bluuebunny/crossref_metadata_2025_split huggingface dataset:
The Effect of Fuel Composition on Combustion Instability Mode Occurrence in a Model Gas Turbine Combustor
Authors: Yoon Jisu , Joo Seongpil , Lee Min Chul , Kim Jeongjin , Oh Jaeyo , Yoon Youngbin 
Published: August 2015

Recently, energy resource depletion and unstable energy prices have become global issues. Worldwide pressure to secure and make more gas and oil available to support global power needs has increased. To meet these needs, alternative fuels composed of various types of fuels have received attention, including biomass, dimethyl ether (DME), and low rank coal. For this reason, the fuel flexibility of the combustion system becomes more important. In this study, H2 and CH4 were selected as the main fuel composition variables and the OH-chemiluminescence measurement technique was also applied. This experimental study was conducted under equivalence ratio and fuel composition variations with a model gas turbine combustor to examine the relation between combustion instability and fuel composition. The combustion instability peak occurs in the H2/CH4 50:50 composed fuel and the combustion instability frequency shifted to higher harmonic of longitudinal mode based on the H2 concentration of the fuel. Based on instability mode and flame length calculation, the effect of the convection time during the instability frequency increasing phenomenon was found in a partially premixed gas turbine combustor. The time-lag analysis showed that the short convection time in a high H2 concentration fuel affects the feedback loop period reduction and, in these conditions, high harmonics of longitudinal mode instability occurs.
This fundamental study on combustion instability frequency shifting characteristics was conducted for H2/CH4 composed fuel and the results contribute key information for the conceptual design of a fuel flexible gas turbine and its optimum operation conditions.
Assessment of Unsteady Pressure Measurement Uncertainty: Part 2 — Virtual Three Hole Probe
Authors: Dell’Era Giulia , Mersinligil Mehmet , Brouckaert Jean-François 
Published: August 2015

With the advancements in miniaturization and temperature capabilities of piezo-resistive pressure sensors, pneumatic probes — which are the long established standard for flow-path pressure measurements in gas turbine environments — are being replaced with unsteady pressure probes. On the other hand, any measured quantity is by definition inherently different from the ‘true’ value, requiring the estimation of the associated errors for determining the validity of the results and establishing respective confidence intervals. In the context of pressure measurements, the calibration uncertainty values, which differ from measurement uncertainties, are typically provided. Even then, the lack of a standard methodology is evident as uncertainties are often reported without appropriate confidence intervals. Moreover, no time-resolved measurement uncertainty analysis has come to the attention of the authors.
The objective of this paper is to present a standard method for the estimation of the uncertainties related to measurements performed using single sensor unsteady pressure probes, with the help of measurements obtained in a one and a half stage low pressure high speed axial compressor test rig as an example. The methodology presented is also valid for similar applications involving the use of steady or unsteady sensors and instruments.
The static calibration uncertainty, steady measurement uncertainties and unsteady measurement uncertainties based on phase-locked and ensemble averages are presented by the authors in [1]. Depending on the number of points used for the averaging, different values for uncertainty have been observed, underlining the importance of having greater number of samples. For unsteady flows, higher uncertainties have been observed at regions of higher unsteadiness such as tip leakage vortices, hub corner vortices and blade wakes.
Unfortunately, the state of the art in single-sensor miniature unsteady pressure probes is comparable to multi-hole pneumatic probes in size, preventing the use of multi-hole unsteady probes in turbomachinery environments. However, the angular calibration properties of a single sensor probe obtained via an aerodynamic calibration may further be exploited as if a three-hole directional probe is employed, yielding corrected total pressure, unsteady yaw angle, static pressure and Mach number distributions based on the phase-locked averages with the expense of losing the time-correlation between the virtual ports. The aerodynamic calibration and derivation process are presented together with the assessment of the uncertainties associated to these derived quantities in this contribution. In the virtual three-hole mode, similar to that of a single-sensor probe, higher uncertainty values are observed at regions of higher unsteadiness.
Steady and Unsteady CFD Calculation of the Laminar-to-Turbulent Transition in a Turning Mid Turbine Frame With Embedded Design
Authors: Bader Pascal , Sanz Wolfgang 
Published: August 2015

