Considering the multiple working conditions creep calculation of single crystal turbine blade under the typical mission profile of engine, combined with the creep constitutive model of single crystal super-alloy under varying loading, a ABAQUS/UMAT subroutine for creep calculation of high temperature structure was developed. The creep of a single crystal turbine blade under the typical design load spectrum was calculated, and the working state of negligible creep damage was identified, so as to simplify the load spectrum. The creep deformations of the turbine blade under 10000 mission cycles of fighter and 40000 mission cycles of transport aircraft were calculated respectively, and the service life was evaluated. The results showed that, under the typical load states used in the calculation, the creep damage of turbine blade under engine cruise state and below was small, which can be ignored in creep calculation. The creep deformation of turbine blade under the simplified creep load spectrum was basically equal to that under the original load spectrum; turbine blades had different creep life under different aircraft mission profiles. Under the typical mission cycle of fighter used in the calculation, the creep life of the turbine blade was about 1/14 of that of transport aircraft, which was related to the duration of high-power state of the engine.

To meet the requirements of virtual flight wind tunnel testing for aircraft with lateral jet, the measurement device composed of two independent four-component wind tunnel balances, transmission shaft and supporting crossbeam, etc, was applied to respectively measure the aerodynamic loads of the front/rear two parts model at the same time. By finite element software, the sensitivity of each balance, the interference of transmission shaft and the influence of high-pressure gas on balance were analyzed. The results showed that the transmission shaft had little effect on the force component of the balance, the interference on the pitch moment was about 2.5% and that on the yaw moment was 8%, the impact of pressure on the front balance was less than 2% and that on the rear balance was less than 9%. Based on the static calibration formula of each balance, an aerodynamic load calculation method suitable for measurement device was generated. The correctness of the method was verified through simulated loading. Finally, the performance of the measurement device for aircraft with lateral jet was verified through the wind tunnel tests. The results of static calibration and wind tunnel tests showed that static calibration data were consistent with finite element analysis results, the measurement device had stable performance and accurate measurement values, so it can meet the requirements of virtual flight test research.

Considering the difficulty of guaranteeing the surface quality of grinding and polishing of the inner hole of the support casing and the low grinding and polishing efficiency, the research on the influence of grinding and polishing process parameters on the surface quality was carried out. Based on Preston's theory, the material removal equation of spiral grinding and polishing was established, and the influence law of spiral grinding and polishing process parameters on surface quality was theoretically revealed. The variance sensitivity analysis was carried out through orthogonal experiments, and the influencing factors such as grinding wheel particle size, grinding wheel feed speed, grinding wheel rotation speed and polishing time on the material removal rate and roughness were found out. On this basis, taking the material removal rate and roughness as the optimization research objectives, a robust optimization design mathematical model based on the Kriging response surface approximation model was constructed, and the particle swarm algorithm was used for calculation and solving, bringing about 2611 sets of optimization solutions. Combined with practical engineering requirements, the optimal process parameters include target particle size 1400, feed speed 3 mm/s, rotational speed 3600 r/min, and polishing time 9 min, providing a technical support for the improvement of process quality in the field of grinding and polishing engineering.

To study the peeling propagation characteristics of outer ring of aero-engine ball bearings, two experiments were carried out. The former was component experiment, which employed outer rings with prefabricated defects, the latter was aero-engine experiment, which adopted those with peeling defects. Atomic emission spectroscopy, portable ferrography, analytical ferrography and energy-dispersive X-ray spectroscopy analysis were conducted to analyze the wear particles in lubricating oil. The results showed that the peeling of the outer ring was progressive. The fatigue wear particles, with scratches along major dimension, continuously increased in quantity and proportion with the peeling. The total amount and size of abrasive particles showed an obvious increase during the peeling development, and also a sharp increase during the rapid propagation. In short, the peeling first appeared at a certain distance behind pits, and then extended along the rolling direction of balls, which can be divided into four stages: cracks initiation, propagation, coalescing, and peeling propagation.

