bluuebunny/crossref_metadata_2025_split huggingface dataset:
Multi-Objective Shape Optimization of Arrowhead-Shaped Film Cooling Hole on Transonic Turbine Blade
Authors: Okita Yoji , Ideta Takeomi , Fujimoto Shu 
Published: January 2021

Abstract
Film cooling has been widely applied to cool hot section components in modern aero-engines, typically turbine airfoils. Blowing cooling air via discrete holes from inside the airfoil realizes coolant film which protects the exposed surface against harsh gas stream. Film cooling holes are characterized in two types, namely, round and shaped. Shaped holes are designed to provide higher film effectiveness with marginal heat transfer enhancement compared to round holes, typically by diffusing and spreading the coolant jet with expanded exits.
A number of factors affect the film cooling performance: hole shape, flow rate parameter (e.g. blowing ratio), density ratio, freestream turbulence, etc. Among these, definitely the hole shape has dominant effects on the performance. Therefore, it has been one of the primary interests in the film cooling research to design/select/optimize the hole geometry, where various ideas, some relatively simple and others even more complicated, were proposed.
The primary aim and contribution of the present paper is to challenge the shaped hole multi-objective optimization on a realistic setup. A medium-loaded modern HPT cascade under transonic condition was employed, where the effects of surface curvature, pressure gradient, Mach number, etc. on the thermal and aerodynamic behavior of film cooling can be reproduced faithfully as in actual operating turbines. Heat transfer enhancement and cascade aerodynamic loss as well as film effectiveness were considered as objective functions. Film effectiveness and heat transfer coefficient were evaluated across almost the entire surface downstream because a shaped hole working well only near the hole often shows very limited overall thermal advantage, based on the authors’ design experience. Any hole design assuming the use of emerging new techniques as AM and/or other breakthrough in manufacturing capabilities in the future is completely out of the present scope. Restrictions in design parameters were carefully set for the resultant hole shape to be well manufacturable with conventional process.
Another important consideration in the present work is to generalize the observations as much as possible from the optimization results. The fundamental risk in optimization work is that whole the logical context led to the optimal solution is often very difficult to follow for designers. As the process being a black box, any minor redesign task needs another full set of optimization, which is not the right course to take, obviously. Even though the present work focused on one specific setup, a factorial effect analysis was carried out in order to make the observations generalized.
Design of Speed Closed-Loop Control With Variable Pressure Difference Valve for Aero Engine
Authors: Jiang Zhen , Wang Xi , Chen Huairong , Wang Yunlai 
Published: January 2021

Abstract
Focusing on the advantages of simple structure, quick response and good static performance when fuel system using variable pressure difference valve (VPDV) and the development trend to full authority digital engine control (FADEC) of modern engine, this paper aims to propose a design method of digital speed closed-loop control system based on fuel metering unit (FMU) using VPDV. Firstly, the working principle of a real engine with VPDV is introduced, and the nonlinear AMESim model of the system is established. Secondly, the static analysis of VPDV and metering valve is carried out, and the principle of speed closed-loop control of the hydraulic mechanical system is revealed. Therefore, we built the structure of digital speed control circuit with VPDV. Thirdly, a PI controller is designed for the model augmented by the engine and VPDV, and the parameters of the controller are optimized by the differential evolution (DE) algorithm. Moreover, the controller of gain scheduling is applied to the engine nonlinear model simulation platform. Simulation results illustrated that the designed digital speed closed-loop control can realize the function of speed regulation, and has the advantages of good servo performance and strong stability in comparison with the original hydraulic mechanical control system.
Analytical Model for the Heat Transfer in Impingement Cooled Spur Gears
Authors: Kromer Christian , von Plehwe Felix C. , Keller Marc C. , Schwitzke Corina , Bauer Hans-Jörg 
Published: January 2021

