bluuebunny/crossref_metadata_2025_split huggingface dataset:
The Pressure Field at the Output From a Low Pressure Exhaust Hood and Condenser Neck of the 1090 MW Steam Turbine: Experimental and Numerical Research
Authors: Hoznedl Michal , Živný Antonín , Macálka Aleš , Kalista Robert , Sedlák Kamil , Bednář Lukáš , Tajč Ladislav 
Published: August 2018

The paper presents the results of measurements of flow parameters behind the last stage of a 1090 MW nominal power steam turbine in a nuclear power plant. The results were obtained by traversing a pneumatic probe at a distance of about 100 mm from the trailing edges of the LSB (Last Stage Blade). Furthermore, both side walls as well as the front wall of one flow of the LP (Low Pressure) exhaust hood were fitted with a dense net of static pressure taps at the level of the flange of the turbine. A total of 26 static pressures were measured on the wall at the output from the LP exhaust hood. Another 14 pressures were measured at the output from the condenser neck. The distribution of static pressures in both cross sections for full power and 600 and 800 MW power is shown. Another experiment was measured pressure and angle distribution using a ball pneumatic probe in the condenser neck area in a total of four holes at a distance up to 5 metres from the neck wall.
The turbine condenser is two-flow design. In one direction perpendicular to the axis of the turbine cold cooling water comes, it heats partially. It then reverses and it heats to the maximum temperature again. The different temperature of cooling water in the different parts of the output cross section should influence the distribution of the output static pressure. Differences in pressures may cause problems with uneven load of the tube bundles of the condenser as well as problems with defining the influential edge output condition in CFD simulations of the flow of the cold end of the steam turbine Due to these reasons an extensive 3D CFD computation, which includes one stator blade as well as all moving blades of the last stage, a complete diffuser, the exhaust hood and the condenser neck, has been carried out. Geometry includes all reinforcing elements, pipes and heaters which could influence the flow behaviour in the exhaust hood and its pressure loss. Inlet boundary conditions were assumed for the case of both computations from the measurement of the flow field behind the penultimate stage. The outlet boundary condition was defined in the first case by an uneven value of the static pressure determined by the change of the temperature of cooling water. In the second case the boundary condition in accordance with the measurement was defined by a constant value of the static pressure along all the cross section of the output from the condenser neck. Results of both CFD computations are compared with experimental measurement by the distribution of pressures and other parameters behind the last stage.
Experimental and Numerical Investigation of Environmental Barrier Coated CMC Turbine Airfoil Erosion
Authors: Okita Yoji , Mizokami Yousuke , Hasegawa Jun 
Published: August 2018

Ceramic matrix composite (CMC) have higher temperature durability and lower density property compared to nickel-based super-alloys which so far have been widely applied to hot section components of aero-engines / gas turbines. One of promising CMC systems, SiC-SiC CMC is able to sustain its mechanical property at higher temperature, though it inherently needs environmental barrier coating (EBC) to avoid oxidation. There are several requirements for EBC. One of such critical requirements is its resistance to particle erosion, whereas this subject hasn’t been well investigated in the past.
The present work presents the results of a combined experimental and numerical research to evaluate the erosion characteristics of CMC+EBC material developed by IHI.
First, experiments were carried out in an erosion test facility using 50 micron diameter silica as erosion media under typical engine conditions with velocity of 225 m/s, temperature of 1311 K, and impingement angles of 30, 60, and 80deg. The data displayed brittle erosion mode in that the erosion rate increased with impact angles.
Next, multi-phase flow simulations were carried out for the experimental setup and the predicted erosion rates were compared with the measured data, and consequently, the applied numerical method was verified.
The validated numerical method was then applied to compute three dimensional flow field, particle trajectories, and erosion profile around a generic turbine airfoil to assess the erosion characteristics of the proposed CMC+EBC material when applied to airfoil. The trajectories indicated that the particles primarily impacted the airfoil leading edge and the pressure surface. Surface erosion patterns were predicted based on the calculated trajectories and the experimentally-based erosion characteristics.
Wall Temperature Measurements in Gas Turbine Combustors With Thermographic Phosphors
Authors: Nau Patrick , Yin Zhiyao , Lammel Oliver , Meier Wolfgang 
Published: August 2018

