A certain type of helicopter experienced high-frequency abnormal vibrations and tail rotor control failure. Through the analysis of flight parameter and vibration data, the fault was located as the fracture of the horizontal short shaft in the tail drive shaft assembly of the transmission system. A fault tree was constructed with the fracture of the horizontal short shaft of the tail drive as the top event, and fault tree analysis and bottom event verification were carried out to determine that the cause of the fault was the divergence of self-excited vibration in the horizontal short shaft of the tail drive. An analysis of the self-excited vibration mechanism of the tail horizontal short shaft was performed. Contact models for mating internal and external spline teeth pairs, as well as a lumped parameter model for the shafting-spline system, were established. A stability analysis of the horizontal short shaft of the tail drive under the radial friction force of the spline was conducted, clarifying the conditions for system instability and measures to avoid it. The dynamic characteristics test of the tail horizontal short shaft was carried out. The results shows that self-excited vibration occurs under conditions of inadequate lubrication at the spline pair position, wear on the spline tooth surfaces and locating section of the flange, etc. Analysis of the failure mechanism and test results indicate that the friction in the spline locating section is the fundamental cause of the self-excited oscillation in the horizontal short shaft of the tail drive. The divergence of this self-excited vibration ultimately leads to the fracture of the shaft.

Based on the response surface method, the surrogate models with comprehensive cooling effectiveness and maximum equivalent thermal stress as responses were constructed respectively. The effects of blow ratio and structural parameters (curvature radius, incidence angle, aspect ratio, splay angle) on the cooling and strength of curved fan-shaped film holes were analyzed. The optimal design is carried out to maximize the comprehensive cooling effectiveness and minimize the maximum equivalent thermal stress. Results indicate that the blowing ratio is the primary factor affecting the average comprehensive cooling effectiveness of curved film holes. When the blowing ratio increases from 0.5 to 1.5, the comprehensive cooling effectiveness increases by more than 43%. And the curvature radius is the main factor influencing the maximum equivalent thermal stress. Specifically, these factors can affect up to 43.15% (concave model) and 48.35% (convex model). Through multi-objective optimization, compared with the reference model, the comprehensive cooling effectiveness of the concave and convex models with the curvature radius of 40 increased by 10.11% and 17.19%, and the maximum equivalent thermal stress decreased by 26.78% and 9.62%, respectively.

In order to achieve effective drag reduction, this paper designed a bionic structure with a circular arc groove based on the dragonfly wing tubular vein structure. By comparing the drag reduction related parameters and turbulence characteristics in the flow field of the incompressible-plate prototype and the modified scheme, the influence of the design parameters of the bionic structure on the resistance characteristics of the flow field and the starting position of the boundary layer transition was determined. The results show that the arrangement of the bionic structure, the geometric size parameters and the number of groove rows affect the flow field regularly. At the same time, it is also found that the groove spacing has a similar effect in the range of size less than or equal to the groove width. The arrangement of the circular arc groove structure with a 0.75 mm deep size scheme at the first 20% flow direction length region of the transition position introduces external disturbances, which makes the original laminar boundary layer velocity profile fuller and the anti-interference ability further enhanced, thus delaying the transition. In this way, the laminar flow coverage area is increased to achieve effective drag reduction. Compared with the prototype, the ideal scheme can increase 10.75% of the flow direction length range covered by laminar flow, and reduce about 6.38% of the friction resistance.

In order to solve the problem that the ingestion of mainstream gases into the wheel-space between the turbine rotor and stator under the influences of the externally-induced and rotationally-induced, which causes overheating of the turbine disk. Three structures and two extended structures with rectangular slots at different positions of the radial rim seal were presented. Reynolds-Averaged Navier-Stokes (RANS) equations were utilized to investigate the influence of setting rectangular slots at different locations on the mainstream ingestion and sealing performance of five kinds of wheel shoulder seal structures. The results show that the numerical calculations are in agreement with the experimental data, the validation of the numerical approach for the sealing effectiveness is reliability. The sealing effectiveness of three improved structures is better than the radial rim seal with different flows, and the best sealing performance is achieved by setting the rectangular slots on the side shoulder of the rotating disc, The sealing effectiveness of this structure is higher than original structure with 21% at the low sealing flow rate, and the invasion area of intrusion section is reduced by 33.8% at the high sealing flow rate. In addition, the two structures expanded based on this structure reduce the heat transfer area at the convex shoulder, but also reduce the sealing effectiveness, and the expansion is not reasonable.

This paper addresses the challenge of low prediction accuracy in cross-model aircraft engine baseline transfer prediction with existing domain adaptation methods by proposing a novel prediction model based on improved adversarial discriminative domain adaptation (ADDA). The proposed approach incorporates Transformer architecture and self-attention mechanisms to extract long-term features from aircraft engine performance data, enhancing the model’s capacity to capture dynamic features. Additionally, in the domain adversarial module, maximum mean discrepancy and information noise contrast estimation optimization structures are introduced to better leverage the information from limited input data and mitigate interference from redundant information. Consequently, the model achieves enhanced accuracy in cross-model aircraft engine baseline prediction. Experimental results demonstrate that the improved ADDA model reduces the mean absolute error and root mean square error of baseline prediction for engine gas temperature by 19.3% and 16.2%, the respectively, while increasing the coefficient of determination by 4.4%. For the baseline of fuel flow, the mean absolute error and root mean square error were reduced by 26.8% and 30.1%, while the coefficient of determination increased by 6.4%. For the baseline of high-pressure rotor speed, the mean absolute error and root mean square error were reduced by 19.6% and 20.0%, while the coefficient of determination increased by 6.5%. This improvement enables more precise prediction of baselines across different aircraft engine models.

A mathematical model was constructed to describe the heat conduction problem of anisotropic pin fins. Through dimensionless analyses, the dimensionless criterion numbers were obtained, which affected the heat transfer process of the fins. Variable separation method, Taylor expansion, and integral averaging were applied to solve the differential equation. Analytical correlations for fin efficiency and heat transfer rate were derived and numerically verified. Based on the proposed analytical solutions, the influence of the direction of the heat conductivity principal axis on anisotropic pin fins was analyzed. The results showed that within the range of radial Bi was 0.05—10, axial Bi was 0.005—10, cross Bi was 0.2—10, and length to diameter ratio of fin was 2—20, the relative error of fin efficiency from the proposed correlation was less than 1.06% compared with the numerical results. Due to the circumferential symmetry of the temperature field, when the heat conductivity principal axis was deflected in the rOφ and φOz planes, it was beneficial to enhance the fins’ heat transfer performance by making the axis, which had a higher principal heat conductivity coefficient along the r and z directions. However, when the axis was deflected in the rOz plane, the optimal deflection angle can be calculated to maximize the heat transfer performance of the fins under given boundary conditions, material properties, and length to diameter ratio. And the maximum heat transfer rate was 2.97 times that of α=0. The results provide a theoretical support for the design of anisotropic pin fins in application.

In view of the highly conformable and adaptive nature of unstructured mesh to complex geometries and the demand in engineering for efficient and versatile computational methods, a new implementation of the N-S (Navier-Stokes)/DSMC (direct simulation Monte Carlo) coupling method using tetrahedral unstructured mesh for complex interfaces was presented with the aim of resolving the fluctuations in the positioning of the coupling interfaces in most N-S/DSMC coupling method for multi-scale transitional flows in the near-continuum regime. This implementation utilized local Knudsen number as a continuum breakdown parameter to partition the continuum/rarefied regions and generate three-dimensional complex N-S/DSMC coupling interfaces. Along each side of the interface, one or multiple layers of interface information transfer meshes were advanced, and information coupling was achieved by the state-based approach. According to this method, there was no need for smoothing or reshaping treatments applied to complex irregular interfaces, with the general applicability to numerical simulations of complex transitional flow regions. Simulations of three-dimensional hypersonic flow around a sphere and a blunt cone showed that, compared with the DSMC method, the shock wave and wall characteristics were in close agreement, with a maximum error of less than 8%. At the same time, computational efficiency was improved by 1.74 and 2.28 times, respectively, validating the method’s accuracy and efficiency.

To reveal the vibration modulation mechanism and envelope spectrum features of an inter-shaft bearing with raceway damage, the modeling, simulations and experimental validations of vibration in a dual-rotor system were conducted. Considering the concurrent action of the rotor unbalance centrifugal loading and the rotor dead weight, the amplitude modulation mechanism of the inter-shaft bearing fault vibration was explored. The results indicated that when the rotor unbalance centrifugal loading was much smaller than the rotor system dead weight, the loading zone of the inter-shaft bearing maintained a traditional fixed pattern. The amplitude modulation of vibration caused by the raceway damage was mainly originated from the change in damage position caused by raceway rotation. When the rotor unbalance centrifugal loading was equivalent to or greater than the rotor system dead weight, the loading zone of the inter-shaft bearing was in a moving pattern. In addition to the modulation due to shaft rotation, a more complex modulation relationship could be caused by the ‘chasing’ motion pattern between the moving loading zone and the raceway damage location. The modulation frequencies encompassed the rotation speeds of the high-pressure rotor, low-pressure rotor and their difference. Moreover, under the action of the rotor unbalance centrifugal loading, the complex modulation relationships altered the energy distribution of the envelope spectrum.

