bluuebunny/crossref_metadata_2025_split huggingface dataset:
Numerical Benchmark of Non-Conventional RANS Turbulence Models for Film and Effusion Cooling
Authors: Andrei L. , Andreini A. , Bianchini C. , Facchini B. 
Published: August 2013

In the course of the years several turbulence models specifically developed to improve the predicting capabilities of conventional two-equations RANS models have been proposed. However they have been mainly tested against experiments only comparing with standard isotropic models, in single hole configuration and for very low blowing ratio. A systematic benchmark of the various non-conventional models exploring a wider range of application is hence missing.
This paper performs a comparison of 3 recently proposed models over three different test cases of increasing computational complexity. The chosen test matrix covers a wide range of blowing ratios (0.5–3.0)including both single row and multi-row cases for which experimental data of reference are available. In particular the well known test by Sinha and Bogard [1] at BR = 0.5 is used in conjuction with two in-house carried out experiments: a single row film-cooling test at BR = 1.5 and a 15 rows test plate designed to study the interaction between slot and effusion cooling at BR = 3.0.
The first two considered models are based on a tensorial definition of the eddy viscosity in which the stream-span position is augmented to overcome the main drawback connected with standard isotropic turbulence models that is the lower lateral spreading of the jet downwards the injection. An anisotropic factor to multiply the off-diagonal position is indeed calculated from an algebraic expression of the turbulent Reynolds number developed by Bergeles [2] from DNS statistics over a flat plate. This correction could be potentially implemented in the framework of any eddy viscosity model. It was chosen to compare the predictions of such modification applied to two among the most common two-equation turbulence models for film-cooling tests, namely the Two-Layer (TL) model and the k–ω Shear Stress Transport (SST), firstly proposed and tested in the past respectively by Azzi and Lakeal [3] and Cottin at al. [4].
The third model, proposed by Holloway et al. [5], involves the unsteady solution of the flow and thermal field to include the short-time response of the stress tensor to rapid strain rates. This model takes advantage of the solution of an additional transport equation for the local effective total stress to trace the strain rate history.
The results are presented in terms of adiabatic effectiveness distribution over the plate as well as spanwise averaged profiles.
Emergence of Kelvin-Helmholtz Instabilities in Gas Turbine Rim Cavities: A Parameter Study
Authors: Rabs M. , Benra F.-K. , Schneider O. 
Published: August 2013

In an earlier paper of the authors, the occurrence of the so called Kelvin-Helmholtz instabilities (KHI) near the rim cavity of a 1.5 stage gas turbine has been examined by the use of CFD methods. It is shown that the KHI’s occur, when the swirl component of the hot gas flow is very strong. Due to the fact, that a high swirl is produced by the guide vanes of the first stage, this matter concerns most common gas turbines. A further paper validated the CFD methods used and derived KHI parameters (vortex appearance, vortex periodicity and vortex velocity) of a splitter plate model.
In the current study, essential parameters revealed by the analysis of a gas turbine rim cavity model are compared to the parameters extracted from the investigation of the splitter plate model and the potential linear theory of Turner. The rim cavity model is derived from a test rig of a 1.5 stage gas turbine. The blades and vanes have been removed from the computations. As main flow boundary conditions, surface averaged parameters are used. It is shown that a description of KHI developing in a rim cavity model is partly possible using splitter plate KHI characteristics and the potential linear theory of Turner as well.
A mathematical approach is formulated, which can predict the vortex velocity of KHI’s in turbine rim cavities.
APU Exhaust Muffler Design Improvements Through Conjugate Heat Transfer CFD Analysis
Authors: Vunnam Kiran , Bouldin Bruce 
Published: August 2013

