bluuebunny/crossref_metadata_2025_split huggingface dataset:
Energy Management of a Hybrid-Electric Aeronautical Propulsion System to Be Used in a Stationary Test Bench
Authors: Rendon Manuel A. , Kyprianidis Konstantinos , Roque Benito Yipsy , de A. Fernandes Daniel , Ferraz Ariele T. , Vieira Luan R. C. 
Published: January 2021

Abstract
Environmental requirements have led the air transportation industry to work towards reducing greenhouse gas emissions and mechanical noise levels. Nowadays, this sector contributes with 2% of the total greenhouse gas emissions, and there is a demand from global aviation regulators for further reducing this percentage. In the last years, the development of Hybrid-Electric Propulsion Systems (HEPSs) has grown. The HEPS combines an Internal Combustion Engine (ICE), for example, Gas Turbine (GT) or reciprocating engine, with an Electric Motor (EM), combining the inherent advantages of both. HEPSs present increased efficiency and operating safety in comparison with conventional ICE-powered systems. Furthermore, they can supply the electrical devices with power. This area of study is multidisciplinary in nature and, therefore, poses research challenges on ICEs, EMs, electronic converters, propeller design, monitoring and control systems, management and supervision systems, energy efficiency and optimization, aerodynamics and aircraft mechanical design. A research project aimed at the characterisation of hybrid-electric aircraft propulsion systems, and the construction of a HEPS prototype, is underway in Brazil. The system is essentially composed of a GT, an EM, three electronic converters, a battery bank and a propeller. It can operate with three different topologies: series, full-electric and turbo-electric. A test bench with all the necessary peripheral and analysis infrastructure is under construction. Present work aims to: (i) develop simplified models for all the test bench components, (ii) given a mission profile, show the results of an initial energy management computing code that determines the optimal hybridization strategy, and (iii) simulate various operating alternatives for the chosen mission profile. The results (i) highlight the impact of critical characteristics of the batteries on the HEPS performance, and (ii) demonstrate the application of the management code on optimizing the aircraft energy consumption for a given mission profile.
Preliminary Design of Variable Nozzle Turbines Based on Sensitive Parameters
Authors: Wittwer Sebastian , Sandor Ivo 
Published: January 2021

Abstract
Recent developments in turbocharged gasoline engines have established new requirements for the turbine. A simple approach of scaling or optimizing existing turbines on component level might not be sufficient in terms of finding an optimal solution according to the multi-point, multi-disciplinary layout target. In the following paper nondimensional functional parameters are derived from turbomachinery analytics and rated on corresponding values of existing turbine stages. The influence of different parameters on aerodynamic performance is discussed based on CFD results and arranged according to their sensitivity for different engine relevant operating conditions. A metamodel for the preliminary design of variable nozzle turbine stages is derived from DoE (Design of Experiments) based CFD results. It is evaluated regarding its predictive quality on several exemplary turbine stages. Both, CFD and experimental results are therefore used while the experimental results are made up of hot gas stand measurements as well as measurements on engine test bench. Thus, not only the influence of functional parameters can be verified on turbine efficiency characteristics, but beyond that also the predictive quality of engine performance can be assessed.
A New Test Facility for Investigating Thermal Behaviour of Effusion Cooling Test Plates for RQL Combustors
Authors: Picchi A. , Andreini A. , Becchi R. , Facchini B. 
Published: January 2021

