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.

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.

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.

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.

Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance

2014 ◽  
Vol 136 (11) ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten H. Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Trailing Edge ◽  
Aerodynamic Loss ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Engine Component ◽  
Annular Sector ◽  
Exit Mach Number ◽  
Nozzle Guide

An experimental investigation on a cooled nozzle guide vane (NGV) has been conducted in an annular sector to quantify aerodynamic influences of shower head (SH) and trailing edge (TE) cooling. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass–flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both TE cooling and SH film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the TE cooling has higher aerodynamic loss compared to the SH cooling. Secondary losses decrease with inserting SH film cooling compared to the uncooled case. The TE cooling appears to have less impact on the secondary loss compared to the SH cooling. Area-averaged exit flow angles around midspan increase for the TE cooling.

scholarly journals The Influence of Film-Cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane

1997 ◽  
Author(s):  
C. Osnaghi ◽  
A. Perdichizzi ◽  
M. Savini ◽  
P. Harasgama ◽  
E. Lutum
Keyword(s):  
Carbon Dioxide ◽  
Mass Flow ◽  
Film Cooling ◽  
Momentum Flux ◽  
Guide Vane ◽  
Flux Ratio ◽  
Aerodynamic Performance ◽  
Trailing Edge ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

The paper presents the results of an investigation on the aerodynamic performance of a full coverage film-cooled nozzle guide vane. The blading is a typical high pressure turbine vane of advanced design, working in the high subsonic regime. Tests have been carried out for a wide range of conditions, including variations in Mach number, coolant to mainstream mass flow rate ratio and location of the coolant injection. Both air and carbon dioxide at ambient conditions have been utilized, as coolant flow. Measurements have been performed in a plane located at 0.5 axial chord downstream of the trailing edge by means of a miniaturized five-hole pressure probe. Performances, in terms of losses, flow angles and profile pressure distributions, for different cooling mass flow rates are presented and compared to the results of the solid blade tests (i.e. with no cooling holes). The results showed a significant increase of the losses with blowing. Test with air and carbon dioxide provided almost equal losses if carried out at the same global momentum flux ratio; however the density ratio was found to influence slightly the share of the coolant fluid among the injection rows and the local momentum flux ratio as well. In order to define the individual contributions of groups of cooling rows on the performance of the blade, three different modes of injection have been tested, namely full, trailing edge and shower head injection. The main trend observed is that trailing edge injection produces the least amount of additional losses at high blowing rates. Full-coverage film-cooling injection did not lead to marked variations in the blade pressure distribution and/or outlet flow angle.

Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance

2014 ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten H. Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Trailing Edge ◽  
Aerodynamic Loss ◽  
Head Cooling ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Annular Sector ◽  
Exit Mach Number ◽  
Nozzle Guide

An experimental investigation on a cooled nozzle guide vane has been conducted in an annular sector to quantify aerodynamic influences of shower head and trailing edge cooling. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass-flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both trailing edge cooling and shower head film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the trailing edge cooling has higher aerodynamic loss compared to the shower head cooling. Secondary losses decrease with inserting shower head film cooling compared to the uncooled case. The trailing edge cooling appears to have less impact on the secondary loss compared to the shower head cooling. Area-averaged exit flow angles around midspan increase for the trailing edge cooling.

Experimental Evaluation of Service-Exposed Nozzle Guide Vane Damage in a Rolls Royce A-250 Gas Turbine

2013 ◽  
Author(s):  
D. Bouchard ◽  
A. Asghar ◽  
M. LaViolette ◽  
W. D. E. Allan ◽  
R. Woodason
Keyword(s):  
Performance Metrics ◽  
Guide Vane ◽  
Trailing Edge ◽  
Root Area ◽  
Trailing Edges ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Edge Damage ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

A unique methodology and test rig was designed to evaluate the degradation of damaged Nozzle Guide Vanes in a transonic annular cascade in the short duration facility at the Royal Military College. A custom test section was designed which featured a novel rotating instrumentation suite. This permitted 360° multi-span traverse measurements downstream of unmodified turbine NGV rings from a Rolls-Royce/Allison A-250 turbo-shaft engine. Downstream total pressure was measured at four span-wise locations on both an undamaged reference and a damaged test article. Three performance metrics were developed in an effort to determine characteristic signatures for common operational damage such as trailing edge bends or cracked trailing edges. The highest average losses were observed in the root area, while the lowest occurred closer to the NGV tips. The results from this study indicated that multiple span-wise traverses were required to detect localized trailing edge damage. Recommendations have been made for future tests, for test rigs and for ideas to develop performance metrics.

