The definition and implication of numerical simulation technology were proposed based on the technical development requirements of advanced aero-engines development.The dimensions of aero-engine numerical simulation technology were studied from specialty,discipline,space,time and tools,while the strategic position and function in promoting the transformation of aero-engine development method were analyzed.The disparity of numerical technology for aero-engine at home and abroad was discussed by analyzing the development status of typical foreign research plans and projects.It is proposed to fully understand the important position and function of numerical simulation technology in the aero-engine development,and to build and develop our own aero-engine numerical system and industry-wide “aero-engine database” as soon as possible.

In view of the difficulties in the design optimization of an overall aircraft power system, such as the considerable calculation scale, complex coupling relationships, sharp discipline conflicts and complicated implementation processes, a variety of advanced technologies such as the high-precision surrogate models, the efficient optimization strategies and the intelligent multi-objective optimization algorithms were developed and applied in three aspects: system decomposition, system modelling and system solution. Finally, the design method of an overall aircraft power system based on multidisciplinary design optimization (MDO) was established. Taking a turboshaft engine, a turbofan engine, a turbojet engine and a main transmission chain of the main reducer of a helicopter transmission system as examples, the engineering application research on MDO of the overall aircraft power system was carried out. Moreover, the performance and strength tests of the optimised compressor, and the optimised overall system test were carried out and verified for the turbojet engine. The research showed that the MDO of the overall aircraft power system could effectively unleash the design potential, significantly improve the comprehensive performance of products and considerably shorten the development cycle. Its engineering application prospect is quite broad, which will change the design and development of an aircraft power system.

The adaptive cycle engine is an important candidate power unit for the next generation of aircraft. After a brief introduction of the research progress at home and abroad, focuses were put on the overall design research of the adaptive cycle engine currently carried out. Firstly, the overall performance scheme and structural form of the typical adaptive cycle engine were introduced. Then, the performance benefits and costs of the adaptive cycle engine were discussed. Finally, the overall design development trend of the future was presented. It is believed that the overall design of adaptive cycle engines in the fu-ture should be carried out under the parallel multi-dimensional and multi-disciplinary optimization system, considering the influences of multi-source uncertainties. Hybrid-dimensional simulation methods can be applied to evaluate the technical characteristics of novel components, and the comprehensive performance optimization of aircraft/engines can be carried out according to the requirements of flight missions.

In order to improve and develop the research system of radially rotating heat pipe, the research status of radially rotating heat pipe along with its engineering application was summarized, including heat transfer performance research, experimental research and numerical simulation research; the mechanism of heat pipe turbine disk temperature and stress control was introduced; other applications of radially rotating heat pipe technology in precision grinding and cooling of high-speed motor were reported. The deficiencies in the study of flow and heat transfer mechanism, numerical simulation model and heat transfer limit of radially rotating heat pipe were presented. The mechanism of heat pipe turbine disk heat transfer and stress control and wider engineering applications of radially rotating heat pipe need to be further explored. Detailed suggestions for the future research of radially rotating heat pipes were proposed, including: combination of experiment and numerical simulation; focus on the flow and heat transfer characteristics of heat pipes under centrifugal and Coriolis force; determination of the heat transfer limit of radially rotating heat pipes, so as to provide a theoretical support for its popularization in engineering application.

Based on the layout and test characteristics of the aviation engine altitude ground test facilities, the simulation system was created, the required equipment was modeled, and application research was conducted. The mathematical modeling of key test equipment such as piping, regulating valves, throttling components, hydraulic control system and Laval nozzle (used to simulate engine flow) was carried out to obtain a model library. A comprehensive digital simulation system and a semi-physical simulation system were built, and their performance was demonstrated by comparing them with real-life test results. The engine’s intake temperature and pressure calculated by the simulation system were compatible with the dynamic change trend when compared with the test data, according to the simulation verification results and analysis. Where temperature relative error was limited to a maximum of 0.5 percent, the pressure relative error was limited to a maximum of 3 percent. The pressure regulation time was shortened to 1/3 by the control technique created based on the model and simulation system of the high-altitude test equipment, the transition state regulation performance was significantly enhanced.

