bluuebunny/crossref_metadata_2025 huggingface dataset:
Effect of Vortex Shedding on the Performance of Scoop Based Lubrication Devices
Authors: Prabhakar Arun , Abakr Yousif Abdalla , Simmons Kathy
Published: August 2017

In civil aircraft aeroengine bearing chambers it is sometimes difficult to feed oil to bearings using the traditional under-race or targeted jet approaches. In such situations one proposed solution is that of a scoop delivery system. Published experimental investigations into scoop performance show that scoop collection efficiency (the percentage of oil delivered by the scoop system to its destination compared to that supplied by the feed jet) is a function of many operational and geometric parameters. However even with high speed imaging it is impossible to experimentally determine in detail the factors that most contribute to reduction in collection efficiency and it is here particularly that a computational fluid dynamics (CFD) investigation has value.

In the work reported here a commercial CFD code (ANSYS Fluent) is used to investigate vortex formation at the scoop tips and the effect these structures have on scoop collection efficiency. The computational domain, a 2D slice through the chosen scoop system, is discretized utilizing ANSYS Meshing. A Volume of fluid (VOF) method is used to model the multiphase flow of oil and air in the system and the RNG k-ε turbulence model is employed.

The results obtained show that the formation of vortices from the tip of the rotating scoops leads to a reduction in pressure in the region near the tip of the oil jet, subsequently causing part of the jet to divert upwards away from the scoop creating a plumed tip. The pluming effect reduces capture efficiency because the oil plume moves outwards under centrifugal effects and this oil is not captured. The frequency of vortex shedding from the scooped rotor was investigated and the Strouhal numbers obtained were around 0.132. This compares well to 0.15 for an inclined flat plate. Two potential methods to reduce the jet pluming effect are investigated one in which the sharp tip of the scoop is blunted and the other in which the jet direction is reversed. The blunt tip increased capture efficiency by almost 2%. Reversing the jet orientation reduces jet pluming but also significantly reduces capture efficiency; it was found to be 10% lower for the case investigated.
Performance Evaluation of a Hydraulic Turbine Used As a Turbodrill for Oil and Gas Applications in Post-Salt Environment
Authors: Monteiro V. G. , Tomita J. T. , Bringhenti C. , Vastenavond A. , Sampaio J. H. B.
Published: August 2017

Turbodrill is a type of hydraulic axial turbomachine that rotates a bit by the action of the drilling fluid on turbine blades, which converts the hydraulic power provided by the high pressure from drilling fluid into mechanical power through turbine stages. The evaluation of hydraulic turbine performance characteristics are important to define feasible rotational speed and mass flow to attend the bit torque requirements during drilling through the post-salt and salt layers. As a result, optimum operational parameters are proposed for gaining the required rotational speed and torque for post-salt environments. The turbine motor presented in this study was established by design methods based on classical aeronautical turbomachinery blade profile to supply 30k Newton-meters (Nm) of torque requested by a polycrystalline diamond compact (PDC) bit to power the complex heterogeneous layer of rock. The performance evaluation of this innovative hydraulic turbine with 200 stages was carried out using computational fluid dynamics (CFD). The simulation considers two different drilling fluid types, sea water and brine. Besides, different flow rates were considered to investigate how velocity vectors, pressure profile, output power and other performance parameters are affected. Due the large amount of data, the first and second stages of the turbine have been used to predict the performance characteristics. This assumption gives interesting results and avoids too heavy computational costs. A commercial CFD solver (ANSYS CFX 15.0®) was used to calculate the governing equations based on Reynolds-Averaged Navier-Stokes (RANS equations) with the addition of turbulence model. The two-equation Shear-Stress Transport (SST) turbulence model was used to account the effects of flow eddy viscosity.
A Parametric Study of Actuator Requirements for Active Turbine Tip Clearance Control of a Modern High Bypass Turbofan Engine
Authors: Kratz Jonathan L. , Chapman Jeffryes W. , Guo Ten-Huei
Published: August 2017

