Interaction Mechanism and Loss Analysis of Mixing between Film Cooling Jet and Passage Vortex (original) (raw)
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LES of Turbulent Mixing In Anti-Vortex Film Cooling Flows
The Anti-vortex film cooling technique is investigated by using large eddy simulation (LES) due to its complex mixture between the mainstream flow, the film cooling flow, and the flow through the anti-vortex holes. A geometry of single row of 30 degree round holes on a flat plate is used as the baseline case. Three different values of velocity ratios (Coolant Jet Velocity/Main Stream Velocity) are studied. Two different positions of the anti-vortex holes are investigated with temperature ratio (main stream temperature / coolant temperature) namely 2. The density ratio is taken in consideration. Use of symmetry boundary condition is avoided to capture three dimensional, unsteady, turbulent nature of the flow. Present simulation is carried out by using FLUENT commercial code. Numerical calculation of film cooling effectiveness is validated with reported experimental results. Results show that the used anti-vortex technique improves the film cooling effectiveness. The numerical boundary layer velocity vectors showed that the anti-vortex holes create reverse vortices against the main vortices that are created by the main hole. These reverse vortices help in keeping the coolant jet flow near the surface.
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.
Flowfield Measurements for Film-Cooling Holes With Expanded Exits
1996
One viable option to improve cooling methods used for gas turbine blades is to optimize the geometry of the film-cooling hole. To optimize that geometry, effects of the hole geometry on the complex jet-in-crossflow interaction need to be understood. This paper presents a comparison of detailed flowfield measurements for three different single, scaled-up hole geometries, all at a blowing ratio and density ratio of unity. The hole geometries include a round hole, a hole with a laterally expanded exit, and a hole with a forward-laterally expanded exit. In addition to the flowfield measurements for expanded cooling hole geometries being unique to the literature, the testing facility used for these measurements was also unique in that both the external mainstream Much number (Ma*, = 0.25) and internal coolant supply Mach number (Ma c = 0.3) were nearly matched. Results show that by expanding the exit of the cooling holes, both the penetration of the cooling jet and the intense shear regions are significantly reduced relative to a round hole. Although the peak turbulence level for all three hole geometries was nominally the same, the source of that turbulence was different. The peak turbulence level for both expanded holes was located at the exit of the cooling hole resulting from the expansion angle being too large. The peak turbulence level for the round hole was located downstream of the hole exit where the velocity gradients were very large.
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...
Volume 5C: Heat Transfer, 2018
Gas turbine components can withstand gas temperatures exceeding the melting point of the alloys they’re made of due to increasingly effective cooling methods. Increasing the operating temperature of a gas turbine is key to improving its power density and exhaust heat for steam or combined-cycle efficiency. In the turbine, the component that experiences the highest gas temperature is the vane directly downstream of the combustor; the most complex flow field in a vane occurs near the endwall. In this study, an experimental investigation is carried out to determine the effect of coolant injection angle and mass flow ratio on film effectiveness on the endwall using the pressure sensitive paint technique for various configurations of jump cooling hole configurations. Two rows of angled holes are upstream of an uncooled vane in a three-vane linear cascade. Injection angle including compound angle is varied from 20 to 60 and coolant to mainstream massflux ratio is varied from 0.5% to 3%. C...
An Experiment Study of Wall Slot Jets Pertinent to Trailing Edge Cooling of Turbine Blades
ASME 2010 3rd Joint US-European Fluids Engineering Summer Meeting: Volume 2, Fora, 2010
An experimental study was conducted to quantify the flow characteristics of wall jets pertinent to trailing edge cooling of turbine blades. A high-resolution stereoscopic PIV system was used to conduct detailed flow field measurements to quantitatively visualize the evolution of the unsteady vortex and turbulent flow structures in cooling wall jet streams and to quantify the dynamic mixing process between the cooling wall jet streams and the main stream flows. The detailed flow field measurements are correlated with the adiabatic cooling effectiveness maps measured by using pressure sensitive paint (PSP) technique to elucidate underlying physics in order to improve cooling effectiveness to protect the critical portions of turbine blades from the harsh ambient conditions.
Heat and Mass Transfer, 2004
Typical film-cooling configuration of a symmetrical turbine blade leading edge is investigated using a three-dimensional finite volume method and a multi-block technique. The computational domain includes the curved blade surface as well as the coolant regions and the plenum. The turbulence is approximated by a two layer k-e model. The computations have been performed using the TLV two-layer and the TLVA models. However, the utilization of the TLV and TLVA models has not improved the prediction of the lateral averaged film cooling effectiveness of gas turbine blades when compared with those obtained using wall function strategy.
Numerical Study of Aerodynamic Performance of Film Cooling With Backward Injection Holes
Volume 4: Heat and Mass Transfer Under Extreme Conditions; Environmental Heat Transfer; Computational Heat Transfer; Visualization of Heat Transfer; Heat Transfer Education and Future Directions in Heat Transfer; Nuclear Energy, 2013
Film cooling has been successfully used in cooling gas turbine components that are exposed to very high temperature environments. One main disadvantage of using film cooling is the aerodynamic losses associated. To address to the needs of obtaining uniform cooling in the downstream regions, backward injection of coolant has proved to be effective. However, there is a need to understand the aerodynamic behaviors of jet and mainstream flows in order to design effective configurations with this scheme of injecting coolant. In this work, the underlying aerodynamic principles of backward injection are studied numerically. All simulations are conducted with Fluent, a commercial CFD software. Results show that the classical counter rotating vortex found in simple cylindrical holes are not seen in the case of backward injections. Backward injection results in reduced coolant requirements and elimination of complex hole designs to avoid jet lift-off.
Mechanical Engineering Journal
The reliability of turbine blades and vanes of modern high temperature gas turbines is assured by turbine blade cooling technologies. Among the various cooling methods, film cooling has been a key technology to ensure the long-term operation of turbine blades and vanes that are exposed to hot gas-path flows. Therefore, many papers have been published aiming at the improvement of film cooling effectiveness by optimization of film cooling hole geometries. Although the turbulence intensity of the mainstream generated in the gas turbine combustor is very high and may reduce the film cooling effect on the turbine vane and blade, there are few papers investigating quantitatively the effect of the mainstream turbulence on the film cooling. For this reason, the influence of mainstream turbulence intensity on film cooling effectiveness was investigated with an active turbulence generator equipped with electric-motor driven propellers for circular and fan-shaped film cooling holes. Spatial distributions of the turbulent mixing field between turbulent mainstream and film coolant jet were measured with quantitative measurement methods, such as PIV and LIF. As a result, it was revealed that when the mainstream turbulence is high the counter-rotating vortex pair is weakened, the film cooling air spreads in the span-wise direction and the lateral-averaged film cooling effectiveness decreases about 10%.