In order to study the convective heat transfer characteristics inside the double-wall composite cooling turbine blade, a narrow semi-closed passage with lateral outflow through sparse holes was designed. Test methods were used to study the effects of different inlet Reynolds numbers and geometric parameters of outflow holes on the convective heat transfer characteristics of the outflow wall surface. The results showed that, the superposition of overflow effect and impact effect occurred in the downstream area of the outlet hole due to the outflow of cool air. A typical droplet cryogenic zone appeared in the downstream area of the outlet hole, and its coverage area increased with the increase of the outlet hole size and inlet Re number; the average Nusselt number of the outlet wall presented four variation characteristics along the flow direction. The average Nusselt number of the inlet section was 80% higher than that of the downstream section; there was an optimal outlet hole span spacing ratio, so that the average convective heat transfer coefficient on the wall reached the maximum. The average wall Nusselt number of the major hole spacing was 20% lower than that of the intermediate hole spacing.

In considering the deficiency of adjusting fuel injection amount or ignition advance angle to suppress knock on two-stroke kerosene engine, the coordinated injection-ignition control strategy for suppressing engine knock was proposed. A two-stroke spark ignition engine was taken as the test engine for comparative study, and its modeling and simulation analysis were carried out by using one-dimensional simulation platform. The MAPs of fuel injection amount and ignition advance angle were obtained by the method of design of experiment (DoE) after adopting the coordinated control strategy. The corresponding experimental investigations were conducted. The results showed that, at engine speed of 4800 r/min, compared with the single parameter of adjusting ignition advance angle to suppress knock intensity of kerosene engine, the optimized strategy was more effective. The specific fuel consumption showed a modest increase with reduced power loss and decreased exhaust temperature. The power recovery of kerosene engine can reach at least 88.3% of the burning gasoline at different loads; meanwhile, the power recovery can reach 96.2% under the condition of full load at engine speeds of 5000—6500 r/min, and the exhaust temperature can be controlled within 475 ℃, effectively promoting the performance of kerosene engine.

The nitrogen-rich air distribution method was studied by using statistical theory, the Entropy-weight improvement technique for order preference by similarity to an ideal solution (TOPSIS) method was applied to establish a comprehensive evaluation method of oxygen concentration decrease rate and uniformity. On this basis, five typical inerting schemes were designed for the wing multi-compartment fuel tank of a transport aircraft. Through numerical simulation method, the above schemes were modeled and calculated, the characteristic indexes of multi-compartment fuel tank were obtained, and the established comprehensive evaluation method was used to evaluate each scheme. The results showed that: (1) the entropy-weight improvement TOPSIS method can effectively evaluate the fuel tank inerting performance and realize the determination of the optimal nitrogen-rich air distribution mode; (2) the semi-uniform inlet distribution mode of nitrogen-rich air was the best when considering the characteristics of oxygen concentration decrease rate and uniformity; (3) when the oxygen concentration decrease rate was taken as the only evaluation index, the non-uniform inlet distribution mode of nitrogen-rich air was the best; when the oxygen concentration uniformity was the only evaluation index, the semi-uniform inlet distribution mode of nitrogen-rich air intake was the best.

To reveal the influence mechanism of temperature measurement error of thermocouple under unsteady operating conditions, one-dimensional unsteady calculation model of temperature measurement of thermocouple was established. A comprehensive evaluation was conducted to reveal the influence of different installation structures, bonding materials and surface heat transfer coefficient conditions on the unsteady temperature measurement error for the thermocouple. The influence of Biot number on unsteady temperature measurement was further explored. Results showed that it is necessary to consider environmental radiation effect. Temperature error was 1.48 K at 300 s temperature measurement time compared with that without consideration of environmental radiation effect. Different installation structures of thermocouple had a greater impact on the temperature measurement error, with a maximum measurement temperature difference of 1.85 K. Moreover, the greater thermal conductivity of the bonding material indicated, the smaller temperature measurement error in the stage of unsteady regular status, and the absolute error was less than 0.5 K after 150 s temperature measurement time. Especially, the surface heat transfer coefficient played an important role in the unsteady temperature measurement error. The larger surface heat transfer coefficient indicated the greater temperature measurement error due to initial temperature field. When the thermal conductivity of bonding material was 2.4 W/(m·K), the variation range of temperature measurement error within the surface heat transfer coefficient range of 50−250 W/(m²·K) was 0.2−0.5 K. Therefore, the research results can provide a reference for the unsteady measurement of wall temperature in the heat transfer experiments of aero-engine.

