bluuebunny/crossref_metadata_2025_split huggingface dataset:
Influence of Geometric Imperfections on Aerodynamic and Aeroelastic Behavior of a Compressor Blisk
Authors: Keller Christian , Kellersmann Andreas , Friedrichs Jens , Seume Joerg R. 
Published: August 2017

Impacts of foreign objects can cause cracks and dents in airfoils, especially in the leading edge. The regeneration of high-pressure compressors blisks with current repair methods is often restricted to a local blending of these edges. This can cause significant changes in the airfoils’ geometrical properties, which in turn influence their aerodynamic and aeroelastic characteristics. Changes at the leading edge have a particularly strong influence on the airfoils’ aerodynamic properties. In order to be able to make an informed decision about if and how a repair should be performed, consequences have to be predicted in advance.
To investigate their influence on the aerodynamic and aeroelastic behavior, typical blend repairs are applied to the geometry of a blisk in a 1.5-stage research axial compressor [1], which are representative in shape and size. Blisks (Blade-Integrated-diSK) are function integrated components, which are expected to have a high life span due to significant costs in design and production. Similar modifications are implemented at different radial heights of the blades, in order to investigate the influence of location and penetration depth of blend repairs. It is assured that only the blend repair region is modified while the rest of the blade stays in the original shape. Thus, a realistic change of the geometry is given.
The numerical study presented here deals with the influence of geometric imperfections, blend repairs in particular, on the aerodynamic and aeroelastic behavior of the high pressure compressors blisks. Results show that blend repairs have an influence on the local pressure distribution as well as on the local flow turning. Even though the leading edge is reshaped during repair, performance degradation can be observed. Furthermore, the working range of the compressor stage is influenced by the blend-repairs, which is of great importance for safe operation. Finally, the local changes in aerodynamics and blade deformation influence the aeroelastic behavior. This influence depends on the investigated mode shape and the location of the modification. The closer the modification is located towards the tip, the more pronounced are the shifts in aerodynamic damping and aerodynamic stiffness. Low torsional mode shapes display the highest sensitivity to the modifications.
Evolution of Reverse Flow in a Transonic Centrifugal Compressor at Near-Surge
Authors: Yamada Kazutoyo , Furukawa Masato , Arai Hiromitsu , Kanzaki Dai 
Published: August 2017

In process centrifugal compressors used in various types of plants, the compressor is the heart of a plant, and it requires high reliability. Therefore, prediction of the surge is important for centrifugal compressors. There have been numerous researches on the surge: study on improvement of surge margin, and study on the rotating stall, which is recognized as a precursor to surge, in impeller or diffuser of the compressor. However, the researches have not focused on the surge inception flow phenomena, namely detailed flow mechanism leading to the surge, although understanding of such flow phenomena is important for prediction of the surge.
The paper describes in detail unsteady flow fields in a transonic centrifugal compressor at near-surge conditions. The flow fields have been investigated by detached eddy simulations (DES) using 400 million grid points. The simulation results show that the huge reverse flow region occupies the flow field near the shroud in the impeller at off-design condition, triggered by the blade stall at the tip of impeller full-blade, and it drastically develops at near-surge. It is also found that the rotating disturbance with reversed flow appears in the diffuser near the endwall at around peak pressure-rise point, and it eventually evolves into the rotating stall cell with a large reverse flow, blocking the flow inside the diffuser at near-surge.
A Fully Coupled Approach for the Integration of 3D-CFD Component Simulation in Overall Engine Performance Analysis
Authors: Klein C. , Reitenbach S. , Schoenweitz D. , Wolters F. 
Published: August 2017

Due to a high degree of complexity and computational effort, overall system simulations of jet engines are typically performed as 0-dimensional thermodynamic performance analysis. Within these simulations and especially in the early cycle design phase, the usage of generic component characteristics is common practice. Of course these characteristics often cannot account for true engine component geometries and operating characteristics which may cause serious deviations between simulated and actual component and overall system performance. This leads to the approach of multi-fidelity simulation, often referred to as zooming, where single components of the thermodynamic cycle model are replaced by higher-order procedures. Hereby the consideration of actual component geometries and performance in an overall system context is enabled and global optimization goals may be considered in the engine design process.
The purpose of this study is to present a fully automated approach for the integration of a 3D-CFD component simulation into a thermodynamic overall system simulation. As a use case, a 0D-performance model of the IAE-V2527 engine is combined with a CFD model of the appropriate fan component.
The methodology is based on the DLR in-house performance synthesis and preliminary design environment GTlab combined with the DLR in-house CFD solver TRACE. Both, the performance calculation as well as the CFD simulation are part of a fully automated process chain within the GTlab environment. The exchange of boundary conditions between the different fidelity levels is accomplished by operating both simulation procedures on a central data model which is one of the essential parts of GTlab. Furthermore iteration management, progress monitoring as well as error handling are part of the GTlab process control environment. Based on the CFD results comprising fan efficiency, pressure ratio and mass flow, a map scaling methodology as it is commonly used for engine condition monitoring purposes is applied within the performance simulation. Hereby the operating behavior of the CFD fan model can be easily transferred into the overall system simulation which consequently leads to a divergent operating characteristic of the fan module. For this reason, all other engine components will see a shift in their operating conditions even in case of otherwise constant boundary conditions. The described simulation procedure is carried out for characteristic operating conditions of the engine.
The Effect of Foreign Object Damage on Compressor Blade High Cycle Fatigue Strength
Authors: Hanschke Benjamin , Klauke Thomas , Kühhorn Arnold 
Published: August 2017

