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This study aims to clarify the mechanism of film hole location at the span-wise direction of an internal cooling channel with crescent ribs on the adiabatic film cooling performance, three configurations are designed to observe the effects of the distance between the center of the ellipse and the side wall(Case 1, l = w/2, Case 2, l = w/3 and for Case 3, l = w/4).

Numerical simulations are conducted under two blowing ratios (i.e. 0.5 and 1) and a fixed cross-flow Reynolds number (Re_c = 100,000) with a verified turbulence model.

It is shown that at low blowing ratio, reducing the distance increases the film cooling effectiveness but keeps the trend of the effectiveness unchanged, while at high blowing ratio, the characteristic is a little bit different in the range of 0 = x/D = 10.

These features could be explained by the fact that shrinking the distance between the hole and side wall induces a much smaller reserved region and vortex downstream the ribs and a lower resistance for cooling air entering the film hole. Furthermore, the spiral flow inside the hole is impaired.

As a result, the kidney-shaped vortices originating from the jet flow are weakened, and the target surface can be well covered, resulting in an enhancement of the adiabatic film cooling performance.

In this study, numerical simulations are performed to compare the adiabatic film cooling effectiveness and reveal the difference of film cooling mechanisms of two models with the same geometries and cross-section areas of film holes’ exits at three typical blowing ratios (M = 0.5, 1 and 1.5). The two models are an elliptical model and a cylindrical model with 90° compound angle, respectively.

Three different cases are considered in this work and the baseline is the model with a cylindrical film hole. The same boundary conditions and a validated turbulence model (realizable k-ε) are adopted for all cases.

The results show that both the elliptical and cylindrical models with 90° compound angle can enhance the film cooling effectiveness compared with the baseline. However, the elliptical model performs well at lower blowing ratios and in the near region at each blowing ratio because of the wider width of the film hole’s exit. The cylindrical model with 90° compound angle provides better film cooling effectiveness in the further downstream area of the film hole at higher blowing ratio because of the less lift-off and better coolant coverage in the larger x/D region along the mainstream direction.

Overall, it can be concluded that although the elliptical and cylindrical models with 90° compound angle have identical hole exits, the different inlet direction and cross-sectional geometry affect the flow structures when the coolant enters, moves through and exits the hole and finally different film cooling results appear.

The present study aims to conduct a numerical investigation of a novel film cooling scheme combining in‐hole impingement cooling and flow turbulators with traditional downstream film cooling, and was originally proposed by Pratt & Whitney Canada for high temperature gas turbine applications.

Steady‐state simulations were performed and the flow was considered incompressible and turbulent. The CFD package FLUENT 6.1 was used to solve the Navier‐Stokes equations numerically, and the preprocessor, Gambit, was used to generate the required grid.

It was determined that the proposed scheme geometry can prevent coolant lift‐off much better than standard round holes, since the cooling jet remains attached to the surface at much higher blowing rates, indicating a superior performance for the proposed scheme.

The present study was concerned only with the downstream effectiveness aspect of performance. The performance related to the heat transfer coefficient is a prospective topic for future studies.

Advanced and innovative cooling techniques are essential in order to improve the efficiency and power output of gas turbines. This scheme combines in‐hole impingement cooling and flow turbulators with traditional downstream film cooling for improved cooling capabilities.

This new advanced cooling scheme both combines the advantages of traditional film cooling with those of impingement cooling, and provides greater airfoil protection than traditional film cooling. This study is of value for those interested in gas turbine cooling.

The purpose of this paper is to report the result of a numerical investigation of film cooling performance on a flat plate for finding optimum blowing ratios.

Steady-state simulations have been performed, and the flow has been considered incompressible. Calculations have been performed with 3D finite-volume method and the k-e turbulence model.

The adiabatic film cooling effectiveness and the effects of density ratio (DR), blowing ratio (M) and main stream turbulence intensity (Tu), coolant penetration, hole incline and diameter are studied. The temperature and film cooling effectiveness contours, centerline and laterally film cooling effectiveness are presented for these cases. Results show that the cases with smaller Tu have better effectiveness. In the console, using the air coolant and in cylindrical hole cases, using CO₂ coolant fluid has higher effectiveness. The results indicated that there is an optimum blowing ratio in the cylindrical hole cases to optimize the performance of new gas turbines.

Investigation of optimum blowing ratio for the convex surfaces and turbine blades is a prospective topic for future studies.

The motivation of this study comes from several industrial applications such as film cooling of gas turbine components. This research gives the best blowing ratio for receiving maximum cooling effectiveness with minimum coolant velocity.

This study optimizes the blowing ratio for film cooling on a flat plate.

To investigate the film cooling effectiveness in a flat plate with a single row of rectangular injection holes.

Three injection holes in model are in a single row. The holes are rectangular cross section and they are 9 × 6.5 mm. The injection holes are inclined at 30° along the mainstream direction. The blowing ratios are from 0.5 to 2.0. The experiments and their computational models are established to investigate its effects at the 330 and 335 and 340 K injection temperatures and the different blowing ratios.

Results show that the blowing ratio and injection temperature and momentum flux ratio affect the film cooling effectiveness and to provide a good film cooling performance in both mainstream and lateral direction a suitable blowing ratio should be selected. In this study, the highest effectiveness is determined at a blowing ratio of 0.5. Further increasing this ratio results in reverse effect on the film cooling effectiveness.

It is the fist time the film cooling effectiveness is compared at the rectangular injection holes as experimental and numerical.

The purpose of this paper is to investigate the film cooling effectiveness and heat transfer coefficient in a flat plate with two rows of rectangular injection holes.