In view of the phenomenon that the secondary flow in the channel causes the deflection of the film trajectory on the pressure surface, the concept of multi region compound angle along the spanwise direction was proposed. Numerical simulation was conducted to investigate the film cooling characteristics of turbine guide vane HS1A. The influences of secondary flow, compound angle and blowing ratio on the film cooling characteristics were analyzed under the condition of turbine outlet Reynolds number of 2.3×10⁵. The results showed that the secondary flow in the near endwall had the ability to promote the spanwise coverage of the film, which improved the cooling effectiveness downstream the film hole outlet compared with the mid blade region, but also intensified the mixing of the jet and the mainstream, and reduced the effective coverage length of the film; according to the secondary flow in different spanwise regions, the film hole compound angle in each region was finely arranged, which can correct all the deflection angles of the film trajectory to 0°, and the average film cooling effectiveness was increased by 10.42%; the compound angle cooling model had good applicability within the range of blowing ratio of 0.5−1.0. When the blowing ratio increased to 1.3, the film trajectory could deflect reversely.

In order to analyze the combustion instability (CI) characteristics of lean premixed pre-vaporized (LPP) combustor, three-dimensional Helmholtz equations were simplified at three different levels. The average temperature field equation without source term, the imported CFD temperature field equation without source term, and the imported combustion flow field characteristics equation with source term, were respectively simulated for the single-head LPP combustor model in three-dimensional frequency domain. The results showed that the temperature distribution in the combustor was an important factor affecting the acoustic eigenfrequency of the combustor, and the source term of heat release rate had no effect on the main frequency. Compared with only setting of average temperature field, importing the 3D temperature field calculated by CFD can obtain more consistent results with the experimental frequency, and the accuracy was increased by 5%. The relationship between the acoustic system and the combustion flow field can be quickly established by solving the frequency domain equation in a decoupling way. The spatial distribution characteristics of heat release rate and hysteresis time were represented by the source term of the Helmholtz equation, which had no effect on the prediction of the natural frequency of the combustion chamber, but the detailed sound pressure distribution characteristics can be obtained.

Self-similarity heat sink (SSHS) has the advantages of compact structure design, prominent heat transfer performance and extendibility, which can be applied in cooling the electronic chips. A new SSHS with the structure of deflector hole was proposed to overcome the maldistribution and improve the overall heat transfer performance. Numerical simulation was carried out to validate the new structure design of the SSHS. For the SSHS with the structure of deflector hole, the coolant was distributed to the deflector holes, subsequently impinged the bottom of the overflow channels, which enhanced the heat transfer of the SSHS. Comparison between the original and optimized SSHS was carried out with respect to flow distribution uniformity, heat transfer capacity and heat dissipation uniformity. For the mass flow rate covering 0.58−1.44 kg/h, the optimized SSHS had a better flow distribution uniformity, the maximum temperature on the heating surface was reduced by 10 K, the heat dissipation uniformity was increased by 57%, and average pressure drop was reduced by 10.4% compared with the original SSHS. The geometric dimensions of the deflector holes were further optimized, and better flow distribution and overall heat dissipation performance were achieved with the sacrifice of increasing the pressure drop of 16.5% compared with the optimized SSHS.

