scholarly journals Investigation of Flow in a Radial Turbine Using Laser Anemometry

1993 ◽  
Author(s):  
Sohail Hamid Zaidi ◽  
Robin L. Elder
Keyword(s):  
Inner Core ◽  
Swirling Flow ◽  
Guide Vane ◽  
Flow Direction ◽  
Flow Region ◽  
Flow Conditions ◽  
Trailing Edge ◽  
Outer Flow ◽  
Laser Anemometry ◽  
Nozzle Guide

A lightweight, high pressure radial inflow turbine was tested and laser anemometry used to measure the flow at various positions within the nozzle guide vanes, immediately upstream of the rotor and at two axial stations downstream of the rotor. The laser anemometry results indicated flow conditions within the nozzle vanes which were largely two dimensional (blade-to-blade with little hub to shroud variation) except at the vane outlet. Unsteadiness due to rotor blade passing effects were detected at the nozzle guide vane trailing edge but had almost entirely decayed at the vane throat. The results also indicate significant variations in flow conditions across the pitch of the nozzles suggesting incidence variations on the rotor of approaching 30 degrees. The laser anemometry results downstream of the turbine show a swirling flow characterised by a turbulent inner core region, a ‘centre annulus’ region of uniform velocity and flow direction and an outer flow region with a similar flow direction but velocity which increases rapidly towards the outer wall. The blade passing unsteadiness (blade-to-blade) is hardly noticeable some 50mm downstream of the rotor trailing edge.

Experimental and Numerical Investigation on the Influence of Trailing Edge Bleeding on the Aerodynamics of a NGV Cascade

2009 ◽  
Author(s):  
Marco Montis ◽  
Reinhard Niehuis ◽  
Mattia Guidi ◽  
Simone Salvadori ◽  
Francesco Martelli ◽  
...  
Keyword(s):  
Guide Vane ◽  
Measured Data ◽  
Flow Conditions ◽  
Trailing Edge ◽  
Pressure Losses ◽  
Ejection Rate ◽  
Nozzle Guide Vane ◽  
Flow Phenomena ◽  
Nozzle Guide ◽  
Good Agreement

A series of tests on a specific designed linear nozzle guide vane (NGV) cascade with trailing edge coolant ejection was carried out to investigate the influence of the trailing edge bleeding (TEB) on the loss behaviour of the profile. Wake traverses with a five-hole probe and measurements of the pressure distribution on the profile were taken varying the ejection rate under reference main flow conditions, namely Re2th = 1.056·106 and Ma2th = 0.8 (Re2th based on the true chord). Wake total pressure losses and isentropic Mach number distributions on the profile were compared to measurements without coolant ejection, showing a significant influence of the TEB both on the wake development and on the flow in the vane passage. Numerical simulations of the experiments showed good agreement with the measured data and provided a deeper understanding of the flow phenomena, revealing the differences in the development of the wake with and without trailing edge coolant ejection and illustrating the blockage effect of the TEB on the flow in the vane passage.

scholarly journals Thermal Fatigue Life Prediction of Thermal Barrier Coat on Nozzle Guide Vane via Master–Slave Model

2019 ◽  
Vol 9 (20) ◽  
pp. 4357 ◽  
Author(s):  
Peng Guan ◽  
Yanting Ai ◽  
Chengwei Fei ◽  
Yudong Yao
Keyword(s):  
Fatigue Life ◽  
Thermal Fatigue ◽  
Life Prediction ◽  
Guide Vane ◽  
Leading Edge ◽  
Fatigue Life Prediction ◽  
Trailing Edge ◽  
Thermal Fatigue Life ◽  
Barrier Coat ◽  
Nozzle Guide

The aim of this paper was to develop a master–slave model with fluid-thermo-structure (FTS) interaction for the thermal fatigue life prediction of a thermal barrier coat (TBC) in a nozzle guide vane (NGV). The master–slave model integrates the phenomenological life model, multilinear kinematic hardening model, fully coupling thermal-elastic element model, and volume element intersection mapping algorithm to improve the prediction precision and efficiency of thermal fatigue life. The simulation results based on the developed model were validated by temperature-sensitive paint (TSP) technology. It was demonstrated that the predicted temperature well catered for the TSP tests with a maximum error of less than 6%, and the maximum thermal life of TBC was 1558 cycles around the trailing edge, which is consistent with the spallation life cycle of the ceramic top coat at 1323 K. With the increase of pre-oxidation time, the life of TBC declined from 1892 cycles to 895 cycles for the leading edge, and 1558 cycles to 536 cycles for the trailing edge. The predicted life of the key points at the leading edge was longer by 17.7–40.1% than the trailing edge. The developed master–slave model was validated to be feasible and accurate in the thermal fatigue life prediction of TBC on NGV. The efforts of this study provide a framework for the thermal fatigue life prediction of NGV with TBC.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

