bluuebunny/crossref_metadata_2025_split huggingface dataset:
Aerodynamic Effects of an Unshrouded Low Pressure Turbine on a Low Aspect Ratio Exit Guide Vane
Authors: Selic Thorsten , Lengani Davide , Marn Andreas , Heitmeir Franz 
Published: July 2013

This paper presents the effects of an unshrouded low pressure turbine (LPT) onto the following exit guide vane row (EGV). The measurement results were obtained in the subsonic test turbine facility at Graz University of Technology by means of a fast response pressure probe in planes downstream of the rotor as well as oil flow visualisation. The test rig was designed in cooperation with MTU Aero Engines and represents the last 1.5 stages of a commercial aero engine. Considerable efforts were put into the adjustment of all relevant model parameters to reproduce the full scale LPT situation.
Different tip clearances were evaluated by means of CFD obtained using a commercial Navier-Stokes code and validated with experimental results. The goal is to evaluate the effect of the varying leakage flow on the flow in the low aspect ratio EGV. Special attention is given to the impact on the development of secondary flows as well as the flow structures downstream of the EGV. The effect of the leakage flow causes a change of the flow structure of the EGV, particularly losses. Considering the largest investigated tip-clearance, the losses increased by 71% when compared to a zero-leakage case.
Experimental and Numerical Study of a Subsonic Aspirated Cascade
Authors: Godard Antoine , Bario François , Burguburu Stéphane , Lebœuf Francis 
Published: July 2013

This paper presents the validation of a design method for aspirated compressor blades, combining a passive separation control by blade shaping with an active flow control by aspiration.
In a first part, a linear aspirated cascade designed according to this method was built and tested at low speed, without and with aspiration. The latter was only applied on the suction surfaces of the blades. Particle Image Velocimetry measurements performed at mid-span of the cascade, in the central passage, showed a complete reattachment of the separated boundary layer on the suction side of the blade. A flow deflection of approximately 65 degrees was achieved requiring an aspirated mass flow rate of 3.3%. However, boundary layer reattachment is effective in a zone centered at mid-span covering 30% of blade span. Flow visualization revealed large corner separation in the presence of aspiration. This is due to the re-establishment of strong pressure gradient on sidewalls of the cascade. No flow control was applied on these zones for optical access purpose. These secondary-flow regions reduced the diffusion occurring within the cascade by nearly 60% in comparison with the design intent. They also increased the expected level of total pressure losses measured by wake traverses downstream of the cascade.
In a second part, numerical simulations of the aforementioned experiment were carried out to help the understanding of the experimental results. The simulations were able to reproduce correctly the characteristic flow features, without and with aspiration, observed and measured during the experiment. Thus, they confirmed the potential of this design method developed for aspirated compressor blades, as well as CFD capabilities to simulate the influence of technological effects like suction slots. A uniform and a non-uniform aspiration distribution along the blade span direction were considered during simulations. Suction distribution was found to have a significant impact on the control by aspiration. This design feature, in addition to flow control on endwalls, has to be taken into account in the three-dimensional design of highly loaded aspirated compressor blades.
Analysis of Combustor/Vane Interaction With Decoupled and Loosely Coupled Approaches
Authors: Salvadori Simone , Riccio Giovanni , Insinna Massimiliano , Martelli Francesco 
Published: July 2013

Numerical techniques are commonly used during both design and analysis processes, mainly considering separated components. Technological progress asks for advanced approaches that allow to analysing the interaction between the components, especially when considering combustor/turbine interaction. Hot spots and inlet swirl profiles generated by the combustor have been demonstrated to affect high-pressure turbine performances and reliability.
This work deals with the investigation of the effects of realistic boundary conditions for the high-pressure turbine vane, also proposing an approach for coupled simulation of the combustor/vane interaction. The method consists in a loosely coupled approach for the data exchange on the combustor/vane interface section. Data from the combustor exit section (stagnation conditions, velocity profile and turbulent quantities) are provided to the vane inlet and vice versa (for the static pressure). The proposed method is applied to a test case consisting of a redesigned combustor and the vane of the MT1 test case from QinetiQ.
A preliminary analysis was dedicated to define the combustor geometry and the operating conditions. Then, the MT1 working conditions have been rescaled and coupled with the combustor, maintaining the stage geometry and the experimental non-dimensional parameters. Second order accurate steady simulations were performed for both combustor and high-pressure turbine vane.
Calculations with a uniform profile and a theoretical nonuniform inlet profile (deriving from the EU funded TATEF2 project) have been considered as representative of commonly used approaches. The results obtained for the stator in terms of isentropic Mach number and Nusselt number along blades surfaces and inner end-wall are compared with each other and with the available experimental data.
Due to the large dimension of computational grids a parallel approach was applied. The activity was carried out using the IBM PLX supercomputer in the frame of the FrUIT project supported by CINECA.
Reconstruction of the Unsteady Pressure Field in a Low Speed Linear Cascade
Authors: Förster Felix , Sims-Williams David , Ingram Grant 
Published: July 2013

