Experimental Survey on Heat Transfer in a Trailing Edge Cooling System: Effects of Rotation in Internal Cooling Ducts

2012 ◽  
Author(s):  
L. Bonanni ◽  
C. Carcasci ◽  
B. Facchini ◽  
L. Tarchi
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Cooling System ◽  
Internal Cooling ◽  
Test Rig ◽  
Trailing Edge ◽  
Tip Mass

The high thermal loads, the heavy structural stresses and the small thickness required for aerodynamic performances make the trailing edge cooling (TE) cooling of high pressure gas turbine blades a critical challenge. The presented paper point out an experimental study focusing the aerothermal performance of a TE internal cooling system of a high pressure gas turbine blade, evaluated under stationary and rotating conditions. The investigated geometry consists of a 30:1 scaled model reproducing the typical wedge shaped discharge duct with one row of enlarged pedestals. The airflow pattern inside the device simulates a highly loaded rotor blade cooling scheme with a 90° turning flow from the radial hub inlet to the tangential TE outlet. Two different tip configurations were tested, the first one with a completely closed section, the second one with 5 holes on the tip outlet surfaces discharging at ambient pressure. To investigate the rotation effects on the trailing edge cooling system performance, a rotating test rig was purposely developed and manufactured. The test rig is composed by a rotating arm that holds the PMMA TE model and the instrumentation. A thin Inconel heating foil and wide band Thermo-chromic Liquid Crystals are used to perform steady state heat transfer measurements. A rotary joint ensures the pneumatic connection between the blower and the rotating apparatus, moreover several slip rings are used for both instrumentation power supply and thermocouple connection. Heat transfer coefficient measurements were made with fixed Reynolds number close to 20k in the hub inlet section and with variable rotating speed in order to set the Rotation number from 0 (non rotational test) up to 0.3. Six different configurations were tested: two different tip mass flow rates (the first one with a completely closed tip, the second one with the 12.5% of the inlet flow discharged from the tip) and three different surface conditions: the first one consists in the flat plate case and the others in two ribbed cases, with different angular orientation (60° and −60° respect to the radial direction). Results are reported in terms of detailed heat transfer coefficient 2D maps on the suction side surface as well as span-wise profiles inside the pedestal ducts. The reported work has been supported by the Italian Ministry of Education, University and Research (MIUR).

Effect of Rotation on a Gas Turbine Blade Internal Cooling System: Numerical Investigation

2016 ◽  
Author(s):  
E. Burberi ◽  
D. Massini ◽  
L. Cocchi ◽  
L. Mazzei ◽  
A. Andreini ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Flow Field ◽  
Transfer Coefficient ◽  
Numerical Investigation ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Cooling System ◽  
Leading Edge ◽  
Internal Cooling

Increasing turbine inlet temperature is one of the main strategies used to accomplish the demands of increased performance of modern gas turbines. As a consequence, optimization of the cooling system is of paramount importance in gas turbine development. Leading edge represents a critical part of cooled nozzles and blades, given the presence of the hot gases stagnation point and the unfavourable geometry for cooling. This paper reports the results of a numerical investigation aimed at assessing the rotation effects on the heat transfer distribution in a realistic leading edge internal cooling system of a high pressure gas turbine blade. The numerical investigation was carried out in order to support and to allow an in-depth understanding of the results obtained in a parallel experimental campaign. The model is composed of a trapezoidal feeding channel which provides air to the cold bridge system by means of three large racetrack-shaped holes, generating coolant impingement on the internal concave leading edge surface, whereas four big fins assure the jets confinement. Air is then extracted through 4 rows of 6 holes reproducing the external cooling system composed of shower-head and film cooling holes. Experiments were performed in static and rotating conditions replicating the typical range of jet Reynolds number (Rej) from 10000 to 40000 and Rotation number (Roj) up to 0.05, for three crossflow cases representative of the working condition that can be found at blade tip, midspan and hub, respectively. Experimental results in terms of flow field measurements on several internal planes and heat transfer coefficient on the LE internal surface have been performed on two analogous experimental campaigns at University of Udine and University of Florence respectively. Hybrid RANS-LES models were used for the simulations, such as Scale Adaptive Simulation (SAS) and Detached Eddy Simulation (DES), given their ability to resolve the complex flow field associated with jet impingement. Numerical flow field results are reported in terms of both jet velocity profiles and 2D vector plots on symmetry and transversal internal planes, while the heat transfer coefficient distributions are presented as detailed 2D maps together with radial and tangential averaged Nusselt number profiles. A fairly good agreement with experimental measurements is observed, which represent a validation of the adopted computational model. As a consequence, the computed aerodynamic and thermal fields also allow an in-depth interpretation of the experimental results.

