ASME Conference Presenter Attendance Policy and Archival Proceedings

2017;():V001T00A001. doi:10.1115/GT2017-NS1.

This online compilation of papers from the ASME Turbo Expo 2017: Turbomachinery Technical Conference and Exposition (GT2017) represents the archival version of the Conference Proceedings. According to ASME’s conference presenter attendance policy, if a paper is not presented at the Conference by an author of the paper, the paper will not be published in the official archival Proceedings, which are registered with the Library of Congress and are submitted for abstracting and indexing. The paper also will not be published in The ASME Digital Collection and may not be cited as a published paper.

Commentary by Dr. Valentin Fuster

Aircraft Engine

2017;():V001T01A001. doi:10.1115/GT2017-63010.

This work details the heat generation analysis of a turbine aero-engine main-shaft bearing using the computer program Advanced Dynamics Of Rolling Elements (ADORE). The empirical models used for traction and churning heat generation are detailed. The predictions of ADORE are shown to demonstrate the differing contributions of traction and churning to total heat generation at different load/speed regimes. These results are then compared with experimental results. In addition, the results of ADORE are also compared with results from the well-known bearing analysis program SHABERTH (Shaft Bearing Thermal Analysis). The comparisons showed good agreement between ADORE and the experimental results for loads between 13.35 and 53.40 kN and speeds between 1.8 and 2.2 MDN. The results also showed under prediction of heat generation by SHABERTH in this regime. Limitations of both programs were identified and speculated to include limitations in the empirical models due to the lack of available experimental traction data at high speeds/loads. Finally, recommendations for future research are provided which will likely provide significant improvements in the ability to predict bearing heat generation in turbine aero-engine applications.

Topics: Heat , Bearings
Commentary by Dr. Valentin Fuster
2017;():V001T01A002. doi:10.1115/GT2017-63072.

The purpose of an aircraft inlet system is to capture airflow from the free-stream and deliver it to an engine at the appropriate Mach number for that system. To meet design constraints, modern fighter aircraft have complex inlets with multiple turns that generally lead to both total pressure and swirl distortion at the engine face. These flow distortions can lead to reduced system performance, operability, and durability introducing issues in the overall success of the weapons system performing its mission. Therefore the integration of the airframe, inlet, and propulsion system is a key design issue in the development of military aircraft. The purpose of this paper is to demonstrate the dynamic (hybrid RANS/DDES) simulation capabilities of the HPCMP CREATE™-AV Kestrel tools by application to a sub-scale airframe/inlet system for a current military aircraft. The computational results were compared to wind tunnel results at various Mach number, angles of attack, angles of sideslip, and corrected flow rates. The inlet pressure recovery and distortion intensities for each case compared well to wind tunnel results. By comparing the computational results and wind tunnel test results, the applicability of these tools to future weapon systems design and development can be assessed.

Commentary by Dr. Valentin Fuster
2017;():V001T01A003. doi:10.1115/GT2017-63082.

In a Boundary Layer Ingesting (BLI) fan system the inlet flow field is highly non-uniform. In this environment, an axisymmetric stator design suffers from a non-uniform distribution of hub separations, increased wake thicknesses and casing losses. These additional loss sources can be reduced using a non-axisymmetric design that is tuned to the radial and circumferential flow variations at exit from the rotor.

In this paper a non-axisymmetric design approach is described for the stator of a low-speed BLI fan. Firstly sectional design changes are applied at each radial and circumferential location. Next, this approach is combined with the application of non-axisymmetric lean. The designs were tested computationally using full-annulus unsteady CFD of the complete fan stage with a representative inlet distortion. The final design has also been manufactured and tested experimentally.

The results show that a 2D sectional approach can be applied non-axisymmetrically to reduce incidence and diffusion factor at each location. This leads to reduced loss, particularly at the casing and midspan, but it does not eliminate the hub separations that are present within highly distorted regions of the annulus. These are relieved by non-axisymmetric lean where the pressure surface is inclined towards the hub. For the final design, the loss in the stator blades operating with BLI was measured to be 10% lower than for the original stator design operating with undistorted inflow. Overall, the results demonstrate that non-axisymmetric design has the potential to eliminate any additional loss in a BLI fan stator caused by the non-uniform ingested flow-field.

Commentary by Dr. Valentin Fuster
2017;():V001T01A004. doi:10.1115/GT2017-63208.

Reducing fluid dynamic power loss as speed increases is crucial to developing highly efficient high-speed aircraft engine gearing. Therefore, in this study, experiments were conducted using a precise friction-loss management technique and a vacuum pump in the gearbox for experimentally classifying fluid dynamic loss. Consequently, it was found that fluid dynamic loss could be classified into “oil-jet acceleration loss and oil reacceleration loss based on the momentum conservation of point mass” and “oil churning loss and windage loss based on the momentum conservation of an incompressible continuum”. Furthermore, the simulation results obtained via appropriate computational fluid dynamics (CFD) modeling of the resultant mechanisms agreed with the experimental results. The results of the present study are expected to improve the efficiency of mechanical systems, e.g., the fan drive gear system of Geared Turbofan™ and the accessory gearbox.

Commentary by Dr. Valentin Fuster
2017;():V001T01A005. doi:10.1115/GT2017-63277.

Maintenance on aircraft engines is usually performed on an on-condition basis. Monitoring the engine condition during operation is an important prerequisite to provide efficient maintenance. Engine Condition Monitoring (ECM) has thus become a standard procedure during operation. One of the most important parameters, the engine thrust, is not directly measured, however, and can therefore not be monitored, which makes it difficult to distinguish whether deteriorating trends e.g. in fuel comsumption must be attributed to the engine (e.g. due to thermodynamic wear) or to the aircraft (e.g. due to increased drag). Being able to make this distinction would improve troubleshooting and maintenance planning and thus help to reduce the cost of ownership of an aircraft. As part of the research project APOSEM (Advanced Prediction of Severity effects on Engine Maintenance), Lufthansa Technik (LHT) and the Institute of Jet Propulsion and Turbomachinery of Technische Universität Braunschweig develop a method for direct measurement of engine thrust during the operation. In this paper, the design process of the On-Wing (OW) Measurement System is presented, including the validation in labratory tests, the mechanical and thermal calibration as well as the final ground test during an engine test run at LHT test cell and the work on the flight test certification.

Commentary by Dr. Valentin Fuster
2017;():V001T01A006. doi:10.1115/GT2017-63320.

More electric and all electric aircraft were already discussed in the eighties of the last century, but recent political and ecological issues now reinforce the electrification of aircraft and engine systems. The development of electric machines and components with increasing power to weight ratio enables the installation of power optimized electric accessories instead of pneumatic and hydraulic systems in order to raise overall efficiency and specific fuel consumption of the engine. While pneumatic and hydraulic components are driven by the aircraft engine, a major challenge is in the supply of electric energy. Storage systems lack in reliability and light weight, fuel cell technology is limited to small aircraft and needs further development in various technical disciplines. An appropriate option is the generation of electric power by engine integrated generators. Performance calculations state increased efficiency by means of split spool power offtake, but have not been validated by a real twin-spool demonstrator yet.

At the ground test facility of the Institute of Jet Propulsion a demonstrator engine has been set up for detailed research on the influence of power extraction from a Larzac 04 C5 jet engine. To facilitate the test vehicle for power offtake of two spools the starter-generator has been complemented by a second generator, which is installed in front of the compressor inlet. It is axially connected to the low pressure spool by a coupling and a special flange mounted onto the low pressure spool. Several subsystems enabling for electric power offtake are integrated into the facilities’ data acquisition system (DAQ) and communication structure. The added components influence the engine in various ways: They manipulate power balance of the spools and alter the inlet pressure distribution and the compressor aerodynamics. Additionally the internal flow distribution is changed as well as the vibration characteristics. Before starting with extensive more electric engine (MEE) power offtake test campaigns, all systems need to be installed and tested successively.

This paper describes the test facility and fundamental more electric engine subsystems, with special focus put on instrumentation and system communication. A first function test demonstrates the operability of the engine after the modification of the low pressure spool. In a further step the influence of the inlet modification onto the compressor inlet aerodynamics, total mass flow, and vibrations of the test vehicle is analyzed. The vibration characteristics are vital for the coupling functionality, which is demonstrated subsequently. Presenting the load system check, special focus is given to communication, load definition, and electromagnetic compatibility. Comparisons to component performance predictions and to the performance of the original engine configuration are drawn for all tests and new limits for the operation of the new more electric configuration are defined. Finally, first data of power offtake of two spools is presented to demonstrate the operability of the MEE test vehicle.

Topics: Engines
Commentary by Dr. Valentin Fuster
2017;():V001T01A007. doi:10.1115/GT2017-63336.

Engine operating cost is a major contributor to the direct operating cost of aircraft. Therefore, the minimization of engine operating cost per flight-hour is a key aspect for airlines to operate successfully under challenging market conditions.

The interaction between maintenance cost, operating cost, asset value, lease and replacement cost describes the area of conflict in which engine fleets can be optimized.

