Showing papers in "International Journal of Turbo & Jet-engines in 2016"
TL;DR: In this paper, an exergetic analysis of a small turbojet engine is presented, including design parameters and performance assessment, by identifying the losses and efficiency of a gas turbine engine.
Abstract: Abstract An exergy analysis is presented including design parameters and performance assessment, by identifying the losses and efficiency of a gas turbine engine. The aim of this paper is to determine the performance of a small turbojet engine with an exergetic analysis based on test data. Experimental data from testing was collected at full-load of small turbojet engine. The turbojet engine exhaust data contains CO2, CO, CH4, H2, H2O, NO, NO2, N2 and O2 with a relative humidity of 35 % for the ambient air of the performed experiments. The evaluated main components of the turbojet engine are the air compressor, the combustion chamber and the gas turbine. As a result of the thermodynamic analysis, exergy efficiencies (based on product/fuel) of the air compressor, the combustion chamber and the gas turbine are 81.57 %, 50.13 % and 97.81 %, respectively. A major proportion of the total exergy destruction was found for the combustion chamber at 167.33 kW. The exergy destruction rates are 8.20 %, 90.70 % and 1.08 % in the compressor, the combustion chamber and the gas turbine, respectively. The rates of exergy destruction within the system components are compared on the basis of the exergy rate of the fuel provided to the engine. Eventually, the exergy rate of the fuel is calculated to be 4.50 % of unusable due to exergy destruction within the compressor, 49.76 % unusable due to exergy destruction within the combustion chamber and 0.59 % unusable due to exergy destruction within the gas turbine. It can be stated that approximately 55 % of the exergy rate of the fuel provided to the engine can not be used by the engine.
34 citations
TL;DR: 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 to maximize loiter duration of UAV or specific power of engine or both.
Abstract: Abstract 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.
21 citations
TL;DR: In this article, the authors analyzed the development of the passage vortex and the horseshoe vortex in a compressor cascade based on delayed detached eddy simulation (DDES) results and explored the effect of passage vortex on corner separation using the RANS method.
Abstract: Abstract The complex flow structures in a linear compressor cascade have been investigated under different incidences using both the Reynolds-averaged Navier–Stokes (RANS) and delayed detached eddy simulation (DDES) methods. The current study analyzes the development of horseshoe vortex and passage vortex in a compressor cascade based on DDES results and explores the effect of the passage vortex on corner separation using the RANS method. Results show that the effect of horseshoe vortex on three-dimensional corner separation is weak, whereas the effect of passage vortex is dominant. A large vortex breaks into many small vortices in the corner separation region, thereby resulting in strong turbulence fluctuation. The passage vortex transports the low-energetic flow near the endwall to the blade suction surface and enlarges corner separation in the cascade. Hence, total pressure loss increases in the cascade.
19 citations
TL;DR: In this article, the effect of fuel spray cone angle in an aerogas turbine combustor has been studied using computational fluid dynamics (CFD) and full-scale combustor testing.
Abstract: Abstract Effect of fuel spray cone angle in an aerogas turbine combustor has been studied using computational fluid dynamics (CFD) and full-scale combustor testing. For CFD analysis, a 22.5° sector of an annular combustor is modeled and the governing equations are solved using the eddy dissipation combustion model in ANSYS CFX computational package. The analysis has been carried out at 125 kPa and 303 K inlet conditions for spray cone angles from 60° to 140°. The lean blowout limits are established by studying the behavior of combustion zone during transient engine operation from an initial steady-state condition. The computational study has been followed by testing the practical full-scale annular combustor in an aerothermal test facility. The experimental result is in a good agreement with the computational predictions. The lean blowout fuel–air ratio increases as the spray cone angle is decreased at constant operating pressure and temperature. At higher spray cone angle, the flame and high-temperature zone moves upstream close to atomizer face and a uniform flame is sustained over a wide region causing better flame stability.
18 citations
TL;DR: In this paper, an integrated model including inlet, engine and nozzle with their internal and external characteristics was built to simulate the propulsion installed performance, and a new performance seeking control scheme under supersonic state was firstly proposed, taking inlet ramp angle as optimizing variable, which is equally important to fuel flow rate, nozzle throat area, guided vane angle of fan and compressor.