Based on the corresponding relationship between the wing pressure distribution and flow separation, a feedback control method for separation flow control by plasma was proposed. Using this method, the separation of wing surface can be distinguished by the pressure of the characteristic points on the model surface, and the plasma flow control can be automatically applied or canceled according to the judgment results. The method was verified on NACA0015 airfoil and flying wing layout models respectively in wind tunnel experiment. The experiment showed that the plasma flow control method based on pressure feedback can realize the automatic suppression of the wing flow separation, thus improving the stall characteristics of the model. The effect of the plasma based on pressure feedback was consistent with that of open-loop control. Both of them can increase the maximum lift coefficient of the flying wing layout model by more than 27% and the stall angle of attack by 4° at a wind speed of 30 m/s.

By taking turbofan engine as the research object, the characteristics and flow field of nacelle and fan integration under the conditions of 0°, 15°, 25° attack angle and 90° crosswind of ±10, ±20 m/s and ±30 m/s were studied by numerical simulation. The results showed that with the increase of attack angle, the negative effect of the reverse crosswind on the performance of the nacelle inlet and fan was greater under the same wind speed. According to the airworthiness clause requirements of maneuverability, wind speed and surge/stall characteristics, when the attack angle was 25°, the forward and reverse crosswind ranges that the nacelle and fan integration performance can withstand were about 0−23 m/s and 0−18 m/s, respectively, and when the crosswind speed was ±30 m/s, the allowable maneuvering attack angle ranges were about 0°−3° and 0°−2°, respectively.

In order to reveal the mechanism of the coupling effect of the flow field induced by the multi dielectric barrier discharge (MDBD) actuator, the numerical simulation method based on the combination of the plasma body force model and Navier-Stokes (N-S) equation was employed to study the actuation characteristics of MDBD under the condition of the static atmosphere. The results indicated that, under the condition of steady actuation, the MDBD plasma actuator can effectively increase the induced flow velocity and thickness, and broaden the actuation area. In addition, under the condition of unsteady actuation, each actuator of MDBD provided a continuous momentum injection to the induced vortex, retarded its dissipation process and enhanced its convective and mixing capabilities. The cyclic actuation frequency had a large impact on MDBD performance. When the actuation frequency was f=20 Hz, the low-pressure region formed by the pulse jet accelerated the induced vortex towards the wall of the plate. The phenomenon of the induced vortex fusion was observed at the condition of f=50 Hz, which can enhance the induced vortex strength and velocity, and reduce the height of the induced vortex core. When the actuation frequency was f=200 Hz, the effect between the induced vortices was weakened, which was convected downstream of the actuator as a format of vortex clusters; in this condition, the effect of MDBD was similar to multiple independent SDBD actuators.

The non-equilibrium effect of high temperature gas in hypersonic vehicle reaction control system’s (RCS) jet interaction flow field was studied. Based on high temperature air and jet gas’s physical and chemical reaction model and by solving three-dimensional non-equilibrium Reynolds-averaged Navier-Stokes (RANS) equations, numerical simulation of typical configuration’s jet interaction non-equilibrium flow field was conducted, the influences of air and jet gas’s non-equilibrium effect were analyzed, and the influences under various flight conditions were also discussed. The results showed that: air’s non-equilibrium effect was significant when flight Mach number was high, which can reduce the additional thrust produced by jet interaction, and decrease the heat flux in jet interaction area; high flight velocity could enhance this effect. Jet gas’s non-equilibrium effect varied under different conditions; when flight altitude was low, jet gas’s components mainly involved in afterburning/recombination reactions, which can increase additional thrust and the heat flux in jet interaction area; when flight altitude was high, jet gas’s components mainly involved in dissociation reactions, which can decrease additional thrust and the heat flux; as the flight altitude increased through the trajectory, jet gas’s dissociation reactions were enhanced and recombination reactions were weakened. To simulate hypersonic vehicle’s jet interaction flow field more precisely, it is necessary to consider high temperature gas non-equilibrium effect.

