential distortion index δ₆₀ significantly dropped from 0.8164 to 0.2880, meeting the requirements for inlet/engine matching.

To investigate the fretting fatigue performance of additively manufactured superalloys, fretting fatigue test specimens were designed, macro fretting fatigue experiments on additively manufactured superalloys under various conditions were conducted, and the fretting fatigue life of specimens under different printing directions and loads was obtained. The fretting fatigue life of additively manufactured superalloy specimens decreased with the increase of peak load in different printing directions, and the fretting fatigue life along the printing direction was higher than that perpendicular to the printing direction. To establish a fretting fatigue life model for additively manufactured superalloys in different printing directions, finite element simulations of fretting fatigue experiments under different conditions were performed, and the stress and strain distributions were obtained. It was found that the maximum Mises stress and maximum strain both occurred in the fretting fatigue contact area, which coincided with the crack initiation position. Furthermore, a fretting fatigue life prediction model was established using the critical plane method. The results showed that the predicted life of SWT and FS parameters for specimens in different printing directions was within ±2 fatigue scatter band.

The unsteady transonic aerodynamic characteristics of a NACA64A010 airfoil in air and heavy gas medium R-134a were simulated by using the Reynolds-averaged Navier-Stokes equations (RANS) and the Spalart-Allmaras (SA) one-equation turbulence model. Under the same Mach number, Reynolds number and reduced frequency, the calculation results of simple harmonic motion in pitch showed that the distribution of modulus and phase of unsteady pressure coefficient in heavy gas medium was obviously different from that in air, the unsteady lift coefficient was not different from that in air medium, and the amplitude and phase of pitch moment coefficient were different from those in air medium. The transonic similarity law was applied to the unsteady aerodynamic force correction, and the amplitude and phase of pitch moment coefficient were transformed into air similarity, but with the increase of reduced frequency, the correction effect of transonic similarity law became worse. Through analysis of pitching moment work, it was shown that if the unsteady aerodynamic force in heavy gas was not corrected by similarity, the flutter characteristics in air and heavy gas could be different, thus affecting the evaluation of flutter characteristics in heavy gas wind tunnel. This study can provide a basic support for the follow-up research on flutter characteristics of aircraft in heavy gas media and the development of correction methods.

Large eddy simulations of hydrogen and hydrocarbon fuel films with boundary-layer combustion were conducted, focusing on the effects of boundary-layer combustion on the near-wall heat and mass transport processes of fuel films. The results showed that boundary layer combustion effectively reduced the heat and mass transport fluxes between the mainstream and fuel film, which enhanced the heat insulating performance of the film. However, the decrease in heat flux within the hydrogen film was not adequate to offset the negative impacts of heat release and consumption of hydrogen with high heat capacity during combustion, leading to a significant deterioration of the heat insulating performance during boundary layer combustion. On the contrary, the reduction of heat flux in the hydrocarbon film had synergistic effect with the heat absorption of the near-wall pyrolysis reactions, leading to a substantial enhancement in the heat insulating performance of the hydrocarbon film.

The ${H_\infty }$ robust stability theory was combined with the classical loop shaping theory to design a two-degree-of-freedom ${H_\infty }$ controller while taking into account both the control performance and robustness requirements of the system; to simplify the control system, the inner-loop controller was reduced to the seventh order, and the outer-loop controller was reduced to the fourth order; for the full envelope wide-range tracking control problem, 18 design points were selected within the flight envelope, the controllers were designed for different operating modes and power levels, and the gain scheduling table was constructed; different flight conditions were selected for simulation, and the simulation results showed that the steady state error was 0 and the settling time was less than 2.5s, which verified the good control performance of the full envelope ${H_\infty }$ controller; finally, the hardware-in-the-loop simulation platform was built to verify the two-degree-of-freedom ${H_\infty }$ control algorithm. The results showed that the control algorithm can complete the calculation within 2ms, and the engine output can track the reference instruction quickly and accurately, verifying the real-time and effectiveness of the algorithm in the hardware environment.

To study the rotational fretting wear characteristic of ball bearing and rotor assemblies for gas turbine, spectroscopy, laser net fines analysis, ferrography and X-ray microanalysis were employed to analyze the wear particles in lubricant. The results showed that the total amount and size of wear particles matched well with the wear process, which varied according to wear patterns and increased sharply at the final period of wear. The rotational fretting wear of ball bearings and rotor assemblies were mainly caused by the rotation of inner ring relative to the shaft, and the flat platelets were the characteristic wear particles for rotational fretting wear. The wear process was not isolated. In short, with the rotation of inner ring relative to shaft caused by friction and high temperature gradient, the wears were founded on spacer sleeve, shaft, and locknut face sequentially, and crushing marks were formed at the bottom of non-load bearing half-ring raceway. To be notice, the analysis process can also provide a reference to the wear mechanism of ball bearings and rotor assemblies of other gas turbines.

An improved Gappy POD (proper orthogonal decomposition) method was developed in response to the low inverse design accuracy and limited operability of the method. Furthermore, improvement strategies were proposed to address the issue of lower inverse design accuracy for transonic blade compared with subsonic blade. Through verification examples, the inverse design of compressor blade based on the improved Gappy POD method was proved simpler and more efficient. Compared with the basic method, the inverse design accuracy of the two blades was improved by 61% and 78%, also the inverse design time was shortened by 77% and 63%, respectively. Further research found that the benefits of the improved method in accuracy of blade fitting mainly came from the improvement of pressure surface reconstruction accuracy, while the geometric deviation of the blade in the second half of the suction surface was relatively large, and the geometric deviation showed a trend of increasing first and then decreasing from the leading edge to the trailing edge.

The conventional Pogo suppression method uses the installation of accumulators in the fluid system, however, the improvement of the carrying capacity of new launch vehicles is a challenge to the conventional passive accumulator, since it is unable to meet the suppression needs of larger flow pulsations. Therefore, the scheme and control method of Pogo active suppression were studied, and simulation validation was achieved. Firstly, a closed-loop coupling model of propulsion system and structural system of liquid rocket was established by using software MATLAB/Simulink, and the Pogo vibration was reproduced in the time domain. Secondly, two active suppression schemes, including piston type control for low-pressure pulsation and relief type control for high pressure pulsation, were proposed. The control method of adaptive filtering algorithm based on secondary path online identification was studied. The simulation validation was modelled with software AMESim and MATLAB/Simulink, and active suppression of Pogo was achieved. The simulation results showed that both relief scheme for high-pressure pulsation and piston scheme for low-pressure pulsation can achieve more than 90% amplitude attenuation, and eliminate the Pogo vibration of liquid rocket, verifying the effectiveness of the active suppression scheme.

From the perspective of flight test, the connotation, research advances, application and problems of aircraft-engine integrated design were discussed from three aspects: inlet/engine compatibility, engine installed performance and aircraft-propulsion integrated control. Airflow matching, inlet pressure/temperature/swirl distortions were still the main reasons for engine instability in flight test. The inlet total pressure distortion during high angle of attack transient maneuver increasing rapidly was much higher than the inlet distortion level during conventional angle of attack flight. The vibration and structural failure induced by inlet distortion cannot be ignored. The installed propulsion force was composed of standard net thrust and drags relating to engine power. Based on the thrust-drag division criterion of “whether it is related to the throttle”, the installed propulsion force could be calculated by combining CFD, wind tunnel test and flight test. A large number of researches carried out by USA military and NASA proved that aircraft-propulsion integrated control technique was able to remarkably improve the aircraft system overall performance and stability. In future, the research on aircraft-propulsion integrated flight test and evaluation technology should focused on aircraft-propulsion integration simulation test, engine installed performance determination, engine installed aerodynamic stability online evaluation and aircraft propulsion integrated control evaluation method, etc.

A small energy spark discharge system based on an RLC discharge circuit was designed for measuring the minimum ignition energy of domestic aviation fuel RP-3. The relationship between spark discharge and charging voltage, as well as capacitor capacity, was studied considering capacitor residual energy and circuit energy loss. The spark discharge current, energy consumption, and efficiency at different energy levels were analyzed, and minimum ignition energy tests for RP-3 fuel were conducted. The results showed that the RLC circuit is suitable for spark discharge systems below 1 J. The larger capacitor meant the greater stored energy, the longer discharge time, and the higher spark discharge efficiency, reaching up to 80%. The minimum ignition energy of fuel vapor varied with temperature, following a “U” shape curve. At 50 ℃ and 101 kPa, the minimum ignition energy was 0.167 mJ.