Modeling of Combustor and Turbine Vane Interaction

2019 ◽  
Author(s):  
Ishan Verma ◽  
Laith Zori ◽  
Jaydeep Basani ◽  
Samir Rida
Keyword(s):  
Swirling Flow ◽  
Guide Vane ◽  
Primary Objective ◽  
Point Of View ◽  
Computationally Efficient ◽  
Joint Simulation ◽  
Chemically Reacting ◽  
Combined Simulation ◽  
Aero Engines ◽  
Nozzle Guide

Abstract Modern aero-engines are characterized by compact components (fan, compressor, combustor, and turbine). Such proximity creates a complex interaction between the components and poses a modeling challenge due to the difficulties in identifying a clear interface between components since they are usually modeled separately. From a numerical point of view, the simulation of a complex compact aero-engine system requires interaction between these individual components, especially the combustor-turbine interaction. The combustor is characterized by a subsonic chemically reacting and swirling flow while the high-pressure turbine (HPT) stage has flow which is transonic. Furthermore, the simulation of combustor-turbine interactions is more challenging due to aggressive flow conditions such as non-uniform temperature, non-uniform total-pressure, strong swirl, and high turbulence intensity. The simulation of aero-engines, where combustor-turbine interactions are important, requires a methodology that can be used in a real engine framework while ensuring numerical requirements of accuracy and stability. Conventionally, such a simulation is carried out using one of the two approaches: a combined simulation (or joint-simulation) of the combustor and the HPT geometries, or a co-simulation between the combustor and the turbine with the exchange of boundary conditions between these two separate domains. The primary objective of this paper is to assess the effectiveness of the joint simulation versus the co-simulation and propose a more practical approach for modeling combustor and turbine interactions. First, a detailed grid independence study with hexahedral and polyhedral meshes is performed to select the required polyhedral mesh. Then, an optimal location of the interface between the combustor and the nozzle guide vane (NGV) is identified. Co-simulations are then performed by exchanging information between the combustor and the NGV at the interface, wherein the combustor is solved using LES while the NGV is solved using RANS. The joint combustor-NGV simulations are solved using LES. The effect of the combustor-NGV interaction on the flow field and hot streak migration is analyzed. The results suggest that the joint simulation is computationally efficient and more accurate since both components are modelled together.

Application of Unsteady CFD Methods to Trailing Edge Cutback Film Cooling

2014 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Navier Stokes ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Unsteady Rans ◽  
Exit Mach Number ◽  
Nozzle Guide

The main purpose of this numerical investigation is to overcome the limitations of the steady modeling in predicting the cooling efficiency over the cutback surface in a high pressure turbine nozzle guide vane. Since discrepancy between Reynolds-averaged Navier–Stokes (RANS) predictions and measured thermal coverage at the trailing edge was attributable to unsteadiness, Unsteady RANS (URANS) modeling was implemented to evaluate improvements in simulating the mixing between the mainstream and the coolant exiting the cutback slot. With the aim of reducing the computation effort, only a portion of the airfoil along the span was simulated at an exit Mach number of Ma2is = 0.2. Three values of the coolant-to-mainstream mass flow ratio were considered: MFR = 0.66%, 1.05%, and 1.44%. Nevertheless the inherent vortex shedding from the cutback lip was somehow captured by the URANS method, the computed mixing was not enough to reproduce the measured drop in adiabatic effectiveness η along the streamwise direction, over the cutback surface. So modeling was taken a step further by using the Scale Adaptive Simulation (SAS) method at MFR = 1.05%. Results from the SAS approach were found to have potential to mimic the experimental measurements. Vortices shedding from the cutback lip were well predicted in shape and magnitude, but with a lower frequency, as compared to PIV data and flow visualizations. Moreover, the simulated reduction in film cooling effectiveness toward the trailing edge was similar to that observed experimentally.

Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

2012 ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Co2 Concentration ◽  
Pressure Side ◽  
Trailing Edge ◽  
Nozzle Guide

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a 5-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side of the vane trailing edge where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge does not reach the pressure side endwall, potentially creating a local hotspot.

scholarly journals Influences of Larger-Scale Vortex Variability on Tornado Pressure Minima in an Outer-Flow Region: Explorations Using a Parametric Tangential Wind Model

2017 ◽  
Vol 145 (5) ◽  
pp. 1597-1614
Author(s):  
Vincent T. Wood ◽  
Robin L. Tanamachi ◽  
Luther W. White
Keyword(s):  
Total Pressure ◽  
Swirling Flow ◽  
Flow Region ◽  
Radial Pressure ◽  
Central Pressure ◽  
Wind Model ◽  
Free Atmosphere ◽  
Outer Flow ◽  
Radial Profiles ◽  
Tangential Wind

