Film Cooling Performance of Tripod Holes on the Endwall Upstream of a First Stage Nozzle Guide Vane

2015 ◽  
Author(s):  
Bharath Viswanath Ravi ◽  
Samruddhi Deshpande ◽  
Sridharan Ramesh ◽  
Prethive Dhilip Dhilipkumar ◽  
Srinath Ekkad
Keyword(s):  
Gas Turbine ◽  
Mass Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Turbine Engines ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Film Cooling Effectiveness ◽  
The Difference ◽  
Nozzle Guide

In view of the growing energy demand, there is an increasing need to augment the thermal efficiency of gas turbine engines. The thermal efficiency and power output of gas turbine engines increase with increasing overall pressure ratio which in turn leads to an increase in turbine inlet temperature. The maximum permissible turbine inlet temperature is limited by the material strength of the components of the gas turbine engines. In this regard, it is important to ensure that the endwalls of the first stage nozzle guide vane, which is one of the critical regions, are adequately cooled. The cooling of the endwall is of particular interest because the leading edge region along the endwall of the stator vane experiences high heat transfer rates resulting from formation of horseshoe vortices. In this paper, the performance of upstream purge slot has been compared against discrete film cooling holes. Three different cooling configurations — slot, cylindrical holes and tripod holes have been investigated by comparing the adiabatic film cooling effectiveness. Furthermore, the effect of coolant to mainstream mass flow ratio on the effectiveness of the different cooling schemes has also been studied. The steady-state experiments were conducted in a low speed, linear cascade wind tunnel. Spatially resolved temperature data was captured using infrared thermography technique to compute adiabatic film cooling effectiveness. Amongst the configurations studied, slot ejection offered the best cooling performance at all mass flow ratios. The performance of tripod ejection was comparable to slot ejection at mass flow ratios between 0.5 and 1.5, with the difference in laterally averaged effectiveness being ∼5%. However, at the highest mass flow ratio (MFR=2.5), the difference increased to ∼20%. Low effectiveness values were observed downstream of cylindrical ejection which could be attributed to jet lift-off.

Simulation of a Multi-Stage Cooling Scheme for Gas Turbine Engines: Part I

2007 ◽  
Author(s):  
M. Ghorab ◽  
I. Hassan ◽  
M. Beauchamp
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Lateral Spreading ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Multi Stage ◽  
Gap Height ◽  
Internal Gap

This paper presents heat transfer characteristics for a Multi-Stage Cooling Scheme (MSCS) design applicable to high temperature gas turbine engines in aerospace and electric power generation. The film cooling and impingement techniques are considered concurrently throughout this study. The proposed design involves passing cooling air from the inside of the turbine blade to the outside through three designed stages. The coolant air is passed through a circular hole into an internal gap creating an impingement of air inside the blade. It then exits through a sequence of two differently shaped holes onto the blade’s external surface. The film cooling effectiveness is enhanced by increasing the internal gap height and offset distance. This effect is significantly diminished however by changing the inclination angle from 90° to 30° at large gap height. The coolant momentum became more uniform by creating the internal gap consequently the coolant air is spread closer to the external blade surface. This reduces jet liftoff as the air exits its hole and also provides internal cooling for the blade. The hole exit positioned on the outer surface of the blade is designed to give a positive and a wide downstream lateral spreading. The MSCS demonstrates greater film cooling effectiveness performance than traditional schemes.

Endwall Film Cooling Using the Staggered Combustor-Turbine Gap Leakage Flow

2014 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Local Pressure ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blind Area ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the Nozzle Guide Vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the pressure side gill region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with upstream staggered slot, simulating the combustor-turbine leakage gap flow. The test cascades model was manufactured according to the GE-E3 nozzle guide vane scaled model, with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. The staggered slots were positioned upstream of the cascades to simulate the combustor-turbine gap leakage. The Pressure Sensitive Painting (PSP) technique was used to detect the film cooling effectiveness distribution on the endwall surface. Through the investigation, the following results could be achieved: 1) the film-cooling effectiveness on the endwall surface downstream the slot and along the pitchwise direction increased, with the highest parameter at Z/Pitch = 0.6; 2) a larger cooled region developed towards the suction side as the blowing ratio increased; 3) the advantage of the staggered slot was apparent on the endwall surface near the inlet area, while the coolant film was obviously weakened along the axial chord at a low blowing ratio. The influence of the staggered slots could only be detected in the downstream area of the endwall surface at the higher blowing ratio.