Aiming at the problem of insufficient consideration of helicopter / engine coupling relationship in the traditional overall parameter design process of turboshaft engines and the lack of boundary restrictions on the feasible region of alternative design parameters, this paper proposes a method for optimizing the overall parameters of turboshaft engines based on the requirements of helicopter flight missions, and reasonably limits the feasible region of design parameter optimization according to the development of turboshaft engines. The method consists of four modules: constraint analysis, task analysis, parameter cycle calculation and optimization iteration. The optimization iteration module involves dynamic adjustment of design parameters, which makes the designed engine more matched with the task requirements. Firstly, the basic principles of the above four modules are derived. Then, based on the given helicopter flight mission profile, the proposed design method is verified by an example. Finally, the results are compared with other design methods for the same flight mission. The research shows that the proposed feasible region boundary of design parameters can reasonably limit the selection range of design parameters and can be used as a basis for the selection of design parameters. The proposed optimization design method of the overall parameters of the turboshaft engine makes up for the shortcomings of other overall design methods of the turboshaft engine, such as the design parameters are not involved in the iterative update and the lack of consideration of the coupling relationship between the helicopter/turboshaft engine performance.

In the view of the existing problems and the application requirements of long endurance, an overall design scheme of long-endurance multi-rotor fuel cell UAVs with fuel cell as the main power supply device and lithium battery as auxiliary power supply device is proposed. Based on the finite element analysis method, the structure of the body was selected, the structural parameters such as the length and diameter of the arm were determined, and a reasonable fuel cell hydrogen storage system platform was designed. The "fuel cell + lithium battery" hybrid power system was constructed, and the topology of the hybrid system was simplified through axiomatic design. According to the design scheme, the UAV system integration was completed, and the test flight tests of the UAV no-load flight, load capacity and endurance were carried out, and the comprehensive performance of the hybrid power system and its control strategy and heat dissipation module was verified. After actual testing, the weight of the UAV is 37.95 kg, the payload capacity is 10 kg, and the endurance time after the payload task system can reach 4 hours and 20 minutes, which meets the expected design effect.

To optimize the dynamic balance layout of high aspect ratio rotor blades, a multi-objective optimization model was established by incorporating radial, axial, and tangential mass moments as key constraints. To address the limitations of traditional genetic, ant colony, and particle swarm optimization algorithms in convergence speed and global search capability, an improved optimization strategy integrating multi-dimensional objective characteristics was proposed. A multi-dimensional fitness function was designed to accommodate multi-directional moment targets, while a symmetry constraint and weight control mechanism were introduced to enhance the structural rationality of the solution. Population initialization was improved, and a local search mechanism was integrated to boost both local convergence precision and global convergence efficiency. Experimental results based on actual blade mass moment data showed that the average fitness of the improved genetic algorithm decreased from approximately 280 000 to 28 000, achieving a reduction of over 90%. Compared to the other two algorithms, optimization efficiency improved by about 15%, and the final solution quality increased by more than 10%.

A composite damping control strategy was proposed to address the control challenges of piezoelectric bimorph actuators for electrohydraulic servo valves in oil environments. For the low-frequency range below 100 Hz, a feedforward compensation strategy based on inverse multiplication was adopted, combined with a feedback controller using a disturbance observer to enhance the system robustness. Experimental results showed that at 1 Hz, the root mean square (RMS) error was reduced from 1.76 μm to 0.83 μm, and at 100 Hz, it was reduced from 3.44 μm to 2.17 μm. For the high-frequency range of 500—700 Hz, a composite strategy based on positive velocity and position feedback (PVPF) damping control was proposed and compared with traditional proportional-integral-derivative (PID), feed-forward (FF)+disturbance observer (DOB), and FF+PVPF control methods. The results demonstrated that the composite control strategy effectively compensated for hydrodynamic disturbances and suppressed the lightly damped characteristics near the first natural frequency, maintaining RMS tracking errors below 2.5 μm under various operating conditions. This study could provide a novel control approach and experimental validation for the application of piezoelectric bimorph actuators in high-precision electrohydraulic servo systems.

A study was conducted to investigate the dynamic characteristics of key components of a cascade thrust reverser mechanism under multiple operating conditions. Based on the working principle of the thrust reverser device, a single-linkage motion mechanism was extracted while retaining the critical contact states of the motion pairs and the dimensions of key components. A simulated test bench for the thrust reverser mechanism was designed and developed. By conducting principle-level experimental tests under different load conditions, the variation patterns of the mechanism's block force and guide rail stiffness were obtained. Meanwhile, the nonlinear dynamic simulation model for the rigid-flexible coupled thrust reverser mechanism with clearance collisions was established. Simultaneously, comparison and verification between experimental tests and simulations were completed. The results indicated that, compared with the lateral loading condition, the maximum block force of the mechanism increased by 19% under longitudinal loading. When the lateral load was below 150N, the peak block force showed an approximately linear increase, with a growth rate of approximately 22%. Under longitudinal loading, the overall deformation of the mechanism's guide rail was relatively stable, while under lateral loading, the deformation was more sensitive to changes. However, as the load increased, the vibration of the guide rail was inhibited. The experimental data were in good agreement with the simulation results, confirming the validity of the experiments and the accuracy of the model.

In view of the unclear lubrication characteristics of high-speed gear transmission under low temperature conditions and the limited improvement of lubrication effect by traditional baffles, a study on the splash lubrication characteristics of high-speed gear transmission was carried out. Firstly, based on the theory of computational fluid dynamics, a two-phase flow analysis model of splash lubrication of high-speed gear transmission was established, and the two-phase flow distribution characteristics were analyzed; then, on this basis, the influences of number of rotations, gear speed, and oil immersion depth on splash lubrication characteristics and torque loss were studied; finally, a bionic honeycomb baffle structure was proposed and its structure was optimized using a multi-island genetic algorithm. The results showed that: in the splash lubrication process of high-speed gear transmission under low temperature conditions, the volume fraction of lubricating oil on the tooth surface decreased with the increase of gear speed, but increased with the increase of oil immersion depth; the torque loss increased with the increase of gear speed and oil immersion depth, and the influence of speed on torque loss was much greater than that of oil immersion depth; under the same working conditions, the average volume fraction of lubricating oil on the tooth surface of the optimized bionic honeycomb baffle structure increased by 68.46% compared with the one without baffle, and increased by 7.88% compared with the original baffle. Research results provide a basis for the study of two-phase flow distribution characteristics of high-speed gear transmission under low-temperature conditions and the optimization design of splash lubrication for high-speed gear transmission.

Aiming at the problem that the fault signal of aero-engine rolling bearing is affected by background noise, which leads to weak characteristics of the fault signal, a fault feature extraction method of rolling bearing based on grey wolf algorithm optimization feature mode decomposition and adaptive window length sliding window noise reduction was proposed. Firstly, the vibration signal was decomposed by feature mode decomposition method, and then the signal components were filtered and reconstructed according to the kurtosis correlation coefficient selection criterion. Secondly, in order to solve the problem that the input parameters need to be manually filtered when using the feature mode decomposition method, the information entropy of the reconstructed signal was selected as the objective function, and the grey wolf algorithm was used to optimize the input parameters to determine the number of modal components and the size of the filter. Then, the noise reduction method of adaptive window length sliding window was used to perform secondary noise reduction and fault feature enhancement on the reconstructed signal, and the noise reduction signal was output. Finally, the noise reduction signal is envelope demodulated to extract fault features. The effectiveness of the proposed method is verified by using the simulation signal, the data set of Case Western Reserve University, the data of the aeroengine intermediate bearing testbed and the data of the main bearing testbed. The results show that the combination of feature mode decomposition and sliding window noise reduction method can effectively filter out the interference noise components . The information entropy has decreased by approximately 40%, making the fault characteristic frequency and its 2-4 times amplitude in the envelope spectrum more prominent. Compared with other existing bearing fault diagnosis methods, it has better noise reduction effect and stronger fault feature extraction ability.

