bluuebunny/crossref_metadata_2025_split huggingface dataset:
Experimental Investigation of the Performance of Different Bleed Air System Designs
Authors: Peltier V. , Dullenkopf K. , Bauer H.-J. 
Published: July 2013

Increasing the turbine inlet temperature to improve thermal efficiency and thus reduce fuel consumption can only be achieved with an optimal supply of cooling air to control the temperature of the turbine blades. The aim of this study is the optimization of bleed air off takes and the reduction of losses caused by the interaction between bleed air and main gas path.
Bleed air for cooling purpose is taken from the main gas path at several locations inside compressor by means of the bleed air system. This system must be designed to pick up to 15% out of the primary mass flow through the engine core. In most of the engine designs bleed air is collected by holes or slots, then lead through a manifold and distributed by one or more tubes to the components. The critical features are perturbations in the main gas path and pressure losses in the bleed system.
Several bleed air system designs are currently in use, each of them causing specific aerodynamic and thermal perturbations which can reduce the overall efficiency of the compressor stage and even lead to thermal distortions in the casing of the compressor and thus reducing the expected life time of the engine component.
To study the efficiency of several bleed air systems a combined experimental and CFD effort was initiated in the framework of the EU project MAGPI. To perform experimental tests, an annular but nonrotating rig has been built at the ITS based on designs chosen by means of CFD studies. The new rig has a modular structure so numerous variation of geometrical und fluid parameters could be studied: different bleed mass flows (4, 9, and 12%), bleed inlet designs, exit tube number and position as well as swirl in the main gas path. First results including pressure losses and flow pattern in the test section will be presented and discussed to improve the understanding of bleed flow systems and to provide a comprehensive data base for validation of CFD methods in engine development.
Systematic Optimisation of a Turbofan Bypass Duct System
Authors: Clemen C. , Albrecht P. , Herzog S. 
Published: July 2013

A reduction of CO2 emissions of aero engines can only be achieved by a reduction in the fuel consumption. For turbofan engines a major key for this is an increase in the engine bypass ratio to enhance the propulsive efficiency.
This leads to an increased mass flow in the bypass duct and hence higher contributions of the bypass duct system to the flow pressure losses and hence the engine fuel consumption. Additionally the turbofan core engine gets smaller in size and the overall diameter of the engine increases. This requires advanced engine mounting concepts, such as core mounted subsystems, which influence the arrangement and the mechanical and aerodynamic design of components in the bypass duct such as fan outlet guide vane, struts, fairings or bifurcations.
This requires an optimized design of the turbofan bypass system with regards to structural, aerodynamic and acoustic criteria. The topology of the bypass duct as well as the position and the design of the individual components have a significant impact on the weight, the efficiency and the noise radiation of the engine and hence need to be investigated in a systematic approach.
Such a design approach was developed in the R&T project OPAL, led by Rolls-Royce Deutschland and funded by the German Federal State of Brandenburg.
The design approach covers the following steps:
- Optimization of the bypass duct shape with regards to minimum pressure loss taking into consideration the mechanical requirements of an engine with core mounted subsystems.
- Optimization of strut, fairing and bifurcation shapes with regards to minimum pressure loss, robustness against flow deviations and minimized upstream flow effects taking into consideration structural and mechanical requirements by engine loads and subsystem routings.
- Optimization of fan outlet guide vane profile and plan form shape with regards to minimum pressure losses and maximum working range taking into consideration structural and acoustic constraints.
- Optimization of fan-outlet guide vane, strut and bifurcation interaction with regards to minimum pressure losses and maximized aerodynamic stability of the fan and outlet guide vanes.
The current paper will present the design approach, the optimization processes and the results of the optimization of the turbofan bypass duct system for the application on modern high-bypass ratio aero engines.
Boundary Layer Transition on a Low Pressure Turbine Blade due to Downstream Potential Interaction
Authors: Penin Véronique , Kulisa Pascale , Bario François 
Published: July 2013

Engine manufacturers wish to reduce the size and weight of their engines, and one way of achieving this is by reducing the rotor-stator gap. It follows that rotor-stator interactions become stronger, especially the influence of the pressure potential, which, despite its rapid spatial decay, becomes significant as the inter-row gap is reduced.
Here we examine the upstream potential effect generated by downstream moving cylindrical rods on an upstream turbine blade. A large scale rectilinear blade cascade was constructed to improve access to the boundary layer. The Reynolds number was 1.6 × 105. Pressure measurements and two-dimensional Laser Doppler Anemometry around the blade were performed to study the boundary layer behavior.
At low turbulence intensity (Tu−in = 1.8%), the laminar boundary layer experiences separation once per rod period. There are two transition modes which alternate during a rod period: separation transition mode and bypass mode.
At high turbulence intensity (Tu−in = 4.0%), no boundary layer separation occurs. The boundary layer follows a bypass transition mode during an entire rod period.
A Physical Model for the Tip Vortex Loss: Experimental Validation and Scaling Method
Authors: Karstadt Sascha , Pelz Peter F. 
Published: July 2013

