A high pressure turbine rotor dynamic stress test method of “telemetry+long span lead wire pipe” was designed for the dual-rotor turbofan aero-engine in the whole engine state. Combined with the simulation design of telemetry cooling amid the dual rotor and customized high temperature strain gauge, the dynamic stress test of the high pressure turbine rotor of the dual-rotor turbofan aero-engine was completed for the first time in China, and the dynamic test data of the high pressure turbine rotor blades in the full speed range and the whole engine state were obtained. The test results showed that the test mechanism and its cooling method designed can effectively solve the problems such as the difficult test signal extraction and poor signal transmission stability of the high pressure turbine rotor in the whole dual-rotor engine, the customized high temperature strain gauge and its installation and protection technology can effectively overcome the harsh installation conditions of the high pressure turbine blades on the whole machine, such as extreme high temperature, airflow erosion and narrow space. The method proposed can provide an important reference for the relevant test tasks.

Focusing on the finite element method for simulating the large deformation of turbine blade under varied loads, a method for calculating the thermal stress coupled with geometric nonlinearity was proposed based on the eight-node hexahedron element. B-bar and hybrid element technologies were used to improve the precision of results; the updated-Lagrangian (UL) incremental approach was presented for the simulation of large deformation, and the Newton-Raphson iterative method was utilized to numerically calculate the thermal stress of turbine blade. Compared with ABAQUS, the relative accuracy of thermal stress and large deformation models reached 99% through examples of notched plate, cube, cantilever beam, and ring; finally, the influence of considering large deformation on thermal stress was discussed. Under temperature, aerodynamic, and centrifugal load conditions, the radial deformation of turbine blades and thermal stress decreased, and relative calculation accuracy improved by 4.67%. This research is of great benefit to the design of radial clearance and assessment of low cycle fatigue life of turbine blade, and furthermore it can provide theoretical and computational support for exquisite design of aeroengine components.

In order to explore the influence of manufacturing errors on rotor modal characteristics, the mechanisms of rotor modal localization, mode step, and frequency steering characteristics were described based on rotor dynamics and perturbation theory. According to actual rotor assembly engineering, the manufacturing error forms of typical mating surfaces were characterized by user-defined functions, and point cloud data were generated. The skin model method was used for the first time to introduce manufacturing errors into the rotor finite element model, and the characteristics of manufacturing errors on its frequency steering, mode step, and mode localization were analyzed for this model. The mode displacement localization factor was used to quantify the degree of rotor vibration mode localization caused by manufacturing errors. The results showed that when the manufacturing error was considered to a certain extent, it could induce the rotor detuning, leading to the change of the rotor system stiffness and aggravating the frequency steering characteristics; at the same time, through analysis of the mode confidence criterion diagram, it can be seen that the mode shape had the phenomenon of dislocation step and sequence step; the detuning effect caused by manufacturing error could make the vibration energy gather in some areas of the rotor, so some frequencies falling in the frequency pass band of the ideal model may fall into the frequency band gap after detuning, and the phenomenon of mode shape localization occurred; further quantitative analysis showed that the mode displacement localization factor can effectively characterize the degree of vibration mode localization. The research methods and results can provide a reference for complex rotor assembly technology.

In view of the problems in condition monitoring of aero-engine main bearing, such as the difficulty in obtaining the real fault samples, the limitation in defining the general alarm threshold under variable conditions and the difficulty to identify the early weak faults, a general alarm method for early faults of rolling bearings was proposed. This method only trained convolutional neural networks based on normal samples, constructed evolution state indicator by the characteristic distance between degraded data and normal data, and unified the alarm thresholds of different working conditions based on training labels; at the same time, the sensitivity of wavelet band envelope signal to early high-frequency fault was used to realize early warning; then, the evolution stages were divided based on the Pauta criterion, according to which the degradation and failure thresholds were determined; finally, the remaining useful life was predicted step by step based on particle filter. Three groups of test results showed that the degradation threshold and failure threshold of wavelet analysis and convolution neural network (Wavelet-CNN) based on different fault test data can be normalized around 0.6 and 1.0, and the predictions of degradation start time were 13.01%, 12.33% and 13.70% earlier than those of non wavelet methods respectively.

In order to solve the problem that the dual-rotor system of aero-engine cannot avoid the critical speed, a model of dual-rotor system with inter-shaft bearing was established. Considering the influence of unbalance sensitivity, damper effect and the load effect of inter-shaft bearing, the tolerability evaluation functions of low-pressure and high-pressure rotor excited modes were constructed, the tolerability objective function and the constraint condition for optimization design on workable mode of dual-rotor were confirmed, and the optimization design method for the workable mode of dual-rotor system with inter-shaft bearing was established. It was found that compared with the traditional critical speed margin design criterion, the maximum amplitude of the disk of the dual-rotor system with inter-shaft bearing was reduced by 39.83%, the total mass of the shafting was reduced by 2.32%, and the external force of the fulcrum was reduced by 64.98% by using the optimization design method for the workable mode, indicating that the optimization design method for the workable mode of dual-rotor system with inter-shaft bearing was effective.

In order to explore the structural reliability analysis method of the intermediate casing under multiple failure modes, a parametric finite element model was established for the deterministic analysis of an aero-engine intermediate casing. Considering the uncertainty of material properties, geometric parameters and external loads of the aero-engine intermediate casing, the limit state functions were constructed for the two most typical failure modes of the intermediate casing: static strength failure and stiffness failure. By constructing an adaptive Kriging (AK) surrogate model for two failure modes and combining with the generalized subset simulation (GSS) method, the failure probability of the intermediate casing structure was predicted. And the correlation of the two failure modes was modeled based on the Copula function theory to determine the mutual influence between them, and the calculation results were compared with AK-GSS method. The results showed that the failure probability of the intermediate casing structure system was in the order of $ 10^{-6} $. Compared with the conventional method, the computational time of the AK-GSS method for solving the failure probability was reduced by 87.7% almost without loss of computational accuracy. In addition, the AK-GSS method still had high accuracy when considering the correlation between the two failure modes of the intermediary magazine.

