Volume 42, Issue 2

Capturing degradation trends from the Condition monitored signals is a proven technique for predicting the Remining Useful Life (RUL) of the equipment, which has gained more prominence in Prognostics and Health Management (PHM) in Industry 4.0. However, this process is tiresome and comprehending all the physical parameters of the system to construct a Health Index that characterize the health state is a complex process, especially if multiple sensors are involved. This work proposes a Deep residual ensemble model which constructs Fused Health Index (FHI) by harnessing temporal property of signals. The proposed Residual network integrates Bi-directional Long Short Term Memory (Bi-LSTM) and Deep Neural Network (DNN) which absorbs individual residuals of both the forward and reverse LSTMs that acts as an important feature to improve the overall prediction process. The work validated using CMAPPS dataset using various unique performance metrics to portray the effectiveness of the model.

The potential of gliding arc plasma-assisted ignition and combustion technology to enhance ignition and combustion performance is attracting increasing attention from the scientific community. A multitude of experimental studies have been conducted by scientists and engineers on its application in gas turbine combustors. This paper presents a review of the research conducted on gliding arc plasma-assisted ignition and combustion over the past five years. Gliding arc plasma exerts a multitude of effects on combustion processes. These effects can be broadly categorized as follows: (1) reduction in ignition delay time; (2) expansion of ignition and lean blowout boundaries; (3) enhancement of ultra-lean burning combustion and low-temperature flame stability; (4) improvement in combustion efficiency; (5) reduction in pollutant emissions; (6) augmentation of stability of unstable fuels such as ammonia. Finally, a prospection on the application of gliding arc plasma assisted ignition and combustion technology in gas turbine combustor is presented.

The cantilevered stator is a typical configuration for axial compressors. To gain insight into the effect of the hub leakage flow, unsteady simulations were performed in a multi-stage cantilevered compressor. The main goal is to gain a better understanding of the interaction between the upstream rotor wake and hub leakage flow, and the effect of the hub leakage flow on the downstream rotor performance. The results show that the upstream rotor wake affects the hub loading of the cantilevered stator periodically and thus changes the leakage flow characteristics, leading to an unsteady benefit in the vicinity of the hub. At the higher span region, the unsteady benefit depends on the reversible recovery and the wake-boundary layer interaction. The increased hub clearance not only causes performance degradation on the cantilevered stator but deteriorates the downstream rotor performance, especially for the absolute angle due to the mass flow redistribution. These results indicate great attention should be paid to the interaction between the upstream rotor wake and the hub leakage flow, and also the stage matching caused by the hub leakage flow in the design process of cantilevered compressors.

In this paper the problem of the gas turbine blades equivalent dynamic stresses and lifetime determination has been studied. The obtained numerical results have been experimentally verified. The equivalent stresses and the lifetime have been determined for the most dangerous forced vibration bending and torsion modes of all three impellers, forming the turbine rotor. The marine gas turbine engine should be as compact as possible. Thus its rotor consists only of three impellers, that causes rather high vibration and thermal loads on the rotor blades. The lowest number of cycles before the blades failure have been observed for blades of the first impeller, that are influenced by the strongest thermal and vibration loads. Obtained results show that for the surface temperature of nearly 1,000 °C the maximum equivalent stresses are approximately equal to the blade material endurance limit. All numerical results are in good correlation with the experimental data, the divergence is less than 7 %. Received results can be used for the next stage of studies, concerning the rotor creep and fatigue crack problems.

The study’s primary goal was to assess the effects of different geometrical configurations of cavity flame holders on the performance characteristics of an H 2 -based scramjet combustor overall efficiency, focusing on mixing efficiency, combustion performance, and total pressure loss. Key findings indicate that incorporating a cavity into the combustor design leads to developing a robust recirculation zone within the cavity area. This recirculation zone is pivotal in enhancing fuel-air mixing and combustion efficiency, with cavity-based combustors showing an earlier onset of combustion and achieving peak combustion efficiencies around 90–95 %. The extent of the recirculation region is notably influenced by the proximity of the strut injector to the cavity’s length. This shock interaction can adversely affect scramjet combustor performance, especially at higher cavity angles ( α ) and specific geometric configurations, such as an L / D ratio of 4 and α  = 30°. This advantage is critical, especially considering the compact length of the combustor, which is a desirable attribute in scramjet design.

