bluuebunny/crossref_metadata_2025 huggingface dataset:
Experimental Measurements of Ingestion Through Turbine Rim Seals: Part 5 — Fluid Dynamics of Wheel-Space
Authors: Sangan Carl M. , Lalwani Yogesh , Owen J. Michael , Lock Gary D.
Published: November 2013

This paper discusses the flow structure in typical rotor-stator systems with ingress and egress. Measurements of concentration, velocity and pressure were made using a rotating-disc rig which experimentally simulated hot gas ingestion into the wheel-space of an axial turbine stage. Externally-induced ingress through rim seals was generated from the non-axisymmetric pressures produced by the flow over the vanes and blades in the external annulus. Measurements were conducted using several single- and double-seal geometries and for a range of sealing flow rates and rotational speeds. The concentration measurements showed that the amount of ingress, which increased with decreasing sealing flow rate, depended on the seal geometry. The swirl velocity in the fluid core increased with decreasing sealing-flow rate but, outside the outer region in the wheel-space, it was largely unaffected by the seal geometry or by the amount of ingress. The radial distribution of static pressure, calculated from the measured swirl velocity in the core, was in good agreement with the pressures measured on the stator. The data for the double seals demonstrated that the ingested gas was predominately confined to the region between the seals near the periphery of the wheel-space; in the inner wheel-space, the effectiveness is shown to be significantly higher.
Application of an Air-Curtain Fluidic Jet Type Seal to Reduce Turbine Shroud Leakage
Authors: Auld Alison , Hilfer Michael , Hogg Simon , Ingram Grant , Messenger Andrew
Published: November 2013

This paper describes the design and optimization of an ‘air-curtain’ type seal using a fluidic jet to reduce tip leakage losses on a small high-speed single stage axial turbine device. The application will essentially demonstrate proof of concept for turbomachinery applications, opening the door for the development of future designs for applications in all scales of turbomachine. CFD is used to develop and optimize the seal design. The performance benefit from applying the new seal is predicted. These calculations illustrate the importance of accurately accounting for the effects of the sealing jet on shroud shear forces, in addition to leakage flow reduction, when determining the overall gain in turbine output power from the improved sealing. It is planned to validate the new seal design in full-scale turbine tests, during the next phase of the work.
Heat Transfer in Rotating Multi-Pass Rectangular Smooth Channel With and Without a Turning Vane in Hub Region
Authors: Lei Jiang , Li Shiou-Jiuan , Han Je-Chin , Zhang Luzeng , Moon Hee-Koo
Published: November 2013

This paper experimentally investigates the effect of a turning vane on hub region heat transfer in a multi-pass rectangular smooth channel at high rotation numbers. The experimental data were taken in the second and the third passages (Aspect Ratio = 2:1) connected by an 180° U-bend. The flow was radial inward in the second passage and was radial outward after the 180° U-bend in the third passage. The Reynolds number ranged 10,000 to 40,000 while the rotation number ranged 0 to 0.42. The density ratio was a constant of 0.12. Results showed that rotation increases heat transfer on leading surface but decreases it on the trailing surface in the second passage. In the third passage, the effect of rotation is reversed. Without a turning vane, rotation reduces heat transfer substantially on all surfaces in the hub 180° turn region. After adding a half-circle-shaped turning vane, heat transfer coefficients do not change in the second passage (before turn) while they are quite different in the turn region and the third passage (after turn). Regional heat transfer coefficients are correlated with rotation numbers for multi-pass rectangular smooth channel with and without a turning vane.
Gas Turbine Cooling: Moving From Macro to Micro Cooling
Authors: Bunker Ronald S.
Published: November 2013

Up to the present time, virtually all production gas turbine engines utilize high-pressure turbine components that are cooled by various modes of bulk, or macro, internal flow passages and networks, many of which are then linked to external film cooling by holes through the walls. Typical examples of macro cooling include serpentine passages, swirl chambers, pin banks, and impingement jets. All cooled turbine airfoils in commercial operation today utilize cooling channels and film holes that are considered to be macroscopic in physical magnitude. The conventional delimiters between macro and micro cooling are (1) whether the feature (internal passages and holes) can be repeatedly manufactured via investment casting methods, and (2) whether the flow passage (film hole, impingement orifice, dust hole) may become plugged by particles in the cooling fluid. The acceptable sizes are relative to the turbine size and the operating environment. This study examines the limitations and weaknesses inherent in macro cooling and the reasoning that demands gas turbine cooling change to micro cooling methods. A brief history of the developments in micro cooling for turbine airfoils is presented, including the reasons why it has not yet become commercially viable. A simple cooling study is used to demonstrate the very significant performance gains that can be obtained with micro cooling, such as cooling flow reductions approaching 40%, and thermal stress (gradient) reductions of 50%. The key to realizing these gains in full lies in cost effective manufacturing and durability in operation.
Flame Density Ratio Effects on Vortex Dynamics of Harmonically Excited Bluff Body Stabilized Flames
Authors: Emerson Benjamin , Murphy Kelvin , Lieuwen Tim
Published: November 2013

