bluuebunny/crossref_metadata_2025 huggingface dataset:
Enhancement of Centrifugal Compressor Operating Range by Control of Inlet Recirculation With Inlet Fins
Authors: Tamaki Hideaki , Unno Masaru , Tanaka Ryuuta , Yamaguchi Satoshi , Ishizu Yohei
Published: August 2015

The operating points of a turbocharger compressor tend to approach or cross its surge line while an engine is accelerating, particularly under low-engine speed conditions, hence the need for an acceptable surge margin under low compressor-speed conditions. A method shifting the stability limit on a compressor low-speed line toward a lower flow rate is expected and inlet recirculation is often observed in a centrifugal compressor with a vaneless diffuser near a surge and under a low compressor-speed condition. First, examples of inlet recirculation were introduced in this paper, whereupon the effect of inlet recirculation on compressor characteristic was discussed by 1-D consideration and the potential shown for growth of inlet recirculation to destabilize compressor operations. Accordingly, this study focused on suppressing the effect of inlet recirculation on compressor characteristics using small fins mounted in a compressor-inlet pipe, and examining whether they enhance the compressor operating range under low-speed conditions. Small fins are known as inlet fins in this paper. According to test results, they showed great promise in enhancing the compressor operating range during inlet recirculation. Besides, attempts were also made to investigate the qualitative effect of inlet fins on flow fields using CFD and the disadvantages of inlet fins were also discussed.
The Quantification of Impact of Geometrical and Operational Parameters Uncertainty on the Numerical Simulation Results of Axial Turbine Working Process
Authors: Kolmakova Daria , Popov Grigorii , Baturin Oleg , Krivcov Alexander
Published: August 2015

The quantification of geometric and physical variables uncertainty impact on turbomachinery row workflow was conducted using several untwisted airfoil cascades of axial turbine nozzle blades with uniform section throughout the channel height.

Profile loss coefficient, mass flow parameter, outlet flow angle were accepted as controlled performance criteria.

The series of computational calculations were carried out for these cascades. The first group of calculations was aimed at the identification of the impact of geometric parameters uncertainty on nozzle blades parameters. The second group of calculations was conducted for identifying the studied parameters depending on the flow parameters changes that are used as boundary conditions in the simulation.

The obtained results showed that initial data uncertainty in CFD calculations has a significant impact on the obtained quantitative estimates. The difference between calculated data modified in accordance with the geometry technological tolerances and workflow parameters measurement error may exceed 5% by value of the considered criteria.
Advanced Numerical Setup for Separation-Induced Transition on High-Lift Low-Pressure Turbine Flows Using the γ-Rθt~ Model
Authors: Minot Alexandre , de Saint Victor Xavier , Marty Julien , Perraud Jean
Published: August 2015

This paper presents RANS computation results on the highly-loaded low-pressure turbine blade T106C using the γ-Rθt~ transition model. First, a setup methodology is described, addressing key points such as inlet values for the turbulence variables and stagnation point anomaly. This setup yields good results overall using the transition model in its standard form. The model is then extended to other turbulence model and other correlation functions and results are discussed. It is found that the best overall compromise between precision and trends is obtained through the use of Langtry et al.’s correlation for the transition Reynolds number and Content et al.’s pair of correlations for transition length and critical Reynolds number. However, this paper is more about giving the reader details on the behavior of the γ-Rθt~ model than finding the best combination of parameters.
Active Control of the Corner Separation on a Highly Loaded Compressor Cascade With Periodic Non-Steady Boundary Conditions by Means of Fluidic Actuators
Authors: Staats Marcel , Nitsche Wolfgang
Published: August 2015

This contribution discusses the impact of a non-steady outflow condition on the compressor stator flow that is forced through a mimic in the wake of a linear low speed cascade to simulate the conditions that would be expected in a pulsed detonation engine. 2D/3C-PIV measurements were made to describe the flow field in the passage. Detailed wake measurements provide information about static pressure rise as well as total pressure loss. The stator profile used for the investigations is highly loaded and operates with three-dimensional flow separations under design conditions and without active flow control. It is shown that side wall actuation helps to stabilize the flow field at every phase angle and extends the operating range of the compressor stator. Furthermore, the static pressure gain can be increased by 6% with a 4% loss reduction in time averaged data.
Scaling 3D Low-Pressure Turbine Blades for Low-Speed Testing
Authors: Giovannini Matteo , Marconcini Michele , Rubechini Filippo , Arnone Andrea , Bertini Francesco
Published: August 2015