Aviation, marine propulsion, and power generation employ gas turbine engines extensively. The combustion process’s high temperatures make gas turbine engine development difficult. For safe turbine blade operation, internal cooling is essential. Ribs, pin-fins, and impinging jets are just examples of the several different types of structures that are typically included in internal cooling systems. The present study investigates the effect of rib placed in between the pin fins in a wedge channel. Nine rib configurations along with base line model without ribs are employed in this study. The SST k-omega turbulence model was employed to simulate the flow field and heat transfer of each case using the commercial software ANSYS Fluent. The investigation demonstrated that the heat transfer distribution and flow field in pin-fin arrays were substantially influenced by the ribs-induced secondary flows, particularly in the spanwise direction, which resulted in a significant heat transfer variation. In comparison to other configurations, the 4H4L rib configuration exhibits the highest thermal performance factor, surpassing the baseline by approximately 16.6 %.

The present study aims to identify the suitable tip clearance and volumetric flow rates for low-speed axial flow fans. The performance of an axial fan depends on various parameters like fan speed, available tip clearance gap, volumetric flow rate, power consumption, and working fluid. Leakage flow occurring at a rotating component such as a blade and solid casing of a ducted axial fan is typically linked to losses and the potential emergence of a rotating stall. The numerical analysis for this study uses the Reynolds Averaged Navier–Stokes equations with the k-omega SST turbulence model to perform the steady-state simulations by varying tip gap and volumetric flow rate. The results proved that the optimum performance was obtained with a tip clearance gap of 1 mm, and maximum fan efficiency was achieved at a volumetric flow rate of 3.9–4.5 m 3 /s. The novelty of this proposed work is to enhance the efficiency of axial flow fans with circular arc-cambered aerofoils using optimum tip clearance and volumetric flow rate through steady-state simulations. This method can be used in turbomachinery to improve fan performance.

Micro Gas Turbine (MGT) designers often have the need for a design tool, capable of rapidly providing feasible design options. The upgrade of such an existing MATLAB ® based mixed flow compressor design application (1D App) is discussed. In MGT engines, diffuser design largely influences the compressor stage performance. Crossover diffusers provide superior efficiency and pressure ratio performance compared to legacy type diffusers, albeit at the expense of operating range. The reduced operating range attributed to single vane crossover diffusers can be alleviated by splitter vane, tandem vane, and combined configurations. Therefore, the 1D App is updated to allow the user the ability to design these additional crossover diffuser configurations. Performance results are verified using Numeca/FINE™ Turbo CFD software. It is shown that the upgraded 1D App provides predicted performance results within 1 % of CFD results for both total-to-total efficiency and pressure ratio.

The effect of implications of the dual cavity with aft wall divergence in a parallel injection reacting flow has been numerically investigated. This research intends to emphasize the behaviour of the supersonic flow under varying divergence angles of the cavity aft wall. A two-dimensional Reynolds-Averaged Navier-Stokes (RANS) equation and an SST k-ω turbulence model with a single-step chemical reaction for hydrogen-air are utilized for the simulation. Followed by the strut injector, the cavities are positioned symmetrically inside the combustor. The bottom cavity aft wall divergence varies, whereas the top wall is mounted with a rectangular cavity. The performance of cavity locations is compared with the baseline DLR model. From the evaluation of numerical outcomes along with different cavity configurations, it has been noted that the 15-degree cavity divergence angle enhances the recirculation zone which leads to improved mixing performance. Also, the cavity improves the combustion stability by increasing the flow residence time. From this numerical analysis, it is associated that an almost 20 % reduction in combustor length is achieved, however, an 18 % rise in the pressure loss is noted because of emanating shock waves from the cavity edges.

