bluuebunny/crossref_metadata_2025_split huggingface dataset:
From Concept to Wind Tunnel Model: Efficient Design Methodology for Innovative Low-Noise Propellers
Authors: Baert Lieven , Grasso Gabriele , Sainvitu Caroline , Lepot Ingrid , van Enkhuizen Marinus J. , Lammers Karel , Bown Nicholas 
Published: October 2022

Abstract
Whereas the market-share for propeller-driven aircraft is expected to rise to 30–40% of Europe’s regional aviation segment, a significant reduction in perceived noise levels remains a critical challenge for the success of further turboprop deployment. Within the framework of the Clean Sky 2 initiative, the Regional Aircraft platform focuses on the integration of technologies and methodologies, and their validation through full scale demonstrators, in order to allow for breakthrough technologies and competitive regional aircraft products.
This paper discusses the challenges for a low-noise propeller design, inherently multidisciplinary, and presents the design methodology employed from concept to wind tunnel model. Critical in the design process of disruptive and innovative concepts is the blade deformation from the manufactured (cold) shape to the running (hot) shape that must be adequately considered. An efficient online surrogate-based optimisation (SBO) approach has been followed where the machine learning models are locally enriched to improve their accuracy in the zones of interest of the design space. Throughout the optimisation feedback was integrated by dynamically strengthening and/or enforcing new constraints, resulting in a propeller design that satisfies all design requirements and achieves a significant near-field noise reduction of 6.4 dB compared to the baseline geometry. After performing a detailed structural assessment, the titanium propeller model has been manufactured and prepared for low-speed and high-speed wind tunnel testing. Preliminary results from the ongoing tests confirm the low-noise propeller’s potential.
Investigation of Tip Leakage Vortex Breakdown in a High-Speed Multistage Axial Compressor
Authors: Seki Ryosuke , Azuma Toshihiko , Iwatani Junji , Nakaniwa Akihiro , Okui Hidetaka , Shibata Takanori 
Published: October 2022

Abstract
In this paper, an unsteady tip leakage flow instability is identified and investigated for an axial compressor at near surge conditions. We describe the results of experimental verification of a new compressor developed by improving the blade geometry based on the criterion for the occurrence of this unsteady phenomenon.
In a high-speed multistage axial flow compressor having a subsonic high stagger rotor blade, a surge test was carried out by changing the tip clearance. Under a condition of large tip clearance, a drastic decrease in static pressure rise coefficient near the surge point was observed. At this operating condition, large, unsteady pressure fluctuation at the blade tip was confirmed, and the occurrence of tip leakage vortex breakdown that was clarified by unsteady multipoint pressure measurement and detailed unsteady numerical simulations. Due to the blockage effect caused by vortex breakdown of the tip leakage, double leakage and axially reversed flow near the trailing edge was observed. It is found that the vortex breakdown region of the tip leakage vortex propagated in the circumferential direction and caused the rotating instability.
In order to investigate the relationship between this unsteady flow phenomenon, tip clearance size and flow pattern, unsteady calculation was conducted by changing the blade tip stagger and tip clearance size. A new concept of tip clearance of staggered pitch reference is proposed, which makes it possible to include the effect of blade loading on the clearance and clarifies that there exists a threshold at which vortex breakdown occurs/does not occur.
On the basis of the above results, a high-speed multi-stage improved compressor was designed and manufactured to prevent tip leakage vortex breakdown. A clearance change test using active clearance control technology was conducted and increase in the static pressure rise coefficient near the surge point was confirmed for each clearance. The design concept of the improved blade which suppressed the unsteady tip leakage flow instability was tested and verified, and the effectiveness of the design guideline in actual gas turbines for power generation was confirmed.
Identification and Stability Analysis of a Turbojet Engine in Start-Up Mode
Authors: Abdulhamitbilal Erkan , Şal Sinan , Jafarov Elbrous M. 
Published: October 2022

Abstract
This paper investigates the system identification and stability analyses of a turbojet engine in start-up mode. Stability analyses are performed for automatic feedback control algorithm which switches engine from start-up to operational mode. This process requires engine rotational speed and acceleration data which are obtained via extensive engine tests. Those tests are carried out on an altitude testbed at sea level static conditions to obtain above mentioned data with respect to the fuel control command. The process consists of construction of a gain scheduled engine dynamics model and Lyapunov stability analyses. Engine dynamics are described as transfer functions. Mathematical model parameters are identified from test data to perform feasible stability analyses and to define feasible stability regions. The contribution of this study for a turbojet engine is the determination of stability maps in start-up mode and stability region or so-called strip for enabling strategy of the automatic feedback control law. The theoretical results are supported by simulations and an engine test example. Results show that the reference speed of the feedback control algorithm should be NREF ≤ NTJ + 10% for a successful stabilization while switching from start-up to operational mode.
Experimental Study of Film Cooling With Favorable and Adverse Pressure Gradients
Authors: Volino Ralph J. , Gillcrist Matthew C. 
Published: October 2022