Heat Transfer of a Chordwise-Oriented Internal Cooling Passage Around a Turbine Airfoil

2001 ◽  
Author(s):  
M. K. Chyu ◽  
O. B. Ojo ◽  
C. H. Yen ◽  
R. S. Nordlund
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Cooling System ◽  
Jet Impingement ◽  
Local Heat Transfer ◽  
Local Heat ◽  
Internal Cooling ◽  
Trailing Edge ◽  
Cooling Passage

Abstract An innovative design of closed-loop cooling system for a stator airfoil consists of a number of internal cooling passages wrapping around both pressure and suction sides of the airfoil. The cooling passages feature (1) jet impingement post a sharp 90-degree turn at the passage inlet, (2) turbulators on the outermost wall, and (3) a nearly 180-degree turn in the trailing edge. In addition, the passage has an irregular cross-section and varies throughout its entire length. A series of heat transfer tests have been performed at Re = 17,000 ∼ 61,000, compared to this tests which uses a new approach, so-called the hybrid liquid crystal technique. The magnitude of local heat transfer coefficient rises sharply in three regions. The first maximum occurs in the region subjected to direct jet impingement as the flow turns into the channel. Compounded with the inlet effect, this maximum, in fact, is the highest heat transfer coefficient over the entire passage. The second and third peaks, both are comparable in magnitude, locate near the trailing edge of the airfoil where the flow experiences a 180-degree turn and near the passage exit with a 90-degree turn. The average value of heat transfer coefficient over the entire passage is about 1.9∼ 2.5 times higher than that with fully developed turbulent flow in a straight channel. This level of enhancement is comparable to that of the conventional ribturbulators with a 90-degree angle-of-attack.

Heat Transfer in an Airfoil Trailing Edge Configuration With Shaped Pedestals Mounted Internal Cooling Channel and Pressure Side Cutback

2006 ◽  
Author(s):  
Shuping P. Chen ◽  
Peiwen W. Li ◽  
Minking K. Chyu ◽  
Frank J. Cunha ◽  
William Abdel-Messeh
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Liquid Crystal ◽  
Transfer Coefficient ◽  
Internal Cooling ◽  
Hybrid Technique ◽  
Pressure Side ◽  
Cooling Channel ◽  
Trailing Edge ◽  
The Heat Transfer Coefficient

Described in this paper is an experimental study of heat transfer over a trailing edge configuration preceded with an internal cooling channel of pedestal array. The pedestal array consists of both circular pedestals and oblong shaped blocks. Downstream to the pedestal array, the trailing edge features pressure side cutback partitioned by the oblong shaped blocks. The local heat transfer coefficient over the entire wetted surface in the internal cooling chamber has been determined by using a “hybrid” measurement technique based on transient liquid crystal imaging. The hybrid technique employs the transient conduction model in a semi-infinite solid for resolving the heat transfer coefficient on the endwall surface uncovered by the pedestals. The heat transfer coefficient over a pedestal can be resolved by the lumped capacitance method with an assumption of low Biot number. The overall heat transfer for both the pedestals and endwalls combined shows a significant enhancement compared to the case with thermally developed smooth channel. Near the downstream most section of the suction side, the land, due to pressure side cutback, is exposed to the stream mixed with hot gas and discharged coolant. Both the adiabatic effectiveness and heat transfer coefficient on the land section are characterized by using the transient liquid crystal technique.

Detailed Experimental Characterization of Heat Transfer Coefficient Over the Internal Cooling Passages of an Additive Manufactured Turbine Airfoil

2016 ◽  
Author(s):  
Sin Chien Siw ◽  
Minking K. Chyu ◽  
Jae Y. Um ◽  
Ching-Pang Lee
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Local Heat Transfer ◽  
Local Heat ◽  
Local Heat Transfer Coefficient ◽  
Scale Model ◽  
Internal Cooling ◽  
Trailing Edge

This report describes the detailed experimental study to characterize the local heat transfer coefficient distribution over the internal cooling passages of a simplified generic airfoil. The airfoil is manufactured through additive manufacturing based on actual geometry and dimensions (1X scale model) of row one airfoil, applicable in large gas turbine system. At the mainbody section, the serpentine channel consists of three passages without any surface features or vortex generators. Both the leading edge and trailing edge sections are subjected to direct impingement. The trailing edge section is divided into three chambers, separated by two rows of blockages. This study employs the well-documented transient liquid crystal technique, where the local heat transfer coefficient on both pressure and suction sides is deduced. The experiments were performed at varying Reynolds number, ranging from approximately 31,000–63,000. The heat transfer distribution on the pressure side and suction side is largely comparable in the first and third pass, except for the second pass. Highest heat transfer occurs at the trailing edge region, which is ultimately dominated by impingement due to the presence of three rows of blockages. A cursory numerical calculation is performed using commercially available software, ANSYS CFX to obtain detailed flow field distribution within the airfoil, which explains the heat transfer behavior at each passage. The flow parameter results revealed that the pressure ratio is strongly proportional with increasing Reynolds number.

