bluuebunny/crossref_metadata_2025_split huggingface dataset:
Validation of a Modified Parallel Compressor Model for Prediction of the Effects of Inlet Swirl on Compressor Performance and Operability
Authors: Floyd Reginald S. , Davis Milton 
Published: August 2015

Engine inlet distortion complications have plagued the turbine engine development community for decades, and engineers have developed countless methods to identify and combat the harmful effects of inlet distortion. One such type of distortion that has gained much attention in recent years is known as inlet swirl, which results in a significant flow angularity at the face of the engine. This flow angularity can affect the pressure rise and flow capacity of the fan or compressor, and subsequently affect compressor and engine performance.
Previous modeling and simulation efforts to predict the effect inlet swirl can have on fan and compressor performance have made great strides, yet still leave a lot to be desired. In particular, a one-dimensional parallel compressor model called DYNTECC (Dynamic Turbine Engine Compressor Code) has been used to analyze the effects of inlet swirl on fan and performance operability of the Honeywell F109 turbofan engine. However, when compared to experimental swirl data gathered at the United States Air Force Academy (USAFA), the model predictions were found to be inaccurate.
This paper documents work done to compare the initial predictions generated by DYNTECC to the latest set of experimental swirl data, analyze the potential shortcomings of the initial model, and modify the existing model to more accurately reflect test data. Extensive work was completed to create a methodology that can calibrate the model to existing clean inlet fan map data. In addition, an in depth study of fan/compressor stalling criteria was conducted, and the model was modified to use an alternate stalling criteria that more accurately predicted the point of stall for various swirl inlet conditions. The prediction of the fan stall pressure ratio for all inlet swirl conditions tested is within 2% of the ground test stall point at the same referred fan speed and referred mass flow.
Adjoint Optimization of an Internal Cooling Channel U-Bend
Authors: Willeke Sebastian , Verstraete Tom 
Published: August 2015

This paper addresses the optimization of a two-dimensional U-bend passage of an internal serpentine cooling channel for reduced total pressure loss by means of a steepest-descent method. A steady-state incompressible flow is considered at a Reynolds number of 40,000 based on the bulk velocity at the domain inlet. The two-equation k-ε model is used for primal turbulence modeling. After only 30 design iterations, the gradient-based optimization results in a reduction of total pressure loss by 46% compared to the baseline geometry.
To obtain the required objective gradients efficiently, a continuous adjoint approach is implemented in the OpenFOAM environment. Adjoint governing equations and boundary conditions are derived from state equations for steady-state, incompressible, turbulent flows under the assumption of frozen turbulence. Two different methods are proposed for modifying the shape of internal and external curves defining the duct geometry. The first method makes use of direct displacement of boundary grid points, allowing for a wide design space. The second, novel parameterization utilizes a projection of the surface sensitivities to an underlying Bézier curve. In this case, the Bézier control points are used as design variables. A comparison of both methods demonstrates a slightly lower performance improvement by the Bézier-based approach due to the reduced design freedom. This approach has, however, several practical advantages.
Previous studies already addressed this optimization problem using gradient-free methods, but were limited in the degrees of freedom given to the shape variation. The present gradient-based optimization allows for a much larger design space and hence is used to compare the different methodologies. It shows that both optimizations result in similar shapes, although the gradient-based method allows for a slightly larger reduction in pressure loss due to the wider design space, while converging faster towards the optimum.
Mechanical Design Features of a Small Gas Turbine for Power Generation in Unmanned Aerial Vehicles
Authors: Cunningham Caleb S. , Ransom David , Wilkes Jason , Bishop John , White Benjamin 
Published: August 2015