Combined with numerical simulation and wind tunnel test technology, the nonlinear coupled constitutive relations (NCCR) model and the simplified generalized hydrodynamic (SGH) model derived by dimensional analysis in hypersonic continuum flows were studied. Based on the hypersonic wind tunnel test system, the aerodynamic force and surface pressure of the type hypervelocity Ballistic 2 (HB2) standard model and the blunt cone were measured under different flow conditions. Meanwhile, under the three-dimensional finite volume framework, the NCCR model with decomposed and undecomposed algorithm and SGH model were used for numerical investigation of the flight models under the test conditions. Result showed that, the aerodynamic force and surface pressure obtained by NCCR model and SGH model were consistent with the solutions of Navier-Stokes (NS) equations as well as the data measured in the wind tunnel. However, the friction/heat flux coefficients predicted by NCCR model with decomposed algorithm at the expansion corner of head of type HB2 were lower than those of NS equations, while the NCCR model with undecomposed algorithm was consistent with NS solver. The computational results and experimental data showed that the accuracy of NCCR model and SGH model in hypersonic continuous flow was validated, and the applicability of decomposed algorithm of NCCR model in three-dimensional high-speed flows shall be further improved.

Axial-force balance of the cracked fuel vapor turbine was studied from the perspective of turbine structure. Partial admission cracked fuel vapor turbines with unshrouded, open, and enclosed impellers were established respectively, and the aerodynamic performance and axial force of turbines under the design condition were compared and analyzed by numerical simulation. Meanwhile, the variation law of the total axial force with pressure ratio under off-design conditions was summarized, which provided guidance and suggestions for the selection of cracked fuel vapor turbine structure forms. The simulation results showed that the aerodynamic performance of unshrouded, open, and enclosed turbines was similar, but the axial force performance was distinctive. When the pressure ratio changed, the axial force stability of the open turbine was the best, the unshrouded one was the worst, and the enclosed turbine axial force was the smallest among the three kinds of cracked fuel vapor turbines. The result analysis showed that when the turbine pressure ratio was lower than 3, the unshrouded turbine should be discarded; if the turbine pressure ratio changed greatly, the open turbine should be adopted. In addition, axial load of bearings should be sufficient when enclosed cracked fuel vapor turbines were exploited.

Based on the self-calibration Kalman filter (SKF) and the multiple-model estimation (MME), considering the influence of unknown inputs (such as gusts, faults, unknown system errors, etc.) on the system state equation in Engineering, the multiple-model self-calibration Kalman filter (MSKF) was proposed. According to the Bayes' theorem, this filtering method used the SKF and the standard Kalman filter (KF) whose weights were assigned automatically to obtain the final filtering result through weight-average way. Compared with the SKF, the MSKF can not only effectively compensate the effects of non-zero unknown inputs, but also improve the estimation accuracy when unknown inputs were zero. A large number of simulation results showed that accuracy can be improved by more than 10%, using the proposed method. In summary, the MSKF has stronger adaptability and robustness.

To study the effects of the connection component loosening fault in aero-engine high-pressure rotor, a Timoshenko beam model considering loosening fault was established. The influences of joint surface stiffness, speed, number of loose bolts, position of loosening fault and tightening torque on rotor dynamics were discussed. The results of theoretical analysis were verified by experiments. The results showed that there was an obvious subcritical resonance in the speed-up process of the fault system, and the critical peak speed was generated in advance. The more number of loose bolts indicated the more obvious period-doubling bifurcation of the system. The loosening fault was located in the middle of the rotor span, which had a great influence on the system dynamics. When the bolt tightening torque was reduced, the second critical peak speed of the system was generated in advance. The experimental data were in good agreement with the results of theoretical analysis, and the proposed theoretical method had certain accuracy and applicability. The research results provide a theoretical and experimental basis for further study on connection component loosening fault of aero-engine high pressure rotor.

The blade mistuning identification and model validation for an academic blisk test piece based on the blade detuning tests were performed. Blade detuning tests were carried out to obtain the ‘blade-alone’ modal frequencies of each individual blade of the blisk. A guideline was proposed for determining the position of the detuning mass onto the blade. Secondly, a novel mistuning identification method was introduced to account for the residual inter-blade coupling effect due to detuning masses and yield more accurate blade mistuning identification results. The impacts of different detuning mass and positions on the blade detuning test and mistuning identification results were investigated. A mistuned blisk finite element model could be further established; the mistuned blisk model was experimentally verified by the conventional modal test. Excellent agreement was achieved between the numerically/experimentally derived natural frequencies and mode shapes of the blisk. The majority of the frequency deviations fell below 0.3%. Results showed that the novel mistuning identification method improved the accuracy of the blade mistuning patterns directly obtained from the blade detuning test results. The resultant mistuned blisk model can effectively reproduce the natural vibration characteristics of the real blisk test piece.

