bluuebunny/crossref_metadata_2025 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.
Numerical Investigation of Effects of Different Hub Tip Diameter Ratios on Aerodynamic Performance of Single Shaft Multistage Centrifugal Compression Systems
Authors: Ceyrowsky Thomas , Hildebrandt Andre
Published: July 2013

Regarding industrial centrifugal compressors in single shaft design, different configurations with e.g. varying numbers of stages or diverse circumferential speeds, necessitate different shaft diameters. Thus the application of impellers with different hub/tip ratios (dh/d2) is daily routine in industrial practice. Increasing hub/tip ratio leads to higher radii and therefore higher relative speeds, to a reduction in the impeller’s meridional length and hence more rapid diffusion, and to a sharper bending from axial to radial direction. In this paper the impact of hub/tip ratio on stage performance is investigated for three different centrifugal compressor stages, by steady state CFD-calculations. The hub/tip ratio is varied between 0.325 < dh/d2 < 0.45. The relation between design stage flow coefficient and hub/tip ratio is also analysed, both at design and off-design. Thermodynamic behaviour is assessed by 1D-data and also by the investigation of secondary flow features. The current analysis shows, that hub/tip ratio’s influence on characteristics is strongly dependent on the particular stage’s design flow coefficient and circumferential Mach-Number. Increasing a high flow stage’s hub/tip ratio is shown to decrease peak efficiency as well, as to limit the operating range. On the contrary, in case of a low flow stage, design point efficiency is hardly affected, but the characteristic curve is tilted around design point, by applying a different hub/tip ratio. However severity of hub/tip ratio’s impact on thermodynamic behaviour shows to decrease together with stage design flow coefficient.
Aerodynamic Investigation on a 3D-Vaned Diffuser Applied on a High Flow 3D Impeller
Authors: Strohmeyer H. , Hildebrandt A.
Published: July 2013

This paper discusses the effect of a diffuser ratio reduction from r4/r2 = 1.55 to r4/r2 = 1.35 of a centrifugal compressor stage and an approach to retain the efficiency by applying a vaned diffuser. Initially, the diffuser ratio of a high flow, high pressure stage is decreased. Following, the stage having the smallest diffuser ratio is used to investigate the performance of a vaned diffuser, whose trailing edge is shifted into the U-turn. The discussion shows a possible improvement by the diffuser vane enlargement. The total stage efficiency of the diffuser ratio of r4/r2 = 1.55 can be retained as long as the separation due to high incidence of the vaned diffuser is low.
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.
An Experimental Verification of a New Design for Cantilevered Stators With Large Hub Clearances
Authors: Lange Martin , Vogeler Konrad , Mailach Ronald , Elorza-Gomez Sergio
Published: July 2013

Proportionally large relative radial clearances can be found within the rear stages of multistage axial compressors of gas turbines and aero engines, with significant impact on their efficiency. A new three-dimensional design for cantilevered stators in axial compressors is presented, with the aim of improving the overall efficiency and losses of rear stage vanes with large relative hub clearances. The new vane design comprises an unconventional dihedral, with special consideration to reduce the losses caused by the hub clearance vortex. The design was tested in a 4-stage low speed axial research compressor under rear stage conditions. The results are compared to the nominal design to validate the reduction of hub clearance losses and blockage. For both designs, the hub clearances over the third and fourth stator were varied from 1.5% up to 6.0% of span. Overall performance data and flow field traverses upstream and downstream of stator 3 and rotor 4 will be presented in this article in comparison with 3D CFD results.
Compressor Efficiency Variation With Rotor Tip Gap From Vanishing to Large Clearance
Authors: Sakulkaew S. , Tan C. S. , Donahoo E. , Cornelius C. , Montgomery M.
Published: July 2013