A method was proposed to design the simulated specimen for fretting fatigue of turbine attachment based on the consistency of damage control parameters. Afterwards, a simulated specimen for the turbine attachment of a certain aero-engine was designed. Firstly, based on geometric similarity, two-dimensional structural dimension of the simulated specimen was determined by ensuring the consistency of the maximum relative slip distance and the equivalent stress distribution within the critical distance; and then, three-dimensional structural dimension of the simulated specimen was determined by optimization analysis with the objective of consistency of the stress intensity factor within the critical crack length. Finally, fretting fatigue test of the simulated specimen was performed. It was found that the crack was initiated at the edge of the contact surface, the direction of the crack growth was perpendicular to the contact surface and the crack length at fracture was within 5% in error with the simulation, thus verifying the rationality of the design method of the simulated specimen.

To solve the problems of high cost, long cycle and difficulty in monitoring in real time the test status of the fatigue test of aero-engine turbine joint structure, the virtual fatigue test technology was studied. The fatigue test of turbine joint structure was carried out to obtain load-displacement data, which were used to construct the NARX (nonlinear auto regressive model with exogenous inputs) neural network to carry out preliminary displacement prediction. On that basis, Kalman filtering was used to correct the predicted state with the measured data, and real-time prediction and updating of virtual fatigue test displacement were realized with the prediction error less than 5%. Finally, based on 3D MAX and Unity 3D, a high-fidelity digital model and virtual environment of the turbine joint structure were constructed to realize the visual display and data visualization of virtual fatigue test process of turbine joint structure.

A multi-frequency elliptic vortex solution model for static and dynamic characteristics of the monolithic floating ring seal with shallow groove dynamic pressure was established. On the basis of verifying the accuracy of the numerical calculation method, the static and dynamic characteristics of the integral floating ring seal with no groove, rectangular groove, spiral groove and T-groove were analyzed. The changes of leakage, buoyancy and dynamic characteristics of the integral floating ring seal under different structures and working condition parameters were studied. The influence of groove type on the stability of the integral floating ring seal rotor was analyzed, and the influence mechanism of dynamic pressure groove type on the dynamic characteristics of the integral floating ring seal was revealed. The results showed that the leakage and the lift force increased with the increase of eccentricity. Compared with the non-slotted floating ring seal, the rectangular groove had the largest leakage, and the T-shaped groove had the largest buoyancy, which was 434.7% of that without groove. Under the same vortex frequency, the effective damping of the rectangular groove was maximum and positive, and the tangential flow force was opposite to the rotor vortex direction, which can restrain the rotor vortex and improve the rotor stability.

The bristles buoyancy effect of brush seal directly affect its sealing performance and service life. The theory of brush seal bristles buoyancy effect was analyzed, and the three-dimensional transient numerical model of brush seal bristles buoyancy effect was established using arbitrary Lagrange Euler (ALE) fluid-structure coupling method. Based on the accuracy of numerical model, the overall deformation characteristics of bristles were studied, and the axial, radial and total deformations of bristle tips were quantitatively analyzed. The effects of structural parameters and working parameters on bristles buoyancy effect of brush seal were studied, and the inducing mechanism of bristles buoyancy effect of brush seal was revealed. The research results showed that the bristles buoyancy effect of brush seal contributed to the disturbance of the front bristles caused by the unsteady radial airflow. The bristles buoyancy effect of brush seal can cause large deformation of the front bristles in the direction of the flow, generate radial clearance between the front plate and the rotor surface, and increase the leakage. When the inlet and outlet pressure ratio increased from 2 to 4, the average deformation of bristle tips increased by 47%, and the protection height of the front plate increased from 1.5 mm to 2.5 mm, and the average deformation of bristle tips increased by 36%, both of which enhanced the bristles buoyancy effect. The radial airflow channel was provided by the clearance between the front plate and the bristle pack, and the radial pressure gradient was generated by the inlet and outlet pressure difference in the clearance between the front plate and the bristle pack as the inducing condition of the bristles buoyancy effect. The bristles buoyancy effect can be reduced by increasing the diameter of the bristles and reducing the clearance between the front plate and the bristle pack.

DD6 single crystal superalloy was taken as the research object to analyse the hardening mechanism of material dislocation movement by describing the microstructure evolution phenomenon, and a multi-scale creep constitutive model considering microstructure evolution was established; then a creep residual life prediction method considering the creep damage by characterizing the creep damage state was proposed. The experimental results showed that the creep model improved the simulation accuracy by 57.6% compared with the θ projection method, and the model parameters were reduced by 1/3 compared with the K-R damage model. The average prediction error of the creep residual life model prediction results was only 5.59%, indicating the validity of the model.

In order to obtain an airfoil with higher aerodynamic performance and lower aerodynamic noise, the RAF-6 airfoil for a general aviation aircraft propeller was optimized. The flow field and sound field of the airfoil were simulated by CFD/FW-H method. The influence laws of four design variables, i.e. the maximum thickness, the position of the maximum thickness, the downbending angle and position of trailing edge, on its aerodynamic performance and aerodynamic noise were studied respectively. Taking the cruise state as the design point, the airfoil multi-objective optimization design was carried out with higher lift-drag ratio and lower aerodynamic noise as the optimization objectives, and the Pareto solution set was obtained. The experimental results verified that the optimized airfoil increased the propeller thrust by 14.7% and reduced the aerodynamic noise by 2.3 dB.

The calculation model of optimal rotor speed and power demand considering flight conditions and the performance analysis model of turboshaft engine with variable speed power turbine were built respectively. Based on these models, the performance analysis method and program of rotor /turboshaft engine power turbine were developed, and the performance analysis under typical flight envelope was carried out for UH60A helicopter and T700 turboshaft engine. The results showed that when the speed of power turbine becames variable, the steady-state matching working line of compressor and high-pressure turbine changed slightly. However, the isentropic efficiency of the power turbine decreased significantly as the speed decreased, indicating that further improvements in its performance were necessary. Compared with the constant speed operation mode, after a completion of a typical flight mission, the coordinated variable speed operation of rotor/turboshaft engine could reduce fuel consumption by nearly 5%.

