bluuebunny/crossref_metadata_2025_split huggingface dataset:
Evaluation of the Fuel Saving Potential Regarding Recuperated Helicopter Flight Conditions
Authors: Zhang Chengyu , Kerler Martin , Gümmer Volker 
Published: August 2018

Incorporating recuperators into turbine engines which enables the heat transfer between exhaust gas and compressed air, indicates considerable potential for lower emissions and SFC (specific fuel consumption). However as a matter of fact, they have not yet found wide acceptance in aircraft applications. One of the main potential disadvantages is because in such situations system overall weight needs to be strictly controlled, and there are concerns that the beneficial fuel saving may not offset the additional bulk and weight of the recuperator.
In this work, aimed at evaluating the performance of the recuperated rotorcraft and investigating the influence of a recupeator on the whole system, a comprehensive simulation framework has been developed which mainly contains three modules for various flight conditions: Helicopter performance module (HPM), GasTurb engine simulation module and the recuperator weight estimation. Specifically, HPM calculates the helicopter power requirement for different flight conditions (altitude, speed, etc.), and GasTurb is used to simulate the performance of conventional and recuperated turbine engine, and compute the corresponding engine operating point to meet the power demand. The recuperator weight estimation is mainly based on previous studies as a function of mass flow and effectiveness. The methodology is adopted for a typical twin engine light helicopter configuration.
In comparison with the conventional non-recuperated cycle, the authors studied the fuel saving potential of the recuperated cycle for different recuperator effectiveness under various flight conditions, and a trade-off analysis was also conducted to identify the flight time required to compensate the additional recuperator weight.
The obtained results suggested that the recuperated cycle possesses great potential, especially for long duration and large range mission, but it may not be necessarily suitable for all types of helicopter missions.
3D Computational Analysis of a Compressor for Heavy Duty Truck Engine Turbochargers
Authors: Abel Matthias , Newton Peter , Martinez-Botas Ricardo F. , Wöhr Michael , Müller Markus , Leweux Johannes 
Published: August 2018

This paper presents a detailed analysis of a centrifugal compressor stage with a vaned diffuser for application in a heavy duty engine turbocharger. The analysis is carried out in order to investigate the capability of a steady-state 3D CFD to predict the pressure distribution resulting from the flow through the stage.
As it is impractical to investigate the whole range of conditions, the first objective of this paper is to define the operating points, which are most relevant to the real operation of the compressor, to be used in this paper. In order to do this, the engine operation is weighted in terms of the total fuel consumption across a reference route: Portland-Pendleton-Portland. After the definition of the primary operating points, a CFD setup is created with regard to previous work on steady-state 3D CFD of a centrifugal compressor stage. The CFD analysis is validated against extensive hot-gas test stand data. The experimental data include the static pressure distribution in the diffuser shroud, determined with an array of 256 pressure transducers. A maximum discrepancy between the CFD and the experiment of around 1.6% difference is achieved for all primary operating points.
Effect of Temperature Ratio on Jet Impingement Heat Transfer in Active Clearance Control Systems
Authors: Da Soghe Riccardo , Bianchini Cosimo , D’Errico Jacopo , Tarchi Lorenzo 
Published: August 2018

Impinging jet arrays are typically used to cool several gas turbine parts. Some examples of such applications can be found in the internal cooling of high pressure turbine airfoils or in the turbine blade tip clearances control of aero-engines. The effect of wall-to-jets temperature ratio on heat transfer is generally neglected by the correlations available in the open literature.
In present contribution, the impact of the temperature ratio on the heat transfer for a real engine Active Clearance Control system, is analyzed by means of validated CFD computations. At different jets Reynolds number and considering several impingement array arrangements, a wide range of target wall-to-jets temperature ratio (TR) is accounted for. Computational results prove that both local and averaged Nusselt numbers reduce with increasing temperature ratios.
An in-depth analysis of the numerical data show that the last mentioned evidence is motivated by both the heat transfer incurring between the spent coolant flow and the fresh jets and the variation of gas properties with temperature through the boundary layer. A scaling procedure, based on TR power law, was proposed to estimate the Nusselt number at different wall temperature levels necessary to correct available open-literature correlations, typically developed with small temperature differences, for real engine applications.
Dynamic Analysis and Reduction of a Cyclic Symmetric System Subjected to Geometric Nonlinearities
Authors: Martin Adrien , Thouverez Fabrice 
Published: August 2018