In order to meet the weight requirements of future jet engines S-shaped intermediate turbine ducts can be supplied with turning load carrying struts in order to save the first vane row of the subsequent low pressure rotor.
In such a duct large flow structures emanating from the outlet of the transonic high pressure stage are transported towards the low pressure rotor and are superimposed by secondary effects generated by the turning struts within the duct. To reduce the fluctuations behind the S-shaped duct and to homogenize the flow, the duct can be equipped with additional splitters. Such an embedded design has the advantage of a more homogenized flow entering the low pressure rotor, thus improving efficiency.
It is crucial for the overall efficiency to optimize the design of such an intermediate turbine duct. To understand all flow effects it is helpful to perform CFD calculations which do not neglect laminar-to-turbulent transition. Therefore this paper presents steady and unsteady CFD results of an S-shaped turbine duct with splitters considering boundary layer transition using Menter’s γ-ReΘ model. The differences between steady and unsteady simulation and the influence of transition are discussed.
CFD Modeling of Low Pressure Steam Turbine Radial Diffuser Flow by Using a Novel Multiple Mixing Plane Based Coupling: Simulation and Validation
Authors: Stein Peter , Pfoster Christoph , Sell Michael , Galpin Paul , Hansen Thorsten 
Published: August 2015

The diffuser and exhaust of low pressure steam turbines shows significant impact on the overall turbine performance. The amount of recovered enthalpy leads to a considerable increase of the turbine power output, and therefore a continuous focus of turbine manufacturers is put on this component. On the one hand, the abilities to aerodynamically design such components is improved, but on the other hand a huge effort is required to properly predict the resulting performance and to enable an accurate modeling of the overall steam turbine and therewith plant heat rate. A wide range of approaches is used to compute the diffuser and exhaust flow, with a wide range of quality. Today it is well known and understood, that there is a strong interaction of rear stage and diffuser flow, and the accuracy of the overall diffuser performance prediction strongly depends on a proper coupling of both domains. The most accurate, but also most expensive method is currently seen in a full annulus and transient coupling. However, for a standard industrial application of diffuser design in a standard development schedule, such a coupling is not feasible and more simplified methods have to be developed.
The paper below presents a CFD modeling of low pressure steam turbine diffusers and exhausts based on a direct coupling of the rear stage and diffuser using a novel multiple mixing plane. It is shown that the approach enables a fast diffuser design process and is still able to accurately predict the flow field and hence the exhaust performance. The method is validated against several turbine designs measured in a scaled low pressure turbine model test rig using steam. The results show a very good agreement of the presented CFD modeling against the measurements.
Design of a Single Stage Turbine for the Quantification of Aerodynamic Damping
Authors: Meinzer Christopher E. , Bittner Simon L. , Schmitt Stefan , Kielb Robert E. , Seume Joerg R. 
Published: August 2015

For an experimental study of aerodynamic damping in an axial low pressure turbine, a single stage subsonic turbine blisk is designed and aerodynamically and aeromechanically assessed. The design goals are a low aerodynamic damping, to allow a long ring down of blade vibration, and aeromechanical stable operating points.
With respect to these requirements, the aeromechanical analysis is executed using a time-linearized RANS method. The results of the final design are compared with another independent implementation of this method. Further, a preliminary flutter design tool is used to predict flutter stability and to cover the full range of current design tools. The rotor blading is manufactured as a blade integrated disk to eliminate friction damping in the planned experiment on the quantification of aerodynamic damping. To complement the study and to discuss the influence of damping, an experimental modal analysis of the manufactured blisk is completed to identify the frequency response and the damping.
Thermo-Mechanical Rotor Analysis Tuned by Engine Operational Data
Authors: Bruzzone Silvia , Mantero Marco , Patrone Stefano , Santolini Benedetto M. , Silingardi Andrea , Traversone Laura 
Published: August 2015

Gas turbine engines must withstand severe thermo-mechanical conditions during present-day load operation, characterized by cyclic transients and long dwell times. Indeed engine components are subject to thermal transient conditions, thermo-mechanical strain and stress fields; those are not easily measurable during operation, making calculations hardly confirmable. All these operational factors can lead a turbine component life reduction, finally increasing lifetime costs.
The developed approach has been based on several calculations, such as thermal and FEM stress evaluation on the rotor components, tuned or validated by different field measurements carried out by thermocouples in the rotor core and the pre-tightening load variation of tie-rod. Transient disks and tie-rod temperatures (calculated by an in-house Secondary Air System code) have been tuned on experimental data. Thus, for rotor thermo-mechanical analysis more reliable boundary conditions have been provided. Rotor FEM analysis has been finally checked comparing the variation of the tie-bolt tension (calculated by FEM analysis) with the experimental behaviour observed during different operating conditions.
Advanced Multi-Cup Fuel Injector Technology for Environmentally Responsible Aviation Gas Turbine Engines
Authors: Steinthorsson Erlendur , Mansour Adel , Hollon Brian , Teter Michael , Chang Clarence 
Published: August 2015