Losses through secondary flows occur in every turbomachine. Between the rotating blades and the casing of a turbomachine there is a secondary flow through the tip clearance caused by the pressure difference between the pressure and the suction side of the blade. This tip leakage flow is not involved in the work done by the rotating blades hence it reduces the aerodynamic efficiency. The flow through the tip clearance rolls up to a spiral vortex on the suction side of the blade and induces drag. Size and circulation of this vortex, according to the Helmholtz vortex theorem, depend on the bound vortex and the width of the tip clearance. Examinations of this structure lead to an idea of describing the tip vortex loss with analytical methods. Therefore an analytical approach is made regarding mainly the circulation at the blade tips.
The method is discussed critically in the context of known loss models. It is shown to be a good summary of earlier methods. Since no explicit geometry data of the turbomachine is needed, it is much easier to use. The most important aspect is the excellent agreement with measurements performed at the Chair of Fluid Systems Technology. In total eleven different fan configurations are measured and analyzed in regard to their tip clearance losses. The measurements are performed at a test rig located at the laboratory of the Chair of Fluid Systems Technology at Technische Universität Darmstadt. Additionally further published measurement data is used to validate the method.
Impact of Tip Clearance Size and Rotation Speed on the Surge Onset in a High Pressure Centrifugal Compressor
Authors: Buffaz Nicolas , Trébinjac Isabelle 
Published: July 2013

The results presented in the paper aim at investigating the impact of tip clearance size and rotation speed on the surge onset in a transonic single-stage centrifugal compressor composed of a backswept splittered unshrouded impeller and a vaned diffuser. For that purpose, various slow throttle ramps into surge were conducted from 100% to 60% design speed of the compressor and two different tip clearance heights were investigated. The 1MW LMFA-ECL test rig was used to carry out the tests in the compressor stage. Unsteady pressure measurements up to 150 KHz were carried out in the inducer (i.e. the entry zone of the impeller between the main blade leading edge and the splitter blade leading edge) and in the diffuser thanks to nine and fifteen static pressure sensors respectively.
At cruise rotation speed (92.7% of the nominal rotation speed), the surge is triggered by a boundary layer separation on the diffuser vane suction side whatever the tip clearance height may be. No precursor of surge or pre-surge activity has been recorded in the diffuser or in the impeller. The surge reveals a spike-type inception and the tip clearance increase does not change the path into instability.
At lower rotation speeds high frequency disturbances (nearly half the BPF) have been recorded in the inducer before surge. These disturbances can be understood as “tip clearance rotating disturbances” because they are generated at the leading edge of the main blades and move along the tip clearance trajectory. These disturbances reveal a very unstable behavior while the compressor runs into a stable operating point even if the flow at the tip of impeller is dramatically affected by these disturbances. But these disturbances do not trigger the surge which always originates in the diffuser.
The Effect of Mach Number on the Loss Generation of LP Turbines
Authors: Vázquez Raúl , Torre Diego 
Published: July 2013

The effect of Mach number on the loss generation of Low Pressure (LP) Turbines has been investigated experimentally in a pair of turbine high-speed rigs. Both rigs consist of a rotor-stator configuration. All the airfoils are high lift, high aspect ratio and high turning blades that are characteristic of state of the art LP Turbines.
Both rigs are identical with exception of the stator. Two sets of stators have been manufactured and tested. The aerodynamic shape of both stators have been designed in order to achieve the same spanwise distribution of Cp (Compressible Pressure coefficient) over the airfoil surface, each one to its corresponding design Mach number (0.61 and 0.88 respectively).
The aim of this experiment is to obtain the sensitivity of profile and endwall losses to Mach number by means of a back-to-back comparison between both sets of airfoils. Because the two sets of stators maintain the same pressure coefficient distribution, Reynolds number and velocity triangles, each one to its corresponding design Mach number; one can state that the results are only affected by the compressibility.
Experimental results are presented and compared in terms of area average, radial pitchwise average distributions and exit plane contours of total pressure losses.
To complete the paper, the impact of the results on the design of LP Turbines is discussed and presented.
Multi-Objective Compressor Design Optimization Using Multi-Design Transfer Between Codes of Different Fidelity
Authors: Poehlmann Fiete , Bestle Dieter 
Published: July 2013