Abstract
In aero engines the combustors are subjected to critical thermal conditions in terms of high temperatures and corrosive environment, which could affect the service life of the entire system. As well known, Thermal Barrier Coatings (TBC) and above all cooling systems represents the state-of-the-art in the nowadays protecting methods: the maximization of this beneficial effect is achieved by defining an optimal cooling arrangement and developing suitable manufacturing technologies for these systems. In modern aero-engine combustors, one of the most effective cooling scheme for liners is composed by an effusion perforation coupled with a slot system to start the film cooling. The cooling performances are deeply influenced by the mutual interactions between swirling and cooling flows. In addition, for typical Rich-Quench-Lean (RQL) combustor architectures, the injection of air provided to promoting the local break-down of the flame mixture fraction, deeply interacts with the swirled flow, generating recirculating structures capable of affecting the development of film cooling and making the design of cooling systems very challenging.
A new test facility for testing effusion test plates for RQL combustors applications has been developed with the final aim of comparing different cooling strategies and at the same time to collect data for numerical model validation. The experimental set-up consists of a non-reactive planar sector rigs with 5 engine-scale swirlers fed with air up to 250 °C and 3 bar. The rig was equipped with outer/inner dilution ports, and a simple inner liner cooling scheme composed of effusion and a slot system: all these features, fed with air at ambient temperature, can be independently controlled in terms of mass flow.
Using dedicated optical accesses, InfraRed (IR) camera tests were performed to retrieve overall effectiveness data imposing a temperature difference between swirling and cooling flows. To better understand those results, Pressure Sensitive Paint (PSP) technique was used to obtain reliable film effectiveness data decoupling the contribution of slot and effusion flows. The thermal characterization was supported by Particle Image Velocimetry (PIV) investigations on the median plane. Tests were performed at different pressure drops across swirler and varying the mass flows of slot and inner/outer liners. The analysis of the data highlighted the influences of the swirling flow on the overall thermal performance and the behaviour of the film cooling system.
Upstream Film Cooling on the Contoured Endwall of a Transonic Turbine Vane in an Annular Cascade
Authors: Salinas Daniel A. , Ullah Izhar , Wright Lesley M. , Han Je-Chin , McClintic John W. , Crites Daniel C. , Riahi Ardeshir 
Published: January 2021

Abstract
The effects of mainstream flow velocity, density ratio (DR), and coolant-to-mainstream mass flow ratio (MFR) were investigated on a vane endwall in a transonic, annular cascade. A blow down facility consisting of five vanes was used. The film cooling effectiveness was measured using binary pressure sensitive paint (BPSP). The mainstream flow was set using isentropic exit Mach numbers of 0.7 and 0.9. The coolant-to-mainstream density ratio varied from 1.0 to 2.0. The coolant to mainstream MFR varied from 0.75% to 1.25%. The endwall was cooled by eighteen discrete holes located upstream of the vane passage to provide cooling to the upstream half of the endwall. Due to the curvature of the vane endwall, the upstream holes provided uniform coverage entering the endwall passage. The coverage was effective leading to the throat of the passage, where the downstream holes could provide additional protection. Increasing the coolant flowrate increased the effectiveness provided by the film cooling holes. Increasing the density of the coolant increases the effectiveness on the endwall while enhancing the lateral spread of the coolant. Finally, increasing the velocity of the mainstream while holding the MFR constant also yields increased protection on the endwall. Over the range of flow conditions considered in this study, the binary pressure sensitive paint proved to be a valuable tool for obtaining detailed pressure and film effectiveness distributions.
Stochastic Modelling and Analysis of Rotating Bladed Discs
Authors: Kumar Rahul , Gupta Sayan , Ali Shaikh Faruque 
Published: January 2021

Abstract
This study focusses on probabilistic modelling of the bladed disc system and numerical estimation of the distributions of the response quantities of the system. Stochastic finite element model of the system consisting of all the assemblies and the hub is developed and reported. The spatial inhomogeneity of mistuned structures is modelled as non-Gaussian random field. Experimentally, the system parameters can be measured at the specified locations of the bladed disk structure. In this analysis, a synthetic data is generated which represent this measured data set. Further, Nataf transformation is implemented to each component of the data set to get the polynomial chaos expansion framework of the system parameters. Since, the random field of the system parameter is approximated as correlated random variables, Spearman’s rank correlation coefficient is used in this manuscript to obtain that correlation among the random parameters across the domain. The approximated probability density function obtained through the aforementioned methodology is compared with the target probability density function of the parameter using Kullback–Liebler (KL) entropy as a metric. Also, the same KL entropy is used as a metric to check the convergence of polynomial chaos terms in the expansion. Next, the proposed polynomial chaos method is integrated with commercial finite element software to quantify the propagation of randomness associated with system parameters into the response quantities. Subsequently, the statistical processing helps in estimating the probabilistic measure of the required response quantities. The results obtained through the conventional Monte Carlo (MC) simulations have been used as the benchmark to compare the response characteristics obtained through the proposed algorithm.
Heat Transfer in a Rotating, Two-Pass, Variable Aspect Ratio Cooling Channel With Profiled V-Shaped Ribs
Authors: Chen I-Lun , Sahin Izzet , Wright Lesley M. , Han Je-Chin , Krewinkel Robert 
Published: January 2021