Based on the life data information from multiple test sources, using the mapping relationship between different data sources, the multi-source data were migrated to the field data source to form a mixed data source, which was used as the basis for the Bayesian statistical analysis of product reliability. For the accelerated life data under different stresses, it was converted to the constant stress level to determine the parameter distribution density function, which was used as the priori condition for the product reliability Bayesian statistical analysis. By combining the Bayesian statistical model with the data migration model, and fusing multi-source data while determining the parameter estimates at the same time, the product density function and the product reliability analysis were obtained. The example showed that this model can effectively achieve data migration by utilizing the mapping relationship between data sources, and can accelerate the synchronous fusion of life data with other types of data sources. The comprehensive evaluation of product reliability after fusing sample data was more comprehensive and objective than that of a single life data source.

A novel passive control method for bio-inspired herringbone riblets was applied to relieve the flow near the blade endwall in a linear cascade, and its effectiveness and mechanism in controlling corner separation were investigated through numerical simulations. The herringbone riblets were placed at the upstream endwall of the blade, and the influence of riblet height and deflection angle on corner separation control was investigated. The results showed that the herringbone riblets can effectively relieve the flow near the blade endwall over the operating range, and the implementation of herringbone riblets with a height of only 0.08 boundary layer thickness and a deflection angle of 30 degrees can reduce the total pressure loss by up to 9.89% and increase the static pressure coefficient by 12.27%. Flow details indicated that small-scale vortices in the riblet channel can accumulate and form high-intensity large-scale vortices close to the bottom of the boundary layer downstream, which effectively reduced additional losses compared with traditional micro vortex generators. Furthermore, the induced vortices enhanced the mixing of the boundary layer and main flow, inhibited the lateral migration of low-energy fluid in the endwall boundary layer, and delayed the formation of separation vortices, eliminating vortex ring in the corner region and effectively improving the flow near the blade endwall.

By numerical calculation of an 8-stage compressor, the characteristics and instability mechanism of marine compressor at design speed were obtained. At the same time, the first 1.5-stage of an 8-stage compressor was chosen to study the influence mechanism of the partial clearance on hub corner stall. The results showed that the flow instability mainly occurred in the hub-corner of stator at the design speed. The main reason is hub corner stall. The closed separation appeared in the corner and block the flow field. The 1.5-stage compressor result showed that the reason for compressor instability at design speed is the same as that of the 8-stage compressor. the partial clearance at different positions can widen effectively the margin. The effect of clearance was best when the clearance was in the middle. The clearance can not only increase the margin from 22.1% to 27.2%, but also cause the minimum loss at the maximum efficiency operation (from 93.9% to 93.21%). The leakage momentum caused by the clearance in other positions was not enough to eliminate the stall vortex in the hub. The optimal clearance was near the vortex core of stall vortex in the hub corner, where the leakage momentum was the maximum.

Based on the traditional variable nozzle turbine guide vane, an end-clearance sealed guide vane (ESGV) was designed. The effect of ESGV on the performance of the turbine was analyzed by numerical simulation under different guide openings. The results showed that ESGV can effectively suppress the end-wall clearance leakage flow of the guide vane and improve the flow state of nozzle ring and rotor channel, and then enhance the turbine efficiency. Finally, the turbine performance tests of the ESGV turbine and the prototype turbine were carried out. The results showed that, when the flow characteristic curves of the two turbines were roughly the same, the efficiency of the ESGV turbine was 5% higher than that of the prototype turbine at the medium opening.