The polynomial response surface model (PRSM) was characterized by its simplicity in modeling and low computational cost, making it widely used in wind tunnel tests based on the modern design of experiments (MDOE) method. However, PRSM cannot process residuals after model fitting, leading to a loss of some model matrix information and increased parameter prediction error. To address this, a residual objective function based on the difference between predicted and measured values was constructed. Then, the crossover and mutation probabilities of the genetic algorithm (GA) were improved by calculating individual fitness values using a roulette wheel selection method. This contributed to the establishment of an MDOE method optimized by an improved adaptive genetic algorithm (IAGA) for PRSM, which was applied to aircraft wind tunnel tests. The results showed that the MDOE method required only about 35% of the experimental points needed by the one factor according to OFAT method. Compared with the traditional GA-based MDOE method, the IAGA-based MDOE method decreased the prediction error of the response surface model by 1.263% and increased the average iteration speed by 3.76 times, which effectively improved the efficiency of wind tunnel tests.

To investigate the impact of nozzle exit area variations on small turbofan engine performance, a 100 daN-class engine across 33 000 r/min to 60 000 r/min speed range under constant speed and constant fuel flow conditions was studied. A mathematical model was developed to quantify the coupling effects between engine parameters, and an optimization strategy was proposed. The results showed that at 54 000 r/min, reducing the equivalent area ratio (EAR) from 1.24 to 0.85 increased the static thrust by 107%, raised the exhaust gas temperature (EGT) by 400 K, and decreased the specific fuel consumption (SFC) by 10.6%. Under constant fuel flow conditions, decreasing EAR led to a 14.8% reduction in speed, a 4% decrease in thrust, a 121 K increase in EGT, and a 6% rise in SFC. Analysis revealed that at EAR of 0.925, SFC was minimized within the 60—80 daN thrust range, indicating the best matching between fuel combustion and gas expansion. Compared with high EAR (1.24), operating with low EAR (0.85—0.925) reduced SFC by 10%—20%, with a 300 K increase in EGT. The optimal operating range was identified as EAR of 0.925—1.0, corresponding to static thrust of 60—80 daN and EGT of 800—950 K. In this range, SFC can be reduced by 15%—20%, static thrust increased by 30%—40%, while EGT remained within material resistance range. These findings could offer valuable insights for adaptive nozzle design and real-time control in unmanned aerial vehicle propulsion systems.

The longitudinal dynamic stability characteristics of a coaxial rigid rotor helicopter in forward flight are greatly affected by the rotor design parameters. The main design parameters of advancing blade concept rotor were calculated and analyzed. CAMRAD II was used to establish XH-59A coaxial high speed helicopter model. The effects of rotor blade flapping frequency, pitch-flapping, blade mass distribution, the position between blade center of gravity-aerodynamic center(CG-AC) and torsional stiffness on the longitudinal long period modes of coaxial helicopter were analyzed. The results showed that: the higher flapping frequency of a rigid blade indicated the stronger rotor angle-of-attack instability and increased with forward flight speed. Positive pitch-flap coupling, large flapping inertia moment and center of gravity(CG) located in front of aerodynamic center(AC) effectively reduced angle of attack instability for rigid rotor and improved the longitudinal dynamic stability of the helicopter; the large torsional stiffness suppressed the blade torsional modes, thus suppressing the influence of the blade CG-AC on the longitudinal long-period modes.

A data-driven coupled prediction method for turbine performance was proposed by incorporating multiple machine learning models and inter-stage flow angle variation patterns. Multiple prediction analysis models were established based on fundamental data, and the entropy weighting method was used to perform comprehensive quantitative analysis of the model predictions. The optimal prediction model was dynamically determined based on data-driven approaches. Verification through multi-stage axial turbine characteristic variation examples demonstrated that under constant rotational speed conditions, compared with the traditional direct prediction method without considering changes in the flow angle, the computational accuracy of the flow-efficiency characteristic and the flow-expansion ratio characteristic was improved by 40.47% and 18.26%, respectively. Furthermore, the maximum error compared with the results of three-dimensional numerical simulations was less than 3%. For variable speed performance prediction, the construction of the turbine characteristic map within a wide flow range of 2 to 8 kg/s was achieved through grid hyper parameter optimization methods. Simulation testing of the constructed map validated that the maximum error in the expansion ratio was close to 3%, meeting the requirements for characteristic estimation accuracy.

Focusing on aerodynamic heat recovery challenges in hypersonic vehicles, the solid-state thermoelectric power generation technology at the leading edge for blunt-body vehicle architectures was investigated. A multi-physics computational model integrating hypersonic vehicles with thermoelectric generators was developed to simulate the electrical output characteristics under actual high-altitude flight conditions. The accuracy of the model was verified through ground platform tests and wind tunnel tests, while the output performance of thermoelectric generators under different environmental conditions was systematically examined. Comparative studies were conducted on thermoelectric devices with various materials and configurations. The results revealed that the thermoelectric generators effectively converted aerodynamic heat into electrical energy, with a single TEG achieving maximum output power of 1.21 W and power density of 1344.44 W/m² under simulated actual flight conditions. Wind tunnel test demonstrated closer approximation to real flight environments compared with ground platform tests, as it yielded a maximum output power of 0.5 W and peak conversion efficiency of 2.58% for individual TEG. Thermoelectric materials in different temperature zones were suited to distinct TEG structures, requiring rational design and arrangement to optimize the performance.

Optimization was conducted to improve the aerodynamic and aeroacoustic performances of a given exhaust volute by particle swarm optimization and Pareto front. By taking the given external dimensions as constraints, a parameterized model of the exhaust volute was established, and aerodynamic parameters such as the total pressure loss, the static pressure recovery and multi-angle-averaged overall sound pressure level were set as the performance indicators. The RANS (Reynolds-averaged Navier-Stokes) method was employed for the prediction of aerodynamic performances. At the same time, based on the sound generation characteristics of the exhaust volute, the exhaust process was analogized to a subsonic jet, and its acoustic performance was predicted using the Tam-Auriault model. The results showed that: ① the particle swarm algorithm was efficient and robust for parameterized optimization. After six rounds of optimization, the total pressure loss coefficient at the outlet of the exhaust volute was reduced by 35%, and the static pressure recovery coefficient was increased by 79%; ② There existed a conflict between the aerodynamic and aeroacoustic performances, the comprehensive aerodynamic and aeroacoustic optimization managed to reduce the total pressure loss coefficient by 20%, and the static pressure recovery coefficient increased by 48%, with a comprehensive noise increment about 1 dB. If the aeroacoustic performance optimization was considered, the noise reduction can be up to 3 dB.

A wall-mounted flexible vortex generator (FVG), a tandem column, and a non-Newtonian fluid were utilized to enhance the heat transfer in a microchannel. The finite element method in arbitrary Lagrangian-Eulerian (ALE) format was used to solve the continuity, momentum, and energy equations describing the flow-solid coupling in microchannels. The flow-induced vibration behavior and heat transfer performance of a power-law fluid in a symmetric wall-mounted FVGs channel were investigated by adjusting the spacing between the FVGs and the columns, and varying the power-law exponent (n) of the fluid under the condition of a fixed distance between the two columns. When the distance between the downstream column and the flexible plate G_x=1.5 and the power-law index n=1.2, Nusselt number Nu increased by 184% and the thermal performance coefficient increased by 45% compared with the straight microchannel.

To improve the lubrication efficiency and application range of bearing jet lubrication, the causes and influencing factors of the air curtain effect within high-speed angular contact ball bearings were investigated. Based on computational fluid dynamics (CFD), a finite element analysis model was constructed for the single-phase flow field characteristics within the bearing chamber. The flow field characteristics of the bearing under different guide/pocket clearances and rotational speeds were analyzed. Based on the Q criterion in the second generation vortex recognition technology, the evaluation method and evaluation standard for the air curtain effect were put forward; and the formation mechanism of air curtain within the bearing chamber was analyzed. The effects of different rotational speeds, guide clearances, and pocket clearances on the air curtain effect were compared. The main influencing factors of the air curtain effect were analyzed using an orthogonal experimental method. The results showed that the air curtain effect within the bearing chamber intensified as the bearing speed increased. Increasing the guide and baffle hole clearances can reduce the air curtain effect. The impact of rotational speed on the air curtain effect was much greater than the impact of cage clearance changes. The baffle hole clearance had a more significant influence on the bearing’s air curtain effect compared with the guide clearance.

