The small gap at stator hub section of 10-stage axial compressor of small power class industrial gas turbine engine was studied to confirm its effect on compressor analysis result. This gap is allowed for manufactural tolerance and thermal expansion during engine operation. For the convenient purpose of CFD geometric modeling, such gap was simplified and the 3D Navier-Stokes code was used to predict the compressor performance then compared the results with the case without a gap. In the case of calculation without a gap, the performance was estimated to be lower than that of test result. It is because of the presence of 3D separation at hub corner of every stator except on the 1st and 2nd stator. The CFD calculation shows that, with a gap, the stall observed at hub corner vanished and the predicted compressor performance agrees well with the test result. From this, it is concluded that the existence of a gap between inner casing and stator brings a considerable effects on the compressor flow distribution and must be taken into account in the design.
Gas turbine engine has been applied to the aircraft and ship propulsion with its advantages of compactness and comparatively short starting time. With a significant improvement in gas turbine efficiency with development of super alloy materials and advancement in cooling technologies in the second half of 1990s, its importance as a source of base load as well as peak load power generation has been increasing. However, with increased demand in nuclear power and renewable energy in the 21st century, there seems to be speculations among the power generation industries that gas turbine will take more or less a buffering role supplementing the irregular inflow of electricity to the grid rather than acting as a base load power source. With the shift in the role of gas turbine from base to supplementary, CHP application based on small powered gas turbine utilizing biogas or syngas as its fuel is expected to increase in the future. In this context, this paper describes the development result of 5MW gas turbine engine for CHP application. It can be operated with LNG or syngas of low LHV fuel. Originally, the engine was designed for LNG as its primary fuel, but since the importance of syngas power generation market will be increasing in the future, a complementary work for modification of combustor part has been carried out and has been tested. However, this paper deals with the parts developed with the use of LNG fuel. The test result of emission characteristics meets the standards required in Korea. The development has been made through the cooperation of Doosan Heavy Industry (DHI, Korea) and Zory-Mashproekt (Ukraine).
Overall efficiency of an axial compressor is largely affected by its front stage when it is operating under transonic flow conditions. For this reason, many manufacturers and researchers are advancing research and development of transonic airfoils in these days. Doosan, in frame of a development of high efficiency gas turbine, developed high efficiency airfoil for a transonic rotor and conducted cascade tests. Therefore, this study deals with a test of two compressor transonic blade cascades at inlet Mach number over 1.1. To improve the efficiency and operating range, two kinds of thickness distribution type based on Enhanced Doosan Airfoil (EDA), which applied unique rule, were applied and assessed. The first airfoil consists of polynomial thickness distribution and the second airfoil consists of new thickness distribution with specially tailored leading edge. In order to ensure accurate geometry of a model, a detailed checkout process upon production of model blades used in the test was performed. This is because, in the case of transonic airfoil, if the inlet leading edge shape differs by more than 0.2% than designed airfoil of leading edge, the result will be completely different. Therefore, not only the tolerance within 0.1% was confirmed but also the shape produced through simulation and 3D CMM scan data. The main parameters for the comparison are an inlet Mach number, an axial velocity density ratio (AVDR) and the kind of thickness distribution. Results of tests and CFD blade to blade analysis using MISES 2.70 are compared. The flow field was visualized using schlieren technique and parameters of the suction side boundary layer were evaluated at several locations based on Pitot probe traverses. The results confirm that a suction peak at the round leading edge disappears in the case of the new thickness type distribution with tailored leading edge. This confirms that the profile shaping without jump in curvature in the leading edge region leads to smooth acceleration without peaks. Nevertheless, results show that the new thickness distribution type is not absolutely good in comparison with the polynomial thickness distribution type with respect to the total pressure loss coefficient. Moreover, bucket range (operating range) is also almost the same. Results of the suction side boundary layer traversing suggest that the transition of the boundary layer takes place beyond the location x/cax > 0.088. The MISES results show that a shock location and the boundary layer parameters are similar to test results. However, values of the loss coefficient show some difference. Therefore, a new correlation in particular transonic flow condition was developed.
Periodic unsteady flow kinematics in a shrouded multistage low-speed axial compressor has been measured for the first time. Data have been acquired at the inlet and exit of a shrouded 3rd-stage stator with a particular focus on the hub flows. The newly found features of the hub flow in a shrouded multistage compressor are different from those at the midspan or in unshrouded (i.e., cantilevered) compressors. First, the merging of the 2nd-stage stator and 3rd-stage rotor wakes causes positive radial migration near the rotor wake pressure surface at the hub of the 3rd-stage stator inlet. Second, the low-momentum labyrinth seal leakage flow of the 3rd-stage stator merges with the 3rd-stage rotor wake to create streamwise vorticity at the 3rd-stage stator inlet hub. Third, contrary to unshrouded stators, suction side hub corner separation in the shrouded 3rd-stage stator reduces rotor wake stretching. Thus, velocity disturbances are attenuated less, and amplitudes of periodic fluctuations in flow angles are larger at the 3rd-stage stator exit hub than at midspan. The positive radial migration of the rotor wake hub flow and wake stretching reduction are expected to decrease efficiency, whereas streamwise vorticity generation is expected to increase efficiency.
Periodic unsteady flow kinematics in a shrouded multistage low-speed axial compressor has been measured for the first time. Data have been acquired at the inlet and exit of a shrouded 3rd- stage stator with a particular focus on the hub flows. The newly found features of the hub flow in a shrouded multistage compressor are different from those at the midspan or in unshrouded (i.e., cantilevered) compressors. First, the merging of the 2nd-stage stator and 3rd-stage rotor wakes causes positive radial migration near the rotor wake pressure surface at the hub of the 3rd-stage stator inlet. Second, the low-momentum labyrinth seal leakage flow of the 3rd-stage stator merges with the 3rd-stage rotor wake to create streamwise vorticity at the 3rd-stage stator inlet hub. Third, contrary to unshrouded stators, suction side hub corner separation in the shrouded 3rd-stage stator reduces rotor wake stretching. Thus, velocity disturbances are attenuated less, and amplitudes of periodic fluctuations in flow angles are larger at the 3rd-stage stator exit hub than at midspan. The positive radial migration of the rotor wake hub flow and wake stretching reduction are expected to decrease efficiency, whereas streamwise vorticity generation is expected to increase efficiency.
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