State-of-the-art fleet management is based on advanced diagnostic and prognostic methods on engine and component level to provide optimized long-term removal and work-scoping forecasts on fleet level based on the individual operation. The key element of these methods is a digital twin of the active engines consisting of multilevel models of the engine and its components. This digital twin can be used to support deterioration and failure analysis, predict life consumption of critical parts and relate the specific operation of a customer to the real and expected condition of the engines on-wing and at induction to the shop. The fleet management data is constantly updated based on operational data sent from the engines as well as line maintenance and shop data.

The approach is illustrated along the real application on the CFM56-5C, a mature commercial two-spool high bypass engine installed on the Airbus A340-300. It can be shown, that the new methodology results in major improvements on the considered fleets.

Topics: Engines
Commentary by Dr. Valentin Fuster
2017;():V001T01A008. doi:10.1115/GT2017-63369.

The compressor aerodynamic design is conducted under the condition of clean inlet in general, but a compressor often operates under the condition of inlet distortion in the practical application. It has been proven by a lot of experimental and numerical investigations that inlet distortion can decrease the performance and stability of compressors. The circumferential or radial distorted inlet in mostly numerical investigations is made by changing the total pressure and total temperature in the inlet ring surface of the compressors. In most of inlet distortion experiments, distorted inlets are usually created by using wire net, flashboards, barriers or the generator of rotating distortion. The fashion of generating distorted inlet for experiment is different from that for numerical simulation. Consequently, the flow mechanism of affecting the flow field and stability of a compressor with distorted inlet for experiment is partly different than that for numerical simulation.

In the numerical work reported here, the inlet distortion is generated by setting some barriers in the inlet ring surface of an axial subsonic compressor rotor. Two kinds of distorted inlet are investigated to exploring the effect of distorted range on the flow field and stability of the compressor with ten-passage unsteady numerical method. The numerical results show that the inlet distortions not only degrade the total pressure and efficiency of the compressor rotor, but also decrease the stability of the rotor. The larger the range of distorted inlet is, the stronger the adverse effect is. The comprehensive stall margin for the inlet distortion of 24 degrees and 48 degrees of ten-passages is reduced about 3.35% and 5.88% respectively. The detailed analysis of the flow field in the compressor indicates that the blockage resulted from tip clearance leakage vortex (TLV) and the flow separation near the suction surfaces of some blades tip for distorted inlet is more serious than that resulted from TLV for clean inlet. Moreover, the larger the range of distorted inlet is, the larger the range of the blockage is. The analysis of unsteady flow shows that during this process, which is that one rotor blade passes through the region affected by the distorted inlet, the range of the blockage in the rotor passage increases first, then reduces, and increases last.

Commentary by Dr. Valentin Fuster
2017;():V001T01A009. doi:10.1115/GT2017-63427.

As an integrated system, turbofan engine airworthiness certification is a complex network because design, operating conditions and multi-disciplines are interlaced. Inlet compatibility specified in FAR regulations is to demonstrate satisfactory of engine operating characteristics throughout the flight envelope, which can be affected by engine installation and operation conditions. One limited operating condition is the high crosswind on the ground. Flow separated at engine inlet, unsteady and non-uniform, passing through the diffuser to the fan face, stimulated the fan blade at a broad frequency range, which could lead to high cycle fatigue.

A ground crosswind test was conducted by an airplane company to demonstrate the engine inlet compatibility with engine mounted on the rear of the aircraft under various crosswind conditions [1] including 90-degree crosswind, quarterly headwind (315-degree) and quarterly tailwind (225-degree). Results showed that among all tested ambient wind conditions, the engine was the least stable under quarterly tailwind (225-degree).

To predict the fan blade response driven by inlet separation, a process of evaluating inlet separation induced stimulus was illustrated in this paper. The stimuli were classified in two parts, i) synchronous stimulus induced by inlet distortion, and ii) non-synchronous stimulus induced by turbulence. Vibration of a wide-chord fan blade was evaluated by modal analysis and Campbell diagram. Test data of total pressure distortion at fan face were analyzed by Fast Fourier Transform (FFT), and the excitations in frequency domain were applied to fan blade for harmonic analysis. Results revealed that the synchronous excitation caused the blade resonating at an elevated stress level, as expected. This study provided a preliminary assessment and a better understanding of fan aeromechanics, when the engine is operating at the unsteady, unstable, and non-uniform flow environment. Discussions of how to control and how to decrease the vibration level were given in the study.

Topics: Engines , Turbofans
Commentary by Dr. Valentin Fuster
2017;():V001T01A010. doi:10.1115/GT2017-63440.

It is anticipated that the contribution of rotorcraft activities to the environmental impact of civil aviation will increase in the forthcoming future. Due to their versatility and robustness, helicopters are often operated in harsh environments with extreme ambient conditions and dusty air. These severe conditions affect not only the engine operation but also the performance of helicopter rotors. This impact is reflected in the fuel burn and pollutants emitted by the helicopter during a mission. The aim of this paper is to introduce an exhaustive methodology to quantify the influence of the environment in the mission fuel consumption and the associated emissions of nitrogen oxides (NOx). An Emergency Medical Service (EMS) and a Search and Rescue (SAR) mission were used as a case study to simulate the effects of extreme temperatures, high altitude and compressor degradation on a representative Twin-Engine Medium (TEM) weight helicopter, the Sikorsky UH-60A Black Hawk. A simulation tool for helicopter mission performance analysis developed and validated at Cranfield University was employed. This software comprises different modules that enable the analysis of helicopter flight dynamics, powerplant performance and exhaust emissions over a user defined flight path profile. The results obtained show that the environmental effects on mission fuel and emissions are mainly driven by the modification of the engine performance for the particular missions simulated. Fluctuations as high as 12% and 40% in mission fuel and NOx emissions, respectively, were observed under the environmental conditions simulated in the present study.

Topics: Pollution , Emissions
Commentary by Dr. Valentin Fuster
2017;():V001T01A011. doi:10.1115/GT2017-63461.

In the present paper, a transient performance code is employed to predict on-wing test data of the IAE-V2500 engine mounted on an Airbus A320-232. The test data was recorded by the engine control system and may serve as an open basis for validation of future transient studies. For the current investigation, the employed code considers the fundamental equations of the constant mass flow method as well as heat transfer effects by a lumped parameter approach.

The study focuses on seven accelerations and one deceleration. Engine test data was gathered with 10Hz sampling rate, imprinting the applied time step of the model. First, the steady-state matching of the test data was conducted. Subsequently, the measurement quantities fuel flow, inlet temperature and inlet pressure were prescribed as time-varying boundary conditions to the transient model. The results of the standard transient model and the model including thermal effects were compared with temperatures, pressures and shaft speeds. The LPT outlet temperature and the working line excursion in the booster map were examined in detail. The outcome concurs with the original statement that thermal effects are mandatory to enhance model accuracy. Lastly, a sensitivity analysis of the thermal input parameters was accomplished and its influence on model prediction investigated.

Commentary by Dr. Valentin Fuster
2017;():V001T01A012. doi:10.1115/GT2017-63523.

Auxiliary power unit (APU) operators face increasingly stricter airport requirements concerning exhaust gas and noise emission levels. To simultaneously reduce exhaust gas and noise emissions and to satisfy the increasing demand of electric power on board, optimization of the current technology is necessary. Prior to any possible demonstration of optimization potential, detailed data of thermodynamic properties and emissions have to be determined. Therefore, the investigations presented in this paper were conducted at a full-scale APU of an operational aircraft. A Pratt & Whitney APS3200, commonly installed in the Airbus A320 aircraft family, was used for measurements of the reference data. In order to describe the APS3200, the full spectrum of feasible power load and bleed air mass flow combinations were adjusted during the study. Their effect on different thermodynamic and performance properties, such as exhaust gas temperature, pressure as well as electric and overall efficiency is described. Furthermore, the mass flows of the inlet air, exhaust gas and fuel input were determined. Additionally, the work reports the exhaust gas emissions regarding the species CO2, CO and NOx as a function of load point. Moreover the acoustic noise emissions are presented and discussed. With the provided data the paper serves as a database for validating numerical simulations and provides a baseline for current APU technology.

Commentary by Dr. Valentin Fuster
2017;():V001T01A013. doi:10.1115/GT2017-63561.

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.

Topics: Bearings , Geometry , Leakage
Commentary by Dr. Valentin Fuster
2017;():V001T01A014. doi:10.1115/GT2017-63591.

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.

Commentary by Dr. Valentin Fuster
2017;():V001T01A015. doi:10.1115/GT2017-63721.

Inlet swirl distortion has become a major area of concern in the gas turbine engine community. Gas turbine engines are increasingly installed with more complicated and tortuous inlet systems, like those found on embedded installations on Unmanned Aerial Vehicles (UAVs) and Auxiliary Power Units (APU). These inlet systems can produce complex swirl distortion in addition to total pressure distortion.

The effect of swirl distortion on engine or compressor performance and operability must be evaluated. The gas turbine community is developing methodologies to measure and characterize swirl distortion. There is a strong need to develop a database containing the impact of a range of swirl distortion and total pressure distortion on compressor operability.