Experimental and numerical investigation of film cooling effectiveness in a flat plate with two rows which are rectangular injection holes. The liquid crystal technique has been used for measuring the heat transfer coefficients on the mixture region. Three injection holes in model are in a single row. The holes are rectangular cross section and they are 9 × 6.5 mm. The injection holes are inclined at 30° along the mainstream direction. The blowing ratios are from 0.5 to 2.0. The experiments and their computational models are established to investigate its effects at the 330 and 340 and 350 K injection temperatures and the different blowing ratios.

The results show that the film cooling effectiveness and heat transfer coefficient of a given flat plate surface, both along the mainstream and lateral direction, depend on the optimum selection of parameters. In this study, the highest effectiveness is determined at a blowing ratio of 0.5. Further increasing the ratio results in reverse effect on the film cooling effectiveness.

It is the fist time the film cooling effectiveness is compared at the rectangular injection holes with two rows as experimental and numerical.

The purpose of this paper is to report a numerical investigation of jet‐cross‐flow interaction in the presence of imperfection inside the injection hole with application to film cooling of turbine blades.

The work includes the prediction of the thermal and hydrodynamic fields by solving the Reynolds Averaged Navier Stokes and energy equations using the finite volume method with a body‐fitted hexahedral unstructured grid. The turbulence field is resolved by use of the k‐epsilon turbulence model.

The computational results show a dramatic and rapid decrease of the film cooling effectiveness when the obstruction is superior to 50 per cent. It is found that when the obstruction is close to the exit hole, the thermal protection is significantly reduced.

The present numerical investigation is simply directed towards a qualitative investigation of hole imperfection effects on film cooling.

The motivation comes from several industrial applications such as film cooling of gas turbine components and fuel injection. One of the main challenges of using film cooling is the blockage of holes by particles ingested by the engine during landing/take off or due to application of thermal barrier coating or due to combustion particles as well as inaccuracies that result from drilling of holes.

The main goal of the present study is to conduct a numerical parametric investigation rather than reproducing the exact Jovanović's experimentation.

Numerical simulations were carried out for two cooling schemes, a circular hole and a louver cooling scheme, at the leading edge of a rotor blade in a complete turbine stage.

Two holes were positioned at the leading edge of a rotating blade, one on the pressure side and the other on the suction side. The methodology was validated with a circular hole case. Numerical results of cooling effectiveness for three blowing ratios at three rotational speeds were successfully obtained. Both blowing ratio and rotating speed of the rotor affect the cooling effectiveness level.

It was shown that for the circular hole, the blowing ratio is the dominant factor at low blowing ratios and the rotational speed is the dominant factor at high blow ratios when jet is prone to lift off in determining the cooling effectiveness level. For the louver scheme, a higher rotational speed leads to a higher level of cooling effectiveness since jet liftoff is avoided.

There are only a few studies of film cooling on a rotational turbine blade and very few studies of film cooling at the leading edge of a rotating turbine blade in the open literature. The present work presents a challenging CFD case. The analysis of film cooling at the leading edge of an airfoil was presented, which sheds light on the physics of film cooling and should prove helpful to the cooling designs of turbine blades.

The purpose of this paper devoted to the application of modal analysis to analyze the flow structure of trailing edge cutback film cooling and the effects of vortex structure on the film cooling effectiveness of the cutback surface.

Large eddy simulation (LES) is used to simulate the trailing edge cutback film cooling. The results of LES are analyzed by proper orthogonal decomposition (POD) method and dynamic mode decomposition (DMD) method. The POD method is used to determine the dominated vortex structure and the energy level of these structures. The DMD method is used to analyze the relationship between vortex structures and wall temperature.

The POD method shows that the flow field consists of three main vortices – streamwise vortex, lip vortex and coolant vortex. The DMD results show that the lip vortex mainly acts on the middle section of the cutback surface, while the streamwise vortex mainly acts on the back section of the cutback surface.

The modal analysis is only based on numerical simulation but the modal analysis of experimental results will be further studied in the future.

This paper presents the powerful ability of the modal analysis method to study complex flows in trailing edge cutback film cooling. Establishing the relationship between vortex and wall temperature by modal analysis method can provide a new idea for studying convective heat transfer problems.

The role of streamwise vortex in the flow of the trailing edge cutback cooling and its effect on the cooling effectiveness of the cutback surface is found.

The purpose of this study is to explore the mist-air film cooling performance on a three-dimensional (3-D) flat plate. In mist-air film cooling technique, a small amount of water droplets is injected along with the coolant air. The objective is to study the influence of shape of the coolant hole and operating conditions on the cooling effectiveness.

In this study, 3-D numerical simulations are performed. To simulate the mist-air film cooling over a flat plate, air is considered as a continuous phase and mist is considered as a discrete phase. Turbulence in the flow is accounted using Reynolds averaged Navier–Stokes equation and is modeled using k–e model with enhanced wall treatment.

The results of this study show that, for cylindrical coolant hole, coolant with 5% mist concentration is not effective for mainstream temperatures above 600 K, whereas for fan-shaped hole, even 2% mist concentration has shown significant impact on cooling effectiveness for temperatures up to 1,000 K. For given mist-air coolant flow conditions, different trend in effectiveness is observed for cylindrical and fan-shaped coolant hole with respect to main stream temperature.

This study is limited to a flat plate geometry with single coolant hole.

The motivation of this study comes from the requirement of high efficiency cooling techniques for cooling of gas turbine blades. This study aims to study the performance of mist-air film cooling at different geometric and operating conditions.

The originality of this study lies in studying the effect of parameters such as mist concentration, droplet size and blowing ratio on cooling performance, particularly at high mainstream temperatures. In addition, a systematic performance comparison is presented between the cylindrical and fan-shaped cooling hole geometries.