Air-assisted fuel injection (AAFI) technology, known for its excellent atomization and fuel adaptability, is a key solution for poor atomization of heavy fuel and mixture formation issues in high-altitude conditions. The applications and progress of AAFI technology were reviewed within low-altitude economy and single-fuel contexts, and a prospective analysis of challenges and potential trends was carried out. Result showed that, aerodynamic forces were crucial for fuel atomization and fragmentation. Spindle-shaped and anchor-type vortex structures in AAFI sprays enhanced atomization through gas-liquid interaction and shear instability, influenced by spray field speed, nozzle structure, and flash boiling. Most studies focused on macroscopic spray formation, while there was little research devoted to micro-scale gas-liquid coupling and system-level adaptability. Air supply posed a major challenge, as existing solutions increased power system weight and limited unmanned aerial vehicle (UAV) endurance. AAFI technology showed potential for dual-fuel engines and low/zero-carbon fuels, with hydrogen, ammonia, and natural gas being taken as possible gaseous media for AAFI.

The coupled modeling problem of hydrogen micro-mixed combustion in water vapor environment was investigated using a combination of laser diagnostic measurements and RANS (Reynolds-averaged Navier-Stokes) numerical simulations. The no-reacting flow field and reacting OH group components distribution in the combustor chamber were experimentally measured. Compared with the experimental results, the deviation of the no-reacting flow field was 12.5% and the distribution of OH group components in the reacting flow was closely matched with the experimental measurements by using the realizable k-ε turbulence model and flamelet-generated manifold-finite-rate (FGM-FR) model. This coupled simulation method demonstrated a relatively accurate method of hydrogen micro-premixed combustion in water vapor environment. The numerical simulation results showed that the change of oxygen mass fraction can affect the flame length, showing a quasi-linear relationship with oxygen mass fraction. The increase in oxygen mass fraction led to an increase in the local combustion rate, reduction in the flame profile area, which improved flame resistance against turbulent stretch and enhanced flame stability as well.

With the rapid development of hypersonic aircraft technology, there are higher requirements for cabin thermal insulation materials. Lightweight and efficient thermal insulation materials are the key to ensure the normal operation of electronic instruments, electrical equipment, and other components inside the aircraft. The working mechanism and thermal conductivity test method of thermal insulation materials in hypersonic vehicle cabin under the modes of heat conduction, heat convection and heat radiation were described. The research status of foam ceramic thermal insulation materials, organic foam thermal insulation materials, fiber thermal insulation materials and aerogel thermal insulation materials was introduced. The results indicated that different thermal insulation materials exhibited different thermal insulation mechanisms and heat transfer forms due to their structural differences. Generally, the thermal conductivity calculated by theoretical mathematical models is not accurate, and steady state and non-steady state methods are often used to measure the thermal conductivity. To solve current difficulties in the research of thermal insulation materials, the future development direction of thermal insulation materials in hypersonic aircraft cabins should focus on optimizing the pore structure, deepening development and design, and innovating special functions.

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.

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.

In order to help achieve the “carbon peaking and carbon neutrality goals” and realize the zero carbon emission of pure hydrogen combustion in gas turbines, a premixed micro-tube burner based on a static mixer was designed to meet the requirements of low emission of pure hydrogen combustion. Under the conditions of normal temperature and pressure, 23.7, 47.3, 71.1 m³/h air flow, the pure hydrogen combustion test was carried out on the 4×4 array of premixed micro-tube burner, with the equivalent ratio ranging from 0.25 to 0.85. The time-average macro flame image was obtained, the NOx emission was measured, the combustion oscillation characteristics in the chamber were measured with a pressure pulsation sensor, and the combustion characteristics of pure hydrogen combustion under different air flows and equivalent ratios were explored. The hydrogen-air mixing in the array micro-tube was analyzed with Fluent software. The results showed that when the equivalent ratio was within 0.7, the volume fraction of NOx emission of the burner under various conditions was within 20×10⁻⁶. The micro-mixing flame was similar to the pure premixed flame. With the increase of air flow, the velocity of the micro-tube nozzle increased, the overall length was stretched, and the flame cone angle became smaller. When the air flow was 23.7 m³/h and the equivalent ratio was less than 0.55, stable hydrogen combustion can be achieved. In other conditions, the starting point of the vibration moved forward with the increase of air flow. With the increase of the equivalent ratio, the amplitude of combustion oscillation first increased and then showed a trend of fluctuation decrease. The oscillation frequency was generally above 700 Hz, and steadily increased with the increase of the equivalent ratio.