The efficiency of aircraft gas turbine engines is sensitive to the distance between the tips of its turbine blades and its shroud, which serves as its containment structure. Maintaining tighter clearance between these components has been shown to increase turbine efficiency, increase fuel efficiency, and reduce the turbine inlet temperature, and this correlates to a longer time-on-wing for the engine. Therefore, there is a desire to maintain a tight clearance in the turbine, which requires fast response active clearance control. Fast response active tip clearance control will require an actuator to modify the physical or effective tip clearance in the turbine. This paper evaluates the requirements of a generic active turbine tip clearance actuator for a modern commercial aircraft engine using the Commercial Modular Aero-Propulsion System Simulation 40k (C-MAPSS40k) software that has previously been integrated with a dynamic tip clearance model. A parametric study was performed in an attempt to evaluate requirements for control actuators in terms of bandwidth, rate limits, saturation limits, and deadband. Constraints on the weight of the actuation system and some considerations as to the force which the actuator must be capable of exerting and maintaining are also investigated. From the results, the relevant range of the evaluated actuator parameters can be extracted. Some additional discussion is provided on the challenges posed by the tip clearance control problem and the implications for future small core aircraft engines.
Integrated Large Eddy Simulation of Combustor and Turbine Interactions: Effect of Turbine Stage Inlet Condition
Authors: Duchaine Florent , Dombard Jérôme , Gicquel Laurent , Koupper Charlie
Published: August 2017

To study the effects of combustion chamber dynamics on a turbine stage aerodynamics and thermal loads, an integrated Large-Eddy Simulation of the FACTOR combustion chamber simulator along with its high pressure turbine stage is performed and compared to a standalone turbine stage computation operated under the same mean conditions. For this specific configuration, results illustrate that the aerodynamic expansion of the turbine stage is almost insensitive to the inlet turbulent conditions. However, the temperature distribution in the turbine passages as well as on the stator vane and rotor blade walls are highly impacted by these inlet conditions: underlying the importance of inlet conditions in turbine stage computations and the potential of integrated combustion chamber / turbine simulations in such a context.
Lean-Premixed, Swirl-Stabilized Flame Response: Flame Structure and Response As a Function of Confinement
Authors: De Rosa Alexander J. , Peluso Stephen J. , Quay Bryan D. , Santavicca Domenic A.
Published: August 2017

The effect of confinement (flame-wall interactions) on the response of a turbulent, swirl-stabilized flame is experimentally examined, with a focus on the shape and structure of these flames. A series of three cylindrical combustors of 0.11, 0.15 and 0.19 m diameter are used to vary the degree of confinement experienced by the flame. Using CH* chemiluminescence images, the shape of the flame in each combustor is described. These images are then further analyzed and reveal marked similarities in the geometry and location of these flames in a defined ‘flame base’ region near the combustor inlet. This similarity in location of the flame base leads to a similarity in the response of this portion of the flame to imposed oscillations. In particular, the phase of the fluctuations in this region are shown to be the same in each confinement. The nature of the fluctuations in the mean flame position are also shown to be similar in each confinement. These results indicate that the geometry of the flame in the base region is not a function of confinement and that the flames are responding to the same convective mechanisms, and in the same manner, in this region of the flame.
As-Built Geometry and Surface Finish Effects on Fatigue and Tensile Properties of Laser Fused Titanium 6Al-4V
Authors: Scott-Emuakpor Onome , George Tommy , Henry Emily , Holycross Casey , Brown Jeff
Published: August 2017

The as-built material behavior of additive manufactured (AM) Titanium (Ti) 6Al-4V is investigated in this study. A solution heat treated, aged, stress relieved, and hot isostatic pressed Laser Powder Bed Fusion (LPBF) AM process was used to manufacture the specimens of interest. The motivation behind this work is based on the ever-growing desire of aerospace system designers to use AM to fabricate components with novel geometries. Specifically, there is keen interest in AM components with complex internal cooling configurations such as turbine blades, nozzle vanes, and heat exchangers that can improve small scale propulsion performance. Though it is feasible to three-dimensionally print parts that meet the Fit portion of a part characteristic description and identification, the Form and Function portions have proven to be more difficult to conquer. This study addresses both the Form and Function characteristics of the LPBF AM process via the investigation of geometry variation and surface roughness effects pertaining to mechanical properties and fatigue behavior of Ti 6Al-4V. Results show that geometry variation may be the cause of increased vibration fatigue life uncertainty. Also, both fatigue and tensile properties show profound discrepancies associated with surface finish. As-built surface finish specimens have lower fatigue and ductility performance, but the results are more consistent than polished data.
The Impact of the Multiple Reference Frame Interface on Modelling the Interaction Between IGVs and the Impeller in Turbocharger Compressors
Authors: Li Xiangjun , Spence Stephen
Published: August 2017