Abstract
Film cooling experiments were conducted on a flat wall subject to favorable and adverse pressure gradients with constant acceleration parameter, K. The test wall included a single row of 5 round holes in line with the flow direction and inclined at 35 degrees to the surface. The hole spacing was 3 diameters. The wall opposite the test wall was moveable, and was set to angles with respect to the test wall that produced K values of −0.5 × 10−6, 0, 1 × 10−6, 2 × 10−6, 2.5 × 10−6, and 3 × 10−6. Blowing ratios of 0, 0.5, 1, and 1.5 were tested at each acceleration. The test wall was equipped with constant flux surface heaters, and data were acquired for each flow condition with the wall both heated and unheated. An infrared camera was used to record wall temperature in a region spanning the three center holes and extending 20D downstream of the holes. From these measurements, local film cooling effectiveness and heat transfer coefficients were determined. In the flow, velocity and temperature profiles were acquired using hot-wire anemometry and a traversing thermocouple probe. Particle image velocimetry was used to acquire velocity fields in a plane perpendicular to the flow direction and 10D downstream of the holes. The pressure gradient had a noticeable effect on the flow, with the favorable pressure gradient reducing liftoff and moving the film cooling jets closer to the wall, and reducing turbulence levels in the boundary layer. The adverse pressure gradient had the opposite effect. Near the wall, however, the effects of the pressure gradient on the film cooling effectiveness and heat transfer were more subtle.
Film Cooling Hole Shape Effects on Turbine Blade Heat Transfer – Part I: Computational Comparison to Experiment
Authors: Sperling Spencer , Christensen Louis , Mathison Randall , Aksoy Hakan , Liu Jong , Nickol Jeremy 
Published: October 2022

Abstract
Gas turbines generate highly unsteady flow fields, which are further complicated by cooling. Film cooling technology is dependent on experimental and computational research performed for simplified geometries, which can be difficult to translate to turbine domains. To help explain disconnects between flat plate experiments and turbine operation, this study performs computational research, grounded with experimental measurements, examining the film cooling performance of several different hole shapes on rotating turbine blades.
Round, fan, and advanced anti-vortex hole geometries are incorporated into experimental hardware and computational models, and pressure and heat flux results form a basis for comparisons between computation and experiment as well as between the cooling geometries. Unsteady and Steady simulations are both evaluated, and results indicate significant model accuracy improvement from the inclusion of unsteady consideration.
Shaped film cooling holes are observed to provide stronger film effectiveness traces on the blade. Both the fan and advanced shaped film cooling holes generate stronger cooling jet cores that remain close to the blade wall on the pressure surface, suction surface, and near the leading edge. Advanced shaped holes provide increased lateral spread and increased resistance to radial migration.
The results of this study help identify benefits of using shaped film cooling on the turbine blade as well as the mechanisms generating the cooling benefits. This will help designers weigh the manufacturing costs of shaped film cooling holes as well as identify areas of the blade where shaped film cooling is needed and where it is not. Additionally, this study observes significant improvements in blade heat transfer predictions through unsteady treatment alone, indicating better computational agreement can be achieved by leveraging lower-cost RANS simulation tools.
Design Optimization of Integrated Anti-Rotation Feature for Power Turbine Nozzles
Authors: Soman Abhimanyu , Colantoni Simone 
Published: October 2022

Abstract
Gas turbine nozzles are static components that are meant to turn and accelerate high temperature, high pressure, and low-velocity flue gas into the downstream turbine row of buckets. During gas turbine operation, nozzles are subjected to high-pressure load due to the expansion of flue gases, in axial and tangential directions. This creates a tendency for nozzle movement in tangential direction which has potential to create flow disturbance and intersegment gap opening. To prevent this movement, it should be held in tangential direction firmly by introducing an anti-rotation feature. A slot is introduced in the nozzle outer sidewall and a pin connected with casing in such a way that the nozzle’s tangential movement is restrained. As the nozzle’ s outer sidewall experiences high thermal gradients in the operating condition, it induces high stress at the nozzle anti-rotation feature. There are many possible design options available to mitigate this challenge. In the present work, anti-rotation feature is integrated with the nozzle’s outer sidewall and a matching slot is provided in the casing. A detailed study is performed to optimize this anti-rotation feature to reduce high thermal-mechanical stress and thereby improve reliability. The low cycle fatigue life is one of the vital requirements in improving reliability. The low cycle fatigue life of the optimized anti-rotation feature is validated using the finite element analysis. This paper describes the process step details in optimizing the anti-rotation feature.
Effect of Vane/Blade Interaction for a Cooled Turbine Blade — Part I: Comparison of Experimental and Numerical Results
Authors: Sperling Spencer , Christensen Louis , Mathison Randall , Aksoy Hakan , Liu Jong , Nickol Jeremy 
Published: October 2022

