bluuebunny/crossref_metadata_2025 huggingface dataset:
The Effect of Chevrons on Crackle: Engine and Scale Model Results
Authors: Martens Steve , Spyropoulos John T. , Nagel Zac
Published: May 2012

GE and the USN continue to work together to find and develop practical techniques to reduce jet noise on tactical aircraft such as the F/A-18 E/F/G. Noise is an important issue for the Navy because of the harsh acoustic environment induced during operations of these aircraft on aircraft carriers and the impact to communities around Naval Air Bases and training sites. The noise generated by these systems is predominantly the noise generated by the exhaust plume due to the low bypass ratio of the engine and very high exhaust jet velocities. The main components of this jet noise are the jet mixing, shock and crackle noise. The present paper reports on progress, following Reference [1] with the F/A-18 E/F/G jet noise reduction program, which is currently focused on the USN near term goal of up to 3 dB reduction in the peak directivity direction. This goal also includes the reduction of the shock and crackle noise components. These goals are currently being pursued with nozzle plume mixing enhancement employing mechanical chevrons. These chevrons can be incorporated in the production version as a redesign of the F414 nozzle seals and do not involve the introduction of additional parts to the nozzle. This paper focuses on the effect of chevrons on the crackle noise component both in full scale on the F404 engine, and in small scale on the F414 engine nozzle in the twin configuration. The paper aims to make the case that this effect, which was first observed during ground engine testing of prototype chevrons, is a beneficial one in reducing/eliminating crackle which continues to be prevalent in high performance tactical aircraft engines today.
Liquid and Gas Flow Behaviour in Highly Rotating Environment
Authors: Chandra Budi , Simmons Kathy , Pickering Stephen , Tittel Marc
Published: May 2012

In a typical aero engine bearing chamber, oil is introduced to lubricate and cool the bearings as well as the bearing chamber wall. The flow of the oil in the bearing chamber is very complex due to the presence of various forces: gravity, windage, capillary, etc. These pose a great challenge for designers, in particular on how to effectively and efficiently scavenge the hot oil out of the bearing chamber. The University of Nottingham Technology Centre in Gas Turbine Transmission Systems is conducting an ongoing experimental program on liquid and gas flow behavior in the highly rotating environment typically found in a bearing chamber. This paper presents results from a study of the thin film on the wall of a generic bearing chamber consisting of an inner rotating shaft and outer stationary cylindrical wall. Thin liquid film behavior in a cylindrical chamber, subject to only gravity or together with the presence of shearing air flow is of interest in many industrial applications and processes. Measurements of film thickness were taken at various axial and angular locations, at various representative inlet flow rates, scavenge ratios and shaft rotational speeds. The experimental data provides insight on the behavior of thin liquid film as it travels down the inner chamber wall subject to both gravity and shearing air flow and provides a data set eminently suitable for CFD validation. The film thickness measurements are discussed in relation to previously published chamber residence volume data [1].
An Acoustic Methodology to Measure Heat Release Rate Fluctuations From Unsteady Laminar Flames
Authors: Li J. , Richecoeur F. , Schuller T. , Huang Y.
Published: May 2012

This paper presents first validations of an alternative method to optical diagnostics to evaluate unsteady heat release rate disturbances. The technique is based on the determination of the propagation time of ultrasonic waves crossing a reacting flow. A train of pulses is synthesized and transmitted to the space by a tweeter. This signal is captured before and after its passage through the flame. The cross-correlation of the two signals yields a narrow compressed pulse response with a main lobe corresponding to the time lag between the incident and transmitted signals. The technique is examined in two generic laminar configurations. For open flames featuring buoyancy effects, it is possible to link the rate of change of fluctuations of the sound travel time dΔt′/dt to heat release rate disturbances Q˙′. For pulsated confined flames, the situation is more complex and a transfer function must be defined between perturbations in the sound travel time Δt′ and heat release rate disturbances Q˙′. Measurements are compared to predictions using different techniques. A general agreement is obtain in term of phasing and amplitudes. This study shows that the proposed technique is very sensitive to small perturbations in the heat release rate and provides the basis for further developments oriented towards more practical configurations.
Elements of Dynamic Characterization of a Bladed Disk by Using the Tip-Timing Method Under Vacuum Conditions
Authors: Kharyton Vsevolod , Gibert Claude , Blanc Laurent , Thouverez Fabrice
Published: May 2012

