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

2017 ◽  
Author(s):  
D. Lengani ◽  
D. Simoni ◽  
M. Ubaldi ◽  
P. Zunino ◽  
F. Bertini ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Region ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Accurate Estimation ◽  
Turbine Cascade ◽  
Loss Mechanism ◽  
Generation Mechanisms

The paper analyzes losses and the loss generation mechanisms in a low-pressure turbine 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 co-efficient 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 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 −9deg), while at positive incidence the losses have a steep increase due to the alteration of the wake path induced by the different loading distribution.

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.

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.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

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

Impact of a Combination of Micro-Vortex Generator and Boundary Layer Suction on Performance in a High-Load Compressor Cascade

2018 ◽  
Author(s):  
Shan Ma ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Chuanle Liu
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Control Methods ◽  
Compressor Cascade ◽  
Boundary Layer Suction ◽  
Micro Vortex Generator

The performance of a compressor cascade is considerably influenced by flow control methods. In this paper, the synergistic effects of combination between micro-vortex generators (MVG) and boundary layer suction (BLS) are discussed in a high-load compressor cascade. Seven cases, which are grouped by a kind of micro-vortex generator and boundary layer suction with three locations, are investigated to control secondary flow effects and enhance the aerodynamic performance of the compressor cascade. The MVG is mounted on the end-wall in front of the passage. The rectangle suction slot with three radial positions is installed on the blade suction surface near the trailing edge. The numerical results show that: at the design condition, the total pressure loss is effectively decreased as well as the static pressure coefficient increase when the combined MVG and SBL method (COM) is used, which is superior to MVG in an aerodynamic performance. At the stall condition, the induced vortex coming from MVG could mix the low-energy fluid and mainstream, which result in the reduced separation, and the total pressure loss decreased by 11.54% when the suction flow ratio is 1.5%. The total pressure loss decreases by 14.59% when the COM control methods are applied.

Optimization of a U-Bend for Minimal Pressure Loss in Internal Cooling Channels: Part I—Numerical Method

2011 ◽  
Author(s):  
Tom Verstraete ◽  
Filippo Coletti ◽  
Je´re´my Bulle ◽  
Timothe´e Vanderwielen ◽  
Tony Arts
Keyword(s):  
Numerical Method ◽  
Pressure Loss ◽  
Total Pressure ◽  
Cooling System ◽  
Differential Evolution Algorithm ◽  
Flow Region ◽  
Total Pressure Loss ◽  
Navier Stokes ◽  
Internal Cooling ◽  
Cooling Channels

This two-parts paper addresses the design of a U-bend for serpentine internal cooling channels optimized for minimal pressure loss. The total pressure loss for the flow in a U-bend is a critical design parameter as it augments the pressure required at the inlet of the cooling system, resulting in a lower global efficiency. In this first part of the paper the design methodology of the cooling channel is presented. The minimization of the total pressure loss is achieved by means of a numerical optimization method that uses a metamodel assisted differential evolution algorithm in combination with an incompressible Navier-Stokes solver. The profiles of the internal and external side of the bend are parameterized using piece-wise Bezier curves. This allows for a wide variety of shapes, respecting the manufacturability constraints of the design. The pressure loss is computed by the Navier-Stokes solver, which is based on a two-equation turbulence model and is available from the open source software OpenFOAM. The numerical method predicts an improvement of 36% in total pressure drop with respect to a circular U-bend, mainly due to the reduction of the separated flow region along the internal side of the bend. The resulting design is subjected to experimental validation, presented in Part II of the paper.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2017 ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake’s path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue 3-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the Stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady Stator 2 performance, time-resolved interrogations of the Stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The Stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between Rotor 1 and Rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the Stator 1 wake and Rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, Stator 2 experiences significantly different inlet blockage and unsteadiness from the Rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of Stator 2.

scholarly journals The Effect of Inlet Boundary Layer Thickness on the Flow Within an Annular S-Shaped Duct

1998 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Toshiyuki Arima ◽  
Mineyasu Oana
Keyword(s):  
Boundary Layer ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Vortex Pair ◽  
Separated Flow ◽  
Total Pressure Loss ◽  
Centrifugal Impeller

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the effect of the inlet boundary layer (IBL) on the flow. A duct with six struts and the same geometry as that used to connect compressor spools on our experimental small two-spool turbofan engine was investigated. A curved downstream annular passage with a similar meridional flow path geometry to that of the centrifugal compressor has been fitted at the exit of S-shaped duct. Two types of the IBL (i.e. thin and thick IBL) were used. Results showed that large differences of flow pattern were observed at the S-Shaped duct exit between two types of the IBL, though the value of “net” total pressure loss has not been remarkably changed. According to “overall” total pressure loss, which includes the IBL loss, the total pressure loss was greatly increased near the hub as compared to that for a thin one. For the thick IBL, a vortex pair related to the hub-side horseshoe vortex and the separated flow found at the strut trailing edge has been clearly captured in the form of the total pressure loss contours and secondary flow vectors, experimentally and numerically. The high-pressure loss regions on either side of the strut wake near the hub may act on a downstream compressor as a large inlet distortion, and strongly affect the downstream compressor performance. There is a much-distorted three-dimensional flow pattern at the exit of S-Shaped duct. This means that the aerodynamic sensitivity of S-Shaped duct to the IBL thickness is very high. Therefore, sufficient carefulness is needed to design not only downstream aerodynamic component (for example centrifugal impeller) but also upstream aerodynamic component (LPC OGV).

Sign in / Sign up

Export Citation Format

Share Document