Abstract
The geared turbofan is a promising concept for civil aircraft jet engines. With the introduction of a gearbox between the low-pressure turbine and the fan, both components can rotate at their respective optimum speed. The geared turbofan enables a lower specific fuel consumption as well as jet engine noise reductions. A planetary gear train is usually chosen for the transmission with the sun gear connected to the low-pressure turbine. This high-speed reduction gear train needs to transmit high loads with a high efficiency in limited installation space. To ensure a safe operation of the gear train, the thermal behavior of the gears needs to be understood. The heat generated by the meshing processes is dissipated by oil impingement cooling. While the field of Elastohydrodynamic lubrication yields good results for the heat generation, no validated model for the impingement cooling process is available in literature. In this study, an analytical model is developed and validated against experimental data. First, the surface area of the oil film on the gear tooth flank formed by the impinging oil jet is calculated. Second, the heat transfer from the gear tooth flank to the oil film is determined. The fluid motion is modeled as an oil film that is flung off the gear tooth flank by centrifugal forces. In addition to the film flow, the presented model takes into account the temperature dependence of the viscosity of the oil and the initial oil film height. The effect of a lubrication oil film on the gear tooth flank before the oil jet impinges is included and its effect on the heat transfer is assessed. The analytical model agrees well with experimental results over the entire range of investigated operating conditions. Finally, a discussion on the effect of several assumptions in the derivation of the analytical model is presented. The validated analytical model can be used as an efficient tool for the design of gear trains with impingement cooled spur gears.
Investigations of Heat Transfer and Film Cooling Effect on a Worn Squealer Tip
Authors: Ye Mingliang , Yan Xin 
Published: January 2021

Abstract
Wear damage commonly occurs in modern gas turbine rotor blade tip due to relative movements and expansions between rotating and stationary parts. Tip wear has a significant impact on the aerodynamic, heat transfer and cooling performance of rotor blades, thus threatening the economy and safety of whole gas turbine system. Based on a simple linear wear model, this paper numerically investigates the aerodynamic, heat transfer and film cooling performance of a worn squealer tip with three starting-locations of wear (sl = 25%Cax, 50%Cax and 75%Cax) and five wear-depths (wd = 0.82%, 1.64%, 2.46%, 3.28% and 4.10%). Firstly, based on the existing experimental data, numerical methods and grid independence are examined carefully. Then, three dimensional flow fields, total pressure loss distributions, heat transfer coefficients and film cooling effectiveness in worn squealer tip region are computed, which are compared with the original design case. The results show that, with the increase of wear depth and the movement of wear starting-location to the leading edge, the scale and intensity of cavity vortex are increased, which results in the extended high heat transfer area on cavity floor near the leading edge. Wear makes more coolant flow out of the cavity, and reduces the area-averaged film cooling effectiveness at the bottom of cavity, but increases the film cooling effectiveness on pressure-side rim. The increase of wear depth makes more flow leak through the tip gap, thus increasing the scale and intensity of leakage vortex and further increasing the total pressure loss in the tip gap. Compared with the original design case, as the wear depth is increased from 0.82% to 4.10%, the mass-averaged total pressure loss in cascade is increased by 0.3–6.7%, the area-averaged heat transfer coefficient on cavity floor is increased by 1.7–29.1% while on squealer rim it is decreased by 3.1–26.3%, and the area-averaged film cooling effectiveness on cavity floor is decreased by 0.035 at most while on squealer rim it is increased by 0.064 at most.
Hybrid RANS/LES Simulation of Combustor/Turbine Interactions
Authors: Xia Guoping , Kalitzin Georgi , Lee Jin , Medic Gorazd , Sharma Om 
Published: January 2021

Abstract
Accurate prediction of thermal field in high pressure turbines is a critical aspect of aerodynamic and durability design. This is particularly true when the flow at turbine inlet exhibits large gradients in temperature, both radially and circumferentially. In other words, in the presence of hot streaks from the combustor. In the numerical study presented in this paper, coupled high-fidelity eddy-resolving simulations of a combustor and a turbine are used to study the differences in the temperature profile at the exit of the first vane and the heat flux on the first blade, resulting from different positioning, or clocking, between the combustor fuel nozzles and turbine vanes. The resolved unsteadiness and turbulence from the combustor impacts mixing and secondary flow in the high pressure turbine. Temperature profiles from both actual combustor CFD simulations, as well as and modulated profiles with more pronounced variation, or pattern factor, are used at the turbine inlet. A threshold of the pattern factor that brings the benefit of clocking is identified. Clocking positioning between the combustor and vanes was studied for the most benefit.
Validation of RANS Turbulence Models for Labyrinth Seal Flows by Means of Particle Image Velocimetry
Authors: Wein Lars , Kluge Tim , Seume Joerg R. , Hain Rainer , Fuchs Thomas , Kähler Christian , Schmierer Roman , Herbst Florian 
Published: January 2021