The present activity was carried out in the framework of the Clean Sky European research project ITURB (“Optimal High-Lift Turbine Blade Aero-Mechanical Design”), aimed at designing and validating a turbine blade for a geared open rotor engine. A cold-flow, large-scale, low-speed (LS) rig was built in order to investigate and validate new design criteria, providing reliable and detailed results while containing costs. This paper presents the design of a LS stage, and describes a general procedure that allows to scale 3D blades for low-speed testing. The design of the stator row was aimed at matching the test-rig inlet conditions and at providing the proper inlet flow field to the blade row. The rotor row was redesigned in order to match the performance of the high-speed one, compensating for both the compressibility effects and different turbine flow paths. The proposed scaling procedure is based on the matching of the 3D blade loading distribution between the real engine environment and the LS facility one, which leads to a comparable behavior of the boundary layer and hence to comparable profile losses. To this end, the datum blade is parameterized, and a neural-network-based methodology is exploited to guide an optimization process based on 3D RANS computations. The LS stage performance were investigated over a range of Reynolds numbers characteristic of modern low-pressure turbines by using a multi-equation, transition-sensitive, turbulence model.
A Three-Regime Based Method for Correlating Film Cooling Effectiveness for Cylindrical and Shaped Holes
Authors: Wang Lieke , Kinell Mats , Najafabadi Hossein N. , Karlsson Matts
Published: August 2015

To cope with high temperature of the gas from combustor, cooling is often used in the hot gas components in gas turbines. Film cooling is one of the effective methods used in this application. Both cylindrical and fan-shaped holes are used in film cooling. There have been a number of correlations published for both cylindrical and fan-shaped holes regarding film cooling effectiveness. Unfortunately there are no definitive correlations for either cylindrical or fan-shaped holes. This is due to the nature of the complexity of film cooling where many factors influence its performance, e.g., blowing ratio, density ratio, surface angle, downstream distance, expansion angle, hole length, turbulence level, etc.

A test rig using infrared camera was built to test the film cooling performance for a scaled geometry from a real nozzle guide vane. Both cylindrical and fan-shaped holes were tested. To correlate the experimental data, a three-regime based method was developed for predicting the film cooling effectiveness. Based on the blowing ratio, the proposed method divides the film cooling performance in three regimes: fully attached (or no jet lift-off), fully jet lift-off, and the transition regime in between. Two separate correlations are developed for fully attached and full jet lift-off regimes, respectively. The method of interpolation from these two regimes is used to predict the film cooling effectiveness for the transition regime, based on the blowing ratio. It has been found this method can give a good correlation to match the experimental data, for both cylindrical and fan-shaped holes. A comparison with literature was also carried out, and it showed a good agreement.
Experimental Study of Sister Hole Film Cooling Performance in a Rotating Flat Plate Model
Authors: Wu Hong , Cheng Huichuan , Li Yulong , Ding Shuiting
Published: August 2015

Film cooling performance of a sister hole was investigated in a flat plate model by applying Thermochromic Liquid Crystal (TLC) technique under the stationary and rotating conditions. The flat plate model is installed in the test section. The sister hole include one main hole and two additional side holes with the smaller diameter in the spanwise direction. The diameter of the main hole is 4 mm and the injection angle is 30°. The density ratio of coolant to mainstream is 1.05. The Reynolds number (ReD) based on the velocity of mainstream and the diameter of the main hole are 2300, 3400 and 4500. Four rotational speeds of 200, 400, 600 and 800 rpm are conducted on both pressure side (trailing wall) and suction side (leading wall) with the blowing ratio varying from 0.14 to 3.5. The effects of blowing ratio, Reynolds number (ReD) and rotation number are mainly analyzed according to film coverage and film cooling effectiveness. The results show that the film performance firstly increases then decreases with the rising of blowing ratio, the optimal blowing ratio is about M=0.5. The film cooling performance is improved with higher Reynolds number (ReD). Under the rotation condition, the film trajectory has an obvious centrifugal deflection which can be enhanced by higher rotation number on the pressure side, and the film deflection moves a little centripetally on the suction side. The film cooling effectiveness on the suction side increases with the rising of rotation number and it is higher than that on the pressure side.
Autoignition of a Jet-A Fuel Spray in a High Temperature Vitiated Air Flow
Authors: Williams Aimee , Shcherbik Dmitriy , Bibik Oleksandr , Lubarsky Eugene , Zinn Ben T.
Published: August 2015

