By adopting the reaction class-based global sensitivity analysis (RC-GSA) methodology and species sensitivity analysis (SSA) methodology, the fuel-relevant skeleton mechanism (containing 25 species and 44 reactions) of methylcyclohexane (MCH) was developed. At the same time, with use of the decoupling method, the skeleton mechanism (containing 72 species and 383 reactions) of MCH was obtained. Furthermore, based on the genetic algorithm, the rate constants of fuel-relevant reactions in the skeleton mechanism of MCH were optimized. The results showed that the optimized skeleton mechanism can greatly improve the prediction accuracy of ignition delay time of MCH and MCH, CO, CO₂ concentrations. By coupling the skeleton reaction mechanisms of n-decane, n-dodecane, iso-hexadecane, MCH, toluene and the reduced reaction mechanism of C₀—C₃, a reduced mechanism (containing 121 species and 469 reactions) of the surrogate fuel for RP-3 kerosene was formed. The errors between the simulated ignition delay time and laminar combustion velocity by the reduced reaction mechanism and the corresponding test values of RP-3 kerosene were less than 5% under multiple working conditions. The calculated values of the main species concentration in the oxidation process were in good agreement with the corresponding test values except for some conditions.

In order to improve the comprehensive combustion performance of trapped vortex cavity afterburner and better understand the effect of radial flameholder installation angle on flow and combustion characteristics, numerical and experimental investigation were employed at 5°, 8°, 10°, 15° under different excess air coefficients. The investigations were performed under the bypass temperature 600 K, the core temperature 1000 K, the core pressure 180 kPa, the bypass ration 0.17, and excess air coefficient 1.27—2. Through investigation, the flow filed, exit temperature, total pressure recovery coefficient, wall temperature of trapped vortex cavity (TVC) and radial flameholder were obtained. The results showed that when the angle was 5°, the total pressure recovery coefficient was 0.002—0.006 lower, the exit temperature was 20—40 K higher, the TVC wall temperature was 63—82 K lower, the radial flameholder wall temperature was 60—70 K higher than that at the angle of 10°.

The influence of hydrocarbon fuel flow deviation on the allowable Mach number upper limit of scramjet was focused and researched. A quasi-one-dimensional coupling model of combustion and cooling was established and verified as an overall analysis tool. The simulation results showed that the working range of scramjet was seriously affected by the flow deviation. Taking the flow deviation $ \,\beta $=−0.5 as an example, the flow deviation lowered the upper limit of the engine flight Mach number from Ma=6 to Ma=5, indicating that the flow deviation seriously restrained the safe operating speed range of the engine and affected the flight mission. The influencing mechanism of different combustion and cooling designs on the upper Mach number limit was compared and analyzed, such as counter flow cooling, multi-stage fuel injection, flow deviation initial position, and maximum wall/fuel temperature. The flow deviation was suppressed and the flight Mach number range was broadened. Finally, based on the experimental data of the cascade-throttle flow deviation suppression method, the upper flight Mach number limit of the engine was extended by 1 Mach number.

In order to improve the cooling performance of ceramic matrix composite combustion liner wall, a new quasi-elliptic multi-inclined hole cooling structure was designed on the basis of the existing circular multi-inclined hole cooling structure. Meanwhile, the cooling effects of two kinds of multi-inclined hole structures were compared by using three-dimensional numerical simulation method. The results showed that: compared with the circular hole feature simulator, the high temperature area on the wall of the new quasi-elliptic inclined hole simulated part was obviously narrower, the hot spots at high temperature were reduced, and the temperature distribution was more uniform. In the warp yarn direction, the high stress area of the simulated part with quasi-elliptic inclined hole features was reduced, and the stress concentration phenomenon was weakened. In the weft yarn direction, the high stress regions in the feature simulators of quasi-elliptic inclined hole and circular hole were found.

In order to understand the effects of sustainable aviation fuel (SAF) on nozzle atomization characteristics and combustor performances, fuel atomization tests were conducted on SAF and aviation kerosene RP-3, as well as combustion performance tests based on annual reverse-flow combustor. The test results indicated that compared with RP-3, under the same fuel supply pressure difference, the mass flow rate and spray cone angle of SAF were slightly smaller, and the Sauter mean diameter was slightly larger. At normal ground temperature, the ignition boundary of SAF was more than 33% narrower than RP-3, and the ignition time was 1.8—2 s longer than RP-3. The ground idle lean blow-out performances of SAF decreased by about 17% compared with RP-3. The exit average temperature of the combustor for SAF was slightly higher than RP-3. The exit temperature distribution of both SAF and RP-3 was essentially the same, with comparable combustion efficiency, overall temperature distribution factor and radial temperature distribution factor. Within the scope of the study, the differences in atomization characteristics and steady-state combustion performances between SAF and RP-3 were not significant. The main differences lied in the poorer ignition and lean blow-out performances of SAF compared with RP-3.

The fine dust particles entering turbine blades are easy to deposit and weaken the heat transfer performance of the internal cooling channels. In order to explore the deposition characteristics of fine dust particles in the cooling channel inside the blade, the effects of fine dust particle diameter and Reynolds number of cooling air on the deposition of fine dust particles were experimentally studied. The experimental results showed that the deposition rate of the internal cooling channel increased firstly and then decreased with the Reynolds number of the cooling air. When the Reynolds number was up to 23810, the deposition rate reached the peak. However, when the Reynolds number of the cooling air continued to increase, the deposition rate showed an obvious downward trend. When the diameter of fine dust particle increased, the deposition rate in the blade channel increased. The deposition rate at the front cavity of the blade was higher than that at the middle chord cavity and the back cavity, and the collection rate at the trailing edge was higher than that at the exit of the back cavity and the dust removal hole. In the blade channel, the amount of particles deposited on the leeward side of the ribs was greater than that on the windward side, and the amount of particles deposited at the corner of the front cavity was greater than that at the corner of the middle chord cavity.

