| 1. |
- Farhanieh, Arman, et al.
(författare)
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Unsteady Effects in the Heat Load Predictions for a Two-Stage Compressor Turbine
- 2016
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Ingår i: PROCEEDINGS OF THE ASME TURBO EXPO: TURBINE TECHNICAL CONFERENCE AND EXPOSITION, 2016, VOL 5A. - : AMER SOC MECHANICAL ENGINEERS. - 9780791849781
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Konferensbidrag (refereegranskat)abstract
- Heat load analysis play an important role in the estimation of hot gas components lifetime. To achieve a high level of accuracy in heat load analysis, predicting the temperature distribution on the vane and blades is one area where further development is needed. Due to strong flow unsteadiness and mixing effects from blade row interactions and cooling injections, accurate heat load predictions have become an engineering challenge. This study uses both steady and time-accurate computational fluid dynamics (CFD) simulations to investigate the unsteady and mixing effects in a two-stage compressor turbine. The commercial code ANSYS CFX-15 is utilized to evaluate the performance of the steady state, mixing plane (MP) method, versus time-accurate, profile transformation (PT) and time transformation (TT) methods. The presence or absence of the rotor-stator cavities from which purge or cooling air is entering the main flowpath can also play an important role in the unsteadiness and mixing properties. Therefore the unsteady effects have been examined for two cases; a simplified model without any cavity and a detailed geometry with all the cavities included. In the simplified case, the cooling has been implemented as local patches. The results are then compared with gas temperature measurements from the real engine tests using thermo-crystals. The measurements include temperature profiles in front of the leading edge of each stator and rotor for both stages. The findings suggest that including cooling cavities may not improve the results in steady state simulations, however their presence in transient simulations can lead to mixing prediction improvements. Moreover, the results indicate that the transient simulations will improve the mixing predictions mainly in the second stage of the turbine. The results also indicate that in transient simulations, number of passages and pitch ratio between the stators of consecutive stages directly affect the results regardless of which transient method is used.
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| 2. |
- Saha, Ranjan, 1984-, et al.
(författare)
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Aerodynamic Implication of Endwall and Profile Film Cooling in a Transonic Annular Cascade
- 2013
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Ingår i: 21st ISABE Conference. - Busan, Korea.
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Konferensbidrag (refereegranskat)abstract
- An experimental study is performed to observe the aerodynamic implications of endwall and profile film cooling on flow structures and aerodynamic losses. The investigated vane is a geometrically similar transonic nozzle guide vane with engine-representative cooling geometry. Furthermore, a new formulation of the cooling aerodynamic loss equation is presented and compared with the conventional methods. Results from a 5-hole pneumatic probe show that the film coolant significantly alters the secondary flow structure. The effect of different assumptions for the loss calculation is shown to significantly change the measured loss.
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| 3. |
- Saha, Ranjan, et al.
(författare)
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Aerodynamic implications of reduced vane count
- 2015
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Ingår i: Proceedings of the ASME Turbo Expo. - : ASME Press. - 9780791856635
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Konferensbidrag (refereegranskat)abstract
- 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.
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| 4. |
- Saha, Ranjan, 1984-, et al.
(författare)
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Experimental studies of leading edge contouring influence on secondary losses in transonic turbines
- 2012
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Ingår i: ASME Turbo Expo 2012. - : ASME Press. - 9780791844748 ; , s. 1109-1119
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Konferensbidrag (refereegranskat)abstract
- An experimental study of the hub leading edge contouring using fillets is performed in an annular sector cascade to observe the influence of secondary flows and aerodynamic losses. The investigated vane is a three dimensional gas turbine guide vane (geometrically similar) with a mid-span aspect ratio of 0.46. The measurements are carried out on the leading edge fillet and baseline cases using pneumatic probes. Significant precautions have been taken to increase the accuracy of the measurements. The investigations are performed for a wide range of operating exit Mach numbers from 0.5 to 0.9 at a design inlet flow angle of 90°. Data presented include the loading, fields of total pressures, exit flow angles, radial flow angles, as well as profile and secondary losses. The vane has a small profile loss of approximately 2.5 % and secondary loss of about 1.1%. Contour plots of vorticity distributions and velocity vectors indicate there is a small influence of the vortex-structure in endwall regions when the leading edge fillet is used. Compared to the baseline case the loss for the filleted case is lower up to 13 % of span and higher from 13% to 20 % of the span for a reference condition with Mach no. of 0.9. For the filleted case, there is a small increase of turning up to 15 % of the span and then a small decrease up to 35 % of the span. Hence, there are no significant influences on the losses and turning for the filleted case. Results lead to the conclusion that one cannot expect a noticeable effect of leading edge contouring on the aerodynamic efficiency for the investigated 1st stage vane of a modern gas turbine.
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| 5. |
- Saha, Ranjan, 1984-, et al.
(författare)
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Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance
- 2014
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Ingår i: ASME Turbo Expo 2014. - : ASME Press. - 9780791845622
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Konferensbidrag (refereegranskat)abstract
- 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 vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass-flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both trailing edge cooling and shower head film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the trailing edge cooling has higher aerodynamic loss compared to the shower head cooling. Secondary losses decrease with inserting shower head film cooling compared to the uncooled case. The trailing edge cooling appears to have less impact on the secondary loss compared to the shower head cooling. Area-averaged exit flow angles around midspan increase for the trailing edge cooling.
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| 6. |
- Saha, Ranjan, 1984-, et al.
(författare)
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Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance
- 2014
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Ingår i: Journal of turbomachinery. - : ASME Press. - 0889-504X .- 1528-8900. ; 136:11, s. 111001-
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Tidskriftsartikel (refereegranskat)abstract
- An experimental investigation on a cooled nozzle guide vane (NGV) has been conducted in an annular sector to quantify aerodynamic influences of shower head (SH) and trailing edge (TE) cooling. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass-flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both TE cooling and SH film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the TE cooling has higher aerodynamic loss compared to the SH cooling. Secondary losses decrease with inserting SH film cooling compared to the uncooled case. The TE cooling appears to have less impact on the secondary loss compared to the SH cooling. Area-averaged exit flow angles around midspan increase for the TE cooling.
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| 7. |
- Saha, Ranjan, 1984-, et al.
(författare)
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Suction and Pressure Side Film Cooling Influence on Vane Aero Performance in a Transonic Annular Cascade
- 2013
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Ingår i: Proceedings of the ASME Turbo Expo. - 9780791855225
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Konferensbidrag (refereegranskat)abstract
- An experimental study on a film cooled nozzle guide vane has been conducted in a transonic annular sector to observe the influence of suction and pressure side film cooling on aerodynamic performance. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at an exit reference Mach number of 0.89. The aerodynamic results using a five hole miniature probe are quantified and compared with the baseline case which is uncooled. Results lead to a conclusion that the aerodynamic loss is influenced substantially with the change of the cooling flow rate regardless the positions of the cooling rows. The aerodynamic loss is very sensitive to the blowing ratio and a value of blowing ratio higher than one leads to a considerable higher loss penalty. The suction side film cooling has larger influence on the aerodynamic loss compared to the pressure side film cooling. Pitch-averaged exit flow angles around midspan remain unaffected at moderate blowing ratio. The secondary loss decreases (greater decrease in the tip region compared to the hub region) with inserting cooling air for all cases compared to the uncooled case.
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