A Large Scale Turbine Test Rig for the Investigation of High Pressure Turbine Aerodynamics and Heat Transfer With Variable Inflow Conditions

2015 ◽  
Author(s):  
Alexander Krichbaum ◽  
Holger Werschnik ◽  
Manuel Wilhelm ◽  
Heinz-Peter Schiffer ◽  
Knut Lehmann
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Large Scale ◽  
Measurement Techniques ◽  
Air Injection ◽  
Test Rig ◽  
High Pressure Turbine ◽  
Film Cooling Effectiveness ◽  
Cooling Air ◽  
Inflow Conditions

Focusing on the experimental analysis of the effect of variable inlet flows on aerodynamics, efficiency and heat transfer of a modern high pressure turbine, the Large Scale Turbine Rig (LSTR) at Technische Universität Darmstadt has been extensively redesigned. The LSTR is a full annular, rotating low speed turbine test rig carrying a scaled 1.5-stage (NGV1 - Rotor - NGV2) axial high-pressure turbine geometry designed by Rolls-Royce Deutschland to match engine-realistic Reynolds numbers. To simulate real turbine inflow conditions, the LSTR is equipped with a combustor simulator module including exchangeable swirlers. Other inflow conditions include axial or turbulent inflow as well as altered relative positions of swirl cores and NGVs by traversing. To investigate combustor-turbine interaction, the LSTR offers a large variety of optical and physical access ports as well as high flexibility to the application of measurement techniques. An elaborate secondary air system enables the simulation of various cooling air flows. The turbine section is equipped with film-cooled NGVs, a hub side seal air injection between NGVs and rotor, as well as a hub side RIDN cooling air injection module designed to provide realistic turbine flow conditions. Exchangeable hub side RIDN-plates allow for investigation of different coolant injection geometries. Measurement capabilities include 5-hole-probes, Pitot and total temperature rakes, as well as static pressure taps distributed along NGV radial sections and at the hub side passage endwall. The NGV passage flow can be visualized by means of Particle Image Velocimetry (PIV). Hot Wire Anemometry (HWA) will be used for time-resolved measurements of the turbulence level at several positions. The distributions of heat transfer and film cooling effectiveness are acquired using infrared thermography and CO2-gas tracing.

Heat Transfer and Aerodynamics of Over-Shroud Leakage Flows in a High-Pressure Turbine

2009 ◽  
Author(s):  
Knut Lehmann ◽  
Richard Thomas ◽  
Howard Hodson ◽  
Vassilis Stefanis
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Large Scale ◽  
Suction Side ◽  
Surface Flow ◽  
Tip Clearance ◽  
Flow Visualisation ◽  
Leakage Flow ◽  
High Pressure Turbine ◽  
High Heat Transfer

An experimental study has been conducted to investigate the distribution of the convective heat transfer on the shroud of a high pressure turbine blade in a large scale rotating rig. A continuous thin heater foil technique has been adapted and implemented on the turbine shroud. Thermochromic Liquid Crystals were employed for the surface temperature measurements to derive the experimental heat transfer data. The heat transfer is presented on the shroud top surfaces and the three fins. The experiments were conducted for a variety of Reynolds numbers and flow coefficients. The effects of different inter-shroud gap sizes and reduced fin tip clearance gaps were also investigated. Details of the shroud flow field were obtained using an advanced Ammonia-Diazo surface flow visualisation technique. CFD predictions are compared with the experimental data and used to aid interpretation. Contour maps of the Nusselt number reveal that regions of highest heat transfer are mostly confined to the suction side of the shroud. Peak values exceed the average by as much as 100 percent. It has been found that the interaction between leakage flow through the inter-shroud gaps and the fin tip leakage jets are responsible for this high heat transfer. The inter-shroud gap leakage flow causes a disruption of the boundary layer on the turbine shroud. Furthermore, the development of the large recirculating shroud cavity vortices is severely altered by this leakage flow.