This article describes the process of identifying the dynamic properties of a blisk by noncontact measurements provided by the tip-timing method. Attention is focused on tip-timing measurements for mode identification in the presence of the Coriolis effect. Firstly, the standing wave response is studied by using tip-timing measurements that provide the amplitude and phase distributions of each blade over a fixed measurement period for identifying the number of nodal diameters of the mode. Secondly, the travelling wave response is identified, followed by the calculation of the frequency response of the blisk, in order to derive the damping ratios. As the mode considered is found in the domain of strong disk-to-blade coupling, it could be subject to the Coriolis effect. Blade tip-timing (BTT) measurements were used to reveal that the mode was split into forward and backward modes with respect to blisk rotation. All the results obtained after processing the tip-timing measurements are compared with data from the strain gauges with which only a few blades were equipped. The correlation of deformations measured by the strain gauges and blade tip tangential displacement obtained through tip-timing measurements is performed by using realistic finite element models of the blisk.
Virtual Gas Turbines: Geometry and Conceptual Description
Authors: di Mare Luca , Kulkarni Davendu Y. , Wang Feng , Romanov Artyom , Ramar Pandia R. , Zachariadis Zacharias I.
Published: May 2012

This paper documents the construction of a Virtual Engine, with particular reference to its geometry and conceptual description.. The phrase Virtual Engine denotes a system which allows simulations of whole gas-turbine engines to be undertaken at any desired level of fidelity or physical modeling. In order to be of any practical use, the system must allow the computations to be setup as automatically as possible and needs to contain provisions for the exchange of boundary data between adjacent computational domains — e.g. solid-gas interfaces. The paper illustrates the application of the system to the representation and analysis of a modern commercial turbofan engine.
The Influence of Turbine Stator Well Coolant Flow Rate and Passage Configuration on Cooling Effectiveness
Authors: Coren D. D. , Atkins N. R. , Long C. A. , Eastwood D. , Childs P. R. N. , Guijarro-Valencia A. , Dixon J. A.
Published: May 2012

Market competitiveness for aero engine power plant dictates that improvements in engine performance and reliability are guaranteed a priori by manufacturers. The requirement to accurately predict the life of engine components makes exacting demands of the internal air system, which must provide effective cooling over the engine duty cycle with the minimum consumption of compressor section air. Tests have been conducted at the University of Sussex using a turbine test facility which comprises a two stage turbine with an individual stage pressure ratio of 1.7:1. Main annulus air is supplied by an adapted Rolls-Royce Dart compressor at up to 440 K and 4.8 kg s−1. Cooling flow rates ranging from 0.71 to 1.46 Cw, ent, a disc entrainment parameter, have been used to allow ingress or egress dominated stator well flow conditions. The mechanical design of the test section allows internal cooling geometry to be rapidly re-configured, allowing the effect of jet momentum and coolant trajectory to be investigated. An important facet to this investigation is the use of CFD to model and analyse the flow structures associated with the cavity conditions tested, as well as to inform the design of cooling path geometry. This paper reports on the effectiveness of stator well coolant flow rate and delivery configurations using experimental data and also CFD analysis to better quantify the effect of stator well flow distribution on component temperatures.
Conceptual Design and Mission Analysis for a Geared Turbofan and an Open Rotor Configuration
Authors: Larsson Linda , Gro¨nstedt Tomas , Kyprianidis Konstantinos G.
Published: May 2012

In this multidisciplinary study a geared open rotor configuration is assessed and compared to an ultra high bypass ratio geared turbofan engine. Both designs assume a 2020 entry into service level of technology. The specific thrust level for minimizing block fuel and the resulting engine emissions for a given mission is sought. The tool used contains models that effectively capture: engine performance, mechanical and aerodynamic design, engine weight, emissions, aircraft design and performance as well as direct operating costs. The choice of specific thrust is a complex optimization problem and several disciplines need to be considered simultaneously. It will be demonstrated, through multidisciplinary analysis, that the open rotor concept can offer a substantial fuel saving potential, compared to ducted fans, for a given set of design considerations and customer requirements.
Impact of a Collector Box on the Pressure Recovery of an Exhaust Diffuser System
Authors: Bernier Bryan C. , Ricklick Mark , Kapat J. S.
Published: May 2012

The effects of an industrial gas turbine’s Exhaust Collector Box (ECB) geometry on static pressure recovery and total pressure loss were investigated in this study. This study aims to further understand how exit boundary conditions affect the performance of a diffuser system. In this investigation, the exhaust diffuser remained constant through each test, with collector box geometries being varied. The same uniform velocity profile was maintained at the diffuser inlet for all geometries considered. The local pressure recovery through the diffuser with 4 axial ports at 4 circumferential locations was reported along with 14 locations in the accompanying ECB. A system performance analysis for each geometry was conducted using the total pressure loss from inlet to exit of the model. Velocity and total pressure profiles obtained with a hotwire anemometer and Kiel probe at the exit of the diffuser and at the exit of the ECB are also presented in this study. Three (3) different ECB geometries are investigated at a Reynolds number of 60,000. Results obtained from these experimental tests are used to validate the accuracy of a 3-dimensional RANS with realizable k-ε turbulence CFD model from the commercial software package Star-CCM+. The study confirms the existence of two strong counter-rotating helical vortices at the exit of the ECB which significantly affect the flow within the diffuser. Evidence of a strong recirculation zone within the ECB was found to force separation within the exhaust diffuser. Extending the length of the ECB proved to decrease the total pressure loss of the system by up to 19% experimentally. Additionally, the realizable k-ε turbulence was able to accurately represent the total pressure loss of the system within 5%. Despite the extremely complex flow field within the ECB, the computational domain reasonably represented the system in both magnitude and trends.
Experimental Investigation on the Effects of a Large Recirculating Area on the Performance of an Effusion Cooled Combustor Liner
Authors: Facchini Bruno , Maiuolo Francesco , Tarchi Lorenzo , Coutandin Daniele
Published: May 2012

