Volume 38, Issue 4

The present paper overviews the works carried out on achieving desired temperature pattern factors at combustor exit in gas turbine engines. These pattern factors are very important from the point of engine performance and life of turbine blades and vanes. They are controlled by a number of geometrical parameters such as liner front-end air passages, primary air holes, atomizer characteristics and air swirl number and dilution zone geometrical configuration. Combustor inlet pressure, Mach number, velocity profile and fuel-air ratio are the major operating parameters that influence the pattern factors. Due to the design uniqueness and importance of pattern factors, it is always a challenge to assess the pattern factors over a wide range of mission points for a gas turbine combustor designed for combat aircraft.

The trench film cooling employs film holes embedded in a slot created on the surface that requires protection from the impact of hot mainstream flow. The present investigation employs Response Surface Methodology (RSM) approach coupled with CFD analysis to develop a regression predictive model and to optimize the trench geometric and flow parameters viz., trench width (w), trench depth (d), film hole compound angle (β) and blowing ratio (M). The Area-averaged film cooling effectiveness (Ƞ AA ) were chosen as a response factor for RSM and with trench design and flow parameters used as input factors for regression analysis. Analysis of variance (ANOVA) analysis was carried out on the regression model to identify the influence of individual parameters. Three dimensional response surfaces that relate the effect of input parameters on the response factor were analyzed. Experimental results of a case identified from the RSM matrix was found to correlate well with computational investigations. Results from the study indicate that the parameters d, β and M have considerable impact on film cooling performance of test surface with trenches.

This study illustrates an deep learning approach supported by a metaheuristic design targeting the foremost features and parameters of artificial neural network (ANN) framework used in predicting relative exergy destruction ( ( f r e l , d e s t * ) (f_{rel, dest}^*) ) of a turboprop components. The development of deep ANN comprising of three-hidden layers using data obtained considering multiple engine input parameters was accomplished. Once the deep learning ANN frameworks were hybridized with a metaheuristic approach, such as genetic algorithms (GAs). The analysis of errors revealed a close fit involving the predicted values of the model and references made on data in f r e l , d e s t ∗ f_{rel, dest}^* for the engine main components. The use of appropriately chosen values in preceding networks weights produced more accurate testing results (Linear correlation coefficient values (R) for the engine components are found to be between 0.996497 and 0.998986) in networks using three hidden layers compared to those using lower hidden layers. Furthermore, optimizing deep ANNs using GAs delivers not only further improved accuracy (R is calculated to be in the range of 0.998929–0.999966 for the engine components) but also an effective utilization of time in the resulting models.

A novel thrust control method based on inverse control is proposed to improve engine response ability. The On Line Sliding Window Deep Neural Network (OL SW DNN) is proposed and adopted as inverse mapping model modeling method of inverse control. The OL SW DNN has deeper layer structure, which makes the inverse mapping model have stronger fitting capacity for nonlinear object than traditional NN. Moreover, due to adopt on-line learning modeling method, the proposed adaptive control method can obtain desired thrust whether engine degrades or not. The comparison simulations of the traditional control method based on PID and the proposed control method are carried out. Compared with the traditional control method, the proposed control method can obtain desired thrust when the engine degradation occurs, but also has fast response ability (the acceleration times for engine thrust increase to 95 % thrust of acceleration object decreases by 1.35 seconds).

This study focuses on for a PW4000 high-bypass turbofan engine using energy, exergo-sustainable and performance viewpoint. For this aim, irreversibility and performance analyses are firstly performed for five main engine components at ≈260 kN maximum take-off thrust force. Besides, overall efficiency of the turbofan is determined to be 33 %, while propulsive and thermal efficiency of the turbofan are 72 % and 46 % respectively at 0.8 M and 288.15 K flight conditions. Secondly, calculation component-based exergetic assessment is carried out using exergetic indicators. According to the calculation, the exergetic efficiency of the engine is 32 %, while its waste exergy ratio is 0.678. Furthermore, exergetic sustainability measure is obtained as 0.473, while enviromental effect factor is 2.112. These indicators are also anticipated to help comprehend the connection between engine performance parameters and worldwide dimensions such as environmental effect and sustainable growth.

Centrifugal compressor performance map prediction is of primary importance for safe and effective operation of turbochargers. This article is a contribution on compressor map prediction using empirical relations based on automotive turbocharger manufacturers’ performance maps. The present method evaluates the minimum and the maximum air flow rates, as well as the maximum compressor pressure ratio by original empirical equations exploiting impeller geometrical data. Newly introduced equations based on the mass flows and the maximum pressure ratio acquired above provide the compressor characteristic lines. The method is validated by applying it to various commercial automotive turbochargers with known performance maps from their manufacturers. At intermediate values of impeller speed, where the turbocharger is expected to match the engine, the computed compressor map agrees to the manufacturer’s data, while, differences are observed at the maximum impeller speed line. From the cases examined, it can be stated that the present model can be applied to predict small diameter, high rotational speed compressor performance, particularly at the high efficiency region that the turbocharger is supposed to match the IC engine.