This work focuses on the laminar combustion characteristics of a certain bio-jet fuel with hydroprocessed esters and fatty acids (HEFA) technology. The laminar burning velocity (LBV) of HEFA bio-jet fuel and HEFA bio-jet fuel/RP-3 mixed fuel were obtained by a constant volume combustion experimental device. The experimental parameters include equivalence ratios (0.8—1.4), initial pressures (0.05, 0.1, 0.15 MPa), initial temperatures (450, 470 K) and HEFA mixing ratios (0, 0.1, 0.2, 0.3, 0.5), et al. The effects of the equivalence ratio, initial pressure and initial temperature on the LBV of HEFA bio-jet fuel and the HEFA mixing ratio on the LBV of HEFA/RP-3 mixed fuel were analyzed. It is found that with the increase of equivalence ratio, the LBV of HEFA bio-jet fuel shows a trend of increasing first and then decreasing, and its peak value appears near the equivalence ratio of 1.1. As the initial pressure increases from 0.05 MPa to 0.15 MPa, the LBV of HEFA bio-jet fuel decreases by 13.54%. As the initial temperature increases from 450 K to 470 K, the LBV of HEFA bio-jet fuel increases slightly. With the increase of HEFA mixing ratio from 0 to 0.5, the LBV of HEFA/RP-3 mixed fuel increases by 3.85%. The results show that compared with RP-3 jet fuel, although the carbon number distribution of HEFA bio-jet fuel is higher, its chemical composition is dominated by alkanes, which leads to its LBV higher than that of RP-3 jet fuel, so that the LBV of the mixed fuel after blending HEFA is slightly increased. This study provides a theoretical basis for the application of bio-jet fuel in aircraft engines.

A ground-connected experiment of a water ramjet engine using magnesium powder as fuel was successfully conducted. By using a continuous adjustment method integrating a variable exit area with a variable-pressure magnesium powder supply, the issue of flameback caused by pressure imbalances between the combustion chamber and the fuel supply system was effectively resolved. This method enabled stable combustion within the metal powder-based water ramjet engine. The engine utilized a dual-stage water injection system arranged in three groups, achieving a total water-to-fuel ratio of 2.5. The combustion chamber maintained an average pressure of 0.181 MPa, producing an average thrust of 42.23 N, with a metal fuel combustion efficiency of 76.98%. The experimental results indicated that an appropriate pressure differential for powder supply is a critical factor for ensuring stable engine operation. Furthermore, the engine’s operating sequence is essential to successful ignition and stable performance. Establishing a suitable thermal environment within the combustion chamber can be identified as the key to achieve continuous and stable combustion in the engine.

The jet behaviors of sub- and super-critical RP-3 aviation kerosene in a subsonic air crossflow were experimentally investigated. Under constant air parameters and fuel mass flow rate conditions, shadowgraph imaging technology was used to visualize the transverse jet structure within the fuel injection temperature range of 352 K to 750 K. The results showed that the breakup/disintegration regime of RP-3 jet gradually changed from mechanical breakup to unchocked flash-boiling, chocked flash-boiling and supercritical regime as the fuel injection temperature increased. During this process, the jet dense core and atomization zone underwent a decrease in length of over 11 and 52 times the nozzle diameter, respectively, and a decrease in coverage of over 6 mm² and 141 mm², respectively. Furthermore, with the increase of the injection temperature, the penetration depth of the jet gradually increased in the mechanical breakup regime, decreased in the unchocked flash-boiling regime, and remained essentially unchanged in both the chocked flash-boiling and supercritical regimes.

To elucidate the variation of stability during transition flight, the XV-15 model was taken as the research object. A comprehensive analysis was conducted to understand how factors such as nacelle angle, flight speed, and rotor/wing aerodynamic interactions, influenced the eigenvalues and eigenvectors of motion modes, and stability derivatives, along with their underlying mechanisms. The findings revealed that nacelle angle primarily influenced the eigenvalues of main motion modes and the eigenvectors of Dutch roll and phugoid modes. The underlying key factors included the body’s angular velocity projection onto the rotor axis, the angle between the rotor disc and flight velocity, and the distance from the rotor hub center to the center of gravity, which changed sine or cosine with the nacelle angle. During low-speed flight, the aerodynamic interference between the rotor and wing exacerbated the instability of phugoid mode. The instability was alleviated as the nacelle angle decreased and the flight speed increased. To facilitate the selection of transition paths for tiltable mechanisms, a comprehensive evaluation metric for assessing motion stability within tilting corridors was introduced. This metric in contour mapping visually represented the aggregate trends in motion stability, offering an intuitive understanding of its variations.

The structural dynamics of integral bladed disks is very sensitive to the blade geometric variances. The methodological development and experimental verification for the high-fidelity digital twin modeling and high-performance dynamic analysis for blisks were presented by leveraging the advance optical geometry measurement technology. Firstly, the real blisk geometries were measured in the form of a high-resolution point cloud by employing a 3D structured blue light scanning system. The point cloud was directly correlated to the tuned blisk model in its nominal design. The high-fidelity digital twin model was constructed by an adaptive mesh deformation technique in an efficient way. Subsequently, high-performance dynamic analysis was realized by a novel model reduction technique specifically proposed for large-sized geometrically mistuned blisks. The basic idea was to construct the reduction mode basis of the full blisk by using the cyclic modes of the individual sectors with geometry mistuning having relatively low memory and computational time cost. The experimental verification results demonstrated that the high-fidelity digital twin model of the real blisk test piece can effectively capture the variations of sector frequencies and mode shapes due to the small blade geometry variances. Moreover, it also allowed to predict the mistuned responses of the rotating blisk under engine-order excitation.

In order to clarify the establishment process of the internal separations for the unstart mode of dual-separation on both body and lip (DSBL), and reveal the prerequisite conditions for forming this unstart mode, two transient processes (increasing the angle of attack and blocking and then releasing downstream channel) for inducing DSBL were studied experimentally. The results were as follows: 1) For the two processes, the first induced separations were all located on the body side, and then large-scale separated flow appeared on the lip side. 2) The ability to maintain the start state of the inlet was not related to DSBL flow mode, but rather depended on choking of the throat. 3) DSBL was induced only when there were separations on the body-side, and the generated separation shock impacted to a specific range of the compression surface of the lip.

In order to study the influence of fuel distribution change on the combustion instability of bluff-body flame under multi-nozzle bluff-body structure in afterburner, the model afterburner was taken as the research object, and the oscillation boundary variation law of bluff-body non-premixed flame instability combustion under different nozzle numbers and inlet velocities was studied. The test results were analyzed by FFT spectrum, kurtosis analysis and deconstruction of flame images. The influence of fuel distribution on the oscillation boundary of the combustion chamber of the bluff body model under different nozzle numbers was studied. The experimental results showed that with the increase of the number of bluff body nozzles, the critical lean equivalent ratio of the system from stable combustion to oscillating state gradually increased, and also increased with the inlet velocity, which was prone to combustion instability. In addition, the pressure fluctuation amplitude and dominant frequency of the oscillation boundary also increased with the increase of the number of nozzles, indicating that the change of the flame structure under the multi-nozzle bluff body mainly caused the change of the oscillation boundary in the model combustion chamber.

Aimed at designing the support structure for the rotor system of the gas generator, the influence of the mechanical characteristics of the support structure on the dynamic characteristics of the rotor is analyzed. On this basis, a metal rubber support structure was proposed, and the mechanical properties of the metal rubber support structure were optimized to ensure the critical speed margin and reduce the dynamic response of the rotor system. Finally, the quasi-static mechanical properties and dynamic mechanical properties of the metal rubber support structure were tested. The results show that: After the optimization of the stiffness and damping of the metal rubber support structure, the amplitude of the fulcrum dynamic load of the rotor system at two critical speeds is reduced by about 40%. In the sweep frequency test, the frequency response function of the optimized metal rubber support structure at the peak is reduced by up to 70%, which verifies the effectiveness of the metal rubber support structure. This research has important engineering reference value for the support structure design of rotor systems.

