To investigate the laws of surface ice accretion and ice shedding characteristics of transonic rotor blades under different conditions, and to obtain the ice shedding position and cross-sectional structure, study takes transonic fan rotor blades as the research object. Based on an ice wind tunnel system with natural low temperature, test studies on ice accretion and shedding of rotor blades were carried out under typical icing conditions. The results show that ice accretion occurs on both the suction surf-ace and pressure surface of the transonic fan rotor blades, and ice shedding starts simultaneously from both surfaces of the blades. As the ambient temperature decreases, the ice shape on the blade surface changes from glaze ice to mixed ice, and then to rime ice. Ice shedding is affected by blade rotational speed, rotational acceleration rate, and ambient temperature; the time required for ice shedding and the degree of ice shedding vary under different conditions. When the ambient temperature is constant, as the ice-accreting rotational speed of the blade increases, the ice-shedding rotational speed increases monotonically, while the ice-shedding time decreases monotonically. Meanwhile, the ice shedding amount on both the pressure surface and suction surface of the blade increases with the increase of blade rotational speed. When ice shedding is conducted by increasing the rotational acceleration rate, as the rotational acceleration rate increases, the degree of ice shedding on the blade surface increases monotonically, while the ice shedding time decreases monotonically. For ice shedding under different temperatures, as the temperature decreases, the degree of ice shedding decreases, and the time requirde for ice to shed from the blade surface is extended. Within a certain range, the prediction model for ice shedding rotational speed established based on test data canaccurately predict the ice shedding rotational speed under different operating conditions.

The flight envelope of Air Turbo Rocket (ATR) engines is broader compared to conventional aero-engines, and designing a full-envelope controller using traditional methods poses significant challenges due to a burdensome design process. To address this issue, this paper proposes a self-learning neural network controller design method based on Reinforcement Learning (RL). Additionally, a method for predictive compensation and stability rewards is proposed to reduce the system oscillation caused by actuator delay. This approach simplifies the actuator to a first-order inertial element exhibiting pure delay. A simulation environment for the ATR engine-actuator system is first established. Based on this environment, a self-learning neural network controller using a predictive compensator and the Proximal Policy Optimization (PPO) algorithm is then developed. Furthermore, the temporal difference signals from the controller output are integrated into the reward function to enhance system stability. The proposed method is validated through numerical simulations and semi-physical experiments. The numerical simulation results demonstrate that the proposed method increases the system’s tolerance to delays from 20 ms to 400 ms. Under an actuator delay of 400 ms, the average steady-state error remains less than 0.1%, the overshoot is limited to 1%, and the settling time does not exceed 3 s. Moreover, compared to the traditional method, the proposed method exhibits higher adaptability to model errors and variations in flight conditions. In the conducted semi-physical simulation experiments, the proposed method achieves stable control of a real electric pump.

To investigate the impact of wake/shock on the aerodynamic load distribution of downstream turbine blades in a multi-stage vaneless counter-rotating turbine, unsteady numerical simulation was utilized to analyze the influence of complex inter-stage flow on pressure fluctuation on the downstream blade surfaces. The study reveals that in the multi-stage vaneless counter-rotating turbine, all guide vanes are eliminated, resulting in suction side of the first-stage moving blade with an extended uncovered section. The extended uncovered suction side forms a converging-diverging wake flow passage resembling a Laval nozzle in interaction with the wake from adjacent blades. Under the unsteady interaction between the stages, the exit Mach number of the first-stage moving blade decreases, leading to the formation of a new compression wave (wake flow passage shock) near the trailing edge of the suction side of the first-stage moving blade superimposed on the existing suction side trailing edge shock. At specific moments within one cycle, both the reflected shock and the wake flow passage shock act on the 28.8% axial position on the suction side of the second-stage moving blade leading to a significant increase in pressure fluctuation peak value at that location. The peak value reaches 81.2% of the peak value induced by the suction side trailing edge shock sweeping, indicating that the suction side trailing edge shock is the primary factor causing pressure load variations on the blade surface of the second-stage moving blade. The main perturbation region for pressure fluctuation induced by the suction side trailing edge shock wave is the leading edge of the second-stage moving blade, with the maximum peak value of pressure fluctuation in this region within one cycle reaching 47.7% of the mean pressure. Due to the dissipation effects of the reflected shock of pressure side trailing edge shock and wake flow passage shock, the wake strength of the first-stage moving blade significantly decreases, resulting in a minor impact on the aerodynamic load distribution on the blade surface of the second-stage moving blade. Frequency analysis results indicate that due to the combined effects of the wake flow passage shock and the reflected shock of pressure side trailing edge shock, the main frequency of pressure fluctuation on the blade surface of the second-stage moving blade is twice the sweeping frequency of the suction side trailing edge shock.

Blade vibration monitoring can ensure the safe operation of aero-engine rotor blades. Among the methods of blade vibration monitoring, Blade Tip Timing (BTT) method has attracted more and more attention because of its advantages of non-contact measurement. However, it is difficult to install the Once-Per-Revolution (OPR) probe in the confined space of aero-engine, and the failure and instability of the OPR signal will reduce the reliability of the blade vibration analysis results, which directly affects the accuracy of the blade vibration parameters identification. The Multi-Probe linear fitting and Time of Arrival (ToA) Linear Correction method based on the BTT (MP-LC-BTT) without OPR is proposed to reduce the errors of single probe linear fitting method for blade vibration displacement analysis. The proposed method can also correct the calculation error of blade vibration displacement due to the nonlinear change of rotation speed, which can improve the analysis accuracy of the blade vibration displacement. A new blade vibration model conforming to the actual vibration characteristics is established, and the effectiveness of the proposed method is verified by numerical simulation. Finally, the reliability and accuracy of the MP-LC-BTT method have been verified by the experiments which include two high-speed blade test-benches and an industrial axial fan. This method can be used in the actual aero-engine monitoring instead of the BTT method with OPR.

Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow, which always accompanies aerodynamic performance penalties. A loss reduction method for smearing the passage shock foot via Shock Control Bump (SCB) located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation. The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model. Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes (RANS) method to reveal flow physics of SCB. Comparisons and analyses on simulation results have also been carried out, showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer. It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100% design rotor speed, and then slightly worsens with decrease of rotatory speed till both equal below 60% rated speed. The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.