The Effects of Blade Passing on the Heat Transfer Coefficient of the Overtip Casing in a Transonic Turbine Stage

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
Steven J. Thorpe ◽  
Roger W. Ainsworth
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Cooling System ◽  
Turbine Stage ◽  
Transfer Coefficients ◽  
Time Resolved ◽  
Turbine Engine ◽  
Transonic Turbine

In a modern gas turbine engine, the outer casing (shroud) of the shroudless high-pressure turbine is exposed to a combination of high flow temperatures and heat transfer coefficients. The casing is consequently subjected to high levels of convective heat transfer, a situation that is complicated by flow unsteadiness caused by periodic blade-passing events. In order to arrive at an overtip casing design that has an acceptable service life, it is essential for manufacturers to have appropriate predictive methods and cooling system configurations. It is known that both the flow temperature and boundary layer conductance on the casing wall vary during the blade-passing cycle. The current article reports the measurement of spatially and temporally resolved heat transfer coefficient (h) on the overtip casing wall of a fully scaled transonic turbine stage experiment. The results indicate that h is a maximum when a blade tip is immediately above the point in question, while the lower values of h are observed when the point is exposed to the rotor passage flow. Time-resolved measurements of static pressure are used to reveal the unsteady aerodynamic situation adjacent to the overtip casing wall. The data obtained from this fully scaled transonic turbine stage experiment are compared to previously published heat transfer data obtained in low-Mach number cascade-style tests of similar high-pressure blade geometries.

scholarly journals Effect of Rotation on a Gas Turbine Blade Internal Cooling System: Experimental Investigation

2016 ◽  
Author(s):  
Daniele Massini ◽  
Emanuele Burberi ◽  
Carlo Carcasci ◽  
Lorenzo Cocchi ◽  
Bruno Facchini ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Flow Field ◽  
Transfer Coefficient ◽  
Cooling System ◽  
Cross Flow ◽  
Leading Edge ◽  
Impinging Jets ◽  
Internal Cooling ◽  
Transfer Pattern

A detailed aerothermal characterization of an advanced leading edge cooling system has been performed by means of experimental measurements. Heat transfer coefficient distribution has been evaluated exploiting a steady-state technique using Thermocromic Liquid Crystals (TLC), while flow field has been investigated by means of Particle Image Velocimetry (PIV). The geometry key features are the multiple impinging jets and the four rows of coolant extraction holes, which mass flow rate distribution is representative of real engine working conditions. Tests have been performed in both static and rotating conditions, replicating a typical range of jet Reynolds number (Rej), from 10000 to 40000, and Rotation number (Roj) up to 0.05. Different cross-flow conditions (CR) have been used to simulate the three main blade regions (i.e. tip, mid and hub). The aerothermal field turned out to be rather complex, but a good agreement between heat transfer coefficient and flow field measurement has been found. In particular, jet bending strongly depends on crossflow intensity, while rotation has a weak effect on both jet velocity core and area-averaged Nusselt number. Rotational effects increase for the lower cross-flow tests. Heat transfer pattern shape has been found to be substantially Reynolds-independent.

The Effects of Blade Passing on the Heat Transfer Coefficient of the Over-Tip Casing in a Transonic Turbine Stage

2006 ◽  
Author(s):  
Steven J. Thorpe ◽  
Roger W. Ainsworth
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Cooling System ◽  
Turbine Stage ◽  
Transfer Coefficients ◽  
Time Resolved ◽  
Turbine Engine ◽  
Transonic Turbine

In a modern gas turbine engine the outer casing (shroud) of the shroudless high-pressure turbine is exposed to a combination of high flow temperatures and heat transfer coefficients. The casing is consequently subjected to high levels of convective heat transfer, a situation that is complicated by flow unsteadiness caused by periodic blade-passing events. In order to arrive at an over-tip casing design that has an acceptable service life it is essential for manfacturers to have appropriate predictive methods and cooling system configurations. It is known that both the flow temperature and boundary layer conductance on the casing wall vary during the blade-passing cycle. The current article reports the measurement of spatially and temporally resolved heat transfer coefficient (h) on the over-tip casing wall of a fully-scaled transonic turbine stage experiment. The results indicate that h is a maximum when a blade-tip is immediately above the point in question, while lower values of h are observed when the point is exposed to the rotor passage flow. Time-resolved measurements of static pressure are used to reveal the unsteady aerodynamic situation adjacent to the over-tip casing wall. The data obtained from this fully-scaled transonic turbine stage experiment are compared to previously published heat transfer data obtained in low-Mach number cascade style tests of similar high pressure blade geometries.