In this paper numerical investigations are presented of how the axial position of the multiple reference frame (MRF) stator-rotor interface between the inlet guide vanes (IGVs) and the impeller would influence the predicted flow field for a turbocharger centrifugal compressor when simulated by the steady RANS method. In the first step, a total of three different axial positions of the MRF IGV-impeller interface were considered and compared with the results of an unsteady simulation to evaluate their accuracy. The results showed that the choice of the MRF interface location significantly influenced the predicted overall performance. At the lower rotational speed, the peak efficiency varied by 1.3% and the corresponding total pressure ratio varied by 0.022. At the high rotational speed, the different axial locations of the MRF interface varied the predicted choke point by 0.012 normalized mass flow rate. The mass flow rate of the near surge (NS) point was over estimated at both the high and low rotational speed by at least 0.038 normalized mass flow rate.

Consideration of the flow field suggested that the MRF interface between the IGV and the impeller should be placed towards the upstream side of the available region to avoid being unphysically influenced by its interaction with the non-uniform pressure in the downstream subsonic flow field and to enable a more accurate prediction of the extent of the inducer shock in transonic operating situations. Based on this understanding, a further improvement was made for the setting of the MRF interface by employing a polyline interface. This achieved a more accurate numerical result for the NS operating point at low rotational speeds. The position of the MRF interface for modelling IGVs in a turbocharger compressor should be suitably chosen according to the objectives of the numerical study.
Effect of Spanwise Variation of Chord on the Performance of a Turbine Cascade
Authors: Deshpande Srikanth , Pradeep A. M. , Thern Marcus , Genrup Magnus
Published: August 2017

This paper compares the aerodynamic performance of two cascade designs, viz.: - constant-chord and varying-chord. The varying-chord design is typical of industrial gas turbines and steam turbine stators in order to reduce manufacturing costs. The present study aims to increase the understanding of the implications of this manufacturing constraint on the aerodynamics of the stator. Experiments are carried out in a linear cascade wind tunnel. Numerical simulations are performed using commercial code CFX. The profile losses and secondary losses in the two designs are compared. The overall total pressure losses indicate better aerodynamic performance of a turbine cascade with constant chord as compared to a turbine cascade of varying-chord design.
Effects of a Circumferential Feed-Back Channel on Aerodynamic Performance of a Single-Stage Transonic Axial Compressor
Authors: Dinh Cong-Truong , Ma Sang-Bum , Kim Kwang-Yong
Published: August 2017

This paper presents an investigation of a circumferential feed-back channel located on shroud surface in rotor domain to find its effects on aerodynamic performance of a single-stage axial compressor, NASA Stage 37, using three-dimensional Reynolds-averaged Navier-Stokes equations. Validation of numerical results was performed using experimental data for both of single rotor and single-stage compressors. A parametric study of the feed-back channel was performed using various geometric parameters related to the locations and shapes of the channel inlet and outlet. The numerical results showed that a reference circumferential feed-back channel increased the stall margin by 26.8% with 0.14% reduction in the peak adiabatic efficiency, compared to the case without the feed-back channel.
Predicting the Temporal Progression of Aircraft Engine Compressor Performance Deterioration due to Particle Deposition
Authors: Döring Felix , Staudacher Stephan , Koch Christian
Published: August 2017