This paper presents an experimental study of acoustically forced bluff body stabilized flames, motivated by the problem of combustion instabilities. The goal of the work is to better understand the flame and flow behavior as functions of the proximity of the acoustic frequency to natural hydrodynamic instability frequencies. It is well known that iso-density, high Re bluff body wakes are globally unstable, exhibiting the Von Karman vortex street. In reacting flows, however, this global mode is suppressed if the density ratio across the flame is sufficiently high. Thus, the density ratio is an important parameter that influences the global mode growthrate. In this study, the flame was longitudinally forced over a range of hydrodynamic global mode to forcing frequency ratios, density ratios, and forcing amplitudes. Longitudinal forcing leads to the symmetric rollup of the two separating shear layers. When the forcing frequency is in the vicinity of the wake’s global mode frequency, the global mode locks into the forcing frequency, and the symmetric shear layers quickly stagger as they convect downstream, leading to a large scale, sinuous flapping of the wake and flame. The axial position at which staggering occurs is a function of the forcing amplitude and the proximity of the forcing frequency to the global mode frequency.

The lock-in phenomenon amplifies the flame’s motion at the forcing frequency. However, if the vortices stagger to a fully sinuous structure, this causes a significant reduction in the flame’s oscillatory heat release through phase cancellation of the upper and lower flame branches. Therefore, if a low density ratio flame is subjected to longitudinal acoustic forcing near its global mode frequency, it will respond with weaker heat release fluctuations than it would away from lock-in. This is true even though the local degree of flame flapping is quite significant. Thus, the results of this study show some phenomena that contradict conventional notions, namely that forcing a globally unstable flow near its global mode frequencies can lead to diminished local heat release oscillations.
Effect of Fuel Staged Proportion on NOX Emission Performance of Centrally Staged Combustor
Authors: Liu Fuqiang , Mu Yong , Liu Cunxi , Yang Jinhu , Mao Yanhui , Xu Gang , Zhu Junqiang
Published: November 2013

The low NOX emission technology has become an important feature of advanced aviation engine. A wide range of applications attempt to take advantage of the fact that staged combustion operating under lean-premixed-prevaporized (LPP) conditions can significantly decrease pollution emissions and improve combustion efficiency. In this paper a scheme with fuel centrally staged and multi-point injection is proposed. The mixing of fuel and air is improved, and the flame temperature is typically low in combustion zone, minimizing the formation of nitrogen oxides (NOX), especially thermal NOX. In terms of the field distribution of equivalence ratio and temperature obtained from Computational Fluid Dynamics (CFD), a chemical reactor network (CRN), including several different ideal reactor, namely perfectly stirred reactor (PSR) and plug flow reactor (PFR), is constructed to simulate the combustion process. The influences of the pilot equivalence ratio and percentage of pilot/main fuel on NOX and carbon monoxide (CO) emissions were studied by Chemical CRN model. Then the NOX emission in the staged combustor was researched experimentally. The effects of the amount of pilot fuel and primary fuel on pollution emissions were obtained by using gas analyzer. Finally, the effects of pilot fuel proportion on NOX emission were discussed in detail by comparing predicts of CRN and experimental results.
Cooling Efficiency for Assessing the Cooling Performance of an Effusion Cooled Combustor Liner
Authors: Wurm B. , Schulz A. , Bauer H.-J. , Gerendas M.
Published: November 2013

Based on experimental results on a liner of a modern direct lean injection combustion chamber using coolant ejection from both effusion cooling holes and a starter film, a method is presented that allows the assessment of the cooling performance of the liner.

As the main focus of the present study is a deeper understanding of the interaction of swirl flow and near wall cooling flow, wall pressure measurements are performed for the calculation of local blowing ratios and local coolant mass fluxes. Thermal investigations allow the calculation of adiabatic film cooling effectiveness and heat transfer coefficients. The pressure drop across the effusion cooled liner is varied between 1% and 3% of the total pressure of the main flow. As experiments are performed without combustion and at low temperature, the influence of radiation is neglected. Results show the impact of the swirled main flow on the stability of the starter film and on the effusion cooling performance. Stagnation areas which could be identified by wall pressure measurements are confirmed by detailed PIV measurements. Thermal investigations reveal reduced cooling performance in the respective stagnation areas. For the definition of the non dimensional cooling efficiency the measurement area is sub divided into rhombic sections, which are located around each effusion cooling hole. Based on the measurement results presented, heat fluxes per unit area can then be calculated and put together to the cooling efficiency.
Analysis and Improvement of a Loosely Coupled Fluid-Solid Heat Transfer Method
Authors: Wang Zhi , Corral Roque , Chaquet Jose M. , Pastor Guillermo
Published: November 2013

