bluuebunny/crossref_metadata_2025 huggingface dataset:
The Design and Model Simulation of a Micro Gas Turbine Combustor Supplied With Methane/Syngas Fuels
Authors: Liu Chi-Rong , Shih Hsin-Yi
Published: September 2016

The design and model simulation of a can combustor has been made for future syngas (mainly H2/CO mixtures) combustion application in a micro gas turbine. In previous modeling studies with methane as the fuel, the analysis indicated the design of the combustor is quite satisfactory for the 60-kW gas turbine; however, the cooling may be the primary concerns as several hot spots were found at the combustor exit. When the combustor is fueled with methane/syngas mixtures, the flames would be pushed to the sides of the combustor with the same fuel injection strategy. In order to sustain the power load, the exit temperature became too high for the turbine blades, which deteriorated the cooling issue of the compact combustor. Therefore, the designs of the fuel injection are modified, and film cooling is employed. Consequently, the simulation of the modified combustor is conducted by the commercial CFD software Fluent. The computational model consists of the three-dimensional, compressible k-ε model for turbulent flows and PPDF (Presumed Probability Density Function) model for combustion process between methane/syngas and air invoking a laminar flamelet assumption. The flamelet is generated by detailed chemical kinetics from GRI 3.0. Thermal and prompt NOx mechanisms are adopted to predict the NO formation.

At the designed operation conditions, the modeling results show that the high temperature flames are stabilized in the center of the primary zone where a recirculation zone is generated for methane combustion. The average exit temperature of the modified can combustor is 1293 K, which is close to the target temperature of 1200 K. Besides, the exit temperatures exhibit a more uniform distribution by coupling film cooling, resulting in a low pattern factor of 0.22. The NO emission is also low with the increased number of the dilution holes. Comparing to the results for the previous combustor, where the chemical equilibrium was assumed for the combustion process, the flame temperatures are predicted lower with laminar flamelet model. The combination of laminar flamelet and detailed chemistry produced more reasonable simulation results. When methane/syngas fuels are applied, the high temperature flames could also be stabilized in the core region of the primary zone by radially injecting the fuel inward instead of outward through the multiple fuel injectors. The cooling issues are also resolved through altering the air holes and the film cooling. The combustion characteristics were then investigated and discussed for future application of methane/syngas fuels in the micro gas turbine. Although further experimental testing is still needed to employ the syngas fuels for the micro gas turbine, the model simulation paves an important step to understand the combustion performance and the satisfactory design of the combustor.
Experimental Investigation of Spray and Combustion Performances of a Fuel-Staged Low Emission Combustor: Part I — Effects of Main Swirl Angle
Authors: Liu Cunxi , Liu Fuqiang , Yang Jinhu , Mu Yong , Hu Chunyan , Xu Gang
Published: September 2016

In order to reduce NOx emissions, modern gas turbines are often equipped with lean-burn combustion systems, where the high-velocity fuel-lean conditions that limit NOx formation in combustors also inhibit the ability of ignition, high altitude relight, and lean combustion stability. To face these issues, internally staged scheme of fuel injection is proposed. Primary and main fuel staging enable fuel distribution control, multi-injections of main fuel leads to a fast and efficient mixing. A fuel-staged low emission combustor in the framework of lean-burn combustion is developed in the present study, the central pilot stage for low power conditions is swirl-cup prefilming atomization, main stage is jet-in-crossflow multi-injection, a combination of primary and main stage injection is provided for higher power output conditions. In lean-burn combustors, the swirling main air naturally tends to entrain the pilot flame and quench it at low power conditions, which is adverse to the operability specifications, such as ignition, lean blow-out, and high-altitude relight. In order to investigate the effects of the main swirl angle on combustion performances, the ignition and lean blow-out performances were evaluated in a single dome rectangular combustor. Furthermore, the spray patterns and flow field are characterized by kerosene-planar laser induced fluorescence and particle image velocimetry to provide insight into spray and combustion performances. Flow-flow interactions between pilot and main air streams, spray-flow interactions between pilot spray and main air streams, and flame-flow interactions between pilot flame and main air streams are comprehensively analyzed. The entrainment of recirculating main air streams on pilot air streams enhances with increase of main swirl angle, because of the upward motion and increasing width of main recirculation zone. A small part of droplets are entrained by the recirculating main air streams at periphery of combustor and a majority of droplets concentrate near the centerline of combustor, making that entrainment of recirculating main air streams on pilot spray and quenching effects of recirculating main air streams on pilot flame is slight, and the quenching effects can be ignored. A follow-on paper will study the effects of venturi angle on the combustion performances.
Design Considerations for Tip Clearance Control and Measurement on a Turbine Rainbow Rotor With Multiple Blade Tip Geometries
Authors: Lavagnoli S. , De Maesschalck C. , Andreoli V.
Published: September 2016

The accurate design, control and monitoring of the running gaps between static and moving components is vital to preserve the mechanical integrity and ensure the correct functioning of any compact rotating machinery. Throughout engine service, the rotor tip clearance undergoes large variations due to installation tolerances or as the result of different thermal expansion rates of the blades, rotor disk and casing during speed transients. Hence, active tip clearance control concepts and engine health monitoring systems rely on precise real-time gap measurements. Moreover, this tip gap information is crucial for engine development programs to verify the mechanical and aerothermal design, and validate numerical predictions.