Aircraft engine performance deterioration due to particle deposition on compressor blading and end walls gradually progresses with increasing time of operation. Deposition effects can be mitigated by on-wing maintenance actions. Application of condition based maintenance strategies in order to minimize operating costs requires high-grade physical deterioration models. In previous work, a model and experimental setup were developed to quantify both magnitudes and timescales of deposition effects on blade row performance as a function of engine operating time. The model and experimental data published therein were now used to predict the deteriorated performance of aircraft engine high-pressure compressors. A given procedure to analytically derive stage characteristics from blade row data was altered in such a way as to account for stage performance deterioration through a set of modifiers. These time-dependent modifiers were calculated from the blade row performance deterioration model. As an example of application, the stage characteristics of the NASA Energy Efficient Engine high-pressure compressor were derived from published data. The stage characteristics were gradually modified according to the aforementioned approach. A stage-stacking procedure was then used to calculate the full map of the deteriorated compressor. Results showed a gradually progressing shift of compressor speed lines to lower mass flows and pressure ratios as well as a collapsing of efficiency isolines. The design point of the new engine and the point of maximum efficiency no longer coincide. Maximum efficiency decreased. The methodology presented enables aircraft engine operators to predict individual engine on-wing recoverable performance deterioration to optimize maintenance scheduling and, thus, reduce operating costs.
Transient Simulation of Critical Flow With Thermal-Hydraulic System Analysis Code for Supercritical CO2 Applications
Authors: Kim Min Seok , Oh Bong Seong , Kwon Jin Su , Jung Hwa-Young , Lee Jeong Ik
Published: August 2017

As a part of Sodium-cooled Fast Reactor (SFR) development, the supercritical CO2 (S-CO2) Brayton cycle is considered as an alternative power conversion system to eliminate sodium-water reaction (SWR) from the current conventional steam Rankine cycle is utilized with SFR. The leakage flow of S-CO2 from turbo-machinery via seal becomes one of important issues since not only it influences the cycle efficiency due to parasitic loss but also it is important for evaluating the system safety under various operating conditions. Thus, a transient simulation for estimating the critical flow in a turbo-machinery seal is essential to predict the leakage flow rate and calculate the required total mass of working fluid in a S-CO2 power system.

This paper briefly reviews the advantages of supercritical CO2 Brayton cycle relative to a typical steam Rankine cycle used in the Sodium-cooled Fast Reactor. In addition, this paper describes the test data using a CO2 critical flow experimental facility with three orifice configurations to model the flow resistance of a rotating shaft labyrinth seal. This data is used for validation of an existing transient analytical tool developed for transient hydraulic system analysis. Contained within this code is the analysis approach of Henry/Fauske from 1971 for two phase critical flow of one-component mixtures. This paper presents prediction of transient pressure, temperature, and flow profiles for critical and sub-critical flows of CO2 relative to blowdown test data to show that reasonable results are obtained. Similar analyses relative to test results of three orifice configurations are conducted and it shows that multiple orifices increase the time to equalize pressure in the blowdown system and therefore equates to higher flow resistance and lower leakage.
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.
Experimental Investigation of the Influence of Chamber Geometry on Bearing Chamber Oil Leakage
Authors: von Plehwe Felix C. , Brox Benedikt , Schwitzke Corina , Bauer Hans-Jörg
Published: August 2017

This paper presents experimental results identifying the influence of chamber geometry on the oil leakage behaviour of an aero engine bearing chamber. The varied geometrical parameters were the size of the bearing chamber, three different drip lip configurations, and a flinger. The position of the flinger was also varied to simulate mechanical and thermal distortions experienced during engine operations. Previous research has shown that the geometry of a bearing chamber influences the complex two phase flow inside it. This flow in turn influences the tendency of a bearing chamber to leak oil under adverse conditions. Since oil leakage from bearing chambers must be avoided for health and safety reasons, this knowledge is of great importance to the designer of a bearing chamber. The influence of the bearing chamber dimensions on oil leakage behaviour through seals has not yet been identified. The results of this work reveal a link between chamber geometry and leakage behaviour. The dependency of oil leakage rates for different chamber dimensions is also influenced by the investigated bearing chamber components. The complex interplay of these influencing factors was investigated and is described in detail.
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.
Effects of Real Gas Model Accuracy and Operating Conditions on Supercritical CO2 Compressor Performance and Flow Field
Authors: Ameli Alireza , Afzalifar Ali , Turunen-Saaresti Teemu , Backman Jari
Published: August 2017