In one aircraft installation of an APU, the compressor experienced a wide range of combined total pressure and swirl distortion levels as a result of the APU inlet geometry. In this unique installation, the inlet system includes a check valve at the entrance of the inlet plenum through which all the airflow to the APU compressor passes. During APU operation, the check valve flaps are pushed open by the force of the airflow entering the plenum. The orientation of the check valve with respect to the compressor centerline introduces a varying amount of combined total pressure and swirl distortion. Consequently, it was necessary to test and map the compressor for a range of check valve positions in order to find an orientation with the lowest surge margin loss.

The check valve comprises of two “D” shaped flaps hinged at the diameter of a circular flow entrance to the plenum. The plenum forms a “U’ shaped volume which allows the flow to turn 90 degrees from the check valve into the compressor annulus. A series of compressor tests was conducted where the check valve was rotated on its base such that the check valve had various clocking orientations with respect to the compressor axis. The check valve orientation was varied by lining the flap hinge line parallel to the compressor centerline (defined as zero degrees) then rotating the check valve in 30 degree increments from this reference all the way up to 150 degrees. Compressor speed lines were mapped at each check valve orientation and the surge margin loss/gain from a nominal undistorted surge line was determined. Fully 3D CFD analyses of each of these check valve orientations were run. The CFD models included the full valve and plenum geometry up to the compressor annulus. From the CFD analysis, the ARP1420 (Reference 1) total pressure distortion descriptors and the AIR5686 (Reference 2) swirl distortion descriptors were calculated on the annular plane at the compressor eye.

Using these total pressure and swirl distortion descriptors determined for each check valve orientation along with the surge margin change available from the test data, a correlation of operability behavior was successfully obtained. This paper discusses the lessons learned and a recommended methodology todetermine an operability correlation for combined total pressure and high swirl distortion inlet systems.

Topics: Pressure , Compressors
Commentary by Dr. Valentin Fuster
2017;():V001T01A016. doi:10.1115/GT2017-63776.

This paper proposes a pulse detonation combustion (PDC) model integrated within Chalmers University’s gas turbine simulation tool GESTPAN (GEneral Stationary and Transient Propulsion ANalsysis). The model will support the development of novel aircraft engine architectures exploiting the synergies between intercooling, aftercooling and PDC. The proposed engine architectures are based on a reference high bypass ratio geared-turbofan engine model with performance levels estimated to be available by year 2050. Parametric studies have been carried out for each proposed advanced architecture, providing engine cycle mid-cruise design point parameters. Design sensitivity studies related to intercooling technology in combination with a PDC are further explored for a number of heat-exchanger design effectiveness values and associated pressure loss levels. The acquired results suggest that the incorporation of PDC technology within a conventional core has the potential to significantly improve engine thermal efficiency. Incorporating intercooling improves the cycle performance for any pre-combustion OPR above 10 and contributes to an increase in specific power over the entire range of OPR. Finally, the results demonstrate that aftercooling the high pressure compressor delivery air further improves core specific power, but cancels out any SFC and thermal efficiency benefits arising from pulse detonation.

Commentary by Dr. Valentin Fuster
2017;():V001T01A017. doi:10.1115/GT2017-63834.

The main purpose of this paper is to discuss the possibility of standard turbofan engine replacement by the turboelectric distributed propulsion system, in future commercial aviation. Paper describes how the distributed propulsion allows to reach significantly greater propulsive efficiency than state-of-the-art high bypass turbofan engines, and presents turboelectric system as the only practical method of distributed propulsion implementation. However, since extra weight of the electric components that would be added can overcome the high propulsive efficiency benefit, a detailed analysis is needed to verify the feasibility of such system. This article shows results of such analysis that was conducted for 90 PAX class regional jet. Thermodynamic cycle calculations, performed for both, turbofan engine and turboelectric distributed propulsion are presented. They prove that distributed propulsion is able to provide great reduction in fuel consumption of uninstalled propulsion system, while performed mission analysis depicts the penalty of extra mass of electric appliances, showing actual profits that are achievable. On this example, advantages and disadvantages of the turboelectric distributed propulsion system in comparison with modern turbofan engines are discuss, taking into account the potential technological development of turbofan engine and additional non-propulsive benefits that turboelectric system is able to provide. Finally, this document also presents mass estimations for different scenarios of electric appliances evolution, which highlight the technology levels that need to be achieved before the system can be introduced in commercial service.

Commentary by Dr. Valentin Fuster
2017;():V001T01A018. doi:10.1115/GT2017-63858.

For the statically indeterminate rotor structure of low-pressure rotor system in aero-engine, mechanics features of bearing deformation and rotor structure were analyzed, combined with the principle of virtual work, and the global stiffness matrix of the rotor system was deduced. According to the Lagrange method, the analytical dynamic model of statically indeterminate rotor system with three flexible supports based on the flexibility analysis was built. An analytical dynamic model of intermediate supporting misalignment was established on the basis of the model. The natural characteristic and unbalance response of aligned and misaligned statically indeterminate rotor system were obtained through the analytical solution, and the experimental verification was conducted. These findings indicate that intermediate bearing misalignment of statically indeterminate rotor system can not only produce the 2X component, but also make the 1X component change. The experimental results are essentially in agreement with the calculated ones. The study laid the foundations for model basis and experimental reference of the statically indeterminate rotor system design, dynamic modeling and vibration control.

Commentary by Dr. Valentin Fuster
2017;():V001T01A019. doi:10.1115/GT2017-63868.

Very high-bypass ratio turbofans with large fan tip diameter are an effective way of improving the propulsive efficiency of civil aero-engines. Such engines, however, require larger and heavier nacelles, which partially offset any gains in specific fuel consumptions. This drawback can be mitigated by adopting thinner walls for the nacelle and by shortening the intake section. This binds the success of very high-bypass ratio technologies to the problem of designing an intake with thin lips and short diffuser section which is well matched to a low speed fan. Consequently the prediction of the mutual influence between the fan and the intake flow represents a crucial step in the design process.

Considerable effort has been devoted in recent years to the study of models for the effects of the fan on the lip stall characteristics and the operability of the whole installation. The study of such models is motivated by the wish to avoid the costs incurred by full, three-dimensional CFD computations. The present contribution documents a fan model for fan-intake computations based on the solution of the double linearization problem for unsteady, transonic flow past a cascade of aerofoils with finite mean load. The computation of the flow in the intake is reduced to a steady problem, whereas the computation of the flow in the fan is reduced to one steady problem and a set of solutions of the linearised model in the frequency domain. The nature of the approximations introduced in the fan representation is such that numerical solutions can be computed inexpensively, whilst the main feature of the flow in the fan passage, namely the shock system and an approximation of the unsteady flow encountered by the fan are retained. The model is applied to a well-documented test case and compares favourably with much more expensive three-dimensional, time domain computations.

Commentary by Dr. Valentin Fuster
2017;():V001T01A020. doi:10.1115/GT2017-63977.

The accurate prediction of drag caused by bluff bodies present in aerospace applications, particularly at high angles of attack, was a challenge. An experimental and numerical investigation of a nacelle intended for fuselage-mounted aircraft engines was completed at several angles of attack between 0 and 45 deg with a Reynolds number of 6 × 105. Steady-flow simulations were conducted on hybrid grids using ANSYS Fluent 15.0 with preference given to the realizable k-ε turbulence model. Both total drag and the pressure-to-viscous drag ratio increased with angle of attack as a consequence of greater flow separation on the suction surface. Near-field and far-field drag predictions had grid uncertainties below 2.5% and were within 10% of experiment, which were less than the uncertainties of the respective force balance and outlet traverse data at all angles of attack. Regions were defined on suction-side x-pressure force plots using the validated CFD data-set that showed where and how much drag could be reduced. At 20 deg angle of attack, there was potential to reduce up to 20% drag contained within the separated flow region.

Commentary by Dr. Valentin Fuster
2017;():V001T01A021. doi:10.1115/GT2017-63978.

Decreasing drag on aircraft components was beneficial towards improving fuel economy and operational range. A generic axisymmetric nacelle-strut configuration typical of those housing fuselage-mounted engines was evaluated at a Reynolds number of 6 × 105 based on the nacelle maximum diameter d = 26.5 cm and an angle of attack of 20 deg. It was estimated that drag could be reduced by 20%. Three case studies were evaluated that added a fillet to the nacelle-strut corner, vortex generating triangular tabs, and flow-path obstructing vanes to improve flow control by reducing suction-side separation. Experimental results showed that a 0.11 d radius of curvature fillet reduced drag by 8% with respect to the baseline case. Numerical results employing the realizable k-ε turbulence model with wall functions predicted no improvements with the tabs and an 8% reduction with the vanes.

Commentary by Dr. Valentin Fuster
2017;():V001T01A022. doi:10.1115/GT2017-64030.

The original sump design of Rolls-Royce AE3007 central sump is a tangential sump. An experimental program at Purdue University was conducted to investigate the characteristics of a tangential sump design. The research employed a bespoke experimental rig consisting of modular transparent chamber to allow unprecedented view of the two-phase flow inside the bearing chamber. It was found that a persistent liquid pooling near the drain entrance of a tangential sump obstructs the outflow to the scavenge off-take. While space requirement is minimal, the scavenge performance of a tangential sump is poor. A prototype advanced sump was proposed and built with features to address various issues found in the tangential sump.