In response to the complex coupling relationship of multiple parameters in the combustion chamber, the influence of the coupling relationship between different head expansion angles of flame tubes and swirle on combustion performance was studied. The research results showed that: to 90° expansion flame tube, with the increase of swirl number, the fuel mixing uniformity gradually deteriorated, the outlet temperature distribution coefficient gradually increased within a wider stable working range; the oil-air ratio of lean blow out gradually decreased, the carbon monoxide (CO) and Unburned hydrocarbon (UHC) emission index decreased, while the nitrogen oxides (NOx) emission index increased. To 45° expanding flame tube, with the increase of swirl number, the fuel mixing uniformity gradually improved, the outlet temperature distribution coefficient gradually decreased, and the NOx emission index decreased.

In view of the solid rocket motor nozzle, the volume of fluid (VOF) method was used to simulate the droplet-gas two-phase flow, and the fuzzy theory and proportional differential control (PD) method were combined to control the droplet to reach a quasi-static state, so as to study the force characteristics of deformable alumina droplet in the flow field. The results showed that the fuzzy theory coupled with PD controller can make the droplet reach the quasi-steady state more efficiently and stably. The drag coefficient of the droplet increased with the increase of deformation degree and the relative Mach number between the flow field and the droplet, and there was no significant coupling between the relative Mach number and Weber number. The effect of droplet deformation and compressibility of flow field on gas-liquid phase interaction should be considered in accurate calculation of two-phase flow field for solid rocket motor. Compared with the drag coefficient model of the rigid sphere, droplet deformation could cause more thrust loss in the gas phase of solid rocket motor nozzle, and also result in smaller particle-free zone in the nozzle.

The influences of different rotating speeds of the rotatable ring of the controllable speed casing on the stability of the low-speed axial flow compressor were studied by numerical simulation. The results showed that the controllable speed casing, which covers the whole axial chord length region of the rotor tip and rotates in the same direction with the rotor, can control the tip leakage flow and realize the stability expansion. By applying external shear stress to the clearance flow, the controllable speed casing increased the mainstream momentum, suppressed the momentum ratio of the leakage flow to the mainstream, improved the deflection of the leakage vortex, and delayed the occurrence of secondary leakage and the forward movement of the mainstream/leakage flow interface, thus broadening the stable operating range of the low-speed compressor stage. While ensuring that the pressure ratio was basically kept unchanged, the stability expansion effect was enhanced with the increase of the rotatable ring speed. When the rotatable ring speed was the design speed of the rotor, the maximum stable operating margin of the low-speed compressor stage can be increased by 30.86%.

In order to predict the dry running time of spiral bevel gear, a coupling prediction method of tooth surface elastohydrodynamic lubrication tooth meshing gear system heat transfer was proposed to reveal the evolution mechanism between micro elastohydrodynamic lubrication and macro flow field / temperature field. The steady-state thermal elastohydrodynamic lubrication model of spiral bevel gear was established by using the elastohydrodynamic lubrication theory, and the friction coefficient of gear tooth surface was obtained; By introducing the oil film retention parameter, the calculation method of time-varying friction coefficient of tooth surface was formed; using computational fluid dynamics(CFD) method, the steady-state temperature distribution of gear meshing and the transient temperature distribution of gear system were simulated and calculated; The dry running time of spiral bevel gear was predicted by parameter transmission and coupling between three dimensions. The results show that the time scale of the model is from O(10⁻⁶) s of the elastohydrodynamic lubrication dimension of the tooth surface to O(10⁻¹) s of the heat transfer dimension of the gear system, and the spatial scale is from O(10⁻⁶) m of the elastohydrodynamic lubrication dimension of the tooth surface to O(10⁻¹) m of the heat transfer dimension of the gear system; The coupled structure-force-lubrication-thermal characteristics set up a prediction method for the dry running time of spiral bevel gear, predicting the failure of gear in about 1.5 hours.

By using backlight high-speed shadowgraphy and liquid film tracking and recognition program, a full time series experimental study was conducted on the instability and breakup process of the liquid film at the exit of a double swirl air-blast atomizer. Based on the phenomenological description, the process of liquid film breakup was analyzed. The characteristics of liquid film deflection, circumferential rotation, radial flapping and spinning under swirling flow were summarized. Based on the time evolution characteristics of liquid structure during the breakup process, three typical breakup modes of liquid film under double swirl airflow were acquired as follow: sheet-bag breakup, ligament breakup, and hybrid breakup modes. By analyzing experimental statistical data and high-speed images, it was found that the average length of the ligaments was significantly affected by operating conditions. The impact of airflow pressure drop was greater than that of liquid flow rate. In addition, the empirical formulas between it and liquid Weber number and liquid Reynolds number were fitted through experimental data. A length deviation of film accumulation deviation was used to measure the probability of the occurrence of breakup modes. It was found that deviation was significant affected by airflow pressure drop and showed a decreasing tendency as airflow pressure drop increased. This indicated that airflow pressure drop dominated the change of liquid film breakup, while liquid flow rate had a less impact on deviation and did not affect the breakup modes.

In view of the unclear lubrication characteristics of high-speed gear transmission under low temperature conditions and the limited improvement of lubrication effect by traditional baffles, a study on the splash lubrication characteristics of high-speed gear transmission was carried out. Firstly, based on the theory of computational fluid dynamics, a two-phase flow analysis model of splash lubrication of high-speed gear transmission was established, and the two-phase flow distribution characteristics were analyzed; then, on this basis, the influences of number of rotations, gear speed, and oil immersion depth on splash lubrication characteristics and torque loss were studied; finally, a bionic honeycomb baffle structure was proposed and its structure was optimized using a multi-island genetic algorithm. The results showed that: in the splash lubrication process of high-speed gear transmission under low temperature conditions, the volume fraction of lubricating oil on the tooth surface decreased with the increase of gear speed, but increased with the increase of oil immersion depth; the torque loss increased with the increase of gear speed and oil immersion depth, and the influence of speed on torque loss was much greater than that of oil immersion depth; under the same working conditions, the average volume fraction of lubricating oil on the tooth surface of the optimized bionic honeycomb baffle structure increased by 68.46% compared with the one without baffle, and increased by 7.88% compared with the original baffle. Research results provide a basis for the study of two-phase flow distribution characteristics of high-speed gear transmission under low-temperature conditions and the optimization design of splash lubrication for high-speed gear transmission.

Aimed at the total pressure recovery coefficient of the combustion chamber at the level of 0.915, the required that the total pressure recovery coefficient of the combustion chamber was increased by 0.01. In order to improve the total pressure recovery coefficient of the combustion chamber, the pre-diffuser and cowling were improved, and the flow mass distribution and flow field of the flame tube were matched design. The component experimental study on the performance of the full-ring combustor show that the total pressure recovery coefficient of the improved scheme is 0.929, which is increased by about 0.014 on the basis of the prototype combustor, meeting the design requirement, and the diffuser designed in the improved scheme has at least contributed not less than 0.005 to increase the total pressure recovery coefficient. The quality of the outlet temperature field can meet the design requirement, the wall temperature of the flame tube can meet the requirement of material use, lean-burn boundary of slow state engine and combustion efficiency are slightly lower than the prototype combustion chamber. The ground experimental of the overall engines shows that the improved scheme under the intermediate thrust index can increase the air flow rate of the compressor by 0.48kg/s, reduce fuel consumption by 2%, and reduce the temperature before the turbine by 6℃ on the basis of the prototype combustion chamber, which can improve engine performance and reliability.

To conduct the vibration environment test of external accessories on aeroengine effectively, a development method of vibration environment test profile was studied. By processing the measured data of the vibration environment of the engine external accessories, it was confirmed that the vibration environment was a mixture of sinusoidal and random excitations, and sinusoidal excitations had order characteristics. A multi-components swept sine-on-random (SoR) vibration profile was developed to accurately capture the real vibration environment. A development process of swept SoR vibration profile using vibration field measurement data was established. Based on the fatigue damage equivalent theory, the multi swept sine components and random components were matched and synthesised during the development process. Four example profiles were developed for different areas on the casing. The result showed that the SoR profile can fully characterize the mixed excitation characteristics of real environmental loads compared with the single component profile. Compared with the profile specified in current standards/specifications, the frequency domain characteristics of the new developed profile were consistent with the real environmental loads, and the vibration level was strictly equal to the statistical tolerance of the real loads.

An innovative method for identifying multiple debris in lubricating oil based on back propagation neural networks was introduced to address the challenge of multi-debris signal overlap. A two-level model framework was proposed, of which the first level network can accurately estimate the number of small debris in overlapping signals, and the second-level network utilized this quantity information to precisely analyze the diameter of these small debris, successfully overcoming the challenges posed by signal overlap. Through sufficient data training and model structure optimization, the model achieved 98.10%, 91.42%, and 92.06% accuracy, respectively, in single, double, and triple debris recognition.

In order to realize the damage analysis of fiber reinforced composites, an efficient multiscale damage analysis method is developed. Firstly, based on the generalized method of cells, a multiscale damage analysis framework is constructed for laminate and plain weave composites, and the damage processes at the mesoscale and microscale under uniaxial tension are analyzed, The results show that the complex weave structure of plain weave composites leads to a more complex mesoscale and microscale damage evolution process, which is significantly different from the damage process of laminates. Based on this, neural networks are introduced, and a data-driven multiscale damage analysis strategy is proposed to realize the efficient damage simulation of plain weave composites. Compared with the experimental and simulant results, the error of predicting tensile strength by the efficient multiscale damage analysis method is less than 7%; and compared with traditional multiscale damage analysis method, the efficiency of macroscale calculations can be improved by about 12.47 times.

