Ranjan Saha

Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes (NGVs) to investigate the aerodynamic performance and the interaction between hub film cooling a ...

The influence of a special fillet between the endwall and airfoil at the leading edge on the performance of the turbine nozzle diaphragm

It is shown from the results of experimental investigations carried out on a nozzle diaphragm’s sector that an enlarged fillet at the vane leading edge does not have an essential effect on the flow and energy losses in the nozzle diaphragm.

Suction and Pressure Side Film Cooling Influence on Vane Aero Performance in a Transonic Annular Cascade

Efficiency improvement in turbomachines is an important aspect in reducing the use of fossil-based fuel and thereby reducing carbon dioxide emissions in order to achieve a sustainable future. Gas turbines are mainly fossil-based turbomachines powering aviation and land-based power plants. In line with the present situation and the vision for the future, gas turbine engines will retain their central importance in coming decades. Though the world has made significant advancements in gas turbine technology development over past few decades, there are yet many design features remaining unexplored or worth further improvement. These features might have a great potential to increase efficiency. The high pressure turbine (HPT) stage is one of the most important elements of the engine where the increased efficiency has a significant influence on the overall efficiency as downstream losses are substantially affected by the prehistory. The overall objective of the thesis is to contribute to the development of gas turbine efficiency improvements in relation to the HPT stage. Hence, this study has been incorporated into a research project that investigates leading edge contouring near endwall by fillet and external cooling on a nozzle guide vane with a common goal to contribute to the development of the HPT stage. In the search for HPT stage efficiency gains, leading edge contouring near the endwall is one of the methods found in the published literature that showed a potential to increase the efficiency by decreasing the amount of secondary losses. However, more attention is necessary regarding the realistic use of the leading edge fillet. On the other hand, external cooling has a significant influence on the HPT stage efficiency and more attention is needed regarding the aerodynamic implication of the external cooling. Therefore, the aerodynamic influence of a leading edge fillet and external cooling, here film cooling at profile and endwall as well as TE cooling, on losses and flow field have been investigated in the present work. The keystone of this research project has been an experimental investigation of a modern nozzle guide vane using a transonic annular sector cascade. Detailed investigations of the annular sector cascade have been presented using a geometric replica of a three dimensional gas turbine nozzle guide vane. Results from this investigation have led to a number of new important findings and also confirmed some conclusions established in previous investigations to enhance the understanding of complex turbine flows and associated losses. The experimental investigations of the leading edge contouring by fillet indicate a unique outcome which is that the leading edge fillet has no significant effect on the flow and secondary losses of the investigated nozzle guide vane. The reason why the leading edge fillet does not affect the losses is due to the use of a three-dimensional vane with an existing typical fillet over the full hub and tip profile. Findings also reveal that the complex secondary flow depends heavily on the incoming boundary layer. The investigation of the external cooling indicates that a coolant discharge leads to an increase of profile losses compared to the uncooled case. Discharges on the profile suction side and through the trailing edge slot are most prone to the increase in profile losses. Results also reveal that individual film cooling rows have a weak mutual effect. A superposition principle of these influences is followed in the midspan region. An important finding is that the discharge through the trailing edge leads to an increase in the exit flow angle in line with an increase of losses and a mixture mass flow. Results also indicate that secondary losses can be reduced by the influence of the coolant discharge. In general, the exit flow angle increases considerably in the secondary flow zone compared to the midspan zone in all cases. Regarding the cooling influence, the distinct change in exit flow angle in the area of secondary flows is not noticeable at any cooling configuration compared to the uncooled case. This interesting zone requires an additional, accurate study. The investigation of a cooled vane, using a tracer gas carbon dioxide (CO2), reveals that the upstream platform film coolant is concentrated along the suction surfaces and does not reach the pressure side of the hub surface, leaving it less protected from the hot gas. This indicates a strong interaction of the secondary flow and cooling showing that the influence of the secondary flow cannot be easily influenced. The overall outcome enhances the understanding of complex turbine flows, loss behaviour of cooled blade, secondary flow and interaction of cooling and secondary flow and provides recommendations to the turbine designers regarding the leading edge contouring and external cooling. Additionally, this study has provided to a number of new significant results and a vast amount of data, especially on profile and secondary losses and exit flow angles, which are believed to be helpful for the gas turbine community and for the validation of analytical and numerical calculations.

Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance

An experimental investigation on a cooled nozzle guide vane has been conducted in an annular sector to quantify aerodynamic influences of shower head and trailing edge cooling. The investigated van ...

Aerodynamic investigation of turbine cooled vane block

The vane block (VB) has been investigated and it gives several important results related to test methods and calculation procedures. The vane block is characterized by a developed film and convective cooling system. Blowing tests demonstrate that there is a weak correlation between cooling type and energy loss. Superposition of these effects is true for the central part over VB height without secondary flows. Coolant discharge increases profile loss and it rises if coolant flow rate is increased. Discharge onto profile convex side through the trailing edge slot influences the most considerably. The discharge through perforation decreases the vane flow capacity and insufficiently influences onto the flow outlet angle, but the trailing edge discharge increases this angle according to loss and mixture flow rate growth. The secondary flows reduce the effect of coolant discharge, which insufficiently changes losses distribution at turbine blades tips and even decreases the secondary losses. The flow outlet angle rises significantly and we are able to calculate it only if we correct the ordinary flow model. In the area of secondary flows, the outlet angle varies insufficiently under any type of cooling. This area should be investigated additionally.

Influence of Prehistory and Leading Edge Contouring on Aero Performance of a Three-Dimensional Nozzle Guide Vane

Experiments are conducted to investigate the effect of the prehistory in the aerodynamic performance of a three-dimensional nozzle guide vane with a hub leading edge contouring. The performance is ...