The development history and research progresses of pneumatic driven piston type and motor driven piston type powder fuel engine propellant feeding systems were introduced. The characteristics, advantages and disadvantages of choked and non-choked powder supply were compared and analyzed. The principle and existing problems of mass flow measurement methods of feeding system, including piston displacement method and weighing method, were illustrated. By summarizing and analyzing related researches on powdered fuel engine feeding system, the following conclusions were drawn: the integration of pneumatic driven piston type was higher, the effect of piston controlling of motor driven piston type was better; the stability of choked powder supply was stronger, yet the utilization rate of air of non-choked powder supply was higher; the precision in measuring flow rate of gravimetric method was higher, yet the scope of application of piston displacement method was wider. The development directions such as optimizing the feeding system structure, exploring the mechanism of choking flow and establishing the relationship model of flow parameters were put forward for further development.

To study the convergence characteristics of the liquid film of a bipropellant pressure-swirl injector, numerical investigation of bipropellant pressure swirl injector was conducted based on the volume of fluid (VOF) coupled Level Set methods under different thruster chamber pressures and pressure drops. The atomization characteristics of the injector were analyzed, and the law of mutual influence between the inner and outer nozzle liquid films was revealed. The results showed that when the pressure drop of outer nozzle was greater than 1.6 MPa, the convergence phenomenon occurred, and the inner and outer spray films were gradually overlapped and converged with each other in the process. After convergence, the downstream speed of the injector outlet relatively decreased. The increase of the thrust chamber pressure led to the shortening of the convergence time, while the influence of the pressure drop on the convergence time was relatively small. With the increase of the pressure drop, the convergence time decreased at first and then increased.

In order to solve the problem of the noise interference and the difficulty in feature extraction in the vibration signal of planetary gearbox, a fault diagnosis method for planetary gearboxes based on jel-lyfish search optimization variational mode extraction (JS-VME), deep belief network (DBN) and supervised Mahalanobis distance uniform manifold approximation and projection algorithms (MS-UMAP) was proposed. The vibration signals of the planetary gearbox were collected, and JS-VME was used to preprocess them to obtain expected IMF （intrinsic mode function）component with strong correlation. Then, DBN was applied to the IMF component to extract feature vectors, and the high-dimensional fault feature set was built. MS-UMAP was used for dimensionality reduc-tion to obtain low-dimensional and sensitive fault features. The low-dimensional fault feature set was applied to the jellyfish search optimization kernel extreme learning machine (JS-KELM) to de-termine fault types. The experiment results of planetary gearbox fault diagnosis showed that com-pared with UMAP, t-SNE, Isomap, LPP, W-Isomap, LLE, LTSA and MDS, the MS-UMAP algorithm had the best dimensionality reduction effect on the feature extraction results of JS-VME-DBN. The fault recognition rate of the proposed method reached 100% with a certain validity in planetary gearbox, such as the cracks, wear and missing teeth.

In order to meet the requirements for high power density, high-quality power supply and AC-DC hybrid power supply for aviation power generation system proposed by aircraft multi-electrification, a variety of AC-DC integrated power generation schemes were compared, and their advantages and disadvantages were analyzed. An AC-DC integrated power generation system based on dual winding induction generator was proposed. The system made full use of the dual-winding induction generator stator with two sets of three-phase windings, so it can achieve a good AC-DC integrated power generation. After preliminary design of the generator, based on the multi-objective optimization algorithm, the dual-winding induction generator was optimized and designed with the objectives of efficiency and power density. An AC-DC integrated power generation control strategy based on active disturbance rejection control was adopted to improve the dynamic performance and load adaptability of the system. Finally, a generator of 60 kW (24 kW AC, 36 kW DC) dual-winding induction generator principle prototype was developed, and the experimental results verified that when the AC side increased and offloaded abruptly, the voltage changed within ±10 V and the recovery time was not more than 25 ms; when the DC side increased and offloaded abruptly, the voltage fluctuated within ±30 V and the recovery time was not more than 45 ms. The system with good dynamic and steady-state performance can realize high-quality AC-DC integrated power generation. Research results showed that the aero-AC-DC integrated power generation system based on dual-winding induction generator is expected to provide a competitive alternative for the multi-electric aircraft power generation system.