In order to investigate Reynolds number effects on the performance of turbofan and improve state performance model calculation accuracy in flight envelope, a computational analysis method based on whole engine test data identification was presented. Identification algorithm convergence and result validity were improved through measurement parameters selection for gas path analysis. Component performance scale factor for each test point was identified by non-linear gas path analysis method. Statistical analysis of the relationship between Reynolds Number and component performance scale factor was performed, and the performance change of each component affected by Reynolds Number was quantitatively obtained. Baseline steady state performance model was modified and calculation accuracy was compared and verified. Compared with test data, modified steady state performance model calculation error was not more than 2.5%. Compared with baseline steady state performance model calculation result for each parameter, the average calculation accuracy was improved 2.3%, and the maximal calculation accuracy was improved 9.2%.

A flow field calculation model for scissors tail-rotor was established based on the computational fluid dynamics (CFD). The nested grid method was used to simulate the blade motion, and the dual-time method was used for time propulsion. A high-efficiency trim strategy based on the “delta method” was adopted to trim the collective pitch for the scissors tail-rotor. On the basis, the aeroacoustic characteristics of the scissors tail-rotor were calculated by using Ffowcs Williams-Hawkings (FW-H) equation. This method was applied to the calculation of scissors tail-rotor (L configuration and U configuration) in different forward velocities. The effects of the scissors angle and vertical space on the aerodynamic force and noise of scissors tail-rotor were compared and analyzed. It was found that the scissors angle had an important effect on aerodynamic force and noise characteristics of the scissor tail rotor, while the vertical space had little effect on the tail rotor within a reasonable range. Compared with the conventional tail rotor, the noise directivity of the scissor tail rotor changed little in forward flight, but the noise amplitude changed significantly.

A calculation method of power balance in CFD 3D machine simulation was presented. The key variables determining the equilibrium relationship such as temperature before turbine and physical speed were deduced and determined. The numerical iterative method based on the steady state flow field and power empirical relationship was proposed to realize the power balance. The power balance calculation method was used to carry out the CFD numerical calculation of the MTJ-80 turbojet, and the coupling and prediction of the performance simulation and control law were realized. Under the condition of constant fuel quantity, the power difference between compressor and turbine can be less than 0.1% through speed iteration, and the steady running speed can be predicted under the condition of fixed fuel supply. Under the condition of constant control speed, the power difference between compressor and turbine was less than 0.15% by adjusting the fuel quantity, and the fuel flow rate under the condition of constant control speed can be predicted. The data verification results showed that the proposed method can be coupled with CFD 3D machine calculation with strong convergence, which solved the power imbalance problem in the previous 3D performance simulation of the machine.

The fluctuation total pressures in aerodynamic interface plane need be measured in inlet experiments. The tube between the pulsating pressure sensor and the measurement point may lead to larger error between the measurement value and the truth value, and also generate larger effect on the measurement of the turbulence. The experiments in Turbine Power Simulation calibration tank of China Aerodynamics Research and Development Center were performed to research the effect of tube on the measurement of fluctuation total pressure and the turbulence. Experiment results showed that the effect of tube seriously affected the measurement of the fluctuating total pressure and the turbulence. The value of fluctuating total pressure in frequency can be magnified more than 10 times and that of turbulence can be magnified more than 2.8 times. A correction method based on the improved dissipation model was proposed and calibrated, which can effectively reduce the measuring error brought by the tube.

For further improving the maneuverability of the flapping-wing micro air vehicle, the movable hinge crank slider flapping mechanism proposed earlier was improved and miniaturized. The kinematics of the flapping mechanism indicated that as the hinges moved, the difference of the bilateral flapping amplitudes increased from 21.3° to 51.2° and the bilateral mean flapping amplitudes decreased to 0°. And a modularized and carbon fiber plate bionic dragonfly prototype was designed. The measurement and flight test of the prototype showed that the lift and thrust of the prototype changed obviously by adjusting the flapping frequency, realizing rapid acceleration and climbing maneuvering flight. By moving the hinge, significant yaw moment and roll moment were generated for maneuvering turn at larger flapping frequency.

With distributed ducted fans propulsion as the background, the Reynolds-averaged Navier-Stokes （RANS） equation was solved by using multiple reference frame （MRF） and momentum source method （MSM） with given force distribution, and numerical simulation analysis of ducted fans with different shapes was carried out. The interaction between the fan panel and the duct wall with different shapes was analyzed and compared. Then numerical calculation of distributed ducted fans with different shapes was carried out. It showed that the circle ducted fan had the best propulsion characteristics, while the square ducted fan had the worst. The propulsion characteristics of the rectangular to circular ducted fan was between them. The existence of corners in non-circular ducted fan could induce separation of the inner flow and generate interference resistance, and the smaller radius of the corner indicated the more significant influence. The non-circular duct wall affected the fan inlet area and the size of the blade tip vortex. Therefore, the ducted fan propulsion efficiency was affected via the fan panel efficiency and lip suction.

