scholarly journals Combustor Flame Radiation and Wall Temperatures for #2 Distillate and a Coal Derived Liquid Fuel

1982 ◽  
Author(s):  
R. J. Kuznar ◽  
E. W. Tobery ◽  
A. Cohn
Keyword(s):  
Heat Flux ◽  
Film Cooling ◽  
Liquid Fuel ◽  
Gas Temperature ◽  
Flame Radiation ◽  
Flame Heat Flux ◽  
Primary Zone ◽  
The Difference ◽  
Process Measurements ◽  
Cooling Air

Flame radiation measurements were made in the primary zone of a film cooled gas turbine combustor, while burning #2 distillate (13 percent H) and a coal derived liquid fuel (8.7 percent H), from the SRC II process. Measurements from three circumferentially located radiometers indicate an average increase in flame heat flux of 55 percent, during combustion of the SRC II blend. Individual radiometer readings measured an increase ranging from 52 to 60 percent. Analytical predictions of the average combustor metal temperature, for the SRC II blend, cannot be explained by the increase in flame heat flux alone. If the same temperature of hot gases which mix in with film cooling air is used for both fuels then the calculated temperatures for SRC II fuel would be higher than those measured. Therefore, a lower value for hot gas temperature is required for the SRC II fuel case. This difference is attributed to the difference inflame shape between the two fuels.

Leading Edge Shielding Concept in Gas Turbines With Can Combustors

2012 ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic ◽  
Vasudevan Kanjirakkad ◽  
Sumiu Uchida
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
High Heat ◽  
Gas Temperature ◽  
Cooling Strategies ◽  
Film Cooling Holes ◽  
Cooling Air ◽  
Heat Loads

The remarkable developments in gas turbine materials and cooling technologies have allowed a steady increase in combustor outlet temperature and hence in gas turbine efficiency over the last half century. However, the efficiency benefits of higher gas temperature, even at the current levels, are significantly offset by the increased losses associated with the required cooling. Additionally, the advancements in gas turbine cooling technology have introduced considerable complexities into turbine design and manufacture. Therefore, a reduction in coolant requirements for the current gas temperature levels is one possible way for gas turbine designers to achieve even higher efficiency levels. The leading edges of the first turbine vane row are exposed to high heat loads. The high coolant requirements and geometry constraints limit the possible arrangement of the multiple rows of film cooling holes in the so called showerhead region. In the past, investigators have tested many different showerhead configurations, varying the number of rows, inclination angle and shape of the cooling holes. However the current leading edge cooling strategies using showerheads have not been shown to allow further increase in turbine temperature without excessive use of coolant air. Therefore new cooling strategies for the first vane have to be explored. In gas turbines with multiple combustor chambers around the annulus, the transition duct walls can be used to shield, i.e. to protect the first vane leading edges from the high heat loads. In this way the stagnation region at the leading edge and the shower-head of film cooling holes can be completely removed, resulting in a significant reduction in the total amount of cooling air that is otherwise required. By eliminating the showerhead the shielding concept significantly simplifies the design and lowers the manufacturing costs. This paper numerically analyses the potential of the leading edge shielding concept for cooling air reduction. The vane shape was modified to allow for the implementation of the concept and non-restrictive relative movement between the combustor and the vane. It has been demonstrated that the coolant flow that was originally used for cooling the combustor wall trailing edge and a fraction of the coolant air used for the vane showerhead cooling can be used to effectively cool both the suction and the pressure surfaces of the vane.

Leading Edge Shielding Concept in Gas Turbines With Can Combustors

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic ◽  
Vasudevan Kanjirakkad ◽  
Sumiu Uchida
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
High Heat ◽  
Gas Temperature ◽  
Cooling Strategies ◽  
Film Cooling Holes ◽  
Cooling Air ◽  
Heat Loads

The remarkable developments in gas turbine materials and cooling technologies have allowed a steady increase in combustor outlet temperature and, hence, in gas turbine efficiency over the last half century. However, the efficiency benefits of higher gas temperature, even at the current levels, are significantly offset by the increased losses associated with the required cooling. Additionally, the advancements in gas turbine cooling technology have introduced considerable complexities into turbine design and manufacture. Therefore, a reduction in coolant requirements for the current gas temperature levels is one possible way for gas turbine designers to achieve even higher efficiency levels. The leading edges of the first turbine vane row are exposed to high heat loads. The high coolant requirements and geometry constraints limit the possible arrangement of the multiple rows of film cooling holes in the so-called showerhead region. In the past, investigators have tested many different showerhead configurations by varying the number of rows, inclination angle, and shape of the cooling holes. However, the current leading edge cooling strategies using showerheads have not been shown to allow a further increase in turbine temperature without the excessive use of coolant air. Therefore, new cooling strategies for the first vane have to be explored. In gas turbines with multiple combustor chambers around the annulus, the transition duct walls can be used to shield, i.e., to protect, the first vane leading edges from the high heat loads. In this way, the stagnation region at the leading edge and the showerhead of film cooling holes can be completely removed, resulting in a significant reduction in the total amount of cooling air that is otherwise required. By eliminating the showerhead the shielding concept significantly simplifies the design and lowers the manufacturing costs. This paper numerically analyzes the potential of the leading edge shielding concept for cooling air reduction. The vane shape was modified to allow for the implementation of the concept and nonrestrictive relative movement between the combustor and the vane. It has been demonstrated that the coolant flow that was originally used for cooling the combustor wall trailing edge and a fraction of the coolant air used for the vane showerhead cooling can be used to effectively cool both the suction and the pressure surfaces of the vane.