Contoured Endwall Flow and Heat Transfer Experiments With Combustor Coolant and Gap Leakage Flows for a Turbine Nozzle Guide Vane

2016 ◽  
Author(s):  
Reema Saxena ◽  
Mahmood H. Alqefl ◽  
Zhao Liu ◽  
Hee-Koo Moon ◽  
Luzeng Zhang ◽  
...  
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Mass Flow ◽  
Film Cooling ◽  
Thermal Loading ◽  
Guide Vane ◽  
Nozzle Guide Vane ◽  
Adiabatic Effectiveness ◽  
Interfacial Gap ◽  
Nozzle Guide

Flow in a high pressure gas turbine passage is complex, involving systems of secondary vortex flows and strong transverse pressure gradients. This complexity causes difficulty in providing film cooling coverage to the hub endwall region, which is subjected to high thermal loading due to combustor exit hot core gases. Therefore, an improved understanding of these flow features and their effects on endwall film cooling is needed to assist designers in developing efficient cooling schemes. The experimental study presented in this paper is performed on a linear, stationary, two-passage cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The sources of film cooling flows are the upstream combustor liner coolant and the leakage flow from the combustor-nozzle guide vane interfacial gap. Measurements are performed on an axisymmetrically-contoured endwall passage under conditions of various leakage mass flow rates to mainstream flow ratios (MFR= 0.5%, 1.0%, 1.5%). Flow migration and mixing are documented by measuring passage thermal fields and adiabatic effectiveness values over the endwall. It is found that, compared to our previous studies with a rotor inlet leakage slot geometry, the thin slot geometry of the nozzle leakage path gives a more uniform coolant spread over the endwall with significant coverage reaching the downstream and pressure-side regions of the passage. Interestingly, the coverage is seen to be only weakly dependent on the leakage mass low ratio and even reduce slightly with an increase in mass flow ratio above 1%, as indicated by lowered endwall adiabatic effectiveness values.

A New Method to Calculate the Coolant Requirements of a High-Temperature Gas Turbine Blade

2005 ◽  
Vol 127 (1) ◽  
pp. 191-199 ◽  
Author(s):  
Leonardo Torbidoni ◽  
J. H. Horlock
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Guide Vane ◽  
Coolant Flow ◽  
Computational Method ◽  
Inlet Temperature ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
High Temperature Gas

Earlier papers by the first author have described a computational method of estimating the cooling flow requirements of blade rows in a high-temperature gas turbine, for convective cooling alone and for convective plus film cooling. This method of analysis and computation, when applied to the whole blade chord was compared to a well-known semi-empirical method. In the current paper, a more sophisticated method is developed from the earlier work and is used to calculate the cooling flow required for a nozzle guide vane (the first blade row) of a high-temperature gas turbine, with given inlet gas temperature and coolant inlet temperature. Now the heat flux through an elementary cross-sectional area of the blade, at given spanwise (y) and chordwise (s) locations, is considered, with a guessed value of the elementary coolant flow [as a fraction dΨs of the external gas flow]. At the given s, integration along the blade length gives the blade metal temperatures at the outer and inner walls, Tbgy and Tbcly. If the value of Tbg at the blade tip y=H is assumed to be limited by material considerations to Tbg,max then the elementary coolant flow rate may be obtained by iteration. Summation along the chord then gives the total coolant flow, for the whole blade. Results using the method are then compared to a simpler calculation applied to the whole blade, which assumes chordwise constant temperatures and constant selected values of cooling efficiency and film-cooling effectiveness.

A New Method to Calculate the Coolant Requirements of a High Temperature Gas Turbine Blade

2004 ◽  
Author(s):  
Leonardo Torbidoni ◽  
J. H. Horlock
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Guide Vane ◽  
Coolant Flow ◽  
Computational Method ◽  
Inlet Temperature ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
High Temperature Gas

Earlier papers by the first author have described a computational method of estimating the cooling flow requirements of blade rows in a high temperature gas turbine, for convective cooling alone and for convective plus film cooling; this method of analysis and computation, when applied to the whole blade chord was compared with a well known semi-empirical method. In the current paper, a more sophisticated method is developed from the earlier work and is used to calculate the cooling flow required for a nozzle guide vane [the first blade row] of a high temperature gas turbine, with given inlet gas temperature and coolant inlet temperature. Now the heat flux through an elementary cross-sectional area of the blade, at given spanwise [y] and chordwise [s] locations, is considered, with a guessed value of the elementary coolant flow [as a fraction dΨ(s) of the external gas flow]. At the given s, integration along the blade length gives the blade metal temperatures at the outer and inner walls, Tbg[y] and Tbcl[y]. If the value of Tbg at the blade tip [y = H] is assumed to be limited by material considerations to Tbg,max then the elementary coolant flow rate may be obtained by iteration. Summation along the chord then gives the total coolant flow, for the whole blade. Results using the method are then compared with a simpler calculation applied to the whole blade, which assumes chordwise constant temperatures and constant selected values of cooling efficiency and film cooling effectiveness.

