Radiative Effectiveness on the Aero- and Thermodynamics in a Highly Thermally Loaded Film Cooling System

2011 ◽  
Author(s):  
Wenping Wang ◽  
Peng Sun ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Heat Transfer ◽  
High Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Cooling System ◽  
Numerical Study ◽  
Discrete Ordinates ◽  
Spanwise Direction ◽  
Inlet Temperature ◽  
Gas Channel

With the increasing of the gas turbine inlet temperature, the radiative heat transfer plays a more important role in the total heat transfer. In this paper, a high temperature test rig has been built to research the radiative effect in high temperature film cooling. The test section is made up of a high temperature hot gas channel and a middle temperature coolant air channel which are separated by a flat plate with a row of film cooling holes. The goal is to analyze the effects of radiation and its interaction between conduction and convection in the internal and film cooling which consider the heat transfer in both gas and solid. Meanwhile, the numerical study on the test cases are also carried out by combining conjugate heat transfer with radiative models. The fluid and solid regions were solved simultaneously. The Discrete Ordinates (DO) model and the Weighted Sum of Gray Gases Model (WSGGM) has been used to solve the radiative transfer equation for the radiation modeling. The results show that the temperature of the plate increase greatly when the radiation is taken into account and the temperature gradient through the plate becomes much larger. The temperature distribution has been changed and become smoother in spanwise direction. The results also indicate that the internal emissivity of the inlet has an influence mainly on the whole temperature of the plate, which suggests that the control of inlet emissivity is a good way for prevent over-high temperature on the first stage gas turbine vane.

Numerical Study on Flow and Heat Transfer of a Cooled First Stage Vane of a Gas Turbine

2016 ◽  
Author(s):  
Lv Ye ◽  
Zhao Liu ◽  
Xiangyu Wang ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Numerical Study ◽  
Coolant Flow ◽  
The Other ◽  
Edge Region ◽  
Flow Rates ◽  
Flow And Heat Transfer ◽  
Film Cooling Holes

This paper presents a numerical simulation of composite cooling on a first stage vane of a gas turbine, in which gas by fixed composition mixture is adopted. To investigate the flow and heat transfer characteristics, two internal chambers which contain multiple arrays of impingement holes are arranged in the vane, several arrays of pin-fins are arranged in the trailing edge region, and a few arrays of film cooling holes are arranged on the vane surfaces to form the cooling film. The coolant enters through the shroud inlet, and then divided into two parts. One part is transferred into the chamber in the leading edge region, and then after impinging on the target surfaces, it proceeds further to go through the film cooling holes distributed on the vane surface, while the other part enters into the second chamber immediately and then exits to the mainstream in two ways to effectively cool the other sections of the vane. In this study, five different coolant flow rates and six different inlet pressure ratios were investigated. All the cases were performed with the same domain grids and same boundary conditions. It can be concluded that for the internal surfaces, the heat transfer coefficient changes gradually with the coolant flow rate and the inlet total pressure ratio, while for the external surfaces, the average cooling effectiveness increases with the increase of coolant mass flow rates while decreases with the increase of the inlet stagnation pressure ratios within the study range.

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

2005 ◽  
Author(s):  
T. Povey ◽  
K. S. Chana ◽  
T. V. Jones ◽  
J. Hurrion
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Rate ◽  
Transfer Rate ◽  
Cooling System ◽  
Numerical Study ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Wall Heat Transfer ◽  
Hot Streak ◽  
End Wall

Pronounced non-uniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of end-wall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the affects of these temperature non-uniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimise the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and end-wall heat transfer rate for an HP nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two non-uniform temperature profiles. The temperature profiles were non-dimensionally similar to profiles measured in an engine. A difference of one half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and end-wall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualisation study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short duration engine-size turbine facility. Mach number, Reynolds number and gas-to-wall temperature ratios were correctly modelled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

2005 ◽  
Vol 129 (1) ◽  
pp. 32-43 ◽  
Author(s):  
T. Povey ◽  
K. S. Chana ◽  
T. V. Jones ◽  
J. Hurrion
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Rate ◽  
Transfer Rate ◽  
Cooling System ◽  
Numerical Study ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Gas Temperature ◽  
Hot Streak

Pronounced nonuniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of endwall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the effects of these temperature nonuniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimize the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and endwall heat transfer rate for a high-pressure (HP) nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two nonuniform temperature profiles. The temperature profiles were nondimensionally similar to profiles measured in an engine. A difference of one-half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and endwall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualization study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short-duration engine-sized turbine facility. Mach number, Reynolds number, and gas-to-wall temperature ratios were correctly modeled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

scholarly journals Cooling and Sealing Air System in Industrial Gas Turbine Engines

1996 ◽  
Author(s):  
A. W. Reichert ◽  
M. Janssen
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
State Of The Art ◽  
Cooling System ◽  
Inlet Temperature ◽  
Turbine Inlet ◽  
Air System ◽  
Calculation System ◽  
Cooling Air