In this paper, further experimental work on the refinements of the advanced sump is presented. This includes experimental study on the effect of a fence and splitter plate, off-take hole diameter and its location, the effect of gravity by tilting, upstream wall fairing, as well as the depth of the sump itself. The experiments showed that the faired upstream wall modification can reduce residence volume significantly across different operating conditions, indicating improvement in scavenge performance. The other modifications help to reduce residence volume only in certain operating conditions. The results of this study helped towards the birth of an optimized sump design, known as Indy sump. The Indy sump has been accepted as a superior sump design and has been used as a benchmark in many sump design studies.

Topics: Design , Optimization
Commentary by Dr. Valentin Fuster
2017;():V001T01A023. doi:10.1115/GT2017-64214.

Enabling high overall pressure ratios, wave rotors and piston concepts seem to be solutions surpassing gas turbine efficiency. Therefore, a comparison of a wave rotor and three piston concepts relative to a reference gas turbine is offered. The piston concepts include a Wankel, a 2-stroke reciprocating engine and a free-piston. All concepts are investigated with and without intercooling. An additional combustion chamber downstream the piston engine is investigated, too. The shaft power chosen corresponds to large civil turbofans. Relative to the reference gas turbine a maximum efficiency increase of 11.2 percent for the piston concepts and 9.8 percent for the intercooled wave rotor is demonstrated. These improvements are contrasted by a 5.8 percent increase in the intercooled reference gas turbine and a 4.2 percent increase due to improved gas turbine component efficiencies. Intercooling the higher component efficiency gas turbine leads to a 9.8 percent efficiency increase. Furthermore, the study demonstrates the high difference between intercooler and piston engine weight and a conflict between piston concept efficiency and chamber volume, highlighting the need for extreme lightweight design in any piston engine solution. Improving piston engine technology parameters is demonstrated to lead to higher efficiency, but not to a chamber volume reduction. Heat loss in the piston engines is identified as the major efficiency limiter.

Commentary by Dr. Valentin Fuster
2017;():V001T01A024. doi:10.1115/GT2017-64379.

With the rapid development of high bypass ratio turbofan engine, the proportion of the nacelle drag increases obviously in the total drag of the aircraft with the increase of nacelle surface area. And the frictional resistance is one of the major contributors of drag. Under the same Reynolds number, the friction resistance in turbulent boundary layer is about 10 times larger as that in laminar boundary layer. Therefore, a correctly profiled engine nacelle will delay the transition in the boundary layer and allow laminar flow to extend back, resulting in a substantial drag reduction. In the previous conference paper (9th reference), a 2D nacelle longitudinal profile-line geometry generator, which allows curvature and slope-of-curvature to be continuous was developed and presented. This established an optimization system to minimize nacelle frictional drag. One of the nacelle profile-line is optimized to achieve minimum drag coefficient, and then is stacked with the other original profile-lines to form the 3D isolated nacelle aerodynamic shape. Finally, a total 23% of nacelle outer surface maintains a laminar flow and its frictional drag coefficient is less than initial shape.

This paper proposes a new 2D nacelle longitudinal profile-line design method, based on PARSEC parameterization, with can generate the profile-line rapidly and precisely. Conservation Full Potential Equation was used to calculate the aerodynamic distribution and obtain the transition location. Then adaptive simulated annealing genetic algorithm was adapted to search 2D profiles of low drag, which would be applied to narrow down design space in 3D nacelle optimization. Second, 2D profiles were stacked circumferentially, by NURBS surface generator, to form the 3D nacelle aerodynamic shape, and an optimization system was established, in combination with the 3D nacelle generator, γ–Reθ transition model, Kriging surrogate model and adaptive simulated annealing algorithm, for natural laminar flow nacelle design. Finally, a total 34% of nacelle surface maintains a laminar flow and its frictional drag coefficient is less than initial shape. The generated optimized loft was evaluated by CFD to determine if the low drag of this optimized nacelle shape can be maintained under different Mach numbers and angles of attack.

Commentary by Dr. Valentin Fuster
2017;():V001T01A025. doi:10.1115/GT2017-64510.

In this paper, a sliding mode (SM) parameter limit regulating system is designed to regulate the fuel flow rate to the turbofan engine. Firstly, a linear engine model is identified using a general engine dynamic nonlinear model. Then based on the one Lyapunov function, one SM parameter limit regulating system is designed mainly including regulators design, selector design and integrator design. After that the feedback gains and coefficient sets (switching gain and boundary thickness for every regulator) of the SM regulators are optimized and chosen. Finally, the global asymptotical stability of the regulating system is demonstrated. The simulation results also show that SM parameter limit regulator functions all the time during engine transient state control process, and the design SM parameter limit regulating system ensures that the target steady speed state or limit steady state can be attained in finite time interval without exceeding critical parameter limits.

Commentary by Dr. Valentin Fuster
2017;():V001T01A026. doi:10.1115/GT2017-64525.

The Inlet distortion, which may lead to the stability reduction or structure failure, is often non-ignorable in an axial compressor. In the paper, the three-dimensional unsteady numerical simulations on the flow in NASA rotor 67 are carried out to investigate the effect of inlet distortion on the performance and flow structure in a transonic axial compressor rotor. A sinusoidal circumferential total pressure distortion with eleven periods per revolution is adopted to study the interaction between the transonic rotor and inlet circumferential distortion. Concerning the computational expense, the flow in two rotor blade passages is calculated. Various intensities of the total pressure distortion are discussed, and the detailed flow structures under different rotating speeds near the peak efficiency condition are analyzed. It is found that the distortion has a positive effect on the flow near the hub. Even though there is no apparent decrease in the rotor efficiency or total pressure ratio, an obvious periodic loading exists over the whole blade. The blade loadings are concentrated in the region near the leading edge of the rotor blade or regions affected by the oscillating shocks near the pressure side. The time averaged location of shock structure changes little with the distortion, and the motion of shocks and the interactions between the shock and the boundary layer make a great contribution to the instability of the blade structure.

Topics: Compressors , Inflow
Commentary by Dr. Valentin Fuster
2017;():V001T01A027. doi:10.1115/GT2017-64545.

Leading-edge vortex flows are often present on propeller blades at take-off, however, their characteristics and aerodynamic impact are still not fully understood. An experimental investigation using Time Resolved Particle Image Velocimetry (TR-PIV) has been performed on a model blade in order to classify this flow with respect to both delta wing leading-edge vortices and the low Reynolds number studies regarding leading-edge vortices on rotating blades. A numerical calculation of the experimental setup has been performed in order to assess usual numerical methods for propeller performance prediction against TR-PIV results. Similar characteristics were found with non slender delta wing vortices at low incidence, which hints that the leading-edge vortex flow may generate vortex lift. The influence of rotation on the characteristics of the leading-edge vortex is compared to that of the pressure gradient caused by the circulation distribution. A discussion on the quality of the PIV reconstruction for close-wall structures is provided.

Commentary by Dr. Valentin Fuster
2017;():V001T01A028. doi:10.1115/GT2017-64612.

The effect of inlet distortion from curved intake ducts on jet engine fan stability is an important consideration for next generation passenger aircraft such as the boundary layer ingestion (BLI) Silent Aircraft. Highly complex inlet flows which occur can significantly affect fan stability. Future aircraft designs are likely to feature more severe inlet distortion, pressing the need to understand the important factors influencing design.

This paper presents the findings from a large CFD investigation into which aspects of inlet distortion cause the most significant reductions in stall margin and, therefore, which flow patterns should be targeted by mitigating technology. The study considers the following aspects of distortion commonly observed in intakes: steady vortical distortion due to secondary flow, unsteady vortical distortion due to vortex shedding and mixing, static pressure distortion due to curved streamlines, and low momentum endwall flow due to thickened boundary layers or separation. Unsteady CFD was used to determine the stall points of a multipassage transonic rotor geometry with each of the inlet distortion patterns applied. Interesting new evidence is provided which suggests that low momentum flow in the tip region, rather than distortion in the main body of the flow, leads to damaging instability.

Topics: Rotors
Commentary by Dr. Valentin Fuster
2017;():V001T01A029. doi:10.1115/GT2017-65031.

For fan/compressor design, quantifying distortion transfer and generation bladerow by bladerow through a fan/compressor is important to understand the flow physics and predict performance. What is needed are descriptors capable of describing distortion profiles with both high and low distortion content and account for the reshaping of distortion profiles. Four key parameters were identified as desirable to quantitatively capture distortion transfer, generation and effects on performance: distortion magnitude, shape, severity and phase. A set of distortion descriptors based on Fourier analysis are shown to quantitatively capture distortion magnitude, shape and phase change across bladerows. These Fourier descriptors are modal amplitude, total amplitude, and phase shift. When used together, these Fourier descriptors can be used to qualitatively describe any conceivable profile shape for any parameter.

Commentary by Dr. Valentin Fuster
2017;():V001T01A030. doi:10.1115/GT2017-65044.