This study investigates the autoignition of a Jet-A fuel spray in a preheated air flow in a 75mm inner diameter quartz tube at atmospheric pressure. Fuel was injected via a pressure-swirl atomizing injector enclosed in an aerodynamically-shaped outer body (7mm diameter) that is installed coaxially with the flow using 3 water-cooled pylons. The air temperature and oxygen content were in the range of 1000–1400 K and 9–12%, respectively by controlling the equivalence ratio in the primary zone of the pre-burner/vitiator and by adding dilution air downstream. The co-flow air velocity ranged from 25–30 m/s for this study (Re = 10,000–12,000), with a trapezoidal profile. Fuel spray was characterized using PDPA. Shape of the spray was transitional between hollow and solid cone with the corner angle of 35° to the axis near injector which reduced to 22° downstream. Spray density varied significantly over cross-section of the tube with the minimum on the axes. Droplets produced have average diameters (D10) of 15–70μm on the axes and periphery, respectively, at 6 cm downstream from the injector. Character of the droplet size distribution was polydisperse. Auto-ignition time delays were captured using a time-averaged camera with CH* (432nm) filter. The measured values agree with delay times previously reported in literature. Two synchronized high-speed cameras with 432nm and 307nm filters were used to investigate dynamics of auto-ignition kernel initiation and convection by capturing of CH* and OH* chemiluminescence at 5000 f ps. This methodology allowed qualitative characterization of the equivalence ratio of kernels in process of their convection and growth. It was shown that kernels are always initiated on the axes of the spray where the average droplet size is minimum. Kernels were formed leaner and become richer as they grow down-stream as indicated by the increase of CH*/OH*intensity ratio. Additionally, kernel behavior depends greatly on air temperature with kernels transitioning from randomly appearing (i.e. single kernel), to periodic, to a constantly auto-igniting flame with the spatial scatter of ignition kernels decreasing with temperature.
Aerodynamic Excitation and Forced Response of Centrifugal Compressor Impeller in Inlet Distortion
Authors: Liu Yixiong , Lao Dazhong , Yang Ce , Wang Leilei , Li Du
Published: August 2015

The forced response of the turbomachinery blade originates from unsteady fluid structure interactions due to the fluctuating aerodynamic load. As one of the primary unsteady issues, inlet distortion flow breaks the uniformity of airflow through the compressor channel which deteriorates the aerodynamic performance of the compressor and intensifies the pressure fluctuation on the blade surface. The work presented here arms to investigate the forced response of a centrifugal impeller induced by compressor inlet distortion. For this purpose, the unsteady flow computation was carried out to provide the temporal evolution of the distorted flow through the compressor and also to quantify the aerodynamic loads acting on the compressor blade due to inlet distortion flow compared with uniform inlet. Meanwhile, the experimental measurement was performed to obtain the transient pressure fluctuation on the blade surface and to validate the accuracy of numeric calculation. The forced response of the compressor impeller based on unsteady excitation was simulated in a finite element (FE) method to gain insight into vibration characteristic of each blade. Time-resolved blade pressure showed the drastic load fluctuation caused by distorted flow mainly located in the blade leading edge region due to the inlet airflow variation. Towards the impeller forced vibration, each blade shows individual response amplitude due to the phase angle difference among the blades. The effect of inlet distortion on the forced response of impeller increases significantly compared with that of the uniform inlet flow.
Experimental and Numerical Investigation of the Flow in a LP Industrial Steam Turbine With Part-Span Connectors
Authors: Häfele M. , Traxinger C. , Grübel M. , Schatz M. , Vogt D. M. , Drozdowski R.
Published: August 2015

An experimental and numerical study on the flow in a three stage low pressure (LP) industrial steam turbine is presented and analyzed. The investigated LP section features conical friction bolts in the last and a lacing wire in the penultimate rotor blade row. These part-span connectors (PSC) allow safe turbine operation over an extremely wide range and even in blade resonance condition. However, additional losses are generated which affect the performance of the turbine. In order to capture their impact on the flow field, extensive measurements with pneumatic multi-hole probes in an industrial steam turbine test rig have been carried out. State-of-the-art three-dimensional CFD applying a non-equilibrium steam (NES) model is used to examine the aero-thermodynamic effects of the PSC on the wet steam flow. A detailed comparison between measurement data and CFD results is performed for several operating conditions. The investigation shows that the applied CFD model is able to capture the three-dimensional flow field in LP steam turbine blading with PSC and the total pressure reduction due to the PSC with a generally good agreement to measured values and is therefore sufficient for engineering practice.
Full-Scale Turbine Vane End-Wall Film-Cooling Effectiveness Distribution Using PSP Technique
Authors: Shiau Chao-Cheng , Chen Andrew F. , Han Je-Chin , Azad Salam , Lee Ching-Pang
Published: August 2015

Researchers in gas turbine field take great interest in the cooling performance on the first-stage vane because the complex flow characteristics and intensive heat load that comes from the exit of the combustion chamber. A better understanding is needed on how the coolant flow interacts with the mainstream and the resulting cooling effect in the real engine especially for the first-stage vane. An authentic flow channel and condition should be achieved. In this study, three full-scale turbine vanes are used to construct an annular-sector cascade. The film-cooling design is attained through numerous layback fan-shaped and cylindrical holes dispersed on the vane and both end-walls. With the three-dimensional vane geometry and corresponding wind tunnel design, the true flow field can thus be simulated as in the engine. This study targets the film-cooling effectiveness on the inner end-wall (hub) of turbine vane. Tests are performed under the mainstream Reynolds number 3.5 × 105; the related inlet Mach number is 0.09 and the free stream turbulence intensity is 8%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, 4%) and density ratios (DR = 1.0, 1.5) are examined. Pressure-sensitive paint (PSP) technique is utilized to capture the detail contour of film-cooling effectiveness on the inner end-wall and demonstrate the coolant trace. The presented results serve a comparison basis for other sets of vanes with different cooling designs. The results are expected to strengthen the promise of PSP technique on evaluating the film-cooling performance of the engine geometries.
Turbulence Field Measurements at the Exit of a Combustor Simulator Dedicated to Hot Streaks Generation
Authors: Bacci Tommaso , Facchini Bruno , Picchi Alessio , Tarchi Lorenzo , Koupper Charlie , Champion Jean-Louis
Published: August 2015