Compressor efficiency variation with rotor tip gap is assessed using numerical simulations on an embedded stage representative of that in a large industrial gas turbine with Reynolds number ∼ 2 to 7×106. The results reveal three distinct behaviors of efficiency variation with tip gap. For relatively small tip gap (less than 0.8% span), the change in efficiency with tip gap is non-monotonic with an optimum tip gap for maximum efficiency. The optimum tip gap is set by two competing flow processes: decreasing tip leakage mixing loss and increasing viscous shear loss at the casing with decreasing tip gap. An optimum tip gap scaling is established and shown to satisfactorily quantify the optimal gap value. For medium tip gap (0.8%–3.4% span), the efficiency decreases approximately on a linear basis with increasing tip clearance. However, for tip gap beyond a threshold value (3.4% span for this rotor), the efficiency becomes less sensitive to tip gap as the blade tip becomes more aft-loaded thus reducing tip flow mixing loss in the rotor passage. The threshold value is set by the competing effects between increasing tip leakage flow and decreasing tip flow induced mixing loss with increasing tip gap. Thus, to desensitize compressor performance variation with blade gap, rotor should be tip aft-loaded and hub fore-loaded while stator should be tip fore-loaded and hub aft-loaded as much as feasible. This reduces the opportunity for clearance flow mixing loss and maximizes the benefits of reversible work from unsteady effects in attenuating the clearance flow through the downstream blade-row. The net effect can be an overall compressor performance enhancement in terms of efficiency, pressure rise capability, robustness to end gap variation and potentially useful operable range broadening.
Performance and Boundary Layer Development of a High Turning Compressor Cascade at Sub- and Supercritical Flow Conditions
Authors: Bode Christoph , Kožulović Dragan , Stark Udo , Hoheisel Heinz
Published: July 2013

Based on current numerical investigations, the present paper reports on new Q2D midspan-calculations and results for the well known high turning (Δβ = 50°) supercritical (Ma1 = 0.85) compressor cascade V2. A Q2D treatment of the problem was chosen in order to avoid the difficult modelling of the porous endwalls in a corresponding 3D approach. All simulations were done with the RANS solver TRACE of the DLR Cologne in combination with modified versions of the Wilcox turbulence model and Langtry/Menter transition model. Existing experimental Q2D midspan-results for the V2 compressor cascade were used to demonstrate the improved ability of the numerical code to determine performance characteristics, blade pressure and Mach number distributions as well as boundary layer parameter and velocity distributions. The loss characteristics show minimum loss regions when plotted against inlet angle or axial velocity density ratio. Within these regions, increasing with decreasing Mach number, the experimental results were adequately predicted. Outside these regions it turned out difficult to reproduce the experimental results due to increasing boundary layer separation. Furthermore, the prediction quality was very good for subsonic conditions (Ma1 = 0.60) and still reasonable for supercritical conditions (Ma1 = 0.85), where shock/boundary layer interaction made the prediction more difficult.
Boundary Layer Control in Turbine Exhaust Diffusers Using Casing Injection and Design Modifications
Authors: Babu Mohan , Shukla Ram Krishan , Maru Arpit , Pradeep A. M. , Roy Bhaskar
Published: July 2013

In this paper, results of CFD analysis of boundary layer flow control in a few representative turbine exhaust diffuser geometries are discussed. The simplest of the above configurations was subjected to rig tests with a few of the parameters simulated to produce realistic turbine exhaust diffuser flow. These rig tests provided matching and validation of the CFD results obtained for the same configuration. The diffuser designs were attempted to create strong diffusion field ahead of the struts, in a controlled manner. On the other hand in some of the configurations, an externally induced casing injection system was introduced to control the diffusing flow field behind the struts. In a separate design attempt, the struts and the rear hub were merged in to a smooth blended body in an attempt to reduce the overall diffuser losses and to maximise diffuser pressure recovery. Both these attempts showed mixed results. The studies were carried out at a Reynolds number of 7.7 × 105 based on the diffuser inlet hydraulic diameter. The performance of the diffusers was assessed in terms of total pressure loss and static pressure coefficient across the diffuser. Significant performance benefits were observed using boundary layer control through casing injection as well as through endwall design modifications. In the best performance configuration, the static pressure at the diffuser exit improved by about 16 % and the total pressure losses reduced by about 30% as compared to the baseline configuration at the same axial location.
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.
Turbomachines Under Periodic Admission: Axiomatic Performance Prediction
Authors: Nakhjiri Mehdi , Pelz Peter F.
Published: July 2013