A new, fast and reliable loosely coupled fluid-solid heat transfer method is described. The approach is based on an improvement of heat transfer coefficient Forward Temperature Backward method. The numerical analysis of the coupling process of an one-dimensional model, shows that the convergence behaviour is influenced by the physical Biot number and a virtual heat transfer coefficient. The effect of using partially converged CFD solutions has also been studied. The novelty of the method is based on the use of dynamic evaluation of the numerical parameters of the coupling, namely the virtual heat transfer coefficient. Two representative models in turbomachinery applications are applied to test the proposed method. It is shown that the computational time can be reduced by a factor of three to five relative to previous existing methods. The new approach only spends around two times the cost of stand-alone CFD simulation of the problem to obtained a coupled fluid/solid thermal analysis.
Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer
Authors: Anto K. , Xue S. , Ng W. F. , Zhang L. J. , Moon H. K.
Published: November 2013

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet.

Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.
An Investigation Into Performance Modeling of a Small Gas Turbine Engine
Authors: Leylek Zafer , Anderson Wesly S. , Rowlinson Glen , Smith Nigel
Published: November 2013

A small gas turbine performance modeling and testing project has been completed as part of a DSTO research program. The main objective of the program was to enhance capability in understanding and modeling the thermodynamic and performance characteristics of gas turbine engines. The secondary objective of the program was development of a simple, low-cost test apparatus for basic thrust augmentation technologies and infrared suppression studies.

Engine performance modeling was conducted using commercial software (Gasturb) and an in-house developed code. Various techniques were used in predicting component performance which included scaling of published performance data, use of standard empirical performance models, mean-line and through flow codes and detailed Computational Fluid Dynamics (CFD) analysis. A comparison of the advantages and disadvantages of each method was made and compared with engine test data.

The AMT Olympus HP turbojet engine was used as the test engine. The engine was instrumented using a bell mouth to measure mass flow rate, load cell to measure thrust, tachometer to measure engine rotational speed, thermocouples to measure flow total temperature and pressure transducers to measure wall static pressure at various stations along the engine.

A pair of Olympus engines were tested for comparison during baseline testing, with consistent results between the two. The first engine was fully instrumented and used in all of the tests. This engine was used to test the engine standard operating line tests to determine bulk performance and establish compressor operating line. Tests were repeated and error analysis conducted to ensure repeatability and validity of the data. The second engine was used as a control engine with only the OEM supplied instrumentation, thrust and bell mouth used for benchmarking purposes. The data from both engines have been compared with the engine performance model and OEM data.

A number of other tests were completed to “stress” the engine and shift its operating line closer to the compressor stall line. This was accomplished through various types of exhaust blockage. Stall behaviour was clearly evident in the initial commissioning tests where a large nozzle blockage resulted in engine stall and incomplete start-up. Engine performance and compressor map results from the tests have been compared to the engine performance model with good agreement.
Automated Combustor Preliminary Design Using Tools of Different Fidelity
Authors: Angersbach Andreas , Bestle Dieter , Eggels Ruud
Published: November 2013

The design of a modern aero-engine combustor is a highly complex and multi-disciplinary task. The combustor design is strongly driven by severe emission regulations and ACARE 2020/2050 goals. Furthermore, new designs have to be developed within short turn-around times. This paper describes a novel approach of an automated preliminary aero-thermal design process of a rich-burn combustor combining 1D, 2D and 3D design tools in order to speed up the design loop and provide improved combustor designs in an early design stage.

The automated design process includes a knowledge-based preliminary design tool, an 1D network solver, a parametric 3D geometry model, a meshing tool, and 3D-CFD analysis. At first, a preliminary combustor design is created based on industrial in-house design rules. The preliminary design tool provides a 2D geometry model and cooling layout. It is coupled with an 1D network solver to calculate the air distribution inside the combustor. The design process includes two state-of-the-art combustor cooling schemes, effusion cooling and impingement effusion cooling. An air flow model for both cooling schemes is created within the network, respectively. The computed air distribution is subsequently used to generate boundary conditions for a 3D-CFD analysis. To perform the CFD calculations, a parametric 3D geometry model of a combustor sector has been developed based on a 2D preliminary design which takes into account mixing port properties, fuel injector, and combustor wall cooling. After an automated meshing 3D-CFD computations are performed. As a result, quick automatic estimation of combustor emissions, size and efficiency can be obtained within the design process. A CFD parameter study of a mixing port variation and their effect on the emissions of NOx and soot is performed using the described layout process.
Effect of Blockage-Ratio on Developing Heat Transfer for a Rectangular Duct With Transverse Ribs
Authors: Ghorbani-Tari Zahra , Wang Lei , Sunden Bengt
Published: November 2013