The Influence of Purge Flow Parameters on Heat Transfer and Film Cooling in Turbine Center Frames

2021 ◽  
Author(s):  
Patrick R. Jagerhofer ◽  
Marios Patinios ◽  
Tobias Glasenapp ◽  
Emil Göttlich ◽  
Federica Farisco
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Convective Heat Transfer Coefficient ◽  
Density Ratio ◽  
Constant Heat Flux ◽  
High Pressure Turbine ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow

Abstract Due to stringent environmental legislation and increasing fuel costs, the efficiencies of modern turbofan engines have to be further improved. Commonly, this is facilitated by increasing the turbine inlet temperatures in excess of the melting point of the turbine components. This trend has reached a point where not only the high-pressure turbine has to be adequately cooled, but also components further downstream in the engine. Such a component is the turbine center frame (TCF), having a complex aerodynamic flow field that is also highly influenced by purge-mainstream interactions. The purge air, being injected through the wheelspace cavities of the upstream high-pressure turbine, bears a significant cooling potential for the TCF. Despite this, fundamental knowledge of the influencing parameters on heat transfer and film cooling in the TCF is still missing. This paper examines the influence of purge-to-mainstream blowing ratio, purge-to-mainstream density ratio and purge flow swirl angle on the convective heat transfer coefficient and the film cooling effectiveness in the TCF. The experiments are conducted in a sector-cascade test rig specifically designed for such heat transfer studies using infrared thermography and tailor-made flexible heating foils with constant heat flux. The inlet flow is characterized by radially traversing a five-hole-probe. Three purge-to-mainstream blowing ratios and an additional no purge case are investigated. The purge flow is injected without swirl and also with engine-similar swirl angles. The purge swirl and blowing ratio significantly impact the magnitude and the spread of film cooling in the TCF. Increasing blowing ratios lead to an intensification of heat transfer. By cooling the purge flow, a moderate variation in purge-to-mainstream density ratio is investigated, and the influence is found to be negligible.

Endwall Effusion Cooling System Behaviour Within a High-Pressure Turbine Cascade: Part 2—Heat Transfer and Effectiveness Measurements

2010 ◽  
Author(s):  
B. Facchini ◽  
L. Tarchi ◽  
L. Toni ◽  
S. Zecchi
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Large Scale ◽  
Cooling System ◽  
Main Flow ◽  
Turbine Cascade ◽  
High Pressure Turbine ◽  
Cooling Flow ◽  
Micro Holes ◽  
Element Approach

The cooling performance of a micro-holed endwall of a large-scale high pressure turbine cascade has been investigated within the European Project AITEB-2. The experimental investigation has been performed for a baseline configuration, with a smooth solid endwall and with a micro-holed endwall providing micro-jets ejection from the wall. A micro-holed endwall made of two modules was adopted in order to reduce the compound angle between the main flow and the micro jets axes. The micro-holed endwall is provided with a total amount of 3294 micro-holes with a diameter of 0.1 per cent of the blade chord. Four different cooling flow rates, from 1.2% to 2.6% of the main flow mass flow rate respectively, were investigated and the experimental results are reported in the paper. Both adiabatic effectiveness and heat transfer coefficient have been measured employing a steady state technique with Thermo-chromic Liquid Crystals (TLC). A thin stainless steel heating foil was used to generate the surface heat flux for the HTC measurements and a data reduction procedure based on a Finite Element approach has been developed to take into account the non uniform heat generation along the endwall.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer, and film cooling effectiveness of advanced high pressure turbine blade tips and endwalls. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with leading edge and trailing edge cutouts. Both blade tip configurations have pressure side film cooling and cooling air extraction through dust holes, which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9×105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aerothermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the midchord region. However, on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2010 ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer and film cooling effectiveness of advanced high-pressure turbine blade tips and endwall. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with a leading edge and trailing edge cut-out. Both blade tip configurations have pressure side film cooling, and cooling air extraction through dust holes which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9 × 105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aero-thermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although, the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the mid-chord region. However on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

scholarly journals Transient Aero-Thermo-Mechanical Multidimensional Analysis of a High Pressure Turbine Assembly Through a Square Cycle

2021 ◽  
Author(s):  
Vladislav Ganine ◽  
John W. Chew ◽  
Nicholas Hills ◽  
Sulfi Noor Mohamed ◽  
Matthew Miller
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Network Code ◽  
Multidimensional Analysis ◽  
High Pressure Turbine ◽  
Time Step ◽  
Local Flow ◽  
Area Of Interest ◽  
Engine Components ◽  
Cooling Air