The effects of groove depth, oil supply pressure, and the number of oil supply holes on leakage and damping performance of SFD with piston ring end-seals were discussed based on the SFD dynamic coefficients identification experimental device. The experiment results showed that the leakage with the groove was 2—7 times that without the groove of SFD. The test results of the damping coefficient under low oil supply pressure were close to the theoretical solution of short bearing, and the test results under high oil supply pressure were between the theoretical solutions of short bearing and long bearing. Compared with the results of SFD without a groove, the damping performance could decrease with the groove structure, but the damping coefficients were close to each other when the oil supply pressure was less than 0.1 MPa. When the groove was shallow (the depth was 5 times the clearance of oil film radius, 5c), the average damping coefficient of single hole & low oil pressure and multiple hole & high oil pressure all can attain 4.0×10⁴—5.0×10⁴ N·s/m, yet when the groove was deep (the depth is 5c), the average damping coefficient of single hole & low oil pressure was large, up to 8.06×10⁴ N·s/m. As a result, the introduction of a groove can significantly improve the damping coefficients of SFD, which can be used to improve the damping performance of SFD under low oil supply pressure.

To solve the lightweight problem of rocket jet vanes, a new jet vane based on high thermal conductivity aluminum nitride (AlN) ceramic was designed. In order to investigate its feasibility, the unsteady numerical simulation method based on fluid-solid thermal coupling was established to research the working process of aluminum nitride ceramics jet vanes under different angles, and their thermal shock resistance was analyzed based on the ceramic strength prediction model in high temperature environment. AlN ceramic jet vanes were processed for the ground static jet test of solid rocket motor at different angles, and the test results were analyzed by SEM. The research showed that, the numerical simulation results were basically consistent with the test results, which verified the effectiveness of the numerical simulation method. For the solid rocket motor with the total temperature of 2284 K, AlN ceramic jet vanes can withstand the maximum mechanical shock and thermal shock caused by the engine gas in 1 s. AlN ceramic had much better high thermal conductivity (320 W/（m∙K） theoretical) and thermal shock resistance than other structural ceramics. AlN ceramic is a good alternative material for small jet vane.

Based on the mesoscopic composition and structural characteristics of particle reinforced aluminum matrix composites, a three-dimensional random meso particle reinforced composites analysis model along with its method considering particle, matrix and interface properties was established. On the meso scale, cubic particle, spherical particle and three-dimensional random polyhedron models were used to characterize the shape of particles respectively. According to the particle size distribution data obtained from particle raw material particle size analysis, a three-dimensional random representative volume element considering the random characteristics of particle spatial distribution and the probability distribution characteristics of particle size was established. Based on Ludwik model, considering the quenching hardening effect, the elastic-plastic constitutive relationship of aluminum matrix was described. The ductile damage of matrix, the elastic-brittle failure of SiC particles and the tensile cracking behavior of interface were considered. The deformation and damage process of material in uniaxial tension were simulated. The uniaxial tensile test verification of SiCp/Al2009 composite standard parts was carried out. The results showed that the maximum errors of elastic modulus, yield strength and tensile strength were less than 5%, 5% and 11%, respectively; the prediction result of elastic modulus was less affected by particle shape; among them, the three-dimensional random polyhedron model had the highest prediction accuracy of tensile strength, and can reflect the failure modes of matrix ductile fracture, particle brittle failure and interface debonding in the tensile fracture process of particle reinforced composites. The model and method can provide a useful reference for the analysis of meso damage mechanism and macro mechanical properties of particle reinforced aluminum matrix composites.