The developing heat transfer characteristics in a rectangular channel (AR = 4) equipped with continuous transverse ribs are experimentally investigated. The ribs were regularly spaced over a section of the channel which was heated by a uniform heat flux. The blockage ratio e/Dh varied from 0.039 to 0.078. Two values of the rib pitch to rib height ratio (10 and 20) were considered, with the Reynolds number from 57,000 to 127, 000. The studied geometry is relevant to turbine structures between high pressure and low pressure turbines in aircraft engines. The maps of local heat transfer coefficient in the inter-rib regions were obtained by using the steady state liquid crystal thermography. The main purpose is to investigate the effect of blockage ratio (e/Dh) on the developing heat transfer behavior. In particular, the heat transfer characteristics between the first repeated ribs, i.e., in the inter-rib regions were studied, where the flow field is fully developed while the thermal field is not yet periodically fully developed.
Large Eddy Simulation of Swirling Kerosene/Air Spray Flame Using Tabulated Chemistry
Authors: Franzelli B. , Vié A. , Fiorina B. , Darabiha N.
Published: November 2013

Accurate characterization of swirled flames is a key point in the development of more efficient and safer aeronautical engines. The task is even more challenging for spray injection systems. On the one side, spray interacts with both turbulence and flame, eventually affecting the flame dynamics. On the other side, spray flame structure is highly complex due to equivalence ratio inhomogeneities caused by the evaporation process. Introducing detailed chemistry in numerical simulations, necessary for the prediction of flame stabilization, ignition and pollutant concentration, is then essential but extremely expensive in terms of CPU time. In this context, tabulated chemistry methods, expressly developed to account for detailed chemistry at a reduced computational cost in Large Eddy Simulation of turbulent gaseous flames, are attractive. The objective of this work is to propose a first computation of a swirled spray flame stabilized in an actual turbojet injection system using tabulated chemistry. A Large Eddy Simulation of an experimental benchmark, representative of an industrial swirl two-phase air/kerosene injection system, is performed using a standard tabulated chemistry method. The numerical results are compared to the experimental database in terms of mean and fluctuating axial velocity. The reactive two-phase flow is deeper investigated focusing on the flame structure and dynamics.
Effect of Jet Shape of Square Array of Multi-Impinging Jets on Heat Transfer
Authors: Yamane Yoshisaburo , Yamamoto Makoto , Motosuke Masahiro , Honami Shinji
Published: November 2013

It is necessary to increase turbine inlet temperature to improve the performance of the aircraft gas turbine engine. Therefore, effective cooling techniques are still required. The purpose of the present study is to clarify the heat transfer characteristics for the high cooling performance of multiple impinging jets. A focus is placed on the effect of the jet ejection shape, since the shape of jet is expected to enhance flow mixing in accordance with a change in vortex structures at the shear layer of the jet. Experiment was made on the wall jet interaction between adjacent impinging jets by changing the jet ejection shape. Both heat transfer and aerodynamic characteristics in 3×3 square arrays of three types of jet hole shapes, which are circle, cross-shape and oblique cross-shape, are investigated at jet diameter Reynolds number of 4,680. Injection distance is ranged from 2D to 6D, and jet-to-jet spacing is 6D where D is a jet hole diameter. Steady state thermochromic liquid crystal technique is employed to measure local and area averaged Nusselt number. A micro flow sensor, which can detect both flow direction and flow intermittency near the wall, is used to clarify the characteristics of the unsteady flow behavior. It is found that higher local Nusselt number area is spread outward in the concave direction of the cross-shaped jet on the target surface. Characteristic flow behavior due to the flow intermittency and the local flow fluctuation induced by the effect of jet shape are observed in the region surrounded by the adjacent impinging jets in the cases of cross-shaped jet and oblique cross-shaped jet. This intermittent flow phenomenon is considered to contribute to the enhancement of the heat transfer in the intermediate region enclosed by surrounded impinging jets.
Investigation of the Effect of Perforated Sheath on Thermal-Flow Characteristics Over a Gas Turbine Reverse-Flow Combustor: Part 1 — Experiment
Authors: Wang Liang , Wang Ting
Published: November 2013

Reverse-flow combustors have been used in heavy, land-based gas turbines for many decades. A sheath is typically installed over the external walls of the combustor and transition piece to provide enhanced cooling through hundreds of small jet impingement cooling, followed by a strong forced convention channel flow. However, this cooling is at the expense of large pressure loss. With the modern advancement in metallurgy and thermal-barrier coating technologies, it may become possible to remove this sheath to recover the pressure loss without causing thermal damage to the combustor chamber and the transition piece walls. However, without the sheath, the flow inside the dump diffuser may exert nonuniformly reduced cooling on the combustion chamber and transition piece walls. The objective of this paper is to investigate the difference in flow pattern, pressure drop, and heat transfer distribution in the dump diffuser and over the outer surface of the combustor with and without a sheath. Both experimental and computational studies are performed and presented in Part 1 and Part 2, respectively. The experiments are conducted under low pressure and temperature laboratory conditions to provide a database to validate the computation model, which is then used to simulate the thermal-flow field surrounding the combustor and transition piece under elevated gas turbine operating conditions.