The effects of different foam metal porosities on the temperature, pressure, adsorption capacity and total adsorption capacity in the hydrogen storage tank during the filling and degassing processes were investigated by numerical simulations with MOF-5 and AX-21 as adsorbents, respectively. In the filling stage, the addition of foam metal into the hydrogen storage tank can significantly improve the effective thermal conductivity in the hydrogen storage tank, reduce the average temperature and increase the average pressure in the tank, contributing to the adsorption reaction. In the degassing stage, the addition of foam metal can increase the average temperature in the hydrogen storage tank, which promoted the desorption of hydrogen. During the degassing process, with the reduction of foam metal porosity, the obvious area of desorption reaction gradually expanded toward the center of the tank, helping to accelerate the desorption and release of hydrogen. There existed optimal filled foam metal porosity leading to the maximum total hydrogen adsorption in the hydrogen storage tank. For MOF adsorbent, the optimal porosity was 0.8; for AX-21 adsorbent, the optimal porosity was 0.9.

To study the characteristics of turbulence fluctuation mass exchange between the concentric staged combustor stages, large eddy simulation was conducted for 9 kinds of swirlers and total pressure loss combination schemes, and partial combination schemes were verified by particle image velocimetry optical experiments. A new quantitative method of turbulence exchange was proposed and the variety rules of turbulence exchange phenomena in 9 kinds of combination schemes were achieved. The results showed that the intensity of turbulence exchange was in direct proportion to the inlet mass flow when the swirlers were kept constant, meanwhile, counter-rotating swirlers caused the relationship between swirlers outlet flow to change from synergism to antagonism, the increased momentum dissipation between main swirlers and pilot swirlers made it difficult for central recirculation zone to maintain a stable state under the repression of first main swirler outlet flow. Finally this made the central recirculation zone contract towards the combustor dome and reduced the intensity of turbulence exchange in the central recirculation zone. In case of counter-rotating, of the 2nd pilot swirler, the intensity reduction of turbulence exchange reached about 70%. And in the slender recirculation zone, the region with the strongest turbulence exchange was in the middle of the recirculation zone.

Using the axisymmetric nozzle under the afterburner as the research object, the effects of flanging methods of heat shield on nozzle cooling were studied numerically. The results show that regardless of whether the heat shield is flanged on both sides or at the outlet, the maximum and average temperatures of the heat shield show a pattern of first decreasing and then increasing with the increase of flanging height; flipping both sides of the heat shield can increase the average and maximum temperatures of the convergent section wall. On the contrary, outlet flanging can reduce the maximum and average temperatures of the convergent section wall; in case of flanging on both sides, the maximum and average wall temperatures of the divergent section can be reduced by up to 2.6% and 1.1%. During outlet flanging, the maximum and average temperatures of the divergent section wall are greatly affected, and can be reduced by up to 14.6% and 23.5%, respectively. When the cooling of the nozzle divergent section wall is taken as the research objective, the outlet flanging is more conducive to the cooling of the nozzle wall.

In order to study the oil distribution characteristics of the under-race lubrication oil passage, the volume of fluid (VOF) method was used to calculate two-phase flow in the axis jet oil receiving scoop, and the formation process of oil film and the flow field characteristics in the passage were obtained. The effects of the oil temperature, shaft speed, oil volume flow rate, and aperture combination on the oil flow distribution were emphatically discussed, and the dimensionless correlation formula of the critical aperture ratio was established. The results showed that, the oil jet impinged on the center of the oil receiving scoop and formed an oil film, and the oil droplets and the oil belts formed by the broken film edge were thrown to the side wall, finally the oil film covered the entire end face. Within the calculation range of operating conditions, the distribution of lubricating oil was mainly affected by the oil volume flow rate and aperture. The flow rate at each outlet increased linearly with the increase of oil volume flow rate, and the oil distribution ratio decreased by an average of 15.05% with the increase of oil volume flow rate. The oil distribution ratio increased with the increase of downstream hole diameter and the aperture ratio. When the dimensionless oil supply flow was larger and the downstream dimensionless hole diameter was smaller, the critical aperture ratio was higher and approached to 1, when the dimensionless oil supply flow decreased or the downstream dimensionless hole diameter increased, the critical aperture ratio declined.

In order to study the influence of the surface gap or grooves on hypersonic aircraft the aerothermal environment and the thermal protection system design, a numerical simulation method coupled with heat transfer of solid field was constructed. The numerical simulation analysis of the perfect gas flow passing a two-dimensional gap, original and reversed “T”-shaped three-dimensional gaps was carried out for isotherm and distributed temperature surfaces coupled with heat transfer of solid field. It can be found from the numerical results that the aero-heating effect at the corner of the rear facade of the two-dimensional and three-dimensional gaps was the worst. There were heat flux peaks at the side of the horizontal and vertical intersections of the original and reversed “T” shaped gaps. The airflow directly impacted the rear facade on the horizontal and vertical intersections of the “T”-shape gap, where a peak of aerodynamic heat flux of 26.18 W/cm² of the entire gap was produced. Due to the vortex structure in the gap, a local heat flux peak value of 6.125 W/cm² was generated on the front facade at the horizontal and vertical intersections of the reversed “T”-shape gap. Considering the influence of the coupled heat transfer of solid field, the wall temperature roses under the action of aero-heating, which reduced the heat conduction effect of the high-temperature gas in the flow field on the wall. Therefore, the overall heat flux distributed inside the gap, on facade at the side of the gap and on the upper surface of the gap was reduced by about 25%, and the peak value at the inflection point on the rear elevation decreased the most by about 32.04%.