Rankine and Brayton cycles are common energy conversion cycles and constitute the basis of a significant proportion of global electricity production. Even a seemingly marginal improvement in the efficiency of these cycles can considerably decrease the annual use of primary energy sources and bring a significant gain in power plant output. Recently, supercritical Brayton cycles using CO2 as the working fluid have attracted much attention, chiefly due to their high efficiency. As with conventional cycles, improving the compressor performance in supercritical cycles is major route to increasing the efficiency of the whole process. This paper numerically investigates the flow field and performance of a supercritical CO2 centrifugal compressor. A thermodynamic look-up table is coupled with the flow solver and the look-up table is systematically refined to take into account the large variation of thermodynamic properties in the vicinity of the critical point. Effects of different boundary and operating conditions are also discussed. It is shown that the compressor performance is highly sensitive to the look-up table resolution as well as the operating and boundary conditions near the critical point. Additionally, a method to overcome the difficulties of simulation close to the critical point is explained.
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.
Impact of Individual High-Pressure Turbine Rotor Purge Flows on Turbine Center Frame Aerodynamics
Authors: Zerobin S. , Aldrian C. , Peters A. , Heitmeir F. , Göttlich E.
Published: August 2017

This paper presents an experimental study of the impact of individual high-pressure turbine purge flows on the main flow in a downstream turbine center frame duct. Measurements were carried out in a product-representative one and a half stage turbine test setup, installed in the Transonic Test Turbine Facility at Graz University of Technology. The rig allows testing at engine-relevant flow conditions, matching Mach, Reynolds, and Strouhal number at the inlet of the turbine center frame. The reference case features four purge flows differing in flow rate, pressure, and temperature, injected through the hub and tip, forward and aft cavities of the high-pressure turbine rotor. To investigate the impact of each individual cooling flow on the flow evolution in the turbine center frame, the different purge flows were switched off one-by-one while holding the other three purge flow conditions. In total, this approach led to six different test conditions when including the reference case and the case without any purge flow ejection. Detailed measurements were carried out at the turbine center frame duct inlet and outlet for all six conditions and the post-processed results show that switching off one of the rotor case purge flows leads to an improved duct performance. In contrast, the duct exit flow is dominated by high pressure loss regions if the forward rotor hub purge flow is turned off. Without the aft rotor hub purge flow, a reduction in duct pressure loss is determined. The purge flows from the rotor aft cavities are demonstrated to play a particularly important role for the turbine center frame aerodynamic performance. In summary, this paper provides a first-time assessment of the impact of four different purge flows on the flow field and loss generation mechanisms in a state-of-the-art turbine center frame configuration. The outcomes of this work indicate that a high-pressure turbine purge flow reduction generally benefits turbine center frame performance. However, the forward rotor hub purge flow actually stabilizes the flow in the turbine center frame duct and reducing this purge flow can penalize turbine center frame performance. These particular high-pressure turbine/turbine center frame interactions should be taken into account whenever high-pressure turbine purge flow reductions are pursued.
Introduction of L12-Ordered Precipitation to Alumina-Forming Austenitic Heat-Resistant Steels With Low Ni Content
Authors: Zhao Bingbing , Dong Xianping , Sun Feng , Zhang Lanting
Published: August 2017

Alumina-forming austenitic (AFA) heat-resistant steels have been reported as a promising new class of steels in recent years with potential applications in advanced ultra-supercritical power plants. It is well known that L12-ordered γ’ phase is the most important precipitate for high-temperature strengthening in Ni-based superalloys and it can be stabilized by increasing the Ni content in heat-resistant steels. In the current work, the evolution of L12-ordered precipitates were compared in the Cu-bearing AFA alloys with 20, 27 and 35 wt.% Ni. After slow tensile tests at 700°C (∼2 × 10−5 s−1), L12-ordered precipitates occurred in all the alloys. Alloy AFA27 displayed the most densely distributed L12-particles in the matrix, whose ultimate tensile strength was also the highest. However, the L12-ordered precipitates were only observed in alloy AFA27 after the slow tensile test at 750°C due to the thermodynamic and kinetic reasons. Flow curves of slow tensile tests indicated different precipitation behaviors at 700°C and 750°C. Chemical composition analysis and thermodynamic calculation revealed that the occurrence of L12-ordered Ni-Cu-Al phase depends on temperature, Ni content and the atomic ratio of Ni/Al. This opens up new opportunities to promote the formation of L12-ordered phase in Fe-based austenitic heat-resistant steels with low Ni content and benefits high-temperature strengthening.
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.
Non-Linear Modeling of Centrifugal Stiffening Effects for Accurate Bladed Component Reduced-Order Models
Authors: Khalifeh Elias , Piollet Elsa , Millecamps Antoine , Batailly Alain
Published: August 2017