In this paper the use of alternative fuels — synthetic kerosene based — were investigated in a small aircraft gas turbine engine to determine their sound and vibrations signature at Georgia Southern University’s Aerospace Engine Laboratory. Three types of fuels were used: a natural gas, a coal derived synthetic kerosene and Jet-A as a reference fuel. The alternative fuels are Fischer-Tropsch process fuels. Synthetic fuels are attractive in the aviation industry because of their potential for reducing energy dependence and the growing need for higher efficiencies, while reducing emissions.

The research SR-30 gas turbine was used to evaluate combustion noise from the selected fuels. The turbine can operate at a maximum speed of 80,000 rpm, produce a maximum thrust of 40 lbf, has a pressure ratio of 3.4 to 1, and a specific fuel consumption of 1.22 lbfuel/thrust per hour. The measurement transducers were state of the art Brüel & Kjaer ¼ inch microphone and a triaxial accelerometer which were interfaced with the analysis software, Brüel & Kjaer Pulse 21. Data was taken at 70,000 rpm for a time span of five minutes at stabilized conditions. One-third octave analysis indicated highest sound difference between the fuels at 400 Hz, and a more pure sound level was obtained. The difference for Jet A was as much as 6 dBA higher for exhaust measurement at 400 Hz, and for FT-NG the exhaust had a higher decibel reading from the exhaust side by as much as 4.4 dBA at the 400 Hz frequency, and FT-CG had as much as 7.3 dBA difference at 315 Hz. The sound from the combustion of these fuels produced the differences of sound when the positions of the microphones changed. However, with this new introduction of sound measurement locations and angles, the combustion differences are still distinct and show a more in depth view of the sound quality that is being received by the microphone.

Commentary by Dr. Valentin Fuster
2017;():V001T01A031. doi:10.1115/GT2017-65218.

An exergy framework was developed taking into consideration a detailed analysis of the heat exchanger (intercooler) component irreversibilities. Moreover, it was further extended to include an adequate formulation for closed systems, e.g. a secondary cycle, moving with the aircraft. Afterwards the proposed framework was employed to study two radical intercooling concepts. The first proposed concept uses already available wetted surfaces, i.e. nacelle surfaces, to reject the core heat and contribute to an overall drag reduction. The second concept uses the rejected core heat to power a secondary organic Rankine cycle and produces useful power to the aircraft-engine system. Both radical concepts are integrated into a high bypass ratio turbofan engine, with technology levels assumed to be available by year 2025. A reference intercooled cycle incorporating a heat exchanger in the bypass duct is established for comparison. Results indicate that the radical intercooling concepts studied in this paper show similar performance levels to the reference cycle. This is mainly due to higher irreversibility rates created during the heat exchange process. A detailed assessment of the irreversibility contributors, including the considered heat exchangers and the secondary cycle major components is made. A striking strength of the present analysis is the assessment of the component irreversibility rate and its contribution to the overall aero-engine losses.

Commentary by Dr. Valentin Fuster

Fans and Blowers

2017;():V001T09A001. doi:10.1115/GT2017-63172.

Fans are main components e.g. in heating, ventilating and air conditioning systems for vehicles or buildings, cooling units of engines and electronic circuits, and household appliances such as kitchen exhaust hoods or vacuum cleaners. End-users increasingly demand a high sound quality of their system or device. The overall objective of a recent research project at the University of Siegen is a multidimensional assessment of fan sound quality.

In a first step an advanced novel semantic differential for the assessment of fan-related sounds is established with the aid of carefully designed jury tests. Eventually, this semantic differential is employed for sound quality jury tests of fans in kitchen exhaust hoods, heat pumps and air purifiers as a first case. Finally, a prediction model is suggested, which relates the outcome from the jury tests to objective metrics.

A principal component analysis is carried out and yields five main assessment criteria with 23 relevant adjective scales. The results show that the perceived sound quality of fan systems is mainly determined by the loudness and tonality of the sound. The spectral content (represented by the sharpness) as well as the time structure (represented by the roughness) have no significant impact on perceived sound quality of the fan systems investigated.

Topics: Sound quality
Commentary by Dr. Valentin Fuster
2017;():V001T09A002. doi:10.1115/GT2017-63331.

An axial flow fan design methodology is developed to design large diameter, low pressure rise, rotor-only fans for large air-cooled heat exchangers. The procedure aims to design highly efficient axial flow fans that perform well when subjected to off design conditions commonly encountered in air-cooled heat exchangers. The procedure makes use of several optimisation steps in order to achieve this. These steps include optimising the hub-tip ratio, vortex distribution, blading and aerofoil camber distributions in order to attain maximum total-to-static efficiency at the design point.

In order to validate the design procedure a 24 ft, 8 bladed axial flow fan is designed to the specifications required for an air-cooled heat exchanger for a concentrated solar power (CSP) plant. The designed fan is numerically evaluated using both a modified version of the actuator disk model and a three dimensional periodic fan blade model. The results of these CFD simulations are used to evaluate the design procedure by comparing the fan performance characteristic data to the design specification and values calculated by the design code. The flow field directly down stream of the fan is also analysed in order to evaluate how closely the numerically predicted flow field matches the designed flow field, as well as determine whether the assumptions made in the design procedure are reasonable.

The fan is found to meet the required pressure rise, however the fan total-to-static efficiency is found to be lower than estimated during the design process. The actuator disk model is found to under estimate the power consumption of the fan, however the actuator disk model does provide a reasonable estimate of the exit flow conditions as well as the total-to-static pressure characteristic of the fan.

Commentary by Dr. Valentin Fuster
2017;():V001T09A003. doi:10.1115/GT2017-63680.

The acoustical characteristics of cooling fans are an essential criterion of product quality in the automotive industry. Fan modules have to suffice growing customer expectations which are reflected in the comfort requirements set by car manufacturers around the world. In order to locate dominant acoustic sources and to reduce the noise emission generated by a shrouded fan configuration, numerical simulations and experimental investigations are performed. The working approach considers variously modified fan geometries and their evaluation regarding arising vortex flow phenomena and their effect on a decreased sound pressure level (SPL) in consideration of an improvement or the constancy of aerodynamic fan performance. Particular emphasis lies on the analysis of secondary flows in the blade tip region by post-processing CFD-results. Due to the large number of geometrical modifications investigated and the importance of highly resolved eddy structures, a hybrid approach is chosen by applying the SAS-SST turbulence model in URANS simulations. The SAS (Scale Adaptive Simulation) delivers LES (Large Eddy Simulation) content in unsteady regions of a RANS-simulation and exhibits not nearly the high computational effort needed to perform a full scale LES. An assessment of the actual propagation of noise emission into the far-field is made by performing experimental investigations on the most promising modifications. The acoustic measurements are carried out in a fan test stand in the anechoic chamber of Duesseldorf University of Applied Sciences. The aerodynamic performance is measured in a fan test rig with an inlet chamber setup in accordance to ISO 5801. The measured acoustical and aerodynamic performances are validated by the industrial partner. The results of the acoustic measurements are in turn utilized to determine indicators of noise radiation in the numerical simulation. Within this work an innovative geometry modification is presented which can be implemented into shrouded fan configurations with backward-skewed blades. The new design exhibits a reduced SPL (A-weighted) of approx. 4 dB over the entire operating range while showing no significant deterioration on the aerodynamic performance. While the design was registered for patent approval cooperatively by the industrial partner and Duesseldorf University of Applied Sciences, further investigations regarding variations of design parameters are performed and presented in this paper. All numerical simulations are performed with ANSYS CFX, a commercial solver widely spread in the industry. Methods similar to those shown in this work can be implemented in the design phase of axial fans in order to develop acoustically optimized fan geometries.

Topics: Simulation , Blades
Commentary by Dr. Valentin Fuster
2017;():V001T09A004. doi:10.1115/GT2017-63795.

The flow path close to the suction side of fan rotor blades mostly affects the overall drag of the blading. The blade lift is affected as well because of the separation of the low energy boundary layer that drives the blade into stall at low fan flow rates.

Forward sweep allows to position the airfoil sections of blades featuring a positive circulation gradient along the span so that they “accompany” the near-wall flow trajectories at the blade suction side. So, rotor efficiency and stall margin of the fan can be improved. On the other hand, blade end effects play a relevant role in high hub-to-tip and low aspect ratio rotors and may compromise the effectiveness of forward sweep. Nevertheless, some authors in the literature stated the beneficial contribution of changing the sweep angle at the ends of the blade both at design and off-design conditions.

The paper studies the end effects on constant-swirl design rotors by means of CFD simulations focusing on the distribution of blade sweep in the near-tip region. In particular, the performance and efficiency calculated for a forward swept tube-axial fan featuring a hub-to-tip ratio equal to 0.4 are compared with those estimated for the corresponding unswept fan at equal duty point. Several modifications of the sweep distribution in the blade tip region are considered in the swept fan to quantify their effect on performance, efficiency and stall margin.

Results show that the addition of up to 6 degrees of local forward sweep at the blade tip to the unswept blading does not affect fan pressure at design operation.

On the other hand, this local increase of the sweep angle allows for a very notable increase of the peak pressure and efficiency at flow rates close to stall inception.