This paper describes investigations into unsteady flow phenomena in a high turning linear cascade with a nominal steady flow. A fast response five hole probe was used to obtain time-accurate measurements. Since a single probe was used to traverse an entire measurement plane, the results were obtained as discrete points and were not synchronous. One method used to address this issue is to traverse the measurement plane point by point and record the signal of a reference probe simultaneously. Synchronization can be performed during post processing by analyzing the reference signal for a common periodic component and phase-shifting all signals with respect to the reference phase of that flow feature.
This paper presents results from the application this unsteady reconstruction technique to the so-called Durham Cascade 28% downstream of the trailing edge in order to provide insight into the nature of the dominant unsteady mode. It was found that there was significant temporal variation in the strength of the passage vortex but the position of the vortex remains fixed.
Design Optimization and Experimental Verification of Centrifugal Compressors With Curvilinear Element Blades
Authors: Sugimura Kazuyuki , Kobayashi Hiromi , Nishida Hideo 
Published: July 2013

This paper describes the optimal design and experimental verification of centrifugal compressors with leaned curvilinear element blades. The design targets were a fully shrouded centrifugal impeller and a low-solidity vaned diffuser.
First, a new method of defining the curvilinear element blade was developed for centrifugal turbomachinery using coordinate transformations between a revolutionary flow-surface coordinate system and a cylindrical coordinate system. All the blade sections in the transferred cylindrical coordinate system were moved and stacked spanwise to form a new leaned blade surface. The inverse transformation results in a curvilinear element blade in the original coordinate system. The direction of movement of the blade section could be any of the existing definitions such as sweep, dihedral, or tangential lean. For simplicity, we have defined a “lean profile” as a general expression for the spanwise distribution profile of movement of the blade sections.
Model compressors with curvilinear element blades were then designed at a suction flow coefficient of 0.073. Optimal lean profiles for the impeller and vaned diffuser that maximized adiabatic efficiency and uniformity of outflow were explored using a multi-objective genetic algorithm, Kriging surrogate model, and steady Reynolds-averaged Navier Stokes simulations. We chose efficiency-weighted solutions since the efficiency and uniformity had a positive correlation. A sensitivity analysis showed that tangential leans near the endwalls are keys for improving the impeller’s efficiency. The chosen optimal impeller had a concave blade suction surface and a concave leading edge. Although clear patterns in geometrical features for optimal diffuser vanes could not be captured, we found that the dihedral profile had a predominant effect on efficiency.
The model compressors was experimentally measured and compared with traditional compressors as to their aerodynamic performance. The experimental apparatus was composed of a suction nozzle, impeller, diffuser, and return channel. The results demonstrated that the models’ adiabatic efficiency was higher by 1.2–1.4% at the design point. Although the stall margins slightly decreased (by 0.7%) in the experiments, the surge margins expanded (by 3.7–6.7%).
Multi Disciplinary Design Optimization and Performance Evaluation of a Single-Stage Transonic Axial Compressor
Authors: Kang Young-Seok , Park Tae-Choon , Yang Soo-Seok , Lee Sae-Il , Lee Dong-Ho 
Published: July 2013

The multi-disciplinary optimization (MDO) method, which integrates aerodynamic performance and structural stability, was utilized in the development of a single-stage transonic axial compressor. Numerical simulations and compressor tests were also carried out to evaluate the aerodynamic performance and safety factor of the optimized compressor. The rotor has 60 design parameters with twelve most sensitive design variables selected for design optimization. The stator was redesigned according to the rotor outlet flow angle variation to match the stator incidence angle by −1∼0 degrees, while maintaining the stage outlet flow angle. The design goal is to maximize both the stage efficiency and the safety factor from the baseline scratch compressor design. The object function is composed of the normalized efficiency and safety factor with weight factors. Initially, an approximation model was created to search for the global optimization within given ranges of variables and considering several design constraints. The genetic algorithm was used to explore the Pareto front of the optimization to find the maximum objective function values. The final design was chosen after a second stage gradient-based optimization process to improve the accuracy of the optimization.
The CFD results showed that more blade loading is burden to the hub region by increasing the incidence angle. The fore part blade loading gradually decreases along the span-wise direction. In addition, normal shock, which spreads along the hub to the blade tip, is confined in the rotor flow passage and pressure surface shock coincidence point moves to be closer to the blade leading edge, indicating an increase in the amount of blade loading. FEA analyses showed that the blade root stress has been drastically relieved, because the optimized blade has trapezoid-shaped hub design relative to the baseline design. The final design achieved efficiency gain of 3.69% and showed a higher safety factor by 2.3 times relative to the baseline model, while maintaining its stage mass flow rate and total-to-total pressure within the design constraints. The compressor performance test data showed good agreement with the optimized design and CFD results. However, there is room for improvement in the optimization process to reflect off-design performance so as to secure more stable compressor operation ranges.
Numerical Investigation of the Effect of Rotor Blade Leading Edge Geometry on the Performance of a Variable Geometry Turbine
Authors: Huang Lei , Zhuge Weilin , Zhang Yangjun , Hu Liaoping , Yang Di , Wang Jiawei 
Published: July 2013

