A simulation of film cooling in the leading edge region of a turbine blade (trench effect on film effectiveness from cylinder in crossflow)

Film cooling is one of the cooling system techniques applied to the turbine blade. Gas turbine use film cooling technique to protect turbine blade from expose directly to a hot gas to avoid the blade from defect. The focus of this investigation is to investigate the effect of embedded three difference depth of trench at cooling holes geometry to the film cooling effectiveness. Comparisons are made under blowing ratio 1.0, 1.25, 1.5 and 2.0. Three configuration leading edge with depth Case A (0.0125D), Case B (0.0350D) and Case C (0.713D) were compared to leading edge without trench. Result shows that as blowing ratio increased from 1.0 to 1.25, the film cooling effectiveness is increase for leading edge without trench and also for all cases. However when the blowing ratio is increase to 1.5, film cooling effectiveness is decrease for all cases. Meanwhile for blowing ratio 2.0, the result shows the effect of depth is too small for all the cases. Overall the Case B with blowing ratio 1.25 has the best film cooling effectiveness with significant improvement compared to leading edge without trench and with trench Case A and Case C

Experimental and simulation study on the effect of geometrical and flow parameters for combined-hole film cooling

Film cooling method was applied to the turbine blades to provide thermal protection from high turbine inlet temperatures in modern gas turbines. Recent literature discovers that combining two cylindrical holes of film cooling is one of the ways to further enhance the film cooling performances. In the present study, a batch of simulations and experiments involving two cylindrical holes with opposite compound angle were carried out and this two cylindrical hole also known as combined-hole film cooling. The main objective of this study is to determine the influence of different blowing ratio, M with a combination of different lateral distance between cooling holes (PoD), a streamwise distance between cooling holes (LoD) and compound angle of cooling hole (1/2) on the film cooling performance. The simulation of the present study had been carried out by using Computational Fluid Dynamic (CFD) with application of Shear Stress Transport (SST) turbulence model analysis from ANSYS CFX. Meanwhile, the experimental approach makes used of open end wind tunnel and the temperature distributions were measured by using infrared thermography camera. The purpose of the experimental approach in the present study is to validate three cases from all cases considered in the simulation approach. As the results shown, the lateral coverage was observed to be increased as PoD and 1/2 increased due to the interaction between two cooling air ejected from both cooling holes. Meanwhile, film cooling performance insignificantly changed when different LoD was applied. As the conclusion, a combination of the different geometrical parameters with various flow parameters produced a pattern of results. Therefore, the best configuration has been determined based on the average area of film cooling effectiveness. For M = 0.5, PoD = 1.0, LoD = 2.5 and 1 / 2 = -45o /+45o case is the most effective configuration. In the case of M = 1.0 and M = 1.5, PoD = 0.0, LoD = 3.5, 1 / 2 = -45o /+45o and PoD = 0.0, LoD = 2.5, 1 / 2 = -45o /+30o are the best configurations based on the overall performance of film cooling

Turbulent mixing film cooling correlation

Film cooling effectiveness correlation predicts air flow requirement for cooling gas turbine combustors. Turbulent mixing model accounts for mixing rate between cooling film and hot gas stream. Resulting equation correlates data within plus or minus 20 percent

Effect of shocks on film cooling of a full scale turbojet exhaust nozzle having an external expansion surface

Experimental film cooling data obtained during exploratory testing with an axisymmetric plug nozzle having external expansion and installed on an afterburning turbojet engine in an altitude test facility is presented. The shocks and local hot gas stream conditions had a marked effect on film cooling effectiveness. An existing film cooling correlation was adequate at some operating conditions but inadequate at other conditions such as in separated flow regions resulting from shock boundary layer interactions

Curved film cooling admission tube

Effective film cooling to protect a wall surface from a hot fluid which impinges on or flows along the surface is provided. A film of cooling fluid having increased area is provided by changing the direction of a stream of cooling fluid through an angle of from 135 deg. to 165 deg. before injecting it through the wall into the hot flowing gas. The 1, cooling fluid is injected from an orifice through a wall into a hot flowing gas at an angle to form a cooling fluid film. Cooling fluid is supplied to the orifice from a cooling fluid source via a turbulence control passageway having a curved portion between two straight portions. The angle through which the direction of the cooling fluid is turned results in less mixing of the cooling fluid with the hot gas, thereby substantially increasing the length of the film in a downstream direction

Turbine airfoil film cooling

The experimental data obtained in this program gives insight into the physical phenomena that occur on a film cooled airfoil, and should provide a relevant data base for verification of new design tools. Results indicate that the downstream film cooling process is a complex function of the thermal dilution and turbulence augmentation parameters with trends actually reversing as blowing strength and coolant-to-gas temperature ratio varied. The pressure surface of the airfoil is shown to exhibit a considerably higher degree of sensitivity to changes in the film cooling parameters and, consequently, should prove to be more of a challenge than the suction surface in accurately predicting heat transfer levels with downsteam film cooling

Film cooling on the pressure surface of a turbine vane

Film-cooling-air ejection from the pressure surface of a turbine vane was investigated, and experimental data are presented. This investigation was conducted in a four-vane cascade on a J75-size turbine vane that had a double row of staggered holes in line with the primary flow and located downstream of the leading edge region. The results showed that: (1) the average effectiveness of film-convection cooling was higher than that of either film cooling or convection cooling separately; (2) the addition of small quantities of film-cooling air always increased the cooling effectiveness relative to the zero-injection case; however, (3) the injected film must exceed a certain threshold value to obtain a beneficial effect of film cooling relative to convection cooling alone

Turbine airfoil film cooling

Emphasis is placed on developing more accurate analytical models for predicting turbine airfoil external heat transfer rates. Performance goals of new engines require highly refined, accurate design tools to meet durability requirements. In order to obtain improvements in analytical capabilities, programs are required which focus on enhancing analytical techniques through verification of new models by comparison with relevant experimental data. The objectives of the current program are to develop an analytical approach, based on boundary layer theory, for predicting the effects of airfoil film cooling on downstream heat transfer rates and to verify the resulting analytical method by comparison of predictions with hot cascade data obtained under this program

Combustion effects on film cooling

The effects of: (1) a reactive environment on film cooling effectiveness, and (2) film cooling on rocket engine performance were determined experimentally in a rocket thrust chamber assembly operating with hydrogen and oxygen propellants at 300 psi chamber pressure. Tests were conducted using hydrogen, helium, and nitrogen film coolants in an instrumented, thin walled, steel thrust chamber. The film cooling, performance loss, and heat transfer coefficient data were correlated with the ALRC entrainment film cooling model which relates film coolant effectiveness and mixture ratio at the wall to the amount of mainstream gases entrained with the film coolant in a mixing layer. In addition, a comprehensive thermal analysis computer program, HOCOOL, was prepared from previously existing ALRC computer programs and analytical techniques

Experimental verification of film-cooling concepts on a turbine vane

Film cooling concepts applied to gas turbine vanes were investigated. The filming cooling air was ejected from a single row of holes on the convex surface and a double row of holes of the concave surface. Tests were conducted at a gas temperature of 1260 K, a gas pressure of 3 atmospheres, and a coolant temperature of 280 K. Mass velocity ratios were varied between 0 and 2.0. Data were taken without film cooling holes, with film cooling holes but without blowing, and with blowing. A small amount of blowing into a nonturbulent boundary layer caused an increase in vane temperatures. Film cooling when combined with convection cooling was verified to be more effective than either film or convection cooling alone