Topics: Fans , Blades
Commentary by Dr. Valentin Fuster
2017;():V001T09A005. doi:10.1115/GT2017-63952.

Here we present a Synthetic Blade Model (SBM) for axial flow fans, derived from Actuator Disk and Actuator Line theories. This new approach is able to model the momentum exchange between the fan and the fluid by adding source terms into momentum equation, like an actuator disk. However, the model accounts for the position of the blades, their rotation and the non-uniform distribution of deflection capability in the blade-to-blade passage, like in an actuator line model.

This approach is derived, described and validated against available data on a reversible tunnel and metro fan.

Commentary by Dr. Valentin Fuster
2017;():V001T09A006. doi:10.1115/GT2017-63965.

The present research focuses on the efficiency improvement at part-load of a centrifugal fan for a 30 kW fuel cell combined heat and power (CHP) unit. For this purpose, the fan stage is equipped with a partially vaned diffuser with a variable cross-sectional area using a moving backplate.

The design and the performance of the partially vaned diffuser with a variable cross-sectional area are described in the first part of this paper. The performance results are compared to measurements of the same centrifugal fan with a vaneless diffuser carried out for the previous investigation. For the second part, the influence of the variable cross-sectional area on the diffuser flow field is investigated using optical PIV (Particle Image Velocimetry) measurements and CFD (Computational Fluid Dynamics) simulations.

The combination of a variable cross-section, partially vaned diffuser was able to achieve a 10 percent increase in pressure ratio, a 5 percentage points increase in part-load efficiency while maintaining the whole operating range of the vaneless, constant cross section reference design.

Topics: Diffusers , Fans
Commentary by Dr. Valentin Fuster
2017;():V001T09A007. doi:10.1115/GT2017-64032.

More than half a century has been spent by technicians involved in fans testing to suggest and fix best practice procedures including also the definitions of several performance and efficiency parameters (e.g., ISO, AMCA, ASHRAE Standards...). However, the huge amount of energy used by ventilation systems and the recent introduction of the stringent European Community regulations on Energy Related Products, suggests to spend additional time on the ancient question about which and how many should be the parameters best suited to classify the performance of fans and their capability to properly use the energy input. This also because there are some attempts to fix the fan total-to-static pressure rise and the related total-to-static efficiency as the only two parameters required to assess the fan quality.

Starting from basic thermodynamic principles, this paper tries to shed light on the parameters that are best suited to assess fan performance and efficiency.

The general layout of a ventilation system is discussed to clearly show which and how many are the parameters required for the optimal matching between fan and system. Finally, some comments on practical advantages and drawbacks in the operation of the fan at the total-to-static best efficiency point are presented as well.

Topics: Fans
Commentary by Dr. Valentin Fuster
2017;():V001T09A008. doi:10.1115/GT2017-64276.

Some methodological differences exist among the design methods for axial-flow fans. These differences generate confusion in the mind of the unexperienced fan designer, who is unaware of which method ensures the achievement of the required pressure rise at highest efficiency. In this work three important differences that appear comparing the classic methods of fan blading design are highlighted and analyzed: i) the choice of the airfoil, ii) the choice of the solidity distribution, and iii) the computation of the stagger angle of the blade elements. A fourth aspect regards the selection of the rotor number of blades. This aspect is treated in relation to the dubious applicability of the drag annulus correlation by Howell to low hub-to-tip ratio fan design and analysis. CFD simulations are performed on three case-study rotor-only fans, comparing blades systematically designed varying the airfoils (British C4 vs American NACA-65), the solidity distribution (Diffusion Factor criterion against the arbitrary selection of the blade-element lift coefficient) and the computation of the stagger angle (with respect to the inlet flow velocity or the mean one). The accuracy of XFOIL-predicted data at low Reynolds number (e.g., 300000) in designing small-to-medium fans is discussed as well. For each of the previous design aspects, results suggest the best indication among those suggested in the classic literature to achieve fan requirements.

Commentary by Dr. Valentin Fuster
2017;():V001T09A009. doi:10.1115/GT2017-64302.

South Africa’s coal-fired power stations use super heated steam to drive generator turbines. In arid regions, air-cooled condensers (ACCs) are used to condense the process steam. These ACCs consists of an array of over 200 axial flow fans, each driven by a motor via a reduction gearbox. Distorted fan inlet air flow conditions cause transient blade loading, which results in variations in output shaft bending and torque. A measurement project was conducted where the input and output shaft of such a gearbox were instrumented with strain gauges and wireless bridge amplifiers. Gearbox shaft speed and vibration were also measured. Torsional and bending strains were measured for a variety of operational conditions, where correlations were seen between gearbox loading and wind conditions. The input side experienced no unexpected loads from the motor or changing wind conditions, whereas output shaft loading was influenced by the latter. Digital filters were applied to identify specific bending components, such as the influence of fan hub misalignment and dynamic blade loading. Reverse loading of the gearbox was measured during the fan stop period, and vibration analysis revealed torsional and gearbox vibrations. This investigation documented reliable full scale ACC gearbox loads.

Commentary by Dr. Valentin Fuster
2017;():V001T09A010. doi:10.1115/GT2017-64417.

One of the main design decisions in the development of low-speed axial fans is the right choice of the blade loading versus rotational speed, since a target pressure rise could either be achieved with a slow spinning fan and high blade loading or a fast spinning fan with less flow turning in the blade passages. Both the blade loading and the fan speed have an influence on the fan performance and the fan acoustics and there is a need to find the optimum choice in order to maximize efficiency while minimizing noise emissions. The present paper addresses this problem by investigating five different fans with the same pressure rise but different rotational speeds in the design point. In the first part of the numerical study, the fan design is described and steady-state Reynolds-averaged Navier-Stokes (RANS) simulations are conducted in order to identify the performance of the fans in the design point and in off-design conditions. The investigations show the existence of an optimum in rotational speed regarding fan efficiency and identify a flow separation on the hub causing a deflection of the outflow in radial direction as the main loss source for slow spinning fans with high blade loadings. Subsequently, Large Eddy Simulations (LES) along with the acoustic analogy of Ffowcs Williams and Hawkings (FW-H) are performed in the design point to identify the main noise sources and to determine the far-field acoustics. The identification of the noise sources within the fans in the near-field is performed with the help of the power spectral density of the pressure. In the far-field, the sound power level is computed using different parts of the fan surface as FW-H sources. Both methods show the same trends regarding noise emissions and allow for a localization of the noise sources. The flow separation on the hub is one of the main noise sources along with the tip vortex with an increase in its strength towards lower rotational speeds and higher loading. Furthermore, a horseshoe vortex detaching from the rotor leading edge and impinging on the pressure side as well as the turbulent boundary layer on the suction side represent significant noise sources. In the present investigation, the maximum in efficiency coincides with the minimum in noise emissions.

Topics: Design , Fans
Commentary by Dr. Valentin Fuster
2017;():V001T09A011. doi:10.1115/GT2017-64517.

The European Union imposed minimum industrial fan efficiency levels in 2013 and then increased them in 2015. In the USA, the Department of Energy (DoE) is also developing regulations aimed at eliminating inefficient industrial fans from the market by 2023. A consequence of this regulatory activity is a need to apply design methods originally developed within the aerospace community to the design of high efficiency industrial fans.

In this paper, we present a process used to design, numerically verify and experimentally test a high-pressure single-stage axial fan. The goal was a fan design capable of working over a range of blade angles in combination with a single fixed cambered plate stator. We present the process used when selecting blade airfoil sections and the vortex distribution along the blade span. The selected methodology is based on a coupling between the aerodynamic response of each blade profile and the chosen vortex distribution, creating a direct link between the load distribution and the aerodynamic capability of the blade profile section. This link is used to develop radial distributions of blade twist and chord for the selected blade profiles that result in the required radial work distribution.

The design method has been enhanced through intermediate verifications using two different numerical methodologies. The methodologies are based on different approaches, in so doing providing confidence in the verification process. The final blade design has been analyzed using a three-dimensional computational fluid dynamic (CFD) code. Results of the CFD analysis indicate that performance of the final blade design is consistent with the design specifications.

The paper concludes with a comparison between predicted and experimentally measured performance. The need is clarified for balance between computational and empirical approaches. When used together the development effort results in a lower cost and higher efficiency design than would have been possible using either approach in isolation.

Topics: Design
Commentary by Dr. Valentin Fuster
2017;():V001T09A012. doi:10.1115/GT2017-64644.

Numerical simulation tools are acquiring a crucial role in the virtual prototyping of new fan rotor blades. This, considering also the possibility in using new advanced materials, is opening the design capability to blades of longer, lighter and more slender structure.

On this perspective, the numerical tools must be improved in order to catch also the non-linear and coupling phenomena which were been treated until now using a weak coupling approach.

Here we present a fluid-structure interaction solver based on the finite element method, through its application to an existing fan blade. The study will show how, even in case of metal structure with little deformation, the coupling between fluid dynamics and structure dynamics can produce effects on both the fluid and the solid involved in the machine main process. The possibility to simulate directly both the aerodynamic and the effective structure response opens the way to improved design capabilities, avoiding time waste and costs due to long experimental testing campaign and over-dimensioned structures.