Variable geometry turbines are more and more widely used in diesel engines to meet the requirements of the stringent emission standards. The VGTs mostly operate at off-design conditions. At highly off-design conditions, there exist complex secondary flow structures and severe flow separation in the rotor passage, which deteriorate the turbine performance largely.
The influence of rotor blade leading edge geometries on the VGT performance was studied by CFD simulations. The blade angle distribution along the leading edge was varied while keeping the radial-fiber rotor construction. The effects of inlet sweep angle distribution and lean angle of the blade leading edge on the turbine flow fields and performance were investigated under different operating conditions.
Results show that the turbine with backswept leading edge has better performance at low U/C, while the turbine with forward swept leading edge has a higher efficiency under high flow rate conditions. With the same sweep angle distribution, the leading edge lean affects the flow fields in the rotor passage as well as the turbine performance significantly. The influence of blade lean on the turbine performance varies according to different swept blading and operating conditions.
A Path Towards the Aerodynamic Robust Design of Low Pressure Turbines
Authors: Bertini Francesco , Credi Martina , Marconcini Michele , Giovannini Matteo 
Published: July 2013

Airline companies are continuously demanding lower-fuel-consuming engines and this leads to investigating innovative configurations and to further improving single module performance. In this framework the Low Pressure Turbine (LPT) is known to be a key component since it has a major effect on specific fuel consumption (SFC).
Modern aerodynamic design of LPTs for civil aircraft engines has reached high levels of quality, but new engine data, after first engine tests, often cannot achieve the expected performance. Further work on the modules is usually required, with additional costs and time spent to reach the quality level needed to enter in service. The reported study is aimed at understanding some of the causes for this deficit and how to solve some of the highlighted problems.
In a real engine, the LPT module works under conditions which differ from those described in the analyzed numerical model: the definition of the geometry cannot be so accurate, a priori unknown values for boundary conditions data are often assumed, complex physical phenomena are seldom taken into account, operating cycle may differ from the design intent due to a non-optimal coupling with other engine components. Moreover, variations are present among different engines of the same family, manufacturing defects increase the uncertainty and, finally, deterioration of the components occurs during service.
Research projects and several studies carried out by the authors lead to the conclusion that being able to design a module whose performance is less sensitive to variations (Robust LPT) brings advantages not only when the engine performs under strong off-design conditions but also, due to the abovementioned unknowns, near the design point as well.
Concept and Preliminary Design phases are herein considered, highlighting the results arising from sensibility studies and their impact on the final designed robust configuration. Module performance is afterward estimated using a statistical approach.
On the Use of Synthetic Jets in Transonic Compressors
Authors: Biollo R. , Staffieri E. , Benini E. 
Published: July 2013

Several fluidic-based methods have been investigated by compressor designers with the aim of improving the aerodynamic behavior of compressor blades. Recently, synthetic jets have been proposed as a promising fluidic-based solution for reducing flow separation inside turbomachines. With the characteristic of zero net mass flux and non-zero momentum flux, synthetic jets could be very effective because at the suction part of the cycle the low-momentum fluid is sucked into the device, whereas in the blowing part a high momentum jet accelerates it. However, to the authors’ knowledge, the use of synthetic jets has never been experimented in transonic compressors. In these machines, synthetic jets could have a great impact on the thickness and stability of the blade suction side boundary layer arising after the interaction with the shock. With the aim to investigate that, a rake of 600 Hz sinusoidal wave form synthetic jets (with no radial velocity component) positioned into a thin radial slot on the blade suction side of the NASA Rotor 37 has been modeled and simulated. Simulations have been carried out using a validated 3D CFD (Computational Fluid Dynamics) unsteady model.
The study gave a clear understanding of positive and negative impacts of the simulated device on the local flow field during the entire jets working cycle. The shock structure and boundary layer resulted significantly affected in the proximity of the synthetic jets, with different effects depending on slot location and jet channel angle. The local perturbation then affected the overall flowfield with impacts on the overall rotor performances. Despite the reported cases were selected for the purpose of the analysis only, and not for the aerodynamic enhancement of the rotor, slight benefits in terms of overall performance were observed for almost all the cases.
Influence of Inlet Guide Vane Wakes on the Passage Flow in a Transonic Axial Compressor
Authors: Leichtfuss S. , Biela C. , Schiffer H.-P. , Heinichen F. 
Published: July 2013

