Experimental investigations on axial temperature profiles, liner wall temperature distribution and exit gas temperature distribution at different equivalence ratios of a tubular type gas turbine combustors.
The combustion is the component in the gas turbine where fuel is added to the airflow and subsequently burnt. A Combustor must satisfy a wide range of requirements whose relative importance may vary depending upon engine type and specific application. The basic performance requirements are pollutant emissions, pressure drop, wall cooling flow, re-light, efficiency, cost, durability and stability limits. The re-lighting requirements and to some extent pollution emission are related to primary zone equivalence ratio. The effect of primary zone equivalence ratio on the temperature distribution in the axial direction, the liner wall temperature distribution and the exit gas temperature distribution is studied for four different equivalence ratios of 0.5, 0.7, 0.9 and 1.1 for tubular type combustion chamber. As the weak limit for kerosene is 0.5, the equivalence ratio of 0.5 is the minimum requirement for the combustion of kerosene type fuel. The kerosene type fuel is similar to that of Jet A type fuel . The pressure distribution along the length of the combustion chamber was evaluated for all the four chambers. All the four chambers were experimentally investigated over an overall air/fuel ratio range of 22.74 to 152.4. The minimum air/fuel ratio of 22.74 is selected. The reason is that usually the air/fuel ratio range of gas turbines is in the range of 60 to 80 . It is worth to mention that the temperatures of near 1400[degrees]C achieved at the centerline of the combustion chamber and the liner wall temperatures in the range of 500[degrees]C for lower air/fuel ratio and 300[degrees]C for higher air/fuel ratio certainly ensures safe and reliable operation of this combustion chamber. The highest pressure drop witnessed along the length of combustion chamber is 6% of the delivery pressure. It is concluded that an equivalence ratio of 0.9 gives the consistent centerline temperature; lower liner wall temperature and uniform exhaust gas temperature.
Experimental Test Rig
The versatile experimental setup has been developed for measuring the pressure loss and temperature distribution across the length of chamber along the centre line, liner wall and exit temperature distribution. The photographic view of the experimental setup is shown in Figure 1.
The air is supplied by the blower at a pressure of 3 bar. The simplex atomizer is used for fuel injection. The four different chambers are designed at primary zone equivalence ratios of 0.5, 0.7, 0.9 and 1.1 keeping the overall equivalence ratio of the combustion chamber same for all the chambers. Venturimeter is used to measure the flow rate of air.
Table below shows the different flow rates of air in different zones of combustion chambers designed at different primary zone equivalence ratios.
Chamber Equivalenc [[??].sub.pz]kg/sec No. e Ratio, [[PHI].sub.pz] [[??].sub.PZF] [[??].sub.PZH] 1 0.5 0.116412 0.132 248 2 0.7 0.103477 0.074 137 3 0.9 0084663 0.534 81 4 1.1 0.071638 0.041 889 Chamber [[??].sub.IZH] kg/ [[??].sub.DZH] kg/sec [[??].sub.WCH] No. kg/sec 1 0.0135 0.29964 0.187275 25 0 2 0.0849 0.26218 0.227430 06 5 3 0.1240 0.22473 0.262185 41 0 4 0.1491 0.18727 0.299640 58 5
[FIGURE 1 OMITTED]
The measurement of temperature along the length of the combustion chamber is measured using K-Type Chromel - Alumel Thermocouples. The signal received from the thermocouple is displayed by the milli-volt meters. The centerline temperature is measured at the centerline of the combustion chamber, i.e. 45 mm from the inner wall of the combustion chamber. The inner diameter of the combustion chamber is 90 mm. The liner wall temperature probes are flushed the skin of the combustion chamber. The pressure measurement along the length of the combustion chamber is measured using 14 manometers. The pressure drop along the length of the combustion chamber in axial direction is measured in mm of diesel.
The accuracy of the system is a function that depends on several variables, such as thermocouple wire material, bead formation, calibration techniques, reference junction and radiation effects. The thermocouples are calibrated by fixed point calibration method at four points, i.e. boiling water, naphthalene balls, Sulphur and gold. The thermocouple is housed in a radiation shield casing with ceramic sleeves covered with inconel tubes. All together, the total error is limited to 5% for liner wall temperature measurements.
Pressure measurement is carried out using a simple U-tube manometer. It is simple, inexpensive and relatively free from errors and also it can be arranged to almost any degree of sensitivity. It is useful for measuring slowly changing pressures.