In order to determine the influence of flow angularity on design parameters of measuring device in low speed wind tunnel flow field calibration, wind tunnel experiment and numerical simulation were conducted to prove the feasibility of the adopted method of assessment. Numerical simulation method was carried out to study the influence of local flow angularity on measuring device blockage, offset, angle of attack and probe length. The law of influence was obtained and the mathematical model of influence was established. The results showed that: the angle of attack, the offset of measuring device and the length of probe have obvious influence on the local flow angularity experiment results, if the measuring device or test method was not properly designed, the error of flow angularity test results could be too large, even leading to wrong experimental conclusion about flow angularity.

A gas turbine guide on a gas turbine starter was taken as the research object, and the effects of throat size of gas turbine guide on the performance of gas turbine starter and components were studied using experimental methods and numerical simulation. The results of experimental research on A, B, C types of gas turbine guide (average outer diameter of the throat was 111.27, 111.94, 112.34 mm) indicated: type C′ starter compared with type A′ starter, normal starting time became shorter by 14%, output shaft disengagement speed of lose efficacy starting increased by 7.1%, maximum output power increased by 11.6%, the performance of starter was improved significantly. Numerical research results indicated: type C″ compared with type A″, the flow rate of turbine stage increased by 3.6%, maximum output power of power turbine increased by 12.2%. The increase in power turbine power was attributed to the combined effect of increased flow rate, increased temperature and increased temperature drop. In summary, the outer diameter of gas turbine guide throat changed the matching point of entire gas turbine by affecting the performance of turbine stage, then the overall performance exhibited a certain degree of dispersion.

The Y-7 aircraft was used as the reference model to carry out the overall design and performance analysis of the distributed electric propulsion aircraft. The power system of the distributed electric propulsion aircraft was designed, including design of propeller parameters, correction of wing parameters, calculation of motor power and type selection and aerodynamic design of propeller, and finally the design of hybrid power system was completed. The mass increase and decrease of each part of the distributed electric propulsion aircraft were completely calculated and its flight performance was analyzed. Compared with the reference model, the range and duration increased by 540 km and 1.2 h respectively, up by 20%. Finally, the three-dimensional modeling was established and aerodynamic characteristics of the distributed electric propulsion aircraft were analyzed. The result provides a theoretical basis for the modeling, simulation and control of distributed electric propulsion aircraft and its engineering applications.

In order to investigate the effect of ship airwake on helicopter trimmed controls, a method of combining numerical simulation and helicopter flight dynamics model was adopted. The ship airwake was obtained by computational fluid dynamics (CFD) numerical simulation method, and the characteristics of ship airwake under active control were explored; at the same time, considering the effect of ship airwake on helicopter, the flight dynamics model coupled with the ship airwake was established. The results of relative hovering trim with/without airwake were calculated, and the effect of active control on helicopter trimmed controls was further compared and analyzed. The results showed that the ship airwake had a significant impact on the helicopter take-off and landing, compared with the interference of the ship airwake on the helicopter controls when there was no control, the addition of the blowing device can effectively suppress the downwash of the ship airwake, reduce the required collective pitch control by 7.8% and pedal control by 7.5%, improve other corresponding controls and alleviate the pilot control load.

In order to improve the lifting efficiency of the flapping wing, the influences of three kinematic parameters, such as reduced frequency, flapping amplitude and pitching amplitude, on the aerodynamic performance of flapping wing were analyzed by the combination of the Taguchi test and numerical solution of three-dimensional N-S equation. The results showed that compared with the worst parameters combination, the average lifting coefficient and lifting efficiency of the flapping wing with best parameters combination increased by 52.1% and 85.52%, respectively. The influences on the aerodynamic performance of flapping wing referred to reduced frequency, flutter amplitude and pitch amplitude in turn. Further, through analysis of the flow field on the flapping wing surface, it was found that the best parameters combination can enhance the intensity of the vortex attached to the flapping wing surface and promote the formation of von Karman vortex street in the wake of the flapping wing, so as to make the flapping wing have better aerodynamic characteristic.

A momentum source method (MSM) for solving the Reynolds average Navier-Stokes (RANS) equations based on the $k {\text{-}} \omega $ SST (shear stress transpot) turbulence model was adopted. For the two-dimensional simplified model of the distributed propulsion wing with lift flaps, research on the aerodynamic-propulsion coupling characteristics and physical mechanism in the vertical take-off, transition and cruise flight state was carried out. The research showed that the suction effect of the duct made the distributed propulsion wing show the phenomenon of increasing lift and reducing drag, and delayed the flow separation of the wing. Compared with the freestream condition, the stall declination angle of the lift flaps in the ducted jet significantly increased from 12° to 34°, and at the same time the lift flaps induced jet deflection, so that the total lift of the distributed propulsion configuration was effectively raised.

A three-dimensional flow field simulation on the outdoor test stand was performed with the CFD numerical simulation method by using crosswind from the crosswind device, then a simplified model was established, and appropriate boundary conditions were obtained. On this basis, joint simulations of the outdoor test stand and the engine under typical crosswind velocities and directions were carried out. From the results, the flow field characteristics of the crosswind device outlet and the engine inlet were analyzed. And the changes of the spatial uniformity of the flow field at the crosswind device outlet and the flow field at the air interface plane of the engine inlet under the influence of the crosswind velocity and direction were revealed. It can be concluded as follows: with the increase of the crosswind velocity, the quality of the flow field at the crosswind device outlet was improved and the flow field distortion at the engine inlet went worse. With the increase of the crosswind angle, the quality of the flow field at the crosswind device was improved and then went worse, and the average total pressure recovery coefficient and the maximum circumferential distortion index at the air interface plane of the engine inlet shared different variations. The maximum circumferential distortion index is more suitable to evaluate the flow field distortion at the engine inlet in this research.