In order to deepen the knowledge of the interaction between modern lean burn combustors and high pressure turbines, a real scale annular three sector combustor simulator has been assembled at University of Florence, with the goal of investigating and characterizing the generated flow field. To generate hot streaks and simulate lean burn combustors behavior, the rig is equipped with axial swirlers, fed by main air flow that is heated up to 531 K, and liners with effusion cooling holes that are fed by air at ambient temperature. The three sector configuration is used to reproduce the periodicity on the central sector.

Ducts of different lengths have been mounted on the swirlers to reduce the interaction of the mainstream with the coolant. Such configurations have been tested, using different measurement techniques, in order to highlight the differences in the resulting flow fields.

The work presented in this paper shows the experimental campaign carried out to investigate the flow turbulence at combustor exit, in isothermal conditions, by means of hot wire anemometry. The goal has been achieved by investigating each test point twice, using an automatic traverse system equipped, in turn, with two split-fiber probes, that allow to measure the velocity components on two planes orthogonal to each other. A method for the time correlation of the signals obtained by the two different tests has been used.

In order to analyse the turbulence decay towards the vanes location, such measurements have been performed on two different planes: one located in correspondence of the combustor exit and the further one placed downstream, in the virtual location of the vanes leading edges.
The Prediction of the Performance and Starting Capability of H-Darrieus Wind Turbines
Authors: Du Longhuan , Dominy Robert G. , Berson Arganthaël
Published: August 2015

The study presented here uses Computational Fluid Dynamics (CFD) models and a development of the Blade Element Momentum (BEM) approach to evaluate the effect of rotor solidity on the performance of a three-bladed H-Darrieus vertical axis wind turbine (VAWT), particularly during its critical starting phase. BEM based models (for example the Double Multiple Stream Tube model) require low computational effort but are unable to provide information about the turbine dynamic start-up process or about the flow structure around the blades. Conversely CFD methods provide detail of the flow structure but are computationally expensive. In this study, an advanced BEM model, the Start-Up model, has been developed which is able to model the turbine dynamic start-up process under different wind and geometric parameters. The developments presented here are supported by 2-D CFD simulations. In this study, the predicted turbine tip speed ratio (TSR) versus power coefficient (Cp) curves from the enhanced BEM approach are compared directly with the CFD results and show good agreement. The study highlights that turbines with larger solidity, which is a function of the ratio of blade chord to turbine rotor radius, (σ = nc/R), demonstrate better self-starting capability at the expense of lower peak efficiency. It is also observed that for a given solidity, turbines with larger blade chord length and correspondingly larger rotor radius have better performance over the entire working range. A comparison between selected aerofoil sections demonstrated that the DU 06 W200 aerofoil section, designed specifically for VAWT application and having 20% thickness with 0.8% camber, has slightly better performance at low tip speed ratio compared with the widely adopted symmetric NACA0018 and NACA0021 sections which themselves have proven self-starting capability. It is also demonstrated using the BEM model that the predicted turbine performance is highly sensitive to the accuracy of aerofoil performance data. The acquisition and publication of extensive, high quality data over the full-range of incidence that is experienced during VAWT start-up and at appropriate Reynolds numbers is shown to be a high priority for predicting turbine performance accurately.
Unsteady Analysis of Inter-Rows Stator-Rotor Spacing Effects on a Transonic, Low-Aspect Ratio Turbine
Authors: Tang Etienne , Leroy Gilles , Philit Mickaël , Demolis Jacques
Published: August 2015

The aerodynamic performances of an axial turbine are affected by the distance between the stator and the rotor. Previous studies have shown different trends, depending mainly on whether the turbine is subsonic or not. The present paper aims at improving the understanding of the effect of rows spacing on the flow through a transonic turbine.

