In order to reasonably analyze the pollutant emission characteristics of the supersonic engine during the landing and take-off (LTO) cycle, a supersonic engine model based on the CFM56-7B27 engine core was built. The emission index (EI) of LTO pollutants was calculated by establishing an emission calculation model, and the emission characteristics were analyzed; the influences of pollutant emission characteristics in climb and idle phases on the LTO supersonic mode standard setting were studied, and then a more representative LTO supersonic mode standard was determined. Results showed that the thrust setting (TS) and time in mode (TIM) in different LTO phases had different effects on pollutant emission characteristics; in the research of LTO standard, the mass of the oxides of nitrogen emitted/rated thrust at 60% of the rated thrust and TIM of 2 min climb point was closer to the supersonic climbing trajectories, and the TS of the idle point can better meet the combustion efficiency requirements limited by the emission characteristics of pollutants (carbon monoxide, unburned hydrocarbons) when it was not less than 10% of the rated thrust. Therefore, it is more reasonable to take 60% of the rated thrust and TIM of 2 min as the climb point standard of the LTO supersonic mode and the TS not less than 10% of the rated thrust as the idle point standard of the LTO supersonic mode.

For the double-wall cooling structure with jet impingement holes and pin fins, in order to study the influences of non-dimansional row spacing and dimensionless hole pitch of the jet impingement holes, the dimensionless diameter of pin fins and the jet impingement Reynolds number on the enhanced heat transfer performance and flow resistance characteristics, experimental studies were conducted based on various geometrical and flow parameters in present study. In addition, the corresponding numerical simulations were also carried out according to the experimental conditions. The results showed that on the whole internal wall surface of the jet impingement target plate with pin fins, the area-averaged Nusselt number monotonically increased with the increase of the jet impingement Reynolds number, and basically presented a linearly increasing tendency. In general, the area-averaged Nusselt number of the wall surface of pin fins was slightly higher than that on the internal wall surface of the jet impingement target plate. With the increase of dimensionless diameter of pin fins, the area-averaged Nusselt number of the entire internal wall surface of the double-wall cooling structure decreased firstly and then increased. In addition, the area-averaged Nusselt number of the entire internal wall surface decreased with the increase of non-dimensionla row spacing. However, the area-averaged Nusselt number of the entire internal wall surface was not sensitive to the variation of non-dimensional hole pitch. The discharge coefficient of the double-wall cooling structure with jet impingement holes and pin fins increased with the increase of the jet impingement Reynolds number, the dimensionless row spacing and non-dimansional hole pitch of jet impingement holes, but decreased with the increase of dimensionless diameter of pin fins.

In response to the characteristics of high-temperature, strong swirl, and non-uniform inflow at the intake of the new generation integrated afterburner, a high-temperature, strong-swirl, non-uniform inlet flow field generation device that can simulate the real intake conditions of an integrated afterburner was designed. At the same time, numerical simulations of the non-uniform flow field coupling inlet Mach number, temperature, and swirl angle were conducted to assess the effectiveness of the design of the non-uniform flow generation device, measurement methods, and evaluation criteria. The accuracy of the numerical simulations was verified through experimental results. The results showed that the numerical calculation of the swirl angle error near the wall was approximately ±2° due to the underestimation of wall dissipation in the numerical calculations, while it was better than ±2° in the central area. Similarly, in the simulation of Mach number non-uniformity, the error in the mainstream area was within 10%; in addition, the simulation of temperature non-uniformity showed larger errors, due to the numerical calculations not considering the heat transfer process from the wall to the outside. Overally, the high-temperature, strong-swirl, non-uniform generation device proposed can generate the real complex non-uniform inlet flow field conditions faced by the intake of the next-generation integrated afterburner. The numerical simulation method used can reveal the flow field's non-uniform characteristics quite accurately.

A three-dimensional simulation method was established to assess the heat exchange performance with methane as coolant for an axial tube-bundled pre-cooler. The flow field of the pre-cooler was obtained at different coolant equivalent ratios and flight conditions, and the control strategy of coolant massflow was proposed from the simulation results. The assessed results indicated that the incoming air could be cooled 131 K at maximum under the condition of three times the equivalent ratio by methane pre-cooler, extending the working velocity spectrum to Ma=3.0 at least for conventional turbine engine. The air field could not completely well-distributed because of the influence of the tube arrangement which led to a maximum temperature distortion of 13.3%. The methane got a maximum temperature rise of 395 K. The temperature difference between the outside and inside tube surfaces was about 15 K, and the methane temperature range at cross section of the tube was about 10 K. The total pressure recovery coefficient of the pre-cooler was 0.715—0.88, the heat transfer efficiency was 0.63—0.9, and the maximum power-weight ratio was 395 kW/kg. The control strategy for coolant was proposed under the condition of engine pre-cooling requirement and coolant consumption limitation. It is recommended that the consumption of coolant should be restricted to 1.5 times the equivalent ratio as flight velocity was lower than Ma=2.5, and 3.0 times the equivalent ratio as flight velocity was higher than Ma=3.0.

In view of the problem that the inlet flow changes dramatically with time during the modal transformation of afterburner/ramjet of turbo-based combined cycle engine, the transverse jet was taken as the research object. Under the conditions of inlet temperature of 300—800 K, inlet velocity of 100—164 m/s and inlet acceleration of 20—100 m/s², the Reynolds mean/discrete phase model was used to discuss the influence of the inlet acceleration on the lateral jet trajectory and Sauter mean diameter (SMD) distribution. The large eddy simulation/fluid volume method was used to discuss the influence of the inlet acceleration on the lateral jet fuel atomization process. The results showed that the inflow acceleration had little effect on the lateral jet trajectory and downstream SMD distribution. The acceleration of the inlet flow may cause the delayed and reverse vortex to be more widely distributed along the jet direction, but the intensity weakened at the edge along the spanwise direction. But the effect was not significant; the time-varying incoming flow had no obvious effect on the fuel crushing and atomization characteristics.