In order to clarify the static/dynamic characteristics of aerostatic bearing under high lifting, the transient model was used to analyze the formation and evolution of supporting gas film turbulence and the energy dissipation process. Then, the flow state evolution, temperature distribution, Mach number variation and vorticity characteristics in the characteristic region of the flow field were described by the method of bidirectional fluid-solid thermal coupling. Furthermore, the theoretical analysis and experimental test results were combined to clarify the characteristics of the internal flow field of the air bearing film, the bearing capacity, stiffness characteristics and micro-vibration characteristics of the system under high lifting. The results showed that there was a phenomenon of aerodynamic heating in the film flow field of air bearing with negative pressure under high lifting. The large pressure drop inside the gas film and the enhanced compressibility of the fluid led to the bad static characteristics of the air bearing. Considering the fluid-solid thermal coupling effect, the calculation accuracy of static characteristics of aerostatic thrust bearings can be effectively guaranteed under high lifting. Especially, the calculation errors of bearing capacity and stiffness were only 2.2% and 2.7%, respectively, when air was supplied with 50 μm film at 1 MPa.

The influence of location of the dry friction damper on dynamics of the shaft/dry friction damper system was studied to ensure that the vibration of the shaft can be suppressed effectively as crossing the first and second critical speeds. Firstly, the nonlinear finite element dynamic model of the shaft/damper system was established based on the theories of Timoshenko beam and nonlinear rub-impact. The responses of the system were obtained by numerical calculation. The typical response of the shaft as crossing critical speeds and the influence of damper location were further analyzed. The results showed that there existed four stages when the shaft passed through critical speeds, including periodic no-rub motion, quasi periodic rub-impact motion, synchronous full annular rub-impact motion, and finally back to periodic no-rub motion. With the damper located at the middle node, the vibration of the shaft can be effectively suppressed when crossing the first critical speed. With the damper located at the one-quarter node and three-eighth node, the vibration of the shaft when crossing the first and second critical speeds can be both effectively suppressed. However, the shaft may not work normally when the damper was located at the three-eighth node.

In order to suppress the vibration of tubular vortex in operation, corresponding damping structure design and damping characteristics research were carried out. For the tubular vortex reducer, the design parameters of the damping structure were summarized. On this basis, the design process of the damping structure of the tubular vortex reducer was summarized. This method was used to design the damping structure and calculate the damping characteristics of an engine vortex reducer, and test verification was carried out. The results showed that the damper designed by this method has good damping effect and can reduce the stress level at key points by more than 80%. The test results and the normal pressure range to achieve the optimal damping effect were close to the calculation results, verifying the accuracy of the method in predicting the damping characteristics.

To address the problems of complex recognition and poor real-time performance in the process of structural health monitoring of aircraft, a digital twin technology-based damage pattern recognition and prediction method for aircraft wings was proposed. The digital twin structural model of the aircraft wing was constructed using modular technology, and the mapping method of sensor data in the structural digital twin model was established based on probabilistic neural network, forming a fast monitoring process of general digital twin aircraft structural damage pattern. Based on an unmanned aerial vehicle, a rapid damage pattern recognition model of its wings was developed. The results showed that the damage pattern identification accuracy of the digital twin recognition model for aircraft structures reached over 96%, which could complete the dynamic trajectory planning task.

In order to improve the efficiency and accuracy of automatic detection, an improved YOLOv8 detection method was proposed. Retinex image enhancement algorithm combing guided filtering was used to improve the contrast of digital radiograph images. Then, the digital radiography images was rotated and flipped to extend the data-set. In the process of model improvement, the Bottleneck module in C2f was replaced by GhostBottleneck module to reduce additional redundant parameters, so the lightweight model was acquired. In addition, spatial attention mechanism was introduced to obtain more spatial information of the defect. The regression range of the prediction box was adjusted to improve the detection accuracy of the proposed model. Several common aluminum alloy weld defects were used for experimental testing and verification. It was verified that the mAP of the improved YOLOv8 was 92.9%, which was better than Faster-RCNN, SSD and YOLOv8. The proposed model can be used for detecting the weld defect.

The double swirler full annual combustor was taken as the experiment object, the outlet temperature distribution under different experiment conditions was compared and analyzed under high temperature/high pressure, high temperature/medium pressure and engine experiment conditions. The experimental schemes with annular bleed and simulated nozzle were designed to simulate the boundary conditions of the engine, and the outlet temperature distribution under different experimental conditions was analyzed. The experimental results showed that the outlet temperature distribution under the medium and high pressure experimental conditions was basically the same, and the hot spot area was basically the same; the outlet temperature distribution level of the middle pressure experiment was obviously better than that of the high- pressure experiment; the temperature distribution curve of high-pressure experiment was in the form of central peak, while that of medium pressure experiment was not obvious. The designed high-pressure experiment outlet temperature distribution law and value were closer to the engine experiment results; compared with the high-pressure experiment, the outlet temperature distribution value of the medium pressure experiment designed regularly had a proportional coefficient of 1.3—1.4.