A rapid vibration response solving method was developed for the fan blade lost rotor system with a fusing structure. The effects of time-varying parameters, such as rotor speed and stiffness, on the vibration response of the rotor system were studied. The intrinsic vibration damping mechanism of the fusing structure was elucidated, and the design principles for fusing structure were proposed based on these findings. The results indicated that the fusing at the #1 support can significantly reduce the first critical speed of the rotor, consequently reducing the resonance response peak when the rotor passed through the critical point. The impact of the speed reduction rate on the dynamic characteristics of rotor with “complete fusing” and “partial fusing” designs differed significantly, necessitating a distinction in design approaches. The stiffness variation duration of the support had a minor effect on the vibration response of the rotor and may be disregarded during the design phase. When the support stiffness reduction rate fell within a certain range, the rotor may approach a resonant state during windmilling operation stage, thus this “resonance zone” should be avoided during design. For complete fusing designs, restoring the stiffness of the support structure during windmilling phase is beneficial for enhancing rotor operational safety, with preferable faster stiffness recovery.

A flow control method in compressor cascade based on blade-attached wedge vortex generator (VG) was proposed, and the numerical simulation of 3D steady Reynolds-averaged Navier-Stokes （RANS） method was carried out. The numerical method was verified, and compared with the experiment results to indicate its correctness. Furthermore, the performance of cascades with four different shape VGs was calculated and the internal flow was analysed. The results showed that the VG with inverted-symmetric shape can effectively improve the cascade performance, and widen the cascade working range. The four-shape schemes can suppress shock wave/boundary layer interaction separation, but the control effects were significantly different. In curved compressor cascade passage, the VG vortex of inverted schemes was close to suction surface, and the control effect was better than that of the forward schemes away from suction surface. The weaker VG vortex pair generated by symmetric scheme can suppress shock wave/boundary layer interaction separation along the axial direction, and the stronger VG vortex from unsymmetric scheme can suppress the separation along the axial and span directions, but with greater static pressure loss. Therefore, considering the separation suppression by VG and its own negative effects, the inverted-symmetric scheme is more suitable for the flow control inside the compressor cascade.

In order to improve the cavitation performance of aviation fuel pump, a centrifugal aviation fuel pump was used as an example to optimize the impeller runner geometry parameters with full parameters, and generate a response sample space with design parameters as inputs and boost pressure value and cavitation volume as outputs. Based on the Meta-Model of optimal Prognosis, the optimal surrogate model was fitted and the global sensitivity analysis was performed. It was found that the impeller inlet edge parameters and the middle profile parameters had a significant influence on the boost pressure value and cavitation volume of the fuel pump. The optimal design parameters were obtained by the genetic algorithm, and the numerical simulation method was used to verify the obtained optimal parameters. The results showed that the pressure coefficient of the optimized fuel pump was improved by 0.035 and the cavitation performance was improved by 22.28%. The impeller cavitation area and cavitation volume decreased significantly, the pressure load distribution between different vanes was improved, and the amplitude of pressure pulsation inside the impeller was reduced.

To achieve smooth mode transition, of turbine-based combined-cycle (TBCC) propulsion system the transition operation modes and transition schemes of TBCC inlet were investigated through wind tunnel test and numerical simulations. Firstly, three transition operation modes of TBCC inlet, including supersonic operation mode, supersonic/subsonic operation mode and subsonic mode, and the criteria for different operation modes were put forward. Then the operation zone, which included the three transition operation modes of inlet mode transition, was established according to the maximum backpressure of turbojet and ramjet flowpath at different modes. Finally, two mode transition schemes were presented based on the operation zone. The results indicated that during TBCC inlet mode transition, the maximum backpressure of the turbojet flowpath of supersonic operation mode decreased from 32 times to 13 times, while it increased from 16 times to 32 times for the ramjet flowpath. The mass flow into two flowpath was changed linearly to the entrance area of turbojet/ramjet flowpaths during supersonic mode transition. Once the inlet operated in supersonic/subsonic mode during mode transition, aerodynamic coupling between two flowpath occurred. During mode transition from 40% to 60%, mass flow ratio of the turbojet flowpath decreased by 12% compared with supersonic operation mode and increased by 7% for the ramjet flowpath.

The prediction of combustion efficiency at different downstream locations of the flame stabilizer is an important problem in the length design of afterburner. Taking the upstream-injection U-shaped bluff-body flame stabilizer as the research object, a prediction model of the combustion efficiency along the downstream of flame holder in the afterburner based on reaction rate controlling was proposed and verified by the combination of numerical simulation and theoretical analysis, under the conditions of incoming flow temperature of 600—900 K, incoming flow velocity of 75—170 m/s and equivalent ratio of 0.22—1.20. At the same time, the prediction model of turbulent flame velocity and the semi-empirical prediction formula of turbulent intensity for afterburner were determined. The results show that compared with the numerical simulation results, the prediction errors of the model for the combustion efficiency at different downstream locations of the flame stabilizer are less than 2.5% at different temperatures and velocities, and less than 20% at different equivalent ratios.

An RP-3 aviation kerosene surrogate fuel capable of both physical and chemical surrogate was proposed in this study. The surrogate fuel was composed of n-dodecane, 2, 5-dimethylhexane, 1, 3, 5-trimethylbenzene and decalin, their molar fractions were 0.54, 0.22, 0.14 and 0.1, respectively. To overcome the limitations of existing skeletal mechanism construction methods, an approach was developed, leading to the successful establishment of a high-precision skeletal mechanism containing 153 species and 858 reactions. Through systematic validation, it was demonstrated that the proposed surrogate fuel accurately predicted the physicochemical behavior of RP-3 fuel, including physical properties (density, viscosity, and spray penetration distance) and fundamental combustion characteristics (ignition delay time, species concentration evolution, laminar flame propagation, and NO emissions). Additionally, numerical simulations confirmed that the spray combustion and ignition processes of RP-3 fuel in a constant-volume combustion chamber at ambient temperatures of 853, 898 K, and 923 K can be effectively replicated, fully verifying its physical and chemical surrogate capabilities. This study could provide valuable insights into the development of high-carbon surrogate fuel formulations and skeletal mechanisms.

In order to solve the dramatic power coastdown of short-range unmanned aerial vehicles at high altitudes, a turbocharging two-stroke aviation piston engine was studied. Based on comprehensive analysis of the turbocharging technology application in the unmanned aerial vehicle platform, a technical scheme including an intake pressure stabilizing system, an exhaust resonant system, a self-lubricating turbocharging system and a closed-loop adaptive control system, was proposed. A one-dimensional simulation model was established and calibrated for a turbocharging two-stroke engine. The DOE （design of experiment） test method was used to complete the coupling relationship analysis of the key parameters of the intake and exhaust system, and the best structural parameters of the intake pressure stabilizing system and exhaust resonance system were obtained. And the coupling and matching of the turbocharging system with the engine performance was performed. At an altitude of 8000 m the power-to-mass ratio before and after supercharging increased from 0.59 kW/kg to 0.96 kW/kg; and the proposed technical scheme was estimated through a flight test of a unmanned aerial vehicle platform. The results showed that, the maximum flight height exceeded 8367 m, which was 85% higher than that without this scheme. The two-stroke supercharging technology has obvious advantages in high-altitude power.

In order to solve the problem of end-wall corner separation in high-load axial compressors, a combined contouring method of suction surface and end-wall with multiple local control channels for secondary flow applicable to multiple working conditions was applied to the final stage static sub of a 2.5-stage compressor, and the effects of the contouring method in terms of efficiency enhancement or loss control in the face of complex flow conditions of the pressurized gas environment were explored. A single-objective optimization of the highest efficiency point resulted in an improvement of about 0.3% in the final stage peak efficiency of the compressor. The flow field analysis showed that when the integrated end wall-suction surface contouring design was applied in a multi-stage compressor environment, the contouring was located in the static sub, but the rotor flux and even the efficiency were also subjected to a certain degree of interstage interference. Due to different flow structures in the end zone of the prototype compressor, the optimal modeling flow control mechanism differed from the conclusions obtained in the previous vane grid study. In this way, the feasibility of the integrated end-wall-suction surface contouring method adopted to the control of various flow structures can be proved to a certain extent, and its application in the compressor environment has been verified.

To study the impact of inter-stage turbine burner (ITB) technology on the overall performance of turboshaft engine, a simulation model of variable specific heat was established using the component-level modeling method on the Visual C++ (VC) platform. Through simulation and comparative analysis, changes in engine performance were investigated under different operating cycle parameter matching conditions. The results indicated that the total temperature at the exit of the main combustion chamber and the ITB reheating had a significant influence on fuel consumption. When the total temperature at the exit of the main combustion chamber increased by 500 K and the ITB transited from closed to fully open state, the relative increase in fuel consumption decreased by 9.24%. Under throttling conditions, as the relative rotational speed of the engine core decreased, the magnitude of power output improvement from activating ITB decreased. Under altitude conditions, with increase of the flight altitude, the reduction in fuel consumption due to ITB was greater than that in a conventional turboshaft engine. Under temperature conditions, as the ambient atmospheric temperature increased, the increase in fuel consumption due to ITB was higher compared with a conventional turboshaft engine.

An experimental system based on damped free vibrations was developed. By applying loads to the blade model to induce a first-order bending mode, followed by releasing the load to create first-order bending vibrations, a comprehensive damping ratio characteristic curve could be acquired from a single test. This was achieved by monitoring and analyzing the time-domain signals of strains and accelerations at crucial positions. Experimental findings showed that the damping ratio curves obtained through experiments aligned closely with those generated through numerical simulations. Furthermore, the maximum damping ratio was kept constant despite of variations in the damper’s inertial load. Additionally, dampers with different contact areas had similar critical damping ratios. However, the damping effect was found to be positively correlated with the length of the shank.

