scholarly journals Uncertainty Quantification Methodologies Applied to the Rotor Tip Clearance Effect in a Twin Scroll Radial Turbine

Fluids ◽  
2020 ◽  
Vol 5 (3) ◽  
pp. 114
Author(s):  
Carlo Cravero ◽  
Andrea Ottonello
Keyword(s):  
Uncertainty Quantification ◽  
Leading Edge ◽  
Expansion Method ◽  
Quantitative Information ◽  
Tip Clearance ◽  
Simulation Tools ◽  
Numerical Predictions ◽  
Physical Experiments ◽  
Comparison Of The Results ◽  
Radial Turbines

In the last three decades computer simulation tools have achieved wide spread use in the design and analysis of engineering devices. This has shortened the overall product design cycle (physical experiments may be impossible during early design stages) and it has also provided better understanding of the operating behavior of the systems under investigation. As a consequence numerical simulation have led to a reduction of physical prototyping and to lower costs for manufacturing production chains. Despite this success, it remains difficult to provide objective confidence levels in quantitative information derived from numerical predictions. The complexity arises from the amount of uncertainties related to the inputs of any computation attempting to represent a physical system. This paper focuses on geometrical sources of uncertainty in the field of CFD applied to twin scroll radial turbines. In particular it has been investigated the effect of uncertainties on tip clearance values at rotor blade leading edge and trailing edge on selected turbine performance parameters. The analysis shows the use of the Surrogate-based uncertainty quantification technique that has been setup by the authors in the Dakota® environment. The polynomial chaos expansion method has been applied to the same case. The comparison of the results coming from the different approaches and the discussion of the pros and cons related to each technique lead to interesting conclusions, which are proposed as guidelines for future UQ applications on the theme of CFD applied to radial turbines.

Rotating Stall Observations in a High Speed Compressor—Part II: Numerical Study

2015 ◽  
Vol 137 (5) ◽  
Author(s):  
J. Dodds ◽  
M. Vahdati
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Numerical Predictions ◽  
Stage 1 ◽  
Variable Stator

In this two-part paper the phenomenon of part span rotating stall is studied. The objective is to improve understanding of the physics by which stable and persistent rotating stall occurs within high speed axial flow compressors. This phenomenon is studied both experimentally (Part I) and numerically (Part II). The experimental observations reported in Part I are now explored through the use of 3D unsteady Reynolds-averaged Navier–Stokes (RANS) simulation. The objective is to both validate the computational model and, where possible, explore some physical aspects of the phenomena. Unsteady simulations are presented, performed at a fixed speed with the three rows of variable stator vanes adjusted to deliberately mismatch the front stages and provoke stall. Two families of rotating stall are identified by the model, consistent with experimental observations from Part I. The first family of rotating stall originates from hub corner separations developing on the stage 1 stator vanes. These gradually coalesce into a multicell rotating stall pattern confined to the hub region of the stator and its downstream rotor. The second family originates from regions of blockage associated with tip clearance flow over the stage 1 rotor blade. These also coalesce into a multicell rotating stall pattern of shorter length scale confined to the leading edge tip region. Some features of each of these two patterns are then explored as the variable stator vanes (VSVs) are mismatched further, pushing each region deeper into stall. The numerical predictions show a credible match with the experimental findings of Part I. This suggests that a RANS modeling approach is sufficient to capture some important aspects of part span rotating stall behavior.

Interaction of Tip Clearance Flow and Three-Dimensional Separations in Axial Compressors

2006 ◽  
Author(s):  
Semiu A. Gbadebo ◽  
Nicholas A. Cumpsty ◽  
Tom P. Hynes
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Tip Clearance ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Tip Clearance Flow ◽  
Separation Line ◽  
Numerical Predictions ◽  
Clearance Flow

This paper considers the interaction of tip clearance flow with three-dimensional (3D) separations in the corner region of a compressor cascade. Three-dimensional numerical computations were carried out using ten levels of tip clearance, ranging from zero to 2.18% of blade chord. The 3D separations on the blade suction surface were largely removed by the clearance flow for clearance about 0.58% of chord. For this cascade, experimental results at zero and 1.7% chord tip clearance were used to assess the validity of the numerical predictions. The removal mechanism was associated with the suppression of the leading edge horseshoe vortex and the interaction of tip clearance flow with the endwall boundary layer, which develops into a secondary flow as it is drifted towards the blade suction surface. Such interaction leads to the formation of a new 3D separation line on the endwall. The separation line forms the base of a separated stream surface which rolls up into the clearance vortex.