In order to obtain the influences of different impingement aperture (IA), impingement spacing of flow direction of impingement hole (IFD), spacing of span direction of impingement hole (ISD) coupling effect on impingement cooling characteristics and structural thermal stress of concave wall in reverse flow combustor, CFD calculation and FEA analysis were carried out. Opt LHD in DOE was selected to determine the sample points in the design space, and a high-precision RBFNN was constructed. Based on NSGA-Ⅱalgorithm, multi-objective optimization was carried out for comprehensive cooling efficiency, non-uniform coefficient of wall temperature distribution and maximum wall thermal stress. The results showed that comprehensive cooling efficiency, non-uniform coefficient of wall temperature distribution and maximum wall thermal stress decreased with the increase of ratio of IFD to ISD, ratio of IFD to IA and ratio of ISD to IA. Through multi-objective NSGA-Ⅱ algorithm, the value range of the three objective functions of the Pareto front of concave wall impingement cooling structure was obtained, i.e.: maximum wall thermal stress was not greater than 5 MPa, comprehensive cooling efficiency was not less than 0.66, and non-uniform coefficient of wall temperature distribution was not greater than 0.16. According to the combination of the optimal structure of concave wall impingement cooling: IA was equal to 0.94 mm, IFD was equal to 4.04 mm, and ISD was equal to 5.45 mm.

To expand the performance of aero-gas turbine engine, a structure of high-pressure turbine inter-vane combustion was proposed. The high-temperature fuel was injected into the inter-vane channel after cooling the vane. And the flame was stabilized by radial vane cavity (RVC). The C3X turbine guide vane was used as the inter-vane combustion model and the effects of radial vane cavity (depth length ratio 0.4−0.6), fuel-air ratio (0.007−0.0105) and fuel temperature (300−500 K) on inter-vane combustion performance were numerically studied. It was observed that the optimization combustion effect was obtained with depth length ratio 0.5. The thermal resistance loss caused by combustion was about 7%, which can realize the approximately isothermal combustion between the vanes. The inter-vane combustion performance decreased with the increase of fuel-air ratio, and the combustion efficiency reached 98.86% at 20 mm away from the blade outlet when the fuel-air ratio was 0.007. The combustion performance of high-temperature fuel in the inter-vane channel was better than that of low-temperature fuel, and the combustion efficiency at the outlet of the blade increased about 13%. The conclusions can provide a reference for the development of inter-vane burner technology.

Three-dimensional fluid-solid coupled numerical simulations were performed for the hydrodynamic gas bearing with an axial throughflow cooling under a stable operating condition, so as to illustrate the multi-parameter effects on the static bearing load and aerodynamic heat. The results showed that the circumferential flow driven by the strong shearing of rotor was dominant in the film-layer gap. The axial throughflow was forced to follow the circumferential flow at its inlet section. Then it moved axially toward the outlet mainly from the larger-thickness film-layer zone, making the three-dimensional flow take on an obvious spiral flow feature. Among the concerned parameters, the eccentricity was identified to be the most important parameter affecting the static bearing load and aerodynamic heat. The film-layer mean gap had a stronger influence than the axial throughflow mass-rate on the static bearing load, but the situation was opposite for the aerodynamic heat. When the static bearing load was kept nearly the same, it was found that the aerodynamic heat effect was weaker in the situation when the hydrodynamic gas bearing operated with a small eccentricity and also a small film-layer man gap. For the situation when the hydrodynamic gas bearing operated with a big eccentricity and also a big film-layer mean gap, the aerodynamic heat effect was stronger, bringing about a more crucial requirement of heat dissipation.