An experimental analysis of a realistic engine cooling scheme was performed on a test article replicating a slot injection and an effusion array with a central large dilution hole. Test section consists of a rectangular cross-section duct with a flat plate comprised of 270 effusion holes arranged in 29 staggered rows (D = 1.65mm, Sx/D = 7.6, Sy/D = 6, L/D = 5.5, α = 30deg) and a dilution hole (D = 18.75mm) located at the 14th row. Both effusion and dilution holes are fed by a channel replicating a combustor annulus, that allows to control cold gas side cross-flow parameters, especially in terms of Reynolds number of both annulus and effusion holes. Upstream the first row, a 6mm high slot, ensure the protection of the very first region of the liner. In order to simulate the combustor flowpath, a backward facing step was installed upstream the slot to generate a large recirculating area. Adiabatic effectiveness, heat transfer coefficient and Net Heat Flux Reduction were evaluated and compared with non recirculating experiments. Measurements were performed by means of a steady-state Thermo-chromic Liquid Crystals (TLC) technique with a thin Inconel heating foil for the heat transfer measurements. A data reduction procedure based on a Finite Element approach has been developed to take into account the non uniform heat generation and conduction due to the large amount of holes. Experiments were carried out considering the combined effects of slot, effusion and dilution holes. Three different effusion blowing ratios (BR = 3-5-7) are investigated, keeping constant the slot flow parameters (BR = 1.3). Results highlight that the presence of the step leads to a general reduction of effectiveness while does not have effects on the heat transfer coefficient.
Modeling of Inlet Distortion Using a Combined Turbofan and Nacelle Inlet Model During Crosswind and Low Speed Forward Operation
Authors: Vunnam Kiran , Hoover Robert
Published: May 2012

Inlet distortion on turbofan nacelles during crosswind and low speed forward operation is an area of concern in the gas turbine engine community. Specifically for aft fuselage mounted nacelles, flow into the inlet is affected by the turning of the airflow at the inlet for crosswind operation, and by fuselage interference effects, such as fuselage based vortices, during low-speed forward operation. A common process of modeling airflow through a turbofan inlet in crosswind and low-speed forward operation is to model the flow boundary, at the fan leading edge location, as a static pressure boundary. For this process, the fan face is represented by a two-dimensional static pressure profile as a function of radius. This process assumes the static pressure at the fan face is uniform circumferentially. In conditions where non-axisymmetric flow effects are present at the fan boundary, such as flow separation in the inlet or local fuselage based ground vortices, an axisymmetric pressure boundary at the fan face is inappropriate. The ingestion of a fuselage based ground vortex will also impact the prediction of inlet distortion into a turbofan inlet. An improved methodology is to model the complete fan and fan stator system using a frozen rotor in a rotating reference frame. This allows the three dimensional flow effects of the fan and stator system to be better modeled within the CFD analysis by allowing the physical geometry of the modeled fan to set the flow characteristics circumferentially in the inlet. The CFD analyses were performed using two methods: (1) with airflow through the nacelle driven by a static pressure boundary at the fan face, and (2) with the fan system modeled as a frozen rotor in a rotating reference frame. The CFD results were evaluated using ARP 1419 circumferential and radial distortion descriptors (Reference 1) at the nacelle’s aerodynamic interface plane location. Results from the fan system CFD analyses are compared to typical values from distortion testing and to CFD results using a static pressure profile boundary condition at the fan face. The goal of this evaluation is to combine the aircraft nacelle and the fan rotor and stator in order to model the impact of a fuselage vortex on inlet distortion where flow through the inlet is set by the fan geometry and fan speed. Initial studies on the isolated nacelle have predicted the effects of ground based vortices on the fan flow at various crosswind velocities. In the current study the effects of fuselage and ground based vortices are studied at various crosswind and head wind velocities at ground idle and takeoff operating conditions.
Active Flow Control Concepts on a Highly Loaded Subsonic Compressor Cascade: Re´sume´ of Experimental and Numerical Results
Authors: Gmelin Christoph , Zander Vincent , Hecklau Martin , Thiele Frank , Nitsche Wolfgang , Huppertz Andre´ , Swoboda Marius
Published: May 2012