A one-stage, low aspect ratio, high pressure turbine case is investigated using CFD. Steady and unsteady phase-lagged RANS computations are performed on this configuration with different inter-blade rows distances. The results are successfully compared with experimental data from a cold air turbine rig. Entropy production balances are used to emphasize the main loss areas and the loss variations caused by changes in inter-blade rows distance. Two techniques are compared for computing these balances, and one of them appears to perform much better. The flow features causing these losses are then identified. Finally, an optimal inter-rows spacing is found. It is a compromise between the losses created by strong stator-rotor interactions at small inter-rows gaps and the losses generated at the endwalls in the inter-rows space at large distances.
Sea Level Performance of a CF-700 Engine Core With Alternative Fuels
Authors: Davison Craig R. , Canteenwalla Pervez , Chalmers Jennifer L. Y. , Chishty Wajid A.
Published: August 2015

The use of alternative fuels has the potential to enhance energy independence and reduce environmental impact of air travel. It is important to characterize gas turbine operation using such fuels under controlled conditions before implementing them in flight. The performance of a CF-700 engine core was examined in a sea level test facility. The following fuels were tested and will be reported on:

1. Jet A-1 – baseline fuel

2. 100% unblended Hydroprocessed Esters and Fatty Acids (HEFA) synthetic kerosene fuel with aromatics (SKA)

3. 100% unblended Fischer-Tropsch (FT) synthetic paraffinic kerosene (SPK)

4. Blended 50% HEFA-SPK and 50% Jet A-1

Fuel 2 above is an alternative fuel that can potentially be used without blending with conventional fuel. One purpose of the static engine testing was to determine if this fuel was suitable for use on subsequent test flights in a Dassault Falcon 20 aircraft. Engine performance testing was conducted at various power settings for each fuel. Relevant plots of performance are presented, compared and discussed. Transient tests were also performed including slams and chops. Observations of the effects of the different fuels on the engine fuel system are presented as some alternative fuels have the potential to cause seals to shrink and leaks to occur. The leaks observed are noted as are the steps taken to mitigate the problem.
Effects of a Vortex Generator Pair on Jet Impingement Heat Transfer in Cross-Flow
Authors: Wang Chenglong , Wang Lei , Sundén Bengt , Revstedt Johan
Published: August 2015

Jet impingement cooling is commonly used in gas turbines. Usually the spent air from the upstream jets forms a cross-flow past the downstream jets, which degrades their heat transfer performance. In the present study, a new method was proposed to promote the jet penetration and enhance the impingement heat transfer. By placing a delta-winglet vortex generator pair (VGP) in the cross-flow upstream of the jet nozzle, it is found that the impingement heat transfer on the target wall is significantly enhanced. The stagnation region shifts upstream and expands compared to the original case. The stagnation and area-averaged Nusselt numbers also increased. The effects of the distance between the VGP and the jet nozzle l1 were also investigated. The optimal spacing l1 is suggested to be 4d, giving the best heat transfer performance. This study sheds new light on the enhancement of jet impingement heat transfer in a cross-flow.
Film Cooling Effectiveness Improvements Using a Non-Diffusing Oval Hole
Authors: Issakhanian Emin , Elkins Christopher J. , Eaton John K.
Published: August 2015

The need for improvements in film cooling effectiveness over traditional cylindrical film cooling holes has led to varied shaped hole and sister hole designs of increasing complexity. This paper presents a simpler shaped-hole design which shows improved film cooling effectiveness over both cylindrical holes and diffusing fan-shaped holes without the geometric complexity of the latter. Magnetic resonance imaging measurement techniques are used to reveal the coupled 3D velocity and coolant mixing from film cooling holes which are of a constant oval cross-section as opposed to round. The oval shaped hole yielded an area-averaged adiabatic effectiveness twice that of the diffusing fan-shaped hole tested. Three component mean velocity measurements within the channel and cooling hole showed the flow features and vorticity fields which explain the improved performance of the oval shaped hole. As compared to the round hole, the oval hole leads to a more complex vorticity field which reduces the strength of the main counter-rotating vortex pair. The counter-rotating vortex pair acts to lift the coolant away from the turbine blade surface and thus strongly reduces the film cooling effectiveness. The weaker vortices allow coolant to stay closer to the blade surface and to remain relatively unmixed with the main flow over a longer distance. Thus, the oval-shaped film cooling hole provides a simpler solution for improving film cooling effectiveness beyond circular hole and diffusing hole designs.
Multi-Objective CFD Optimisation of Shaped Hole Film Cooling With Mesh Morphing
Authors: Proietti D. , Pranzitelli A. , Andrews G. E. , Biancolini M. E. , Ingham D. B. , Pourkashanian M.
Published: August 2015