Abstract
Flow fields in axial gas turbines are highly complex and unsteady due to the interactions of rotating and stationary components. The current study incorporates experimental and computational analyses of a modern, rotating high-pressure turbine stage to examine the nature and effects of unsteadiness at vane passing scales. Increased knowledge of the unsteady flow field and its impacts on heat transfer will help engineers to produce higher quality thermal predictions of gas turbine components with efficient use of computing resources.
Computational research was carried out using single sector simulations using steady, sliding mesh, and harmonic balance treatments. Simulation results are compared to measurements of heat flux and pressure on the rotating airfoils of a cooled turbine stage operating at design corrected conditions.
Unsteady analysis of experimental and computational results highlights the importance of unsteady considerations in gas turbine research. Simulation results of unsteady pressure and heat transfer linked to vane passing perform well in some regions of the blade, but single sector computations cannot predict frequencies below vane passing frequency. It is therefore important to perform experiments and simulations in turbine representative environments.
Comparisons between results from the various computational models and experimental data form a basis for evaluating the benefits and drawbacks for each type of model treatment. Harmonic balance and sliding mesh simulations both increase the level of agreement between heat transfer data and results to within unsteadiness bounds, and each tends to have specific advantages. Harmonic balance simulations are particularly well equipped for comparisons in the frequency domain, but nearly double the memory requirement, making them more expensive to perform.
Part I of this paper focuses on the computational approaches implemented and how they compare to experimental results. Part II investigates the unsteady mechanisms that cause differences between steady and time-accurate solutions, as well as how they impact the flow over the rest of the airfoil.
Effect of Vane/Blade Interaction for a Cooled Turbine Blade — Part II: Unsteady Pressure and Heat Transfer Distributions
Authors: Sperling Spencer , Christensen Louis , Mathison Randall , Aksoy Hakan , Liu Jong , Nickol Jeremy 
Published: October 2022

Abstract
Unsteady flow fields are generated in gas turbines due to interactions between rotating and stationary components. Increased understanding of unsteadiness and its effect on blade heat transfer is crucial to thermal durability predictions for turbine components. This study examines unsteady pressure, heat transfer, and film cooling on a rotating high-pressure turbine stage using experimental and unsteady computational tools. Comparisons between experimental and computational results identify several key elements necessary for successful computational models. Models benchmarked against experimental data provide valuable insight into the nature of flow field unsteadiness and its impact on blade heat transfer.
Comparisons presented in Part I of this paper identified a region near the leading edge of the suction surface of the airfoil where the computational and experimental results did not agree. Part II of the paper will examine the mechanisms driving unsteadiness in this region and how they can be better predicted. The source of this disagreement is that vane blade interaction causes large swings in incidence angle at the leading edge, which causes periodic separation on the suction surface. In addition, pressure fluctuations couple with unsteadiness in cooling flow at the leading edge to further amplify unsteadiness.
Examining the links between pressure and heat transfer fluctuations shows that the pressure surface and suction surface each had strong but unique correlations, and the leading edge is complicated by coupling between pressure and shower head cooling unsteadiness. This indicates a different turbulent Prandtl number may be required in different regions of the blade to accurately track heat transfer. Further complicating heat transfer predictions, small changes in the velocity field cause large levels of unsteadiness in film cooling. The relative movement of large-scale vortex structures over a vane pass increase time-averaged dissipation of cooling jets beyond steady prediction levels.
The results of this study shed light on the nature of unsteadiness, highlighting specific components of unsteadiness observed to augment heat transfer. Better characterizations of unsteadiness can help designers utilize computationally efficient design tools. With increased understanding of how unsteadiness impacts time-averaged performance, a more accurate representation of time-averaged performance can be leveraged from low-cost RANS computational tools.
Experimental Evaluation of a Wavy Trailing Edge Cooling Design as an Alternative to Pressure Side Cutback Cooling
Authors: Ullah Izhar , Burdett Timothy A. , Wright Lesley M. , Han Je-Chin , Lee Ching-Pang 
Published: October 2022