Participating in NASA’s Environmentally Responsible Aviation (ERA) Project, Parker Hannifin built and tested multipoint Lean Direct Injection (LDI) fuel injectors designed for NASA’s N+2 55:1 Overall Pressure-Ratio (OPR) gas turbine engine cycles. The injectors are based on Parker’s earlier three-zone injector (3ZI) which was conceived to enable practical implementation of multipoint LDI schemes in conventional aviation gas turbine engines. The new injectors offer significant aerodynamic design flexibility, excellent thermal performance, and scalability to various engine sizes.
The injectors built for this project contain 15 injection points and incorporate staging to enable operation at low power conditions. Ignition and flame stability were demonstrated at ambient conditions with ignition air pressure drop as low as 0.3% and fuel-to-air ratio (FAR) as low as 0.011. Lean Blowout (LBO) occurred at FAR as low as 0.005 with air at 460 K and atmospheric pressure. A high pressure combustion testing campaign was conducted in the CE-5 test facility at NASA Glenn Research Center at pressures up to 250 psi and combustor exit temperatures up to 2,033 K (3,200 °F). The tests demonstrated estimated LTO cycle emissions that are about 30% of CAEP/6 for a reference 60,000 lbf thrust, 54.8-OPR engine. This paper presents some details of the injector design along with results from ignition, LBO and emissions testing.
Influence of Geometry and Test Method on Fatigue Crack Propagation Behavior of a Ni-Base Jet Engine Alloy
Authors: Lenets Youri N. , Peralta Alonso , Miller Ross , Neumann James 
Published: August 2015

The majority of fatigue crack growth tests worldwide makes use of the compact tension (CT) specimens that are not necessarily representative of cracks developing under service conditions in highly stressed components of a jet engine. Over the years other geometries have been designed to facilitate a study of relatively small, semi- or quarter-elliptically shaped surface flaws subjected to high tensile and compressive stresses. Despite an extensive use by the aerospace community, practical aspects of testing and data analysis relevant to the complex surface- or corner-flawed geometries are not regulated by a commonly accepted set of rules.
Two types of test specimens — CT and surface-crack tension (SCT) — were machined from a forged and heat treated Inconel 718. For both geometries the crack orientation and propagation direction with respect to the original forging were the same.
The CT specimens were tested in accordance with the ASTM Standard E647 as well as using an alternative compression pre-cracking procedure. After correct application of the compression pre-cracking to the CT geometry both approaches had yielded reasonably consistent results. At high ΔK values both studied geometries also produced similar results, however, as the ΔK values decreased, a trend towards slower crack growth rates in the SCT specimens became evident.
In order to address a so-called small-crack effect, several SCT specimens received much smaller crack starter notches produced by the focused ion beam (FIB) technology.
The results of the present study highlight the importance of the appropriate material properties for accurate and reliable service life prediction of the critical aerospace propulsion hardware with particular emphasis on the influence of specimen/crack geometry and test method on the fatigue crack propagation behavior of jet engine alloys.
A Random Sampling Strategy for Tuning Contact Parameters of Underplatform Dampers
Authors: Gastaldi Chiara , Gola Muzio M. 
Published: August 2015

All numerical models of friction damped bladed arrays require knowledge of contact-friction parameters, which are established either through direct frictional measurements, done with the help of a separate single contact test arrangement, or by fine tuning the parameters in the numerical model of the real damping device and comparing the experimental response of a damped blade against its computed response. Recent results from direct measurements on underplatform dampers and the subsequent cross-comparison of experimental and numerical results have put into evidence several features which are usually neglected in FE models of damper-blade systems:
– markedly different friction coefficients at different contact points;
– friction coefficients evolving with time and cycle number towards a stable shape, in a systematic and repeatable manner with dramatic consequences on the shape of the hysteresis cycle and on dissipated energy;
– particular cases where minimal variations in the friction coefficients lead to gross changes of the damper behavior;
Identifying the contact parameters to assure the best match between model and experimental results becomes crucial to guarantee that the validated damper model will produce the correct cyclic forces on the blades during vibrational motion. While the tuning process described in the previous papers was a search based on the progressive refinement whose results depended on the operator’s ability to match different patterns of the model and experimental results, this paper dwells on a more objective and controllable method based on properly chosen indicators. The latter method is based on a sampling technique (Latin Hyper-cube) which produces a large number of solutions (in the present case 5000) on the basis of randomized extractions of contact parameters (in the present case 5) between given boundaries.
An advantage of this method is a systematic exploration of the influence of each input contact parameter on the collection of output indicators, considered acceptable according to predetermined criteria.
The main indicators suited for the purpose are three, i.e., the relative errors on the real and imaginary parts of the HBM complex spring equivalent to the hysteresis cycle and a measure of shape similarity between the experimental and simulated cycle. The paper shows the selection procedure which has been adopted to produce the final set of eligible solutions, which are further reduced by applying a secondary indicator based on the similarity of a kinematical parameter. In this paper this parameter is chosen to be a measure of the shape similarity of the damper rotation during the hysteresis cycle.
Fuel Composition Influence on Gas Turbine Ignition and Combustion Performance
Authors: Mosbach Thomas , Burger Victor , Gunasekaran Barani 
Published: August 2015