An Auxiliary Power Unit (APU) is an additional gas turbine engine located in the tail cone section of an aircraft which can be operated while the aircraft is on the ground or in flight. It is used to generate electricity for the aircraft’s electrical systems and to provide air to the environmental control units (ECU’s) when the main engines are not operating or there is a desire to unload the main engines, such as in an engine out situation. The APU is also used for main engine starting. An APU typically has an exhaust system that vents out of the rear of the tail cone. When the APU is in operation, the exhaust emits a very loud noise which, if not muffled, could be an irritant to the members of the ground crew.
To reduce the impact of this exhaust noise, the APU is commonly fitted with a muffler. The muffler is placed around the APU’s exhaust pipe so that all of the APU’s exhaust is channeled through the muffler. The muffler is designed and constructed to substantially reduce the intensity of the noise emitted by the APU exhaust. The muffler is made of metal and has a tendency to get very hot during operation of the APU because of the high temperature of the exhaust gasses generated by the APU. It has been observed that the temperatures on the outer skin of the muffler commonly reach above 1,000 degrees Fahrenheit. If the muffler is not insulated, this heat will radiate outward from the muffler to the tail cone.
Modern aircraft tail cones are commonly made from composite materials to help keep the overall weight of the aircraft low. Such materials cannot tolerate the high temperatures radiating from the muffler and if exposed to such temperatures for any length of time, may experience some form of failure. Accordingly, aircraft manufactures commonly mandate that the heat radiating from the muffler not exceed a predetermined limit. So it is essential to design a muffler which not only attenuates the noise levels of the APU exhaust, but also need to be insulated with a low conductive insulated blanket around it.
A conjugate heat transfer CFD analysis was performed on a new APU exhaust system to optimize the exhaust muffler blanket design. Several different blanket material types and configurations were analyzed to minimize the heat transfer through the muffler. The CFD analysis included many geometric details including the exhaust eductor, to better model the exhaust gas profile, the muffler baffles and the tail cone structure. Muffler blanket outer surface temperature values and tail cone skin temperatures were monitored to determine the optimum design. This paper discusses the CFD model and analysis and discusses the results. It explains the different design iterations needed to arrive at acceptable muffler blanket outer surface temperatures and tail cone skin temperatures.
Comparison of Gas Turbine Vane Pressure Side and Suction Side Film Cooling Performance and the Applicability of Superposition
Authors: Kinell Mats , Utriainen Esa , Najafabadi Hossein Nadali , Karlsson Matts , Barabas Botond 
Published: August 2013

In order to protect a solid surface exposed to high temperature gaseous flows, e.g. gas turbines and rocket engines, a second gas at lower temperature may be introduced into the hot boundary layer, i.e. one obtains a three temperature problem. The impact of the film cooling on a prototype vane due to variation in blowing ratio, the shape of the hole-outlet and position has been experimentally investigated. The semi-infinite and low conductive test object, initially at a uniform temperature, was exposed to a sudden step change in main flow temperature and a time-resolved surface temperature was measured using an IR camera. By assuming constant values of the heat transfer coefficient and the film cooling effectiveness over time, the heat equation was solved using least squares.
The prototype vane was tested for different film cooling row positions on the pressure and suction side. Both cylindrical as well as fan shaped holes were investigated with and without showerhead cooling.
The resulting heat transfer coefficient and film cooling effectiveness on the pressure side is compared to flat plate studies and to the results from the suction side. Also, the applicability of using superposition on showerhead cooling and on single/double rows is investigated. Furthermore, the results are compared to other published airfoil film cooling experiments and to CFD analysis for which conclusions are drawn on quantitative and qualitative capabilities of this tool.
Experimental Study on Racetrack-Shaped Holes Impingement Cooling With Bump Type Roughening Element
Authors: Nakamata Chiyuki , Okita Yoji , Yamane Takashi , Fukuyama Yoshitaka , Yoshida Toyoaki 
Published: August 2013