Numerical simulation methods were employed to quantitatively and qualitatively explore the impact of endwall movement on the aerodynamic performance and tip flow characteristics of tandem diffusion cascades from the perspectives of entropy production rate, blockage factor and kinetic energy component of leakage flow. The main conclusions were as follows: firstly, after the endwall movement, the overall loss of tandem cascade was reduced, the blockage in tip area was intensified, and the flow turning angle was decreased, while the lag angle was increased. In the range of −4° to 4° angle of incidence, the loss was reduced by more than 3.9%, and the blockage was increased by more than 30.4%. Secondly, the endwall movement increased the leakage flow of the front and rear blades by 12.3% and 9.9%, respectively, but decreased the gap jet flow. Furthermore, the endwall movement enlarged the ratio of the kinetic energy of the leakage flows of the front and rear blades, with the secondary flow kinetic energy becoming dominant. It also increased the loads on the front and rear blades and made the position of the maximum pressure difference move forward in advance. As a result, the morphology and development of the leakage vortices were changed and the formation of the jet vortices was suppressed. In addition, the endwall movement significantly weakened the endwall shear effect and the mixing effect of the gap jet, expanded the circumferential influence range of the leakage flow of the front blade and caused secondary leakage. This resulted in a significant reduction in the entropy production of the front blade, while the change in the entropy production of the rear blade was relatively small. Moreover, the blockages of both the front and rear blades were intensified. The impact of the endwall movement and the increase in the angle of incidence on the front blade of the tandem cascades was greater than that on the rear blade, which was mainly the result of the regulating effect of the gap jet.

An experimental system was developed to evaluate the dynamic response of the blade. This system modulated the standard pressure between the triangular prism damper and the platform by regulating the quantity of weights, so as to validate the variation in damping ratios produced by dry friction dampers under structural vibration stress. By sweeping frequencies around the resonant frequency of the simulated blade test specimen, the amplitude-frequency response curve was obtained, from which the critical damping ratio was computed using the half-power bandwidth method. By measuring the vibration stress, the damping ratio characteristic curve was also acquired. The experimental results demonstrated that the two contact surfaces of the triangular prism damper had different damping effects. Moreover, the computed damping ratio characteristic curve agreed well with the experimental data. It was found that: the inertial load of the damper had negligible influence on the maximum damping ratio; when the actual contact area was small, the damping performance of the damper was unstable; and the length of the shank of the blade was positively correlated with the damping effect of the damper.

To reveal the transient properties of two-phase flow and boiling heat transfer in cryogenic tube chill-down process, a computational fluid model (CFD) model to account for transient variations in the cryogenic chill-down event was developed. In this CFD model, different heat transfer models were selected to calculate fluid-solid coupling heat transfer rates at different stages of chill-down. The results showed that the flow patterns inside the tube could be divided into four stages, including liquid level rising, film boiling, transition and nucleate boiling, and pipe flooding. Due to the influence of gravity in the development of the interior two-phase flow, the distribution of a general stratified flow with local “upwarped” appeared at the liquid surface position near the tube wall. The thicknesses of the vapor film at the bottom and side wall decreased fluctuatingly from 0.093 mm and 0.124 mm, respectively. Moreover, it was found that the maximum temperature difference at the cross section was about 90 K, and the temperature difference between the tube inlet and outlet reached about 50 K. Owing to the transition from film boiling to transition and nucleate boiling, the local void fraction could increase significantly.

The service conditions of the aero-engine accessory transmission system are complex, the internal and external excitations are coupled with each other, and the excessive vibration often causes local damage to the accessory structure. However, there are few system-level dynamics considerations, and direct test measurement is difficult and has large errors. Considering an aero-engine accessory transmission gear system, the multi-excitation (mesh stiffness excitation, error excitation and friction excitation) dynamic modeling analysis under different working conditions was carried out. By introducing the position angle and displacement projection vector, and using the idea of finite element matrix assembly, a 51-degree-of-freedom dynamic model of a certain type of aero-engine accessory gear transmission system was established; using the potential energy analysis method, the meshing stiffness excitation characterization of alternating single and double teeth was completed. The random function was employed to complete the meshing error excitation characterization, and a friction excitation characterization method was established based on the elastohydrodynamic lubrication model; the variable step size Runge-Kutta method was used to solve the dynamic response of the transmission system, and the dynamic characteristics of the transmission system were analyzed by various internal and external excitations. The results showed that the coupling effect of high speed and low stiffness increased the risk of instantaneous overload, tooth surface detachment and fatigue damage between gear pairs; lubrication conditions, precision grades and gear structures had a greater impact on system dynamics; the gear was the most dangerous, and its vibration margin should be paid attention to. This could provide a basis for the follow-up dynamics research of the central transmission system.

In the numerical simulation of aircraft icing, the Lagrange method has the problems of low efficiency and poor adaptability in calculating complex configurations. Therefore, this paper proposes a fast algorithm for calculating the droplet collection efficiency of complex configurations. Based on the calculation of droplet collection efficiency by the particle statistics method, the Monte Carlo method was coupled to improve the adaptability of the algorithm to complex configurations. The initial droplet release array that can cover all impact components was determined by block tracking and adaptive analysis based on object surface distance. Furthermore, the drip-grid array encryption technology and variable-step acceleration technology based on error control are proposed to improve the computational efficiency, and the relationship between computational efficiency and control variables in the acceleration method is analyzed and obtained. The calculation results show that the developed algorithm can accurately and efficiently obtain the three-dimensional (3D) droplet collection efficiency, which is suitable for complex configurations. It provides a new method and idea for calculating the droplet collection efficiency of aircraft and provides a technical reference for the study of aircraft icing characteristics and the design of anti-icing systems.

In order to improve the accuracy of engine starting performance calculation based on the component method, a starting throttling inverse calculation method based on the power extraction method was proposed. It clearly defined the power balance relationship of the starting process, and established the nonlinear equations of the engine starting, so as to reversely evaluate the combustion efficiency based on the fuel flow, rotor acceleration rate and starter characteristics during the starting process. According to the similarity principle, the calculation method of converted torque scaling factor was proposed, the conversion between efficiency characteristics and torque characteristics was completed, and the effectiveness of the method was verified. Furthermore, a pre-ignition starting model was established, enabling the simulation of starting from zero to idle speed. The simulation model was validated based on the experimental data of a micro-turbojet engine, and the results showed that the maximum errors in the rotational speed and P3 during the starting process were 2.81% and 2.22%. This method allowed for an approximate simulation of the starting process and enabled the reverse calculation of combustion efficiency during that process. In addition, the plausibility of the sub-idle component characteristics can be tested. This method can provide a reference for modeling and validation of starting performance for other types of aerospace engines.

The potential coupling mechanism of transverse combustion instability in an O₂/CH₄ rectangle rocket combustor was investigated experimentally and numerically. High-frequency pressure sensors were employed to capture the dynamic pressure characteristics within the combustion chamber. Numerical simulations were conducted using the stress-blended eddy simulation (SBES) and flamelet-generated manifolds (FGM) methods. The results showed that the numerical simulation successfully predicted the transverse combustion instability observed experimentally. The numerical results, including pressure waveform, main frequency, and root mean square amplitude, were in good agreement with the experimental data and theoretical analysis. A comprehensive analysis of the driving mechanism and Rayleigh index of the transverse combustion instability was presented. Examination of the flow field characteristics in the combustion chamber revealed that the transverse pressure wave induced periodic ‘stripping’ of the oxygen jet and vortex disruption. Heat release pulses were generated during vortex fuel combustion and coupled with the first width (1W) mode. Rayleigh index analysis indicated that edge injectors exhibited stronger driving characteristics for transverse combustion instability compared with central injectors.

To further improve the energy management efficiency of the solid divert and attitude control system of the multi-objective kill vehicle (MOKV), and obtain the minimum overall mass scheme, a propulsion system design method for MOKV based on the guidance strategy was proposed. Firstly, a rapid evaluation method for the multi-pulse guidance strategy based on linear covariance analysis was proposed to determine the maximum velocity increment required for each maneuver. Then, the layout scheme of the three-pulse solid divert and attitude control propulsion system with a symmetrical dual-combustion chamber was determined, which featured a simple structure and high controllability. Subsequently, performance parameter and mass models for the propulsion system were established, and an optimal design process of the propulsion system based on the guidance strategy was developed. Through iterative optimization using phase diagram analysis, the mass-optimized solution satisfying size constraints can be obtained, and the design criteria for the sub-interceptors were further provided. Finally, a case design was carried out for typical application scenarios, and the miss distance was evaluated through Monte Carlo random test. The case results showed that the MOKV can carry 12 sub-interceptors, with a total mass of 49.43 kg, an axial dimension of 540.7 mm, and a radial dimension of 231.9 mm. After three pulse ignitions, the maximum miss distance of the MOKV was reduced from the kilometer level to the hundred-meter level and even the ten-meter level, which can achieve precise interception of the target by the sub-interceptors. The theory and method proposed can provide a strong support for the high-efficiency energy management and lightweight design of MOKV.