Experimental Studies of Leading Edge Contouring Influence on Secondary Losses in Transonic Turbines

Efficiency improvement in turbomachines is an important aspect in reducing the use of fossil-based fuel and thereby reducing carbon dioxide emissions in order to achieve a sustainable future. Gas turbines are mainly fossil-based turbomachines powering aviation and land-based power plants. In line with the present situation and the vision for the future, gas turbine engines will retain their central importance in coming decades. Though the world has made significant advancements in gas turbine technology development over past few decades, there are yet many design features remaining unexplored or worth further improvement. These features might have a great potential to increase efficiency. The high pressure turbine (HPT) stage is one of the most important elements of the engine where the increased efficiency has a significant influence on the overall efficiency as downstream losses are substantially affected by the prehistory. The overall objective of the thesis is to contribute to the development of gas turbine efficiency improvements in relation to the HPT stage. Hence, this study has been incorporated into a research project that investigates leading edge contouring near endwall by fillet and external cooling on a nozzle guide vane with a common goal to contribute to the development of the HPT stage. In the search for HPT stage efficiency gains, leading edge contouring near the endwall is one of the methods found in the published literature that showed a potential to increase the efficiency by decreasing the amount of secondary losses. However, more attention is necessary regarding the realistic use of the leading edge fillet. On the other hand, external cooling has a significant influence on the HPT stage efficiency and more attention is needed regarding the aerodynamic implication of the external cooling. Therefore, the aerodynamic influence of a leading edge fillet and external cooling, here film cooling at profile and endwall as well as TE cooling, on losses and flow field have been investigated in the present work. The keystone of this research project has been an experimental investigation of a modern nozzle guide vane using a transonic annular sector cascade. Detailed investigations of the annular sector cascade have been presented using a geometric replica of a three dimensional gas turbine nozzle guide vane. Results from this investigation have led to a number of new important findings and also confirmed some conclusions established in previous investigations to enhance the understanding of complex turbine flows and associated losses. The experimental investigations of the leading edge contouring by fillet indicate a unique outcome which is that the leading edge fillet has no significant effect on the flow and secondary losses of the investigated nozzle guide vane. The reason why the leading edge fillet does not affect the losses is due to the use of a three-dimensional vane with an existing typical fillet over the full hub and tip profile. Findings also reveal that the complex secondary flow depends heavily on the incoming boundary layer. The investigation of the external cooling indicates that a coolant discharge leads to an increase of profile losses compared to the uncooled case. Discharges on the profile suction side and through the trailing edge slot are most prone to the increase in profile losses. Results also reveal that individual film cooling rows have a weak mutual effect. A superposition principle of these influences is followed in the midspan region. An important finding is that the discharge through the trailing edge leads to an increase in the exit flow angle in line with an increase of losses and a mixture mass flow. Results also indicate that secondary losses can be reduced by the influence of the coolant discharge. In general, the exit flow angle increases considerably in the secondary flow zone compared to the midspan zone in all cases. Regarding the cooling influence, the distinct change in exit flow angle in the area of secondary flows is not noticeable at any cooling configuration compared to the uncooled case. This interesting zone requires an additional, accurate study. The investigation of a cooled vane, using a tracer gas carbon dioxide (CO2), reveals that the upstream platform film coolant is concentrated along the suction surfaces and does not reach the pressure side of the hub surface, leaving it less protected from the hot gas. This indicates a strong interaction of the secondary flow and cooling showing that the influence of the secondary flow cannot be easily influenced. The overall outcome enhances the understanding of complex turbine flows, loss behaviour of cooled blade, secondary flow and interaction of cooling and secondary flow and provides recommendations to the turbine designers regarding the leading edge contouring and external cooling. Additionally, this study has provided to a number of new significant results and a vast amount of data, especially on profile and secondary losses and exit flow angles, which are believed to be helpful for the gas turbine community and for the validation of analytical and numerical calculations.

Aerodynamic Implications of Reduced Vane Count

Given the shortage of fossil fuels and the growing greenhouse effect, one strive in modern gas turbines is to make maximum usage of the burnt fuel. By reducing the number of vanes or blades and thereby increasing the loading per vane (or blade) it is possible to spend less cooling air, which will have a positive impact on the combined cycle efficiency. It also reduces the number of components and usage of metal and thereby also the cost of the engine. These savings should be achieved without any efficiency deficit in aerodynamic efficiency. Based on the fact, aerodynamic investigations were performed to see the aerodynamic implications of reduced vane number in a transonic annular sector cascade. The number of new nozzle guide vane was reduced with 24% compared to a previous design with higher vane count. The investigated vanes were two typical high pressure gas turbine vanes. Results regarding the loading indicated an expected increase with the reduced vane case. The minimum static pressure at the suction side is lower and at an earlier location for the reduced vane case and therefore, an extension of the trailing edge deceleration zone is observed for the reduced vane case. Results regarding losses indicate that even though the losses produced per vane significantly increases for the reduced vane case, a comparison of mass averaged losses between the reduced vane case and previous vane case show similar spanwise loss distributions. Assessing results leads to a conclusion that the reduction of the number of vanes in the first stage seems to be a useful method to save cooling flow as well as material costs without any significant deficit in overall efficiency.Copyright

Influence of pre-history and leading edge contouring on aero-performance of a 3D nozzle guide vane

Experiments are conducted to investigate the effect of the pre-history in the aerodynamic performance of a threedimensional nozzle guide vane with a hub leading edge contouring. The performance is ...