A Computational Fluid Dynamics (CFD) optimisation of a single row of film cooling holes was performed. The aim was to achieve the highest adiabatic cooling effectiveness while minimising the coolant mass flow rate. The geometry investigated by Gritsch et al. [1] was the baseline model. It consisted of a row of cylindrical, 30° inclined holes, with a mainstream inlet Mach number of 0.6, a blowing ratio of 1 and a plenum for the upstream cooling air flow. The predictions agreed with the experimental data with a maximum deviation of 6%. The geometry was then optimised by varying three shape parameters: the injection angle, the lateral hole expansion angle and the downstream compound hole angle. A goal driven optimisation approach was based on a design of experiments table. The minimisation of the coolant mass flow together with the maximisation of the minimum and average cooling effectiveness were the optimisation objectives. The shape modifications were performed directly in the ANSYS Fluent CFD solver by using the software RBF Morph in the commercial software platform ANSYS Workbench. There was no need to generate a new geometry and a new computational mesh for each configuration investigated. The dependency of the average effectiveness along the plane centreline on the three geometrical parameters was investigated based on the metamodel generated from the design of experiments results. The goal driven optimisation led to the optimal combination of the three shape parameters to minimise the coolant flow without reducing the cooling effectiveness. The best results were obtained for a geometry with 20° hole angle and 7.5° compound angle injection, leading to a reduction of 15% in the coolant mass flow rate for an enhanced adiabatic cooling effectiveness. The results also showed the preponderance of the centreline angle over the other two parameters.
Aeroelastic Analysis of Supersonic Intake Under Downstream Perturbation
Authors: Yao Cheng , Liu Zhansheng , Yan Jiajia , Zhang Guanghui
Published: August 2015

Smart flexible structure is used to reduce moving equipments of modern supersonic intake with variable geometry design. As a result, aero-elasticity of flexible intake need to be considered in aircraft design. A fully coupled fluid structure interaction (FSI) computation framework was developed in present work to simulate aero-elasticity of a two-dimensional supersonic intake model with flexible cowl wall under downstream back pressure perturbation. Transonic viscous flow in the intake was solved using finite volume method. Finite element method was used to model dynamic response of flexible wall through, consideration of geometric non-linearity. Unsteady flow in the intake with rigid wall under back pressure perturbation shows a large amplitude oscillation of shock train, which includes a period of shock train appearing and disappearing. It is found in FSI results that flexible cowl wall vibrates with downstream perturbation frequency, although multiple frequency components excitation loads are found in spectrogram. Wall flexibility plays the role of ‘buffer zone’, which limits the formation of large amplitude shock train oscillation during FSI process. Structural damping could be used to enlarge ‘buffer zone’ effect.
Turbine Platform Cooling and Blade Suction Surface Phantom Cooling From Simulated Swirl Purge Flow
Authors: Li Shiou-Jiuan , Lee Jiyeon , Han Je-Chin , Zhang Luzeng , Moon Hee-Koo
Published: August 2015

The paper presents the swirl purge flow on platform and a modeled land-based turbine rotor blade suction surface. Pressure sensitive paint (PSP) mass transfer technique provides detailed film cooling effectiveness distribution on platform and phantom cooling effectiveness on blade suction surface. Experiments have completed in a low speed wind tunnel facility with a five blade linear cascade. The inlet Reynolds number based on the chord length is 250,000. Swirl purge flow is simulated by coolant injection through fifty inclined cylindrical holes ahead of the blade leading edge row. Coolant injections from cylindrical holes go through nozzle endwall and a dolphin nose axisymmetric contour before reach platform and blade suction surface. Different “coolant injection angles” and “coolant injection velocity to cascade inlet velocity” results in various swirl ratios to simulate real engine conditions. Simulated swirl purge flow uses coolant injection angles of 30, 45, and 60 degrees to produce swirl ratios of 0.4, 0.6, and 0.8, respectively. Traditional purge flow has coolant injection angle of 90 degree to generate swirl ratio of 1. Coolant to mainstream mass flow rate ratio (MFR) is 0.5%, 1.0% and 1.5% for all swirl ratios. Coolant to mainstream density ratio maintains at 1.5 to match engine conditions. Most of the swirl purge and purge coolant approaches platform, but small amount of the coolant migrates to blade suction surface. Swirl ratio of 0.4 has highest relative motion between rotor and coolant and severely decreases film cooling and phantom cooling effectiveness. Higher MFR of 1% and 1.5% cases suffer from apparent decrement of the effectiveness while increasing relative motion.
A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part I — Introduction and Aerodynamic Measurements
Authors: Puetz Franz , Kneer Johannes , Schulz Achmed , Bauer Hans-Joerg
Published: August 2015

An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a flattened temperature profile at the outlet. As a consequence, the heat load on the endwall of the first nozzle guide vane is increased. The secondary flow field dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and film cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and film cooling on both a flat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and influence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall flow field. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall flow field. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of flow field in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.
Operational Strategies of Wet Cycle Micro Gas Turbines and Their Economic Evaluation
Authors: Stathopoulos Panagiotis , Paschereit Christian Oliver
Published: August 2015

The simultaneous expansion of variable renewables and combined heat and power (CHP) plants in Europe has given rise to a discussion about their compatibility. Due to the concurrence of high wind power generation and high heating loads, it has been argued that only the flexible, electricity oriented operation of CHP plants could go along with the extended penetration of renewables in the European energy system.