Numerical simulation methods were used to compare and analyze the effects of low-pressure turbine blade vibration at different frequencies on the separation and transition of suction surface boundary layer and flow losses at low Reynolds number (Re=25000), with its aim of exploring the influence mechanism of low-pressure turbine blade vibration on separation and transition. The research showed that the relative motion between the fluid and the blade due to the blade vibration made the separation flow meet the main flow in advance. The advanced transition caused by this was able to limit the development of the separation bubble, reduce the size of the separation bubble and weaken the backflow mixing inside the separation bubble. The blade vibration thinned the boundary layer, weakened the flow blockage and wake mixing near the trailing edge and substantially lowered the turbulent pulsation level in the separation and transition process. Furthermore, the impacts above reduced the total pressure loss up to 23.02%, and improved the aerodynamic performance significantly.

Based on the development trend and characteristics of aero-engine and combustor, the demand of ceramic composite materials for new generation aero-engine combustor was analyzed, the composition, type characteristics and high temperature corrosion resistance of ceramic composite liners were expounded, the process principle, advantages and disadvantages of ceramic composite liners under different fabrication methods were summarized, and the application status of ceramic composite materials in combustor liner was discussed in detail, finally, combined with the significant anisotropic characteristics and challenges of ceramic composite materials, the key technologies for the design of ceramic composite liner of combustor were proposed. The result showed that the ceramic composite liner had obvious technical advantages in the application of combustor, and its service status on multiple advanced military and civil aviation engines verified the feasibility of developing the technical route of ceramic composite liner. However, due to the high sensitivity of ceramic composite to thermal stress and its complex heat transfer and mechanical characteristics, its engineering application still faced great technical challenges; the key technologies that need to be solved urgently in the engineering application of ceramic composite liner include the cooling design of ceramic composite liner, the combustion organization design of ceramic composite combustor, the connection design of ceramic composite liner components, and the forward strength design of ceramic composite liner.

In response to the combustion instability of a coaxial staged combustor, the addition of lobe to the first and second main stage combustion hubs of a tower swirlers was studied. The differences in cold velocity field of combustors under different structures were compared through experimental results. The large eddy simulation method was applied to obtain the global heat release rate fluctuation spectrum of the combustor, as well as the vorticity and heat release rate within a pulsation period. Finally, the velocity and heat release rates of different combustors were analyzed using dynamic mode decomposition. Adding lobes to the first main stage hub could reduce the frequency of strong vortices, and the pulsation amplitude of the heat release rate was 4% of the global average heat release rate, about 45% lower than the prototype combustor. The heat release rate mode exhibited high-frequency small amplitude. Adding lobes to the second main stage led to an increase in the maximum vortex intensity. The pulsation amplitude of the heat release rate with a main frequency of 471 Hz reached more than 30% of the global average heat release rate, and periodic flashbacks occurred, which was not conducive to the stable operation of the combustion chamber.

In view of the problem of steady flow intake control of engine intake components in icing wind tunnel, and by combining the working principle of engine intake simulation system available with intake conditions, the Kalman filter model-free adaptive control method was adopted; by establishing the system dynamic linearization data model, and using the Kalman filter to make a true value estimate of the actual dynamic flow output, the difference between the estimate and the expected value was calculated through the dynamic data model, the input amount of the speed of the pumping equipment of the system was obtained, and the steady flow intake control was carried out. Finally, it was applied to the anti-icing test of the intake components of a certain type of engine. The test results showed that the engine intake simulation system performed stably and reliably, the output intake flow met the test requirements, and the control accuracy of the intake air steady flow reached 0.1 kg/s.

Based on the liquid atomization instability theory, a semiempirical model was established to predict the Sauter mean diameter (SMD) of pressure swirl atomizers, considering the Kelvin-Helmholtz (K-H) instability and Rayleigh-Taylor (R-T) instability generated by the interactions between air and liquid film. The experiments were also conducted using phase Doppler particle analyzer (PDPA) technique and digital off-axis holography with liquid temperature ranging from 240—300 K and liquid pressure ranging from 0.5—3 MPa. The results showed that: there were circumferential waves and axial waves on the surface of liquid sheet. With the decrease of liquid pressure and liquid temperature, the instability of liquid film was inhibited, which led to the increase of the SMD. Compared with the K-H instability, the effect of the liquid pressure on R-T instability was more significant; the effects of liquid physical properties, geometrical structures and operating conditions were included in the semiempirical correlation. The predictions showed good agreement with the experimental results, and the maximum uncertainty of the semiempirical correlation to predict the SMD was about ±15% for the available experimental data, making it valuable for the prediction of atomization performance and the optimization of the structure of pressure swirl atomizers.

To investigate the effect of fuel/oxidizer mixture filling rate on the flame acceleration and deflagration-to-detonation transition (DDT) in a pulse detonation engine, an experimental study was conducted with ethylene taken as fuel and oxygen-rich air of 40% oxygen volume fraction as the oxidizer. Fully developed detonation waves were successfully obtained using different combustion chamber configurations, different ignition positions, and different numbers of obstacles at the mixture filling speeds of 0, 2.5, 5.7, 8.9 m/s, and 14.1 m/s. The results showed that the greater filling speed of the mixture indicated the faster flame development for the ability to detonate under working conditions. The deflagration to detonation transition time can be reduced to 38.9% of the filling speed of 0 m/s when the filling speed was 8.9 m/s. Meanwhile, the number of obstacles required for DDT can be reduced to 2 pairs from 3 pairs. Shortening the DDT section and the ignition section length but increasing the mixture fill rate still allowed to successfully organize the detonation, providing some guidance for optimizing the combustion chamber configuration of pulse detonation engines, and reducing the engine length and weight for improving propulsion performance.