Prediction of the Influence of Upstream Wake Passing on Downstream Blade Row External Heat Transfer Coefficient: A New Physically-Based Method

2010 ◽  
Author(s):  
K. S. Chana ◽  
B. R. Haller
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbines ◽  
Cooling System ◽  
Guide Vane ◽  
External Heat ◽  
Test Facility ◽  
Internal Cooling ◽  
External Heat Transfer

For gas turbines, accurate prediction of the external heat transfer coefficient on the high pressure (HP) turbine rotor blades is of immense importance, as this component is critical and operates at material limits. Furthermore the external heat load is the governing boundary condition for the design of the internal cooling system of the blade. There is a continuous drive to increase the turbine entry temperature to increase the cycle efficiency, whilst developing blade cooling systems with higher efficiency (i.e. using less cooling air). A new systematic procedure has been developed and validated to predict the external heat transfer to a blade surface. The procedure allows for the unsteady effects caused by the passing of upstream nozzle guide vane (NGV) wakes. The early part of the suction surface was shown to have a pessimistic prediction of external heat transfer coefficient which resulted in unnecessary over-cooling of the blade in this region. The heat transfer aspect is found from the well-known TEXSTAN differential boundary layer method, developed by Mike Crawford at Texas University from the original approach of Spalding & Patankar. The method is validated against the MT1 turbine tested in the QinetiQ Turbine Test Facility. Predictions and comparisons have also been carried out on the VKI turbine stage. The level of agreement with the test data is shown to be good.

Numerical Investigation of the Blade Tip and Overtip Casing Aerothermal Performance in a High Pressure Turbine Stage

2016 ◽  
Author(s):  
Kun Du ◽  
Zhigang Li ◽  
Jun Li
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Thermal Load ◽  
Rotor Blades ◽  
Turbine Stage ◽  
Cavity Depth ◽  
Blade Tip

In modern transonic gas turbine engines, the blade tip and overtip casing endures high thermal load, therefore these components are always subjected to thermal failures due to large unsteady heat flux. The unsteadiness is induced by the interaction of the rotor blades and periodic upstream wake of the vanes. The present study adopts a typical high pressure gas turbine stage (GE-E3 engine), and the computational domain consists of 1 high pressure stator vane and 2 rotor blades. The rotor blade in question has a squealer tip with a clearance gap about 1% of the blade height. This study focuses on the physics of the heat transfer characteristic of the blade tip and overtip casing regions. The present simulations were conducted using three-dimensional unsteady Reynolds-averaged Navier-Stokes (URANS) commercial code at the real engine conditions ( Mexit = 1.07, n = 8450rpm ). The standard k–ω turbulence model was utilized to model the turbulence. The accuracy of CFD predictions has been validated by comparison with the experimental data. The steady, unsteady and time-averaged results on the blade tip and overtip casing have been observed and discussed. Results indicate that the depth of the cavity has great influence on the blade tip and overtip casing. The averaged heat transfer coefficient on the blade tip is reduced with the increase of the cavity depth, however, the thermal load on the blade tip presents a contrary tendency. Moreover, the largest unsteadiness was observed for the case with D = 3.0 among the cases investigated, especially near the suction side squealer. In addition, the variation of the cavity depth has little effect on the heat transfer coefficient and thermal load on the overtip casing.

scholarly journals Effect of Rotation on a Gas Turbine Blade Internal Cooling System: Experimental Investigation

2017 ◽  
Vol 139 (10) ◽  
Author(s):  
Daniele Massini ◽  
Emanuele Burberi ◽  
Carlo Carcasci ◽  
Lorenzo Cocchi ◽  
Bruno Facchini ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Flow Field ◽  
Transfer Coefficient ◽  
Cooling System ◽  
Leading Edge ◽  
Impinging Jets ◽  
Internal Cooling ◽  
Coefficient Distribution ◽  
Transfer Pattern

A detailed aerothermal characterization of an advanced leading edge (LE) cooling system has been performed by means of experimental measurements. Heat transfer coefficient distribution has been evaluated exploiting a steady-state technique using thermochromic liquid crystals (TLCs), while flow field has been investigated by means of particle image velocimetry (PIV). The geometry key features are the multiple impinging jets and the four rows of coolant extraction holes, and their mass flow rate distribution is representative of real engine working conditions. Tests have been performed in both static and rotating conditions, replicating a typical range of jet Reynolds number (Rej), from 10,000 to 40,000, and rotation number (Roj) up to 0.05. Different crossflow conditions (CR) have been used to simulate the three main blade regions (i.e., tip, mid, and hub). The aerothermal field turned out to be rather complex, but a good agreement between heat transfer coefficient and flow field measurement has been found. In particular, jet bending strongly depends on crossflow intensity, while rotation has a weak effect on both jet velocity core and area-averaged Nusselt number. Rotational effects increase for the lower crossflow tests. Heat transfer pattern shape has been found to be substantially Reynolds independent.

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