Abstract
The thermal performance of two V-type rib configurations is measured in a rotating, two-pass cooling channel. Modeling modern, high pressure, turbine blades, the cross-section of the cooling channel varies from the first pass to the second pass. The coolant travels radially outward in the rectangular first pass with an aspect ratio of 4:1. Near the tip region, the coolant turns 180°, and travels radially inward in a 2:1 rectangular channel. The serpentine passage is positioned such that both the first and second passes are oriented 90° to the direction of rotation. The leading and trailing surfaces of both the first and second pass of the channel are roughened with V-type rib turbulators. The thermal performance of two V-type configurations is measured in this two-pass channel. The first V-shaped configuration is similar to a traditional V-shaped turbulator with a narrow gap at the apex of the V. The configuration is modified by off-setting one leg of the V to create a staggered discrete, V-shaped configuration. The ribs are oriented 45° relative to the streamwise coolant direction. In both passes, the rib spacing is P/e = 10 and the rib height – to – channel height is e/H = 0.16. The heat transfer enhancement and frictional losses are measured for both rib configurations with varying Reynolds and rotation numbers. The Reynolds number varies from 10,000 to 45,000 in the AR = 4:1 first pass; this corresponds to 16,000 to 73,500 in the AR = 2:1 second pass. Considering the effect of rotation, the rotational speed of the channel varies from 0–400 rpm with maximum rotation numbers of 0.39 and 0.16 in the first and second passes, respectively. The heat transfer enhancement on both the leading and trailing surfaces of the first pass of the 45° V-shaped channel is slightly reduced with rotation. In the second pass, the heat transfer increases on the leading surface while it decreases on the trailing surface. The 45° staggered, discrete V-shaped ribs provide increased heat transfer and thermal performance compared to the traditional V-shaped and standard, 45° angled rib turbulators.
Ekman Layer Scrubbing and Shroud Heat Transfer in Centrifugal Buoyancy-Driven Convection
Authors: Gao Feng , Chew John W. 
Published: January 2021

Abstract
This paper presents large-eddy and direct numerical simulations of buoyancy-driven convection in sealed and open rapidly rotating cavities for Rayleigh numbers in the range 107–109, and axial throughflow Reynolds numbers 2000 and 5600. Viscous heating due to the Ekman layer scrubbing effect, which has previously been found responsible for the difference in sealed cavity shroud Nusselt number predictions between a compressible N-S solver and an incompressible counterpart using the Boussinesq approximation, is discussed and scaled up to engine conditions. For the open cavity with an axial throughflow, laminar Ekman layer behaviour of the mean flow statistics is confirmed up to the highest condition in this paper. The Buoyancy number Bo is found useful to indicate the influence of an axial throughflow. For the conditions studied the mean velocities are subject to Ra, while the velocity fluctuations are affected by Bo. A correlation, Nu′ = 0.169(Ra′)0.318, obtained with both the sealed and open cavity shroud heat transfer solutions, agrees with that for free gravitational convection between horizontal plates within 16% for the range of Ra′ considered.
An Improved Reverse Monte Carlo Method for the Investigation of Aerodynamic and Infrared Radiation Characteristics of a Flying Wing UAV
Authors: Gao Xiang , Yang Qingzhen , Yang Huicheng , Bai Jin , He Yubo 
Published: January 2021