Results and Discussions
The experimental investigations were carried out on four different combustion chambers designed at different primary zone equivalence ratios of 0.5, 0.7, 0.9 and 1.1 for kerosene type fuel. The combustion chambers are designated as Chamber 1 for primary zone equivalence ratio 0.5, Chamber 2 for primary zone equivalence ratio 0.7, Chamber 3 for primary zone equivalence ratio 0.9, Chamber 4 for primary zone equivalence ratio 1.1. This four designed chambers are experimentally investigated for different air/fuel ratios ranging from 22.74 to 152.4 for on Axial Temperature Profiles, Liner Wall Temperature Distribution and Exit Gas Temperature Distribution. Figure 2 shows the locations of axial temperature probes, liner wall temperature probes and exit gas temperature probes. The thermocouples are spaced at a distance of 50 mm from each other for centerline and liner wall temperature measurements, while for exit gas temperature distribution the thermocouples are placed at a distance of 16 mm from each other.
[FIGURE 2 OMITTED]
Figures 3 to 11 shows the graphs at centerline, liner wall and exit gas distribution for all the four chambers at different air/fuel ratios at different locations as given in Figure 2. All the four chambers were investigated for fuel rich to fuel lean air/fuel ratios. The air/fuel ratio range is from 22.74 to 152.4. The designed air/fuel ratio is 122.02. In all, the combustion chambers were checked for nine different air/fuel ratios.
At low air/fuel ratios, the liner wall temperatures encountered are in the range of 600[degrees]C, while the center line temperatures are in the range of 1300-1400[degrees]C. The exit gas distribution for chamber 2, chamber 3 and chamber 4 is more or less uniform. Also, chamber 3 gives the lower and gradually varying liner wall temperatures and consistent exit gas temperature as well as consistent center line temperatures. The centerline temperatures are higher for higher equivalence ratios in the primary zone.
Similar, trend is observed as in Figure 3. Here again, the chamber 3 gives the lower liner wall temperatures and consistent centerline and exit gas temperatures.
[FIGURE 3 OMITTED]
Similar, trend is observed as in Figures 3 and 4. Here again, the chamber 3 gives the lower liner wall temperatures and consistent exit gas temperatures. But the centerline temperature decreases at location 2 and there after increases in Figure 5. But that trend is not witnessed in Figure 6.
From figures 3 to 6, it can be concluded that combustion chamber 3, designed at an equivalence ratio of 0.9 gives consistent centerline temperatures, lower liner wall temperatures and uniform exit gas temperature.
At higher air/fuel ratio of 59.73, the liner wall temperatures encountered by chamber 2 is lowest, but the exit gas temperature quality is poor. Also, the centerline distribution is not uniform; rather it is increasing at first, then decreasing, and again increases. This trend can be explained by the fact that may be dissociated products are combined in the intermediate zone.
[FIGURE 4 OMITTED]
[FIGURE 5 OMITTED]
[FIGURE 6 OMITTED]
Similar trend is observed for yet higher air/fuel ratio of 77.08. Here chambers 1 to 4 shows the similar trend for centerline temperatures as discussed in Figure 7 for chamber 2. For liner wall and exit gas temperatures, chamber 3 gives consistent temperatures which are more or less uniform. Although, lower liner wall temperatures are encountered for chamber 1 and chamber 2 compared to chamber 3, but the trend is of sinusoidal form.
[FIGURE 7 OMITTED]
Similar trend as discussed for Figure 8 is observes for higher air/fuel ratios. The liner wall temperature for chamber 3 in Figures 9, 10 and 11 are uniformly increasing and then remains constant. Also, the exit gas temperatures are uniform for chamber 3. At designed air/fuel ratio of 122.02, all the four chambers gives consistent trend for liner wall, centerline and exit gas temperature distribution.
[FIGURE 8 OMITTED]
[FIGURE 9 OMITTED]
[FIGURE 10 OMITTED]
[FIGURE 11 OMITTED]
From figures 3 to 11, it can be concluded that chamber 3 designed at primary zone equivalence ratio gives consistent centerline temperatures and lower liner wall temperatures. Also the exit temperature distribution is uniform; with the pattern factor not more that 25% for any air/fuel ratios. Thus, it can be stated that primary zone equivalence ratio of 0.9 is most suited for combustion chamber at almost all working conditions of air/fuel ratios.