In order to meet the lightweight requirements of high-power density spiral bevel gear transmission, a multi-parameter optimization technology for spiral bevel gear blank was proposed. Taking the tooth number, modulus, and tooth width of the gear as design variables, the minimum sum of the volume of the gear pair as the objective function, and the working condition limitations, installation conditions, and multiple strength requirements as constraints, the genetic algorithm was synthesized to realize the optimization design of spiral bevel gear blank parameters. Examples with power of 0.75 MW and 5 MW were used to verify the effectiveness of the proposed technique. The results showed that: compared with traditional design methods such as adaptive method and nonlinear mathematical programming method, the proposed technique could achieve global optimization with a success rate of 100% thanks to the adaptability and parallel ability of genetic algorithm; in addition, compared with the design requirements, the redundancy of pitch diameter and tooth width in the 0.75 MW case was 31% and 19.23%, respectively, while that in the 5 MW case was slightly more than 4.69% and 2.65%, showing that the proposed technique could provide effective design margin for designers under different design conditions.

To address the problem of poor stability in the cage of needle roller bearings under high-speed working conditions, the HK0608 needle roller bearing was taken as the research object, and a bearing dynamics analysis model was established based on multi-body dynamics analysis software. Meanwhile, the deviation ratio of the cage center of mass vortex radius and the cage slip rate were selected as optimization objectives, and the structure of the bearing was optimized by using orthogonal experimental method, multiple regression, principal component analysis method and non-dominated sorting genetic algorithm Ⅱ (NSGA-Ⅱ) multi-objective optimization genetic algorithm. The results showed that the rotational speed had the greatest impact on the deviation ratio of the center of mass vortex radius, and the radial load had the greatest impact on the slip rate of the cage; when the radial load was 1084 N, the rotational speed was 21036 r/min, and the wall inclination angle was 5.8°, the stability of the retainer was the best. At the same time, the stability of the cage under different rotational speeds and radial loads before and after optimization was compared and analyzed. It was found that the stability of the two cages first increased and then decreased with the increase of rotational speed and radial load, and the stability of the cage of the optimized bearing was improved. The outcomes of the research can serve as a guide for designing the structure of needle roller bearings under high-speed working conditions.

Taking a two-stage lubricating oil pump for a specific aero-engine as the object, a testing system was developed to evaluate the flow performance under pure oil medium, meeting the test requirements of the oil pump. Performance tests under pure oil medium were conducted to explore the high-altitude performance of the pump and analyze the impact of various operating parameters on the flow performance of the lubricating oil pump. The results indicated that the volumetric efficiency of the two-stage pump fluctuated by less than 5% with changes in outlet pressure, with the rotational speed ranging from 1003 to 4014 r/min, and outlet pressure below 600 kPa in the ground tests. The oil flow rate demonstrated a linear relationship with rotational speed, increasing in conjunction with the rotational speed, while volumetric efficiency declined. During high-altitude tests, the oil flow rate and volumetric efficiency of the second-stage pump exhibited a slight increase compared with those of the first-stage pump at varying inlet pressures. The oil flow rate and volumetric efficiency of the two-stage pump significantly declined with the reduction in inlet pressure. The oil flow rate and volumetric efficiency of the two-stage pump exhibited a significant decrease at flight altitudes below 8 km. With the increase of rotational speed, the decrease in oil flow rate and volumetric efficiency accelerated. The volumetric efficiency remained below 50% at all rotational speeds until reaching a flight altitude of 12 km.

An amplitude-identified multiple signal classification (MUSIC) algorithm for blade tip timing (BTT) signals of rotor blades was proposed for identifying surge characteristics of compressors. By analyzing the real-valued sinusoidal signal model of BTT, the relationship between the amplitude and frequency of asynchronous vibrations and the eigenvalues and eigenvectors of the autocorrelation matrix was explored, allowing for compensation of the amplitude of the pseudo-spectrum within the framework of the MUSIC algorithm. The experiment of rotor surge identification using BTT technique was conducted. By comparing with fast Fourier transform and the least squares fitting method, the effectiveness of proposed method was verified. The results indicated that surge faults exhibited a significant increase in amplitude and speed fluctuation in the time domain. In the frequency domain, surge faults exhibited low-frequency asynchronous vibration with large amplitude. Compared with Fourier transform and least squares fitting, the proposed method can accurately identify the surge characteristic frequency of 5.3 Hz, and the focusing capability of its amplitude identification was more than three times that of the Fourier transform. The proposed method with higher frequency resolution and amplitude recognition accuracy can effectively extract the characteristics of surge faults in the frequency domain.

To investigate the effect of real coupling deviation on the aerodynamic performance of a blade cascade, chord length deviation, thickness deviation, and leading edge radius deviation were applied to a high load compressor blade cascade for research based on the arbitrary polynomial chaos method of moments. The research results showed that compared with the prototype, the probability of an increase in total pressure loss coefficient and a decrease in static pressure coefficient under negative angle of attack conditions was about 79.47% and 92.83%, respectively; under positive angle of attack conditions, the probability of an increase in total pressure loss coefficient was about 91.11%, and the probability of a decrease in static pressure coefficient was about 86.42%. The aerodynamic performance under different operating conditions was most sensitive to the deviation of the leading edge radius. Combined with the analysis of the loss source, it was found that the leading edge loss played a dominant role. Therefore, strict control of the leading edge machining accuracy was required during machining. Compared with the prototype blade, under negative attack angle conditions, the degree of separation in the blade angle region with added coupling error changed significantly, and the separation point, recirculation area, and axial length of the recirculation area increased significantly; under positive angle of attack conditions, the degree of separation in the blade angle region with added coupling error was not significant.

Electric propulsion, with its high specific impulse, efficiency and long lifespan, has emerged as one of the main space propulsion methods following cold gas and chemical propulsion. Water-fueled electric propulsion, thanks to its non-toxic, pollution-free and low-cost benefits, has increasingly gained attention. A comprehensive summary and evaluation of the current development status of water-fueled electric propulsion technology were conducted both domestically and internationally. Based on the method of thrust generation, it can be classified into three types: electrothermal, electromagnetic, and electrostatic. Result showed that, electrothermal propulsion achieved significant progress in ground experiments and in-orbit demonstrations, making it the most mature water-based propulsion technology. Water-based electromagnetic and electrostatic propulsion technologies are still in the developmental stage but are progressing steadily. Future advancements in water-fueled propulsion technology are expected to focus on the integration of chemical and electric propulsion systems, multimodal propulsion designs, and their application in deep space exploration. With technologies becoming more mature, water-fueled propulsion technology will develop in directions that integrate chemical and electric propulsions.

The fatigue fracture occurred during the low speed performance test of swept-forward fan blades. The cause was investigated to improve the design. The damage caused by blade resonance and flutter was excluded through simulation analysis. The high-cycle fatigue analysis showed that the second-order modal danger point of the fan blades was consistent with the crack initiation position. The analysis suggested that the unsteady flow at the tip generated 6.5 times the rotational frequency exciting force, which stimulated the second-order natural mode of the fan blades. Blade tip amplitude and dynamic pressure measurement experiments verified the analysis results. At the fracture speed, the dynamic pressure signal exhibited multiple non-synchronous frequencies. The amplitude of the fan blade tip was 3.7 mm, and the vibration stress at the crack initiation position was 671.2 MPa. The fan blade skimmed forward greatly, resulting in low local bending mode frequency of the forward-swept part, which was easy to be resonated. Blade improvement measures were taken to remove partial forward-swept of the blade and eliminate the local bending mode. The experiment verified the effective improvement measures. This scheme is of reference significance for the design of forward swept small aspect ratio fan blades to avoid asynchronous vibration.

With the growing use of throttleable liquid rocket engines, a simplified physico-mathematical model for a plane thruster chamber’s unit, equipped with liquid-liquid throttleable pintle injectors, was formulated and numerically simulated. Spray, flow, and mixing characteristics of the injector spray were analyzed. The impact of film angle and momentum ratio on these characteristic was determined. Results revealed a distinct oxygen-rich low-velocity zone, with fluid crushing influenced by total momentum. Flow exhibited a unique recirculation zone beneath the injector sleeve, where vorticity was intensified with momentum ratio rise, unaffected by film angle. Variations in film angle and momentum ratio altered the radial velocity near the impact, affecting the flow rates regionally. Mixing efficiency declined with film angle increase, while kerosene space utilization initially rose then fell with momentum ratio from 1.5 to 3.0, peaking at about 2.37.

In the reliability design process of rotating machinery, the assumption of failure independence between different components may lead to unsatisfactory results in reliability allocation. Therefore, a reliability allocation method based on failure correlation and fuzzy analytic hierarchy process (FAHP) was proposed. Based on the analytic hierarchy process model and initial evaluations, a non-fuzzy evaluation matrix was constructed for all system components, aiming to reduce the subjective influence from the initial evaluations. By introducing the load dispersion coefficient, a failure correlation factor reflecting the failure correlation between components was devised to adjust the coefficients of complexity, severity, and occurrence within the evaluation matrix, and the reliability weights of components were calculated. A reliability allocation model was constructed by inputting reliability weights into the Gumbel Copula function. Taking a liquid rocket engine turbopump as an example, a comparative analysis was conducted. Since taking into account failure correlation, the proposed method can provide a more reasonable and lower cost reliability allocation scheme, where the turbine and shell with higher failure hazard were reasonably allocated higher reliability indexes of 0.99982 and 0.99992, and the dimensionless cost of system was 95.21305, much lower than the results under the assumptions of failure independence and negative correlation.

The algorithm for reconstructing the profile of transparent liquids from a single viewpoint was investigated. A correction factor was introduced in the iterative error calculation steps to refine the reconstruction process. The influencing factors of the correction factor were explored, and a method for determining it was provided. The validity of the algorithm was verified, and the optimized algorithm was applied to analyze the reconstruction of multiple liquid surface profiles and 3D liquid surfaces. The results showed that the correction factor should be segmented based on the curvature of the curve and the step size of feature point placement during the reconstruction process. After applying segmented correction, the overall accuracy was improved by 30.88%, and the maximum error was reduced by 45.72%. The optimized algorithm enhanced the reconstruction accuracy for synchronized long feature points. For the standard liquid surface profile and the standard 3D liquid surface reconstructed, the number of feature points required to achieve the same accuracy was reduced by 38.46% and 20%, respectively. Additionally, the algorithm demonstrated effective control of cumulative errors and exhibited broad applicability.