Abstract
This study measures the film cooling effectiveness along the trailing edge of a turbine blade in a linear cascade. The film cooling effectiveness is measured and analyzed using the pressure sensitive paint (PSP) technique. Two different trailing edge designs are investigated including the standard pressure side cutback and the new alternating discharge design (referred to as a wavy trailing edge design). The alternating discharge design is a new design with a wavy structure between the pressure and suction surfaces at the trailing edge. The new wavy structure allows the coolant to discharge from the trailing edge so that it alternates between the pressure and suction surfaces. Testing is carried out in a five-blade, linear, steady state cascade with inlet and exit Mach numbers of 0.20 and 0.30, respectively. The freestream turbulence intensity is measured to be 10.5% upstream of the blade leading edge. Coolant-to-mainstream mass flow ratios (MFR) vary from 0.30% to 1.25% and coolant to mainstream density ratios (DR) of 1.0, 1.5 and 2.0 are examined. The total pressure loss coefficients are also acquired to compare the aerodynamic loss between the two trailing edge designs. The pressure loss coefficient is acquired by traversing a group of pitot static probes across the blade span in a plane downstream of the trailing edge, resulting in a pressure map at the exit plane. A positive MFR and DR effect is witnessed with almost no change in aerodynamic loss. The results provide the potential of the alternating discharge as a promising trailing edge design and provide gas turbine designers with an improved trailing edge cooling scheme having acceptable aerodynamic loss.
The Validation of a Parametric Leading Edge Model for Probabilistic CFD Analyses of Post-Service Compressor Airfoils
Authors: Schlüter Lukas , Voigt Paul , Voigt Matthias , Mailach Ronald , Schmidt Robin , Rostamian Mirco , Becker Bernd 
Published: October 2022

Abstract
The impact of geometric variation of compressor airfoils due to manufacturing scatter and operational effects on the aerodynamic performance is a well-addressed topic in the research field of turbomachinery. Thereby, the airfoils’ leading edge (LE) shapes are highly relevant to their aerodynamic behavior, potentially causing increased losses and separation. In previous probabilistic computational fluid dynamics (CFD) analyses, the LEs were usually modeled as semi-circles or semi-ellipses, whereas their noticeable deviation from these simplified shapes is given comparatively little attention. The study at hand aims to fill this gap by introducing a parametric model for the quantification and parametric recreation of the variability of high pressure compressor (HPC) LE shapes. A set of digitized mid stage HPC rotor blades originating from several jet engines with differing fields of operation are used for the demonstration of the introduced approach. The individual LE 2D-contours are extracted on multiple sections over the blade’s height and compared to a symmetric reference shape. This yields characteristic difference curves enabling the detailed quantification of the variability with a minimum amount of parameters. The derived parametric model allows the statistical reproduction of the measured variability, which is geometrically and aerodynamically validated in the following. The approach’s capabilities in regard to probabilistic CFD analyses are demonstrated by conducting and evaluating two- and three-dimensional flow simulations. Furthermore, the observed geometric variability is linked to the rotor row’s aerodynamic behavior highlighting its importance for the performance variance of HPC rotors.
Interface Reduction in an Equivalent Linearization Algorithm for Nonlinearly Coupled Systems Under Random Excitation
Authors: Förster Alwin , Panning-von Scheidt Lars 
Published: October 2022

Abstract
Stochastic excitation is a rarely discussed topic regarding the vibrational behavior of turbine blades. While the calculation of stochastic quantities describing the stochastic steady-state vibration response is comparatively straightforward in the linear case, it becomes much more challenging in the case of nonlinear couplings. A method suitable for calculating approximations of the stochastic steady-state vibration response is the equivalent linearization method. However, the efficiency of this method decreases with an increasing number of degrees of freedom. This is especially challenging if the number of nonlinearly coupled degrees of freedom is large, as in the case of an extended contact interface, such as a shroud contact. While the uncoupled part of the system can be reduced using a component mode synthesis, the remaining nonlinear interface degrees of freedom have to remain unreduced to evaluate the nonlinear forces. To address this problem, this paper presents an approach to reduce the interface degrees of freedom within the framework of the equivalent linearization method. The presented method is based on a representation of the dynamics of the contact interface by means of a reduced set of Legendre polynomials. However, the evaluation of the nonlinear force still takes place in physical coordinates. The presented procedure is demonstrated using a beam model with different contact pairings as well as a more realistic model of a bladed disk assembly.