A Computational Validation of Turbine Nozzle Guide Vane Aerodynamic Experiments in an HP Turbine Stage

2011 ◽  
Author(s):  
O¨zhan H. Turgut ◽  
Cengiz Camcı
Keyword(s):  
Validation Study ◽  
Total Pressure ◽  
Guide Vane ◽  
Pressure Coefficient ◽  
Axial Flow ◽  
Turbine Stage ◽  
Aerodynamic Loss ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Computational Validation

A computational validation study related to aerodynamic loss generation mechanisms is performed in an axial flow turbine nozzle guide vane (NGV). The 91.66 cm diameter axial flow turbine research facility has a stationary nozzle guide vane assembly and a 29 bladed HP turbine rotor. The NGV inlet and exit Reynolds numbers based on midspan axial chord are around 300000 and 900000, respectively. The effect of grid structure on aerodynamic loss generation is investigated. GAMBIT and TGRID combination is used for unstructured grid, whereas GRIDPRO is the structured grid generator. For both cases, y+ values are kept below unity. The finite-volume flow solver ANSYS CFX with SST k–ω turbulence model is employed. Experimental flow conditions are imposed at the boundaries. The flow transition effect and the influence of corner fillets at the vane-endwall junction are also studied in this paper. Grid independence study is performed with static pressure coefficient distribution at the mid-span of the vane and the total pressure coefficient at the NGV exit. The velocity distributions and the total pressure coefficient at the NGV exit plane are in very good agreement with the experimental data. This validation study shows that the effect of future geometrical modifications on the endwalls and the vane will be predicted reasonably accurately. The current study shows that an accurately measured turbine stage geometry, a properly prepared block structured/body fitted grid, a state of the art transitional flow implementation, and realistic boundary conditions coming from high resolution turbine experiments are all essential ingredients of a successful NGV aerodynamic loss quantification via computations.

Factors Influencing Computational Predictability of Aerodynamic Losses in a Turbine Nozzle Guide Vane Flow

2016 ◽  
Vol 138 (5) ◽  
Author(s):  
Özhan H. Turgut ◽  
Cengiz Camci
Keyword(s):  
Validation Study ◽  
Total Pressure ◽  
Guide Vane ◽  
Pressure Coefficient ◽  
Transitional Flow ◽  
Turbine Stage ◽  
Aerodynamic Loss ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Aerodynamic Losses

This paper deals with the computational predictability of aerodynamic losses in a turbine nozzle guide vane (NGV) flow. The paper shows that three-dimensional (3D) computations of Reynolds-Averaged Navier Stokes (RANS) equations have the ability to adequately represent viscous losses in the presence of laminar flows, transitional regions, and fully turbulent flow areas in the NGV of an high pressure (HP) turbine stage. The Axial Flow Turbine Research Facility (AFTRF) used for the present experimental results has an annular NGV assembly and a 29-bladed HP turbine rotor spinning at 1330 rpm. The NGV inlet and exit Reynolds numbers based on midspan axial chord are around 300,000 and 900,000, respectively. A general purpose finite-volume 3D flow solver with a shear stress transport (SST) k–ω turbulence model is employed. The current computational study benefits from these carefully executed aerodynamic experiments in the NGV of the AFTRF. The grid independence study is performed with static pressure coefficient distribution at the midspan of the vane and the total pressure coefficient at the NGV exit. The effect of grid structure on aerodynamic loss generation is emphasized. The flow transition effect and the influence of corner fillets at the vane–endwall junction are also studied. The velocity distributions and the total pressure coefficient at the NGV exit plane are in very good agreement with the experimental data. This validation study shows that the effect of future geometrical modifications on the turbine endwall surfaces will be predicted reasonably accurately. The current study also indicates that an accurately defined turbine stage geometry, a properly prepared block-structured/body-fitted grid, a state-of-the-art transitional flow implementation, inclusion of fillets, and realistic boundary conditions coming from high-resolution turbine experiments are all essential ingredients of a successful turbine NGV aerodynamic loss quantification via computations. This validation study forms the basis for the successful future generation of nonaxisymmetric endwall surface modifications in AFTRF research efforts.

