Blade and vane leading edge fillet on endwall cooling in linear turbine cascades. (original) (raw)

Influences of large fillets on endwall flows in a vane cascade with upstream slot film-cooling

Experimental Thermal and Fluid Science, 2019

Investigations in cascade employ filleted blades to influence the near endwall secondary flows and total-pressure losses. The secondary flows aggravate the aerodynamic losses and endwall thermal stresses in the gas turbine passages. Investigations of different configurations of the slot-bleed flow from the endwall of cascade upstream show significant influences on the near endwall flows. In the present paper, the near endwall flow-field in a linear cascade employing filleted vanes and bleed flow from the upstream endwall slots is measured experimentally. Two fillet profiles are tested, one is larger than the other. The objectives are to quantify the combined effects of endwall fillet, fillet profiles, and film-cooling flow on the endwall region flow-field. The fillet covers the junction of vane and endwall upstream of the cascade throat region. The discontinuous bleed-slots near the cascade entrance provide the film-cooling flow on endwall and simulate the gaps between combustor/nozzlevane discs or stator/rotor discs in the gas turbine. The inlet Reynolds number of 2.0E+05 is based on the chord length of vane profile. The density and temperature ratios of the coolant flow to mainstream are both about 1.0. The inlet blowing ratio of the film-cooling flow varies between 1.0 and 2.8. The flow-field is measured through the distributions of flow temperature, yaw angle, axial vorticity, and total-pressure losses along the vane passage. The effects on the flow-field are then presented by comparing the cases of filleted vanes with the cases of un-filleted vanes. The results of flow yaw angle and axial vorticity in the filleted passage are smaller than those in the passage without the fillet. The yaw angles responsible for the endwall secondary flows are the smallest for the smaller fillet (Fillet-2). The temperature field indicates the pitchwise distributions of the coolant on endwall specially near the pressure side are better when the smaller fillet is employed. Also, the weakened passage vortex of the endwall secondary flows in the filleted passage reduces the total-pressure losses. Although the totalpressure losses decrease at very high coolant mass flux with and without the fillet, the losses are always smaller for the filleted passage than for the un-filleted passage. The present investigation is beneficial for improving and optimizing the endwall film-cooling in the gas turbine passage.

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

Journal of Turbomachinery, 2012

This paper investigates the influence of coolant injection on the aerodynamic and thermal performance of a rotor blade cascade with endwall film cooling. A seven 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 10 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 five-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 behavior has been also analyzed by using thermochromic liquid crystals technique to obtain film cooling effectiveness distributions. All this information, in...

Effects of non-axisymmetric endwall contouring and film cooling on the passage flowfield in a linear turbine cascade

Experiments in Fluids, 2015

List of symbols c Coolant upstream of impingement plate C ax Axial chord length D Hole diameter H Impingement gap height M Blowing ratio (ρ c U c /ρ ∞ U ∞) Ma Mach number P Pressure PIV Particle image velocimetry p Pitch length Re Reynolds number (ρ ∞ U ∞ C ax /μ ∞) S Blade span s Static or streamwise coordinate T Temperature tke Turbulent kinetic energy 3/4 u ′ 2 + v ′ 2 U Velocity U inviscid Average inviscid velocity vector U meas Average measured velocity vector U sec Secondary velocity vector (U meas − U inviscid) u ′ Fluctuating velocity x Blade axial coordinate y Blade pitchwise coordinate z Blade spanwise coordinate δ Boundary layer thickness (99 %) μ Dynamic viscosity ρ Density ∞ Mainstream conditions at the cascade inlet

Heat Transfer Measurements on the Endwall of a Variable Speed Power Turbine Blade Cascade

2018

Heat transfer measurements were obtained on the endwall of a 2-D section of a variable speed power turbine (VSPT) rotor blade linear cascade. Infrared thermography was used to help determine the transition of flow from laminar to turbulent as well as determine regions of flow separation. Steady state data was obtained for six incidence angles ranging from +15.8° to -51°, and at five flow conditions for each angle. Nusselt number was used as a method to visualize flow transition and separation on the endwall surface and showed the effects of secondary flows on the surface. Nusselt correlation with Reynolds number from multiple flow conditions was used to plot local values of the correlation exponent and indicated the state of the local boundary layer as the flow transitioned from laminar to turbulent as well as secondary flow features.

Heat transfer and film-cooling for the endwall of a first stage turbine vane

International Journal of Heat and Mass Transfer, 2005

Secondary flows that result in turbomachines from inherent pressure gradients in airfoil passages, are the main contributors to aerodynamic losses and high heat transfer to the airfoil endwalls. The endwalls present a challenge to durability engineers in maintaining the integrity of the airfoils. One means of preventing degradation in the turbine is to film-cool components whereby coolant is extracted from the compressor and injected through small cooling holes in the airfoil surfaces. In addition to film-cooling, leakage flows from component interfaces, such as the combustor and turbine, can provide cooling in localized areas but also provide a change to the inlet boundary condition to the passage. This paper presents measurements relevant to the endwall region of a vane, which indicate the importance of considering the inlet flow condition.