To reduce the pressure loss in a compressor disk cavity, an anti-swirling waist circular drum hole structure was designed. Large Eddy simulation (LES) and the RNG k-ε model were used to investigate the vortex evolution and pressure loss characteristics in the cavity, respectively, revealing the vortex breakdown mechanism and drag reduction mechanism of the anti-swirling waist circular drum hole. The results showed that the anti-swirling waist circular drum hole can reduce the pressure loss in the cavity. The high-speed vortex was intensified rapidly in the low region of the cavity with waist circular drum, and the vortex scale increased rapidly, resulting in the rapid increase of pressure loss coefficient with the radial height decrease. Compared with the waist circular drum hole, the anti-swirling waist circular drum hole can effectively restrain the vortex scale increase and reduce the pressure loss in the cavity by 15.6%. With simple structure and linear airflow characteristic, the anti-swirling waist circular drum hole has high engineering application value.

By using an integrated model with disc cavity and turbine blade, the endwall film cooling effectiveness was investigated numerically by using the shear stress transfer (SST) model to solve the Reynolds-averaged Naiver-Stokes (RANS) equation. The rotational Reynolds number at the outlet of the disc cavity was 1.5×10⁵. Carbon dioxide was chosen as the coolant to maintain the coolant-to-mainstream density ratio. The diffusion process of coolant was characterized by solving the turbulent transport equation. The effect of coolant inlet angle (−45°, 0°, +45°) on the endwall film cooling effectiveness was investigated. It was found that the inlet angle of coolant had a significant impact on the endwall film cooling effectiveness, and the −45° inlet angle can significantly improve the endwall film cooling effectiveness at various coolant-to-mainstream mass flow rates.

Compared with the gaseous fuels, liquid fuels have better detonability and wider explosive limits. Endothermic cracking reactions of a liquid hydrocarbon fuel occurring at the beginning of combustion under specific temperature and pressure can produce lighter gaseous fuels with a smaller cell size, which has the potential of dramatically reducing the critical initiation energy of the mixture and increasing its detonability. Study on the influence of the gaseous product components and content from the liquid fuel cracking reactions on the deflagration-to-detonation transition (DDT) time and DDT distance was carried out to help obtain better operation condition for generating more detonable mixtures and guide the design of detonation combustion chambers for liquid fuels. An experimental study on the detonation-initiation characteristics of bi-component gaseous fuel acquired from the products of RP-3 aviation kerosene thermal cracking reactions was also conducted via optical method. A comparison of flame propagation velocity in the process of detonation-initiation with different component fuels was carried out. The results showed that the detonation may fail when the methane molar fraction was greater than 60%, and the gaseous unsaturated hydrocarbons such as olefins can enhance the deniability of the mixture fuel. Meanwhile, increasing the equivalence ratio appropriately can enlarge the explosive limit of the mixture fuel.

In the combustion instability experiment, oscillation pressure was usually sampled by the semi-infinite pressure tube, the amplitude of the sampled oscillation pressure was attenuated and the phase was delayed. Therefore, the method of combining theoretical analysis of acoustic propagation in tube, experimental research and numerical simulation was adopted to study the influences of the diameter of the semi-infinite pressure tube and the structure of the mounting seat on the dynamic response characteristics of the semi-infinite pressure tube, and the parameters such as gain and phase difference of the oscillation pressure measured by the semi-infinite pressure tube were analyzed. The results showed that when the length of the semi-infinite pressure tube was insufficient, it caused oscillation in the response curve produced by terminal reflection. Reducing the diameter of the semi-infinite pressure tube could increase the pressure amplitude attenuation, so the length of the semi-infinite pressure tube can be appropriately shortened. Mounting seat structure could lead to the oscillation in the amplitude-frequency characteristics and phase-frequency characteristics of the semi-infinite pressure tube due to the mounting seat reflection, and the larger volume of mounting seat cavity indicated the more stronger oscillation. In addition, the reason of response curve oscillation caused by mounting seat was analyzed by theoretical analysis and numerical simulation. The research content can accurately predict and correct the measurement deviation of the semi-infinite pressure tube, which has reference significance for the measurement of oscillation pressure in the gas turbine combustor.

A new type of corrugated channel cooling structure was proposed and designed for the key scientific problem of strengthening the cooling and heat dissipation of cascades. A refined numerical simulation was carried out to analyze the effect of the cold air inlet Reynolds number and corrugated shape parameters on the heat transfer performance, and the flow and heat transfer mechanism of the corrugated channel cooling structure of the high Reynolds number turbine blade was studied. The calculation results showed that the alternating peaks and troughs of the corrugated channels had a strong perturbation effect on the flow field, and the local heat transfer coefficient was 2—3 times stronger than the smooth channels; the heat transfer effect varied with positions on the same corrugation, and the heat transfer coefficient was the largest at the contraction of the channel; the heat transfer capacity of the corrugated channel was closely related to the shape of the corrugation, and the heat transfer effect was the best near H/L=0.115 when the Reynolds number of cold air inlet was large. The physical mechanism of the corrugated channel to enhance heat transfer was revealed, providing a technical support for the design of the cooling structure of aero-engine cascades.

The heat load of 2-D plug nozzle sharply increases under the afterburning condition, therefore, it is necessary to introduce cold air to cool the plug. Based on subscale model experiment verification, a numerical simulation of full-scale model was performed to investigate the effect of plug film cooling on the aerodynamic and cooling performance of 2-D plug nozzle. The effects of total pressure ratio of inlets, diameter of film holes and perforated percentage of film holes were analyzed by contrast. The calculation results indicated that: within the range of the study parameters, film cooling made the plug surface temperature significantly drop and had a little effect on the thrust coefficient of the nozzle. Compared with the case without cooling, the total pressure ratio of inlets increased from 1.02 to 1.20, the surface temperature of the plug decreased by 20% to 45% and the total pressure recovery coefficient decreased by 0.22% to 1.26%. A lower channel pressure and a bigger film outflow resistance were caused by a higher perforated percentage of film holes, which even caused the hot gas backflow in the tail of plug. Considering the cooling air discharge condition of the whole plug surface, a small perforated percentage of film holes showed more advantages. Reducing the diameter of film holes meant increasing the amount of film holes and the film coverage, thus making the cooling effect slightly enhanced.