The species transport model, discrete phase model (DPM) and Eulerian wall film model were used to simulate the heat transfer characteristics of the spray cooling on circular chemically patterned modified surfaces under uniform heating conditions. The effect of wettability gradient on heat transfer performance on surfaces was studied by comparing four modified surfaces’ liquid film thickness, liquid film flow rate, surface temperature inhomogeneity and average surface heat transfer coefficient. The results showed that the heat transfer performance of the wettability gradient surfaces was better than that of uniformly wet table surfaces, which can promote the liquid drainage and optimize the liquid management. Surfaces with large wettability gradients had faster liquid film flow rate and smaller liquid film thickness, and hydrophobic surfaces may have better liquid drainage than hydrophilic-hydrophobic hybrid surfaces, yet hydrophilic-hydrophobic hybrid surfaces had the best heat transfer performance through a good hydrophobic design to promote nucleation and hydrophilic performance to retard drying, and can improve surface temperature inhomogeneity. Increasing the number of wettability and increasing the gradient can enhance the heat transfer.

In order to deeply understand the importance of the flow control of supercooled water film on the aircraft surface for anti-icing/deicing, the influence of supercooled water film on aircraft icing was elaborated from supercooled water film formation process and water film icing model, and then the main ideas and methods of supercooled water film flow control were proposed. The influential factors of supercooled water film flow were introduced, and the effects of environmental parameters such as wind speed and intrinsic characteristics of material surface, including roughness on the flow of supercooled water film, were systematically analyzed. On this basis, the research status of indirect control of supercooled water film flow by regulating the movement of droplets was summarized, a new idea of using gradient non-wetting surface to directly control the movement of supercooled water film was proposed, and the important problems and development trends to be solved urgently in the development of supercooled water film flow control were comprehensively prospected.

To further understand the drag reduction characteristics of tube-fin vortex reducers (TFVRs), large eddy simulation (LES) was used to compare and analyze the turbulent pulsation and vortex scales in a tube vortex reducer (TVR) and a TFVR, and the vortex breaking mechanism was revealed by coherent structures and power spectrums. Results showed that the tube in TVR reduced the pressure loss by breaking the large-scale vortex, while the fin in TFVR can further destroy the upstream Ekman boundary layer and inhibit the development of large-scale vortex, and the reduced swirl ratio effectively weakened the surge phenomenon of small scale vortex at the inlet of tube, increased the entropy in this region, and achieved a higher degree of vortex suppression, and this effect became more significant as the installation height of the lower end of the fins decreased. The energy integral length scale first increased and then decreased with the decrease of the radial height in TFVR. The peak value in TFVR moved towards the high radius direction, and the energy integral length scale was lower downstream of the cavity, compared with TVR.

In order to obtain the interaction mechanism between different nozzles and pulse detonation duct burner and improve the thrust enhancement performance of pulse detonation duct burner, the pulse detonation duct burners without nozzle, with seven convergent nozzles, seven convergent and divergent nozzles and five fluidic nozzles were numerically calculated. The results showed that the convergent nozzle could reflect the compression wave, slow down the discharge of high-pressure working substance, and then improve the axial force gain. The convergent divergent nozzle could further accelerate the flow on the basis of the convergent section, but it could make the shock wave enter the divergent section some time in a cycle, resulting in partial thrust loss; The fluidic nozzle could adjust the working substance parameters in the expansion section of the nozzle, but it could also bring some mixing loss. When the fluidic nozzle was used under these three working conditions, the axial force gain increase rate was the highest, which was 22.11%, 15.06% and 15.23%, respectively.

The multi-stage brush seal has the problem of premature failure due to the uneven distribution of pressure drop between different stages. A new two-stage pressure equalizing brush seal structure with the opening of the downstream backsplash was proposed, and a three-dimensional entity numerical calculation model of the new two-stage pressure equalizing brush seal was established. On the basis of verifying the accuracy of the numerical model, the leakage flow characteristics of the new two-stage pressure equalizing brush seal were analyzed numerically. The influence of pore parameters on the leakage characteristics of the new structure and the balance of pressure drop between stages was studied. The results showed that the new two-stage pressure equalizing brush seal with pressure equalizing hole improved the pressure drop equalization and reduced the friction heat effect of the brush seal compared with the traditional structure. The pressure drop equilibrium coefficient of the traditional two-stage brush seal was obviously greater than that of the new two-stage brush seal, and the pressure drop equilibrium coefficient of the new structure was improved by 45.6%—67.9% compared with the traditional structure. When the diameter of pressure equalizing hole increased from 0.2 mm to 0.8 mm, the height of pressure equalizing hole increased from 3.55 mm to 8.35 mm, and when the number of pressure equalizing hole rows was set to 3, the pressure drop equilibrium coefficients decreased by 38.5%, 7.7% and 25.1%, respectively. The leakage amount of the new structure was slightly larger than that of the traditional structure, and the leakage amount increased with the increase of the diameter of the pressure equalizing hole, but had no obvious change with the increase of the height of the pressure equalizing hole, and increased with the increase of the number of pressure equalizing hole rows.