Phosphor thermometry has been developed for wall temperature measurements in gas turbines and gas turbine model combustors. An array of phosphors has been examined in detail for spatially and temporally resolved surface temperature measurements. Two examples are provided, one at high pressure (8 bar) and high temperature and one at atmospheric pressure with high time resolution.
To study the feasibility of this technique for full scale gas turbine applications a high momentum confined jet combustor at 8 bar was used. Successful measurements up to 1700 K on a ceramic surface are shown with good accuracy. In the same combustor, temperatures on the combustor quartz walls were measured, which can be used as boundary conditions for numerical simulations.
An atmospheric swirl-stabilized flame was used to study transient temperature changes on the bluff body. For this purpose, a high-speed setup (1 kHz) was used to measure the wall temperatures at an operating condition where the flame switches between being attached (M-flame) and being lifted (V-flame) (bistable). The influence of a precessing vortex core (PVC) present during M-flame periods is identified on the bluff body tip, but not at positions further inside the nozzle.
Experimental Study of the Leakage and Rotordynamic Coefficients of a Long Smooth Seal With Two-Phase, Mainly-Oil Mixtures
Authors: Zhang Min , Childs Dara W. , Mclean James E. , Tran Dung L. , Shrestha Hari 
Published: August 2018

Tests are conducted on a newly developed 2-phase annular seal stand (2PASS) at the Turbomachinery Laboratory of Texas A&M University. The test fluid is a mixture of silicone oil (PSF-5cSt) and air. Two spargers are used to produce mainly-oil mixtures by injecting air bubbles into the oil flow.
The test seal is a smooth annular seal with inner diameter D = 89.306 mm, length-to-diameter ratio L/D = 0.65, and radial clearance Cr = 0.188 mm. Tests are performed with inlet gas-volume-fraction GVF = 0%, 2%, 4%, 6%, and 10%, rotor speed ω = 5, 7.5, 10, and 15 krpm, inlet temperature Ti = 39.4 °C, exit pressure Pe = 6.9 bars, and pressure drop PD = 31, 37.9, and 48.3 bars. The test seal is centered, and there is no intentional prerotation of the fluid at the seal inlet.
The complex dynamic stiffness coefficients of the test seal are measured and fitted by the frequency-independent stiffness Kij, damping Cij, and virtual-mass Mij coefficients.
Test results show that adding air into the oil flow does not change the seal’s mass flow leakage ṁ discernibly but significantly impacts the seal’s rotordynamic characteristics.
Some planned 5 krpm cases with low inlet GVFs at PD = 31 and 37.9 bars are not accomplished due to stator instabilities, which are likely caused by negative stiffness of the test seals. For ω = 5 krpm when PD = 31 and 37.9 bars, direct stiffness K decreases from positive to negative as inlet GVF decreases.
For all PDs and speeds, K increases as inlet GVF increases from zero to 10% except for 6% ≤ inlet GVF ≤ 10% when PD = 48.3 bars, where K decreases as inlet GVF increases. The K increment will increase a pump rotor’s natural frequency and critical speed. Increasing the rotor’s natural frequency would also increase the onset speed of instability (OSI) and improve the stability of the rotor. Adding air into the oil flow has little effect on cross-coupled stiffness k and direct damping C. Increasing inlet GVF has negligible effects on direct virtualmass M when ω ≤ 10 krpm and PD ≤ 37.9 bars, but generally decreases M when ω = 15 krpm or PD = 48.3 bars.
Increasing inlet GVF has little effect on effective damping Ceff and does not change the seal’s resultant stabilizing force discernibly. Ceff = C − k/ω + mqω, where mq is the cross-coupled virtual-mass.
Test results are compared to predictions from San Andrés’s [1] model. The model is based on a bulk-flow model and the Moody friction formula assuming that the liquid-gas mixture is isothermal and homogenous. The model reasonably predicts ṁ, C, and Ceff. All predicted K values are positive, while measured K values are negative for some test cases. Predicted k values are close to measurements when ω = 5 krpm and are larger than test results when 7.5 ≤ ω ≤ 15 krpm. M predictions are smaller than measurements, and the discrepancy between predicted and measured M values generally increases as inlet GVF increases.
Mistuning Identification Approach With Focus on High-Speed Centrifugal Compressors
Authors: Weber Robby , Kühhorn Arnold 
Published: August 2018