scholarly journals Full Coverage Discrete Hole Film Cooling: The Influence of Hole Size

1985 ◽  
Author(s):  
G. E. Andrews ◽  
A. A. Asere ◽  
M. L. Gupta ◽  
M. C. Mkpadi
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Flame Temperature ◽  
Large Temperature ◽  
Hole Size ◽  
Gas Temperature ◽  
Cooling Effectiveness ◽  
Large Hole ◽  
Full Coverage ◽  
Primary Zone

The influence of hole size and hence blowing rate on full coverage discrete hole wall cooling for gas turbine combustion chamber applications was investigated. Two temperature conditions were used firstly a 750K gas temperature and 300K coolant, and secondly a realistic combustor primary zone conditions of 2100K flame temperature and 700K coolant. It was shown that a large hole size resulted in a significant improvement in the overall cooling effectiveness due to a reduced film heat transfer coefficient. At high temperature the cooling effectiveness was reduced due to radiative heat transfer from the flame gases. At low coolant flow large temperature increases of the coolant occurred within the wall and approached the transpiration situation.

scholarly journals METHOD FOR DETERMINING ADIABATIC FILM EFFECTIVENESS IN PRESENCE OF THERMAL BOUNDARY LAYER

2021 ◽  
pp. 1-20
Author(s):  
James Parker ◽  
Thomas Povey
Keyword(s):  
Boundary Layer ◽  
Film Cooling ◽  
Thermal Boundary Layer ◽  
Complex Function ◽  
Mixing Layer ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Point Of Interest ◽  
The Difference ◽  
Processing Techniques

Abstract In this paper we present a new method for determining adiabatic film effectiveness in film-cooling experiments with non-uniform inlet temperature distributions, in particular the situation of an inlet thermal boundary layer. This might arise in a quasi-steady experiment due to loss of heat from the mainstream flow to the inlet contraction walls, for example. In this situation the thermal boundary layer would be time varying. Adiabatic film effectiveness is generally normalised by the difference between mainstream and coolant gas temperatures. Most importantly these temperatures are generally assumed to be spatially—and, possibly temporally—uniform at the system inlet. In experiments with non-uniform inlet temperature, the relevant hot-gas temperature for a particular point of interest on a surface is not easily determined, being a complex function of both the inlet temperature profile and the flow-field between the inlet and the point of interest. In this situation, adiabatic film effectiveness cannot be uniquely defined using conventional processing techniques. We solve this problem by introducing the concept of equivalent mainstream effectiveness, a non-dimensional temperature for the mainstream that can be used to represent the thermal boundary layer profile at the inlet plane, or the effective temperature of the mainstream gas—which we refer to as the equivalent mainstream temperature—entrained into the mixing layer affecting the wall temperature at a particular point of interest.

Numerical Investigation on the Differences Between Temperature Ratio and Density Ratio in Film-Cooled Endwall

2018 ◽  
Author(s):  
Ran Yao ◽  
Jianhua Wang ◽  
Ming Wang ◽  
Wei Song
Keyword(s):  
Numerical Investigation ◽  
Film Cooling ◽  
Gas Turbines ◽  
Density Ratio ◽  
Influence Factors ◽  
Ambient Air ◽  
High Density ◽  
Temperature Ratio ◽  
The Difference ◽  
Cooling Air

This paper presents a numerical investigation on the rationality and reliability of film cooling experiments, in which the realistic temperature ratio (TR) of mainstream to cooling air are substituted by the density ratio (DR) of ambient air to a foreign gas with high density. Today, advanced gas turbines operate at much higher temperatures than the allowable temperature of turbine component materials, which makes it very difficult to achieve real TR in the most of laboratory environments, because the real TRs are usually larger than 2.0. Therefore, the foreign gas (for example CO2) with high density was widely used to obtain a proper density ratio for simulating the mixing process of cooling air with high temperature mainstream thereby. However, the TR effect on film cooling is not completely replaced by DR, because the influence factors of film cooling performances are not only DR, but also the thermal properties of cooling gas and mainstream, such as specific heat capacity, viscosity and conductivity. In this work, a film-cooled endwall is used as specimen, and DRs are controlled by two ways, i.e. changing mainstream temperatures, and using a foreign gas, respectively. The numerical results of film cooling performances at the same DR obtained by the two ways are analyzed and compared. The analysis reveals the difference of film cooling performances between TR and DR on the film-cooled endwall, and the comparison indicated that the errors caused by the substitution of foreign gas are acceptable, only when BR, DR, Re∞ and T∞ are all small, but when BR, DR or Re∞ increases, the relative error cannot be neglected, and it may reach 30% in real running conditions.