It is generally believed that the error of gas temperature calculation based on enthalpy conservation method is caused by the error of direct measurement and the influence of high temperature decomposition. However, in the experiment of a certain high temperature combustion chamber, it was found that, taking thermocouple measurement results as the reference standard to evaluate the performance of gas analysis method under high temperature environment may introduce a large response error, so the turbulent flame experimental data of Sandia laboratory were used as reference standard to compare and analyze the temperature of each measuring point. The results showed that the overall average error of the temperature calculated by the enthalpy conservation method was small compared with the temperature measured in the experiment, but there were large errors in the initial reaction region and the combustion region. Further research found that: the uncertainty of temperature measurement in gas analysis was mainly derived from reaction progress. The uncertainty of temperature calculation results can be quantified by using reaction progress variable, and the applicable range of gas analysis method can be judged and limited, providing a reference for further research on the application of the enthalpy conservation temperature measurement method in the temperature measurement of turbulent flame and engine combustion chamber and the evaluation of uncertainty.

Wind tunnel test, flight experiment, e^N method, transition mode and transition criterion were introduced, and the improvements of these methods in the sensitivity principle and transition criterion were pointed out. On this basis, hypersonic transition characteristics of compression corner, swept-back leading edge and swept-back rudder surface were analyzed in detail, and the key problems in Görtler vortex growth mechanism and cross-flow instability research were sorted out. This indicated that on the basis of deepening the understanding of transition mechanism, it is necessary to improve the simulation capability of quiet wind tunnel and the accuracy of CFD simulation to meet the needs of engineering on boundary layer transition prediction.

Considering the design requirements of hypersonic aircraft’s binary variable geometry inlet regulating mechanism for light weight, low energy consumption and high accuracy, multi-objective optimization was carried out for the mechanism, the design scheme with the best comprehensive performance was obtained and the feasibility of the scheme was verified. Firstly, through force analysis and ADAMS software simulation, the minimum driving force required for the wedge plate and throat plate was obtained. Then the mathematical model of the mass, energy consumption and stiffness of the mechanism was established, the design variables and constraints of the mechanism were determined, and the Pareto solution set was obtained by multi-objective optimization using NSGA-Ⅱ optimization algorithm. The visualization of Pareto frontier was realized by drawing the level diagram, and a group of optimal solutions were selected for the design scheme. Finally, the feasibility of the scheme was verified by mechatronics concept designer (MCD) kinematics simulation analysis. The results showed that the weight of the mechanism was reduced by 6.48% and the energy consumption reduced by 8.35% compared with that before optimization, and also it can meet the displacement demand of actuation.

The supersonic civil aircraft has objective market prospect in the future, and the key lies in development of propulsion system with economy, environmental protection and comfort. The overall development of supersonic civil aircraft was analyzed. The design differences and challenges of supersonic civil aircraft propulsion system with low fuel consumption, low noise and low emission were compared and analyzed with those of subsonic civil aircraft propulsion system. The advantages and bottlenecks of turbojet configuration, medium-bypass turbofan configuration and variable cycle configuration engines as supersonic civil aircraft propulsion systems were discussed in detail. The internal principle of difficulty in taking into full account the fuel consumption, emission and noise was analyzed. An idea of taking fuel consumption, noise and emission into full account by improving engine flow capacity was put forward, and the main reasons for limiting the current engine flow capacity were analyzed from two aspects: low-pressure turbine work limitation and inner bypass flow area limitation. Finally, based on the original innovation of aero-thermal layout of variable cycle engine, for the purpose of providing reference and support for the development of supersonic civil aircraft propulsion system, the integrated design and evaluation methods for performance, emission and noise should be developed to support the development idea of realizing simultaneous development of future supersonic civil aircraft propulsion system’s fuel consumption level, emission level and noise level.

The computational model of a solar array on an irregularly curved surface for its application to a stratospheric airship was established using numerical discrete method, and parameter investigations under different heading angles of the airship were carried out. A layout optimization platform based on the solar radiation and solar array models was developed to optimize the area of the solar array, and the optimum layouts of the solar array under different heading angles were obtained. Results indicated that flat and curved surfaces can lead to large differences in the performance of the solar array, and the heading angles strongly affected the output power of solar cells. The proposed optimization method can help effectively reduce the area and weight of the solar array under the condition of sufficient output energy of the solar array. Moreover, the area of the solar array was reduced by 21.9% when the heading angle was 90°.

In order to study the turbine aerodynamic characteristic variation when the fuel changed from natural gas to mixed-hydrogen fuel, an aerodynamic characteristic definition for turbine was given, in which the definition of corrected mass flow and corrected rotation speed can integrated the gas property variation. The turbine aerodynamic characteristics 2-stage cooled high pressure turbine and 6-stage uncooled power turbine of LM2500+SAC gas turbine was investigated based on the RANS (Reynolds averaged Navier-Stokes) simulation method and the new characteristic definitions. Also, the fuels variations from natural gas fuel to mixed-hydrogen fuel are studied. The results shows that, the two work characteristic curves can be classified using recommended aerodynamic parameters definition (the corrected mass-flow distinction between the two fuels was under 0.2% at the same corrected rotating speed and total pressure ratio conditions), and it was indicated that the real turbine aerodynamic characteristic of mixed-hydrogen fuel could be gotten from the turbine aerodynamic characteristic of original natural gas fuel using the characteristic definition. This provides convenience for the application turbine aerodynamic characteristic data.