The current work focuses on the wet cycle simulation of a Turbec T-100. Three operational strategies are applied on the heat and electricity demand data of a public building, to assess the economic and environmental performance of the wet cycle. The operation of the mGT is modeled in ASPEN Plus and the model is validated with data found in literature. The economic aspects of the operational strategies are assessed with a financial model, which takes into account the current CHP policy incentives and price levels. Furthermore, the advantages and drawbacks of wet operation are highlighted by its comparison to the typical heat-driven operation of dry-cycle mGTs, with a reference to the same case study. It is shown that the wet cycle turbines have a higher number of full load equivalent operating hours and can achieve higher investment payback, with minor drawbacks to their overall environmental performance.
Experimental Analysis of the Pressure Field Inside a Vaneless Diffuser From Rotating Stall Inception to Surge
Authors: Bianchini Alessandro , Biliotti Davide , Rubino Dante Tommaso , Ferrari Lorenzo , Ferrara Giovanni
Published: August 2015

An accurate estimation of rotating stall is one of the key points for high-pressure centrifugal compressors, as it is often connected with the onset of detrimental subsynchronous vibrations which can prevent the machine from operating beyond this limit.

With particular reference to vaneless diffuser rotating stall, the most common practice in industrial machines is to make use of a limited number of dynamic pressure probes to reconstruct the stall characteristics after an ensemble averaging approach.

In this study, a 1:1 model of an industrial compressor stage was tested in a dedicated test rig and equipped with 24 pressure probes properly distributed along the diffuser circumference with the scope of providing a real-time visualization of the spatial pressure distribution within the diffuser. The results allowed the assessment of some important characteristics of the stall cells that were historically supposed based on averaged data, e.g. the cells rigidity. Moreover, the present study confirmed the existence of a stall pattern with two almost axisymmetric lobes. Finally, the transient analysis of both the stall inception and the surge onset was carried out, highlighting the flow field evolution in the diffuser under these conditions.
An Experimental and Numerical Investigation on the Effects of Aerothermal Mixing in a Confined Oblique Jet Impingement Configuration
Authors: Schulz Sebastian , Schindler Alexander , von Wolfersdorf Jens
Published: August 2015

An investigation to characterize the effect of entrainment in a confined jet impingement arrangement is presented. The investigated configuration models an impingement cooled turbine blade passage and holds two staggered rows of inclined impingement jets. In order to distinctly promote thermal entrainment phenomena the jets were heated separately. A steady-state liquid crystal technique was used to obtain near-wall fluid temperature distributions for the impingement surfaces at adiabatic conditions. Additionally, flow field measurements were undertaken using Particle Image Velocimetry (PIV). Furthermore, compressible RANS simulations were carried out with ANSYS CFX using Menter’s SST turbulence model to accompany the experiments. Distributions of effectiveness, velocity, and turbulent kinetic energy detail the complexity of the aerothermal situation. The study was conducted for a jet Reynolds number range from 10,000 to 45,000.

The experimental and numerical results are generally in good agreement. Nevertheless, the simulations predict flow features in particular regions of the geometry that are not as prominent in the experiments. These affect the effectiveness distributions, locally. The investigations revealed that the effectiveness is independent of the temperature difference between the heated and cold jet as well as the jet Reynolds number.
Velocity and Turbulence Intensity Profiles Downstream of a Long Reach Endwall Double Row of Film Cooling Holes in a Gas Turbine Combustor Representative Environment
Authors: Cresci Irene , Ireland Peter T. , Bacic Marko , Tibbott Ian , Rawlinson Anton
Published: August 2015

The continuous demand from the airlines for reduced jet engine fuel consumption results in increasingly challenging high pressure turbine nozzle guide vane (NGV) working conditions. The capability to reproduce realistic boundary conditions in a rig at the combustor-turbine interaction plane is a key feature when testing NGVs in an engine-representative environment. A large scale linear cascade rig to investigate NGV leading edge cooling systems has been designed with particular attention being paid to creating engine representative conditions at the inlet to the NGVs. The combustor simulator replicates the main features of a rich-burn design including large dilution jets and extensive endwall film cooling. A three-dimensional computational domain including the entire combustor simulator has been created and RANS CFD simulations have been run in order to match Reynolds number and mainstream-to-coolant momentum flux ratio; velocity and turbulence measurements have been acquired at the NGV inlet plane at ambient temperature. In this engine-representative environment the authors focused their attention on the flow field downstream of different endwall film cooling holes configurations: three arrangements of a double row of staggered cylindrical holes (lateral pitch-to-diameter ratio of 2–3–6) and one with intersecting holes (intersecting angle of 90°) are experimentally and numerically analyzed. Velocity, turbulence intensity and integral length scales are predicted and measured for a density ratio of 1 and coolant-to-mainstream momentum flux of 6. A hot wire sensor was mounted on a two-axis traverse mechanism able to move the probe in the spanwise and lateral directions. Three slots allowed to reposition the traverse and take measurements at three downstream locations (stream-wise distance-to-diameter ratio of 4.2–9.2–14.2). The research confirmed the strong influence of the endwall coolant on the flow field at the NGV inlet plane and the hole spacing results a key parameter in managing the film development. Closer-spaced hole configurations can assure an effective film coverage. The integral length scales are strongly connected to the hole diameter and spacing. Intersecting holes can potentially reduce the amount of required coolant at a fixed pressure ratio, but they offer worst film performance than cylindrical holes. RANS simulations proved to be able to get the main trends shown by the measurements.
Enhanced Monitoring Capabilities for Legacy Engine Programs
Authors: Banjac Ratko , Lecerf Nicolas , Stouky Yacine , Benomar Rabie , Louza Rachid
Published: August 2015