An experimental investigation on the trailing edge cooling of turbine blades

Propulsion and Power Research, 2012

An experimental study was conducted to quantify the flow characteristics of the wall jets pertinent to trailing edge cooling of turbine blades. A high-resolution stereoscopic particle image velocimetry (PIV) system was used to conduct detailed flow field measurements to quantitatively visualize the evolution of the unsteady vortices and turbulent flow structures in the cooling wall jet streams and to quantify the dynamic mixing process between the cooling jet stream and the mainstream flows. The detailed flow field measurements were correlated with the adiabatic cooling effectiveness maps measured by using pressure sensitive paint (PSP) technique to elucidate underlying physics in order to explore/optimize design paradigms for improved cooling effectiveness to protect the critical portions of turbine blades from harsh environments.

Experimental Acquisition Of An Endwall Heat Distribution In A Linear Turbine Cascade

2020

The focus of this research is the experimental acquisition of endwall heat transfer distributions for an aft loaded vane with a large leading edge. This study will investigate endwall heat transfer distributions over five inlet turbulence levels ranging from intensities of 0.7% through very high turbulence levels as high as 17.4%. The investigation will be conducted at three varying Reynolds numbers based on true chord length and exit conditions ranging from 500,000 to 2000,000. The infrared thermography technique will be applied to the acquisition of the endwall heat transfer data due to the full surface image which can be developed from the acquired thermographs. An in situ calibration technique will be used to enhance the measurement accuracy. The experiment was conducted in a linear cascade test section, consisting of four turbine vanes with upper and lower bleed flows. Linear Cascade can reflect most of the flow characteristics in real gas turbine nozzles. This experiment exhib...

Effects of Upstream Endwall Film Cooling on a Vane Cascade Flowfield

Journal of Propulsion and Power, 2017

The effects of film-cooling on the endwall region flow and aerodynamic losses are investigated experimentally as the film-flow is delivered from the slots in the endwall upstream of a linear vane cascade. Four slots inclined at 30° deliver the film-jet parallel to the main flow at four blowing ratios between 1.1 and 2.3 and at a temperature ratio of 1.0. The slots are employed in two configurations pitchwise-all four slots open (case-1) and two middle slots open (case-2). The inlet Reynolds number to the cascade is 2.0E+05. Measurements of the blade surface pressure, axial vorticities, yaw angles, and total pressure loss distributions along the cascade are reported with and without (Baseline) the film-cooling flow. The results show the film-flow changes the orientations, distributions, and strength of the endwall secondary flows and boundary layer. The case-1 of film-cooling provides more massflux and momentum than the case-2 affecting the passage vortex legs. The overall total pressure losses at the cascade exit are always lower for the film-cooling cases than for the Baseline. The overall losses are also lower at the low blowing ratios, but higher at the high blowing ratios for the film-cooling case-1 than for the case-2.

The effect of jet angle and velocity ratio on turbine blade film cooling

WIT transactions on modelling and simulation, 2013

This work is concerned with the thermal effect of the turbine blade by film cooling. The jet flow penetration and flow structure was simulated and solved numerically. This investigation presents a tool to obtain qualitative information about the penetration area and flow structure of the mixing flow at different jet angle configurations. Two models of airfoil (NACA 0021) with hole diameters, d=0.1 and 0.2 cm were studied. The simulation was carried out using the commercial code (FLUENT 6.3). Several cases were considered; including three-velocity ratio, VR=0.3, 0.7, 1,1.5, 2 and two jet suction angles, ß of 37.5 0 and 90 0 , and four different angles of attack ,α= 0 0 , 5 0 , 10 0 , and 15 0. The cooling films build up on the blade surface was simulated. The work gives a clear picture of the penetration area for the normal and angled-injection cases. The results at different sections, x/s= 0.5 and 0.9 for the first and second model provided the optimum holes rows spacing and the low suction angle for maximum film cooling effectiveness.

Computational Predictions of Heat Transfer and Film-Cooling for a Turbine Blade With Nonaxisymmetric Endwall Contouring

Journal of turbomachinery, 2011

Three-dimensional contouring of the compressor and turbine endwalls in a gas turbine engine has been shown to be an effective method of reducing aerodynamic losses by mitigating the strength of the complex vortical structures generated at the endwall. Reductions in endwall heat transfer in the turbine have been also previously measured and reported in literature. In this study, computational fluid dynamics simulations of a turbine blade with and without nonaxisymmetric endwall contouring were compared to experimental measurements of the exit flowfield, endwall heat transfer, and endwall filmcooling. Secondary kinetic energy at the cascade exit was closely predicted with a simulation using the SST k-turbulence model. Endwall heat transfer was overpredicted in the passage for both the SST k-and realizable k-turbulence models, but heat transfer augmentation for a nonaxisymmetric contour relative to a flat endwall showed fair agreement to the experiment. Measured and predicted film-cooling results indicated that the nonaxisymmetric contouring limits the spread of film-cooling flow over the endwall depending on the interaction of the film with the contour geometry.