Blade vibrations are one of the main cost drivers in turbo-machinery. Computational blade vibration analysis facilitates an enormous potential to increase the productivity in the design of bladed components. Increasing computing power as well as improved modeling and simulation methods lead to comprehensive calculation results. This allows for a more precise prediction and assessment of experimental data.
Usually, in the field of turbomachinery, identical blades are assumed to lower the required computational resources. However, mistuning is unavoidable, since small deviations due to the manufacturing process will lead to slightly different blade behavior. Potential effects such as mode localization and amplification can be treated statistically and have been thoroughly studied in the past. Since then, several reduced order models (ROMs) have been invented in order to calculate the maximum vibration amplitude of a fleet of mistuned blisks. Most commonly, mistuning is thereby modeled by small material deviations from blade to blade, e.g. Young’s modulus or density.
Nowadays, it is common knowledge that the level of manufacturing imperfection (referred as level of mistuning) significantly influence mode localization as well as vibration amplification effects. Optical measurements of the geometric deviations of manufactured blades and converting to a high-fidelity finite element model make huge progress. However, to the knowledge of the authors, there is no reliable method, that derives a characteristic quantity from the geometric mistuning, that fits into the mentioned statistically approaches.
Therefore, experimental data is needed to quantify the level of mistuning. Several approaches, which isolate blade individual parameters, are used to identify the dynamic behavior of axial compressors and turbines. These methods can be applied to medium-speed centrifugal turbine wheels but tend to fail to evaluate high-speed compressor with splitter blades. This paper briefly presents the original approach and discusses the reasons for failure. Thereafter, a new approach is proposed. Finally the level of mistuning and important quantities to perform a statistical evaluation of a high-speed compressor is shown.
Through-Flow Modeling of Single- and Two-Shaft Gas Turbines at Wide Operating Range
Authors: Wiedermann Alexander , Petrovic Milan V. 
Published: August 2018

A program suite of coupled analysis tools for all components of industrial gas turbines developed by the authors was applied to off-design operation of single- and two-shaft gas turbines. The through-flow methods are based on a stream function approach and a finite element solution procedure. They include high-fidelity loss and deviation models with improved correlations. Advanced radial mixing and end-wall boundary layer models were applied to simulate 3D flow effects. For air-cooled turbine analysis, various types of cooling air injection were encompassed: film cooling, trailing edge injection and disc/end-wall coolant flow.
Several improvements and extensions of the meridional solvers involved were employed. For axial compressor analysis, the original perfect gas treatment was replaced with the same real gas library that had already been implemented for flow computations in air-cooled expansion turbines and combustion modules. Ambient conditions with wet inflow of the compressor can now be taken into account. Compared to the previous model, better closure of the thermodynamic cycle can be achieved with the new approach. The meridional flow solver applied to the cooled turbine was upgraded and has become more stable for transonic outlet Mach numbers. With regard to the cooling flow supply, the codes offer greater freedom at the compressor extraction and turbine injection stations, allowing more realistic modeling of the secondary flow bifurcation.
Off-design operating conditions and sensitivity studies for both single- and two-shaft gas turbines covering the entire range of interest are presented and are compared with experimental rig-engine data where available.
The Effect of Tip Shroud Geometries on Last Turbine Stage Efficiency
Authors: Granovskiy Andrey , Afanasiev Igor 
Published: August 2018