This paper presents an overview of the critical design requirements for testing engine-representative blade tip flows in a rotating turbine facility. The manuscript specifically focuses on the challenges related with the design, verification and monitoring of the running tip clearance during a turbine experiment.

In the large-scale turbine facility of the von Karman Institute, a rainbow rotor was mounted for simultaneous aerothermal testing of multiple blade tip geometries. The tip shapes are a selection of high-performance squealer-like and contoured blade tip designs. On the rotor disc, the blades are arranged in seven sectors operating at different clearance levels from 0.5 up to 1.5% of the blade span.

Prior to manufacturing, the blade geometry was modified to compensate for the radial deformation of the rotating assembly under centrifugal loads. A numerical procedure was implemented to minimize the residual unbalance of the rotor in rainbow configuration, and to optimize the placement of every single airfoil within each sector. Subsequently, the rotor was balanced in-situ to reduce the vibrations and satisfy the international standards for high balance quality. The single-blade tip clearance in rotation was measured by three fast-response capacitive probes located at three distinct circumferential locations around the rotor annulus. Additionally, the minimum running blade clearance is captured with wear gauges located at five axial positions along the blades chord. The capacitance probes are self-calibrated using a multi-test strategy at several rotational speeds. The in-situ calibration methodology and dedicated data reduction techniques allow the accurate measurement of the distance between the turbine casing and the local blade tip features (rims and cavities) for each rotating airfoil separately. General guidelines are given for the design and calibration of a tip clearance measurement system that meets the required measurement accuracy and resolution in function of the sensor uncertainty, nominal tip clearance levels and tip seal geometry.
Experimental and Numerical Verification of an Optimization of a Fast Rotating High Performance Radial Compressor Impeller
Authors: Elfert M. , Weber A. , Wittrock D. , Peters A. , Voss C. , Nicke E.
Published: September 2016

Outgoing from a well-proven radial compressor design which has been extensively being tested in the past known as SRV4 impeller (Krain impeller), an optimization has been performed using the AutoOpti tool developed at DLR’s Institute of Propulsion Technology. This tool has shown its capability in several tasks, mainly for axial compressor and fan design as well as for turbine design. The optimization package AutoOpti was applied to the redesign and optimization of a radial compressor stage with a vaneless diffusor. The numerical results of this optimization were presented by Voss et al. [1] and by Raitor et al. [2]. The optimization was performed for the SRV4 compressor geometry without fillets using a relatively coarse structured mesh in combination with wall functions. The impeller geometry deduced by the optimization had to be slightly modified due to manufacturing constraints. In order to filter out the improvements of the new so-called SRV5 radial compressor design, two work packages were conducted: The first one was the manufacturing of the new impeller and its installation on a test rig to investigate the complex flow inside the machine. The aim was, first of all, the evaluation of a classical performance map and the efficiency chart achieved by the new compressor design. The efficiencies realized in the performance chart were enhanced by nearly 1.5 %. A 5 % higher maximum mass flow rate was measured in agreement with the RANS simulations during the design process. The second work package comprises the CFD analysis. The numerical investigations were conducted with the exact geometries of both, the baseline SRV4 as well as the optimized SRV5 impeller including the exact fillet geometries. To enhance the prediction accuracy of pressure ratio and impeller efficiency the geometries were discretized by high resolution meshes of approximately 5 million cells. For the blade walls as well as for the hub region the mesh resolution allows a low-Reynolds approach in order to get high quality results. The comparison of the numerical predictions and the experimental results shows a very good agreement and confirms the improvement of the compressor performance using the optimization tool AutoOpti.
Effect of Aromatics in Jet Fuels on Spray Characteristics Downstream of an Aeronautical Pressure Swirl Atomizer
Authors: Leparoux Julien , Lecourt Renaud , Penanhoat Olivier
Published: September 2016

Standard aeronautic fuels have a lower limit of aromatics of 8% (by volume) with about 18% for regular Jet A1. It has been shown that aromatics contained in Jet fuel have an impact on the fine particle emissions. In order to reduce these emissions, alternative fuels with lower aromatic content have been identified as a promising solution. Change Jet fuel composition can have several effects on spray and combustion behaviors, among others: atomization process, droplet evaporation, flame structure, pollutant and particle emissions. Then, it is necessary to evaluate the impact of this change on gas turbine performance and operability.

The present study is focused on the spray behavior investigation with different aromatic content. Four Jet fuels are investigated including conventional Jet A1 kerosene, drop-in fuel with a mixture of half conventional Jet fuel and synthetic paraffinic kerosene (SPK), SPK with 8% of aromatics and pure SPK. The tests are performed at atmospheric conditions on the MERCATO testbed located at ONERA (FR). Phase Doppler Anemometry (PDA) measurements are carried out for the four fuels on an injection system composed of a pressure swirl atomizer and an air swirler.

