Exploitation of Acoustic Effects in Film Cooling

2015 ◽  
Vol 137 (10) ◽  
Author(s):  
Matthew Collins ◽  
Thomas Povey
Keyword(s):  
Film Cooling ◽  
Pressure Wave ◽  
Cooling System ◽  
Large Pressure ◽  
Cooling Hole ◽  
Cooling Design ◽  
Computational Fluid Dynamics Cfd ◽  
Rate Profile ◽  
Film Cooling Holes ◽  
First Time

There have been numerous studies of the behavior of shaped film cooling holes for turbine applications. It is known that the introduction of coolant is an unsteady process, and a handful of studies have described and characterized the unsteadiness. To the best of our knowledge, there are no studies in which unsteady acoustic effects have been actively exploited such that they have led to novel designs with improved cooling performance. This paper discusses the fundamental mechanism of pressure wave propagation through cooling holes and describes systems in which holes which have been acoustically shaped have led to a direct improvement in film cooling hole performance. The mechanism relies on sequential pressure wave reflection within an acoustically shaped hole and is therefore applicable in regions where the external surface is subject to large pressure wave fluctuations at high frequency. The principle is developed analytically, and then demonstrated with a number of computational fluid dynamics (CFD) simulations. We demonstrate that a desired temporal mass flow rate profile can be achieved by appropriate acoustic shaping of the cooling hole. The purpose of this paper is to describe the fundamental design considerations relevant to acoustic shaping. The discussion is developed with reference to a film cooling system for the over-tip region of an unshrouded rotor. The performance benefit of the system in terms of modulation of unsteady mass flux and ingestion characteristics is quantified. It is believed that this is the first time this significant effect has been exploited in film cooling design.

Exploitation of Acoustic Effects in Film Cooling

2014 ◽  
Author(s):  
Matthew Collins ◽  
Thomas Povey
Keyword(s):  
Film Cooling ◽  
Pressure Wave ◽  
Cooling System ◽  
Large Pressure ◽  
Cooling Hole ◽  
Cooling Design ◽  
Rate Profile ◽  
Shrouded Rotor ◽  
Film Cooling Holes ◽  
First Time

There have been numerous studies of the behavior of shaped film cooling holes for turbine applications. It is known that the introduction of coolant is an unsteady process, and a handful of studies have described and characterized the unsteadiness. To the authors’ knowledge there are no studies in which unsteady acoustic effects have been actively exploited such that they have led to novel designs with improved cooling performance. This paper discusses the fundamental mechanism of pressure wave propagation through cooling holes, and describes systems in which holes which have been acoustically shaped have led to a direct improvement in film cooling hole performance. The mechanism relies on sequential pressure wave reflection within an acoustically shaped hole, and is therefore applicable in regions where the external surface is subject to large pressure wave fluctuations at high frequency. The principle is developed analytically, and then demonstrated with a number of CFD simulations. We demonstrate that a desired temporal mass flow rate profile can be achieved by appropriate acoustic shaping of the cooling hole. The purpose of this paper is to describe the fundamental design considerations relevant to acoustic shaping. The discussion is developed with reference to a film cooling system for the over-tip region of an un-shrouded rotor. The performance benefit of the system in terms of modulation of unsteady mass flux and ingestion characteristics is quantified. It is believe that this is the first time this significant effect has been exploited in film cooling design.

Enhanced Film Cooling Effectiveness With Surface Trenches

2013 ◽  
Author(s):  
K. Vighneswara Rao ◽  
Jong S. Liu ◽  
Daniel C. Crites ◽  
Luis A. Tapia ◽  
Malak F. Malak ◽  
...  
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Operating Conditions ◽  
Airfoil Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Ansys Cfx ◽  
Cooling Hole ◽  
Computational Fluid Dynamics Cfd ◽  
Film Cooling Holes

In this study, cylindrical and fan shaped film cooling holes are evaluated on the blade surface numerically, using the Computational Fluid Dynamics (CFD) tool ANSYS-CFX, with the objective of improving cooling effectiveness by understanding the flow pattern at the cooling hole exit. The coolant flow rates are adjusted for blowing ratios of 0.5, 1.0 & 1.5 (momentum flux ratios of 0.125, 0.5 & 1.125 respectively). The density ratio is maintained at 2.0. New shaped holes viz. straight, concave and convex trench holes are introduced and are evaluated under similar operating conditions. Results are presented in terms of surface temperatures and adiabatic effectiveness at three different blowing ratios for the different film cooling hole shapes analyzed. Comparison is made with reference to the fan shaped film cooling hole to bring out relative merits of different shapes. The new trench holes improved the film cooling effectiveness by allowing more residence time for coolant to spread laterally while directing smoothly onto the airfoil surface. While convex trench improved the centre-line effectiveness, straight trench improved the laterally-averaged and overall effectiveness at all blowing ratios. Concave trench improved the effectiveness at blowing ratios 0.5 and 1.0.