Abstract: Abstract An integrated model including inlet, engine and nozzle with their internal and external characteristics was built to simulate the propulsion installed performance. With the integrated model, a new performance seeking control scheme under supersonic state is firstly proposed, taking inlet ramp angle as optimizing variable, which is equally important to fuel flow rate, nozzle throat area, guided vane angle of fan and compressor. Specially, engine installed thrust replaces its total thrust as one crucial factor for performance seeking control. Installed performances under supersonic state are significantly improved with the new scheme, as installed thrust increases of up to 4.9% in the maximum thrust mode, installed specific fuel consumption improvements of up to 3.8% in the minimum fuel consumption mode, and turbine temperature decreases of up to 0.6% in the minimum turbine temperature mode. The simulation results also indicates that, the performance seeking control scheme proposed shows superiority in restraining of the increasing of rotational speed and turbine temperature in performance seeking control.
15 citations
TL;DR: In this paper, the relationship between load spectra and the fatigue life of main components of the aero-engines is obtained based on distribution function and a generalized stress-strength interference model.
Abstract: Abstract Load spectra of aero-engines reflect the process of operating aircrafts as well as the changes of parameters of aircrafts. According to flight hours and speed cycle numbers of the aero-engines, the relationship between load spectra and the fatigue life of main components of the aero-engines is obtained. Based on distribution function and a generalized stress–strength interference model, the cumulative fatigue damage of aero-engines is then calculated. After applying the analysis of load spectra and the cumulative fatigue damage theory, the fatigue life of the first-stage turbine disks of the aero-engines is evaluated by using the S–N curve and Miner’s rule in this paper.
15 citations
TL;DR: In this paper, the authors used the standard k-ε model (SKE) and the Reynolds stress model (RSM) to predict the tip leakage flow in a low-speed large-scale axial compressor rotor.
Abstract: Abstract The standard k-ε model (SKE) and the Reynolds stress model (RSM) are employed to predict the tip leakage flow (TLF) in a low-speed large-scale axial compressor rotor. Then, a new research method is adopted to “freeze” the turbulent kinetic energy and dissipation rate of the flow field derived from the RSM, and obtain the turbulent viscosity using the Boussinesq hypothesis. The Reynolds stresses and mean flow field computed on the basis of the frozen viscosity are compared with the results of the SKE and the RSM. The flow field in the tip region based on the frozen viscosity is more similar to the results of the RSM than those of the SKE, although certain differences can be observed. This finding indicates that the non-equilibrium turbulence transport nature plays an important role in predicting the TLF, as well as the turbulence anisotropy.
14 citations
TL;DR: In this article, a real-time optimization control method is proposed to extend turbo-fan engine service life, which is based on an on-board engine mode devised by a MRR-LSSVR (multi-input multi-output recursive reduced least squares support vector regression method).
Abstract: Abstract A real-time optimization control method is proposed to extend turbo-fan engine service life. This real-time optimization control is based on an on-board engine mode, which is devised by a MRR-LSSVR (multi-input multi-output recursive reduced least squares support vector regression method). To solve the optimization problem, a FSQP (feasible sequential quadratic programming) algorithm is utilized. The thermal mechanical fatigue is taken into account during the optimization process. Furthermore, to describe the engine life decaying, a thermal mechanical fatigue model of engine acceleration process is established. The optimization objective function not only contains the sub-item which can get fast response of the engine, but also concludes the sub-item of the total mechanical strain range which has positive relationship to engine fatigue life. Finally, the simulations of the conventional optimization control which just consider engine acceleration performance or the proposed optimization method have been conducted. The simulations demonstrate that the time of the two control methods from idle to 99.5 % of the maximum power are equal. However, the engine life using the proposed optimization method could be surprisingly increased by 36.17 % compared with that using conventional optimization control.
14 citations
TL;DR: The optimization results indicated that the rotor vibration reached the minimum value and the design efficiency and quality were improved by the multidisciplinary design optimization in the case of meeting the design requirements, which provided the reference to improve the design Efficiency and reliability of the aero-engine rotor.