In this paper, a spray analysis of liquid velocity and droplet diameter measurements is described and linked to the variations of fuel properties. In the range of parameters covered by the four different fuels, it is shown that the spray behavior of each fuel is similar to the conventional Jet A1.
Numerical and Experimental Studies of Reynolds Number and Stator Clocking Effect on Performance of a High-Loaded Two-Stage Compressor With 3.7 Total Pressure Ratio
Authors: Mileshin V. , Druzhinin I. , Savin N. , Kozhemyako P.
Published: September 2016

One of the vital tasks related to the improvement of the efficiency of turbomachines is an increase in their operating altitude by reducing the influence of Reynolds numbers, Re, on the turbomachine parameters. Therefore, the results presented in this work on the effect of Re on parameters of an axial compressor composed of two high-loaded stages are attractive both from scientific and practical points of view.

This work presents the results of experimental investigations of Re effect on gas-dynamic characteristics of a highly-loaded two-stage compressor (HPC-2), simulating the first two stages of the High Pressure Compressor, (HPC), for an advanced engine. The compressor has the following key gasdynamic parameters:

– corrected mass flow rate: 31.8 kg/s

– total pressure ratio: 3.7

– coefficient of aerodynamic load: 0.421

The experimental study of HPC-2 is carried out at the Central Institute of Aviation Motors, (CIAM), C-3 test facility. Reynolds number is changed by decreasing pressure at the inlet from 1 bar to 0.2 bar and changing the clocking position of stator rows at optimal points of performance for two rotational speeds, 0.7 and 0.88.

Tests of HPC-2 compressor show that a stepwise decrease of the inlet pressure from 1.0 0.7, 0.4 down to 0.2 bar (a decrease in Re from 1785000 to 294000) leads to a smooth decrease in maximum compressor efficiency by 1% and a shift of compressor characteristics towards lower air flow rates (by 2%).

The experimental studies of compressors for present-day engines show that unsteady processes play a key role in compressor efficiency and stability.

One way to control unsteady flow in compressors is clocking effect. Recently clocking of stators and rotors has been investigated in details using the HPC-2 two-stage compressor (total pressure ratio 3.7) in H = 0; M = 0 conditions. The effect of Reynolds number on compressor characteristics is studied in this work for the HPC-2 compressor while also investigating stator clocking effect. The rotor clocking effect is not studied. Tests of HPC-2 compressor show that a decrease in the inlet pressure and changes in Re results in a stronger clocking effect, which is 0.5% in terms of efficiency under atmospheric conditions at the inlet and reaches 1% with a decrease in the inlet pressure from atmospheric pressure down to 0.2 bar.
An Acoustic Time-of-Flight Approach for Unsteady Temperature Measurements: Characterization of Entropy Waves in a Model Gas Turbine Combustor
Authors: Wassmer Dominik , Schuermans Bruno , Paschereit Christian Oliver , Moeck Jonas P.
Published: September 2016

Lean premixed combustion promotes the occurrence of thermoacoustic phenomena in gas turbine combustors. One mechanism that contributes to the flame-acoustic interaction is entropy noise. Fluctuations of the equivalence ratio in the mixing section cause the generation of hot spots in the flame. These so called entropy waves are convectively transported to the first stage of the turbine and generate acoustic waves that travel back to the flame; a thermoacoustic loop is closed. However, due to the lack of experimental tools, a detailed investigation of entropy waves in gas turbine combustion systems has not been possible up to now. This work presents an acoustic time-of-flight based temperature measurement method which allows the detection of temperature fluctuations in the relevant frequency range. A narrow acoustic pulse is generated with an electric spark discharge close to the combustor wall. The acoustic response is measured at the same axial location with an array of microphones circumferentially distributed around the combustion chamber. The delay in the pulse arrival times corresponds to the line-integrated inverse speed of sound. For validation of this new method an experimental setup was developed capable of generating well defined entropy waves. As a reference temperature measurement technique a hot-wire anemometer is employed. For the measurement of entropy waves in an atmospheric combustion test rig, fuel is periodically injected into the mixing tube of a premixed combustor. The subsequently generated entropy waves are detected for different forcing frequencies of the fuel injection and for different mean flow velocities in the combustor. The amplitude decay and phase lag of the entropy waves adheres well to a Strouhal number scaling for different mean flow velocities. In addition, simultaneously to the entropy wave measurement, the equivalence ratio fluctuations in the mixing tube are detected using the Tunable Diode Laser Absorption Spectroscopy (TDLAS) technique.
Experimental Analysis of High-Amplitude Temporal Equivalence Ratio Oscillations in the Mixing Section of a Swirl-Stabilized Burner
Authors: Blümner Richard , Paschereit Christian Oliver , Oberleithner Kilian , Ćosić Bernhard
Published: September 2016