To address the issue of graphite dust deposition within the helium turbine of high-temperature reactors (HTGR) impacting performance and maintenance, the motion and deposition behaviors of graphite dust particles on the helium turbine walls with film cooling were investigated by numerical simulations. The rebound effects between the particles and the wall, as well as the resuspension of deposited particles caused by the flow field were considered. A deposition-resuspension model for graphite dust particles was established. The results indicated that due to the shape of the turbine blades, almost no particles impacted the suction surface, leading to no particle deposition. Due to higher shear stress at the trailing edge, the deposition on the pressure surface mainly concentrated at the middle. Small particles with lower inertia were more prone to deposition under the influence of counter-rotating vortices, and their deposition rate was primarily influenced by the collision rate. Larger particles with a lower critical deposition velocity exhibited a significantly lower deposition rate, which was mainly affected by the adhesion rate. As the blowing ratio increased, the entrainment effect of the cooling film intensified, resulting in a continuous increase in the deposition rate of smaller particles.

The light-round pattern of kerosene spray flame in a triple-swirler combustor was investigated experimentally using simultaneous 10 kHz Mie scattering of droplet and OH* chemiluminescence (CL) measurements. It was shown that light-round pattern occurred in two different sub-phases, including the quasi-ignition phase and the adjacent swirling flame development phase. The linear and zigzag propagation patterns were identified in the former, and the upstream-entrainment (UE), downstream-hysteresis (DH), and multi-clusters-expansion (MCE) propagation patterns were recognized in the latter. The possibility of the light-round pattern was analyzed statistically. It was found that zigzag-MCE was the primary light-round pattern at different conditions with varying inlet air velocity, combustor pressure, and fuel-air ratio. However, under conditions near the light-round boundary, the zigzag-DH pattern appeared with a high probability of 77.8%. Similarly, under low air velocity or high fuel-air ratio, linear MCE patterns predominated.

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.

To systematically study the flame propagation characteristics and flame instability of natural gas/air premixed flames at different initial temperatures, initial pressures, and equivalence ratios, high-precision numerical simulations were carried out by using a 3D, transient large eddy simulation method with adaptive mesh refinement modeling. Firstly, simulation results at atmospheric temperature and pressure were validated. The simulated flame morphology was in good agreement with experimental data, and the simulated critical flame radius was also in good agreement with that obtained by theoretical prediction. Secondly, key parameters of natural gas/air flames under different initial conditions were analyzed, revealing the dominant mechanism by which initial conditions influenced the inherent instability of flames. Finally, perturbation analysis of the flame structure was conducted based on the Fast Fourier Transform of cross-sections, revealing the complex interaction between perturbations and stretch rates. The results indicated that different initial temperatures had little effect on the flame morphology, but as the initial pressure increased, the crack splitting on the flame front gradually intensified, and the number of cells on the flame surface increased. Increasing the initial pressure had a greater effect on the reduction of flame thickness, while the density ratios were not sensitive to changes of initial pressure. Therefore, the initial pressure had a more significant effect on the flame stability. The disturbance energy of the wave numbers ranging from 4 to 15 increased significantly with initial pressure, and the disturbances with smaller wavelengths developed consequently. By contrast, the initial temperature and equivalence ratio had little effect on the wave number range of disturbances.

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.

Based on existing dry-out heat flux model, the research model of dry-out heat flux of horizontal microcolumn evaporator with a relative error of about 7.6% was obtained by combining the capillary pressure and permeability calculation methods under the hot spot effects. The evaporator surface was divided into evaporative, non-evaporative and hot spot areas, and the effects of evaporator geometry, hot spot area and superheat on the heat flux to dryness were investigated. Calculated results showed that the maximum heat transfer capacity depended on the equilibrium relationship between capillary pressure and permeability; microcolumns with geometries close to the optimal pitch ratio (～0.25) and higher corresponded to better heat dissipation, and microcolumns with the smallest receding contact angle corresponded to an enhancement of the heat flux of desiccation by a factor of about two. The modelled superheat was negatively correlated with the hotspot area, with larger hotspot areas corresponding to smaller superheat, and larger local pressures corresponding to larger Nu and smaller superheat.