As part of the Intelligence Advanced Research Projects Activity (iARPA) Great Horned Owl (GHO) program, Southwest Research Institute® (SwRI®) developed and tested a small gas turbine for power generation in Unmanned Aerial Vehicles (UAV). This development program focused on advancing the state of the art in UAV power systems by meeting key metrics in weight, fuel efficiency, and noise generation.
Design, assembly, and testing of the gas turbine were completed in-house at SwRI. Fundamental mechanical design features of the gas turbine include an integrated 7 kW motor-generator, minimal oil lubrication system, cantilevered compressor/turbine assembly, and can combustor with air-atomizing fuel nozzles. The compressor/turbine assembly is cantilevered directly off of the motor-generator shaft, which spins on hybrid ceramic bearings. Due to potential rotor natural frequencies in the design operating range, the rotor-dynamic design of this configuration was a special design challenge. The outboard rotor bearing is softly supported on O-rings to provide compliance and drive shaft natural frequencies below the operating range.
The lube oil system is another interesting design feature of the GHO gas turbine. It is based on a minimal oil lubrication system previously used at SwRI. The minimal oil lubrication system relies on low oil flow rates and cooling air to pull droplets of oil through the bearing. The oil passes through the machine and is consumed during combustion. This system eliminates traditional oil recirculation hardware for simplicity and weight savings.
The can combustor features a modular design and uses additive manufacturing techniques to facilitate easy and cost effective prototyping. All combustor components are manufactured from Inconel 718 using direct metal laser sintering (DMLS) with additional post-machining. These parts are particularly challenging for DMLS because of their thin walls and high aspect ratio. The custom air-atomizing fuel nozzles also highlight one of the exciting advantages of the DMLS process. Each nozzle would be difficult to machine using traditional techniques because of the tight internal flow passages; however, they are simple to construct using additive manufacturing.
The GHO turbine developed by SwRI demonstrates interesting design features including a minimal oil lubrication system, a cantilever shaft with softly supported bearing, and combustor components built using additive manufacturing techniques. This design provides a platform for further development, testing, and demonstration of small gas turbine technology for UAV power generation.
Modelling Non-Uniform Bleed in Axial Compressors
Authors: Grimshaw S. D. , Pullan G. , Hynes T. P. 
Published: August 2015

The coupling between the bleed system and the flowfield of a downstream compressor stage is studied using two approaches.
In the first, three-dimensional, full annulus, unsteady computations simulate the flow in a low speed research compressor with non-uniform bleed extraction. Comparisons with experimental data show that the flow prediction in the main annulus is accurate to within 0.005 of flow coefficient and 0.5° of flow angle. The CFD is then used to provide a description of flow within the bleed system itself.
In the second approach, a two-dimensional mean radius model, similar to that adopted by Hynes and Greitzer in previous work on compressor stability, is used to simulate the response of the compressor to non-uniform bleed. This model is validated against experimental data for a single stage compressor and despite the inherent assumptions (two dimensional flow and simplified compressor response) provides a satisfactory prediction of the flow for preliminary design purposes with orders of magnitude less computational cost than full 3D CFD. The model is then used to investigate the effect of different levels of bleed non-uniformity and of varying the axial distance between the bleed and the downstream stage. Reducing bleed non-uniformity and moving the stage away from the bleed slot are predicted to reduce the circumferential non-uniformity of the flow entering the stage.
Unsteady Wet Steam Flow Field Measurements in the Last Stage of Low Pressure Steam Turbine
Authors: Bosdas Ilias , Mansour Michel , Kalfas Anestis I. , Abhari Reza S. , Senoo Shigeki 
Published: August 2015