Abstract: Abstract An integrated analytical method based on multidisciplinary optimization software Isight and general finite element software ANSYS was proposed in this paper. Firstly, a two-disk rotor system was established and the mode, humorous response and transient response at acceleration condition were analyzed with ANSYS. The dynamic characteristics of the two-disk rotor system were achieved. On this basis, the two-disk rotor model was integrated to the multidisciplinary design optimization software Isight. According to the design of experiment (DOE) and the dynamic characteristics, the optimization variables, optimization objectives and constraints were confirmed. After that, the multi-objective design optimization of the transient process was carried out with three different global optimization algorithms including Evolutionary Optimization Algorithm, Multi-Island Genetic Algorithm and Pointer Automatic Optimizer. The optimum position of the two-disk rotor system was obtained at the specified constraints. Meanwhile, the accuracy and calculation numbers of different optimization algorithms were compared. The optimization results indicated that the rotor vibration reached the minimum value and the design efficiency and quality were improved by the multidisciplinary design optimization in the case of meeting the design requirements, which provided the reference to improve the design efficiency and reliability of the aero-engine rotor.
9 citations
TL;DR: In this article, numerical simulation is performed for the three-dimensional turbulent flow field in a centrifugal pump by solving the Reynolds-averaged Navier-Stokes equations and the RNG k-epsilon turbulent model.
Abstract: Abstract Numerical simulation is performed for the three-dimensional turbulent flow field in a centrifugal pump by solving the Reynolds-averaged Navier-Stokes equations and the RNG k-epsilon turbulent model. The finite volume method and the SIMPLE algorithm are employed for the solution of the system. All the parameters in the centrifugal pump at different blade angular positions are obtained by simulation. The flow structure is analyzed and the distributions of the energy gradient function K are calculated at different blade angular positions based on the energy gradient method. According to the energy gradient method, the location which has larger value of K is easier to cause instability and to be of high turbulence intensity. The result shows that the flow instability is easier to be excited nearing the tongue where the value of K is large. The unstable flow area nearing the tongue is also in agreement with the zone where the velocity decreases rapidly. The sudden variation of velocity contributes to the large value of K. The research result also indicates that the tongue has large impact only on the impeller passages passing the tongue.
9 citations
TL;DR: The coupled problem of VCE (Variable Cycle Engine) conceptual design is solved by utilizing improvedCSSO strategy, and the design parameter given by improved CSSO strategy is better than the original one and the integrated performance of V CE is significantly improved.
Abstract: Abstract In order to obtain better integrated performance of aero-engine during the conceptual design stage, multiple disciplines such as aerodynamics, structure, weight, and aircraft mission are required. Unfortunately, the couplings between these disciplines make it difficult to model or solve by conventional method. MDO (Multidisciplinary Design Optimization) methodology which can well deal with couplings of disciplines is considered to solve this coupled problem. Approximation method, optimization method, coordination method, and modeling method for MDO framework are deeply analyzed. For obtaining the more efficient MDO framework, an improved CSSO (Concurrent Subspace Optimization) strategy which is based on DOE (Design Of Experiment) and RSM (Response Surface Model) methods is proposed in this paper; and an improved DE (Differential Evolution) algorithm is recommended to solve the system-level and discipline-level optimization problems in MDO framework. The improved CSSO strategy and DE algorithm are evaluated by utilizing the numerical test problem. The result shows that the efficiency of improved methods proposed by this paper is significantly increased. The coupled problem of VCE (Variable Cycle Engine) conceptual design is solved by utilizing improved CSSO strategy, and the design parameter given by improved CSSO strategy is better than the original one. The integrated performance of VCE is significantly improved.
TL;DR: In this article, the authors present dynamic leakage rate and coupled interaction for variable speed rotor-labyrinth (LABY) seal, with rotating speed from 18 to 30 krpm.
Abstract: Abstract The following paper presents dynamic leakage rate and coupled interaction for variable speed rotor-labyrinth (LABY) seal, with rotating speed from 18 to 30 krpm. Variable speed rotor vibration characteristics are incorporated into transient computational fluid dynamic (CFD) calculations as boundary conditions of seal flow field to show the real-time effect of rotordynamic in seal flow field. Leakage rate across a variable speed rotor-seal increases with rotor vibration, but this effect is prominent at lower speed than at higher speed. Leakage characteristic is determined by differences in rotor vibration amplitude rather than rotating speed. The results also reveal that aerodynamic forces of labyrinth seal flow field can improve rotor stability, and this interaction between rotor and seal decreases with the increase of rotating speed.