In calibration tasks under engine conditions, steady-state manufacturer performance maps are applied to periodic turbocharger operation. This procedure is exposed to considerable uncertainties. In this work the axiomatic form of the energy equation as well as Euler turbomachinery equation are used to generate a general form of the respective equations which allow for periodicity. Thus, the concept of apparent speed and apparent efficiency is introduced. The latter can attain values greater than unity.
Study of Convective Flow Effects in Endwall Casing Treatments in Transonic Compressor Rotors
Authors: Hah Chunill , Mueller Martin , Schiffer Heinz-Peter
Published: July 2013

The unsteady convective flow effects in a transonic compressor rotor with a circumferential-groove casing treatment are investigated in this paper. Experimental results show that the circumferential-groove casing treatment increases the compressor stall margin by almost 50% for the current transonic compressor rotor. Steady flow simulation of the current casing treatment, however, yields only a 15% gain in stall margin. The flow field at near-stall operation is highly unsteady due to several self-induced flow phenomena. These include shock oscillation, vortex shedding at the trailing edge, and interaction between the passage shock and the tip clearance vortex. The primary focus of the current investigation is to assess the effects of flow unsteadiness and unsteady flow convection on the circumferential-groove casing treatment. Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) techniques were applied in addition to steady Reynolds-averaged Navier-Stokes (RANS) to simulate the flow field at near-stall operation and to determine changes in stall margin. The current investigation reveals that unsteady flow effects are as important as steady flow effects on the performance of the circumferential grooves casing treatment in extending the stall margin of the current transonic compressor rotor. The primary unsteady flow mechanism is unsteady flow injection from the grooves into the main flow near the casing. Flows moving into and out of the grooves are caused due to local pressure difference near the grooves. As the pressure field becomes transient due to self-induced flow oscillation, flow injection from the grooves also becomes unsteady. The unsteady flow simulation shows that this unsteady flow injection from the grooves is substantial and contributes significantly to extending the compressor stall margin. Unsteady flows into and out of the grooves have as large a role as steady flows in the circumferential grooves. While the circumferential-groove casing treatment seems to be a steady flow device, unsteady flow effects should be included to accurately assess its performance as the flow is transient at near-stall operation.
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.
High Pressure Compressor Stabilization by Controlled Pulsed Injection
Authors: Hiller Sven-Jürgen , Hermann Jakob , Widhopf-Fenk Robert
Published: July 2013

Steady tip injection with re-circulated air as a reliable method for high pressure compressor stabilization has the disadvantage to reduce the efficiency due to the usage of already compressed air. Therefore, the minimization of recirculation air and/or a large extension of the stable operation range is mandatory. The paper describes an alternative approach by using an active compressor stability control system with modulating valves. The system was tested at a high pressure compressor. It recognized evolving stall cells reliable. The achieved available stable throttling range of the compressor (in terms of mass flow) was more than doubled by active controlled injection compared to steady tip injection at 90% aerodynamic speed. The other way round, depending on the tip flow structure and the aerodynamic speed injection mass flow saving up to 50% at a similar surge margin compared to steady injection was achieved.
Characterization of Aerodynamic Performance of Boundary-Layer-Ingesting Inlet Under Crosswind
Authors: Liou Meng-Sing , Lee Byung Joon
Published: July 2013