Last stages of steam turbines and heavy-duty power gas turbines contribute significantly to output power and efficiency of whole turbine. Moreover, radial distribution of parameters downstream of the last stage provides boundary conditions for diffuser design. Thus, the increase of the last stage efficiency and obtainment of favorable radial distribution downstream of the last rotor blade is very important. Due to the long blades of last stages, resonance might occur. To avoid dangerous frequencies a damping wire or damping bolts are used. Such damping elements result in additional losses, so to minimize these losses a damping shroud is used instead. In general, the full damping shroud has to provide both the aerodynamic loss reduction and the resonance frequency offset. However, in most cases due to mechanical integrity limits instead of the full shroud a partial shroud is used. In this case the loss reduction feature of the partial shroud is diminished as compared with the full shroud. Sometimes, the use of the partial shroud results in the decrease of the efficiency compared with a stage with unshrouded rotor blades at small tip clearances. In this paper, a numerical investigation of the flow structure around full and partial shrouds with various geometries as well as the effect of the various shroud geometries on the turbine stage efficiency is carried out. Eight geometries with different number of fins of various heights are studied. Moreover, stage efficiencies for both shrouded and unshrouded blade are compared. Based on this comparison, reasonable design recommendations aimed to reduce the losses within the radial gap over the shroud are developed. In particular, filling the space in the gap with the additional honeycombs is considered and the effect on the flow structure and the last stage efficiency investigated.
Numerical results obtained in the paper correspond well to the published test data.
Inlet Air Filter Elements Site and Laboratory Tests of Four Different Manufacturers
Authors: Ingistov Steve , Kohn Joshua 
Published: August 2018

This Paper describes three-month onsite testing utilizing travelling laboratory furnished with four dedicated air ducts and required instrumentation needed to determine efficiency of filtration for submicron size particles. It also describes the subsequent testing of seasoned filters in offsite Laboratory furnished with instrumentation and equipment to conduct the dry and wet tests on previously Watson-field tested filters.
The opportunity to simultaneously test four different IAFE, provided by four different Manufacturers, was unique and helpful in discerning the differences between these filters.
Engine Testing for the Demonstration of a 3D-Woven Based Ceramic Matrix Composite Turbine Vane Design Concept
Authors: Watanabe Fumiaki , Manabe Takashi 
Published: August 2018

In this paper, the authors present a solid vane concept which uses 3D woven fabric based SiC/SiC CMC and the result of engine testing which was conducted to demonstrate this concept.
The authors devised a design concept, in which a vane is made from a folded woven, for the uncooled turbine vanes of the aircraft engines. The advantage of this concept is that there is no structural weak point because a whole vane including outer / inner platform consists of one continuous woven.
In order to demonstrate the applicability of this concept to the actual engine parts, CMC turbine vanes were tested with IHI-IM270 that is a small industrial gas turbine engine. After 862 hour testing, there was no defect such as cracks, wear, and deformation.
Effects of Inlet Disturbances on Fan Stability
Authors: Lee Kuen-Bae , Wilson Mark , Vahdati Mehdi 
Published: August 2018