The use of optimization has become a standard in turbomachinery design. Within several design processes different analysis tools with increasing fidelity are combined to provide more and more reliable information on the compressor behavior, however, at the cost of increased evaluation time. Therefore, optimization processes have to exploit the full potential of low-fidelity codes with low computation time and have to use the optimization results as good starting points for more costly optimization steps based on high-fidelity codes. In case of multi-objective optimization, the intermediate results are not just single solutions but complete Pareto-fronts which have to be transferred from one design code to the next. Performing independent optimizations for each low-fidelity Pareto-solution with a high-fidelity code would be rather time consuming and inefficient. In order to decrease optimization time, a multi-design transfer approach is proposed in this paper. An additional integer parameter is introduced to identify the low-fidelity Pareto-solutions and serves as additional design variable for the high-fidelity optimization sub-process. This results in a mixed-integer optimization problem where genetic algorithms typically have the tendency of premature convergence to local optima due to an early selection of few specific values for the integer parameter. As a solution to this issue an extended problem formulation with respect to the integer parameter is introduced, which allows to firstly improve each of the low-fidelity inputs during high-fidelity optimization and then to converge to the final Pareto-front.
The method is applied to a compressor design problem where a low-fidelity 1D meanline code is combined with a higher-fidelity 2D streamline curvature method. Different strategies of integrating these codes in a common process are discussed and assessed with respect to their solution sets and effort.
Analysis of the Unsteady Flow Field in Turbines by Means of Modal Decomposition
Authors: Lengani Davide , Selic Thorsten , Spataro Rosario , Marn Andreas , Göttlich Emil 
Published: July 2013

This paper presents the results of a modal decomposition method applied to the time resolved data of two different test turbines. The analysis is carried out on the measurements performed by fast response aerodynamic pressure probes as well as on CFD simulations. As shown in the earlier aeroacoustic theory, a plurality of rotating patterns, also called spinning modes, are generated by the rotor-stator interactions. The modes may be computed from the flow quantities, such as total pressure, velocity and flow angles, through Fourier decompositions performed in time and space. The deterministic unsteadiness is then simplified to a limited number of Fourier coefficients. At a fixed radial position, circumferential lobes are identified for any multiple of the blade passing frequency. Therefore, the flow may be described as the superposition of rotating patterns, the spatial characteristics of which are correlated to the linear combinations of blade/vane number.
This analysis has been applied to a one and a half stage low pressure turbine and to a two-stage counter-rotating transonic turbine. In the former test case there is a limited number of modes that characterize the flow field. Hence, the decomposition in modes simplifies considerably the evaluation of the sources of unsteadiness and deterministic stresses. The second test case presents more complex interactions. In fact, the presence of two rotors induces oscillations at frequencies that corresponds to the linear combinations of the two blade passing frequencies. Circumferential modes are identified for the most characteristic frequencies and their physical meaning is discussed.
Numerical Investigation of the Flow in a Radial Vaned Diffuser Without and With Aspiration
Authors: Marsan A. , Trébinjac I. , Coste S. , Leroy G. 
Published: July 2013

An analysis of the flow in a centrifugal compressor vaned diffuser from the nominal operating point of the compressor stage to a point near surge was conducted. Study of performance coefficients, and then of the skin-friction pattern, reveals the growth of a corner stall between the hub wall and the suction side of the vane when moving the operating point towards surge.
Considering the location of the skin-friction pattern singular elements, a boundary layer suction technique has been developed and then numerically tested. The hub wall corner stall was controlled, and performances predicted near surge have been significantly modified, as well as the flow structures when reaching the limit of numerical stable operating range: the major change in the topology of the flow occurs now within the impeller, in the splitter leading edge region, and let think about a leading role of the main blade tip clearance vortex in the instabilities release. The surge massflow seems to have been significantly reduced.
Reverse Flow Turbine-Like Operation of an Axial Flow Compressor
Authors: Gill A. , von Backström T. W. , Harms T. M. 
Published: July 2013

This article describes an experimental investigation of the flow structures occurring in an axial flow compressor during second quadrant operation for reversed rotor rotation in the incompressible flow regime. In second quadrant operation, the flow direction is reversed, but the pressure is lower at the compressor inlet than at the outlet. The compressor thus acts as an axial flow turbine.
A three stage axial flow compressor, with a mass flow rate of 2.7 kg/s and a pressure ratio of 1.022 was investigated. The design rotor tip Mach number is 0.2. Three operational points within the second quadrant were investigated, at flow coefficients of −0.482, −0.553 and −0.843. A five hole conical probe and a 50 micron diameter inclined hot film anemometer were used in this investigation. Radial traverses downstream of rotor rows and a time-dependent area traverse downstream of the first stage stator were performed. Three-dimensional time-dependent numerical Navier-Stokes solutions using the non-linear harmonic approximation for single blade passages in each blade row for each of the cases are compared with experimental work.
The compressor has already been show to be capable of attaining relatively high turbine efficiency (76%) when operating in this mode. Examination of the flow field shows that little to no flow separation occurs on the rotor or stator blades. The wakes of all blades are found to be thin and sharp, and the area between wakes is large and approximately uniform. Numerical results agree relatively well with experimental results.
Hub Extension in an Axial Gas Turbine Diffuser
Authors: Hirschmann A. , Volkmer S. , Casey M. , Montgomery M. 
Published: July 2013