A life assessment method based on nonlinear ultrasonic detection signals was proposed for blade structures with corrosion defects. First, the finite element method was used to simulate the ultrasonic non-destructive testing process of the corrosion defect specimens, and the relationship between the characterization parameters of the corrosion pit and the nonlinear coefficient was explored. Then, by establishing a fatigue life prediction model of a single corrosion pit test piece, a fatigue life assessment method for blade structures with corrosion pits based on nonlinear ultrasound was proposed. Finally, taking a compressor blade as an example, the detailed process of the assessment method was introduced, and the feasibility was verified. The results showed that the corrosion pit diameter was positively correlated with the nonlinear coefficient; Under the same corrosion conditions, changes in the maximum stress and nonlinear parameter were consistent; the life prediction results of specimens with single corrosion pit fell in the factor of-two scatter band, and the error range of life prediction results of blade structure with corrosion pits was within 25%. The study showed that the nonlinear coefficient could characterize the damage caused by corrosion pits effectively, and the nonlinear ultrasonic detection methodology could provide a feasible means for fatigue life assessment of blade structures with corrosion pits.

An arrayed microtube nozzle burner was employed in achieving efficient fuel and oxidizer mixing through cross-jet flows. Research on the combustion characteristics of hydrogen and oxygen in a steam environment was conducted. Through numerical simulation, the flame stability of the micro tube in steam environment was verified. The combustion efficiency of arrayed model burner with different microtube spacings was also contrasted. Experiments on the combustion of pure hydrogen and oxygen in a steam environment were carried out under atmospheric conditions. Flame images, ignition and extinction boundaries, combustion efficiency, and combustion oscillation characteristics of the burner were obtained. The results indicated that increasing steam temperature can significantly widen the flame ignition and extinction boundaries of the burner. The combustion efficiency of hydrogen was kept above 99.9% under various conditions, and the pressure fluctuations of the burner were at relatively low levels. This initially validated that the proposed burner can ensure efficient combustion of hydrogen and oxygen in a steam environment, with minimal combustion oscillations.

The criterion of the farthest transmission path minimum energy loss for complex space pipelines was proposed. Based on the clarification of the connotation and determination basis of such a criterion, a finite element model of complex space pipeline system was created, and the division of vibration transmission path and the prediction of vibration power flow curve corresponding to different transmission paths were also achieved. A complex pipeline vibration transmission testing platform was established. By comparing the calculation results between tests with finite element predictions, it was found that the maximum calculation error for the first three natural frequencies was 3.5%, both mode shape results matched well, and the variation trend of the vibration power flow curves at the output end of the pipeline system under different transmission paths obtained by the two methods demonstrated a good consistency, with the maximum error of the power flow peak only reaching 12.9%, which verified the correctness of the model. In addition, such a criterion can be employed to rank the farthest transmission path of the vibration of space pipelines, and after combining with the power flow loss results obtained under different resonance states, the main vibration transmission path can be effectively identified. The research results can provide a new ideal and approach for the vibration reduction, isolation, and avoidance of complex space piping systems in aero-engines.

Given the lack of analysis methods for thermal-fluid coupling in high-speed gear transmission systems with complex oil multi-branch lubrication, a thermal-fluid analysis model based on the multiphase flow finite volume method was proposed. The flow field and temperature field within the gearbox were studied with this model. Bending and abrupt contraction of the oil route caused significant pressure drops. This affected the jet velocity and oil-air ratio at the outlet, leading to fragmentation of high-speed gear meshing jets and lower lubrication and cooling performance. It was found that the initial structure of the gearbox suffered from jet fragmentation and oil storage issues, resulting in a windage loss of 12.30 kW, which was 85.20% of the total loss. Improvements to the oil system increased the jet velocity and oil-air ratio. By adding oil outlets and installing shrouding, the system’s windage loss was significantly reduced to 5.22 kW, increasing the transmission efficiency from 93.54% to 96.67%. It provids methodological support for the high-power-density design of high-speed aviation transmission.

In response to the problem of the limited inter-stage buffer volume of the airborne micro high-pressure compressor leading to a decrease in compressor discharge flow, the effects of the inter-stage buffer volume and the phase angle between the front and rear stages on the compressor discharge flow were explored based on a multi-stage compressor thermodynamic simulation model. A calculation method for the compressor cylinder volume coefficient considering the effect of the inter-stage buffer volume was proposed, and the relative inter-stage buffer volume was defined to describe the variation pattern of volume coefficient with the inter-stage buffer volume. The existence of an optimal size of inter-stage buffer volume was proved, which supported the design process of airborne micro compressors. The research results were as follows: the main reason for the decrease in compressor discharge flow came from the inter-stage buffer volume between the 1st and 2nd stages, which affected the volume coefficient of the 1st stage cylinder; and the extent of the influence was determined by the phase angle between the 1st and 2nd stages. In the case of the constant inter-stage buffer volume, as the phase angle between the 1st and 2nd stages increased, the volume coefficient of the 1st stage cylinder initially increased and then decreased. When the phase angle between the 1st and 2nd stages was 0°—45° or 270°—360°, the volume coefficient of the 1st stage cylinder was 0.626 at the minimum. And when the phase angle was closer to 180°, the volume coefficient of the 1st stage cylinder was closer to the designed value of 0.733; with the enlargement of the inter-stage buffer volume, the volume coefficient of the front stage cylinder increased rapidly at first and then slowly, gradually approaching the designed value; the optimal range for the relative inter-stage buffer volume was between 1 and 2, which can minimize the compressor design costs while ensuring the volume coefficient of the cylinder.

To address the challenges of collaborative difficulties and multi-disciplinary simulation integration in design, a Model-Based Systems Engineering (MBSE) methodology was adopted to construct a Requirements-Function-Logic-Physical (RFLP) systems engineering framework and develop an efficient decomposition-synthesis model for system-component relationships, enabling streamlined modeling and management. Through the development of requirement item management, Systems Modeling Language (SysML) visualization modeling tools, and component-embedded simulation solutions supported by integrated performance solvers, a multi-disciplinary collaborative design platform was established, which achieved a complete forward design process for aero-engines based on MBSE and built a Browser/Server (B/S) architecture environment supporting multi-institute, multi-disciplinary, and multi-team collaboration, providing an efficient and reliable solution for aero-engine design. The platform supports lightweight visualization, management, and traceability of three-dimensional structural models, and implements a forward design workflow through departmental permission management and approval processes. Using a high-bypass dual-variable cycle engine as a validation example, the platform improved design accuracy and reduced redundant work; SysML visual modeling provided a solid foundation for subsequent simulations; multi-disciplinary simulation integration and multi-scheme comparison accelerated design iteration speed; and the three-dimensional structure management module enhanced model review and optimization efficiency.

To solve the problems of poor robustness and difficulty in adding constraints of the inverse design method based on Gappy POD （proper orthogonal decomposition）, an inverse design method combining particle swarm optimization algorithm and Gappy POD was developed. For reducing the accuracy dominated by the overconstrained state, a pressure surface correction iteration method based on CFD calculation was proposed to make the inverse design out of the overconstrained state. The verification of subsonic and transonic compressor blade profiles showed that this method can achieve high-precision inverse design under the maximum thickness constraint, and the calculation time was shorter than the forward optimization design method. The modified design results combined with the control of key flow areas showed that the overall incidence angle performance of the two blade profiles were significantly improved compared with the original blade profiles. The static pressure rise of the blade profiles was basically unchanged at the design incidence angle, and the total pressure loss was reduced by 9.53% and 12.7%, respectively. At the same time, the available incidence angle range of subsonic blade profile remained unchanged, while that of transonic blade profile was expanded by 14.3%.

To study the influence of combustor length on swirl combustion instability, experiments were conducted in a model combustor based on a gas turbine design under different equivalence ratios. High-frequency pressure sensors and photomultiplier tubes were used to simultaneously measure the pressure oscillations within the combustor and heat release oscillations characterized by CH*. High-speed photography and laser-induced fluorescence were employed to capture the flame structures. Results showed that the combustor length had a minor effect on the time-averaged shape of the flames; stable flames tended to form a “V” shape, while unstable flames tended towards an “M” shape, with the “M” shaped flames covering a larger area. The combustor length affected the main resonant frequency and pressure amplitude of unstable combustion. Shorter combustors had higher main resonant frequencies and lower pressure amplitudes, while pressure oscillations were coupled with and heat release oscillations. Unstable flames experience flashback on both sides, with swirls in the outer shear layers exerting a “pulling” effect on the flame, leading to surface wrinkling and ultimately forming “island” distributions. The overall heat release intensity was associated with the flame shape, and periodic thermo-acoustic oscillations caused the flame shape to switch between “V” and “M” forms.