Enhanced Cooling Effectiveness in Full-Coverage Film Cooling System With Impingement Jets

2008 ◽  
Author(s):  
Sang Hyun Oh ◽  
Dong Hyun Lee ◽  
Kyung Min Kim ◽  
Moon Young Kim ◽  
Hyung Hee Cho
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Hole Diameter ◽  
Cooling Effectiveness ◽  
Cooling Plate ◽  
Film Cooling Effectiveness ◽  
Impingement Jet ◽  
Full Coverage ◽  
Cooling Hole ◽  
Film Cooling Holes

An experimental investigation is conducted on the cooling effectiveness of full-coverage film cooled wall with impingement jets. Film cooling plate is made of stainless steel, thus the adiabatic film cooling effectiveness and the cooling effect of impingement jet underneath the film cooling plate are comprised in the cooling effectiveness. Infra-red camera is used to measure the temperature of film cooled surfaces. Experiments are conducted with different film cooling hole angles, such as 35° and 90°. Diameters of both film cooling holes and impinging jet holes are 5 mm. The jet Reynolds number base on the hole diameter (Red) ranges from 3,000 to 5,000 and equivalent blowing ratios (M) varies from 0.3 to 0.5, respectively. The distance between the injection plate and the film cooling plate is 1, 3 and 5 times of the hole diameter. The streamwise and spanwise hole spacing to the hole diameter ratio (p/d) are 3 for both the film cooling hole plate and the impingement jet hole plate. The 35° angled film cooling hole arrangement shows higher film cooling effectiveness than the 90° film cooling hole arrangement. As the blowing ratio increases, the cooling effectiveness is enhanced for both the 35° almost constant regardless of H/d, while H/d = 1 shows a minimum value for the angled film cooling hole.

Novel Turbine Rotor Shroud Film-Cooling Design and Validation: Part 1

2009 ◽  
Author(s):  
K. S. Chana ◽  
B. Haller
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Turbine Rotor ◽  
Test Facility ◽  
Pressure Probe ◽  
Dimensional Approach ◽  
Flow Angle ◽  
Cooling Hole ◽  
Cooling Design

This paper is part one of a two part paper which considers a shroud film-cooling system designed using a two-dimensional approach. Heat transfer to rotor-casings has reached levels that are causing in-service difficulties to be experienced. Future designs are likely to need to employ film-cooling of some form. There is currently very little information available for film-cooling on shroudless turbine rotor-casing liners. Heat transfer literature on uncooled configurations is not extensive and in particular, spatially-detailed, time-accurate data are rare. This paper describes the aero-thermodynamic design and validation of a rotor casing film-cooling system for a transonic, high-pressure shroudless turbine stage. The design was carried out using a boundary layer code with the film-cooling hole geometry representative of an engine configuration and, has been subjected to mechanical constraints similar to those for an engine component. The design consists of two double rows of cooling holes and two ‘cooling-hole’ shape configurations, cylindrical and fan shaped. The design was tested in the QinetiQ short duration turbine test facility (TTF). Measurements taken include casing heat transfer using thin film gauges and stage exit total pressure, Mach number and flow angle using a three-hole pressure probe. Results showed that while the cooling produced a reduction in the heat transfer rate close to the injection point, the film was stripped off the casing and entrained in nozzle guide vane secondary and rotor overtip flow, where it was transported spanwise towards the hub in the rotor passage. Using the results obtained from this deign a second cooling design was carried out, using a three-dimensional approach this gave significantly better cooling performance. The thee-dimensional design and validation is reported in GT2009-60246 as part 2 of this paper.