TL;DR: In this paper, the effect of inlet clearance on the performance of a centrifugal blower, with parallel wall volute, over its full operating range was reported. But, the performance was not analyzed in terms of efficiency and specific work for all mass flow rates.
Abstract: Abstract 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.
TL;DR: In this article, 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.
Abstract: Abstract 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.
TL;DR: The comparison result between the simulation result and the experimental data shows that, the full-range model built by the computation method of low-speed component characteristics is of a certain accuracy, which can meet the needs of a turboshaft engine semi-physical simulation.
Abstract: Abstract In this paper, an approximate computation method of low-speed component characteristics in aeroengine is used and full-range component characteristics is obtained by combining experimental data above idle. Moreover, based on components matching method and variable specific heat method, a full-range static and dynamic mathematical model of turboshaft engine is built, including start-up state. And the numerical simulation result of the engine whole working process is also showed in this paper. The comparison result between the simulation result and the experimental data shows that, the full-range model built by the computation method of low-speed component characteristics is of a certain accuracy, which can meet the needs of a turboshaft engine semi-physical simulation.
TL;DR: In this article, the efficiency of tab of two geometries in promoting the mixing of a Mach 2 elliptic jet has been studied, and it was found that the triangular tab is a better mixing promoter than the crosswire.
Abstract: Abstract The efficiency of tabs of two geometries in promoting the mixing of a Mach 2 elliptic jet has been studied. Limiting tab of triangular and circular geometry (crosswire) of 5 % blockage placed along major and minor axis at the nozzle exit, are tested for nozzle pressure ratio from 4 to 8, in steps of one. Both tabs are efficient mixing promoters, at all the tested NPRs, when placed along the minor axis. But along major axis the crosswire retards the mixing, at all the NPRs. The triangular tab along the major axis is also found to retard the mixing at NPRs 4 and 5, but for nozzle pressure ratios above 5 it causes mixing enhancement even when placed along the major axis. The triangular tab is found to be a better mixing promoter than the crosswire. The maximum core length reduction of 88 % is caused by triangular tab along the minor axis is at NPR4. The corresponding core length reduction for the crosswire is only 72 %. Shadowgraph pictures of controlled jets show that both tabs weaken the waves in jet core. The geometry and orientation of the tab and the expansion level influence the mixing caused by the tab.
TL;DR: In this article, the authors investigated the thrust augmentation and acoustic performance of a Pulse Detonation Engine (PDE) with ejector system, and found that the divergent ejector performed the best based on thrust augmentation, while the reduction effect for OASPL and Peak Sound Pressure Level (PSPL) at 60° was most prominent for the convergent ejectors.
Abstract: Abstract Thrust augmentation and acoustic performance of a Pulse Detonation Engine (PDE) with ejector system is experimentally investigated. For these tests the LEjector/DEjector is varied from 1.18 to 4 and the axial placement of the ejector relative to the PDE exhaust is varied from an x/DPDE of −3 to 3. Results from the tests show that the optimum LEjector/DEjector based on thrust augmentation and Overall Sound Pressure Level (OASPL) is found to be 2.61. The divergent ejector performed the best based on thrust augmentation, while the reduction effect for OASPL and Peak Sound Pressure Level (PSPL) at 60° is most prominent for the convergent ejector. The optimum axial position based on thrust augmentation is determined to be x/DPDE = 2, while, x/DPDE = 0 based on OASPL and PSPL.
TL;DR: To reduce the conservativeness caused by a single LPV controller of the whole flight envelope and the common Lyapunov function method, a new method is proposed to design a family of LPV switching controllers.
Abstract: Abstract To cover the whole range of operating conditions of aero-engine, a double-layer LPV model is built so as to take into account of the variability due to the flight altitude, Mach number and the rotational speed. With this framework, the problem of designing LPV state-feedback robust controller that guarantees desired bounds on both H∞$${H_\infty}$$ and H2$${H_2}$$ performances is considered. Besides this, to reduce the conservativeness caused by a single LPV controller of the whole flight envelope and the common Lyapunov function method, a new method is proposed to design a family of LPV switching controllers. The switching LPV controllers can ensure that the closed-loop system remains stable in the sense of Lyapunov under arbitrary switching logic. Meanwhile, the switching LPV controllers can ensure the parameters change smoothly. The validity and performance of the theoretical results are demonstrated through a numerical example.