Cooling effectiveness of an impingement cooling with array of racetrack-shaped impingement holes is investigated. Two types of specimens are investigated. One is a plain target plate and the other is a plate roughened with bump type elements. Sensitivity of relative location of bump to impingement hole on the cooling effectiveness is also investigated.
Experiments are conducted under three different mainflow Reynolds numbers ranging from 2.6×105 to 4.7×105, with four different cooling air Reynolds numbers for each main flow condition. The cooling air Reynolds numbers are in the range from 1.2×103 to 1.3×104.
Heat Transfer and Pressure Drop Measurements in High Solidity Pin Fin Cooling Arrays With Incremental Replenishment
Authors: Busche M. L. , Moualeu L. P. , Chowdhury N. , Tang C. , Ames F. E. 
Published: August 2013

Leading edge heat loads on turbine airfoils can be reduced by increasing the diameter of the leading edge. The lower external heat transfer and more generous curvature may allow for cooling this region internally. Large diameter leading edge regions are expected to exhibit a relatively broad region with nearly constant heat transfer. However, the ability of internal passages to cool a surface diminishes with distance as cooling air picks up thermal energy within the passage. Two novel internal cooling geometries which incrementally replenish cooling air, using impingement holes distributed along the array, have been designed and tested. These cooling methods have been compared to a baseline high solidity passage in terms of both array heat transfer and pressure drop. Heat transfer rates and pressure drop have been determined on a row by row basis to provide a means to assess their ability to sustain adequate cooling levels across the entire leading edge region. The authors believe turbine airfoil designs integrating large diameter leading edge regions with properly designed internal passages have the potential to eliminate showerhead cooling arrays in many industrial gas turbine applications. This change is especially beneficial in environments where fuel or air impurities have the potential to clog leading edge showerhead cooling arrays.
Heat transfer and pressure drop measurements were acquired in a bench scale test rig. Reynolds numbers ranged from approximately 5000 to 60,000 for the constant height channel arrays based on the pin diameter and the local maximum average velocity across a row. The high solidity pin fin arrays have an axial spacing (X/D) of 1.074 and a cross channel spacing (S/D) of 1.625. The constant section pin fin arrays have channel height to diameter ratios of 0.5. Each array has eight rows of pins with six pins per row in a staggered arrangement. Heat transfer testing was conducted using a constant temperature boundary condition.
Numerical Investigation of Inlet Swirl in a Turbine Cascade
Authors: Schmid Gregor , Schiffer Heinz-Peter 
Published: August 2013

New combustion concepts towards lean burn aim at reducing peak temperatures and therefore emissions, especially nitrogen oxides. High swirl is required in order to enhance the mixing of fuel and air and thus, improve combustion and flame stability. In a numerical investigation of a turbine vane cascade the effect of such inlet swirl on aerodynamic losses, secondary flow pattern and heat transfer is investigated.
The computations are conducted prior to particle image velocimetry and five-hole-probe measurements in a cascade of six vane passages and swirl generators upstream of each passage. The analysis covers three constituent parts: First, different swirl intensities are simulated which resemble the situation in a real combustion chamber. Second, different clocking positions are investigated — the swirl cores are either aligned with the vane leading edge or with midpassage — and finally, swirl orientation as clockwise, anticlockwise and counter rotating swirl is analysed. Two-dimensional inlet boundary conditions are applied to model the discrete swirl cores. Furthermore, a comparison with circumferentially averaged as well as with axial inflow conditions is made.
Increasing the swirl number at the inlet boundary results in reduced heat transfer coefficient within the vane passage and higher pressure loss. Heat transfer through vanes and endwalls is maximal if the swirl generators are aligned with the vane leading edge and counter rotating swirl.
Coupled CFD-ANN Procedure for Extending Heat Transfer Correlations Out of Their Range of Validity
Authors: Viano Andrea , Ottino Gabriele , Ratto Luca , Spataro Giuseppe 
Published: August 2013