Variable inlet guide vanes enhance the efficiency and stability of modern transonic compressors. The current quest for compact and highly efficient aero-engines requires for higher stage-loading and small axial gaps between adjacent blade rows, increasing interaction between blade rows and introducing further unsteadiness into the flow. Modeling these interactions is relevant to jet engine implementation.
Recent advancements in numerical simulation of unsteady flow in multiple blade rows create an additional way to investigate the flow patterns formed by the unsteady interaction. In the case of transonic compressors, the experimental and numerical database is small. The current article will contribute to this database by investigating the flow downstream of an inlet guide vane and the influence this flow has on the passage in a transonic compressor.
Unsteady flow phenomena are resolved by piezoresistive wall pressure tappings. Numerical and experimental data show that the interaction of blade rows influences the formation of the tip leakage vortex. Analyzing the vortex structure within the transonic compressor stage it is possible to show how the blade rows interact and how geometric modifications in the VIGV tip gap model influence simulation results.
On the Effect of Volute Tongue Design on Radial Turbine Performance
Authors: Suhrmann Jan F. , Peitsch Dieter , Gugau Marc , Heuer Tom 
Published: July 2013

With an increasing need for gas turbines with rather low flow rates in many industrial applications, e.g. decentralized power generation, aircrafts or automotive turbochargers, the development of small size radial turbines becomes more and more important.
A major step in the development of a radial turbine stage is the preliminary design, which is the definition of basic geometrical features and the calculation of general turbine flow parameters at the design point and within the operating range. These are mainly the rotational speed, the expansion ratio, the flow rate and in particular the expected turbine efficiency.
In a radial turbine stage, the volute component delivers the flow to the rotor wheel and according to the geometrical form it defines major flow parameters like the mass flow parameter or the absolute rotor inlet flow angle. Amongst others, the way the flow enters the turbine wheel represents one of the most important loss generating factors. Thus, on the one hand an approach is necessary for the calculation of the optimum rotor inlet flow angle, in order to avoid dispensable losses due to secondary flow in the turbine wheel region. On the other hand, the volute tongue generates flow non-uniformity which has an effect on the overall circumferential averaged rotor inlet flow angle. Furthermore, the local flow pattern downstream of the volute tongue can generate suboptimal flow conditions for the turbine wheel. Hussain and Bhinder [1] measured the flow field at the outlet of a vaneless volute at different circumferential positions and detected a variation of the outlet angle of about Δα = 10°. The authors conclusion was, that the influence on the stage performance of flow non-uniformity generated by the volute could exceed the one of pressure losses through the volute.
In this paper, the effect of different geometrical volute parameters on the flow condition especially at the turbine wheel inlet area is investigated. Experimental data of the influence of different volute tongue geometries on the flow field is difficult to generate. Hence, comprehensive numerical investigations are made using steady 3D-CFD calculations of the turbine volute as well as calculations of complete turbine stages including a turbine wheel geometry.
Based on the numerical results, a design guideline is developed to estimate the influence of the geometric volute parameters on the flow and to raise the quality of the preliminary design process.
The Effect of Four Part Gap Geometry Configurations for Variable Stator Vanes in a Compressor Cascade
Authors: Gottschall Marcel , Vogeler Konrad , Mailach Ronald 
Published: July 2013

The article describes numerical investigations on the influence of four different endwall clearance topologies for variable stator vanes to secondary flow field development and the performance of high pressure compressors. The aim of this work is to quantify the characteristics of different clearance configurations depending on the penny-axis position and the penny diameter for a typical operating range. All clearance configurations were implemented to a linear cascade of modern stator profiles. The analysis was introduced using a relative clearance size of 1.3% chord at three stagger angles and two characteristic Reynolds numbers to model the operating range on aircraft engines. 3D numerical calculations were carried out to gain information about the flow field inside the cascade. They were compared with measurements of a 5-hole-probe as well as pressure tappings on the airfoil and the endwall.
The CFD shows the clearance characteristics in good agreement with the measurements for the lower and the nominal stagger angle. Small gaps in the rear part of the vane have a beneficial effect on the flow field. In contrast, a clearance in the higher loaded front part of the vane always resulted in increased losses. Otherwise, the significant enhanced performance of a rear part gap, which was measured at the higher stagger angle, was not reflected by the CFD. The reduced mixing losses and the higher averaged flow turning even compared to a configuration without a clearance are not verified with the calculations. Large flow separations at the high stagger angle result in a two to four times higher underturning of the CFD in comparison to the experiments. The clearance effects to the characteristic radial loss distribution up to 40 % bladeheight also deviate from the measurements due to heavy mixing of clearance and reversed separated flow.