TL;DR: In this article, the failure mode is found to be fatigue with multiple cracks inside the turbine blade having crack origin at metal carbides, which may be defined as inclusions in the form of highly elongated along deformation direction.
Abstract: Abstract This paper deals with failure analysis of a low-pressure turbine blade of a straight flow turbojet engine. The blade is made of a wrought precipitation hardened Nickel base superalloy with oxidation-resistant diffusion aluminizing coating. The failure mode is found to be fatigue with multiple cracks inside the blade having crack origin at metal carbides. In addition to the damage in the coating, carbide banding has been observed in few blades. Carbide banding may be defined as inclusions in the form of highly elongated along deformation direction. The size, shape and banding of carbides and their location critically affect the failure of blades. Carbon content needs to be optimized to reduce interdendritic segregation and thereby provide improved fatigue and stress rupture life. Hence, optimization of size, shape and distribution of carbides in the billet and forging parameters during manufacturing of blade play a vital role to eliminate/reduce extent of banding. Reference micrographs as acceptance criteria are essential for evaluation of raw material and blade. There is a need to define the acceptance criteria for carbide bandings and introduce more sensitive ultrasonic check during billet and on finished blade inspection.
TL;DR: In this paper, the authors present a self-study of the JES-STVS revolution, where STVS stands for wing-engine-airframe-integrated, embedded stealthy-jet engine-inlets, restructured engines inside Stealth, Tailless, canard-less, Thrust Vectoring IFPC Systems.
Abstract: Abstract To stay relevantly supported, jet-engine researchers, designers and operators should follow changing uses of small and large jet engines, especially those anticipated to be used by/in the next generation, JET-ENGINE-STEERED (“JES”) fleets of jet drones but fewer, JES-Stealth-Fighter/Strike Aircraft. [1–19, Figure 5 on WIN-WIN VOLL-ROLL-TARGETING & APPENDIX]. In addition, some diminishing returns from isolated, non-integrating, jet-engine component studies, vs. relevant, supersonic, shock waves control in fluidic-JES-side-effects on compressor stall dynamics within Integrated Propulsion Flight Control (“IPFC”), and/or mechanical JES, constitute key relevant methods that currently move to China, India, South Korea and Japan. [4, 5, 20–31]. The central roles of the jet engine as primary or backup flight controller also constitute key relevant issues, especially under post stall conditions involving induced engine-stress while participating in crash prevention or minimal path-time maneuvers to target. [32–45]. And when proper instructors are absent, self-study of the JES-STVS REVOLUTION is an updating must, where STVS stands for wing-engine-airframe-integrated, embedded stealthy-jet-engine-inlets, restructured engines inside Stealth, Tailless, canard-less, Thrust Vectoring IFPC Systems. [4, 5]. Anti-terror and Airliners Super-Flight-Safety are anticipated to overcome US legislation red-tape that obstructs JES-add-on-emergency-kits-use. [1, 2, 13, 16, 17, 19, 32, 33, 36, 39–41, 43–45].
TL;DR: In this article, an integrated performance-CFD-chemical reactor network (CRN) methodology was developed to investigate the impact of steam injection on combustion and NOx emissions, and the established CRN model was then used for steam addition study and the results showed under air/steam mixture atmosphere, high steam content could push the NOx formation region to the post-flame zone and a large amount of NOx emission could be reduced when the steam mass fraction is quite high.
Abstract: Abstract The steam injection technology for aircraft engines is gaining rising importance because of the strong limitations imposed by the legislation for NOx reduction in airports. In order to investigate the impact of steam addition on combustion and NOx emissions, an integrated performance-CFD-chemical reactor network (CRN) methodology was developed. The CFD results showed steam addition reduced the high temperature size and the radical pool moved downstream. Then different post-processing techniques are employed and CRN is generated to predict NOx emissions. This network consists of 14 chemical reactor elements and the results were in close agreement with the ICAO databank. The established CRN model was then used for steam addition study and the results showed under air/steam mixture atmosphere, high steam content could push the NOx formation region to the post-flame zone and a large amount of the NOx emission could be reduced when the steam mass fraction is quite high.