Unsteady temporal fluctuations of the equivalence ratio in lean premixed gas turbine combustors are one of the most important driving mechanisms for thermoacoustic instabilities. In this work, high-amplitude equivalence ratio fluctuations in the mixing section of a swirl-stabilized burner are assessed for the first time. The applied non-intrusive sensor is based on fixed-wavelength modulation spectroscopy of methane at 1653 nm using a near-infrared tunable diode laser. The measurements are performed at isothermal operating conditions without the presence of a flame at 25°C and at atmospheric pressure. The equivalence ratio fluctuations are generated by acoustic forcing of the air flow while the fuel injection flow rate is kept constant. Acoustic forcing amplitudes up to 220% of the mean flow velocity are assessed. Measurements are conducted at different axial distances from the fuel injection point to study the spatio-temporal evolution of the equivalence ratio fluctuations. The results show a frequency-dependent saturation of temporal equivalence ratio fluctuations with increasing forcing amplitude, which can not be described through the available model. These results are in good agreement with preceding studies and indicate the saturation of the flame response due to a saturation of equivalence ratio fluctuations. Furthermore, a decreased attenuation of temporal mixture inhomogeneities for small forcing amplitudes is found.
Advanced Identification of Coherent Structures in Swirl-Stabilized Combustors
Authors: Sieber Moritz , Paschereit Christian Oliver , Oberleithner Kilian
Published: September 2016

We present an application of a newly introduced method to analyze the time-resolved experimental data from the flow field of a swirl-stabilized combustor. This method is based on classic proper orthogonal decomposition (POD) extended by a temporal constraint. The filter operation embedded in this method allows for continuous fading from the classic POD to the Fourier mode decomposition. This new method — called spectral proper orthogonal decomposition (SPOD) — allows for a clearer separation of the dominant mechanisms due to a clean spectral separation of phenomena. In this paper, the fundamentals of SPOD are shortly introduced. The actual focus is put on the application to a combustor flow. We analyze high-speed PIV measurements from flow fields in a combustor at different operation conditions. In these measurements, we consider externally actuated, as well as natural dynamics and reveal how the natural and actuated modes interact with each other. As shown in the paper, SPOD provides detailed insight into coherent structures in swirl flames. Two distinct PVC structures are found that are very differently affected by acoustic actuation. The coherent structures are related to heat release fluctuations, which are derived from simultaneously acquired OH* chemiluminescence measurements. Besides the actuated modes, a low frequency mode was found that significantly contribute to the global heat release fluctuations.
Effusion Cooled Combustor Liner Tiles With Modern Cooling Concepts: A Comparative Experimental Study
Authors: Jackowski Thomas , Schulz Achmed , Bauer Hans-Jörg , Gerendás Miklós , Behrendt Thomas
Published: September 2016

Advanced combustion techniques in aero engines require highly effective cooling schemes of combustor liners. One parameter affecting the cooling performance is the geometry of the cooling holes themselves. So far, the freedom in the design of cooling holes was limited due to the manufacturing techniques. With emerging additive manufacturing methods, e.g. Direct Metal Laser Sintering, however, the geometry of the cooling holes is virtually unlimited. Especially the entrance and the curvature of the cooling holes determines the through-flow of the hole and consequently the cooling performance of the ejected cooling film.

In this study a set of combustor liner tiles with two innovative and four traditional cooling hole geometries will be analyzed and compared to each other in terms of cooling performance. The innovative geometries have bent cooling holes with a nearly horizontal outlet. All specimens have the same cooling hole pattern. The cooling performance is determined by comparing the total cooling effectiveness for a given pressure difference across the combustor liner tiles. The coolant mass flow rate is gained from experimentally determined discharge coefficients for the respective pressure difference. The first set of measurements is conducted in an atmospheric open-loop test rig at reduced temperatures but realistic density ratios between hot gas and coolant.

The specimen with the best cooling performance has been selected for an investigation in a high pressure test rig at realistic combustor conditions (pressure, temperature) including fluctuations of the cooling air to simulate combustion instabilities. The cooling performance again is determined by the total cooling effectiveness for a given pressure difference across the combustor liner tiles.
Analytical Uncertainty Analysis for Hot-Wire Measurements
Authors: Hösgen Christian , Behre Stephan , Hönen Herwart , Jeschke Peter
Published: September 2016

This paper presents an analytical approach to quantify the uncertainty of velocity-, angle- and turbulence intensity measurements obtained by hot-wire anemometry. Based on the given results of the model, the paper points out the most critical parameters influencing the quality of the flow measurement.

