In order to meet the requirements of low-flow intake anti-icing test of domestic aero-engines, based on the icing wind tunnel engine intake simulation system, the air intake characteristics of typical low-flow engine intake components were analyzed, and a small-flow engine intake simulation method was developed. A water ring vacuum pump was proposed to solve the problem that the pressure of the suction pipe was greatly reduced due to the large flow resistance of the intake. On this basis, the engine intake simulation system was optimized, and the process of conducting the anti-icing test of the low-flow engine intake components in the icing wind tunnel was established, which was applied to the anti-icing test of a certain type of low-flow engine intake components. The results showed that the engine intake simulation system met the requirements of the anti-icing test of the low-flow engine intake components, and the control accuracy of the intake flow reached ±0.05 kg/s(±0.33%FS). The flow change during the test can be used to identify the anti-icing effect. The relevant research can provide a reference for the design and optimization of the engine intake simulation system and the assessment of the anti-icing characteristics of the low-flow engine intake components.

Considering the shortcomings in the design and installation of the reheat gas-driven distributed propulsion system, a partial turbine-electric distributed propulsion system was proposed by combining with the turbine-electric driven method. A design point calculation model was established based on the component model. The energy flow mechanism of the propulsion system was analyzed, and the design method of key energy transfer parameters was proposed. Based on the research, the influence of design parameters on the propulsion system was analyzed, and the performance and design parameters of different distributed propulsion systems were compared and analyzed. The results showed that the fuel consumption of the partial turbo-electric distributed propulsion system was more sensitive to the turbine inlet temperature, and the total pressure ratio had less influence. Compared with the gas-driven distributed propulsion system, the partial turbo-electric distributed propulsion system had a fuel consumption advantage of 1.7%, and when the power ratio was reasonably selected, it could improve the gas-driven distributed propulsion system. Focusing on the fuel consumption of the propulsion system, the performance suitability of the partial turbo-electric distributed propulsion system based on the gas-driven propulsion system was demonstrated.

In order to realize the movement of the micro flapping-wing aircrafts’ wings along the complex trajectory, a flap-sweep multi-degree-of-freedom flapping-wing driving mechanism was designed. In view of the problem that the inertial force and elastic deformation of the mechanism’s transmission components affect the actuators’ driving force during the high-frequency motion, a rigid-flexible coupling dynamic model of the mechanism was established. At the same time, a dynamic performance factor that can quantify the difference between the driving force required by the actuators and its ideal value was proposed. Finally, combined with the orthogonal experiment, the influence of each thin-plate-component’s thickness on the dynamic performance of the mechanism was studied. The research results showed that the thin-plate-component subjecting to the out-of-plane load had a great influence on the dynamic performance of the driving mechanism, and the longer transmission chain’s force-transmission-path and the closer position to the actuators indicated the more significant impact; the flapping motion performance of the driving mechanism was stronger than the sweeping motion performance; In addition, the dynamic performance of the driving mechanism was not positively related to the thin-plate-component’s thickness.

An efficient aerodynamic prediction method applicable for contra-rotating propellers in axial flight was developed based on Blade Element Momentum Theory. Firstly, the accurate decomposition of aerodynamic forces and Prandtl tip loss were employed to eliminate errors introduced by the small angle assumptions at high inflow angles. Secondly, based on the helical tip vortex evolution of the contra-rotating system, a wake superposition model was developed, and the interaction pattern was built, which took into account the axial aerodynamic interactions and circumferential aerodynamic interference that the front propeller exerted on the rear propeller. Finally, the aerodynamic model of contra-rotating propellers in axial flight was established based on inflow angle solutions. The method was applied to predict the aerodynamic performance variation with advance ratio for a contra-rotating propeller under different flight speeds. The calculated thrust and power in non-stall conditions were consistent with the measurements, and the whole propulsive efficiency agreed well with measurements. Aerodynamic prediction comparisons of the single propeller, the coaxial rotor and the contra-rotating propeller revealed that the established model can provide more reasonable performance predictions including thrusts, powers and efficiencies for contra-rotating propellers compared with conventional BEMT models, which assume small angles, neglect some of the interactions in contra-rotating system and rely on inflow velocity solutions.