Simulations were employed to study the film cooling effectiveness of ellipse conical holes on the leading-edge of turbine vane. The influences of two geometry parameters, forward and lateral expansion angle, on the adiabatic film cooling effectiveness were studied comparatively. And optimization was also conducted within the range of forward and lateral expansion angle, 0°—18° and 0°—16°, respectively. Results showed that the ellipse conical hole with a forward expansion angle of 1.4° and a lateral expansion angle of 11.1° presented the highest film cooling effectiveness, which was 147.5% higher than that of the cylindrical hole. Moreover, the relationship between the two geometry parameters and the film cooling effectiveness can be fitted to quartic function. When lateral expansion angle was low, the film cooling effectiveness increased with forward expansion angle. Otherwise, the film cooling effectiveness decreased with forward expansion angle. Additionally, when forward expansion angle was lower, the film cooling effectiveness increased and then decreased with lateral expansion angle. And the film cooling effectiveness was roughly constant or showed a small increase with lateral expansion angle.

The existing dryout threshold model was optimized by adding gravity and combining with capillary force and permeability solution methods, so as to obtain the best combined model (Darcy_avg(S)+SE) for characterizing the heat transfer performance of a vertical micropillar evaporator with an average error of about 7%. The effect of micropillar geometry was investigated using this model. Model predictions indicated that the maximum heat transfer capacity of the evaporator was balanced between permeability and capillary pressure; those geometries close to the optimal pitch ratio (d/l≈0.35) and higher micropillars correspond to greater heat dissipation capacity; and that micropillar arrays with smaller receding contact angles correspond to greater dryout thresholds. The increase of the dryout length under gravity led to a significant decrease of the dryout threshold, and the genetic algorithm can be effectively used to solve for the optimal size at different dryout lengths. The arrangement method affected the dryout threshold, the forked-row arrangement of the micropillar arrays increased the heat transfer capacity by nearly 13% compared with the smooth-row arrangement at optimal spacing ratio.

According to the requirements of low pollution emission of gas turbine in recent years, in order to take advantage of an elliptical combustor and improve poor matching between its outlet and the gas turbine, a concept of shape morphing was proposed, which made the inlet of the jet-stabilized combustor elliptical and the outlet circular. Using numerical simulation methods, the combustion characteristics and flow and emission characteristics at the outlet of the combustor were studied. The influence of shape morphing on the emission and flow characteristics of the combustor was investigated. Shape morphing combustor reduced NO emission by 51.26% compared with circular combustor, maintaining the advantage of low emissions from the elliptical combustor; compared with the elliptical combustor, 2.85% of NO emission was sacrificed, but the uniformity of outlet temperature was improved by 4.27%. At the same time, it can provide a more matched temperature distribution for the blades in the rear gas turbine, and prove the feasibility of the cross section gradual change concept, which can serve as a reference for further research on cross section gradual change technology.

In order to analyze the influence of interstage turbine burner pressure recovery factor on the net thrust and specific fuel consumption of mixed exhaust turbofan engine with medium bypass ratio under different flight conditions, the interstage burning was added, and the steady-state performance calculation model of component level was established based on the original engine without interstage turbine burner cycle parameters. The simulation results showed that: when the turbofan aeroengine flight height was 5 km and flight Mach number was 0.8, and the interstage turbine burner pressure recovery factor changed from 0.92 to 0.8, resulting in the increase of specific fuel consumption by 12.2%. When the flight height was 5 km, and flight Mach number was 1.8, the interstage turbine burner pressure recovery factor changed from 0.92 to 0.8, which increased the specific fuel consumption of the engine by 20.3%. When the calculation program was used in model simulation, the pressure recovery factor of interstage turbine burner was basically unchanged. However, existing studies on interstage turbine burner showed that: the interstage turbine burner pressure recovery factor could aggrandize with the increase of flight Mach number. When the flight Mach number changed from 0.8 to 1.8, the interstage turbine burner pressure recovery factor increased by more than 2%. Therefore, the calculation results were adopted from the perspective of variable interstage turbine burner pressure recovery factor to study its influence on engine performance.

The dynamic characteristics of lost motion based on transmission errors were explored. By considering the influences of torque, loading rate, and rotational speed, the relationship between lost motion and transmission errors was analytically derived. The dynamic characteristics of lost motion were substantiated, revealing that lost motion can be obtained by subtracting reverse transmission errors. The classification of lost motion was examined, and lost motion was divided into dynamic lost motion, quasi-static lost motion, and static lost motion. The testing methods for lost motion were provided, by which dynamic lost motion can be obtained through bidirectional transmission error analysis, and quasi-static lost motion and static lost motion can be acquired through hysteresis loop analysis. Experimental investigations on lost motion were conducted using examples of small and large reducers, confirming the dynamic characteristics of lost motion. It was observed that dynamic lost motion and quasi-static lost motion exhibited dynamic variations, while static lost motion was kept unaffected by loading rate dependence, aligning with theoretical analyses. In conclusion, the theoretical and engineering values of the research were highlighted.

In view of the problem of coupling multiple excitation sources of rolling bearing under conventional monitoring methods, the electrostatic monitoring technology was introduced, and a fault feature extraction method based on short-time Fourier transform and cepstrum was proposed. A test platform for electrostatic monitoring of rolling bearing was designed and built to collect the electrostatic, vibration signals of rolling bearing under normal and fault conditions; from the perspective of time domain, frequency domain and time-frequency domain, it was proved that the method of time-frequency analysis combined with cepstrum can accurately extract the eigenvalues matching with the actual bearing fault location; by comparing with the fault characteristics of synchronous vibration signal, the low-frequency characteristics of electrostatic signals were prominent and the high-frequency attenuation was fast. The test results showed that the early wear failure of rolling bearing could be accompanied by strong electrostatic phenomenon. The short-time Fourier transform and cepstrum analysis of electrostatic signals can effectively remove the high-frequency excitation sources and highlight the bearing fault characteristics in low-frequency. Compared with vibration detection, the signal source collected by electrostatic detection can reflect the bearing fault information more directly, providing an idea for equipment fault diagnosis.