This paper investigates the effects of inlet disturbance caused by crosswind on a fan blade operation and addresses the possible aerodynamic instabilities which can arise for such a fan-intake system. The work is carried out by using a three-dimensional unsteady CFD model (AU3D), and for a modern low-speed fan rig for which extensive measured data is available. The computational domain includes the fan, with outlet guide vanes (OGV) for the bypass flow and engine-section stators (ESS) for the core and a symmetric intake upstream of the fan (a whole low-pressure domain). The unsteady full annulus simulations under crosswind are performed to analyze the effects of inlet disturbances on the operation of this blade. It was observed that, for sufficiently high amplitudes of crosswind, the intake lip separates, and results in a significant loss of stall margin. Moreover, even in the absence of lip separation, the blade can still stall prematurely due to non-homogeneous flow caused by the two contra-rotating trailing vortices.
In the 2nd phase of this study, the effects of fan loading on the suppression of flow separation in the intake, and the consequent stall margin of the fan blade was explored. The results indicated that, as the fan speed increases it becomes more capable of reducing the inlet distortion levels, and consequently it can extend the stall margin of the fan blade.
Dual Drive Booster for a Two-Spool Turbofan: High Shaft Power Offtake Capability for MEA and Hybrid Aircraft Concepts
Authors: Kloos Vadim , Speak Trevor H. , Sellick Robert J. , Jeschke Peter 
Published: August 2018

The effects of high shaft power offtake in a direct drive, a geared drive, and a novel turbofan configuration are investigated. A design and off-design performance analysis shows the configuration specific limitations and advantages.
The more electric aircraft (MEA) concept promises to offer advantages with respect to aircraft performance, maintenance and operating costs. The engines for the MEA concept are based on conventional turbofan architectures. These engines are designed for significantly increased shaft power offtake that is required by the airframe, and the shaft power is usually taken off the high-pressure spool. This can impair the off-design performance of the engine and lead to compromises during engine design and to operability limitations. Taking the power off the low-pressure spool mitigates some of the problems but has other limitations. In this work, an alternative novel turbofan architecture is investigated for its potential to avoid the problems related to high shaft power offtakes. This architecture is called the dual drive booster because it uses a summation gearbox to drive the booster from both the low- and high-pressure spool. The shaft power, if taken off the booster spool, is effectively provided by both the low- and high-pressure spools, which allows the provision of very high power levels. This new concept is benchmarked against a two-spool direct drive and a geared drive turbofan. Furthermore, it is described, how the new architecture can incorporate an embedded motor generator. The presented concept mitigates some of the problems which are encountered during high power offtake in conventional configurations. In particular, the core compressors are less affected by a change in shaft power offtake. This allows higher power offtakes and gives more flexibility during engine design and operation. Additionally, the potential to use the new configuration as a gas turbine-electric hybrid engine is assessed, where electrical power boost is applied during critical flight phases. The ability to convert additional shaft power is compared with conventional configurations. Here, the new configuration also shows superior behavior because the core compressors are significantly less affected by power input than in conventional configurations. The spool speed and its variation is more suitable for electrical machines than in conventional configuration with low-pressure spool power transfer.
The dual drive booster concept is particularly suited for applications with high shaft power offtakes and inputs, and should be considered for propulsion of more electric aircrafts.
Multi-Objective Optimization of the Cooling Configuration of a High Pressure Turbine Blade
Authors: Wagner Frank , Kühhorn Arnold , Janetzke Timm , Gerstberger Ulf 
Published: August 2018

Due to the increasing turbine inlet temperature and in order to improve the overall efficiency it is necessary to optimize the cooling design of the hot gas components of an aero engine.
The current paper discusses the strategy of optimizing a rotor blade cooling configuration of a small civil aero engine, comprising of films and internal turbulators (ribs). An insight into the parametrization is given including the location of the films and ribs as well as the number of the films and ribs. The parameter reduction results in 18 input parameters for the optimizations to limit the number of parameters to an acceptable level.
Two optimizations are carried out with the primary objectives of non-dimensional mass flow and overall cooling effectiveness. Different optimization algorithms are used, namely AMGA and NSGA-II, and compared afterwards. A further optimization is carried out with direct objectives of mass flow and mean surface temperature using the AMGA algorithm.
The outputs from the optimizations are presented as a pareto-front. These plots are used for a comparison of the optimization algorithms and formulations respectively. Finally, the differences are discussed and the advantages and disadvantages of the algorithms used are highlighted.