The heat transfer coefficient and pressure losses are among the main parameters to be evaluated in gas turbine cooling network design. Due to the complexity of these estimates, correlation-based computations are typically used as a result of time-consuming and expensive experimental activities. One of the main problems that the industry has to face is that these correlations, based on non-dimensional experimental data, produce reliable results in a range of validity typically different from that encountered in gas turbine applications.
This paper will present preliminary results of an innovative procedure based on CFD analyses and Artificial Neural Networks, able to extend correlation predictions out of their range of validity, without any additional experimental data. Well-known test cases were replicated by building corresponding CAD geometries which were discretized by means of appropriate meshes, resulting from grid-independence studies. CFD analyses, based on the RANS approach, were performed to overlay the computations of the Nusselt number obtained from experimental activities. A preliminary comparison among turbulence models was carried out to find one leading to a good agreement with the experimental data. Then, an optimization method, based on Evolutionary Algorithms, was applied to the CFD analyses in order to find the best set of constant values for the chosen turbulence model, leading to the most accurate prediction of the experimental dataset. The resulting ad hoc CFD model was adopted in order to analyse test case configurations characterized by parameters within and external to the correlation validity field, building a sufficiently wide feeding database. A feed-forward multi-layer neural network was selected among network architectures typically used in engineering applications for prediction analyses. ANNs were chosen because they enable the solution of these complex nonlinear problems by using simple computational operations. The selected Artificial Neural Network was trained by a back-propagation procedure on the CFD results regarding Nusselt number. The validation of the resulting ANN was performed comparing its outputs with experimental data external to the correlation range of validity, which had not been used in the training session. Good agreement has been found. Results are presented and discussed.
Numerical Investigation of Heat Transfer and Cooling Effectiveness Within a LPT-Vane Cascade and its Comparison to Experimental Results
Authors: Beuermann Annette , Pahs Andreas , Rochhausen Stefan 
Published: August 2013

Gaps between stationary and rotating parts in turbines have to be fed with cooling air to keep metal temperatures below material limits. Reducing the coolant mass flow and analysing its impact on the flow field concerning aerodynamic and thermal data were the objectives of experiments, performed within the European research project AITEB.
As part of this project, measurements of cooling effectiveness and heat transfer on the endwall of a low pressure turbine nozzle guide vane were performed at a low speed cascade wind tunnel at DLR Göttingen.
Higher cooling mass flow rates increase secondary flows and subsequently heat transfer whilst metal temperatures are reduced due to larger coverage with coolant. It was also shown that heat transfer varies significantly with different flow angles.
According to the experimentally investigated flow fields numerical studies were performed using the DLR code TRACE, a RANS-Solver for turbomachinery flows. TRACE simulations were done using the Wilcox k-ω turbulence modelling. The boundary conditions were taken from the experimental setup.
In this paper the numerical data was analysed and compared with the experimental results regarding thermodynamics. The simulations confirmed a high influence of the flow angle. Within the flow regime affected by the injected coolant a good agreement between the numerical heat transfer results and the experimental data was observed. The qualitative and quantitative values were met after finding the optimum calculation parameters. Only in the region downstream of the throat area a different behaviour became obvious. Concerning film-cooling effectiveness quantitative differences between simulation and experiment were found whilst qualitative good agreement was observed.
Full Thermal Experimental Assessment of a Dendritic Turbine Vane Cooling Scheme
Authors: Luque S. , Batstone J. , Gillespie D. R. H. , Povey T. , Romero E. 
Published: August 2013