Rotating Stall Observations in a High Speed Compressor: Part 2 — Numerical Study

2014 ◽  
Author(s):  
J. Dodds ◽  
M. Vahdati
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Numerical Predictions ◽  
Stage 1 ◽  
Variable Stator

In this two part paper the phenomenon of part span rotating stall is studied. The objective is to improve understanding of the physics by which stable and persistent rotating stall occurs within high speed axial flow compressors. This phenomenon is studied both experimentally (part 1) and numerically (part 2). The experimental observations reported in Part 1 are now explored through the use of 3D unsteady RANS simulation. The objective is to both to validate the computational model and, where possible, explore some physical aspects of the phenomena. Unsteady simulations are presented, performed at a fixed speed with the three rows of variable stator stagger vanes adjusted to deliberately mismatch the front stages and provoke stall. Two families of rotating stall are identified by the model, consistent with experimental observations from Part 1. The first family of rotating stall originates from hub corner separations developing on the stage 1 stator vanes. These gradually coalesce into a multi-cell rotating stall pattern confined to the hub region of the stator and its downstream rotor. The second family originates from regions of blockage associated with tip clearance flow over the stage 1 rotor blade. These also coalesce into a multi-cell rotating stall pattern of shorter length scale confined to the leading edge tip region. Some features of each of these two patterns are then explored as the variable stator vanes are mismatched further, pushing each region deeper into stall. The numerical predictions show a credible match with the experimental findings of Part 1. This suggests that a RANS modelling approach is sufficient to capture some important aspects of part span rotating stall behavior.

Interaction of Tip Clearance Flow and Three-Dimensional Separations in Axial Compressors

2006 ◽  
Vol 129 (4) ◽  
pp. 679-685 ◽  
Author(s):  
Semiu A. Gbadebo ◽  
Nicholas A. Cumpsty ◽  
Tom P. Hynes
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Tip Clearance ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Tip Clearance Flow ◽  
Separation Line ◽  
Numerical Predictions ◽  
Clearance Flow

This paper considers the interaction of tip clearance flow with three-dimensional (3D) separations in the corner region of a compressor cascade. Three-dimensional numerical computations were carried out using ten levels of tip clearance, ranging from zero to 2.18% of blade chord. The 3D separations on the blade suction surface were largely removed by the clearance flow for clearance about 0.58% of chord. For this cascade, experimental results at zero and 1.7% chord tip clearance were used to assess the validity of the numerical predictions. The removal mechanism was associated with the suppression of the leading edge horseshoe vortex and the interaction of tip clearance flow with the endwall boundary layer, which develops into a secondary flow as it is driven towards the blade suction surface. Such interaction leads to the formation of a new 3D separation line on the endwall. The separation line forms the base of a separated stream surface which rolls up into the clearance vortex.

scholarly journals Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

scholarly journals Depth-Integrated Two-Phase Modeling of Two Real Cases: A Comparison between r.avaflow and GeoFlow-SPH Codes

2021 ◽  
Vol 11 (12) ◽  
pp. 5751
Author(s):  
Seyed Ali Mousavi Tayebi ◽  
Saeid Moussavi Tayyebi ◽  
Manuel Pastor
Keyword(s):  
Rock Avalanche ◽  
Two Phase ◽  
Simulation Code ◽  
Simulation Tools ◽  
Mesh Free ◽  
Particle Hydrodynamics ◽  
Landslide Evolution ◽  
Smoothed Particle ◽  
Hazard And Risk ◽  
Comparison Of The Results

Due to the growing populations in areas at high risk of natural disasters, hazard and risk assessments of landslides have attracted significant attention from researchers worldwide. In order to assess potential risks and design possible countermeasures, it is necessary to have a better understanding of this phenomenon and its mechanism. As a result, the prediction of landslide evolution using continuum dynamic modeling implemented in advanced simulation tools is becoming more important. We analyzed a depth-integrated, two-phase model implemented in two different sets of code to stimulate rapid landslides, such as debris flows and rock avalanches. The first set of code, r.avaflow, represents a GIS-based computational framework and employs the NOC-TVD numerical scheme. The second set of code, GeoFlow-SPH, is based on the mesh-free numerical method of smoothed particle hydrodynamics (SPH) with the capability of describing pore pressure’s evolution along the vertical distribution of flowing mass. Two real cases of an Acheron rock avalanche and Sham Tseng San Tsuen debris flow were used with the best fit values of geotechnical parameters obtained in the prior modeling to investigate the capabilities of the sets of code. Comparison of the results evidenced that both sets of code were capable of properly reproducing the run-out distance, deposition thickness, and deposition shape in the benchmark exercises. However, the values of maximum propagation velocities and thickness were considerably different, suggesting that using more than one set of simulation code allows us to predict more accurately the possible scenarios and design more effective countermeasures.