Modern steam turbines need to operate efficiently and safely over a wide range of operating conditions. This paper presents a unique unprecedented set of time-resolved steam flowfield measurements from the exit of the last two stages of a low pressure (LP) steam turbine under various volumetric massflow conditions.
The measurements were performed in the steam turbine test facility in Hitachi city in Japan. A newly developed fast response probe equipped with a heated tip to operate in wet steam flows was used. The probe tip is heated through an active control system using a miniature high-power cartridge heater developed in-house.
Three different operating points, including two reduced massflow conditions, are compared and a detailed analysis of the unsteady flow structures under various blade loads and wetness mass fractions is presented. The measurements show that at the exit of the second to last stage the flow field is highly three dimensional. The measurements also show that the secondary flow structures at the tip region (shroud leakage and tip passage vortices) are the predominant sources of unsteadiness at 85% span. The high massflow operating condition exhibits the highest level of periodical total pressure fluctuation compared to the reduced massflow conditions at the inlet of the last stage. In contrast at the exit of the last stage, the reduced massflow operating condition exhibits the largest aerodynamic losses near the tip. This is due to the onset of the ventilation process at the exit of the LP steam turbine. This phenomenon results in 3 times larger levels of relative total pressure unsteadiness at 93% span, compared to the high massflow condition. This implies that at low volumetric flow conditions the blades will be subjected to higher dynamic load fluctuations at the tip region.
Capabilities of Thermal Wall Functions to Predict Heat Transfer on the NGVs of a Gas Turbine With Multiple Can Combustors
Authors: Mazzoni Cosimo Maria , Luque Salvador , Rosic Budimir 
Published: August 2015

Modern gas turbine combustion systems are characterised by enhanced air/fuel mixing and this results in more diffused and redistributed hot streaks leading to higher thermal loads established on the vane surface and endwalls. Thus, a detailed aero-thermal characterization of the near-wall region has become crucial both for the analysis of turbine performances and for the subsequent design of special features, such as the use of advanced materials and/or novel efficient cooling concepts. In order to investigate complex combustor-turbine flow interactions high fidelity experimental facilities and numerical tools are required.
In particular, detailed prediction of heat transfer at the gas turbine hot walls is computationally very demanding since it requires fully resolved boundary layers and very fine computational meshes. The present work investigates the capabilities of computationally less demanding near wall treatments (thermal wall functions) to predict heat transfer in gas turbine first stage vanes. This paper first summarizes the recent progress in the implementation of heat transfer capabilities into the CFD solver TBLOCK, by describing the near-wall treatments for forced thermal convection adopted in combination with standard turbulence models and aerodynamic wall functions. In order to assess the capabilities of the flow solver TBLOCK to predict heat transfer under engine realistic conditions an experimental cascade is then modelled numerically. The test case is the new linear cascade built at Oxford’s Osney Thermofluids Laboratory to simulate combustor-vane interactions in gas turbines for power generation; it consists of four first stage vane passages downstream a contracting inlet duct divided in two by a transition splitter that acts as the wall between two can combustors. Both aerodynamic and heat transfer numerical results are compared with available experimental data.
Numerical predictions show good agreement with experiments and well reproduce the aero-thermal influence of the combustor wall, showing the reliability of standard CFD tools in simulating these flow regimes without demanding CPU costs.
Experimental Study of Non-Reacting Low NOx Combustor Simulator for Scaled Turbine Experiments
Authors: Hall Benjamin F. , Povey Thomas 
Published: August 2015

This paper presents the first experimental characterisation of a fully annular combustor simulator incorporating both swirl and temperature pattern. The simulator has been designed to non-dimensionally replicate the conditions of next-generation low NOx combustors, with high swirl and high near wall temperature gradients, making it suitable for use in fully scaled (correct Re, M, N/√T, TG/TW) turbine facilities, to investigate combustor turbine interaction effects.
In reacting combustor flows with high swirl, combustion has a significant influence on vortex stability and therefore flow structure. In non-reacting simulators, with approximately isothermal flow, it is common to find instabilities that are not present in the equivalent reacting flow. Where this is the case, artificial methods must be used to stabilise the vortex core. It has previously been shown numerically and in simple experimental studies that a low-momentum axial jet at the vortex centre suppresses precession of the vortex core.
This paper reports experimental data for highly swirling flows both with and without stabilizing jets. This investigation is the first of its type, and the results have implications for the design of simulators for next-generation scaled turbine experiments both in industry and academia.
Improving Lifetime and Manufacturability of an RQL Combustor for Microturbines: Design and Numerical Validation
Authors: Laranci Paolo , Bidini Gianni , D’Alessandro Bruno , Zampilli Mauro , Forcella Fabio , Fantozzi Francesco 
Published: August 2015