The paper presents experimental and numerical results for a highly loaded, low speed, linear compressor cascade with active flow control. Three active flow control concepts by means of steady jets, pulsed jets, and zero mass flow jets (synthetic jets) are investigated at two different forcing locations, i.e. at the end walls and the blade suction side. Investigations are performed at the design incidence for jet-to-inlet velocity ratios from approximately 0.7 to 3.0 and two different Reynolds numbers. Detailed flow field data are collected using a five-hole pressure probe, pressure tabs on the blade surfaces, and time-resolved particle image velocimetry. Unsteady Reynolds-Averaged Navier-Stokes simulations are performed for a wide range of flow control parameters. The experimental and numerical results are used to understand the interaction between the jet and the passage flow. Variation of jet amplitude, forcing frequency, and blowing angle of the different control concepts at both locations allows determination of beneficial control parameters and offers a comparison between similar control approaches. The paper combines the advantages of an expensive but accurate experiment and a fast but limited numerical simulation.
Anisotropic Creep Damage and Elastic Damage of Notched Directionally Solidified Materials
Authors: Stewart Calvin M. , Gordon Ali P.
Published: May 2012

Drives to improve gas turbines efficiency have lead to an increase in firing temperatures. This increase in exhaust temperature has a negative impact upon turbine blade life. Both engineers and material scientists have produced methods to improve turbine blade life under these conditions. Cooling holes have become commonplace and use relatively cool gas to create a lower temperature barrier around a turbine blade. These cooling holes creating internal and external surfaces; a common sight of crack initiation. Directionally-solidified (DS) turbine blades have also become commonplace. These turbine blades exhibit a transversely-isotropic grain structure that improves creep strength in a desired direction. To model a component under such conditions, anisotropic constitutive models are required. In this paper, an anisotropic tertiary creep damage constitutive model for transversely-isotropic materials is given. The influence of creep-damage on general linear elasticity (elastic damage) is described by a modified Hooke’s compliance tensor. Finite element simulations of a V-notched tensile specimen are conducted to replicate a crack initiation site. A discussion on stress triaxiality, stress redistribution, and damage distribution due to anisotropy is provided.
Investigations of Secondary Flow Suction in a High Speed Compressor Cascade
Authors: Gmelin Christoph , Thiele Frank , Liesner Karsten , Meyer Robert
Published: May 2012

Numerical and experimental results for a high-speed compressor cascade with secondary flow suction are presented. Steady flow suction of low momentum fluid from the back flow region in the corner between end wall and vane is considered in order to diminish the corner separation. Investigations are performed at the design point with an inlet Mach number of 0.67 and a Reynolds number of 560,000 based on axial chord and inlet velocity. The steady Reynolds-Averaged Navier-Stokes simulations are evaluated against data from the accompanying experiment collected with pitot tubes and Conrad angle probes. Laminar separation bubbles on both suction and pressure surface are observed. Thus, transition from laminar to turbulent flow is respected in the simulations. The uncontrolled base flow case and various suction ratios (ratio of drawn to passage mass flow) are exploited. Additionally, the position of the slot is varied numerically. It is found that relocation of the slot slightly away from the suction surface improves the performance of the flow suction.
Dual Time Stepping Algorithms With the High Order Harmonic Balance Method for Contact Interfaces With Fretting-Wear
Authors: Salles Loi¨c , Blanc Laurent , Thouverez Fabrice , Gouskov Alexander M. , Jean Pierrick
Published: May 2012

Contact interfaces with dry friction are frequently used in turbomachinery. Dry friction damping produced by the sliding surfaces of these interfaces reduces the amplitude of bladed-disk vibration. The relative displacements at these interfaces lead to fretting-wear which reduces the average life expectancy of the structure. Frequency response functions are calculated numerically by using the multi-Harmonic Balance Method (mHBM). The Dynamic Lagrangian Frequency-Time method is used to calculate contact forces in the frequency domain. A new strategy for solving non-linear systems based on dual time stepping is applied. This method is faster than using Newton solvers. It was used successfully for solving Nonlinear CFD equations in the frequency domain. This new approach allows identifying the steady state of worn systems by integrating wear rate equations a on dual time scale. The dual time equations are integrated by an implicit scheme. Of the different orders tested, the first order scheme provided the best results.
Numerical and Experimental Investigations of Combustion Instability Phenomena in Gas Turbine Burners for Heavy Duty and Aero-Engine Applications
Authors: Accornero Daniele , Caruggi Mario , Nilberto Alessandro , Pittaluga Ferruccio
Published: May 2012