Influences of Tip Leakage Flows Discharged From Main and Splitter Blades on Flow Field in Transonic Centrifugal Compressor Stage

2018 ◽  
Author(s):  
Masanao Kaneko ◽  
Hoshio Tsujita
Keyword(s):  
Flow Field ◽  
Centrifugal Compressor ◽  
Leading Edge ◽  
Tip Clearance ◽  
Blade Loading ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Leakage Flows ◽  
Transonic Centrifugal Compressor ◽  
Tip Leakage Flows

A transonic centrifugal compressor impeller is generally composed of the main and the splitter blades which are different in chord length. As a result, the tip leakage flows from the main and the splitter blades interact with each other and then complicate the flow field in the compressor. In this study, in order to clarify the individual influences of these leakage flows on the flow field in the transonic centrifugal compressor stage at near-choke to near-stall condition, the flows in the compressor at four conditions prescribed by the presence and the absence of the tip clearances were analyzed numerically. The computed results clarified the following noticeable phenomena. The tip clearance of the main blade induces the tip leakage vortex from the leading edge of the main blade. This vortex decreases the blade loading of the main blade to the negative value by the increase of the flow acceleration along the suction surface of the splitter blade, and consequently induces the tip leakage vortex caused by the negative blade loading of the main blade at any operating points. These phenomena decline the impeller efficiency. On the other hand, the tip clearance of the splitter blade decreases the afore mentioned acceleration by the formation of the tip leakage vortex from the leading edge of the splitter blade and the decrease of the incidence angle for the splitter blade caused by the suction of the flow into the tip clearance. These phenomena reduce the loss generated by the negative blade loading of the main blade and consequently reduce the decline of the impeller efficiency. Moreover, the tip clearances enlarge the flow separation around the diffuser inlet and then decline the diffuser performance independently of the operating points.

Experiments on an Axial Fan Stage: Time-Resolved Analysis of Rotating Instability Modes

2015 ◽  
Vol 137 (6) ◽  
Author(s):  
Benjamin Pardowitz ◽  
Ulf Tapken ◽  
Lars Neuhaus ◽  
Lars Enghardt
Keyword(s):  
Leading Edge ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Axial Fan ◽  
Time Resolved ◽  
Instability Waves ◽  
Instability Modes ◽  
Mode Decomposition ◽  
Rotating Instability ◽  
Annular Cascade

Rotating instability (RI) occurs at off-design conditions in axial compressors, predominantly in rotor configurations with large tip clearances. Characteristic spectral signatures with side-by-side peaks below the blade passing frequency (BPF) are typically referred to RI located in the clearance region next to the leading edge (LE). Each peak can be assigned to a dominant circumferential mode. RI is the source of the clearance noise (CN) and an indicator for critical operating conditions. Earlier studies at an annular cascade pointed out that RI modes of different circumferential orders occur stochastically distributed in time and independently from each other, which is contradictory to existing explanations of RI. Purpose of the present study is to verify this generally with regard to axial rotor configurations. Experiments were conducted on a laboratory axial fan stage mainly using unsteady pressure measurements in a sensor ring near the rotor LE. A mode decomposition based on cross spectral matrices was used to analyze the spectral and modal RI patterns upstream of the rotor. Additionally, a time-resolved analysis based on a spatial discrete-Fourier-transform (DFT) was applied to clarify the temporal characteristics of the RI modes and their potential interrelations. The results and a comparison with the previous findings on the annular cascade corroborate a new hypothesis about the basic RI mechanism. This hypothesis implies that instability waves of different wavelengths are generated stochastically in a shear layer resulting from a backflow in the tip clearance region.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

scholarly journals Heat Transfer on a Film-Cooled Rotating Blade

1999 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Three Dimensional ◽  
Leading Edge ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

A multi-block, three-dimensional Navier-Stokes code has been used to compute heat transfer coefficient on the blade, hub and shroud for a rotating high-pressure turbine blade with 172 film-cooling holes in eight rows. Film cooling effectiveness is also computed on the adiabatic blade. Wilcox’s k-ω model is used for modeling the turbulence. Of the eight rows of holes, three are staggered on the shower-head with compound-angled holes. With so many holes on the blade it was somewhat of a challenge to get a good quality grid on and around the blade and in the tip clearance region. The final multi-block grid consists of 4784 elementary blocks which were merged into 276 super blocks. The viscous grid has over 2.2 million cells. Each hole exit, in its true oval shape, has 80 cells within it so that coolant velocity, temperature, k and ω distributions can be specified at these hole exits. It is found that for the given parameters, heat transfer coefficient on the cooled, isothermal blade is highest in the leading edge region and in the tip region. Also, the effectiveness over the cooled, adiabatic blade is the lowest in these regions. Results for an uncooled blade are also shown, providing a direct comparison with those for the cooled blade. Also, the heat transfer coefficient is much higher on the shroud as compared to that on the hub for both the cooled and the uncooled cases.

Sign in / Sign up

Export Citation Format

Share Document