The dynamic pressure analysis capability of PSP (pressure sensitive paint) technology at high spatial resolution was constrained by multiple factors. Reliable dynamic data feedback was urgently required for its technological iteration. A multi-frequency pressure generator was designed and constructed based on the theory of rectangulawar resonant cavity. The spatiotemporal pressure of the first five low-frequency standing wave modes was numerically and experimentally studied using frequency domain analysis and PSP technology. The results showed that the installation position of the sound source could affect the excitation effectiveness of the standing wave mode. All theoretical standing wave modes can be excited while the installation position was located near the top corner of the wall. The real-time dynamic PSP results were severely affected by time-domain noise, resulting in poor reliability of quantitative results. The pressure measurement results based on phase averaging demonstrated that the constructed multi-frequency pressure generation system effectively achieved the excitation of standing wave modes and the formation of corresponding spatiotemporal pressures at the design frequency. Under single sound source input, spatiotemporal dynamic pressure with an amplitude of approximately 0.8 kPa can be generated. The ability of the typical PSP dynamic measurement system was verified, achieving a spatial resolution of 4 points per mm² and a measurement accuracy of 50 Pa in a dynamic pressure environment with a frequency of 1192 Hz.

Numerical simulation methods were employed to quantitatively and qualitatively explore the impact of endwall movement on the aerodynamic performance and tip flow characteristics of tandem diffusion cascades from the perspectives of entropy production rate, blockage factor and kinetic energy component of leakage flow. The main conclusions were as follows: firstly, after the endwall movement, the overall loss of tandem cascade was reduced, the blockage in tip area was intensified, and the flow turning angle was decreased, while the lag angle was increased. In the range of −4° to 4° angle of incidence, the loss was reduced by more than 3.9%, and the blockage was increased by more than 30.4%. Secondly, the endwall movement increased the leakage flow of the front and rear blades by 12.3% and 9.9%, respectively, but decreased the gap jet flow. Furthermore, the endwall movement enlarged the ratio of the kinetic energy of the leakage flows of the front and rear blades, with the secondary flow kinetic energy becoming dominant. It also increased the loads on the front and rear blades and made the position of the maximum pressure difference move forward in advance. As a result, the morphology and development of the leakage vortices were changed and the formation of the jet vortices was suppressed. In addition, the endwall movement significantly weakened the endwall shear effect and the mixing effect of the gap jet, expanded the circumferential influence range of the leakage flow of the front blade and caused secondary leakage. This resulted in a significant reduction in the entropy production of the front blade, while the change in the entropy production of the rear blade was relatively small. Moreover, the blockages of both the front and rear blades were intensified. The impact of the endwall movement and the increase in the angle of incidence on the front blade of the tandem cascades was greater than that on the rear blade, which was mainly the result of the regulating effect of the gap jet.

The steady transonic aerodynamic characteristics of supercritical airfoil RAE2822 and CHN-T2 in heavy gas and air were calculated by CFD. Based on the transonic similarity law and critical specific heat ratio, aerodynamic data in heavy gas medium were converted to data in equivalent air medium. The results showed that the surface pressure distributions in these two media were in good agreement with each other. For flutter tests, the slope of the lift coefficient curve and pitch moment coefficient curve at small angles of attack raised concern, after similarity correction, the slope values obtained in heavy gas medium was less than 2% different from that in air medium, which met the requirements of flutter test.

To investigate the application of aluminum nano-particles (Al NPs) as high-energy additives in hydrocarbon fuels, experiments were conducted using n-decane, oleic acid (OA), and Al NPs as fuel, surfactant, and metal additive, respectively, to prepare single n-Al/decane droplet. The effects of Al NPs mass fraction (1%—15%) and laser power (9.1—100.0 W) on the droplet evaporation and combustion characteristics were studied under infrared laser excitation at a wavelength of 1064 nm. The results showed that the addition of Al NPs significantly accelerated droplet evaporation and resulted in micro-explosions and combustion phenomena. Micro-explosions were the key to ignite the n-Al/decane droplet. The ejected Al NPs were ignited by laser in the air, forming a flame that promoted the combustion of fuel droplet. By increasing the laser power and aluminum content, the temperature rise of droplet can be accelerated and the frequency of micro-explosions can be increased, thereby improving the evaporation rate of droplet. As the mass fraction of Al NPs increased from 2.5% to 15%, the ignition sensitivity of droplet was significantly enhanced, and the minimum laser power required to ignite droplet was reduced from 83.3 W to 28.0 W.

Cracks may be produced on the exhaust windows of turbine trailing edge. In order to solve the crack failure of the single crystal alloy turbine blade in practical project application, multiple manufacturing factors leading to premature cracking of exhaust windows were summarized. The crack fracture properties were identified by macro and micro observation utilizing video microscope, SEM and energy spectrum analysis. The results showed that the exhaust window cracks were originated from the partition wall. The groove defect on the partition wall was due to the ceramic core which was not polished before casting. The higher strength of ceramic core had a strong impact on the recrystallization grain of the partition wall, which was related to the excessive casting residual stress. The “groove crack” on the partition wall was resulted from the corrosion liquid attacking the interdendritic region. The difference in surface potential between the dendrite and interdendritic region played a critical role on the surface morphology and corrosion mechanism. Casting flaw on the partition wall was due to the interface reaction between superalloy and ceramic core. The fatigue crack was produced due to lack of auxiliary support during the grinding process of the saw tooth shaped shroud.

To realize strain measurement accuracy at different scales, a multi-scale full-field strain measurement method based on digital image correlation (DIC) was proposed. A multilayer speckle preparation method was developed, and the quality of speckle was verified by numerical simulation test. Then, the out-of-plane displacement was quantified by wedge plane model, and a false strain correction method based on lens imaging model was proposed. Finally, the calculation methods of out-of-plane displacement and the method of false strain correction were verified by relevant experiments. The results indicated that the average error of the surface displacement calculation method was less than 4.20%; compared with the traditional DIC method, the accuracy of strain measurement by using the false strain correction algorithm was improved by more than 8.16%, achieving high-precision measurement of multiscale full-field strain.

In order to meet the needs of active control of rotor vibration of high performance aero-engine, an active control model was established based on the active elastic support/dry friction damper, and a switch controller and a PI (proportional-integral) controller were designed according to the modal control strategy. Then, simulation analysis and experimental verification were carried out with a certain type of aero-engine high pressure rotor tester as the object. The results showed that the switch control and PI control methods designed according to the modal control strategy effectively reduced the vibration of the rotor passing through critical speed by up to 96.17%. In addition, compared with switch control method, PI control method was more conducive to rotor stability, and it can achieve control objectives at lower cost, thanks to its control ability to react in real time to vibration under various working conditions.

In response to the need for crack monitoring in high-speed rotor blades, the blade tip timing technology for high-speed rotor blades was investigated using non-contact measurement. A method for reconstruction under sampled blade tip timing signals based on compressed sensing was proposed. Based on the time-frequency sparse characteristics of blade dynamic frequency, the under sampling signal model was improved, and adaptive block orthogonal matching tracking method was adopted to solve it, in order to monitor the change rule of blade dynamic frequency with rotational speed under the condition of variable rotational speed of the blade. High-cycle fatigue tests on high-speed rotor blades were conducted to measure both blade dynamic strain and blade tip vibration signals. The time-frequency characteristics of vibration signals from both intact and cracked blades were compared at different speeds. The presence of cracks led to a shift in blade vibration frequency, and the frequency shift allowed for early crack diagnosis. When a crack occurred, the natural frequency decreased by 24.5 Hz. The proposed monitoring method showed a dynamic frequency error of less than 0.50% when compared with strain gauge results, demonstrating high signal reconstruction accuracy and crack identification rate. This method can provide an effective solution for health monitoring and early fault diagnosis of rotating blades.