A low-cost, high-precision experimental verification platform of film cooling effect was designed. The high-speed and high-temperature flame impact experiment was carried out for the film holes with different structures of plate samples. The infrared temperature measurement technology was used to collect and analyze the temperature field of the plate. The results showed that the temperature gradient in x direction of the two hole shapes was smaller than that in y direction, but the cat-ear hole had better cooling effect in y direction; The x-direction comprehensive cooling efficiency of cat-ear hole was also better than that of cylindrical hole, and the cooling air flow was inclined to form a stable film cover; In addition, the overall cooling effect of cat-ear hole in unit area was better than that of cylindrical hole, but greater thermal stress may occur near the outlet of cold air. The cooling area of cat-ear hole film cooling was also larger than that of cylindrical hole.

In order to meet the requirements of supercritical fuel injection for Advanced Aero Engines in the future, the schlieren system was used to conduct experimental research on the injection of near/supercritical RP-3 aviation kerosene into the still atmosphere, and the process was analyzed by combining with the 10-component substitute of RP-3 aviation kerosene. The results revealed that near/supercritical fuel injection could produce shock wave structure and phase transition near the nozzle exit. However, there were differences between near critical and supercritical jets in the overall jet structure or the jet structure near the nozzle. Compared with the near critical injection, the supercritical injection was larger in the gas-phase region/liquid-phase region in the Mach disk, and had a longer reliquefication distance; at the same time, with the increase of injection temperature, Mach-disk diameter, longitudinal distance and jet expansion angle all decreased. But with the increase of injection pressure, Mach-disk diameter, longitudinal distance and jet expansion angle increased.

A rocket liquid oxygen tank was taken as an example to simulate the temperature-drop characteristics of the liquid oxygen tank with subcooled cycle. Using liquid nitrogen as the simulated working medium, the principle scaling cryogenic propellant subcooled cycle experimental system was established, and the accuracy of the numerical mode was verified by the experimental data. The effects of the subcooled cycle flow rate and return form on the temperature-drop rate and thermal stratification characteristics of the tank were simulated and analyzed. Results showed that, due to the long cylinder section of the first stage liquid oxygen tank, the fluid in the tank was more fully mixed and the temperature uniformity of the liquid was better. For the secondary tank, due to its short axial length, part of the subcooled liquid oxygen was directly sucked out through the upper suction port of the tank, such that the tank temperature cannot be reduced to 70 K. The temperature in the tank was obviously stratified, and the temperature uniformity was poor. After the optimization of the tank structure, the cooling rate and temperature uniformity of the tank were significantly improved.

In view of the obvious pressure oscillations of 200−230 Hz in the hot test of a hydrogen oxygen rocket engine gas generator, a low frequency combustion stability simulation mathematical model was established to analyze whether there is low-frequency unstable combustion phenomenon of limit cycle related to combustion delay. Simulation results under different combustion time delay, pressure drop of injector and combustion chamber volume showed that the oscillations frequency related to the combustion time delay was significantly lower than the test data. The low-frequency fluctuation in the test data may be excited by the acoustic frequency of the feed line. Further analysis showed that the key parameter to determine the stability of the system is the ratio of combustion delay to gas residence time. When the ratio was greater than a critical value, the system became unstable, otherwise the system became stable. Based on the simulation data fitting, a semi empirical formula for calculating the inherent frequency of the system was formed. The inherent frequency of the system decreased with the increase of the sum of combustion delay and gas residence time. The stable boundary of the system under different injector pressure drops was obtained. With the increase of pressure drop ratio, the critical ratio of combustion delay and gas residence time from stable to unstable became larger.

The staged startup is an important measure to improve the startup quality and launch reliability of large thrust liquid-oxygen (LOX)/kerosene staged combustion rocket engines. The staged startup characteristics of the engine were studied by system dynamics simulation. The modular universal model library of transient simulation of liquid propulsion system (Tulips) was developed by using the Modelica language. Based on the model library, the system dynamics simulation models of the engine were built and validated layer-by-layer from bottom to top. The simulation results showed that, the startup quality was preferable when the prestage was set at the level where the thrust was higher than 40% of the rated value. The delay of entering the prestage was longer if the level of the prestage was lower. When the level of the prestage was low, the startup quality can be improved by advancing the moment of igniting the thrust chamber and the moment of transferring the resistance condition of the fuel throttle valve. If the turbopump rotational inertia was larger, the system became more stable when entering the prestage.

The thermodynamic calculation method, normal shock wave theory, heat transfer theory, and strength theory were adopted to carry out the matching design and theoretical analysis of the performance and structural scheme of liquid attitude and orbit control engines and cylindrical diffusers. A design calculation method and procedure for determining the performance and structural scheme of diffusers was presented from the aspects of its performance, main dimensions, heat transfer mode, strength and stability verification. Influence curves of the area ratio of cylindrical diffusers, combustion chamber pressure ratio, vacuum chamber pressure ratio, and gas isentropic index on the main dimensions and operating characteristics of diffusers were obtained. A cylindrical diffuser with an inner diameter of 1.1 m, a length of 8 m, and an inner wall thickness and cooling jacket width of 10 mm was designed for the testing requirements of a certain NTO/MMH liquid attitude and orbit control engine with a thrust of 5000 N. The working envelope demonstrated the ability to conduct relevant tests within the engine combustion chamber pressure range of 1—5 MPa, maximum flowrate range of 1—3 kg/s, and simulated working altitude range of 30—60 km. The research showed that using the design method for gas thermodynamic calculation introduced into the diffuser inflow before and after the normal shock wave, more accurate gas thermodynamic parameters required for cooling calculations can be obtained. In consideration of the chemical reaction of high-temperature gas, a more reasonable performance and structural matching scheme can be obtained using the method, making it convenient to determine the test capacity range of existing diffusers.