The search for ever lighter weight has become a major goal in the aeronautical industry as it has a direct impact on fuel consumption. It also implies the design of increasingly thin structures made of sophisticated and flexible materials. This may result in nonlinear behaviours due to large structural displacements. Stator vanes can be affected by such phenomena, and as they are a critical part of turbojets, it is crucial to predict these behaviours during the design process in order to eliminate them.
This paper presents a reduced order modelling process suited for the study of geometric nonlinearities. The method is derived from a classical Component Mode Synthesis with fixed interfaces, in which the reduced nonlinear terms are obtained through a STEP procedure using an adapted basis composed of linear modes completed by modal derivatives. The whole system is solved using a harmonic balance procedure and a classic iterative nonlinear solver. The application is implemented on a schematic stator vane model composed of nonlinear Euler-Bernoulli beams under von Kàrmàn assumptions.
Robust Fluid Topology Optimization Using Polynomial Chaos Expansions: TOffee
Authors: Gaymann A. , Montomoli F. , Pietropaoli M. 
Published: August 2018

The paper presents an innovative solution to robust topology optimization developed for components that can be manufactured by additive manufacturing. Topology optimization has been used in fluid dynamics to optimize geometries based on a target set of performances required for the flow paths. These target performances can be defined as pressure losses or heat exchanges for example, and multiple optimized geometries can be found in the literature.
However, none of these cases considered the impact of stochastic variations and are based on a deterministic optimization. It means the optimization has been done for a single boundary condition value. Would this boundary be random, as it is the case in real life gas turbines, then the optimized geometry, optimized for a single set of boundary conditions, will underperform.
Robust topology optimization obtains a geometry able to cope with these random variations. The robust optimization method has been implemented in an in-house solver TOffee and relies on a multi-objective function.
2D and 3D robust optimized geometries are obtained and their performance compared to deterministic cases over a range of boundary conditions. Superiority of robust geometries as compared to deterministic geometries is shown. Robust topology optimization presents a great interest in the gas turbine industry due to the greater performance obtained by the optimized geometries while taking into consideration random variations of boundary conditions, making the simulations closer to real life conditions.
For the first time in this work it is shown a fluid topology optimization solution with sedimentation that are inherently able to cope with uncertainty.
Reduction of the Design Space to Optimize Blade Fir-Tree Attachments
Authors: Alinejad F. , Botto D. , Gola M. , Bessone A. 
Published: August 2018

The blade attachment, both dovetail or fir-tree, transfers the centrifugal load from the blade to the disc, generating high mean and peak stresses in notches as well as on contact surfaces. Hence, the strength of the attachment is one of the main concern of the designers for improving the performance of the engine and several optimization procedure have been put forward to minimize the state of stress in the attachment for a given centrifugal load. The optimization process is generally driven by a parametric model. The selection of the proper parameters and their variation ranges represent one of the main issues for the process to converge in a reasonable amount of time. Simulation methods and optimization algorithms have been improved a lot in the past years. Nevertheless, the computational effort of the finite element analysis involved in the optimization procedure of complex geometries remains a critical task. Moreover, an accurate evaluation of the local contact stresses is highly dependent on the mesh refinement, increasing the computing time of the whole optimization process. Moreover, a multi-objective optimization, in addition to robustness design approach, is the designer tool to improve the attachment performance.
The searching domain reduction of the optimization process improves the computational performance reducing the convergence time of the solution. To achieve this goal, a preliminary selection of the design space has been performed by means of an analytical approach.
This paper describes a new design criterion based on one dimensional approach. The criterion has been implemented in an in-house tool that takes faster decisions, if compared with a two or a three dimensional model, about the number of possible feasible solutions. During the geometrical optimization phase of the blade fir-tree attachment, in which a parametric model is used, the authors try to handle the geometrical non-feasibility with a combination of Latin Hypercube Sampling (LHS) and an adaptive penalty method. The optimization is done via the genetic algorithm and the computational time of the reduced domain is compared with the original one.
Utilization of a Thermo-Mechanical Model Coupled With Multi-Objective Optimization to Enhance the Start-Up Process of Solar Steam Turbines
Authors: Topel Monika , Vitrano Andrea , Laumert Björn 
Published: August 2018

The need to mitigate the climate change has brought in the last years to a fast rise of renewable technologies. The inherent fluctuations of the solar resource make concentrating solar power technologies an application that demands full flexibility of the steam turbine component. A key aspect of this sought steam turbine flexibility is the capability for fast starts, in order to harvest the solar energy as soon as it is available. However, turbine start-up time is constrained by the risk of low cycle fatigue damage due to thermal stress, which may bring the machine to failure. Given that the thermal limitations related to fatigue are temperature dependent, a transient thermal analysis of the steam turbine during start process is thus necessary in order to improve the start-up operation.
This work focuses on the calculation of turbine thermo-mechanical properties and the optimization of different start-up cases in order to identify the best solution in terms of guaranteeing reliable and fast start-ups. In order to achieve this, a finite element thermal model of a turbine installed in a concentrating solar power plant was developed and validated against measured data. Results showed relative errors of temperature evolutions below 2%, making valid the assumptions and simplifications made. Since there is trade-off between start-up speed and turbine lifetime consumption, the model was then implemented within a multi-objective optimization scheme in order to test and design faster start-ups while ensuring safe operation of the machine. Significant improvements came up in terms of start-up time reduction up to 30% less than the standard start-up process.
Thermomechanical Model Reduction for Efficient Simulations of Rotor-Stator Contact Interaction
Authors: Guérin Nicolas , Thorin Anders , Thouverez Fabrice , Legrand Mathias , Almeida Patricio 
Published: August 2018