Computational Validation of the Flow Through a Turbine Stage and the Effects of Rim Seal Cavity Leakage on Secondary Flows

2012 ◽  
Author(s):  
Özhan H. Turgut ◽  
Cengiz Camcı
Keyword(s):  
Validation Study ◽  
Guide Vane ◽  
Pressure Coefficient ◽  
Axial Flow ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Transitional Flow ◽  
Interface Plane ◽  
Exit Plane ◽  
Computational Validation

A computational validation study related to aerodynamic loss generation mechanisms is performed in an axial flow turbine. The 91.66 cm diameter axial flow turbine research facility has a stationary nozzle guide vane assembly and a 29 bladed HP turbine rotor. The NGV inlet and exit Reynolds numbers based on midspan axial chord are around 300000 and 900000, respectively. GRIDPRO is used as the structured grid generator. y+ values are kept below unity. The finite-volume flow solver ANSYS CFX with SST k–ω turbulence model together with the transitional flow model is employed. Experimental flow conditions are imposed at the boundaries. The computational predictions are compared to experimental data at NGV exit plane and rotor inlet plane. NGV exit plane measurements come from a previous experimental study with a five-hole probe and the data at rotor inlet plane is taken by the current authors using a Kiel probe with 3.175mm head diameter. The comparison of rotor-stator interface models shows that the stage model, which calculates the circumferentially averaged fluxes and uses as the boundary condition at the interface plane, agrees well with the experimental total pressure coefficient data at the NGV exit. The difference between the NGV only simulation and the rotor-stator simulation is emphasized. The effect of rim seal flow on the mainstream aerodynamics is investigated. This validation study shows that the effect of future geometrical modifications on the endwalls and the vane will be predicted reasonably accurately.

Influence of Leading Edge Fillet and Nonaxisymmetric Contoured Endwall on Turbine NGV Exit Flow Structure and Interactions With the Rim Seal Flow

2013 ◽  
Author(s):  
Özhan H. Turgut ◽  
Cengiz Camcı
Keyword(s):  
Flow Structure ◽  
Total Pressure ◽  
Guide Vane ◽  
Axial Flow ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Rotor Blades ◽  
Constant Thickness ◽  
Computational Evaluation ◽  
Nozzle Guide

Three different ways are employed in the present paper to reduce the secondary flow related total pressure loss. These are nonaxisymmetric endwall contouring, leading edge (LE) fillet, and the combination of these two approaches. Experimental investigation and computational simulations are applied for the performance assessments. The experiments are carried out in the Axial Flow Turbine Research Facility (AFTRF) having a diameter of 91.66cm. The NGV exit flow structure was examined under the influence of a 29 bladed high pressure turbine rotor assembly operating at 1300 rpm. For the experimental measurement comparison, a reference Flat Insert endwall is installed in the nozzle guide vane (NGV) passage. It has a constant thickness with a cylindrical surface and is manufactured by a stereolithography (SLA) method. Four different LE fillets are designed, and they are attached to both cylindrical Flat Insert and the contoured endwall. Total pressure measurements are taken at rotor inlet plane with Kiel probe. The probe traversing is completed with one vane pitch and from 8% to 38% span. For one of the designs, area averaged loss is reduced by 15.06%. The simulation estimated this reduction as 7.11%. Computational evaluation is performed with the rotating domain and the rim seal flow between the NGV and the rotor blades. The most effective design reduced the mass averaged loss by 1.28% over the whole passage at the NGV exit.

Some Aspects of Wake-Wake Interactions Regarding Turbine Stator Clocking

2000 ◽  
Vol 123 (3) ◽  
pp. 526-533 ◽  
Author(s):  
Maik Tiedemann ◽  
Friedrich Kost
Keyword(s):  
Total Pressure ◽  
Guide Vane ◽  
Pressure Fluctuations ◽  
Fast Response ◽  
Turbine Stage ◽  
Flow Angle ◽  
Wake Interactions ◽  
Turbulence Data ◽  
Number Flow ◽  
Nozzle Guide