Abstract Previous studies have neglected to distinguish between a central pressure deficit due to a tornado itself and due to a parent mesocyclone in which the tornado is embedded. To obtain improved understanding of the influences of larger-scale vortex variability on smaller-scale tornado pressure deficits, a parametric tangential wind model supplemented with a cyclostrophic speed equation was used to explore the role that the variability plays in influencing radial pressure deficits by deducing radial pressure deficit distributions from radial profiles of hypothetically superpositioned, dual-maxima tangential velocities in the free atmosphere, where a dominant swirling flow was in approximate cyclostrophic balance. The cyclostrophic approximation was partitioned into two separate components, allowing one to scrutinize and determine which of the concentric vortices contributes most significantly to the tornado pressure minima. The model parametrically constructed a smaller-scale, stronger vortex (rapidly swirling flow) that was centered within a larger-scale, weaker vortex (slowly swirling flow) to represent a tornado centered within a supercell, low-level, parent mesocyclone above a tornado boundary layer. The radial pressure deficit fluctuations were varied by changing one of five key velocity-controlling parameters assigned to one vortex to represent a variety of vortex strengths. Based on eight experiments, the larger-scale, weaker (smaller scale, stronger) vortex contributed less (more) to the total pressure deficit than the smaller-scale, stronger (larger scale, weaker) vortex. The stronger vortex centered within the larger-scale, weaker vortex has a larger central pressure minimum than it does in the absence of the larger-scale vortex.

The Influence of the Boundary Layer State and Reynolds Number on Film Cooling and Heat Transfer on a Cooled Nozzle Guide Vane

2000 ◽  
Author(s):  
Hans Reiss ◽  
Albin Bölcs
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Flow Conditions ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Film cooling and heat transfer measurements were carried out on a cooled nozzle guide vane in a linear cascade, using a transient liquid crystal technique. Three flow conditions were realized: the nominal operating condition of the vane with an exit Reynolds number of 1.47e6, as well as two lower flow conditions: Re2L = 1.0e6 and 7.5e5. The vane model was equipped with a single row of inclined round film cooling holes with compound angle orientation on the suction side. Blowing ratios ranging form 0.3 to 1.5 were covered, all using foreign gas injection (CO2) yielding an engine-representative density ratio of 1.6. Two distinct states of the incoming boundary layer onto the injection station were compared, an undisturbed laminar boundary layer as it forms naturally on the suction side, and a fully turbulent boundary layer which was triggered with a trip wire upstream of injection. The aerodynamic flow field is characterized in terms of profile Mach number distribution, and the associated heat transfer coefficients around the uncooled airfoil are presented. Both detailed and spanwise averaged results of film cooling effectiveness and heat transfer coefficients are shown on the suction side, which indicate considerable influence of the state of the incoming boundary layer on the performance of a film cooling row. The influence of the mainstream flow condition on the film cooling behavior at constant blowing ratio is discussed for three chosen injection regimes.

EXPERIMENTAL STUDY OF THE ENDWALL HEAT TRANSFER OF A TRANSONIC NOZZLE GUIDE VANE WITH UPSTREAM JET PURGE COOLING PART 1 - EFFECT OF DENSITY RATIO

2021 ◽  
pp. 1-36
Author(s):  
Shuo Mao ◽  
Ridge A. Sibold ◽  
Wing Ng ◽  
Zhigang LI ◽  
Bo Bai ◽  
...  
Keyword(s):  
Large Scale ◽  
Guide Vane ◽  
Flow Direction ◽  
Density Ratio ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Vital Role ◽  
Blowing Ratio ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract Nozzle guide vane platforms often employ complex cooling schemes to mitigate the ever-increasing thermal loads on endwall. This study analyzes, experimentally and numerically, and describes the effect of coolant to mainstream blowing ratio, momentum ratio and density ratio for a typical axisymmetric converging nozzle guide vane platform with an upstream doublet staggered, steep-injection, cylindrical hole purge cooling scheme. Nominal flow conditions were engine-representative and as follows: Maexit = 0.85, Reexit,Cax = 1.5×106 and an inlet large-scale freestream turbulence intensity of 16%. Two blowing ratios were investigated, each corresponding to the design condition and its upper extrema at M = 2.5 and 3.5, respectively. For each blowing ratio, the coolant to mainstream density ratio was varied between DR=1.2, representing typical experimental neglect of coolant density, and DR=1.95, representative of typical engine conditions. The results show that with a fixed coolant-to-mainstream blowing ratio, the density ratio plays a vital role in the coolant-mainstream mixing and the interaction between coolant and horseshoe vortex near the vane leading edge. A higher density ratio leads to a better coolant coverage immediately downstream of the cooling holes but exposes the in-passage endwall near the pressure side. It also causes the in-passage coolant coverage to decay at a higher rate in the flow direction. From the results gathered, both density ratio and blowing ratio should be considered for accurate testing, analysis, and prediction of purge jet cooling scheme performance.

Experimental Investigation of Innovative Cooling Schemes on an Additively Manufactured Engine Scale Turbine Nozzle Guide Vane

2020 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Jet Impingement ◽  
Development Effort ◽  
Suction Surface ◽  
Trailing Edge ◽  
Pin Fin ◽  
Nozzle Guide ◽  
Film Cooling Holes

Abstract This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

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