Turbomachinery rotor-stator unilateral contact induced interactions play a growing role in lifecycle analysis and thus motivate the use of accurate numerical prediction tools. Recent literature confirmed by ongoing in-house experiments have shown the importance of thermomechanical coupling effects in such interactions. However, most available (possibly reduced-order) models are restricted to the sole mechanical aspects.
This work describes a reduction technique of thermomechanical models involving unilateral contact and frictional contact occurrences between rotor and stator components. The proposed methodology is grounded on Guyan and Craig–Bampton methods for the reduction of the structural dynamics in conjunction with Krylov subspace techniques, and specifically the Craig–Hale approach, for the reduction of the thermal equations.
The method has the capability to drastically reduce the size of the model while preserving accuracy. It stands as a reliable strategy to perform simulations of thermomechanical models with localized mechanical and thermal loads.
Near and Far Field Noise Decay From a Quadcopter Propeller With and Without a Leading Edge Notch
Authors: Van Treuren Kenneth , Wisniewski Charles , Cinnamon Emily 
Published: August 2018

Electric propulsion is being considered for a wide range of airframes from large commercial transports to the small Unmanned Aerial Systems (UASs). These electric systems, especially for small fixed wing UASs and quadcopters, need to be both efficient and quiet if they are to operate in an urban/populated environment or used in an Intelligence, Surveillance, and Reconnaissance (ISR) scenario.
A propeller test facility was developed to record propeller performance and sound generation in the near field behind UAS propellers. The question of defining near and far field noise was studied by characterizing sound decay with distance from a UAS propeller. Defining near and far field noise is a subject that is not addressed well in the literature. Far field noise generally follows the 1/r decay rate and near field does not. Behind the propeller there are other flow field interactions that also change the decay rate, which this study illustrates. The data presented in this paper shows the difficulty in measuring sound around a UAS propeller and begins to resolve this topic. Previous UAS propeller design work by the authors resulted in propellers that were quieter in the near field and at the same time more efficient. Their studies showed RPM and tip vortex formation both contribute significantly to propeller sound generation. Disrupting the tip vortex formation should decrease the noise being generated. The current work extends these initial findings and examines the noise generation of a stock quadcopter propeller from a DJI Phantom 2 platform. One inch aft of the plane of rotation, this propeller, a 9.4 × 5.0, has a peak sound pressure level (SPL) of approximately 118 dBA under normal static operation producing 0.7 lbf of thrust at approximately 5900 RPM. Modifications were made to four stock propellers by cutting a notch perpendicular to the leading edge of the propeller at the 0.75 r/R and 0.87 r/R locations. The notches were of different depths and widths. Of the modifications, three of the configurations did not noticeably decrease the sound. However; the final configuration reduced the peak near field SPL to 111 dBA, a 6% reduction in dBA over the stock configuration corresponding to a greater than 50% reduction in sound generation. Smoke visualization confirms that a notch located at 0.87 r/R effectively disrupts the tip vortex formation, causing the tip vortices to dissipate much earlier than the stock propeller without the notch. Examining the noise frequency spectrums associated with both the stock and the modified propeller also confirm that the notch changes the magnitude and frequency distribution of the sound being generated.
The Use of Blended Blade and End Wall in Compressor Cascade: Optimization Design and Flow Mechanism
Authors: Li Jiabin , Ji Lucheng , Yi Weilin 
Published: August 2018