The matching of the double bypass engine compression system was investigated to reveal the operating mechanism of the multi-connected compression system of the variable cycle engine. Three-dimensional numerical simulation showed that there was a strong coupling between the compression components, the bypass ducts, and the bypass regulation mechanisms. A well-organized bypass flow was a key for the variable cycle engine to exhibit its aerodynamic advantage. The adjustment of the bypass regulation mechanism not only changed the flow state of the bypass but also influenced the matching state of the compression components, hence the mode transition required for the cooperation of all components. To realize fast matching analysis of the multi-connected compression system, an integrated variable-dimensional analysis method for variable cycle compression system was proposed by combining the component through flow program with the zero-dimensional bypass program. The feasible region of the compression system for the single bypass mode was deduced based on the variable-dimensional analysis method, thus providing the theoretical basis for the matching design of variable cycle engines.

Under the self-developed optimal transport meshless (OTM) framework, a thickness growth algorithm was developed and the anisotropic oxidation growth process of thermal oxide layer (TGO) was simulated. Using this method, the typical transition section of TGO layer was taken as an object to study the change of stress and displacement under thermal cycling load. The simulation results coincided with the test. The results showed that this method can well simulate the wrinkle phenomenon in the process of interface growth. Compared with the finite element method, the deformation of the element was uniform, making it suitable for numerical simulation of the growth process of TGO. The maximum stress of the thermal barrier coating mainly occurred at the convex position, and the lateral change of the convex position had a tendency to aggravate the large deformation of the oxide layer interface.

The application status of additive manufacturing technology in aero-engine was introduced, and the key technologies of strength, lifetime assessment and design methods for additive manufacturing structures in aero-engine were emphatically reviewed. The shortcomings and development trends of existing research were discussed in terms of nondestructive detection methods and equivalence criteria for additive manufacturing defects, strength and lifetime prediction methods considering the effects of defects, and the integrated innovative structure design technology of aero-engine additive manufacturing structures, respectively. The results showed that strength and life assessment of additive manufacturing structures in aero-engine was at the initial stage. Most of them were geared for the additive manufacturing materials and simple structures under a single failure mode, making it necessary to conduct research on compound fatigue failure, multi-factor failure coupled of structural characteristics-widespread defects-surface morphology, so as to develop data-driven evaluation methods, damage tolerance design methods and special test techniques for aero-engine additive manufacturing structures.

In order to better utilize the outstanding mechanical properties of shape memory alloys (SMA) in both shape memory effect and superelasticity in aerospace field, the research progress of SMA in terms of materials and process, constitutive model, shape memory effect application and superelasticity application was reviewed. The research features of the ternary high-temperature alloys, such as NiTiHf and NiTiAu, SMA heat treatment process and 3D printing process, shape memory effect and superelasticity constitutive model, SMA wire, tube, spring and belt actuators designed by using shape memory effect, vibration dampers and adaptive structures designed by using superelasticity in aerospace field were mainly discussed, and their current shortcomings were presented. The future trend of SMA application was proposed: with the integrated development of materials, process, control, and information technology, the structures with SMA could show more structure diversity and work in a wider range of working temperature with more intelligence.

By projecting the aero-elastic modes of the mistuned bladed disk to the modal space spanned by the tuned modes, a closed-form expression of the mistuned aerodynamic damping ratio as a linear superposition of the tuned damping was obtained. It was theoretically demonstrated that: a mistuned aero-elastic mode was constructed by several tuned and independent modes; and the contribution of the tuned modes with high aeroelastic damping can increase the aeroelastic damping of the mistuned mode. A method to predict the aerodynamic damping ratio of the mistuned bladed disk was proposed. It started with respective analysis of aero-elastic coupling and mistuning, and the aerodynamic damping ratio of the mistuned modes can be predicted. A single-time aeroelastic analysis was required for two significant benefits. One is reducing the experimental measurements and the other is decreasing the calculation cost to accelerate the optimization process of mistuning pattern. The bladed disk with NASA-Rotor37 blade profile was considered, and several typical mistuning patterns were applied with different levels of mistuning strength. Results showed that the closed-form expression had an error lower than 0.1%. The prediction method would overrate the stability boundary with error lower than 5%, so it is capable to capture the overall trends of the accurate results.