The experimental results show that the pressure loss between the dump diffuser inlet and exit is 1.15% of the total inlet pressure for the non-sheathed case and 1.9% for the sheathed case. This gives a 0.75 percentage point (or 40%) reduction in pressure losses. When the sheath is removed in the laboratory, the maximum increase of surface temperature is about 35%, with an average increase of 13%–22% based on the temperature scale of 23 K, which is the temperature difference of bulk inlet and outlet temperature.
An Experimental Study of Mist/Air Film Cooling on a Flat Plate With Application to Gas Turbine Airfoils: Part 2 — Two-Phase Flow Measurements and Droplet Dynamics
Authors: Zhao Lei , Wang Ting
Published: November 2013

A Phase Doppler Particle Analyzer (PDPA) system is employed to measure the two-phase mist flow behavior including flow velocity field, droplet size distribution, droplet dynamics, and turbulence characteristics.

Based on the droplet measurements made through PDPA, a projected profile describing how the air-mist coolant jet flow spreads and eventually blends into the hot main flow is proposed. This proposed profile is found to be well supported by the measurement results of the turbulent Reynolds stresses. The coolant film envelope is identified with shear layers characterized by higher magnitudes of turbulent Reynolds stresses. In addition, the separation between the mist droplet layer and the coolant air film is identified through the droplet measurements — large droplets penetrate through the air coolant film layer and travel further into the main flow.

With the proposed air-mist film profile, the heat transfer results on the wall presented in Part 1 are re-examined and more in-depth physics is revealed. It is found that the location of optimum cooling effectiveness is coincided with the point where the air-mist coolant stream starts to bend back towards the surface. Thus, the data suggests that the “bending back” film pattern is critical in keeping the mist droplets close to the surface, which improves the cooling effectiveness for mist cooling.
Turbulent Consumption Speeds of High Hydrogen Content Fuels From 1–20 atm
Authors: Venkateswaran Prabhakar , Marshall Andrew D. , Seitzman Jerry M. , Lieuwen Tim C.
Published: November 2013

This work describes measurements and analysis of the turbulent consumption speeds, ST,GC, of H2/CO fuel blends. We report measurements of ST,GC at pressures and normalized turbulence intensities, u′rms/SL,0 up to 20 atm and 1800, respectively for a variety of H2/CO mixtures and equivalence ratios. In addition, we present correlations of these data using laminar burning velocities of highly stretched flames, SL,max, derived from quasi-steady leading points models. These analyses show that SL,max can be used to correlate data over a broad range of fuel compositions, but do not capture the pressure sensitivity of ST,GC. We suggest that these pressure effects are more fundamentally a manifestation of non-quasi-steady behavior in the mass burning rate at the flame leading points.
Influence of Nonlinear Flame Models on Sustained Thermoacoustic Oscillations in Gas Turbine Combustion Chambers
Authors: Campa Giovanni , Cinquepalmi Marina , Camporeale Sergio Mario
Published: November 2013

In the recent years a great interest has been devoted to the understanding of the nonlinear dynamics characterizing the thermoacoustic combustion instabilities. Although linear techniques are able to predict whether the non-oscillating steady state of a thermoacoustic system is “asymptotically” stable (without oscillations) or unstable (increasing oscillations), a thermoacoustic system can reach a permanent oscillating state (the so called “limit cycle”), even when it is linearly stable, if a sufficiently large impulse occurs. A nonlinear analysis is able to predict the existence of this oscillating state and the nature of the bifurcation process.

The aim of this work is to investigate the behavior of gas turbine combustion chambers in presence of nonlinear flame models. The bifurcation diagrams, obtained by using a continuation technique in the frequency domain, give the amplitude of the oscillations as a function of a chosen flame parameter. The Helmholtz equation is used to model the combustion chamber and nonlinear terms are introduced in the flame model, starting from the classical k–τ formulation. A three-dimensional finite element method (FEM) is used for discretization of the computational domain and a solver of quadratic eigenvalue problems is combined with Newton technique in order to identify the points of the bifurcation diagram. First, a simple Rijke tube configuration, as can be found in the literature, is examined in order to obtain bifurcation diagrams. Then, the nonlinear analysis is extended to simplified annular configurations. The obtained results show how the nonlinear behavior is influenced by varying some control parameters, such as the time delay, yielding useful indications to designers and experimentalists.
Influence of the Supply Circuit on Oil Pump Performance in an Aircraft Engine Lubrication System
Authors: Ippoliti Laurent , Hendrick Patrick
Published: November 2013