In the recent years, as the technical developments in the field of GT related technology are more and more driven by regulations on environmental pollution control, a whole series of different industrial evolution and innovation lines are investigated so to make combustion processes ever “cleaner”. Among those, there is for sure the adoption of lean and ultra lean combustion processes to be pursued by means of air-fuel premixing combustion technologies. Within this scenario, at DIMSET/SCL (Savona Combustion Laboratory, Dept. of Thermal Machines, Energy Systems and Transportation, Univ. of Genoa) since several years research activities are carried out, mainly within the frame of EC-funded Research Programmes (ICLEAC, MUSCLES, TLC, H2-IGCC) and cooperation with industrial companies of the energy sector (Ansaldo Energia S.p.A.) and aero-propulsion (Avio Group) sectors. Research activities can take advantage of a close integration between experimental facilities, such as several reactive and non-reactive dedicated burner test-rigs, instrumented with LDV, PDA and PIV laser-based equipment, as well as of in-house continuously improved reactive Navier-Stokes solvers for combustor analysis (NastComb solver) and design (TPM method). The paper deals with the stability characterisation of the different combustion-processes taking place within several GT power plants, namely, the heavy duty AE64-3A heavy duty gas turbine (Ansaldo Energia), already present on the market, the so-called Liquid and Gas Rapid Pre-Mix burners, LRPM and GRPM, designed at DIMSET/SCL and still prototypical, and the Avio-designed LPP (Lean Premixed Prevaporised) burner, for aero-engine applications. The research has been addressed at in-depth characterising the stability behaviour of the burner’s operation. In particular, those aspects have been investigated deemed of greatest importance in affecting a stable performance profile, such as swirlers’ design, burner’s internal aerodynamics, premixing duct configuration, fuel typology and injection modalities, etc. The paper gives a synoptic view both of the research approaches (experimental, instrumental, numerical analysis and design) jointly pursued by DIMSET/SCL team in investigating the combustion instability, as well as of the obtained results, which help in pointing out those burner design and operational parameters which appear as most critical in affecting instability insurgence and self-sustainment.
Turbulence and Heat Transfer Measurements in an Inclined Large Scale Film Cooling Array: Part II—Temperature and Heat Transfer Measurements
Authors: Thurman Douglas R. , El-Gabry Lamyaa A. , Poinsatte Philip E. , Heidmann James D.
Published: May 2012

The second of a two-part paper, this study focuses on the temperature field and surface heat transfer measurements on a large-scale models of an inclined row of film cooling holes. Detailed surface and flow field measurements were taken and presented in Part I. The model consists of three holes of 1.9-cm diameter that are spaced 3 hole diameters apart and inclined 30° from the surface. Additionally, another model with an anti-vortex adaptation to the film cooling holes is also tested. The coolant stream is metered and cooled to 20°C below the mainstream temperature. A thermocouple is used to obtain the flow temperatures along the jet centerline and at various streamwise locations. Steady state liquid crystal thermography is used to obtain surface heat transfer coefficients. Results are obtained for blowing ratios of up to 2 in order to capture off-design conditions in which the jet is lifted. Film cooling effectiveness values of 0.4 and 0.15 were found along the centerline for blowing ratios of 1 and 2 respectively. In addition, an anti-vortex design was tested and found to have improved film effectiveness. This paper presents the detailed temperature contours showing the extent of mixing between the coolant and freestream and the local heat transfer results.
Thermal Performance of a Bump Type Gas Foil Bearing Floating on a Hollow Shaft for Increasing Rotating Speed and Static Load
Authors: Sim Kyuho , Song Jin Woo , Lee Yong-Bok , Kim Tae Ho
Published: May 2012

Identifying thermal characteristics of gas foil bearings (GFBs) provides an insight for successful implementation into high speed oil-free turbomachinery. The paper presents temperature measurements of a bump type GFB floating on a hollow shaft for various operating conditions. Two angular ball bearings support the hollow shaft at one end (right), and the other end (left) is free. Test GFB has the outer diameter of 100 mm and the axial length of 45 mm, and the hollow shaft has the outer and inner diameters of 60 mm and 40 mm, respectively. An electric motor drives the hollow shaft using a spline coupling connection. A mechanical loading device provides static loads on test GFB upward via a metal wire, and a strain gauge type load cell placed in the middle of the wire indicates the applied loads. During experiments for shaft speeds of 5 krpm, 10 krpm, and 15 krpm and with static loads of 58.86 N (6 kgf), 78.48 N (8 kgf), and 98.1 N (10 kgf), twelve thermocouples measure the outer surface temperatures of test GFB at four angular locations of 45 deg, 135 deg, 215 deg, and 315 deg, with an origin at the top foil free end, and three axial locations of bearing centerline and both side edges at each angle. Two infrared thermometers measure the outer surface temperature of the hollow shaft at free and supported ends close to test GFB. Test results show that GFB temperatures increase as the shaft speed increases and as the static load increases, with higher temperatures in the loading zone (135 deg and 215 deg) than those in the unloading zone (45 deg and 315 deg). In general, the recorded temperatures are highest at 225 deg where a highest hydrodynamic pressure is expected to build up. Measured temperatures at the bearing centerline are higher than those at the side edges, as expected. In addition, large thermal gradients are recorded in the hollow shaft along the axial direction with higher temperatures at the supported end. Angular ball bearings and lip seal supporting the hollow shaft might produce significant heat generation due to mechanical contact as the shaft speed increases. The axial thermal gradient of the shaft is thought to cause higher temperatures at the bearing right edge facing the ball bearing support than those at the left edge. The present test data along with detailed test GFB/shaft geometries and material properties benchmark thermohydrodynamic (THD) model predictions of test GFB with a rotating hollow shaft.
Dependency of Unsteady Time-Linearized Flutter Investigations on the Steady State Flow Field
Authors: Blocher Michael , May Markus , Schoenenborn Harald
Published: May 2012