For a considerable amount of time blade integrated disks (blisks) are established as a standard component of high pressure compressors (HPCs) in aero engines. Due to the steady requirement to increase the efficiency of modern HPCs, blade profiles get thinned out and aerodynamic stage loading increases. Ever since, aerofoil design has to balance structural and aerodynamic requirements. One particularity of aero engines is the possibility to ingest different kinds of debris during operation and some of those particles are hard enough to seriously damage the aerofoil. Lately, a growing number of blisk-equipped aero engines entered service and the question of foreign object damage (FOD) sensitivity relating to compressor blade high cycle fatigue (HCF) has emerged. Correct prediction of fatigue strength drop due to a FOD provides a huge chance for cost cutting in the service sector as on-wing repairs (e.g. borescope blending) are much more convenient than the replacement of whole blisks and corresponding engine strips.
The aim of this paper is to identify critical FOD-areas of a modern HPC stage and to analyze the effects of stress concentrations — caused by FOD — on the fatigue strength. A process chain has been developed, that automatically creates damaged geometries, meshes the parts and analyses the fatigue strength. Amplitude frequency strength (af-strength) has been chosen as fatigue strength indicator owing to the fact, that amplitudes and frequencies of blade vibrations are commonly measured either by blade tip timing or strain gauges. Furthermore, static and dynamic stress concentrations in damaged geometries compared to the reference design were computed. A random variation of input parameters was performed, such as the radial damage position at blade leading edge and damage diameter.
Based on results of the different samples, correlations of input parameters and the fatigue strength drop have been investigated. Evaluation shows a significant mode dependence of critical blade areas with a large scatter between drops in fatigue strength visible for mode to mode comparison. Keeping in mind the necessity of fast response times in the in-service sector, FOD sensitivity computations could be performed for all blade rows of the HPC — including the analysis of possible borescope blending geometries — in the design stage. Finally, the actual amplitude frequency levels (af-levels) of the modes excited during operation have to be appropriately taken into consideration. For example, a pronounced af-strength drop due to a FOD may not be critical for safe engine operations because the observed mode is excited by small af-levels during operation. Hence, the endurance ratio — a quotient of af-level and af-strength — is used as assessment criterion.
Heat Resistant Probe Combining Optic and Acoustic Sensors for Advanced Combustion Monitoring Including Detection of Flame Instabilities
Authors: Kraft Gerhard E. , Giuliani Fabrice , Pfefferkorn Lukas , Paulitsch Nina , Andracher Lukas 
Published: August 2017

Jet engines have remained almost entirely mechanical machines for fail-safe reasons, despite the increasing sophistication of modern gas turbines. However, the trend goes toward more electronic devices for a better operation monitoring. This is the late approach called system of systems in aeronautics.
New regulations such as the ICAO/CAEP/10 nvPM Standard set limitations on soot emissions. CO reduction is also an issue. One possible strategy toward more efficient combustion and less pollutant emissions is an advanced management of the safety margins. This is combined with an obligation to reduce operation costs. Therefore new measurement techniques are required for precision combustion monitoring during operation. The specific data requested covers the success of ignition, the margin before the lean-blow-out limit, the effective burner load conditions and the stability of combustion. Many optical measurement techniques are available for advanced combustion diagnostics (Warnatz et al 2001). Their main features are precision and non-intrusivity. However, if these techniques are commonly used in a combustion laboratory or on a test-bench, no application had a breakthrough so far on a flying system. The implementation of optical devices in the aggressive environment of a combustor is challenging. Some critical details are for instance the need for a permanently transparent optical interface or the thermal protection of the sensitive parts.
In the scope of the project “emotion” subsidised by the FFG, a heat resistant probe combining optic and acoustic sensors was developed for this purpose. This probe will make advanced combustion monitoring possible. It will comply with the above mentioned rules or constraints. It could be mounted on the pressure casing with a view on the liner. It will monitor the presence or absence of a flame, it will report on the ignition success or failure, it will compare the observed flame power to the expected load, and detect the presence of a combustion instability.
In this paper, several sensors are considered. Three different circuits for optical light intensity measurement are assessed. A combined optical-acoustic sensor arrangement called the Rayleigh-Criterion probe is introduced. This most promising configuration is tested and validated on an atmospheric combustion test rig. The presented results support the further development of this probe, first for use on test benches where this technology can achieve maturity, then towards deployment first in power gas turbines and eventually in aeroengines.
Mechanical Design and Validation Testing for a High-Performance Supercritical Carbon Dioxide Heat Exchanger
Authors: Sullivan Shaun D. , Farias Jason , Kesseli James , Nash James 
Published: August 2017