The subject of this paper is the study of the influence of inlet piping conditions on the volumetric efficiency of the oil feeding pumps by experimental and numerical ways. The major issue with the operating of pumps is the phenomenon of cavitation which decreases the efficiency and may even lead to damages. In particular, the work was focused on the influence of the fluid velocity in the inlet piping. For the experimental study, an inlet piping geometry has been used with varying diameters to prove the variation of performance with velocity. A numerical study has been conducted on a shortened fluid domain and shows promising results but still needs some improvements.
Cooling Structure Optimization for a Rib-Roughed Channel in a Turbine Rotor Blade
Authors: Chi Zhongran , Ren Jing , Jiang Hongde
Published: November 2013

Cooling design for the air-cooled turbine blades is a critical issue in modern gas turbine engineering. Advances in CFD technology is providing new prospects for turbine cooling design, as the optimum cooling structures of the blades could be designed through the optimization search coupled with the Conjugate Heat Transfer (CHT) analysis. In this paper, the optimization study for the rib arrangement of a rib-roughed channel in a rotor blade is discussed.

The optimization study introduced is realized utilizing a parametric analysis platform, which consists of the parametric design and mesh generation tool and the commercial CHT solver ANSYS CFX. For the optimization study, firstly a group of Design of Experiments (DoE) analysis of a rib-roughed rectangular channel is performed in order to find the optimum rib arrangement and to explore the objective of the optimization search. Then, the optimization search of the optimum rib arrangement is performed for a rib-roughed channel within a rotor blade based on the multi-island Genetic Algorithms (GA) of iSIGHT. During optimization search, a constant pressure drop is assumed within the cooling system, and the CHT simulations are approached for the interior only in order to make the search computationally faster.

According to the DoE analysis, minimizing the averaged wall temperature on blade surface is chosen as the optimization objective for the design of rib arrangement. The results of the GA search shows that the optimal rib arrangement with best cooling performance can be decided, and the optimal mass flow rate for the cooling channel is found simultaneously. The optimum schemes of the rib arrangement found by the DoE analysis and GA search are quite identical, which further validates the feasibility of design optimization for the blade cooling structure with the GA and CHT simulations.
Computational Investigation of Flow and Heat Transfer Characteristics of Impingement Cooling Channel
Authors: Ramakumar B. V. N. , Joshi D. S. , Sridhar Murari , Liu Jong S. , Crites Daniel C.
Published: November 2013

Impingement cooling offers very high heat transfer coefficients. Flow field, involved in impingement cooling is dominated by stagnation zone, transition zone and developing zone. Understanding of complex flow phenomenon and its effects on heat transfer characteristics is useful for efficient designing of impingement channels. Computational fluid dynamics (CFD) has emerged as a powerful tool for the analysis of flow and heat transfer systems. Honeywell has been investigating the use of CFD to determine the characteristics of various complex turbine blade cooling heat transfer augmentation methods such as impingement. The objective of this study is to develop CFD methodology which is suitable for computational investigation of flow and heat transfer analysis of impingement cooling through validation. Single row of circular jets impinging on concave (curved) surface has been considered for this study. The validation was accomplished with the test results of Bunker and Metzger [10] and with the correlations of Chupp et al. [7]. The parameters which are varied in this study include jet Reynolds number (Re2B = 6750–10200), target plate distance to jet diameter ratio (Z/d = 3 and 4), and target surface sharpness (i.e. radius ratio, r* = 0.2, 0.4 and 1) the simulations are performed under steady state conditions. Predicted results are compared for local endwall heat transfer results along the curve length of the mid span target wall. Flow field results obtained at different locations are presented to understand the heat transfer behavior.
Effect of Upstream Purge Slot on a Transonic Turbine Blade Passage: Part 2 — Heat Transfer Performance
Authors: Roy Arnab , Blot Dorian M. , Ekkad Srinath V. , Ng Wing F. , Lohaus Andrew S. , Crawford Michael E.
Published: November 2013

Heat transfer experiments with and without upstream purge cooling were carried out for a high turning (127°) airfoil passage in the presence of an upstream purge slot in a transonic linear cascade. Different coolant to mainstream mass flow ratios (MFR) were investigated at cascade design exit Mach number (0.88) and design incidence angle. The experiments were performed at Virginia Tech’s transient transonic blow down facility. A transient Infrared thermography technique was employed to measure the endwall surface temperature. Heat transfer coefficient (HTC) and film cooling effectiveness (ETA) were calculated from measured temperature assuming a 1-D semi-infinite transient conduction through a solid with convective boundary condition. In this experiment, the blade span increases in the mainstream flow direction in order to match realistic inlet/exit airfoil surface Mach number distribution. Results indicate strong interactions between coolant flow and cross passage secondary flow where significant coolant coverage is observed at higher leakage flow rates through the purge slot. The backward facing step created by the purge slot seems to be the driving factor on influencing endwall HTC compared to with or without blowing cases. Three-dimensional viscous CFD has also been performed for further insight of flow characteristics and to support experimental data. Aerodynamic measurements at cascade exit plane are provided in the companion paper GT2013-94951 [1], “Effect of upstream purge slot on a transonic turbine blade passage: Part – 1 – Aerodynamic performance”.
Variable Asymmetric Turbine for Heavy Duty Truck Engines
Authors: Schmidt Sebastian , Rose Martin G. , Müller Markus , Sumser Siegfried , Chebli Elias , Streule Thomas , Stiller Michael , Leweux Johannes
Published: November 2013

