Measurement and Simulation of a Transonic Innovative Cooling System (ICS) for High-Temperature Transonic Gas Turbine Stages

2004 ◽  
Author(s):  
E. Go¨ttlich ◽  
L. Innocenti ◽  
A. Vacca ◽  
W. Sanz ◽  
J. Woisetschla¨ger ◽  
...  
Keyword(s):  
Temperature Distribution ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Numerical Comparison ◽  
Blade Surface ◽  
Trailing Edge ◽  
Advantages And Disadvantages

Gas turbine design technology requires the development of transonic turbine stages capable of carrying high stage load and of handling hot gas temperatures at turbine inlet. A reliable cooling system is necessary to cope with shocks emanating from preceding blade rows and impinging on the blade especially in the leading edge region. In order to fulfill these requirements researchers at Graz University of Technology have been working on an Innovative Cooling System (ICS) since 1995. The ICS is able to cover large areas of the blade surface with an effective cooling film and to reduce the metal temperature without a shower head cooling arrangement at the leading edge and any trailing edge cooling air ejection. In this paper the authors present a numerical comparison of the ICS to a conventional modern film cooling system both implemented in the same industrial transonic gas turbine blade. An experimental determination of the adiabatic film cooling effectiveness distribution around the blades surface was necessary for the ICS because of its uncommon design. The measurements were done on a cylindrical blade in a linear cascade arrangement. An infrared camera system was used to determine the effectiveness of this newly designed cooling system by measuring the temperature distribution on the blade surface. Then a numerical simulation of heat transfer and of internal and external cooling for the turbine blade at test rig conditions was performed. The ICS showed a lower outer wall temperature distribution of the blade compared to a standard film cooling system. The heavily loaded leading edge as well as the trailing edge are well cooled. Further conclusions on the advantages and disadvantages of the ICS are drawn.

Recent Advances in the Study of Film Cooling on the Gas Turbine Blade

2014 ◽  
Vol 971-973 ◽  
pp. 143-147 ◽  
Author(s):  
Ping Dai ◽  
Shuang Xiu Li
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Performance ◽  
Leading Edge ◽  
Development Trend ◽  
Research Progress ◽  
Gas Turbine Blade ◽  
New Generation

The development of a new generation of high performance gas turbine engines requires gas turbines to be operated at very high inlet temperatures, which are much higher than the allowable metal temperatures. Consequently, this necessitates the need for advanced cooling techniques. Among the numerous cooling technologies, the film cooling technology has superior advantages and relatively favorable application prospect. The recent research progress of film cooling techniques for gas turbine blade is reviewed and basic principle of film cooling is also illustrated. Progress on rotor blade and stationary blade of film cooling are introduced. Film cooling development of leading-edge was also generalized. Effect of various factor on cooling effectiveness and effect of the shape of the injection holes on plate film cooling are discussed. In addition, with respect to progress of discharge coefficient is presented. In the last, the future development trend and future investigation direction of film cooling are prospected.

Improvement of Film Cooling Effectiveness in the Gas Turbine Endwall by Use of Optimization Framework

2019 ◽  
Author(s):  
Daisuke Hata ◽  
Kazuto Kakio ◽  
Yutaka Kawata ◽  
Masahiro Miyabe
Keyword(s):  
Pressure Gradient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness

Abstract Recently, the number of gas turbine combined cycle plants is rapidly increasing in substitution of nuclear power plants. The turbine inlet temperature (TIT) is constantly being increased in order to achieve higher effectiveness. Therefore, the improvement of the cooling technology for high temperature gas turbine blades is one of the most important issue to be solved. In a gas turbine, the main flow impinging at the leading edge of the turbine blade generates a so called horseshoe vortex by the interaction of its boundary layer and generated pressure gradient at the leading edge. The pressure surface leg of this horseshoe vortex crosses the passage and reaches the blade suction surface, driven by the pressure gradient existing between two consecutive blades. In addition, this pressure gradient generates a cross-flow along the endwall. This all results into a very complex flow field in proximity of the endwall. For this reason, burnouts tend to occur at a specific position in the vicinity of the leading edge. In this research, a methodology to cool the endwall of the turbine blade by means of film cooling jets from the blade surface and the endwall is proposed. The cooling performance is investigated using the transient thermography method. CFD analysis is also conducted to investigate the phenomena occurring at the endwall and calculate the film cooling effectiveness.

Simulation Film Cooling in the Leading Edge Region of a Turbine Blade (Trench Effect on Film Effectiveness from Cylinder in Crossflow)

2014 ◽  
Vol 554 ◽  
pp. 317-321
Author(s):  
Mohamad Rasidi Bin Pairan ◽  
Norzelawati Binti Asmuin ◽  
Hamidon bin Salleh
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Edge Region ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Hot Gas ◽  
Cooling Technique

Film cooling is one of the cooling techniques applied to the turbine blade. Gas turbine used film cooling technique to protect turbine blade from directly expose to the 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 coolant holes geometry. Comparisons are made at four difference blowing ratios which are 1.0, 1.25 and 1.5. 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. 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.