The influence of the steady state flow solution on the aero-elastic stability behaviour of an annular compressor cascade shall be studied in order to determine sensitivities of the aero-dynamic damping with respect to characteristic flow parameters. In this context two different flow regimes — a subsonic and a transonic case — are subject to the analysis. The pressure distributions, steady as well as unsteady, on the blade surface of the NACA3506 profile are compared to experimental data that has been gained by the Institute of Aeroelasticity of the German Aerospace Center (DLR) during several wind tunnel tests at the annular compressor cascade facility RGP-400 of the Ecole Polytechnique Fe´de´rale de Lausanne (EPFL). Whereas a certain robustness of the unsteady CFD results can be stated for the subsonic flow regime, the transonic regime proves to be very sensitive with respect to the steady state solution.
Open-Loop Control of Combustion Instabilities and the Role of the Flame Response to Two-Frequency Forcing
Authors: C´osic´ Bernhard , Bobusch Bernhard C. , Moeck Jonas P. , Paschereit Christian Oliver
Published: May 2012

Controlling combustion instabilities by means of open-loop forcing at non-resonant frequencies is attractive because neither a dynamic sensor signal nor a signal processor is required. On the other hand, since the mechanism by which this type of control suppresses an unstable thermoacoustic mode is inherently nonlinear, a comprehensive explanation for success (or failure) of open-loop control has not been found. The present work contributes to the understanding of this process in that it interprets open-loop forcing at non-resonant frequencies in terms of the flame’s nonlinear response to a superposition of two approximately sinusoidal input signals. For a saturation-type nonlinearity, the fundamental gain at one frequency may be decreased by increasing the amplitude of a secondary frequency component in the input signal. This effect is first illustrated on the basis of an elementary model problem. In addition, an experimental investigation is conducted at an atmospheric combustor test-rig to corroborate the proposed explanation. Open-loop acoustic and fuel-flow forcing at various frequencies and amplitudes is applied at unstable operating conditions that exhibit high-amplitude limit-cycle oscillations. The effectiveness of specific forcing parameters in suppressing self-excited oscillations is correlated with flame response measurements that include a secondary forcing frequency. The results demonstrate that a reduction in the fundamental harmonic gain at the instability frequency through the additional forcing at a non-resonant frequency is one possible indicator of successful open-loop control. Since this mechanism is independent of the system acoustics, an assessment of favorable forcing parameters, which stabilize thermoacoustic oscillations, may be based solely on an investigation of burner and flame.
Optimization of 6.2:1 Pressure Ratio Centrifugal Compressor Impeller by 3D Inverse Design
Authors: Zangeneh M. , Amarel N. , Daneshkhah K. , Krain H.
Published: May 2012

In this work, the redesign of a centrifugal transonic compressor impeller with splitter blades by means of the three-dimensional inverse design code TURBOdesign-1 is presented. The basic design methodology for impellers with splitter blades is outlined and is applied in a systematic way to improve the aero/mechanical performance of a transonic 6.2:1 pressure ratio centrifugal compressor impeller. The primary design variables are the main and splitter blades loading and their thickness distributions, the splitter to main blade work ratio, as well as the span-wise swirl distribution. The flow in the original and redesigned impellers are then analyzed by means of a commercial CFD code (ANSYS CFX). The predicted flow field for the original impeller is compared with detailed L2F measurements inside and outside the impeller. The validated CFD results are used to compare the flow field in the optimized and original impeller. It is shown that the inverse design method could be effectively used to control the position and strength of the shock waves, eliminate flow separation and hence obtain a more uniform impeller exit flow in order to improve the aerodynamic performance. In addition, some results are presented on the comparison of stress and vibration in both impellers.
Application of RANS and LES to the Prediction of Flows in High Pressure Turbine Components
Authors: Gourdain Nicolas , Gicquel Laurent Y. M. , Fransen Remy , Collado Elena , Arts Tony
Published: May 2012