Figure 12 shows the locations of centerline pressure probes. The probes are spaced at a distance of 50 mm from each other.
[FIGURE 12 OMITTED]
Figures 13 to 21 show the graphs of pressure distribution for all the four chambers at different air/fuel ratios. The pressure probes are at the centerline of the liner and the pressure distribution along the length of the combustion liner is investigated using simple U tube diesel manometers.
[FIGURE 13 OMITTED]
[FIGURE 14 OMITTED]
[FIGURE 15 OMITTED]
[FIGURE 16 OMITTED]
At lower air/fuel ratios, as illustrated in Figures 13 to 16, there is a slight pressure rise at locations 2 and 3 and there after the pressure decreases. This is obviously due to higher temperature levels at lower air/fuel ratios which offer more expansion of gases in a smaller volume of combustion chamber resulting in higher pressure levels.
With increase in the air/fuel ratio, the pressure levels do not increase at locations 2 and 3, but the pressure is gradually decreasing. This may be due to lower temperatures encountered at higher air/fuel ratios compared to lower air/fuel ratios. The pressure drop witnessed for all the four chambers at different air/fuel ratios is not more that 6% of the delivery pressure. These suggest the aerodynamic superiority for the designed combustion chambers.
[FIGURE 17 OMITTED]
[FIGURE 18 OMITTED]
[FIGURE 19 OMITTED]
[FIGURE 20 OMITTED]
[FIGURE 21 OMITTED]
The emission levels for the different chambers were analytically evaluated using thermodynamic equilibrium model. The emissions levels of OH, H2, H2O, CO, CO2, NO and N2 are evaluated for different equivalence ratios of the primary zone. Figure 22 shows the emission levels of the different combustion chambers evaluated by thermodynamic equilibrium model.
[FIGURE 22 OMITTED]
It can be concluded that chamber 3 designed at primary zone equivalence ratio gives consistent centerline temperatures and lower liner wall temperatures. Also the exit temperature distribution is uniform; with the pattern factor not more than 25% for any air/fuel ratios. Thus, it can be stated that primary zone equivalence ratio of 0.9 is most suited for combustion chamber at almost all working conditions of air/fuel ratios. Also the pressure drop across the combustion chamber for any chamber is not more than 6% of delivery pressure. This clearly suggests the aerodynamic superiority of the designed chamber. The NO emissions are lower for primary zone equivalence ratio of 0.9 compared to an equivalence ratio of 1.1. The emissions levels of other components are also lower at an equivalence ratio of 0.9 in the primary zone.
In nutshell, it is concluded that while designing the combustion chamber for gas turbine applications, the primary zone equivalence ratio should be kept 0.9 to get better exit gas temperature distribution and uniform exit gas temperature for fuel rich to fuel lean air/fuel ratios.
 Mellor A. M., Design of Modern Turbine Combustors, ACADEMIC PRESS INC, 1990.
 Arthur W. Lefebvre, Gas Turbine Combustion, Taylor and Francis, 2nd ed. 1999.
We thank the staff of Fluid and Hydraulic Machines laboratory of S. V. National Institute of Technology, Surat for their help with the experimental setup. Many thanks are due to Hitesh Solanki from the conceptualization of this topic to the experimental evaluation. I record my permanent gratitude for the faith and support of the people with whom I really worked and lived - my parents and my wife. They gave up much more than they expected, but they contributed much more than they can imagine.
Digvijay B. Kulshreshtha (1), S. A. Channiwala (2), Saurabh B. Dikshit (3) and Kamlesh V. Chaudhari (4)
(1) Lecturer, Mechanical Engineering Department, C. K. Pithawalla College of Engineering and Technology, Surat--395007, India E-mail: firstname.lastname@example.org
(2) Professor, Mechanical Engineering Department, S. V. National Institute of Technology, Surat--395007, India E-mail: email@example.com
(3) Lecturer, Mechanical Engineering Department, C. K. Pithawalla College of Engineering and Technology, Surat--395007, India. E-mail: firstname.lastname@example.org
(4) Ph. D. Student, Mechanical Engineering Department, S. V. National Institute of Technology, Surat--395007, India. E-mail: email@example.com
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|Author:||Kulshreshtha, Digvijay B.; Channiwala, S.A.; Dikshit, Saurabh B.; Chaudhari, Kamlesh V.|
|Publication:||International Journal of Applied Engineering Research|
|Date:||Mar 1, 2009|
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