Siemens heavy duty Gas Turbines have been well known for their high power output combined with high efficiency and reliability for more than 3 decades. Offering state of the art technology at all times, the requirements concerning the cooling and sealing air system have increased with technological development over the years. In particular the increase of the turbine inlet temperature and reduced NOx requirements demand a highly efficient cooling and sealing air system. The new Vx4.3A family of Siemens gas turbines with ISO turbine inlet temperatures of 1190°C in the power range of 70 to 240 MW uses an effective film cooling technique for the turbine stages 1 and 2 to ensure the minimum cooling air requirement possible. In addition, the application of film cooling enables the cooling system to be simplified. For example, in the new gas turbine family no intercooler and no cooling air booster for the first turbine vane are needed. This paper deals with the internal air system of Siemens gas turbines which supplies cooling and sealing air. A general overview is given and some problems and their technical solutions are discussed. Furthermore a state of the art calculation system for the prediction of the thermodynamic states of the cooling and sealing air is introduced. The calculation system is based on the flow calculation package Flowmaster (Flowmaster International Ltd.), which has been modified for the requirements of the internal air system. The comparison of computational results with measurements give a good impression of the high accuracy of the calculation method used.

Numerical Study of Heat Transfer in Novel Wavy Trailing Edge Design for Gas Turbine Airfoils

2019 ◽  
Author(s):  
Sarwesh Parbat ◽  
Li Yang ◽  
Minking Chyu ◽  
Sin Chien Siw ◽  
Ching-Pang Lee
Keyword(s):  
Heat Transfer ◽  
Steady State ◽  
Gas Turbine ◽  
Numerical Study ◽  
Turbine Blades ◽  
Suction Side ◽  
Inlet Temperature ◽  
Protective Film ◽  
Pressure Side ◽  
Trailing Edge

Abstract The strive to achieve increasingly higher efficiencies in gas turbine power generation has led to a continued rise in the turbine inlet temperature. As a result, novel cooling approaches for gas turbine blades are necessary to maintain them within the material’s thermal mechanical performance envelope. Various internal and external cooling technologies are used in different parts of the blade airfoil to provide the desired levels of cooling. Among the different regions of the blade profile, the trailing edge (TE) presents additional cooling challenges due to the thin cross section and high thermal loads. In this study, a new wavy geometry for the TE has been proposed and analyzed using steady state numerical simulations. The wavy TE structure resembled a sinusoidal wave running along the span of the blade. The troughs on both pressure side and suction side contained the coolant exit slots. As a result, consecutive coolant exit slots provided an alternating discharge between the suction side and the pressure side of the blade. Steady state conjugate heat transfer simulations were carried out using CFX solver for four coolant to mainstream mass flow ratios of 0.45%, 1%, 1.5% and 3%. The temperature distribution and film cooling effectiveness in the TE region were compared to two conventional geometries, pressure side cutback and centerline ejection which are widely used in vanes and blades for both land-based and aviation gas turbine engines. Unstructured mesh was generated for both fluid and solid domains and interfaces were defined between the two domains. For turbulence closer, SST-kω model was used. The wall y+ was maintained < 1 by using inflation layers at all the solid fluid interfaces. The numerical results depicted that the alternating discharge from the wavy TE was able to form protective film coverage on both the pressure and suction side of the blade. As a result, significant reduction in the TE metal was observed which was up to 14% lower in volume averaged temperature in the TE region when compared to the two baseline conventional configurations.

High-Temperature Potential of Uncooled Radial Turbines

1978 ◽  
Vol 100 (2) ◽  
pp. 294-302 ◽  
Author(s):  
D. J. Arnold ◽  
O. E. Balje
Keyword(s):  
Heat Transfer ◽  
High Temperature ◽  
Gas Turbine ◽  
Thermal Efficiency ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
High Reaction ◽  
Temperature Potential ◽  
Axial Turbines ◽  
Radial Turbines

Radial turbines are used predominantly for turbo-charges where the geometry is frequently compromised to favor low fabrications costs. Theoretical as well as experimental investigations have shown that the efficiency potential of radial turbines is as high as the efficiency potential of high reaction axial turbines. Structural and heat transfer studies on radial turbines show that the highest stresses in “deep scalloped” radial rotors occur at locations where the metal temperature is considerably lower than at rotor inlet. Thus the maximum allowable gas inlet temperature for radial turbines is several hundred degrees higher than for high-reaction axial turbines. This difference tends to increase with increasing expansion ratios, at least up to expansion ratios of 10:1. Since the thermal efficiency of typical gas turbine cycles increases with increasing gas temperatures and increasing expansion ratios, it results that the application of uncooled radial turbines will yield cycle efficiencies which are not obtainable with uncooled axial turbines.

Gas Turbine Combustor Cooling by Augmented Backside Convection

1979 ◽  
Vol 101 (1) ◽  
pp. 109-115 ◽  
Author(s):  
D. M. Evans ◽  
M. L. Noble
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Cooling System ◽  
Exhaust Gas ◽  
Confined Space ◽  
Gas Turbine Combustor ◽  
Exhaust Gas Emission ◽  
Cooling Air ◽  
Cooling Methods

Traditionally, gas turbine combustor walls have been cooled by one or more of the various film cooling methods. The current motivation to control exhaust gas emission composition has led to the serious consideration of backside convection wall cooling, where the cooling air is introduced to the main gas stream not prior to the dilution zone. Due to the confined space and the severe nature of the wall cooling problem, it is essential to maximize the heat transfer/pumping power characteristic, which suggests an augmented convection technique. A particular heat transfer design of a combustor cooled by means of transverse rib turbulence promoters applied to the exterior wall of the annular spaces surrounding the primary and secondary zones is described. Analytical methods for designing such a cooling system are reviewed and a comparison between analytical and experimental results is presented.