In view of the difficulty to accurately extract and separate the features of the early fault signals of rolling bearings, a compound fault feature extraction method of rolling bearing based on parameters adaptive maximum second-order cyclostationarity blind deconvolution (CYCBD) was proposed. Based on different fault types, the harmonics energy ratio index was used as the fitness function, and the sparrow search algorithm was used to adaptively obtain the optimal filter length and cycle frequency of deconvolution. The obtained optimal parameters combination was used to extract the fault components in the original signal one by one, and the envelope spectrum analysis of the deconvolution signal was carried out to realize the diagnosis of compound fault of the bearing. The analysis results showed that the proposed method can clearly and accurately separate 1—4 times of the inner ring characteristic frequency and 1—6 times harmonic component of the outer ring fault from the measured signal of bearing fault under the background of strong noise, while other common methods can only extract a few fault frequencies with low resolution. The proposed method has obvious diagnostic effect, higher application value and promotion performance.

In view of the fact that the multi-scale sample entropy is greatly affected by the sample length, the coarse graining process is relatively rough, and the shortage of effective information may be easily ignored, based on the composite multi-scale sample entropy, the energy distribution between sampling points was used as the weight for coarse graining calculation, and an improved composite multi-scale sample entropy was proposed and applied to the fault diagnosis of planetary gearbox. The influences of different parameters and noise characteristics on the improved composite multi-scale sample entropy algorithm were studied through simulation signals. The stability of the improved algorithm was verified by comparing it with multi-scale sample entropy, generalized multi-scale sample entropy and composite multi-scale sample entropy. Combined with variational mode decomposition, principal component analysis and support vector machine, the fault diagnosis of planetary gearbox experimental signals was carried out. The comparison results showed that the method can effectively realize the common fault diagnosis of the sun gear of the planetary gearbox under different working conditions and structures, and the fault identification rate was more than 95%, with certain effectiveness.

A simulation model was established based on the finite difference method and the thick plate element to clarify the fluid-solid coupling mechanism of multi-layer thrust foil bearing. The compressible Reynolds equation based on the first-order slip model was linearized by the Newton-Raphson method, and the evolution law of the static characteristics with the operating condition parameters was obtained by iterative solution. The lift-off speed test rig of thrust foil bearing was built, and the judgment basis of lift-off speed was clarified. The numerical results were in good agreement with the experimental results. The results showed that when the speed was not more than 20000 r/min, the bearing capacity generally decreased by about 3% after considering the slip boundary, which should be considered. Due to the secondary wedge effect, the pressure distribution under small clearance presented an apparent “double peak” shape. The change of the friction state in the bearing clearance could lead to the fluctuation of the experimental data, based on which the lift-off speed could be determined. The comparison results with literature data and experimental data indicate that the simulation model is more suitable for analyzing the bearing under a heavy load.

A quasi-static mechanical model of the duplex angular contact ball bearings (DACBBs) with local defects on the outer raceway was established. The defect depth, circumferential extent and the time-varying characteristic of the bearing were considered. On this basis, the influences of the size and location of local defects on the internal contact load of DACBBs were systematically investigated under pure radial and combined loads. The results showed that the contact load was very sensitive to the local defect. When the left or the right row contained defect or both of them had defects, the load distribution of the two rows was quite different. The load distribution of the row with defect produced a local mutation, and the amplitude and width of the mutation increased with the increase in the defect size. However, the mutation of the other row without defects was very small. In addition, the DACBBs should be considered as a whole for research and analysis. Under different arrangements, the impact of defects was also quite different. The current study has important implications for revealing the failure mechanism, reliability analysis and design of DACBBs.

Considering the problem of empirical wavelet transform (EWT) in extracting optimal frequency band of the rolling bearing fault signal, an improved EWT method based on extracting energy envelope trend line to adaptively divide frequency band was proposed and applied to rolling bearing fault diagnosis. The Teager energy operator was used to convert the spectrum into energy spectrum, and the energy envelope was obtained by repeated Hilbert transform. Local maximum values were extracted and smoothed to obtain the energy envelope trend line, and the first-order difference was performed to select effective extreme points to adaptively divide the frequency band. A normalized fault characteristic frequency saliency index was constructed as an effective criterion for fault diagnosis and optimal resonance frequency band selection. The algorithm was verified by rolling bearing fault simulation and experiment data. The results showed that compared with the original EWT, the proposed method can effectively identify the early faults of rolling bearings and reasonably select the optimal resonance frequency band. The proposed indexes for the outer and inner race fault data can be increased by 48.0% and 174.1% on average.

Ignition experiments and numerical simulation were conducted to study a combustion heater with hydrogen and oxygen. Two different types of combustion heater, i.e: bipropellant combustion heater and tripropellant combustion heater, were studied. Both of them achieved self-ingition. Compared with the tripropellant combustion heater, the bipropellant combustion heater had shorter ignition delay time by 17.0%. The bipropellant combustion heater and tripropellant combustion heater worked stably and lasted for 190 ms and 189 ms, separately, in the experiments. The high temperature zones in the experiments and numerical simulations were similar. The high temperature zone in the tripropellant combustion heater located behind the hydrogen ignition hole, while the bipropellant combustion heater had a V-shape high temperature zone. From the experimental images of the injectors, the tripropellant combustion heater had a bigger ablation area, which was in consonance with the numerical results.

In order to study the effect of the valve opening time difference on the working process of the space liquid rocket engine using hypergolic propellants, the high-altitude simulated thermal test of a 150 N engine was carried out. The engine working stability, ignition thrust peak and response time were investigated when the oxidant valve and fuel valve were opened 40, 100, 500, 1000 ms in advance, respectively. The test results showed that the engine can be successfully ignited, and the thrust after stabilization was basically unchanged. When the oxidant valve and the fuel valve were opened first, the ignition thrust peak of the engine was about 1.01—1.05 times and 1.04—1.07 times of the stable thrust, respectively, which was equivalent to the opening of the two-way valve synchronous signal. When the fuel valve was opened first, the startup response time was extended by about 16 ms. When the valve was opened alone, the oxidant was fully vaporized, the ice particles were ejected in the middle area of the flow field, and the fuel was partially vaporized. When the valve opening time difference reached 500 ms and 1000 ms, while the oxidant valve and the fuel valve were opened separately, the output thrust of the engine was about 11 N and 6 N, respectively, accounting for 7% and 4% of the stable thrust respectively, and the output thrust of the latter fluctuated and kept falling.