In order to investigate the flow field and soot emission characteristics of staged combustor, the flow field characteristics were studied by experimental and numerical simulation methods, and the flow development process was revealed. The results of large eddy simulation (LES) showed that there was a 1820 Hz periodic velocity oscillation at the outlet of the pilot stage, but without obvious flow pulsation at the outlet of the main mode, and there was a precession vortex core (PVC) downstream the swirler. Numerical studies of soot emission characteristics showed that soot was abundant in primary recirculation zone. The soot concentration increased significantly with the increase of ratio of fuel and air in the lean combustion state. Along axial direction, the concentration of soot all showed a trend of decrease after rising first, and the peak corresponding to the axial position gradually moved back, ultimately resulting in a difference in soot emission from the combustor exit.

At present, few studies are devoted to the effect of structural parameters of converging diverging mixing ducts on circularly lobed nozzle ejector. Therefore, several converging diverging mixing ducts with different geometric parameters were designed firstly, and an experimental study on the pumping performance of circularly lobed nozzle and circular nozzle exhaust-ejector scaled-down models was developed. The results showed that when the throat diameter and length of the mixing duct were smaller, the pumping ratio of the circular nozzle was higher than that of the lobed nozzle within lower main flow range, but the situation was reversed as the mass flow increased. When the throat diameter and length of the mixing duct were larger, under the condition of wall-attached main flow, the pumping ratio of the lobed nozzle was higher than that of the circular nozzle within the experimental mass flow range. With the increase of the mass flow, a maximum value of the pumping ratios appeared. Furthermore, as the throat diameter and length increased, the maximum value also gradually increased, and the maximum growth rate was 58.5%. Secondly, a numerical calculation model was established and verified by experimental data, with the error not more than 4.5%. The simulation results showed that the total pressure recovery coefficient increased as the throat diameter and length of the mixing duct increased. Therefore, larger mixing ducts throat diameter and length had an improved effect on flow loss.

The three-dimensional flow field of the thrust vectoring nozzle was simulated by detached eddy simulation method, and the pressure coefficient and density gradient distribution of flow field were analyzed. The dynamic modal decomposition (DMD) technology was applied to modal decomposition of the pressure coefficient of the z=0 section, the obtained modal recombination flow field was selected to evolve in time. The results indicated that the first five-order modes obtained by dynamic modal decomposition can be used to reconstruct the pressure coefficient field of the dual-throat control vector nozzle more completely. The first-order mode mainly illustrated the swing phenomenon of the separated shock wave and its influence on the pressure pulsation of the shear layer between the recirculation zone and the main flow; the second-order mode mainly illustrated the separation of the vortex system in the shear layer. The third-order mode mainly illustrated the change in the intensity of the separated shock wave, the fourth and fifth-order modes were mainly manifested as high-order oscillations in the position and intensity of the separated shock wave and the shear layer near the recirculation zone.

In this article, full cavity squealer tip and several pressure side and suction side cutback blade tips were designed, their detailed flow field and aerodynamic performance were compared with full cavity squealer tip numerically, including vortices inside cavity and flows around blade tip trailing edge, differences in tip leakage flow rate, downstream total pressure loss and entropy rise. The result indicated that pressure side cutback designs caused the cavity vortex leak earlier compared with full cavity tip, a small vortex formed at cut region leading to increase in total pressure loss of 7.1% at most; flow inside cavity left at cut region for suction side cutback designs instead of crossing over cavity wall, downstream total pressure loss was relatively smaller, with the reduction of 4.6% at most. Compared with full cavity and pressure side cutback squealer tip, suction side cutback designs have better aerodynamic performance.