The Effects of Ambient Conditions on Gas Turbine Emissions—Generalized Correction Factors

1978 ◽  
Vol 100 (4) ◽  
pp. 640-646 ◽  
Author(s):  
P. Donovan ◽  
T. Cackette
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Turbine Engines ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Correction Factors ◽  
Oxides Of Nitrogen ◽  
Ambient Conditions ◽  
Nitrogen Emission ◽  
Wide Range

A set of factors which reduces the variability due to ambient conditions of the hydrocarbon, carbon monoxide, and oxides of nitrogen emission indices has been developed. These factors can be used to correct an emission index to reference day ambient conditions. The correction factors, which vary with engine rated pressure ratio for NOx and idle pressure ratio for HC and CO, can be applied to a wide range of current technology gas turbine engines. The factors are a function of only the combustor inlet temperature and ambient humidity.

scholarly journals Adiabatic Effectiveness and Heat Transfer Coefficient of Shaped Film Cooling Holes on a Scaled Guide Vane Pressure Side Model

2004 ◽  
Vol 10 (5) ◽  
pp. 345-354 ◽  
Author(s):  
Jan Dittmar ◽  
Achmed Schulz ◽  
Sigmar Wittig
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Test Model ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness

The demand of improved thermal efficiency and high power output of modern gas turbine engines leads to extremely high turbine inlet temperature and pressure ratios. Sophisticated cooling schemes including film cooling are widely used to protect the vanes and blades of the first stages from failure and to achieve high component lifetimes. In film cooling applications, injection from discrete holes is commonly used to generate a coolant film on the blade's surface.In the present experimental study, the film cooling performance in terms of the adiabatic film cooling effectiveness and the heat transfer coefficient of two different injection configurations are investigated. Measurements have been made using a single row of fanshaped holes and a double row of cylindrical holes in staggered arrangement. A scaled test model was designed in order to simulate a realistic distribution of Reynolds number and acceleration parameter along the pressure side surface of an actual turbine guide vane. An infrared thermography measurement system is used to determine highly resolved distribution of the models surface temperature. Anin-situcalibration procedure is applied using single embedded thermocouples inside the measuring plate in order to acquire accurate local temperature data.All holes are inclined 35° with respect to the model's surface and are oriented in a streamwise direction with no compound angle applied. During the measurements, the influence of blowing ratio and mainstream turbulence level on the adiabatic film cooling effectiveness and heat transfer coefficient is investigated for both of the injection configurations.

Thermo-Economic Optimization in the Design of Small-Scale and Residential Cogeneration Systems

2009 ◽  
Author(s):  
C. P. Lea˜o ◽  
S. F. C. F. Teixeira ◽  
A. M. Silva ◽  
M. L. Nunes ◽  
L. A. S. B. Martins
Keyword(s):  
Gas Turbine ◽  
Urban Areas ◽  
Pressure Ratio ◽  
Optimization Method ◽  
Turbine Engines ◽  
Small Scale ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Economic Optimization ◽  
Micro Turbine

In recent years, gas-turbine engines have undergone major improvements both in efficiency and cost reductions. Several inexpensive models are available in the range of 30 to 250 kWe, with electrical efficiencies already approaching 30%, due to the use of a basic air-compressor associated to an internal air pre-heater. Gas-turbine engines offer significant advantages over Diesel or IC engines, particularly when Natural Gas (NG) is used as fuel. With the current market trends toward Distributed Generation (DG) and the increased substitution of boilers by NG-fuelled cogeneration installations for CO2 emissions reduction, small-scale gas turbine units can be the ideal solution for energy systems located in urban areas. A numerical optimization method was applied to a small-scale unit delivering 100 kW of power and 0.86 kg/s of water, heated from 318 to 353K. In this academic study, the unit is based on a micro gas-turbine and includes an internal pre-heater, typical of these low pressure-ratio turbines, and an external heat recovery system. The problem was formulated as a non-linear optimisation model with the minimisation of costs subject to the physical and thermodynamic constraints. Despite difficulties in obtaining data for some of the components cost-equations, the preliminary results indicate that the optimal compressor pressure ratio is about half of the usual values found in large installations, but higher than those of the currently available micro-turbine models, while the turbine inlet temperature remains virtually unchanged.

