Volume 39, Issue 3

In this paper, global optimization design of a transonic compressor 3D blade (Rotor 37) has been carried out by a self-developed aerodynamic shape optimization (ASO) platform based on improved parallel synchronous particle swarm optimization (PSPSO). To improve the performance of PSPSO, coefficient of variation (COV) based attenuation method with new parameters is proposed and then validated by optimization tests. Flow field of blade is calculated by an in-house computational fluid dynamic (CFD) code called PMB3D-Turbo, which is validated by Rotor 37. Choosing Rotor 37 as the case, optimization object is to maximize the peak adiabatic efficiency, meanwhile constraining mass flow and total pressure ratio. The solutions show that, the ASO platform is effective to transonic compressor blade and variations of thickness distribution near the trailing edge can help improve the adiabatic efficiency of a transonic compressor blade.

The corner stall is known as a thorny problem for further increase of aerodynamic load and stability of compressors because of the notable instability separation flow in 3D space. The unsteady flow field in 6 channel rectangle cascades with clearance is studied by numerical simulation of RANS with SST turbulence model and γ-Re θ transition model. The analysis of results shows that with the increase of attack angle, the low energy fluid in the corner region chooses to move along spanwise direction firstly and fluctuates periodically. At critical angle of attack, the low energy fluid adapts to the too large flow drag by breaking through the limitation of local space in identified cascade corner and propagates as a way of unstable stall cells at lateral speed of 27.47% of inlet velocity.

In this study, the thickness of water film, which is frequently formed on compressor blade surface, was investigated. Simultaneously, its movement and extra torque demand were provided by a simulation model of the code. Having taken into account the transformation in blade’s profile and the thickness feature of the films, the working medium characteristic transformation in compressor stage was analyzed. In addition, formation, development and movement of the water film on a compressor stage were simulated by using unsteady numerical methods in various water injected conditions. Movement characteristics of water droplets in cascade were investigated to understand the flow mechanisms responsible for water film formation process. The forming and the tearing process of water films, which formed on the blade surface, were analyzed at different injection conditions. In consideration of the “Scoop effect” in actual situation, the maximum quantity of injected water can reach 12 %. Results indicated that continuity and region of the film would be developed with the increment of droplet size and injection rate. It was also found that water film’s tearing process brought about flow losses near blade surface due to the increment of surface roughness.

Fuzzy and probability-box ( p -box) variables exit widely in aerospace engineering. To evaluate the reliability of turbine discs under the mixture of these two types of variables and guarantee safety, the critical point lies in how to deal with the fuzzy variables. In this paper, a novel method based on equivalent transformation of entropy and saddlepoint approximation (SPA) is proposed to estimate the reliability of turbine discs with the mixture of fuzzy and p -box variables. The advantage of the proposed method is that it can transform fuzzy variables whose memberships are non-normal into normal random variables through entropy invariability; meanwhile, using the SPA, the required sample size and corresponding computational cost decreases greatly. An example is used to illustrate the proposed method and a comparison is also made with the interval Monte Carlo simulation (IMCS). The results indicate that the proposed method is promising and has higher efficiency with almost the same accuracy.

Unsteady pulsed suction (UPS) was applied as an unsteady flow control (UFC) technique in a certain highly loaded compressor cascade to control the flow separations. Only two bleed holes symmetrically mounted on the endwalls (one on the upper endwall and another on the lower endwall) were set up to achieve steady constant suction (SCS) and UPS. The improvements in aerodynamic performance by SCS and UPS under different time-averaged suction flow rates are firstly investigated and compared. The related unsteady aerodynamic parameters of UPS such as excitation frequency, excitation location, pitch angle, and skew angle are discussed and analyzed in detail. The results show that UPS can provide a better flow control effect than SCS with the same time-averaged dimensionless suction flow rate in the control of flow separation. The aerodynamic performance of compressor cascades can be significantly enhanced by UPS when unsteady aerodynamic parameters are in their optimum ranges. Based on the optimum parameters for UPS, the total pressure loss coefficient is reduced by 19 % only with the time-averaged dimensionless suction flow rate m s =0.4 %.

In order to solve the problem of torsional vibration instability of control unit in helicopter/engine system during maneuver flight, this paper has conducted the research on the integrated torsional vibration suppression of helicopter/engine system. Firstly, an integrated helicopter/engine model combined with torsional vibration is established based on the UH-60 helicopter model. Then the torsional vibration filter that is composed of a notch filter and a low-pass filter is applied to speed control loop of the engine to suppress the torsional vibration of power turbine speed. It is shown that the integrated helicopter/engine model combined with torsional vibration can simulate the torsional vibration characteristics under variable rotor speed, and both the low-order and high-order amplitudes of torsional vibration can be effectively eliminated below 0.04 % at constant rotor speed through utilizing the torsional vibration filter. Meanwhile, when the helicopter operates under variable rotor speed, the fundamental frequency of torsional vibration varies from 1.30 Hz to 2.30 Hz. However, all torsional vibration amplitudes within this frequency range can be damped less than 0.05 % through the notch filter whose bandwidth is 2.3 Hz, which proves the remarkable effect of torsional vibration suppression.

