• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
• /
• 2003.05a
• /
• pp.297-302
• /
• 2003

To ensure a safe operation of the prototype KGT-74 ㎾ small gas turbine, vibrational reliabilities of the compressor 1st, 2nd. and 3rd stages and turbine blades have been estimated and reviewed. FE analyses have been tarried out to obtain the natural vibration characteristics of the blades, and impact modal testings have been performed on every each one of the blades to measure their 1st natural frequencies. Then, the Campbell diagram analyses have been carried out to Judge the safety of the blades from resonant failures up to 6k harmonics. Results show that the compressor 1st stage blade is exposed to a potential resonant failure with 3k harmonic around a rated speed of 30,000rpm but that the other compressor 2nd and 3rd stages and turbine blades are safe from resonant failures. Finally. 27,900 rpm Is selected as the safe operation limit for the KGT-74 ㎾ gas turbine relative to the blade vibrations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.11a
• /
• pp.94-97
• /
• 2008

As the partial admission turbine has a intrinsically unsteady and three dimensional flow region, numerical calculation time of these study has been too long time. The numerical analysis for gas temperature of turbine blade surface has been performed to investigate development of temperature with various pyro start pressure. Computations have been carried out several turbine rotational speeds in the range from 0 to 16000 rpm and inlet conditions with 1423K, 7.2MPa. As a result, the more rotational speed and pyro starter pressure of turbine increased, the more turbine blade's temperature decreased. It is also found that flow field of turbine blade inlet area at pyro starter pressure of 5.75MPa and rotational speed of 12100 rpm formed surface temperature uniformly.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.12 no.6
• /
• pp.48-55
• /
• 2008

We have simulated the performance of a simple engine model with a gas turbine engine simulation program based on CFD. 2-dimensional Navier-Stokes code for the viscous flow was applied to simulate a compressor and a turbine, and the chemical equilibrium code with the lumped method was applied to simulate the combustor. Unsteady-flow phenomenon between rotor and stator of the compressor and the turbine was analyzed by steady mixing-plane method. In this way, the influence of the turbine blade pitch on the engine was investigated. It was shown that the compressor is operated at more higher pressure conditions as narrower the pitch distance of the turbine.

• Transactions of the Korean Society of Automotive Engineers
• /
• v.24 no.1
• /
• pp.74-81
• /
• 2016

The aim of this study was to find the initial design value of turbine blade for electrical type turbocompound system generating 10 kW. Turbocompound is one of the waste heat recovery system applying to internal combustion engine to recover exhaust gas energy that was about 30 % of total input energy. To design the turbine blade, 1-D mean line flow model was used. Exhaust gas temperature, pressure, flow rate and turbine rotating speed was fixed as primary boundary conditions. The velocity triangles was defined and used to determine the rotor inlet radius and width, the rotor outlet radius at shroud and radius at hub, the rotor flow angles and the number of blades.

• Plant Journal
• /
• v.18 no.4
• /
• pp.48-57
• /
• 2023

In this study, I discussed a way of the improved operation criteria to detect combustion instability in advance F-Class Gas Turbine, which adopts lean pre-mixed combustion system. The data of 16 blades path temperature thermocouple installed radially at the gas Turbine exit were collected to analyze the variation of individual blade path temperature. The cumulative variation in individual blade path temperature for one week under normal combustion conditions was confirmed to be up to 26℃. On the other hand, in the event of combustion instability, the symptoms of increased temperature variations in the individual thermocouple were mostly seen from a few days ago. Based on the results of this study, it is deemed appropriate to inspect and maintain in Ulsan Thermal Power Gas Turbine when the individual blade path temperature exceeds 50℃ of the cumulative variation for 10 days.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.04a
• /
• pp.414-417
• /
• 2011

On the basis of the paper described herein, rotor blade failure in the compressor, gas generator turbine, and power turbine and the resulting internal damage is contained within the peripheral hardware and engine casings. For the safety reason, the blade containment was regulated by aviation authority. For reducing the weight of the case, a heaviest single component of a jet engine, the blade containment capability was analyzed by engine manufacturer. The procedure established for containment design involves an energy balance method based on the comparison of the kinetic energy of released blade and the strain energy of the containment zone. The LS-DYNA simulation can also be introduced to predict behavior of released blade and case. All of the analytic and numerical result are described ${\ldots}$.

• Plant Journal
• /
• v.9 no.4
• /
• pp.31-36
• /
• 2013

In order to extend a hot gas parts replacement cycle of a gas turbine, blade row 1 from low pressure turbine, which has a significant impact on the cycle, has been selected from stored set after one cycle use. Taking into account the status of the first stage moving blade in LP turbine operated more than 27,000 equivalent operating hours(EOH) and the replacement cycle in the same type of gas turbine, the replacement of the high temperature components installed on the GT, a study subject, can be extended from 24,000 to 27,000 EOH.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.542-545
• /
• 2009

In this study, the effect of turbine geometry on the overall performance of a gas turbine was investigated by computational fluid dynamics. Overall engine performance was predicted through a full engine simulation program which can predict the interactions of the compressor, the combustor and the turbine. The compressor and the turbine analysis code solves 2D and 3D Navier-Stokes equations respectively. The chemical equilibrium code was applied to simulate the combustor. The computations were performed for two different shapes of turbine nozzle. The nozzle shapes adopted a baseline blade and a blade with fillet.

• Proceedings of the Korean Society of Precision Engineering Conference
• /
• 2006.10a
• /
• pp.625-626
• /
• 2006

• Proceedings of the Korean Society of Precision Engineering Conference
• /
• 2006.10a
• /
• pp.623-624
• /
• 2006