In order to investigate the influence of different injection structures on the fuel mixing characteristics inside the annular combustion chamber of a rotating detonation engine, numerical simulation analysis of the cold-state mixing flow field was conducted using the commercial software FLUENT, and the three-dimensional N-S equations were solved using a density-based scheme. Based on the annular gap-small orifice injection structure taken as the basic model, and under the same inlet conditions, the impacts of fuel injection angle and position on the mixing effect of the cold-state flow field were analyzed. The research results indicated that changes in the injection angle affected the mixing effect of methane-oxygen and total pressure loss. Within the range of 60° to 180° for the injection angle, 60° exhibited the best mixing effect and minimum total pressure loss. As the methane injection position moved forward, the distance between the injection point and the combustion chamber increased, and the mixing time lengthened, leading to an enhancement in the mixing effect. Under non-premixed conditions, due to the influence of lateral methane jets, a complex and varied vortex structure was generated in the expansion section, enhancing the mixing effect of methane-oxygen.

To investigate the fatigue properties of nickel-based alloy fabricated by additive manufacturing and heat treatment as standard in the aerospace field, the microstructural tests, tensile tests, and high-cycle and very high-cycle fatigue tests at stress ratios R=−1 and 0.1 were carried out on as-built and solution aging specimens. The test results showed that the tensile and fatigue properties of the additive manufactured nickel-based alloy were significantly improved after solution aging treatment. Surface failure was the main fatigue failure mode of additive manufactured nickel-based alloy at room temperature. The location of crack nucleation could be shifted from the surface of the material to the interior at the long-life regime of R=0.1. The improvement of tensile and fatigue properties was mainly attributed to the precipitation of γ', γ'' and δ phases. Under the synergistic effects of these strengthening phases, the matrix was significantly strengthened, dislocation movement was inhibited and crack growth was hindered. Defects and incompatible local grain orientation can lead to internal crack nucleation. Microcracks in the internal crack nucleation area propagated under shear force in a grain transgranular manner, and the grains fracked to form facets. Based on the self-strain energy theory, a life prediction method was proposed, and the prediction results were in good agreement with the test results.

Taking the subsonic airfoil of a turbofan engine fan blade as the research object, numerical simulation of the original airfoil and two leading edge erosion airfoils at five atmospheric heights was carried out with reference to the Committee on Extension to the Standard Atmosphere (COESA) standard atmospheric model, so as to explore the influence of leading edge erosion on the flow characteristics of subsonic cascade in high altitude and low Reynolds number environment. The results showed that at 0° angle of attack, the low Reynolds number condition weakened the sensitivity of the airflow to the leading edge morphology, so that the pressure distribution at the leading edge, the flow separation and transition of the boundary layer on the suction surface, and the separation degree of the trailing edge of the three airfoils tended to be consistent. Finally, the increase of total pressure loss caused by the leading edge erosion became smaller with the decrease of Reynolds number, and the total pressure loss coefficient and pressure ratio of the three airfoils tended to be consistent. At the angle of attack of 4°, the leading edge erosion can promote the occurrence of transition at low Reynolds number and reduce the degree of trailing edge separation, so that the total pressure loss was less than 0° angle of attack; when the inlet Mach number was 0.6, it reduced by 23.6%, and when 0.8, it reduced by 41.2%.

To reveal the distribution pattern of residual stress in high-temperature alloys after electron beam welding and heat treatment, simulation and experimental research of the residual stress field during the electron beam welding and subsequent heat treatment process were conducted. A dual ellipsoidal and conical heat source model was used to simulate the thermal input of the electron beam during welding, which matched well with the actual weld pool morphology. The simulated residual stress in the joint exhibited a symmetrical “M”-shaped double-peak distribution along the weld center, with a maximum residual tensile stress of 268 MPa occurring approximately 3 mm from the weld centerline. Based on an ABAQUS UMAT subroutine, a heat treatment simulation method considering solid-phase transformations was developed to simulate the residual stress distribution after heat treatment. The results indicated that the peak residual tensile stress in the welded plate decreased by 48% after heat treatment, while the location of the peak remained unchanged. Furthermore, XRD measurements were conducted on the actual welded plates after welding and heat treatment. A comparison with the simulation results showed that the distribution pattern of residual stress near the weld matched well with the experimental results, with a simulation peak error of no more than 3% for welding residual stress and no more than 11% for post-heat treatment residual stress, thereby validating the effectiveness of the simulation method.

The needle gas valve is a critical component for energy management in solid-propellant attitude and orbit control rocket engines. Needle gas valves can be classified into subsonic and supersonic configurations. These two valve types primarily operate in subsonic and supersonic flow regimes, respectively, and exhibit distinct working characteristics and application scenarios. The CFD numerical simulation method was employed to analyze flow losses, valve stem loads, and thermochemical ablation of both valve configurations under multiple opening levels. A comparative analysis of the advantages and disadvantages between these two structural designs was also conducted. The results indicated that flow losses for both valve configurations decreased with the increasing opening levels. Under identical operating conditions, the supersonic gas valve exhibited higher flow losses compared with the subsonic configuration. At different opening levels, the supersonic valve demonstrated more stable load forces on the valve stem, with significantly lower loads at large openings. Additionally, the structural dimensions of the valve stem exerted a substantial influence on the valve load distribution. The thermochemical ablation rate at the subsonic valve stem was five times higher than that of the supersonic valve stem. Conversely, the ablation rate at the throat of the supersonic valve chamber was 1.5 times greater than that observed in the subsonic valve.

Based on the principle of turbine vanes and the mechanism of flow and heat transfer, a fluid-thermal-solid coupling calculation system is developed through secondary development, incorporating ANSYS Fluent, ANSYS APDL, and Flowmaster. This system achieves the invocating of different programs, automatic calculating, data exchanging between different computational domains, and monitoring of calculation results. Based on this integrated simulation software, the flow phenomena in the blade channels and heat transfer phenomena of oil-cooled turbine vanes with different cooling channels aspect ratios are analyzed. The results show that the structural and layout of cooling channels significantly influence the cooling effectiveness of oil-cooled turbine vanes. Moreover, as the aspect ratio of U-shaped cooling channels increases, the oil-cooled turbine vanes have a great impact on the temperature distribution.With the increase of the aspect ratio of U-shaped channels, the cooling efficiency of the turbine vanes decreases, but the reduction is minor; oil cooling structures with smaller aspect ratios exhibit more uniform temperature distribution on the turbine vanes outer walls and a larger cooling coverage area, while oil cooling structures with larger aspect ratios have a more concentrated cooling effect and a smaller cooling coverage area.

In response to the complex coupling relationship of multiple parameters in the combustion chamber, the influence of the coupling relationship between different head expansion angles of flame tubes and swirle on combustion performance was studied. The research results showed that: to 90° expansion flame tube, with the increase of swirl number, the fuel mixing uniformity gradually deteriorated, the outlet temperature distribution coefficient gradually increased within a wider stable working range; the oil-air ratio of lean blow out gradually decreased, the carbon monoxide (CO) and Unburned hydrocarbon (UHC) emission index decreased, while the nitrogen oxides (NOx) emission index increased. To 45° expanding flame tube, with the increase of swirl number, the fuel mixing uniformity gradually improved, the outlet temperature distribution coefficient gradually decreased, and the NOx emission index decreased.

The influences of different rotating speeds of the rotatable ring of the controllable speed casing on the stability of the low-speed axial flow compressor were studied by numerical simulation. The results showed that the controllable speed casing, which covers the whole axial chord length region of the rotor tip and rotates in the same direction with the rotor, can control the tip leakage flow and realize the stability expansion. By applying external shear stress to the clearance flow, the controllable speed casing increased the mainstream momentum, suppressed the momentum ratio of the leakage flow to the mainstream, improved the deflection of the leakage vortex, and delayed the occurrence of secondary leakage and the forward movement of the mainstream/leakage flow interface, thus broadening the stable operating range of the low-speed compressor stage. While ensuring that the pressure ratio was basically kept unchanged, the stability expansion effect was enhanced with the increase of the rotatable ring speed. When the rotatable ring speed was the design speed of the rotor, the maximum stable operating margin of the low-speed compressor stage can be increased by 30.86%.

In order to improve the accuracy of engine starting performance calculation based on the component method, a starting throttling inverse calculation method based on the power extraction method was proposed. It clearly defined the power balance relationship of the starting process, and established the nonlinear equations of the engine starting, so as to reversely evaluate the combustion efficiency based on the fuel flow, rotor acceleration rate and starter characteristics during the starting process. According to the similarity principle, the calculation method of converted torque scaling factor was proposed, the conversion between efficiency characteristics and torque characteristics was completed, and the effectiveness of the method was verified. Furthermore, a pre-ignition starting model was established, enabling the simulation of starting from zero to idle speed. The simulation model was validated based on the experimental data of a micro-turbojet engine, and the results showed that the maximum errors in the rotational speed and P3 during the starting process were 2.81% and 2.22%. This method allowed for an approximate simulation of the starting process and enabled the reverse calculation of combustion efficiency during that process. In addition, the plausibility of the sub-idle component characteristics can be tested. This method can provide a reference for modeling and validation of starting performance for other types of aerospace engines.

