Effects of Unsteady Wakes on Flow Through High Pressure Turbine Passages With and Without Tip Gaps

2014 ◽  
Author(s):  
Ralph J. Volino ◽  
Christopher D. Galvin ◽  
Cody J. Brownell
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Direction ◽  
Velocity Fields ◽  
Total Pressure Loss ◽  
High Pressure Turbine ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Wake Passing

Experiments were conducted in a linear high pressure turbine cascade with wakes generated by moving upstream rods. The cascade included an adjustable top endwall that could be raised and lowered above the airfoils to change the tip gap. Conditions were considered with no tip gap, and gaps of 1.5% and 3.8% of axial chord. For each of these, cases were documented both with and without upstream wakes. The pressure distributions on the airfoils were acquired at the midspan and near the tip for each case. The total pressure loss was measured in the endwall region. Velocity fields were acquired in two planes normal to the flow direction using particle image velocimetry (PIV). For the case with no tip gap, the passage vortex and other vortices were clearly visible in the velocity fields. For the cases with a tip gap, the tip leakage vortex was the dominant flow feature, and it became stronger as the gap size increased. The other vortices were still present, but were moved by the tip leakage vortex. For the cases with unsteady wakes, the PIV data were ensemble-averaged based on phase within the wake passing cycle, to show the motion and change in strength of the vortices in response to the wake passing. The regions of high total pressure loss can be explained in terms of the secondary velocity field.

Numerical Simulation of Deposition Phenomena on High-Pressure Turbine Vane Using UPACS

2019 ◽  
Author(s):  
Kenta Mizutori ◽  
Koji Fukudome ◽  
Makoto Yamamoto ◽  
Masaya Suzuki
Keyword(s):  
Numerical Simulation ◽  
High Pressure ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
High Pressure Turbine ◽  
Pressure Loss Coefficient ◽  
Turbine Vane ◽  
Total Pressure Loss Coefficient

Abstract We performed numerical simulation to understand deposition phenomena on high-pressure turbine vane. Several deposition models were compared and the OSU model showed good adaptation to any flow field and material, so it was implemented on UPACS. After the implementation, the simulations of deposition phenomenon in several cases of the flow field were conducted. From the results, particles adhere on the leading edge and the trailing edge side of the pressure surface. Also, the calculation of the total pressure loss coefficient was conducted after computing the flow field after deposition. The total pressure loss coefficient increased after deposition and it was revealed that the deposition deteriorates aerodynamic performance.

Characteristics of Tip-Leakage Flow at High Stagger-Angle Setting for Rotor Blade in an Axial Flow Fan

2012 ◽  
Author(s):  
Takahiro Nishioka ◽  
Masayoshi Joko
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Maximum Efficiency ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Stagger Angle

Rotor-tip flow fields at high stagger-angle setting were investigated to clarify the loss generation mechanism in a high specific-speed axial-flow fan. The tip clearance flow in the cases of large and small clearances, which are 2.0% and 1.0% of the rotor tip chord length respectively, are experimentally and numerically evaluated at the maximum efficiency point and the operating limit. At the maximum efficiency point, the tip leakage vortex reached to the rotor exit in both cases of large and small tip clearances. However, the leakage vortex in the case of large tip-clearance passed closer to the pressure side of the adjacent blade than that in the case of small one. Moreover, in the case of large tip clearance, the tip leakage vortex generated the large total pressure loss in the blade passage, and the interaction between the tip leakage vortex and the wake also generated the large total pressure loss at the rotor exit. Therefore, the maximum efficiency of the rotor and the fan was lower than that in the case of small tip clearance. At the operating limit, the tip-leakage vortex extended inside the blade passage and reached to the front part of the pressure side of the next blade in the case of small tip-clearance. Moreover, the double leakage flow occurred in the case of small tip clearance. In contrast, the leakage vortex reached to the leading edge of the next blade, and the spillage of the tip leakage flow from the leading edge occurred in the case of large tip clearance. The spillage of the tip leakage flow induced the larger total pressure loss than that induced by the double leakage flow. Therefore, the pressure rise in the case of large tip clearance is lower than that in the case of small tip clearance at the operating limit. It was concluded from the experimental and numerical results at the high stagger-angle setting for rotor blade that the loss generation mechanism depended on the behavior of tip-leakage vortex and that this behavior also depended on the tip-clearance.

Effects of Endwall Boundary Layer Thickness and Blade Tip Geometry on Flow Through High Pressure Turbine Passages

2014 ◽  
Author(s):  
Ralph J. Volino
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Layer Thickness ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Thickness ◽  
Total Pressure Loss ◽  
High Pressure Turbine ◽  
Image Velocimetry ◽  
Blade Tip

Experiments were conducted in a linear high pressure turbine cascade with an adjustable tip gap at one endwall. The cascade included a wake generator with moving rods that simulated the effect of an upstream vane row. Cases were documented with no tip gap, a gap of 1.5% of axial chord, and a gap of 3.8% of chord. Cases with flat blade tips were considered with thick and thin endwall boundary layers. Cases with flat tips and squealer tips were documented with the thin endwall boundary layer. For all cases data were acquired both with and without upstream wakes. Documentation included total pressure loss fields in the endwall region and corresponding velocity fields acquired using particle image velocimetry (PIV). The PIV measurements showed the various vortices in the flow field and their response to unsteady wakes. The strength and position of the vortices were directly related to regions of high total pressure loss. Reducing the endwall boundary layer thickness tended to reduce losses, but also resulted in increased leakage flow, which increased losses, particularly in cases with a large tip gap. The squealer tip reduced losses compared to the flat tip cases.

