Roberto Abram

Forced Convection Heat Transfer in Channels With Rib Turbulators Inclined at 45 deg

Local and average Nusselt numbers and friction factors are presented for rectangular channels with an aspect ratio of 5 and angled rib turbulators inclined at 45 deg with parallel orientations on one and two surfaces of the channel. The convective fluid was air, and the Reynolds number varied from 9000 to 35,500. The ratio of rib height to hydraulic diameter was 0.09, with the rib pitch-to-height ratio equal to 13.33 or 6.66. Experiments were based on the use of heating foils (for the attainment of uniform heat flux condition) and of the steady-state liquid crystal thermography (for the identification of isotherm lines and the reconstruction of local heat transfer coefficient). Local results showed quasiperiodic profiles of Nusselt number in the streamwise direction, whose features were strongly affected by the value of rib pitch and by the spanwise coordinate. For all the investigated geometries a heat transfer augmentation, relative to the fully developed smooth channel, was found; when inclined rib turbulators were placed on two opposite surfaces of the channel, the full-surface Nusselt number was higher (by 10–19%) than that for the one-ribbed wall channel, but pressure drop penalties also increased by a factor of about 3. For both the one- and two-ribbed wall channels, the best heat transfer performance for a constant pumping power, in the explored range of Reynolds number, was generally achieved by the larger rib pitch-to-height ratio (=13.33).

Influence of Purge Flow Injection Angle on the Aerothermal Performance of a Rotor Blade Cascade

This paper is focused on the influence of stator-rotor purge flow injection angle on the aerodynamic and thermal performance of a rotor blade cascade. Tests were performed in a seven-blade cascade of a high-pressure gas turbine rotor at low Mach number (Ma2is = 0.3) under different blowing conditions. A number of fins were installed inside the upstream slot to simulate the effect of rotation on the seal flow exiting the gap in a linear cascade environment. The resulting coolant flow is ejected with the correct angle in the tangential direction. Purge flow injection angle and blowing conditions were changed in order to identify the best configuration in terms of end wall thermal protection and secondary flows reduction. The 3D flow field was surveyed by traversing a five-hole miniaturized pressure probe in a downstream plane. Secondary flow velocities, loss coefficient, and vorticity distributions are presented for the most significant test conditions. Film cooling effectiveness distributions on the platform were obtained by thermochromic liquid crystals (TLC) technique. Results show that purge flow injection angle has an impact on secondary flows development and, thus, on the end wall thermal protection, especially at high injection rates. Passage vortex is enhanced by a negative injection angle, which simulates the real counter rotating purge flow direction.

Design, Validation and Verification of Film Cooling on Gas Turbine Rotor Endwall

This paper describes the design process carried out to introduce film cooling coverage on the hub platform of a first turbine blade in order to protect from hot gas corrosion the trailing edge platform region. The different steps described consist in a design phase, a validation by experimental tests, the production follow up and verification feedbacks. The first stage endwall is a critical region due to secondary flows overheating; critical areas have been identified by 3D thermal analysis and confirmed through damage reports. A design phase has been carried out to determine the cooling holes configuration in terms of position, number, inclination and diameter. A CFD analysis allowed to define the hot gas streamlines near the pressure side fillet region in order to identify the best holes arrangement to ensure a proper thermal coverage; the thermal effectiveness and coverage length has been subsequently verified by means of an experimental activity developed by University of Bergamo. On the basis of the experimental results a 3D thermal analysis of the new holes configuration has highlighted the improvement in terms of local wall temperature reduction. Finally the new film holes have been introduced in the machining cycle of the blade and realized by electrochemical drilling. During a maintenance inspection after 20000 operating hours an endoscope investigation has confirmed the improvement obtained, showing no substantial signs of overheating.