In order to study the influence of different injector structure details on the cold mixing characteristics of GH₂/GO₂ continuous detonation engine, the cold flow field of GH₂/GO₂ continuous detonation engine was numerically simulated with commercial software Fluent. Based on the annular-hole injector structure, 12 different engines with different injector structures were designed. The influences of different oxygen injector outlet expansion angles, different hydrogen injection angles, different hydrogen injector outlet expansion angles, and unilateral injection on the mixing characteristics were studied under the same inlet conditions. The results showed that the mixing effect decreased first and then increased when the expansion angle of oxygen injector outlet was within the range of 0°−20°, and the optimum expansion angle was 20°. When the hydrogen injection angle was within the range of 30°−90°, the mixing effect increased first and then decreased, and the optimal injection angle was 45°. Within the range of 0°−10°, increasing the expansion angle of the hydrogen injector outlet reduced the mixing effect. The mixing effect of gas-hydrogen bilateral injection was obviously better than that of unilateral injection.

A powder feeding device with a built-in intake channel was designed for the piston-driven powder fuel supply system in powder engines, the piston movement was realized by using the User Defined Function (UDF), and the action of the coupling of the gas-powder-piston was established. Numerical simulation was carried out to investigate the powder fuel supply characteristics under different initial operating pressures (0.6, 1.2, 1.8, 2.4, 3.0 and 3.6 MPa) in the powder storage tank based on the Eulerian-Eulerian two-fluid model. The results showed that the gas-solid interface mainly fluctuated around the intake under different initial operating pressures. With the increase of the initial operating pressure, the fluctuation amplitude of the powder flow rate decreased, the mean powder flow rate within the stable conveying stage was closer to the theoretical value, and the fluctuation amplitude of the powder layer (powder volume fraction of 0.1) area decreased; the area-averaged volume fraction of solid phase at the two-phase nozzle throat section increased with the increase of the initial operating pressure in the storage tank, but the fluctuation amplitude of the granular temperature decreased. The pressure in the storage tank at an initial operating pressure of 3.6 MPa was kept stable for a longer period compared with 0.6 MPa, and the pressure fluctuation in the storage tank was reduced by 59.1%.

In order to study the effect of acceleration field on the combustion characteristics of HTPB/AP/Al propellant, an overload test rocket with both burning rate measurement and condensate product collection functions was designed. The effects of different overload magnitudes (−50g to +50g) on combustion chamber pressure, mean combustion velocity, transient combustion velocity, condensed phase products and rocket plume were investigated by using the overload test method. The research showed that: (1) the reverse overload had little effect on the pressure of the combustion chamber and the burning rate of the propellant. With the increase of forward overload, the combustion chamber pressure and propellant burning rate increased. (2) Compared with that without overload, the combustion chamber pressure increased by 35.8% under +10g overload, 69.9% under +30g overload, and 76.8% under +50g overload, and there was a “hump phenomenon” under +30g overload and +50g overload. (3) Compared with 0g, the burning rate increased by 21% under +10g overload, 40% under +30g overload, and 44% under +50g overload. At +30g and +50g overload, the transient combustion velocity increased and then decreased, the maximum value increased with the increasing overload, the time to reach the maximum value decreased with the increasing overload. (4) The content of carbon and elemental aluminum in the medicine cup decreased by 100% and 82.28%, respectively, with the increase of positive overload, and the alumina increased by 402.17% with the increase of positive overload. There was almost no carbon and aluminum in the collection device. And the particle size of the condensed-phase product in the collection device decreased with the increase of positive overload. (5) Acceleration field had a significant effect on the color of rocket plume; under reverse overload, the engine plume flame appeared yellow accompanied by bright sparks; under forward overload, it appeared purple.

For a solid rocket ramjet with design Mach number of 3, two-dimensional numerical simulation was used to study the effects of suction holes at different locations on the unstart Mach number, anti-back pressure capability and total pressure recovery coefficient of an inlet with adjustable primary gas flow, i.e. a large range of back pressure changes. By setting suction holes in the inner compression section, the unstart Mach number of the inlet was reduced from 2.7 to 2.4, and the anti-back pressure capability was increased by 12.28%; The unstart Mach number of the inlet decreased from 2.7 to 2.6 due to the suction of the outer compression section, and the anti-back pressure capability was not improved; however, the anti-back pressure capability of throat suction inlet was increased by 11.24%, and the unstart Mach number did not change. The internal compression section and throat section suction can improve the total pressure recovery under certain working conditions, especially the total pressure recovery coefficient under over rated working conditions. Finally, a scheme to improve the performance of the inlet under Mach number of 3—5 was proposed, and the total pressure recovery under Mach number of 4—5 was increased by about 5% on average.

Taking the fault conditions of automatic actuator failure in a low-thrust liquid rocket engine test system as the research object, the models of the pneumatic liquid valve, the pressure regulator, the combustion chamber and the propellant tank were conducted. And the simulation system was built on Amesim platform. Through comparison between calculated data and test data, the reliability of new models was verified respectively. Furthermore, the simulation effect on the normal working condition of the whole test system was demonstrated, showing that the steady state error of main parameters was less than 2.8%. Based on the normal working condition, the prediction results of fault conditions under failure of the pneumatic liquid valve control cavity leakage and the pressure regulator internal leakage were calculated. The results showed that the chamber pressure can deviate from the normal conditions by 14.3% and the mixing ratio by 37.3% after the failure of pressure regulator internal leakage for 30 s. And when the leakage area of the pneumatic valve reached 24 mm², the propellant valve was completely closed. The prediction results can be accumulated for further work on fault detection, fault analysis and intelligent control of the test system.