According to the finite element simulation results of the turbine disc, the stress state at the inspection position of the disk center was determined, and the ratio of the circumferential stress and axial stress and the radial stress gradient of the circumferential stress at the disc center hole were taken as the design indexes to ensure the consistency of stress state of the simulated specimen with the actual disc. Two kinds of simulators, namely, the multi axial simulators reflecting the biaxial stress state, and the plate notch simulators reflecting the stress gradient, were designed for the inspection position of the center hole of the wheel disc. Low cycle fatigue tests were carried out by using these two kinds of specimens under actual load conditions of the disc, and the test results were statistically analyzed. Furthermore, the multi-axial fatigue life prediction model and the life model considering the influence of stress gradient were used to evaluate the fatigue life of these two simulated specimens. The scatter bands of the prediction results of biaxial simulated specimen were within 2 and those of the flat notch simulated specimen were within 3. The results could provide engineering reference for the life evaluation of aeroengine structures.

The definitions of the nonlinear modes and the numerical methods to solve a nonlinear dynamic problem were reviewed. The progress in application of the damped nonlinear normal modes for structures with frictional damping was summarized. The related researches in the last decade for different friction damping devices were also summarized, including the tip shrouded blades, underplatform dampers, blade root damping and friction ring dampers. Furthermore, test works related to nonlinear modal testing were also reviewed. Open problems and future directions were highlighted. It was concluded that the nonlinear mode gradually developed from the theoretical stage to engineering stage. Combined with ad-hoc reduced order modeling techniques, nonlinear modal analyses were implemented to reveal the modal characteristics of high-fidelity finite element models of frictionally damped blades/blisks. The extended energy balance method and nonlinear modal synthesis were employed to build a bridge between nonlinear modes and steady-state response, which can significantly improve the efficiency of response-based parameter analysis. Nonlinear modal testing still in its infancy cannot be applied to engineering structures yet.

The stress distributions of the thermal barrier coating-substrate system with a hole under thermal mismatch were calculated using finite element method. Calculation results indicated that the interfacial normal/shear stress concentration occurred at the edge of the film hole, and higher circumferential stress existed near the edge of the hole. The coating near the hole edge was prone to cracking and spalling. The effects of the thickness of the ceramic layer, the outer diameter of the model, the thickness of the oxide layer, the hole diameter, and the temperature distribution on the local stress near the hole edge were calculated and analyzed. Results revealed that the region with interfacial normal and shear stress became larger while the thickness of the ceramic layer increased. The outer diameter of the sub-model, which was used to calculate the stress near the edge of the film hole, should be greater than the sum of four times of the ceramic layer’s thickness plus half of the diameter of the hole. Under the cooling condition, the interfacial normal and shear stress increased with the thickness of the oxide layer. At high temperature, the circumferential stress at the edge of the ceramic layer near the hole decreased with the diameter of the hole. The interfacial stress and circumferential stress near the hole edge increased in the case of non-uniform temperature distribution.

A macroscopic phenomenological model and a finite element subroutine were used to simulate the creep behavior of bi-crystal structural material with different orientations. The selected creep model had a clear physical significance, and three terms of the creep model can be used to describe the three stages of creep deformation process respectively. The effects of casting deviation angle, secondary orientation and grain boundary of bi-crystal structural material on the creep deformation characteristics were also analyzed. These results indicated that the casting deviation angle can significantly reduce the creep properties of the material based on two rotations of the material coordinate system. The secondary orientation had insignificant effect on the creep behavior of the material, and the inclined grain boundary can change the creep behavior of the material near the grain boundary. The change of the incline angle of grain boundary of the material caused the change of creep deformation behavior near the grain boundaries.

The mechanical model and finite element model of the high speed flexible rotor system of GTF (geared turbofan) engine were established. Based on analysis of the composition and characteristics of bearing dynamic load, the critical speed bearing dynamic load control method and modal shape control evaluation parameters were proposed. In this method, the angular motion of the large inertial component in the critical speed modal shape was controlled, the sensitivity of the rotor to couple unbalance excitation was reduced, and the node position of the bending modal shape was controlled close to the bearing, so as to realize the bearing dynamic load control under the lateral bending critical speeds of the high speed flexible rotor system. The results showed that the above critical speed modal shape control can be achieved by adopting the double supporting layout around the low-pressure turbine and optimizing the conical shell stiffness of the low-pressure compressor. The effectiveness of bearing dynamic load control method was verified by the 14% − 65% reduction of bearing dynamic load under the lateral bending critical speeds.