Diffusers downstream of axial gas turbines typically support a hub carrying the turbine rotor bearing. The blunt end of the hub is one of the main causes of loss production. However, a long hub can theoretically reduce the losses and improve the pressure recovery of the diffuser and thus the overall performance of the gas turbine. At the end of the hub the flow experiences a sudden expansion, which can be described by the Borda Carnot equation for incompressible fluid. The pressure loss at the end of the hub depends on the dynamic head and the area ratio of the expansion. With a long hub the blunt end moves further downstream into the diffuser and therefore the flow velocity and the area ratio of the step are reduced, resulting in a lower pressure loss.
Two different casing geometries, each with a short and a long hub configuration are examined experimentally and numerically using CFD (Computational Fluid Dynamics). The geometrical configurations lead to a separation at the casing for a uniform total pressure profile at the inlet which shows the highly loaded state of the diffuser configuration. Experiments have been conducted at a high subsonic inlet Mach number comparable to real turbine exit flows and a low Mach number representing incompressible flow. It is shown that in some cases the long hub has a benefit to pressure recovery. However, the Mach number at the inlet of the diffuser influences the intensity of this effect. A high inlet Mach number increases the losses in the inlet section of the diffuser and at the struts supporting the hub and hence decreases the diffuser performance. Consequently the performance depending on the particular diffuser design can be decreased with an extended hub configuration.
Applications of EARSM Turbulence Models to Internal Flows
Authors: Mehdizadeh O. Z. , Temmerman L. , Tartinville B. , Hirsch Ch. 
Published: July 2013

Turbulence modeling remains an active CFD development front for turbomachinery as well as for general industrial applications. While DNS and even LES still seem out of reach within the typical industrial design cycle due to their high computational cost, RANS-based models remain the workhorse of CFD. Currently, the most widely used models are Linear Eddy-Viscosity Models (LEVM), despite their known limitations for certain flow complexities. Therefore, extending the reliability of eddy-viscosity models to more complex flows without significantly increasing the computational cost can immediately contribute to more reliable CFD results for wider range of applications. This, in turn, can further reduce the need for costly tests and consequently can reduce the product development cost.
A promising approach to achieve this goal is using Explicit Algebraic Reynolds Stress Models (EARSM), obtained through a simplification of the full Differential Reynolds Stress Models (DRSM), and can be perceived as an extension of LEVMs by including the non-linear relation between the turbulence stress tensor, the mean-flow gradient and the turbulence scales. These models are thus less demanding than DRSM, yet capable of capturing more complex turbulence features, compared to LEVM, such as anisotropy in the normal stresses. This may be particularly important in corner flows, for instance, in the hub-blade regions or in diffusers.
This work explores the application of EARSM models to a double diffuser and a high-performance centrifugal compressor stage (HPCC). The results are compared to available experimental data [1,2] showing the importance of including the anisotropy of turbulence in the model, particularly in presence of turbulent corner flows in a diffuser. Furthermore, the EARSM results are also compared to results from the commonly used SST turbulence model. The CFD comparison includes details of the flow structure in the diffuser, where the most noticeable impact from the use of EARSM turbulence models is expected.
The Effect of Tip Leakage Vortex for Operating Range Enhancement of Centrifugal Compressor
Authors: Tomita Isao , Ibaraki Seiichi , Furukawa Masato , Yamada Kazutoyo 
Published: July 2013

Recently, the application of turbochargers is increasing because they are effective in improving fuel consumption of engines. One of the most important turbocharger characteristics is compressor operating range, since it has been used in various driving patterns with the advent of variable geometry turbochargers. Owing to the complicated phenomena such as rotating stall occurring at low flow rate condition, flow analysis is very difficult and details of flow structure have not been fully understood for a long time since the early 1970’s.
In this study, two compressors with different operating range width were investigated with experimental and computational flow analysis. In the compressor with narrow operating range, the amplitude of blade passing pressure fluctuation decreases rapidly and rotating stall occurs near surging. On the other hand, in the compressor with wide operating range, the blockage by the tip leakage vortex breakdown play a role in stabilizing the flow filed and keeping up a high performance even at low flow rates.