The proposed approach is divided into three subsections, each accounting for the uncertainties due to the measurement equipment, probe calibration and measurements in the test-rig respectively. For every subsection the propagation of uncertainties is stated in the form of analytical formulas starting with the measurands, e.g. voltage, pressure, temperature, and ending with velocity, angle and turbulence intensity. The approach is applied to data, obtained by a triple hot-wire probe, on two-dimensional traverses in a 1.5-stage cold air turbine. The calibration process and, in particular, the temperature and pressure deviation between calibration and measurement are identified as having a major impact on the measurement uncertainty. With respect to the Aachen test-rig the uncertainty in velocity measurements is lower than 3.5%. The investigation of the turbulence intensity uncertainty yields an increase in uncertainty with increasing turbulence intensity but an insensitivity towards pitch- and yaw-angle uncertainty.
Numerical Investigation of Heat Transfer in a High Aspect Ratio Double Wall Channel With Pin Fin and Jet Array Impingement
Authors: Kan Rui , Tian Shuqing
Published: September 2016

A combined impingement-pedestal geometry for turbomachinery double wall cooling application is studied numerically with the shear stress transport turbulence model. Conjugated CFD simulation is performed to investigate the cooling effectiveness distribution. The configuration consists of a high aspect ratio cooling duct, with jet array impinging onto the pin fin-roughed wall. The jet Reynolds number varies from 8,000 to 80,000, jet-to-target wall spacing is kept constant at Z/Dj=0.8. Three main parameters are investigated, including the jet Reynolds number, pin fin shapes (circular and elongated) and the relative location between jets and pin fins (the jet placed uniformly inside the duct or more densely at the front of the duct). For more detailed investigations, the pin fin diameter and impingement hole diameter are varied independently, and a total of 26 configurations are studied. The results show that the double wall configuration with pin fins significantly increases the heat transfer coefficients, compared to that with only impingement. Non-uniform jet arrangement results in a stronger crossflow and enhances heat transfer on the pins, which brings an increase of cooling effectiveness and more uniform temperature distribution.
Detailed Numerical Characterization of the Suction Side Laminar Separation Bubble for a High-Lift Low Pressure Turbine Blade by Means of RANS and LES
Authors: Bigoni Fabio , Vagnoli Stefano , Arts Tony , Verstraete Tom
Published: September 2016

The scope of this work is to obtain a deep insight of the occurrence, development and evolution of the laminar separation bubble which occurs on the suction side of the high-lift T106-C low pressure turbine blade operated at correct engine Mach and Reynolds numbers. The commercial codes Numeca FINE/Turbo and FINE/Open were used for the numerical investigation of a set of three different Reynolds numbers. Two different CFD approaches, characterized by a progressively increasing level of complexity and detail in the solution, have been employed, starting from a steady state RANS analysis and ending with a Large Eddy Simulation. Particular attention was paid to the study of the open separation occurring at the lowest Reynolds number, for which a Large Eddy Simulation was performed in order to try to correctly capture the involved phenomena and their characteristic frequencies. In addition, the potentialities of the codes employed for the analysis have been assessed.
A Numerical Study on the Reactivity of Biogas/Reformed-Gas/Air and Methane/Reformed-Gas/Air Mixtures at Gas Turbine Relevant Conditions
Authors: Diaz Gomez Maqueo Pablo , Versailles Philippe , Bourque Gilles , Bergthorson Jeffrey M.
Published: September 2016

This study investigates the increase in methane and biogas flame reactivity enabled by the addition of syngas produced through fuel reforming. To isolate thermodynamic and chemical effects on the reactivity of the mixture, the burner simulations are performed with a constant adiabatic flame temperature of 1800 K. Compositions and temperatures are calculated with the chemical equilibrium solver of CANTERA® and the reactivity of the mixture is quantified using the adiabatic, freely-propagating premixed flame, and perfectly-stirred reactors of the CHEMKIN-Pro® software package. The results show that the produced syngas has a content of up to 30 % H2 with a temperature up to 950 K. When added to the fuel, it increases the laminar flame speed while maintaining a burning temperature of 1800 K. Even when cooled to 300 K, the laminar flame speed increases up to 30 % from the baseline of pure biogas. Hence, a system can be developed that controls and improves biogas flame stability under low reactivity conditions by varying the fraction of added syngas to the mixture. This motivates future experimental work on reforming technologies coupled with gas turbine exhausts to validate this numerical work.
Numerical Simulation of the Flow in the Row Labyrinth Seal of an Axial Turbine Using a Low-Mach Preconditioning Technique
Authors: Fiedler Jens , Weber Anton , Bosco Arianna , Engel Karl , di Mare Francesca
Published: September 2016

Labyrinth seals on shrouded blades are an effective way for reducing efficiency penalties, as compared to free ended blades. Due to the difficulties of gaining optical access to cavity regions, mostly pressure measurements are available in the literature, from which the details of the flow must be inferred. The use of numerical tools can provide insight in the flow topology and therefore help obtaining a better understanding of the factors (geometric, thermodynamic and aerodynamic) which can affect the performance of the machine. Whilst in the main passage relatively high Mach numbers are to be found (0.3–1.3), the flow field in the cavities is dominated by extremely low flow speeds with strong recirculation patterns. The treatments of such flows, where large disparities between the acoustic and convective speed exist, are known to be highly problematic if density-based solution methods are employed. As the flow conditions approach the incompressibility limit a degradation of the convergence behaviour can be observed, leading, potentially to incorrect solutions. In order to overcome these problems preconditioning methods can be conveniently applied to the Navier-Stokes equations. In the current work the formulation of a fully implicit local preconditioning method with domain control of the Mach number dependency is presented. Numerical simulations of turbomachine components are generally performed on truncated domains. In order to prevent unphysical reflections at open boundaries and interfaces non-reflecting boundary conditions have been developed, e.g. [6, 8]. As reported in the available literature, low Mach preconditioning can cause stability problems and strongly impair the quality of the results especially in proximity of the domain’s boundaries. As shown in [14, 21] an appropriate scaling treatment of the boundary conditions is also required to alleviate such issues.