The flow field structure and characteristics in a double stage axial swirl combustion chamber of an aero-engine were experimentally tested and numerically calculated, and the flow field structure, streamline and velocity distribution characteristics on the different longitudinal and cross sections in the combustion chamber under different conditions were obtained. The results showed that there were two symmetric distribution angle recirculation zones and one central recirculation zone on the longitudinal section in the combustion chamber. With the increase of distance along the positive X-axis or the decrease of intake flow rate, the air velocity in the recirculation zone decreased. With the increase of distance along the positive Y-axis, the airflow rotation direction tended to be the same and the air flow velocity decreased on the cross section in the combustion chamber. Meanwhile, with the increase of intake flow rate, the flow field shape and flow velocity distribution characteristics of the air on the same cross section were basically the same, but the air flow velocity increased gradually. Under different conditions, on the X=0 mm longitudinal section, the axial velocity distribution at each axial distance was basically consistent. At Z=0 mm, the axial velocity reached the maximum negative velocity, and gradually decreased with the increase of radial distance. With the axial distance increasing, the recirculation zone area was gradually reduced, the axial velocity increased first, then decreased, and reached its maximum at Y=50 mm. Meanwhile, with the increase of intake flow rate, the axial velocity gradually increased at different axial distances. The calculated distribution characteristics of axial and radial velocities on the X=0 mm longitudinal section and at an axial distance of Y=30 mm were in good agreement with the corresponding experimental data under different conditions, and the relative error was less than 10%.

To address the challenge of reliably monitoring aviation engine bearing faults under complex operating conditions with limited information from a single sensor, the STFT-IncepNext model was proposed for bearing fault diagnosis. Initially, sensor data from different positions within the same time window were concatenated to enrich the vibration information of bearings across various spatial dimensions. Subsequently, to capture the transient changes of fault components in the vibration signal, the short-time Fourier transform (STFT) was applied to convert multi-sensor vibration signals into time-frequency representations. Finally, a lightweight IncepNext network extracted global features of fault information embedded in the time-frequency representations, and a Softmax classifier identified the fault category. Experimental results demonstrated that the proposed approach effectively enhanced signal fault characteristics, and improved the discriminability of vibration features under various bearing states. Under specific experimental conditions, the method achieved an accuracy rate of 100% for diagnosing vibration faults in aeroengine bearings. Compared with STFT-EdgeNeXt, STFT-ResNeXt, STFT-ShuffleNet, and STFT-ResNet18, the method exhibited superior performance, providing a feasible approach for diagnosing faults in aeroengine bearings.

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.

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 vibration displacement 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.

The contact of the spline pair between the cylinder and shaft of piston pumps affects their operation reliability. It is important to analyze the contact mechanism of the spline pair and the vibration responses of the cylinder and shaft. In this study, a finite element contact dynamic model of the cylinder and shaft was developed, and the vibration responses of the cylinder and shaft were analyzed. The finite element contact dynamic model was described. The forces acting on the cylinder were analyzed, and the excitation forces were obtained. The contact mechanism and the vibration responses at various outlet pressure conditions were simulated and analyzed. The results showed that: the contact dynamic model can simulate the contact mechanism of the spline pair and the vibration responses of the cylinder and shaft. Under the influence of time-varying piston chamber pressure, the tooth contact position of the spline changed with the varying rotational angle. The increase of the outlet pressure led to the increase in the tooth contact stress, vibration responses and shaft deflection. The deflection at the spline of the shaft was much larger than that at other positions, and the deflection decreased from the spline side to the two end sides.

A new flow tube grate tooth sealing structure model of high-altitude table was built based on the heat-fluid-solid coupling method. Simulation and experimental studies of the sealing characteristics were carried out. The influence of the design gap on the sealing performance and contact wear of grate teeth was revealed. The gap was optimized by response surface methodology. Finally, the structural performance before and after optimization was analyzed comparatively. The study showed that the error between the simulation and experimental results of the leakage volume did not exceed 9.2%. Taking the remaining clearance not exceeding 0.1 mm as the optimization condition, the front-end design clearance of 1.9 mm was obtained, the seal design clearance was 0.75 mm, and the back-end design clearance was 2.1 mm. After optimization, the leakage of grate tooth seal structure was 0.00456 kg/s, and the sealing performance was improved by 82.6% compared with that before optimization. After optimization, the maximum flow rate of the grate teeth, the maximum temperature of the structure and the maximum deformation of the structure were reduced by 18%, 2.1% and 3.6%, respectively. This study can provide a technical support for the design, installation and commissioning, and safe use of the grate seal structure of the high-altitude simulation experiment table.