This paper presents the numerical simulation of Mach 1.5 supersonic jet with perforated tabs. The jet with straight perforation tab was compared with jets having slanted perforated tabs of different diameters. The perforation angles were kept as 0° and 10° with respect to the axis of the nozzle. The blockage areas of the tabs were 4.9 %, 4.9 % and 2.4 % for straight perforation, 10° slanted perforation ( Φ   {{{\Phi }}_{\ }}  = 1.3 mm) and 10° slanted perforation ( Φ   {{{\Phi }}_{\ }}  = 1.65 mm) respectively. The 3-D numerical simulations were carried out using the software. The mixing enhancements caused by these tabs were studied in the presence of adverse and favourable pressure gradients, corresponding to nozzle pressure ratio (NPR) of 3, 3.7 and 5. For Mach number 1.5 jet, NPR 3 corresponds to 18.92 % adverse pressure gradients and NPR 5 corresponds to 35.13 % favourable pressure gradients. The centerline Mach number of the jet with slanted perforations is found to decay at a faster rate than uncontrolled nozzle and jet with straight perforation tab. Mach number plots were obtained at both near-field and far field downstream locations. There is 25 % and 65 % reduction in jet core length were observed for the 0° and 10° perforated tabs respectively in comparison to uncontrolled jet.

Two dimensional computations were conducted to investigate the performance of NACA 0010 cascades with Gurney Flaps. The cascades with three different stagger angles (β B ) viz. 45 ° , 90 ° and 135 ° were studied for the aerodynamic performance characteristics of the blades with and without Gurney Flaps and the results were compared. The Gurney Flap’s length is 2 % of the chord length and has a 45 ° inclination with respect to the chord line. A significant lift augmentation is observed in all the three blade angle configurations due to the Gurney Flap. In terms of aerodynamic efficiency, it was observed that the presence of Gurney Flap was detrimental when β B  = 45 ° , while the same was highly advantageous when β B  = 135 ° . The streamline visualizations indicated the existence of counter rotating vortex pairs behind the Gurney Flap and that their rate of decay with respect to the Angle of attack was affected by the stagger angle at which the blades were arranged.

A novel control method combining nonlinear model predictive control (NMPC) with linear kalman filter (LKF) and adaptive notch filter (ANF) for an integrated helicopter/engine system with variable rotor speed is proposed to enhance the response ability of turboshaft engine. Firstly, based on the integrated helicopter/engine model with variable rotor speed, an ANF combined with frequency estimation is introduced. Then, in order to estimate the significant and unmeasurable performance parameters utilized in model predictive control, such as temperature before turbine, a nonlinear model predictive controller based on LKF is presented. The simulation verifications demonstrate that the fundamental frequency of torsional vibration changes from 1.30 Hz to 2.70 Hz when the rotor speed varies continuously by 50 %. In this case, compared with the notch filter, all torsional vibration amplitudes are damped below 0.1 % by ANF. Meanwhile, in comparison with the PID controller, the NMPC can reduce the overshoot and droop of the power turbine speed to less than 1 % with steady-state error no more than 0.5 % at the off-design point of NMPC based on adaptive torsional vibration suppression. The applications of LKF and similar transformation improve the control accuracy and robustness performance of the NMPC designed at a single operating point.

The flow field in a compressor cascade is very complex owing to the highly 3D, turbulent, and viscous properties. However, in the through-flow analysis method, the viscosity effects are taken into account using empirical models. These models were based on experimental results for early blades. However, the blade types in modern compressors are quite different from those in older compressors. Therefore, it is not possible to predict the performance of modern compressors using the old empirical models. In this study, several multiple circular-arc (MCA) blades commonly used in modern compressors were simulated. After the simulations, a database of the cascade performance was built. Based on this database, some new models were established to predict the performance of modern cascades using regression analysis methods and an artificial neural network (ANN) method. The accuracy of all these new models is high enough for use in engineering applications.