The influence of different jet fuel compositions on aviation gas turbine combustion performance was investigated. Eight fuels including conventional crude-derived Jet A-1 kerosene, fully synthetic Jet fuel, synthetic paraffinic kerosenes, linear paraffinic solvents, aromatic solvents and pure compounds were tested. The tests were performed in the altitude relight test facility located at the Rolls-Royce Strategic Research Centre in Derby (UK). The combustor employed was a twin-sector representation of an RQL gas turbine combustor. The combustor was operated at sub-atmospheric air pressure of 41 kPa and air temperature of 265 K. The temperature of the fuels was regulated to 288 K. The combustor operating conditions corresponded to a simulated low stratospheric flight altitude near 9,000 metres.
The experimental work at the Rolls-Royce (RR) test-rig consisted of classical relight envelope ignition and extinction tests, and ancillary optical measurements: Simultaneous high-speed imaging of the OH* chemiluminescence and of the soot luminescence was applied to obtain spatial and temporal resolved insight into the ongoing processes. Optical emission spectroscopy was also applied simultaneously to obtain spectral and temporal resolved insight into the flame luminescence.
First results from the analysis of the OH* chemiluminescence and detailed fuel analysis results were presented in previous papers [1, 2]. This article presents further results from the analysis of the soot luminescence imaging and flame spectra.
It was found in general that the combustion performance of all test fuel formulations was comparable to regular Jet A-1 kerosene. Fuel related deviations, if existent, are found to be small.
Development of a Blade Parametric Modeling Methodology for Design and Analysis of Computer Experiments
Authors: Leylek Zafer , Neely A. J. 
Published: August 2015

This paper will present an enhanced parametric modeling technique for gas turbine stator and rotor blades. The enhanced blade parametric modeling system has been developed as part of a wider research program into global surrogate modeling of compressor and turbine aerodynamic performance using Design and Analysis of Computer Experiments (DACE) based techniques.
The proposed method is based on a hybrid of geometric feature and Non-uniform Rational B-Spline (NURBS) based techniques. A base-line geometry is defined using the physical parameters and represented using NURBS curves and surfaces. A number of constraints are then imposed on the parametric model to ensure that DACE techniques can be effectively utilized. This is accomplished by mapping the geometric feature based parameters from the physical space to an alternative parametric space so that all feasible and numerically stable blade configurations can be represented using a unit hyper-cube. This method ensures a one-to-one mapping between the parametric sub-space and the geometric feature based system. The mapping is geometrically and numerically stable and does not produce ill-conditioned and unrealistic blade geometries.
The development of the blade parametric modeling process allows the application of the complete suit of DACE tools and techniques. The method is valid for all axial blade profiles which include compressor and turbine stator and rotor blades.
A Method for Forecasting the Condition of Several HPT Parts by Using Bayesian Belief Networks
Authors: Giesecke Daniel , Wehking Moritz , Friedrichs Jens , Binner Matthias 
Published: August 2015

The competitive ability of jet engine maintenance companies depends mainly on turn around time and overhaul costs. Both airline and maintenance companies need the best possible accuracy regarding the prediction of emerging costs and time of engine maintenance process to secure their operation. Estimating the deterioration status of engine modules prior to disassembling is one of the greatest challenges for the maintenance process. In a pilot project a Bayesian belief network (BBN) has been developed to determine the deterioration condition of the General Electric CF6-80C2 first stage high pressure turbine (HPT) nozzle guide vane (NGV). The aim of this paper is to extend the used BBN techniques to the HPT first and second stage rotor blades and the second stage vanes. Thereby, its objective is to prove the successful application of the developed method for constructing a BBN for component hardware forecast.
The BBN is composed of following parameters: component repair history, region, on-wing cycles, airfoil material, thrust rating, engine wing position and customer segment. Performing statistical data analysis and combining these parameters with expert knowledge result in component specific BBNs. These nets provide a moderate forecast accuracy of 59 percent for the first stage rotor blades, 65 percent for the second stage rotor blades and promising 89 percent for the second stage NGVs.
The paper concludes that a BBN has very good qualities to forecast the hardware condition of HPT components impressively shown by virtue of the nozzles. Therefore, it is worth to transfer the developed method to other modules in order to accurately predict the degradation of the components in an unconventional way.