A predictive model for temperature and stress at critical locations of turbine disks was developed using engine service data, with its aim to provide essential inputs for service damage assessment and life analysis of hot-section components. The research derived gas path parameters related to turbine disk loading through engine thermodynamic cycle analysis. Based on numerical simulation results, a Long Short-Term Memory (LSTM) neural network model was constructed and trained to predict turbine disk temperature and stress. The generalization of proposed model capability was evaluated through cross-validation and noise injection. Result demonstrated that the established model achieved a normalized mean square error of 5.4×10⁻⁴ on the test set. The developed methodology allowed to analyze temperature and stress evolution patterns during typical flight maneuvers, providing crucial inputs for engine mission design and turbine disk service damage assessment.

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 demonstrated that both control methods effectively suppressed rotor vibration during critical speed transitions, achieving a maximum vibration reduction of 96.17%. Furthermore, compared to the switching control method, the PI controller proved more conducive to rotor stability. It achieved the control objectives at a lower cost, owing to its superior ability to adapt in real-time to vibrations under various operational conditions.

To address high missed detection rates for small-sized damages and significant environmental variations in aero-engine blade inspection during service-induced damage, an enhanced YOLOv5-based detection method was developed. A data augmentation strategy integrating random affine transformation, HSV space enhancement, and bidirectional flipping was proposed, which effectively mitigated sample imbalance across damage categories. Second, a multi-kernel adaptive convolution module and multi-scale feature aggregation module were introduced to improve network capability in extracting fine-grained damage features and discriminating subtle characteristics, effectively addressing the technical bottlenecks of insufficient detection performance for small-sized damages with indistinct morphological features. The experimental results showed that compared with the baseline model, the detection accuracy (mAP@0.5) of the improved model increased by 16.8%. In the comparison, the detection accuracy of the improved model was higher than that of mainstream object detection frameworks such as Faster-RCNN, SSD, Deformable DETR, YOLOv8, YOLOv11, and YOLOv12. Furthermore, the proposed method exhibited 13.9% and 15.3% improvements in mAP@0.5 over existing defect detection models SW-YOLO and YOLOv8-EMA, respectively.

Foreign object damage is an important factor affecting the service of aeroengine fan and compressor blades. Small foreign object damaged notch can be typically repaired by machining or polishing. To evaluate the fatigue properties of the machined and polished blades, the high-cycle fatigue strength of the machined and polished blades was tested, and the fatigue limit prediction model was built based on the theory of critical distance, and compared with the classical Peterson model. The influences of different stress selection methods on the accuracy of the model were analyzed. The results showed that the model based on the theory of critical distance can accurately predict the fatigue limit of repaired notched blades, including a variety of repair processes, such as machining and polishing. There was little difference between the prediction using the maximum absolute value of the principal stress and the normal stress along the blade length. The model accuracy from theory of critical distance (average error less than 11%) was much higher than the classical Peterson’s model (average error 26.03%).

The dynamic characteristics of scallop damper seals play a crucial role in influencing the operational stability of the system in the presence of fluid-induced excitation forces acting on the rotor. The impedance method was employed to experimentally identify the dynamic characteristic coefficients of scallop damper seals. Experiments were conducted to investigate the variations in these coefficients under different inlet pressures (0.141, 0.181 MPa) and speeds (1 800, 2 400, 3 000 r/min). The obtained findings were subjected to a comparative analysis with numerical computations derived from computational fluid dynamics (CFD). The findings showed that the numerical results agreed well with the experimental results. The numerical errors were less than 15% under variable speed and 50% under variable pressure. The experimental damping coefficient results were slightly higher than the theoretical results, while the stiffness coefficient exhibited the opposite behavior. The direct stiffness exhibited a positive correlation with rotational speed and inlet pressure, whereas it displayed a negative correlation with vortex frequency. Conversely, effective damping showed a positive correlation with inlet pressure and vortex frequency at low frequencies (f<100 Hz) and a negative correlation with rotational speed. At higher frequencies (f>100 Hz), it became stable. Moreover, the effective damping at 0.181 MPa inlet pressure was approximately 1.6 times greater than that at 0.141 MPa inlet pressure. It was observed that rotational speed had a relatively lesser impact on the dynamic characteristics of scallop-damper seals if compared with inlet pressure.

Composite design is an important way to improve the thrust to weight ratio of aero-engines. However, due to the variability in lay-up angles and sequences, it is difficult to determine the optimal lay-up scheme for each structure in the cascade. Therefore, the design of composite material reverse thrust cascades was carried out. The sample points were designed by genetic algorithm, and finite element analysis of composite structures with different lay-up schemes was performed. The surrogate model was established, the deformation, strength and stiffness under the working load of each characteristic structure in the reverse thrust cascade were taken as the targets, and finite element analysis was carried out for different lay-up schemes after optimization, and the optimal lay-up scheme was obtained by comprehensive comparison. The composite reverse thrust cascade models were modeled according to the optimal lay-up scheme and compared with the aluminum alloy reverse thrust cascade model. Result showed that the gap with the aluminum alloy cascades was only about 11%, verifying the superiority of the selected mixed lay-up 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.