A fusion algorithm of multi design points (FAMDP) was proposed for the design of turboshaft-turbofan variable cycle engine (TSFVCE) with two operating modes, including: turboshaft mode and turbofan mode. A turbojet engine was taken as an example to introduce the establishment of a FAMDP model. The FAMDP model for the TSFVCE was established. The typical operating points of the TSFVCE, including high-altitude cruise turbofan mode operating points, ground turboshaft mode oper-ating points, and modal conversion operating points, were determined, and the design of the engine scheme was completed. The performance requirements of the aircraft for the TSFVCE at each typical operating point were used as input for the FAMDP. Compared with the parameters obtained by conventional methods, the results showed a difference of less than 1%. The fusion algorithm of multi design points proposed can provide a reference for the overall design of variable cycle engines.

In order to explore a better derived configuration of the aero-spike for drag and heat reduction of the supersonic blunt aircraft, numerical simulation was used to study the flow characteristics of three kinds of typical configurations with single flow interferent, six kinds of configurations with double flow interferents, two types of configurations with multiple flow interferents and blunted cone-spike; it was believed that the geometric essence of adding an aero-spike was equivalent to a “hollowed-out” conical blunt body. The simulation results showed that when the relative diameter of the flow interferent at the head was large, the variation curve of drag reduction rate with the relative length of the spike didn’t have an obvious peak point, instead, there was a peak stage with very little variation, and the drag reduction effect was best when the relative diameter was about 0.3−0.4; the maximum drag reduction rates of the typical spikes were about 50% and slightly increased by using configurations with double flow interferents; when multiple flow interferents were added in the middle, the drag reduction rates increased with the growing number of flow interferents, the maximums were more than 60%, but the aerodynamic heating problem was more serious. In comparison, the blunted cone-spike had the best comprehensive performances of drag and heat reduction, its maximum drag reduction rate could reach about 60%, and the heat reduction rate was about 7%.

To facilitate the large-scale exploration of design space and the design of aero-structural integration in the conceptual design stage, a three-dimensional parametric modeling and optimization method for wing aero-structural analysis was proposed by using the class-shape function transformation (CST) method. Based on the two-dimensional CST, the analytical function form of the three-dimensional CST parametric geometric model was deduced. Established by using the mesh adaptive discrete and structural feature extraction technology, the aero-structural analysis parametric model of the three-dimensional wing can support the rapid modeling and optimization of the aerodynamic structure integration simultaneously, including parameters of the wing geometry configuration, structural layout, structural size, material properties, etc. Besides, it can automate the modeling process for the geometric models with a wide range of parameterization and aero-structural models. By applying this method, the aerodynamic structure of a large scale wing can be optimized from a multidisciplinary perspective. This method was adopted to work out the design of optimal aerodynamic structure integration for a large aspect ratio wing. According to the optimization result, the number of beams was reduced from 2 to 1, the number of wing ribs reduced from 20 to 15, and the mass reduced by 26.1%.

To accomplish the trade-off design of high-altitude propellers with high efficiency and light mass, a multi-disciplinary and multi-objective optimization design method was proposed by considering both aero-structure of the propeller. Theoretically, Pareto solution set with the objective of maximum thrust and minimum mass can be obtained. However, in practical engineering applications, due to large amount of optimization variables, only the fitting trend of the Pareto solution set was obtained in acceptable time. To avoid overlong optimization period, the staged optimization approach was proposed. Stage 1: the optimal propeller diameter was decided by the Pareto solution set fitting trend and constraints; Stage 2: the aerodynamic shape was obtained by aerodynamic optimization based on the optimal propeller diameter; the structural scheme was obtained through structural optimization. This approach was used to optimize the propeller for a solar-powered unmanned aerial vehicle, the two stages took 96 h and 4 h, respectively, the propeller was manufactured, simulated and tested. The comparison results showed that the maximum thrust error was 10.9%, the mass error was 6.9%, the stiffness error was 15.2%, and the natural frequency error was 15.4%; so the test results also demonstrated its rationality and validity.