In order to evaluate the sizing of all-electric aircraft for conceptual design precisely, a total mass evaluation module was established, which was combined with its own design features, mission requirements and energy consumption analysis. A sizing matrix plot module was established to enclose the equations and fulfill some performance constraints. The design system can assess take-off mass, wing loading, power loading and some crucial sizing parameters, like wing area and wingspan. Based on this system, three domestic and international all-electric aircrafts were adopted for design and comparison: for different aircraft types, the mass, wing area, wingspan, wing load and other design parameters obtained by the program were all close to the actual data; the absolute numerical difference of the data comparison indicated that the simulation and energy consumption of the mission were within a reasonable range. And the relative numerical differences were also acceptable for the overall parameter evaluation in the conceptual design stage. The results show that it can verify the validity and feasibility of the developed method, and can also support the all-electric aircraft design with parameter selection and performance evaluation.

In view of at three random typical pressure jump phenomena due to complex pipeline layout and high flow velocity, the nonlinear multi-solution flow characteristics in right-angle multi-way pipeline were proposed by numerical simulation. Under certain disturbance conditions, a large eddy in the multi-way cavity and spiral flow along the branch pipeline were formed, which caused additional flow resistance and local pressure jump in complex pipeline. On the basis, in order to eliminate the large eddy in the multi-way cavity, the suppression scheme by appending guide devices or baffles was proposed. With the actual pipeline layout, the ground recurrence test and suppression improvement test of random pressure jump phenomena were carried out. The experimental results indicated that the multi-solution characteristics of the three complex multi-way pipeline were validated, and the improved scheme can effectively eliminate the random pressure jump phenomena.

Accurately mastering the working characteristics and margin of liquid rocket engine under different parameters is crucial to its reliability. High-altitude simulation and ground hot fire tests of the second generation 490 N engine under off-rated conditions were carried out, and the influence of thrust and mixing ratio changes on the engine working characteristics was studied. Results indicated that, the engine can work normally within the wide envelope scope of mixing ratio 1.54—1.80 and vacuum thrust 372—584 N under high-altitude simulation environment. With the increase of thrust, the vacuum specific impulse and throat temperature both rose, the combustion chamber efficiency increased, kept stable and decreased in turns, and the nozzle efficiency increased slightly. As the mixing ratio increased, the vacuum specific impulse and throat temperature also increased, the combustion chamber efficiency did not change obviously, and the nozzle efficiency decreased slightly. For the rated mixing ratio, the combustion chamber pressure fluctuated smoothly within the range of 0.61—1.56 MPa, enabling to work with the vacuum thrust of 345—900 N. However, it produced combustion oscillation of low/middle frequency (207 Hz) coupled with the feed system at 0.51 MPa. The intensifying erosion of hot flow and increasing temperature in the throat under high working conditions accelerated the loss of coating, which reduced the long-term working life of engine. Nonetheless, it can still meet the working life requirements of 25 000 s for satellite identification within a certain off-rated range and single ignition duration.

Based on the introduction of the principle of Hall thruster, the influencing factors of Hall thruster thrust were deduced and analyzed, summarizing how to change Hall thruster thrust. The types of space missions applicable to different thrust levels, such as μN-level level, mN-level and N-level were summarized. The models of Hall thrusters at home and abroad were classified according to three levels, and the development and trend of each series of models of the three thrust levels were sorted out and analyzed respectively. The technical bottleneck and development trend were summarized and analyzed. Hall thrusters of various thrust stages were comprehensively and systematically reviewed. The results indicated that the mN-level was the most mature. With more abundant types of space missions, the μN-level level and N-level Hall thrusters had considerable development potential. The key technologies for the development of Hall Thrusters in each thrust stage were prospected. The results suggested that Hall Thrusters should improve the overall performance, expand the thrust coverage, strengthen the multi-mode working ability, develop hollow cathode and explore different propellants, thus providing a reference for the further development of Hall electric propulsion.

In order to accurately evaluate the effect of different cooling schemes on the heat transfer characteristics of thrust chamber of high-pressure liquid oxygen and hydrocarbon rocket engine, a complete model combining regenerative channel, liquid film shield, thermal barrier coating and radiation heat transfer was established. The Ievlev semi-empirical model was adopted for calculation of convection on gas wall. The Shruvik safety margin evaluation criterion was used in analysis. The radial direction multi-layer temperature and heat flux of the thrust chamber were calculated. Based on a large thrust liquid oxygen and kerosene rocket engine, the upper limit of heat transfer capacity of different cooling schemes was studied, and the influence of different thrust chamber pressures on cooling design was analyzed. Results showed that, thrust chamber can mainly rely on regenerative cooling technology to meet the cooling requirements at 12 MPa or below; film cooling slots should be further used at 16 MPa or below; thermal barrier coating should be further installed at 18 MPa or below; several intensive cooling measures must be adopted at 20 MPa or higher.