In order to study the vibration characteristics and contact stress of an inter-shaft bearing with local defects, a two-dimensional fully flexible dynamic model of the inter-shaft bearing with single and compound faults was established based on the multi-body dynamics method. With this model, the time-frequency distribution characteristics and contact stress variation law of the inter-shaft bearing with local defects were analyzed. Compared with the lumped parameter model based on nonlinear Hertz contact theory and the experimental results, the accuracy of the multi-body dynamics model was verified. The research results showed that the error between the fault characteristic frequency obtained based on the established model and the lumped parameter model and the experimental results was less than 1%. The maximum value of the equivalent contact stress of the bearing appeared at the fault crack, spread around and reduced; the contact stress in the bearing area was significantly higher than that in the non bearing area, and the time interval of the change of the equivalent stress was the reciprocal of the fundamental frequency of the inter-shaft bearing fault characteristic frequency.

In view of the problem in identifying the damage extent of typical rolling bearing faults, based on acoustic emission parameter analysis and waveform flow analysis methods, the time arrival feature index (TAFI), count, number of hits and energy were combined, and the non-dimensional parameter fault factor as the ratio of the peak power of the fault characteristic frequency to the average power in the frequency band between its adjacent octave was introduced. An identification method of the damage extent for typical rolling bearing faults based on acoustic emission multi-parameter fusion was proposed. In order to verify the identification effect of this method on the damage extent of typical rolling bearing faults, a test bench for simulating typical rolling bearing faults was built. Test was carried out and the acoustic emission signals of rolling bearing with two defects of serious damage caused by wire-cutting and weak damage caused by pitting caused were collected. The identification method proposed was applied to the acoustic emission signals of typical faulty bearings and healthy bearings measured under the same speed conditions. The results showed that the acoustic emission characteristic parameter TAFI can preliminarily determine whether the bearing was faulty, the count and number of hits can effectively identify the fault types of rolling bearings. The characteristic parameter energy can effectively identify different damage extents of the rolling bearing with outer ring fault and rolling element fault. The fault factor parameter was introduced to characterize the damage degree of rolling bearings with different defects, and the different damage extents of rolling bearing with typical faults were effectively identified through the numerical differences of the fault factors of 1−5 times fault characteristic frequency between wire-cutting and pitting defects, which made up for the lack of fault damage extent identification of inner ring fault with characteristic parameter energy. This method can effectively identify different damage extents of typical rolling bearing faults.

Smooth particle hydrodynamics (SPH) method and Lagrange method were used to simulate the process of hail impacting and the experiments of a Hail impacting on a plate, a titanium alloy sheet and a real fan disk were carried out. The results showed that the high strain rate constitutive model and Smooth particle hydrodynamics (SPH) method could simulate the process of hail impacting precisely; the simulation results had a good agreement with the experiment results. Based on this constitutive model and method, simulation of a hail of 1.1 cm diameter impacting into an aero engine at the state of climbing was carried out and the deformation of the blades after impacting was revealed. The result showed that impacting on the tip site of the blade would have the worst situation of plastic strain. An aero engine hail impacting test was carried out, and the deformation error of the impacted blade and simulation results was less than 10%. The front edge radius and the maximum thickness distribution of the blade were improved, and the anti-hail impacting ability of the blade was enhanced, and the improved design experiences had been accumulated for anti-hail impacting.

In the double propeller rotor system of the counter rotating propeller fan, the vibration signals are transmitted between the inner and outer rotors through the intermediate bearings to generate complex coupling vibration, which usually takes the form of beat vibration of basic frequency and multiple frequency, resulting in difficult identification of vibration signals and low efficiency of dynamic balance. The conventional least square identification method only considered the basic frequency signal, but neglected the interference of harmonics and noise, resulting in low identification accuracy. In order to solve this problem, a harmonic least square method (HLSM) signal was proposed. Through simulation experiments, the identification effects of various identification methods and HLSM were compared, verifying that the identification accuracy of HLSM was higher. The results of HLSM identification were applied to double rotor dynamic balancing on the built coaxial counter rotating double rotor test-bed. The results showed that the unbalanced vibration of inner and outer rotors was reduced by 70% and 96%, respectively.

In order to study oil film force characteristics of the elastic ring squeeze film damper (ERSFD) and analyze the characteristic parameters of the dynamic system of ERSFD, based on the radial deformation of the elastic ring, the differential calculation model of the inner and outer oil film pressure of the ERSFD was established. The differential calculation model was solved by using the finite difference method. The influence of the height, number, width of the elastic ring boss, and the inner and outer connecting holes of the elastic ring on the oil film force characteristics of ERSFD was studied through numerical simulation. The results showed that compared with the traditional squeeze film damper, the radial deformation of the elastic ring changed the internal and external oil film clearances, helping to restrain the nonlinear change of the oil film stiffness with the height of eccentricity. The elastic ring boss in ERSFD regulated the distribution pattern of the internal and external oil film pressure of the damper. The existence of the boss reduced the squeeze oil film area, reducing the stiffness and damping of the internal and external oil films of ERSFD. The oil holes connected inside and outside the elastic ring reduced the stiffness of the inner oil film and increased the stiffness of the outer oil film.

Volume of fluid (VOF) method was used to simulate the free level change of fuel sloshing in a multi-compartment wing fuel tank. Fuel sloshing characteristics of the tank under multiple operating conditions were analyzed. According to the analysis results, the rib spill port configuration, the diameter of oil strings and the number of ribs were selected as design variables, and the relative displacement amplitude of fuel mass center was taken as the optimization objective to design the orthogonal experimental design. The rib configuration of the tank was optimized by range analysis and ANOVA, and the optimal rib configuration was obtained. The results showed that VOF method can accurately and effectively simulate the fuel flow characteristics, and the fuel tank rib can effectively restrain fuel sloshing. By comparing the optimized rib tank with the initial tank, the average relative displacement of the fuel mass center was reduced by 66.54% and the maximum displacement by 46.43%. The displacement of fuel mass center was effectively controlled and the anti-sloshing effect of wing tank was improved.

