| 1. |
- El-Gabry, Lamyaa, et al.
(författare)
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Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade
- 2012
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Ingår i: Proceedings of the ASME Turbo Expo. - : ASME Press. - 9780791844748
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Konferensbidrag (refereegranskat)abstract
- An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge can be used to visualize the mixing of the coolant flow with the mainstream.Flow field measurements are performed in the downstream plane with a 5-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side of the vane trailing edge where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge does not reach the pressure side endwall, potentially creating a local hotspot.
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| 2. |
- Ellbrant, Lars R, 1984, et al.
(författare)
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Design of compressor blades considering efficiency and stability using CFD based optimization
- 2012
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Ingår i: ASME Turbo-expo 2012. - 9780791844748 ; 8:PARTS A, B, AND C, s. 371-382
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Konferensbidrag (refereegranskat)abstract
- To design a highly loaded axial transonic compressor several objectives need to be considered simultaneously. From an aerodynamic perspective, one of the major requirements is high efficiency at a specific operating condition where the fuel consumption is of main interest. Furthermore, the compressor needs to have a sufficient stall-margin along the entire flight envelope to ensure a stable operating range. This work is focused on creating an efficient design method which produces a trade-off between high stall margin and high efficiency. The design method is based on an automatic multiobjective optimization process divided into two steps. In the first step, 2D blade profiles are optimized where both efficiency and stall margin are considered. Once the optimization is finished the selected profiles are stacked together to be further optimized in 3D. When going to the second step, i.e. a 3D optimization, one can focus on a smaller set of design variables thereby reducing the time to get what is considered the optimal solution. The results show that it is possible to rate designs with potential of having high stall margin and high efficiency both in the 2D and 3D optimization. The main contribution in this work is the design method, which offers an efficient way of designing robust blades where the designer can decide the best trade off between stall margin at part speed and efficiency at the design point. Copyright © 2012 by ASME.
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| 3. |
- Noor, Hina, et al.
(författare)
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Selection of one-dimensional design parameter 'reaction degree' for 1 st stage of a cooled gas turbine
- 2012
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Ingår i: Proc. ASME Turbo Expo. - 9780791844748 ; , s. 2345-2354
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Konferensbidrag (refereegranskat)abstract
- The recommendations available today in open literature for the choice of design parameter such as flow coefficient, stage loading and reaction degree incorporates mainly the influence of aerodynamics loss on efficiency. However, it is difficult to find the recommendation relating the influence of not only the aerodynamics loss but also cooling mass flow and cooling losses on varying most influential design parameters. In this paper, preliminary design and performance guidelines are presented for a cooled turbine stage using the 1D design tool LUAXT. The intention is to provide recommendations on the selection of design parameters, mainly reaction degree, which is found to be highly influenced by not only the aerodynamics loss but also the cooling mass flow and cooling loss such as in 1st stage of a High Pressure Turbines (HPT). The One-Dimensional (1D) design methods used to perform this task are verified and validated against experimental test data. A comparison of different loss models has been performed to provide most accurate outcomes for certain tested ranges. Based on the outcomes of this study, 'Craig & Cox' loss model has been considered to perform subsequent investigations for HPT design and performance estimation while formulating a parametric study. From this study, the design recommendations for the selection of performance parameter reaction degree are developed for cooled turbines. The results shows that for a HPT 1st stage, the recommended reaction degree range of 0.20 to 0.37 seems to provide the optimum stage design when chosen for stage loading in between 1.40 to 1.80 along with the stator exit flow angle in range of 74° to 78°.
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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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