Considering the difficult problems existing in current extremely low speed calibration facilities that the calibration accuracy for anemometers is not high enough and the calibration is vulnerable to the contamination of temperature and humidity discrepancies, an extremely low speed calibration facility with high control precision of velocity was designed with air temperature and humidity being able to be controlled to required values. The facility had a velocity range of 0.1 to 1.0 m/s, an air temperature range of ambient atmospheric temperature to 60 ℃ and a humidity range of 20%RH to 80%RH. The layout arrangement was discussed. The mechanical structure and its components and accessories were designed and illustrated. The motion velocity control system and temperature and humidity control system were designed, tested and adjusted. The results of the adjustment tests showed that the maximum velocity control error and the maximum relative velocity control error were 0.000989 m/s and 0.241%, respectively, satisfying the design target error 0.003 m/s and 0.4%, better than the international optimal accuracy 0.02 m/s and 0.5% (wind tunnel type), and 0.001 m/s and 0.82% (probe-moving type); the maximum temperature control error was 0.9 ℃, meeting the design target error 1 ℃. The maximum humidity control error was 2.9%RH, less than the design target error 4%RH. When the facility was applied to the calibration of a hot-wire anemometer, the maximum error of the fitting curves of the calibration results was 0.02236m/s, which satisfied the design target accuracy 0.03 m/s and was within 0.018—0.03377 m/s, the error range of the international wind tunnel calibration facilities, and close to 0.014 m/s, the error of the international probe-moving facilities; the maximum relative error of fitting curves was 5.214%, close to 4%, the error of the international wind tunnel calibration facilities and close to the lower limit of 2.42%—15.04%, the international probe-moving facilities. The estimated velocity uncertainty of the facility was about 0.0159 m/s, close to the lower limit of 0.014—0.06 m/s, the uncertainty of the international wind tunnel calibration facilities, and close to 0.012 m/s, the uncertainty of the international probe-moving calibration facilities.

For the improvement of jet noise modeling, comprehensive analysis of experimental data was conducted to study the normalization characteristics of jet impingement noise. The experiment of subsonic jet impingement on an inclined plate was carried out in a full anechoic chamber. A microphone arc array was applied to collect the far field sound data. The nth power law of jet velocities was utilized to analyze the jet impingement noise. The results of overall sound pressure level (OASPL) showed a good agreement with the nth power law, but the velocity factors (n) varied in the observation directions (θ). The velocity factors of OASPL in directions of 30°, 90° and 120° were 6.8, 8.0 and 9.7, respectively. After normalization, the narrow-band sound pressure level spectra collapsed well and the velocity factors of main band were 7.5 and 10.0, respectively in the directions of 30° and 120°. It can be established that the growth rates of impingement noise were significant higher than the 8th power law in the upstream directions. Moreover, the normalization results and velocity factors indicated that the impingement noise had a little effect on downstream directions. It was found that wall jet-related noise dominated in downstream direction, and the spectra presented dipole noise and quadrupole noise in band of St<0.2 and 0.2≤St≤2 respectively in the direction of 30°.

Considering the scheme of the rectangular nozzle of the engine embedded on the wing of the blended wing-body layout aircraft, a control strategy based on the modification of the engine nozzle was proposed. The flow fields of blended wing-body layout aircraft were simulated numerically based on Reynolds average Navier-Stokes equation. The effects of jet pressure ratio, flap deflection angle, flap leading edge radius, upward deflection flaps and combined flaps on the lift-enhancing effect of blended wing-body layout aircraft were analyzed. The results showed that when the flap deflection angle was 40° and the jet drop pressure ratio was large, the negative pressure peak of the leading edge of the flaps decreased and the jet flow separated prematurely at the leading edge of the flap. Increasing the leading edge radius of the flap could help reduce the centripetal force required for deflection and promote jet adhesion. The design of upward deflection flaps and combined flaps at the engine nozzle weakened the influence of the right vortex and surface cross flow on the jet adhesion, which can promote the adhesion of the jet under large drop pressure ratio and large flap angle. Compared with the unmodified nozzle configuration, the combined flap design can achieve an average net thrust deflection angle of 56.10° and the lift coefficient increased by 0.16 when jet drop pressure ratio was 1.45 and angle of attack was 0°, and the lift coefficient increment was kept stable within the range of the calculated angle of attack.

An end-to-end deep learning model for reverse airfoil design based on attention mechanism was proposed to simplify the reverse airfoil design process. The model could learn the relationship between the airfoil curve and the pressure distribution, and the matching airfoil picture could be generated by directly inputting the pressure distribution image. 6561 samples were generated, with 6000 used for training and 561 for validation. The model's root mean square error on the verification set was 0.0023, its average relative deviation was 2.53%, its training and verification time was 743.4 s and 12.18 s, respectively, and its average prediction time for an airfoil curve was 0.0217 s. The results demonstrated the model's high accuracy and efficiency.

In view of the looseness of bearing supports in rotor systems, the mathematical representation of bearing impact process was established based on the geometric characteristics of loose supports. The bearing force of the loose end was calculated by Lankarani-Nikaravesh model. By using Newton's law, the dynamic model of the rotor system with bearing support looseness was established. The Runge-Kutta method was adopted to solve the dynamic equations. The influences of the factors, such as looseness degree, and working speed and damping, on the dynamic performance of the rotor system with bearing support looseness were studied. The results showed that: when the support of the rotor system was loose, the odd frequencies, such as f_s, 3f_s and 5f_s etc., could show up on its spectrum. Besides, the service life of the bearing could be significantly reduced by the increase of the bearing force and the frequency of load alternation caused by the looseness of the support. With the looseness being heavier, the accuracy of the rotor system in the looseness direction decreased gradually. The rotor system might become more unstable as the degree of looseness increased, the operating speed increased, and the system damping decreased. The research results can provide a reference for the fault analysis of the bearing support looseness. It also can be applied into the looseness diagnosis of bearing pedestal.