The flow field of an anti-clockwise main rotor helicopter with two engines in steady forward flight, side flight and hovering was simulated by numerical simulation method, and the total inlet pressure loss and inlet temperature rise under different flight states were obtained. The results showed that the average total pressure loss of the engine intake surface increased with the increase of steady forward flight speed, and the maximum was about 1.61%. During side flight, there was a large total pressure loss in the outer area of the downstream engine inlet and a large air temperature rise in the downstream engine inlet. At the same speed, the inlet temperature rise of the downstream engine was higher when flying on the right wind, which had greater influence on the engine performance. The maximum average total pressure loss of the engine inlet surface was about 1.14% under the hovering state. Compared with the flight data of this helicopter, the effectiveness of the numerical simulation results was verified. The result can provide a reference for domestic flight test subject planning and engine installation loss evaluation of the same type of helicopter.

Hybrid power system for UAV (unmanned aerial vehicle) based on a two-stroke engine with a maximum power of 14.9 kW and a power-to-mass ratio of 2.8 was investigated. An in-house quasi-static UAV model including the series hybrid powertrain system and different energy management strategies was developed to characterize the performance of a multi-rotor UAV with 80 kg maximum takeoff weight. The UAV performance was compared using two different energy management strategies and dynamic-programing optimal solutions, then the effects of payloads and battery energies on the fuel consumption and flight duration of an UAV with defined flight profiles were investigated. It was found that an ideal energy management strategy should avoid high-power battery charging and discharging resulting in less power loss of the battery system, and that Equivalent Consumption Minimum Strategy exhibited lower fuel consumption than Rule-Based Strategy. The fuel consumption increased by using more battery energy in a long flight mission due to the low energy density of the battery, while larger payload and shorter flight duration led to higher system efficiency. Despite of anticipated increment in battery power density in the future, hybrid UAVs still exhibit longer flight duration capability than the electrical ones.

Discussions were performed to determine whether the accuracy of finished bearing could meet the design requirements, as the assembly and disassembly as inevitable processes not only consume working hours, but also may cause scratches on the rolling element and ring, resulting in the decline or even loss of product accuracy. According to the kinematic and geometric relationship of the angular contact ball bearing at condition in motion, based on the accuracy elements of parts, the analytical models of contact angle, radial runout and axial runout of inner ring of assembled bearing were established, the calculation method of dynamic accuracy was proposed, the variation trends of bearing accuracy in different processing stages were studied, and the relationship between groove curvature radius, groove bottom diameter, ball diameter and related accuracy elements and dynamic accuracy was systematically analyzed. The calculation results showed that the groove bottom diameter had a significant influence on the contact angle. In order to meet the design requirements, selection in groups was required after machining; When the error amplitude increased, the radial and axial runout increased almost linearly; The number of balls had little effect on the rotation accuracy, but it affected the stability of operation. The dynamic accuracy test data of the finished bearing were consistent with the calculation results, which indicated that the model was accurate and reasonable. Therefore, whether the dynamic accuracy meets the operation requirements can be obtained through the test of the parts’ design parameters. There is no need for the sleeve assembly and disassembly, which is conducive to improving the qualification rate and production efficiency.

In order to deal with the shortcomings of the second-order contact analysis method of spiral bevel gears and the complex problems of its high-order contact theory realization, based on the combination of ease-off topological surface equation and tooth surface equation of spiral bevel gears and the analytical relationship between transmission error and contact trace and ease-off, a high-order contact analysis method for discrete tooth surfaces with high-order contact parameters of high-order derivative of transmission ratio and short-range curvature of contact trace was proposed, and a simple calculation method based on finite difference was established. The results showed that the fluctuation values of the high-order derivative of the transmission ratio of the high-order tooth surface were 0.0031, 0.0019 and 0.001, respectively, which reflected the global characteristics of the tooth surface morphology. The short-range curvature fluctuation values of the contact line were 0.0000769, 0.000586 and 0.000127, respectively, indicating the complexity of the tooth surface contact process along the contact trace. The results not only verified the correctness and effectiveness of the discrete tooth surface high-order contact analysis method, but also showed that the method reduced the calculation difficulty of high-order contact parameters, providing the possibility for the global design of tooth surface.