Endwall Contouring Optimization in a High Pressure Turbine Vane With Consideration of Rim Seal Flow

2014 ◽  
Author(s):  
Shuzhen Hu ◽  
Hualing Luo
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
High Pressure Turbine ◽  
Pressure Loss Coefficient ◽  
Turbine Vane ◽  
Endwall Contouring ◽  
Total Pressure Loss Coefficient

In this paper, non-axisymmetric endwall contouring optimizations in a high pressure turbine second vane passage were performed with and without including the rim seal flow. Aerodynamic performances of two contoured endwalls were studied and compared with the planar endwall case. The optimization was carried out with a commercial software package NUMECA Fine/Design 3D. A genetic algorithm was chosen as the optimization method. The objective was to minimize the total pressure loss in the high pressure turbine vane. Without the rim seal flow, the contoured endwall decreased the strength of the passage vortex, and reduced the total pressure loss coefficient up to 10.4%. However, the benefit achieved by this contoured endwall was lost when the rim seal flow was included by adding a under platform section in the computation model. With the existence of the rim seal flow, the flow field near the endwall was significantly changed. After recontouring the endwall using the optimization procedure with including rim seal flow, a 9.8% decrease in total pressure loss coefficient was achieved with the reduction of the vortex strength.

Conjugate Heat Transfer Characteristics of Double Wall Cooling With Gradient Diameter of Film and Impingement Holes

2021 ◽  
Author(s):  
Juan He ◽  
Qinghua Deng ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Direction ◽  
Turbine Blades ◽  
Total Pressure Loss ◽  
Advanced Technique ◽  
Cooling Effectiveness ◽  
Uniform Pattern ◽  
Double Wall

Abstract Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.

Aerodynamic optimisation of a winglet-cavity tip in a high-pressure axial turbine cascade

2016 ◽  
Vol 232 (4) ◽  
pp. 649-663 ◽  
Author(s):  
Zhihua Zhou ◽  
Shaowen Chen ◽  
Songtao Wang
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Pressure Side ◽  
Tip Clearance Flow ◽  
Clearance Flow

Tip clearance flow between rotating blades and the stationary casing in high-pressure turbines is very complex and is one of the most important factors influencing turbine performance. The rotor with a winglet-cavity tip is often used as an effective method to improve the loss resulting from the tip clearance flow. In this study, an aerodynamic geometric optimisation of a winglet-cavity tip was carried out in a linear unshrouded high-pressure axial turbine cascade. For the purpose of shaping the efficient winglet geometry of the rotor tip, a novel parameterisation method has been introduced in the optimisation procedure based on the computational fluid dynamics simulation and analysis. The reliability of a commercial computational fluid dynamics code with different turbulence models was first validated by contrasting with the experimental results, and the numerical total pressure loss and flow angle using the Baseline k-omega Model (BSL κ-ω model) shows a better agreement with the test data. Geometric parameterisation of blade tips along the pressure side and suction side was adopted to optimise the tip clearance flow, and an optimal winglet-cavity tip was proven to achieve lower tip leakage mass flow rate and total pressure loss than the flat tip and cavity tip. Compared to the numerical results of flat tip and cavity tip, the optimised winglet-cavity design, with the winglet along the pressure side and suction side, had lower tip leakage mass flow rate and total pressure loss. It offered a 35.7% reduction in the change ratio [Formula: see text]. In addition, the optimised winglet along pressure side and suction side, respectively, by using the parameterisation method was studied for investigating the individual effect of the pressure-side winglet and suction-side winglet on the tip clearance flow. It was found that the suction-side extension of the optimal winglet resulted in a greater reduction of aerodynamic loss and leakage mass flow than the pressure-side extension of the optimal winglet. Moreover, with the analysis based on the tip flow pattern, the numerical results show that the pressure-side winglet reduced the contraction coefficient, and the suction-side winglet reduced the aerodynamic loss effectively by decreasing the driving pressure difference near the blade tips, the leakage flow velocity, and the interaction between the leakage flow and the main flow. Overall, a better aerodynamic performance can be obtained by adopting the pressure-side and suction-side winglet-cavity simultaneously.