In the current work non-reflecting boundary conditions, based on the formulation of Giles [6, 8], have been suitably modified to work reliably also in the limit of incompressible flows. To prove the robustness and accuracy of the algorithm implemented in the DLR’s CFD code TRACE, a canonical testcase representing an abstraction of the flow topology found in a labyrinth seal, such as a lid-driven cavity, is shown. Finally, the simulation of the steady flow in a multistage, shrouded low-pressure turbine is presented. For this, a classic RANS approach has been adopted using the k-ω model to illustrate the effectiveness of the developed method in a typical industrial application. Of particular significance and interest is the analysis of the mass conservation properties of the numerical scheme attained at mixing planes between rotor and stator and at non-matching grid interfaces, denoted as “zonal” and “zonal-mixed” interfaces.
Modulation and Radial Migration of Turbine Hub Cavity Modes by the Rim Seal Purge Flow
Authors: Schädler R. , Kalfas A. I. , Abhari R. S. , Schmid G. , Voelker S.
Published: September 2016

In the present paper, the results of an experimental and numerical investigation of the hub cavity modes and their migration into the main annulus flow are presented. A one-and-a-half stage, unshrouded and highly loaded axial turbine configuration with 3-dimensionally shaped blades and cylindrical end walls has been tested in an axial turbine facility. Both, the blade design and the rim seal purge flow path are representative to modern high pressure gas turbines.

The unsteady flow field at the hub cavity exit region has been measured with the fast-response aerodynamic probe (FRAP) for two different rim seal purge flow rates. Furthermore, fast-response wall mounted pressure transducers have been installed inside the cavity. Unsteady full-annular CFD simulations have been employed in order to complement the experimental work.

The time-resolved pressure measurements inside the hub cavity reveal clear cavity modes which show a strong dependency on the injected amount of rim seal purge flow. The numerical predictions provide information on the origin of these modes and relate them to pronounced ingestion spots around the circumference.

The unsteady probe measurements at the rim seal interface show that the signature of the hub cavity induced modes migrates into the main annulus flow up to 30% blade span. Based on that, an aerodynamic loss mechanism has been found, showing that the benefit in loss reduction by decreasing the rim seal purge flow rate is weakened by the presence of turbine hub cavity modes.
Clocking in Low-Pressure Turbines
Authors: Evans K. R. , Longley J. P.
Published: September 2016

The effect of stator clocking has been experimentally and computationally investigated using a low-speed, two-stage, Low-Pressure Turbine which was specifically designed to maximise the clocking potential by aligning the Stator 1 wake segments with the Stator 2 leading edge along the span. It was verified that the wake segments are aligned to within 10% of stator pitch across the span. The measured clocking effect on the work extraction is 0.12% and on efficiency is 0.08%. Although the effect of clocking is small, it is repeatable, periodic across four stator pitches and consistent between independent measurements. Furthermore, factors to consider for a reliable clocking investigation are discussed.

The measurements revealed that the majority of the clocking effect on the work extraction occurs in Stage 2 and it originates at Stator 2 exit. This indicates that the flow is being processed differently within Stator 2. There is also an effect on the Stage 1 work. In each blade row the measured clocking effect on the lost work is similar across the span.

The computations with meshed cavities do not capture any clocking effects in Stage 1. This indicates that an unsteady viscid phenomenon within Rotor 1 is not captured by the fully turbulent calculation e.g. unsteady transition. However, the computations do capture the measured clocking effect on the Stage 2 work extraction.

It is hypothesised that the clocking effect on Stator 2 flow turning is dominated by a steady, inviscid process.
Accounting for Eccentricity in Compressor Performance Predictions
Authors: Young Anna M. , Cao Teng , Day Ivor J. , Longley John P.
Published: September 2016

It is known that the tip-clearance gap plays a pivotal role in determining the performance of an aero-engine compressor, both in terms of stability and pressure rise. However, the exact effect of tip-clearance size on stall margin over the range of clearances experienced during engine operation is not known, and most designers substitute rules of thumb for real knowledge when developing new compressors.

Eccentricity in the tip-clearance is also known to affect compressor performance, though again little work has been done to quantify the penalties. It is generally assumed that the stall margin of an eccentric machine will be approximately equal to that of a concentric machine with a clearance equal to the maximum eccentric clearance. Results given in this paper show a stabilising effect of the small tip-clearance sector of an eccentric compressor on the large tip-clearance sector, so the penalty on stall margin is not as large as commonly assumed. It is shown that the stall margin penalty in a single-stage eccentric machine is only 50–60% of that which a concentric compressor with clearance equal to the maximum clearance would exhibit. In addition, the effect of eccentricity is shown to diminish as the average clearance is increased.