Abstract
Accurate prediction of labyrinth seal flows is important for the design and optimisation of turbomachinery. However, the prediction of such flows with RANS turbulence models is still lacking. The identification of modelling deficits and the development of improved turbulence models requires detailed experimental data. Consequently, a new test rig for straight labyrinth seals was built at the Institute for Turbomachinery and Fluid Dynamics which allows for non-intrusive measurements of the three dimensional velocity field in the cavities.
Two linear eddy viscosity models and one algebraic Reynolds stress turbulence model have been tested and validated against global parameters, local pressure measurements, and non-intrusive measurements of the velocity field. While some models accurately predict the discharge coefficient, large local errors occurred in the prediction of the wall static pressure in the seal. Although improved predictions were possible by using model extensions, significant errors in the prediction of vortex systems remained in the solution. These were identified with the help of PIV results. All turbulence models struggled to accurately predict the size of separations and the swirl imposed by viscous effects at the rotor surface. Additionally, the expansion of the leakage jet in the outlet cavity is not modelled correctly by the numerical models. This is caused by a wrong prediction of turbulent kinetic energy and, presumably, its rate of dissipation.
Squealer Tip Treatment Design for Axial Compressors
Authors: Bruni Giuseppe , Taylor James , Krishnababu Senthil , Miller Robert , Wells Roger 
Published: January 2021

Abstract
End-wall flows are amongst the main sources of losses in the rear stages of a typical multi-stage axial compressor. Reducing the tip leakage losses in the rotor blades and vanes can provide an increased efficiency and stall margin of a given axial compressor stage. One approach is to use squealer tips, which are traditionally designed to minimize the effect of tip rubbing. However, squealers can also provide a significant performance benefit, when designed considering aerodynamics from the beginning, as shown in this paper. A CFD based methodology, in which the blade or vane thickness distribution is varied in a controlled manner was developed. This design methodology was used to create different types of squealer tip geometry for a representative stage in a low speed compressor rig. Three different tip concepts were designed, based on a Suction Side Squealer, on a Pressure Side Squealer and on the combination of the two being merged between the leading edge and trailing edge, this new design is called the SuPr Tip. Subsequent experimental tests carried out agreed with the predicted relative ranking of the different squealer designs and on the superior performance of the SuPr tip design over the others, thus validating the methodology and the design process.
Aerodynamic Investigation of a Composite Low-Speed Fan for UHBR Application
Authors: Rodrigues Martin , Soulat Laurent , Paoletti Benoit , Ottavy Xavier , Brandstetter Christoph 
Published: January 2021

Abstract
A composite fan stage representative of a modern UHBR architecture has been investigated experimentally on a novel test facility at Ecole Centrale de Lyon. These measurements show indications for strong overloading of the tip region resulting in extensive blockage of the blade passage.
The performance of the fan is analyzed with extensive instrumentation including radial profiles upstream and downstream of the rotor. Unsteady pressure measurements help to interpret the flow structure in the tip region. The results are presented across a range of operating points on the design speedline. At the stability limit, the machine suffers from Non-Synchronous Vibrations which result from small scale aerodynamic disturbances propagating between the leading edges. A detailed analysis on the occurring waveforms is presented for two operating speeds.
In order to further analyze the observed phenomena, a numerical study has been conducted using the RANS solver elsA. The results of steady calculations are discussed in comparison with the detailed experiments. Unsteady simulations near the stability limit accurately predict the aerodynamic disturbances observed during NSV.
The obtained results are unusual for typical state-of-the-art transonic fans, as they show the same behavior as high-pressure compressor front stages, dominated by blockage caused by tip leakage flow. Even though flutter is not observed, the observed Non-Synchronous Vibration mechanism is a critical aeroelastic phenomenon which is of great interest for future designs of low speed fans.
Pumping Loss of Shrouded Meshed Spur Gears
Authors: Hurrell Michael J. , Sawicki Jerzy T. 
Published: January 2021

Abstract
High speed rotorcraft transmissions are subject to load-independent power losses consisting of drag loss and pumping loss. Tightly conforming shrouds enclosing the transmission gears are often incorporated to reduce the drag component of the total load-independent losses. However, tightly conforming axial shrouds can result in an increase in the pumping loss component. Quantifying the pumping loss of shrouded gear transmissions has been the subject of many studies. This study presents a new approach for estimating pumping loss based on the concept of swept volume and examines the applicability of the approach to various shroud configurations.
The drag loss and pumping loss of a shrouded spur gear pair has been determined through testing using the NASA Glenn Research Center (GRC) Gear Windage Test Facility. The results from this testing have been compared to theoretical results using the formulations presented in this study. In addition, computational fluid dynamic (CFD) analysis has been conducted for the various shroud configurations tested at NASA GRC. The results from the CFD analysis confirm the theoretical and empirical results and provide insight into the applicability of the swept volume approach for estimating pumping power loss of shrouded gear transmissions.
Nonlinear Stability Analysis of a Generic Fan With Distorted Inflow Using Passage-Spectral Method
Authors: Romera David , Corral Roque 
Published: January 2021