TL;DR: In this paper, a pusher-type Light Transport Aircraft (LTA) with twin turbo prop engines mounted at the aft end of the fuselage was evaluated with and without an oil cooler duct attached to the nacelle.
Abstract: Aircraft engines need a cooling system to keep the engine oil well within the temperature limits for continuous operation. The aircraft selected for this study is a typical pusher type Light Transport Aircraft (LTA) having twin turbo prop engines mounted at the aft end of the fuselage. Due to the pusher propeller configuration, effective oil cooling is a critical issue, especially during low-speed ground operations like engine idling and also in taxiing and initial climb. However, the possibility of utilizing the inflow induced by the propeller for oil cooling is the subject matter of investigation in this work. The oil cooler duct was designed to accommodate the required mass flow, estimated using the oil cooler performance graph. A series of experiments were carried out with and without oil cooler duct attached to the nacelle, in order to investigate the mass flow induced by the propeller and its adequacy to cool the engine oil. Experimental results show that the oil cooler positioned at roughly 25 % of the propeller radius from the nacelle center line leads to adequate cooling, without incorporating additional means. Furthermore, it is suggested to install a NACA scoop to minimize spillage drag by increasing pressure recovery.
TL;DR: Dynamic study both in a simplified and a multi-degree of freedom, comprehensive helicopter model is done to reveal the law of torsional vibration and a simple filter is designed with some simulations.
Abstract: Abstract With both the advantages like attacking close targets and the disadvantages especially like dynamic coupling, helicopter deserves more investigations these days. This paper did dynamic study both in a simplified and a multi-degree of freedom, comprehensive helicopter model, so that to reveal the law of torsional vibration. In the simplified model, the law how arbitrary parameter affects the first-order vibration mode, is discussed. Then, the validation is done in a multi-degree of freedom model by means of the fast Fourier transformation (FFT) method. In this case, how the low-frequency vibration mode relates with the first-order vibration mode is clearly presented, as well as the research direction to design a filter. Lastly, a simple filter is designed with some simulations.
TL;DR: A rapid prototyping controller with fault-tolerant control ability based on NI CompactRIO platform is designed and verified on the semi-physical simulation test platform and shows that the controller can realize the aero engine control safely and reliably with little influence on controller performance in the event of fault on sensor.
Abstract: Abstract A new fault-tolerant control method for aero engine is proposed, which can accurately diagnose the sensor fault by Kalman filter banks and reconstruct the signal by real-time on-board adaptive model combing with a simplified real-time model and an improved Kalman filter. In order to verify the feasibility of the method proposed, a semi-physical simulation experiment has been carried out. Besides the real I/O interfaces, controller hardware and the virtual plant model, semi-physical simulation system also contains real fuel system. Compared with the hardware-in-the-loop (HIL) simulation, semi-physical simulation system has a higher degree of confidence. In order to meet the needs of semi-physical simulation, a rapid prototyping controller with fault-tolerant control ability based on NI CompactRIO platform is designed and verified on the semi-physical simulation test platform. The result shows that the controller can realize the aero engine control safely and reliably with little influence on controller performance in the event of fault on sensor.
TL;DR: In this article, the authors proposed a new method that simulates the engine starting by replacing the non-linear thermodynamic model with a linear dynamic model, which is supplemented with a simplified static model.
Abstract: The steady-state and transient engine performances of gas turbine control system development are usually evaluated by applying full thermodynamic engine models. Most models only address the operating range between the idle and maximum power points, but more recently, they also address a sub-idle operating range.The lack of information about the component maps at the sub-idle modes creates major challenges for the starting system and control system designers. A common method to cope with the problem extrapolates the performances of the engine components to the sub-idle operation range. Precise extrapolation is a challenge to be studied by many scientists. As a rule, many scientists are only concerned about 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 known reports about a model that considers all of these mentioned aspects and simulates the engine starting.To synthesize a thermodynamic model of starting, most known methods require the performance of the components in the sub-idle range. The proposed paper addresses a new method that simulates the engine 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, the static model consists of simplified relations between the gas path parameters and the corrected rotational speed.The paper also describes an algorithm for model synthesis and its practical application to real data.Copyright © 2014 by ASME
TL;DR: In this article, the authors investigated the performance of plasma-assisted ignition in a propane/air mixture and found that a shorter ignition delay 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%.
