scholarly journals Numerical Investigations of Film Cooling and Particle Impact on the Blade Leading Edge

Energies ◽  
2021 ◽  
Vol 14 (4) ◽  
pp. 1102
Author(s):  
Ke Tian ◽  
Zicheng Tang ◽  
Jin Wang ◽  
Milan Vujanović ◽  
Min Zeng ◽  
...  
Keyword(s):  
Particle Size ◽  
Film Cooling ◽  
Gas Turbines ◽  
Three Dimensional ◽  
Leading Edge ◽  
Turbine Engines ◽  
Small Particles ◽  
The Impact ◽  
First Time ◽  
Blade Leading Edge

As a vital power propulsion device, gas turbines have been widely applied in aircraft. However, fly ash is easily ingested by turbine engines, causing blade abrasion or even film hole blockage. In this study, a three-dimensional turbine cascade model is conducted to analyze particle trajectories at the blade leading edge, under a film-cooled protection. A deposition mechanism, based on the particle sticking model and the particle detachment model, was numerically investigated in this research. Additionally, the invasion efficiency of the AGTB-B1 turbine blade cascade was investigated for the first time. The results indicate that the majority of the impact region is located at the leading edge and on the pressure side. In addition, small particles (1 μm and 5 μm) hardly impact the blade’s surface, and most of the impacted particles are captured by the blade. With particle size increasing, the impact efficiency increases rapidly, and this value exceeds 400% when the particle size is 50 μm. Invasion efficiencies of small particles (1 μm and 5 μm) are almost zero, and the invasion efficiency approaches 12% when the particle size is 50 μm.

Numerical Predictions of Film Cooled NGV Blades

2003 ◽  
Author(s):  
P. Adami ◽  
F. Martelli ◽  
K. S. Chana ◽  
F. Montomoli
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Cooling System ◽  
Critical Evaluation ◽  
Three Dimensional ◽  
Leading Edge ◽  
Test Case ◽  
Numerical Predictions ◽  
Turbine Vanes ◽  
Cylindrical Holes

Film-cooling is commonly used in modern gas turbines to increase inlet temperatures without compromising the mechanical strength of the hot components. The main objective of the study reported here is the critical evaluation of the capability of CFD, to predict film-cooling on three-dimensional engine realistic turbine aerofoil geometries. To achieve this aim two different film-cooling systems for NGV aerofoils are predicted and compared against experiments. The application concerns the following turbine vanes: • the AGTB-B1 blade investigated by the “Institut fur Strahlantriebe of the Universitat der Bundeswehr Munchen (Germany)”; • the MT1 HP NGV investigated by QinetiQ (ex DERA, UK). In the first test case the application mainly focuses on the interaction between the main flow and the coolant jets on the leading edge of the cooled aerofoil. In the second case, vane heat transfer rate is predicted with the film-cooling system made of six rows of cylindrical holes in single and staggered configuration.

Aerothermal Investigations of Mixing Flow Phenomena in Case of Radially Inclined Ejection Holes at the Leading Edge

1999 ◽  
Vol 122 (2) ◽  
pp. 334-339 ◽  
Author(s):  
Dieter E. Bohn ◽  
Karsten A. Kusterer
Keyword(s):  
Secondary Flow ◽  
Gas Turbines ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Back Side ◽  
Stagnation Line ◽  
Flow Phenomena ◽  
The Impact

A leading edge cooling configuration is investigated numerically by application of a three-dimensional conjugate fluid flow and heat transfer solver, CHT-flow. The code has been developed at the Institute of Steam and Gas Turbines, Aachen University of Technology. It works on the basis of an implicit finite volume method combined with a multiblock technique. The cooling configuration is an axial turbine blade cascade with leading edge ejection through two rows of cooling holes. The rows are located in the vicinity of the stagnation line, one row on the suction side, the other row is on the pressure side. The cooling holes have a radial ejection angle of 45 deg. This configuration has been investigated experimentally by other authors and the results have been documented as a test case for numerical calculations of ejection flow phenomena. The numerical investigations focus on the aerothermal mixing process in the cooling jets and the impact on the temperature distribution on the blade surface. The radial ejection angles lead to a fully three-dimensional and asymmetric jet flow field. Within a secondary flow analysis, the cooling fluid jets are investigated in detail. The secondary flow fields include asymmetric kidney vortex systems with one dominating vortex on the back side of the jets. The numerical and experimental data show a respectable agreement concerning the vortex development. [S0889-504X(00)00102-1]

A Systematic Experimental Study on the Aerodynamics of Leading Edge Film Cooling on a Large Scale High Pressure Turbine Cascade

1998 ◽  
Author(s):  
Sabine Ardey ◽  
Leonhard Fottner
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Injection Site ◽  
Film Cooling ◽  
Gas Turbines ◽  
Large Scale ◽  
Three Dimensional ◽  
Leading Edge ◽  
Turbine Cascade ◽  
High Pressure Turbine

Systematic isothermal investigations on the aerodynamic effects of leading edge film cooling were carried out on a large scale high pressure turbine cascade named AGTB. In the vicinity of the stagnation point the AGTB turbine cascade has one injection site on the suction side and one on the pressure side. Three injection geometries were tested: Slots (two dimensional geometry), streamwise inclined holes (symmetrical three dimensional geometry) and compound angle holes (fully three dimensional geometry). The injection angle in streamwise direction, the blowing ratio, the inlet turbulence intensity, the inlet Mach number, and the inlet Reynolds number were kept constant at values typically found in modern gas turbines. The measured data comprise the coolant plenum state, the cascade inlet conditions, the flow field in the cascade exit plane including secondary flows, the static pressure distribution in the mid span section of the blade and in the near hole region, the coolant flow field close to the injection site on the leading edge, Schlieren images of the coolant penetration height and oil-and-dye flow visualizations of the blade surface. The experimental data are summarized and documented as a test case that can be used for validation purposes of prediction methods.