In view of the existing problems and 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 was 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, 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 were verified. After actual testing, the weight of the UAV was 37.95 kg, the payload capacity was 10 kg, and the endurance time after loading the payload task system can reach 4 hours and 20 minutes, which can meet the expected design effect.

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.

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.

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.

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.

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.

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.

In view of 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, a method for optimizing the overall parameters of turboshaft engines was proposed based on the requirements of helicopter flight missions, and the feasible region of design parameter optimization was reasonably limited according to the development of turboshaft engines. The method consisted of four modules: constraint analysis, task analysis, parameter cycle calculation and optimization iteration. The optimization iteration module involved dynamic adjustment of design parameters, which made the designed engine more matched with the task requirements. Firstly, the basic principles of the above four modules were derived. Then, based on the given helicopter flight mission profile, the proposed design method was verified by an example. Finally, the results were compared with other design methods for the same flight mission. The research showed 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 made up for the shortcomings of other overall design methods of the turboshaft engine, for example, the design parameters were not involved in the iterative update and the lack of consideration of the coupling relationship between the helicopter / turboshaft engine performance.

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 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.

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%.

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.

To address the accuracy problems of standard SA (Spalart-Allmaras) turbulence model when predicting complex rotating flows in aviation compressors, how the SA-helicity modified turbulence model performed in rotating flows was analyzed. It showed how the helicity correction helped compensate for strain rate through eddy viscosity mechanisms, leading to better predictions of corner separation and tip leakage flow while maintaining good boundary layer accuracy. Three turbulence models: standard SA, SA-helicity, and SA-QCR (quadratic constitutive relation), were tested, using NASA Rotor 67, a 3.5-stage axial compressor, and a publicly released high-load single-stage axial compressor from AECC Shenyang Engine Research Institute. Their performance at different rotational speeds was analyzed by focusing on performance characteristic curves, tip stall prediction, secondary flow prediction, and near-stall point prediction. Results showed that both the SA-helicity and SA-QCR models accurately predicted the trends of the characteristic curves. The SA-helicity model increased eddy viscosity in the corner separation region under adverse pressure gradients, thereby suppressing excessive corner separation. It also reasonably predicted the channel secondary flow induced by root separation, leading to an increased pressure ratio in the lower half of the blade. Compared with the QCR model, the helicity model better predicted the development and extent of tip leakage flow, significantly expanding the flow range, pressure ratio and efficiency, and showing better agreement with experimental results. Furthermore, the boundary layer flow predicted by the helicity model resisted adverse pressure gradients better, resulting in separation bubble reattaching earlier, thereby enhancing diffusion capability near the tip region and expanding the pressure ratio and flow range. Overall, the SA-helicity model overcame the limitations of traditional turbulence models in predicting how rotating secondary flows interacted with main flows in complex compressor environments.

To carry out the aerodynamic optimization design and the low-noise-characteristic design of the contra-rotating propfans in the aerodynamic design inverse problem phase, a program FODNOPPa2, specifically for predicting the discrete frequency noise, was recently developed. Based on the formulation Formu 1A derived by Farassat, this program offered the time-domain solutions of the Ffowcs Williams-Hawkings equation. By utilizing this program, the discrete frequency noise of a model contra-rotating propfans called CRPFt, which was designed using compressible lifting surface theory, was acquired under the take-off condition with a flying altitude 300 m and a Mach number 0.24. Observation points were synchronized with the movement of the fuselage and distributed across five representative lines or curves. Utilizing the nonlinear harmonic solver available in the computational fluid dynamics software NUMECA FINE™/Turbo, it tried to acquire a three-dimensional viscous quasi-unsteady flow field and to extract the aerodynamic loads as the input of the noise calculation. This study demonstrated that the nonlinear harmonic approach was capable of effectively simulating the three-dimensional viscous quasi-unsteady flow field of CRPFt under the take-off condition. Typically, the sound pressure level (SPL) of the rotor harmonics held a greater significance than the SPL of the interference harmonics at the observation points situated in close proximity to the middle plane between the two propfans. Nevertheless, the SPL of the interference harmonics had more significance when measured at the observation points located further away from the middle plane.

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%.

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.

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.