Sensitivity of Cascade Pressure Distribution for Inverse Design of Turbine Blade

2019 ◽  
Author(s):  
R. Nanthini ◽  
B. V. S. S. S. Prasad ◽  
Y. V. S. S. Sanyasiraju
Keyword(s):  
Pressure Distribution ◽  
Turbine Blade ◽  
Wedge Angle ◽  
Leading Edge ◽  
Small Variation ◽  
Suction Side ◽  
Inviscid Flow ◽  
Initial Guess ◽  
Inverse Design ◽  
Trailing Edge

Abstract In an iterative inverse design of a turbine blade, choice of initial guess profile is crucial. As the pressure distribution is very sensitive to the leading and trailing edge shapes and the profile slope and curvature, a good initial guess profile will help in faster convergence. In this paper, the sensitivity of the pressure distribution is determined by carrying out numerical simulations with ANSYS Fluent 17.2 for the inviscid flow. The flow domain comprises of a two dimensional transonic turbine cascade. It consists of a turbine blade enclosed by inlet, outlet and periodic boundaries. Inlet total pressure, total temperature and inlet angle are given as the boundary conditions at the inlet and static pressure is imposed at the outlet boundary. The flow is solved for continuity, momentum and energy equations. Sensitivity of different parameters — leading edge thickness, trailing edge thickness, leading edge shape, inlet and outlet wedge angle on the pressure distribution is demonstrated for VKI blade cascade. It is found that the pressure side of the profile is less sensitive and that even a small variation in suction side of the profile geometry can affect the performance of the blade significantly. It is shown that, with the proposed methodology and sequence of steps, the final guess blade is quite close to the original blade.

Turbine Blade Trailing Edge Flow Characteristics at High Subsonic Outlet Mach Number

2003 ◽  
Vol 125 (2) ◽  
pp. 298-309 ◽  
Author(s):  
Claus H. Sieverding ◽  
Hugues Richard ◽  
Jean-Michel Desse
Keyword(s):  
Mach Number ◽  
Pressure Distribution ◽  
Turbine Blade ◽  
Vortex Shedding ◽  
High Speed ◽  
Pressure Coefficient ◽  
Vortex Formation ◽  
Flow Characteristics ◽  
Suction Side ◽  
Trailing Edge

The paper presents an experimental investigationof the effect of the trailing edge vortex shedding on the steady and unsteady trailing blade pressure distribution of a turbine blade at high subsonic Mach number M2,is=0.79 and high Reynolds number RE=2.8×106. The vortex formation and shedding process is visualized using a high-speed schlieren camera and a holographic interferometric density measuring technique. The blade is equipped with a rotatable trailing edge cylinder instrumented side-by-side with a pneumatic pressure tap and a fast response pressure sensor for detailed measurements of the trailing edge pressure distribution. The experiments demonstrate that contrary to the isobaric dead air region demonstrated at low subsonic Mach numbers the data reveal the existence of a highly nonuniform trailing edge pressure distribution with a strong pressure minimum at the center of the trailing edge. This finding is significant for the determination of the base pressure coefficient that is in general measured with a single pressure-sensing hole at the trailing edge center. The paper investigates further the effect of the vortex shedding on the blade rear suction side and discusses the superposition of unsteady effects emanating from the trailing edge and from the neighboring blade. The experimental data are a unique source for the validation of unsteady Navier-Stokes codes.

Three-Dimensional and Rotational Aerodynamics on the NREL Phase VI Wind Turbine Blade

2008 ◽  
Vol 130 (3) ◽  
Author(s):  
Alvaro Gonzalez ◽  
Xabier Munduate
Keyword(s):  
Wind Turbine ◽  
Turbine Blade ◽  
Three Dimensional ◽  
Separated Flow ◽  
Leading Edge ◽  
Wind Turbine Blade ◽  
Trailing Edge ◽  
Rotating Blade ◽  
Nrel Phase Vi ◽  
Edge Separation

This work undertakes an aerodynamic analysis over the parked and the rotating NREL Phase VI wind turbine blade. The experimental sequences from NASA Ames wind tunnel selected for this study respond to the parked blade and the rotating configuration, both for the upwind, two-bladed wind turbine operating at nonyawed conditions. The objective is to bring some light into the nature of the flow field and especially the type of stall behavior observed when 2D aerofoil steady measurements are compared to the parked blade and the latter to the rotating one. From averaged pressure coefficients together with their standard deviation values, trailing and leading edge separated flow regions have been found, with the limitations of the repeatability of the flow encountered on the blade. Results for the parked blade show the progressive delay from tip to root of the trailing edge separation process, with respect to the 2D profile, and also reveal a local region of leading edge separated flow or bubble at the inner, 30% and 47% of the blade. For the rotating blade, results at inboard 30% and 47% stations show a dramatic suppression of the trailing edge separation, and the development of a leading edge separation structure connected with the extra lift.