The multi-field coupling instantaneous evolution law of laser quenching process was quantitatively revealed, then the significance analysis of 40Cr laser quenching process parameters was carried out. Based on the calculation of phase diagram (CALPHAD) method, the temperature changes physical properties of the material were calculated, and the coupled numerical model of 40Cr gear steel during laser quenching was established. Numerical calculations of the transient temperature, phase transition, and stress distribution of 40Cr laser quenching revealed the interaction mechanism between phase transition behavior and plastic stress. Axio Vert.A1 microscope, scanning electron microscope (SEM), super depth of field 3D microscope, and microhardness tester were used for analysis. The significant effects of laser radius, laser power, and scanning speed on quenching quality were analyzed based on orthogonal experiments. The results showed that the significant process parameters affecting the maximum temperature and phase transformation depth were spot diameter, scanning speed and laser power. The distribution of residual stress was “hump”, and the significant process parameters affecting the residual tensile stress were spot diameter, laser power and scanning speed. This research provides an essential theoretical basis for effectively controlling residual quenching stress and optimizing quenching process parameters.

Two brush seal friction coefficient experimental identification models were established based on cylindrical circumferential friction, direct force measurement and indirect torque method, an experimental device for friction and wear characteristics of brush seals was designed and built, and eight brush seal test pieces with different structural parameters were designed and processed. The experimental results of two kinds of friction coefficient identification models were compared and analyzed, and the influence of structural parameters on the positive pressure of the brush bristle and rotor surface, the influence of working conditions on the friction coefficient of the brush bristle, and the influence of interference on the wear of the brush bristle were studied. The research results showed that the friction coefficients of brush bristle measured by the direct force measurement method and the indirect torque method were less different from each other, and the stability of the direct force measurement method was better than that of the indirect torque method. Both the static and dynamic positive pressures on the surface of the brush seal and the rotor increased with the increase of the protection height of the baffle, decreased with the increase of the radial length of the brush, and increased with the increase of the thickness of the brush pack; at the same interference under the condition of measurement, the static positive pressure was greater than the dynamic positive pressure; the positive pressure of the brush seal filament and the rotor surface was greater in the interference loading phase than in the interference unloading phase, and the speed increase phase was greater than the speed reduction phase, the brush filament exhibited a hysteresis effect. The friction coefficient of the brush seal bristle decreased with the increase of the interference between the brush bristle and the rotor, and also decreased with the increase of the rotor speed; with the increase of the friction time, the friction coefficient decreased rapidly at the initial stage of wear, and then kept basically stable. The wear amount of the brush bristle of the brush seal increased with the increase of the interference between the brush bristle and the rotor. When the interference between the brush bristle and the rotor increased from 0.3 mm to 0.4 mm, the wear amount of the brush bristle increased by 296.66%.

In order to consider the influence of stress gradient on the fatigue strength of SiCp/Al composite structure, a notched fatigue strength prediction method of SiCp/Al composite was developed based on the three-dimensional space vector stress field intensity method (TSVFM) and the fatigue strength of smooth specimen. In the TSVFM, the equivalent stress integral forms of classic one-dimensional stress field intensity method and effective distance point stress field intensity method were applied respectively, which avoided the construction of three-dimensional weight function and artificial determination of fatigue damage region. The fatigue test scheme of SiCp/2009Al composite smooth specimen was determined by up-and-down method, and the axial (R=−1) fatigue strength corresponding to 10⁷ cycles of SiCp/2009Al composite was 180.91 MPa. The fatigue life distribution of SiCp/2009Al was obtained by the scattered point method. The fatigue test results of smooth specimen showed that there was an obvious platform area in the stress-life relationship of SiCp/2009Al composite. The axial (R=−1) fatigue test of notched specimen of SiCp/2009Al composite was carried out by step-by-step loading method, and the fatigue strength of notched specimen was 82.2 MPa. The fatigue strength prediction results of notched specimen were in good agreement with the test results, and the maximum error was within 10%. The notched fatigue strength prediction method of SiCp/Al composites based on the ETSVFM was better than the CTSVFM.