Commentary by Dr. Valentin Fuster
2017;():V001T09A013. doi:10.1115/GT2017-64679.

The concept of morphing geometry to control and stabilize the flow has been proposed and applied in several aeronautic and wind turbine applications. We studied the effect of a similar passive system applied on an axial fan blade, analysing potential benefits and disadvantages associated to the passive coupling between fluid and structure dynamics.

The present work completes a previous study made at the section level, giving a view also on the three-dimensional effects.

We use the numerical computation to simulate the system which defines a complex fluid-structure interaction problem. In order to do that, an in-house finite element solver, already used in the previous study, is applied to solve the coupled dynamics.

Topics: Testing , Blades
Commentary by Dr. Valentin Fuster


2017;():V001T25A001. doi:10.1115/GT2017-63043.

The marine gas turbine exhaust volute is an important component that connects a power turbine and an exhaust system, and it is of great importance to the overall performance of the gas turbine. Gases exhausted from the power turbine are expanded and deflected 90 degrees in the exhaust volute, and then discharge radially into the exhaust system. The flows in the power turbine and the nonaxisymmetric exhaust volute are closely coupled and inherently unsteady. The flow interactions between the power turbine and the exhaust volute have a significant influence on the shrouded rotor blade aerodynamic forces. However, the interactions have not been taken into account properly in current power turbine design approaches.

The present study aims to investigate the flow interactions between the last stage of a shrouded power turbine and the nonaxisymmetric exhaust volute with struts. Special attention is given to the coupled aerodynamics and pressure response studies. This work was carried out using coupled computational fluid dynamics (CFD) simulations with the computational domain including a stator vane, 76 shrouded rotor blades, 9 struts and an exhaust volute. Three-dimensional (3D) unsteady and steady Reynolds-averaged Navier-Stokes (RANS) solutions in conjunction with a Spalart-Allmaras turbulence model are utilized to investigate the aerodynamic characteristics of shrouded rotors and an exhaust volute using a commercial CFD software ANSYS Fluent 14.0. The asymmetric flow fields are analyzed in detail; as are the unsteady pressures on the shrouded rotor blade. In addition, the unsteady total pressures at the volute outlet is also analyzed without consideration of the upstream turbine effects.

Results show that the flows in the nonaxisymmetric exhaust volute are inherently unsteady; for the studied turbine-exhaust configuration the nonaxisymmetric back-pressure induced by the downstream volute leads to the local flow varying for each shrouded blade and low frequency fluctuations in the blade force. Detailed results from this investigation are presented and discussed in this paper.

Commentary by Dr. Valentin Fuster
2017;():V001T25A002. doi:10.1115/GT2017-63176.

Gas turbine engines are widely used as the marine main power system. However, they can’t reverse like diesel engine. If the reversal is realized, other ways must be adopted, for example, controllable pitch propeller (CPP) and reversible gearing.

Although CPP has widespread use, the actuator installation inside the hub of the propeller lead to the decrease in efficiency, and it takes one minute to switch “full speed ahead” to “full speed astern”. In addition, some devices need to be added for the reversible gearing, and it takes five minutes to switch from “full speed ahead” “to “full speed astern”.

Based on the gas turbine engine itself, a reversible gas turbine engine is proposed, which can rotate positively or reversely. Most important of all, reversible gas turbine engine can realize operating states of “full speed ahead”, “full speed astern“ and “stop propeller”. And, it just takes half of one minute to switch “full speed ahead” to “full speed astern”.

Since reversible gas turbine engines have compensating advantages, and especially in recent years computational fluid dynamics (CFD) technology and turbine gas-dynamics design level develop rapidly, reversible gas turbine engines will be a good direction for ship astern.

In this paper, the power turbine of a marine gas turbine engine was redesigned by three dimensional shape modification, and the flow field is analyzed using CFD, in order to redesign into a reverse turbine. The last stage vanes and blades of this power turbine were changed to double-layer structure. That is, the outer one is reversible turbine, while the inner is the ahead one. Note that their rotational directions are opposite. In order to realize switching between rotation ahead and rotation astern, switching devices were designed, which locate in the duct between the low pressure turbine and power turbine. Moreover, In order to reduce the blade windage loss caused by the reversible turbine during working ahead, baffle plates were used before and after the reversible rotor blades.

This paper mainly studied how to increase the efficiency of the reversible turbine stage, the torque change under different operating conditions, rotational speed and rotational directions, and flow field under typical operating conditions. A perfect profile is expected to provide for reversible power turbine, and it can decrease the blade windage loss, and increase the efficiency of the whole gas turbine engine. Overall, the efficiency of the newly designed reversible turbine is up to 85.7%, and the output power is more than 10 MW, which can meet requirements of no less than 30% power of rated condition. Most importantly, the shaft is not over torque under all ahead and astern conditions. Detailed results about these are presented and discussed in the paper.

Commentary by Dr. Valentin Fuster
2017;():V001T25A003. doi:10.1115/GT2017-63346.

The Auxiliary Ships and New Acquisition Support Branch (Code 425) of the Naval Surface Warfare Center, Philadelphia Division conducted a study to assist the Marine Corps Systems Command in assessing the feasibility of using a gas turbine engine as a propulsion system on future United States Marine Corps Amphibious Combat Vehicles (ACV). The study was focused on developing and testing a gas turbine intake solution for the ACV that can remove saltwater from the intake airstream of a notional 3,000 horsepower ACV engine.

Code 425 developed a two-part solution for the intake of the ACV. The first part of the solution is an intake shroud designed to elevate the intake to protect the engine from deck water wash. The second part of the solution is the Combustion Air Separation System (CASS), a gas turbine intake filtration system designed to remove marine contaminants that enter the intake. Code 425 tested a CASS prototype for its efficiency at removing saltwater spray and bulk water up to 10 gallons per minute. Test results showed that the CASS met each requirement and that an ACV intake system incorporating both the intake shroud and the CASS should protect the gas turbine engine from saltwater ingestion.

Commentary by Dr. Valentin Fuster
2017;():V001T25A004. doi:10.1115/GT2017-63503.

As a high performance gas turbine, GT28 combines with a two-spool gas generator and a free power turbine. Under the condition of ISO, its power and efficiency are 28MW and 37% for marine mechanical propulsion, respectively. Considering the design characteristics and operating performance of GT28 gas turbine can meet the requirements of many marine propulsion, mechanical driven and electrical power generation, and this paper introduces the potential application of GT28 gas turbine in different industrial and marine fields. On this basis, the related key technologies are discussed briefly. Finally, a derivative network is presented to describe the relationships of different application and development of GT28 gas turbine.

Topics: Gas turbines
Commentary by Dr. Valentin Fuster
2017;():V001T25A005. doi:10.1115/GT2017-63651.

The United States Navy has successfully operated their Landing Craft Air Cushion (LCAC) with Vericor’s ETF40B engines since 2001. The engines interface with the craft drivetrain through sprag clutches, which engage when the engine output is greater than the drive shaft speed. Historically, sprag clutch failures have been observed on multiple LCACs, resulting in craft downtime and associated repair & replacement costs. Subsequently, a US Navy investigation revealed the presence of high frequency, low level fluctuations in the shaft rotational speed when operated at steady state conditions. The study also suggested that the speed governor fuel control could excite these fluctuations and induce sprag clutch failure. The original speed governor gains were tuned for maximum transient performance, but not necessarily steady state stability. The aggressive gain selection resulted in a governor response that could be characterized as “hyper-reactive” where the system was willing to respond to even the slightest disturbance, including common drivetrain noise. The US Navy requested that Vericor modify the Full Authority Digital Engine Controller (FADEC) speed governor logic to improve steady state stability while maintaining an acceptable transient response. This paper summarizes the basics of ETF40B operation on the LCAC and describes the effort that improved the governor response. As of the writing of this paper, two (2) years have passed in which no sprag clutch failures have been observed on US Navy LCACs operating with the optimized speed governor.

Commentary by Dr. Valentin Fuster
2017;():V001T25A006. doi:10.1115/GT2017-63718.

Electric starter development programs have been the subject of ASME technical papers for over two decades. Offering significant advantages over hydraulic or pneumatic starters, electric starters are now poised to be the preferred choice amongst gas turbine customers. That they are not now the dominant starter in the field after decades of investment and experimentation is attributable to many factors. As with any new technology, progress is often unsteady, depending on budgets, market conditions, customer buy-in, etc. Additionally, technological advances in the parent technologies, in this case electric motors, can abruptly and rapidly change, further disturbing the best laid introduction plans. It is therefore not too surprising that only recently, is the industry beginning to see the deployment of electric starters on production gas turbines. The earliest adoption occurred on smaller gas turbine units, generally less than 10 MW in power. More recently, gas turbines greater than 10 MWs are being sold with electric starters. The authors expect that regardless of their size or fuel supply, most all future gas turbine users will opt for electric starters. This may even include the “larger” frame machines with power greater than 100 MW. Starting with some past history, this paper will not only summarize past development efforts, it will attempt to examine the current deployment of electric starters throughout the marine and industrial gas turbine landscapes. The large-scale acceptance of electric start systems for both new production and retrofit will depend on the favorable cost/benefit assessment when weighing both first cost and life cycle cost. The current and intense activity in electric vehicle applications is giving rise to even more power dense motors. The paper will look at some of these exciting applications, the installed products, and the technologies behind the products. To what extent these new products may serve the needs of the gas turbine community will be the central question this paper attempts to answer.