Influence of Film-Hole Shape and Angle on Showerhead Film Cooling Using PSP Technique

2009 ◽  
Vol 131 (6) ◽  
Author(s):  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Cooling Design ◽  
Cylindrical Holes ◽  
Film Cooling Holes ◽  
Compound Angle

The effect of film-hole geometry and angle on turbine blade leading edge film cooling has been experimentally studied using the pressure sensitive paint technique. The leading edge is modeled by a blunt body with a semicylinder and an after-body. Two film cooling designs are considered: a heavily film cooled leading edge featured with seven rows of film cooling holes and a moderately film cooled leading edge with three rows. For the seven-row design, the film holes are located at 0 deg (stagnation line), ±15 deg, ±30 deg, and ±45 deg on the model surface. For the three-row design, the film holes are located at 0 deg and ±30 deg. Four different film cooling hole configurations are applied to each design: radial angle cylindrical holes, compound angle cylindrical holes, radial angle shaped holes, and compound angle shaped holes. Testing was done in a low speed wind tunnel. The Reynolds number, based on mainstream velocity and diameter of the cylinder, is 100,900. The mainstream turbulence intensity is about 7% near of leading edge model and the turbulence integral length scale is about 1.5 cm. Five averaged blowing ratios are tested ranging from M=0.5 to M=2.0. The results show that the shaped holes provide higher film cooling effectiveness than the cylindrical holes, particularly at higher average blowing ratios. The radial angle holes give better effectiveness than the compound angle holes at M=1.0–2.0. The seven-row film cooling design results in much higher effectiveness on the leading edge region than the three-row design at the same average blowing ratio or same amount coolant flow.

Experimental Evaluations of the Relative Contributions to Overall Effectiveness in Turbine Blade Leading Edge Cooling

2019 ◽  
Vol 141 (4) ◽  
Author(s):  
Carol E. Bryant ◽  
Connor J. Wiese ◽  
James L. Rutledge ◽  
Marc D. Polanka
Keyword(s):  
Biot Number ◽  
Film Cooling ◽  
Leading Edge ◽  
Internal Surface ◽  
Internal Cooling ◽  
Surface Cooling ◽  
External Cooling ◽  
Cooling Design ◽  
Computational Fluid Dynamics Cfd ◽  
Film Cooling Holes

Gas turbine components are protected through a combination of internal cooling and external film cooling. Efforts aimed at improving cooling are often focused on either the internal cooling or the film cooling; however, the common coolant flow means the internal and external cooling schemes are linked and the coolant holes themselves provide another convective path for heat transfer to the coolant. Measurements of overall cooling effectiveness, ϕ, using matched Biot number models allow evaluation of fully cooled components; however, the relative contributions of internal cooling, external cooling, and convection within the film cooling holes are not well understood. Matched Biot number experiments, complemented by computational fluid dynamics (CFD) simulations, were performed on a fully film cooled cylindrical leading edge model to quantify the effects of alterations in the cooling design. The relative influence of film cooling and cooling within the holes was evaluated by selectively disabling individual holes and quantifying how ϕ changed. Testing of several impingement cooling schemes revealed that impingement has a negligible influence on ϕ in the showerhead region. This indicates that the pressure drop penalties with impingement may not always be compensated by an increase in ϕ. Instead, internal cooling from convection within the holes and film cooling were shown to be the dominant contributors to ϕ. Indeed, the numerous holes within the showerhead region impede the ability of internal surface cooling schemes to influence the outside surface temperature. These results may allow improved focus of efforts on the forms of cooling with the greatest potential to improve performance.

Combined Experimental and CFD Study of a HP Blade Multi-Pass Cooling System

2009 ◽  
Author(s):  
D. Jackson ◽  
P. Ireland ◽  
B. Cheong
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Detailed Comparison ◽  
Bulk Temperature ◽  
Internal Cooling ◽  
3 Dimensional ◽  
Turbine Cooling ◽  
Cooling Hole ◽  
The Difference

Progress in the computing power available for CFD predictions now means that full geometry, 3 dimensional predictions are now routinely used in internal cooling system design. This paper reports recent work at Rolls-Royce which has compared the flow and htc predictions in a modern HP turbine cooling system to experiments. The triple pass cooling system includes film cooling vents and inclined ribs. The high resolution heat transfer experiments show that different cooling performance features are predicted with different levels of fidelity by the CFD. The research also revealed the sensitivity of the prediction to accurate modelling of the film cooling hole discharge coefficients and a detailed comparison of the authors’ computer predictions to data available in the literature is reported. Mixed bulk temperature is frequently used in the determination of heat transfer coefficient from experimental data. The current CFD data is used to compare the mixed bulk temperature to the duct centreline temperature. The latter is measured experimentally and the effect of the difference between mixed bulk and centreline temperature is considered in detail.