Steady CFD Simulation of a High Pressure Turbine Rotor With Hub and Shroud Purge Flow

2018 ◽  
Author(s):  
Wolfgang Sanz ◽  
Stefan Zerobin ◽  
Manfred Egger ◽  
Pascal Bader ◽  
Paul Pieringer ◽  
...  
Keyword(s):  
High Pressure ◽  
Secondary Flow ◽  
Cfd Simulation ◽  
Turbine Rotor ◽  
High Pressure Turbine ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Contour Plots ◽  
Purge Flow ◽  
Layer Flow

Purge air is injected at the hub and shroud of axial turbines in order to avoid hot gas entering the gaps between stationary and rotating blade rows. The purge flows considerably interact with the main flow and influence the secondary flow like the tip leakage vortex. Therefore, at Graz University of Technology the flow in a product-representative one-and-a-half stage test turbine under the influence of purge flows was investigated. Four individual purge mass flows differing in flow rate, pressure, and temperature were injected through hub and tip cavities before and after the unshrouded high-pressure turbine rotor. In order to get more insight into the cavity flows and the flow evolution in the rotor this configuration is studied with a steady CFD simulation with and without purge flows. It was found that the secondary flow and especially the tip leakage vortex is significantly influenced by the purge flow which varies in circumferential direction. The differences between purge and zero-purge flow conditions are discussed with the help of radial distributions and contour plots of stream-wise vorticity. Streamlines allow to follow the path of the purge flows in the rotor and show the radial displacement of the secondary flow vortices. Wall streamlines describe the changes in the boundary layer flow and their effect on the vorticity after the trailing edge.

Aerothermodynamics of a High-Pressure Turbine Blade With Very High Loading and Vortex Generators

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
Reinaldo A. Gomes ◽  
Reinhard Niehuis
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Flow Separation ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Vortex Generators ◽  
High Pressure Turbine ◽  
Very High

AITEB-2 is a project where aerothermal challenges of modern high pressure turbine designs are analyzed. One of the scopes of the project is to allow for new gas turbine designs with less parts and lighter jet engines by increasing the blade pitch and therefore the aerodynamic blade loading. For transonic profiles, this leads to very high velocities on the suction side and shock induced separation is likely to occur. The total pressure loss increase due to flow separation and strong shocks, as well as the underturning of the flow, limits the increase of the blade pitch. In this paper, experiments using a linear turbine blade cascade with high aerodynamic loading are presented. The blade pitch is increased such that at design conditions, a strong separation occurs on the suction side. The experiments were run at high subsonic exit Mach numbers and at Reynolds numbers of 390,000 and 800,000. In order to reduce the flow separation and the aerodynamic losses, air jet vortex generators are used, which create streamwise vortices prior to the separation start. Since in high pressure turbine blades film cooling is widely used, also the influence of film cooling both with and without using vortex generators is analyzed. Film cooling is provided on the suction side by two rows of cylindrical holes. This paper provides an analysis of the influence of different main flow conditions, film cooling, and vortex generators on total pressure loss, heat transfer and film cooling effectiveness. The experiments show that the vortex generators, as well as the film cooling reduce flow separation and total pressure losses. The effects are also seen in the local heat transfer, especially with enhanced heat transport in the region with flow separation. The cases presented in this paper deal with complex flow phenomena, which are challenging to be predicted with modern numerical tools correctly. Therefore, the experimental data serve as a comprehensive database for validation of simulation tools in the AITEB-2 project.

scholarly journals A Compact Diffuser System for Annular Combustors

1983 ◽  
Author(s):  
R. C. Adkins ◽  
J. O. Yost
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Pressure Recovery ◽  
Analytical Relationship ◽  
Exit Mach Number ◽  
Aerodynamic Simulation

Airflow tests have been conducted on an aerodynamic simulation of a combustor with pre-diffuser of compact configuration. The inlet Mach number throughout the tests was 0.35. The configuration was successful because of the attainment of a high pressure recovery, (Cp = 0.80), coupled with an exceptionally low total pressure loss (λ = 0.04). A useful analytical relationship is derived between the aerodynamic performance of combustor, compressor exit Mach number and diffuser performance.

scholarly journals The Influence of Downstream Passage on the Flow Within an Annular S-Shaped Duct

1997 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Toshiyuki Arima ◽  
Mineyasu Oana
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Horseshoe Vortex ◽  
Meridional Flow ◽  
Total Pressure Loss ◽  
Navier Stokes ◽  
Compressor Performance

Experimental and numerical investigations were carried out to gain a better understanding or the flow characteristics within an annular S-shaped duct, including the influence of the shape of the downstream passage located at the exit of the duct on the flow. A duct with six struts and the same geometry as that used to connect the compressor spools on our new experimental small two-spool turbofan engine was investigated. Two types of downstream passage were used. One type had a straight annular passage and the other a curved annular passage with a similar meridional flow path geometry to that of the centrifugal compressor. Results showed that the total pressure loss near the hub is large due to instability of the flow, as compared with that near the casing. Also, a vortex related to the horseshoe vortex was observed near the casing, in the case of the curved annular passage, the total pressure loss near the hub was greatly increased compared with the case of the straight annular passage, and the spatial position of the above vortex depends on the passage core pressure gradient. Furthermore, results of calculation using an in-house-developed three-dimensional Navier-Stokes code with a low Reynolds number k-ε turbulence model were in good qualitative agreement with experimental results. According to the simulation results, a region of very high pressure loss is observed near the hub at the duct exit with the increase of inlet boundary layer thickness. Such regions of high pressure loss may act on the downstream compressor as a large inlet distortion, and strongly affect downstream compressor performance.

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