NASA has been studying future transport concepts, envisioned to be technically realizable in the timeframe of 2020–2030, to meet environmental and performance goals. One concept receiving considerable interest involves a propulsion system embedded into a hybrid wingbody aircraft. While offering significant advantages in fuel savings and noise reduction by this concept, there are several technical challenges that are not encountered in the current fleet and must be overcome so as to deliver target performance and operability. One of these challenges is associated with an inlet system that ingests a significantly thick boundary layer, developing along the wingbody surface, into a serpentine diffuser before the flow meeting fan blades. The flow is subject to considerable total pressure loss and distorted at the fan face, much more significantly than in the inlet system of conventional aircraft. In our previous studies [1, 2], we have shown that through innovative design changes on the airframe surface, it is possible to simultaneously increase total pressure recovery and decrease distortion in the flow, without resorting to conventional penalty-ridden flow control concepts, such as vortex generator or boundary layer bleeding/suction. In the current study, we are interested in understanding the following issues: how the embedded propulsion system performs under a crosswind condition by studying in detail the flow characteristics of two inlets, the baseline and another optimized previously under the cruise condition. With the insight, it is hoped that it can help in the follow-on study by devising effective strategies to minimize flow distortion arising from the integration of an embedded-engine system into an airframe to the level acceptable to the operation of engine fan.
Numerical Studies on Improving Performance of Rotor-67 by Blended Blade and EndWall Technique
Authors: Ji Lucheng , Tian Yong , Li Weiwei , Yi Weilin , Wen Quan
Published: July 2013

The paper describes an advanced blading concept for highly loaded turbomachinery, named as Blended Blade and EndWall (BBEW). It implies the following three separate aspects or any of their combinations, i.e. increasing the dihedral between blade surface and endwall as much as possible, increasing the minimum curvature radius of curved transition surface that connects the blade with the endwall, and decreasing the streamwise gradient of the dihedral as the dihedral reduces streamwise. To the knowledge of the authors this is the first approach reported in the open literature that try to control corner separations and horse-shoe vortex (HSV) systematically under the guidance of the Rules of Dihedral (RD). In this paper, BBEW is applied to NASA rotor 67 which is featured by hub corner separation at the suction side and intensive HSV originating from blade leading edge at the hub. Two successive BBEW modifications are made to promote the performance of rotor 67. First, only a curved transition surface is designed to cover the hub corner at the suction side. Emphasis is paid on the streamwise distribution of the minimum curvature radius. Then, the second curved transition surface is added at the leading edge in near hub part. The nearby streamwise gradient of the dihedral is changed from infinity to a finite. Numerical tests showed that the implementation of BBEW had not only nearly eliminated the hub corner separation, but also weakened the leading edge HSV. Performance improvements are observed throughout its operation range. The work presented makes a contribution to a fully 3-D blading methodology for higher aerodynamic loading, engine efficiency and specific thrust.
Multi-Stage Turbomachinery Blades Optimization Design Using Adjoint Method and Thin Shear-Layer N-S Equations
Authors: Ji Lucheng , Li Weiwei , Tian Yong , Yi Weilin , Chen Jiang
Published: July 2013

Traditionally, 3D aerodynamic shape design with the aid of optimization algorithm in an analysis mode has provided a rational and direct search through design space, but it is usually too time-consuming. Further improvement to reduce design cycle is probably a necessary concern in turbomachinery community. Due to less computational cost, adjoint method has received considerable attention in recent years. This paper focuses on continuous adjoint method, and couples with thin shear-layer N-S equations to formulate an efficient sensitivity analysis model for multi-stage turbomachinery blades in the specified objective function. This model includes adjoint equations/boundary conditions, and the sensitivity of objective function to design variable vector. Integrating a 3D blade perturbation parameterization and the simple steepest decent method, a frame of a gradient-based aerodynamic shape design system is constructed. Numerical implementation to solve flow equations and adjoint equations is very similar, and once they are converged respectively, the sensitivity can be calculated by complex method and mesh perturbation efficiently. Thus, a fast Automatic-CFD-Design tool is developed, including three sub-solvers to solve flow equations, adjoint equations and calculate sensitivity respectively. Flow surface design of a 1-1/2 compressor stage in the specified target pressure distribution is used to validate the present approach. Flow field design of NASA transonic compressor stage 35 aiming to increase efficiency and remain mass flow rate and pressure ratio unchanged is taken.
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.
Optimization of Non-Axisymmetric End Wall Contours for the Rotor of a Low Speed, 1 1/2 Stage Research Turbine With Unshrouded Blades
Authors: Bergh Jonathan , Snedden Glen , Meyer Christiaan
Published: July 2013