Based on the extended Kalman filter (EKF), the self-calibration extended Kalman filter (SEKF) and the multiple-model estimation (MME), and considering the influence of unknown inputs (such as gusts, faults, unknown system errors, etc.) on the nonlinear system state equation in Engineering, the multiple-model self-calibration extended Kalman filter (MSEKF) was proposed to expand the application scope of the multiple-model self-calibration Kalman filter (MSKF). According to the Bayes’ theorem, this filtering method used the EKF and the SEKF whose weights were assigned automatically to obtain the final filtering result through weight-average way. The MSEKF can not only effectively compensate the effects of non-zero unknown inputs on the nonlinear system, but also improve the estimation accuracy compared with the SEKF when these effects were zero. A large number of simulation results using the proposed method showed that the accuracy can be improved by more than 4%, showing stronger adaptability and robustness.

In order to identify the characteristics of a turboshaft engine during surge, the bench test of a turboshaft engine was carried out by inlet distortion method, and the sound pressure signals on both sides of axial compressor and centrifugal compressor were measured and collected by sound pressure sensor. The test environment and background noise of sound pressure signal were corrected, the characteristics of compressor blade stall mass and low-frequency surge caused by intake reduction were detected by time-frequency analysis method, and the surge signal of a turboshaft engine was extracted and identified by wavelet low-frequency reconstruction of sound pressure signal. The results showed that with the increase of intake air, the amplitude of sound pressure signal at the rotor frequency of axial flow compressor and centrifugal compressor decreased, and stall mass was generated. Surge can be first detected on the right side of axial flow compressor, with the surge frequency about 60 Hz.

In view of vulnerability to cracks and pore defects due to lacking of defect detection probability data in radiographic examination of additive manufacturing, the linear defects and pore defects of GH3625 superalloy additive parts were researched. The CIVA2020 simulation platform was used to simulate X-ray inspection and obtain the probability of detection (POD) curve of defects, research the influence of different size changes of the two types of defects on the POD of defects. The detectable size of defects and POD under different influencing factors were determined, and the POD curve equation of linear defect affected by depth and the POD curve equation of pore defect affected by radius were fitted by Sgompertz function, the defect detection probability model of additive manufacturing line defects and pore defects was established respectively. The detectable length size of linear defects was 0.211 mm, the detectable width size was 0.213 mm, the detectable depth size was 0.178 mm at a probability of 90% under 95% confidence level, the detectable diameter size of pore defects was 0.188 mm, and the detectable height size was 0.190 mm. The simulation results were compared and verified by micro-focus radiography and film radiography of actual specimens. The research results showed that the defect detection probability model is more accurate and can provide a basis for reliability analysis of crack and pore defect detection in additive manufacturing.

In view of the problem of determining the modal number and quadratic penalty factor of variational mode decomposition (VMD), a parameter optimization algorithm based on orthogonality index, energy ratio and variational energy entropy (VEE) was proposed. For the decomposed single component signal, the instantaneous frequency identification method based on polynomial chirplet transform (PCT) and the instantaneous damping ratio identification method based on energy method were developed. The simulation research of 3-DOF (degree of freedom) time-varying structure and the experimental research of time-varying steel beam were carried out. The results showed that the optimized VMD method can accurately separate the time-varying components of the multi-DOF system with strong anti-noise performance. The instantaneous frequency identification method based on PCT had strong time-varying frequency tracking performance, strong anti-noise ability, and high accuracy of time-varying frequency identification, and the average error was less than 1%. The energy method can accurately identify the instantaneous damping ratio of the structure with obvious anti-noise advantage, and the identification error was maintained at about 10%.