This paper investigates the capability of numerical simulations to estimate unsteady flows and wall heat fluxes in turbine components with both structured and unstructured flow solvers. Different numerical approaches are assessed, from steady-state methods based on the Reynolds Averaged Navier-Stokes (RANS) equations to more sophisticated methods such as the Large Eddy Simulation (LES) technique. Three test cases are investigated: the vortex shedding induced by a turbine guide vane, the wall heat transfer in another turbine guide vane and a separated flow phenomenon in an internal turbine cooling channel. Steady flow simulations usually fail to predict the mean effects of unsteady flows (such as vortex shedding) and wall heat transfer, mainly because laminar-to turbulent transition and the inlet turbulent intensity are not correctly taken into account. Actually, only the LES (partially) succeeds to accurately estimate unsteady flows and wall heat fluxes in complex configurations. The results presented in this paper indicate that this method considerably improves the level of physical description (including boundary layer transition). However, the LES still requires developments and validations for such complex flows. This study also points out the dependency of results to parameters such as the freestream turbulence intensity. When feasible solutions obtained with both structured and unstructured flow solvers are compared to experimental data.
Effect of Fuel Distribution on Spray Dynamics in a Two-Staged Multi-Injection Burner
Authors: Providakis T. , Zimmer L. , Scouflaire P. , Ducruix S.
Published: May 2012

Burners operating in lean premixed prevaporized (LPP) regimes are considered as good candidates to reduce pollutant emissions from gas turbines. Lean combustion regimes result in lower burnt gas temperatures and therefore a reduction on the NOx emissions, one of the main pollutant species. However, these burners usually show strong flame dynamics, making them prone to various stabilization problems (combustion instabilities, flashback, flame extinction). To face this issue, multi-injection staged combustion can be envisaged. Staging procedures enable fuel distribution control, while multipoint injections can lead to a fast and efficient mixing. A laboratory-scale staged multipoint combustor is developed in the present study, in the framework of LPP combustion, with an injection device close to the industrial one. Using a staging procedure between the primary pilot stage and the secondary multipoint one, droplet and velocity field distributions can be varied in the spray that is formed at the entrance of the combustion chamber. Non-reactive and reactive flows are characterized through an extensive Phase Doppler Anemometry (PDA) campaign. Three staging values, corresponding to three different flame stabilization processes, are analyzed, while power is kept constant. It is shown that mean values and droplet distributions are affected by the staging procedure in the non-reactive as in the reactive situations. Using adequate post-processing, it is also possible to study non-reactive and reactive flow/flame dynamics. Spectral analysis shows that the non-reactive flow is strongly structured by a high frequency rotating structure that can clearly be associated with a precessing vortex core (PVC), while the reactive situation encounters a strong acoustic-flame coupling leading to a low frequency oscillation of both the velocity field and the spray droplet distribution. In this last situation, high frequency phenomena, which may be due to PVC, are still visible.
Prognostics and Health Management of Aircraft Engine EMA Systems
Authors: Watson Matthew J. , Smith Matthew J. , Kloda Jared , Byington Carl S. , Semega Kenneth
Published: May 2012

In this study, the authors conducted a model-based, engine system analysis of Electro-Mechanical Actuators (EMAs). This effort employed an existing, NASA developed, aircraft engine model. A critical engine actuator within the model was replaced by a dynamic, physics based EMA model that includes: controller, motor, drivetrain and feedback sensor sub-models. The actuator model includes simulation of the electrical, mechanical and thermal response of the system. The resulting platform was used to simulate a range of critical actuator fault conditions including: feedback resolver fault, ball-screw degradation, motor winding short, and LVDT non-linearity. Since the available experimental data from propulsion system EMAs is very limited, this platform provides an ideal opportunity to evaluate and enhance prognostic capability for critical engine applications. The model fault tests were used to demonstrate a prototype prognostics and health management (PHM) system for engine EMAs. First, the system response was used to develop an appropriate mode detection algorithm to identify the ideal system conditions for collection of diagnostic evidence. Then, using the acquired transient and steady-state system response, diagnostic data features were derived from EMA related sensors and engine performance parameters. Using these features as a starting point, a system level reasoner was created using multiple classification techniques including LDA, QDA and SVM. Using model generated data with simulated system variance, it was demonstrated that the reasoner provides excellent fault detection, isolation and severity assessment capability for all considered fault modes. Finally, a suitable actuator life model was developed and a probabilistic prognostic approach was used to determine the remaining useful life of the system. The demonstrated PHM system will significantly enhance the ability to safely operate aircraft, schedule maintenance activities, optimize operational life cycles, and reduce support costs.
Boundary Layer Control for a Vertical Axis Wind Turbine Using a Secondary-Flow Path System
Authors: Untaroiu Alexandrina , Raval Archie , Wood Houston G. , Allaire Paul E.
Published: May 2012