The potential coupling mechanism of transverse combustion instability in an O₂/CH₄ rectangle rocket combustor was investigated experimentally and numerically. High-frequency pressure sensors were employed to capture the dynamic pressure characteristics within the combustion chamber. Numerical simulations were conducted using the stress-blended eddy simulation (SBES) and flamelet-generated manifolds (FGM) methods. The results showed that the numerical simulation successfully predicted the transverse combustion instability observed experimentally. The numerical results, including pressure waveform, main frequency, and root mean square amplitude, were in good agreement with the experimental data and theoretical analysis. A comprehensive analysis of the driving mechanism and Rayleigh index of the transverse combustion instability was presented. Examination of the flow field characteristics in the combustion chamber revealed that the transverse pressure wave induced periodic ‘stripping’ of the oxygen jet and vortex disruption. Heat release pulses were generated during vortex fuel combustion and coupled with the first width (1W) mode. Rayleigh index analysis indicated that edge injectors exhibited stronger driving characteristics for transverse combustion instability compared with central injectors.

An amplitude-identified multiple signal classification (MUSIC) algorithm for blade tip timing (BTT) signals of rotor blades was proposed for identifying surge characteristics of compressors. By analyzing the real-valued sinusoidal signal model of BTT, the relationship between the amplitude and frequency of asynchronous vibrations and the eigenvalues and eigenvectors of the autocorrelation matrix was explored, allowing for compensation of the amplitude of the pseudo-spectrum within the framework of the MUSIC algorithm. The experiment of rotor surge identification using BTT technique was conducted. By comparing with fast Fourier transform and the least squares fitting method, the effectiveness of proposed method was verified. The results indicated that surge faults exhibited a significant increase in amplitude and speed fluctuation in the time domain. In the frequency domain, surge faults exhibited low-frequency asynchronous vibration with large amplitude. Compared with Fourier transform and least squares fitting, the proposed method can accurately identify the surge characteristic frequency of 5.3 Hz, and the focusing capability of its amplitude identification was more than three times that of the Fourier transform. The proposed method with higher frequency resolution and amplitude recognition accuracy can effectively extract the characteristics of surge faults in the frequency domain.

To investigate the effect of real coupling deviation on the aerodynamic performance of a blade cascade, chord length deviation, thickness deviation, and leading edge radius deviation were applied to a high load compressor blade cascade for research based on the arbitrary polynomial chaos method of moments. The research results showed that compared with the prototype, the probability of an increase in total pressure loss coefficient and a decrease in static pressure coefficient under negative angle of attack conditions was about 79.47% and 92.83%, respectively; under positive angle of attack conditions, the probability of an increase in total pressure loss coefficient was about 91.11%, and the probability of a decrease in static pressure coefficient was about 86.42%. The aerodynamic performance under different operating conditions was most sensitive to the deviation of the leading edge radius. Combined with the analysis of the loss source, it was found that the leading edge loss played a dominant role. Therefore, strict control of the leading edge machining accuracy was required during machining. Compared with the prototype blade, under negative attack angle conditions, the degree of separation in the blade angle region with added coupling error changed significantly, and the separation point, recirculation area, and axial length of the recirculation area increased significantly; under positive angle of attack conditions, the degree of separation in the blade angle region with added coupling error was not significant.

Electric propulsion, with its high specific impulse, efficiency and long lifespan, has emerged as one of the main space propulsion methods following cold gas and chemical propulsion. Water-fueled electric propulsion, thanks to its non-toxic, pollution-free and low-cost benefits, has increasingly gained attention. A comprehensive summary and evaluation of the current development status of water-fueled electric propulsion technology were conducted both domestically and internationally. Based on the method of thrust generation, it can be classified into three types: electrothermal, electromagnetic, and electrostatic. Result showed that, electrothermal propulsion achieved significant progress in ground experiments and in-orbit demonstrations, making it the most mature water-based propulsion technology. Water-based electromagnetic and electrostatic propulsion technologies are still in the developmental stage but are progressing steadily. Future advancements in water-fueled propulsion technology are expected to focus on the integration of chemical and electric propulsion systems, multimodal propulsion designs, and their application in deep space exploration. With technologies becoming more mature, water-fueled propulsion technology will develop in directions that integrate chemical and electric propulsions.

The fatigue fracture occurred during the low speed performance test of swept-forward fan blades. The cause was investigated to improve the design. The damage caused by blade resonance and flutter was excluded through simulation analysis. The high-cycle fatigue analysis showed that the second-order modal danger point of the fan blades was consistent with the crack initiation position. The analysis suggested that the unsteady flow at the tip generated 6.5 times the rotational frequency exciting force, which stimulated the second-order natural mode of the fan blades. Blade tip amplitude and dynamic pressure measurement experiments verified the analysis results. At the fracture speed, the dynamic pressure signal exhibited multiple non-synchronous frequencies. The amplitude of the fan blade tip was 3.7 mm, and the vibration stress at the crack initiation position was 671.2 MPa. The fan blade skimmed forward greatly, resulting in low local bending mode frequency of the forward-swept part, which was easy to be resonated. Blade improvement measures were taken to remove partial forward-swept of the blade and eliminate the local bending mode. The experiment verified the effective improvement measures. This scheme is of reference significance for the design of forward swept small aspect ratio fan blades to avoid asynchronous vibration.

The theoretical model of the windage resistance heating characteristics of the bolt structure in the rotor-stator cavity was analyzed. A numerical solution model for windage resistance heating characteristics of the bolt structure in the rotor-stator cavity was established to analyze the internal flow field characteristics of the rotor-stator cavity and study the influences of structural and operating parameters on windage resistance heating of the bolt structure in the rotor-stator cavity on the basis of verifying the accuracy of the solution model. The theoretical formula for windage resistance heating of the bolt structure in the rotor-stator cavity was constructed using the correction coefficient method. The research results indicated that when the bolt rotated with the rotating disc, the surface of the bolt interacted with the viscous airflow through friction, and the temperature of the gas increased, resulting in windage resistance heating effect. Under the working conditions studied in this article, when the speed increased from 6000 r/min to 15000 r/min, the wind resistance temperature rise coefficient of the bolt structure increased by 24.2%; the inlet flow rate increased from 2.2 g/s to 8.0 g/s, and the wind resistance temperature rise coefficient of the bolt structure decreased by 51.5%; when the number of bolts increased from 6 to 36, the wind resistance temperature rise coefficient of the bolt structure increased by 65.1%; the theoretical formula for windage resistance heating was caused by the bolt structure of the rotor-stator cavity, which can accurately calculate windage resistance heating caused by the bolt structure of the rotor-stator cavity. This article has provided a theoretical basis for the analysis of windage resistance heating characteristics of the bolt structure of the rotor-stator cavity.

An automatic transition control method considering aircraft lift and thrust characteristic was proposed, and the effectiveness was evaluated by measurement of pilot workload. The lift and thrust between rotor and wing in conversion corridor were analyzed through steady-state calculation, and a desired conversion path and nacelle rate distribution in different phases were planned. The collective pitch and angle of attack were introduced to adjust the rotor lift and thrust and wing lift respectively to match the desired path. In the control structures of collective pitch and angle of attack, feedforward and command schedule of static path parameters were introduced to adapt to the lift and thrust characteristics of the aircraft, and the error between static planning and dynamic tilting was eliminated by combining conquer altitude feedback. A pilot model was introduced and the workload was quantified by wavelet analysis. Compared with the maneuver executed by a pilot, the automatic transition control can effectively reduce the pilot workload, the maximum energy amplitudes of the collective and longitudinal inputs were reduced by 33% and 27%, respectively. And the frequency components of both controls were reduced to below 0.8 rad/s, indicating level 1 handling quality rating. Meanwhile, the automatic transition control eliminated the dynamic error effectively, with smaller tracking error of desired path and height variation.