Radial flow variable nozzle turbine (VNT) enables better matching between a turbocharger and engine and can improve the engine performance as well as decrease the engine emissions, especially when the engine works at low-end operation points. With increased nozzle loading, stronger shock wave and clearance leakage flow may be generated and consequently introduces strong rotor–stator interaction between turbine nozzle and rotor, which is a key concern of rotor high-cycle fatigue (HCF) failure. With the purpose of developing a low shock wave intensity turbine nozzle, the influence of grooved vane on the shock wave characteristics is investigated in the present paper. A Schlieren visualization experiment was first carried out on a linear turbine nozzle with smooth surface and the behavior of the shock wave was studied. Numerical simulations were also performed on the turbine nozzle. Guided by the visualization and numerical simulation, grooves were designed on the nozzle surface where the shock wave was originated and numerical simulations were performed to investigate the influence of grooves on the shock wave characteristics. Results indicate that for a smooth nozzle configuration, the intensity of the shock wave increases as the expansion ratios increase, while the onset position is shifted downstream to the nozzle trailing edge. For a nozzle configuration with grooved vane, the position of the shock wave onset is shifted upstream compared to the one with a smooth surface configuration, and the intensity of the shock wave and the static pressure (Ps) distortion at the nozzle vane exit plane are significantly depressed.

The present paper numerically and experimentally investigates the stall inception mechanisms in a centrifugal compressor with volute. Current studies about stall inception pay more attention on the axial compressors than the centrifugal compressors; especially, the circumferential position of stall inception onset and the stall process in the centrifugal compressor with asymmetric volute structure have not been studied sufficiently yet. In this work, the compressor performance experiment was conducted and the casing wall static pressure distributions were obtained by 72 static pressure sensors first. Then, the full annular unsteady simulations were carried out at different stable operating points, and the time-averaged static pressure distributions were compared with the experimental results. Finally, the stall process of the compressor was investigated by unsteady simulations in detail. Results show that the stall inception onset is determined by the impeller leading edge (LE) spillage flow, and the occurrence time of trailing edge (TE) backflow is prior to the LE spillage. The nonuniform static pressure circumferential distribution at impeller outlet induced by volute tongue causes the two stall inception regions locating at certain circumferential positions, which are 120 deg and 300 deg circumferential positions at impeller LE, corresponding to the circumferential static pressure peak (PP) and bulge regions at impeller outlet, respectively. In detail, at rotor revolution 2.86, a small disturbance that the incoming/tip clearance flow interface is perpendicular to axial direction occurs at 120 deg position, but this disturbance did not cause the compressor stall. Then at revolution 7, the first stall inception zone (spillage region) occurs at 120 deg position, causing the compressor stall with positive pressure ratio performance. At approximately revolution 23, the second stall inception zone occurs at about 300 deg position; however, both the intensity and size of this stall inception zone are smaller than those of the first stall inception zone. These two stall inception zones are not moving along circumferential direction because the stall inception circumferential position is dominated by the impeller outlet static pressure distribution. Even then, the obvious low frequency signals appear after the spillage crossing two blade LEs, because at this moment, the spillage vortex caused by the tip leakage flow begins to shed. However, due to the asymmetric structure limitation, this vortex cannot move across full annular. Furthermore, the spillage vortexes cause the local low static pressure zone ahead of blade LE in the centrifugal compressor with volute, suggesting that the spillage can be predicted by the steady casing wall static pressure measuring. The development of blockage zones at impeller LE is also investigated quantitatively by analyzing the stall blockage effect.

The turbines used in turbochargers naturally experience unsteadiness caused by inlet pulsating flow conditions and stator–rotor interaction. The unsteadiness has an influence on turbine performance. Meanwhile, under certain small-nozzle opening conditions, strong shock waves can be generated. The synergistic effect of turbine inlet pulsation and shock waves has a significant influence on the turbine performance, rotor blade loading as well as the excitation force exerted on the turbine rotor, which is responsible for turbine rotor high cycle fatigue. In order to understand the influence of pulsating flows on turbine performance and the shock wave characteristic at nozzle trailing edge as well as the incidence angle characteristic of the rotor blade, unsteady numerical simulations were performed to investigate the effect of pulsating flow conditions on the performance, flow characteristics in frequency domain and shock wave behavior in a variable nozzle turbine. The results indicate that the turbine inlet pressure pulsation has strong influence on the turbine performances. Meanwhile, the turbine inlet pulsation flow has a strong influence on the intensity of the shock wave and clearance leakage flow in the nozzle, which causes significant flow losses in the turbine. In addition, at the turbine rotor inlet, the unsteadiness caused by the turbine inlet pulsation varies significantly along the circumferential direction and spanwise. Up to two-thirds of the unsteadiness caused by the turbine inlet pulsation dissipates before entering the rotor due to the flow dissipation and mixing process along the nozzle streamwise. The excitation force exerted on the rotor blade leading edge caused by the turbine inlet pulsation is about the same level as that caused by the stator–rotor interaction.

A passive shock wave control method, using a grooved surface instead of the original smooth surface of a gas turbine nozzle vane to alter a single shock wave into a multiple shock wave structure, is investigated in this paper, so as to gain insight into the flow characteristics of a multiple shock wave system and its variations with various grooved surface geometry parameters. With the combination of numerical and experimental approaches, the shock wave structure and the flow behavior in a linear turbine nozzle channel with different grooved surface configurations were compared and analyzed in details. The numerical and experimental results indicate that the multiple shock wave structure induced by the grooved surface is beneficial for mitigating the intensity of the shock wave, reducing the potential excitation force of the shock wave and decreasing the shock wave loss as well. It was also found that the benefits are related to the geometry of the grooved surface, such as groove width, depth, and number. However, the presence of the grooved surface inevitably causes more viscous boundary layer loss and wake loss, which maybe a bottleneck for general engineering application of such a passive shock wave mitigation method.

The opposed rotary piston (ORP) engine, distinguished by its exceptional power-to-weight ratio and uncomplicated design, serves as an optimal power system for Unmanned Aerial Vehicles (UAVs). Based on the three-dimensional simulation platform, the engine performance, combustion, and emission characteristics of the ORP engine at different speeds and ignition timings are clearly clarified. A larger angle of the spark plug position corresponds to a wider ignition timing range and higher power output. However, this increases the likelihood of engine knock. The optimal position of the spark plug is 18 deg before top dead center 2 (TDC2). As the ignition timing is advanced, both the pressure and temperature within the cylinder rise, and the crank angle associated with the peak values shifts nearer to TDC2. As the ignition timing shifts from −13.4 °CA to −22.8 °CA, the maximum in-cylinder pressure rises from 35.5 bar to 59.6 bar at 3000 r/min. The delayed ignition at a given ignition timing range accelerates flame formation due to a higher in-cylinder pressure at ignition. Advanced ignition can significantly enhance engine power and lower fuel consumption, substantially improving the endurance of UAVs. At 3000 r/min, the peak power, 36.3 kW, and minimal ISFC, 231.1 g/kWh, are achieved at an ignition timing of −22.8 °CA. Advanced ignition results in a wider flame propagation region, effectively avoiding incomplete combustion in the combustion chamber corners under high-speed engine conditions. The distribution of NO_x closely follows the high-temperature region, with more accumulation observed in the opposite direction of rotation. Advanced ignition contributes substantially to HC emission reduction in the combustion chamber.