Abstract: Abstract 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.
TL;DR: In this article, a simplified pin/disc model was conducted to simulate the practical working condition under applied loads using finite element (FE) analysis, and life prediction for the turbine fan disc was conducted by combining the material S-N curve, mean stress effects and concentration stress factor obtained by means of FE method.
Abstract: Abstract Fan discs are critical components of an aero engine. In this paper, contact stress and life prediction of a turbine fan disc were investigated. A simplified pin/disc model was conducted to simulate the practical working condition under applied loads using finite element (FE) analysis. This study is devoted to examining the effects of interface condition of pin/disc such as gap and coefficient upon the maximum stress. The FE model indicated that the maximum stress occurs at the top right corner in the second pin hole, and larger gap or friction coefficient has a significant effect on the maximum stress. In addition, FE analysis without considering friction is also conducted. The results show that the dangerous point is similar to the result which considers friction and the stress state is relatively larger than that of considering friction. Finally, based on FE analysis result, life prediction for the fan disc is conducted to combine the material S-N curve, mean stress effects and concentration stress factor obtained by means of FE method.
TL;DR: In this article, a hydrogen-fueled dual-mode scramjet combustor was investigated experimentally, where room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted.
Abstract: Abstract In this work, a hydrogen-fueled dual-mode scramjet combustor was investigated experimentally. Clean and dry air was supplied to the combustor through a Mach 2 nozzle with a total temperature of 800 K and a total pressure of 800 kPa. The high enthalpy air was provided by an electricity resistance heater. Room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted. Except wall pressure profiles, velocity and temperature profiles in and at exit of the combustor were also measured using hydroxyl tagging velocimetry (HTV) and tunable diode laser absorption spectroscopy (TDLAS), respectively. Results showed that combustion occurred mainly at the bottom side of the combustor. And there were also an extreme disparity of the velocity and temperature profiles along the Y-direction, i.e. the transverse direction.
TL;DR: The integrated optimization design method in which uniform design, response surface methodology and genetic algorithm are used in combination performs well in improving the duct aerodynamic performance and can be also applied to wider fields of mechanical design.
Abstract: Abstract This paper presents an integrated optimization design method in which uniform design, response surface methodology and genetic algorithm are used in combination. In detail, uniform design is used to select the experimental sampling points in the experimental domain and the system performance is evaluated by means of computational fluid dynamics to construct a database. After that, response surface methodology is employed to generate a surrogate mathematical model relating the optimization objective and the design variables. Subsequently, genetic algorithm is adopted and applied to the surrogate model to acquire the optimal solution in the case of satisfying some constraints. The method has been applied to the optimization design of an axisymmetric diverging duct, dealing with three design variables including one qualitative variable and two quantitative variables. The method of modeling and optimization design performs well in improving the duct aerodynamic performance and can be also applied to wider fields of mechanical design and seen as a useful tool for engineering designers, by reducing the design time and computation consumption.
TL;DR: In this article, an experimental study of the recirculation zone and mixing lengths for bluff-body stabilized flames is conducted at non-reactive conditions and the auxiliary turbulence created from the wall of the combustor is also studied and maintained to levels as low as 5%.
Abstract: Abstract An experimental study of the recirculation zone and mixing lengths for bluff-body stabilized flames is conducted at non-reactive conditions. This paper reports the prediction of recirculation zone length from dynamic pressure measurements. The auxiliary turbulence created from the wall of the combustor is also studied and maintained to levels as low as 5%. The experiments are conducted by varying the velocity from 5 m/s to 8 m/s for V-Gutters bluff-body with induced angles of 60, 90 and 120o. These gutters are maintained at same blockage ratio so that gutter angle to flow velocity is studied. It is inferred from the experiment that as the velocity in the duct increases, the length of the recirculation zone varies 5 mm for all V-Gutter angle. However, an increase in the V-gutter angle is observed to greater effect than an increase in the velocity, recirculation zone length which varied from 70 mm for 60o V-gutter to 150 mm for 120o V-gutter. Simultaneously a sharp reduction in shear distribution along the length of the combustor are observed, it influences in understanding the mixing characteristics in combustion.