One of the main objectives of this research is to construct a high-fidelity aerothermal model for controlling the turboshaft engine power turbine rotational speed. Firstly, a high-fidelity aerothermal model of General Electric T700 is formed. The turboshaft engine aerothermal model is simulated and compared to engine design point data on MATLAB/Simulink environment. Thermodynamic equations and some algebraic equations are used. In predicting compressor mass flow rate, the adaptive neuro-fuzzy inference system method is preferred. A propeller model is also added to the turbo shaft engine aerothermal model and to keep the power turbine rotational speed at a constant value, a Proportional Integral Derivative controller is designed and applied. Open-loop and closed-loop simulations are conducted and successful outcomes are achieved. Results indicate that the differences between simulation and actual engine data are within acceptable limits. The suggested model can be readily updated and expanded to control additional parameters of the engine. PACS 2010 Classification: Control theory in mathematical physics, 02.30. Yy, Computer modeling and simulation, 07.05.Tp, 05.70.Ce Thermodynamic functions and equations of state.

A multi-nozzle thrust matching control method is proposed in this paper to provide the thrust vector required of lift-fan type short takeoff and vertical landing (STOVL) aircrafts during hover. Attitude-matched thrust commands for each nozzle of STOVL engine are generated based on a six-degree-of-freedom aircraft model. The thrust controller is divided into two blocks according to the coupling analysis. Based on the combination of linear matrix inequality and differential evolution algorithm, a multi-target controller parameters optimization method is proposed to achieve stable disturbance suppression and fast command tracking. Simulations results show that the proposed command generated model is effective and the decoupling control method can effectively suppress the coupling between the loops and realize the matching control of thrust.

By the finite element analysis software ABAQUS, utilizing General Contact and Contact damping algorithm, the brush seal hysteresis numerical model with pressure differential considered was established. The hysteresis was quantified by the hysteresis energy, and the hysteresis energy and maximum stress are obtained by numerical calculations. An orthogonal test was conducted to study the effects of bristle diameter, bristle cant angle, fence height, bristle length, upstream and downstream differential pressure, and rotor radial displacement on the hysteresis and maximum stress of the brush seal. Results show that the primary and secondary orders of parameters affecting hysteresis are: rotor radial displacement, bristle diameter, upstream and down-stream differential pressure, bristle cant angle, bristle length, and fence height. The primary and secondary orders of parameters affecting maximum stress are: upstream and downstream differential pressure, fence height, bristle cant angle, bristle diameter, bristle length, and rotor radial displacement. Finally, based on the numerical results, a fitted correlation was developed. Comprehensive effects of the six parameters on both performances were analyzed, and structural design optimization methods focusing on single performance and comprehensive performance were proposed, providing references for the design of brush seals in practical engineering applications.

Ribbed-swirl cooling performs excellently in enhancing heat transfer, but previous studies were almost conducted in simple circular tubes under static conditions. The influence of actual mainstream flow, heat conduction of blade as well as the rotation effect was ignored. To explore its application potential in a real environment, the thermal performance of swirling cooling after adding annular ribs to its target surface in a real turbine blade was analyzed, and the impacts of centrifugal force along with Coriolis force were also included. From the results, the ribs affected the distribution of nozzle mass flow rate. For nozzles near the blade root, the mass flow rates in ribbed-swirl cooling were smaller than those in smooth-swirl chamber, while near blade tip, those were higher. Besides, the influence of Coriolis force from SS (suction side) to PS (pressure side) on the development of cross-flow and swirl in smooth and ribbed chamber exerted a significant difference. For ribbed-swirl cooling, cross-flow passing through ribs induced flow separation, which was conducted to cross-flow suppression. The combined action of the above two aspects led to an obvious heat transfer advantage of ribbed-swirl cooling. In comparison to smooth-swirl chamber, its blade surface temperature was further reduced to more than 25 K.

Drawing the inspiration from Full Engine Simulation performed by the scientific community, the methodology for performing an integration analysis of single stage turbine with convergent exhaust nozzle of a small gas turbine engine is developed in the present study. Initially, a numerical simulation on exhaust system supported with experimental validation has been discussed. A scope of improvement in the exhaust system study has been identified and these principles are applied to improve the accuracy of prediction of CFD analysis. By taking the confidence from the validation, a 3D steady pressure-based analysis is carried out to understand the flow patterns and aero-thermal properties of the single stage turbine with convergent type nozzle influenced by the integration. The integration resulted in increased specific work output of the turbine and increased the thrust coefficient of the nozzle. Further important outcomes of this study highlight the changes observed in the performance parameters with and without integration. The thermal pattern across the rotor blades from CFD analysis is observed to be in-line with the used blades of the gas turbine engine thereby reinforcing the accuracy of the methodology developed for the integration study.