To address the problem that the installation error in the encoder signal makes it difficult to extract the bearing outer race fault features, an error suppression scheme for encoder installation was proposed. According to the method, the envelope on the original instantaneous angular speed signal was obtained based on the encoder mounting error characteristics combined with the root mean square envelope technique. The encoder installation error components were further fitted based on the segmented stepwise approximation technique and the absolute mean difference index combined with the inverse slope correction method. The optimized envelope window length was determined adaptively by improving the energy ratio index, and the corresponding residual signal was obtained. Through the envelope spectrum of the residual signal, the fault characteristics order of the bearing outer race was obtained and the fault characteristics were revealed. Compared with the traditional cepstral pre-whitening method which suppressed periodic components, the first 3 fault characteristic orders of the outer ring of rolling bearing were clearly extracted by this method, proving that the proposed method has better suppression effect on encoder error. The effectiveness of the proposed scheme was verified by simulation and experiment.

In order to improve the wear resistance of rolling linear guide, combined with the wear test and theoretical analysis, the maximum impact force, maximum contact stress and friction resistance are taken as optimization goals. The constraint function of slider reaction force and the range of structural parameters were determined by desgin requirement. By analyzing the parameter sensitivity, initial contact angle, curve ratio of the slider raceway, ball diameter and the increment of the rail height were selected as variables; The optimization used CCRD and Kriging response surface to construct the agent model. Though the NSGA-Ⅱ algorithm, the friction resistance and maximum impact force of the structure were reduced by 2.09% and 15.00% respectively, while the maximum contact stress increased by only 4.33%. The ball diameter decreased from 5.56 to 5.1744 mm, and the initial contact angle decreased from 45 to 38.878°. Though Spearman correlation analysis, it is known that the coefficient value between the ball diameter and the slider reaction force is 0.23, and the coefficient value between the initial contact angle and slider reaction force are −0.82. These two variables balance each other to achieve the optimal solution in optimization.

A method of CFD simulation was used to calculate the discharge flow ratio of different fillets, hinge forms, and aspect ratios. Impact mechanism of different pressure relief flow was also researched. These results could support the engineering design of nacelle pressure relief door. The research showed that the fillet only increased the discharge flow ratio slightly. In addition, pressure relief doors of different hinges had different discharge flow ratios, the goose neck type was superior to the pivot hinge type and the page hinge type, due to the hinge installation which caused the changes of the outside heading flow and the exhaust channel. Specifically, the pressure relief door with goose neck hinge had the strongest discharge capacity owing to the largest exhaust channel, and some air flowed from the installation slot. However, the external flow hit the page hinge type and then blocked the flow from both sides, which caused the weakest discharge capacity. The external flow of the pivot hinge type had stronger ejection effect on the discharge flow, causing stronger discharge capacity than the page hinge type. However, the discharge capacity of the pivot hinge type was weaker than the goose neck type because of the smaller exhaust channel. The pressure relief door with larger aspect ratio had larger discharge flow ratio, due to the reduced winding flow of the external air and the increased exhaust flow owing to the coiled vortex.

Secondary power system (SPS) is a complicated airborne system which undertakes varieties of key tasks. Three typical SPSs of mechanical link, pneumatic link and electric link types were introduced with examples of U.S. fighters, and their advantages and disadvantages were summarized. The research on SPSs developed by U.S. force and defense contractors was reviewed and analysis on technical feature of these SPSs was conducted. Besides, four key technologies in advanced SPS were summarized and the extended applications of SPS were briefed. Finally, the development trend of SPS was prospected. Four key technical characteristics of advanced SPS in future were presented: multiple electric architecture, high integration of structure, comprehensive utilization capability of onboard energy and synergies between main engines and itself.

To establish reasonable turbine design parameter calculation method and guarantee the accuracy of design work on the basis of quadratic function, an approximate computation method of S2 stream surface circumferential angle and blockage coefficient inside blade row passage was proposed for Euler equation time-marching through design method. The through-flow design method utilized finite volume method to solve 2D unconservative Euler equation in orthogonal curvilinear coordinate. Exact Riemann solution was used to calculate interface parameters of grid element. And third order Godunov scheme with TVD property was implemented. While semi-implicit scheme was employed in temporal discretization., the blade profile loss, secondary loss, and blade tip clearance leakage loss were computed by empirical loss model. Then secondary loss and tip clearance leakage loss could be redistributed radially. On the other hand, shock wave loss was regarded as accurate after solution. A single stage turbine was then designed with the through flow design method. After that, 3D blade shapes were profiled according to through flow result. Next, 3D viscous CFD software was used to simulate turbine and verify the effectiveness of through flow design method. Given the same inlet and outlet conditions, compared with 3D result, through flow mass flow result was about 1.46% higher, expansion ratio was 0.005 lower, and isentropic efficiency was 0.0077 higher. Finally, it can be concluded that through flow design method required fewer number of grid points, featuring higher computation efficiency with favorable convergence. Besides, the calculation method of blade row passage S2 stream surface circumferential angle and blockage coefficient was rational. And also, accurate results of turbine overall performance, and dimensionless parameter like flow coefficient, loading coefficient, and reaction can be obtained.

In order to study the impact of different fuels on the overall performance of rocket based combined cycle (RBCC) engines, a quasi-one-dimensional overall performance simulation model of RBCC engines was built. The thrust and specific impulse performance of RBCC engines fueled with liquid oxygen kerosene, hydrogen peroxide kerosene, liquid oxygen methane, and liquid oxygen liquid hydrogen were studied. Combined with the integrated aircraft engine performance analysis model, the effects of different fuel engine performance on flight mission capability were studied. The results showed that at ejector mode, the thrust of hydrogen fuel was 1.3 times greater than kerosene fuel; the hydrogen fueled RBCC engine had the farthest cruising distance of 4470 km; under the same aircraft parameters, hydrogen peroxide kerosene fuel RBCC powered aircraft had the highest maneuverability. This method can provide a reference for the overall performance scheme design and fuel selection of RBCC engines.