A physically enhanced prediction method based on particle swarm optimization (PSO) optimization support vector regression (SVR) was proposed to address the issue of low accuracy and discontinuous isentropic efficiency of compressor performance at low speed. This method combined several physical extrapolation methods and utilized particle swarm optimization to optimize the support vector regression model parameters for the regression prediction of compressor’s low speed characteristics. The traditional single-stage characteristic parameters of the compressor were converted into corrected torque, and the compressor zero speed characteristic line was extrapolated. The high-speed characteristics were employed as the training set to establish the support vector regression model, PSO was used to obtain the high-precision regression SVR model, and the model was applied to predict the low speed characteristics of compressor. Results showed that the minimum determination coefficient was 0.976, the maximum mean square error was 0.06%, and the maximum average relative error was 8.96% in the prediction results based on PSO-SVR model. The three working states of compressor were predicted at low speed based on PSO-SVR model, including compressor, stirrer and turbine states. Therefore, the prediction method can fully mine the information contained in the data, improve the prediction accuracy of compressor at low-speed characteristics and provide data support for engine starting simulation.

In order to quickly obtain the combustion efficiency of different downstream positions of the integrated stabilizer under different conditions in the integrated afterburner length design stage. A prediction model of the combustion efficiency along the downstream of the integrated flameholder in the afterburner was proposed and verified based on the combination of numerical simulation and theoretical analysis, under the conditions of 600—1250 K incoming flow temperature, 75—170 m/s incoming flow velocity and 0.08—0.16 equivalent ratio, taking U-shaped integrated flameholder as the research object. At the same time, the empirical prediction formula of turbulent pulsation velocity for an integrated afterburner were determined. It found that the predicted error of the model is less than 4% in the non-fuel spontaneous combustion condition and less than 12% in the fuel spontaneous combustion condition.

The coupling relationship between the configuration state parameters of the interface-connected rotor system has an important impact on the aero-engine vibration. Taking into account the distribution of component center of mass deviation and skewness of the principal axis of inertial along the axial direction, it was proposed that the essence of rotor configuration state quality control is to optimize and control the consistency of the “three axes”. So that the rotational inertia internal load of the interface was reduced. A virtual assembly prediction model based on measurable geometric and mechanical parameters of rotor components was established. The collaborative optimization of rotor coaxiality, unbalance, and internal load of the interface has been achieved. A closed-loop control scheme covering the entire process of manufacturing-assembly-disassembly has been developed. Engineering applications demonstrate that the engine test first-pass qualification rate was increased by over 15%, while the fluctuation amplitude of the majority of key parameters was reduced by more than 30%.

To make the mechanical model better match the physical model, a dynamic model of the helicopter main transmission system considering the structural flexibility was constructed by using the hybrid lumped parameter/finite element modeling method. Based on the established model, the inherent characteristics were analyzed, and the resonance energy transfer characteristics of the system were comprehensively studied using Campbell diagram and modal energy method. The results showed that the vibration modes of the helicopter main transmission system can be classified into global pure torsional vibration mode, fixed shaft-planetary gear coupling vibration mode, local planetary gear vibration mode and local fixed shaft transmission vibration mode. When the input speed was close to 5210 r/min, the meshing frequency of low-speed stage was easy to cause the resonance of the system, and the vibration energy may be mainly transmitted through gear 6, shaft 5, bearing 10 and bearing seat 10. Among them, gear 6 and bearing 10 were considered dangerous components due to their large modal energy.

The noise generated by impinging jets is a critical issue during the takeoff and landing of fighter jets and carrier-based aircraft. However, the mechanism behind the generation of tone in subsonic impinging jets remains poorly understood. This study focuses on an impinging jet with a nozzle pressure ratio of 1.69 and an impingement distance of three nozzle diameters. Large eddy simulation is employed to capture the near-field flow, and an acoustic analogy model is used to predict the far-field noise. The extended Helmholtz decomposition and principal correlation decomposition methods are applied to analyze the modal characteristics and their correlation between the near-field flow and far-field noise. The results reveal that, due to the influence of the impingement plate, a supersonic region appears in the flow field, forming shock cells. The noise modal of the impinging jet exhibit an axisymmetric structure. The tone in the far field is closely related to the shock waves in the near field. The associated modal structure manifests as waves that gradually propagate outward from the impingement region, and the transmission process connects the vortex structures on the flat plate with the shock structures in the free jet region. The pressure field in the near field dominates noise generation compared to the velocity field. The alternating positive and negative amplitude of the modal amplitudes in the impingement region is a key factor in the generation of tone. This study elucidates the mechanisms of impinging jet noise and provides a theoretical foundation for the optimization and control of impinging jet noise.

The matching problem between pulse detonation combustor and turbine components is one of the difficulties restricting the development of pulse detonation turbine engine. In order to clarify the research progress on turbine performance and optimization strategies under pulse detonation gas impact, the flow characteristics and evolution process of exhaust flow field were briefly introduced with the pulse detonation combustor exhaust characteristics as the starting point. From the aspects of theoretical analysis, experimental testing and numerical simulation, the research status of key issues such as turbine performance evaluation methods, turbine operating characteristics, turbine internal flow characteristics and loss mechanism under pulse detonation gas impact was summarized. The research achievements of Northwestern Polytechnical University in the field of pulse detonation combustor and turbine component matching were reviewed, and the key problems to be solved in this field were prospected. Optimization strategies for turbine performance under pulse detonation gas impact were proposed, including stabilized pressure device, shock attenuation device, and turbine blade optimization design, etc. This indicates that the organic integration of various optimization strategies is one of the future research directions.

The permanent magnet synchronous motor is a critical component of electric propulsion systems, and its operational status is integral to the system’s safe functioning. Failure of the permanent magnet synchronous motor (PMSM) can trigger multiple physical quantity changes, making it challenging to achieve accurate fault diagnosis relying on a single signal source. To address this issue, a decision-level multi-channel information fusion fault diagnosis method was proposed, by combining convolutional neural networks and gated recurrent units (CNN-GRU) and improved Dempster-Shafer (D-S) evidence argumentation. Initially, analytical and finite element methods were employed to quantitatively analyze the vibration and current frequency domain characteristics of local demagnetization and rotor eccentricity faults in permanent magnet synchronous motors, thereby enhancing the interpretability of diagnostic results. Subsequently, a decision-level fusion diagnostic model was established, by integrating CNN-GRU and improved D-S evidence theory based on Pignistic probability distance and weighted Deng entropy. Finally, a motor fault simulation tester was constructed, and the model was validated using experimental data. The results demonstrated that multi-channel fault diagnosis is superior to single-channel diagnosis results. The decision-level fault diagnosis based on multi-source data, with the fusion of 4 channels, achieved diagnostic accuracy of 100%, 100%, and 99.3%, respectively, under three operating conditions. The proposed method accurately identified the types of permanent magnet synchronous motor faults, providing a reference for fault diagnosis in electric propulsion systems and offering potential value for engineering applications.

Given the lack of analysis methods for thermal-fluid coupling in high-speed gear transmission systems with complex oil branch lubrication, a thermal flow analysis model based on the Multiphase Flow Finite Volume Method was proposed. The flow field and temperature field within the gearbox were studied with this model. Bending and abrupt contraction of the oil route caused significant pressure drops. This affected the jet velocity and oil-air ratio at the outlet, leading to fragmentation of high-speed gear meshing jets and lower lubrication and cooling performance. It was found that the initial structure of the gearbox suffered from jet fragmentation and oil storage issues, resulting in a windage loss of 12.3 kW, which was 85.20% of the total loss. Improvements to the oil system increased the jet velocity and oil-air ratio. By adding oil outlets and installing shrouding, the system’s windage loss was significantly reduced to 5.22 kW, increasing the transmission efficiency from 93.54% to 96.67%.

The aerodynamic instability inception of aero-engine multistage high pressure ratio compressor has weak feature, complicated mechanism and difficulty in detection and prediction. A systematic experimental investigation of rotating stall and surge inception in a full-size multistage high pressure ratio axial flow compressor was developed. The rotating stall and surge characteristics at different rotate speed and angel of variable vanes were obtained by using a multi-section, multi-phase, high sample rate test method combined with the whole ring and sector uniformed measuring point distributions. These efforts aimed at providing an important research basis for compressor aerodynamic stability research and aero engine active safety control. The results showed that at the relative corresponding speed of 0.8 and 1.0, the instability was manifested as surge; before entering the surge state, the amplitude and order of the circumferential modal wave of inception characteristic frequency increased gradually; the stage leading to instability was different at different speeds. The influence of variable vanes angle was reflected in the modal wave amplitude change before surge.