Numerical Study of Aerodynamic Performance of Film Cooling With Backward Injection Holes

2013 ◽  
Author(s):  
Ganesh Subbuswamy ◽  
Xianchang Li ◽  
Kunal Gharat
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Numerical Study ◽  
Aerodynamic Performance ◽  
Main Disadvantage ◽  
Cylindrical Holes ◽  
Lift Off ◽  
Very High ◽  
Aerodynamic Losses

Film cooling has been successfully used in cooling gas turbine components that are exposed to very high temperature environments. One main disadvantage of using film cooling is the aerodynamic losses associated. To address to the needs of obtaining uniform cooling in the downstream regions, backward injection of coolant has proved to be effective. However, there is a need to understand the aerodynamic behaviors of jet and mainstream flows in order to design effective configurations with this scheme of injecting coolant. In this work, the underlying aerodynamic principles of backward injection are studied numerically. All simulations are conducted with Fluent, a commercial CFD software. Results show that the classical counter rotating vortex found in simple cylindrical holes are not seen in the case of backward injections. Backward injection results in reduced coolant requirements and elimination of complex hole designs to avoid jet lift-off.

scholarly journals Numerical Heat Transfer Analysis of an Innovative Gas Turbine Combustor: Coupled Study of Radiation and Cooling in the Upper Part of the Liner

2005 ◽  
Author(s):  
A. Andreini ◽  
A. Bacci ◽  
C. Carcasci ◽  
B. Facchini ◽  
A. Asti ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Temperature Distribution ◽  
Gas Turbine ◽  
Cooling System ◽  
Numerical Study ◽  
Radiative Properties ◽  
Heat Transfer Analysis ◽  
Transfer Model ◽  
Cold Side ◽  
Additional Tool

A numerical study of a single can combustor for the GE10 heavy-duty gas turbine, which is being developed at GE-Energy (Oil & Gas), is performed using the STAR-CD CFD package. The topic of the present study is the analysis of the cooling system of the combustor liner’s upper part, named “cap”. The study was developed in three steps, using two different computational models. As first model, the flow field and the temperature distribution inside the chamber were determined by meshing the inner part of the liner. As second model, the impingement cooling system of the cold side of the cap was meshed to evaluate heat transfer distribution. For the reactive calculations, a closure of the BML (Bray-Moss-Libby) approach based on Kolmogorov-Petrovskii-Piskunov theorem was used. The model was implemented in the STAR-CD code using its user coding features. Then the radiative thermal load on the liner walls was evaluated by means of the STAR-CD-native Discrete Transfer model. The selection of the radiative properties of the flame was performed using a correlation procedure involving the total emissivity of the gas, the mean beam length and the gas temperature. The estimated heat flux on the cap was finally used as boundary condition for the calculation of the cooling system, consisting of 68 staggered impingement jet lines on the cold side of the cap. The resulting temperature distribution shows a good agreement with the experimental values measured by thermocouples. The results confirm the validity of the implemented procedure, and point out the importance of a full CFD computation as an additional tool to support classic correlation design procedures.

Numerical Study of Mist Film Cooling in Combustor at Operating Conditions

2009 ◽  
Author(s):  
Srinivasa Rao Para ◽  
Xianchang Li ◽  
Ganesh Subbuswamy
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Numerical Study ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Cooling Effectiveness ◽  
Wall Functions ◽  
Film Cooling Effectiveness ◽  
Combustor Liner

To improve the gas turbine thermal performance, apart from using a high compression ratio, the turbine inlet temperature must be increased. Therefore, the gas temperature inside the combustion chamber needs to be maintained at a very high level. Hence, cooling of the combustor liner becomes critical. Among all the cooling techniques, film cooling has been successfully applied to cool the combustor liner. In film cooling, coolant air is introduced through discrete holes and forms a thin film between the hot gases and the inner surface of the liner, so that the inner wall can be protected from overheating. The film will be destroyed in the downstream flow because of mixing of hot and cold gases. The present work focuses on numerical study of film cooling under operating conditions, i.e., high temperature and pressure. The effect of coolant injection angles and blowing ratios on film cooling effectiveness is studied. A promising technology, cooling with mist injection, is studied under operating conditions. The effect of droplet size and mist concentration is also analyzed. The results of this study indicate that the film cooling effectiveness can increase ∼11% at gas turbine operating conditions with mist injection of 2% coolant air when droplets of 10μm and a blowing ratio of 1.0 are applied. The cooling performance can be further improved by higher mist concentration. The commercial CFD software, Fluent 6.3.26, is used in this study and the standard k-ε model with enhanced wall functions is adopted as the turbulence model.

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