Band 33, Heft 3 -

This paper investigates the ignition performances of plasma-assisted ignition in propane/air mixture. The results show that a shorter ignition delay time is obtained for the plasma ignition than the spark ignition and the average ignition delay time of plasma-assisted ignition can be reduced at least by 50%. The influence of air flow rate of combustor, the arc current and argon flow rate of plasma igniter on ignition delay time are also investigated. The ignition delay time of plasma-assisted ignition increases with increasing air flow rate in the combustor. By increasing the arc current, the plasma ignition will gain more ignition energy to ignite the mixture more easily. The influence of plasma ignition argon flow rates on the ignition delay time is quite minor.

The present work reports the effect of inlet clearance on the performance of a centrifugal blower, with parallel wall volute, over its full operating range. For a particular impeller configuration, four volutes based on constant angular momentum principle, have been designed and analysed numerically for varying inlet clearances ranging from 0 mm (ideal clearance) to 5 mm. The computational methodology is validated using experimental data. The results indicate that as the clearance increases, the impeller performance in terms of both static and total pressure rise deteriorate. Further, the stage performances deteriorate in terms of efficiency and specific work for all mass flow rates. However, the performance of volute improves at lower mass flow rates compared to the Best Efficiency Point (BEP). A set of correlations have been developed to predict the change in stage performance as a function of clearance ratio. The non-dimensional values of change in specific work, isentropic efficiency and static pressure are found to be same irrespective of the shape of the volute.

The steady-state and transient engine performances in control systems are usually evaluated by applying thermodynamic engine models. Most models operate between the idle and maximum power points, only recently, they sometimes address a sub-idle operating range. The lack of information about the component maps at the sub-idle modes presents a challenging problem. A common method to cope with the problem is to extrapolate the component performances to the sub-idle range. Precise extrapolation is also a challenge. As a rule, many scientists concern only particular aspects of the problem such as the lighting combustion chamber or the turbine operation under the turned-off conditions of the combustion chamber. However, there are no reports about a model that considers all of these aspects and simulates the engine starting. The proposed paper addresses a new method to simulate the starting. The method substitutes the non-linear thermodynamic model with a linear dynamic model, which is supplemented with a simplified static model. The latter model is the set of direct relations between parameters that are used in the control algorithms instead of commonly used component performances. Specifically, this model consists of simplified relations between the gas path parameters and the corrected rotational speed.

Supersonic turbine is widely used in the turbo pump of modern rocket. A preliminary design method for supersonic turbine has been developed considering the coupling effects of turbine and nozzle. Numerical simulation has been proceeded to validate the feasibility of the design method. As the strong shockwave reflected on the mixing plane, additional numerical simulated error would be produced by the mixing plane model in the steady CFD. So unsteady CFD is employed to investigate the aerodynamic performance of the turbine and flow field in passage. Results showed that the preliminary design method developed in this paper is suitable for designing supersonic turbine. This periodical variation of complex shockwave system influences the development of secondary flow, wake and shock-boundary layer interaction, which obviously affect the secondary loss in vane passage. The periodical variation also influences the strength of reflecting shockwave, which affects the profile loss in vane passage. Besides, high circumferential velocity at vane outlet and short blade lead to high radial pressure gradient, which makes the low kinetic energy fluid moves towards hub region and produces additional loss.

A hybrid diagnostic method utilizing Extended Kalman Filter (EKF) and Adaptive Genetic Algorithm (AGA) is presented for performance degradation estimation and sensor anomaly detection of turbofan engine. The EKF is used to estimate engine component performance degradation for gas path fault diagnosis. The AGA is introduced in the integrated architecture and applied for sensor bias detection. The contributions of this work are the comparisons of Kalman Filters (KF)-AGA algorithms and Neural Networks (NN)-AGA algorithms with a unified framework for gas path fault diagnosis. The NN needs to be trained off-line with a large number of prior fault mode data. When new fault mode occurs, estimation accuracy by the NN evidently decreases. However, the application of the Linearized Kalman Filter (LKF) and EKF will not be restricted in such case. The crossover factor and the mutation factor are adapted to the fitness function at each generation in the AGA, and it consumes less time to search for the optimal sensor bias value compared to the Genetic Algorithm (GA). In a word, we conclude that the hybrid EKF-AGA algorithm is the best choice for gas path fault diagnosis of turbofan engine among the algorithms discussed.

In this study, a genuine code was developed for optimization of selected parameters of a turboprop engine for an unmanned aerial vehicle (UAV) by employing elitist genetic algorithm. First, preliminary sizing of a UAV and its turboprop engine was done, by the code in a given mission profile. Secondly, single and multi-objective optimization were done for selected engine parameters to maximize loiter duration of UAV or specific power of engine or both. In single objective optimization, as first case, UAV loiter time was improved with an increase of 17.5% from baseline in given boundaries or constraints of compressor pressure ratio and burner exit temperature. In second case, specific power was enhanced by 12.3% from baseline. In multi-objective optimization case, where previous two objectives are considered together, loiter time and specific power were increased by 14.2% and 9.7% from baseline respectively, for the same constraints.

Numerical investigations were conducted on a turbine cascade, with end-wall cooling by a single row of cylindrical holes, inclined at 30°. The mainstream fluid was hot air and the coolant was CO 2 gas. Based on the Reynolds number, the flow was turbulent at the inlet. The film hole row position, its pitch and blowing ratio was varied with five different values. Taguchi approach was used in designing a L 25 orthogonal array (OA) for these parameters. The end-wall averaged film cooling effectiveness (ηˉ$$\bar \eta $$) was chosen as the quality characteristic. CFD analyses were carried out using Ansys Fluent on computational domains designed with inputs from OA. Experiments were conducted for one chosen OA configuration and the computational results were found to correlate well with experimental measurements. The responses from the CFD analyses were fed to the statistical tool to develop a correlation for ηˉ$$\bar \eta $$ using regression analysis.

Numerical integration simulations were performed on a ready-made central strut-based rocket-based combined-cycle (RBCC) engine operating in the ejector mode during the takeoff regime. The effective principles of various cowl lip positions and shapes on the inlet operation and the overall performance of the entire engine were investigated in detail. Under the static condition, reverse cowl lip rotation in a certain range was found to contribute comprehensive improvement to the RBCC inlet and the entire engine. However, the reverse rotation of the cowl lip contributed very little enhancement of the RBCC inlet under the low subsonic flight regime and induced extremely negative impacts in the high subsonic flight regime, especially in terms of a significant increase in the drag of the inlet. Changes to the cowl lip shape provided little improvement to the overall performance of the RBCC engine, merely shifting the location of the leeward area inside the RBCC inlet, as well as the flow separation and eddy, but not relieving or eliminating those phenomena. The results of this study indicate that proper cowl lip rotation offers an efficient variable geometry scheme for a RBCC inlet in the takeoff regime.