Abstract Better understanding and more accurate prediction of heat transfer and cooling flows in aero engine components in steady and transient operating regimes are essential to modern engine designs aiming at reduced cooling air consumption and improved engine efficiencies. This paper presents a simplified coupled transient analysis methodology that allows assessment of the aerothermal and thermomechanical responses of engine components together with cooling air mass flow, pressure and temperature distributions in an automatic fully integrated way. This is achieved by assembling a fluid network with contribution of components of different geometrical dimensions coupled to each other through dimensionally heterogeneous interfaces. More accurate local flow conditions, heat transfer and structural displacement are resolved on a smaller area of interest with multidimensional surface coupled CFD/FE codes. Contributions of the whole engine air-system are predicted with a faster mono dimensional flow network code. Matching conditions at the common interfaces are enforced at each time step exactly by employing an efficient iterative scheme. The coupled simulation is performed on an industrial high pressure turbine disk component run through a square cycle. Predictions are compared against the available experimental data. The paper proves the reliability and performance of the multidimensional coupling technique in a realistic industrial setting. The results underline the importance of including more physical details into transient thermal modelling of turbine engine components.

scholarly journals Transient Aero-Thermo-Mechanical Multidimensional Analysis of a High Pressure Turbine Assembly Through a Square Cycle

2020 ◽  
Author(s):  
Vlad Ganine ◽  
John W. Chew ◽  
Nicholas J. Hills ◽  
Sulfi N. Mohamed ◽  
Matthew M. Miller
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Network Code ◽  
Multidimensional Analysis ◽  
High Pressure Turbine ◽  
Time Step ◽  
Local Flow ◽  
Area Of Interest ◽  
Engine Components ◽  
Cooling Air

Abstract Better understanding and more accurate prediction of heat transfer and cooling flows in aero engine components in steady and transient operating regimes are essential to modern engine designs aiming at reduced cooling air consumption and improved engine efficiencies. This paper presents a simplified coupled transient analysis methodology that allows assessment of the aerothermal and thermomechanical responses of engine components together with cooling air mass flow, pressure and temperature distributions in an automatic fully integrated way. This is achieved by assembling a fluid network with contribution of components of different geometrical dimensions coupled to each other through dimensionally heterogeneous interfaces. More accurate local flow conditions, heat transfer and structural displacement are resolved on a smaller area of interest with multidimensional surface coupled CFD/FE codes. Contributions of the whole engine air-system are predicted with a faster mono dimensional flow network code. Matching conditions at the common interfaces are enforced at each time step exactly by employing an efficient iterative scheme. The coupled simulation is performed on an industrial high pressure turbine disk component run through a square cycle. Predictions are compared against the available experimental data. The paper proves the reliability and performance of the multidimensional coupling technique in a realistic industrial setting. The results underline the importance of including more physical details into transient thermal modelling of turbine engine components.

Large-Scale Simulation of the Clocking Impact of 2D Combustor Profile on a Two Stage High Pressure Turbine

2014 ◽  
Author(s):  
Thomas E. Dyson ◽  
David B. Helmer ◽  
James A. Tallman
Keyword(s):  
High Pressure ◽  
Large Scale ◽  
High Pressure Turbine ◽  
Cfd Simulations ◽  
Two Stage ◽  
Substantial Impact ◽  
Sliding Mesh ◽  
Wall Flow ◽  
End Wall ◽  
Film Flows

This paper presents sliding-mesh unsteady CFD simulations of high-pressure turbine sections of a modern aviation engine in an extension of previously presented work [1]. The simulation included both the first and second stages of a two-stage high-pressure turbine. Half-wheel domains were used, with source terms representing purge and film flows. The end-wall flow-path cavities were incorporated in the domain to a limited extent. The passage-to-passage variation in thermal predictions was compared for a 1D and 2D turbine inlet boundary condition. Substantial impact was observed on both first and second stage vanes despite the mixing from the first stage blade. Qualitative and quantitative differences in surface temperature distributions were observed due to different ratios between airfoil counts in the two domains.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

2020 ◽  
Author(s):  
Joao Vieira ◽  
John Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Reynolds Numbers ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

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