Abstract
The dependence of the aerodynamic stability of fan blades with the nodal diameter and amplitude of the inlet perturbations is studied. The analysis is conducted using a block-wise spatial Fourier decomposition of reduced-passages to reconstruct the full-annulus solution. The method represents very efficiently the unsteady flow generated by inlet non-uniformities. The explicit spatial Fourier approximation is exploited to characterize the relevance of each nodal diameter of the inlet perturbation in the fan stall process and study the nonlinear stability in a harmonic by harmonic basis. This approximation allows studying the contribution to stall of each circumferential mode separately. The methodology has been assessed for the NASA rotor 67. The maximum amplitude of total pressure distortion at which the compressor becomes unstable and triggers a stall process has been mapped. It has been proven that despite the complexity of a screen-induced total pressure distortion the only relevant parameter for the nonlinear stability of the fan is the most unstable nodal diameter. The equivalence in terms of stability between realistic distortion screens and single harmonic distortions has been assessed. Full-annulus simulations have been conducted to assess the accuracy of the simplified nonlinear stability limit. It is concluded that performing a nonlinear simulation with the proper single harmonic perturbation is enough to assess fan stability. The total pressure error with respect the full annulus simulation including a screen-induced pressure deficit at the intake is below 10%. It is shown that for the NASA rotor 67 running at the nominal speed the most unstable nodal diameter is the first. This study not only shows a reduction in computational time to assess nonlinear fan stability by a factor of seven but also creates an efficient methodology for understanding the nonlinear instability of fans due to inlet distortion profiles.
Wall-Modelled Large Eddy Simulations of Axial Turbine Rim Sealing
Authors: Palermo Donato M. , Gao Feng , Amirante Dario , Chew John W. , Bru Revert Anna , Beard Paul F. 
Published: January 2021

Abstract
This paper presents WMLES simulations of a chute type turbine rim seal. Configurations with an axisymmetric annulus flow and with nozzle guide vanes fitted (but without rotor blades) are considered. The passive scalar concentration solution and WMLES are validated against available data in the literature for uniform convection and a rotor-stator cavity flow. The WMLES approach is shown to be effective, giving significant improvements over an eddy viscosity turbulence model, in prediction of rim seal effectiveness compared to research rig measurements. WMLES requires considerably less computational time than wall-resolved LES, and has the potential for extension to engine conditions. All WMLES solutions show rotating inertial waves in the chute seal. Good agreement between WMLES and measurements for sealing effectiveness in the configuration without vanes is found. For cases with vanes fitted the WMLES simulation shows less ingestion than the measurements, and possible reasons are discussed.
Experimental Investigation on the Effect of Varying Purge Flow in a Newly Commissioned Single Stage Turbine Test Facility
Authors: Payne Daniel , Kanjirakkad Vasudevan 
Published: January 2021

Abstract
In order to produce efficient engines it is essential for gas tur-bine designers to understand the interaction between the primary and secondary air systems in critical parts of the engine. One of these is the first stage turbine, where the ingress of the hot an-nulus air into the rotor stator cavity could be catastrophic due to the increased heat load on the disc posts and on the rotor blades themselves (through reduced cooling). To ensure that this does not happen, contactless seals (rim seals) are built into the outer radius of the rotating disc. Additionally, a secondary air flow rate must be appropriately set in order to ‘purge’ the hot air that could be ingested into the rim seal cavity. However, this purge airflow could cause deterioration of the turbine performance as it re-joins the main annulus flow at the interface between the rim seal cavity and the main annulus. The deterioration in performance is pri-marily due to the difference in kinematic (flow velocity and mass flow) and thermodynamic (density, enthalpy) properties of the two stream of air. It is therefore essential to understand the optimum seal geometry and purge flow rates required to prevent the ingestion of the hot annulus air while maintaining the required turbine performance. In this paper we present experimental test results from a single stage turbine facility, the Rim Seal (RiSe) rig, at the University of Sussex. The turbine stage incorporates a model rotor-stator cavity system that is representative of the first stage turbine in a gas turbine engine. The facility is capable of generating disc cavity rotational Reynolds numbers of the order of 2.2 × 106 and axial Reynolds number of the order of 0.7 × 106, while operating at a pressure ratio of 2.5. The paper will present the salient features of the test facility, the various instrumentation employed, and the operating specifications of the stage. The paper will discuss the effect of varying the purge flow for a fixed operating point of the turbine. Results presented will include typical mission profiles, cavity radial temperature distribution, and the measured cavity sealing effectiveness.