Non-axisymmetric end wall contouring has become an established method for the reduction of the losses associated with secondary flow. To date, the majority of designs have been produced by the aeroengine manufacturing community and as a result access to specific design and methodological details is often limited [1]. In contrast, whilst the details of non-embargoed work are more freely available, much of this work has been carried out in simplified environments, with the most common of these being 2-dimensional, linear cascades, and therefore do not include a number of features which are present in the flow field of a real turbine [2]. Recent work by Snedden et al [3] involved the introduction of “generic”, non-axisymmetric end wall contours, originally designed for a linear cascade (the so-called Durham cascade), into the rotor row of a low speed, 1 1/2 stage research turbine. While an increase in rotor performance was noted, a detailed inspection of the flow results suggested that even greater improvements could be obtained through the design of custom end walls for the turbine. This investigation therefore covers the design of custom non-axisymmetric end wall contours for the rotor row of an annular turbine rig with unshrouded blades (the same rig as that used by Snedden), using a modified version of an end wall design routine originally developed for the production of non-axisymmetric end walls for a linear cascade environment.
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.
Performance Improvement of Return Channel in Multistage Centrifugal Compressor Using Multi-Objective Optimization
Authors: Nishida Y. , Kobayashi H. , Nishida H. , Sugimura K.
Published: July 2013

The effect of the design parameters of a return channel on the performance of a multistage centrifugal compressor was numerically investigated, and the shape of the return channel was optimized using a multi-objective optimization method based on a genetic algorithm to improve the performance of the centrifugal compressor. The results of sensitivity analysis using Latin hypercube sampling suggested that the inlet-to-outlet area ratio of the return vane affected the pressure loss in the return channel, and that the inlet-to-outlet radius ratio of the return vane affected the outlet flow angle from the return vane. Moreover, this analysis suggested that the number of return vanes affected both the loss and the flow angle at the outlet. As a result of optimization, the number of return vanes was increased from 14 to 22 and the area ratio was decreased from 0.71 to 0.66. The radius ratio was also decreased from 2.1 to 2.0. Performance tests on a centrifugal compressor with two return channels (the original design and optimized design) were carried out using two-stage test apparatus. The measured flow distribution exhibited a swirl flow in the center region and a reversed swirl flow near the hub and shroud sides. The exit flow of the optimized design was more uniform than that of the original design. For the optimized design, the overall two-stage efficiency and pressure coefficient were increased by 0.7% and 1.5%, respectively. Moreover, the second-stage efficiency and pressure coefficient were respectively increased by 1.0% and 3.2%, It is considered that the increase in the second-stage efficiency was caused by the increased uniformity of the flow, and the rise in the pressure coefficient was caused by a decrease in the residual swirl flow. It was thus concluded from the numerical and experimental results that the optimized return channel improved the performance of the multistage centrifugal compressor.
Aerodynamic Design of High Endwall Angle Turbine Stages: Part 2—Experimental Verification
Authors: Cranstone A. W. , Pullan G. , Curtis E. M. , Bather S.
Published: July 2013

An experimental investigation of a turbine stage featuring very high endwall angles is presented. The initial turbine design did not achieve a satisfactory performance and the difference between the design predictions and the test results was traced to a large separated region on the rear suction-surface. To improve the agreement between CFD and experiment, it was found necessary to modify the turbulence modelling employed. The modified CFD code was then used to redesign the vane, and the changes made are described. When tested, the performance of the redesigned vane was found to have much closer agreement with the predictions than the initial vane. Finally, the flowfield and performance of the redesigned stage are compared to a similar turbine, designed to perform the same duty, which lies in an annulus of moderate endwall angles. A reduction in stage efficiency of at least 2.4% was estimated for the very high endwall angle design.
Endwall Loss and Mixing Analysis of a High Lift Low Pressure Turbine Cascade
Authors: Lyall M. Eric , King Paul I. , Sondergaard Rolf
Published: July 2013