Aero-Thermal Performance of a Rotor Blade Cascade With Stator-Rotor Seal Purge Flow

The present paper investigates the effects of purge flow from a stator-rotor seal gap on the aerodynamic and thermal performance of a rotor blade cascade. Particular attention is paid to thermal results in the leading edge area that is typically difficult to protect. Experimental tests have been performed on a seven-blade cascade of a high-pressure rotor stage of a real gas turbine at low Mach number (Ma2is = 0.3). To simulate the rotational effect in a linear cascade environment, a number of inclined fins have been installed inside the stator-rotor gap, making the coolant flow to exit with the right tangential velocity component. Tests have been carried out at different blowing conditions, with mass flow rate ratios up to 2.0%. Aerodynamic effects of purge flow on secondary flow structures were surveyed by traversing a 5-hole miniaturized pressure probe in a plane 0.08cax downstream of the trailing edge. Film cooling effectiveness distributions on the end wall platform were obtained by using Thermochromic Liquid Crystals technique. Results allowed to investigate the effect of purge flow injection from the upstream gap on the secondary flows development and on the thermal protection capability. Purge flow injection of 1.0% reduced secondary flow losses and was found to effectively protect the front end wall region, up to about 0.5cax downstream of the leading edge. Increasing the purge flow up to 1.5%–2.0% provided a better thermal protection not only stream wise, but also in the region close to the leading edge because of the weakened washing activity of the horseshoe vortex.Copyright

Improving the Film Cooling of a Rotor Blade Platform

This paper shows the results of an experimental activity developed in cooperation between Ansaldo Energia and the Department of Engineering and Applied Science of Bergamo University with the aim of assessing the impact of newly designed holes on the thermal protection of a rotor blade platform. The original rotor blade platform featured 10 cylindrical holes located along the blade pressure side. Moreover, the channel front side was cooled exploiting the seal purge flow exiting the stator to rotor interface gap. The front mid channel, and particularly the region around the inter-platform gap, remained uncooled. To protect this region two sets of cylindrical holes were designed and manufactured on a 7 blade cascade model for experimental verification. Aerodynamic and thermal tests were carried out at low Mach number. To evaluate the interaction of injected flow with secondary flows a 5hole probe was traversed downstream of the trailing edge plane. The thermal behavior was analyzed by using Thermochromic Liquid Crystals technique, so to obtain film cooling effectiveness distributions. The 7-hole configuration coupled with a low blowing ratio of about 1.0 provided the best thermal protection without any impact on the aerodynamic performance.

Computational Predictions of Aero-Thermal Performance of a Turbine Filleted Blade Cascade With Endwall Film Cooling

In this study computational fluid dynamic simulations of a turbine blade with endwall film cooling were compared to measurements of both aerodynamic and thermal performance. The experimental data were collected at low Mach number (Ma2is = 0.3) in a linear cascade arrangement with 7 blades which geometry is typical of first stage high pressure turbine. A junction between the blade hub and the platform is provided by a 3D fillet. Coolant is injected through ten cylindrical holes distributed along the blade pressure side. Coolant to mainstream mass flow ratio was set to assure an inlet blowing ratio of M1 = 2.4 and M1 = 3.2. The simulations were carried out using the Shear Stress Transport (SST) k-turbulence model. Numerical predictions were compared against experimentally measured secondary flows and endwall film cooling effectiveness, at different injection conditions. Simulation results agreed with the experiments for what concerns the general shape and the location of secondary flows. However, some limitations in the modeling were highlighted when going into the details of loss computation and vortex structure. Predictions overestimated both secondary and midspan blade wake losses. Moreover, the effect of the fillet on the aerodynamic flow features was not fully captured. Predicted film cooling results showed the sweeping of coolant across the passage in agreement with experiments even though jets persistency was higher than that measured. Levels of adiabatic effectiveness were generally well simulated.

Influence of Coolant Flow Rate on Aero-Thermal Performance of a Rotor Blade Cascade With Endwall Film Cooling

This paper investigates the influence of coolant injection on the aerodynamic and thermal performance of a rotor blade cascade with endwall film cooling. A 7 blade cascade of a high-pressure-rotor stage of a real gas turbine has been tested in a low speed wind tunnel for linear cascades. Coolant is injected through ten cylindrical holes distributed along the blade pressure side. Tests have been preliminarily carried out at low Mach number (Ma2is=0.3). Coolant-to-mainstream mass flow ratio has been varied in a range of values corresponding to inlet blowing ratios M1 = 0 – 4.0. Secondary flows have been surveyed by traversing a 5-hole miniaturized aerodynamic probe in two downstream planes. Local and overall mixed-out secondary loss coefficient and vorticity distributions have been calculated from measured data. The thermal behaviour has been also analysed by using Thermochromic Liquid Crystals technique, so to obtain film cooling effectiveness distributions. All this information, including overall loss production for variable injection conditions, allow to draw a comprehensive picture of the aero-thermal flow field in the endwall region of a high pressure rotor blade cascade.