Supercritical carbon dioxide (sCO2) Brayton cycles hold great promise as they can achieve high efficiencies — in excess of 50% — even at relatively moderate temperatures of 700–800 K. However, this high performance is contingent upon high-effectiveness recuperating and heat rejection heat exchangers within the cycle. A great deal of work has gone into development of these heat exchangers as they must operate not only at elevated temperatures and very high pressures (20–30 MPa), but they must also be compact, low-cost, and long-life components in order to fully leverage the benefits of the sCO2 power cycle.
This paper discusses the mechanical design and qualification for a novel plate-fin compact heat exchanger designed for sCO2 cycle recuperators and waste heat rejection heat exchangers, as well as direct sCO2 solar absorber applications. The architecture may furthermore be extended to other very high pressure heat exchanger applications such as pipeline natural gas and transcritical cooling cycles. The basic heat exchanger construction is described, with attention given to those details which have a direct impact on the durability of the unit. Modeling and analysis of various mechanical failure modes — including burst strength, creep, and fatigue — are discussed. The design and construction of test sections, test rigs, and testing procedures are described, along with the test results that demonstrate that the tested design has an operating life well in excess of the 100,000 cycles/90,000 hour targets. Finally, the application of these findings to a set of design tools for future units is demonstrated.
A System Integration Approach for Heavy-Duty Gas Turbine Upgrades Using Improved Rotor Thrust Predictions and Application of Advanced Thrust Bearing Designs
Authors: Bavassano Francesco , Mantero Marco , Traverso Riccardo , Livermore-Hardy Richard , Blair Barry 
Published: August 2017

The progressive upgrading of heavy-duty gas turbines, aimed at increased performance, can ultimately introduce more onerous operating conditions, to the point that original design limits can be approached.
An increased gas turbine pressure ratio together with compression and expansion line adjustments can directly affect the rotor axial thrust. Other than the individual forces acting on the rotor, a key component to be taken into account is the fluid film thrust bearing, which should assure safe and reliable operation during the worst case operating conditions. Typically, such bearings are designed with large safety margins, yet it is possible that the new and more challenging conditions require a bearing capability upgrade, especially when field retrofit needs pose additional constraints.
A succession of performance upgrades have been carried out on Ansaldo Energia’s AE94.2 E-Class GT. An accurate understanding of the thrust-related phenomena proved necessary and drove improvements in the thrust bearing design along with hardware adjustments to lower the rotor thrust.
This paper addresses calculations and experimental arrangements for the rotor axial thrust evaluation on the aforementioned GT and considers both the matters related to the secondary air system for the thrust generation and the mechanical/functional matters for the bearing upgrade.
It is shown that issues such as uneven load sharing across the thrust bearing, or the variability of rotor thrust from engine to engine within the fleet, strongly affect the maximum thrust assessment and thus the requirements used in the process of selecting a suitable bearing. A predictive calculation method is described considering the main thrust contributions. Field experimental setups and main observations are reported. Measurements have been carried out using thermocouples and load cells placed on many of the thrust bearing pads. Moreover, the engine cavities carrying the highest and/or the most uncertain thrust share have been instrumented and characterized by pressure sensors.
The development of an upgraded thrust bearing is finally depicted through the main issues addressed, such as improved thrust pad lining material, increased load sharing efficiency and enlarged thrust bearing active surface area. Waukesha Bearings test results on the upgraded lining material, a high-tin aluminium alloy are reported as well.
A multidisciplinary approach is presented as necessary to manage the crucial challenge of improving the thrust balancing system, especially in the case of a formerly designed engine which receives a powerful upgrade.
Tip-Clearance Measurements on an Engine High Pressure Turbine Using an Eddy Current Sensor
Authors: Sridhar V. , Chana K. S. 
Published: August 2017