Turbochargers with variable turbine geometry (VGT) are established in diesel engines for passenger cars because of the beneficial effect on transient operation. The variability permits the reduction of exhaust back pressure, resulting in lower fuel consumption. There are only a few applications in heavy duty truck engines due to increased mechanical complexity and vulnerability to failure.

This paper presents a turbine concept with a simple variability developed for a heavy duty engine. The variability is achieved upstream of the rotor by changing the sectional area of the volute. This can be done through a rotationally movable ring which shifts the circumferential position of the volute tongues. These separate both scrolls of a double segment turbine and can be rotated by an electric actuator.

The performance maps measured at the hot gas test stand show the large variability of the flow parameter and the high efficiency levels over the operating range of the variable asymmetric turbine (VAT).

The flow field is computed by the use of 3D-CFD simulations in order to analyze the loss-generating mechanisms that occur within the machine.

Test runs on an engine test stand demonstrate the high potential of the concept concerning reduction of fuel consumption and a wide scope of realizable EGR rates in order to reduce NOx emissions in a cost-effective way. The resultant large mass flow variability allows the deletion of the waste gate and enables efficiency improvements.
Leading Edge Impingement With Racetrack Shaped Jets and Varying Inlet Supply Conditions
Authors: Jordan C. Neil , Elston Cassius A. , Wright Lesley M. , Crites Daniel C.
Published: November 2013

Impinging jets are often employed within the leading edge of turbine blades and vanes to combat the tremendous heat loads incurred as the hot exhaust gases stagnate along the exterior of the airfoil. Relative to traditional cylindrical jets, racetrack shaped impinging jets have been shown to produce favorable cooling characteristics within the turbine airfoil. This investigation experimentally and numerically quantifies the cooling characteristics associated with a row of racetrack shaped jets impinging on a concave, cylindrical surface. Detailed Nusselt number distributions are obtained using both a transient liquid crystal technique and commercially available CFD software (Star CCM+ from CD-Adapco). Three geometrical jet inlet and exit conditions are experimentally investigated: a square edge, a partially filleted edge (r/dH,Jet = 0.25), and a fully filleted edge (r/dH,Jet = 0.667). Additionally, to investigate the effect of high crossflow velocities at the inlet of the jet, a portion of the flow supplied to the test apparatus radially bypasses the impingement section. Thus, the mass flow rate into the test section is varied to achieve the desired inlet crossflow conditions and jet Reynolds numbers. As a result, jet Reynolds numbers (ReJet) of 11500 and 23000 are investigated at supply duct Reynolds numbers (ReDuct) of 20000 and 30000. The results are compared to baseline cases where no mass bypasses the test section. Additionally, the relative jet – to – jet spacing (s/dH,Jet) is maintained at 8, the relative jet – to – target surface spacing (z/dH,Jet) is 4, the target surface curvature – to – jet hydraulic diameter (D/dH,Jet) is 5.33, and the relative thickness of the jet plate (t/dH,Jet) is 1.33. Measurements indicate that the addition of fillets at the edges of the jet orifice and the introduction of significant crossflow velocity at the inlet of the jet can significantly degrade the cooling characteristics on the leading edge of the turbine blade. The magnitude of such degradation generally increases with increasing fillet size and inlet crossflow velocity. The V2F model is adequate for predicting the flow field and target surface heat transfer in the absence of inlet crossflow; however, it is believed the turbulence within the jet is overpredicted by the CFD leading to elevated heat transfer coefficients (compared to the experimental results).
Shock-Flame Interaction Modeling in a Constant-Volume Combustion Channel Using Detailed Chemical Kinetics and Automatic Mesh Refinement
Authors: Wijeyakulasuriya Sameera D. , Rajagopal Manikanda , Nalim Razi
Published: November 2013