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.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

scholarly journals Numerical Study of Sediment Erosion Analysis in Francis Turbine

2019 ◽  
Vol 11 (5) ◽  
pp. 1423 ◽  
Author(s):  
Md Rakibuzzaman ◽  
Hyoung-Ho Kim ◽  
Kyungwuk Kim ◽  
Sang-Ho Suh ◽  
Kyung Kim
Keyword(s):  
Erosion Rate ◽  
Cavitation Erosion ◽  
Numerical Study ◽  
Leading Edge ◽  
Suction Side ◽  
Hydraulic Turbine ◽  
Francis Turbine ◽  
Trailing Edge ◽  
Sand Erosion ◽  
Turbine Design

Effective hydraulic turbine design prevents sediment and cavitation erosion from impacting the performance and reliability of the machine. Using computational fluid dynamics (CFD) techniques, this study investigated the performance characteristics of sediment and cavitation erosion on a hydraulic Francis turbine by ANSYS-CFX software. For the erosion rate calculation, the particle trajectory Tabakoff–Grant erosion model was used. To predict the cavitation characteristics, the study’s source term for interphase mass transfer was the Rayleigh–Plesset cavitation model. The experimental data acquired by this study were used to validate the existing evaluations of the Francis turbine. Hydraulic results revealed that the maximum difference was only 0.958% compared with the CFD data, and 0.547% compared with the experiment (Korea Institute of Machinery and Materials (KIMM)). The turbine blade region was affected by the erosion rate at the trailing edge because of their high velocity. Furthermore, in the cavitation–erosion simulation, it was observed that abrasion propagation began from the pressure side of the leading edge and continued along to the trailing edge of the runner. Additionally, as sediment flow rates grew within the area of the attached cavitation, they increased from the trailing edge at the suction side, and efficiency was reduced. Cavitation–sand erosion results then revealed a higher erosion rate than of those of the sand erosion condition.

Measurement and Simulation of a Transonic Innovative Cooling System (ICS) for High-Temperature Transonic Gas Turbine Stages

2004 ◽  
Author(s):  
E. Go¨ttlich ◽  
L. Innocenti ◽  
A. Vacca ◽  
W. Sanz ◽  
J. Woisetschla¨ger ◽  
...  
Keyword(s):  
Temperature Distribution ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Numerical Comparison ◽  
Blade Surface ◽  
Trailing Edge ◽  
Advantages And Disadvantages

Gas turbine design technology requires the development of transonic turbine stages capable of carrying high stage load and of handling hot gas temperatures at turbine inlet. A reliable cooling system is necessary to cope with shocks emanating from preceding blade rows and impinging on the blade especially in the leading edge region. In order to fulfill these requirements researchers at Graz University of Technology have been working on an Innovative Cooling System (ICS) since 1995. The ICS is able to cover large areas of the blade surface with an effective cooling film and to reduce the metal temperature without a shower head cooling arrangement at the leading edge and any trailing edge cooling air ejection. In this paper the authors present a numerical comparison of the ICS to a conventional modern film cooling system both implemented in the same industrial transonic gas turbine blade. An experimental determination of the adiabatic film cooling effectiveness distribution around the blades surface was necessary for the ICS because of its uncommon design. The measurements were done on a cylindrical blade in a linear cascade arrangement. An infrared camera system was used to determine the effectiveness of this newly designed cooling system by measuring the temperature distribution on the blade surface. Then a numerical simulation of heat transfer and of internal and external cooling for the turbine blade at test rig conditions was performed. The ICS showed a lower outer wall temperature distribution of the blade compared to a standard film cooling system. The heavily loaded leading edge as well as the trailing edge are well cooled. Further conclusions on the advantages and disadvantages of the ICS are drawn.