A new annular RQL combustion chamber of an 80 kWel Elliott TA80R micro gas turbine was designed and built with a modified geometry to overcome known failures at low running hours (around 2500 hrs) caused by overheating. Design considered simplified manufacturability and flow optimization to reduce emission while maintaining similar temperatures and efficiencies. A preliminary geometry was analyzed and also built to verify manufacturability and economics. It was easily built with overall brute costs around 3000 €. It has also run continuously for over 27.000 hrs.
An optimized geometry, however, guaranteed similar TIT with respect to the original geometry with a considerable reduction in CO and NOx emissions.
Given the installation of the mGT at the IRP (Integrated Pyrolysis Regenerated Plant) the modified geometry was tested through CFD analysis on syngas from biomass thermochemical processes. The results show that further modifications of the liner are required for optimal operation and to reach adequate values for Turbine Inlet Temperature.
CFD Investigation of Labyrinth Seal Leakage Performance Depending on Mushroom Shaped Tooth Wear
Authors: Doğu Yahya , Sertçakan Mustafa C. , Bahar Ahmet S. , Pişkin Altuğ , Arıcan Ercan , Kocagül Mustafa 
Published: August 2015

Conventional labyrinth seal applications in turbomachinery encounter a permanent teeth tip damage and wear during transitional operations. This is the dominant issue that causes unpredictable seal leakage performance degradation. Since the gap between the rotor and the stator changes depending on engine transitional operations, labyrinth teeth located on the rotor/stator wear against the stator/rotor. This wear is observed mostly in the form of the labyrinth teeth becoming a mushroom shape. It is known that as a result of this tooth tip wear, leakage performance permanently decreases, which negatively affects the engine’s overall efficiency. However, very limited information about leakage performance degradation caused by mushroom wear is available in open literature.
This paper presents a study that numerically quantifies leakage values for various radii of mushroom shaped labyrinth teeth by changing tooth-surface clearance, pressure ratio, number of teeth, and rotor speed. Analyzed parameters and their ranges are mushroom radius (R=0–0.508mm), clearance (cr=0.254–2.032mm), pressure ratio (Rp=1.5–3.5), number of teeth (nt=1–12), and rotor speed (n=0–80krpm).
CFD analyses were carried out by employing compressible turbulent flow in 2-D axi-symmetrical coordinate system. CFD leakage results were also compared with well-known labyrinth seal semi-empirical correlations.
Given a constant clearance, leakage increases with the size of the mushroom radius that forms on the tooth. This behavior is caused by less flow separation and flow disturbance and the vena contracta effect for flow over the smoothly shaped mushroom tooth tip compared to the sharp-edged tooth tip. This leakage increase is higher when the tooth tip wear is considered as an addition to the unworn physical clearance, since the clearance dominates the leakage.
The leakage affected by the number of teeth was also quantified with respect to the mushroom radius. The rotational effect was also studied as a secondary parameter.
Self-Excited Blade Vibration Experimentally Investigated in Transonic Compressors: Rotating Instabilities and Flutter
Authors: Holzinger F. , Wartzek F. , Jüngst M. , Schiffer H.-P. , Leichtfuß S. 
Published: August 2015