In this paper, experiments and computational modelling are used to examine the effects of eccentric tip-clearance on flowfield redistribution and compressor performance. In particular, the three-dimensional nature of the flowfield generated by an eccentric gap is shown for the first time. The purpose of this work is not to provide ‘hard-and-fast’ design rules for eccentric compressors (this cannot be done on the basis of single-stage measurements), but to provide a starting point for a better physical understanding of the problem.
A High-Dimensional Model Representation Guided PSO Methodology With Application on Compressor Airfoil Shape Optimization
Authors: Safari A. , Hajikolaei K. H. , Lemu H. G. , Wang G. G.
Published: September 2016

Although metaheuristic techniques have recently become popular in optimization, still they are not suitable for computationally expensive real-world problems, specifically when the problems have many input variables. Among these techniques, particle swarm optimization (PSO) is one of the most well-known population-based nature-inspired algorithms which can intelligently search huge spaces of possible arrangements of design variables to solve various complex problems. The candidate solutions and accordingly the required number of evaluated particles, however, dramatically increase with the number of design variables or the dimension of the problem. This study is a major modification to an original PSO for using all previously evaluated points aiming to increase the computational efficiency. For this purpose, a metamodeling methodology appropriate for so-called high-dimensional, expensive, black-box (HEB) problems is used to efficiently generate an approximate function from all particles calculated during the optimization process. Following the metamodel construction, a term named metamodeling acceleration is added to the velocity update formula in the original PSO algorithm using the minimum of the metamodel. The proposed strategy is called the metamodel guided particle swarm optimization (MGPSO) algorithm. The superior performance of the approach is compared with original PSO using several benchmark problems with different numbers of variables. The developed algorithm is then used to optimize the aerodynamic design of a gas turbine compressor blade airfoil as a challenging HEB problem. The simulation results illustrated the MGPSO’s capability to achieve more accurate results with a considerably smaller number of function evaluations.
3D CFD Compressor Map Computation of a Multi-Stage Axial Compressor With Off-Design Adjusted Rotor Geometries
Authors: Janke Christian , Goller Markus , Martin Ivo , Gaun Lilia , Bestle Dieter
Published: September 2016

Compressor maps of aero engines show the relation between corrected mass flow, corrected shaft speed, pressure ratio, and efficiency, where different operating conditions of the compressor are represented by different speed lines. These speed lines are an important information for the compressor design process, since they show important operation bounds like surge and choke. Typically, 3D CFD compressor maps are computed with the so called hot geometry given by the aerodynamic design point. But in reality aerofoil shapes change depending on engine speeds and gas loads resulting in twist of the blades and changes of tip clearance. In order to obtain a higher quality compressor map, all these effects must be taken into account. Therefore, a process is utilized which uses coupled CFD and FE analyses to account for load adjusted geometries aside the design point. For transformation of FE results into the CFD model a cold-to-hot blade morphing technique is used. The studies are performed for a 4.5 stage high speed axial compressor, where effects of varying tip clearance and geometry deformation are considered separately from each other. Finally, their combined effects are studied.
The Effects on Stability, Performance, and Tip Leakage Flow of Recirculating Casing Treatment in a Subsonic Axial Flow Compressor
Authors: Wang Wei , Chu Wuli , Zhang Haoguang , Wu Yanhui
Published: September 2016

Recirculating casing treatment (RCT) was studied in a subsonic axial flow compressor experimentally and numerically. The RCT was parameterized with the injector throat height and circumferential coverage percentage (ccp) to investigate its influence on compressor stability and on the overall performance in the experimentation. The injector throat height varied from 2 to 6 times the height of the rotor tip clearance, and the ccp ranged from 8.3% to 25% of the casing perimeter. Various RCT configurations were achieved with a modular design procedure. The rotor casing was instrumented with fast-response pressure transducers to detect the stall inception, rotational speed of stall cells, and pressure flow fields. Whole-passage unsteady simulations were also implemented for the RCT and solid casing to understand the flow details. Results indicate that both the compressor stability and overall performance can be improved through RCT with appropriate geometrical parameters. The effect of injector throat height on the stability depends on the choice of ccp, i.e., interaction effect exists. In general, the RCT with a moderate injector throat height and a large circumferential coverage is the optimal choice. Phase-locked pattern of the casing wall pressure reveals a weakened tip leakage vortex under the effect of RCT compared with the solid casing. The numerical results show that the RCT has a substantial effect on tip blockage even when the blade passages break away from the domain of RCT. The reduction of tip blockage induced by the tip leakage vortex is the main reason for the extension of stable operation range. The unsteadiness of double-leakage flow is detected both in the experiment and in numerical simulations. The pressure fluctuations caused by double-leakage flow are depressed with RCT. This observation indicates reduced losses related with the double-leakage flow. Although the stall inception is not changed by implementing RCT, the stall pattern is altered. The stall with two cells is detected in RCT compared with the solid casing with only one stall cell.
Numerical and Experimental Investigation of Noise Radiation for Double-Suction Centrifugal Fans of Air-Conditioner Indoor Unit
Authors: Zhao Xifeng , Sun Jinju , He Youwei , Lin Kang
Published: September 2016

Aerodynamic performance and noise level of AC (Air-Conditioner) indoor unit fan system has always been receiving much concern. But most previous studies have been focused on aerodynamic performance and noise level, and less attention is paid to the noise radiation problem. It is essential to identify and quantify the noise radiation, so as to control and reduce the noise level effectively.