Based on the analysis of the development status of aviation heavy fuel piston engines at home and abroad, the core technical indexes of engines with different power levels were obtained by comparing and analyzing the existing mature models, and the important significance of developing aviation heavy fuel piston engines in China was put forward. According to the power level, the technical path of the development of aviation heavy fuel piston engine was analyzed, the technical difficulties faced by the heavy fuel engine with high power-to-weight ratio were summarized, and the key technologies of the spark-ignition aviation heavy fuel piston engine, namely, the rapid atomization technology, the rapid cold start technology, the knock suppression technology and the high-altitude supercharging power recovery technology of heavy fuel, were presented. A breakthrough over the above key technologies will be of great significance to the development of aviation heavy fuel piston engine technology.

A simulation model of the solid oxide fuel cells-gas turbine (SOFC-GT) hybrid power system on the basis of direct ammonia fuel was established, to developed an efficient power generation system with a high power-mass ratio optimized by architecture, and studied the effects of fuel utilization and system fuel allocation on system power allocation, mass of various subcomponents, and energy losses. Then, the performance of the established SOFC-GT hybrid system was evaluated by changing parameters such as the compressor pressure ratio, fuel flow rate and air flow rate. The power-mass ratio analysis of the system was also carried out under the optimal performance condition. The simulation results showed that, the net power generation efficiency of the system can reach 56.85%, and the exergy efficiency can reach 50.71% at the design conditions. Meanwhile, the net power generation and the power-mass ratio reached 213 kW and 0.7303 kW/kg, respectively. So, this result can meet the power-mass ratio standard given by the Pacific Northwest National Laboratory (PNNL) for the SOFC-GT hybrid system used in the aerospace field. Finally, the application of the system on commercial aircraft as both main power system and auxiliary power unit was discussed, and the designed SOFC-GT hybrid system showed good aviation application prospects.

The ignition delay time, the laminar combustion speed of natural gas and the main species concentrations in the natural gas oxidation process were simulated by six detailed reaction mechanisms, and also compared with corresponding experimental data. The results showed that, compared with the other five detailed reaction mechanisms, Aramco 2.0 mechanism had the highest accuracy in predicting the ignition delay, laminar combustion and oxidation characteristics of natural gas. Based on Aramco 2.0 mechanism, an initial reduced reaction mechanism (including 21 species and 150 reactions) of natural gas (including CH₄, C₂H₆ and C₃H₈) was formed using the path sensitivity analysis, production rate analysis and reaction path analysis. At the same time, based on decoupling method, through combining the reaction mechanism of C₁−C₃ in the initial reduced reaction mechanism with the detailed reaction mechanism of H₂/CO and the reduced reaction mechanism of NOx, the reduced reaction mechanism of natural gas was constructed (including 40 species and 189 reactions). Compared with the corresponding experimental data, it was found that the reduced reaction mechanism can well predict the ignition delay, laminar combustion and oxidation characteristics of natural gas under various conditions.

To explore the impact of endwall movement on the performance of cantilevered stators, the effects of hub movement, hub movement speed and the gap size of cantilevered stators on the internal flow and performance of a low-speed axial compressor were researched in detail. The results showed the flow blockage/loss of hub area was reduced and the compressor performance was improved with the increase of the endwall movement speed or the decrease of the gap size in case of endwall movement. This phenomenon was mainly attributable to the disappearance of the secondary flow and the reduction of the leakage flow involved in the mixing. The endwall movement also affected the inlet flow field of the cantilevered stators, which in turn affected the characteristics of the rotor. However, compared with the effect on the characteristics of the stator, the change amplitude of the rotor characteristics was relatively small.