Vertical axis wind turbines (VAWTs) have typically lower efficiency compared to their horizontal counterparts (HAWTs), but are attractive for places where taller structures are prohibited, as well as for regions where available wind speeds are lower. For HAWTs, the blades are always perpendicular to the incoming wind, providing a continuous thrust throughout the rotation. Contrary to HAWTs, VAWTs have advancing blades and retreating blades, where blades backtrack against the wind, causing lower efficiency. Hence, any modifications that can be made to improve the efficiency of VAWTs can be beneficial to the wind industry. Passive flow control permits the airfoil geometry to be modified by means of grooves or slots without requiring heavy mechanisms or actuators. Hence, this form of boundary layer control seems advantageous for wind turbines, so that minimal amount of maintenance is required, while complexity of the turbine is not significantly increased. Such modification changes the boundary layer over an airfoil reducing flow separation and reversed flow. This study introduces a new form of passive flow control: Secondary-flow control system, which works on the principle of mass removal, eliminating flow separation at different apparent angles of attack in a VAWT. CFD analysis is used to investigate passive flow control for the airfoils NACA8H12 and LS0417 in a three-bladed VAWT configuration. A secondary flow path is initially designed and optimized in a single airfoil configuration, and then used to adjust the wind turbine blade design. The effects of secondary-flow control system in a VAWT design configuration are investigated by comparison with the non-modified airfoil design. The CFD results indicate that secondary-flow path system can be used to modify and control the boundary layer for a wind turbine. It is believed that secondary-flow control system incorporated in VAWT design has potential for improving turbine efficiency. Further research should be conducted to optimize the secondary-flow path system according to the shape of the airfoil in a 3D VAWT configuration, so that blades interference can be captured.
A Preliminary Numerical CFD Analysis of Transonic Compressor Rotors When Subjected to Inlet Swirl Distortion
Authors: Castaneda Javier , Mehdi Ahad , di Cugno Domenico , Pachidis Vassilios
Published: May 2012

A preliminary investigation of a CFD capability to assess the impact of inlet swirl distortion on transonic compressor rotors has been carried out. In the late 1960s with the advent of turbo fan engines, industry and government agencies became increasingly aware of the inlet total pressure distortion problem. Since then, the inlet/engine compatibility assessment has become a significant issue within the propulsion system life cycle. Nowadays the development of high-performance military aircraft and UAV with maneuvers before unthinkable, entail considerable levels of inlet flow angularity. The importance of developing a rigorous methodology to understand the effect of inlet swirl distortion on turbomachinery has also become one of the major concerns of present day. NASA rotor 67 and 37 were selected for this investigation having different hub to tip radius and aspect ratios. The steady state CFD simulations were carried out on two types of inlet swirl distortion scenarios: Bulk swirl (both Rotor 37 and 67) and Ground vortex (only Rotor 67). A parametric study to define the swirl angle distribution for ground vortex cases was also accomplished. The non-dimensional ground clearance, wind conditions and core vortex location at the inlet/engine AIP were the parameters taken into account. The study carried out suggests that ground vortex core location and vortex rotational direction greatly affect the shift of the speedline. This emphasizes the importance of identifying the radial location of ingested vortex core at the AIP as the turbomachinery response differs depending on it. Similar shift in speedlines for the bulk swirl cases were also observed.
Influence of Variable Geometry Transients on the Gas Turbine Performance
Authors: Barbosa Joa˜o Roberto , Silva Franco Jefferds dos Santos , Tomita Jesuino Takachi , Bringhenti Cleverson
Published: May 2012

During the design of a gas turbine it is required the analysis of all possible operating points in the gas turbine operational envelope, for the sake of verification of whether or not the established performance might be achieved. In order to achieve the design requirements and to improve the engine off-design operation, a number of specific analyses must be carried out. This paper deals with the characterization of a small gas turbine under development with assistance from ITA (Technological Institute of Aeronautics), concerning the compressor variable geometry and its transient operation during accelerations and decelerations. The gas turbine is being prepared for the transient tests with the gas generator, whose results will be used for the final specification of the turboshaft power section. The gas turbine design has been carried out using indigenous software, developed specially to fulfill the requirements of the design of engines, as well as the support for validation of research work. The engine under construction is a small gas turbine in the range of 5 kN thrust / 1.2 MW shaft power, aiming at distributed power generation using combined cycle. The work reported in this paper deals with the variable inlet guide vane (VIGV) transients and the engine transients. A five stage 5:1 pressure ratio axial-flow compressor, delivering 8.1 kg/s air mass flow at design-point, is the basis for the study. The compressor was designed using computer programs developed at ITA for the preliminary design (meanline), for the axisymmetric analysis to calculate the full blade geometry (streamline curvature) and for the final compressor geometry definition (3-D RANS and turbulence models). The programs have been used interatively. After the final channel and blade geometry definition, the compressor map was generated and fed to the gas turbine performance simulation program. The transient study was carried out for a number of blade settings, using different VIGV geometry scheduling, giving indication that simulations needed to study the control strategy can be easily achieved. The results could not be validated yet, but are in agreement with the expected engine response when such configuration is used.