Based on the computational fluid dynamics (CFD) method and noise solved FW-H (Ffowcs Williams-Hawkings) equation, the aeroacoustic noise of rigid coaxial rotor was calculated in hover, and the influence of aerodynamic interference on noise characteristics was analyzed. The study indicated that: in hover stage, the surface pressure of upper and lower rotor blades changed periodically with the rotation of the blade, and there were eight small cycles in one revolution; the periodic variation of the sound source led to inconsistent “superposition effect” produced by the sound pressure at observation points in different directions in the plane of the rotor disk, and the noise level at the intersection of upper and lower rotors was higher than that at other azimuths; due to the time-varying load, there was still significant radiation noise in the direction of the rotation axis, so the noise sound pressure level in this direction was much larger than that of single rotor. With the increase of rotor spacing, aerodynamic interference between twin rotors was reduced, and the normal force fluctuation of the maximum noise sound pressure level on the sound radiation sphere also decreased.

In order to improve the mixing efficiency of the air flow between the inner and outer bypasses of the rear variable area bypass injector (RVABI) in the variable cycle engine, and to improve the velocity uniformity between the mixed air flows, a method of adjusting the outer bypass area of the RVABI with lobed structure was proposed. By means of numerical simulation, the total pressure loss, thermal mixing efficiency, velocity distribution and vortex evolution were studied and analyzed at the bypass ratio of 0.11−0.23, and the two reference models were compared. The results showed that the method of adjusting the outer bypass area of the variable area bypass injector by introducing the lobed structure can significantly improve the uniformity of the flow field velocity and greatly improve the thermal mixing efficiency of the inner and outer bypasses. And with the increase of the bypass ratio, compared with the reference configuration band, the thermal mixing efficiency of the area adjustment method of the lobed structure was improved more significantly. The mixing degree depended on the scale and influence range of the flow vortex; the lobed shape and corresponding area can be optimized synchronously to further improve the regulating performance.

To improve the prediction accuracy of exhaust gas temperature (EGT), the complexity of the data should be reduced. A combined empirical modal decomposition (EMD) and long short-term memory neural network (LSTM) method was proposed to predict EGT. First, EGT data with time series characteristics were decomposed into intrinsic mode function (IMF) and residual (RES) containing the same characteristics using EMD; the components were predicted using LSTM model; and the results predicted from all components were superimposed to obtain the predicted values of EGT. The prediction results of EMD-LSTM model and single LSTM model were compared and analyzed. The results showed that the former had 35% and 42% lower root mean square error and average relative error than the latter. It indicated that this model has better prediction accuracy in predicting the EGT value of APU.

In order to solve the problem of performance degradation that consequently reduces the engine control quality and even endangers the engine operation safety, a method of estimating the real-time state and performance variation trend for the actuator was proposed. Considering the measurable signals of practical aero-engine servo actuator were less than the performance parameters, a adaptive estimation method of combined state was put forward by means of recognition and classification of the actuation pattern, the balanced current of electro-hydraulic servo in the steady state was estimated by unscented Kalman filter, the actuation gain and actuation time delay in the dynamic state were estimated by the method of Broyden-Fletcher-Goldforb-Shanno (BFGS) , the performance parameters were updated in real time, and the adaptive model of servo actuator was established. A vane servo actuator loop of turbofan engine was simulated. The simulation results showed that, when single servo parameter can be measured, the absolute error of balance current estimation was less than ±0.2 mA, the relative error of actuation gain estimation was less than 4%, and the absolute error of actuation delay period estimation was less than one control period in different actuation states, and the adaptive model can estimate the state of actuator accurately and track the performance variation trend in real time, so the method can provide technical support of the control loop design and fault diagnosis for aero-engine servo actuator.

Based on the conventional dual redundancy digital electronic control system, a scheme of engine control system with digital electronic backup was proposed. When the main control system failed, the electronic backup system can take over the main control system to realize various functions of the engine, and improve the task reliability of the system. The backup system was built into the hydraulic and mechanical components to enhance the system’s electronic anti-interference and electromagnetic resistance. The results of semi-physical and engine tests verification showed that the main and backup systems, as well as the system and the engine worked well; the main and backup switching functions were normal, and smooth transition of the switching process parameters was realized. When switching in cruise state, the thrust disturbance was less than 1.5%; each control function of the backup system was normal, the control accuracy of the main control parameter pressure ratio was less than 0.1, the swing was less than 0.1, the control accuracy of the guide vane was less than 0.15°, and the swing was less than 0.2°, the fuel and guide vane followed well during acceleration and deceleration, and the steady-state and transient performance met the engine operating requirements; the heat dissipation design also met the operating requirements of the device.

Based on the unscented Kalman filter (UKF), the self-calibration unscented Kalman filter (SUKF) and the multiple-model estimation (MME), considering the influences of unknown inputs (such as drift error of the IMU in medical manipulator, gust encountered by the running train and failure of onboard components) on the strongly nonlinear system state equation in engineering, the multiple-model self-calibration unscented Kalman filter (MSUKF) was proposed to expand the application scope of the multiple-model self-calibration Kalman filter (MSKF). According to the Bayes' theorem, this filtering method used the UKF and the SUKF whose weights were assigned automatically to obtain the final filtering result through weight-average way. A large number of simulation results showed that the accuracy of MSUKF was 50% higher than that of divergent UKF, and 4% higher than that of unbiased SUKF, presenting stronger adaptability and robustness.

In view of the characteristics of low dynamic viscosity and hard film forming of liquid oxygen on the sealing face, a hydrodynamic-hydrostatic hybrid mechanical seal was proposed. Considering the hydrostatic effect of the orifice and the cavitation effect of the liquid film, the finite element analysis model of the liquid film was established to study the working principle of the hydrodynamic-hydrostatic hybrid mechanical seal and analyze the influences of the geometrical parameters of the hydrodynamic and hydrostatic structure on the sealing performance. Within the research scope, the optimum range of the geometrical parameters of the spiral groove and hydrostatic structure was as follows: the groove-weir ratio was 0.6—0.7, the groove-dam ratio was 0.6—0.7, the groove depth was 12—15 μm, the spiral angle was 15°—18°, the orifice diameter was 0.4—0.5 mm, and the orifice number was 6—12.