Taking the launch pads consisting of two-nozzle launch vehicle power system as the research object, the thermal environment of launch pads during the rocket launching was studied by numerical simulation method. Based on three-dimensional compressible Navier-Stokes equations, k-ε turbulence model and second order TVD (total variation diminishing) upwind scheme were used to establish the gas jet model of two-nozzle launch vehicle. The research showed that the intersections of the nozzle central axis and the flame deflectors withstood the greatest impact of the guide surface, where the temperature and pressure were extremely high. The back splash caused by the rocket tail jet impinging on the flame deflectors may further deteriorate the thermal environment of the launch pad. The drift of the rocket during launching could increase the impact of the jet on the launch pad, rapidly increase its surface temperature and pressure, and reduce its service life. The research method provides an effective method for the thermal environment assessment of launch pads during launch vehicle takeoff and has important engineering application value for the safety design of thermal protection system.

In order to quantitatively analyze the measuring error source of pivot point excursion for flexible joint, the pivot point excursion at different vector angles under 10 MPa vessel pressure was calculated using ANSYS software. According to the experimental results of flexible joint, the influences of pendulum rod deformation, horizontal displacement sensor push rod measurement error, push rod measurement error and horizontal deviation of vertical displacement sensor on the measurement results of pivot point excursion were studied respectively. In addition, the influence of the structural error of flexible joint was studied by simulations. Results showed that the measured pivot point excursion was in good agreement with the simulation after correcting four kinds of errors. The error sources in the measurement of pivot point excursion were confirmed, in which the push rod measurement errors of horizontal and vertical displacement sensors accounted for 65.98% of the cylindrical envelope height error, and 77.32% of the cylindrical envelope radius error, respectively. And these led to a great influence on the axial and radial drifts of flexible joint. Moreover, the structural error of flexible joint also affected the measurement of pivot point excursion. The distribution of pivot point excursion caused by the thickness error of elastomers and reinforcements was consistent. The results can provide a theoretical guidance for the error analysis of pivot point excursion measurement of flexible joint.

To study the effects of pintle structural parameters on the combustion efficiency in transcritical combustion, a standard k-ε turbulence model and a non-adiabatic stable diffusion flamelet model were utilized to numerically study the transcritical combustion of an LOX/CH₄ pintle engine considering the real gas properties of fluid. The study analyzed the impact of various methods for calculating physical properties on the flow field within the thrust chamber, and the effects of radial and axial annular seam width of pintle injector on engine combustion efficiency were analyzed. The results showed that the size of both the central recirculation zone and the high temperature zone was reduced when considering the transcritical effect. Within a certain range, as the radial annular seam width increased, the combustion efficiency initially decreased and then increased, as the axial annular seam width increased, the combustion efficiency decreased. A greater combustion efficiency can be achieved by reducing the axial annular seam width and increasing the radial annular seam width. When the total momentum ratio was less than 1, increasing axial momentum can effectively improve mixing. However, if the total momentum ratio was greater than 1, increasing axial momentum can hinder the improvement of mixing. The combustion efficiency decreased with the increase of the total momentum ratio. When the total momentum ratio for different operating conditions was similar, the operating condition with a higher momentum ratio exhibited a higher combustion efficiency.

Based on the image data collected from the solid rocket motor mooring experiment, the image mean and variance analysis, the fast Fourier transform (FFT) of the time series of the brightness of the typical position of the flame, and the improved Hilbert-Huang transform (improved HHT) were used to analyze the data. Transient characteristics were analyzed and processed. The analysis results showed that: in the obtained frequency band, there was no obvious main pulsation frequency in the entire exhaust plume, and there were small pulsation components at each frequency, and the pulsation amplitude in the core area was smaller than that in the turbulent mixing zone. There was a correlation between the pulsations in the core region and that in the turbulent mixing zone, and the improved HHT based on the improved adaptive noise complete set empirical mode decomposition algorithm (ICEEMDAN) can be used as an effective means to analyze the transient characteristic change mechanism of non-stationary signals such as solid rocket engine flames.

In order to study the effect of equivalence ratios and hydrogen mass fractions on the propagation characteristics of a rotating detonation wave, the evolution process was numerically simulated by utilizing the kerosene and hydrogen as fuel and air as oxidant. The propagation characteristics of rotating detonation wave, the component distribution characteristics of internal flow field and the stability of detonation wave were analyzed. The simulated results showed that with the increase of hydrogen mass fractions, the range of equivalence ratios for successful initiation of a rotating detonation in the combustor was narrowed gradually. At the same time, the recorded peak pressures of detonations decreased, while the propagation speed increased with the velocity deficits showing a non-monotonous variation with the equivalence ratio. Under the fuel-lean condition, the fuel and oxidant in the fuel-rich zone were not evenly mixed and the oxygen was distributed in strips; under the fuel-rich condition, the deflagration zone increased and its reaction strength was intensified, and the edge of the oxygen-rich zone was in wavy shapes. The propagation stability of rotating detonations was higher when the equivalence ratio was 1.0 to 1.2, and the time interval from ignition to the formation of a stable rotating detonation wave reached the minimum at the stoichiometric equivalence ratio, but increased with the increase of hydrogen mass fractions.