The forced response of the mistuned bladed disk was analyzed by use of the fundamental mistuning model (FMM) combined with the influence coefficient method. The finite element method and computational fluid dynamics method were used to obtain the modal results and aerodynamic influence coefficients of the tuned bladed disk, and then the FMM considering aerodynamic damping was constructed. The modal frequency, mode shape and damping ratio of the mistuned bladed disk were obtained by solving the matrix eigenvalues, and then the modal superposition method was used to calculate the forced response of the mistuned bladed disk under traveling wave excitation. The above method was used to calculate the forced responses of various intentional mistuned bladed disks before and after the superposition of random mistuning. Results indicated that the forced response of the mistuned bladed disk decreased significantly after considering aerodynamic damping. Taking the alternating mistuned bladed disk as an example, the forced response can be reduced by increasing the mistuning value within an appropriate range. The frequency difference between adjacent blades of the model should be 6%, and the corresponding mistuning value was 4.32. In addition, increasing the frequency difference between adjacent blades with an interval of one blade can further reduce the forced response.

In view of coupled vibration of boosters and labyrinth seals rotor/stator structure, and for overcoming the drawback of the prevailing method which disregards the effects of the nonlinearity of rub-impact and fails to identify the influence of key parameters, a coupled vibration dynamic model of booster rotor/stator was established, and the coupled vibration was calculated. Furthermore, the motion stability was analyzed and the influence of key parameters was revealed. The results showed three vibration modes of the coupled vibration: damped mode, sustained mode and divergent mode. Consequently, primary and secondary instability regions can be detected. The location of the instability region varied linearly with the modal frequencies, and changed inversely with the nodal diameters. Accordingly, the length and the number of the instability region reduced as the rotor/stator damping increased or the rubbing force and the initial excitation decreased. The higher angular acceleration indicated the lower coupled vibration amplitude. The results also showed the drawback of the traditional method based on the coupled resonance margin when predicting the dynamic response of the rotor/stator coupled vibration. Only the primary instability regions can be identified and the results were lower than the numerical results.

Based on Rotor37 single rotor compressor performance experiment data, Monte Carlo method and guide to the uncertainty in measurement method were used for uncertainty evaluation of measuring the compressor isentropic efficiency by temperature rise method, the measurement uncertainty of isentropic efficiency under different rotational speed and flow rate conditions were analyzed, and the allocation scheme of isentropic efficiency measurement uncertainty was studied. Results showed that the best estimate value and standard uncertainty evaluated by these two kinds of methods were basically the same, but the shortest inclusion interval of 95% inclusion probability evaluated by Monte Carlo method was narrower than that evaluated by guide to the uncertainty in measurement method, and the gap was bigger in case of non-normal distribution of input, so guide to the uncertainty in measurement method should be carefully used in such case. At the same speed, the measurement uncertainty of isentropic efficiency increased with the increase of flow rate. At different rotational speeds, the measurement uncertainty of isentropic efficiency increased with the decrease of rotational speed. It became a lot more difficult to measure the compressor isentropic efficiency of high accuracy by temperature rise method at low rotational speed. The results of isentropic efficiency measurement uncertainty allocation showed that higher accuracy of total temperature measurement was required for equal action allocation, making it more difficult for realization. Under the maximum flow condition of 70% designed rotational speed of the compressor studied, given that the relative measurement uncertainty of isentropic efficiency was 0.5%, the measurement uncertainty of total pressure and total temperature should reach 35 Pa and 0.083 K, respectively, according to equal accuracy allocation. Increasing radial measurement points, and using platinum resistance total temperature probe and gas flow total temperature calibration method can improve the measurement accuracy of both total temperature and isentropic efficiency.