Taking the measures parameters of the engine as the initial characteristics, the MLP （multilayer perceptron） model of aero-engine compressor disk stress and temperature prediction was established by using artificial neural network technology, and BP （back propagation） neural network algorithm was used for training. The results showed that the prediction results of this method were in good agreement with the traditional finite element calculation results. The relative deviations were all within 1%, the determination coefficients were above 0.95, and the root mean squared error was within 5. Moreover, the calculation speed increased from hour level to minute second level, providing a basis for subsequent engineering applications.

Simulations were carried out to investigate the difference of film cooling effectiveness distribution at different blade heights under different blowing ratios and rotational Reynolds numbers. Five cylindrical holes with a diameter of 0.8 mm were located at 17.8% streamwise location and at 10%, 30%, 50%, 70% and 90% blade height, respectively. The study was conducted under three rotational speeds of 400 , 600 r/min and 800 r/min, correspondingly to rotational Reynolds numbers of 357 000, 536 000 and 715 000, respectively. Five blowing ratios of 0.50, 0.75, 1.00, 1.25 and 1.50 were involved. Results showed that film trajectories in the near hub region were deflected a lot toward the tip under the effect of passage vortex. The distribution of film cooling effectiveness at different blade heights was different. Under the effect of centrifugal force and Coriolis force caused by rotation, the increase of blow ratio and rotating Reynolds number yielded different effects on film deflection at different blade heights. The effect of rotational Reynolds number on film cooling at different blade heights was different.

In order to solve the difficulty of determining the key heat transfer parameters in the design of rocket pressurization system, a heat transfer parameter identification method based on adaptive chaotic particle swarm optimization (ACPSO) algorithm was proposed. The mathematical model of the pressurization system considering the heat transfer terms was established. The parameters to be identified were selected by sensitivity analysis, and then identified by particle swarm optimization method with local minimum prevention and adaptive weight strategy. The root mean square function with weight decay was optimized. The results showed that the identified simulation value was in good agreement with the measured value. The deviation between the simulation and the measured value of the opening time of the pressurization electric valve of the oxygen tank was less than 3%, and the deviation between the simulation value and the measured value of the temperature at the end of a flight of the gas bottle was only 2.4 K. If the identification results were applied to a similar newly developed pressurization system, the volume and weight of the gas bottle was reduced by 32 L and 11.6 kg compared with the design under the adiabatic assumption, which effectively saved cost caused by redundant design.

The Helmholtz equation with the mean flow source term was solved by the acoustic finite element method. Then, on the basis of considering the high temperature and average flow field of the combustion chamber, the influence of baffle structure parameters on the acoustic mode characteristics of the liquid rocket engine combustion chamber was analyzed. The results showed that: increasing the number or length of baffle reduced the eigenfrequency of the first-order tangential mode of the combustion chamber; when the number of baffle was 4, the damping rate of the first-order tangential mode of the combustion chamber was greatest; the longer length of the baffle indicated the smaller distribution area of the first-order tangential mode acoustic pressure and the larger damping rate; the type of baffle had an insignificant effect on the first-order tangential mode characteristics of the combustion chamber.

In order to avoid propellant gasification, cryogenic liquid rocket engine needs pipeline precooling during propellant filling. To reveal the two-phase flow characteristics of cryogenic fluid during pipeline precooling, study was carried out for the precooling process of liquid methane pipeline in a small liquid oxygen/methane engine. The turbulent heat transfer process under different inlet flow rates was simulated and analyzed by Lee evaporation model, and the variation rules of methane volume fraction, temperature, pressure and velocity during the pipeline precooling process were obtained. The results showed that the flash of liquid methane occurred during the precooling process, and the change of methane temperature and pressure was the main factor affecting the flash. At low flow rate, the precooling time decreased with the increase of mass flow rate; when the mass flow rate increased to a certain extent, the precooling time was prone to be stable. The results can predict the optimal precooling flow rate in the allowable time, which has a guiding role in improving the precooling efficiency and the filling process of cryogenic propellant.