A high lift LPT profile designated L2A is used as a test bed for studying the origin of endwall mixing loss and the role of vortical structures in loss development. It is shown analytically and experimentally that the mixing forces within the endwall wake can be decoupled into either mean flow or turbulent forces, and can be further classified as either reversible or irreversible. Among the irreversible forces, mean flow shear is negligible compared to turbulent shear, suggesting that turbulence dissipation is the dominant cause of loss generation. As a result, the mean flow components of the vortical structures do not generate significant mixing losses. Rather than mixing effects, the mean flow of the vortices cause the suction surface boundary layer to separate inside the passage, thereby generating the large low energy regions typical of endwall flows. Losses are generated as the low energy regions mix out. This vortex separation effect is demonstrated with an experiment using a profile fence and pressure surface modification near the endwall. The findings in this paper suggest that profile modifications near the endwall that suppress suction surface separation may provide loss reductions additive to those that weaken vortical structures, such as endwall contouring.
Investigation of the Influence of Hub Gap Size to the Performance of a Subsonic Compressor Stator
Authors: Shi Ke , Chen Haixin , Fu Song , van Rennings Ruben , Thiele Frank
Published: July 2013

This paper presents a RANS study of the hub clearance effects on the performance of a subsonic compressor stator. The inlet boundary conditions are from the calculation of inlet guide vanes. The k-ω SST turbulence model is adapted to resolve the Reynolds stresses. The present numerical results are compared with the experiment carried out at Technical University Berlin. The circumferentially averaged total pressure has a strong decrease in the lower span region from hub to nearly 50% channel height, while the tangential flow angle reduces from approximately 40% channel height to the hub, linearly. The above phenomenon indicates that the leakage flow in the gap between stator blade and the hub does not turn sufficiently. This leads to a smaller incidence angle of the flow to the stator, thus, the lower span of the stator works in smaller attack angle, 0 to 13 degrees lower than the higher span. Surface flow patterns on the hub and both side of the blade surfaces are compared with the oil flow visualization in the experiment. The compressor stator is shown to operate under large separation and strong back flow conditions. The hub leakage flow is studied together with the endwall flow phenomenon for full gap configuration. Two separation lines are observed on the hub. One is lying in front of the blade leading edge plane indicating the separation due to the leading edge leakage flow which spills out of the passage before the flow enters the passage. The other is caused by the interaction between the strong hub leakage flow and the incoming flow. This separation line undergoes an abrupt turning just after the flow leaving the stator passage. The effect of the hub gap size on the leakage flow and the whole flow passage in the stator, including the strength and location of the vortex structure, the location and size of the separation bubble, as well as the back flow behavior, is analyzed. With the help of a novel vortex identification method, the flow field of this subsonic compressor stator and the inlet guide vanes can be visualized illustrating the behavior at the operation point when rotating instability occurs. The parameter η4 can help identifying the stretching and relaxation of the vortex. This approach reveals significant flow details [1]. Combined with DPH (Dynamic Pressure Head) contour and streamlines, the detailed vortices structures and topology in a subsonic compressor can also be further elucidated. The study illustrates different vortices structures in the compressor, as well as their behavior in different gap size configurations.
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.
The Effect of Tip Clearance on the Performance of Contra-Rotating Compressor
Authors: Gao Limin , Li Xiaojun , Feng Xudong , Liu Bo
Published: July 2013

Contra-rotating technology has been considered as an important approach to further improve the thrust-weight ratio of aircraft engine because of its structure and aerodynamic superiority. In the present work, the impact of tip clearance on the performance of a CRAC which consist of two counter-rotating rotors is investigated numerically. To detect the stall margin of CRAC exactly, the back pressure dichotomy method is developed, the grid indepence is verified and the performance is measured. A large number computation is carried out to explore the influence of tip clearance on the operating range of contra-rotating compressor. Finally the flow filed near the tip clearance is analyzed to find the relation between the tip clearance and the first stall rotor. The result shows: (1) Efficiency and pressure ratio decrease with the tip clearance size increased, but there is an optimal tip clearance size corresponding to a relative wide operating range. (2) The first stall stage of contra-rotating compressor varies with the tip clearance size increases. For the present CRAC, ROT2 is the first stall stage with the tip clearance size no greater than 0.5mm, while the ROT1 is the first stall stage if the tip clearance size greater than 0.5mm.