The Effects of Axisymmetric Convergent Contouring and Blowing Ratio On Endwall Film Cooling and Vane Pressure Side Surface Phantom Cooling Performance

In this paper, a detailed numerical investigation on the endwall film cooling and vane pressure side surface phantom cooling was performed, at the simulated realistic gas turbine operating conditions (high inlet freestream turbulence level of 16 %, exit Mach number of 0.85 and exit Reynolds number of 1.7×106). Based on a double coolant temperature model, a novel numerical method for the predictions of adiabatic wall film cooling effectiveness was proposed. This numerical method was validated by comparing the predicted results with experimental data of endwall Nusselt number, endwall film cooling effectiveness and near endwall flow visualization. The results indicate that the present numerical method can accurately predict endwall thermal load distributions and endwall film cooling distributions, and vane surface phantom cooling distributions. The endwall heat transfer coefficient, endwall film cooling effectiveness, phantom cooling effectiveness of the vane pressure side surface and total pressure loss coefficients (TPLC) were predicted and compared for two endwall contouring shapes (flat endwall and axisymmetric convergent contoured endwall) at three different blowing ratios (low blowing ratio of BR=1.0, design blowing ratio of BR=2.5 and high blowing ratio of BR=3.5) with a constant density ratio of DR=1.2, based on the present novel numerical method.

Numerical investigations of endwall film cooling design of a turbine vane using four-holes pattern

Purpose Endwall film cooling protects vane endwall by coolant coverage, especially at the leading edge (LE) region and vane-pressure side (PS) junction region. Strong flow impingement and complex vortexaa structures on the vane endwall cause difficulties for coolant flows to cover properly. This work aims at a full-scale arrangement of film cooling holes on the endwall which improves coolant efficiency in the LE region and vane-PS junction region. Design/methodology/approach The endwall film holes are grouped in four-holes constructal patterns. Three ways of arranging the groups are studied: based on the pressure field, the streamlines or the heat transfer field. The computational analysis is done with the k-ω SST model after validating the turbulence model properly. Findings By clustering the film cooling holes in four-holes patterns, the ejection of the coolant flow is stronger. The four-holes constructal patterns also improve the local coolant coverage in the “tough” regions, such as the junction region of the PS and the endwall. The arrangement based on streamlines distribution can effectively improve the coolant coverage and the arrangement based on the heat transfer distribution (HTD) has benefits by reducing high-temperature regions on the endwall. Originality/value A full-scale endwall film cooling design is presented considering interactions of different film cooling holes. A comprehensive model validation and mesh independence study are provided. The cooling holes pattern on the endwall is designed as four-holes constructal patterns combined with several arrangement choices, i.e. by pressure, by heat transfer and by streamline distributions.

AERO-THERMAL ASPECTS OF FILM COOLED NOZZLE GUIDE VANE ENDWALLS - PART 2: THERMAL MEASUREMENTS

Flow over gas turbine endwalls is complex and highly three-dimensional. As boundaries for modern engine designs are pushed, this already-complex flow is affected by aggressive application of film cooling flows that actively interact. This two-part study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The approach flow conditions represent flow exiting a low-NOx combustor. The test section includes geometric and cooling details of a combustor-turbine interface in addition to endwall film cooling flows injected upstream of the passage. The first part of this study describes in detail, the passage aerodynamics as affected by injection of cooling flows. It reveals a system of secondary flows, including the newly-discovered Impingement Vortex, which redefines our understanding of the aerodynamics of flow in a modern, film-cooled, first-stage vane row. The second part investigates, through thermal measurements, the distribution, mixing and disruption of cooling flows over the endwall. Measurements are made with and without active endwall film cooling. Descriptions are made through adiabatic surface effectiveness measurements and correlations with in-passage velocity (presented in part one) and thermal fields. Results show that the newly-discovered impingement vortex has a positive effect on coolant distribution through passage vortex suppression and by carrying the coolant to hard-to-cool regions in the passage, including the pressure surface near the endwall.

Implementation of Vortex Generator and Ramp to Improve Film Cooling Effectiveness on Blade Endwall

With the arrangements of vortex generators (VG) and ramp, film cooling effects on endwall near leading edge were numerically investigated at two blowing ratios (i.e. M = 0.5 and M = 1). To determine suitable numerical methods, mesh independency analysis and turbulence model selection were carried out based on the existing experimental data and LES results. With the numerical methods, flow fields near the leading edge were visualized to illustrate the influence of VG and ramp on coolant coverage on blade endwall. Film cooling effectiveness distributions on endwall and coolant trajectories near leading edge were compared among five different configurations with VG and ramp. The results show that the attachment of coolant on blade endwall is improved with the implement of VG between shaped-hole and leading edge. With the implementation of ramp on endwall between cooling hole and leading edge, the coolant spreads wider on endwall along pitchwise direction than the baseline case. With the implementation of VG and ramp, film cooling effect on endwall near leading edge is significantly improved as compared with the only ramp and only VG cases. Compared with the baseline case, pitchwise-averaged film cooling effectiveness on blade endwall near leading edge is increased by about 9%, and the film cooling effectiveness distributions on endwall along pitchwise direction become much uniform, for the case with both ramp and VG at M = 1.