The most important flow behaviour of supersonic compressor cascades is the shock wave boundary layer interaction (SWBLI). Large eddy simulation (LES) and multiple analysing methods are applied in current study to capture more details of the flow field. It is noted that the LES can catch the dual peaks feature near the SWBLI region with respect to the experimental results. Besides, SWBLI is not only the main losses source in the cascade, but also the most important origin of the unsteadiness behaviour. The high frequency signals correspond to the coherent structure in the boundary layer and dissipate downstream in the cascade, while the low frequency signals relate to the motion of the reflection point of the passage oblique shock wave and dominate the frequency spectrum downstream.

Conjugate heat transfer analysis is carried out on the internal surface of the first-stage nozzle guide vane of a gas turbine, which has both impingement and film cooling holes. The mainstream flow Reynolds number and internal coolant flow Reynolds number systematically changed and its effect on internal local surface temperature variation is studied. It is found that an increase in the coolant mass flow rate causes a non-uniform decrease in the local internal surface temperature. The external film coolant jet-lift off and internal impingement cross-flow are significant contributors to the non-uniform variation in surface temperature. It is also observed that the leading edge regions are prone to jet lift-off, whereas the tip regions of the suction surface are prone to self-induced cross-flow, due to which hot patches are formed in these regions. Hot patches are observed near the hub regions of a pressure surface due to the reduced film thickness on the external surface. From these observations it is concluded that local values of internal surface temperature are differently affected in different regions of the vane surface for a given combination of mainstream and coolant flow rates. Therefore, the conventional method of obtaining the internal temperature distributions by considering generalized geometries may not yield accurate solutions, in predicting the life of the nozzle guide vane.

The effect of aspect ratio (AS = 0.5, 1.0, 2.0) and Reynolds number (Re = 1, 2, 4, 8 × 10 5 ) on the flow and performance characteristics of a 90° curved rectangular diffuser was examined with the aid of the commercial CFD code ANSYS FLUENT. The results from computational analysis revealed the presence of stream vorticity indicative of secondary flow brought about by the curvature of the diffusing duct. As the fluid flow developed, the inertial core flow was pushed to the outer concave wall by overbearing centrifugal forces. The fluid flow sustained marginal changes with increasing aspect ratio and Reynolds number. However, the flow uniformity at the exit was enhanced. The rise in coefficient of static pressure (C p ) recovery and the drop in coefficient of total pressure loss (C loss ) with increasing Reynolds number suggested improved diffuser performance. The flow uniformity index was found to be the highest for the diffuser of aspect ratio 1.

The lean blowout (LBO) limits are critical for the performance of multipoint lean direct injection (MPLDI) combustors. The recessed pilot stage is a promising method to improve the LBO limits for the MPLDI combustors. In the present study, experiments are conducted to investigate the effects of the recess length of the pilot stage on the LBO limits for the MPLDI combustors. Five types of configurations with different recess lengths (0 mm, 7.5 mm, 15 mm, 22.5 mm and 30 mm) of the pilot stage are utilized. It could be concluded from the experimental results that the recessed pilot stage is an efficient method to improve the LBO limits for the MPLDI combustors. Compared with the configuration whose pilot stage is flush mounted, the maximum improvement of the LBO FAR could reach about 42 % by recessing the pilot stage for 22.5 mm. When the recess length is short, the LBO FAR decreases quickly with the increasing of the recess length. When the recess length is long, the LBO FAR decreases slowly with the increasing of the recess length. The mechanism for the function of the recessed pilot stage is revealed by carefully designed numerical simulations to be the suppression of the effect of the disturbance air.

Pulse detonation engines work similar to pulsejets, with the major difference taking place in the combustion mechanism. Having detonation, instead of conventional combustion comes with certain benefits, but also, a couple of major issues which will be addressed in this paper. These changes come as a result of the increased amount of energy released by the explosive process, with much higher fluid velocities and temperatures that easily surpass those of conventional engines. These high values of the thermodynamic parameters prove to be a tough technological problem when it comes to producing this type of engines, so this paper aims at analyzing the evolution of these parameters throughout a working cycle of the Tangential Impulse Detonation Engine (TIDE) and how they affect the internal structure of the detonation chamber. The non-steady character of the thermodynamic parameters from a Computational Fluid Dynamics (CFD) simulation will be used a as a starting point for the Finite Element Analysis (FEA), which will help determine whether or not, the structure can withstand the immense load that comes with such violent processes. The final part of this article will provide optimization criteria to decrease the loads generated by the pulsed detonations.

Investigations on the performance of a rear wall angled cavity with upstream transverse fuel injection in a Mach 1.8 flow field is experimentally studied in a non-reacting flow facility. The high speed flow is directed to a circular cross sectional supersonic combustor and proceeded towards the cavities having two consecutive angles being inclined towards the downstream flow direction. Wall mounted injector is positioned at a distance of 10 mm upstream from the cavity. Air is used as the injectant to simulate the gaseous fuel. The experiments are performed to explore the effect of the increase in injection pressures within various rear wall angled cavities by comparing with the ‘no-injection’ case and to finally assess the mixing performance of the flow. Transverse injection through upstream wall orifice of the cavities outlines a more uniform mixing compared to ‘no-injection’ configuration. Increase in injection pressures enhances mixing and stagnation pressure loss values.