Introducing a New Test Rig for Film Cooling Measurements With Realistic Hole Inflow Conditions

2017 ◽  
Author(s):  
Marc Fraas ◽  
Tobias Glasenapp ◽  
Achmed Schulz ◽  
Hans-Jörg Bauer
Keyword(s):  
Film Cooling ◽  
Gas Flow ◽  
Test Rig ◽  
Hot Gas ◽  
Cooling Hole ◽  
Measurement Principle ◽  
Density Ratios ◽  
Operational Range ◽  
Film Cooling Holes

Further improvements in film cooling require an in-depth understanding of the influencing parameters. Therefore, a new test rig has been designed and commissioned for the assessment of novel film cooling holes under realistic conditions. The test rig is designed for generic film cooling studies. External hot gas flow as well as internal coolant passage flow are simulated by two individual flow channels connected to each other by the cooling holes. Based on a similarity analysis, the geometry of the test rig is scaled up by a factor of about 20. It furthermore offers the possibility to conduct experiments at high density ratios and realistic approach flow conditions at both cooling hole exit and inlet. The operational range of the new test rig is presented and compared to real engine conditions. It is shown that the important parameters are met and the transfer-ability of the results is ensured. Special effort is put onto the uniformity of the approaching hot gas flow, which will be demonstrated by temperature and velocity profiles. A first measurement of the heat transfer coefficient without film cooling is used to demonstrate the quality of the measurement principle.

Computational Predictions of Endwall Film Cooling for a Turbine Nozzle Vane With an Asymmetric Contoured Passage

2008 ◽  
Author(s):  
Yoji Okita ◽  
Chiyuki Nakamata
Keyword(s):  
Film Cooling ◽  
Computational Study ◽  
Suction Side ◽  
Geometrical Parameters ◽  
Passage Vortex ◽  
Rear Part ◽  
Cooling Hole ◽  
Secondary Vortices ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

This paper presents results of a computational study for the endwall film cooling of an annular nozzle cascade employing a circumferentially asymmetric contoured passage. The investigated geometrical parameters and the flow conditions are set consistent with a generic modern HP-turbine nozzle. Rows of cylindrical film cooling holes on the contoured endwall are arranged with a design practice for the ordinary axisymmetric endwall. The solution domain, which includes the mainflow, cooling hole paths, and the coolant plenum, is discretized in the RANS equations with the realizable k-epsilon model. The calculated flow field shows that the pressure gradients across the passage between the pressure and the suction side are reduced with the asymmetric endwall, and consequently, the rolling up of the inlet boundary layer into the passage vortex is delayed and the separation line has moved further downstream. With the asymmetric endwall, because of the effective suppression of the secondary flow, more uniform film coverage is achieved especially in the rear part of the passage and the laterally averaged effectiveness is also significantly improved in this region. The closer inspection of the calculated thermal field reveals that, with the asymmetric passage, the coolant ejected from the holes are less deflected by the secondary vortices, and it attaches better to the endwall in this rear part.

Computational Analysis of Conjugate Heat Transfer and Particulate Deposition on a High Pressure Turbine Vane

2009 ◽  
Author(s):  
Weiguo Ai ◽  
Thomas H. Fletcher
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Particle Method ◽  
Experimental Results ◽  
Turbine Vane ◽  
Particulate Deposition ◽  
Cooling Hole ◽  
Lagrangian Particle Method ◽  
Film Cooling Holes

Numerical computations were conducted to simulate flyash deposition experiments on gas turbine disk samples with internal impingement and film cooling using a CFD code (FLUENT). The standard k-ω turbulence model and RANS were employed to compute the flow field and heat transfer. The boundary conditions were specified to be in agreement with the conditions measured in experiments performed in the BYU Turbine Accelerated Deposition Facility (TADF). A Lagrangian particle method was utilized to predict the ash particulate deposition. User-defined subroutines were linked with FLUENT to build the deposition model. The model includes particle sticking/rebounding and particle detachment, which are applied to the interaction of particles with the impinged wall surface to describe the particle behavior. Conjugate heat transfer calculations were performed to determine the temperature distribution and heat transfer coefficient in the region close to the film-cooling hole and in the regions further downstream of a row of film-cooling holes. Computational and experimental results were compared to understand the effect of film hole spacing, hole size and TBC on surface heat transfer. Calculated capture efficiencies compare well with experimental results.

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