For the conformal contact pairs such as the roller-circular surface of the pocket, the roller-end face of the pocket, and the roller-rib with highly close contact profiles, based on the Winkler's elastic foundation model, a conformal contact modeling method was proposed; by combining the conformal contact modeling method with the Newton-Euler's dynamics theory, the six degree-of-freedom dynamics modeling method of the bearing with circular pockets considering conformal contact was proposed; and by comparing with both the reference results and the experiment data, the proposed dynamics modeling method was validated. On this basis, influences of both the mass eccentricity of the cage and the angle misalignment of the inner and outer rings on the cage slip/whirl, the roller slip/skew/tilt, the roller-cage collision and the roller-rib collision were studied by use of dynamics simulation. The results showed that the radial mass eccentricity of the cage could aggravate the cage slip and the roller slip, but contributed to improve the cage whirl stability, and the axial mass eccentricity may cause the roller skew/tilt; when the axial mass eccentricity was combined with the radial mass eccentricity, the roller-pocket peripheral collision and the roller-rib peripheral collision easily appeared. Secondly, the angle misalignment of the inner and outer rings was more likely to cause the roller skew/tilt, and the roller-pocket peripheral collision and the roller-rib peripheral collision, but it can significantly reduce the cage slip and the roller slip.

Based on Hertz elastic contact theory and raceless control theory, a quasi-static analysis model of angular contact ball bearings with five degrees of freedom under different preloading mechanisms was established. Newton Raphson method was used to numerically solve the established model for exploring the influences of preloading mechanisms and interference amount on the dynamic characteristics of angular contact ball bearings. The effectiveness of the model was verified by comparing the results of contact angle and bearing stiffness of the established model with those of existing literature. The influences of the interference amount on the assembly pressure of the inner and outer rings and the contact deformation of the raceway as well as the influences of the interference amount and the preloading mechanisms on the bearing stiffness were explored. The results showed that with the increase of interference, the assembly pressure of inner and outer rings decreased and the initial contact angle decreased; with the increase of interference, the radial stiffness of the bearing increased, while the axial stiffness and angular stiffness decreased; compared with constant pressure preloading, positioning preloading can improve the bearing stiffness; ghe influence of rotation speed on contact deformation and assembly pressure was investigated, and the release rotation speed was determined to be about 40000 r/min.

With strong nonlinearity, high uncertainty and strong external disturbance for flight altitude simulation of altitude test facility, a robust cascade linear active disturbance rejection control (LADRC) algorithm via the reduced order linear extended state observer (RLESO) was proposed. In particular, the main characteristics and control issues of the controlled plant were analyzed, and the generalized controlled plant can be included by the butterfly valve position loop and flight altitude loop. The RLESO and corresponding controller were designed for the two loops respectively, meanwhile the cascade control system was constructed. The robust cascade LADRC was realized in the simulation environment and compared with the classical PID control scheme. When the engine was subject to the thrust transient test, the maximum fluctuation value of the controlled pressure was reduced from 3.5 kPa to 0.8 kPa. This indicated that the robust cascade LADRC technology via RLESO can significantly improve the dynamic control quality. The ideal robust control performance and anti-disturbance ability for flight altitude simulation were obtained.

In order to reduce the impact of uncertain factors in manufacturing on the overall performance of aero-engines, a robust design model was constructed by considering the uncertainty of high-pressure turbine performance in the design stage. Monte Carlo simulation was used in this model to quantify the effect of uncertainty. A global optimization algorithm was developed to solve the robust design model. Numerical results validated the advantages of the robust design model. Its optimal solution can reduce the variability of the engine performance by an average of 15.97%, which showed better robustness in uncertain environment than the solution obtained by generic deterministic design method.