The modeling of centrifugal stiffening effects on bladed components is of primary importance in order to accurately capture their dynamics depending on the rotor angular speed. Centrifugal effects impact both the stiffness of the component and its geometry. In the context of the small perturbation framework, when considering a linear finite element model of the component, an assumption typically made in the scientific literature involves a fourth-order polynomial development of the stiffness matrix in terms of the angular speed. This polynomial development may fail to provide an accurate representation of the geometry evolution of a blade. Indeed, the error on the blade-tip displacement associated to the use of a linear finite element model quickly reaches the same order of magnitude as the blade-tip/casing clearance itself thus yielding a 100 % error on the blade-tip/casing clearance configuration. This article focuses on the presentation of a methodology that allows for creating accurate reduced order models of a 3D finite element model accounting for centrifugal stiffening with a very precise description of the blade-tip/casing clearance configuration throughout a given angular speed range. The quality of the obtained reduced order model is underlined before its numerical behaviour in the context of non-linear dynamic simulations be investigated. It is evidenced that the new reduced order model features specific interactions that could not be predicted with a linear model. In addition, results highlight the limitations of numerical predictions made for high angular speeds with a linear model. Finally, a particular attention is paid to the numerical sensitivity of the proposed model. As a downside of its increased accuracy, it is underlined that its computation must be done carefully in order to avoid numerical instabilities.
Development of a Last Stage Blade Row Coupled by Damping Elements: Numerical Assessment of its Vibrational Behavior and its Experimental Validation During Spin Pit Measurements
Authors: Siewert Christian , Sieverding Frank , McDonald William J. , Kumar Manish , McCracken James R.
Published: August 2017

Last stage blade rows of modern low pressure steam turbines are subjected to high static and dynamic loads. The static loads are primarily caused by the centrifugal forces due to the steam turbine’s rotational speed. Dynamic loads can be caused by instationary steam forces, for example. A primary goal in the design of modern and robust blade rows is to prevent High Cycle Fatigue caused by dynamic loads due to synchronous or non-synchronous excitation mechanisms. Therefore, it is important for the mechanical design process to predict the blade row’s vibration response. The vibration response level of a blade row can be limited by means of a damping element coupling concept. Damping elements are loosely assembled into pockets attached to the airfoils. The improvement in the blade row’s structural integrity is the key aspect in the use of a damping element blade coupling concept. In this paper, the vibrational behavior of a last stage blade row with damping elements is analyzed numerically. The calculation results are compared to results obtained from spin pit measurements for this last stage blade row coupled by damping elements.
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.
Pitfalls of Fan-Shaped Hole Design: Insights From Experimental Measurement of In-Hole Flow Through MRV
Authors: Issakhanian Emin , Elkins Christopher J. , Eaton John K.
Published: August 2017

Film cooling jets from discrete round holes are very susceptible to jet lift-off which reduces surface effectiveness. Since the experiments of Goldstein et al. (1974), shaped holes have become prominent for improved coolant coverage. Fan-shaped holes are the most common design and have shown good improvement over round holes. However, fan-shaped holes introduce additional parameters to the already complex task of modeling cooling effectiveness. This study presents velocity and vorticity fields measured using high-resolution magnetic resonance velocimetry (MRV) to study three different fan-shaped hole geome tries at two blowing ratios. Because MRV does not require line of sight, it provides otherwise hard to obtain experimental data of the flow within the film cooling hole in addition to the mainflow measurements. By allowing measurement within the cooling hole, MRV shows how poor choice of diffuser start point and angle can be detrimental to film cooling if overall hole length and cooling flow velocity are not properly accounted for in the design. The downstream effect of these choices on the jet height and counter-rotating vortex pair is also observed.