Computations of Film Cooling for the Leading Edge Region of a Turbine Blade Model

1995 ◽  
Author(s):  
Pingfan He ◽  
Martha Salcudean ◽  
Ian S. Gartshore
Keyword(s):  
Turbine Blade ◽  
Mass Flow ◽  
Film Cooling ◽  
Leading Edge ◽  
Relative Mass ◽  
Computational Domain ◽  
Blade Surface ◽  
Stagnation Line ◽  
Curved Blade ◽  
Blade Model

Computations of film cooling are presented based on the geometry of a UBC experimental turbine blade model. This model has a semi-circlar leading edge with four rows of laterally-inclined film cooling orifices positioned symmetrically about the stagnation line. The computational domain follows the physical domain and includes the curved blade surface as well as the coolant regions in the circular coolant orifices. The injection orifices are inclined spanwise at 30° to the blade surface. A multi-zone curvilinear grid is used to simulate the complex configuration. Grids are generated by a block-structured elliptic grid generation method which represents exactly the curved blade surface as well as the circular injection orifices. Computations over the cooled turbine blade model are carried out for overall mass flow ratios of 0.52 and 0.97. The relative mass flow ratios from each orifice are specified to match experimental values. Density ratios of coolant to free stream were taken to be unity (constant density). Comparison of predicted film cooling effectiveness with experimental data showed reasonable agreement.

A210 Experimental and Numerical Studies on Leading Edge Film Cooling of a HP Turbine Blade Model(Gas Turbine-6)

2009 ◽  
Vol 2009.2 (0) ◽  
pp. _2-55_-_2-60_
Author(s):  
Ken-ichi FUNAZAKI ◽  
Naota YAMACHI ◽  
Hamidon BIN SALLEH ◽  
Toshihiko TAKAHASHI ◽  
Eiji SAKAI ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Numerical Studies ◽  
Blade Model

Comparative Study Between a Cut-Back and Conventional Trailing Edge Film Cooling System

2006 ◽  
Author(s):  
Janendra C. Telisinghe ◽  
Peter T. Ireland ◽  
Terry V. Jones ◽  
David Barrett ◽  
Changmin Son
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Total Pressure ◽  
Cooling System ◽  
Pressure Measurements ◽  
Flat Plates ◽  
Trailing Edge ◽  
Scaled Model ◽  
Cooling Hole ◽  
Low Speed Wind Tunnel

The present experimental study investigates the aero performance differences between a conventional turbine blade trailing edge and a trailing edge with a sharp cut-back. Both geometries include trailing edge film cooling. A scaled model of a conventional turbine blade trailing edge and the trailing edge with a sharp cut-back including the scaled film cooling hole geometries were incorporated into flat plates. Experiments were conducted in a low speed wind tunnel to establish the performance change caused by introducing such a cut-back system on the trailing edge. Experiments were conducted at a Reynolds number of 1.9 × 106 at the trailing edge and for blowing ratios from 0.9 to 2.4. The experimental data presented include; static pressure variation on the plate surfaces; the change in discharge coefficient due to the cut-back; detailed total pressure measurements via a 2D traverse mechanism at a plane 40mm downstream of the flat plates. The total pressure measurements were used to establish the mixed out loss for both configurations.

scholarly journals Quasi-3D Numerical Computations on a Film-Cooled Gas Turbine Nozzle

1998 ◽  
Author(s):  
F. Bassi ◽  
S. Rebay ◽  
M. Savini
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Complex Geometry ◽  
Computational Effort ◽  
Navier Stokes ◽  
Gas Turbine Blade ◽  
Trailing Edge ◽  
Density Ratios ◽  
The Impact

The aim of this work is to assess the accuracy of a “quasi-3d” Navier-Stokes solver equipped with the k-ω turbulence model in the computation of a film-cooled gas turbine blade under a variety of flow conditions. The “quasi-3d” formulation was chosen as a cheap approach to investigate a large number of test conditions for a nozzle of complex geometry (around 400 cooling holes) which would require a large computational effort for a truly 3d simulation. The developed code has been used to investigate the influence of various cooling geometries and blowing conditions (mass flow rate and/or density ratios) on the aerodynamic behaviour of the cascade (in terms of loading, losses and flow angles) and their impact on the mixing process downstream of the trailing edge. The investigated nozzle is an advanced design turbine vane working in high subsonic regime. It is characterized by a marked endwall contouring at the casing and by the presence of 12 rows of holes (including a trailing edge row of slots) so as to obtain full-coverage film-cooling of the solid surfaces. This vane has been extensively tested in the Politecnico di Milano Fluid Dynamics Laboratory (formerly C.N.P.M.) blowdown transonic wind tunnel and a great amount of data are therefore available for validation purposes. The uselfulness of the proposed approach is fully analyzed and discussed throughout the paper and it is shown that the relation between the cascade performance and the variation of the investigated parameters is correctly described. In addition we address and discuss which ejection boundary conditions and which loss definitions are best suited for a meaningful comparison with the experimental measurements. In conclusion, in the case considered the developed code seems to be a valuable tool to determine the impact of film-cooling on the aerodynamic performance of a gas turbine blade.

Sign in / Sign up

Export Citation Format

Share Document