Nowadays, the flow field at the compressor is more and more complex with the increasing of the aerodynamic loading. The complex flow in the endwall regions is thus key to aerodynamic blockage, loss production, and finally its performance deterioration. The design of Blended Blade and End Wall (BBEW) contouring technology had been proved to be useful in delaying, reducing, and eliminating the corner separation in the compressor. The BBEW technology can adjust the dihedral angle between the suction and the endwall in 30% of the spanwise easily, which is different with the fillet. However, the design of the BBEW always relies on the experiences of the designers, and the effective design results cannot be the optimal result. This paper presents an optimization design method for the BBEW technology, and analyses the flow mechanism of the BBEW design.
Firstly, the parameters for the BBEW design is simplified as two, one is the maximum blended width, the other is the axial position of the maximum blended width. The optimal result can be obtained through the response surface method.
Secondly, based on the optimization method, this paper make an optimization BBEW design at the suction side of a NACA65 linear compressor cascade with the turning angle 42 degrees. The numerical results show that the optimal BBEW design can eliminate the boundary layer separation at the corner intersection region, and reduce the suction side separation. When the incidence angle is 0 degrees, the BBEW technology can reduce the total pressure loss coefficient by 5%, and reduce the aerodynamic blockage coefficient by 14%. The aerodynamic performance of the cascade shows a more obvious improvement with the BBEW design at a larger incidence. The total pressure loss coefficient of the cascade is reduced by 20% at 15 degrees incidence.
The numerical study shows that the design with the BBEW can control the axial development of the dihedral angle between the suction side and the endwall, which can eliminate the boundary layer separation at the corner intersection region. What’s more, the BBEW technology can produce a pressure gradient at the axial position of the maximum blended width, and value of the pressure gradient in proportion to the maximum blended width. This pressure gradient enhance the kinetic energy of the low energy fluid at the endwall region, which is consist of the secondary cross flow, thus elevating the capability to withstand the adverse pressure gradient, and improve the suction side separation around the trailing edge.
An Experimental Investigation of Full-Coverage Film Cooling Characteristics of a Turbine Guide Vane
Authors: Wu Jin , Zhang Li , Cheng Li-jian , Jiang Ru , Fu Zhong-yi , Zhu Hui-ren 
Published: August 2018

This paper researches on the effects of Reynolds number and mass flow ratio on the film cooling characteristics at high turbulence intensity (Tu = 15%). The experiment adopted an actual three-dimensional twisted vane and presents the film cooling characteristics on full-coverage film surface in a two-passage, linear cascade. The cooling effectiveness and heat transfer coefficient of the vane’s whole surface were obtained by using transient liquid crystal measurement technique. The transient liquid crystal is SPN/R35C1W, whose bandwidth is 2°C. There are fifteen rows of film cooling holes which have different diameter, injection angle and yaw angle. The secondary flow was supplied by two cavities. The front cavity supplied the secondary flow to thirteen rows of film cooling holes that were arranged in the suction surface, the leading edge and the front half of the pressure surface. The rear cavity supplied the secondary flow to the rear half of pressure surface which included two rows of film cooling holes. The investigated parameters are Reynolds number of 1 × 105, 1.3 × 105 and 1.6 × 105 and the mass flow ratio of MFR = 5.5%∼12.5% (6 cases). The data recorded in the experiment was analyzed with MATLAB.
Results show that the combined effects of mass flow ratio and channel vortex are the maintain reasons that influence the distribution of cooling effectiveness in the contour. Increasing the mass flow ratio can improve the film cooling effectiveness on leading edge and pressure surface, while that presents complex rule on suction surface. Increasing the Reynolds number can improve the heat transfer coefficient at the same mass flow ratio. When increasing the mass flow ratio, the heat transfer coefficient increases on leading edge and pressure surface at Re = 1.6 × 105. However, the decreases at film hole outlet region on the suction side, and not obviously changes at the film hole downstream region.
Impact Test for the Leading Edge of CMC Vane Based on Actual Aircraft Engine Field Data
Authors: Obuchi Kenro , Watanabe Fumiaki , Kuroki Hiroshi , Yagi Hiroyuki , Arai Kazuyoshi 
Published: August 2018

Ceramic matrix composites (CMCs) have lower density and a higher service temperature limit than nickel based alloys which have been used for turbine components of aircraft engines. These properties of CMCs have the potential to reduce the weight of turbine components and improve turbine thermal efficiency with a higher turbine inlet temperature (TIT). One of the technical issues of the CMC turbine vane is a relatively lower impact resistance than nickel based alloy turbine vanes. There are various previous works about impact resistance of CMCs, but there is little work that assumed actual engine conditions.
The objective of this work was to verify the resistance of SiC/SiC CMC turbine vane to the impact phenomena that occur in the actual aircraft engine. The field damage survey was conducted on actual metal turbine vanes of commercial engines overhauled in IHI. The survey made it clear that the typical damage was less-than-0.127-mm-dent at the leading edge. In addition, the dropped weight impact test using the actual turbine airfoil which is made from a nickel based alloy was conducted at ambient temperature. The amount of energy required to make the dent of a certain size that was observed in actual metal turbine vanes was estimated. Then, the dropped weight impact test using the CMC test piece with a leading edge shape was conducted at the impact energy estimated by the metal turbine airfoil. The results showed that the failure mode of the CMC test piece was local damage with dents of a certain size and not a catastrophic failure mode.
From this work, the damage to be assumed on CMC vane in actual aircraft engines was identified. As a future task, the effect of the damage to the fatigue capability of CMC turbine vanes needs to be investigated.