More efficient and powerful gas turbine engines can be designed using constant-volume combustors that may involve ignition of a combustible mixture using a hot gas jet, subsequent flame and pressure-wave propagation, and their interactions. Accurate prediction of three-dimensional transient turbulent combustion is computationally challenging. To resolve propagating turbulent combustion, predict ignition, and track pressure waves accurately requires techniques to minimize the numerical cell count and kinetics calculation times. This study of shock-flame interaction (SFI) used detailed chemistry that includes low-temperature ignition reactions. Computational cells with similar temperatures and composition were grouped as ‘zones’ where kinetics are solved only once per zone per time step, using average values of species concentrations and thermodynamic properties for that zone. This avoids expensive kinetic calculations in every computational cell, with considerable speedup. A relatively coarser initial mesh was refined selectively and automatically, based on predicted velocity and temperature gradients, tracking propagating pressure waves and flames. The time step is variable, limited by the local speed of sound, to ensure accurate wave propagation. These techniques, previously validated for non-premixed, premixed and multiple-fuel turbulent combustion in industrial IC engines, are applied to study SFI during premixed combustion in a long constant-volume combustor.
Numerical Investigation of Flow and Heat Transfer in Gas Turbine Serpentine Passage Cooling and Comparison With Experimental Data
Authors: Kathiravan S. , De Prosperis Roberto , Ciani Alessandro
Published: November 2013

Due to recent advancements made in computational technology, CFD tools are capable of accurately capturing complex physical phenomenon. The proposed novel CFD methodology improves the prediction reliability and capability of Gas Turbine Blade heat transfer and secondary flow behaviour.

This paper discusses a robust CFD based methodology to validate the complex gas turbine blade cooling design using detailed 3D flow & conjugate heat transfer analysis. Both primary and secondary flow domains along with blade metal are considered in one single integrated CFD model. This will capture the coupled heat transfer and tip vortices mixing effects and hence accurately predict the secondary cooling flow. The secondary flow path geometry consists of serpentine passages with turbulator features in the flow path to improve the effective heat transfer. Several sensitivity studies were performed using the above model to understand the impact of turbulator fillets, tip hole coating thickness, domain interface and suitably accounted for in the full scale simulation. The numerical simulation results were extensively validated with GE industrial Frame5 gas turbine prototype test thermocouple data and thermal profiles (span-wise) obtained from metallographic images. This novel method gives a thorough understanding of flow-thermal physics involved in serpentine cooling and helps to optimize effective cooling flow usage.
Experimental and Theoretical Investigation of Thermal Effectiveness in Multi-Perforated Plates for Combustor Liner Effusion Cooling
Authors: Andreini Antonio , Facchini Bruno , Picchi Alessio , Tarchi Lorenzo , Turrini Fabio
Published: November 2013

State-of-the-art liner cooling technology for modern combustors is represented by effusion cooling (or full-coverage film cooling). Effusion is a very efficient cooling strategy based on the use of multi-perforated liners, where metal temperature is lowered by the combined protective effect of coolant film and heat removal through forced convection inside each hole. The aim of this experimental campaign is the evaluation of the thermal performance of multi-perforated liners with geometrical and fluid-dynamic parameters ranging among typical combustor engine values. Results were obtained as adiabatic film effectiveness following the mass transfer analogy by the use of Pressure Sensitive Paint, while local values of overall effectiveness were obtained by eight thermocouples housed in as many dead holes about 2 mm below the investigated surface. Concerning the tested geometries, different porosity levels were considered: such values were obtained both increasing the hole diameter and pattern spacing. Then the effect of hole inclination and aspect ratio pattern shape were tested to assess the impact of typical cooling system features. Seven multi perforated planar plates, reproducing the effusion arrays of real combustor liners, were tested imposing 6 blowing ratios in the range 0.5–5. Test samples were made of stainless steel (AISI304) in order to achieve Biot number similitude for overall effectiveness tests.

To extend the validity of the survey a correlative analysis was performed to point out, in an indirect way, the augmentation of hot side heat transfer coefficient due to effusion jets. Finally, to address the thermal behaviour of the different geometries in presence of gas side radiation, additional simulations were performed considering different levels of radiative heat flux.
On the Breakup Process of Round Liquid Jets in Gaseous Crossflows at Low Weber Number
Authors: Wang Shao-lin , Huang Yong , Wang Fang , Liu Zhi-lin , Liu Lei
Published: November 2013

Liquid jets in cross air flows are widely used and play an important role in propulsion systems, such as ramjet combustors. In this paper, experiments were carried out to investigate the properties of the primary breakup of liquid jets in subsonic transverse airflows at low Weber number. The test ranges included crossflow Weber numbers of 0.5–6.7, liquid-to-gas momentum ratios of 3–120, and Ohnesorge number of 0.0086. Four different injectors with diameter 0.4mm, 0.5mm, 0.6mm and 1mm have been used. A high speed camera was used to observe the jet column breakup process. Results show that the surface wavelength decreases not only with the increase of the gas Weber number but also with the increase of the momentum ratio. The breakup length decreases with the increase of the gas Weber number, in addition to its increase with the increase of the momentum ratio. The droplet diameter decreases with the increase of both the gas Weber number and momentum ratio, although the gas Weber number will dominate the breakup process. The surface wavelength, breakup length, and droplet diameter were also analyzed with to obtain semi-theoretical correlations.