Volume 28, Issue 4

Recently, the health monitoring system of major gas path components of gas turbine uses mostly the model based method like the Gas Path Analysis (GPA). This method is to find quantity changes of component performance characteristic parameters such as isentropic efficiency and mass flow parameter by comparing between measured engine performance parameters such as temperatures, pressures, rotational speeds, fuel consumption, etc. and clean engine performance parameters without any engine faults which are calculated by the base engine performance model. Currently, the expert engine diagnostic systems using the artificial intelligent methods such as Neural Networks (NNs), Fuzzy Logic and Genetic Algorithms (GAs) have been studied to improve the model based method. Among them the NNs are mostly used to the engine fault diagnostic system due to its good learning performance, but it has a drawback due to low accuracy and long learning time to build learning data base if there are large amount of learning data. In addition, it has a very complex structure for finding effectively single type faults or multiple type faults of gas path components. This work builds inversely a base performance model of a turboprop engine to be used for a high altitude operation UAV using measured performance data, and proposes a fault diagnostic system using the base engine performance model and the artificial intelligent methods such as Fuzzy logic and Neural Network. The proposed diagnostic system isolates firstly the faulted components using Fuzzy Logic, then quantifies faults of the identified components using the NN leaned by fault learning data base, which are obtained from the developed base performance model. In leaning the NN, the Feed Forward Back Propagation (FFBP) method is used. Finally, it is verified through several test examples that the component faults implanted arbitrarily in the engine are well isolated and quantified by the proposed diagnostic system.

Recently, an aero gas turbine health monitoring system has been developed for precaution and maintenance action against faults or performance degradations of the advanced propulsion system which occurs in severe environments such as high altitude, foreign object damage particles, hot and heavy rain and snowy atmospheric conditions. However to establish this health monitoring system, the online condition monitoring program is firstly required, and the program must monitor the engine performance trend through comparison between measured engine performance data and base performance results calculated by base engine performance model. This work aims to develop a GUI type on-line condition monitoring program for the PT6A-67 turboprop engine of a high altitude and long endurance operation UAV using LabVIEW. The base engine performance of the on-line condition monitoring program is simulated using component maps inversely generated from the limited performance deck data provided by engine manufacturer. The base engine performance simulation program is evaluated because analysis results by this program agree well with the performance deck data. The proposed on-line condition program can monitor the real engine performance as well as the trend through precise comparison between clean engine performance results calculated by the base performance simulation program and measured engine performance signals. In the development phase of this monitoring system, a signal generation module is proposed to evaluate the proposed online monitoring system. For user friendly purpose, all monitoring program are coded by LabVIEW, and monitoring examples are demonstrated using the proposed GUI type on-condition monitoring program.

This is a study of the multi-variable optimization of the forward-skewed blade by means of the Genetic Algorithms combined with an improved Hierarchical Fair Competition model. In this optimization scheme, the Bezier curves were used to describe the airfoil profiles, numerical simulations were adopted to estimate the aerodynamic performance of the axial-flow machine and a specific fitness function was put forwards to evaluate the fitness level of every individual. In order to testify the optimization algorithms and to study the effects of the forward-skewed blades on the fan performance, a series of comparative experiments and calculations were carried out to illustrate the performance difference between the optimized and the other three fans.

Recently, an analytical model to calculate the stability of an axial-flow compressor rotor has been presented in the scientific literature. The range of validity of that theoretical characterization was supported by several lemmas and theorems. One of the main results was the definition of a dimensionless coefficient for determining the location of the stability line in the rotor map. In this work the mathematical structure of that solution is studied. As a result of this detailed study, a new stability theorem and a new stability coefficient are obtained. This stability coefficient is an improvement of the previous one since it is physically and mathematically well defined in all the operational points of the compressor map. As a consequence, the new model is able to capture the stall inception for rotors and stators as well as the full characteristic curve (pressure rise versus mass flow rate) including rotating stall and possibly reverse flow. It is proved, as a consequence of the restriction imposed by the Stability Theorem, that each local component (rotor or stator) has its own instability point and its own post-stall characteristic curve. This theoretical criterion for predicting the averaged characteristic curve is in good accord with the experimental data. The stability coefficient is also verified for a compressor stage. Finally, the model is shown to provide an adequate quantitative and qualitative description of the averaged stall line giving a physical explanation of the mechanism involved in the instable region of the compressor map.

In this study, a mixed-type turbine was designed and tested with the double or single stage to improve the specific torque when it operates at a low partial admission rate. The turbine consists of double stages and the outer diameter of its rotor is 108 mm. The turbine rotor blades were designed as an axial-type blade along the mixed flow direction because the partial admission rate was 1.7–2.0% depending on the flow direction. Performance characteristics were measured at the double and single stage rotors to investigate the effect of the second stage on the low partial admission. In addition, when the flow direction was radially inward or outward at the nozzle, turbine performances were studied. In this experiment, the specific power, torque, and total-to-static efficiency were measured at various rotational speeds to compare with the turbine performance according to different operating condition. The tested results showed that the second stage should be adopted to increase the operating torque when the operating rotational speed was less than the critical rotational speed. The specific torque was improved by 7.8% using the second stage at a radially inward flow direction turbine

To accomplish the research goals of capturing the effects of blade row interactions on compressor performance, small changes in performance must be measurable. This also requires axi-symmetric flow so that measuring one passage accurately captures the phenomena occurring in all passages. Thus, uniform inlet flow is a necessity. The original front-driven compressor had non-uniform temperature at the inlet. Additional challenges in controlling shaft speed to within tight tolerances were associated with the use of a viscous fluid coupling. Thus, a new electric motor, with variable frequency drive speed control was implemented. To address the issues with the inlet flow, the compressor is now driven from the rear resulting in improved inlet flow uniformity. This paper presents the design choices of the new layout in addition to the preliminary performance data of the compressor and an uncertainty analysis.

Numerical simulation of two-dimensional transverse water jet into subsonic air stream is presented in this study. Two-dimensional Reynolds Averaged Navier–Stokes equations along with the standard k - ε turbulence model are used to model this problem using a software, FLUENT. Results include deformation of the boundaries of the water jet and flow fields for both water and air. It is found that jet break-up length and the penetration height increase with an increase in the water to air momentum ratio. However, the free-stream Mach number affects these parameters only marginally. In addition, distribution of water to air mass ratio indicates there is a decrease in mass of water as the liquid jet propagates downstream.

An experimental visualization has been carried out to study the dependence of the pressure hill height and the influence zone expanse for flow past rectangular blocks of flat square face and varying length, over a Reynolds number range from 1364 to 4931. It is found that, the pressure hill length and the influence zone expanse decrease with the length to width ratio of the block, up to about L / W = 1, for Reynolds number up to 1586. For higher Reynolds numbers, both H / W and Z / W increase with the model length, till L / W = 1. For L / W more than 1, both H / W and Z / W gradually become independent of L / W . The ratio of Z / H is influenced only marginally by L / W up to 1, and for greater values of L / W , Z / H is almost a constant at all Reynolds numbers of the present study.