Operational availability and reliability have never been that high than on current single aisle aircraft engines, close to 100%. Engine Condition Monitoring (ECM) contributes to these good statistics and aims at making it even better. This engines generation will still be operated for at least twenty years. Meanwhile, new engine generation, with new monitoring capabilities, will enter into service within a couple of years. Regardless of engine generation, Operators’ expectations on operational availability and maintenance cost reduction will continue to grow. This statement leads to propose monitoring capability update for current engine generation. Historically powerplant monitoring was mainly dedicated to engine removal planning and full power demonstration. With time and product knowledge increase, usage of trend monitoring, on a restricted set of parameters, provided early warning detection capability to assess some systems health and identify line maintenance actions to preserve engine time on-wing. Unfortunately, legacy monitoring process does not take advantage of full avionics capabilities so that there is an opportunity to bring new monitoring functions on current engine fleet. Consequently, constraints for such Project are the adaptation of software without adding any sensor and fitting the existing operational process. New introduced functions will enable to monitor oil consumption, start capability, enhanced performances, sensors intermittency and control system. This paper describes, in a systemic approach, the implementation of the operational process, the adaptation of both embedded and ground system, the outcomes of the monitoring functions and finally the outcomes on the data and skill management on the customer support side. As a consequence, new services for new business can emerge in a new customer support paradigm.
Prediction and Measurement of Unsteady Blade Surface Pressures on an Open Rotor
Authors: Sohoni N. G. , Hall C. A. , Brandvik T. , Parry A. B.
Published: August 2015

In this paper we present experimental measurements from high frequency unsteady pressure transducers on the blade surfaces of a counter-rotating open rotor and results obtained with full-annulus URANS simulations. A quantitative metric is introduced which can be used to evaluate the quality of the transmission of the front rotor flow features through the interpolating interface between the rotors. Comparisons are made for the open rotor operating at two angles of attack: 0° and 12°. The former setting is a datum case for ‘uninstalled’ operation; the latter setting represents ‘installed’ operation at an extreme angle of attack (AoA). Operation at 12° AoA causes the front rotor wake to vary around the annulus, and leads to shaft-order side-bands for interaction tones. The computational predictions of the interaction tones and shaft-order side-bands are of the correct frequencies and order of magnitude for both the low and the high AoA cases. An analysis of the propagation of the front rotor wakes and tip vortices shows that the use of a sufficiently fine grid is essential in order to accurately predict the unsteady pressure on the rear rotor surface.
Ion Current Measurements as a Method for the Detection of the Reaction Rate in Combustion With Swirl Stabilized Airblast Injection Systems
Authors: Wollgarten J. Christopher , Zarzalis Nikolaos , Turrini Fabio , Peschiulli Antonio
Published: August 2015

Due to strict emission legislation, the trend in the development of aero-engine gas turbine combustion is heading towards lean burning approaches. Lean combustion reduces the combustion temperatures and therefore also the nitrogen oxides emissions. Unfortunately, lean combustion suffers from instabilities and the operation close to the point of lean blowout increases the risk of imminent blowoff. Active stability control is therefore inevitable. The objective of this work is to evaluate the signal obtained from an ion current measurement technique to enable combustion control for aircraft propulsion applications in the near future.

In the past ion current measurements have been used in several studies as flame turbulence analyzer and to detect the reaction rate. However, investigations in lean burning and swirl stabilized airblast injection combustors for future propulsion concepts are rare. The signal obtained from an ion current detector inside a combustor depends strongly on the measurement position. In this experimental investigation field measurements at atmospheric conditions of the ion concentration in a tubular combustor with a sampling rate of 8 kHz are compared with 4 kHz time resolved temperature and OH* chemiluminescence measurements in order to determine the position of the reaction inside the combustor. Variations were performed of the air to fuel ratio (AFR), the air preheating temperature and the pressure drop across the injection system to clarify the interpretation of the ion current signal. The results indicate a strong dependence of the ion current signal on the AFR and that the technique has distinct advantages compared to OH* chemiluminescence measurements: The measurement equipment is comparable non-expensive and the results reveal that the reaction rate is measured directly and are not interpreted from a 3D image. A transition in flame shape from a compact to a tornado flame can be clearly identified with the applied probe. Furthermore, regions with high temperature fluctuations do not necessarily reveal the reaction zone in a recirculating flow field.