Gas turbine health monitoring is an important area of research. As the performance of aircraft and power plants increase, they will require better sensors for health monitoring systems to prevent failures. Health monitoring systems help in preventive maintenance reducing unnecessary downtime and maintenance costs.
Gas turbine blades are subjected to dynamic loads caused by rotor imbalances, distortions in the intake flows etc. These loads cause low or high cycle fatigues and the blades can fail over time. Tip-timing and tip-clearance systems makes it possible to assess turbomachinery blade vibration by using non-contact measurement systems such as optical, eddy current, hall effect, capacitve etc. The most widely used systems in industry are optical, however, these systems are still largely prone to contamination problems from dust, dirt, oil, water etc. Further development of these systems for in-service use is problematic because of the difficulty in eliminating contamination of the optics. Other systems, although immune to contamination, may not be able to measure both tip-clearance and tip-timing at the same time due to their operating principle. Another limitation is that they cannot be used in high temperature applications such as in a high pressure turbine where the temperatures can reach 1400°C. Eddy current sensors are found to be quite robust and can measure both tip-timing and tip-clearance. They are currently being used for gas turbine health monitoring applications at low temperatures such as in the compressor stage and last stage of a steam turbine.
A new high temperature eddy current sensor has been developed in-house at the University of Oxford for application in gas turbine tip-timing and tip clearance measurements to assess blade vibrations. The current sensor is a modified version of the existing eddy current sensor that is able to operate at high temperatures of about 1400°C. The paper presents the development of the sensor and experimental results of tip clearance measurements in the high pressure turbine stage of a jet engine. In the engine tests, two blades were reduced in height to increase the tip-clearance and the measurements were taken at both idle and max operating speeds. The sensor was found to work in these harsh environments and was sufficiently sensitive to accurately determine the tip clearance at these elevated temperatures. Tests were carried out mainly to demonstrate the technique of obtaining good tip clearance measurements and the survivability of the sensors in the high temperature and pressure environment.
Experimental Investigation of Oil Shedding From an Aero-Engine Ball Bearing at Moderate Speeds
Authors: Santhosh R. , Hee Jee Loong , Simmons Kathy , Johnson Graham , Hann David , Walsh Michael 
Published: August 2017

In civil aero-engine transmission system bearings are used for shaft location and load support. An experimental test facility in the University of Nottingham’s Gas Turbine Transmissions Research Centre (G2TRC) was designed and commissioned to investigate oil behaviour as it exits an engine-representative ball bearing. In the rig, oil is delivered to the bearing inner race and cage via under-race feed at three delivery locations i.e. front, mid and rear of the bearing assembly. An electromagnetic load system is designed and implemented to allow engine representative axial loads up to 35 kN to be applied to the bearing. This paper details the rig design including the load and under-race lubrication systems and gives information about bearing oil shedding mechanisms observed.
In this phase of testing high speed images are acquired at shaft speeds between 1000 and 7000 rpm at an oil flowrate of 5.2 litres per minute and bearing axial load of 10 kN. The work presented here focusses on oil shedding from the bearing cage. Oil shedding behaviour from aeroengine ball bearing is identified to share many similarities to that observed in the past for shedding from rotating disks and cups. However, it is shown that it not possible to predict the conditions at which transition in flow regimes will occur for the aeroengine bearing on the basis of correlations for simpler geometries (spinning disks and cups). The work presented here is the first observation of flow regimes in an aeroengine ball bearing involving high-resolution highspeed imaging.
Effect of Blade Aspiration Slot Configuration on the Aerodynamic Performance of a Highly Loaded Aspirated Compressor Cascade
Authors: Zhang Longxin , Cai Le , Liu Bao , Ding Jun , Wang Songtao 
Published: August 2017

As a promising active flow control method, boundary layer suction (BLS) can be used to enhance the aerodynamic performance of the highly-loaded compressor effectively, and due to this reason, extensive studies have been carried out on it. However, contrast to those abundant studies focusing on the flow control effects of BLS, little attention has been paid on the design method of the aspiration flow path.
This work presents a 3-D steady numerical simulation on a highly-loaded aspirated compressor cascade. The aspiration slot is implemented at its best location based on the previous experimental studies and the aspiration flow rate is fix to 1.5% of the inlet massflow. The plenum configuration follows the blade shape and remains unchanged. One-side-aspiration manner is adopted to simplify the aspiration devices. Two critical geometry parameters, slot angle and slot width, are varied to study the effects of blade aspiration slot configuration on the cascade loss, radial distribution of the aspiration flow rate and inner flow structures within the aspiration flow path. Results show that the slot configuration does affect the cascade performance. In comparison with the throughflow performance, it is especially true once the flow loss caused by the aspiration flow path is also taken into account, and higher flow loss will be generated within the aspiration flow path if an inappropriate scheme is adopted. In the present investigation, apart from the cases with larger negative slot angle, a wider slot is more preferable to a narrower one, since it could enhance the aspiration capacity near the endwall regions and lower the dissipation loss within the aspiration flow path. In terms of the slot angle, a larger negative value, i.e., the slot direction more aligned with the incoming flow, is not beneficial to improve the throughflow performance, while concerning the flow loss yield by the aspiration flow path, a proper negative slot angle is always optimal.