A method for measuring 3-D unsteady flow at an exit of a transonic multistage compressor rotor was presented by using combined high-response pressure probes consisting of a cylinder 1-hole probe and a tip-wedged 1-hole probe, which were calibrated in a wind tunnel. The probes were put in the inter-stage at the same axial and radial but different tangential positions with the help of two traverse mechanisms, and unsteady pressure data were obtained utilizing a phase-locked technique at different angular positions. Distributions of 3-D flow parameters were obtained through data processing. Result showed that the combined probes were characterized by small size, high frequency, wide measurement range and weak interrupt compared with other current 3-D high-response probes, while the method can obtain flow pitch angle and more accurate total pressure, static pressure and Mach number compared with 2-D high-response probe techniques. The combined-probe technique provides a feasible 3-D flow measurement method for flow diagnosis in small multistage compressors.

Starting from the working principle and characteristics of turboprop engine and the design requirements of full authority digital engine and propeller control (FADEPC), the univariate regulation methods of equal turbine front total temperature regulation, equal rotor speed regulation and equal propeller power regulation, as well as the bivariate regulation methods of keeping the total turbine front temperature unchanged or propeller power unchanged by adjusting the fuel flow, and keeping the rotor speed unchanged by adjusting the blade installation angle were analyzed. According to the working characteristics of turboprop engine and from the perspective of control realizability, an open-loop + closed-loop structured piecewise scheduling bivariate combined control method was proposed, in which the propeller engine was adjusted according to equal propeller power, equal rotor speed and equal total temperature behind the turbine within the whole flight envelope, and switched according to the power limit height designed. The feasibility of the control method was verified by simulation under three different flight speeds, and the maximum relative error of rotating speed was 0.1%.

Based on the specific design requirements of heat exchanger in aviation and aerospace field, existing data were investigated and analyzed. Combined with the engineering practice, the main methods to improve the performance of heat exchanger, including performance optimization method and the interdisciplinary exploration of new types of heat exchangers, were introduced. The branch and distribution of current research on various heat transfer enhancement were summarized. Structure-based enhanced heat transfer was still taken as one of the main research directions, and some novel structures were introduced. In the exploration of new types of heat exchangers, the concepts of smart heat exchanger and chemical reaction heat exchanger were introduced, elaborating the design of heat exchange under multiple working conditions and high heat flow. Among them, the heat transfer enhancement coefficient of the smart heat exchangers can reach up to 2. Some new concepts, including evaluation indicator, uniform-mixed flow and operating weight produced in the design of heat exchanger, were analyzed and explored, which can supplement and improve the design system. Finally, the application prospect of heat exchanger in aviation and aerospace was presented.

To investigate the influences of swirler structures on the ignition performance of aero-engine combustors, the large eddy simulation combined with the wall-adapting local eddy-viscosity （WALE） sub-grid model, the dynamic thickened flame model, and a single spark was used to simulate the ignition processes of an axial-radial swirler combustor and a dual-axial swirler combustor. Results indicated that the flow fields at the ignition position changed over time due to the turbulent fluctuation under the same structures and conditions, so the ignition time affected the simulation result. To avoid the use of a single spark in simulations against the ignition results on the condition that the combustor can be ignited successfully in experiments, the spark should be started when the instantaneous velocity pointed to the recirculation zone and the velocity magnitude was smaller than the average value. Comparing the dynamic evolution process of the flow field of the axial-radial swirler combustor with the dual-radial swirler combustor, it was found that the ignition position of the dual-axial swirler combustor was directly opposite to the swirler jet, and the probability of the velocity pointing towards the recirculation zone was lower. The flame was prone to propagate downstream rather than to the recirculation zone. Therefore, the ignition performance of the double-axial swirler combustor was worse than the axial-radial swirler combustor.

Three aspects including plasma-assisted ignition and combustion experiments, plasma-assisted combustion diagnosis, and plasma-assisted combustion mechanism were reviewed respectively, and the latest results of the research on non-equilibrium plasma ignition and combustion were summarized. New advances in plasma-assisted combustion include: basic combustion research on plasma assisted cold flame, experimental studies of plasma assisted low-calorific fuel and ammonia combustion, studies of plasma-enhanced flame stability, simulation of hybrid discharge plasma-assisted ignition and combustion, and the establishment of new models of plasma-assisted combustion simulation, etc. It was believed that the relevant research should be centered on the experimental studies, then a wealth of data and experience could be accumulated for further research and applications; and different types of plasma should be considered carefully and their advantages should be fully utilized in experiments, because using the hybrid discharge plasma could maximize their effects on assisting ignition and combustion.