This paper investigates the vibrations that occurred on the blisk rotor of a 1.5-stage transonic research compressor designed for aerodynamic performance validation and tested in various configurations at Technische Universität Darmstadt.
During the experimental test campaign self-excited blade vibrations were found near the aerodynamic stability limit of the compressor. The vibration was identified as flutter of the first torsion mode and occurred at design speed as well as in the part-speed region. Numerical investigations of the flutter event at design speed confirmed negative aerodynamic damping for the first torsion mode, but showed a strong dependency of aerodynamic damping on blade tip clearance.
In order to experimentally validate the relation between blade tip clearance and aerodynamic damping, the compressor tests were repeated with enlarged blade tip clearance for which stability of the torsion mode was predicted.
During this second experimental campaign, strong vibrations of a different mode limited compressor operation. An investigation of this second type of vibration found rotating instabilities to be the source of the vibration. The rotating instabilities first occur as an aerodynamic phenomenon and then develop into self-excited vibration of critical amplitude.
In a third experimental campaign, the same compressor was tested with reference blade tip clearance and a non-axisymmetric casing treatment. Performance evaluation of this configuration repeatedly showed a significant gain in operating range and pressure ratio. The gain in operating range means that the casing treatment successfully suppresses the previously encountered flutter onset. The aeroelastic potential of the non-axisymmetric casing treatment is validated by means of the unsteady compressor data.
By giving a description of all of above configurations and the corresponding vibratory behavior, this paper contains a comprehensive summary of the different types of blade vibration encountered with a single transonic compressor rotor. By investigating the mechanisms behind the vibrations, this paper contributes to the understanding of flow induced blade vibration. It also gives evidence to the dominant role of the tip clearance vortex in the fluid-structure-interaction of tip critical transonic compressors. The aeroelastic evaluation of the non-axisymmetric casing treatment is beneficial for the design of next generation casing treatments for vibration control.
Numerical Modeling of Soot Production in Aero-Engine Combustors Using Large Eddy Simulations
Authors: Franzelli B. , Riber E. , Cuenot B. , Ihme M. 
Published: August 2015

Numerical simulations are regarded as an essential tool for improving the design of combustion systems since they can provide information that is complementary to experiments. However, although numerical simulations have already been successfully applied to the prediction of temperature and species concentration in turbulent flames, the production of soot is far from being conclusive due to the complexity of the processes involved in soot production. In this context, first Large Eddy Simulations (LES) of soot production in turbulent flames are reported in the literature in laboratory-scale configurations, thereby confirming the feasibility of the approach. However numerous modeling and numerical issues have not been completely solved. Moreover, validation of the models through comparisons with measurements in realistic complex flows typical of aero-engines is still rare.
This work therefore proposes to evaluate the LES approach for the prediction of soot production in an experimental swirl-stabilized non-premixed ethylene/air aero-engine combustor, for which soot and flame data are available. Two simulations are carried out using a two-equation soot model to compare the performance of a hybrid chemical description (reduced chemistry for the flame structure/tabulated chemistry for soot precursor chemistry) to a classical full tabulation method. Discrepancies of soot concentration between the two LES calculations will be analyzed and the sensitivity to the chemical models will be investigated.
A Semi-Empirical Model to Predict Aircraft Soot Emission in Rich Zone of RQL Combustor
Authors: Choo Kyung Hak , Lee Sangmin , Denney Russell K. , Mavris Dimitri N. 
Published: August 2015

The reduction of particulate matter emissions is becoming a requirement for aircraft turbine engine combustor design. This requirement leads to the need to estimate particulate emissions during the conceptual design phase. Current prediction methods are based on detailed numerical simulation techniques such as CFD, which are unsuitable for conceptual design due to high computational cost.
This paper introduces a new approach employing a semi-empirical model for prediction of the soot emission indices of RQL combustors. The proposed approach dramatically improves computational efficiency by avoiding complex numerical calculations.
The model is based on the response surface developed from experimental data for turbulent non-premixed flames. The data has been extracted from the literature, employing statistical methods such as machine learning techniques and polynomial regressions to apply the turbulent flame data to the actual operating conditions in the primary zone of aircraft engine combustors.
The model is developed by first representing the combustor primary zone by chemical reactor networks constructed in CHEMKIN based on a statistical PDF approach to simulate the non-uniform distribution of time-evolving local mixture fraction with a beta distribution. The reactor networks are used to estimate the concentrations of soot precursor species in the rich zone. The empirical equations are then used with the predicted concentrations to predict the soot formation rate. Finally, these results are used along with the turbulent non-premixed flame data to develop the final model through a model calibration process.