Abstract
Using the improved ray tracing method to improve the reverse Monte Carlo (RMC) method, which is used to analyze the infrared radiation (IR) characteristics of the exhaust system, can greatly improve the computational efficiency and accuracy, and the calculation accuracy is improved by more than 8% compared with the RMC using ray tracing algorithm. For a flying wing unmanned aerial vehicle (UAV), the influence of the geometry of the double S-shaped nozzles and the single S-shaped nozzles on the internal flow field were analyzed, then the influence of the internal and external flow on the aerodynamic performance of the S-shaped exhaust system, and the IR characteristics of the aircraft with different inlet and exhaust system in the 3μm-5μm and 8μm-14μm bands were studied.
The aerodynamic performance of the S-shaped nozzles are studied by numerical simulations. The IR characteristics of the rear hemisphere on the single S-shaped nozzles and the double S-shaped nozzles were obtained, those nozzles have the similar engine matching performance. The double S-shaped nozzle, resulting in a decrease of the radiation intensity of the nozzle by at least 65%, compared to the single S-shaped nozzle in the 3μm-5μm band. The aerodynamic characteristics of the flying wing UAV with the two S-shaped exhaust systems are also compared, and the changes in lift and resistance are analyzed.
The forward IR intensity of the flying wing UAV is significantly lower than that of the backward direction, and the upper direction IR intensity is higher than that of the lateral direction and the downward direction. Compared to the flying wing UAV with the single S-shaped exhaust system, the flying wing of the double S-shaped exhaust system has a lower IR intensity, for the peak intensity of the rear hemisphere in the 3μm-5μm band is reduced at least 80%, the maximum value of the locked distance is reduced to 25% of the UAV using the single S curved exhaust system. The ratio of the amount of aircraft skin radiation to the total radiation increases from 30% in the 3μm-5μm band to more than 70% in the 8μm-14μm band. From the results of spectral analysis of UAV from 13μm-14μm, the spectral radiance of the carbon dioxide absorption and emission band in the detector image is significantly less than that of the UAV with a single S-shaped nozzle.
A Stochastic Immersed Model of Breakup Characteristics in Dense Air Atomization Flow Field
Authors: Deng Tian , Ren Xingming , Li Yaxuan 
Published: January 2021

Abstract
For the low-speed liquid injected into the high-speed strong turbulent gas flow in the same direction, the atomization is a transient-intensive spray, and there are many factors affecting and controlling the atomization. In this paper, the distribution and characteristics of the liquid breakup in the air atomized flow field are analyzed. A stochastic immersed model to simulate the liquid core is developed, in which, the liquid core is regarded as an immersed porous medium with a random structure, and the probability of existence is used to simulate the position of the liquid core. The initial fragmentation mechanism of the air blast atomization is applied as the global variables of the stochastic process. Using the above stochastic immersed model, combined with the Large Eddy Simulation method, the numerical simulation of the downstream flow field of a coaxial jet air atomizing nozzle is carried out. Additional force is added to the momentum equation in the LES model. Instantaneous air velocity at the air-liquid interface is characterized by instantaneous liquid phase velocity at the same time. The size of the initial atomized droplet satisfies a probability distribution, and once the large droplets are formed, the Lagrangian method is used to track the droplets. The comparison between the simulation results and the experimental results shows that this stochastic immersed model can quickly capture the information of length and position of the liquid nucleus. When the gas-liquid momentum ratio M is 3∼10000, the liquid core length can be predicted more accurately. When M&gt;10, the prediction result is much better than phenomenological model. This model is capable of capturing flow field structures such as recirculation zones and large-scale vortices. The results of initial spray angle from experiment expression give slightly better agreement with this model. Increasing the momentum ratio leads to decreasing of the initial spray angle. The particle size of the droplets near the nozzle can be accurately predicted, especially when the gas velocity is large (bigger than 60 m/s), and the average diameter prediction error of the droplets is less than 10%.
Prediction of Rotor Burst Using Strain-Based Fracture Criteria to Comply With the Engine Airworthiness Regulation
Authors: Li Huihui , Wang Kaiming , Zhang Chuncheng , Wang Weiguo , Chen Guoguang 
Published: January 2021