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.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

2017 ◽  
Author(s):  
Mingjie Zhang ◽  
Nian Wang ◽  
Andrew F. Chen ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Profile ◽  
Elliptical Cylinders ◽  
Blade Leading Edge

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio and blowing ratio are studied. Computational simulations are performed using the realizable k-ε turbulence model. Effectiveness obtained by CFD simulations are compared with experiments. Three leading edge profiles, including one semi-cylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semi-cylinder model, shaped holes are located at 0 degrees (stagnation line) and ± 30 degrees. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,900 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile and on turbine blade leading edge region film cooling with shaped-hole designs.

scholarly journals State-of-the-Art Cooling Technology for a Turbine Rotor Blade

2017 ◽  
Author(s):  
Jason Town ◽  
Doug Straub ◽  
James Black ◽  
Karen Thole ◽  
Tom Shih
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
State Of The Art ◽  
Internal Heat ◽  
Leading Edge ◽  
Jet Impingement ◽  
Test Facility ◽  
Turbine Wheel ◽  
Baseline Design ◽  
External Cooling

Effective internal and external cooling of airfoils is key to maintaining component life for efficient gas turbines. Cooling designs have spanned the range from simple internal convective channels to more advanced double-walls with shaped film-cooling holes. This paper describes the development of an internal and external cooling concept for a state-of-the-art cooled turbine blade. These cooling concepts are based on a review of literature and patents, as well as, interactions with academic and industry turbine cooling experts. The cooling configuration selected and described in this paper is referred to as the “baseline” design, since this design will simultaneously be tested with other more advanced blade cooling designs in a rotating turbine test facility using a “rainbow turbine wheel” configuration. For the baseline design, the leading edge is cooled by internal jet impingement and showerhead film cooling. The mid-chord region of the blade contains a three-pass serpentine passage with internal discrete V-shaped trip strips to enhance the internal heat transfer coefficient. The film cooling along the mid-chord of the blade uses multiple rows of shaped diffusion holes. The trailing edge is internally cooled using jet impingement and externally film cooled through partitioned cuts on the pressure side of the blade.

Experimental Study of the Effects of an Oscillating Approach Flow on Overall Cooling Performance of a Simulated Turbine Blade Leading Edge

2009 ◽  
Author(s):  
Ross Johnson ◽  
Jonathan Maikell ◽  
David Bogard ◽  
Justin Piggush ◽  
Atul Kohli ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Oscillation Frequencies ◽  
Overall Effectiveness ◽  
Blade Leading Edge

When a turbine blade passes through wakes from upstream vanes it is subjected to an oscillation of the direction of the approach flow resulting in the oscillation of the position of the stagnation line on the leading edge of the blade. In this study an experimental facility was developed that induced a similar oscillation of the stagnation line position on a simulated turbine blade leading edge. The overall effectiveness was evaluated at various blowing ratios and stagnation line oscillation frequencies. The location of the stagnation line on the leading edge was oscillated to simulate a change in angle of attack between α = ± 5° at a range of frequencies from 2 to 20 Hz. These frequencies were chosen based on matching a range of Strouhal numbers typically seen in an engine due to oscillations caused by passing wakes. The blowing ratio was varied between M = 1, M = 2, and M = 3. These experiments were carried out at a density ratio of DR = 1.5 and mainstream turbulence levels of Tu ≈ 6%. The leading edge model was made of high conductivity epoxy in order to match the Biot number of an actual engine airfoil. Results of these tests showed that the film cooling performance with an oscillating stagnation line was degraded by as much as 25% compared to the performance of a steady flow with the stagnation line aligned with the row of holes at the leading edge.

Optimization of Trenched Film Cooling Using RSM Coupled CFD

2018 ◽  
Vol 0 (0) ◽  
Author(s):  
V.G. Krishna Anand ◽  
K.M. Parammasivam
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Response Surfaces ◽  
Test Surface ◽  
Response Factor ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Flow Parameters ◽  
Considerable Impact ◽  
The Impact

AbstractThe trench film cooling employs film holes embedded in a slot created on the surface that requires protection from the impact of hot mainstream flow. The present investigation employs Response Surface Methodology (RSM) approach coupled with CFD analysis to develop a regression predictive model and to optimize the trench geometric and flow parameters viz., trench width (w), trench depth (d), film hole compound angle (β) and blowing ratio (M). The Area-averaged film cooling effectiveness (ȠAA) were chosen as a response factor for RSM and with trench design and flow parameters used as input factors for regression analysis. Analysis of variance (ANOVA) analysis was carried out on the regression model to identify the influence of individual parameters. Three dimensional response surfaces that relate the effect of input parameters on the response factor were analyzed. Experimental results of a case identified from the RSM matrix was found to correlate well with computational investigations. Results from the study indicate that the parameters d, β and M have considerable impact on film cooling performance of test surface with trenches.

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