Liquid Fuelled Low NOx Radial Swirlers With Central Pilot Combustion

2007 ◽  
Author(s):  
Gordon E. Andrews ◽  
M. N. Kim ◽  
Mike C. Mkpadi ◽  
Sheriff A. Akande
Keyword(s):  
Natural Gas ◽  
Liquid Fuel ◽  
Fuel Injection ◽  
Liquid Fuels ◽  
Lean Combustion ◽  
Auto Ignition ◽  
Primary Zone ◽  
Low Nox Combustion ◽  
The Difference ◽  
Injection Location

Radial swirlers have proved effective in achieving low NOx using natural gas and this work investigates their use with kerosene with and without a central NG pilot. Two kerosene fuel injection locations were compared: at the inlet to the vane passages on the centreline with co-flow injection and 20mm downstream of the 76mm diameter swirler exit through the wall of a 76mm diameter 40mm long discharge duct. Flash back and auto ignition problems cannot occur with the downstream wall fuel injection location. All configurations were also tested with natural gas so that the difference in emissions due to the change from gas to liquid fuel could be established. The results show that ultra low NOx emissions can be achieved for kerosene with vane passage injection and that the use of a central pilot increases the NOx but improves the flame stability and power turndown. However, on liquid fuels the pilot to main flame propagation was not as good as when natural gas was used as the main fuel. Liquid fuel injection at the radial swirler wall outlet was effective but had slightly higher NOx for lean mixtures and worse HC and CO emissions. However, for richer primary zone mixtures the NOx was lower than for vane passage injection and this indicated that rich/lean combustion was occurring, without the uniform mixing and low NOx combustion that occurred with natural gas injection at this location.

Prediction of Low Temperature Hot Corrosion Rate in Film Cooled Coal Fired Gas Turbines

2003 ◽  
Author(s):  
Vaidyanathan Krishnan ◽  
Sanjeev Bharani ◽  
J. S. Kapat ◽  
Y. H. Sohn ◽  
V. H. Desai
Keyword(s):  
Corrosion Rate ◽  
Hot Corrosion ◽  
Film Cooling ◽  
Gas Turbines ◽  
Power Plants ◽  
Gas Temperature ◽  
Air Temperatures ◽  
Operational Life ◽  
Cooling Hole ◽  
Cooling Air

The concept of coal based gas turbine power plants has drawn considerable interest in recent years. Coal or syngas based power plants like IGCC have shown significant potential for meeting the ever-increasing power demands as well as stricter environmental regulations. The trouble free operational life of such power plants is limited by a major factor namely hot corrosion of the turbine components. Hitherto, the mechanism of hot corrosion has been investigated in a simpler context, which is not directly applicable to gas turbines in the presence of film cooling techniques. The present paper is an attempt to model hot corrosion in the presence of film cooling relevant to gas turbines, using a simple resistance model and the inherent analogy between heat and mass transfer. This paper considers film cooling air temperatures in the range of 450°C to 550°C, and a free stream gas temperature of 1425°C, with 0.5% sulfur in the fuel. For lower cooling air temperatures (less than 500°C), film cooling air suppresses corrosion, whereas for higher cooling air temperature corrosion rate is more in the presence of film cooling. With film cooling, there is a sharp peak in corrosion rate close to the cooling hole (within 10 slot widths). Due to the possibility that the base superalloy may be exposed in this region, designers should consider the high corrosion rate seriously. However, the present model is limited in its prediction because of its simplicity. Further improvement of the model is essential for optimization purposes.

Film Cooling Experimental Study on Sinusoidal Corrugated Liner for Afterburner

2020 ◽  
Vol 37 (2) ◽  
pp. 123-134 ◽  
Author(s):  
Shan Yong ◽  
Tan Xiao Ming ◽  
Zhang Jing Zhou ◽  
Wu Yan Hua
Keyword(s):  
Film Cooling ◽  
Discharge Coefficient ◽  
Jet Engines ◽  
Cooling Effectiveness ◽  
Blowing Ratio ◽  
Flow Loss ◽  
Opening Ratio ◽  
Ram Effect ◽  
The Difference ◽  
Cooling Air

AbstractThe need of improved cooling effectiveness for hot components in jet engines has led to new designs of afterburner liners. In the present paper, experiments were performed to reveal the cooling characteristics and flow loss of various sinusoidal corrugated liners for an advanced afterburner. It is found that there are alternate high temperature and low temperature zones corresponding to the wave crests and troughs of the corrugated liner, respectively. Compared to the flat liner, the corrugated liner increases cooling effectiveness by 10 % at the blowing ratio of 0.5, by 4.5 % at the blowing ratio of 3.2. However, the difference in discharge coefficients for these two kinds of liners is only 4.3 %. The increased opening ratio of film holes from 1.42 % to 3.72 % for corrugated liners is able to improve the cooling effectiveness by 9.8 %. However, the discharge coefficient is decreased by 34.1 %. The augment of amplitude of liners can enhance the ram effect of cooling air, which has an advantage to increase the local blowing ratio, and to form good cooling film at the windward of wave troughs. The flow loss is bigger while the cooling effectiveness is enhanced due to the change of amplitude of the liner.

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