Numerical investigation of blowing ratio, density ratio and axial position of film holes on the vane endwall film cooling effectiveness with upstream step

The aerothermal performance of interrupted slot and film holes was numerically investigated. Previous study indicates that the interrupted slot performs better compared to the conventional slot. In the meanwhile, the step formed along with the interrupted slot affects the film cooling characteristics. In this article, a row of film holes is arranged downstream of the step, and the mass flow rate for the interrupted slot is constant at 1%. Blowing ratio (BR) from 0.5 to 1.5 and density ratio from 1 to 2 were studied for the film holes. Endwall film cooling effectiveness distribution indicates that film cooling is easily affected by the secondary flow inside passage and the upstream step. Coolant traces are split into two parts due to the effects of step vortex and transverse flow. For different density ratios, increasing BR shows a different trend of film cooling effectiveness due to the variation of coolant momentum. The coolant jet is easily affected by the secondary flow when its momentum is low, but tends to liftoff when its momentum is too high. As a result, it is better to position the film holes far away from the upstream step. The total pressure loss coefficient distribution at the passage exit indicates that the coolant injection increases the total pressure loss. But density ratio has smaller effect on the loss variation. Besides, two axial positions of cooling holes were studied to improve the endwall cooling performance. Without the effect of step vortex, the film effectiveness of cooling holes is improved.

Implementation of Rectangular Vortex Generator Pairs to Improve Film Cooling Effectiveness On Transonic Rotor Blade Endwall

The rectangular vortex generator pairs (RVGPs) are arranged upstream the film cooling holes to achieve a better coolant coverage on endwall near pressure-side corner area. The endwall film cooling effectiveness distributions under transonic flow conditions are numerically calculated for the single RVGP and double rows of RVGPs cases. At first, the effects of three geometrical parameters (i.e. distance between RVGP and cooling hole, height of RVGP and attack angle of RVGP) on endwall film cooling effectiveness are studied with a single hole and RVGP at different mainstream inlet Reynolds numbers and blowing ratios. Then, the double rows of RVGPs are applied to further enhance the overall film cooling effectiveness on blade endwall. The results show that the implementation of RVGPs significantly enhances the film cooling effect on transonic blade endwall at pressure-side corner area. With the increase of RVGP height, the lateral coolant coverage on endwall corner area is improved. However, by decreasing the distance between vortex generator pair and cooling hole, the film cooling effectiveness downstream of the cooling holes is increased. The attack angle of RVGP mainly affects the shape of coolant spreading on endwall surface. The RVGP with optimum dimensions and arrangement is able to suppress the coolant from lifting off the endwall and increase the coolant diffusion near endwall. Compared with no vortex generator case, the area-averaged film cooling effectiveness on endwall with double rows of RVGPs is improved by 13.16%.

Aero-thermal Aspects of Film Cooled Nozzle Guide Vane Endwall - Part 1: Aerodynamics

The first stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics are not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The test section involves an engine-representative combustor-turbine interface geometry, combustor coolant flow and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Effects of Endwall Film Coolant Flowrate on Secondary Flows and Coolant Mixing in a First Stage Nozzle Guide Vane

Modern gas turbines are subjected to very high thermal loading. This leads to a need for aggressive cooling to protect components from damage. Endwalls are particularly challenging to cool due to a complex system of secondary flows near them that wash and disrupt the protective coolant films. This highly three-dimensional flow not only affects but is also affected by the momentum of film cooling flows, whether injected just upstream of the passage to intentionally cool the endwall or as combustor cooling flows injected further upstream in the engine. This complex interaction between the different cooling flows and passage aerodynamics has been recently studied in a first stage nozzle guide vane. The present paper presents a detailed study on the sensitivity of aero-thermal interactions to endwall film cooling mass flow to mainstream flow ratio. The test section represents a first stage nozzle guide vane with a contoured endwall and endwall film cooling injected just upstream of it. The test section also includes an engine-representative combustor–turbine interface geometry with combustor cooling flows injected at a constant rate. The approach flow conditions represent flow exiting a low-NOx combustor. Adiabatic surface thermal measurements and in-passage velocity and thermal field measurements are presented and discussed. The results show the dynamics of passage vortex suppression and the increase of impingement vortex strength as MFR changes. The effects of these changes of secondary flows on coolant distribution are presented.