Commentary by Dr. Valentin Fuster
2017;():V001T25A007. doi:10.1115/GT2017-63751.

Axial flow cyclone separator with guide blade has been widely used, due to its low resistance, huge gas processing and small volume. Although its structure is simple, three-dimension strong rotating turbulent flow forms which involves many complex interactions such as dual-phase separation, adsorption and electrostatic interference. This paper is focused on studying the resistance performance of the axial flow cyclone separator. Numerical simulation methods are carried out to acquire the internal flow field characteristics under different operating pressure and temperature conditions. The result shows that the pressure drop decreases under the same operating pressure, as the operating temperature increases. When the operating temperature is the same, the higher operating pressure enhances the value of the pressure drop. Velocity distribution, pressure contours and turbulent viscosity contours have been presented, to analyze the characteristics of the internal airflow, so as to help optimize the design. Experiments are intended to verify the results of numerical simulation and explore the internal flow field of the cyclone separator further. The cyclone separator has 8 rotary blades which are split into 8 parts, namely one blade is 45° in the tangential direction. 0° and 22.5° are chosen in the experiment. The dimensionless pressure distribution is shown. A comparison of the CFD results and the experimental results is made to prove that the numerical simulation methods are correct and accurate. The curve of the numerical simulation results is very close to that of the experimental results with the similar trend. It is concluded that the methods can predict the internal flow field characteristics of the axial flow cyclone separator.

Commentary by Dr. Valentin Fuster
2017;():V001T25A008. doi:10.1115/GT2017-64048.

Following the successful development of aircraft jet engines during World War II (WWII), the United States Navy began exploring the advantages of gas turbine engines for ship and boat propulsion. Early development soon focused on aircraft derivative (aero derivative) gas turbines for use in the United States Navy (USN) Fleet rather than engines developed specifically for marine and industrial applications due to poor results from a few of the early marine and industrial developments. Some of the new commercial jet engine powered aircraft that had emerged at the time were the Boeing 707 and the Douglas DC-8. It was from these early aircraft engine successes (both commercial and military) that engine cores such as the JT4-FT4 and others became available for USN ship and boat programs. The task of adapting the jet engine to the marine environment turned out to be a substantial task because USN ships were operated in a completely different environment than that of aircraft which caused different forms of turbine corrosion than that seen in aircraft jet engines. Furthermore, shipboard engines were expected to perform tens of thousands of hours before overhaul compared with a few thousand hours mean time between overhaul usually experienced in aircraft applications. To address the concerns of shipboard applications, standards were created for marine gas turbine shipboard qualification and installation. One of those standards was the development of a USN Standard Day for gas turbines. This paper addresses the topic of a Navy Standard Day as it relates to the introduction of marine gas turbines into the United States Navy Fleet and why it differs from other rating approaches. Lastly, this paper will address examples of issues encountered with early requirements and whether current requirements for the Navy Standard Day should be changed. Concerning other rating approaches, the paper will also address the issue of using an International Organization for Standardization, that is, an International Standard Day. It is important to address an ISO STD DAY because many original equipment manufacturers and commercial operators prefer to rate their aero derivative gas turbines based on an ISO STD DAY with no losses. The argument is that the ISO approach fully utilizes the power capability of the engine. This paper will discuss the advantages and disadvantages of the ISO STD DAY approach and how the USN STD DAY approach has benefitted the USN. For the future, with the advance of engine controllers and electronics, utilizing some of the features of an ISO STD DAY approach may be possible while maintaining the advantages of the USN STD DAY.

Commentary by Dr. Valentin Fuster
2017;():V001T25A009. doi:10.1115/GT2017-64381.

In order to facilitate the application of special structural ejectors, which improve the ability of pumping the secondary flow without additional power consumption, reducing the flue gas temperature at the export and enhancing the ship viability under the threat of infrared guided weapons, this paper regardes the 90 ° bend tabs ejector as the research object according to the actual situation of our country’s ships, focuses on the inner effect of the existence of tabs on the flow field in the bent channel, and mainly revealed the transformation of the vortex around the tabs, for providing an explanation to a certain extent about how the tabs affect the macro performance of ejector. With ANSYS software, ring 8 equilateral triangles tabs were designed with 120 ° wall surface mounting angle. With adjusting the blocking ratio of the main outlet area based on the similar zoom, setting inlet swirl angle, and building a hybrid grid to compute, the vortex structure distribution and the development around tabs were observed. The maximum vorticity of vortex at different distances in the mixing tube to the mix tube exit had been calculated to reflect the change of vortex intensity. The final results show that although the streamwise vortices are still located in an axial symmetrical distribution, the swirl angle leads to an uneven distribution of the flow on both sides of a single tab. The inlet swirl angle can make the symmetry of the steamwise vortex vaguer, but the effect of the convection to the vortex is enhanced. The blocking area ratio of the nozzle cross-sectional surface has a large effect on the vorticity of the streamwise vortex. The calculation results show that the larger the blocking area is, the greater the vorticity of streamwise vortex is, which also shows that when the tab shape is fixed, the tab surface area will increase the streamwise vorticity. Through the above research, the shape and the change of the streamwise vortex generated by the tabs in the bent ejector are clearly demonstrated, which can be a reference for the design of high performance bent ejector.

Commentary by Dr. Valentin Fuster
2017;():V001T25A010. doi:10.1115/GT2017-65281.

It has been conjectured that if sulfur in fuel is removed, engine materials will cease to experience attack from hot corrosion, since this sulfur has been viewed as the primary cause of hot corrosion and sulfidation. Historically, hot corrosion has been defined as an accelerated degradation process that generally involves deposition of corrosive species (e.g., sulfates) from the surrounding environment (e.g., combustion gas) onto the surface of hot components, resulting in destruction of the protective oxide scale. Most papers in the literature, since the 1970s, consider sodium sulfate salt as the single specie contributing to hot corrosion.

Recent Navy standards for Navy F-76 and similar fuels have dropped the sulfur content down to 15 parts per million (ppm). Most observers believe that the removal of sulfur will end hot corrosion events in the Fleet. However, the deposit chemistry influencing hot corrosion is known to be much more complex consisting of multiple sulfates and silicates. Sulfur species may still enter the combustion chamber via ship’s air intake, which may include seawater entrained in the air. In addition to sodium sulfate, seawater contains magnesium, calcium and potassium salts, and atmospheric contaminants that may contribute to hot corrosion. This paper will cover some of the revised understanding of hot corrosion and consider other possible contaminants that could further complicate a full understanding of hot corrosion.

Topics: Corrosion
Commentary by Dr. Valentin Fuster

Honors and Awards

2017;():V001T51A001. doi:10.1115/GT2017-63205.

Turbine cooling is a battle between the desire for greater hot section component life and the techno-economic demands of the marketplace. Surprisingly little separates the haves from the have nots. The evolution of turbine cooling is loosely analogous to that of the Darwinian theory of evolution for animals, starting from highly simplistic forms and progressing to increasingly more complex designs having greater capabilities. Yet even with the several generations of design advances, limitations are becoming apparent as complexity sometimes leads to less robust outcomes in operation. Furthermore, the changing environment for operation and servicing of cooled components, both the natural and the imposed environments, are resulting in new failure modes, higher sensitivities, and more variability in life. The present paper treats the evolution of turbine cooling in three broad aspects including the background development, the current state-of-the-art, and the prospects for the future. Unlike the Darwinian theory of evolution however, it is not feasible to implement thousands of small incremental design changes, random or not, to determine the fittest for survival and advancement. Instead, innovation and experience are utilized to direct the evolution.

Over the last approximately 50 years, advances have led to an overall increase in component cooling effectiveness from 0.1 to 0.7. Innovation and invention aside, the performance of the engine has always dictated which technologies advance and which do not. Cooling technologies have been aided by complimentary and substantial advancements in materials and manufacturing. The state-of-the-art now contains dozens of internal component cooling methods with their many variations, yet still relies mainly on only a handful of basic film cooling forms that have been known for 40 years. Even so, large decreases in coolant usage, up to 50%, have been realized over time in the face of increasing turbine firing temperatures. The primary areas of greatest impact for the future of turbine cooling are discussed, these being new engine operating environments, component and systems integration effects, revolutionary turbine cooling, revolutionary manufacturing, and the quantification of unknowns. One key will be the marriage of design and manufacturing to bring about the concurrent use of engineered micro cooling or transpiration, with the ability of additive manufacturing. If successful, this combination could see a further 50% reduction in coolant usage for turbines. The other key element concerns the quantification of unknowns, which directly impacts validation and verification of current state-of-the-art and future turbine cooling. Addressing the entire scope of the challenges will require future turbine cooling to be of robust simplicity and stability, with freeform design, much as observed in the “designs” of nature.

Topics: Cooling , Turbines
Commentary by Dr. Valentin Fuster

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