For the improvement of the flow field similarity in multi-stage turbine modeling tests operating at medium temperature level, a three-dimensional numerical model of a two-stage turbine was constructed and validated. The influence of the test modeling criteria selection on the flow field similarity was quantitatively studied and a new method for adjusting the specific heat ratio of working fluid was proposed. Results indicated that, the specific heat ratio of working fluid under medium temperature modeling conditions was not equal to the design value, resulting in a gradual decrease of the flow field similarity along the flow direction in the multi-stage turbine; When ensuring the similarity criterion of expansion ratio, the aerodynamic load coefficient of the last stage turbine blade was reduced by 4.78%, and the blade outlet flow angle deviated by 3°—5°; When ensuring the similarity criterion of corrected power, the corrected speed of the last turbine stage is 1.29% higher, and the back pressure for the film cooling flow of the blade decreased by 2.5%—6.5%; The test working fluid generated by combusting air into a gas and mixing with vapor had the same values of specific heat ratio to the real gas used in the different classes of heavy duty gas turbine. As this working fluid was used in the test condition at medium temperature level, the three modeling criteria including the corrected speed, the corrected power and the expansion ratio were satisfied together, and the deviation of aerodynamic parameters from the design values were within the range of −0.54% to 0.52%.

An integrated computational approach was introduced for prediction of single-stage fan tonal noise. The nonlinear harmonic (NLH) method was used for computation of acoustic sources, the incidence wave in the free field was based on acoustic analogy, and the scattering wave generated by boundaries was solved by boundary integral element method (BIEM). The sound propagation inside the duct and radiation outside the duct can be predicted simultaneously. Then, the present method was applied to tonal noise prediction of the Advanced Noise Control Fan (ANCF) developed by NASA Glenn Research Center. The effect of interface locations was investigated at first as the experimental settings. The results indicated that the interface location at 0.5 times the axial spacing was most suitable for the computation of rotor-stator interaction, which was closest to the experimental results of tonal noise. Then, the effect of axial spacing on tonal noise was calculated. With the decrease of axial spacing, harmonic loading on the stator vanes and sound radiation in the far field showed an increasing trend. The shape of noise directivity at the first blade passing frequency (BPF) was similar, and the noise directivity at the second BPF showed big difference in shape and amplitude.

In order to investigate the influence of the bridge holes on flow and thermal behavior of impingement double-wall cooling for gas turbine leading blades, the impingement double-wall cooling configurations with 0°, 5°, 10°, 15° and 20° tapered bridge holes were calculated by using ANSYS CFX numerical simulation. Results indicated that the tapered bridge holes can significantly enhance the comprehensive heat transfer capacity. As the angle of bridge holes increased from 0° to 20°, the angle of holes did not have a significant influence on flow loss of inner and outer chambers for the blades. Likewise, the cooling performance of the inner chamber target wall was not sensitive to the change of the bridge hole angle. With the growing angle of bridge holes, the thermal behavior of the target wall for the outer chamber increased at first and then decreased, reaching the biggest value at 15°. When the angle of bridge holes increased to 15°, its average heat transfer intensity was 19.7% higher than the case with 0° bridge holes. Similarly, the thermal performance factor of the whole configuration also became larger at first and then became smaller. The largest thermal performance factor occurred at 15° too, which was 12.15% higher than the case with 0° bridge holes.

To reduce the residual unevenness after installation by optimizing the installation sequence of turbine blades, a threshold iterative local search algorithm (TILS) was proposed. Based on the iterative local search algorithm (ILS), this algorithm adopted the combination of threshold limited disturbance and random disturbance to escape from local optimum, which reduced the number of iteration steps required to reach the local optimal solution on average. Experiments demonstrated that this method can find the approximately optimal combination of blade sequences in a relatively short time, which improved the search efficiency by more than 20% compared with the ILS algorithm. Compared with existing group sorting methods, genetic algorithms, and cloud adaptive genetic algorithm (CAGA) algorithms, the approximately optimal solution of synthetic mass product calculated by this algorithm was reduced to 0.33%—31%, and the computation time was significantly shorter.

Motivated by the demand of airworthiness compliance of civil aeroengine components fabricated using selective laser melting (SLM) technique, the applications of SLMed aeroengine components were summarized. Additionally, reference related to airworthiness compliance of SLMed components were analyzed. Meanwhile, airworthiness compliance demonstration method of SLMed aeroengine components was investigated. Presently, some SLMed aeroengine components have passed the airworthiness certification in foreign countries. However, China is still in the initial stage in this field. According to the related reference and airworthiness compliance demonstration method of aeroengine components manufactured through traditional processing techniques, the airworthiness compliance demonstration method of SLMed aeroengine components should mainly contain the following parts: establishment of selective laser melted material specifications, certification of selective laser melting process, identification of material quality and determination of material properties.

The effective area coefficient K was calculated using the concept of effective projected area. In Comsol software, the Spalart-Allmaras (S-A) turbulence model and particle drag model were used to improve the calculation formula in the standard, and a certain type of aircraft was used as an example to simulate the flow field and particle tracking. It was found that the larger particle diameter indicated the larger effective projected area; the higher flight speed indicated the more number of particle collisions; the charging current density increased with the rise of the cruising altitude. Finally, it was concluded that the maximum charging current density of the aircraft was 395 μA/m², which was very close to the actual observation value and the error was within 1.25%.

To realize the engineering requirement for fast automatic correction of model component characteristics, an enhanced automatic correction strategy of the identification model component characteristics was proposed. Taking steady-state test data as input, the designed correction strategy allowed to analyze and select suitable characteristic correction coefficient combinations based on sensor measurements. The proposed method also coupled individual rig test data of engine to design the equilibrium equations, and used the double-loop strategy of multi-point model correction to quickly and automatically correct component characteristics. Finally, the fast automatic correction of the identification model of a certain variable cycle engine was realized. The inverse flow path disturbance, damping coefficient self-adjustment method of Newton-Raphson method and characteristic map interpolation protection logic were adopted to improve the operation rate and stability of the algorithm. The simulation results showed that the maximum output error of the corrected model was less than 0.1%, and the consuming-time was reduced by more than 98.6% compared with the common component-level model, which was simulated in single and double bypass modes on a computation with 2.10 GHz processor. The corrected model can be used for control law design and also provide a reference for determining the current real state of the engine.