The criterion of the farthest transmission path minimum energy loss for complex space pipelines was proposed. Based on the clarification of the connotation and determination basis of such a criterion, a finite element model of complex space pipeline system was created, and the division of vibration transmission path and the prediction of vibration power flow curve corresponding to different transmission paths were also achieved. A complex pipeline vibration transmission testing platform was established. By comparing the calculation results between experimental tests with finite element predictions, it was found that the maximum calculation error for the first three natural frequencies was 3.5%, both mode shape results matched well, and the variation trend of the vibration power flow curves at the output end of the pipeline system under different transmission paths obtained by the two methods demonstrated a good consistency, with the maximum error of the power flow peak only reaching 12.9%, which verified the correctness of the model. In addition, such a criterion can be employed to rank the farthest transmission path of the vibration of space pipelines, and after combining with the power flow loss results obtained under different resonance states, the main vibration transmission path can be effectively identified. The research results can provide a new ideal and approach for the vibration reduction, isolation, and avoidance of complex space piping systems in aero-engines.

Large eddy simulations of hydrogen and hydrocarbon fuel films with boundary-layer combustion were conducted, focusing on the effects of boundary-layer combustion on the near-wall heat and mass transport processes of fuel films. The results showed that boundary layer combustion effectively reduced the heat and mass transport fluxes between the film and mainstream, which enhanced the heat insulating performance of the film. However, the decrease in heat flux within the hydrogen film was not adequate to offset the negative impacts of heat release and consumption of hydrogen with high heat capacity during combustion, leading to a significant deterioration of the heat insulating performance during boundary layer combustion. On the contrary, the reduction of heat flux in the hydrocarbon film had synergistic effect with the heat absorption of the near-wall pyrolysis reactions, leading to a substantial enhancement in the heat insulating performance of the hydrocarbon film.

Based on the measured data of throat area and blade surface geometry, the uncertainty distribution of throat area was analyzed by numerical simulation method on the basis of quantitative modeling of surface geometry deviation. Using this as a characteristic correction variable, polynomial fitting was used to establish a turbine characteristic correction method considering the deviation of throat area. The results showed that the dispersion of the high-pressure turbine throat area was between −1.08% and 1.4%, and the average mass flow increased by about 1.813%; when measuring three cross-sections, the maximum measurement area error can reach 3%. Increasing the number of measurement cross-sections can effectively reduce the measurement error; the mass flow rate and efficiency prediction error of the turbine characteristic correction model can be controlled below 2% through stepwise fitting and training verification.

The theoretical model of the windage resistance heating characteristics of the bolt structure in the rotor-stator cavity was analyzed. A numerical solution model for windage resistance heating characteristics of the bolt structure in the rotor-stator cavity was established to analyze the internal flow field characteristics of the rotor-stator cavity and study the influences of structural and operating parameters on windage resistance heating of the bolt structure in the rotor-stator cavity on the basis of verifying the accuracy of the solution model. The theoretical formula for windage resistance heating of the bolt structure in the rotor-stator cavity was constructed using the correction coefficient method. The research results indicated that when the bolt rotated with the rotating disc, the surface of the bolt interacted with the viscous airflow through friction, and the temperature of the gas increased, resulting in windage resistance heating effect. Under the working conditions studied in this article, when the speed increased from 6000 r/min to 15000 r/min, the wind resistance temperature rise coefficient of the bolt structure increased by 24.2%; the inlet flow rate increased from 2.2 g/s to 8.0 g/s, and the wind resistance temperature rise coefficient of the bolt structure decreased by 51.5%; when the number of bolts increased from 6 to 36, the wind resistance temperature rise coefficient of the bolt structure increased by 65.1%; the theoretical formula for windage resistance heating was caused by the bolt structure of the rotor-stator cavity, which can accurately calculate windage resistance heating caused by the bolt structure of the rotor-stator cavity. This article has provided a theoretical basis for the analysis of windage resistance heating characteristics of the bolt structure of the rotor-stator cavity.

The flight principle of the compound high-speed helicopter was analyzed, the control strategies for different flight modes were designed, and the flight dynamics model was constructed. An attitude active disturbance rejection control algorithm based on high-order sliding mode differentiator was proposed to address the problem in accurately measuring system states caused by sensor noises and complex flight environment. Then an attitude tracking control system for the compound high-speed helicopter was constructed in MATLAB/Simulink, and simulation comparison was conducted with PID (proportion integral differential) controller and LQG (linear quadratic Gaussian) controller. The research results indicated that the proposed controller can quickly and stably track the target attitude angle without overshooting. Compared with PID controller and LQG controller, the error convergence speed increased by 31.5% and 64.2%, respectively. Moreover, under noise disturbance, the maximum tracking error was 34.2% smaller than LQG controller and 39.5% smaller than PID controller.

To obtain health indicators that can accurately describe the degradation process, a new health indicator based on envelope spectrum statistical features and Pearson correlation coefficient was proposed for remaining useful life prediction. Firstly, a first prediction time identification method was proposed based on the Boostrap sampling method and the 3 sigma principle to obtain a suitable full-life degradation threshold. Secondly, the envelope spectrum probability distributions at different time points were calculated and the health index was obtained based on the Pearson correlation coefficient. Finally, the remaining useful life of bearing was predicted by a hybrid model of exponential and linear regression. The experimental results showed that the proposed health index can effectively reflect the bearing degradation trend, and the prediction accuracy of the hybrid exponential and linear regression model was improved by 23.7% compared with other prediction models.

The unsteady transonic aerodynamic characteristics of a NACA64A010 airfoil in air and heavy gas medium R-134a were simulated by using the Reynolds-averaged Navier-Stokes equations (RANS) and the Spalart-Allmaras (SA) one-equation turbulence model. Under the same Mach number, Reynolds number and reduced frequency, the calculation results of simple harmonic motion in pitch showed that the distribution of modulus and phase of unsteady pressure coefficient C_p in heavy gas medium was obviously different from that in air, the unsteady lift coefficient C_l was not different from that in air medium, and the amplitude and phase of pitch moment coefficient C_m were different from those in air medium. The transonic similarity law was applied to the unsteady aerodynamic force correction, and the amplitude and phase of pitch moment coefficient C_m were transformed into air similarity, but with the increase of reduced frequency, the correction effect of transonic similarity law became worse. Through analysis of pitching moment work, it was shown that if the unsteady aerodynamic force in heavy gas was not corrected by similarity, the flutter characteristics in air and heavy gas could be different, thus affecting the evaluation of flutter characteristics in heavy gas wind tunnel. This study can provide a basic support for the follow-up research on flutter characteristics of aircraft in heavy gas media and the development of correction methods.

Three distribution forms of uniform fouling thickness, leading edge fouling thickness and trailing edge fouling thickness were established for the typical thickness distribution of compressor blade fouling, a diffuser cascade was taken as the research object, and numerical simulation was carried out to study the influences of different distribution forms and sizes of fouling on the degradation of aerodynamic performance of the compressor cascade. The results showed that the uniform fouling thickness overestimated the degradation of aerodynamic performance caused by the non-uniform thickness distribution of fouling on real compressor blades. The results of all three thickness distribution forms showed that the degradation of aerodynamic performance was more pronounced under negative incidence. Compared with the trailing edge fouling thickness, the effect of leading edge fouling thickness on aerodynamic performance degradation was more significant. When the maximum thickness of fouling was similar, the loss caused by the leading edge fouling thickness increased by 84.35% and the pressure ratio decreased by 0.98% compared with the trailing edge fouling thickness, which further illustrated the fact that the modelling of the compressor blade fouling should take into account its non-uniform distribution location and size to obtain more reliable results in the assessment of the aerodynamic performance degradation.

Under actual operating conditions of compressor, factors such as shaft eccentricity, casing deformation or machining and assembly errors may lead to circumferential non-uniform characteristics of blade tip clearance, yielding an important impact on its performance, especially its stability. Taking NASA Rotor 67 as the research object, the circumferential non-uniform conditions caused by shaft eccentricity were studied numerically. The circumferential non-uniform distribution conditions of tip clearance were established by means of casing-rotor separated-domain modeling, and the three-dimensional compressor flow field under three eccentricity conditions was simulated to obtain the circumferential non-uniform distribution rules of flow field, and the stall criteria of compressor with circumferential non-uniform clearance were established based on the "equivalent clearance principle". The results showed that the increase of tip leakage flow in the large clearance area of the eccentric rotor led to the intensified blocking effect, and the local tip flow coefficient decreased and got closer to the stall boundary, while the small clearance area was opposite. The extremum position of flow coefficient did not coincide with that of gap and had a certain phase delay. The stalling flow coefficient of the eccentric rotor was close to that of the concentric rotor with the average clearance of large gap sectors. Therefore, the equivalent principle of "average clearance of large gap sectors" can be used as the criterion for estimating the stalling point of the circumferential non-uniform gap compressor.

The wing/rotor aerodynamic interference of multi-rotor tilting wing aircraft in helicopter mode was analyzed. Computational model for the aerodynamic interference analysis was established based on CFD method. The aerodynamic interference of the wing/rotor in hover and forward flight was numerically calculated, the influence of aerodynamic interference on the aerodynamic characteristics of the wing was analyzed, and the influence law of flap and aileron deflections was further studied. The results showed that, under the influence of rotor wake, the wing generated a large backward force, accounting for 5.4% of the total rotor force. The flap and aileron deflections during hovering can achieve effective longitudinal and heading maneuvers, and the flap and aileron up deflections during forward flight can reduce the backward wing force by 30%, thereby increasing the maximum forward flight speed.