In order to study the relationship between blade machining deviation and compressor stability, and establish the quantitative correlation between common machining accuracy and compressor working range, the middle section of two-stage stator blades of multi-stage high-load axial compressor was taken initially as the research object, a reduced order model of blade surface geometric uncertainty was constructed, and the deviation blade profile database under three kinds of common machining accuracy was generated. Combined with the relationship between the neural network prediction uncertainty input variables and the range of critical cascade angle of attack, pseudo-Monte Carlo method was used to generate a large number of samples and carry out statistical analysis. The results showed that, compared with the prototype, the introduction of machining deviation decreased the cascade positive critical angle of attack, and increased the negative critical angle of attack. Therefore, the range of critical angle of attack decreased, and the aerodynamic performance deteriorated more than the prototype. Taking the positive critical angle of attack as an example, when the machining accuracy increased from level 2 to level 1, the mean value of the positive critical angle of attack decreased from 7.4858° to 7.5571°, and the sensitive part of the blade changed from the whole blade area to the leading edge and the front half chord length, and the influence trend of the increase or decrease of the profile on the critical Angle of attack also changed. Given the statistical quantification from above analysis results, the research conclusions could provide a theoretical basis for the design and optimization of 3D cascades or compressor rotors in the future, thus further saving the machining cost.

In order to numerically study the impact of distorted inflow on a transonic compressor with less computational costs, a three-dimensional simulation method was developed based on body-force model (BFM). Firstly, the modeling method of building body-force source term was studied. The test data of the Darmstadt transonic compressor were used to verify the body-force model in performance prediction of the blade rows with clean inflow. Furthermore, steady simulations with a 180° circumferential total temperature and total pressure distortion were performed using BFM. This showed that the BFM can reproduce the distorted compressor characteristics. The evolution of the inlet distortion through the compressor and the circumferential phase change of the total temperature across the rotor were in line with unsteady Reynolds averaged Navier-Stockes (URANS). Under inlet total temperature and pressure distortion, the coupling trend of compressor and upstream distorted flow field was opposite. Compared with URANS, the BFM can significantly reduce the computational costs by 5 orders of magnitude while effectively capturing the main flow features inside the compressor.

In order to reveal the effects of the slip state on the flow field evolution under different working conditions of the dynamic pressure gas foil bearing, taking the dynamic pressure gas foil bearing as the research object, the modified Reynolds equation for circumferential partial slip was established and solved by the over-relaxation iteration method. The influence of slip on the flow field in the bearing clearance was analyzed under the change of rotating speed, eccentricity, clearance height and foil deformation. The results showed that with the increase of rotating speed or the decrease of clearance height, the fluid-solid interface changed from no slip state to the stator side slip state, and finally reached the slip state at both sides. The area and velocity of the slip region on the rotor side from the pressure rise region and the stator side from the pressure drop region to both sides gradually increased. However, as the eccentricity increased, the slip velocity on both sides increased gradually, and the slip area was almost unchanged. The deformation of the foil caused the slip velocity field to exhibit a stepped distribution. Partial slip had a maximum effect of 17% on the highest film pressure. The reason for the above law is that the changes of velocity gradient and pressure field make the interface shear stress and ultimate shear stress change complexly.

In order to predict the three-point contact phenomenon and reduce the risk of abnormal wear, with the help of Newton-Euler vectorial mechanics, combining with the force and structure features, relative location and velocity relationship of internal element of the arched inner race ball bearing were described, the force of bearing parts, inertia force, gyroscopic torque, fluid drag force and extrusion force were considered in process of the arched inner race ball bearing quasi-dynamic modeling considering the three-point contact condition, the optimization of algorithm for solving nonlinear equations of the model was completed, and the accuracy of the model was validated by comparing with foreign mature software and test of the cage rotational speed. The results indicated that the proposed model had a maximum error of 12% for contact angle and contact stress distribution, and a maximum error of 2% for maximum contact stress, and also had high precision for the contact traces prediction, at the same time, the model had a maximum error of 10% for the cage rotational speed by the test validation, so that the model can be used for improving optimization design of the arched inner race ball bearing.

In view of the failure characteristics of high temperature and high speed bearings, such as adhesive wear and cage fracture, the bearing dynamic analysis model and bearing testing rig were built. The optimized design, performance analysis and test verification of the bearing were carried out. The research results showed that with the increase of the speed, the contact stress of the inner ring increased, and the contact stress of the outer ring decreased. The maximum contact stress of the hybrid ceramic ball bearing was greater than that of the all-steel bearing; with the increase of the speed, the rolling elements collided with the cage. The force, cage slip rate and spin-to-roll ratio increased, and the cage stability was reduced. The cage slip rate and spin-to-roll ratio of the all-steel bearing were both greater than those of the hybrid ceramic ball bearing, and the collision force of the rolling elements and the cage was equivalent; at the oil supply temperature of 110 °C, the speed was 120000 r/min, the temperature of the hybrid ceramic ball bearing was lower than that of the all-steel bearing, the vibration acceleration of the hybrid ceramic ball bearing was lower than 2.0g, and the vibration acceleration of the all-steel bearing was lower than 4.0g. It can be judged that the temperature rise of hybrid ceramic ball bearings was lower than that of all-steel bearings through the temperature difference between supply and return oil; the test verified that hybrid ceramic bearings were more suitable for high-temperature and high-speed working conditions than all-steel bearings.

In view of many degradation features of aero-engine performance and their mutual influence, the RUL（remaining useful life） prediction method of aero-engine based on Copula similarity was proposed considering the nonlinear correlations of the degradation features. The working state of the aero-engine was classified through K-means clustering, and a degradation model was established to select three sets of sensors with the most obvious degradation performance trend. Based on the Copula function, the correlation modeling and analysis of the selected three sets of sensors were carried out to build the Copula structure between engine sensors. The prediction of the remaining life of aero-engine was realized based on Copula similarity. The results showed that compared with traditional methods, the prediction errors of the aero-engine RUL based on Copula similarity were reduced by 13.053%, 31.328% and 74.602% respectively, and the prediction accuracy was improved.