Thermodynamic Performance of Complex Gas Turbine Cycles

2002 ◽  
Author(s):  
Sanjay ◽  
Onkar Singh ◽  
B. N. Prasad
Keyword(s):  
Gas Turbine ◽  
Specific Power ◽  
Turbine Engines ◽  
Inlet Temperature ◽  
Air Cooling ◽  
Gas Turbine Engines ◽  
Current State ◽  
Thermodynamic Performance ◽  
Internal Convection ◽  
Plant Efficiency

This paper deals with the thermodynamic performance of complex gas turbine cycles involving inter-cooling, re-heating and regeneration. The performance has been evaluated based on the mathematical modeling of various elements of gas turbine for the real situation. The fuel selected happens to be natural gas and the internal convection / film / transpiration air cooling of turbine bladings have been assumed. The analysis has been applied to the current state-of-the-art gas turbine technology and cycle parameters in four classes: Large industrial, Medium industrial, Aero-derivative and Small industrial. The results conform with the performance of actual gas turbine engines. It has been observed that the plant efficiency is higher at lower inter-cooling (surface), reheating and regeneration yields much higher efficiency and specific power as compared to simple cycle. There exists an optimum overall compression ratio and turbine inlet temperature in all types of complex configuration. The advanced turbine blade materials and coating withstand high blade temperature, yields higher efficiency as compared to lower blade temperature materials.

A Method for Correlating the Influence of External Crossflow on the Discharge Coefficients of Film Cooling Holes

2000 ◽  
Vol 123 (2) ◽  
pp. 258-265 ◽  
Author(s):  
D. A. Rowbury ◽  
M. L. G. Oldfield ◽  
G. D. Lock
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Discharge Coefficient ◽  
Guide Vane ◽  
Discharge Coefficients ◽  
Nozzle Guide Vane ◽  
Aero Engine ◽  
Nozzle Guide ◽  
Film Cooling Holes ◽  
Asme Paper

An empirical means of predicting the discharge coefficients of film cooling holes in an operating engine has been developed. The method quantifies the influence of the major dimensionless parameters, namely hole geometry, pressure ratio across the hole, coolant Reynolds number, and the freestream Mach number. The method utilizes discharge coefficient data measured on both a first-stage high-pressure nozzle guide vane from a modern aero-engine and a scale (1.4 times) replica of the vane. The vane has over 300 film cooling holes, arranged in 14 rows. Data was collected for both vanes in the absence of external flow. These noncrossflow experiments were conducted in a pressurized vessel in order to cover the wide range of pressure ratios and coolant Reynolds numbers found in the engine. Regrettably, the proprietary nature of the data collected on the engine vane prevents its publication, although its input to the derived correlation is discussed. Experiments were also conducted using the replica vanes in an annular blowdown cascade which models the external flow patterns found in the engine. The coolant system used a heavy foreign gas (SF6 /Ar mixture) at ambient temperatures which allowed the coolant-to-mainstream density ratio and blowing parameters to be matched to engine values. These experiments matched the mainstream Reynolds and Mach numbers and the coolant Mach number to engine values, but the coolant Reynolds number was not engine representative (Rowbury, D. A., Oldfield, M. L. G., and Lock, G. D., 1997, “Engine-Representative Discharge Coefficients Measured in an Annular Nozzle Guide Vane Cascade,” ASME Paper No. 97-GT-99, International Gas Turbine and Aero-Engine Congress & Exhibition, Orlando, Florida, June 1997; Rowbury, D. A., Oldfield, M. L. G., Lock, G. D., and Dancer, S. N., 1998, “Scaling of Film Cooling Discharge Coefficient Measurements to Engine Conditions,” ASME Paper No. 98-GT-79, International Gas Turbine and Aero-Engine Congress & Exhibition, Stockholm, Sweden, June 1998). A correlation for discharge coefficients in the absence of external crossflow has been derived from this data and other published data. An additive loss coefficient method is subsequently applied to the cascade data in order to assess the effect of the external crossflow. The correlation is used successfully to reconstruct the experimental data. It is further validated by successfully predicting data published by other researchers. The work presented is of considerable value to gas turbine design engineers as it provides an improved means of predicting the discharge coefficients of engine film cooling holes.

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