A fault diagnosis model based on feature extraction was proposed for aeroengine rolling bearing fault diagnosis. The original vibration signals were preprocessed by envelope demodulation, and only 512 points of each segment of data were taken as fault features, and used as input of bidirectional long short-term memory (BiLSTM) model to diagnose the inner ring faults, outer ring faults, rolling body faults and three different fault degrees corresponding to each fault. The model made up for not only the disadvantages of long input data and obscure features caused by the original vibration signal input, but also the uncertainty of fault diagnosis by extracting vibration features manually. Experiments on the open data set of rolling bearings showed that the accuracy of fault identification was above 99.8%. An aero-bearing tester was built to verify the method and model. BiLSTM based on feature extraction can diagnose the bearing faults more accurately. The proposed method has important engineering value for the fault diagnosis of aeroengine rolling bearings.

In order to explore the influence of slot casing treatment on the first stall stage of counter-rotating compressor, a two-stage counter-rotating compressor was taken as the research object. The influence of slot casing treatment on the first stall stage of the counter-rotating compressor was studied by numerical simulation method. The results showed that the slot casing treatment did not change the initial stall disturbance type of the counter-rotating compressor, which was still a spike stall. The first stall stage of the compressor was converted from rotor R2 to rotor R1 by the forward movement of the casing treatment, while the backward movement of the casing treatment did not change the stall stage of the compressor. After casing treatment, the leading edge spillage of the rotor R2 was inhibited, and the unsteady fluctuation intensity in the blade passage was reduced. However, the leading edge spillage of the rotor R1 was intensified under the near-stall condition. The interface between the mainstream and the leakage flow was pushed out of the blade passage. At the same time, the unsteady oscillation intensity in the blade passage was increased, which finally made the rotor R1 first enter the stall state. It was difficult to suppress the occurrence of leading edge spillage when the casing treatment was processed to move downstream of rotor R2. Although the leading edge spillage phenomenon also occurred at rotor R1 at this time, the leading edge spillage of rotor R2 was more intensified, and the position of the interface of the mainstream and leakage flow was further away from the leading edge of the blade, making it easier for the compressor to stall.

The three-dimensional numerical analysis models of teeth-on-stator labyrinth seal (TOS LS), teeth-on-rotor labyrinth seal (TOR LS) and interlocking labyrinth seal (ILS) were established. The effects of four tooth profile angles (θ=0°, 15°, 30°, 45°) on the leakage and dynamic characteristics of three labyrinth seals were investigated by applying multi-frequency elliptical whirling orbit model and computational fluid dynamics method. Under the operating conditions of 15000 r/min rotational speed and 6.9×10⁵ Pa inlet pressure, although ILS leaked the least, it was easy to cause rotor instability if the tooth profile angle was too small (θ=0°, 15°); TOR LS had the worst sealing performance, while TOS LS had the best system stability. As the tooth profile angle increased from 0° to 45°: the leakage flow rate of TOS LS, TOR LS and ILS decreased by 5.6%, 5.1% and 16.8%; the absolute value of negative flow-induced tangential force for the middle cavity increased by 60.2%, 133.9% and 470.3%; the effective damping of the whole seal section increased by 44.9%−61.9%, 30.7%−53.6% and 90.4%−445.3%, respectively, indicating the system stability was significantly enhanced.

During rocket launching, the secondary combustions between the fuel-rich exhaust gas and the oxygen of air were made, leading to a temperature rise. Based on three-dimensional compressible Navier-Stokes equation, hybrid RANS/LES turbulence model, DOM model, and finite-rate chemical kinetics, the reaction model of multi-nozzle rocket was established. And the validity of model was verified by comparing with the wind tunnel experimental data. Then, a comparison study between reaction and frozen flows of two-/four-nozzle rockets was developed. The results showed that the secondary combustion mainly occurred in the mixed layer. With the increase of the flight altitudes, the increase of the peak temperature caused by afterburning decreased, while the maximum was 10.16% and the minimum was 0.86%. At the same height, the afterburning effect was strengthened with increasing distance from nozzle exit. Comparing with two-nozzle rocket, the afterburning had less effect on the four-nozzle rocket base thermal environment. In addition, the peak heat flux of the rocket base increased first and then decreased with height.