Abstract
Relative to the rotor overspeed compliance governed by civil aviation airworthiness regulation, nowadays Area-Average Stress method is commonly used approach. However, in order to effectively apply the Area-Average Stress method in analyzing burst speed, large amount of testing data is needed to define an important element of this method: a correction factor. This prerequisite hinders the use of this method for many companies which have limited test data. Meanwhile, analysis of rotor burst speed based on Strain-based Fracture Criteria using true stress-strain curves and burst tests has been done on the LPT rotor, and a work procedure obtaining the most critical burst speed for certification is proposed. The analysis results, which had a good correlation with test results, showed that Strain-based Fracture Criteria can accurately predict the burst speed considering the most adverse combination of dimensional tolerances, temperature, and material properties, and rotor dimensional growth under the overspeed condition. Both are required by the aircraft engine airworthiness overspeed regulation. Compared to the Area-Average Stress method, Strain-based Fracture Criteria reflects the physical essence of the rotor burst more realistically and can be simply verified without requiring too much test data, therefore it has a good application prospect in the aircraft engine airworthiness.
Design Space Exploration of Stagnation Temperature Probes via Dimension Reduction
Authors: Scillitoe Ashley D. , Ubald Bryn , Seshadri Pranay , Shahpar Shahrokh 
Published: January 2021

Abstract
The measurement of stagnation temperature is important for turbomachinery applications as it is used in the calculation of component efficiency and engine specific fuel consumption. This paper examines the use of polynomial variable projection to identify dimension reducing subspaces for stagnation temperature probes. As an example application we focus on a simplified Kiel probe geometry, but the proposed data-centric approach could be readily applied to new datasets with different geometries, boundary conditions and design objectives.
The design of Kiel probes is non-trivial, with a large design space, complex flow physics, and competing design objectives. Two design objectives are considered: (1) the stagnation pressure loss, to reduce instrumentation losses; (2) the change in recovery ratio with respect to Mach number, to reduce temperature measurement uncertainty.
Subspaces are obtained for the two design objectives, allowing the influence of seven design parameters to be understood. The entropy generation rate is used to provide physical insights into loss mechanisms. The recovery ratio subspace indicates that for the present probe there is an optimum vent-to-inlet area which minimises the change in recovery ratio with respect to Mach number, and design modifications that yield further small improvements are explored.
Finally, the uncertainty in recovery ratio due to manufacturing variability is shown to be important. In comparison to global sensitivity measures, the use of an active subspace is shown to provide important information on what manufacturing tolerances are important for specific designs. New designs can also be selected that are insensitive to given manufacturing tolerances.
Large-Eddy Simulation of an Integrated High-Pressure Compressor and Combustion Chamber of a Typical Turbine Engine Architecture
Authors: Pérez Arroyo Carlos , Dombard Jérôme , Duchaine Florent , Gicquel Laurent , Odier Nicolas , Exilard Gorka , Richard Stéphane , Buffaz Nicolas , Démolis Jacques 
Published: January 2021

Abstract
The design optimization of aviation propulsion systems by means of computational fluid dynamics is key to increase their efficiency and reduce pollutant and noise emissions. The recurrent increase in available computing power allows nowadays to perform unsteady high-fidelity computations of the different components of a gas turbine. However, these simulations are often made independently of each other and they only share average quantities at interfaces. In this work, the methodology and first results for a sectoral large-eddy simulation of an integrated high-pressure compressor and combustion chamber of a typical turbine engine architecture is proposed. In the simulation, the compressor is composed of one main blade and one splitter blade, two radial diffuser vanes and six axial diffuser vanes. The combustion chamber is composed of the contouring casing, the flame-tube and a T-shaped vaporizer. This integrated computation considers a good trade-off between accuracy of the simulation and affordable CPU cost. Results are compared between the stand-alone combustion chamber simulation and the integrated one in terms of global, integral and average quantities. It is shown that pressure perturbations generated by the interaction of the impeller blades with the diffuser vanes are propagated through the axial diffuser and enter the combustion chamber through the dilution holes and the vaporizer. Due to the high amplitude of the pressure perturbations, several variables are perturbed at the blade-passing frequency and multiples. This is also reflected on combustion where two broadband peaks appear for the global heat release.