Based on the energy method for numerical calculation of aeroelastic stability of turbo-mechanical blades, a simulation method for solving the aeroelastic stability of labyrinth seal rings was developed. Based on the geometrical parameters of the grate tooth sealing ring in the standard literature, a numerical model of the aeroelastic stability was established, and the accuracy of the numerical model was verified. The effects of cavity width, wall thickness and sealing clearance on the aeroelastic stability were studied. The results showed that the calculation results of the aeroelastic stability of the labyrinth seal ring had the same instability nodal diameter and the same aeroelastic stability trend as the standard literature. Therefore, the accuracy of numerical method to solve the aeroelastic stability of labyrinth seal ring was verified. With the increase of the width of the tooth cavity, the aeroelastic stability was worsened, and the influence of the width of the tooth cavity on the aeroelastic stability increased with the pressure ratio. With the increase of wall thickness and sealing clearance, the aeroelastic instability was improved, and the influence of wall thickness and sealing clearance on the aeroelastic stability decreased with the pressure ratio.

Electronic components are prone to overheating and failure during operation. Microchannel heat sinks (MCHS) has the advantages of compact structure, good heat exchange performance, and broad application prospect. A wall-mounted flexible vortex generator (FVG), a tandem column, and a non-Newtonian fluid were utilized to enhance the heat transfer in a microchannel. The finite element method in Arbitrary Lagrangian-Eulerian (ALE) format was used to solve the continuity, momentum, and energy equations describing the flow-solid coupling in microchannels. The Flow-induced Vibration behavior and heat transfer performance of a power-law fluid in a symmetric wall-mounted FVGs channel were investigated by adjusting the spacing between the FVGs and the columns, and varying the power-law exponent (n) of the fluid under the condition of a fixed distance between the two columns. When the distance between the downstream column and the flexible plate G_x=1.5 and the power-law index n=1.2, Nusselt number Nu increased by 184% and the thermal performance coefficient increased by 45% compared with the straight microchannel.

Taking the subsonic airfoil of a turbofan engine fan blade as the research object, numerical simulation of the original airfoil and two leading edge erosion airfoils at five atmospheric heights was carried out with reference to the Committee on Extension to the Standard Atmosphere(COESA) standard atmospheric model, so as to explore the influence of leading edge erosion on the flow characteristics of subsonic cascade in high altitude and low Reynolds number environment. The results showed that at 0° angle of attack, the low Reynolds number condition weakened the sensitivity of the airflow to the leading edge morphology, so that the pressure distribution at the leading edge, the flow separation and transition of the boundary layer on the suction surface, and the separation degree of the trailing edge of the three airfoils tended to be consistent. Finally, the increase of total pressure loss caused by the leading edge erosion became smaller with the decrease of Reynolds number, and the total pressure loss coefficient and pressure ratio of the three airfoils tended to be consistent. At the angle of attack of 4°, the leading edge erosion can promote the occurrence of transition at low Reynolds number and reduce the degree of trailing edge separation, so that the total pressure loss was less than 0° angle of attack; when the inlet Mach number was 0.6, it reduced by 23.6%, and when 0.8, it reduced by 41.2%.

This work focuses on the laminar combustion characteristics of a certain bio-jet fuel with hydroprocessed esters and fatty acids (HEFA) technology. The laminar burning velocity (LBV) of HEFA bio-jet fuel and HEFA bio-jet fuel/RP-3 mixed fuel were obtained by a constant volume combustion experimental device. The experimental parameters include equivalence ratios (0.8—1.4), initial pressures (0.05, 0.1, 0.15 MPa), initial temperatures (450, 470 K) and HEFA mixing ratios (0, 0.1, 0.2, 0.3, 0.5), et al. The effects of the equivalence ratio, initial pressure and initial temperature on the LBV of HEFA bio-jet fuel and the HEFA mixing ratio on the LBV of HEFA/RP-3 mixed fuel were analyzed. It is found that with the increase of equivalence ratio, the LBV of HEFA bio-jet fuel shows a trend of increasing first and then decreasing, and its peak value appears near the equivalence ratio of 1.1. As the initial pressure increases from 0.05 to 0.15 MPa, the LBV of HEFA bio-jet fuel decreases by 13.54 %. As the initial temperature increases from 450 to 470 K, the LBV of HEFA bio-jet fuel increases slightly. With the increase of HEFA mixing ratio from 0 to 0.5, the LBV of HEFA/RP-3 mixed fuel increases by 3.85 %. The results show that compared with RP-3 jet fuel, although the carbon number distribution of HEFA bio-jet fuel is higher, its chemical composition is dominated by alkanes, which leads to its LBV higher than that of RP-3 jet fuel, so that the LBV of the mixed fuel after blending HEFA is slightly increased. This study provides a theoretical basis for the application of bio-jet fuel in aircraft engines.

To broaden the effective attack angle range of the blade cascade under the influence of low Reynolds number and achieve the purpose of blade cascade expansion and loss reduction, numerical and experimental researches on a small deflection angle diffuser blade cascade were conducted. Three riblet-structures with different angles were selected to control it by studying the main characteristics of turbulence after blade stall. The research results indicated that diffusion type surface ridges are an effective means to expand the effective attack angle of the blade cascade and reduce losses. Its characteristic was represented by expanding the positive stall boundary without sacrificing the loss of the design angle of attack, with a maximum reduction of 17.7%. No-angle riblets achieved reduction at a high angle of attack, with a maximum reduction rate of 19.58%. Overall, the reduction effect of diffusion type riblets was better than that of non-angle riblets and better than that of convergence type riblets. In addition, it was found that the width and peak loss of the wake decreased after reduction, and the diffusion line of the wake shifted towards the pressure surface side.

A method for real-time monitoring of rolling bearings in aircraft engines based on characteristic energy was proposed to address the challenging issue of real-time monitoring of rolling bearings in aircraft engines. This method first decomposed the original vibration signal using CEEMDAN to obtain several components, and then calculated the kurtosis and correlation coefficient of each component. Subsequently, based on the kurtosis-correlation coefficient criterion, it selected strong impact components for reconstruction and performed envelope demodulation to maximally retain effective information related to bearing fault impact components. Finally, the method calculated the characteristic energy of faulty and normal bearings from the information in the envelope spectrum, established a diagnostic baseline and characteristic energy belt, and achieved monitoring of bearing operating status. The effectiveness of this approach was validated using data from the Case Western Reserve University deep groove ball bearing test rig, a constructed rolling bearing test rig, and a test rig for a certain type of turbofan aircraft engine bearing component. The results showed that the proportion of the characteristic energy of the outer ring fault bearing in the whole envelope spectrum energy was 59.5%—75.9%, and the method can provide an effective means for the fault diagnosis and online monitoring of the main bearing of aircraft engine.

Considering the complex aerodynamic interference phenomena in the helicopter and fixed-wing modes of the tilt-quadrotor, a set of numerical methods suitable for simulation of the interference flow field of the tilt-quadrotor were established based on the CFD method. The aerodynamic interference between the rotor multi-vortex system and the wing and fuselage was systematically simulated, and the occurrence and evolution mechanisms of unsteady aerodynamic characteristics were revealed. The results showed that in the helicopter mode, the blade root vortex generated by the front rotor could act on the upper surface of the front wing and roll up the secondary vortex below the wing. The blade root vortex generated by the rear rotor could be attracted by the tip vortex of the front rotor, and the obvious left shift phenomenon may occur and act on the leading edge of the rear wing. In the fixed-wing mode, the tip vortex dragged out by the front rotor could move backward rapidly, and the aerodynamic interference with the rear rotor was less. The front/rear rotor had less influence on the fuselage pressure distribution, and the fuselage as a whole showed typical fixed-wing fuselage pressure characteristics.

Longitudinal tensile tests were carried out on SiC_f/TC17 composites at room/high temperature to investigate the tensile behaviors. The damage evolution and failure mechanisms were revealed based on microscopic fracture morphology analysis. Afterwards, a constitutive model was developed to describe the tensile behaviors of SiC_f/TC17 composites. The results showed that the ultimate tensile strength of SiC_f/TC17 composites decreased with the increasing temperature, while the nonlinear segment of the stress-strain curve increased. The major failure mechanisms at room temperature lied in multiple fractures of the interfacial reaction layer and random breakage of weak fibers, whereas large-scale interface debonding and fiber pullout, matrix cracking and fiber breakage were more common at high temperatures. The results of different strength-predicted models demonstrated that the failure mode of SiC_f/TC17 composites at room temperature was controlled by local loading sharing, while the high-temperature ultimate tensile strength was more consistent with the global loading sharing model. The stress-strain curve of SiC_f/TC17 composites was simulated by the proposed constitutive model with coupling fiber cumulative damage. The simulation results exhibited a trend similar to that of the experimental data at 25 ℃ and 450 ℃. Finally, based on the tensile properties obtained from the tests, finite element stress-strain analysis and static strength calibration of the TMCs blade-ring structure were carried out. The result indicated that the blade-ring structure exhibited a significantly elevated strength reserve factor at the typical service temperature.

Research on the dispersion control methods during the engine acceptance test process is a key to control performance dispersion throughout the entire life cycle of aero turbofan engine. Based on the deviation of component characteristics and geometric area, carry out research on between engine matching and performance dispersion. The control methods of fan pressure ratio tolerance ±0.75% and rotors relationship ±0.5% is proposed. Tests on 10 new engines, thrust dispersion reduces from ±4.9% to ±2.2% and exhaust gas temperature dispersion reduces from ±2.0% to ±0.7%. Based on the control methods of fan pressure and rotors relationship, an engine performance degradation model which characterizes the relationships among thrust degradation, working time, exhaust gas temperature and repair is established to solve thrust dispersion during the whole life cycle. The control method of exhaust temperature margin 20—30 ℃ of new engine is proposed. Endurance test on 1 engine, intermediate state thrust is degraded 1.2%. The superimposed thrust dispersion of the other 9 new engines is ±3.4%, Which is solving the thrust dispersion problem during the whole life cycle.