In order to clarify the fatigue characteristics of the rigid rocker arm of the aero-engine and meet the requirements of the rigid rocker arm during actual operation, a rigid rocker fatigue test bench of the VSV mechanism was designed based on a certain aero-engine. In order to simulate the actual working conditions, the equivalent aerodynamic load was applied to the equivalent blade by the loading device. The stroke of the actuator was controlled by PID and operated according to the preset stroke-time curve. The rigid rocker arm was subjected to 1000 fatigue cycles under the working condition, and the changes of blade angle and mechanism hysteresis force during the fatigue process were analyzed. The test results showed that the hysteresis force of the left piston rod was greater than that of the right one. The hysteresis force of the first stage connecting rod was much smaller than that of the other stages, and the blade rotation angle had higher accuracy after 1000 fatigue cycles.

A promising indicator was presented to identify connection states. A numerical method was proposed to compute the wave conversion induced by the local change of the connection joint. The cylinder with a flange was used as an example. The Zhong-Williams scheme was extended to resolve the repeated-root wave shapes by developing an iterative method based on the inverse power method. In combination with the finite element model of the flange joint, the general steps were introduced to construct a diffusion matrix. The dispersion curves and the forced responses were subsequently employed to verify the numerical accuracy of the proposed method and the diffusion matrix. Finally, propagating waves with a circumferent wavenumber of 0—8 were entered into the joint. The local changes allowed for reducing the elastic modulus of the bolts for simplicity. Results showed that the variations in a single bolt, adjacent bolts, and diagonal bolts can convert the energy from the incident waves to some propagating and evanescent waves. The amplitudes of the converted waves increased up to 36 times. It indicated that wave conversion is a promising indicator to monitor the local statuses of joint structures.

In view of the complex working conditions and prominent vibration of the variable speed turbine rotor, the optimization design technology of the variable speed power turbine rotor support layout and squeeze film damper was established. Based on the theory of similar structure and dynamics, a rotor experimental system simulating a certain type of engine was designed and built to verify the critical speed distribution, vibration mode and rotor vibration response characteristics under each scheme. The experimental analysis of the impact of damper structural parameters was carried out to provide support for the design of squeeze film damper of variable speed power turbine rotor. The research results showed that the maximum error in calculating the critical speed of the rotor experimental system was below 5%, and the maximum error in modal vibration mode was below 7%, which verified the accuracy of the dynamic characteristics calculation method. Through optimization of the rotor support structure and damper, the vibration amplitude of the rotor was reduced by 67.3%, and the operation was more stable within the working range of variable speed, which verified the effectiveness of the proposed variable speed turbine rotor vibration reduction scheme, thus laying a foundation for the dynamic design experimental verification research.

In order to research the combined cycle fatigue properties of composite, this paper comprehensively analyzes the existing design ideas and research progress of existing combined cycle fatigue test platforms, aiming to building a combined cycle fatigue test platform for composite materials. Based on the fatigue failure mode of composite, the 2D braided composite specimen is designed as the asymmetric dumbbell shape for the fatigue test. The test result shows that failure occurred in the fatigue assessment area of the specimen. The strain curve of the failure site shows the characteristics of low frequency and high amplitude caused by low cycle fatigue load, and the characteristics of high frequency and low amplitude caused by high cycle fatigue load. The completion of fatigue test demonstrates a test method for the research of combined cycle fatigue test of composite materials.

Using the method of Python’s secondary development of ABAQUS, a refined shell back head structure model was established. According to the characteristics of the fiber winding angle and thickness of the composite shell head, the calculation program was compiled to accurately calculate the winding angle of each longitudinal winding layer and the thickness of the winding layer varying with the radius. The finite element analysis model of the rear head structure of the solid rocket motor composite shell was established, and the strain was analyzed under certain internal pressure load. The accuracy of the model was verified by comparing the analysis results with the experimental results. The UMAT （user-defined material） subroutine was used to introduce the damage failure criterion to analyze the shell structure, and the failure position and failure form of the structure were further obtained when the internal pressure load gradually increased. Finally, it was concluded that the final failure form of the winding layer structure was the fiber fracture failure near the equatorial position of the head, helping to provide a basis for the design of the composite shell structure in the future.

For the random censored data in the reliability engineering test, a nonparametric reliability evaluation of the product was realized based on the Nelson-Aalen (NA) estimation theory from the perspective of the analysis of cumulative failure rate function. The nonparametric maximum likelihood estimation of the cumulative failure rate in continuous and discrete forms was given by using the obtained discrete samples, and the NA estimation form of the cumulative failure rate function under the random censored samples was derived. A completely nonparametric confidence evaluation model was derived from the reliability by NA estimation. A generalized weighted moving average model was constructed to estimate the reliability after the maximum observation time of the sample. Finally, example analysis showed that when the life distribution information was completely unknown, the NA model realized the effective confidence evaluation of product reliability based on random censored life data, and the estimated relative bias rate was controlled below 0.9787%, and the estimation accuracy was improved significantly with the increase of sample size and the decrease of censoring ratio. The results verified the validity and evaluation accuracy of NA reliability calculation.

