Jin Taek Chung

Blade tip leakage flow in gas turbines is associated with aerodynamic loss and local heat transfer variation in the tip region. In this study, the flow structure, total pressure loss coefficient, heat transfer coefficient (HTC), and film cooling effectiveness (FCE) were numerically investigated for a plane tip (PLN) and five squealer tip geometries: a conventional squealer tip (SQR), cutback squealer tip (CBS), multi-cavity squealer tip (MCS), triangular-grooved suction-side squealer tip (GSS), and multi-cavity triangular-grooved suction-side squealer tip (MGS). All configurations were compared under the same cascade geometry, tip-clearance condition, and inlet/outlet boundary conditions to examine the geometry-dependent relationship among aerodynamic loss, heat transfer, and film cooling performance. Film cooling was evaluated at blowing ratios of M = 1 and 2 using a camber-line hole arrangement, and the effect of hole rearrangement was further examined at the same blowing ratio and with the same number of cooling holes. The results indicate that the aerodynamic and thermal characteristics of the tip region vary with the leakage-flow path, cavity recirculation, and reattachment behavior formed by each tip geometry. Under the present conditions, SQR showed the lowest downstream total pressure loss coefficient, with a 7.27% reduction relative to PLN, whereas MGS showed the lowest geometry-normalized heat transfer rate among the tested geometries. Increasing the blowing ratio tended to increase FCE, although local cooling performance was affected by high-pressure or reattachment-dominated regions where coolant ejection, surface attachment, or lateral spreading was limited. Compared with the camber-line arrangement, the rearranged hole configuration increased local FCE by up to 29.6% for CBS and 23.3% for MGS at the same blowing ratio. These results may be used as comparative data for evaluating squealer tip geometries and cooling-hole placement during preliminary blade tip cooling design.

To increase film cooling effectiveness levels downstream of film cooling holes, trenches manufactured in the thermal barrier coating can be adopted. The performance of this solution depends on the trench geometrical characteristics, namely its depth and width. A large eddy simulation (LES)-based numerical study has been performed at Korea University to investigate the effects of trench configuration on thermal protection, resulting in a 22% increase in cooling performance compared to the reference case without a trench. The present paper reports the results of an experimental investigation carried out at Bergamo University on a Plexiglass flat plate model with a set of three fan-shaped holes incorporated into an existing wind tunnel, replicating the numerical setup. Pressure-sensitive paint (PSP) technique was used to measure the adiabatic film cooling effectiveness. Besides the standard shaped hole case, the best and worst trench configurations coming from a DoE approach have been examined at various coolant to mainstream blowing ratios M in the range between 0.5 and 3. Cases at M = 1.5 were used to cross-check the prediction capability of the LES numerical simulation for the selected trenched cases. PSP and LES results are in good agreement, also with the literature data. The high depth/low width trench was shown to give an improved performance for M larger than 1.

This study experimentally investigated the effects of cascade inlet velocity on the distribution and the level of the heat transfer coefficient on a gas turbine blade tip. The tests were conducted in a transient turbine test facility at Korea Aerospace University, and three cascade inlet velocities—30, 60, and 90 m/s—were considered. The heat transfer coefficient was measured using the transient IR camera technique with a linear regression method, and both the squealer and plane tips were investigated. The results showed that the overall averaged heat transfer coefficient was generally proportional to the inlet velocity. As the inlet velocity is increased from 30 m/s to 60 m/s and 90 m/s, the heat transfer coefficient increased by 11.4% and 25.0% for plane tip, and 26.6% and 64.1% for squealer tip, respectively. However, the heat transfer coefficient near the leading edge of the squealer tip and the reattachment region of the plane tip was greatly affected by the cascade inlet velocity. Therefore, heat transfer experiments for a gas turbine blade tip should be performed under engine simulating conditions.

In this study, we numerically investigated the effects of design parameters, such as the strut geometry or diffusion angle, on the performance of an industrial turbine exhaust diffuser. Turbine exhaust diffusers are commonly used to change the kinetic energy of exhaust gases from the outlet of turbine stages into the static pressure. The turbine exhaust diffuser investigated in this work consisted of an annular diffuser with five identical struts equally spaced around the front circumference and a conical diffuser with a hub extension at the rear. Four design parameters were considered and several values for each parameter were tested in this study. The aerodynamic performances of the studied diffusers were evaluated according to their pressure recovery coefficients and rates of total pressure loss. Contours for the velocity, pressure, and entropy increase were plotted and compared for the various diffuser shapes. The numerical results showed that the strut thickness and the axially swept angle of the strut significantly influence the aerodynamic performance of the turbine exhaust diffuser, whereas the strut lean angle and the diffuser hade angle are less important.

Gas turbines are widely used in various industrial applications particularly in the aero-propulsion of aircraft engines. To improve the efficiency and thermal performance of gas turbines, a working fluid with a higher temperature at the inlet of the gas turbine engine is required and the higher inlet temperature often exceeds the melting temperature of the material used in the

The demands to apply transonic centrifugal compressor have increased in the advanced gas turbine engines. Various techniques are used to increase the aerodynamic performance of the centrifugal compressor. The effects of the inclined leading edges in diffuser vanes of a transonic centrifugal compressor on the flow-field unsteadiness and noise generation are investigated by solving the compressible, three-dimensional, transient Navier–Stokes equations. Diffuser vanes with various inclination angles of the leading edge from shroud-to-hub and hub-to-shroud are numerically modeled. The results show that the hub-to-shroud inclined leading edge improves the compressor performance (2.6%), and the proper inclination angle is effective to increase the stall margin (3.88%). In addition, in this study, the transient pressure variations and radiated noise prediction at the design operating point of the compressor are emphasized. The influences of the inclined leading edges on the pressure waves were captured in time/space domain with different convective velocities. The pressure fluctuation spectra are calculated to investigate the tonal blade passing frequency (BPF) noise, and it is shown that the applied inclination angles in the diffuser blades are effective, not only to improve the aerodynamic performance and stall margin, but also to reduce the BPF noise (7.6 dB sound pressure level reduction). Moreover, it is found that the diffuser vanes with inclination angles could suppress the separation regions and eddy structures inside the passages of the diffuser, which results in reduction of the overall sound pressure level and the broadband noise radiated from the compressor.

In this study, the influences of the flow cut and axial lift of the impeller on the aerodynamic performance of a transonic centrifugal compressor were analyzed. The flow cut is a method to reduce the flow rate by decreasing the impeller passage height. The axial lift is a method of increasing the impeller passage height in the axial direction, which increases the impeller exit width (B2) and increases the total pressure. A NASA CC3 transonic centrifugal compressor with a backswept angle was used as a base compressor. After applying the flow cut, the total pressure at the target flow rate was lower than the total pressure at the design point due to the increase in the relative velocity at the impeller exit. After applying the axial lift, the total pressure at the design flow rate was increased, which was caused by the reduction in the relative velocity as the passage area at the impeller exit was increased. By applying the flow cut and axial lift methods, it was shown that the variation in relative velocity at the impeller exit has a significant effect on the variation in total pressure. In addition, it was found that the relative velocity at the impeller exit of the target flow rate is maintained similar to the base impeller when the flow cut and the axial lift are combined. Therefore, by combining the flow cut and the axial lift, three transonic centrifugal impellers with flow fractions of 0.7, 0.8, and 0.9 compared to the design flow rate were newly designed.