This investigation is aimed at an experimental determination of the unsteady flowfield downstream of a transonic high pressure turbine stage. The single stage measurements, which were part of a joined European project, were conducted in the windtunnel for rotating cascades of the DLR Go¨ttingen. Laser-2-focus (L2F) measurements were carried out in order to determine the Mach number, flow angle, and turbulence distributions. Furthermore, a fast response pitot probe was utilized to determine the total pressure distribution. The measurement position for both systems was 0.5 axial rotor chord downstream of the rotor trailing edge at midspan. While the measurement position remained fixed, the nozzle guide vane (NGV) was “clocked” to 12 positions covering one NGV pitch. The periodic fluctuations of the total pressure downstream of the turbine stage indicate that the NGV wake damps the total pressure fluctuations caused by the rotor wakes. Furthermore, the random fluctuations are significantly lower in the NGV wake affected region. Similar conclusions were drawn from the L2F turbulence data. Since the location of the interaction between NGV wake and rotor wake is determined by the NGV position, the described effects are potential causes for the benefits of “stator clocking” which have been observed by many researchers.

Aerothermal Performance of Shielded Vane Design

2017 ◽  
Vol 139 (11) ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Total Pressure ◽  
High Speed ◽  
Guide Vane ◽  
Leading Edge ◽  
Experimental Facility ◽  
Numerical Studies ◽  
Nozzle Guide Vane ◽  
Inlet Conditions ◽  
Nozzle Guide

This paper presents an experimental investigation of the concept of using the combustor transition duct wall to shield the nozzle guide vane leading edge. The new vane is tested in a high-speed experimental facility, demonstrating the improved aerodynamic and thermal performance of the shielded vane. The new design is shown to have a lower average total pressure loss than the original vane, and the heat transfer on the vane surface is overall reduced. The peak heat transfer on the vane leading edge–endwall junction is moved further upstream, to a region that can be effectively cooled as shown in previously published numerical studies. Experimental results under engine-representative inlet conditions showed that the better performance of the shielded vane is maintained under a variety of inlet conditions.

Influence of the Gas-to-Wall Temperature Ratio on the Boundary Layer Transition: Investigation of the Wake Behind a Turbine Nozzle Guide Vane

2019 ◽  
Author(s):  
Pietro Formisano ◽  
Tânia S. Cação Ferreira ◽  
Tony Arts
Keyword(s):  
Wall Temperature ◽  
Total Pressure ◽  
Guide Vane ◽  
Boundary Layer Transition ◽  
Upstream Region ◽  
Bypass Transition ◽  
Temperature Ratio ◽  
Layer Transition ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract Previous investigations performed at the von Karman Institute for Fluid Dynamics (VKI) have shown an influence of the gas-to-wall temperature ratio on the bypass transition development along the VKI LS89 blade suction side. In the present work, the influence of this quantity on the flow field downstream of this highly-loaded nozzle guide vane is studied through the evaluation of the aerodynamic losses. The investigation is organized in three sections with different combinations of exit Mach numbers and freestream turbulence intensity (FSTI) while Tgas/Twall is varied between 1.1 and 1.3 for all the tests. The Isentropic Compression Tube facility (CT-2) at VKI allowed the determination of the total pressure loss across the cascade by means of a Pitot tube in the upstream region and a downstream three-hole needle probe. The latter is traversed in the pitch-wise direction by a pneumatic traversing system. Finally, the cascade aerodynamic efficiency is quantified by means of the kinetic energy loss coefficient ζ and the total pressure drop profile distortions in the wake region.

Mainstream Aerodynamic Effects due to Wheelspace Coolant Injection in a High-Pressure Turbine Stage: Part II — Aerodynamic Measurements in the Rotational Frame

2001 ◽  
Author(s):  
Christopher McLean ◽  
Cengiz Camci ◽  
Boris Glezer
Keyword(s):  
High Pressure ◽  
Guide Vane ◽  
Pressure Coefficient ◽  
Three Dimensional ◽  
Axial Flow ◽  
Main Stream ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Coolant Injection ◽  
Cooling Air