In order to monitor the vibration status of aeroengines and acquire warning signals in real-time, an enhanced SENet (squeeze-and-excitation network) model was proposed based on gas path and vibration parameters. Compared with the previous research which used datasets generated from specific lab situations and simulation data, actual QAR （quick access recorder） data were adopted for random sampling of the datasets. This technique could characterize the real operation status and the interaction of parameters better in vibration systems. The results showed that it is possible to forecast the vibration of aeroengines, and the SENet model could effectively and timely track sudden changes and the fluctuation of vibration. In addition, the applicability of this method into other vibration parameters and different types of aeroengines was tested. Furthermore, compared with other classical learning algorithms , the SENet model may obtain a smaller error in vibration forecasting. At the same time, the experiments showed that compared with previous research only focusing on the vibration, using the fusion of multi parameters could improve the accuracy of the forecast.

The principle and data processing method of the non-contact vibration measurement system based on blade tip timing (BTT) were introduced, and applied for analyzing losing-corner fault of aero-engine compressor blades of the fifth stage. The axial and circumferential positions of timing-tip sensors were optimized. Leading and trailing edges of three typical blades (original profile, optimized profile and chamfered profile blades) were measured by non-contact vibration measurement, and the natural frequency and magnitude of the whole blades were obtained. Comparing the vibration parameters of the leading edge with the trailing edges of original profile and optimized profile blades, the resonant amplitudes were almost same. Comparing the vibration parameters of the leading edge of three typical blades, the resonance speeds of the optimized profile blades were a little higher than those of original profile blades, and those of the chamfered blades were larger than the optimized profile blades. Comparing the vibration parameters of the trailing edge of original profile and optimized profile blades, the resonance speeds of the optimized profile blades were a little higher than those of original profile blades. It can be judged that the main reason of the original profile blades occurring losing-corner fault was that the blades were resonant at the working speeds of 4200 r/min. The resonant speeds of the optimized profile blades were about 40–80 r/min higher than the original profile blades, leading to the reduction of the losing-corner fault. The resonant speeds of the chamfered blades were about 140–180 r/min higher than the original profile blades, and the fault can be eliminated.

In order to evaluate the PV value of the carbon graphite sealing material used in the aero-engines of some current models at home, the dynamic test bench was independently designed according to the actual service conditions of the fulcrum flexible circumferential carbon graphite sealing material for engine at high altitude-long endurance, under the condition of high friction linear velocity (60 − 130 m/s), light load (30 − 400 kPa) and temperature (25 − 600 ℃). The test results showed that the allowable [PV] values of XX1#-1, XX2#-1, XX3#-1, XX4#-1 and XX5#-1, using a constant speed of 80 m/s and gradually increasing the load, were 4.8, 5.9, 6.9, 9.1, 12.3 MPa·m/s, respectively. The allowable [PV] values of XX2#-2, XX3#-2, XX4#-2 and XX5#-2, using a permanent load of 300 kPa and increasing the linear velocity step by step, were 6.0, 8.0, 10.0 and 12.0 MPa·m/s, respectively. The installation demonstration of an institute found that the test samples were prone to abnormal wear and gas leakage, when the allowable [PV] value of the test sample was less than the allowable [PV] value of the engine. When the allowable [PV] value of the test sample was greater than the allowable [PV] value of the engine, this method can not only provide material characteristic data for selection of the engine fulcrum carbon graphite sealing material, but also contribute to the research of the engine carbon graphite sealing material identification method. In addition, the time cost, labor cost and installation demonstration cost of graphite sealing material selection for engine will be also greatly reduced through this test method.

Based on the measurement of tip clearance, the prediction method of aeroengine rotor vibration displacement and axis trajectory was studied. Based on the dynamic model of tip clearance change, the numerical calculation and finite element simulation of tip clearance change were carried out, and the experimental research was carried out for the rotor tester with casing. The tip clearance of two measuring points perpendicular to each other was measured by eddy current sensor. The signal was processed by Hilbert Huang transform to extract its low-frequency component. The cross-correlation analysis method was used to extract the frequency conversion signal, and the axis trajectory was drawn from the frequency conversion component. Compared with the direct measurement of the axis trajectory on the rotating shaft of the same section, the coincidence degree of these two was more than 90%. The test results fully showed the correctness and effectiveness of the research method, providing an effective technical way to indirectly obtain the rotor vibration displacement through the blade tip clearance test.

