To predict the performance parameter changing trend of an aeroengine, a novel double-extremum learning particle swarm optimization (DELPSO) algorithm is proposed. Inspired by human learning behavior, this algorithm simulates this behavior, including the strategies of collective learning, private tutoring, and research behavior, so that obtained final solutions would be in the global optimal area or its neighbor area as close as possible. Meanwhile, to improve the prediction performance, a nonlinear mapping function is designed to describe the feature relationship between inputs and outputs of historical data. Based on the DELPSO, the fitness function synthetically considers the changing trend and the prediction error and can adaptively select optimal parameters of the nonlinear mapping function. The experimental results demonstrate that the DELPSO has globally stable and reliable performance. To validate the prediction performance of the proposed DELPSO, it is also applied to an aeroengine. Its good prediction performance indicates that the proposed DELPSO is an important reference for maintenance decision-making of aeroengines.
The vibration optimum design for a simulated power turbine rotor without blades on the disk by using an optimization method based on the finite element method is described in this paper. The installation position of the two-stage turbine disks is chosen as design variables under the constraints of feasible regions for the critical speeds of the rotor. The objective functions are to minimize the transient response of the acceleration at the bearings and the amplitude of the disks. Predictions of the dynamic characteristics of the rotor are obtained by using ANSYS code. The optimization problem is solved by using commercial optimization code ISIGHT. The optimum installation position of the two-stage turbine disks is determined after optimization design. Experimental tests under the optimized structure show that the amplitude of the two-stage turbine disks which are recognized as the most concerned optimization objectives are reduced by 59 % and 56 % respectively in comparison with the comparative structure. The encouraging results demonstrate the potential of the presented method as an engineering design tool and also lay a foundation for the design of the real power turbine rotor used in turbo-shaft engine.
Thrust augmentation in ejector mode is important to improve engine performance so as to enable more propulsion applications of RBCC. Usually, the internal flow-path is configured mainly for combustion organization in ramjet and scramjet modes, thus the RBCC performance in ejector mode will mainly depend on the primary rocket operating parameters and jet expansion state. Considering such restrictions and requirements on the primary rocket, this paper studies the effects of the primary rocket jet on the thermodynamic cycle performance of ejector mode, in which incoming air is dominated by ejector-suction by the primary rocket jet at low flight Mach numbers. It is found that the engine performance in ejector mode will be improved by increasing the primary rocket chamber pressure and nozzle expansion ratio. Furthermore, for better design of primary rocket, primary rocket chamber pressure should be maximized on the basis of ensuring a complete expansion of primary rocket to the ambient pressure at the design point of flight conditions after 3-D CFD simulations for a full flow-path of RBCC including fore-airframe and aft-airframe of a flight vehicle. The results of 3-D numerical simulations show that the bypass ratio and specific impulse increase by 35.5 % and 12.5 % respectively at sea-level static condition.
This paper aims in assessing the effect of biofuel blend such as butanol, jatropha methyl ester, soya methyl ester and rapeseed methyl ester as an additive for the aviation fuel. In addition to the blends, the nanoparticle TiO2 of 3 % is added to the biofuel. The nanoparticle mixed at the concentration of 300ppm by ultrasonication process. The fuel Jet A, B27T, J27T, S27T and R27T are investigated for combustion and emission characteristics for various throttle settings in micro gas turbine engine. Addition of additives improves the ultimate property of the fuel by reducing the kinematic viscosity. The fuel blend B27T reports 25 % increase in total static thrust and 22 % reduction in thrust specific fuel consumption. From the results it is evident that, all fuel blends showed a significant reduction in emission values owing to high oxygen content. In addition, the thermal efficiency of the B27T and J27T is improved appreciably to 30 % and 10 % higher than Jet A fuel owing to the influence of the nanoparticle TiO2. On the other hand, the emissions like CO and NOx reduced drastically up to 70 % and 45 % respectively.
Ribs effects on the heat transfer performance and cooling air flow characteristics in various aspect ratios () U-shaped channels under different working conditions are numerically investigated. The ribs angle and channel orientation are 45° and 90°, respectively, and the aspect ratios are 1:2, 1:1, 2:1. The inlet Reynolds number changes from 1e4 to 4e4 and rotational speeds include 0, 550 rpm, 1,100 rpm. Local heat transfer coefficient, endwall surface heat transfer coefficient ratio and augmentation factor are the three primary criteria to measure channel heat transfer. Ribs increase the heat transfer area and improve heat transfer coefficient of ribbed surfaces significantly, especially in the 1st pass, while the endwall surface contributes more to channel heat transfer because of the larger area and relatively smaller heat transfer coefficient. The wide channel (=2:1) owns the better augmentation factor than the narrow channel (=1:2) and ribs heat transfer weight increases with an increase of the inlet Reynolds number. Rotating slightly reduces the ribs heat transfer weight in channel and the trailing surface in 1st pass is the main influence object of rotating.
The onset of spike stall induced by the interaction of hub corner separation flow with the tip leakage flow is investigated in detail by numerical method in this paper. The time resolved results indicate that the remarkable radial secondary flow from hub to tip near the trailing edge is formed when the compressor approaching rotating stall. The radial secondary flow is unstable and cross-passages propagates, which flows in and away out of the tip region periodically. The disturbance caused by radial secondary flow will influence the tip leakage flow directly by reforming the vortexes in blade tip region. A secondary vortex which comes from the radial migration of corner separation and is induced by the tip leakage vortex appears in the tip region. The simulation result demonstrates that the generation of the secondary vortex is an important symbol of blockage growth in the tip region at the stall inception phase. The disturbance produced by secondary vortex is an incentive of the leading edge overflow and the intensity of secondary vortex could be used as a criterion of rotating stall before leading edge spillage.
The synthetic jet actuator was applied to the nozzle jet vector control by experimental and numerical simulation methods to study fluid thrust vector control in this paper. The nozzle jet produced a steady and continuous deflection under the action of synthetic jet actuators. Through the analysis of the flow structure and control mechanism of the interaction between the synthetic jet and the nozzle jet, it was found that the force formed by the pressure gradient in the nozzle duct, the entrainment and ejection of the synthetic jet to the mainstream, and the momentum synthesis of the vertical synthetic jet were able to cause the deflection of the nozzle jet. The actuator excitation voltage can adjust synthetic jet velocity, which in turn affects the nozzle jet deflection angle.
The experiment system of pulse detonation engine is set up to investigate the acoustic signature of detonation acoustic. The research on detonation acoustic characteristic of pulse detonation engine in three kinds of working states, deflagration stage, detonation stage and external detonation stage, is carried out. Results from the test show that the maximal PImpact, PJet1, τ+ and τInterval are obtained in external detonation stage. In deflagration stage, the concentrate of PDE acoustic signal energy is obvious. The energy of impulsive rise of impact noise and jet noise is an important part of PDE acoustic in deflagration stage. In detonation stage, a concentration of energy distribution is appeared in low frequency, and the energy of medium-high frequency is decreased. In external detonation stage, energy distribution is more concentrated in low frequency. The value of η11/η5 can be used to detect the working state of pulse detonation engine. When η11/η5 is less than 5, pulse detonation engine is in deflagration stage, however, when η11/η5 is large than 5, pulse detonation engine is in detonation stage or external detonation stage.