The current paper deals with the aerodynamic measurements in the rotational frame of reference of the Axial Flow Turbine Research Facility (AFTRF) at the Pennsylvania State University. Stationary frame measurements of “Mainstream Aerodynamic Effects Due to Wheelspace Coolant Injection in a High Pressure Turbine Stage” were presented in part-I of this paper. The relative aerodynamic effects associated with rotor – nozzle guide vane (NGV) gap coolant injections were investigated in the rotating frame. Three-dimensional velocity vectors including exit flow angles were measured at the rotor exit. This study quantifies the secondary effects of the coolant injection on the aerodynamic and performance character of the stage main stream flow for root injection, radial cooling and impingement cooling. Current measurements show that even a small quantity (1%) of cooling air can have significant effects on the performance and exit conditions of the high pressure turbine stage. Parameters such as the total pressure coefficient, wake width, and three-dimensional velocity field show significant local changes. It is clear that the cooling air disturbs the inlet end-wall boundary layer to the rotor and modifies secondary flow development thereby resulting in large changes in turbine exit conditions. Effects are the strongest from the hub to midspan. Negligible effect of the cooling flow can be seen in the tip region.

Multi-Fidelity Modelling of a Fully-Featured HP Turbine Stage

2017 ◽  
Author(s):  
Giorgio Occhioni ◽  
Shahrokh Shahpar ◽  
Haidong Li
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
State Of The Art ◽  
Guide Vane ◽  
Phase Lag ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Film Cooling Holes

An improvement in overall efficiency and power output for gas turbine engines can be obtained by increasing the combustor exit temperature, but the thermal management of metal parts exposed to hot gases is challenging. Discrete film cooling, combined with internal convective cooling is the current state-of-the-art available to aerothermal designers of these components. To simplify the simulation problem in the aerodynamic design phase, it is common practice to replace the cooling holes with source strips applied to the blade. This could lead to inaccuracies in high pressure turbine performance prediction. This study has been carried out on a fully-featured high pressure turbine stage using high-fidelity simulations. The film cooling holes on the nozzle guide vane and on the rotor are initially modelled using a strip model approach. Then, to increase the model fidelity, the strips on the suction side of the rotor are replaced with discrete fan shaped film cooling holes. A rigid body rotation is also applied to the nozzle guide vane to vary the stage capacity and reaction. The effects of the mesh topology & resolution are also taken into account. The results obtained with these two approaches are then compared, giving the designers a better understanding on film cooling modelling and relationship between capacity, reaction and performance. The accurate prediction of the complex interaction between cavity inflows and the main-flow, still represent a challenge for the state of the art RANS solvers. Hence, an unsteady phase-lag approach has been used to overcome the RANS limitations. A validation of the unsteady solutions has been carried out with respect to experimental data.

scholarly journals Efficiency Measurements of an Annular Nozzle Guide Vane Cascade With Different Film Cooling Geometries

1998 ◽  
Author(s):  
Charles R. B. Day ◽  
Martin L. G. Oldfield ◽  
Gary D. Lock ◽  
Stephen N. Dancer
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Guide Vane ◽  
Cumulative Effect ◽  
Aerodynamic Efficiency ◽  
Nozzle Guide Vane ◽  
Density Ratios ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

This paper further extends the research reported by Day et al. (1997), which reported aerodynamic efficiency measurements on an annular cascade of engine representative transonic nozzle guide vanes with extensive film cooling. This work compares the measured aerodynamic efficiencies of blades with 14 rows of cylindrical cooling holes with a new geometry in which 8 of the rows have been replaced by holes having a fan-shaped exit geometry. The effects of adding trailing edge slot ejection are also presented. By selectively blocking rows of holes, the cumulative effect on the mid-span efficiency of adding rows of cooling holes has also been determined. A dense foreign gas (SF6/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios, momentum ratios and blowing rates under ambient temperature conditions. The flowfield measurements have been obtained using a four-hole pyramid probe in a short duration blowdown facility which correctly models engine Reynolds and Mach numbers, as well as the inlet turbulence intensity. Experimental results are presented as area traverse maps (total pressure, isentropic Mach number and flow angles), from which the incremental changes in efficiency due to film cooling have been calculated. The effects of different assumptions for the coolant total pressure are shown. Experimental data agrees reasonably well with loss predictions using a Hartsel model.

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.

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