Experimental Evaluation of Service-Exposed Nozzle Guide Vane Damage in a Rolls Royce A-250 Gas Turbine

2014 ◽  
Vol 136 (10) ◽  
Author(s):  
D. Bouchard ◽  
A. Asghar ◽  
M. LaViolette ◽  
W. D. E. Allan ◽  
R. Woodason
Keyword(s):  
Performance Metrics ◽  
Guide Vane ◽  
Trailing Edge ◽  
Root Area ◽  
Trailing Edges ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Edge Damage ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

A unique methodology and test rig was designed to evaluate the degradation of damaged nozzle guide vanes (NGVs) in a transonic annular cascade in the short duration facility at the Royal Military College. A custom test section was designed which featured a novel rotating instrumentation suite. This permitted 360 deg multispan traverse measurements downstream from unmodified turbine NGV rings from a Rolls-Royce/Allison A-250 turbo-shaft engine. The downstream total pressure was measured at four spanwise locations on both an undamaged reference and a damaged test article. Three performance metrics were developed in an effort to determine characteristic signatures for common operational damage such as trailing edge bends or cracked trailing edges. The highest average losses were observed in the root area, while the lowest occurred closer to the NGV tips. The results from this study indicated that multiple spanwise traverses were required to detect localized trailing edge damage. Recommendations are made for future testing and to further develop performance metrics.

scholarly journals Clocking and Potential Effects in Combustor–Turbine Stator Interactions

Aerospace ◽  
2021 ◽  
Vol 8 (10) ◽  
pp. 285
Author(s):  
Pawel Flaszynski ◽  
Michal Piotrowicz ◽  
Tommaso Bacci
Keyword(s):  
Guide Vane ◽  
Turbulence Models ◽  
Flow Distribution ◽  
Aircraft Engine ◽  
Potential Effect ◽  
Flow Conditions ◽  
Azimuthal Position ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Stress Models

Investigations of combustors and turbines separately have been carried out for years by research institutes and aircraft engine companies, but there are still many questions about the interaction effect. In this paper, a prediction of a turbine stator’s potential effect on flow in a combustor and the clocking effect on temperature distribution in a nozzle guide vane are discussed. Numerical simulation results for the combustor simulator and the nozzle guide vane (NGV) of the first turbine stage are presented. The geometry and flow conditions were defined according to measurements carried out on a test section within the framework of the EU FACTOR (full aerothermal combustor–turbine interactions research) project. The numerical model was validated by a comparison of results against experimental data in the plane at a combustor outlet. Two turbulence models were employed: the Spalart–Allmaras and Explicit Algebraic Reynolds Stress models. It was shown that the NGV potential effect on flow distribution at the combustor–turbine interface located at 42.5% of the axial chord is weak. The clocking effect due to the azimuthal position of guide vanes downstream of the swirlers strongly affects the temperature and flow conditions in a stator cascade.

Influence of Upstream Flow Conditions on the Heat Transfer to Nozzle Guide Vanes

1988 ◽  
Vol 110 (3) ◽  
pp. 412-416 ◽  
Author(s):  
V. Krishnamoorthy ◽  
B. R. Pai ◽  
S. P. Sukhatme
Keyword(s):  
Guide Vane ◽  
Gas Temperature ◽  
Flow Conditions ◽  
Cold Flow ◽  
Hot Gas ◽  
Nozzle Guide Vane ◽  
High Turbulence ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Upstream Flow

The influence of a combustor located just upstream of a nozzle guide vane cascade on the heat flux distribution to the nozzle guide vane was experimentally investigated. The surface temperature distribution around the convectively cooled vane of the cascade was obtained by locating the cascade, firstly in a low-turbulence uniform hot gas stream, secondly in a high-turbulence, uniform hot gas stream, and thirdly in a high-turbulence, nonuniform hot gas stream present just downstream of the combustor exit. The results indicate that the increased blade surface temperatures observed for the cascade placed just downstream of the combustor can be accounted for by the prevailing turbulence level measured at cascade inlet in cold-flow conditions and the average gas temperature at the cascade inlet.

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