To meet the needs of high-fidelity numerical simulation of aero-engine combustors, the AECSC-IBM software was developed based on the immersion boundary method (IBM) and the large-eddy simulation-transported probability density function combustion model (LES-TPDF). The original geometric structure of the combustor was mapped with grid markers. The simulation accuracy of turbulent flow and combustion was verified by a simulation example of a twin-swirl combustor and Sandia jet flame. In the simulation of twin cyclone combustors, the average errors of axial,radial and tangential velocities at the outlet of the swirler were 15.7%, 23.8%, and 15.0%, respectively. The average relative errors of temperature and fuel mass fraction for flame-E, flame-F were 14.69% and 5.22%, 14.18% and 5.54%, respectively. Furthermore, AECSC-IBM software was applied to a single head combustor of a real structure, and the root mean square error of the simulated exit temperature was 11.66% compared with the experimental data. Example tests showed that AECSC-IBM software can map geometric models quickly and accurately, reduce the workload of complex geometry high-quality mesh generation greatly, and simulate the two-phase turbulent combustion phenomenon in the aero-engine combustion chamber efficiently and accurately. The simulation results can provide a reference for combustion field data in combustion chamber refinement research and development,presenting practical engineering value.

To precisely design multiphase fuel nozzles, so that they can be applied to the aero-engines with the technology of cooling the turbine cooling air, the fuel mass flow rate of sub-/supercritical kerosene in an axisymmetric model nozzle was studied. By adjusting the fuel temperature under multiple sets of fixed injection pressure, the mass flow rate change of sub-/supercritical fuel was obtained. According to the phase conditions of kerosene, the flow rate curves were divided into three phase regimes, i.e., liquid, gas-liquid, and supercritical, and the mechanisms in the process of flow rate changing with kerosene injection status under different phase conditions were discussed. The flow rate prediction methods in the liquid and the supercritical regime, as well as the fitting correlation of the discharge coefficient in the gas-liquid regime, were proposed. Results showed that, as injection temperature increased from normal temperature to 750 K, the flow rates slowly decreased in the liquid regime, sharply declined in the gas-liquid regime, and then gradually declined in the supercritical regime beyond the “inflection point” temperature. The maximum error between the calculated and experimental results was 3.8%, and the square of the correlation coefficient of the fitting correlation was 0.9478, which can serve multiphase nozzle design. Meanwhile, the acquired mass flow rate data can provide the boundary conditions for the research of injection structures downstream the nozzles, supporting advanced aero-engine design and development.

In order to investigate the influence of pilot stage swirl intensity on ignition performance, ignition tests using a centrally-staged combustor equipped with three different pilot stages with different inner/outer swirl numbers were carried out at atmospheric temperature and sub-and atmospheric pressures, and the ignition boundaries were obtained. The ignition test results showed that the ignition performance of inner weak swirl and outer strong swirl pilot stage had the best ignition performance, and increasing the inner swirl intensity was not conducive to ignition. The spatial distributions of the fuel spray at the outlet of the three configurations of pilot stage were measured by using the Mie scattering technology. The measurement results showed that increasing the inner swirl intensity can increase the spray angle, and reducing the outer swirl intensity can reduce the spray angle. The pilot stage with inner weak swirl and outer strong swirl led to proper fuel distribution near the central axis of the combustor and near the ignitor, thereby improving the ignition performance. The results are useful for improving ignition performance of centrally-staged combustors.

For the model combustor with a single dome, large eddy simulation was carried out to simulate the dynamic propagation process of flames at different ignition timing. In a pulsation cycle of the flow field, the effect of spark phases on ignition reliability was studied. The simulation results revealed that for the pulsating flow field in the model combustor, even within the range of ignition limits, the single spark pulse can not cause successful ignition at any spark phase. By comparing the frequency of the flow field with the characteristics of the single spark pulse, it was found that the duration of the single spark pulse was much shorter than the pulsation period of the flow field. This is the fundamental reason why spark phases seriously affect the ignition reliability. In order to eliminate the effect of spark phases on ignition reliability, it is suggested that a single spark pulse should be replaced by a series of spark pulses when using large eddy simulation to study ignition problems. And the duration of spark pulses should cover at least one complete pulsation cycle of the flow field.