An Experimental Investigation of Flow Characteristics Downstream of Discrete Film Cooling Holes on Turbine Blade Leading Edge

2011 ◽  
Vol 383-390 ◽  
pp. 5553-5560
Author(s):  
Shao Hua Li ◽  
Hong Wei Qu ◽  
Mei Li Wang ◽  
Ting Ting Guo
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Suction Side ◽  
Wall Region ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Curvature Gradient ◽  
Film Cooling Holes

The gas turbine blade was studied on the condition that the mainstream velocity was 10m/s and the Renolds number based on the chord length of the blade was 160000.The Hot-film anemometer was used to measure the two-dimension speed distribution along the downstream of the film cooling holes on the suction side and the pressure side. The conclusions are as follows: When the blowing ratio of the suction side and the pressure side increasing, the the mainstream and the jet injection mixing center raising. Entrainment flow occurs at the position where the blade surface with great curvature gradient, simultaneously the mixing flow has a wicked adhere to the wall. The velocity gradient of the u direction that on the suction side increase obviously, also the level of the wall adherence is better than the pressure side. With the x/d increasing, the velocity u that on the pressure side gradually become irregularly, also the secondary flow emerged near the wall region where the curvature is great. The blowing ratio on the suction side has a little influence on velocity v than that on the pressure side.

A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part I — Introduction and Aerodynamic Measurements

2015 ◽  
Author(s):  
Franz Puetz ◽  
Johannes Kneer ◽  
Achmed Schulz ◽  
Hans-Joerg Bauer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Film Cooling ◽  
Gas Turbines ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Test Facility ◽  
Endwall Contouring

An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a flattened temperature profile at the outlet. As a consequence, the heat load on the endwall of the first nozzle guide vane is increased. The secondary flow field dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and film cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and film cooling on both a flat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and influence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall flow field. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall flow field. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of flow field in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.

NUMERICAL INVESTIGATIONS OF A CENTRIFUGAL BLOOD PUMP

2004 ◽  
Vol 04 (03) ◽  
pp. 237-255 ◽  
Author(s):  
W. K. CHAN ◽  
Y. W. WONG ◽  
Y. DING
Keyword(s):  
Leading Edge ◽  
Suction Side ◽  
Small Region ◽  
Inverse Design ◽  
Blood Pump ◽  
Impeller Blade ◽  
Centrifugal Blood Pump ◽  
Computational Fluid Dynamics Cfd ◽  
Inlet Angle ◽  
Side Flow

This paper presents computational fluid dynamics (CFD) studies of a centrifugal blood pump. 3-D models of five different blade geometries are investigated numerically using CFX-TASCflow. The impellers were designed using an inverse design technique where the swirl distributions were prescribed. The results showed the flow in the impeller passages is highly dependent on the impeller blade profiles. The flow in the radial blade impeller is unsatisfactory as flow separates at the leading edge of the suction side. Flow is confined mainly to the pressure side. Design 2, with an inlet angle of 6.7° and outlet angle of 30°, offers the greatest potential as only a small region of flow reversal is detected. Further optimization is necessary to completely eliminate regions of flow reversals. The highest scalar shear stress in both designs is 240 Pa and 120 Pa respectively. In addition, this paper demonstrates that the use of inverse design can help the designer to better design and analyze the flow field in centrifugal blood pumps.

Study of Particle Ingestion Through Two-Stage Gas Turbine

2014 ◽  
Author(s):  
Adel Ghenaiet
Keyword(s):  
Gas Flow ◽  
Numerical Study ◽  
Leading Edge ◽  
Suction Side ◽  
Solid Particles ◽  
Two Stage ◽  
Trailing Edge ◽  
Erosion Rates ◽  
Tracking Model ◽  
Over The Top

This paper presents a numerical study of particle laden gas flow through a two-stage hp axial turbine, by means of an in-house code based on the Lagrangian tracking model and the finite element method. As fly-ash solid particles trajectories and locations of impacts are predicted, the local erosion rates and the deteriorations of blades are assessed. The computed trajectories provide a detailed description of particles behaviors and reveal that particle impacts on the aft of vane pressure side usually lead to significant variations in the directions of particles to the next rotor blade, and subsequently particles impact the suction side. The plots of equivalent erosion rates indicate the vanes and blades locations which suffer more erosion. The first vane pressure surface is impacted more than any other component, but higher rates are seen at the top corner from trailing edge. The critical regions of erosion wear in the first rotor are observed over the top of blade leading edge extending along the tip as well as a rounding of the top corner from trailing edge. In the second vane, the regions of higher erosion are revealed over the last third of leading edge and the top corner extending along tip. The erosion in the second rotor is over a large area of suction side till the tip corner. The predicted areas of extreme erosion, also shown by the deteriorated profiles, are indicators for anticipated vanes and blades failures.

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