The mechanical process of the interaction between the deformation of the positioning end face and the centering cylinder was considered for the rotor flange-bolted connection structure and its load environment, and a quantitative evaluation method for slip damage was proposed. Based on this, the mechanical process of interface slip damage under complex load and its key influential factors were analyzed. The results showed that due to the interface discontinuity of the rotor bolted connection structure, the deformation in different directions between the positioning end face and the centering cylinder could interact with each other, and could be affected by the process and assembly characteristic parameters.

Considering the deficiency that the classic models of M-C (Manson-Coffin) and SWT (Smith-Watson-Topper) can only predict the low cycle fatigue life at a fixed temperature, a high-accuracy, strong-generalization modified model for predicting life at various temperatures based on the total strain, plastic strain, maximum stress, and other parameters was proposed. The life prediction results for the nickel-based alloy from room temperature to 923.15 K low cycle fatigue test points showed that 94.74% of the prediction results of the proposed model were distributed in the ±1.5 times scatter band, which were more accurate than the M-C model (82.89%) and the SWT model (86.84%). Simultaneously the modified model was widely generalizable, which was not only capable of accurately predicting the low cycle fatigue life at non-test data points, but also applicable to other hot component materials. Additionally, the proposed model can accurately predict the full-field life of hot components at various temperatures. Full-field life prediction experiments on a gas turbine radial turbine showed that the prediction results were distributed reasonably just like expected trends.

The fluid flow force on the deflector is of great significance for analyzing the vibration and whistling faults of the deflector-jet servo valve. The existing models of fluid flow force only consider the jet impact force while neglecting the static pressure on the walls of the V-shaped groove. For this, it was proposed that the fluid flow force was mainly composed of two parts: the impact force generated by the oil impacting the walls of the V-shaped groove in the primary jet phase, and the static pressure on the walls of the V-shaped groove in the pressure recovery phase. Based on the mass conservation law, the momentum conservation law, and the continuum hypothesis, the variation of jet impact force and static pressure with aspect to the deflector displacement was investigated, and a novel mathematical model of flow force was derived. Then, experiment and CFD simulation for the pressure characteristic of the deflector-jet amplifier were conducted to verify the accuracy of the proposed flow force model. The results showed that the larger deflector displacement indicated the larger difference of pressure area between the two sides of the V-shaped groove, and the more obvious effect of static pressure on the flow force. Validated by the CFD simulation, the mathematical model of high accuracy can provide the theoretical support for the vibration and resonance analysis of deflector-jet servo valve.

The flow characteristics of large diameter butterfly valves are one of the foundations for fast and steady control of air-in total pressure, air-in total temperature, and air-out static pressure in aero-engine altitude simulating test. The numerical simulation was adopted to accurately describe its flow characteristics. The influences of pressure ratio, opening angle, valve diameter, and valve upstream static pressure on flow coefficients of large diameter butterfly valves were researched. The results showed that the flow coefficients firstly increased then remained constant with decrease of pressure ratio. And it had a critical pressure ratio. The inherent flow characteristics of large diameter butterfly valves had approximately equal percentage. The flow coefficients decreased sharply at the beginning, then slowly along with the opening angle. The influences of valve diameter and valve upstream static pressure were smaller and can be ignored. On this basis, the influence significance of pressure ratio, opening angle, and valve diameter may be obtained. The primary and secondary order was opening angle>pressure ratio>valve diameter. The response surface analysis showed that there were significant interaction influences on the flow coefficients between pressure ratio and opening angle. And the flow coefficients varied dramatically when the opening angle was equal or lesser than 55° and pressure ratio was above or equal 0.62, indicating the adjustment sensitivity of the large diameter butterfly valves was good. Finally, the table on flow coefficient of large diameter butterfly valve was obtained on the basis of pressure ratio and opening angle. Using the test results, the maximum relative errors of the simulated values and the corresponding test values were 5.07%.

For the difficulty to carry out testability modeling and evaluation due to the lack of information, the testability modeling and evaluation method for multisource information fusion were studied. Firstly, referring to testability structure model and Bayesian networks model, a layered hybrid testability model consisting of system component unit, failure, mode, test, directed edge and node conditional probability was constructed. Then, the method of determining the node conditional probability was proposed by combining expert experience and experiment data, the testability evaluation reasoning model was established, and accurate description of probability information under small sample condition was realized; Lastly, it was verified that the testability evaluation result based on hierarchical hybrid model simulation has the highest accuracy, compared with the traditional modeling evaluation method, the fault detection rate was reduced by 0.79%.

In view of the problem that the conventional PI decoupling control method of turboprop engine has insufficient robustness in the whole envelope, a PI decoupling control method based on model reference adaptive compensation was proposed. Based on the rotor dynamics and order characteristics of turboprop engine, a method to obtain the transfer function matrix of turboprop engine was also proposed. On this basis, combined with the zero pole elimination method, the decoupling link and double loop PI controller were designed respectively. By establishing the expected state space equation of the closed-loop system, a model reference adaptive compensation method was designed to effectively improve the robust performance of the controller. In addition, in order to further improve the dynamic tracking performance of the control method, a fast adaptive law based on adaptive rate covariance adjustment was adopted. Based on the nonlinear simulation model of turboprop engine, simulation verification of the control method was carried out. The results showed that at different flight envelope points, compared with the conventional PI decoupling controller, the PI decoupling control method based on model reference adaptive compensation can shorten the maximum regulation time of gas turbine speed by about 15%, reduce the overshoot by more than 90%, and cut down the maximum overshoot or droop of power turbine speed by more than 60%. It exhibited higher dynamic control effect, superior robustness and disturbance rejection ability.