A full thermal experimental assessment of a novel dendritic cooling scheme for high-pressure turbine vanes has been conducted and is presented in this paper, including comparisons to the conventional cooling arrangement for these components. The dendritic cooling system consists of cooling holes with multiple internal branches which enhance internal heat transfer and reduce the blowing ratio at hole exit. Three sets of measurements are presented, which describe, first, the local internal heat transfer coefficient of these structures, and, secondly, the cooling flow capacity requirements and overall cooling effectiveness of a highly engine-representative dendritic geometry.
Full-coverage surface maps of overall cooling effectiveness were acquired for both dendritic and baseline vanes in the Annular Sector Heat Transfer Facility, where scaled near-engine conditions of Mach number, Reynolds number, inlet turbulence intensity and coolant-to-mainstream pressure ratio (or momentum flux ratio) are achieved. Engine hardware was used, with laser-sintered metal counterparts for the novel cooling geometry (their detailed configuration, design, and manufacture are discussed). The dendritic system will be shown to offer improved overall cooling effectiveness at a reduced cooling mass flow rate due to a more uniform film cooling effectiveness, a decreased tendency for films to lift off in regions of low external cross flow, improved through-wall heat transfer and internal cooling efficiency, increased internal wetted surface area of the cooling holes, and the enhanced turbulence induced in them.
Coupled Simulation of the Secondary Air Flow, Heat Transfer, and Structural Deflection of a Gas Turbine Engine
Authors: Tondello Guilherme , Boruszewski Wolodymir , Mengele Fernando , Assato Marcelo , Shimizu Silvio , Ziegler Shayne 
Published: August 2013

In secondary air flow in gas turbines, labyrinth seals are used to control the flow to and from each cavity and to the rotor blades for cooling purposes.
Those components and the final flow rate are very sensitive to gap clearance and displacement due to structural and thermal loads during operation, therefore designing those seals and knowing the resultant flow rates in each part of the circuit during the design phase is not an easy task, and tuning those gap values may bring significant increase in turbine efficiency.
This paper describes the application of coupled commercial codes for secondary air flow and structural simulation for better evaluating temperature profiles and labyrinth seal behavior during operation.
Flowmaster V7 was used for building a one dimensional model of the complete secondary air flow path including swirl effects and heat transfer phenomena, and ANSYS was used for building a structural model, taking into account both rotational and thermal loads.
The labyrinth seals clearances, and thermal interactions between solid and fluid were coupled bi-directionally between the two simulation software.
This simulation focused in the system, including the effects of each region, passage, seal and cavity in the calculations.
The turbine model simulated was a VSE’s gas turbine under development, having a nominal rotation of 22600 rpm.
This paper presents the numerical characteristics of each model, the details about the 1D fluid and 3D structural coupling, and the results obtained.
Comparison of Harmonic and Time-Marching Unsteady CFD Solutions With Measurements for a Single Stage High-Pressure Turbine
Authors: Green Brian R. , Mathison Randall M. , Dunn Michael G. 
Published: August 2013

The unsteady aerodynamics of a single-stage high-pressure turbine has been the subject of a study involving detailed measurements and computations. Data and predictions for this experiment have been presented previously, but the current study compares predictions obtained using the nonlinear harmonic simulation method to results obtained using a time-marching simulation with phase-lag boundary conditions.
The experimental configuration consisted of a single-stage high-pressure turbine (HPT) and the adjacent, downstream, low-pressure turbine nozzle row (LPV) with an aerodynamic design that is typical to that of a commercial high-pressure ratio HPT and LPV. The flow path geometry was equivalent to engine hardware, and operated at the proper design-corrected conditions to match cruise conditions. The high-pressure vane and blade were un-cooled for these comparisons. All three blade-rows are instrumented with flush-mounted, high frequency response pressure transducers on the airfoil surfaces, and the inner and outer flow path surfaces, which include the rotating blade platform and the stationary shroud above the rotating blade.
Predictions of the time-dependent flow field for the turbine flow path were obtained using a three-dimensional, Reynolds-Averaged Navier-Stokes CFD code. Using a two blade row computational model of the turbine flow path, the unsteady surface pressure for the high-pressure vane and rotor was calculated using both unsteady methods.
The two sets of predictions are then compared to the measurements looking at both time-averaged and time-accurate results, which show good correlation between the two methods and the measurements. This paper concentrates on the similarities and differences between the two unsteady methods, and how the predictions compare with the measurements since the faster harmonic solution could allow turbomachinery designers to incorporate unsteady calculations in the design process without sacrificing accuracy when compared to the phase-lag method.