Proper design optimization and precise error characterization of temperature sensing-elements are rudimentary in the field of high accuracy gas temperature measurements. The intricate flow structures and heat transfer mechanisms related to a shielded thermocouple probe are studied in this paper using computational fluid dynamics and conjugate heat transfer (CFD/CHT) simulation tools. Owing to the probe’s novel geometry, it was a pre-requisite to characterize the metrological fidelity of the probe and the steady-state errors were determined from the simulations. The influence of certain geometrical parameters on the error characteristics of the probe was also investigated. The characteristic variation of the metrological fidelity of the probe with respect to those geometrical parameters should facilitate the optimization of the probe design.

A supersonic kerosene ignition and flame stabilization experiments were conducted on a directly connected supersonic combustion test bench. The kerosene fuel was jetted by wall jet at Ma 2.5 airflow. High-speed photography was used to record the CH* emission during ignition, extinguishing and evolution of the flame. Experiments of different equivalence ratios were performed. The processes of ignition, flame holding, and extinguishing were observed as well. The experiment showed the characteristics of ignition core initiation and extension. The time of ignition increased with the increase of equivalence ratios. Flame stability during the process of Ma 2.5 at the entrance of the combustion chamber was also studied. An equilibrium flame pattern of shock wave and flame was discovered in the experiment. In the stable flame state, shock waves near the kerosene jet orifice promote atomization and blending, and the combustion chamber pressure with stable flame makes the shock waves stable near the kerosene jet orifice, thus forming the flame stability model. The whole process and characteristics of kerosene ignition, flame holding and extinguishing are revealed in the experiment.

The purpose of this study is to develop a metaheuristic design for primary parameters and architectures of two models of artificial neural network (ANN) in predicting a cargo aircraft’s exergo-emissions (exergy destruction ratio, r ex,dest , and waste exergy ratio, r wex ) at different flight stages. Hybrid genetic algorithm (GA)-ANN models have been accomplished utilizing real databases of r ex,dest and r wex at various powers. Implementing a metaheuristics-based optimization on multilayer perceptron (MLP)-ANNs has produced the most favourable initial weights, step-size, biases, and training algorithm’s back-propagation (BP) momentum rate in addition to optimum number of neurons in the hidden layer(s). In accordance with an error assessment, a close fit linking real data and r wex (linear correlation ratio, R, value of 0.999851) as well as r ex,dest (R value of 0.999985) predicted values is found. In the r ex, dest estimation model, the accuracy among single-hidden-layer networks has been confirmed to be higher; whereas, highly accurate testing outcomes have been obtained in two-hidden-layer networks as far as modeling of r wex is concerned. ANN models’ optimization by GAs has increased the accuracy of the resulting models (R value of 0.999987 and 0.999869 for r ex,dest and r wex , in that order ascertaining a drop-off in the testing stage errors).

A Min-Max structure combined with several linear limit controllers are utilized for conventional aero-engines safety control, which is complex and conservative, therefore a new nonlinear model predictive control strategy is proposed that owns the ability to manage time-varying input/output limits during acceleration/deceleration transient states of aero-engines. Penalty function items are constructed according to the characteristics of input/output limits, and are incorporated in the cost function of online receding horizon optimization for limit management of aero-engines. Considering the large computational burden of nonlinear component-level aero-engine model directly used as prediction model, a piecewise linear model is built to describe the steady-state and dynamic properties of aero-engines and its design process is explained in detail. Further, computational efforts are investigated as control horizon, prediction horizon are varied in the improved nonlinear predictive controller. Particularly, when penalty function items are activated, control performance and computational burden are examined, thus providing guidelines for key parameters tuning. Simulations show two clear results. One is that computational efforts become larger for bigger control horizon or prediction horizon, but it depends more on control horizon than prediction horizon. The other is that the proposed method can handle time-varying input/output limits effectively without resulting in a great increase of computational burden.

The effects of back pressure and cavity L/D ratio on the shock wave structure in the cold flow field of a typical cavity-based scramjet combustor with combined inlet and isolator is investigated numerically in the selected scramjet models. The scramjet with a throat ratio of TR 0.0 and cavity L/D 6.04 was analyzed. To perform such analysis, steady, 2-D RANS was used with SST k-ω. From the analysis, the value of static pressure along the cowl surface, contours of Mach number and pressure were obtained. The scramjet was modeled with different TR 0.1, 0.2, 0.25 and 0.3 with the same cavity L/D 6.04 and different cavity L/D 4.04, 9.04 and 12.04 with the same TR 0.25. All the models were analyzed with the same inlet conditions and the results were obtained. From the analysis, it was observed that the increase in back pressure moves the shock train towards the inlet of the isolator which leads to ‘engine unstart’ after the throat ratio of TR 0.1. Also, it is observed that there is an optimal L/D ratio of the cavity L/D 9.04 which restricts the propagation of high-pressure waves obtained in the combustor.