To study the influence of starter power characteristics on aero-engine starting performance and determine the optimal starting power demand for engine starting, the matching mathematical model of aero-engine air turbine starting system was established, and the influence of the starter with different rated powers and power curves on the starting process of the engine was simulated and analyzed. Finally, taking the minimum rated power of the starter as the objective function, the mathematical model of engine starting power optimization was established, and the genetic algorithm was used for optimization. The simulation results showed that for the same engine, the engine starting time was shortened by 10.32% when the rated power increased by 20%, and the starting time of the engine was 3.44% different when the rated power was the same but the power characteristic curve was different. And the optimized engine starting demand power was reduced by 12.59% under the condition of keeping the engine starting time unchanged. The results can provide a theoretical reference for the starter selection and parameter design of aero-engine air turbine starting system.

To investigate the detailed design of turbine rim seal, a high-radial sealing hole structure was presented on the basis of the radial rim seal configuration, and the hole was considered as the main sealing flow path. The effects of circumferential injection angle and the number of holes on the sealing efficiency and unsteady flow characteristics in the wheel-space were studied using the experimentally validated unsteady Reynolds-average Navier-Stokes (URANS) numerical method, and the influence mechanisms of the hole structure on the instability flow structure and the degree of gas ingestion in the wheel-space were analyzed. The results showed that the change of circumferential injection angle of the hole can optimize the sealing efficiency by 10.15%, and also can reduce it by 8.32% in the worst case. Negative injection angle of the hole enhanced the unsteady effect in the wheel-space, accompanied by stronger fluid shear, which led to an increase in the scale of the Kelvin-Helmholtz instability vortex structure, the increase of gas ingestion and decrease of sealing efficiency. Positive injection angle of the hole made the relative velocity direction of the mixed fluid more uniform, which reduced the scale of the Kelvin-Helmholtz instability vortex structure; the degree of gas ingestion was weakened, and the sealing efficiency was improved. Increasing the number of holes reduced the jet momentum of sealing flow, which led to an increase in the scale of the Kelvin-Helmholtz instability vortex structure in the wheel-space, intensified the depth of gas ingestion and also reduced the sealing efficiency, which, however, was higher than that of the rim seal without holes.

Using delayed detached eddy simulation/analogy acoustics (DDES/AA) hybrid simulation method, the noise reduction effects and physical mechanism of upstream stator with “arch” distributed trailing edge (Bionic S) and downstream rotor with “arch” distributed leading edge (Bionic R) on turbine turbulent wake interaction broadband noise were studied. The research found that, except for individual frequencies, the upstream “arched” trailing edge stator blades (Bionic S) and the downstream “arched” leading edge rotor blades (Bionic R) can reduce the turbulent wake interaction broadband noise at most of frequencies below 10 kHz. Based on this, the wake characteristics and physical mechanism of noise reduction of Bionic S were preliminarily explored. Bionic S can change the characteristics of the stator wake, making it show periodic changes in the span direction, and then periodical distribution of the time-space correlation coefficient of pressure fluctuation, indicating that the bionic configuration greatly enhanced the phase-delay of pressure fluctuation at different spanwise positions. The influence existed in the entire interaction between wake and suction surface during downstream transport. And the discontinuous characteristics of spatial correlation coefficient distribution of the pressure fluctuation showed the necessity of noise reduction study of bionic configuration in the future to be carried out in the three-dimensional flow.

In order to investigate the influence of high subsonic compressor blade vibration on the flow state of the boundary layer under low Reynolds number (Re) conditions, numerical simulations were conducted to analyze the rule of separation, reattachment, transition and flow loss of compressor blade surface boundary layer under different vibration frequencies. According to the simulation results, under the condition of Re=1.5×10⁵, the increase in the normal relative velocity of the vertical wall of the boundary layer caused by blade vibration could trigger the mixing of the separated boundary layer with the mainstream in advance, which promoted the advance of transition. In this case, the normal velocity pattern at normal position near the wall was fuller, which improved the ability of the boundary layer to resist separation and limited the development of the separation bubble. In addition, the boundary layer and separation bubble thickness became “thin” due to blade vibration, which reduced the accumulation of low-energy fluid at the trailing edge, and weakened flow blockage near the trailing edge and wakes mixing, thereby reducing flow loss and improving the aerodynamic performance of the high subsonic compressor blade under low Re conditions.

To explore the possibility of improving the efficiency of variable geometry turbines, the effects of the pivot diameter coefficient, pivot cross-sectional shape, end clearance, and spacing between stator and rotor on turbine aerodynamic performance were investigated by numerical simulation at rotor speed of 2700 r/min and guide vane rotation angles of −5.0°, −2.5°, 0°, +2.5°, and +5.0°. And the sensitivity of isentropic efficiency to the above design parameters and the field mechanism were analyzed. The results showed that the isentropic efficiency could be improved by 0.064% for each 1.0% increase in the diameter coefficient D/L. The rhombic section had the best improvement in turbine efficiency, and the isentropic efficiency improvement was above 0.8% at different guide vane rotation angles. The end clearance was the most sensitive parameter to the turbine efficiency. When the end clearance increased from 0.5% to 1.0% span, the isentropic efficiency decreased by a minimum of 2.08% and a maximum of 4.54%. Increasing the spacing between the stator and rotor could improve the turbine performance to some extent, and decreasing the spacing was detrimental to turbine efficiency. Suitable design parameters can effectively improve the efficiency of variable geometry turbines at low speeds.