Self-excited instability experiments were performed to investigate the thermo-acoustic instabilities and their relationship with heat release rate fluctuations in swirled partially premixed combustors. The reconstructed phase space was used to figure out the pressure fluctuations, and their relationship was analyzed by using Rayleigh index. The proper orthogonal decomposition method was employed to extract flame coherent structures. Results indicated that pressure fluctuations presented evolutions of low-amplitude, intermittent, limit cycle and low-amplitude oscillations, with the increase of equivalence ratio. The fluctuation frequency was controlled by the first order pure acoustic mode of the cavity (eigenfrequency was around 80.8 Hz). The Rayleigh index was around zero value when oscillations occurred, while this parameter maintained positive at the thermo-acoustic instability state. The proper orthogonal decomposition results showed that the flame distributions reflected by the first two order modes (the total energy of these two mode was above 55%) became bright and dark alternately along the longitudinal direction at the limit cycle, resulting from continuous vortex sheddings. The frequency of time coefficients (83.2 Hz) was consistent with that of pressure pulsations (83.3 Hz). At the low equivalence ratio, the mode spatial distribution had no regular pattern. When oscillations were intermittent, the dominant mode was of an axisymmetric pattern. Only the outer edge of the flame had large-scale longitudinal motions at a high equivalence ratio.

Due to the irregular shape of the high-temperature hot components in the aero-engine, the heat radiation process has obvious discontinuous and shielding effects, and how to achieve accurate simulation of this process is a challenge. Using discontinuous finite element to discretize the angle and space domain in the radiation transfer equation, and based on the hierarchical limiting strategy, the Barth-Jespersen limiter was introduced in the angle and space domain to suppress the non-physical oscillation of the radiation intensity numerical solution in the angle and space. By comparing with the results of the finite volume method and the Monte Carlo method in the literature, the validity of the method was proved. In addition, by comparing with the angular analytical solution of radiation intensity, it was verified that the limiter can effectively suppress the numerical oscillation caused by the radiation discontinuity effect, and can eliminate the nonphysical solution of radiation intensity. The three-dimensional characterization of the radiation intensity in the angular domain at any spatial position was realized.

In order to study the verification method of aero-engine’s compliance with the airworthiness terms of surge/stall characteristics under distortion conditions, the distortion test method for engine stability evaluation was applied to the airworthiness field. And a turbofan engine with large bypass ratio was taken as the object, the component characteristics at 0°—30° attack angles were simulated, and the distortion generators with different structures were designed according to the obtained distortion map. The results showed that the designed distortion generator can better simulate the distortion map, aerodynamic performance and flow field characteristics of the fan at different attack angles to a certain extent, and the maximum deviation of stability margin was 4.32%. The relationship between engine surge margin, different attack angles and distortion generator was summarized to determine the stable working range required by airworthiness terms. According to China Civil Aviation Regulations, a method for verifying the compliance of aero-engine surge/stall characteristics with inlet distortion was developed, and its task stage in the model qualification process was defined.

Nonlinear spring element and damping element were used to simulate the collision contact between the cage and the rolling element, and a dynamic model of rolling bearing was proposed considering some nonlinear factors such as collision contact, friction, cage pocket clearance, etc. Collision contact characteristics between the cage and the rolling element were investigated by dynamic simulation, the influences of bearing load and rotational speed on the cage contact characteristics were also analyzed. The results showed that discontinuous collision contacts appeared between the rolling element and cage pocket front bar or rear bar alternately, and the rotational speed of the cage fluctuate d irregularly because of the contact forces. The maximum contact force and the fluctuation range of cage rotatioanal speed decreased with the increase of bearing radial load, and increased with the increase of rotational speed of bearing. The proposed model can provide a theoretical basis for the design of rolling bearing and failure analysis of the cage.

In order to facilitate the design of the transition state control law of the turbofan engine, a fixed states method based on the model was proposed. In this method, the speed states were fixed during the engine acceleration and deceleration process, and then the optimal fuel meeting the physical constraints was solved reversely. Finally, the optimal acceleration and deceleration control law was built. Using a turbofan engine as an example, the open-loop fuel ratio and closed-loop n-dot control laws were designed by this method based on component level model. The simulation results showed that the results of two control laws were basically consistent, and also compliant with the requirement of minimum acceleration and deceleration time; the simulation of high and low compressor speed was consistent with the design status and the main parameters of engine, such as turbine outlet total temperature, combustor excess air coefficient, surge margin, didn’t exceed the limit. It verified the correctness and effectiveness of the proposed fixed states method of designing turbofan engine acceleration and deceleration control law.