The objective of the present investigation is to identify the influential factors of the Blade Passing Frequency (BPF) noise radiation and investigate noise directivity distribution of two double-suction centrifugal flow fans with different fan blades used in AC indoor unit. A hybrid numerical approach is developed for predicting the unsteady fan flow and noise radiation. Unsteady fan flow is solved by Detached Eddy Simulation (DES) method. Then the areoacoustic calculation is performed based on the predicted unsteady flow field, and the obtained results are used as dipole source inputs in acoustic Finite Element Model (FEM) for predicting the noise radiation of the centrifugal fans.

Experimental tests of aerodynamic and aeroacoustic performance are conducted respectively for the fan. A comparison is made between the experimental and measured results. It is demonstrated that the predicted aeroacoustic and noise radiation both agree with the measured ones and the fan blade geometry is much influential to acoustic prediction. Finally, directivity of BPF noise radiation to noise level is evaluated for two different fans.
A One-Dimensional Analytical Method for Turbine Blade Preliminary Cooling Design
Authors: Li Chen , Liu Jian-jun
Published: September 2016

The turbine blade cooling design is a complex procedure including one-dimensional preliminary cooling design, detailed two-dimensional design and fluid network analyses, and three-dimensional conjugate heat transfer and FEM predictions. Frequent alteration and modification of the cooling configurations make it unpractical to obtain all of three-dimensional design results quickly. Preliminary cooling design deals mainly with the coolant requirements and can be knitted into fluid network to look up the expected cooling structural style to promote three-dimensional geometry design.

Previous methods to estimate the coolant requirements of the whole turbine blade in the preliminary cooling design were usually based on the semi-empirical air-cooled blade data. This paper combines turbine blade internal and external cooling, and presents a one-dimensional theoretical analytical method to investigate blade cooling performance, assuming that the coolant temperature increases along the blade span. Firstly, a function of non-dimensional cooling mass flow rate is derived to describe the new relationship between adiabatic film cooling effectiveness and overall cooling effectiveness. Secondly, a new variable related to film cooling is found to estimate the required adiabatic film cooling effectiveness without using the empirical correlations. Finally, a theoretical calculation about the relationship between non-dimensional cooling mass flow rate and overall cooling effectiveness well corresponds to semi-empirical air-cooled blade data within regular range of cooling efficiency.

The currently proposed method is also a useful tool for the blade thermal analysis and the sensitivity analysis of coolant requirements to various design parameters. It not only can provide all the possible options at the given gas and coolant inlet temperatures to meet the design requirement, but also can give the third boundary conditions for calculating the blade temperature field. It’s convenient to use the heat transfer characteristic of internal cooling structures to estimate the coolant mass flow rate and the channel hydraulic diameter for both convection cooling and film cooling.
Evaluating Effects of Fuel Properties on Smoke Emissions
Authors: Ruslan Meor , Ahmed Ihab , Khandelwal Bhupendra
Published: September 2016

Volatile oil price and environmental impact of conventional jet fuel are key motivators towards the proposing of alternative jet fuels. This article introduces and establishes a relationship between jet fuel properties/composition and smoke emission. It is an important and comprehensive task as it underlines the base references and scientific reasoning on fuel compositions / properties; very few, if any, studies have investigated the effects of each of the properties/ compositions on smoke emissions. Two sets of fuels were tested on small re-commissioned Honeywell GTCP85 APU gas turbine Engine. The first set was consisted of 8 novel fuels, while the second was a blend of varied percentages of Jet A-1 and other alternative fuel. This is to provide a wide range of properties and compositions. The results were compared to those of Jet A-1on the same platform (Honeywell GTCP85 APU). It was observed that not all fuel compositions/properties have the same effects on the smoke number. Some of them such as: Specific Energy, Kinematic, viscosity, Biphenyls, monocycloparaffin, AlkylBenzene, Fluorenes, Distillation temp (90%), Carbon (%mass), Naphthalene, Composite Density, Benzocycloparaffin, Density at 15C°, Aromatics (%Vol) and Net heat of Combustion have a clear direct effect on the smoke number, while others such as iso-paraffin and flashpoint have a reduced impact on smoke number. This data shall be used to predict the effect of certain composition/ property on the smoke emission, thus it could be avoided or to be taking into considerations when producing or using new alternative fuels.