To investigate the optimal flight trajectory for hypersonic vehicle and Turbine-Based Combined Cycle (TBCC) engine to enhance mission efficacy, a performance calculation model of a variable-cycle TBCC engine is developed utilizing the method of airframe/engine integrated performance analysis. Employing the Radau pseudospectral method for trajectory optimization, the study of the hypersonic vehicle involves the angle of attack, the scramjet engine equivalence ratio, and the core driven fan stage (CDFS) stator vane angle as control variables. Two optimization scenarios were considered: minimizing climb time and maximizing range. The findings indicate that by optimally adjusting the angle of attack and the scramjet engine’s equivalence ratio, the vehicle’s climb time can be reduced by up to 33.68 %, and the total flight duration by 8.07 %. Moreover, optimizing these variables, along with CDFS stator vane angle, can decrease fuel consumption during climb and potentially extend cruising distance by 8.56 %, increasing overall range by 3.63 %.

This research focuses on identifying isotherms on the surface of a gas turbine combustion chamber using a colour-changing solvent known as temperature-indicating paint (TIP). The study employs a Multi Colour Change (MCC) 350-8 TIP to detect surface temperature variations. An innovative algorithm is proposed to identify these variations through irreversible colour changes, utilizing a colour-matching technique. The surface temperature data obtained from this method is compared with experimental calibration and computational analysis, showing not just agreement, but a high level of agreement, reinforcing the reliability of the method. This approach aims to enhance the accuracy of temperature detection in combustion chambers, which is crucial for optimizing performance and reducing emissions. The findings contribute to developing more efficient gas turbine systems by providing a reliable method for monitoring and controlling combustion temperatures. This study’s novelty lies in its use of colour-changing technology to achieve precise temperature mapping on complex surfaces.

The performance seeking control (PSC) method is aimed at improving the performance of aircraft engine without changing hardware devices. However, the complexity of the system makes it difficult to achieve its application in engineering. Thus, the key issues in PSC of turbofan engine are studied based on global sensitivity analysis, and the improvements are made to the optimization system according to the analysis. Firstly, the principle of PSC is analyzed based on traditional frameworks. And an improved linear programming method based on task scheduling is proposed to reduce the time consumption within a single closed-loop simulation. The impact of the estimation model accuracy on optimization results is studied adopting a sensitivity analysis method based on Monte Carlo sampling to address the optimization lose efficacy. Finally, the hardware-in-the-loop simulation is carried out utilizing model-based design method in a certain type of turbofan engine electronic controller.

Model predictive control (MPC) can be applied to aero-engine for optimal command tracking and constraint handling. However, the performance of MPC is highly dependent on the accuracy of the predictive model. Therefore, a new structure MPC with an additional modeling error compensation loop is proposed, which can substantially reduce the dependence of MPC on the accuracy of the predictive model, and enhance the performance and real-time property of the system simultaneously. The new MPC replaces the traditional feedback correction loop with a compensation control loop to realize high-performance tracking over the large envelope with only one predictive model. The compensation loop controller employs an augmented discrete linear quadratic method for robust tracking capability. Simulations on the turboprop engine show that, compared with the standard MPC, the new MPC shows greater robustness over the envelope with smaller settling time, overshoot, and fluctuation of power turbine speed.

Experimental studies were performed to explore the ignition performance of rotating gliding arc. Dual-axial swirler and beveled holes combined radial vanes swirler were used to produce rotating gliding arc. The ignition test with dual-axial swirler at normal temperature and pressure showed that the ignition fuel to air ratio (FAR) was reduced by more than 30 % compared with conventional ignition method. The ignition performance of four kinds of beveled holes combined radial vanes swirlers was compared. The ignition tests were also performed at the ground low temperature and high altitude conditions. The results showed that the ignition FAR was reduced by more than 20 % at ground low temperature and 3.0 km altitude and was higher at 4.5 km and 5.5 km altitude compared with the conventional ignition method. The result indicated that the low pressure had a more obvious influence on the discharge and ignition performance compared with the low temperature condition.