Residual unbalance in aero-engine blade-disk systems is strongly influenced by the coupled transmission of manufacturing and assembly errors. This study proposes a measurement-informed blade arrangement optimization method that integrates blade mass moments, disk eccentricity, and assembly-angle deviations into a unified unbalance model. Coordinate measurements, blade mass-moment measurements, and assembly-position error identification are used to establish an integrated error transmission model, in which multiple error sources are represented as resultant unbalance vectors. The blade arrangement is formulated as a permutation optimization problem that minimizes resultant unbalance. A robust hybrid genetic algorithm with variable-neighborhood search is developed, combining population-based global exploration with local refinement through swap, insertion, and reversal operations. Unlike conventional strategies mainly relying on blade mass moments or empirical heavy-light alternation, the proposed method incorporates disk eccentricity and assembly-angle deviations into the unbalance-vector calculation, enabling phase compensation among measured error-induced unbalance components. Dynamic balancing experiments using actual machined blades and a disk show that the optimized arrangement reduces vibration amplitude from 0.0114 m/s2 to 0.0069 m/s2, a 39.43 % reduction, demonstrating improved arrangement effectiveness and dynamic balancing performance.
To mitigate turbine performance degradation caused by the highly pulsating exhaust from a pulse detonation combustor (PDC) in pulse detonation turbine engines, a shock divider was employed in this study to regulate the detonation gas. An asymmetric dual-channel shock divider was designed and manufactured, and a coupled experimental test platform consisting of the PDC, the shock divider, and a turbocharger was established. Comparative experiments with and without the shock divider were conducted over a PDC operating frequency range of 10 Hz-30 Hz. The experimental results indicate that the shock divider can effectively reduce the pressure pulsation amplitude of the pulse detonation gas, thereby enabling stable coupled operation of the three components across the entire range of operating conditions. Furthermore, the shock divider can enhance the turbine rotational speed, output power, and efficiency under all operating conditions. Specifically, the relative increase in turbine power reaches up to 58.1 %, and the absolute increase in turbine efficiency reaches up to 13.7 %, with the optimal enhancement effect observed at 20 Hz. The findings of this study demonstrate that the shock divider is an effective approach to improving turbine performance under pulse detonation gas impact.
This study presents a three-dimensional numerical simulation investigation into two distinct swirl configurations (axial and radial) of hydrogen fuel recirculation combustor under a lean equivalence ratio of 0.3, focusing on the effects of swirler blade angle and micro-mixing hole diameter on combustion performance and pollutant emissions. The results demonstrate that: swirler blade angle significantly influences combustion efficiency, total pressure recovery coefficient, and NOx formation. For axial swirlers, the combustion efficiency peaks at 99.92 % when the blade angle is 30 degrees, while the total pressure recovery coefficient decreases from 98.61 % to 97.26 %. The Emission Index (EI) exhibits an approximately linear increase with increasing blade angles. For radial swirlers, combustion efficiency rises from 99.56 % to 99.90 % as the blade angle increases from 30 degrees to 60 degrees, while the total pressure recovery coefficient drops from 98.68 % to 97.55 %. Temperature distribution factors (OTDF/RTDF) decrease continuously, and EI shows nonlinear growth. Micro-mixing hole diameter has a minimal impact on the total pressure recovery coefficient and temperature distribution metrics but significantly affects combustion efficiency. For axial swirlers, combustion efficiency reaches its maximum at 0.8 mm hole diameter, with negligible changes in total pressure recovery (<1 %) and temperature distribution coefficients. The EI index exhibits nonlinear growth. For radial swirlers, combustion efficiency peaks at 99.90 % when the micro-mixing hole diameter is 0.8 mm, with near-constant total pressure recovery and insignificant variations in OTDF/RTDF. The EI index shows nearly linear growth under these conditions.
Abstract This article presents an experimental study to investigate the effects of fluidic injection on the mixing characteristics of a slightly over-expanded jet with NPR of 4 and Mach number of 1.6. The experiments were performed with two injectors placed at the nozzle exit and run at an injection pressure (IP) of 1.0, 1.5 and 2.0 bar. The free jet (FJ) without injection was regarded as reference case. To determine the development of the jet, measurements of normalized total-pressure were made along the centreline and in the radial directions at different downstream positions. The measured potential-core lengths were 8.36D e , 8.86D e , 8.28D e , and 6.35D e for FJ, IP 1.0, IP 1.5, and IP 2.0, respectively. The core length is increased by 5.98 % on IP 1.0, decreased by 0.95 % on IP 1.5 and by 24.04 % on IP 2.0 compared to FJ. This is due to weak interaction, with the relatively low injection momentum, which is the reason for the long core in IP 1.0. As the injection pressure increases, the shear-layer is more developed, more entrainment and more spreading and mixing are achieved.
To clarify the low-temperature ignition characteristics of a gas-turbine can-annular combustor, a coupled numerical model of a torch igniter and combustor was established to examine ignition limits, torch-jet behavior, internal flow, fuel-droplet evaporation, and transient flame propagation. With an ignition energy of 4 J, the isolated torch igniter achieved successful ignition at 264 K but failed at 259 K and 234 K. In the coupled configuration, combustor ignition succeeded at 269 K but failed at 264 K. As the inlet-air temperature decreased from 334 K to 264 K, the area of the 1,500 K isosurface decreased from 0.0355 m2 to 0.0235 m2, a reduction of approximately 33.8 %, while the maximum flow velocity decreased only from 42 m/s to 41 m/s. Although the overall flow topology remained similar, lower temperature suppressed droplet evaporation, reduced vapor-phase fuel concentration, and increased the retention of relatively large droplets near the flame-tube head, primary holes, and walls. At 264 K, an initial ignition kernel still formed near the torch-igniter outlet, but insufficient thermal coverage, weakened evaporation, and deteriorated local mixing prevented sustained flame growth and propagation into the primary combustion zone.
