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.

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.

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 Investigation of Dump Diffuser Combustor Performance at Uniform and Non-Uniform Inlet Conditions

2020 ◽  
Author(s):  
Heyu Wang ◽  
Kai Hong Luo
Keyword(s):  
Boundary Layer ◽  
Numerical Investigation ◽  
Layer Thickness ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Thickness ◽  
Static Pressure ◽  
Aerodynamic Performance ◽  
Aero Engine ◽  
Inlet Conditions

Abstract A numerical investigation has been conducted for an axisymmetric dump diffuser combustor, which is a simplified geometry of a typical lean-burn combustor in a modern civil aero-engine gas turbine. The aerodynamic performance of the combustor is analyzed with an emphasis on two common performance parameters: static pressure recovery and total pressure loss. The former is essential in maintaining high-pressure air flow across the liner, whereas the latter involves the specific fuel consumption of the aero-engine. At first, the effects of geometrical parameters of the dump diffuser combustor are investigated. A high diffuser angle seems to be detrimental to both static pressure recovery and total pressure loss. On the other hand, a high dump gap ratio is beneficial from the aerodynamic performance point of view. However, all these desired characteristics are subject to mechanical constraints and their implications for specific consumption. Optimum values of those parameters should exist for a given desired aerodynamics performance. The majority of previous researches, including the first part of this study, have been carried out with uniform inlet conditions due to a typical independent design cycle of each component. The effects of compressor exit conditions are usually not considered in the early stage design process. In the second part of this study, various inlet conditions representing a more realistic compressor exit condition such as inlet symmetrical and asymmetrical boundary layer thickness are investigated. The performance of an asymmetrical configuration with a thin boundary layer thickness near the outer annulus is almost comparable to that of its uniform counterpart. Findings of this study provide useful input for combustor designers to improve the combustor’s performance based on the compressor exit conditions.

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.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

Location Effect of Boundary Layer Suction on Compressor Hub-Corner Separation

2014 ◽  
Author(s):  
Ping-Ping Chen ◽  
Wei-Yang Qiao ◽  
Karsten Liesner ◽  
Robert Meyer
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Base Flow ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Corner Separation ◽  
Boundary Layer Suction

The large secondary flow area in the compressor hub-corner region usually leads to three-dimensional separation in the passage with large amounts of total pressure loss. In this paper numerical simulations of a linear high-speed compressor cascade, consisting of five NACA 65-K48 stator profiles, were performed to analyze the flow mechanism of hub-corner separation for the base flow. Experimental validation is used to verify the numerical results. Active control of the hub-corner separation was investigated by using boundary layer suction. The influence of the selected locations of the endwall suction slot was investigated in an effort to quantify the gains of the compressor cascade performance. The results show that the optimal chordwise location should contain the development section of the three-dimensional corner separation downstream of the 3D corner separation onset. The best pitchwise location should be close enough to the vanes’ suction surface. Therefore the optimal endwall suction location is the MTE slot, the one from 50% to 75% chord at the hub, close to the blade suction surface. By use of the MTE slot with 1% suction flow ratio, the total-pressure loss is substantially decreased by about 15.2% in the CFD calculations and 9.7% in the measurement at the design operating condition.

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.

Effects of axial length and integrated design on the aggressive intermediate turbine duct

2018 ◽  
Vol 233 (4) ◽  
pp. 443-456 ◽  
Author(s):  
Qingzong Xu ◽  
Pei Wang ◽  
Qiang Du ◽  
Jun Liu ◽  
Guang Liu
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Secondary Flow ◽  
Axial Length ◽  
Pressure Loss ◽  
Total Pressure ◽  
Guide Vane ◽  
Three Dimensional ◽  
Intermediate Turbine Duct ◽  
Aero Engine

With the increasing demand of high bypass ratio and thrust-to-weight ratio in civil aero-engine, the intermediate turbine duct between the high pressure and low pressure turbines of a modern gas turbine tends to shorter axial length, larger outlet-to-inlet area ratio and high pressure-to-low pressure radial offset. This paper experimentally and numerically investigated the three-dimensional flow characteristics of traditional (ITD1) and aggressive intermediate turbine duct (ITD2) at low Reynolds number. The baseline case of ITD1 is representative of a traditional intermediate turbine duct of aero-engine design with non-dimensional length of L/dR = 2.79 and middle angle of 20.12°. The detailed flow fields inside ITD1 and flow visualization were measured. Results showed the migration of boundary layer and a pair of counter-rotating vortexes were formed due to the radial migration of low momentum fluid. With the decreasing axial length of intermediate turbine duct, the radial and streamwise reverse pressure gradient in aggressive intermediate turbine duct (ITD2) were increased resulting in severe three-dimensional separation of boundary layer near casing surface and higher total pressure loss. The secondary flow and separation of boundary layer near the endwall were deeply analyzed to figure out the main source of high total pressure loss in the aggressive intermediate turbine duct (ITD2). Based on that, employing wide-chord guide vane to substitute “strut + guide vane”, this paper designed the super-aggressive intermediate turbine duct and realized the suppression of the three-dimensional separation and secondary flow.

Effect of Turbine Blade tip shape on the Total Pressure Loss of a Turbine Cascade

2009 ◽  
Vol 12 (2) ◽  
pp. 39-45 ◽  
Author(s):  
Ki-Seon Lee ◽  
Seoung-Duck Park ◽  
Young-Chul Noh ◽  
Hak-Bong Kim ◽  
Jae-Su Kwak ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Turbine Cascade ◽  
Blade Tip

Accurate Estimation of Profile Losses and Analysis of Loss Generation Mechanisms in a Turbine Cascade

2017 ◽  
Vol 139 (12) ◽  
Author(s):  
D. Lengani ◽  
D. Simoni ◽  
M. Ubaldi ◽  
P. Zunino ◽  
F. Bertini ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Accurate Estimation ◽  
Loss Mechanism ◽  
Pressure Loss Coefficient ◽  
Generation Mechanisms ◽  
Total Pressure Loss Coefficient

The paper analyzes losses and the loss generation mechanisms in a low-pressure turbine (LPT) cascade by proper orthogonal decomposition (POD) applied to measurements. Total pressure probes and time-resolved particle image velocimetry (TR-PIV) are used to determine the flow field and performance of the blade with steady and unsteady inflow conditions varying the flow incidence. The total pressure loss coefficient is computed by traversing two Kiel probes upstream and downstream of the cascade simultaneously. This procedure allows a very accurate estimation of the total pressure loss coefficient also in the potential flow region affected by incoming wake migration. The TR-PIV investigation concentrates on the aft portion of the suction side boundary layer downstream of peak suction. In this adverse pressure gradient region, the interaction between the wake and the boundary layer is the strongest, and it leads to the largest deviation from a steady loss mechanism. POD applied to this portion of the domain provides a statistical representation of the flow oscillations by splitting the effects induced by the different dynamics. The paper also describes how POD can dissect the loss generation mechanisms by separating the contributions to the Reynolds stress tensor from the different modes. The steady condition loss generation, driven by boundary layer streaks and separation, is augmented in the presence of incoming wakes by the wake–boundary layer interaction and by the wake dilation mechanism. Wake migration losses have been found to be almost insensitive to incidence variation between nominal and negative (up to −9 deg) while at positive incidence, the losses have a steep increase due to the alteration of the wake path induced by the different loading distribution.

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