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1 – 10 of over 2000
Article
Publication date: 8 June 2012

C.X‐Z. Zhang and I. Hassan

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.

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Originality/value

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.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 22 no. 5
Type: Research Article
ISSN: 0961-5539

Keywords

Article
Publication date: 3 January 2017

Mojtaba Kazemi Kelishami and Esmail Lakzian

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.

272

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Research limitations/implications

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

Practical implications

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.

Originality/value

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

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 27 no. 1
Type: Research Article
ISSN: 0961-5539

Keywords

Article
Publication date: 30 October 2007

İbrahim Koç

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

Abstract

Purpose

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

Design/methodology/approach

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.

Findings

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.

Originality/value

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

Details

Aircraft Engineering and Aerospace Technology, vol. 79 no. 6
Type: Research Article
ISSN: 0002-2667

Keywords

Article
Publication date: 23 January 2009

İbrahim Koç, Yaşar İslamoğlu and Ünal Akdağ

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.

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Originality/value

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

Details

Aircraft Engineering and Aerospace Technology, vol. 81 no. 2
Type: Research Article
ISSN: 0002-2667

Keywords

Article
Publication date: 1 December 2006

X.‐Z. Zhang and I. Hassan

To develop a reliable methodology and procedure of simulating the jet‐in‐crossflow using the current turbulence models and numerically investigate the cooling performance…

Abstract

Purpose

To develop a reliable methodology and procedure of simulating the jet‐in‐crossflow using the current turbulence models and numerically investigate the cooling performance of a new scheme for the engines of next generation.

Design/methodology/approach

A new advanced film cooling scheme is proposed based on the literature survey and a systematic methodology developed to successfully predict the right level of heat transfer in the CFD simulation of film cooling.

Findings

The proposed cooling scheme gives considerable lower heat transfer coefficient at the centerline in the near hole region than the traditional cylindrical hole, especially at a high blowing ratio when traditional cylindrical hole undergoes liftoff.

Research limitations/implications

The number of cooling holes in the computational domain is limited by the speed of the computers used.

Practical implications

The new methodology can be used to numerically test new cooling schemes in the design of turbine blades and to provide useful information/data under actual working conditions to design engineers.

Originality/value

This paper provides some useful information on the simulation of film cooling in terms of the performance of different turbulence models and wall treatments and also sends some valuable messages regarding the design of cooling scheme of turbine blades to the technical community.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 16 no. 8
Type: Research Article
ISSN: 0961-5539

Keywords

Article
Publication date: 23 June 2020

Guohua Zhang, Gongnan Xie and Bengt Ake Sunden

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…

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Originality/value

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.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 31 no. 1
Type: Research Article
ISSN: 0961-5539

Keywords

Article
Publication date: 1 June 2004

N. Qin, Y. Zhu and S.T. Shaw

In this paper, the effectiveness of a number of active devices for the control of shock waves on transonic aerofoils is investigated using numerical solutions of the…

1273

Abstract

In this paper, the effectiveness of a number of active devices for the control of shock waves on transonic aerofoils is investigated using numerical solutions of the Reynolds‐averaged Navier‐Stokes equations. A brief description of the flow model and the numerical method is presented including, in particular, the boundary condition modelling and the numerical treatment for surface mass transfer. Comparisons with experimental data have been made where possible to validate the numerical study before some systematic numerical simulations for a parametric study. The effects of surface suction, blowing, and local modification of the surface contour (bump) on aerofoil aerodynamic performance have been studied extensively regarding the control location, the mass flow strength and the bump height. The numerical simulations highlight the benefits and drawbacks of the various control devices for transonic aerodynamic performance and identify the key design parameters for optimisation.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 14 no. 4
Type: Research Article
ISSN: 0961-5539

Keywords

Article
Publication date: 1 June 2006

A. Immarigeon and I. Hassan

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…

1084

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Research limitations/implications

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.

Practical implications

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.

Originality/value

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.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 16 no. 4
Type: Research Article
ISSN: 0961-5539

Keywords

Article
Publication date: 18 October 2018

Dong Liang, Wenjie Wang and Peter J. Thomas

Numerical and experimental results for different oncoming base-flow conditions indicate that nonuniform trailing edge blowing (NTEB) can expand the performance range of…

108

Abstract

Purpose

Numerical and experimental results for different oncoming base-flow conditions indicate that nonuniform trailing edge blowing (NTEB) can expand the performance range of compressors and reduce the thrust on the rotor, while the efficiency of the compressor can be improved by more than 2 per cent.

Design/methodology/approach

Relevant aerodynamic parameters, such as total pressure, ratio of efficiency and axial thrust, are calculated and analyzed under conditions with and without NTEB. Measurements are performed downstream of two adjacent stator blades, at seven equidistantly spaced reference locations. The experimental measurement of the interstage flow field used a dynamic four-hole probe with phase lock technique.

Findings

An axial low-speed single-stage compressor was established with flow field measurement system and nonuniform blowing system. NTEB was studied by means of numerical simulations and experiments, and it is found that the efficiency of the tested compressor can be improved by more than 2 per cent.

Originality/value

Unlike most of the previous research studies which mainly focused on the rotor/stator interaction and trailing edge uniform blowing, the research results summarized in the current paper on the stator/rotor interaction used inlet guide vanes for steady and unsteady calculations. An active control of the interstage flow field in a low-speed compressor was used to widen the working range and improve the performance of the compressor.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 1
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 31 July 2019

Guohua Zhang, Xueting Liu, Bengt Ake Sundén and Gongnan Xie

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…

Abstract

Purpose

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).

Design/methodology/approach

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

Findings

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.

Research limitations/implications

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.

Originality/value

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.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 29 no. 8
Type: Research Article
ISSN: 0961-5539

Keywords

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